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Honeywell

Commercial Flight Systems Group

Business and Commuter Aviation Systems Division

Honeywell Inc.

60X

29000

Phoenix, Arizona 85038

SPZ-8000 Digital Integrated Flight

Control System (DIFCS)

Cessna Citation Vll

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PROPRIETARY NOTICE

This document and the information disclosed herein are proprietary data of Honeywell Inc. Neither this

document nor the information contained herein shall be used, reproduced, or disclosed to others

without the written authorization of Honeywell Inc., except to the extent required for installation or

maintenance of recipient’s equipment,

NOTICE - FREEDOM OF INFORMATION ACT (5 USC 552) AND

DISCLOSURE OF CONFIDENTIAL INFORMATION GENERALLY (18USC 1905)

This document is being furnished in confidence by Honeywell Inc. The information disclosed herein

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Date Received

Honeywell’s Continuous Quality Process

READER COMMENTS

(Mail or FAX this form to [602] 436-4100)

Honeywell welcomes all comments and recommendations to improve future editions of this publication.

Your Name

Company/Airline

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Country

Zip

Telephone No. FAX Date

Honeywell Pub. No.

ATA

No.

Manual

COMMENTS/RECOMMEN DATIONS:

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FOLD

FOLD

-----------------------------------------------------------------------------------------------------------------------------------------------------------

From

Honeywell

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Date

REPORT OF POSSIBLE DATA ERROR

(Mail or FAX this form to [602] 436-4100)

Your Name

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Received

Address

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country Zip

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FAX

Date

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Manual

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—

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NO.

PARA-

GRAPH

FIGURE

NO.

—

TABLE

NO.

PROBLEM

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RECORD OF REVISIONS - VOLUME I

For each revision, put the revised pages in your manual and discard the superseded pages. Write the

revision number and date, date put in manual, and the incorporator’s initials in the applicable columns on

the Record of Revisions. The initials HI show Honeywell Inc. is the incorporator.

Revision

Revision

Date Put

Number

Date

In Manual

By

Revision Date Put

Insertion

Number In Manual Date

By

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Hone~ell

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System Description (cent)

49

50

51

52

53

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57

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Component Description

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Component Description (cent)

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F 197/198

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198.2

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F 198.71198.8

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F 198.13/198.14

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REVISION

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198.113

198.114

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198.118

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198.122

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REVISION

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198.163/1 98.164

System Operation

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1

TABLE OF CONTENTS - VOLUME 1

Para~raph

Svstem Description

1.

General

2. System Description

A. AHZ-600 Attitude and Heading Reference System (AHRS)

B.

ADZ-81O Air Data System

c.

AA-300 Radio Altimeter System (Optional)

D. EDZ-816 Electronic Flight Instrument System (EFIS)

E. DFZ-800 Dual Flight Guidance System

F. PRIMUS@ 870 Digital Weather Radar System

G. MDZ-816 Multifunction Display System (Optional)

H.

SRZ-850 Integrated Radio System

1. FMZ-800/900 Flight Management System (Optional)

J. LSZ-850 Lightning Sensor System (Optional)

K. TCAS II (Optional)

L. Global Positioning System (Optional)

M. LASEREF@ Ill Inertial Reference System (Optional)

3.

Digital Information Transfer Systems

Paae

1

1

23

24

25

25

26

26

27

27

28

32

33

34

35

35

36

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TABLE OF CONTENTS - VOLUME I (cent)

Section

Para~raDh

2

Comwnent Descri~tion (cent)

4.

Optional AA-300 Radio Altimeter System

A. RT-300 Radio Altimeter Receiver/Transmitter

B. AT-300 Radio Altimeter Antenna

5. Paragraph 5 is not applicable to this system.

6.

EDZ-816 Electronic Flight Instrument System (EFIS)

and Optional MDZ-816 Multi function Display (MFD) System

A.

B.

c.

D.

E.

F.

G.

H.

1.

J.

ED-800 Electronic Display

ED-800 Used As An Electronic Attiiude

Director Indicator (EADI)

ED-800 Used As An Electronic Horizontal

Situation Indicator (EHSI)

EFIS Reversionary Controls and Annunciators

ED-800 Used As A Multifunction Display (MFD)

SG-816 Symbol Generator

MG-816 MFD Symbol Generator

DC-81 O Display Controller

MC-800 MFD Control ler

RI-206S Instrument Remote Controller

~

136

136

139

140

140

140

145

158

177

179

187

193

198.2

198.9

198.15

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Section

2

TABLE OF CONTENTS - VOLUME I (cent)

Paragraph

Component DescrirXion (cent)

D.

RM-850 Radio Management Unit (RMU)

E. AV-850A Audio Control Unit

F. CD-850 Clearance Delivery ControVDisplay Unit

G.

DI-851 DME Indicator

H. AT-860 ADF Antenna

1. AT-851 MLS Antenna

11. Optional LSZ-850 Lightning Sensor System

A. LP-850 Lightning Sensor Processor

B. LU-850 Lightning Sensor Controller

c.

AT-850 Lightning Sensor (Teardrop) Antenna

D. AT-855 Lightning Sensor (Brick) Antenna

12. Optional Traffic Alert and Collision Avoidance

System (TCAS 11)

A. RT-91 O TCAS Computer Unit

B. DV-91 O VSi/TRA Display

c.

TCAS/RMU Control

D. AT-91 O Directional Antenna

E.

Typical Bottom Omnidirectional Antenna

PaJp

198.86

198.91

198.95

198.99

198.102

198.104

198.105

198.105

198.111

198.113

198.115

198.116

198.116

198.122

198.128

198.134

198.135

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Section

4

5

6

7

8

TABLE OF CONTENTS - VOLUME II

Paragraph

Ground Check

1< General

2. Equipment and Materials

3.

Procedure

Fault Isolation

1. General

2.

Procedure

Interconnects

Table 501 - Interconnect Information

Table 502- Optional System Interconnect Information

Svstem Schematics

Removal/Reinstallation and Adjustment

1.

2.

3.

General

Equipment and Materials

Procedure for Displays and Indicators

Page

301

301

301

301

401

401

401

501

503

598.259

601

701

701

701

702

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LIST OF ILLUSTRATIONS - VOLUME I

3wE

1-1

1-2

1-3

1-4

1-5

1-6

1-7

1-8

1-9

1-1o

1-11

1-12

System Flow Diagram

SRZ-850 Integrated Radio System Flow Diagram

SPZ-8000 Wdh MFD System Flow Diagram

Optional FMZ-800/900 Flight Management System

Optional LSZ-850 Lightning Sensor System

TCAS/Mode S Flow Diagram

Global Positioning SysterrVlnertial Reference System Flow Diagram

Standard Component Locations

Optional Component Locations

Radio Management System Bus Diagram

Radio System Bus (RSB) Network

Lightning Symbols

E w

7

9

11

13

14

15

17

19

21

29

31

33

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Em E

1-24

1-25

1-26

1-27

1-28

1-29

1-30

1-31

1-32

1-33

1-34

2-1

LIST OF ILLUSTRATIONS - VOLUME I (cent)

ASCB Waveform

RSB Data Field Structure

Audio System Bus Network

Digital Audio Data Sequence

Octal Label 274

Data Bits 11 thru 29

BCD Bit Assignments

BCD Data for Selected Course

Five-Character DME Word

Six-Chacacter DME Word

ARINC Data Transmission

AH-600 Strapdown AHRU

E aw

53

55

59

60

61

62

62

62

63

63

64

102

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Eh E

3-5

3-6

3-7

3-8

4-1

4-2

4-3

6-1

6-2

6-3

6-4

6-5


AL-801 Altitude Preselect Controller

AL-801 Altitude Preselect Controller Block Diagram

DS-1 25A TAS Temperature Indicator

DS-1 25A TAS Temperature Indicator Block Diagram

RT-300 Radio Altimeter ReceiverTfransmitter

RT-300 Radio Altimeter Receiverfiransmitter Block Diagram

AT-300 Radio Altimeter Antenna


ED-800 Electronic Display Block Diagram

EADI Displays and Annunciators

EADI - Amber Caution and Failure Annunciations

Red EADI Failure Annunciations

PaJgg

130

132

133

135

136

138

139

140

143

149

154

157

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m E

6-18

6-19

6-20

6-21

6-22

6-23

6-24

6-25

6-26

6-27

7-1

7-2



SG-816 Symbol Generator Block Diagram

MG-816 MFD Symbol Generator

MG-816 MFD Symbol Generator Block Diagram


DC-81 O Display Controller Block Diagram

MC-800 MFD Controller

MC-800 MFD Controller Block Diagram


RI-206S Instrument Remote Controller Schematic

FZ-800 Flight Guidance Computer

FZ-800 Flight Guidance Computer Block Diagram

Page

187

191

193

197

198.2

198.7

198.9

198.13

198.15

198.16

198.17

198.20

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EWE

9-3

9-4

9-5

9-6

9-7

9-8

9-9

10-1

10-2

10-3

104

10-5


CD-800/81 O Control Display Unit

CD-800/81 O Control Display Unit Block Diagram

DL-900 Data Loader

DL-900 Data Loader Block Diagram

OZ-800 Receiver Processor Unit

OZ-800 Receiver Processor Unit Block Diagram

AT-801 H-Field Brick Antenna

RNZ-850 Integrated Navigation Unit

RNZ-850 Integrated Navigation Unit Block Diagram

RCZ-850/851A Integrated Communication Unit

RCZ-850 Integrated Communication Unit Block Diagram

RCZ-851 A Integrated Communication Unit Block Diagram

?s9 2

198.54

198.61

198.63

198.64

198.65

198.67

198.68

198.69

198.74

198.75

198.78

198.79

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E 9 &

10-18

11-1

11-2

11-3

11-4

11-5

11-6

12-1

12-2

12-3

12-4

12-5

LIST OF ILLUSTRATIONS - VOLUME 1(cent)

AT-851 MLS Antenna

LP-850 Lightning Sensor Processor

LP-850 Lightning Sensor Processor Block Diagram

LU-850 Lightning Sensor Controller

LU-850 Lightning Sensor Controller Schematic

AT-850 Antenna

AT-855 (Brick) Antenna

RT-91O TCAS Computer

TCAS CU Panel Layout

RT-91 O TCAS Computer Block Diagram

DV-91 O VS1/TRA Display

DV-91 O VSlflFIA Display Formats

Page

198.104

198.105

198.107

198.111

198.112

198.113

198.115

198.116

198.118

198.121

198.122

198.124

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Ew ?

14-2

14-3

14-4

14-5

14-6

14-7

14-0

201

202

203

204

205

IRU

IRU

IRU


Mounting Tray and Blower Kit

Rear Connector Layout

Signal Interface Diagram

Mode Select Unit (MSU)

MSU Schematic

Battery Backup Unit

Typical IRU Battery Power Operating Time

AP, YD, HSI SEL, and GA Mode Select Diagram

ED-8oo EADI Display Flow Diagram

Dual EFIS Display System EADI Interconnects

ED-800 EHSI Display Fiow Diagram

Dual EFIS Display System EHSI Interconnects

Page

198.154

198.156

198.157

198.158

198.161

198.162

198.163

213

215

217

219

221

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LIST OF TABLES - VOLUME I

Table

1-1

1-2

1-3

1-4

1-5

1-6

1-7

2-1

2-2

2-3

2-4

Standard System Components

Optional System Components

Equipment Required But Not Supplied by Honeywell

ASCB Frame Structure Allowing 40, 20, and 10

Update Rates

RSB Message Numbers (NORMAL MODE)

SSM Bit Assignments

Differential Output Voltages

AH-600 Strapdown AH RU Leading Particulars

AH-600 AHRU Dip Angle Compensation Programming

CS-412 Dual Remote Compensator Leading Particulars

FX-600 Thin Flux Valve Leading Particulars

PaJ&

2

4

6

48

56

63

65

103

106

115

117

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LIST OF TABLES - VOLUME I (cent)

Table

6-6

6-7

7-1

7-2

7-3

7-4

8-1

8-2

8-3

9-1

9-2

9-3

MC-800 MFD Controller Leading Particulars

RI-206S Instrument Remote Controller Leading Patiiculars

FZ-800 Flight Guidance Computer Leading Particulars

GC-81O Flight Guidance Controller Leading Particulars

SM-200 Servo Drive and SB-201 Bracket Leading Particulars

SM-200 Servo Drive Dash No. Differences

WU-870 Antenna and Receiver/Transmitter Leading Particulars

WC-870 Weather Radar Controller Leading Particulars

WI-870 Weather Radar Indicator Leading Particular

NZ-820/920 Navigation Computer Leading Particulars

CD-800/81 O Control Display Unit Leading Particulars

DL-900 Data Loader Leading Particulars

?s9 2

198.9

198.15

198.18

198.21

198.28

198.28

198.32

198.35

198.44

198.50

198.55

198.63

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Table

10-10

11-1

11-2

11-3

11-4

11-5

12-1

12-2

12-3

12-4

12-5

12-6

LIST OF TABLES - VOLUME I (cent)

AT-851 MLS Antenna Leading Particulars

LP-850 Lightning Sensor Processor Leading Particulars

LP-650 Configuration Strap (CS) Jumpers

LU-850 Lightning Sensor Controller Leading Particulars

AT-850 Antenna Leading Particulars

AT-855 Antenna Leading Particulars

RT-91 O TCAS Computer Leading Particulars

RT-91 O TCAS Computer ARINC 429 Output Data

RT-91 O TCAS Computer-To-Mode S Transponder Data

XS-91 O Mode S Transponder-To-TCAS Computer Data

DX-91 O VSl~RA Display Leading Particulars

TCAS Symbology

PaJ&

198.104

198.105

198.109

198.111

198.113

198.115

198.116

198.119

198.120

198.120

198.122

198.126

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LIST OF TABLES - VOLUME II

Table

201

System Pedormance/Operating Limits

301 Ground Maintenance Test Procedure

501 Intermnnect Information

502 Interconnect Information for LASERE@ Ill, TCAS 11,AA-300 Radio

Aftitude Systems and the DS-1 25A TAS/TEMP Indioator

Paae

202

303

503

598.259

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INTRODUCTION

This manual provides general system maintenance instructions and theory of operation for the SPZ-8000

Digital Integrated Flight Control System (DIFCS) for Cessna Citation Vll aircraft.

This manual provides block diagram information and intemonnect diagrams to permit a general

understanding of System interface.

Common system maintenance procedures are not presented in this manual. The best established shop

and flight l ine practices should be used.

Reference Documents

System checkouts, operational testdchecks, fault isolation, and repair are made only during ground

maintenance. Detailed instructions for these ground maintenance procedures are presented in the following

Honeywell Description and Installations manuals listed below.

SVstem

Honeywell Pub. No.

AA-300 Radio Altimeter System

15-3321-06

Global Positioning System Sensor Unit

95-8698

LASERE@ Ill Ineriial Reference System

15-3343-011

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Abbreviation

AHRS

AHRU

AIL

ALT

ANN, ANNUN

ANT

AOSS

AP, AIP

APE

APP, APR

APS

APSB, APSBK

ARM

AS

ASCB

All

AUX

AZ

BARO

WA

BC

BCD

BRG

BRK

CAP

Description

Attitude and Heading Reference System

Attitude and Heading Reference Unit

Aileron

Altitude

Annunciator

Antenna

After Over Station Sensor

Autopilot

Autopilot Engage

Approach

Altitude Preselect

APS Bracket

Armed

Airspeed

Av”kmicsStandard Communications Bus

Attitude

Auxiliary

Azimuth

Barometric

Bank Angle

Back Course

Binary-Coded-Decimal

Bearing

Brake

Capture

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Abbreviation

DEFL

DEG

DEMOD

DET

DEV, DEVN

DG

DH

DIFCS

DIFF

DISPL

DMA

DME

DN

DRC

DSR

DUP

EFIS

EL, ELEV

EMI

ENG

EO

E OFF

EX LOC

EXT

FD, F/D

Description

Deflection

Degree

Demodulator

Detector, Detent

Deviation

Directional Gyro

Decision Height

Digital Integrated Flight Control System

Differential, Difference

Displacement

Direct Memory Access

Distance Measuring Equipment

Down

Dual Remote Compensator

Desired

Duplicate

Elect ronic Flight Instrument System

Elevator, Elevation

Elect romagnetic interference

Engage

Easy-On

Easy-Off

Expanded Localizer

Extend, External

Flight Director

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Abbreviation

ILS

INC-DEC

IND

INS

INTLK

INTGL

INV

1/0

IRC

IRS

IRU

1s0

Ivv

KN

L

LAT

LBS

L/c

LH

LOC

LP

LPV

LRN

LSS

LTG

Description

Instrument Landing System

Increase-Decrease

Indicator

Inertial Navigation System

Interlock

Integral

Invert

Input/Output

Instrument Remote Controller

Inertial Reference System

Inertial Reference Unit

Isolation

Instantaneous Vertical Velocity

Knots

Left

Lateral

Lateral Beam Sensor

Inductive/Capacitive

Left Hand

Localizer

Lightning Processor

Latched Power Valid

Long-Range Nav

Lightning Sensor System

Lighting

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Abbreviation

Pews

Plso

PITCH SYNC

POR

PRI, PRIM

PROC

PROG

P/s

Pv

Pw

PWM

PWR

R

RA

RA, R/A, RAD ALT

RAM

RCB

RCT

RCVR

RCWS

REF

REL

RET

RETR

REV

Description

Pitch Control Wheel Steering

Parallel In Serial Out

Pitch Synchronization

Power On Reset

Primary

Processor

Programmer, Programming

Pitot Switch

Power Valid

Pitch Wheel or Pulse Width

Pulse Width Modulated

Power

Right

Resolution Advisories

Radio Altimeter

Random Access Memory

Radio Communication Bus

React

Receiver

Roll Control Wheel Steering

Reference

Release

Return

Retract

Reverse Course (Same as Back Course)

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Abbreviation

SRN

SSEC

STAB

STAEL

STC

STCS

STP

STR, STRG

SYNC

TA

TAS

TAT

TCAS

TCS

TGT

TK

TKE

TLA

TLE

TP

TRK

TSO

lTL

UART

Description

Short-Range NAV

Static Source Error Correction

Stabilization

Station Elevation

Sensitivity Time Control

Single Trim Channel Select

Steep

Steering

Synchronization

Traffic Advisories

True Airspeed

True Air Temperature

Traffic Alert and Collision Avoidance

System

Touch Control Steering

Target Alert

Turn Knob

Track Error

Torque Limit Aileron

Torque Limit Elevator

Test Point

Track

Technical Standard Order

Tuned to Localizer

Universal Asynchronous Receiver Transmitter

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NOTICE

[

CRITICAL ITEMS

COMPLIANCE REQUIRED

/

Honeywell has an Airworthiness Analysis procedure performed for all its airborne products to ensure that

equipment designed by Honeywell will not create a hazardous in-flight condition. As a result of the

Analysis, certain installations have been designated INSTALLATION CRITICAL, and 100 percent

compliance with those installations is required.

INSTALLATION CRITICAL is defined as:

Specific methods of installation are required to ensure that either the failure of the assembly or part is

extremely improbable or that its failure could not create a hazardous condition. The clearance

(distance) between the keeper pins and the drum brackets, and the diameter of the aircraft control

cables are designated INSTALLATION CRITICAL.

Measuring the distance between the keeper pins and the servo drum bracket for proper clearance,

and verifying the diameter of the aircraft control cables are critical to avoiding failures that could cause

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Aircraft

System Component

Qty Parl No. Ref Des.

Attitude and Headina Reference System

AH-600 Strapdown Atti tude and Heading

2

Reference Unit (AHRU)

FX-600 Flux Valve

2

CS-412 Dual Remote Compensator

1

Electronic Flight Instrument System (EFIS)

ED-800 Electronic Display (EHSI) 2

ED-800 Electronic Display (EADI)

2

Inclinometer Kit

2


1



DC-8 10 Display Controller

Air Data System

2

(see note)

2

2

AZ-81 O Digital Ak Data Computer


2

1

7003360-932

7010133

2593379-002

7003110-921

7003110-921

7005400-901

4026206-974

7011674-316

7011674-416

7005819-729

7000700-954

7004577-903

1

4

5

2

3

N/A

23

65

65

115

9

16

8/11/2019 spz-8000 difcs


Aircraft

Svstem Component

@

Part No.

Ref Des.

Integrated Radio Svstem

ML-850 MLS Receiver

AT-851 MLS Antenna (Fore)

AT-851 MLS Antenna (Aft)

RCZ-850 Integrated Communication Unit

RM-850 Radio Management Unit

AT-860 ADF Antenna

AV-850A Audio Control Unit



CD-850 Clearance Delivery Control/Display

Unit

2

2

2

2

2

2

2

2

2

1

7510600-901

7510638-901

7510638-901

7510700-901

7012100-983

7510300-901

7511001 -9XX

7513006-911

7510100-911

7513000-805

116

118

119

143

144

158

160

163

164

165

Standard System Components

Table 1-1 (cent)

8/11/2019 spz-8000 difcs


Aircraft

System Component

m

Part No.

Ref Des.

Radio Altimeter System

RT-300 Radio Altimeter Receiver~ransmitter

AT-300 Antenna (Receiver)

AT-300 Antenna (Transmitter)

1

1

1

7001840-926

7003586

7003586

20

21

22

Air Data Svstem

DS- 125A TAS/SAT/TAT Indicator

1

7002638-906

27

Weather Radar System

WC-870 Weather Radar Controller (note 1)

WC-870 Weather Radar Controller (note 2)

WI-870 Weather Radar Indicator (note 2)

2

1

1

7008471-803

7008471-801

7007700-801

61

61

63

Fliaht Management System

CD-81 O Control Display Unit (Color option)

NZ-820/920 Navigation Computer (note 3)

DL-900 Data Loader (note 4)

2

2

1

7007549-901

7004402-VAR

7016600-901

120

121

123

8/11/2019 spz-8000 difcs


Aircraft

System Component

Qtv Part No.

Ref Des.

LicahtningSensor System

LP-850 Lightning Sensor Processor

LU-850 Lightning Sensor Controller (note 2)

AT-850 Lightning Antenna (Teardrop)

AT-855 Lightning Antenna (Brick)

Global Positionirm System

GZ-81 O Global Positioning System Sensor Unit

(GPSSU)

AT-81 O GPS Antenna

Inertial Reference System

Inertial Reference Unit (IRU)

Mode Select Unit

TCAS II

XS-91 O Mode S Transponder

RT-91 O TCAS R/T Computer Unit

AT-91 O TCAS Directional Antenna

1

1

1

1

1

1

2

2

1

1

1

7011822-903

7011865-903

4057697-901

7014062-901

HG2021AB02

26002806-201

HG2001 ABXX

CG1042AB04

4061400-903

4066010-903

7514060-902

145

146

147

147

149

150

170

172

191

193

194

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Function and Designation

Qtv DescrirXion

AP Disconnect Switches - S1, S2

2 Normally closed, momentary pushbutton,

1 pole; Switch rated at 28 V dc at 100 rnA

Touch Control Steering Switches - S3, S4

11 Normally open, momenta~ pushbutton, 1 pole;

Go-Around Switches - S5, S6

28 Vdcat 100mA

AHRS Switches - S7, S8

Annunciator Reset Switch - S9

Reversion Switches - S1O, S11, S12, S13

Maintenance Test SEL Switch - S14

ASCB Test Connectors:

Pl, CPI

Jl, CJ1

P2, CP2

J2, CJ2

Over Temp/System Maintenance

Annunciators; FB-1 through FB-8

Aural Alefi Horn; Horn-1, Horn-2, Horn-3

Relays; K-1, K-2

Resettable Annunciators; SA-1, SA-2

1

2

2

2

2

8

3

2

2

2-pole single throw; toggle; rated at 100 rnA

Trompeter 3105-0032-2 (4 lug)




Magnetic Latching Annunciators, Minelco Part

No. BHGD21T-28-BLK/YS-209, Annunciator

Mfg, Minelco Inc, Sub. of Talley Industries Inc,

Thomastom, CT

Mallory Sonalert, Model SC-628, or equivalent

1-pole double throw, Coil -28 V dc, 100 mA,

Contacts - c 1 ampere

Momentary SPST switches with integral

8/11/2019 spz-8000 difcs


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Figure 1-3

22-05=07

Pages 11/12

Jun 1/93

8/11/2019 spz-8000 difcs


429S

-

.

RECEIVER

PROCESSOR UNIT

RECEIVER

PROCESSOR UNIT

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TO sPZ-8W0

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8/11/2019 spz-8000 difcs


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8/11/2019 spz-8000 difcs


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DS-125A TAS/TEMP

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AUDIO CONTROL UNIT

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Pages 15/16

RT-300 RADIO ALTIMETER

RECEIVER/TRANSMITTER

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8/11/2019 spz-8000 difcs


8/11/2019 spz-8000 difcs


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Standard Component Locat”nns

Figure 1-8

ITEM NO.

1.

2.

3.

4.

5.

6.

7.

6.

9.

10.

11.

12.

13.

14.

15.

16.

t7.

NOMENCLATURE

AH-600STRAPDOWNAHRU

FZ-600FLIGHTGUIDANCECOMPUTER

SG-616SYMBOLGENERATOR

AZ-61ODIGITALAIRDATACOMPUTER

WU-870ANTENNAANDRCVFVXMTR

RCZ-650INTEGRATEDCOMMUNIT

RNZ-650INTEGRATEDNAVUNIT

MSL-650RECEIVER

AT-851MLSAANTENNA

SM-200SERVO DRIVE(ELEVATOR)

SM-200SERVO DRIVE(RUDDER)

FX-600THIN FLUXVALVES

SM-200SERVO DRIVE(AILERON)

CS412 DUALREMOTECOMPENSATOR

INSTRUMENTPANELANDPEDESTAL

MOUNTEDCOMPONENTS

ED-6ooELECTRONICDISPIAYS (EADI

GC-61OFLIGHTGUIDANCECONTROL

W1470 WXtNOICATOR

DC-61ODISPLAYCONTROLLER

01-651DME INDICATOR

AL-601ALTITUDEPRESELECTCONTR

RM-650RADIOMANAGEMENTUNIT

RI-206SINSTRUMENTREMOTECONT

AV-650AAUDIOCONTROLUNIT

SI-225N225S MACHAIRSPEEDINDIC

CD-65OCLEARANCEDELIVERYCON

DISPLAYUNIT

AILERONSURFACEPOSITIONFEEDBAC

SYNCHRO(CESSNAFURNISHEDITEM)

AT-6&1A13FANTENNA

AO

22-05-07

Pages 19/20

Jun 1193

8/11/2019 spz-8000 difcs


c

2 (PILOT’S)

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Optional Component Locations

Figure 1-9

/

/

4 ,6 , 1 5

ITEMNO

-

1.

2.

3.

4.

5.

8.

7.

8.

9.

10.

11.

12.

13,

14.

15.

16.

NOMEMCIATURE

MO-616MFD SYMBOLGENERATOR

LASERIllINERTIALREF UNIT

NZ-S20/920NAVCOMPUTER

GZ-61OGLOBALPOSITIONINGSYSTEM

SENSORUNIT

AT-61OGPSANTENNA

XS91OKK)DES TRANSPONDER

AT-3LMRADIOALTIMETERANTENNA(REC

AT-3ooRADIOALTIMETERANTENNA(TRA

RT-300RADIOALTIMETERRECEIVER/TRA

LP-650LIGHTNINGSENSOR PROCESSOR

02-600 RECEIVERPROCESSORUNIT

AT-S5LU8S5IGHTNINGANTENNA

AT-SWH-FIELDBRICKANTENNA

RT-910TCASDIRECTIONALANTENNA

RT-91OTCASR/l COMPUTERUNIT

INSTRUMENTPANELANDPEDESTAL

MOUNTEDCOMPONENTS

EO-600ELECTRONICDISPLAY(MFD)

LU-650LIGHTNINGSENSORCONTROLL

WC-670WX CONTROLLER

WI-670WX INDICATOR

DS-125ATASTEMP INDICATOR

WDE SELECTUNIT

DV-91OVS1/TRADISPLAY

MC-600MFDCONTROLLER

cD._10 CONTROLDISPLAYUNIT

Pages 21/22

8/11/2019 spz-8000 difcs


2.

System Description

The SPZ-8000 DIFCS consists of the following subsystems, which are described in paragraphs 2.A

thru 2.M.

AHZ-600 Atti tude and Heading Reference System (AHRS)

ADZ-81 O Air Data System (ADS)

AA-300 Radio Altimeter System


DFZ-800 Dual Fl ight Guidance System

PRIMUS@ 870 Weather Radar System

MDZ-816 Multifunction Display (MFD) System


FMZ-800/900 Flight Management System (FMS)

LSZ-850 Lightning Sensor System

Traffic Alert and Coll ision Avoidance System (TCAS 11)

Global Positioning System (GPS)

LASEREF@ Ill Inertial Reference System (IRS)

The SPZ~8000 is a complete automatic flight control system providing complete fail-operational

execution of fl ight director guidance, autopilot, yaw damper, and trim functions. The automatic path

mode commands are generated by the FZ-800 fl ight guidance computer, which integrates the attitude

and heading reference, air data, and EFIS into a complete aircraft control system. As a control

system, the SPZ-8000 DI FCS provides the stabilization and control needed to ensure optimum

8/11/2019 spz-8000 difcs


2,

The system data communication is split between the main system bus (ASCB) and priiate line paths

provided for specific sensitive data for which fault isolation is required. These specific private line

paths include the following:

AH-600 AHRU attitude and heading to EFIS

AZ-81 O DADC Mach airspeed to SI-225W225S Mach Airspeed Indicator

AZ-81 O DADC vertical speed to EHSI Vertical Speed Indicator

GC-81 O Controller to FZ-800 Flight Guidance Computer

DC-81 O Display Controller to SG-816 Symbol Generator

c SG-816 Symbol Generator to ED-800 EFIS displays

MC-800 MFD Controller to MG-816 MFD Symbol Generator

. MG-816 MFD Symbol Generator to ED-800 MFD display

Also, switched NAV data is input directly to the AFCS and flight instruments to ensure that both

subsystems may independently assess ILS and MLS data during approaches.

The system displays heading, course, radio bearing, pitch and roll attitude, barometric altitude,

selected alert altitude, radio altitude, rate-of-turn, course deviation, glideslope deviation, to-from

indications, and DME indications. Lighted annunciators denote selected flight mode, altitude alert,

decision height, and go-around mode engagement. Pitch and roll steering commands developed by

the FZ-800 Flight Guidance Computer in conjunction with the GC-81 O Flight Guidance Control ler are

displayed by steering pointers to enable the pilot to reach and/or maintain the desired flight path or

attitude.

A. AHZ-600 Attitude and Heading Reference System (AHRS)

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External switches enable the pilot to manually slew the AH-800 AHRU heading to

output and to make mode control inputs to the AHRU. The CS-412 Dual Remote

provides single cycle N-S, E-W compensation for the flux valve.

2. B. ADZ-81 O Air Data System

The ADZ-81 O Air Data System comprises the following components:

AZ-81 O Digital Air Data Computer

.

DS-125A TAWTEMP Indicator (Pilot’s side)

.

S1-225S Mach Airspeed Indicator (Pilot’s Side)

.

SI-225A Mach Airspeed Indicator (Copilot’s Side)


any desired

Co~ensator

The AZ-81 O Digital Air Data Computer (DADC) is a microprocessor-based digital computer that

accepts both digital and analog inputs, performs digital computations, and supplies both digital

and analog outputs. It receives pitot-static pressures and total air temperature inputs for

mmputing the standard air data functions. The DADC provides outputs suitable for driving the

S1-225/U225S and DS-1 25A indicators, transfxmder, flight recorder, flight director, and autopilot,

as well as other elements of the flight mntrol system.

The AL-801 Altitude Preselect Controller provides displays for altiiude alefting and attitude

preselect. The amputations for these functions are performed by the AZ-81 O DADC.

c.


8/11/2019 spz-8000 difcs


2. D.


The EFIS comprises the following components:

ED-800 Electronic Display (EADI and EHSI)




The EFIS displays pitch and roll attitude, heading, course orientation, flight path commands,

weather presentations, and mode and source annunciations, air data parameters, and fault

warning information.

The primary features the EFIS brings to the fl ight control system are display integration,

flexibility, and redundancy. Essential display information from sensor systems, and automatic

flight control, navigation, performance, and caution-warning systems are integrated into the

pilot’s prime viewing area. Each symbol generator is capable of driving four ED-600 displays,

such that in case of a symbol generator failure, the remaining symbol generator drives the

displays on both sides. In the case of a display failure, a composite attitude/heading display

format can be displayed on the remaining display.

The switching of attitude and navigation sensor data to be displayed is provided electronical ly.

All comparison monitoring of criiical display information is done within the EFIS.

The primary attitude data from the AHRS is sent to the EFIS symbol generator over a dedicated

serial bus to meet the certification requirements for isolation of the primary data to the pilot’s

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The single GC-81 O Flight Guidance Controller is used to engage the system, select the

operating modes, and select the HSI and DADC being used to interface with the flight guidance

mmputer. The pitch wheel is also part of this unit.

2. F.

PRIMUS@ 870 Digital Weather Radar System

The PRIMUS@ 870 Weather Radar system is an X-Band digital radar, designed for weather

detection and analysis, and ground mapping. The system consists of the following components:

.

WU-870 Antenna and Receiver/Transmitter Unit

WC-870 Weather Radar Controller

WI-870 Weather Radar Indicator (Optional)

The PRIMUS@ 870 system detects storms along the flight path of the aircraft and gives the flight

crew a visual indication, in color, of storm intensity. In the weather detection mode, target

returns are displayed at one of five video levels (O, 1, 2, 3, or 4), with O represented by a black

screen because of weak or no returns, and levels 1, 2, 3, and 4 represented by green, yellow,

red, ati magenta respectively, to show progressively stronger returns. In ground mapping

mode, video levels of increasing reflectivity are displayed as black, cyan, yellow, and magenta.

When the PRIMUS@ 870 is operated in con@ction with the EFIS, radar video is provided for

display on the EHSI. Radar information may also be displayed on the Multifunction Displays

(MFDs). The radar range, radar operating mode, and antenna tilt functions are all controlled by

pushbuttons on the WC-870 (if installed the WI-870), or menu selections on the MFD. The pilot

side display and the copilot side display are independently controllable. When both displays are

active, the left side EFIS or MFD displays the data received during the time that the antenna is

sweeping from left to right, and the right side EFIS and MFD displays the data received during

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2. H.


The SRZ-850 Integrated Radio System includes the following components:

.

.


RCZ-850 Integrated Communication Unit

RCZ-851 A Integrated Communication Unit (for use with TCAS only)

ML-850 MLS Receiver

RM-850 Radio Management Unit (RMU)


CD-850 Clearance Delivery ControlKlisplay Unit


AT-860 ADF Antenna

AT-851 MLS Antennas (two per system)

The SRZ-850 Integrated Radio System is a dual, remote-mounted, digital radio system that

enmmpasses all standard navigation and communication functions, including VOR, DME, ILS,

MLS (optional), VHF communication with extended frequency range, MARKERS, and Modes

A/C/S Transponders, al l of which are operated from two (Pilot and Copilot) Radm Management

Units (RMUS). The radio system also interfaces with the optional Traffic Alert and Collision

Avoidance System (TCAS). The RMUS also provide backup navigation display capabil ities.

The ML-850 Microwave Landng System (MLS) Receiver is used (as an ancillary NAV unit) in

conjunction with the RNZ-850 Integrated Navigation Unit. The RNZ-850 Integrated Navigation

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SECONDARY

PRIMARY

RSB

RM-B50

RADIO SY

RMu

J

I

w

AV-850A

AUDIO CONTROL UNIT

RCZ.850

INTEGRATED

COM UNIT

(

(

DIGITAL

AUDIO

RNZ-850

SECONDARY

EMBUS

RM.850

RSB

RMu

GiiiiiJ

AV-850A

AUDIO CONTROL UNIT

-m–

8/11/2019 spz-8000 difcs


With the exception of a DME Indicator, all the navigation data is displayed on the EFIS. A

separate DME indicator is provided, that displays the distance, groundspeed, time-to-station, and

station identifier.

Basic to the overall system design are cluster modules in the COM and NAV remote units. The

cluster module is an interfacing element that collects data from the RSB, distributes this data to

the respective functional modules (ADF, DME, etc) via the MLS Radio Communication Bus

(RCB), and also collects data through the MLS RCB from the functional modules to be broadcast

on the RSB.

The RM-850 RMU broadcasts messages addressed to radio functional modules and receives

data from the radios via the RSB. Three major functions of the RMU are to output tuning

(channel or frequency) control data, output operational mode control data for the radios, and

display the tuned active channel or frequency and operational mode.

The RSB is a high-speed (667 kHz) multi-user bus that allows all radios and control heads to

broadcast data on the bus for the purpose of tuning radios to the desired channel or frequency

for aircraft communication and navigation. Three buses are used for redundancy in the event

that one or more buses become inoperative for any reason.

Physically, RSB consists of three separate multi-user serial halfduplex, digital communications

buses. Each bus is electrically isolated from the others, and all buses are electrically isolated

(transformer coupled) from the circuitry inside the units installed on the bus. Each bus is

terminated at each end with an appropriate termination (resistor) network as shown in figure

1-11.

The bus is connected in a manner that one primary bus serves all of the radios. Two secondary

8/11/2019 spz-8000 difcs


r

———.

PILOT RMU

I

I

(

SECONDARY +

RSB

1-

1

(

+

I

PRIMARY

RSB _

L ———.

r

———.

PILOT OME

lNOfCATOR

I

(

SECONDARY +

RSB

1-

1

I

I

(

RIMARY +

RSB _

L ———.

pTmGv-

1

{

ECONDARY +

RSB

1-

1

BOBBIN

,—— —

COPILOT RMU

1

I

1

+

SECONDARY

- RSB

1

I

-1

I

+

PRIMARY

RSB

I

)———

J

1

I

I

I

I

I

-l

1

I

I

I

8/11/2019 spz-8000 difcs


2. 1.

FMZ-800/900 Flight Management System (Optional)

The FMZ-800/900 Flight Management System (FMS) consists of the fol lowing components:

CD-800/81 O Control Display Unit

NZ-820/920 Navigation Computer

DL-900 Data Loader

OZ-800 Receiver Processor Unit


The FMS has many varied functions such as; remote radio tuning, flight plan building and

storage, waypoint creation and storage, and information on navaids and earth reference points,

such as airports, intersections, runways, and routes. However, the prirnaty function of the FMS

is accurate short- and long-range lateral and vertical navigation. Although the FMS interfaces

with a variety of short-range and long-range sensors, the sensors themselves are not part of the

FMS. The FMS provides lateral and vertical navigation guidance for display and coupling to the

DIFCS. The CD-800/CD-81 O Control Display Unit (CDU) is the primary means for pilot interface

with the system.

The Navigation Computer can intefface with three long-range sensors, one via an ARINC 429

bus and two over the ASCB bus. Each Navigation Computer can also connect to dual Collins

Proline 2 or Bendix/King DME Receivers and a single VOR Receiver. The intetface to the

AHRS, Air Data, MFD, EFIS, and DIFCS is over the Avionics Standard Communications Bus

(ASCB). Flight plans are also transferred between Navigation Computers over the ASCB, while

the link to the CDU is over an RS-422 private-line interface. To provide high accuracy long-

range navigation, the Navigation Computer is designed to connect to AHRS, Omega/VLF

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8/11/2019 spz-8000 difcs


2. K.

TCAS II (Optional)

The following components comprise TCAS 11:

. RT-91 O TCAS Computer Unit (CU)

.

AT-9 10 TCAS Directional Antenna(s)

XS-91 O Dual Diversity Mode S Transponder

.

DV-91 O Vertical Speed Indicator/Traffic and Resolution Advisory

(VSI/TRA) Display

.

Parl of the two RM-850 Radio Management Units

.

Parl of the aircraft audio system

NOTE- The trans rider, the RMUS, and the audio system are pafi of the

‘“ (PRIMU

8

11)SRZ-850 Integrated Radio System.

TCAS is designed to act as a backup to the Air Traffic Control (ATC) system and the “see and

avoid” concept. TCAS computes closure rate and altitude of all transponder-equipped aircraft in

the surrounding airspace. Surveillance volume is defined by a minimum horizontal radius of 14

nautical miles, and a minimum vertical range of *12,700 feet. TCAS continually interrogates

transponders in that airspace, processes their replies, and tracks their fliihtpaths. Flightpaths

that are predicted to penetrate a collision area surrounding the TCAS aircraft are annunciated by

TCAS. The physical dimensions of the collision area are time-based and vary as a function of

horizontal and veftical closure speeds (Range Rate and Altitude Rate) and horizontal and

vettical distances (Range and Altitude) between the TCAS aircraft and the intruder aircraft.

TCAS operational displays are divided into two distinct advisories:

8/11/2019 spz-8000 difcs


2. L.

Global Positioning System (Optional)

The Global Positioning System (GPS) tracks a minimum of four satel li tes, processes the

received signals, and determines the system latitude, longitude, altitude, time and velocity.

When less than four satel li tes remain trackable, the system uses inertial information from the

IRS and air data computers to continue determinatiin of position. When a fourth satellite is

acquired, the system revetts to normal tracking mode. The GPS mnsists of the following

components:

Global Positioning Sensor System Unit (GPSSU)

.

AT-8 10 GPS Antenna

M. LASERE@ Ill Inertial Reference System (Optional)

The IRS is an all attitude inertial sensor system. Typical installation is normalfy a dual system

configuration comprising the following components:

.

Two Inertial Reference Units and Mounting Trays

.

Two Mode Select Units

24 V dc battery backup (not supplied)

The LASEREl@ Il l IRS senses movement and rotations using inertial accelerometer sensors

and laser ring gyros (within the IRU). From this information, the system calculates present

aircraft position, velocity, heading, and atti tude. The IRS then outputs this information digitally to

the Dual Flight Control System, Weather Radar System, DADC, FMS, EFIS, and MFD System (if

installed).

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8/11/2019 spz-8000 difcs


FZ-800

BUS

CC) NTROLLER

(ACTIVE)

$

AFCS 1

10Hz

AHRS 1

40Hz

B

AFCS 2

10Hz

8/11/2019 spz-8000 difcs


All other units such as Digital Air Data Computers, Attitude and Heading Reference Systems,

etc, that are connected to the ASCB are defined as “users” or “subsystems.” The bus users and

the bus control lers are al l transformer coupled and impedance-matched to the data and clock

transmission lines. Data transmitted onto the bus drives one line more positive, and the other

line more negative. This interface method provides protection from faults, transients, and RF

interference. By design, the ASCB interfaces are virtually immune to lightning-induced

transients, hot shorts, ground shorts, and RF threats. The design precludes any fault

propagation between the bus and various interconnected users and/or bus control lers At the

same time, the ASCB interconnect structure provides superior RF emissions characteristics,

ensuring that ASCB wil l not interfere with sensitive receivers onboad the aircraft.

The users and bus controllers are connected to the buses via a splicing arrangement (using

solder rings). Figure 1-14 shows the network for a standard SPZ-8000 system configuration,

Figure 1-15 shows the network for a system that has an optional MFD symbol generator (SG)

installed. Other network options (in addition to the MFD SG), include an FMS NAV computer

and LASERE@ IRU, as shown in figure 1-16.

The ASCB also has private-line networks with AHRS and the EFIS/MFD symbol generators.

Figure 1-17 shows the priiate-line netwofi for AHRS and the EFIS symbol generators, and

figure 1-18 shows the private-line network for a system that includes an optional MFD SG.

Physical characteristics common to all of the ASCB networks are listed as fol lows

There are two independent ASCBS denoted “A” and “B,” each consisting of two

wire pairs denoted “Data” and “Clock.”

The ASCB transmission lines are Raychem2524E0114 with a therrnorad

jacket.

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8/11/2019 spz-8000 difcs


c

(0

m

o

0

-.

w

m

w

N

Cb

o

0

0

PILOT’S FMSNAv

121J1A

C121J1B

COPILOTS FMS NAV

{

{

L>{;c

;

H

e

I

26

H

1 z 1

DATA

DATA

c

9

27 c

I

,

&

}

“B”

..A..

{

->;; 1

<29

CLOCK H

12

+

}

H CLOCK

c

13

<30+

c

121.IIB

{

{

z;:~

C121J1A

H

26

1

<8

,

I

g

1

H DATA

DATA

c

27

<9

c

A

1

1

}

“A-

..B..

{

u;;

29

<12

CLOCK ;

z

}

H CLOCK

30

<13

c

9J1A

C9J1B

PILOTS DADC

COflLOrS DADC

{

(

+>)1

H

11

I

<13

H

DATA

I

z

1

OATA

c

12

<14

c

T

I

,

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““B”

..A..

{

*P

24 15

CLOCK ;

~ }

25

H CLOCK

<16

c

9J1B

C9J1A

.{

{

n;~

H

13

I

I

1

<11

DATA

I

1

1

z )

c

14

H DATA

A

I

I

12

c

I

v

}

“A-

“0”

{

G]’

15

<24

CLOCK H $= 1

c

H CLOCK

16

<25

c

8/11/2019 spz-8000 difcs


PILOT’S AHRS

I

lJIB

{

H

m

DATA

c

‘6A”

{

I

H

CLOCK

c

I

64>,

65

1

‘1

66>,

67>1

———

(~

- -) f“ -—–—

—-—

—-—4

PfLOT’S SG

:;q “

65J1A

[

{

H

75

DATA

c

76

“A”

{ ‘)

I

I

H 78

CLOCK

c

79~

J

CIJIB

~64*

1

>67 >~

_——

)( ‘––

‘)(-

- -)

———

———

.

4

4 *

COPILOTS SG

C65J1A

i

)75

I

D }

H

DATA

>76

c

I

I

;,’ }1

“A”’

>78

H

I

CLOCK

)79

c

COPILOT’S

AHRS

}

H

DATA

c

1

A’

H

CLOCK

c

1

I

AD-29772

8/11/2019 spz-8000 difcs


PfLOT% AHRS ]

I

lJIB

{

{

t-l

DATA

c

A’”

{

H

CLOCK

c

64

65

%

I

66

67

{

H

75

[ F+

DATA

c

76

A’

{

I I

H

78

CLOCK

c

79

—

—

)

(,

*

—

—

,_—-

--—.

H

DATA

-——-

--—.

CIJIB

F

q

1

“ “ DATA

65

}

1

I

}

.A.

H

+: c cLOcK

[:

—

(

——-.

-——

D

I

:)

I COPILOT’SSG

q}

C65J1A

}

-+75 H

DATA

+76 C

)

+78 ‘ H

“A”

+7, c CLOCK

L

8/11/2019 spz-8000 difcs


Figure 1-19 illustrates an example of a typical user subsystem. It shows a

the DADC. The defined message content is shown in the box to the right.

user address defined for

Other data in front and in

back of the actual data is control and error checking information required in all user messages.

Forty times per second (every 25 ms) the active bus controller begins a series of interrogations of the

users on the ASCB. Each 25-rns time block is known as a “Frame.” There are a total of eight

different frames defined, with different groups of subsystems transmitting in each frame. Some

subsystems will reply in each and every frame, some will reply in alternate frames, some in every

fourth frame, and some only every eighth frame. This allows update rates of 40, 20, 10, and 5 times

per second or slower. Individual subsystem requirements dictate a 40 Hz update rate for AHRS and

10 Hz for the AFCS, air data, and EFIS. Refer to table 1-4 for specific frame content.

WRu%w{p+-.

FLAG RESPONSE

ADDRESS 06

t

DADC

TRANSMISSION

PRESSURE ALTITUDE

BARO ALTITUDE

ALTITUDE RATE

INDICATED AIRSPEED

TRUE AIRSPEED

MACH

c

8/11/2019 spz-8000 difcs


3

0

I

FRAnEo FRAKEl FRAME2

FRAnE3

FRAKS5

FRAUS6 FRAIIE7

PM

SOTS

BOTH

START

CONTROL

FTIU

AFCS,L

START

CONTROL

START

~TART

iWrROL

START

CONTROL

SPARE,2

START

START

CONTROL CONTROL

START

CONTROL

BUS A

BUS B

PMC-P,L Ft4CS,*

FIKS,L

PHC-P,MCS,R

iPARE1

FMCS,R

KFo

AFCS,R

BUS A

BUS B

AFCS,L

FD

AFCS,R

SPARS,3

DADC,L

ANRS,L

ANRS,R

EFIS,L

DADC,R

BUS A

BUS B

EFIS,L DADC,L

DADC,R

ANRS,L

AHRS,R

XFER EFIS

EFIS,R

EFIS,R

ANRS,L

ARRS,R

XFEREFIS

BUS

A

BUS B

ANRS,L

ANRS,R

XFI?REFIS

UiRS,

ANRS,R

ASRS,L

ANRS,R

XFEREFIS

ABRS,L

ARRS,L

ANRS,R

ARRS,R

XFER

EFIs

EFIsXFER

EFIS,L

EFIS, R

KFD

FMCS,L

PMc-P,L

NAVW,L

XPER

EFIS

XFERBFISors

FER

EFIS

BOTN

EFIS

XFER

SFIS ~=

EFIS,L

EFIS,R

MFD

EFISXFER

EFIS,L

EFIS,R

nFD

EFIS XFER IEFIsxmt EFIS XFER

EFIS,L EFIS,L

EFIS,R EFIS,R

BUS A

BUS B

EFIS,L

EFIS,R

FIS,

EFIS,R

EFIS,L

EFIS,R

UAIL

HASL

BUS A

BUS B

UFO

Mm

FD

MFLl

HAIL

BUS A

BUS B

FMCS,L

FMcS,L

F14CS,

FMCS,R

MCS,R

PK-P,R

FMS,R

PMC-P,R

P?CS,R

PUC-P,R

BIWD

BUS A

BUS B

=.

=

m

v

+

o

5.

m

P14C-PL P 4c-P,

IPUC-P,L

PMC-P,R

BRGD

BltGD

BUS A

BUS B

BUS

A

BUS B

NAV CDU,L

NAV

m4, L

NAV CDU,R

NAV RE?4,R

CDKM CDU,L

COF04CDU,R

NAV RD4,L

NAV l@l,R

:UIRI REt4,

CCS4MRF24,

NAV IW?4,

NAVR124,R

C(M4

REM,L

CCM REM,R

I

NAVRD4,R

1

*

opt

10M1

8/11/2019 spz-8000 difcs


When the active bus controller starts a frame, it does so with two short messages; a FRAME

START message, and a CONTRO~EST message. The FRAME START message is simply a

‘wake-up’ call to all users, announcing that a new frame is starting. The CONTROUTEST

message is resetved for functions such as identifying frame number (1, 2, 3, or 4 in this

example) and controlling maintenance test activity. Following the CONTRO~EST message,

the bus controller requests all users to transmit for that patiicular frame.

Figure 1-20 illustrates typical bus requests and responses. Following the FRAME START and

CONTROUTEST, a request for AHRS 1 is transmitted on both buses. AHRS 1 responds with its

data on Bus A. AHRS 2 request is transmitted on both buses. AHRS 2 responds with data on

Bus B. This process continues as shown in figure 1-20 until all subsystems have transmitted

their messages. Both buses then go inactive until the beginning of the nexl bus frame.

The bus controller repetitively transmits user subsystem requests at the proper times,

independent of whether the subsystems actually respond with their data messages. User

subsystems need not all be in existence on the bus. Requests may be transmitted for

subsystems that are optional and not installed in a particular application. The bus controller

database defines the length of each user message so that the bus controller may request

transmission at the proper times, independent of responses. Table 14 shows the complement

of subsystems requested to transmit in each of eight sequential frames. After frame seven is

complete, the sequence repeats, stafting again with frame zero.

In Version A of the bus (table 1-4), control tasks are alternated between the FZ-800 and the

SG-816. Primary bus control is contained in the FZ-800, with secondary bus control

accomplished by the SG-816.

The FZ-800 has control of the bus for approximatefy the first 7.2 ms. After this, the FZ-800

8/11/2019 spz-8000 difcs


BUS A ACTIVITY

m

-

IE KREJ

AHRS 1 MSG.

-

*

AFCS 1 MSG.

BUS B ACTIVITY

I

FRAME START

I

I CONTROIJTEST I

AHRS 2 MSG.

-

8/11/2019 spz-8000 difcs


To be compatible with the transformer-coupled data bus, all ASCB messages are Manchester II

encoded before being applied to the bus. Unlike Non-Return-to-Zero (NRZ) data, which requires

a bandwidth of dc to fC(clock frequency), Manchester encoded data is limited to the frequency

range of f~2 to f=, Also, since Manchester data must transition in the middle of each bti period,

the data clock is contained within the data and is easily extracted at each receiver for data

decoding. This feature avoids having to send a synchronous clock on sepamte lines along with

the data. Manchester II encoding is illustrated in figure 1-21.

Referring to the timing diagram (figure 1-21), the NRZ data are encoded into Manchester II

format by a Manchester II ertcoder/decoder chip. The phase relationship is as defined in the

timing diagram. The clock frequency (f~ is 667 MHz. To reduce timing problems associated

with data skew, jitter, and settling time, the circuit device providing the Manchester II

encoder/decoder element with data uses the trailing edge of the transmit clock for its data

shifting function. In the receiver mode, the encoder/decoder chip provides the next circuit device

with the NRZ data and a properly phased clock for shifting the data into the system. Again, the

nexl circuit device uses the same trailing clock transition as that used by the encoder/decoder

for data transitions.

TRANSMITTER d

ENABLED

,, ,,

NRZDATA

~

CLOCK

8/11/2019 spz-8000 difcs


8/11/2019 spz-8000 difcs


Figure 1-24 shows the start of data decoding as a result of the receiver recognizing the 1-1/2

HIGH followed by the 1-1/2 LOW after the series of ZEROS. This shows the ASCB bus data

waveform as seen on the bus at approximately 5 volts peak-to-peak. This amplitude indicates

that there is no load on the bus whatsoever (open circuit). Typical waveform amplitudes are

between 3 and 4 volts peak-to-peak, and are dependent upon the actual number of users that

are connected to the bus. Since all users are essentially connected in parallel, more users lowe

the bus impedance, and consequently, the data waveform amplitudes. Amplitudes below 2.5

volts peak-to-peak indicate an abnormally low impedance or abnormally low resistance

somewhere on the bus.

o

1

I

S13

00101

1

MSB

- —

\\

hi-////

111111

1

1111118 1

1111

‘ 1

81

I

I

Ill I

I

111 I

Ii

Ill

. . . .. .

II

II

8/11/2019 spz-8000 difcs


3. B. Radio System Bus (RSB)

The Honeywell Radio System Bus (RSB) is the principal communications network

interconnecting the LRUS in the SRZ-850 Integrated Radio System. All the LRUS in the radio

system are connected to the RSB. Reliable transfers of data via RSB are ensured by designed-

in redundancy and predefine protection and isolation mechanisms. Control and data protocols

are also predefine to ensure consistent application of the databus. It is a fail-operational

databus system, and actually consists of three shielded twisted pairs; denoted as the PRIMARY

bus, LEH-SIDE SECONDARY bus, and RIGHT-SIDE SECONDARY bus. Again, “Fail-

operational” means that if any device connected to the bus fai ls, the bus remains operational.

All units that are connected to the RSB, such as the RM-850 Radio Management Unit (RMU),

RNZ-f350 Integrated Navigation (NAV) Unit, RCZ-850 Integrated Communciation (COMM) Unit

etc, are defined as users. The RSB bus users are all transformer-coupled and irnpedance-

matched to the databus transmission lines. The bus is a shielded-twisted-pair which is

differential ly driven. Data transmitted onto the bus drives one line more positive, and the other

l ine more negative. This interface method provides protection from faults, transients, and RF

interference. By design, the RSB interfaces are virtually immune to lightning-induced transients,

hot shorts, ground shorts, and RF threats. The design precludes any fault propagation (via

RSB) between the various interconnected users. At the same time, the RSB interconnect

structure provides superior RF emissions characteristics, ensuring that RSB will not intetfere with

sensitive receivers onboard the aircraft. The users are connected to the data buses via a

splicing arrangement (using solder rings) as shown in figure 1-11.

Data flow on RSB is bidirectional with a bit transmission rate of 667 kHz (1.5 pdbti). Data traffic

flow on RSB does not require a bus controller, All users receive and identify all bus data. Since

each user knows its own user number, it sets up an internal timer, based upon the last message

8/11/2019 spz-8000 difcs


As shown in figure 1-25, in the message Ot ime slot, the left side NAV Unit transmits on both the

PRIMARY and LEIT-SIDE SECONDARY buses. Then, in the message 1 time slot, the right

side NAV Unit transmits on both the PRIMARY and RIGHT-SIDE SECONDARY buses. Then

there is a spare time slot (message 2) for future expansion. Since some messages combine

data from more than one radio function, RMU, COM, Transponder, VOWLOC, Glideslope,

Marker, DME, ADF, and MLS require eight messages per system side. Left-side system = 8,

Right-side system =

8, and spare time slots = 8 more, totaling 24.

t--

192 MSEC PER FIELD

--1

MESSAGE NO.

012345

1819202122230

PRIMARY BUS

~~

LEFT-SIDESECONDARY

~~

RIGHT-SIDE SECONDARY

~~

EXPANSIONTIMESLOT

a=

NAV RMU

DATA

DATA

AD-34583@

RSB Data Field Structure

Figure 1-25

Honeywell

MAINTENANCE

MANUAL

CITATIONVll

8/11/2019 spz-8000 difcs


;CI:D

o,

1. 2 3, 4. 5 6, 7, 8

NAV

REM

RMU COM

NAV REM

9,

10, 11 12,

13.

14 ~;fi16t 17 & Z..h

20 21. F;:, 23

COM REM NAV REM

1 Low

MSG. NO.

I

MSG. NO.

I

14;:4Rg~.

1

High CONTROL CONTROL

4

Low

MLS

I

COM VOR/ILS

4 High AZ DEV

PRESET

BRGILOC DEV

5 Low MLS

I

ATC LEFT

I

VORIILS

5 High GP DEV

OP MODE

GS OEV

6 LOW OME OIST

ATC LEFT

VOR/ILS

6 High RT-SIDE REPLY COOE MARKER

7 Low

ATC RIGHT

DME DIST

7 High OP MODE

DME DIST

RT-SIDE

8 LOU

PRESET

FMS ma

ATC RIGHT

8 High

REPLY CODE

9 Low

ONE OIST

ATC/TCAS

LFT-SIOE

9 High DME DIST OP MODE PRESET

F t s ..bfl

O Low

ATC/TCAS

DME STATUS

O High

ALT/RANGE

R-S PRESET

DME DIST

1 Low

LFT-SIOE COM STRAPS OME CHAN

1 High WORD 1

R-S PRESET

2 Low DME STATUS ::I)OS;RAPS DME GS

2 High LFT-SIDE R-S PRESET

*

MSG. NO. MSG. NO.

CONTROL CONTROL

COM MLS

STATUS

OUTPUT AZ

COM

MLS

CHAN

OUTPUT GP

COM

I

MLS

PRESET

AZ DEV

*

I

--i

ATC

STATUS

:;: D:ST

ATC

,, ,,

REPLY CODE

4

TC

ALTITUDE

OME DIST

ATC LFT-SIDE

-- . . . . . . --

I

ATC/TCAS

~~E TUS

STATUS

ATC/TCAS :14:54;:N

ALT/RANGE

I [

ADF ADF

ATC

CHAN MAG ERG

REPLY CODE

ADF VOR/ILS

PRESET

MISC.

BRG/LOC OEV STATUS

I

%+%--RN

VOR/ILS

I

~t4:

:ATUS VHF COM

PRESET

, ,,

CHANNEL

I

1

I

MLS

DME TTS

VOR-DME

OP MODE

FMS ma,,

OPMODE

MLS

I

:::;:~TUS VOR-DME

CHAN CHANNEL

I

MLS FWO.

::: $; N

MLS-DME –

SEL. AZ

OPMODE

M~LS5S;~wDGP

DUE GS

MLS-OME

FMS . .~?.

CHANNEL

8/11/2019 spz-8000 difcs


WORO

o.

1, 2

3, 4, 5

Pos

NAV

REM

RMU COM

6, 7, 8 9,

10. 11 12, 13, 14 15, 16. 17 18, 19, 20 21. 22, 23

NAV REM

COM REM

NAV REM RMU NAV NAV REM FMS/RMU

%%%-1--

18

LOW VOR/ILS

18 High STATUS

k

2

Low VOR/ILS

I

AUX1

22 High IOENT

OP MOOE

26 LOW NAV CLUSTER AUX2

26 High

STRAPS

OPMOOE

OME STATUS AUX1 DME IOENT

MLS-DME MLS

~Ms , b .

L-S PRESET

. . . ---- --

RT-SIDE CHAN CHAN

OME OPMOOE

OME CHAN AUX1 ~~:);:fNT

NAV STRAPS

m:

[ye.

~~s .~?.

L-S PRESET

------ ---

WORO 1

OME CHAN

OME GS

AUX2 MLS AUX NAV STRAPS MLS SEL. GP MLS

L-S PRESET STATUS OATA MORO 1 WORO 2 MLS GSTATUS OPMODE

OME TTS

AUX2 MLS AUX NAV STRAPS

)44; 8:;D. MLS

L-S PRESET --------- OATA WORD 1

WORO 3

. .

CHANNEL

DME IDENT AUX2 MLS AUX NAV STRAPS

MLS BASIC MLS

L-S PRESET

--------- DATA WORO 2

UORO 4

1.3,4,5,6 FORW/BACK

AZIMUTH

OME IOENT CO~T:~J:TER MLS AUX AHRS-A429

L-S PRESET

MLS 8ASIC

DATA UORO 2 NAV HEAOING 1,3,4,5,6 MLS GP

AOF ATC MLS AUX MLS 8ASIC

STATUS

AOF

CONFIG OATA WORO 3 WORO 2

OPMOOE

AOF ATC MLS AUX

CHAN

MLS BASIC

AOF

CONFIG

OATA WORO 3

WORD 2

CHANNEL

ADF ATC MLS AUX

PRESET

MLS C:~MCI~:STEF

CONFIG OATA WORD 4 GEN OATA

AOF C:~A3~:STER

MLS AUX

IOENT

MLS N:JMN&STEF

OATA WORD 4

GEN OATA

AOF

COM CLUSTER

IOENT STRAPS

COM CLUSTER SYSTEM

STRAPS

ONIOFF

CO: TMCj;:TER POST SYS

POST RAOIOS

3.

c.

Digital Audio Bus

8/11/2019 spz-8000 difcs


The audio bus network shown in figure 1-26 is a dual system configuration, with system No. 1

being the pilot’s side and system No. 2 being the copilot’s side. Digitizing the audio offers the

advantage of complete independence from grounding problems within the aircraft and the

absolute elimination of ground noise pick-up, whine and cross-talk.

Each side has a “One-Way” digital audio bus, consisting of a differential ly driven, shielded

twisted-pair. Data transmitted onto the bus drives one line more positive, and the other line

more negative. This interface method provides protection from faults, transients, and RF

interference. By design, the interfaces are virtually immune to lightning-induced transients, hot

shorts, ground shorts, and RF threats. The design precludes any fault propagation (via digital

audio bus) between the various interconnected users. At the same time, the digital audio bus

interconnect structure provides superior RF emissions characteristics, ensuring that the digital

audio bus will not interfere with sensitive receivers onboard the aircraft. The users are

connected to the data buses via a splicing arrangement (using solder rings) which experience

has shown to be extremely reliable and damage resistant. The type of cable that is specified for

use meets regulatory guidel ines for flammabWy and smoke, and is resistant to hydraulic fluids

and fuel.

Each remote LRU contains a Cluster Module, which, in turn, contains five digitizer chips. These

are standard “Off-The-Shelf” chips (called CODECS - for COder/DECoder) that are used by most

telephone companies. The five digitizers are sampled in sequence, their digital outputs are

assembled into a digital data message, and the message is transmitted on the digital audio bus.

The remote COM Units provide digitized COM receive audio, and the remote NAV Units provide

digitized VOWLOC, ADF, and MARKER BEACON audio. The NAV Units also feed discrete

8/11/2019 spz-8000 difcs


mm

f

——

PILOT AUDIO PANEL

--

‘~ h

COPILOT AUDIO PANEL

AUDIO

{

+ 160J1-4 .

T

;%l

II

. C160J1-~ +

}

AUDIO

BUS 1

- 160J1-p

.

—

“+ , ‘

‘Y

.

C160J1 -p -

BUS 1

—

—-

-—

J

II

4

160J1 -N ‘—

——

AUDIO

{

160J1 -q

BUS 2

160J1 -g

PILOT COMM

I

{

+ 143J1 -56 ‘1’

AUDIO

m

BUS 1

II

- 143J170

II

4

—-

143JI -42

——

I

II

A

\J

—

.-

--

—

t-i

C160J1-N

COPILOT COMM

I

q}

r~

C143J1-56

d

AUDIO

II

BUS2

Y

C143J1-70

I

4

C143JI-42

NOTE: The Digital Audio Bus is very similar to both the RSB and ASCB

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described earlier in this section, except the clock frequency is 1 MHz

instead of 667 kHz, and the data bit assignments are different. Refer

to the explanation associated with figures 1-21 thru 1-24.

As shown in figure 1-27, in each transmitted message, the Preamble consists of 8 + 1

Manchester one bits; and the sync consists of 1-1/2 bits of “HIGH” followed by 1-1/2 bits of

“LOW, which the receiver uses for synchronization. The remaining six bytes contain eight bits

each, at 1.0 @it. The Status byte identifies the message as COM or NAV. The digital audio

panel then decodes and processes the individual bytes as appropriate to the flight crew

selections.

r’28’sEc7

c

I

N

II

c

DATA BUS

llNl,~

AMBLE

NAvMEs~AG. ~ 8BITS.CHWORD

PRE SYNC STATUS VOFl

ADF MARKERAUX1 AUX2

AMBLE LOC

AD-345s4@

Bits 9 and 10 may also be used as data bits in high resolution data words. Bits 11 thru 29

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compose the data field. Bit 11 is the least significant bd (LSB), and bit 29 is the most significant

bit (MSB). In most cases, bits 30 and 31 form the Sign Status Matrix (SSM), which identifies the

sign and validity of the data. Like bits 9 and 10 above, bits 30 and 31 may also be used as data

btis in high resolution data words. Bit 32 is used for parity.

In the Octal Label, bds 1 thru 8 are used to represent numbers Othru 377. The eight btis are

broken into two groups of three and one group of two, as shown in figure 1-28. Each group

represents a digit encoded in binary with the least significant bd (LSB) having a value of one.

The Octal Label is transmitted with the most significant bit (MSB) of the most signif~ant digit

first. This “reversed label” characteristic is a legacy from past systems in which the octal coding

of the label field was, apparently, of no paflicular significance,

BIT NUMBER

BINARY VALUE

LSB

CHARACTER VALUE

<

8 7 6 5 4 3 2

1

1 2 4 1 2 4 1 2

0 0 1 1 1 1 0 1

4

7 2

MSB

AD-34565@

Octal Label 274

Figure 1-28

Units, ranges, resolution, refresh rate, and number of significant bits for information transferred

are encoded in either Binafy Coded Decimal (BCD) or Binary (BNR) (two’s con@ement)

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29

I

I

I

MSB

DATA

LSB

AD-34566@

Data Bits 11 thru 29

Figure 1-29

If bits 11 thru 29 contain data bits in a Binary Coded Decimal (BCD) format (see figure 1-30), the

data is grouped into four bit bytes, each byte denoting a decimal column. The 19 data bits are

broken up into four groups of four bits and one group of three bfls. Each group of four can

represent a number from Oto 9; the ffih group can represent a number from Oto 7. Refer to

the following examples of BCD data fields. Data bit number 11 (the eleventh bti transmitted in a

word) has the binary value of 1. Data bits numbered 12, 13, and 14 have the binary value of 2,

4, and 8 respectively. Each group of bits 15 thru 29 have similarly assigned values as shown

below. Using ths format, decimal numbers (or characters) between O and 9 can be assembled

using combinations of these four binary values.

29

28 27 26

25 24 23

22 21 20 19 18

17 16 15 14 13 12

11

4 2 1 8 4 2 1

8 4 2 1 8 4 2 1 8 4 2

1

MSB

DATA

LSB

BCD Bit Assignments

AD-34567@

Figure 1-30

Figure 1-32 shows an example of a DME data word that requires five characters.

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Parameter:DME Distance OctalLabel:201

Value:257S6 NM

29

28

27 26

25

24 23

22

21

a 19 18

17

16 15

14

13 12 11

0 1

0 0 1

0 1 ~ 1

1 1 1 0

0 0 0

1 1 0

2

5

7

8

6

AD-34569@

Five-Character DME Word

Figure 1-32

Figure 1-33 shows an example of position data words requiring six characters. As can be seen,

bits 9 and 10 are used, and the format is changed slightly.

Paramek PresentPos.Long.

octal Labef:011

ValuaE 175°59.9’

29

28 27

26

25 24

23 22 21 20

19 18 17 16 15 14 13 12 11 10 9

1 0

1 1 1 0

1 0 1 0

1 0 1

1 0 0 1

1 0 0 ‘

1

7 5 5

9

9

AtH4570@

Six-Character DME Word

Figure 1-33

For angular range, O thru 359.xxx degrees is encoded as O thru plus or minus 179.XXX degrees.

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The sign bits (30 and 31) determine the semicircle being referenced. The positive portion of the

semicircle includes O thru 179.XXX degrees. The negative portion includes 180 thru 359.xxx

degrees. An all zeros configuration represents O and 180 degrees. All ones represents 179.XXX

and 359.xxx degrees. Two’s complement notation is used for the negative half.

Pariiy is one of the simplest of all the error checking methods used in data handling. There are

two basic parii configurations, “ODD” and “EVEN.” ARINC 429 transmissions are always odd

parity, and bit 32 is the pariiy bit. ARINC 429 receivers are programmed to always expect an

odd number of ones in each 32-bit word. Bit 32 is set to one when there are an even number of

binay 1s in the word, and set to a zero when there are an odd number of binary 1s in the word.

This creates a word that always contains an overall odd number of 1s.

To be compatible with the transformer-coupled data bus, all ARINC 429 messages are

Manchester II encoded before being applied to the bus. Unlike Non-Return-to-Zero (NRZ) data,

which requires a bandwidth of dc to fC(clock frequency), Manchester encoded data is limited to

the frequency range of f~2 to fC. Also, since Manchester data must transition in the middle of

each bfl period, the data clock is contained within the data and is easily extracted at each

receiver for data decoding. This feature avoids having to send a synchronous clock on separate

lines along with the data. Figure 1-34 illustrates Manchester II encoding.

ARINC 429 transmissions return to the zero voltage condition at the end of each bd period. As

can be seen below, a high on Line A, and a low on Line B is a binary one. In addition, a low on

Line A, and a high on Line B is a binary zero, When both Line A and Line B are at zero volts,

there is no data bit being transmitted. ARINC 429 transmitters must provide a minimum dead

time of four bits between messages because the receivers synchronize to the transmitted data

by recognizing the four-bn dead time as the synchronizing command.

Tri-level bipolar modulation consisting of “HI” (binary one), “LO” (binary zero) and “NULL” (no

8/11/2019 spz-8000 difcs


data) states are used in the transmission of data. The differential output signal voltage across

the specified odput terminals (balanced to ground at the transmitter) should be as given in table

1-7, when the transmitter is open circuit: -

HI (1)

NULL (V)

LO (0)

Line A

to

+Iot 1.0

0 f 0.5

-10* 1.0

Line B

Line A

to +5 to 0.5

0 f o.25

.5 f ().5

Ground

Line B

to -5 * 0.5

0 f 0.25

+5 f O. j

Ground

Differential Output Voltages

Table 1-7

The differential voltage presented at the receiver is dependent upon line length and the number

of receivers connected to a transmitter. The nominal voltage range at the terminals is likely to

be between 6.5 and 13 volts peak-to-peak. Receiver input common mode voltages (Line A to

Ground and Line B to Ground) are not specified because of the dtilculties of defining ground

with any satisfactory degree of precision.

3. E.

Collins Commercial Standard Digital Bus (CSDB)

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There are several RS-422 pods on each SRZ-850 Radio Management Unit (RMU). These

RS-422 ports emulate Collins radios by outputting completely compatible Pro-Line II (PLII) di@al

bus information for all radm functions. For example, backup VOWLOC/GS/MKR navigation

display data is sent to the Radio Management Units from the No. 1 Navigation Unit on PLII

digital data buses.

The data format is in accordance with Collins Commercial Digital Bus (CSDB) standard

523-0772774-00611 R. This data bus is frequently referred to as the Collins Pro-Line II Serial

Data bus, or PLII.

The PLII bus system is made up of transmitters and receivers connected by shielded twisted

wire pairs. Data is transmitted by a single transmitter to either a single receiver or to a group of

up to 20 receivers connected in parallel. Each PLII bus carries data in one direction only.

Bid@ctional transmission between two LRUS must be accomplished by using two sets of

transmitters, receivers, and twisted wire pair buses.

F. RS-422 (Electrical Specification)

Strictly speaking, RS-422 is an electrical specification, as defined by Electronics Institute of

America (EIA). Nonetheless, the term RS-422 is used throughout this manual to describe any

data bus consisting of a shielded-twisted-pair that is not described so far in this section.

Examples are:

The bus that carries data from the GC-81 O Flight Guidance Panel Controller to the

FZ-800 Flight Guidance Computers.

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1.

SECTION 2

COMPONENT DESCRIPTION

General

This section provides an illustration, leading particulars, a brief description, and a bfock diagram or

schematic of each component used in the System. The information is only for the specific

components l isted in Section 1, table 1-1. When a component picture uti iiies cai louts, the description

for each callout is presented with upper left-hand caliout described first and proceeding clockwise

unless it is a minor item grouped with a major callout.

The components are separated into the following subsystems:

Subsystem

ParaaraDh

AHZ-600 Atti tude and Heading Reference System

2

ADZ-81O Air Data System

3


4

Reserved Subsystem Not Applicable to Citation WI Aircraft

5

EDZ-816 Electronic Flight Instrument System and Optional MDZ-816

6

Multifunction Display System

DFZ-800 Dual Fliiht Guidance System

7

2.

AHZ-600 Atti tude and Headina Reference System (AHRS)

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The SPZ-8000 System may include one or two AHZ-600 Attitude Heading Reference Systems

(AHRS). Each AHRS consists of one AH-600 Attitude Heading Reference Unit (AHRU), a CS-412

Dual Remote Compensation Valve, and one FX-600 Flux Valve. The AHRU measures the inertial

motion of the aircraft, the flux valve provides long-term magnetic heading information, and the AH RS

then computes attitude, magnetic heading, angular rates, and linear accelerations.

A. AH-600 Strapdown Attitude and Heading Reference Unit (AHRU) (See figures 2-1 thru 2-6, and

table 2-1.)

The AHRU contains the necessary power supplies, sensors, and electronics to compute attitude

and magnetic heading, and provides the necessary digital signals for the primary fl ight displays,

flight guidance, flight management, and other aircraft systems as rquired. The sensors within

the AHRU include fiber optic gyros, which sense angular motion around all three axes; and

accelerometers, which sense linear motion along all three axes. It is capable of 360-degree

displacement in the roll and heading axes, and M% degree displacement in the pitch axis.

The AHRU provides the excitation, current feedback control, and signal demodulation interfaces

for the dual remote compensator and flux valve.

Dimensions (maximum):

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Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..16.63 inches (427.5 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.91 inches (124.7 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..7.62 inches (193.5 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 13.51b(6.12 kg)

Power Requirements:

Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . .

+28Vdc, 60 Watts starl and 40 Watts run

Mating Connectors:

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

(Cannon) DPX2MA-67S-67S-33B-OOO0

Mounting Tray . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Honeywell Part No. 7004651

AH-600Strapdown AHRU

Leading Particulars

Table 2-1

The AHRSisprovided with twopower source inputs. Theprimary power forthe No. l AHRSis

the No. 1 avionics bus; for the No. 2 AHRS it is the No. 2 avionics bus. The auxiliary power for

each AHRS is provided from a continuously charged standby battery pack. The AHRS standby

battery pack is controlled by the standby vertical gyro ON-OFF switch. Separate circuit breakers

are provided for each of these power circuits. AHRS shutdown in flight due to power load or bus

switching transients is prevented by automatic power transfer within the AHRS to the auxiliary

The AH-600 also computes true airspeed (TAS) as a monitor function to verify the

reasonableness of the data received. Under a no failure condition, the AHRS prima~

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TAS data is received from the on-side ASCB TAS. If this source is unavailable or is

unreasonable, the AHRS will automatically revert to the cross-side ASCB TAS. If this

source is unavailable or is unreasonable, the AHRS will automatically revert to the IAS

input. The AHRS also automatically reverts to using indicated airspeed (IAS) data when

the TAS is less than 150 knots.

In the slaved mode, the difference between the indicated heading and the flux valve

heading is displayed on the slave error indicator (heading sync indicator) located on the

EHSI. The card has two symbols: a cross (+) and a dot (D). During straight and level

flight the indicator will generally be centered with excursions toward the cross or dot

occurring over a 20- to 30-second time period. This activity is normal and indicative of

good magnetic heading data. In turns, the display can show a steady dot or cross.

Following return to straight and level flight, the indicator will return to the centered

condition within 2 minutes.

The verticality of the AHRS may be checked or corrected during unaccelerated flight by

pressing a remote vertical gyro FAST/NORM switch shown in figure 2-3, to FAST for a

minimum of 10 seconds. This causes the ATT flag to be displayed on the EADI, the

autopilot to disengage, and the flight director modes to reset. Upon releasing the

FAST/NORM switch, the ATT flag will clear, and pitch and rolt attitude will become active.

The autopilot and flight director can be re-engaged at this time.

VERT GYRO 1 ~LH GYRO SLAVE=

FAST 1 MAN LH J

NOTES: 1. During basic mode, the vertical gyro FAST/NORM function should be

used frequently, in level unaccelerated fl ight, to correct for drift and

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acceleration errors.

2. On the ground, the on-side DADC must be supplying its own AHRS

true airspeed or the AHRS will enter the basic mode.

The DG mode disables the automatic slaving of the heading outputs. Entty into this mode

can only be achieved by momentarily pressing a remote GYRO SLAVE MAN/AUTO

pushbutton (figure 2-3). Entty into the DG mode occurs when the pushbutton is released

and is confirmed by the DG1 (DG2) annunciator on the EHSI. AHRS operation in the DG

mode results in a heading system similar to a free directional gyro, and which is subject to

drift and turn error. For this reason, AHRS operation in the DG mode results in reduced

heading accuracy. In the DG mode, the compass sync annunciator is removed on the

EHSI.

While in the DG mode, the heading card can be manually set to any heading using a

remote GYRO SLAVE LH/RH switch (figure 2-3). The control is inactive in the slaved

mode. When the switch is being used, the EHSI will display the HDG fail flag. The

switch will automatical ly center to the OFF position when released.

Upon exit from the DG mode, the AHRS performs an automatic synchronization of the

heading outputs to the present flux valve magnetic heading. This feature can also be

used if a heading error should develop. While in the slaved mode, the error can be

quickly removed by momentarily entering the DG mode and returning to the slaved mode.

This is performed by pressing the GYRO SLAVE MAN/AUTO button (figure 2-3).

2. A.

(4) Ground Initialization

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The AHRS requires approximately 3 minutes to initialize following application of power.

The initialization is complete when the AIT and HDG flags clear on the EADI and EHSI.

During the initialization, the aircraft must remain stationary. Wind gusts and aircraft

buffeting are not limiting in this respect, Similarly, all normal preflight operations, including

engine starts and passenger loading, may be carried out while the AHRS is initializing. If

the initialization requires more than 3 minutes, the AHRS could have detected excessive

aircraft motion. If aircraft movement has occurred during initialization, the AHRS must be

recycled and a new initialization commenced.

The initialization time-out can be observed if the vertical gyro FAST/NORM switch is

momentarily moved to FAST after power is applied and the AHRS AUX PWR and BASIC

annunciators extinguished. The EHSI heading card will slew to approximately south (180

degrees). The heading will decrease at the rate of 1 degree/second until the heading

card indicates nofih (O degrees). At this time, the 3-minute initialization period is complete

and all indications return to normal.

If the heading card stops and does not step to an indication of Odegree, the initialization

of that AHRS has not been completed in a satisfactory manner. The main and auxiliary

dc power to that AHRS should be removed by opening the appropriate circuit breakers

and then reapplied to restaft the initialization. Press the VG ERECT button and observe

the time-out sequence.

NOTE: It is necessary that both breakers (primary and AUX) be pulled out.

Resetting each breaker individually will not reset the AHRS.

2. B. AHRU Functional Description

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The fol lowing paragraphs describe the operation of the AH-600 AHRU with reference to block

diagram, figure 2-5.

(1)

Power Supply

The power switching function selects 28-volt input power from either of the two sources

that are usually tied to each side of the aircraft power distribution system. If the PWR 1

input suffers a transient or fails, the AHRS switches to PWR 2 if it is present and of

sufficient level to power the system. Energy storage within the AHRS Power Supply

Subsystem is sufficient to survive the bus switching interrupt without foming a

reinitialization or loss of data.

The power supply is protected for short circuits, transformer saturation, regulator loop

faults, and high- and low-line transients. The AHRS will reinitialize itseff i f a transient

condition causes the power inputs to simultaneously drop below 18 volts for more than

200 ms,

(2)

Flux Valve Drive

The flux valve drive provides 400 Hz excitation through the mmpensator to the flux valve

primary winding. The driie signal is generated by a dc-to-ac converter powered from the

28-volt aircraft ~wer to the AHRS. The drive frequency is derived from the 4831 Hz

system clock. An 800 Hz demodulator reference is also generated.

The compensator function is totally isolated from other circuit functions and derives its

2. B. (4)

Central Processor Unit (CPU)

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At the heart of the CPU are the microprocessors and their support functions. Two

processors are used in the design: a fixed point, 16-bit, general purpose processor, and

a floating point, math coprocessor. These two devices perform the control, logic, and

computational functions of the AHRS. Their 5 MHz clock speed is derived from a

dedicated 15 MHz crystal oscillator.

The control functions provide the system interface to the processors for interrupt control

(ASCB and Real-Time Clock), as well as wait state generation for insuring correct

address, data, and control signal timing (Ready). Processor and master resets are

generated in response to the power reset signal. Additionally, a time down signal voltage

is provided that is proportional to the length of a power interrupt. The processors are also

reset in response to an invalid heartbeat valid (HBV) signal.

Program code is stored in read-only memory (ROM). Random access memory (RAM)

provides nonvolatile storage for computational results. RAM access is disabled on power

interrupts and must be re-enabled by the processor when power is restored.

The processor control, address, and data buses are confined to the CPU to prevent

external faults from crippling the program execution. Access to the system buses for

input/output and ancillary functions is through the system bus drivers. The system control,

address, and data buses provide CPU access to the hardware input/output (1/0) functions.

The system buses also provide CPU access to the IMU ROM, which contains calibration

coefficients for the sensor elements.

The processors themselves are continuously checked for proper operation by performing

IMU

-CALU3RAT10N MEMorw

—M B ll “~E SS OR

-PITCH

-NARHE MONITORS

I

w -3

POWERSUPPLV

– WI(

SVTES(3MMEM

JIA

–ROLL

–EXCITATIONREP

–LINE FILTER

-4K BvTESPwRINTERRUF7ABLERAM

–VAW

-ACCEL TEMPS

–ISOLATION

–M2CELERATION

3

-HAR~flE MONITOR

-FIEW1.ATEDGC

-GvRO TEMPS

- SO- RE MO NI TOR

–REFVOLTAGES

4P

JIB

-MONITOR

5

VALIO

JIA

INPUTMTA

CONVERTER

—ATOO

–OISCRETE

INPuTS

. OUTPUT OATA

CONVERTER

–O TO A

–OISCRETE OUTPUTS

–MONITORS

J18

8

9

m

}

P IT CH S VNCHRO NO

1

8/11/2019 spz-8000 difcs


30 ‘ 1

ANALOG

El

IRSPEEO 31

32

TAS V ALIO 33<

I

I

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F AS T E RR EC T C M O 1 5

M AI NT T ES T C MO 1 6

a

MOOEELECTCOMMON 17

( {G]

2 SVGC C ON TRO L 2 1 ;

-R

23

LAMPTESTINPUT 23<1

I

I

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ERIAL

GAT A 45

INPUT

~;

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{ 18

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Ml 51

SWJ2TEM

AOOREBS

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21

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JIB

{

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ATT

REF INPUT 21

26 VAC. 400 Hz 22 L

t--

-MONITORS

4

ANALOG ANO OIGITAL

$

ENOAROUNOS

——

—.

h

DIGITAL BUSINTERFACE

—ASCB BUS CONTROLLER

–MONITORING

52

53

54

1

s

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w

51

}

PITCHSVNCt-iRONO 2

ROLLSVNCHRONO 1

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~ 41 ALIXPOWERANN

+42 MAINTTESTMOIJEACKNOWLEOGE

-+

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BuS N O 1

I I I I

/

1

I

1

I

A SC B B uS N O 2

ATTITUOEUNCTIONS

11

EFISRIvATEIUS

AH-600 Strapdown AHRS Block Diagram

F igure 2- 5 (Sheet 1 o f 2 )

Pages 109/1 10

Ati Ru OAT ABUS

IMU

-CALIBRATIONMEMORY

CPU

-PITCH

I

‘m%

‘ALO “m;

–16 BIT“ I%wcESSOR

-MARWE MONIT ORS

–ROLL –EXCITATIONREF

AHRUAODRESSI

—VAW

-ACCEL TEMPS

CONTROL BUS

–ACCELERATION

–GYRO TEMPS

JIB

OUTPUT DATA

(x]

CONVERTER

30

(Y)

–OTOA

31

$ }

HOG~t40 1

–DISCRETEOUTPUTS

(z)

JIA

32

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(x)

40

INPUT MTA

CONVERTER

–A TO D

–DISCRETE

INPuTS

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24

VOCN0 1

PwRGN13

2 8 VO CN0 2

8/11/2019 spz-8000 difcs


28Voc

SOUAllE WAVE

FLUX VALVEEXC

(

FLUX VALVE COMMON

FLUX VALVE INPUT

{

FLUXvALVESHIELD

MANUAL

HOG SLEW

COOING

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00 MOOESELECTNPUT

18 <

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(z) ::

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–FV GRIVE

–FV CURRENT

SERVO

A

ANALOG ANO OIGITAL

~

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w

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4

v

I

DIGITAL BUS INTERFACE

—ASCO BUS CONTROLLER

-MONITORING

Y

3 l NSTR UL 4~ H OG V ALI O

37 AUX H OG W

P

+ia

H 24

}

HOG RAIE

I

2 5 HOG RAT E W

13 OVER T EW~N ANN

1

1

SIGNAA

GNO

2

6

W&Is

7

b}

IA

H

12

WI

SLAVING ~

13

l+> 37 WVIN.FA4LUREANN

L

1

\

A SC O B US N o I

I I I I

I I

I

I

ASC O B us No 2

tiEADiNOWCTIONS

I EFIS nWATEuS

AH-600 Strapdown AHRS Block Diagram

Figure 2-5 (Sheet 2)

22=05-07

Pages 111/112

Jun 1/93

8/11/2019 spz-8000 difcs


2. B.

(6) Analog and Discrete Outputs

8/11/2019 spz-8000 difcs


Although the AHZ-600 is principal ly digital in nature, the unit is configured to provide

analog output signals to those auxiliary systems not participating on the ASCB bus. Two

pitch and two roll synchros, as well as a set of two-wire ac pitch and roll signals for the

weather radar antenna system, are provided with a common phase reference input. Two

heading synchros with a common phase reference input are also provided for use with an

RMI or other remote indicator. An additional heading synchro with a separate phase

reference input is also available. All synchro outputs are isolated by Scott-T transformers

and are capable of driving three 500-ohm synchro loads. Analog outputs of normal

acceleration and rate of turn are provided as dc voltage signals while slaving error is

available as a dc current for driving an ammeter or other indicator.

Discrete outputs are provided to indicate specific AHZ-600 operational modes and analog

data valid status. Several of these outputs are redundant with data available on the

ASCB and provide system protection from certain types of failure modes, Moreover, the

data valid discrete outputs are controlled asynchronously with respect to the digital buses

so that an invalid indication is annunciated immediately upon recognition of a failure.

(7) Monitoring

The AHRU performs extensive system self-checking during all modes of operation. These

monitors are implemented in both software and hardware and provide protection from

undetected multi-axis hardovers.

The majoriiy of monitoring is performed by the CPU under software control. Its internal

monitors provide memory checksums, data in-range and reasonableness checks; and

2. c.

CS-412 Dual Remote Compensator (See figures 2-7 and 2-8, and table 2-3.)

8/11/2019 spz-8000 difcs


The CS-412 Dual Remote Compensator minimizes system deviation caused by local magnetic

disturbances.

The Dual Remote Compensator insefts small dc voltages on the flux valve output to minimize

comoass svstem deviation caused by local magnetic disturbances from the airframe and the

ele%cal s~stems onboard the aircraft.

CS-412 Dual Remote Compensator

Figure 2-7

Honeywell

MAINTENANCE

MANUAL

CITATIONVll

SCREWDRIVER

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ADJUSTMENTS

Pw

TEST

(

– DC

OUTPUT + Dc

N-S TEST

E-W TEST

I

FROM

FLUX

VALVE

kohl

(x)

(Y)

FLUX VALVE

OUTPUT 10

AH-SW AHRU

F

I

I

A

,

c

f

s

T

v

A

I

-S,8VDC

4)

4

7

CURRENT

LIMITING

RESISTORS

I

I

SJl

I

I

~ ‘

I>K

&

Dc

BLOCKING

I

>

H

CAPACITORS

1

IJ

(E<

1

Z) AUXILIARY

(x)

FLUX

VALVE

(Y)) OUTPUT

CHASSISND

2. D. FX-600 Thin Flux Valve (See figures 2-9 and 2-10, and table 2-4.)

8/11/2019 spz-8000 difcs


The flux valve detects the magnitude and direction of the earth’s magnetic field and converts it to

electrical information that is used to align the AH-600 AHRU to magnetic north.

AD-32728@

FX-600 Thin Flux Valve

Figure 2-9


Length . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.40 inches (111 .76 mm)

8/11/2019 spz-8000 difcs


GRN (X) B101

\

I

A

~)

BLK (Y)

I

B

RED (Z)

I

p ~

,

-4

ma:}

I

WHT

>F

800 t-iz

SIGNAL

OUTPUT

28 VDC

SQUARE WAVE

STATOR

COMMON

AD-8661-R1

3.

ADZ-81 O Air Data System

8/11/2019 spz-8000 difcs


A. AZ-81 O Digital Air Data Computer (See figures 3-1 and 3-2, and table 3-1.)


Length (including handle) . . . . . . . . . . . . . . . . . . . . . . . . . ...15.76 inches (400.3 mm)

8/11/2019 spz-8000 difcs


Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3.59 inches (91.2 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.62 inches (193.5 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..9.71b(4.4kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 28 Vdc. l.l AMaxirnum

26VacRef,60mA

Mating Connector:

J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. DPX2MA-A106P-A1O6P-33B-OO24

Mating Pneumatic Connectors:

Pitot (straight) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40007-2B24*

Static (straight) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40007-2A26*

Pitot(90° elbow) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..40007-2B24E*

Static (90°elbow) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..40007-2A26E*

All part numbers are American Safety Flight Systems.

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

Tray, Barry Part No. 404A-38-S-l/DPX2-0

AZ-8 10 Digital Air Data Computer

Leading Particulars

Table 3-1

Internal

Update Rate

Parameter Units

Data Range

Times/See

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Pressure Altitude*

Baro Corrected Alt*

Altitude Rate*

Attitude Valid

Indicated Airspeed

IAS Valid

True Airspeed*

True Airspeed Valid

Total Air Temp

Temp Valid

Static Air Temp

Preselect Altitude

Crank-In-Motion

Preselect Alt Valid

VMO*

VMO Warning

Dynamic Pressure

DME Range

Feet

Feet

Ft/Min

--

Knots

. .

Knots

-.

‘c

.-

‘C

Feet

. .

-.

Knots

--

lnHg

Nm

-1000 to 60000

-1000 to 60000

-20000 to +20000

30 to 450

. .

30 to 599

. .

-50 to +99

-.

-99 to +50

O to 60000

. .

. .

30 to 450

0 to 22

.-

10

10

10

5

5

5

3

5

2

2

2

10

5

5

2

5

5

3

The DADC incorporates a self-test mode. When activated via a cockpit test switch, the DADC

outputs static data on the ASCB and ARINC 429 buses as specified in table 3-3. The

appropriate data will be displayed on the DS-1 25A TAS/Temperature Indicator and the EADI.

8/11/2019 spz-8000 difcs


Static air temperature (SAT), total air temperature (TAT), and true airspeed (TAS) are displayed

as dashes on the DS-1 25A TAS/Temperature Indicator for an invalid condition.

Parameter

Data Output

Pressure Altitude

Baro Altitude

Altitude Rate

IAS

TAS

Mach

TAT

SAT

Preselect Altitude

VMO

Dynamic Pressure

MMO

4000 Ft

Present Altitude

5000 Ft/Min

325 Kts (290 Kts*)

301 Kts

0.790 M

-16 “C

-45 “c

12,000

Ft

335 Kts (280 Kts*)

9.0 lnHg

0.80 M

JIB

(s<,

I

8/11/2019 spz-8000 difcs


18

IGITAL 10

RANGE ,1

Ps

“=~’ ~

Ps

FREOUENCY

SENSOR TEMP

SERIAL

DI;I’?AL

CONVERTER

PAR;?LEL

cONVERTER

[AR(NC 419,

[

DIGITAL BUS

BUS[TO SH2)

A

CHARACTERIZATION

PROM

‘T

PITOT

‘T

PRESSURE

‘T

-

SENSOR

sENsOR

TEMP

SENsOR TEMp

INTERFACE

{ f~

JIA

HOLD

60

pILOT

SLEW

HOLD

CONTROL ‘:

COPILOT

SLEW

INTERFACE

JIA

‘lr

BARO OISABLE 38

SSEC DISABLE 39

SELFTEST 52

TA TSAT SEL 7 ,

ALERTERSEL 93

PILOT,

COPILOTSEL lm

{

ALT 55

B

PILOT ALTRATE 56

HOLD

ENG

A,s 57

MA CH 58

, , . +4

ILOT SLEWENG 59

(s

LT 55

COPILOT ALTRATE %

HOLD

ENG

A/S 57

M4Cti 58

COPILOT SLEWENG 59

JIA

10I

102

AIRCRAFT ,OS

DISCRETE

3

ARITHMETIC

PROCESSOR

[s511)

=Pzl

MAINTENANCE

MANUAL

CITATIONVll

Jlh

WE”

3)

Q ) ~

AMTA

JIA

ll’ola

VNAV

[z

x

HI

,)

(0 ?= AM

OISPLAV

DATA

2XVAC

tNPUT

I

nEALcE

m

J1@

LO

OENEnATw

(w

lY

I

n

(cl

I

;)

BOATA

M

(Ml

I

s

(cl

1

I

;)

Bux

m

I

1

1

SWITCNEOMl

DIOITAL BUS

)- ALm LION-T

1

JIO

1

m

\

‘)1

=OL77Pu’

m

AAImc

~ Not

4s

m

m~

m

~ N02

JIEI

1

w

11

xx

V74AV

OWPUV

u

OUTPUT

RW31STER

E4

JiE

IASSEL@

I

“=”w’’’””o’

EE _

8/11/2019 spz-8000 difcs


77

D1OWAL

JIA

Xs

A&%0

FINE

L

*

n&oLLm

11

ii

‘1}

.Oos: n ~~

CONVERTER

nEsOtvEn

Corn; ~

M’nlvos

I

{“”-l

c wix7uY

JIA

4

IS

m

17

la

ALTITUOE

nEPonTINo m

RlEo19TEa

m

21

‘ ~

IIX5xow,,

VERTICAL

PATHER~

PRESELECT

ALT ERROR

D

II

JIA

a

v

1

INS

a

Zm

WNc44n0

X%?i%li

OUTPUTS

x

xx

}

vm~~

z

ALW77JOX

m

dlT(

Ull

—

—

Cw&u

CUTPUT

CONVERTER

,,. ~

1

w

47

IAS

la

MACM

4s

}

PILOT

OUTPUTS

4s

AAACW

COM

Ax

AIRSPEEO

OUTPUTS

*E*

Jlm

~:

1

OPILOT

4s

WTPUT9

COM

4s

JIA

?a IAS NO 3

JIB

~4

‘ “”ELmcvAL’o

TASAC

JIA

~ X41

SAAuLW

Cl%%s

H=i

)

1O

urtnm

VAUOS

54

JIA

4s

VAUO

}

Jtm

nvorm

a

37

NOLO OUTPUT

AWSPEEO VALIM

WENTPATN

VAJ.10

PRESELECTLTVALIO

I t

)7s

JIA

E

ES

77

JIB

m

AAlsc

OUTWT

1s

FIEOl lTER

I

7e

74

7?

m

WAoWAANINO

ALTIWOE SWITCN

T O/ MOM NULL

ALTALEnT LIGHT

ALTALERT NORN

PnESELECT AL S00FT

PRESELECT ALTCRANK

WAYPOIN11 ANN

WAYPOINT 2ANN

JIA

P RE SS A LT N O 1

P RE SS A LT N O 2

AZ-81 ODigital Air Data Computer

Block Diagram

Figure 3-2 (Sheet 2)

Pages 125/1 26

Hone~eU

3. B.

Sl-225A/Sl-225S Mach Airspeed Indicator

MAINTENANCE

MANUAL

CITATIONVll

(See figures 3-3 and 3-4, and table 3-3.)

8/11/2019 spz-8000 difcs


MACH COUNTER

VMO POINTER

G

IAS POINTER

YELLOW- INOE)(

AD 3073

S1-225A/Sl-225S Mach Airspeed Indicator

Figure 3-3


Honeywell

The Mach Airspeed Indicator provides a

displays of indicated airspeed (IAS) and

MAINTENANCE

MANUAL

CITATIONVll

servoed counter display of MACH, and servoed pointer

maximum allowable speed (VMO). The ADC provides

the instrument with the driving signals for all three functions. Four indices of various colors are

8/11/2019 spz-8000 difcs


also provided that allow the pilot to mark critical airspeeds.

A failure monitor controls the maximum allowable airspeed (VMO) failure flag and airspeed

failure (OFF) flag. The VMO flag and the OFF flag are operated by a common permanent

magnet motor. A failure affecting maximum allowable airspeed only causes display of the VMO

legend. A failure affecting airspeed only or both airspeed and maximum allowable airspeed

causes display of the OFF legend. The failure condition and the resultant flag displayed are

summarized below.

Failure Condition

Flag Dis@wed

Absence of primary instrument OFF

Internal power supply failure

OFF

Loss of reference voltage

OFF

Persistence of excessive IAS servo null signal

OFF

Absence of exiernal IAS data valid signal

OFF

Persistence of excessive Vw null signal

VW

Absence of external altitude data valid signal

v

MO

J1

I

A

5 VAC

INSTRUMENT

LIGHTING

I

LIGHTING

8/11/2019 spz-8000 difcs


{:’, - —

B

I

+

CHASSIS “ 1

GND

I

SINGLE

GND

c+

=

INPUT

POWER

26 VAC

400

Hz

{:= ‘“

I

28 VDC

I

v

ALTITUDE

(FOR vMO

AIRSPEED/VMO

MONITORING)

I

FAILUREFLAG

28 VDC IASI

I

+

FAILURE

VALID INPUT K < “

MONITOR

T

b

+12 VDC REF LO M

1-

1

I

IASSINGLE INPUT R

-22.2 mVDC/KN

b

I

4)

I

I

3. c.

AL-801 Altitude Preselect Controller (See figures 3-5 and 3-6, and table 3-5.)

8/11/2019 spz-8000 difcs


>

1:1 1:1 FEET

[-1

l-l

‘J

@J

ALTSET

Q

SET

AD-4126


Figure 3-5


Length from rear of bezel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.90 inches (226.1 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..3.28 inches

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.54 irtches

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.31b

(83.4 mmj

(39.1 mm)

(0.590 kg)

Power Requirements:

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I <,

J1

DDSONO 1

I

““’0”02+3’

8/11/2019 spz-8000 difcs


““s’””’4’

DDS*NO-----4+.

00s2N01~ 7

I

‘“’’””’-r”

SEWN04+W-=A

5LEWTACH(H,N0 ,—.+-< :;

SLEWTACH(H] NO 2 + —< 46

I I

WI

SLEWTACH(L1NO 2

++,, * 1

I

10

I

I

J1

i“ I r

.vDcN02~31~

‘;”

“”’””’’----i+’

DCGNDNO I

‘2 I

‘cGNDNO-’ T_

-mm

3. D.

DS-1 25A TAS Temperature Indicator (See figures 3-7 and 3ft, amd table 3-6.)

8/11/2019 spz-8000 difcs


TAT “C

TAS KNOTS

~Bm>

‘ “

DIM KNOB

@

\o

SAT ‘C

DIM

(Qy

AD-35321@

DS-1 25A TAS Temperature Indicator

Figure 3-7


Length from rear of bezei . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.18 inches (182.4 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3.28 inches (83.4 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.56 inches (39.8 mm)

Weight (maxirrwm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.Olb(0.454 kg)

Power Requirements:

The DS-1 25A TAS Temperature Indicator receives true airspeed (TAS), static air temperature

(SAT), and total air temperature (TAT) signals from the air data computer. TAS is displayed in

knots from 150 to 599 and SAT and TAT in ‘C from -99 to +50. The indicator normally displays

TAT and then displays SAT when the SAT switch is pressed. Display dimming is controlled by

8/11/2019 spz-8000 difcs


the DIM knob on the front panel. If the air data valid signal goes invalid, both displays will be

blank except for a single dash in the middle digit of each display.

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4.

Honeywell

AA-300 Radio Altimeter Svstem

MAINTENANCE

MANUAL

CITATIONVll

A.

RT-300 Radio Altimeter Receiver/Transmitter (See figures 4-1 and 4-2, and table 4-1.)

8/11/2019 spz-8000 difcs


RT-300 Radio Altimeter Receiverflransmitter

Figure 4-1

Dimensions (maxinwm):

Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..ll.07inches (281.2 mm)

Wtih . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Operational Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . O-2500ft

Data Outputs/Accuracy:

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Precision Output . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DC analog voItage(O-2500ft)

Gradient: -4. OmVdc/ft

Oalt=O volt

Accuracy:

o-looft,f3ft

lf)()-50()ft,*3°0

500-2500ft,*4°o

Auxiliary Output . . . . . . . . . . . . . . . . . . . . . DC analog voltage (O-2500ft)

Gradient: Per ARINCcharacteristic 552,

ALT = (0.02h + 0.4) V dc for

altitudes below 480 ft and

(10 + 10Ln h + 20) Vdc

500

for attitudes above 460 ft

Accuracy: o-loo ft, *4ft

100-500 n, +4%

500-2500 ft, ?5°0

Altitude Trips . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

100 mA current sink provided at and

below trip points indicated below:

TriD Point

Accuracy

8/11/2019 spz-8000 difcs


; TO TRANSMIT

; TO RECEIVE

r?

NTENNA

-—-

r?

NTENNA

--—

I

1

TRANSMITTER/ To ‘“LsE

VIDEO

MODULATOR

RECEIVER

AGCVOLTAGES

PROCESSOR

STCVOLTAGE

TRACK

VALID

J1

J1

I

I

SELF TEST T

1]

I

I

I

I

1;

I

I

1)

TEST INHIBIT NO. 1 D

OUTPUT

I

ASSEMBLY

=

I

W ALT OUTPUT (EH)

X AUX OUTPUT

Y

FLAG WARNING

F TRACK INVALID

N OUTPUT COMMON

8/11/2019 spz-8000 difcs


5.

6.

Paracwaah 5 is not applicable to this svstem.

EDZ-816 Electronic FlifXl Instrument Svstem (EFIS) and MDZ-816 Multi function Displav (MFD) System

8/11/2019 spz-8000 difcs


A.

ED-800 Electronic Display (See figures 6-1 and 6-2, and table 6-1.)

The ED-800 depicted in fgure 6-1, is a large format, 5 by 6 inch, high resolution cathode-ray-tube

(CRT) display. This unit presents ADI and HSI, or MFD information compiled by any of the

symbol generators. A single EDZ-816 EFIS uses two (2) ED-800 Electronic Displays, one as an

Electronic Attitude Director Indicator (EADI) and the other as an Electronic Horizontal Situation

Indicator (EHSI). The MDZ-816 MFD system uses one ED-800 Electronic Display.

All ED-800 display units are identical and interchangeable, except when used as an ADI. In this

case, an inclinometer is attached to the bezel. Leading padiculars of the ED-800 are listed in

table 6-1 and a block diagram is shown in figure 6-2.

Refer to paragraphs 6. A. (1) thru 6. A. (4) for a functional description of the ED-800. Paragraph

6. B describes the display features of the ED-800 when used as an AD I. Paragraph 6. C

describes HSI features and the composite display. Paragraph 6. D describes EFIS reversionary

controls and annunciators, and paragraph 6. E describes the display features of the ED-800 when

used as an MFD.


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..10.50 inches (266.70 mm)

8/11/2019 spz-8000 difcs


Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6.08 inches (154.48 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..5.08 inches (129.03 mm)

Weight (maxinwm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 10.31b(4.67kg)

Power Requirements:

Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 28Vdc.65Watts maxinwm

Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..5Vacordc. l.2Watts maximum

and28Vdc, l.2 Watts maximum

Mating Connectors:

J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. MS27473E20B35S

with strain relief MS27506-B20-2

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

Clamp, Honeyuvell Part No. 7000066-6

or MSPlnc, Part No. 64440

User Serviceable Parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NONE


Leading Particulars

Table 6-1

6.

A. (2)

Video and Dimming System

The video system provides the individual drive signals to each of the three (red, blue, and

green) electron guns in the CRT. Amplitude of the gun drives are adjusted to provide the

required color selection menu. Four bits of color selection data are used providing for a

8/11/2019 spz-8000 difcs


possible 8 colors on both raster and stroke operation (for a total of 16).

The overall intensity of the display output is controlled in the video system by a signal

from the autodimming system. In the auto-dimming system, the pilot-selected intensity

(from the dimming control on the DC-81O cmtroller) is modulated by a control signal

generated from two strategical ly located ambient light sensors. This al lows the pilot to

select a different intensity level for the weather radar display and the remainder of the HSI

functions.

(3) System Monitor

A system monitor is incorporated in the ED-800 to provide CRT phosphor protection and a

system invalid signal to the symbol generator whenever the following conditions are

detected:

Loss of deflection in either axis

c Abnormal power supply outputs

“ Improper CRT filament current

The circuitry also provides a 5-second time delay between application of CRT filament

current and high-voltage power turn-on. This al lows the system to stabil ize quickly and

also protects the CRT catbodes from the effects of excessive initial anode current.

r

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POWER SUPPLY I

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STROKE READY

(

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(SEC)

X DEFLECTION

{

(PRt)

(SEC)

Y DEFLECTION

(

(PRI)

(SEC)

RASTER/STROKE

(

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ED-800 Electronic Display Block Oiagram

Figure 6-2

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Pages 143J144

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ORRECTION

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AD s272

22-05-07

6. B. ED-800 Used As An Electronic Attitude Director Indicator (EADI)

The EADI combines the famil iar true sphere-type atti tude display with lateral and vertical

computed steering signals to provide the pilot commands required to intercept and maintain a

desired flight path. The EADI provides the following display information:

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.

.

.

.

.

.

Glideslope or Glidepath Deviation

Expanded Localizer or Azimuth Deviation

Radio Altitude

Rising Runway

- Digits/ Readout

- Decision Height

Marker Beacon Annunciation

Cross-side Sym&J Generator Switching

Rate-of -Turn

Fast/Sbw AOA Command

Attitude Source

FD Mode Annunciations

Airspeed Trend Error

Autopilot Engage Status

Air Data Command

Airspeed Display

Digital Airspeed

Analog Vertical Speed

(1) ED-800 EADI Displays and Annunciators (See figure 6-3.)

(a) Decision Height Display

6. B.

(1) (d) Glideslope, Vertical Navigation, or Glidepath Deviation Pointer

The glideslope pointer and scale are in view when tuned to an ILS frequency to

display aircraft deviation from glideslope beam center. Aircraft is befow glidepath if

pointer is displaced upward. Each glideslope dot represents displacement from the

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beam center-line. During ILS approaches, if decision height (DH) is set below 160

feet with the AP engaged, a Categofy II window (green) appears on the glideslope

deviation scale. If the VNAV mode is selected, the pointer indicates the VNAV

computer path center to which the aircraft is to be fbwn. Vertical track aletl (VTA) is

annunciated 1 minute prior to VNAV capture and is removed at VNAV capture. If

MLS is selected, the pointer would indicate deviation from the selected glidepath

angle.

The vettical deviation scale is identified to show its current function: ILS, MLS, and

VNAV. The glideslope and glidepath pointer is on the right side of the scale; the

VNAV pointer is on the outside of the scale.

(e) Flight Director Mode Annunciators

Flight director vertical and lateral modes are annunciated along the top of the EADI.

Armed vertical and lateral modes are annunciated in white to the Iefl of the captured

vertical and lateral mode annunciators. Capture mode annunciators are displayed in

green and are located to the left of top center for lateral modes and in the upper

right corner for vetiical modes.

As the modes transition as specified bebw, a white

box is drawn around the capture or hold mode annunciator for 5 seconds.

Lateral Transitions

6. B.

(1) (f)

Flight Guidance Computer Status

A green AP ENG is annunciated whenever the autopilot is engaged. If touch control

steering is being used, the AP ENG annunciation is replaced with an amber TCS

ENG.

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(9)

Vertical Speed Indicator

A therornometer-type scale and pointer to the right of the attitude sphere indicates

current aircraft vertical climb or descent. The green vertical speed indicator has a

range of M999 ft/min with a resolution of 50 fl/min for speeds under 1000 ft/min and

100 ft/min for speeds over 1000 ft/min. The green vertical speed pointer indicates

the approximate verlical speed and the direction of vertical speed travel. The

vertical speed scale has a range of t7000 ft/min with scale markings at 1,000,

2,000, and 6,000 ft/min.

(h)

Radio Altitude Display

The four-digit display indicates the aircraft’s radio altitude from O to 2500 feet. The

resolution above 200 feet of altitude is 10 feet; below 200 feet, the resolution is 5

feet. The display is blanked for altitudes greater than 2500 feet. When the radio

altitude data is invalid, the display indicates a dash in each of the digits.

(i)

Marker Beacon

Marker beacon information is displayed to the right of the expanded localizer/azirrwth

scale. The markers are of the specified mlors of blue for outer, amber for middle,

6. B.

(1) (1) Expanded Localizer or Azimuth Pointer

Expanded Iocalizer is displayed by the Iocalizer pointer whenever a valid Iocalizer

sgnal is available. Raw Iocalizer displacement data from the navigation receiver is

amplified approximately 7-1/2 times to permit the bcalizer pointer to be used as a

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sensitive reference indicator of the aircraft’s position with respect to the center of the

Iocalizer. It is normally used for assessment only.

During final approach, the pointer serves as an indicator of the Category II window.

Keeping the expanded Iocalizer pointer within its full-scale marks ensures the pibt

that he will touch down within *33 feet of the certterfine of the runway when using a

Category II ILS system. When tuned to other than an ILS frequency, the expanded

Iocalizer display is replaced by the rate-of-turn display. When MLS is selected, the

expanded Iocalizer pointer displays deviation from the selected azimuth angle.

(m) Attitude Sphere

The sphere moves with respect to symbolic aircraft reference to display actual pitch

and roll attitude. Pitch attitude marks are in 5degree increments.

(n) Aircraft Symbol

The symbol serves as a stationary representation of the aircraft. Aircraft pitch and

roll attitudes are displayed by the relationship between the fixed miniature aircraft

and the movable sphere. The miniature aircraft is fbwn to align the command cue

to the aircraft symkd in order to satisfy the commands of the selected flight director

mode.

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6. B.

(1) (q)

Pitch and Roll Command Cue

The pitch and roll command cue displays computed steering commands to capture

and maintain a desired flight path. The aircraft symbol is always flown to the flight

director cue. The cue is biased out-of-view if an invalid condition occurs in the fl ight

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director pitch or roll channel.

NOTE: Unless deselected with DC-81 O Display controller, the command bars

are shown in both EADIs and reflect fl ight director guidance to the selected

modes based on the selected EHSI data.

(r) Rate-Of-Turn Display

When tuned to other than an ILS frequency, the rate-of-turn is displayed by a pointer

and scale at the same location as the expanded Iocalizer. The rate-of-turn of the

aircraft is indicated by the position of the pointer against scale indices. The marks at

the extreme left and right sides of the scale represent a standard rate of turn

(2-minute or 3-degree per second turn rate).

(s)

Air Data Command Display

When selecting a flight director mode of either fl ight level change (FLC) or vertical

speed (VS), the command reference will appear in the lower left comer. The

guidance controller pitch wheel may be used to change the air data command

reference. For other vertical modes, the air data command display will be removed.

6, B.

(2)

ED-800 EADI Amber Caution and Failure Annunciation (See figure 6-4.)

(a)

Same Attitude Source

There is no attitude source annunciated if the pilot and copilot are using their normal

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(b)

(c)

(d)

attitude sources. Selecting other attitude sources causes the new source to be

annunciated in white. If the pilot and copilot have selected the same attitude source,

that atti tude source is annunciated in amber on both EADIs.

Flight Director Failure

An amber FD FAIL warning is displayed at the top left of the EADI in the event of a

flight director fai lure. Also, the flight director cue and all FD mode annunciators are

removed. During self-test, if the FD mode annunciator test is valid, the word TEST

is annunciated in magenta at the same location as FD FAIL.

Decision Height Warning

When the radio altitude is within 100 feet of the decision height, a white box will

appear to the left of the radio altitude display. When at or bebw the decision height,

an amber DH will appear inside the white box.

Comparison Monitor

Selected pilot and copilot input data is compared in the symbol generator. If the

difference between the data exceeds predetermined levels, an out-of-tolerance

symbol will be displayed. A list of the compared signals and the displayed

Compared Signals

Tolerance

Displayed Symbol

1. Pitch Attitude

*6 Deg

PIT

2. Roll Attitude

M Deg

ROL

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3. Heading *6

Deg

HDG

‘ 4. Localizer

~40 mv

LOC (AZ**)

* 5. Glideslope

*50 mv

GS (GP*’)

6. Pitch and Roll Attitude f% Deg

AIT

* 7.

Localizer and Glideslope

*4o mv (LOC)

ILS (MLS**)

+50 mv (GS)

NOTES:

These comparisons are only active during flight director Iocalizer and glideslope

carXure with both NAV receivers tuned to a LOC frecmencv.

** When MLS is selected on both NAV receiver’s (pilot’s and copilot’s), Iocalizer (LOC)

becomes azimuth (AZ), glideslope (GS) becomes glidepath (G P), and ILS becomes

MLS. The tolerance for AZ is approximately 1/2 dot and GP is approximately 2/3 dot.

** The heading monitor threshold is *6 degrees for bank angles up to 6 degrees. When

bank angles exceed 6 degrees the threshold is fl 2 degrees heading and remains at

*I 2 degrees for 45 to 90 seconds after bank angle is reduced below 6 degrees. If

the compared heading sources are not the same (both MAG or TRU), the comparison

monitor is disabled.

3EClS10ti

SAME

FLIGHT

ATTITUDE

DIRECTOR

HEIGHT

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COMMON SYMBOL

GENERATOR

ANNUNCIATOR

F1

L IGHT DIRECTOR

COUPLED TO THE

CROSSIOE HSI

BLANK

< HSI

HSI >

NOTE

AIRSPEED

DISPLAY

FAILURE

HEADING

COMPARISON

MONITOR

AIR DATA

COMMAND

FAILURE

DECISION

HEIGHT

FAILURE

m

TATUS

MESSAGES

AP ENG

TCS ENG

FA4S MSG

m

S

ILS

AZ

GP

MLS

I

6. B.

(2) (h) Decision Height Failure

In the event of an open DH potentiometer, or during self-test, amber dashes will

replace the numerical values of the decision height display.

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(i)

0)

(k)

(1)

Air Data Command Failure

In the event of an air data computer failure, amber dashes will replace the

numerical value of the specific air data command displayed.

Airspeed Display Failure

In the event of an air data computer failure, amber dashes will replace the

numerical airspeed value.

HSI Couple Symbol (Cross-Side Command Cue)

Normally both flight guidance computers are coupled to the left EHSI. The copilot’s

EADI will usually have an amber <HSI symbol to indicate that the left EHSI is

supplying information for the command bars. The HSI SEL arrow on the GC-81 O

Flight Guidance Controller will display the same selection. The HSI couple symbol

will move between EADIs, depending on which HSI is selected.

Common Symbol Generator

When in the reversionary mode and one symbol generator is driving both pilot and

copilot display tubes, a reversionary warning is given in amber, which indicates the

information source. This display appears next to the upper left corner and will

6. B. (3) (b) Excessive Deviation

Between 300 feet and 100 feet radio altitude, when the flight guidance APP mode

is selected, and the autopilot is engaged, this feature is enabled. The green CAT II

window on the glideslope scale will then be displayed.

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With the excessive deviation monitor enabled, exceeding either of the following

thresholds will cause the expanded Iocalizer and glideslope scales to flash.

Signal Threshold

Localizer

25 mV (- 1/4 dot)

Glideslope

75 mV (- 1/2 dot)

(4)

ED-800 EADI Red Failure Annunciations

(a)

(b)

Attitude Failure (See figure 6-5.)

In the event of a failure of the attitude display, the pitch scale and roll pointer will

be removed, the sphere will be painted blue, and a rd ATT FAIL will be displayed

in the middle of the sphere upper half.

Glideslope, VNAV or Glidepath, Expanded Localizer or Azimuth, Fast/Slow

Command, and Rate-Of-Turn Failures (figure 6-5)

In the event of a failure of any of these systems, the pointer is removed and a red

X is drawn through the scale. The annunciation ILS, MLS, or VNV remains at the

bottom of the vertical deviation scale to identify the invalid information.

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6. C.

ED-800 Used As An Electronic Horizontal Situation Indicator (EHSI)

The EHSI combines numerous displays to provide a map-l ike display of the aircraft position.

The indicator displays aircraft displacement relative to VOR radials, Iocalizer, and glideslope

beam. At power-up, the EHSI presents a full compass display. By pressing the DC-81 O

FULIJMAP button, the full compass display is changed to a partial compass format. Also, if

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weather radar returns or map information is desired, pressing the WX or MAP button on the

DC-81 O changes the full compass display to a partial compass displaying weather radar returns

or map information. The EHSI provides the following ful l and patiial compass display

information:

Full Compass Displays

Heading

“ Heading Sync

Heading Select

Heading and NAV Source Annunciators

Course Select

Course or Azimuth Deviation

Distance

Ground Speed or Time-To-Go

To/From

Desired Track

Bearing 1 and 2

. Vertical, Glideslope, or Glidepath Deviation

Partial Cornpa ss Displays Only

Weather Radar

DRl~

HEADING

HEADING

FORE

;$l:LE SOURCE

:~ECT

LUBBER

BEARING

ANNUNCIATOR

LINE

WAYPOINT

POINTERS

ANNUNCIATOR

FMS APPROACH

ANNUbKIATOR

\

COURSWDESIRED

\

I // /

TRACKDISPIAY

J $ /// ,

CONPASSSYNC

ANNUNCIATOR

COURSE

SELECT

POINTER

NAVIGATION

<SOURCE

ANNUNCIATOR

= DME HOLD

RADAR MODE

(NOTE 2)=

M

VOR 1 lLSl

~::: ILS2

FMS1

MLS2 FMS2

\

\

Honeywell \

\

h

%

Y

+

CRS

MAG1

WPT

ILS1

::5

\:<\’g: ‘ J ‘1’/>/ 2“’ ‘M

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A

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WEATHER RADAR ml

ANNUNCIATOR I 11-”’=

SWRCE “

ANNUNCIATOR

(~,

GUDESLOPW

GUDEPATH

%vERTICAL

DEVIATION

POINTER

_GROUND SPEED

DISPLAY(NOTE1)

II

@

\

@

/

HEADING

AIRCRAFT Af=f

SELECT

SYMBOL

RECIPROCAL

COURSE OR

LUBBER

DISPLAY

COURSE

AZIMUTH

LINE

POINTER

DEVIATION BAR

NOTES:

1. TIME-TOGO ISALSO DISPIAY ATTHIS LCCATION.

2. lX APPEARS WHEREVER WEATHER RADAR ISTRANSMllTER

AND WX ISNOT SELECTED ON THE DC-81O.

EHSI Displays and Annunciators

Figure 6-7

o

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0 ‘+

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AD453

22=05-07

Pages 159/1 60

6. C. (1) (c)

Drift Angle Bug (INS only)

If available, the drift angle bug with respect to the lubber line represents drift angle

left or right of the desired track. The drift angle bug with respect to the compass

card represents aircraft actual track. The bug is displayed as a green triangle that

moves around the outside of the compass card (either partial or full).

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(d) Heading Source Annunciator

The current heading source is annunciated in white (top left side of the EHSI) when

the pilot and copilot sources are not the same. As other heading sources are

selected, they are also annunciated in white at the top left side of the EHSI. If the

heading sources (pilot and copilot) are the same, the annunciation is in amber.

(e) Heading Select Bug and Heading Select Display

The notched blue heading select bug is positioned on the rotating heading dial by a

remote heading knob to select and display preselected compass heading. The bug

rotates with the heading dial; therefore, the clifference between the bug and the fore

lubber line index is the amount of heading error applied to the flight director

computer. A digital heading select display is provided for convenience in setting

the heading bug.

(f) Heading Display and Dial and Fore and Aft Lubber Lines

Gyro stabilized magnetic compass information is displayed on the heading dial,

which rotates with the aircraft throughout 360 degrees. The azimuth ring is

graduated in 5-degree increments. Fixed heading marks are at the fore and aft

lubber line positions and at 45-degree bearings.

6. C. (1) (h) Waypoint Annunciator

This amber annunciator indicates waypoint passage for the long-range navigation

system displayed on the EHSI. The annunciator l ights 2 minutes prior to waypoint

passage.

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(i) Navigation Source Annunciators

Annunciation of the navigation source is displayed in the upper right hand corner.

Long-range navigation sources (INS, VLF, RNAV, FMS) are displayed in blue to

distinguish them from short-range sources annunciated in white.

(0

Distance Display

The distance display indicates the nautical miles to the selected DME station or

waypoint. Depending on equipment, the distance wil l be displayed in a O-399.9 or

a o-3999 nautical mile format. DME HOLD is indicated by an amber H adjacent to

the distance readout.

(k) Vertical Navigation, Glideslope, or Glidepath Deviation Pointer

The vertical navigation display and annunciator come into view when the VNAV

mode on the flight director is selected. The deviation pointer then indicates the

VNAV’S computed path center to which the aircraft is to be flown.

NOTE: Consult the appropriate documentation for the installed LRN to verii its

VNAV capability.

The glideslope display and annunciator come into view when a VHF NAV source is

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6. C.

(1) (q)

Weather Radar Annunciators (Full Compass Display)

Weather radar modes and antenna tilt angle are annunciated on the left side of the

EHSI. Target Alert annunciators (all formats) are only displayed when partial

compass format is selected.

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Tilt Angle - The angle of the weather radar antenna position is displayed in

positive degrees for up-tilt and negative degrees for down-titt.

Radar Mode - The TX annunciation appears whenever the weather radar is

transmitting, and WX mode is not selected on the DC-810 Display Controller.

(r)

Course Select Pointer

The yellow course pointer is positioned on the rotating heading dial by a remote

course knob to select a magnetic bearing that coincides with the desired VOR radial

or Iocalizer course. The course pointer rotates with the rotating heading dial to

provide a continuous readout of course error to the flight director computer.

When long-range navigation is selected, the course pointer becomes a desired track

pointer. The position of the desired track pointer is mrttrolled by the long- range

navigation system. A digital display of desired track (DTRK) is displayed in the

upper left hand corner.

(s) Compass Sync Annunciator

The compass sync annunciator indicates the state of the compass system in the

slaved (AUTO) mode. The bar represents commands to the directional gyro to slew

to the indicated direction (+ for increased heading and o for decreased heading).

6. C. (2) EHSI Partial Compass ARC Format (See figure 6-8.)

The partial compass format displays the same information as the full compass format,

except for the following differences:

The paftial compass mode displays a 90° arc (+45°) of the compass card.

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Pressing the FULIJMAP button once on the DC-81 O Display Controller causes the

heading dial to change to the partial compass format allowing one waypoint for

each bearing pointer, wind vectors, and VOR/DME ground station positions to be

displayed. (EFIS also has the capability of displaying multiple waypoints.)

Digital Heading readout - For convenience, a display of the aircraft’s current

heading is provided at the top of the compass card.

Drift Bug - The drift bug will be displayed when FMS is the selected navigation

source. The drii bug indicates the angular difference between FMS calculated

track and aircraft current track.

Range Rings - Range rings are displayed to aid in the use of radar returns and

position of navaids. The outer range ring is the compass card boundafy and

represents the selected range on the radar. The range annunciation on the inner

ring represents one-half the range setting of the weather radar. When the

weather radar is off, the display indicates the 100 mile range.

Wind Vector Display -

Wind

vector information is displayed left of bottom center,

The wind can be shown with velocity and direction or broken into headhail

component and crosswind mmponent. In both cases, the arrow shows the

direction and the number indicates velocity of the wind. The type of display is

determined during installation. The wind vectors are available from long-range

HEADING READOUT

DRIFT BUG

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NAVAID=

\ /

\ I

/ Iioneywell

/

/

\

/

I

\

/

/ / \

\

ADF

HDG

25 i

.o+o\ 50>

319

7

15+

260 KTS

—

/

f-l

u

----

~

RANGE

7RINGS

I

RANGE ANNUNCIATION

(NAUTICAL MILES)*

2.5 25

50

1:

100

12.5

150

*RANGE ANNUNCIATION ON

INNER RING IS 1/2 THE

RANGE SETTING OF THE

Hone~eII

6. C.

(3) EHSI Weather Radar Displays (See

MAINTENANCE

MANUAL

CITATION Vll

figure 6-9.)

Pressing the WX button on the DC-810 Display Controller when the EHSI in the pattial

compass format selects the radar return (storm intensity levels) display. Weather radar

antenna titt angle, modes, and target alert station are annunciated on the left side of the

EHSI.

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(a)

Heading (HDG) OFF Scale Arrow

Any time the heading select bug is moved more than 45° from the fore lubber line,

this arrow will come into view above the compass card boundary. The arrow can be

on the left or right side and indicates the closest direction to the bug.

Note that while an ILS frequency is tuned as the EHSI navigation source, the MAP

format is inhibited. Toggling is allowed between FULL and ARC formats only.

(b) Target Alerl (TGT) and Variable Gain (VAR) Annunciator

The target alerl annunciator warns of level 3 targets 7.5 degrees either side of

aircraft heading within a 60 to 120 nautical mile (NM) range. A green TGT indicates

an armed condition (target alert selected) while an amber TGT indicates a weather

alerl condition (e.g., level 3 WX return detected within 7.5 degrees of the aircraft

heading, but beyond the selected radar range). For target alerl to be operable, the

gain must be in the preset position. An amber VAR indicates the radar is operating

in the variable gain mode.

(c) Range Ring and Annunciator (WR, NAV, and NAV/WR Formats Only)

Range is selected on the weather radar controller. One-half the selected range is

TGT OR VAR ANNUNCIATOR

I

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M %,

ANNUNCIATIONS

wAIT (GREEN)

STBY (GREEN)

TEST (GREEN)

wx (GREEN)

RcT (GREEN)

GcR (AMBER)

GMAP (GREEN)

FAIL (AMBER)

TIJRB (GREEN)

FFILN

K

Honeywel l

l \

\

\

o

/

WX ANTENNA/

TILT ANGLE

,

VOR 1

,/ \

+

VOR2

r

L

A n

/ ;::L:FF

ARROW

+ WEATHER

RADAR

RETURN

6. C.

(3) (f)

Weather Radar (WX) Mode Annunciations

The following radar operating modes are annunciated on the EHSI.

Operating Modes

Wait

Annunciation

WAIT

Color

Green

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Standby

Test Weather

REACT

Ground Map

Ground Clutter

Reduction

Turbulence

Fail

Transmitting-but not

selected for display

STBY

TEST

RCT

GMAP

GCR

T

FAIL

TX

Green

Green

Green

Green

Amber

Green

Amber

Magenta

The TX annunciation appears whenever the weather radar is transmitting, and WX

mode is not selected on the DC-81 O Display Controller.

(9)

Lightning Sensor (LX) Mode Annunciation

The following mode annunciations may a~~ear on the weather radar indicator, EHSL

or MFD disp~ays.

Annunciation

. .

ODerating Modes

6. C.

(4) EHSI Map Mode Wdh VOR Selected For Display

In the map format, when VOR is selected for display, the normal ARC course select

display (pointer, scale, and deviation) is removed and replaced by the following display

(figure 6-1 O).

The VOR or VOR-DME station is displayed at its geographical position with the

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corresponding symbol (if display range permits).

“ The course information is indicated by a digital readout (same as ARC) and by a

course line centered on the VOR station. The TO information is represented by a

continuous line, the FROM information being represented by a dashed line.

If the VOR station is out of the display range, an arrow is added to the course line to

indicate the direction of the selected course to be followed.

The deviation is displayed as a digital cross-track distance readout.

If the selected VOR bearing (1or 2) is different from the VOR NAV source (VOR 2 or

VOR 1), a magenta navaid symbol is displayed at the geographic location. If the

symbol for the selected bearing is out of map range, then the appropriate VOR

bearing pointer (1 or 2) is displayed.

If the selected source VOR bearing (1 or 2) is the same as the VOR NAV source

(VOR 1 or VOR 2), a magenta number corresponding to the VOR bearing number is

displayed to the left of the green or yellow VOR symbol. If the selected bearing

symbol is out of map range, a magenta bearing pointer is displayed.

“ Map range is controkf from the installed weather radar range control.

6. C. (5)

EHSI Map Mode Wtih FMS Selected For Display (See figure 6-11.)

With the EHSI in the full compass format, pushing the FULUMAP button twice on the

DC-81 O Display Controller wil l display the paftial compass map format. I f the instalJed FMS

has the capability, up to six waypoints are displayed, along with the desired track between

waypoirtts. This assumes that the displayed range has been selected accordingly with the

weather radar range control.

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(a) Upon selection of the MAP mode, the following will occur:

.

.

The course select pointer and HS1-type course deviation bar displays are

removed.

Both bearing pointers are removed.

A digital course deviation display wil l be present on the bottom of the EHSI

display. This wil l show the position of the aircraft with respect to the desired

track,

Muttiple waypoints will appear on the EHSI.

A white track line connects waypoint to waypoirtt.

Most map symbology is a function of the the installed FMS. Each waypoint is

identified by a number 01 thru 99 or when the FMS communicates by the GAMA

standard bus for alphanumerics, the waypoint is identified by name.

The waypoint to which the aircraft is flying is magenta in cofor. All other

waypoints are white.

If the EFIS is receiving valid VOR station and DME distance, the navaids for the

two VOR stations will be available for display on the EHSI, no matter where the

bearing selector switches are set. The blue navaid will be VOR 1 (NAV 1), and

the green navaid will be VOR 2 (NAV 2).

6. C.

(5) (b) Heading Select Bug Out-Of-View Arrow

Any time the heading select bug is moved more than 45° from the fore lubber line,

this arrow will come into view above the compass card boundary. The arrow can

be on the left or right side and indicates the closest direction to the bug.

Note that while an ILS freauencv is tuned as the EHSI navigation source. the MAP

.

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format is inhibited. Toggling is allowed between FULL and’

ARC formats only.

NAVIGATION

TO-FROM

HEADING SOURCE

HEADING

SOURCE

ANNUNCIATOR

ANNUNCIATOR

DISPLAY

ANNUNCIATOR

,,,,,,g,,,:D+s2ii\A,l

J-J%J::E

HEADING 315T0

SELECT

30 NM

RINGS

BUG

+10

H*3

;Z:T

<~

-

4

F

6. C. (6) EHSI Full Compass - Amber CautiodFailure Annunciations (See figure 6-12.)

(a)

Heading Source or Navigation Source Annunciators

When the pilot and copilot have selected the same heading or navigation source,

the applicable source is annunciated in amber; otherwise, the annunciation is in

white. For SRN sources, if the pilot and copilot have both cross switched to the

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(b)

(c)

(d)

other’s source, the annunciator would be amber even though they would be from

different sources.

If the pilot selects the copilot’s (VOR 2) navigation source and the copilot selects

the pilot’s (VOR 1) navigation source, both annunciators are amber to indicate

cross-switched sources.

DME Hold Annunciator

When DME is set in the hold position, an amber H is displayed

to

the left of the

numerical DME readout.

Waypoint (WPT) Alert Annunciator

An amber WPT annunciation from a long-range navigation system indicates

waypoint passage. The annunciator l ights 2 minutes prior to waypoint passage.

Display Failure Annunciators

When any of the following systems fail, the digital display is replaced by amber

dashes.

6. C. (6) (f)

Weather Radar Target Alerts

Weather radar target alerts are annunciated on the left side of the EHSI. An

amber TGT indicates a weather alert condition and an amber VAR indicates the

radar is operating in the variable gain mode.

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(9) Course Select and Heading Select Annunciators

Failure of the course or heading select signals will cause these displays to be

replaced by amber dashes. They are also dashed when the heading display is

invalid.

SAME HEADING

OR NAV SOURCE

(AMBER)

COURSE SELECT

FAILURE

(AMBER DASHES) ,

(NOTE 2)

w

—-

- <,

I WEATHER TARGET I

I

TGT VAR

1

II /“’ >.?

WX FAILURE

(AMBER)

\ DMEHOLD

ANNUNCIATOR

(AMBE+)

\ WAYPOINT

6. C. (7) EHSI Full Compass—Red Failure Annunciations (See figure 6-13.)

(a)

Heading Failure

A failure of the heading system valid results in the removal of drii angle, bearing

pointers, To-From arrow, select course pointer, selected heading bug, course

deviation pointer, and course scale. The digital select course and digital heading

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(b)

select readouts are dashed, and a red HDG FAIL is displayed at the top of the

heading dial.

Course, Vertical, Glideslope Deviation, or MLS GlidePath Deviation Failure

A failure of the course, vertical, glideslope, or glidepath deviation systems results in

the removal of the course, vertical, glideslope, or glidepath deviation pointer, and

paints a red X through the scale.

(8) EHSI Partial Compass Failure Annunciations

The partial compass failure annunciations are identical to those of the full compass format

with the exception of Course Select/Desired Track Deviation failure. Should this failure

occur, the deviation bar is removed from the display and a red X is drawn through the

scale.

HEADING FAILURE

AMBER DASHES DISPLAY

(NOTE 2)

(RED)

6. C. (9) Composite Mode Symbology (See figure 6-14.)

In the event of a display unit failure, the EAD1/EHSl DIM control on the DC-81 O Display

Controller is turned to the OFF position to display a composite attitude and NAV format on

the other ED-800 display. Figure 6-14 defines the location and form of the composite

display elements. As in normal EADI and EHSI presentations, all elements are not

displayed at the same time, The presence or absence of each display element is

determined by flight phase, NAV radio tuning, selected flight director mode, absolute

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altitude, etc. The failure, caution, and warning annunciations function is much the same

as for the normal display mode.

The composite mode deviation functions as a simple, fixed card CDI (course deviation

indicator) for VOR data. As long as the aircraft is headed within 90 degrees of the

selected course or selected radial, as long as the TO-FROM annunciation is correct, the

CDI is directional; othetwise, it displays reverse sensing and the techniques required for

reverse sensing apply.

For Iocalizer (LOC) data, this CDI display contains some additional capability. When the

aircraft has a heading greater than 90 degrees to the selected inboard Iocalizer course,

the CDI will reverse polarity. In this case, it will remain directional.

ROLL

ATTITUDE

AITITUDE POINTER NAVIGATION

TO-FROM

SOURCE

DECISION AND

SOURCE

DISTANCE

ANNUNCIATOR ANNUNCIATOR HEIGHT SCALE ANNUNCIATOR DISPLAY

COURSEIDESIRED

TRACKDISPLAY

6. D. EFIS Reversionary Controls and Annunciators (See figure 6-15.)

The EDZ-816 EFIS allows pilot selection of alternate source data inputs, EADI or EHSI displays,

or symbol generators. The extent of the reversionary switching capabil ity depends on the

installed options. Control of these functions is mostly done with external cockpit mounted

switches. These alternate selections allow the pilot to maintain usable fl ight displays even after

multiple failures.

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The EADI and EHSI displays are normally reconfigured with the DC-81 O dimming controls.

Additional controls for selection of alternate EFIS sources are located on the outer portion of

each pilot’s instrument panel (figure 6-1 5). Switches on the pilot’s side control the pilot’s EFIS

and the copilot’s switches control the copilot’s EFIS. Alternate source selection is described

below.

(1) Heading Reversion (HDG REV) Button

The HDG REV button selects alternate heading sources for display

on

the EHSI as listed

below.

Action

Pilot

@E&l

Power-up

MAG 1

MAG 2

First Push

MAG 2

MAG 1

Second Push

MAG 1

MAG 2

(2) Attitude Reversion (All REV) Button

The ATT REV button selects alternate atti tude sources for display

below.

on the EADI as listed

6.

D. (3)

Symbol Generator Reversion (SYM GEN REV) Button

Pressing the SYM GEN REV button selects the opposite side symbol generator as an

alternate source of information display on the EADI and EHSI. The sequence for

reversionary source selection is listed below.

Action ~ -

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Power-up Blank

Blank

First Push

SG 2

SG 1

Second Push

SG 1

SG 2

The selected symbol generator is annunciated in amber on the pilot and copilot EADIs as

SG 1 or SG 2, depending on whether the source is the number 1 (pilot) or number 2

(copilot) symbol generator. Pressing the SG GEN REV button a second time reverts the

EADI and EHSI displays back to the original failed condition. If both symbol generators

fail , the multi function symbol generator (MG) may be used as a source of information

display.

If the above sequence is used, the failed side becomes a slave to the remaining operating

EFIS. If the MG is used, the failed side retains full operational capability but the MFD unit

is unusable (blank). The MG reversionary selection is accomplished with the MFD

controller.

When the MG is used as a pilot or copilot symbol generator, all flight director

modes

are

reset and the on-side VOFVLOC is selected. However, this only occurs on the side

selected by the HSI SEL button on the GC-81 O Flight Guidance Controller. If the MG is

used as a backup on the side not selected by the HSI select button, all modes and

6, E. ED-800 Used As A Multifunction Display (MFD)

The MDZ-816 Multifunction Display System has four major functions:

Weather radar

Navigation data

Q Checklist

EFIS Reversionary

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One of the most signif~ant of these functions is its ability to back up either of the two EFIS

symbol generators or one of the EHSI displays. Should one SG fail, the pilot can select the

MFD symbol generator to take over operation of the failed side’s displays with al l functions and

operations unchanged. The EFIS DC-81 O controller on that side will continue to operate the

display formats as before. The MFD display can also be used as a backup in the event of an

EHSI display failure.

The MFD system greatly expands on the navigation mapping capabilities of the EFIS, This is

primarily due to the fact that the MFD display area can be used exclusively for map formats

without the need for the essential heading and NAV data that the EHSI also has to contain.

Some of the additional information that can be added to the traditional map display of waypoint

locations includes waypoint and VOR identifiers, aircraft present position in LAT/LON

coordinates, and the TO waypoint time to go. This additonal data is supplied by the optional

Flight Management System for display on the MFD. The MFD system also has a north-up plan

function in addition to the usual heading-up map display. Both formats make use of a designator

controlled by the MC-800 MFD Controller joystick. The position of the designator can be

automatical ly transmitted to the FMS to be used in defining a new waypoint.

NOTE:

The display capability of the MFD depends on the FMS installed in the aircraft.

The display formats shown in this section assume that a Honeywell FMZ

Color weather radar information from the PRIMUS@ 870 Receiver-Transmitter is

presented in the form of an overlay by raster techniques on the stroke written display. A

white outer range ring is provided. An inner range ring is also provided with its associated

label also stroke written in white characters on the right side of the display. Weather

intensity levels are differentiated by the standard convention of red, yel low, green, and

blue areas.

The blue field is generated by the rain-echo attenuation compensation circuitry to warn the

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pilot that the weather levels in this area cannot be calibrated and are therefore of

unknown precipitation density. Ground mapping may also be displayed on the standard

convention of magenta, yellow, and cyan areas. The radar mode is annunciated in the

upper right side of the display.

A target alert annunciator is provided to warn of level 3 targets 7.5 degrees on either side

of the aircraft flight path 60 to 120 miles in range. A green TGT annunciates this feature,

changing to an amber TGT when active. When the gain is not calibrated, an amber VAR

annunciator will be displayed in the target alert area. Annunciators below the outer range

label display RCT in green characters for Rain Echo Attenuation Compensation Technique

(REACT).

A weather radar failure will remove the raster weather display and force the mode

annunciator to display WX in amber characters.

A magenta TX is displayed where WX is annunciated when the P-870 is ON and weather

is not selected for MFD.

If the installation is equipped with a dual weather radar controller, the green arrow (+)

over the WX annunciation indicates which weather radar control ler has control of the

weather display.

When coupled to a compatible LRN, the NAV route with up to six waypoints can be

displayed to the range limit of 1200 miles, or the next route segment can be displayed.

When weather returns are selected, the maximum selectable range is slaved to the

WC-870 WX Controller. With a compatible NAV source, such as the Honeywell FMS with

stored database, other pertinent navigation data beyond route mapping such as VOR

station locations, and time-to-go to the next waypoint, can be selected and displayed.

A movable designator can aid in relocating the next waypoint. When the designator is

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6.

moved from its “home” position, the LAT-LON display reflects the designator position,

which then can be automatically loaded as the next waypoint into compatible LNAV or

FMS sources. The map mode displays shown on figure 6-16 are described below.

E. (2) (a)

(b)

(c)

(d)

Heading Display

The HDG display indicates the actual heading of the aircraft, It is the same heading

information displayed on the EHSI.

VOFUDME Symbols

These symbols are added upon actuation of the VOR button on the MC-800. They

represent the nearby VOR stations stored in the LRN database.

WX Target Alert

This annunciator warns the pilot of level 3 targets 7.5 degrees on either side of the

aircraft and 60 to 120 miles in range.

Selected NAV Source

6.

E. (2) (h)

WX Tilt Angle

The angle the weather radar antenna is positioned is displayed in positive degrees

for up-tilt and negative degrees for down-tilt.

Airport Annunciator

The airports are identified as a function of the APT button on the MC-800.

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Designator Annunciations

These annunciations display the course and distance tot he designator from present

position.

MAGITRU

HEADING

TUNED

SELECTED NAV

DESIRED TRACK

ANNUNCIATIONS

DTRK LINE

WX TARGET

ALERT

6. E.

(2) (k) Waypoint and Waypoint Data

The number of available waypoints is dependent upon the LRN, which is providing

the data, while the MFD can only display six waypoints depending on the selected

range. The waypoint to which the aircratl is flying is magenta in color. All other

waypoints are white. The DAT button on the MC-800 will add the following

information to the display if it is avai lable from the long-range NAV system.

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Waypoint identification (either number or name)

Distance to TO waypoint in nautical miles

ETA in GMT at the TO waypoint, if available, or lTG

(1) Aircraft Symbol

The aircraft symbol provides a visual cue as to the aircraft position in relation to the

desired track.

(m) Crosstrack Deviation

Crosstrack deviation indicates the deviation in nautical miles to the right (R) or left

(L) of the desired track.

(n) Displacement Line

Displacement Line indicates the position of the designator relative to the nose of the

aircraft.

(o) Designator

6. E.

(3)

MFD Plan Mode (See figure 6-17.)

A unique NAV PLAN format features a “true north-up” orientation in which the aircraft is

positoned with respect to the NAV route and progresses along the route, while the

maximum range is depicted by a circle around the outer perimeter. The north-up

orientation enhances the flight planning function and further clarifies the aircraft

relationship to the programmed route. In this display, the designator is homed to the TO

waypoint and both appear in the center of the display. The aircraft symbol is still plotted

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at present position (if present position is on the display) and is oriented with respect to

heading.

if the designator is moved from the TO waypoint, the designator symbol will remain in the

center of the display while the designator course/distance annunciation in the lower right

corner will be from the waypoint. The designator remains in the center during SKP and

joystick operations. Weather is not available in the PLAN mode, so range is controlled

solely from the MC-800. Other operations are the same as for MAP mode.

AIRPORT

SELECTEDNAV

SELECTEDNAV

ANNUNCIATOR

SOURCE

SOURCE

“NORTH-UP’

DISTANCETO

IDENTIFER

6. E.

(4) MFD Checklist Display

The MFD Symbol Generator is capable of storing and displaying 200, 400, or 800 pages

of text. These pages are stored in controlled internal PROM with content as defined by

the aircraft operator. Page composition is 12 lines with a maximum of 24 characters per

line, All text is stroke-written for sunlight readability.

The NORM button on the MC-800 MFD controller provides entry into the normal checklist

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display function. The normal checklists are arranged in the order of standard flight

operations. Use these buttons to access the normal checklist index page that contains

the lowest order incomplete and unskipped checklist with the active selection at that

checklist. The SKP, RCL PAG, and ENT buttons and the joystick on the MC-800 MFD

controller provide centrol of this function.

The EMER button on the MC-800 provides entry into the abnormal and emergency

checklist display. Actuation of EMER results in the presentation of the first page of the

abnormaVemergency master index. When a selection is made, an index, arranged by

aircraft systems, is presented. The crew can then select the l isting for the malfunctioning

system area, which in turn wil l provide access to the specific malfunction checklist. The

format of the MFD checklist very closely follows the aircraft’s approved abbreviated

checklist.

Under EMER conditions the SKP, RCL, PAG, and ENT buttons and the joystick perform

as described for NORM with the exception of the action taken upon completion of the

checklist. All checklist i tems are removed from the page and “EMERGENCY

PROCEDURE COMPL~E” is written below the amber checklist title. This is cleared

when the index is selected.

6. E. (5)

EFIS Backup Modes

EFIS backup is provided by the MFD as an addition to the existing EFIS reversionary

modes. This method has the following advantages:

The pilot can cope with EFIS failures through the EFIS controller and maintain the

MFD for checklists, weather radar, and enhanced mapping. Alternately, the pilot

can satisfy dispatch requirements for certain flight regimes through MFD backup of

the EFIS failures and, in this instance, forego the normal MFD functions.

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The MFD does not itself become a dispatch critical item.

Selection of EFIS backup by the MFD is accomplished by the rotary MOD selector switch

on the MC-8OO MFD Cent roller. Normal MFD functions are available in the MFD position,

and EFIS backup modes are obtained by selecting the HSI or SG positions. The HSI and

SG positions are spatial ly oriented to the side of the cockpit concerned.

HSI - Selection of this position will result in an HSI display on the MFD.

Composition of the HSI will be determined by the EFIS DC-81 O Display Controller.

SG - Selection of this position will result in replacement of the EFIS symbol

generator by the MFD symbol generator for the EFIS displays. In this case the

MFD CRT will be blanked. Composition of the EFIS displays will be determined by

the EFIS DC-81 O Display Controller.

6. F. SG-816 Symbol Generator (See figures 6-18 and 6-19, and table 6-3.)

The SG-816 Symbol Generator (figure 6-1 8) is the heart of the EFIS. It receives heading,

attitude, and short- and long-range navigation sensor and weather radar inputs. It also receives

mode logic inputs from the flight guidance computer. All inputs are processed and transmitted to

the ED-800 Electronic Displays as a function of the selections made on the DC-81 O and GC-81 O

Controllers. Leading particulars for the SG-816 Symbol generator are listed in table 6-3.

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The following paragraphs describe the operation of the symbol generator with reference to block

diagram, figure 6-19.


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15.78 inches (400.81 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.91 inches (124.71 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...7.62 inches (193.55 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15.01b,20z(6.86 kg)

8/11/2019 spz-8000 difcs


Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 Vdc,70Watts (maximum)

Mating Connectors:

Jland J2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DPX2MA-106S-106P-33B-OOO2

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

Tray Model MT-800 ,Part No. 7003272-90 1,

Boxmount Tray, Part No. 8314,

or Barry Tray, Part No. 93995-1


Leading Particulars

Table 6-3

6. F. (1)

Display Interface

The display interface generates the signals for both the EADI and EHSI simultaneously.

That is, one of the displays is in the stroke or vector mode while the other display is being

6,

F. (2) Vector Generator

The vector generator responds to commands from the display CPU to create the digital

deflection and video signals used by the display interface in the stroke mode. The heart

of the vector generator is a microprogrammable state machine that controls the action of

this circuit’s hardware. Included in the vector controller’s repertory of instructions is

character creation, character rotation, character initial position, and vector length. The

vector accumulator is a combination of registers, adders, multiplexer, and memory

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necessary to carry out the instructions of the vector controller.

The ping-pong RAM is two identical storage areas used alternately by the vector controller

and the display CPU. The display CPU writes into one area while the vector controller

reads from the other. At the end of a frame the storage areas “ping-pong.” This circuit

al lows the display CPU and the vector control ler to use this memory simultaneously

without interfering with each other, affording much higher operating speeds. The SG-816

uses two vector generators because it has to perform more functions in the stroke mode,

such as filling in the pointers on the EHSI.

(3) Raster Generator

The raster generator creates the attitude sphere (Horizon) used on the EADI and the

weather radar (WX) overlay that appears on the partial compass HSI display. The WX

memory includes the WX interface circuits and the weather radar refresh memory. The

horizon generator accepts pitch and roll information from the display CPU and generates

the color coded (blue for sky, brown for ground) horizon video signals. The timing

generator creates the clocks and the start and stop pulses that synchronize the video

pulse trains to the deflection ramps.

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Honeywell

phl”~JANCE

CITATIONVll

O V. 1 T HR U W /. ? - >

v o T HRu v -F , DC-O T HRU Do .F , 0 1. 0 ‘ >

THFVJ D1. F. DV-El T li RU DV F

012- 0 THRU 012-F->

P PO-O Tii RU P PO .F - ~

\

1

I

+L

M 1/0 CPU A4 DISPLAY CPU

>

*

m

EEl

G

h

E

OI VIOEO

ORIVERS

AOI

O/A CONVERTOR

ANO

DEFLECTION

ORIVER

1

-+

RS 1

IRS

2

v

~Ao’

‘“’ O’F

[

AOIFLYBACK

AOICLK

A OI R S

~ ;::;SR

AOI~

CS AO I ~

47ASCB

SI VIOEO

ORIVER

HSI O,A

CONVERTER

ANo

DEFLECTION

ORIVER

Esl

vECTOR

LE-1

- - +&&d

8/11/2019 spz-8000 difcs


E

4

ACCUMULATOR

1

A2RASTER

GENERATOR

=3

wx INTERFACE

ANO MEMORY

TIMING

GENERATOR

HORIZON

GENERATOF

WX SERIALOATA—>

WXVIOEOOATA—--j

Oti

INNERMKR

MIOOLEMKR

LRN2 TO/FROM

L RN 2 M T O EV

>

LRN2 V ERT OEV

)

OUTER MKR

HOG SYNC

NAV TOFROM

NAV

CRS

OEV

GS OEV

v N Av v ER T O EV

RAO ALT

sPEEO

CMO

PITCH CMO

ROLL

CMO

RATE OF TURN

LRN1 TOFROM

LRN I LAT DEV

LRN1 vERT OEV

A9 POWER

SUPPLY

~ +28 VOC INPUT

~ +28 VOC COMMON

~ +5 Voc

~ -5 Voc

~ 15VOC

~ -15VOC

B

N O 2 P IT CH

N O 2 R OLL

No 2 1- 10 0

LR2 BRGIACT

LRN20SRTRK

1>

m

NO

1

PITCH

N O 1 R OL L

NO 1 H OG

LR1 BRG ACT

LRN1 OSRITRK

AOF BRG

NAV ERG

CRS S EL

HOG SEL

m

m

SG-816 Symlml Generator Block Diagram

Figure 6-19

pages 191/1 92

6. G. MG-816 MFD Symbol Generator (See figures 6-20 and 6-21, and table 6-4.)

The MG-816 Symbol Generator (figure 6-20) receives heading, attitude, short- and long-range

navigation sensor, and weather radar inputs, as well as mode logic inputs from the flight

guidance computer. The MG-816 symbol generator input ports are connected in parallel with

the pilot and copilot SG-816 Symbol Generator input ports. All inputs are processed and

transmitted to the ED-800 MFD display as a function of the MC-800 MFD Controller when in the

MFD mode. When in the EFIS backup modes, the display functions are controlled by the

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DC-81 O Display Controller.

The MFD Symbol Generator also has a removable checklist module located on the front of the

MG-816 Symbol Generator. This module is programmed to each customer’s operating

requirements as an available option; in this case, a standard Citation VII checklist. Leading

particulars for the MG-816 Symbol Generator are listed in table 6-4.


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..16.03 inches (407.16 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7.53 inches (191.26 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..7.62 inches (193.55 mm)

Weight (maximum), . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 25.Olb(ll.34kg)

8/11/2019 spz-8000 difcs


Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...28 Vdc,120Watts (maximum)

Mating Connectors Jl, J2, andJ3. . . . . . . . . . . . . . . . . . . . DPX2MA-A106P-A1O6P-33B-OOO1

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

Tray Model MT-61 O, Pati No. 7007246-901

MG-816MFD Symbol Generator

Leading Particulars

Table 6-4

The following paragraphs describe operation of the Symbol Generator with reference to the

block diagram, figure 6-21.

6.

G. (1)

Display Interface (A2)

The display interface supplies the signals for both the EADI and EHSI at the same time.

Once every 1/60 of a second, the EADI starts its vectoring format. At the same time, the

raster generator begins its EHSI raster format. At the end of the EHSI raster format, a

In the stroke mode, 12-bit digital signals from the vector generator are converted to

analog signals and steered by an analog multiplexer to the proper X and Y output

amplifiers. At the same time, the video drive sets the corresponding video signals from

the vector generator and directs them through line drivers to the display that is in the

stroke mode. In the raster mode, timing signals from the raster generator start the high-

and low-speed ramps which, are steered to the other set of X and Y output amplifiers by

the analog multiplexer. The digital multiplexer in the video driver presents the raster video

signals to the line drive connected to the display that is in the raster mode.

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6. G. (2)

Vector Generator (Al )

Each of the two identical vector generators responds to commands from the display CPU

to create the digital deflection and video signals used by the display interface in the stroke

mode. The heart of the vector generator is a microprogrammable vector controller that

controls the action of this circuit’s hardware. Included in the vector controller’s instructions

is character creation, character rotation, character initial position, dash generator, variable

writing speed controller, and vector length. The vector accumulator is a combination of

registers, adders, multiplexer, and memory necessary to carry out the instructions of the

vector controller. In the ping-pong RAM, two identical storage areas are used alternately

by the vector controller and the display CPU. The display CPU writes into one area while

the vector controller reads from the other. At the end of a frame the storage areas “ping-

pong.” This circuit allows the display CPU and the vector controller to use this memory at

the same time without interfering with each other, allowing for much higher operating

speeds.

(3) Raster Generator (A4)

6. G. (5) 1/0

CPU (A6)

The essential element of this ACA is a 16-bit monolithic microprocessor. This processor

receives data from the triple shared RAM found on the ARINC ACA and also from the

shared RAM found on the ASCB ACA. The data is then transferred to the DISPLAY

processor, ARINC processor, and the ASCB processor for display ancYor output to other

systems.

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(6)

DC Variables (A9)

This ACA converts the synchro, resolver, and variable dc inputs to a digital format for use

by the remainder of the system. The synchm sample and hold circuit has buffers and

analog multiplexer sufficient to send 11 synchro signals in two channels. The two

channels of synchm signals are converted from three-wire to sine/cosine signals by an

electronic Scott Tee. The sine and cosine signals are then synchronously peak detected

and converted to digital numbers by the 12-bit A to D. In addition, the buffer and MUX

circuitry isolates and multiplexes up to 11 variable dc signals, which are subsequently

converted to digital signals by the 12-bit A to D.

(7) ASCB (A7)

The major function of this ACA is to service two system ASCB buses, two private line

ASCB buses, and the Proline II intetface. This is accomplished by a 4-to-2 input

multiplexer drii ing dual ASCB processing channels, which include Manchester coding and

HDLC protocol encoder/decoder. This circuitry also includes a DMA control ler working

with an 8086 microprocessor and a part of RAM shared with the 1/0 CPU. In addition, a

Honeywell

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MG-816 MFD Symbol Generator Block Diagram

Figure 6-21

Pages 197/1 98

.

6. G. (1O) Input Multiplexer ACA (Al O)

The input multiplexer selects pilot or copilot side systems and directs them to the various

EFIS cards contained within the Symbol Generator,

(11) Checklist Driver ACA (Al 1)

The checklist driver provides the interface to the checklist module and contains additional

switching circuitry.

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(a)

Checklist Module

A feature of the MFD/WX system is the easily removable checklist module. This

module may be removed without breaking the unit’s seal and is not required for the

continued operation of other MFD functions. This allows the checklist to be

modified without the need to return the entire Symbol Generator. The checklist

module contains enough memofy and control circuitry for up to 800 pages of

checklist text.

NOTE: If the MG-816 MFD Symbol Generator is sent in for repair

or exchange, the checklist module has to be removed and

retained for installation in the replacement MG-816.

(b) Symbol Generator Removal

The MG-816 Symbol Generator provides continuity for EFIS CRT secondary signals

and EFIS outputs in the event of either an MFD failure or loss of power. However,

6. H.

DC-81 O Display Controller (See figures 6-22 and 6-23, and table

6-5.)

The DC-81 O Display Controller (figure 6-22) provides the pilot with a convenient method of

controlling EFIS display formatting modes, such as:

displayed sensors

display dimming

self-test

radio altitude decision height setting

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c full or partial compass display

c single cue (SC) or cross pointer (CP) selection.

The DC-81 O has two bearing pointer source selectors, decision height knob, separate EADI and

EHSI master dim controls, self-test switch, and seven momentary pushbuttons located on the

front panel. Leading particulars are provided in table 6-5.

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6.

H. (3) Ground SpeeWTime-To-Go (GSPD~G) Button

By pressing the GS/TTG button, ground speed or time-to-go will alternately be displayed in

the lower right comer of the EHSI.

(4) Single Cue/Cross Pointer (SC/CP) Button

By pressing the SC/CP button, the Ilight director command cue(s) maybe toggled back

and forih from single cue configuration to the cross pointer configuration.

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(5)

Fl ight Director Command (FD CMD) Button

Command cues are displayed on both EADIs whenever a flight director mode is selected.

The command cue on the non-HSl selected side may be biased from view by pressing the

FD CMD button on that side. Pressing the button a second time will restore the cue to the

display.

Pressing the HSI SEL button on the GC-81 O flight guidance controller resets everything to

the initial state. In addition, during GA and dual HSI approaches, command cues are

displayed on both sides regardless of the previous state, and the FD CMD buttons on both

DC-81 0s are locked out during the dual HSI approach.

(6)

Navigation (NAV) Button

By pressing the NAV button, VOR/LOC information is selected for display on the EHSI.

(7) Flight Management System (FMS) Button

6.

H. (9) Dim Controls

The dimming system employed by the EFIS is semiautomatic. Two inputs contribute to

the overal l display brightness of each ED-800 Electronic Display:

Ambient light sensed by the photosensors on each ED-800

Setting of the dimming controls

The DIM pot sets the nominal intensity for each display. The photosensors located on

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each ED-800 cause the light output of each display to be modulated about the nominal

intensity as a function of the l ight incident on each display.

(a)

(b)

(c)

ADI DIM Control

The ADI DIM control dims the raster and stroke wriing on the EADI. Turning the

control to the OFF position causes the EADI to go blank and the composite mode

to be displayed on the EHSI.

HSI DIM Control

The HSI DIM control dims stroke writing and the raster on the EHSI. Turning the

control to the OFF position causes the EHSI to go blank and the composite mode

to be displayed on the EADI.

WX DIM Control

The WX DIM control dims only the raster on the EHSI that contains weather radar

6.

H. (12) TEST (TST) Button

By pressing the test button, the displays will enter the test mode. In the test mode, flags

and cautions are presented along with a check of the radio altimeter. The following test

routine is displayed:

NOTE” Test of the EFIS is only functional on the ground. Radio altimeter test is functional at all

‘“ times except during GS CAP/TRK or GP CAP/TRK.

.

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.

Course select, heading select, and DH set digital displays are replaced by amber

dashes.

AIT and HDG displays are flagged.

All pointers/scales are flagged with a red X.

All heading related bugs/pointers are removed.

Command bars are biased from view.

Radio altimeter digital readout displays radio altimeter self-test value. (Slews to 100

feet for Honeywell radio altimeter.)

Comparator monitor annunciates AIT, HDG, and ILS (if ILS sources are seleded

on both sides) or MLS (if MLS sources are selected on both sides).

The word TEST (in magenta color) is annunciated in the lateral capture location on

Honeywell

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STRAPPING

ALUM3K+3

Pages 198.TI1 98.8

6. 1. MC-800 MFD Controller (See figures 6-24 and 6-25, and table 6-6.)

The MC-800 MFD Controller provides the means by which the pilot can control the MFD display

modes and format. The following paragraphs describe the controller functions.

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AD-10336


Figure 6-24

Dimensions (maxirmm):

Length . . . . . .

Width . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6.59 inches (167.39 mm)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5.75 inches (146.05 mm)

Honeywell ~$gvf”c’

6. 1.

(1)

(2)

MAP/PLAN Button

The MAP/PLAN button alternately selects the heading up MAP display or the North up

PLAN mode for display.

Source (SRC) Button

The SRC button alternately selects the source of long-range navigation data for mapping.

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(3)

(4)

(5)

Weather (WX) Button

The WX button is used to call up weather radar returns on the MFD map display. When

weather is displayed, the map range is controlled by the WC-870 Weather Radar

Controller.

Normal (NORM) Button

The NORM button provides enty into the MFD’s normal checklist display function. The

normal checklist is arranged in the order of standard flight operations. Button actuations

cause presentation of the normal checklist index page that contains the lowest onder

incomplete and unskipped checklist with the active selection at that checklist. The SKP,

RCL, PAG, and ENT buttons and the joystick provide control of this function.

Emergency (EMER) Button

The EMER button provides entry into the MFD’s emergency checklist display function.

Actuation of EMER results in the presentation of the first page of the highest priority callup

On an index page - actuation results in display of the checklist corresponding to the

active index line selection. The checklist is presented at the page containing the

lowest order incomplete item with the active selection at that item. If the checkfist

had previously been completed, the system forces all items in the checklist to

incomplete and presents the first page of the checklist with the active selection at the

first item.

On a checklist page - actuation forces the active selection to complete and advance

the active selection to the next incomplete item. If ENT is actuated with the active

selection at the last item in a checklist, the operation depends upon the completion

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status of the checklist.

If the checklist is not complete (one or more items skipped) the system presents the

page containing the lowest order incomplete item with the active selection at that

item.

If the checklist is complete (all items complete) the system presents the index page

containing the next higher order checklist with the active selection at that checklist.

Joystick - The joystick provides additional paging and cursor control. Each

actuation results in the action described:

UP moves the active selection to the lower order item

.

DOWN moves the active selection to the next higher order item (this is

identical to SKP)

6. 1.

(7)

VHF Omni Range (VOR) Button

The VOR button is used to add VOWDME symbols to the map and plan displays.

(8)

Data (DAT) Button

The DAT button is used to add long-range navigation information to the map and plan

displays.

First actuation will add the following data to the lower right corner of the display:

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- Waypoint identification

- Estimated Time of Arrival (ETA) in Greenwich Mean Time (GMT) at the TO

waypoint if known; otherwise, Time-To-Go (lTG) to the TO waypoint.

Second actuation - If no destination information is known, this step shall be to data

OFF. However, if destination identification, ETA, or lTG is known, this step shall

replace the TO waypoint data as described above with the destination data.

I f some destination data is known but the waypoint identification is not, the mnemonic

DEST shall be used in place of the waypoint identification.

(9)

Airport (APT) Button (Not applicable to -925 units)

The APT button is used to add airport designators to the map and plan displays.

r

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8/11/2019 spz-8000 difcs


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Block Diagram

Figure 6-25

Pages 198.1 3/198.14

6. J.

RI-206S Instrument Remote Controller (See figures 6-26 and 6-27, and table 6-7.)

The Instrument Remote Controller interfaces with the Symbol Generator to provide heading and

course selection. Activation of the PULL SYNC switch causes synchronization of the heading

bug to present heading (lubber line). The PULL DIR switch allows automatic selection of a TO

direction desired VOR course having zero deviation.

e

COURSE 1

HANDING COURSE 2’

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Figure 6-26


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..4.31 inches (109.5 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 inches (146.1 mm)

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7. DFZ-800 Dual Flight Guidance System

A.

FZ-800 Flight Guidance Computer (See figures 7-1 and 7-2, and table 7-1.)

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Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..15.13 inches (384.3 mm)

Wdh . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.91 inches (124.7 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..7.62 inches (193.5 mm)

Weight (approximate) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.llb(6.04kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 28 Vdc.40Watts

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Mating Connector:

J1, J2 . . . . . . . . . . . . . . . . . . . . . . . . . .

Cannon Part No. DPX2MA-67S- I06P-33B-0002

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

Tray, Honeywell Part NO. 7003272-901

FZ-800 Flight Guidance Computer

Leading Particulars

Table 7-1

The FZ-800 Flight Guidance Computer (FGC) processes information about the aircraft actual

attitude versus a desired atti tude as a function of selected flight mode to produce autopilot pitch,

roll, andyawcontrol outputs and flight director pitch and roll steering command outputs. In

addition to the modes selectable on the GC-81 O Control ler, the computer wil l produce pitch and

roll control outputs for any flight director mode except go-around.

The FGC has a dual processor architecture, each processor performing different control and

All the 1/0 is memory mapped, and each processor individually controls its own analog and

discrete input/output transfers with the exception of the serialized discretes. Discretes fall into

two categories: direct and serial ized. The latching of the serialized discrete inputs is under the

control of the A-processor only. Once the inputs are latched, however, each processor has

independent access to them. The serialized discrete outputs (to the control panel) are solely

under the control of the A-processor.

The Heartbeat Monitor and Power Supply Monitor Interlocks ensure disengagement of the FGC

in case of a processor failure, a software failure, a power supply failure, or a power outage. The

Sewo Drive Engage Interlocks ensure that the flight control functions can be activated only if al l

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the monitors are satisfied. The flight controls are output through the Trim, A/P, and Y/D servo

drives.

The Flight Director Interface outputs the analog bar commands and validity annunciations

computed by the A-processor.

Honeywell

MAINTENANCE

MANUAL

CITATION Vll

SCRATCH

PAD

4

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CLOCK 2

8/11/2019 spz-8000 difcs


t ‘

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8/11/2019 spz-8000 difcs


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GC-81 O Flight Guidance Controller

Figure 7-3


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..6.50 inches (165.l mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5.75 inches (146.l mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 2.63 inches (66.7 mm)

The GC-81 O Flight Guidance Control ler is used to engage/disengage the system, select the

operating modes, and select the HSI and DADC being used to interface with the flight guidance

computer. The pitch wheel is also part of this unit. The function of each switch or control is

described in the following paragraphs.

7. B. (1) AP and YD Buttons

The AP button engages autopilot and yaw darnper functions simultaneously, but

disengages only the autopilot functions. The YD button engages the yaw damper only

and disengages the yaw damper and autopilot. The active channel is annunciated by the

lighted A and B located on either side of the AP and YD buttons. When the autopilot and

8/11/2019 spz-8000 difcs


(2)

yaw damper systems are in a normal no failure condition, the pilot channel is

automatically selected as the active channel and the (A) annunciator on the AP and YD

engage switches will be lighted. If the pilot wishes to select the copilot channel (right

FGS) as the active channel, he can press the AFCS B button on the instrument panel.

When the system is engaged, the (B) annunciator on the AP and YD switches will be

lighted indicating that the right channel is active. The AFCS A or AFCS B buttons can be

used to select the active FGC.

NOTE: The autopilot cannot be engaged on the ground.

HSI SEL Button

The HSI SEL button alternately selects either the pilot’s or copilot’s HSI and DADC data

for lateral and vertical guidance to both flight guidance computers. The DAFCS power-up

logic selects data from the pilot’s HSI and DADC. When the system is transferred to the

alternate HSI and DADC, all flight director modes are cancelled. Operating modes must

7. B.

(5)

(6)

(7)

APP (Approach) Button

The APP button selects the appropriate gains to arm and capture the lateral deviation

sgnal for VOR, LOC, and AZ, and both lateral and vertical navigation signals for ILS and

MLS to meet approach criteria.

BC (Back Course) Button

The BC button commands the flight director computer to track the Iocalizer back course.

ALT (Altitude) Button

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(8)

(9)

The ALT button commands the system to hold the current altitude. Capturing the altitude

displayed on the AL-801 Alti tude Preselect Control ler wil l al low the system to maintain that

altitude.

VNAV (Vettical Navigation) Button

The VNAV button commands the system to follow the vertical flight path guidance from a

compatible long- range navigation system, when selected.

BANK Button

The BANK button commands the guidance computer to use reduced bank angle (17

degrees) when in the HDG mode. Automatic bank angle change occurs at 34,275 feet

MSL. During a climb, bank switches to half bank; during descent, bank returns to full bank

values.

r

———. ———— ———. ———. ———— ———— ————

1

EGGE-LIGHTINO

UTTON BACK4.IGHTINO BUL13S

POWER

(NOT OUALI

““””-

1

::=

FDMOOE

SWITCHES

~

I

i

AP

PARALLEL

I

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(+

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INPUTS

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PARALLEL

SHIFTERS INPUTS

ANNJN

(Pf)TE)

VALI,Y1

I

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I

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LAMP TEST 1

ANNuN VALlO1

AOVISORV OISPLAY

1AND2 PSDISCRETES I

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SERIAL

OATA OUT

SERIAL DATA

T O F GC 1

t

5V—

ISC L21SCFIETES

———

ISC

LSCRSTES

TOIFF40M

8/11/2019 spz-8000 difcs


-’;’L’GHTS

I

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HANNEL

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1

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CHANNEL

i

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INPUT

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1 A NO 2 P B A RM 1

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SHIFTED TO PROVIOE

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INPUT TO TNE PISO

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-—

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L ———— ———— ———— —. —.. ———— ———— ———— J

m *% U

GC-81O Flight Guidance Controller Block Diagram

Figure 7-4

Pages 198.25/1 98.26

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8/11/2019 spz-8000 difcs


F

——— ——— ——— ——

-1

: SERVO ORIVE ASSEMBLYPOWERLOOP

,.

I

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EXCITATION

(CONTROLLEDBY

AUTO PILOT

ENGAGEANOIOR

TOUCHCONTROL

=EERING)

.- . ...- ------ . ..

?ROPRIATE COMBINATION

VE ASSEMBLYDASH

SHIELDGNO K 1’

8/11/2019 spz-8000 difcs


i=

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——— ——— .—

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8.

PRIMUSXB 870 Weather Radar System

WARNING: HEATING AND RADIATION EFFECTS OF WEATHER RADAR CAN BE HAZARDOUS

TO LIFE.

Maximum Permissible Exposure Level (MPEL)—Personnel should remain at a distance greater than R

(figure 8-3) from the radiating antenna in order to be outside of the envelope in which radiation

exposure levels equal or exceed 10 mW/cm2, the limit recommended in FAA Advisory Circular AC No.

20-686, August 8, 1980, Subject: “Recommended Radiation Safety Precautions for Ground Operation

of Airborne Weather Radar.” The radius R to the MPEL bounda~ is calculated for the radar system

on the basis of radiator diameter, rated peak-power output, and duty cycle. The greater of the

8/11/2019 spz-8000 difcs


distances calculated for either the far-field or near-field is based on the recommendations outlined in

AC No. 20-686.

The American National Standards Institute (ANSI) in their document ANSI C95.1 -1982, recommends

an exposure level of no more than 5 mW/cm2.

Honeywell, Inc. recommends that operators follow the 5 mW/cm2 standard. Figure 8-1 shows MPEL

for k-th exposure levels.

A~R~~EA&~

RADOME

AIRCRA~ LU6BERLINE

I

12

IN

I 1~

9

FT

3U3CKl

*

Honeywell

8. A.

WU-870 Antenna and Receiver~ransmitter

MAINTENANCE

MANUAL

CITATION Vll

Unit (See figures 8-2 and 8-3, and table 8-1.)

The WU-870 Antenna and Receiver/Transmitter Unit is an integrated unit that inmrporates

transmitter, receiver, and antenna into a single unit. The Antenna and Receiver/Transm”Mer Unit

accepts either a 10- or 12-inch flat plate radiator with transmitter and receiver components

mounted on the rear of the antenna. The remainder of the circuit~ is contained in the

electronics package that forms the Antenna and Receiver~ransmitter Unit base. The Antenna

and Receiver/Transmitter Unit transmits and receives X-band radio frequency energy for the

purposes of weather detection and ground mapping. The 9345*30 MHz transmitted signals

are sent directly to the antenna from the transmitter circuitty, which is mounted on the rear of the

antenna. Echo signals received by the antenna are applied directly to the receiver, which is also

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mounted on the rear of the antenna. The receiver and processing system processes these

signals by encoding them into one of four levels depending on their intensity, scan converts

them, and outputs the scan converted data to the various display systems.


Base Diameter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8.66 inches (22.0 mm)

Height (Antenna flat) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.00 inches (30.5 mm)

Height (Antenna full ac) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.00 inches (30.5 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.01b(6.40 kg)

Prime Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

+22to +32 Vdc, 90 Watts (maximum)

Antenna

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Size . . . . . . . . . . . . . . . . . . . ..i . . . . . . . . . . . . . . . . . . . . . . .

12-inch flat plate radiator

Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Line-of-sight, +30 degrees

Tilt . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. fls degrees

Scan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Full -120 degrees (~60 degrees)

Sector- 60 degrees (*30 degrees)

Roll Axis Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Recessed onfront panel

Transmitter

Frequency .

Power . . . .

Pulse Widths

PRF . . . . . .

Receiver

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9345*30 MHZ

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1.3kW, nominal, magnetron

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1.2, 1.5,2.4,4.8,9, 18, and27ys,

determined by selected range

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

120, 240, 360, and 480 Hz,

determined by selected range

MODULATOR

t

VIDEO

TR. LIMITER

DETECTOR AMpL

Pw

CIRCULATOR

MAGNE~ON

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GND FOR 203 Mll

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+,

STC

XSTC

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GNO FOR DUAL CONTROL OPERAW3N

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~F44;70R INVERTEO

ON,OFF

PROCESSOR

—

—

lb

r

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BuS

8/11/2019 spz-8000 difcs


Dimensions (maxinwm):

Length {from rear of bezel) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.0inches(177.8 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.75 inches (146.1 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.87 inches (47.5 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..l.91b(0.86 kg)

Power Requirements:

Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

+22to+32Vdc, 5.8 Watts (maximum)

Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..28 Vdcat6.O Watts (nominal)

8/11/2019 spz-8000 difcs


or 5 V ackic at 4.6 Watts (nominal)

Mating Connector (Jo) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS27473E14B-18S

with strain relief MS27506-B14-2

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Dzus Fasteners

WC-870 Weather Radar Controller

Leading Particulars

Table 8-2

8. B.

(1)

Range Buttons

Radar operating range selections are made with two momentary-contact pushbutton

The GCR mode has the fol lowing display l imitations.

Q Will not remove all of the ground return.

Removes some of the weather returns.

Effectivity is reduced as the antenna scans away from dead ahead.

The scintil lation frequency of the ground radar returns is fewer than that of rainfall radar

returns. A digital frequency filter is used to separate ground returns from the rainfal l

returns, and only the rainfal l returns are displayed when the GCR mode is selected.

Since some of the rainfall returns fall into the same spectrum as the ground returns, there

is some loss of weather return in the GCR mode. As a resuft, the weather presentation in

this mode cannot be considered calibrated. However, the GCR mode gives the pilot a

8/11/2019 spz-8000 difcs


dramatically improved look at weather in terminal areas or mountainous terrain where it

may be necessary to titt the antenna toward the ground to see weather ahead. GCR is

operational in WX mode and selected ranges of 50 NM or less.

Selecting the 100, 200, or 300 mile range or the tuttxdence detection (TRB) mode turns

off the ground clutter reduction. The GCR legend is deleted from the mode annunciation

and variable gain is engaged if previously selected. Subsequent selection or ranges of

50-miles or less re-engages GCR. If not already selected, GCR forces the radar into

preset gain.

8. B.

(4) TGT (Target Alert) Button

The TGT button is a momentary alternating action pushbutton switch that enables and

disables the target alert mode of the radar system. Target alert is selectable in any WX

range except 300 NM. When selected, target aleft monitors beyond the selected range

and 7.5 degrees on each side of the aircraft heading. Also the target must have the

8. B.

(5)

The TGT button can also be used to override (turn off) the radar attitude stabilization.

The radar is normally atti tude stabil ized and automatically compensates for rol l and pitch

maneuvers. Attitude stabil ization is turned off by pressing the TGT button four times

within 3 seconds. Stabilization is turned back on by again pressing the TGT button four

times within 3 seconds.

SECT (Sector) Button

The SECT button is a momentary alternating action pushbutton switch that selects either

full azimuth scan (120 degrees) or sector scan (60 degrees).

8/11/2019 spz-8000 difcs


(6)

TILT/PULL AUTO Knob

The TILT/PULL AUTO knob is a single turn rotary control that varies antenna tilt between

15 degrees up and 15 degrees down. (Clockwise rotation tilts beam upward Oto 15

degrees; counterclockwise rotation ti lts beam downward O to -15 degress.) The range

between +5 degrees and -5 degrees is expanded for ease of setabiliiy. A digital readout

of the antenna tilt angle is displayed on the EFIS.

Pulling out on the TILT knob causes the system to enter the Auto Tilt mode. In Auto Tilt

the antenna tift is automatically adjusted with regard to the selected range and barometric

altitude. The antenna tilt will automatically readjust with changes in altitude anctlor

selected range. Afso note that while the radar system is in Auto Titt, the tilt control can

fine-tune the titt setting by *2 degrees. The digital readout will always show the

commanded til t of the antenna regardless of the til t command source (auto tiff command

or manual tilt command).

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8/11/2019 spz-8000 difcs


OEI

A

SECT

TGT

“z%) ““E?l -

>

/

g B:3F:1E4RTBo- TB6

-6o1 THRU -S04 AUTO TILT -

I

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1

13

BUFFER

/

A3U13

TBO - TB6

-601 THRU -604, - 622 RCT

“EL

1

‘A’Nw’pER~~+-’

I

-601 THRU

TILTWIPER

A2ul

AID

T

CONV

16

“ P AM P

TILT

/

-

A3U19

BUFFER

OE5-

A3U 16 TBO - TB6

OE3

T

A3U17

oE6-

-60S, -622 “X

I

I

If

96 KHZ

J

-0

.144

MHZ

Osc

A3U7.

1

v

A3Y1

,5

,5

t’

[

i

f

CONTROL

IABEL

192KHZ

PROM

PROM

A3U15 A3U12

A v

UART

‘ u

LINE

ERNALBUS DRIVER Scl

A3U6

A3U2

TORTA

I

OEI

DECODER

0E2

MUX

OE3

10F6

OE4

-E

4

A3U3

8/11/2019 spz-8000 difcs


SBY

m

PROM

I

MAP

3

(MODE)

I

FP

A2U7

TBo - TB2

TST

(U14)

I

I

rB2 - TB6

OE3-

1

BUFFER

A3U22

I

I

I

OCTAL

PROM

BUFFER +

( RN G) -

r

1

A3U26

A3U29

FP

+

*

L

L

26V DC~

r

PANELLIGHT

FP/RNG

OSC CIRCUIT

TST

A2U12,A2U9,A2U16

(U20, U16)

x

I

PULSE

+

+15VDC

I

1

POWER SUPPLY

26V-TO-5V

CONVERTER

CIRCUITS

t-

+ 5 V A C

AI

NOTE:

I

1. REFERENCE DESIGNATIONS ARE AS FOLLOWS:

PANELLAMP

CIRUIT

Al a Lo” voLTAGE powER suPPLY ccA

A2 = LOGIC 1 CCA

I

1

A3 = LCM31C CCA

2. PART REFERENCE DESIGNATIONS IN PARENTHESIS

ARE FOR CCA A2 PART NO. 7012655 .

AD-e9?3.m

WC-870 Weather Radar Controller Block Diagram

Figure 8-5

22-05-07

Pages 198.41/1 98.42

Jun 1/93

8. C.

WI-870 Weather Radar Indicator (See figures 8-6 and 8-7, and table 8-3,)

The WI-870 Indicator is a weather radar controller and electronic display integrated into a single

panel-mounted LRU (figure 8-6). Operation of all controls and switches on the Indicator are

identical to the controls and switches on the WC470 Controller. The display is a large format

five-inch diagonal color CRT similar to the one used in the MFD. When installed in place of the

MFD, the WI-870 Indicator provides all control functions for the weather radar system and

displays scan converted data processed by the Antenna and Receiverflransmitter Unit. In

addition, high-speed video input capability is provided through a separate Universal Digital

I ntertaceUDI) port to permit display from auxiliafy systems-such as Data NAV and th~ Lightning

Sensor System. Leading particulars are listed in table 8-3.

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Honeywell

Q

./” -- “-”

.— . .— ..—.- —-=

TRB

m

N $’

RANGE

a

GCR

n

4

i

I


Length (from rearof bezel) . . . . . . . . . . . . . . . . . . . . . . . . ...11.49 inches (291.85 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.83 inches (122.68 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..6.265 inches (159.13 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.01b(4.54 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . .

+22to +32Vdc, 36Watts (maximum)

Panel LQhting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..+28 Vdc, 0.200A (nominal)

8/11/2019 spz-8000 difcs


or5Vac/dc,O.750 A (nominal)

Mating Connectors:

J101 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. MS3126F22-21S

J102 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. KJGF14A35SN

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

Tray, Honeywell Part No. 7011359-901

W1-870 Weather Radar indicator

Leading Padiculam

Table 8-3

Table 8-4 describes the operation of the W1-870 Controller.

1. BRT

Single turn display brightness control that adjusts

4. Mode Switch

OFF

SBY

Wx

GMAP

Seven-position rotary switch which selects

primary radar modes.

Removes power from system.

Standby. Places system in non-operational

mode.

Places system in the operational Weather (WX)

mode.

Places system in the operational Ground Map

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FP

TST

5. TGT

6. RCT

(GMAP) mode.

Flight Plan. Permits extended range display of

navigational data provided through the Universal

Digital Interface (UDI) UDI pat.

Activates the system self test mode.

Momentary alternate-action pushbutton which

enables the Target Alerl function. This button

also disables STAB if pressed once and then

three more times within 4 seconds. To enable

STAB, repeat. When active, gain is forced to

preset.

Momentaty alternate-action pushbutton which

enables the REACT (RCT) function. RCT is

9. RANGE

A two-pushbutton range selection system with

permits range selection from 5 to 300 NM full

scale in WX, RCT, or GMAP mode or 5 to 1000

NM full scale in the Flight Plan mode. The up

arrow pushbutton selects increasing ranges while

the down arrow pushbutton selects decreasing

ranges. The last range is remembered when

switching between WX, RCT, or GMAP and FP.

Upon reaching maximum or minimum range,

further pressing of the pushbutton causes the

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range to rollover to minimum or maximum range,

respectively. If FSBY is wired to a

weight-on-wheels switch, the unit wil l be in

Forced Standby on the ground unless both

RANGE pushbuttons are pressed simultaneously.

10. AZ

Momentary alternate-action pushbutton which

permits displaying and removing azimuth marks

from the display.

11. SCT

Alternate-action pushbutton which selects either

full azimuth scan angle (120 degrees) or sector

azimuth scan angle (60 degrees).

WI-870 Control Functions

UDI

HORIZ

TIMING

REF

VERT

GENERATOR SWEEP

)Dl

-

UDI

;ELECT

CIRCUITS

B

STBY ~

1----

HoRlz

SWEEP

8/11/2019 spz-8000 difcs


‘f’uRE=c-

+zEl-

-4 1-

X-Y

COLOR

MEMORY

RANGEIAZ

CONTROL

.

MARK

BUS

CONTROL

GENERATOR

BUS

ENCODERI

DECODER

f

ALPHA-

NUMERIC DECODER/

GENERATOR PRIORTIZER

‘ORcEDsTBy~

ONOFF---La

WI-1370 Weather Radar Indicator Block Diagram

Figure 8-7

VIDEO

BfANK -

COLOR

VIDEO

AMPL

k

rl

TC

I H.v. I

.

0

PWER

SUPPLY

AD-6873 - m

22=05-07

Pages 198.471198.48

Jun 1193

9.

FMZ-800/900 Flight Management System (Optional~

A. NZ-820/920 Navigation Computer (See figures 9-1 and 9-2, and table 9-1.)

The NZ-820/920 Navigation Computer (figure 9-1) receives its FMS command data from the

CD-81 O Control Display Unit (CDU), and its FMS input data from the Avionics Standard

Communications Bus (ASCB), Radio Systems Bus (RSB), and DL-900 Data Loader. The

navigation computer contains the necessary power supplies, electronics, and database memory

to receive and process

sensor input information, while providing highly accurate position

information to the flight crew. Additionally, the navigation mmputer has the ability to remotely

tune all the radios on the aircraft, as well as provide a means for the flight crew to create and

store waypoints and flight plans. Leading particulars are listed in table 9-1.

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The NZ-820/920 Navigation Computer provides both lateral and vertical navigation guidance.

The navigation computer has a 320k byte (NZ-820) or 1.2 megabyte (NZ-920) internal navigation

database that is used for storage of waypoints, navaids, routes, airports, and other NAV data for

easy access by the pilot. The NZ-820 database only allows loading one of the available four

regions of data at a time. The NZ-920, with the expanded database, enables loading all four

regions of data to allow international operation without changing databases. An internal

keeps the clock and calendar running when power is removed.

battery

Honeywell

MAINTENANCE

MANUAL

CITATIONVll


Length . . . . .

Width . . . . . .

Height . . . . . .

Weight (approximate)

Power Requirements

Mating Connector:

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

17.03 inches (432.6 mm)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4.91 inches (124.7 mm)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7.62 inches (193.5 mm)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

14.8 lb (6.71 kg)

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

28 V dc, 65 Watts

8/11/2019 spz-8000 difcs


J1 . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

Cannon Part No. DPX2-67S-106P-33B-0089

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Tray. Paft No. 7003272-903

NZ-820/920 Navigation Computer

Leading Particulars

Table 9-1

The navgation computer can interface with three long-term sensors via ARINC 429 buses and

the ASCB. Each navigation computer can also connect to dual Proline II or Bendix/King DME

receivers and a single VOR receiver. The interface to the AH-600 AHRU, AZ-810 DADC,

FZ-800 FGC, SG-816 (EFIS), and MG-816 (MFD), is over the ASCB. Flight plans are also

transferred between navigation computers over the ASCB while the link to the CDU is over a

RS-422 ‘private-line’ interface. To provide high-accuracy long-range navigation, the navigation

The aircraft position is computed as a function of logic switches called navigation update modes.

The four position update modes are radio/ineti lal, radio only, inertial only, and dead reckoning.

The navigation mode hierarchy is a function of sensor and data availabi lity. The radidinertial

position is computed by using radio position.

The radio position is then combined with the calculated inertial velocity for computing

radlofinertial position. If valid rho/theta or omega data is being received on the ground, a radio

position wil l be computed. Sensor and radio data availabi lity defines the priority for each

navigation mode. The highest priority nav mode is radioiinertial, folfowed by inertial only, radio

only, and dead reckoning. These priorities are based upon sensor accuracies.

When the navigation update mode changes from one of the four modes listed to the no

navigation mode and the aircraft is airborne, the present aircraft position is frozen for display on

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the CDU PROG page for 2 minutes. After that, if the aircraft is airborne and the no navigation

mode is stil l active, the aircraft position is invalid for display. If an initial start-up power transient

occurs when the no navigation mode is active, the above display logic is invalidated and the

aircraft posit-on is not val id for display unless a new position can be mmputed.

The radm/inedial NAV mode is active when the following conditions are true:

Valid acceleration and angular data are available.

A valid radio position is computed from at least two DMEs, one VOR/DME pair, or an omega

sensor.

The radio position and inertial acceleration are combined in the AHRS velocity fi lter to compute

the north and east inertial ground speed components. These components are then combined

with radio position in a complementary filter to compute the radiofinertial position. The inertial

only NAV mode is active when at least one AHRS is providing valid acceleration and angular

Veri ical Navigation (VNAV) within the FMS allows the operator to define vertical path information,

which is then flown by the aircraft automatically when the proper ffight director mode has been

selected. FMS VNAV may be used throughout the flight. The VNAV can be utilized to climb on

the optimum IAS and automatically transition to the optimum MACH. Descents can be set up for

a path mode (pseudo glideslope) or programmed MacWIAS letdowns.

VNAV controls the vertical flight path by sending speed (CAS or Mach) or vertical speed targets

to the flight director. The fl ight director will limit the vertical and along path accelerations for

passenger comfort. Alti tude targets are also sent to the fl ight director, which uses its internal

alti tude capture logic to determine when to sequence to the attitude capture and hold modes.

The flight level change (FLC) function engages a speed on elevator climb or descent based on

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aircraft position relatiie to the altitude preselector. The operation of FLC is the same for VNAV

or fl ight director operation, the only difference being the selection of the speed target. VNAV

FLC uses the prestored MacWIAS values from the navigation computer; whereas, the basic flight

director FLC synchronizes to the current airspeed.

The VNAV function of the FMS is integrated into the various pages of the CDU display. The

primary locations of VNAV information are on the ACTIVE FLIGHT PLAN and PROGRESS

pages. The vertical definition of the flight plan includes speed, angle, and altitude constraints at

waypoints. VNAV will not function until all PERF INIT information has been programmed into the

CDU. If VNAV is not desired, simply omit the PERF INIT step of preflight. tt shoukt be noted

that the altitude preselector provides an active input into the FMS VNAV function. Since the

VNAV preflight computations for each waypoint are done with regard only to the alti tude

preselector, the cruise altitude from PERF INIT page 3 and altitude constraints in the flight plan,

it is suggested that all VNAV constraints should be defined in the ACTIVE FLIGHT PLAN priir to

PERF INIT. After PERF INIT, compliance with subsequently entered altitude constrains is

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a

GMT

CLOCK

CDU

INTERFACE

=?EiEl+

DATA BASE

E2 PROM

I

w

ARINC

429

XMIT

l)

S422

RCIIRS

9. B. CD-800/810 Control Display Unit (See figures 9-3 and 9-4, and table 9-2.)

The CD-800/81 O Control Display Unit (CDU) provides the primary means for pilot input to the

flight management system. It also provides output display for the navigation computer. The

CDU utilizes a full alphanumeric keyboard, as well as decimal, dash, and slash. Four line

selection keys are provided on each side of the CRT. Seven function keys are provided to allow

direct access to specific display pages. Annunciators are located in the top of the bezel to

advise the pilot of the system’s status.

The CRT in the CDU has nine lines of text 24 characters long. l%e top line of the CDU display

is dedicated as a title line and the bottom line is used as a scratchpad and to display messages.

A manual dimming knob is used for long-term dimming adjustments, while ambient light sensors

are used for short-term display brightness adjustments under varying cloudhunlight conditions.

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PHOTO

PHOTO

SENSOR

ANNUNCIATORS

/

SENSOR

/ \

CRT

DISPLAY

~ACllVl I I I PI AN l/bj~>l

LE~

LINE

SELECT

H

17it

Ml I

02(7)0

93.(3NM

1)1SI

KSI C

HI

61(31

RIGHT

LINE

SELECT


Length (CD-800 ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..7.5Oi5ches(l9O.5 mm)

Length (CD-810 ) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..10.00 inches (254.O mm)

Width (both) . . . . . . . . . . . . . . . . . . . . ... . . . . . . . . . . . . . ..5.7linches(l46.l mm)

Height (CD-800) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . , . . ..6.75 inches (15mm)mm)

Height (CD-810) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..7.50 inches (190.5 mm)

b

Weight:

CD-800 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9.4 lb (4.27 kg)

CD-810 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

12.7 lb (5.76 kg)

Power Requirements:

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Primary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 Vdc,40 Watts

Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..5Vac

User Replaceable Parts

Knob, Brightness Control . . . . . . . . . . . . . . . . . . . . . . . . . Honeywell Part No. 7008508

Setscrew(.112-40xl/8”) . . . . . . . . . . . . . . . .

MS51021-9, Honeywell Part No. 0455-128

Mating Connector:

J1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. MS3126F22-55SX

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Standard Dzus rail

9. B.

(3)

Annunciators

The six annunciators located at the top of the CDLJ keyboard panel operate independently

from the CRT and keyboard. Lighting of the annunciators is initiated by the Navigation

Computer via the RS-422 serial data link. The two mbrs used for annunciations are white

and amber. White indicates an advisory annunciation, and amber indicates an alerting

annunciation. The following paragraphs describe each annunciator:

(a) Display (DSPLY) Annunciator

The DSPLY annunciator is an advisory (white) that lights when the CDU is

displaying a page that is not relative to the current aircraft lateral or vertical fl ight

path. The DSPLY annunciator wil l light under the following conditions:

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When displaying a flight plan page other than page 1.

When displaying a stored fl ight 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 an alert (amber) that lights when the FMS is navigating via

the DR mode, which is defined to be the loss of radio updating and the loss of all

position sensors. The DR annunciator will l ight under the following conditions:

When the FMS has been operating in the DR mode for bnger than 3 minutes.

9. B.

(3) (d)

(e)

Message (MSG) Annunciator

The MSG annunciator is an advisory type (white) that lights when the FMS is

displaying a message in the scratchpad to the flight crew. The annunciator shall

extinguish after the message(s) have been cleared from the scratchpad.

OFFSET Annunciator

The OFFSET annunciator is an advisory type (white) that lights when a laterally

offset path has been entered into the FMS using the progress page. The

annunciator turns off when the offset has been removed. If there is an offset when

the APRCH annunciator is Ighted, the offset will be removed and the annunciator

turned off.

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(f) Approach (APRCH) Annunciator

The APRCH annunciator is an advisoty type (white) that lights when in approach

mode. The NAV Computer output sensitivity of lateral deviation to the EHSI will be

ramped to a higher sensitivity when the annunciator is lighted. The APRCH

annunciator will light under the following conditions:

If the destination elevation is specified, distance to destination is less than 15

NM, altitude is less than 2500 feet ative the destination elevation, and the speed

is less than 200 knots.

If the destination elevation is not specified, distance to destination is less than 15

NM, and speed is less than 200 knots.

9. B.

(5)

Function Keys

There are four function keys, and the function of each is described in the following

paragraphs:

(a)

Previous (PREV) and NEXT Page Keys

The number of pages in a particular mode or menu display are shown in the upper

right hand corner of the display. The format is AAIBB. AA signifies the number of

the current page that is displayed. BB signifies the total number of pages that are

available for pilot viewing/modification. Page changes shall be done by selecting

the PREV and NEXT keys. When in the PLAN mode, these keys will increment or

decrement the map center waypoint.

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(b)

Clear (CLR) Key

The CLR key performs the following functions:

When a message is present in the scratchpad,

delete that message.

pressing the CLR key shall

When an alphanumeric entry resides in the scratchpad, one character shall be

cleared from the scratchpad (from riiht to left) for each time the button is

pressed.

When an alphanumeric entry resides in the scratchpad and the CLR key is hekf

down, the first character is cleared within 100 ms. After 400 ms have elapsed,

9.

B. (6) (b)

(c)

(d)

Navigation (NAV) Mode Key

Pressing the NAV mode key shall enable the pilot to access the NAV index page.

The pilot may select any of the submodes by pressing the line select key.

Flight Plan (FPL) Mode Key

Pressing the FPL mode key shall display the first page of the flight plan. If there is

no flight plan currently entered, the pilot may manually enter a flight plan, select a

stored fliiht plan, or create a stored flight plan.

Progress (PROG) Mode Key

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(7)

(e)

Pressing the PROG mode key shall display the first page of the progress pages.

The purpose of this mode is to show the current status of the flight. This first

progress page shall display the ‘to’ waypoint, the destination, the navaids that are

currently tuned for radio updating, and the update status of each navigation

computer.

Direct To/intercept (DIR) Mode Key

Pressing the DIR mode key shall display the active fight plan with the DIRECT and

INTERCEPT prompts.

Alphanumeric Keys

The Control Display Unit provides a full alphanumeric keyboard to enable pilot inputs to

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9. c. DL-900 Data Loader (See figures 9-5 and 9-6, and table 9-3.)

The DL-900 Data Loader is used to transfer navigation related data to the NZ-820~20

Navigation Computer. The DL-900 uses 3.5-inch diskettes and has an RS-422 interface with the

navigation computer.

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AD-m793@l

9. c. (1) Navigation Database Loading

The DL-900 Data Loader provides transfer of data derived from the Jeppesen database

from a 3-1/2 inch floppy disk to the NAV Computer local EEPROM memory. This data

includes navaids, vuaypoints, airports, airport runways, airport procedures, and jet routes

organized in regional partitions of the entire Jeppesen data source. The database is

updated every 28 days. The data transfer rate is 312K baud. The total time required to

load an NZ-820 full database is approximately 3 minutes. The NZ-920 requires

approximately 8 minutes to load.

(2) Flight Plan Loading

The DL-900 Data Loader also has the capability of interfacing with a ground-based

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Lockheed Jet Plan computer or equivalent. I t is capable of transferring an optimized flight

plan from the ground-based computer to the navigation computer via a 3-1/2 inch floppy

disk. For each flight plan, the fol lowing data wilt be stored: lateral waypoints, origin,

destination, winds, and temperatures at each waypoint.

‘1 (H)

RS232 DATA IN a ~

RS232 SIG GND V < ;

(c)

(H)J1

1>

Z RS232 DATA OUT

CENTRAL PROCESSING

[ <t

(c)

UNIT (CPU)

DATA BUS IN G

- Z80 CPU I

FROM NAV

IX, ~

I

(H)

-8 K BYTES RAM

COMPUTER H

-32 K BYTES EPROM

(c) ] ~

- RS422 INTERFACE

DATA BUS OUT

9. D.

OZ-800 Receiver Processor Unit (See figures 9-7 and 9-8, and table 9-4.)

I

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AD-15616


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.58 inches (370.33 mm)

Wtih . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2.25 inches (57.15 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.60 inches (193.04 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 6.51b(2.95 kg)

Power Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 Vdc,40Watts (maximum)

Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.2to13.6kHz

Mating Connector . . . . . . . . . . . . . . . . . . . . . . . . . . Cannon Part No. DPXBMA-57-335-OOOl

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Mounting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tray(l/4ATRShort Box or Equivalent)

0Z-800 Receiver Processor Unit

Leading Particulars

Table9-4

The 0Z-800 Receiver Processor Unit (RPU) receives and processes data from the ground-based

OMEGAWLF stations to provide latitude, longitude, N-S veloci iy, E-W velocity, and station

information to the flight management system. The antenna receives the OMEGNVLF signals

and converts them for processing by the RPU. All signals from or to the RPU are transmitted

1

OMEGFUVLF

ANTENNA

ANALOG i lIVLF DIGITAL DATA

*

*

w

INPUT

RCVR

DIGITAL

t

BUFFERED CLK

<

PROCESSOR

RCVR

1

SELECT

CONTROL

4

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GNLYOPEN

DISCRETE

DISCRETEd

INPUTS

INTERFACE

HDG~AS

--i

SYNCHRO

INTERFACE

I

NPUTS

9. E.

AT-801 H-Field Brick Antenna (See figure 9-9 and table 9-5.)

The AT-801 Antenna receives the OMEGAVVLF signals and mnverts them for processing by the

02-800 RPU. The antenna also incorporates built-in test circuitry to monitor its own operation.

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AD.156 7


Figure 9-9

10.

ODtional SRZ-850 Integrated Radio System

A.

RNZ-850 integrated Navigation Unit (See figures 10-1 and 10-2, and table 10-1.)

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AD-1 3743-R2@


The RNZ-850 Integrated Navigation Unit is a completely seff-contained navigation system. It

contains the NV-850 VHF NAV Receiver module, the DF-850 Automatic Direction Finder (ADF)

module and a six-channel scanning DM-850 Distance Measurement Equipment (DME) module.

Also within the navigation unit is a cluster module that contains the circuitry necessaty to handle

all of the digital outputs of the navigation unit modules and place them on the digital audio and

radio system buses.

Another function of the cluster module is an MLS Interface. The cluster module has circuitry and

drivers to feed the information coming from the RSB to the ML-850 MLS Receiver in the same

manner as it feeds information to one of the internal modules. This makes the external MLS, in

effect, a module of the NAV unit; however, it is housed separately so it may be used for

independent applications where a full NAV unit is not needed.

Each one of the modules is self-contained within its own housing and connects to the cfuster

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module via ribbon cabfe. For bus utilization considerations, it is most efficiint to have several

units feeding their information to the central cluster module and that module placing the data

onto the bus rather than having each unit irrdhridually address the bus, By this packaging

approach, functions are gathered together resufling in a considerable savings in weight, volume,

and installation labor.

As previously mentioned, the NAV unit is physically divided internally into four modules, the

VOR, DME, ADF, and cluster, and are all provided power on an independent basis through the

rear connector and through ribbon cables that feed each unit. Therefore, each module has its

own power supply, and is contained within a cast housing that has covers providing shiekfing

and protection for the module. Removal of heat generated within the modules is provided for by

sinking power devices to a special heat sinking structure within the module for transfer to the

outer surfaces of the unit. Air flow provisions within each module also assist in the heat removal

process, aided by a noncritical fan located on the mounting rack at the rear of the unit.

10. A.

(1)

NV-850 VHF NAV Receiver

The VHF NAV receiver portion is a self-contained module housed within a die cast

assembly comprising three printed circuit boards internal to the casting and one external.

The NAV receiver houses four major functions: the VOR/Localizer Receiver, Glideslope

Receiver, Marker Beacon Receiver, and Power Supply/Processor.

The navigation receiver has extensive buitt-in test circuitry. This BITE operation includes a

setf-test signal generator and modulator buil t into the unit. When energized by flight crew

or power-up command, the injected signal is identical to a VOR/lLS signal and starts the

testing at the earl iest possible stages of the various receivers, just after the antenna.

BITE commands an extensive check of the various circuitry within the NAV receiver and

wil l cause the various outputs to move in a very specified and regular sequence, allowing

the pilot to confirm that the entire navigation receiver is operating pmperfy. Additionally,

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there are other maintenance monitor and signal monitors associated with the NAV module,

which continuously check for proper operation and valid signals within the unit. These

circuits continuously watch over the operation of the unit and, shoufd any operating

parameter move outside of its nominal range, this condition wil l be stored in nonvolati le

memory for subsequent maintenance readout.

(a)

VOR Receiver

The VOR portion of the NAV receiver is used to intercept a VOR radial in the radio

frequency range of 108.00 to 117.95 MHz on channels spaced 50 kHz apart. The

VOR receiver provides radio deviation, To-From, bearing, and flag outputs to the

EFIS Symbol Generator for display on the EHSI and to the FZ-800 Flight Guidance

Computer for automatic capture and traddng of the selected VOR radial.

10. A.

(1) (d)

Power Supply

The self-contained power supply is fed from a dedicated pin on the main connector

of the unit allowing for independent application of aircraft supply vottages to the

NAV receiver.

(2) DM-850 DME

The DM-850 Distance Measuring Module is

a

six=hannel scanning DME that

simultaneously tracks four selected DME channels for distance, ground speed, and time-

to-station as well as tracking two additional channels for the IDENT functions. This gives

the system the capability of tracking four channels and having the decoded identifier

readily available from two additional channels. The unit dedicates two of its four channels

to a Flight Management System when installed. Thus, with an FMS, the flight crew has

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two remaining channels to control and display IDENT, distance, time-to-station, and

ground speed. The IDENT only channels will have decoded IDENT ready, so when the

crew selects the preset VOR channel, the instant search capability of its DME will allow

the functionality of 4 full-time channels for the crew.

The ranging capability of the DME is up to 300 miles, ground speed capability up to 1000

knots, and time-to-station capability up to 999 minutes. These signals are sent from the

DME in several formats, one of which is the normal radio system format appearing on the

digital bus via the cluster module, another is an RS-422 format. Also provided is an

ARINC-568 standard output of six wires on which data, sync,

andcl ock

re provided to

output the distance. In addition to the digital outputs, a 40 millivolts per nautical mile

analog output and audio capable of driving two 600-ohm audio loads with the IDENT

signal is also provided. Self testing is accomplished via a built-in signal generator, which

The ADF also has a dual bandwidth operating mode. In order to meet the requirements o

the regulatory agencies for bandwkfth, current ADF receivers must be designed so that the

audio fidelity of the receiver is severely degraded. The ADF receiver has a voice mode of

operation so that, when desired, the ADF audio qualiiy may be inqmved to allow voice

and other types of signals to be clearly received.

For self-test, the ADF has a built-in oscil lator circuit~ located in the antenna which, when

energized, couples directly into the antenna circuitry and provides a complete test of the

entire ADF system.

The ADF has an input function allocated for HF COM keying information. Previous

units

are quite susceptible to onboard HF transmitters, which can cause the ADF pointing cirmit

to be disturtmd during transmissions. The DF-8S0 ADF module provides an input signal

so that when the HF is transmitting, the ADF processor will reject the false influences

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created.

10. A.

(4) Cluster Module

The cluster module is attached to the 212-pin rear connector of the NAV unit. All of the

signals from the aircraft wiring harness, with the exception of the antennas, come through

this rear connector and onto the cluster module. They are then distributed to the various

modules over ribbon cabfes that plug into the edges of the cluster module. The cluster

module power supply receives its power via a diode OR connection through each one of

the modules, assuring that power wil l be available even with several individual module

power supplies turned off.

The cluster module also contains circuitry associated with the digital audio system. All of

RADIO

SYSTEM ~

BUS

CLUSTER

MODULE

T

R

s

:

BB

I

k;

T

E

k

VOR/lLS/MKRMODULE

r

RFAND

~ RS-422

SIGNAL

~ ANALOG

POWER

PROCESSING

SUPPLY

~ AUXCONTROL

I

d

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DIGITAL

AUDIO ~

BUS

:

FF

AA

c:

E

DIGITAL

AUDIO

INTERFACE

DMEMODULE

RCB

INTERFACE

RF AND

~

SIGNAL

~

+

POWER PROCESSING

SUPPLY

~

I

*

I

II

ARING 588

RS-422

ANALOG

-

I

4

ADF MODULE

e

10. B.

RCZ-850/851 A Integrated Communication Unit (See figures 10-3, 10-4, 10-5, and table 10-2.)

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AD-13744-Rl@

RCZ-850/851 A Integrated Communication Unit

Figure 10-3

The RCZ-850/851 A Integrated Communication Unit, also known as the COM unit, is identical in

concept to the RNZ-850 Integrated NAV unit in that it contains internal modules that feed their

signals through a cluster module and have their signals placed on the RSB (radio system bus)

for operation. The modules within the COM unit are the TR-850 VHF communication transceiver

and the XS-850 Mode S air traffic control transponder or an XI-851 TCAS interface module

(RCZ-851 A only). Each one of the modules, again like the NAV unit, is selfantained within its

own housing, has its own internal power supply (except for the TCAS interface module), and its

own interface to ttre cluster module. The operation and cooling of the COM unit is also identical

to the NAV unit with a fan mounted on the mount and controlled by signals from the individual

modules according to their internal cooling requirements.

The cluster module has its own onboard power supply and receives its primary 28-vott input

power from both the VHF COM Transceiver Module andhe Mode S Transponder Module so

that in the event either of them is energized, the cluster module will be energized. The COM

cluster module, like the NAV cluster module, contains audio interface circuit~ for the signals from

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the COM unit. Because of the nature of the function, the Mode S transponder has no audio

output circuit~.

The COM cluster module has an additional function in that there are four undedicated audio

inputs available on the connector of the unit so that non-Honeywell products that provide analog

audio sgnals can gain access to the digital audio bus.

tO. B.

(1)

TR-850 VHF Communication Transceiver

The TR-850 VHF COM Transceiver module provides air-toground and air-to-air voice and

data communications in the radio frequency range of 118.00 to 136.975 MHz (or from

118.00 to 151.975 MHz in extended frequency range) on channels spaced 25 kHz apart

Built into the COM is a self-test oscillator which, when energized, will cause a signal to

appear in the receiver and wil l verify its operation.

Other features of the COM include the standard Radio System nonvolatile maintenance

log and internal monitoring to verify circuitry performance and to record any deviation from

nominal olxwation for later recall by maintenance personnel.

The COM transmitter power output is a nominal 20 Watts, a guaranteed minimum of

16 Watts, and is applicable across the entire frequency range of 118.000 to 151.975 MHz.

The receiver sensitivity of the COM is a nominal 2.5 (hard) microvotts.

10. B. (2) XS-850 Mode S Transponder

The XS-850 Mode S transponder module works with the Air Traffic Control Radar Beacon

System (ATCRBS) to provide enhanced surveillance and communication capability

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required for air traffic control automation. The Mode S Transponder rmduie uses 1030

MHz to receive interrogations and 1090 MHz to transmit replies. It is fully functional with

ATCRBS Modes A and C and capable of providing Basic Mode S operation. Mode S

albws digital addressing of individual aircraft and the communication of messages back

and forth between the air and the ground and is a fundamental portion of the FAA

proposed Traffic Alert and Collision Avoidance System.

When the transponder senses a change in the reply code commanded by the control

head, it will hold the current reply code until the new code has remained constant for

approximately 3 seconds. Then it will begin to use the newly selected reply code. This is

done in an effort to avoid transmitting false alarms and false emergency signals when the

emergency codes are inadvertently used during the process of tuning.

RADIO

SYSTEM ~

BUS

CLUSTER

MODULE

R

s

B

*

I

N

T

E

R

F

t

E

COMMODULE

I

11

MODE S MODULE

RCB

INTERFACE

RF AND

SIGNAL

-ALTIMETER

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DIGITAL

AUDIO -

Bus

DIGITAL

AUDIO

NTERFACE

POWER

SUPPLY

POWER

PROCESSING

SUPPLY

w

I

I

1

I

- [ AUX AUDIO

Hone~ell

-————

I

I

COM

AUXCONTROLBUS

COM AIRCRAFTINTERFACES<

a

TCAS AIRCRAFT lNTERFACES&

I

I

I

I

COM CLUSTER

MODULE

MODULE

NTERCONNECT

R

s

B

R

c

B

MAINTENANCE

MANUAL

CITATION Vll

———— —

lNTEGRAT;C~M~NlT7

I

I

I

I

++

4

RCB

i

b INTERFACE

VHF C43M

I

+

POWER

MODULE

SUPPLY

I

,COM

ANTENNA

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RSB PRIMARY BUS-

RSB SECONDARY BUS

I

I

I

I

I

DIGITAL AUDIO BUS

,4

N

T

E

R

F

A

c

E

I

N

T

E

R

F

A

c

E

DIGITAL

AUDIO

INTERFACE

+

t

RCB

b INTERFACE

TCAS

INTERFACE

MODULE

I

I

I

I

t

POWER

I*

e

FROM CLUSTER P.S.

I

AUX RCB

BUSSES

AUXAUDIO

INPUTS

AUXDISCRETE

AUDIO STATUS

10. B.

(3) (a)

Interface Functions

The TCAS interface module supports fwr interface functions:

Radio communication bus (RCB) interiace

Low-speed ARINC 429 interface

c Discrete inpuffoutput (1/0) interface

Remote programming interface.

The processor RCB interface function cxmverts one data format to another (for

example, RCB to ARINC data), as shown in the examples below.

~

m

Serial RCB/RSB data from COM Discretes and serial ARINC data to

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unit transponder

Discretes from aircraft and Serial RCB data to COM unit

transponder

The RCB communication format is the same as that used in COM units with an

internal XS-850 transponder, except that it includes additional data words to support

TCAS. Depending on the operational mode, the processor initiates an RCB data

block transmission every 50 ms.

The ARINC 429 interface function outputs ARINC labels 013, 015, and 016 on a

low-speed ARINC 429 data link. The processor formats each label from data

10. B.

(3) (b)

Operational Modes

The module operates in five basic modes, depending on which portion of the

program is being executed. Each operational mode is defined by the module RCB

control byte.

A brief description of the five operational modes follows.

The power-on mode is the module initial state at power-up, Following

power-on tests, the processor sends the null control byte until it receives the

configuration data byte from the COM cluster module.

In the normal mode, the module outputs ARINC 429 data and discrete signals

to the transponder based on data received on the RCB. It also monitors the

signal interfaces and records malfunctions.

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In the POST/PAST mode, the module executes the fol lowing checkout

sequence:

Sets the transponder in the test mode by raising the FUNCTIONAL TEST

OUTPUT discrete to the HIGH state

Reads the XPDR FAIL INPUT discrete to determine if the transponder

passed its self-test

Reads the module self-test status to determine if any internal fai lure has

occurred

Returns a FAIL condition to the RCB if either the transponder or the module

10. c.

ML-850 Microwave Landing System (MLS) Receiver

(See figures 10-6 and 10-7, and table 10-3.)

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AD-1 3745-R2Q

ML-850 MLS Receiver

Figure 10-6

The ML-850 is housed in a self-contained package and, unlike

the other units, does not have a

cluster module. It is intended to be used with the Radio System as another module associated

with the RNZ-850 Integrated Navigation Unit. In its operation, it is fully integrated with the NAV

unit and may be thought of as simply another module for the NAV unit with the only exception

being that its package is separate. By doing this, the MLS receiver is allowed to take full use of

all the bus and internal system operation information. For example, selecting the MLS mode on

the EFIS control will cause the DME to select the channel commanded for the MLS and will pair

up for the approach operation.

The MLS receiver is similar to the other radio receiver products. It has a receiver, synthesizer,

signal processing, and a power supply. The receiver is a triple conversion super heterodyne that

begins at the 5 GHz MLS operating frequency and converts down to several stages before finally

presenting the signal to the signal processing cimuit~. The frequency synthesizer provides the

internal RF signals necessary for the operation of the receiver. The power supply is a

conventional switch mode supply and provides the voltages necessary for the MLS operation. It

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is operated from the aircraft 28 volt dc supply line, totally independent of the other units in the

aircraft.

The ML-850 MLS Receiver decodes and processes data from an MLS ground station and

provides an accurate indication in both azimuth (equivalent to bcalizer) and elevation (equivalent

to glideslope) of the deviation from the desired flight path. The deviation data is displayed on the

EADI and EHSI.

The ML-850 operates in the frequency range of 5031.0 to 5090.7 MHz on 200 channels spaced

3000 kHz apart. Selection of the desired azimuth and elevation angle and tuning is

accomplished with the RM-850 RMU. In its operation, the ML-850 is fully integrated with the

RNZ-850 Integrated Navigation Unit and may be thought of as simply another module for the

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MLs ANTENNAS

TT

‘ ———

———— ———

1

I

DPSK

I

II

?

ANTENNA

SWITCH

280

MAIN

I

I

b

+

PROCESSOR

I

I

SYNTHESIZER

-

1

I

4.

t

I

b

I

t

TCXO

1;

1{0+ BITE

I

PROCESSOR

RCB

BITE

COMM

RCB

TO

NAV UNIT

OR

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1,

d

8031

L——————

———— ——

JL++—————

J

MLS

CDU

‘----5--iF--

“

+5 +28 +15 -15

F

r

I

I

POWER

SUPPLY

I

L

————

———— ——

——

———— ———— —

A

28 hX

10. D.

RM-850 Radio Management Unit (RMU) (See figures 10-8 and 10-9, and table 10-5.)

CURSOR

\

TRANSFER

KEY (LEFT —

SIDE)

LINE SELECT

KEYS (LEFT —

PHOTO

CA

- ~ SENSOR

II 1

II 1

II -1

‘cm’TN’v’l

,123.20 110.25

p=q

109.35

b w wI ‘EMORy-’ I

1471

II

162.5

1ATCON ANT

~MLSl -MDU

RANSFER

f-II

KEY (RIGHT

SIDE)

[ II

[ II

p

II 11

LINE SELECT

KEYS (RIGHT

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SIDE)

FUNCTION

KEYS

1

MAN. MODE C“: 600 ~

W 300” E

GP

3.8

BAZ 200

I

[I

SIDE)

f II

The Radio Management Unit (RMU) is the central control unit for the entire radio system. It

provides complete capabWty for controlling the operating mode, frequencies, and codes within al

the units of the Radio System. Additionally, the RMU has the capability to switch its operation

from its primary radio system to the cross-side system. The RMU is amorCRT-based

controller featuring the proven concept of selecting a function by pushing a line select key

adj acento the parameter that is to be changed. Any selectable parameter, such as a VOR

frequency, may be changed by pressing the corresponding line key next to the displayed

parameter and then rotating the controller tuning knob.

For ease of operation, the RMU screen is divided into five dedicated windows. Each window

groups the data associated with a pafiicular function of the radio system. The five windows

(COM, NAV, Transponder, ADF, and MLS) each provide for co~lete control of frequency ancVo

operating mode of the associated function. The RMU also has other display modes, called

pages, which perform additional features and functions for the control of the radio system.

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The RMU is also the input to the radio system for external FMS tuning in that digital signals from

the FMS come into the RMU where they act in much the same manner as if the front tuning

knob was being operated. This allows the FMS to enter into the system in an organized manner

and will appear to the system as if the flight crew is tuning the receiver.

As a safety feature of the RMU, should any of the components of the radio system fail to

respond to commands from the RMU, the frequencies or operating commands associated with

that patiicular function will be removed from the RMU and replaced with dashes. This will alerl

the crew to the fact that the radio system operation is not normal.

Also available in the RMU is a maintenance mode of operation, when not in flight. During this

mode, various pages are utilized to allow maintenance personnel access to the maintenance log

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10. D.

(5) (f)

Page (PGE) Key

Pressing the PGE key once will change the RMU display to the COM preset

frequency memory page. Pressing it a second time will move the display to the

NAV preset frequencies page. Keying PGE a third time calls up the discrete

RADIO ON-OFF page, which is the last page of the RMU program. A fourth push

of the PGE key will return the display to the Main Page. All of these back pages

assign a “Return” function to the lower feft line select key. Pressing this key will

bring back the “Front” page.

(9)

Test (TST) Key

Pressing the TST key causes the ccmponent associated with the yellow cursor’s

present position to activate its internal seff-test cirwits for a complete end-to-end

test of the function. Hold the TST key down for the duration of the test, about 2

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seconds for COM transceiver, 5 to 7 seconds for DME, ATC, ADF, and about 20

seconds for NAV (VOFVILS). Releasing the TST key at any time immediately

returns the function to normal operation.

(h)

DME Key

The DME key deslaves the DME from the active VOR frequency, and alfows tuning

of a different DME channel without changing active VOR. Successive presses of

the DME key enable display and selection of the DME channels in VHF and TACAN

formats.

(6) Cursor

DISPLAY

RADIO

SYSTEM

-i

RSB

BUS

INTERFACE

DISPLAY

DISPLAY

CONTROL

-

DRIVERS

n

1

1

RS-422

FMS

RS-422

INPUT

INTERFACE

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ATC IDENT

COM FREQXFR

COM MEM SEL

VOR MKR SENS

SIDE SEL BITS

DISCRETE

TEST ENABLE

INPUTS

WEIGHT-ON-WHEELS

WOW POIARllY

FRONT

-

CONTROL

)

PROCESSOR *

1

BEZEL

DISCRETE

POWER OFF

OUTPUTS

CONTROLS

1

CONTROLS

Hone~ell ~f~NcE

10. E. AV-850A Audio Control Unit (See figures 10-10 and 10-11, and table 10-6.)

H

r MICROPHM —

?545

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( - J g g. l&~l(-J

SPEAKER

HEAUWUINE

AD-18840


Figure 10-10

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10. E.

(3)

(4)

Emergency (EMER) Switch

When the EMER switch is pressed, the microphone is connected directly to a

predetermined VHF COM transceiver, and the transceivers received audio is connected

directly to the aircraft’s headphone. The system may be wired to simultaneously route a

single NAV audio to headphones. When EMER is selected, headphone volume is

controlled by the master headphone volume control . Al l electronic circuitry is el iminated in

the EMER position. This mode also disables all other audio paneI modes.

Audio Source Control

Each control (COM, NAV, ADF, DME, MLS) mmbines the function of switch andvolume

control. The control energizes a particular channel’s audio when unlatched (out position)

and de-energizes the audio when latched (in position). Rotation of this control will adjust

audio level from minimum at the fully CCW position to maximum at the fully CW position.

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(5)

(6)

lD/Voice Switch

The lD/Voice switch is operated by setting a rotary switch and is used to filter the VOR

and ADF audio signals. In the ID mode (CCWposition), the VOR or ADF audio is filtered in

such a way as to enhance the Morse Code identification. [n the VOICE mode (CW

position), the audio is filtered to reduce the Morse Code signal for received ADF and

VOR/lLS audio. In the BOTH (center position), the VOR and ADF signals are not

subjected to any fil tering in the audio frequency band.

Speaker and Headphone Controls

These controls are used to adjust the speaker and headphone amplifier’s volume. They

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10. F.

CD-850 Clearance Delivery Control/Display Unit (See figures 10-12 and 10-13, and table 10-7.)

SYSTEM INSTALLATION

REMOTE TUNE

TUNING

NAV AUDIO ON

ANNUNCIATOR \

ANNUNCIATOR

CURSOR ANNUNCIATOR

RADIO TUNING

ANNUNCIATORS

TRANSFER KEY

EMERGENCY

MODE

ANNUNCIATOR

SQUELCH

ANNUNCIATOR

TRANSMIT

ANNUNCIATOR

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NORM/EMERGENCY

MODE SWITCH

NAV AUDIO

ON/OFF SWITCH

SQUELCH

ON/OFF SWITCH

TUNING KNOBS

AD-29837

CD-850 Clearance Delivery ControVDisplay Unit

Figure 10-12

The CD-850 Clearance Delivery CDU provides an alternate or emergency backup capability for

tuning the remote mounted VHF Communications transceiver ancVor VHF Navigation Receiver in

the event that the primary Radio System Bus (RSB) tuning is not available, or if the pilotloperator

wishes to override the bus tuning for any reason.

The CD-850 can be used before engine start for initial mmmunications with low-power drain. It

can act as a stand-alone control unit or a backup thifd control. The CD-850 has several

operating modes, which are selected by either the mode knob or by installation strapping on the

rear connector. The modes selected by installation strapping are:

Clearance Delivery mode, which is the normal operating mode.

COM only mode, which makes the unit dedicated to COM tuning only.

NAV only mode, which makes the unit dedicated to NAV tuning only.

The normal and emergency modes are submodes that are selected by the mode knob and are

used if the unit is strapped for clearance delivery.

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10. F

The following paragraphs describe each function or control:

(1)

(2)

System Installation Annunciator

Either the 1, 2, or 3 annunciator is ON to indicate to which system the CD-850 is

connected.

Remote Tune Annunciator

This annunciator is active only when the CD-850 is strapped for NAV only or COM only

tuning. RMT is annunciated when the radio is tuned from some source other than the CD-

10. F.

(7) Transmit (TX) Annunciator

This annunciator indicates when the COM transmitter is ON.

(8)

NAV AUDIO OnK3ff Switch

This atternate action pushbutton is used to toggle NAV audio ON or OFF.

(9) squelch (SQ) Or’VOffSwitch

This alternate action pushbutton is used to toggle the COM squelch ON or OFF.

(10) Tuning Knobs

The tuning knobs are used to change the frequency indicated by the tuning cursor.

(11) Normal/Emergency Mode Switch

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This rota~ switch knob provides alternate selection of the Normal and Emergency modes

when the CD-850 is strapped for operation as a clearance delivery head. This switch is

nonoperating in the COM only or NAV only modes.

(12) Transfer Key

In the clearance delivery mode, the transfer key alternately selects either the COM

frequency (top) or the NAV frequency (bottom) to be mnnected to the tuning knobs.

In the NAV only or COM only configuration, the transfer key toggles the active (top)

AUX PORT

Y f

COM AUX _ SHIFT

DATA

*

PORT

REGISTER

-

CLOCK

DISpLAy

/

,9

DISPLAY

AND DRIVER

ENA6LE

DRIVERS

DATA ~

-

A

4 )

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{

NAV AUX

PORT

{ PROCESSOR

LOAD

)

i

PRIMARY RSB

~

RSB

RECEIVER

+

“n”T’ ZKb$==l

i

J

SQUAT SW SENSE

J--=%- I _,

I

? r

-1

10. G. DI-851 DME Indicator (See figures 10-14 and 10-15, and table 10-8.)

RMU

FREQ SELECT

PILOT’S/COPILOT’S MLS TUNE

ANNUNCIATOR (1/2) ANNUNCIATOR ANNUNCIATOR

DME HOLD

ANNUNCIATOR

3A–LA

+7.s:lLR5&

HLD NAV PRE 12 MLS

DME DISTANCE

\D>~ y_,M/~Ei’Z~,ATOR

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DME CHANNEL

PARAMETER

SELECT SELECT


Figure 10-14

AD-15823-R1


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10. G. (5)

Parameter Display and Select

The DME station identifier, the computed ground speed of the aircraft in knots, or time-to

go (time to reach the ground station) in minutes is displayed as a function of the

parameter select (SEL) button. The KTS/MIN annunciator identifies which parameter is

being displayed. Each time the SEL button is pressed, the display changes as follows:

SEL Button

Parameter Annunciator

Power Up

Identifier

Blank

1st Push

Ground Speed KTS

2nd Push

Time-To-Go MIN

3rd Push

Identifier Blank

(6)

DME Hold Annunciator

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HLD is annunciated if the DME frequency is split from the VOR.

OATA

e

CLOCK DISPLAY

,9

ORIVERS

/

OISPUY

PRIMARY

ENAE L5 d

mm

lNPuT~

RS8

Y )

R.%

PROCESSOR

10. H. AT-860 ADF Antenna (See figures 10-16 and 10-17, and table 10-9.)

AD-141W

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AT-860 ADF Antenna

Figure 10-16


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..18.3 inches (414.8 mm)

Width . . . . . . . . . .’ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..8.33 inches (211.6 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..l.51inches (38.3 mm)

The AT-860 ADF Antenna performs the functions of reception, amplification, and combination of

RF signals so as to yield low-frequency reception and directional information. Normal reception

of AM signals is performed by the E-field element, or vertically polarized antenna; while bearing

information is provided by H-field antennas in the form of a pair of loop antennas mounted at

right angles to each other. By carefully combining the amplified signals, bearing information is

obtained in the form of phase modulation on the received RF, which is demodulated and

processed in the receiver.

The antenna also contains a self-test circuit that radiates a 120 kHz signal into the sense and

loop antennas. This checks the operation of both the AT-880 ADF Antenna and the DF-850

ADF Receiver Module. Proper operation is indicated by a 1 kHz tone and a bearing indication o

135 degrees.

0[ >=

‘

OOP Cos Cos

BALANCED

ANTENNA LOOP

AMP

MOOULATOR

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SIN

LOOP

SIN

50 OHM

AMPL

OUTPUT

v

90 DEG

PHASE

EQUALIZER

10. 1.

AT-851 MLSAntenna (See figure 10-18, andtable 10-10.)

AD-15804@

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AT-851 MLS Antenna

Figure 10-18


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..2.50 inches (62.5 mm)

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1.50 inches (38.l mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 0.75 inches (19.05 mm)

The Processor receives signals that are naturally generated by lightning activity, and determines

their range from this energy distribution. At the same time, bearing is computed by means of

Antenna crossed loops in a manner similar to an ADF.

The LP-850 Processor outputs display data (lightning symbols) directly to the WI-870 Weather

Radar Indicator through a universal digital interface (UDI). The UDI port permits lightning data

encoded in a raster format to be overlaid on a radar display. When no DATA NAV computer is

in use and the radar system is in the Standby mode, the LP-850 Processor takes over the entire

radar display and creates a 360-degree display of lightning data.

Lightning data is also converted to ARINC 429 low-speed digital messages (mode, strike rate,

location, fault, and test page information) for display on EFIS or MFD-type systems. Label

assignments do not conform to ARINC 429. The data stream contains range, bearing, and

severiiy data for up to 3 alert and 50 rate symbols. The 429 data also contains all information

available for a 360 degree area with a radius of 125 NM around the aircraft. The information is

prioritized so that symbols in the forward 120° sector are sent first (in order from the closest

symbol). It is the task of the display device to determine which symbols fall within its display

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area. Some displays may limit their lightning data to the forward scan sector and may not utilize

the full 360 degree information.

The LP-850 Processor has a basic maintenance and operational philosophy; condition

monitoring. Multilevel built-in software and hardware tests (BIT) are used by the LP-850

Processor to monitor itself and other system components for proper operation in order to detect

and record faults. The purpose of BIT is to detect and isolate failures internal or external to the

LP-850 wherever possible. All BIT capabilities are executed by the LP-850 Processor in

software. The first occurrence of a failure and a single indication of the first repeat of that failure

is recorded in nonvolatile (maintenance) memory as a fault code. Subsequent repeats on the

11. LSZ-850 Lightning Sensor System (OptionaI~

A.

LP-850 Lightning Sensor Processor (See figures 11-1 and 11-2, and table 11-1.)

~O.O

>~,ooo

O.OO

,.OO

,>

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AD-15286

28 VDC *

POWER

f12vDcTo

SUPPLY

* ANTENNA

I

HN, HP, E-FIELD

ANALOG

FROM ANT

+ TO DIGITAL

CONVERTER

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DATA

PROCESSOR

11. A. (1) Configuration Straps

The lightning system has a number of options that may be selected by means of jumper

pins (Configuration Straps CS1 to CS16) located at connector J101 of the LP-850

Processor. These pins are jumpered to ground to configure the lightning system (LX

mode) and are only read by the processor during power-up initialization. If it is necessary

to change the jumper pins, the system must be shut down (for at least 5 seconds) in order

to force the LP-850 Processor to read the straps again.

The pins are read as a hexadecimal representation of the jumper configuration. A jumper

to ground equals logic “O.”

The use of each jumper pin is defined in table 11-2.

Configuration strap 16 is the MSB of the left-most digit, and CSI is the LSB of the right-

most digit. The procedures for field adjustment of the configuration straps are found in

the LSZ-850 System Description and Installation Manual, Honeywell Pub. No.

A09-3950-01 .

(2) ARINC Transmit Data

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The LP-850 Processor encodes and transmits the following nonstandard ARINC Labels in

the order listed below. All labels are presented in OCTAL format. The specific labels

being transmitted depend upon mode of system operation and amount of lightning data

present. A data stream consisting of all applicable words is transmitted each 265* 50

ms, beginning with label 001 fol lowed by other applicable labels in numeric order.

Label

001

Assignments

Initial transmission with discretes and fault data.

Lightning System Options

RESERVED

1 = oRen, O = shorkd to J101 Pins A47 or A64

DISPLAY AZIMUTH ANGLE

90 Degrees

120 Degrees

160 Degrees

180 Degrees

ARINC INPUT (J101 PINS B28-B29) FORMAT

ARINC 419 (561)

ARINC 429

ARINC INPUT DATA SPEED

Configuration Straps

@l_

1

CS-2

CS-3

—.

1 0

1

1

0

1

0 0

CS-4

T

1

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High (100 Kbytes per seoond)

Low (12.5 Kbytes per second)

HIGH/LOW RECEIVER GAIN

CS-6 is deactivated and the receiver is always

in high gain.

POSITIONING MODE

o

1

CS-6

T

Qs-J CS-9

Cs-1o

S-8 _ ,_

1

7- 1 1

Lightning System Options

HEADING DATA TYPE

ARINC 407 Synchro

ARINC 419/429 Digital

DATA OUT SPEED

High (100 Kbytes per second)

Low (12.5 Kbytes per second)

SPARE JUMPERS

No Connection

Configuration Straps

CS-12

1

0

CS-13

o

1

CS-14

CS-15

CS-16

—— —

1 1 1

LP-850 Configuration Strap (CS) Jumpers

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Table 11-2 (cent)

11.

B. LU-850 Lightning Sensor Controller (See figures 11-3 and 11-4, and table 11-3.)

The LU-850 Lightning Sensor Controller contains a simple rotary mode switch for the Lightning

Sensor System and requires no power apart from its panel lamps. This controller is supplied

optionally. Mode control of the Lightning Sensor System can be accomplished in one of three

ways: through the LU-850 Lightning Sensor Controller, the WC-870 Weather Radar Controller, or

the WI-870 Weather Radar Indicator. The method used depends upon which components are

installed in the aircraft. If an MFD system is installed, the LU-850 can be replaced by using the

controls on the WC-870. If an MFD is not installed, the LU-850 can be replaced by using the

controls on the WI-870.

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AD-3531 7(Q

Operation of the rotary knob is identical to the LSS rotary knob on the WC-870 Weather Radar

Controller and/or WI-870 Weather Radar Indicator. The LSS knob is a four-position rotay switch

that allows the LSZ-850 Lightning Sensor System to be operated in the following modes:

Mode

Function

OFF

Removes power from the Lightning Sensor System.

SBY (Standby) Places the Lightning Sensor System in nonoperational

mode. Display of data from the system is inhibited, but

data is still accumulated.

LX (Lightning) Lightning Sensor System is fully operational. Lightning

strike data is collected, processed, and displayed

CLR/TST (Clear/lest) Accumulated data is cleared from memory of the Lightning

Sensor System. After 3 seconds the test mode is initiated.

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JI

—

A

c

B

E

+ 2EIV LIGHTING

LIGHTING COMMON

+ 5V LIGHTING

PWR ON COMMAND

11. c. AT-850 Lightning Sensor (Teardrop) Antenna (See figure 11-5,

and

table 11-4.)

NOTE: This antenna is intended for external mounting only, on the top or

bottom surface of the aircraft. An AT-855 (Brick) Antenna mav

used if the location for installation is protected from wind -

turbulence. Also, the AT-850 antenna is encapsulated and not

repairable.

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AD-13956

AT-850 Antenna

Figure 11-5


The Antenna contains crossed loop H-field antennae and an E-field antenna similar to an ADF

antenna.

Preamplifier stages are also built into the Antenna in order to enhance the system’s immunity to

noise originating in aircraft wiring.

The H-field loop antennae are designated Hn and Hw and are orientated in such a manner that

Hn will be most sensitive to signals originating ahead or behind the aircraft, and the Hw antenna

will be most sensitive to signals originating abeam the aircraft.

The E-field antenna is constructed such as to be most sensitive to vertical E-fields. ‘

~1 o 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 winding is driven with a

simulated lightning signal and couples with the E- and H-elements of the Antenna. Thus, the

test mode is able to provide an evaluation of the performance of the Antenna, its preamplifiers,

and cabling to the Processor.

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Honeywell

11. D. AT-855 Lightning Sensor (Brick) Antenna

MAINTENANCE

MANUAL

CITATIONVll

(See figure 11-6, and table 11-5.)

The AT-855 Antenna is functionally interchangeable with the AT-850 (Teardrop) Antenna.

However, the AT-855 antenna is not aerodynamic and is intended for instal lations that are

protected from wind turbulence. The AT-855 Antenna circuit~ (crossed loop H-fiekt antenna, E-

field antenna, and test windings) are also encapsulated and is not repairable.

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AT-855 (Brick) Antenna

Figure 11-6


1

l

X

lW

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tw

The TCAS Computer Unit (CU) is packaged in a modular concept unit (MCU) outline referred to

as a 6 MCU short. Six plug-in assemblies, removable from the top of the unit for shop

maintenance, and interconnected by a motherboard, make up the CU. Its rear panel is a size-3

AFUNC-600 connector with six cavities (A thru F). Its front panel has a carrying handle, a self-

test command switch, and 11 annunciators used for maintenance. Two NA622CE2 hooks

secure the CU to the mounting tray.

The CU requires external cooling air in accordance with ARINC 600 or ARINC 404 in order to

maintain the highest possible mean-time-between-failures (MTBF). In those installations where

external cooling is not available, a mounting tray with an integral fan is required. The mounting

tray is not supplied by Honeywell.

12. A. (1)

CU Functional Description

The TCAS CU interrogates airborne transponders, processes their replies, and produces

video graphic data for use by the VS1/lRA display. The CU contains the RF transmitter

and receivers necessary to send and receive replies from transponder-equipped aircraft,

Dual microprocessors process the surveillance and collision avoidance system (CAS)

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data. The CU uses that data to determine which aircraft represents a potential collision

threat and the appropriate vertical response to avoid a midair coll ision or near midair

incident.

If the threat persists, the CU supplies visual and aural advisories to the flight crew to

assure progressive vertical separation. Progressive vertical separation avoids the threat

while causing the least deviation of the TCAS aircraft from its current rate of climb or

descent. An interface is provided with an onboard Mode S Tran.sfxmder in order to

coordinate avoidance maneuvers with other TCAS equipped aircraft.

12. A. (4) CU Rear Connector Layout

External plug-in connectors on the rear panel interface the CU to the TCAS and

transponder system LRUS. The CU rear connector has six cavities (designated A, B, C,

D, E, and F), which provide the following interface functions:

Cavity A - Left top plug (LTP) connects the CU to the top directional antenna

Cavity B - Left middle plug (LMP) connects the CU to the bottom omnidirectional or

directional antenna

Cavity C - Left bottom plug (LBP) connects the CU to the aircraft mutual suppression

bus and the 115 V, 400 Hz power bus

Cavity D - Right top plug (RTP) is not installed

Cavity E - Right middle plug (RMP) connects the CU to the transponder(s), radio

altimeter(s), and RA/TA display(s)

Cavity F - Right bottom plug (RBP) connects the CU to various aircraft discrete lines.

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A

complete list of interface connector pins is provided in SECTION 6, INTERCONNECTS.

Software updates can be incorporated into the CU program memory by means of an

ARINC 615 data loader port. This serial port is accessible either through the front panel

connector or the rear panel RMP connector.

~

12. A. (5) CU ARINC 1/0

The CU transmits ARINC 429 high-speed output data to the VS1/TRA displays as listed in

table 12-2.

The CU also transmits ARINC 429 high-speed output data to the Mode S Transponder as

listed in table 12-3. This data is for coordination with other TCAS II equipped aircraft.

The Mode S Transponder transmits ARINC 429 high-speed output data to the CU as

listed in table 12-4. This data is for coordination with other TCAS II equipped aircraft.

Parameter/Signal Name

Label Rate

Control Panel Set

Altitude Select

TCAS Mode/Sens

013

015

016

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Intruder Range

Intruder Altitude

Intruder Bearing *

Own Aircraft Altitude

Vertioal RA

Horizontal RA

130

131

132

203

270

271

2-3 Hz

2-3 Hz

2-3 Hz

2 Hz

2-3 Hz

2-3 Hz


Label Rate

Control Panel Set

013

Altitude Select

015

TCAS Mode/Sens 016

Own Aircrafl Altitude

203

Vertical RA

270

Horizontal RA 271

272

273

Select TCAS Sensitivity 274

2 Hz

2-3 Hz

2-3 Hz

2-3 Hz

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275

276

277

Maintenance 350

RT-91 O TCAS Computer-To-Mode S Transponder Data

Table 12-3

SUPPRESSION

Bus

TOP ANT

COAX

SOTTOM ANT

COAX

PARTOF

ARINC 600

CONNECTOR

1

I

4:

4

I

9

Ic

A4

4

RF 1/0

A3

-;

A5 ~

SURVEILLANCE

I

RF CPLURF

RECEIVER

-

1/0

I

PRoCESSOR

PARTOF

ARINC 600

CONNECTOR

i

I

I

I

I

I

8/11/2019 spz-8000 difcs


I

A

I

A6

<: :

RF

I

TRANSMITTER

4

I

t__

115VAC

i

400

Hz

4 POtiE R

1

4*

4

POWER

I

SUPPLY

1-

OUTPUT DISCRETES

2*

1

~ AuRAL OUTPUTS

I

I

I

I

I

1~

Honeywell

12. B. DV-91 O VSVTRA Display (See figures 12-4,

MAINTENANCE

MANUAL

CITATION Vll

12-5, and 12-6, and table 12-5.)

Major components of the DV-91 O VS1/TRA include a full-color high resolution Iquid crystal

display (LCD) and bezel assembly, backlight assembly, four plug-in circuit card assemblies,

motherboard, and a circular 41-pin connector as the sole electrical interconnect to the aircraft

system. The unit interfaces with the same high-speed ARINC 429 bus used by the TCAS CU. It

also accepts ARINC 565 and ARINC 429 vertical sped inputs.

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12. B. (1) Functional Description

The DV-91 O VSUTRA is a microprocessor-controlled electronic flight instrument. Its basic

function is to provide an indication of the aircraft’s vertical speed, TCAS traffic information,

and warning advisories. The display symbology used for advisories is described in the

TCZ-91 O System Description and Installation Manual, Honeywell Pub. No. Al 5-3840-001.

The VS1/TRA is also displays system mode, status and test annunciations. When the CU

is operating in an extended test mode, the VS1/lRA dk+play presents lest pages containing

system diagnostic information. Test pages with diagnostic examples are also shown in the

TCZ-91 O System Description and Instal lation Manual. Figure 12-5 shows typical displays.

The electronic vertical speed indicator (VSI) portion of the DV-91O display presents rate of

climb or rate of descent on a scale centered around zero vertical speed. The vertical

speed display is derived from signals input directly to the VS1/lRA. Three possible

sources exist for vertical speed data including:

.

ARINC 429 data

.

DC analog vottage according to ARINC 575 (approximately 500 mV per 1000 ft/min)

AC analog voltage according to ARINC 585 (approximately 250 mV per 1000 ft/min).

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These three selectable sources provide compatibil”~ with most aircraft.

Advisory information is received from the CU on a dedicated high-speed ARINC 429 bus.

The ARINC bus carries bearing, altitude, and range data for each threat. The VS1/TRA

uses that information to give an indication of the proximity of the threat and of the vertical

speed required to avoid the threat. A green band overlaying the VSI scale points to the

desired vertical speed; and a red band indicates the vertical speed range to be avoided.

TYPICAL

DISPLAY

X’ ””-

“\

f

\

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FAILURE

ANNUNCIATIONS

These pins are jumpered to ground to configure the DV-91 O VS1/TRA Display and are

only read by the microprocessor during power-up initialization. If it is necessary to change

the jumper pins, the unit must be shut down in order to force the microprocessor to read

the straps again.

The pins are read as a hexadecimal representation of the jumper configuration. A jumper

to ground equals logic “O.” Configuration strap CS7 is the MSB, and CSO is the LSB. A

complete list of connector pins with strap options is provided in the TCZ-91 O System

Description and Installation Manual.

12. B. (3) Built-in Tests

The microprocessor performs multilevel built-in software and hardware tests (BIT) to

monitor itself and other system components for proper operation in order to detect and

record faults. The purpose of BIT is to detect and isolate failures internal or external to

the DV-91 O VS1/TRA wherever possible. All BIT capabilities are executed by the

microprocessor in software. The first occurrence of a failure and a single indication of the

first repeat of that failure is recorded in nonvolatile (maintenance) memory as a fault code.

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INTERNAL

FEEDBACKLIGHTSENSOR

BACKLIGHT

LIGHT

m

ASSY

SENSOR

PIO Al

r I

I

FEEDBACK LIGHT SENSOR

I

VALID

BOOT

3

ANALOG

CONTROL

ECA

I

A5

12. B. (4) TCAS Displays

The TCAS modes use color-coded symbols and data tags to map air traffic and local

threat aircraft on the VS1/TRA Display.

Four traffic symbols are used: solid circle, sol id square, solid diamond, and hollow

diamond. See figure 12-5 for examples. A different color is assigned to each symbol

type, as listed in table 12-6.

Graphic Symbol

Color

Display Function

Solid Square

Red

Resolution Advisory (RA)

Solid Circle

I

Amber

I

Traffic Adviso~ (TA)

Solid Diamond

I

Blue

I

Proximate Traffic

Hollow Diamond

Blue

Other Traffic

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TCAS Symbology

Table 12-6

(a)

Colors

Red

Represents an immediate threat to a TCAS-equipped aircraft. Prompt action is

12. B. (4) (b) Traffic Identification

Resolution Advisory

Intruder aircraft entering the warning area, 20 to 30 seconds from the TCAS II

collision area are represented as a solid red square. This type of traffic will

result in an RA.

Traffic Advisory

Intruder aircraft entering the caution area, 35 to 45 seconds from the TCAS II

collision area are represented as a solid amber circle. This type of traffic will

result in a TA.

Proximate Traffic

Aircraft within display range, and within the selected vertical window, are

represented as a solid cyan diamond. Proximate traffic is shown to improve

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situational awareness in the event of a potential conflict with higher prioriiy RA

or TA aircraft.

Other traffic

Any transponder-replying traffic that is not classified as an intruder or

proximate traffic, and is within the display range, and is within the selected

vertical window, are represented as hollow cyan diamonds (only in view when

no RA or TA is in progress). The predicted flight paths of proximate traffic and

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12. C. (1) Line Select Keys

The first push of the line select key moves the yellow cursor to surround the data field

associated with that particular line select key. This then electronically connects that data

field to the tuning knobs so that the mode or code may be changed.

(2) Code Select Key

Press this key to place the cursor around the transponder code data line. Now the large

outer tuning knob controls the left two digits, and the smaller inner knob controls the right

two digits. Figure 12-7 shows a selected code of 1471.

Since only one transponder can operate at a time, both RMUS will be displaying the same

transponder information. Therefore, if a code or mode is changed on one RMU, the other

RMU will track it. Since the other RMU is being tuned by a remote source, the data

changed wil l appear in yel low.

Press and hold this key for more than 2 seconds to change the code to that which was

stored in the memory. To store a code in memory, dial the desired code within the

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cursor, and then press the STO key. The stored code will remain in memory during recall

and during power-down.

(3)

Mode Select Key

Pressing this key moves the cursor to the mode line, and enables several functions.

Press this key again to toggle between standby and the last active mode,

Honeywell ~$$ ””

12. c. (4)

(5)

Altitude Display Key

Press this key to move the cursor to this line. Press this key again, or twist either tuning

knob, to toggle the TCAS intruder alti tude display between relative alti tude (REL) and

uncorrected altitude (FL).

Surveillance Window Key

Press this key to move the cursor to this line. Press this key again, or twist either tuning

knob, to select one of the following survei llance window sizes:

NORMAL -2700 feet above own aircraft and 2700 feet below own aircraft

+

ABOVE -7000 feet above own aircraft and 2700 feet below own aircraft

BELOW -2700 feet above own aircraft and 7000 feet below own aircraft

These selections are determined by the flight crew, depending on the vertical path of the

aircraft. NORMAL would be selected during level flight. ABOVE or BELOW would be

selected during high rate climbs or descents.

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(6)

The results of these two key selections will be seen on each TCAS display separately. In

normal operation, RMU No. 1 will select these functions for the left side TCAS display,

and RMU No. 2 will select for the right side TCAS display. If either RMU is in the cross-

side control mode (with magenta banner l ines) that RMU wil l control the cross-side

display, just as all other cross-side controls.

PGE Key

On the RMU Page Menu, press the MAINTENANCE line select key, and then the RMU

SETUP line select key. The RMU SETUP page will be displayed, as shown in

figure 12-10. On this page ATC FLIGHT ID may be disabled, and therefore not

transmitted in the Mode S replies. When ATC FLIGHT ID is disabled, the FLIGHT ID

legend is not shown on the ATC/TCAS Control Page, nor on the Main Operating page.

II41

II

~:

PAGE MENU SYSTEM 1

RADIO PAGE

SYS OWOFF

COM MEMORY

NAVIGATION

o

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II

[1

II

II

NAV MEMORY ENGINE PG1

ATC/TCAS

ENGINE PG2

MLS

RETURN

MAINTENANCE /

Ho-

0

ATC/TCAS CONTROL PAGE -

INTRUDER ALTITUDE: REL

TA DISPLAY: AUTO

FLIGHT ID AA 125B

FLIGHT LEVEL I 22500

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RETURN

Hol l e wOM

o

1(

RMU SETUP SYSTEM 1

IIh]

MLS DISPLAY -ON

~1

II

[I

II

1

TCAS DISPLAY -ON

1

ATC FLIGHT ID - ENABLE

1

1

II

II

II

II

8/11/2019 spz-8000 difcs


II

RETURN FOR

1

NORMAL OP.

II

TUNE

12. D. AT-91O Directional Antenna (See figure 12-11 and table 12-7.)

AD-32826@

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AT-91 O Directional Antenna

Figure 12-11


Height Outside Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..806 inches (20.47 mm)

Height Inside Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..l.56inches (39.62 mm)

The antenna is capable of receiving replies from all directions simultaneously with bearing

information using amplitude-ratio monopulse techniques. Insertion loss differences in coaxial

cable lengths from the antenna to the TCAS computer need only be matched to within 0.5 dB,

which corresponds to a 5 to 10 foot difference in length depending on the specific cable type.

Losses between the antenna and the computer unit must be 2.5 f 0.5 dB, including line

connections.

12. E. Typical Bottom Omnidirectional Antenna (See figure 12-1 2.)

.— ——

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u AtI=’827@

Typical Omnidirectional Antenna

Figure 12-12

12. F.

XS-91O Mode S Transponder (See figures 12-13 and 12-14, and table 12-8.)

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The transponder is packaged in a 4 modular concept unti (MCU)

outl ine as defined in

ARINC 600-7. The’basic ‘mechanical chassis is constructed from of l ightweight aluminum alloy

sheet metal with extruded side panels for additional strength. Four plug-in assemblies,

removable from the top of the unit for shop maintenance, and interconnected by a mothedxmrd,

make up the transponder. Its rear panel is a size-2 ARINC-600 connector with three cavities (A,

B, and C). The front panel has a carrying handle, a seff-test command switch, and

six annunciators used for maintenance.

The transponder requires external cooling air in accordance with ARINC 600 or ARINC 404 in

order to maintain the highest possible mean-time-between-failures (MTBF). In those installations

where external cooling is not available, a mounting tray with an integral fan is required. The

mounting tray is not supplied by Honeywell.

12. F. (1) Mode S Functional Description

The transponder is a surveillance and communication system required for operation of the

TCAS. The data link capability of the Mode S Transponder allows it to setve as an

essential element of the TCAS. TCAS-equipped aircraft are airborne interrogators,

communicating with other TCAS-equipped aircraft through their Mode S Transponders.

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All TCAS avoidance maneuvers are coordinated through the transponder. In addition, the

Mode S Transponder is responsible for providing alti tude data and RM-850 RMU control

inputs to the TCAS CU.

The name Mode S comes from its direct-selectable address format, thus mode select

(Mode S). Each Mode S equipped aircrafl has an individual, airframe-specific, assigned

address code. No two aircraft have the same two address codes. Address codes are

12. F. (2)

Mode S Built-in Tests

The transponder performs multilevel built-in software and hardware tests (BIT) to monitor

itself and other system components for proper operation in order to detect and record

faults. The purpose of BIT is to detect and isolate failures internal or external to the

transponder wherever possible. All BIT capabil ities are executed by the transponder in

software. The first occurrence of a failure and a single indication of the first repeat of that

failure is recorded in nonvolati le (maintenance) memory as a fault code. The transponder

maintains a log of the last ten flights.

(3) Mode S Maintenance Indicators

PASS/FAIL indicator lamps and a PUSH TO TEST button on the transponder front panel

(figure 12-14) supply system status for maintenance purposes. By momentarily pressing

PUSH TO TEST, maintenance or engineering personnel can activate a self-test cycle and

monitor fautt data for the current and preceding fl ight legs on the indicator lamps.

(4) Mode S Program Pins (Configuration Straps)

The transponder is designed to accommodate various system configurations. Program

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pins on the rear transponder connector inserts al low the user to select airframe specific

functions that satisfy a particular instal lation. These pins are jumpered to ground to

configure TCAS and are only read by the transponder during power-up initialization. If i t

is necessary to change the program pin jumpers, the system must be shut down in order

to force the transponder to read the pins again.

(5) Mode S Rear Connector Layout

Honeywell

Honeywell

I

MAINTENANCE

MANUAL

CITATION Vll

A

\

OHM METER

LEADS

/

B

4

0000000000

~oooooooooo

~oooooooooo

4

00000000007

~oooooooooo

40000000000

~oooooooooo

80000000000

700000000007

w

_ TOP

ANTENNA

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XPDR PASS @ @ TOP ANT

XPDR FAIL @ @ BOT ANT

CNTL PNL @ @ ALT SIG

D

- BOITOM

ANTENNA

13. Global Positioning System (Optiona~

A.

Global Positioning System Sensor Unit (See figures 13-1 and 13-2, and table 13-1.)

The GPSSU consists of a flange-mounted device with two connectors. It contains five internal

circuits; a power supply circuit, a navigation processor circuit, an RF circuit, and two signal

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The GPSSU receives signals transmitted by the NAVSTAR satelli tes, computes present position,

altitude, true track, and groundspeed, and outputs this data on an ARINC 429 databus. If the

GPSSU is not able to maintain track of at least four satellites, it uses pressure altitude from the

DADC, and received data from the remaining satelli te(s) to compute present position. If the

GPSSU is not able to track any satellites for 30 seconds, it reverts to the Acquisition Mode.

During this mode, the GPSSU accepts position data from the FMS, and transmits that data

(which is identified as FMS data) until it has acquired at least four satel lites, when it re-enters

the Navigation Mode.

The GPSSU is a two-channel, single-frequency GPS receiver capable of receiving the L1

frequency transmissions (1575.42 MHz) from NAVSTAR satell ites. The GPSSU performs the

following functions:

Tracks the L1 coarsdacquisition (C/A) code transmitted by the NAVSTAR

global positioning system (GPS) satellites.

Locks onto the satellite signal.

Computes the Pseudo-range to the satelli te. Pseudo-range consists of the actual range

modified by receiver clock errors.

Computes the Pseudo-range rate from the satellite (Doppler). Pseudo-range

rate consists of the actual range rate modified by receiver clock errors.

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Decodes the satellite data.

Computes the aircraft position, (this is referred to as the navigation solution).

ARINC 429 standard communication buses provide the interface for direct data exchanges with

the Flight Management System (FMS) NZ-8201920 Navigation Computers and AZ-81 O Digital Air

Data Computers. ARINC outputs include aircraft position, velocity, and satellite information. The

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13 A. (1) (d)

(e)

(f)

Nav Mode

In the Nav mode, the GPSSU updates and transmits data on the ARINC 429 data

bus to its interfaces. The data, which includes latitude, longitude, altitude, time,

and velocity, are derived from pseudo range and pseudo range rate measurements.

These measurements are performed seven times a second. The G PSSU remains

in the nav mode as long as it is able to track four satellites. If it is unable to track

four satellites, the GPSSU enters the altitude-aiding/clock coasting submode.

Aftitude-Aidin~Clock-Coasting Submode

The GPSSU enters the altitude-aiding/clock-coasting subrnode from the nav mode

when it is unable to track four satellites. In this submode, the GPSSU uses inertial

or pressure altitude inputs to determine position and other data. The GPSSU

remains in this submode as long as one to three satellites are being tracked.

When the GPSSU has acquired four satellites, the GPSSU re-enters the nav mode.

If the GPSSU cannot track any satellites for 30 seconds, the GPSSU revetts to the

acquisition mode.

Fauft Mode

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The fault mode occurs when built-in test equipment (BITE) detects a critical failure.

In this mode, all outputs are invalid.

(2)

Signal Processor Modes of Operation

The signal processor has two modes of operation: the continuous tracking mode and the

Hone~ell ~~~~NcE

Binary (BNR) Data Format

Parameter/Sianal Name

Label Units Digital Range Resolution

Pseudo Range

Pseudo Range Fine

Pseudo Range Rate

Delta Range

Satellite Position X

Satellite Position X Fine

Satellite Position Y

Satellite Position Y Fine

Satellite Position Z

Satellite Position Z Fine

UTC Measure Time

GPS Altitude (MSL)

HDOP

VDOP

061

062

063

064

065

066

070

071

072

073

074

076

101

102

Meters

Meters

Meters/Second

Meters

Meters

Meters

Meters

Meters

Meters

Meters

Seconds

Feet

*268435456

256

f4096

f4096

*671 08864

64

f671 08864

64

+671 08864

64

10

~131072

1024

1024

256

0.125

0.0039

0.0039

64

0.0039

64

0.0039

64

0.0039

9.5367 E-6

0.125

0.031

0.031

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Track Angle - True

GPS Latitude

GPS Longitude

GPS Ground Speed

Latitude Fractions

103

110

111

112

120

Degrees

Degrees

Degrees

Knots

Degrees

-E180

+180

f180

4096

1.716E-4

0.0055

1.716 E-4

1.716 E-4

0.125

8.38E-8

MAINTENANCE

MANUAL

CITATION Vll

Binary Coded Decimal BCD

Data Format

Label

Units

Digital

Range

Resolution


UTC

Date

Equipment

125

HR:MIN

23:59.9 0.1 Min.

260

D:M:Y

1

ID

377

GPSSU Binary Coded

Decimal (ARINC 429) Output Data

Table 13-4

Discrete (DIS) Data Format

Label

Units

Digital

Parameter/Signal Name Range

Resolution

8/11/2019 spz-8000 difcs


GPSSU Status

273

Maintenance Discrete No. 2

352

System Time Counter

354

Seconds

262144

1

Maintenance Discrete No. 1

355

‘HG~21~G~U— — — — —

h

—1

I

+}

DADC 1 2

ARINC419 (575)

Q

OR 429

12.5 KHz

DADC2 2

[+

2

TIME MARK NO. 1

REAL TIME

CLOCK 1 Hz

D

2 TIME MARK NO, 2

I

Q

MC/lRSl 2

Mc/lRs2 2

D

}

TIME MARK NO, 3

ARINC 429

12.5 KHzOR

100 KHz

ARINC 429

I

429 OUT NO. 1

8/11/2019 spz-8000 difcs


a

I

DADCINPUT

419/429 SELECT

-1

429 OUTPUT

HS/LSSELECT

-1

12.5 Kt-lzOR

\

100 KHz

OPENIGROUND

DISCRETE

INPUTS

Q

b

429 OUT NO. 2

+

2

429 OUT NO. 3

BEGINATP

-q]

13. B. AT-81 O GPSSU (Dome) Antenna

TOP

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14. LASEREF@ 111nertial Reference System (iF?S~

A.

HG2W1 A602/HG200fAC02 inertial Relerence Unit (See figures 14-1 and 14-2, and tables 14-1

and 14-2).

The HG2001 At302 or HG2001 A(XJ2IRU isa $tr~wn, &t iler -tuned navigation system. The

IRU contains the necessary power supplies, sensors, and electronics to compute attitude and

true heading. h turther computes present position, inertial veiociiy vectors, magnetic heading,

sensor systematic error compensation, arid provides the necessa~ d~ita~ signals for the

EFIS/MFD flight displays, f~iht guidance, ?Iightmanagement, weather radar, and other aircraft

systems as required, Leading particulars are listed intable 14-1.

8/11/2019 spz-8000 difcs


\


Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...7.64

Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...4.88

Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...13.12

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

inches (194.1 mm)

irtches(124.0 mm)

inches (333.0 mm)

27.0 lb (12.25 kg)

Power Requirements:

Primary AC . . . . . . . . . . . . . . . . . . . . . . . . . .

115Vrms, single-phase, 400 Hz (nominal)

Primary/SecondaryDC . . . . . . . . . . . . . . . . . . . . . . . . . +28 Vdc,80Watts (maximum)

Backup (battery) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..+24Vdc. 4 ampere-hour

Mating Connector (Jo) . . . . . . . . . . . . . . . . . . . . . . .

llTCannon PatiNo. BKAD2-313-30001

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

Tray, MT-260, 4MCU (with Blower Kit)

Honeywell Part No. 26006092-101

Blower Kit:

Optional DC Blower Kit . . . . . . . . . . . . . . . . . . . . . .


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Optional AC Blower Kit . . . . . . . . . . . . . . . . . . . . . .


Fan, FN-260 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Fan filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Inertial Reference Unit

Leading Particular

Information is transmitted to and from the DIFCS, Weather Radar System, DADC, FMS,

EFIS, and the MFD System through multiple input and output (1/0) communications ports

(ARINC 429 and ASCB). The HG2001 AB02 uses Versions A and B of the ASCB word

format to transmit data and the HG2001 AC02 uses Version C. Specific information about

ASCB and ARINC word formats can be found in the Installation Manual for the

LASERE~ Ill, Pub. No. Ml 5-3343-011. The accuracy and resolution of the data

provided by the IRU is listed in table 14-2.

NOTE:

Accuracy of the IRU atti tude angle outputs is directly dependent upon

the accuracy with which the mounting tray is al igned with the aircraft

axes during installation.

14. A. (2) IRU Power Transfer

The IRU power supplies can accept either 115 V ac or +28 V dc power from the aircraft

and backup battery as primary power. Power switching to primary ac, primary dc, or

backup battery power is handle by the IRU,

The +28 V dc aircraft power can be connected as a secondaty power source (pin C-7) to

protect the IRU from ac power interruptions and line transients. When using 115 V ac as

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the primaty power (pin C-1), the IRU automatically transfers to +28 V dc power whenever

the primary ac power drops below 85 V rrns. If the voltage level on the +28 V dc power

(pin C-7) also drops below 18 V dc, the IRU automatically transfers to the +24 V dc (pin

C-2) backup battery power.

If the IRU is operating with +28 V dc aircraft power as the primary source, then the IRU

automatically transfers to the +24 V dc backup battery power. The backup battery power

Parameter Limitation Navigation Mode Attitude Mode

1.

Present position

FAR 121,

NA

Appendix G

O.1OO

RES: 0.010

2.

Pitch angle

3.

Roll angle

None

TSO-C4C

RES: 0.010

None, except for the following

TSO-C4C

.1 0“

RES: O.O1°

ondition: When cos pitch -

c 0.087, then roll angle = last

computed value

None, except for the following

condition: When cos pitch

c 0.087, then heading = last

computed value

1) Computed between

latitudes (Iat) 73”N and

60°S only:

a) Between *50° tat

RES: O.O1°

4.

True heading

0.4°

RES: O.1°

NA

5.

Magnetic heading

2° Initial tracking

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3°

SET HDG 1°

RES: O.1°

Operational

accuracy 15° hr

drift fl?aX

b) Greater than 50° Iat

RES: O.1°

) When cos pitch c 0.087;

then heading = last

commtted value

6. Groundspeed

None

12 kts

NA

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Parameter

24. Track angle

magnetic

25. Drii angle

26. Flight path angle

27. Track angle rate

28. North-south (N-S)

velocitv

29. East-west (E-W)

velocity

Limitation

Not computed when velocity

<20 kts; accuracy based on

120 kts ground speed










None

None

Navigation Mode

6°

RES: O.1°

5.0°

RES: 0.10 BCD

RES: 0.005° BNR

0.4°

RES: O.1°

0.25”tsec

RES: O.1° BCD

RES: 0.005° BNR

flz Ids

RES: 0.125 ktS

flz Ids

RES: 0.125 ktS

Attitude Mode

NA

NA

NA

NA

NA

NA

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30. inertial altitude

Barometric altitude input

required; accuracy specified

with constant altitude input;

fil ter at steady state; no error

assumed in air data input;

resolution as specified with

5ft

RES: 1 ft

NA

FAN FILTER

sEE VIE

REMOVALTA

I

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14. A. (5)

IRU Built-in Tests

The IRU performs multilevel built-in software and hardware tests (BIT) to monitor itsetf

and other system components for proper operation in order to detect and record faults.

The purpose of BIT is to detect and isolate failures internal or external to the IRS

wherever possible. All BIT capabilities are executed by the IRU microprocessor in

software. The first occurrence of a failure and a single indication of the first repeat of that

failure is recorded in nonvolatile (maintenance) memory as a fault code. Fault codes are

defined in the Installation Manual for the LASERE~ Ill, Pub. No. Ml 5-3343-011.

(6) IRU Rear Connector Layout (See figure 14-3.)

External plug-in connectors on the rear panel interface the unit to the system LRUS. The

rear connector has three groups of connector inserts (designated A, B, and C), which

provide the following interface functions:

Cavity A - the top insert group provides the test and signal

interface

Cavity B - the middle insert group provides the system signal

interface

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Cavity C - the bottom insert group provides the power interface.

This group connects the IRU to the aircraft 115 V, 400 Hz power

bus and the +28 V dc essential bus

A complete listing of the connector pins is provided in the INTERCONNECTS section.

See figure 14-4 for a diagram of the IRU interface.

Honeywell

MAINTENANCE

MANUAL

CITATION W

o

0

/

ABC DE FGHJK

10000000000

20000000000

30000000000

40000000000

50000000000

600 00 00 0 0 0 0

70000000000

80000000000

90000000000

100 0 0 0 0 0 0 0 0 0

110000000000

120000000000

130000000000

140000000000

150000000000

A BCD E F GH J K

10000000000

20000000000

30000000000

40 000 00 0 0 0 0

50 000 00 0 0 0 0

~ TOP

INSERT

8/11/2019 spz-8000 difcs


60 000 00 0 0 0 0

70 000 00 0 0 0 0

.90000000000

90 0 00 0 0 0 0 0 0

100 0 0 0 0 0 0 0 0 0

110 0 0 0 0 0 0 0 0 0

20 o 0 0 0 0 0 0 0 0

130 0 0 0 0 0 0 e 0 0

140 0 0 0 0 0 0 0 0 0

B MIDDLE

INSERT

FMC 1

FMC 2

NDU

ARINC 429

INITIALIZATION

12.5 KHz

DADC 1

d

2

d]

RINC 575/429

AIRDATA

DADC2

12.5 KHz

2

I

.

a

CLOCK

1

ASCB NO. 2

d

667 KHz

2

DATA

J

ARINC429

100 KHz

ASCBNO, 1

667 KHz

b

429-1

B

2 429-2

2

429-3

D

2 429-4

b

429-5

+

2

429-6

CLOCK

R

2

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1.

AIR DATA575/429 SEL ~

SDI 1

SDI 2

3

SDI 3

SDI 4

=1

L

ON DC ANNUNCIATOR

14. B. Mode Select Unit (See figures 14-5 and 14-6, and table 14-4.)

The Mode Select Unit (MSU)

provides

mode

selection, status indication, and test initiation for

one IRU. The MSU mnsists of a mode switch, six annunciators, and a test switch.

TEST

SWITCH

ALIGN FAULT

NAVRDY

NO AIR

ON BAIT

BAIT FAIL

h

4

\

tiDE

SELECT

ANNUNCIATORS

AD-3534e

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SWITCH

Mode Select Unit (MSU)

Figure 14-5

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14. B. (1) (h) ATT-, NAV-, or ALIGN-TO-OFF

After a 3-second delay, the IRU enters the power-off submode for approximately 7

seconds. At the end of 10 seconds, the IRU enters the Off mode.

(i) AlT-, NAV-, or ALIGN-TO-OFF-TO-ALIGN, -NAV, or -AIT

If the mode select switch is reset to Align, NAV, or AIT after 3 seconds in the OFF

position, but before the 10 second powerdown procedure has been completed, the

IRU completes the power-down procedure and then restarts power-on procedures.

(2) Annunciators

All of the MSU panel annunciators are driven by discrete outputs (OPEN/GROUND) from

the IRU.

(a) ALIGN

This indicates that the IRU is in the ALIGN mode. A flashing ALIGN annunciator

indicates that an incorrect latitude/longitude entry, or excessive aircraft motion

during alignment.

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(b)

(c)

FAULT

This indicates an IRS fault.

ON BATT

J1

u

e

f

M

N

A

T

s

h

LOGIC GND 4*

*

OFF

MODE SEL 1

5 ALN

T

ATT

OFF

o

I

ALN

MODE SEL 2

3 NAV

Q

T

ATT

o

+20 v Dc

BLOWER

CONTROL

NOTES:

~ C“RRENTNOTTOEXCEED 250MA

~ TWO PARALLEL LAMPS ARE USEDON

CG1042ABO3, MOOS O.

1,

ND 2.

~ FJIN.11-KNOT USED ON CCi1042AB03.

MOD 0,

~ PIN P1-14USED0NCGt042AB03

AND CG1042ABO4, MOD 2 (-128

AND -129 ASSEMBLIES).

ANNUNCIATOR TEST

1

2

3

I

8/11/2019 spz-8000 difcs


a

x

ALIGN

NAV RDY

4

5

\

I

I

I

I

14. c. Battery Backup (See figures 14-7 and 14-8, and table 14-5.)

The battery backup provides an alternate +24 V dc power source to the IRU when aircraft

~rifyfafy&wer is I&t.

1-

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Each IRU requires a battery backup source. The battew charger begins operation when the IRU

is powered on and continues until fully charged. When fully charged, the backup battery will

supply al l required power for at least 90 minutes.

The maximum charger input current with +28

V dc applied is 10 amperes.

The time required to charge the battery depends upon the battery charge level and temperature.

A fully discharged battery at a temperature of 75 “F (23.9 “C) may require as much as 1 hour to

charge.

The IRU transmits a CHARGER INHIBIT discrete signal (pin A-G9) to the battery during power-

up and BITE submode. This output discrete signal is used to turn off the battery charging

circuitw for 15 seconds while the IRU is being operated on battery power. Figure 14-8 shows

the typ:kal IRU operating time on backup power.

Z,llllllllllll

11..L.._l

l l –rrrrrnTl l :

_ . _ . l .

; I

“lvr_rl_rrrrrn I

I

1

‘T —~-”

~24

ml, m

—.., ,.

I ___ l

II

-L

i

8/11/2019 spz-8000 difcs


II

I

I

+

$ z&l

. ,.-,,.

I

I

I

1

Lu

I

22

,.-

~

1111

BATTERYFAILMONITORING LEVEL

.1, ,. 1

I

I

SECTION 3

SYSTEM OPERATION

1.

General

This section describes the operation of the System by separating the flight director/autopilot

description into the roll (paragraph 3. D), pitch (paragraph 3. E), and yaw (paragraph 3.F) channels of

operation. A table listing the system limits is contained in paragraph 2.

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Mode Control or Sensor Parameter Value

VOR,

Track

ourse Knob and

VOR APP, or

NAV Receiver

Roll Angle Limit

LNAV (cent)

Roll Rate Limit

Crosswind

Correction

Over Station:

Course Change

Roll Angle Limit

APR (LOC or

Course Knob and NAV Lateral Capture:

AZ) or BC Receiver Beam Intercept Angle

(HDG SEL)

f27

deg

4.0 deg/sec VOR APP

4.0 deg/sec VOR

Up to &J5 deg

Course Error

Up to *9O deg (VOR)

*3O deg (VOR APP)

*27 deg

Up to *9O deg

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Capture Point

Function of Beam,

Beam Closure Rate,

and Course Error

Min Trip Point

*= mv *

Mode Control or Parameter Value

APR (GS or

NAV or MLS

G P) Receiver

GA

Control Switches

on Throttles

(Disengage A/P)

Pitch Hold TCS Switch

GS/GP Capture:

Capture Point

Pitch Command Limit

Pitch Rate Limit:

Gain Programming

Fixed Flight Director

Pitch-Up Command;

Wings Level in Roll

Pitch Attitude

<150 mV GS Beam

Deviation TAS,

and VS

+10 deg, -15 deg

Preset

Minimum 2.0 deg/sec

Starts at 1500 ft radio

altitude for GS or

6 NM DME for GP

10.0 deg nose up

+ZO deg Maximum

8/11/2019 spz-8000 difcs


Depressed

ALT Hold or ADC or FMS

VALT Hold

Command

ALT Hold Engage

Range

o to 65,536 ft

Mode Control or Sensor Parameter Value

FLC or

ADC or FMS

Engage Range

VFLC

Hold Engage

Error

Pitch Limit

Pitch Rate Limit

ALT

ADC and Altitude

Preselect Capture

Preselect Preselect

Range

Controller

Maximum Vertical

Speed for Capture

Capture Maneuver

Damping

80 to 350 kn

0.4 to 0.85 MACH

*5 kn

t.01 MACH

,2*xlQQWl)

sec TAS

o to 65,536 f

Complemented

Vert Acceleration

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Pitch Limit

Pitch Rate

Limit Limiter

Preset

3.

Flight Director/Autopilot Functional Description

A. General

Paragraph 3.B discusses conditions and functions that are referred to in the text accompanying

each mode of operation in paragraphs 3.D, 3.E, and 3.F. Paragraphs 3.D, 3.E, and 3.F discuss

the signal flow thru the flight guidance computer for each flight path mode and the associated

roll, pitch, or yaw flight control axis. Figure 208 (sheets 1 thru 5), figure 209 (sheets 1 thru 9),

and figure 210 (sheets 1 and 2) are simplified diagrams that show the signal flow and

interconnect wiring for the applicable selected flight director mode and autopilot axis. Figures

201 thru 209 are mode select, switching, and AP engage logic diagrams that are used in

conjunction with figures 208, 209, and 210 to aid in understanding the system operation,

B. Control Functions

(1) Lateral Beam Sensor (LBS)

When flying to intercept the VOR or LOC beam in the heading select mode, the LBS will

be tripped as a function of beam deviation, course error, TAS, and DME. In the LOC

mode, the course error is compared with the beam deviation signal and rate of crossing

the beam to determine the LBS trip point.

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When the LBS trips, the flight director commands a turn toward the desired VOR radial or

runway at the optimum point for a smooth capture of the beam. If the intercept angle to

the beam center is very shallow, the LBS will not trip until the aircraft is near beam center.

For this reason, an override on the LBS occurs when the beam deviation reaches a

specified minimum. The minimum beam sensor trip point is *35 mV. The maximum LBS

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3. B. 7.

(8)

The VOR approach over station sensor trips when the following conditions are satisfied:

(a) VAPP TRACK has occurred plus 3 seconds of elapsed time.

(b) Either of the following conditions occurs:

Distance to the VOR station less than 3000 feet and DME valid.

Lateral deviation is greater than 75 mV and the rate of deviation is greater than 8 mV

per second and the DME not valid.

GS Track

Glideslope track occurs after the aircraft has captured the glideslope and is now tracking

the beam. The track phase provides for tighter flying of the beam. The following

conditions 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.

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(9)

c The vertical deviation must be changing at a rate of less than 3 ftkec.

GS CAP

3. B.

(10) (a)

VOR AOSS 1 and VAPP AOSS 1 will occur when the following conditions are all

satisfied:

VOR 0SS or VAPP 0SS has occurred dependent on the active lateral mode.

A calculated period of time has elapsed since the last to/from transition on the

HSI in order for AOSS 1 to trip. The period of time elapsed is calculated

using true airspeed and altitude. The higher the altitude, the longer it takes to

get through the cone of erratic radio information, therefore the longer the time

period must be. Likewise, the lower the aircraft altitude, the smaller the cone

of erratic radio information, and the shorter the time period must be to trip

AOSS 1. The required elapsed time period is also affected by the aircraft’s

true airspeed. The faster the airspeed, the quicker the aircraft will be through

the cone. The slower the airspeed, the longer it will take to pass through the

cone, and a longer time period is requirai to trip AOSS 1.

(b) VOR AOSS 2 and VAPP AOSS 2 will occur when the following conditions are all

sat isfied:

The respective VOR AOSS 1 or VAPP AOSS 1 has tripped plus 3 seconds.

Beam deviation is less than 75 mV.

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. Beam rate is less than 25 feet per second

Once VOR AOSS 2 or VAPP AOSS 2 trip, beam deviation will again be part of the

control signal.

The track condition is identified when the green asterisk extinguishes. At this time course

error is eliminated from the command signal, leaving beam deviation and inettial damping

from AHRS to maintain the aircraft on beam center.

3. B.

(13) LOC CAP 1 and BC CAP 1

Localizer and back course capture 1 are the initial capture phases of their respective

modes. Localizer capture 1 and back course capture 1 will occur when the following

conditions are all satisfied:

(a)

LOC armed plus 3 seconds.

(b) Beam deviation is less than 175 mV.

(c) Either of the following occurs:

Lateral beam sensor trips.

Beam deviation less than 35 mV.

(14) LOC CAP 2 and BC CAP 2

Localizer and back course capture 2 are capture phases which indicate the aircraft is now

8/11/2019 spz-8000 difcs


flying closer to the center of the beam. The capture 2 phase will occur for each mode

when the following conditions are all satisfied:

(a) LOC CAP 1 plus 3 seconds.

3. B.

(16) LOC Track 2 and BC Track 2

The track 2 submode will occur only after track 1 has been satisfied. There is no visual

indication to the pilot that the track 2 mode has been activated. Radio altitude, distance to

the transmitter, and a vertical velocity indicating the aircraft is descending are the factors

involved in determining the track 2 condition. When these conditions reach ceriain levels,

track 2 is tripped so as to provide tighter control during the final stages of an approach.

The track 2 phase will occur when the following conditions are all satisfied:

(a) LOC track 1 has been tripped.

(b) The aircraft is descending at a vertical speed which would indicate a runway

approach.

(c) Either of the following conditions has occurred:

Distance to the transmitter is less than approximately 5 miles and the radio

altimeter is invalid.

Radio altitude less than 1200 feet with the radio altimeter valid.

C. Flight Director Mode Selection (See figure 206.)

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There are nine mode select pushbutton switches located on the GC-81 O Flight Guidance

Controller as shown on sheet 1. When one of these switches is pushed, a ground (PB ARM) is

provided at 1lJ1 -47 to the FZ-800 to interrupt the “A”processor. Also, when a switch is pushed,

Honeywell

If VOR is annunciated in white indicating the

fol lowing inputs are present to the “OR” gate

GA

APP SEL

VOR CAP

HSI SEL

NAV SOURCE CHANGE

STANDBY

MAINTENANCE

MANUAL

CITATION Vll

mode is armed, the mode will be reset if any of the

to reset (R) the flip-flop:

Also, the mode will be cleared if any of the following inputs are present to the OR gate to clear

(CLR) the flip-flop:

. Flight Guidance Computer Valid not present (FGC VALID)

. Voted Automated Heading Reference System Valid not present (VAHRS VALID)

Selected Digital Air Data Computer Valid not present (SDADC VALID)

Tuned to a Localizer Frequency (lTL)

Selected HSI valid not present for two seconds (SEL HSI VALID “2 SEC)

8/11/2019 spz-8000 difcs


MAINTENANCE

MANUAL

CITATIONVll

11

1

2

3

5

B

il

71

Is

n

1

m

17

14

19

2(

4

4

-—— ——— ——— ——— ——— ——— .—— ——— -—,

1

Z-EWFL121NlOUIOANCE COUPU7CR

m-slo PLIONT3UIOANCE coNmoLLEn

I

I

I

I

I

I

I

I

I

I

}

ANNUNCIATOR OATA

7

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

I

i

+

(H)

cl

I

SElllAL

DAT AT OGC4 1 O

L

}

1

1--o

LOCK

}

)

STROBE

<

(H)

(cl

_ LAMP

T ES T I

‘k

z

_ T OCROS S

CHANNEL

PARALLEL

r

OUTPUTS

CLOCK

c=

—

ANNUN.

DRIVERS

ANNVLD

t

END AROUNO

LAMP LAMP TEST

T ES T 1 *

- DRIvE 1

k

+

CLOCK

WRALLEL

*

STR08E

-1

INPUTS

I *

( :

HslSEL

1

ERIALOUT

I I ~

L EF Ti RI GH T \ 2

-------- .

n,,,-c ,,,,.,..

4iL-

@-

Yo

I

NOTIES,

I

1.THE TNREESWITC

LASELEO AR VO,A

2. TNEINPUTIS LEV

TOmowos A5v

INPuTTOTNEPIS4

o&oPl

wm- -

I

f

I

+1

MRALLEL IN.

ANN Vi D ..–

I II -1 “Q,

<c,

cl-r

II

SVGCANNUNCIATOR VALIO

TCS *

1

1.

{

I+) 7

(c) o

?4

7s

7s

Ss

PILOT’S

1}

ARALLEL

I

NWTS

SEalAL

f I

OATAOUT

A

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Autopilot Engage Logic Diagram


Pages 243/244

3. D. Roll Channel Functional Operation

(1) Heading Select

HDG mode See

figure 208, sheet 1.)

The heading select mode is used to intercept and maintain a magnetic heading. The

mode is engaged by pressing the HOG button on the GC-81 O Flight Guidance Controller.

HOG will be annunciated on the EADI. Engaging the heading select mode will reset all

previously selected lateral modes. The flight guidance computer will now generate the

proper roll command to bank the aircraft to intercept and maintain the pilot selected

heading.

The heading cursor on the EHSI is positioned around the compass card to the heading

the pilot desires to intercept, using the heading knob on the RI-206S Instrument Remote

Controller (IRC). The heading select signal from the IRC to the SG-816 Symbol

Generator represents the desired ainxaft heading. In the symbol generator, the desired

aircraft heading is compared against actual aircraft heading and the resultant heading

select signal is routed to the FZ-800 Flight Guidance Computer through the Avionics

Standard Communications Bus (ASCB).

In the flight guidance computer, the heading error signal is TAS (True Airspeed) gain

programmed. TAS gain programming is performed on the heading error signal to achieve

approximately the same aircraft turn radius, regardless of the aircraft’s airspeed and

altitude. The TAS computation is derived from airspeed and barometric altitude

information provided from the AZ-81 O Digital Air Data Computer, through the ASCB.

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From the TAS gain programmer, the heading select command is routed to figure 208,

sheet 4, and is processed as discussed in paragraph 3. D.(8).

When reaching the lateral beam sensor (LBS) trip point, the system automatically drops

the heading select mode and switches to the VOR capture phase.

The following is

observed on the EADI:

The white VOR annunciator extinguishes

The green HDG annunciator extinguishes

A green VOR* is annunciated

The asterisk indicates the system is now in the capture phase of operation. The FZ-800

now generates the proper roll command to bank the aircraft to capture and track the

selected VOR radial.

When the course select pointer was set on the EHSI using the course knob on the

RI-206S Instrument Remote Controller, the course select error signal was established.

This signal represents the difference between the actual aircraft heading and the desired

aircraft course. The course error signal is then sent from the SG-816 Symbol Generator

to the FZ-800 through the Avionics Standard Communications Bus (ASCB). Next, the

course error signal is TAS (True Airspeed) gain programmed. TAS gain programming of

the course error signal is performed to achieve approximately the same aircraft turn radius

for a given command, regardless of the aircraft’s airspeed and attitude. The TAS

computation is derived from airspeed and barometric altitude information provided from

the AZ-81 O Digital Air Data Computer through the ASCB. From the TAS gain

programmer, the course error signal is summed with radio deviation.

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The radio deviation signal is routed from the navigation receiver to the SG-816 Symbol

Generator on the Collins Proline II bus. From the symbol generator, the radio deviation

As the aircraft approaches the VOR station, it will enter a zone of unstable radio signal.

This zone of confusion radiates upward from the station in the shape of a truncated cone.

In this area, the radio signal becomes highly erratic and it is desirable to remove it from

the roll command. The over station sensor monitors for entry into the zone of confusion

and opens the 0SS switch, removing radio deviation from the roll command. When the

aircraft exits the zone of confusion, the system displays VOR* on the EADI, again

indicating it is in the capture mode. When track conditions are again satisfied, the

asterisk is removed.

From the course cut limiter, the VOR SEL command is routed to figure 208, sheet 4 and

is processed as discussed in paragraph 3.D.(8).

3. D. (3) VOR Approach (VOR APP) Mode (See figure 208, sheet 2.)

The VOR Approach mode provides for intercept, capture, and tracking of a selected VOR

radial when less than 25 DME miles from the VOR station, or when using the VOR as an

approach reference to land. The VOR approach mode is set up and flown exactly like the

VOR mode, with the following differences.

Select the APP pushbutton on the GC-81 O Flight Guidance Controller.

Capture and track annunciators on the EADI will identify VOR APP.

Selected gains in the FZ-800 Flight Guidance Computer are changed to optimize

system performance in the VOR APP mode.

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(4) LNAV Mode (See figure 208, sheet 2.)

c Tune the navigation receiver to the published front course Iocalizer frequency for the

runway in use.

c Set the course pointer on the EHSI for the inbound runway heading.

c Set the heading cursor on the EHSI for the desired heading to perform a course

intercept.

Q Select NAV as the navigation source on the DC-81 O Display Controller.

The EHSI now displays the relative position of the aircraft to the center of the Iocalizer

beam and the desired inbound course. With the heading cursor set for course intercept,

the heading select mode will be used to perform the intercept. Outside the normal

capture range of the Iocalizer signal (between one and two dots on the EHSI), pressing

the NAV button on the GC-81 O Flight Guidance Controller will cause the EADI to

annunciate:

LOC in white

c HDG SEL in green

The aircraft is now flying the desired heading intercept and the system is armed for

automatic Iocaiizer beam capture.

With the aircraft approaching the selected course intercept, the lateral beam sensor (LBS)

is monitoring kxalizer beam deviation, beam rate, and TAS. At the computed time, the

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LBS will trip and capture the Iocalizer signal. The flight guidance computer now drops the

heading select mode and generates the proper roll command to bank the aircraft toward

The radio deviation signal is routed from the navigation receiver to the SG-816 Symbol

Generator on the Collins Proline II bus. From the symbol generator, the radio deviation

signal is routed to the FZ-800 through the ASCB, where the signal is lateral gained

programmed.

Lateral gain programming is required to adjust the gain applied to the Iocalizer signal due

to the aircrafi approaching the Iocalizer transmitter and beam convergence caused by the

directional qualities of the Iocalizer transmitter. The lateral gain programmer is controlled

by a distance from transmitter estimator. The distance estimator is actually a low 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, then distance is calculated using

radio altitude and glideslope deviation data. If only radio altitude is valid, distance is first

estimated for capture and then when

in

the

final track 2 mode, it is assumed that an

approach to the runway is being made without glideslope, and distance is calculated

based on radio altitude only.

If radio altitude information is not valid, then distance is estimated as a function of

glideslope deviation and TAS. If neither radio altitude nor glideslope data is valid, then

distance is estimated as a function of TAS.

From the lateral gain programmer, the Iocalizer signal is filtered, amplified, and summed

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with the course error signal. The resultant Iocalizer command signal is then course cut

limited. The course cut limiter functions primarily when approaching Iocalizer beam center

With the aircraft outside the normal Iocalizer capture limits, the EADI wil l annunciate the

fol lowing modes at this time:

HDG SEL in green

LOC in white

GS in white

Any other vertical mode in use at this time will also be annunciated on the EADI. At

Iocalizer capture, the EADI will annunciate:

LOC* in green

GS in white

Any other vertical mode in use at the time

The fl ight guidance computer now generates a roll command to smoothly capture and

track the Iocalizer signal. Wfih the Iocalizer signal captured, the ainxaft proceeds inbound

and at the computed time, wil l automatically capture and track the glideslope signal.

At this time, the EADI will annunciate:

LOC* in green

c GS* in green

The aircraft is now flying a fully coupled

ILS approach.

3. D.

(7) Back Course (BC) Mode (See figure 208, sheet 3.)

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Honeywell

3. D. (8) Roll Autopilot Mode Flow (See figure

The roll autopilot diagram shows two

MAINTENANCE

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208, sheet 4.)

signal paths for the lateral steering command. The

first path is with the autopilot disengaged and routes the lateral steering command to the

SG-816 Symbol Generator only. This path is discussed in paragraph 3.D.(8)(a). The

second path is with the autopilot engaged, and routes the lateral steering command to

both the symbol generators and to the aileron sefvo drive motor. This path is discussed

in paragraph 3.D.(8)(b).

When the autopilot is engaged and no lateral flight director mode is selected, the system

will automatically roll the aircraft wings level, and then revefi to the basic autopilot mode

of heading hold. This is discussed in paragraph 3.D.(8)(c). The roll hold mode of opera-

tion is discussed in paragraph 3. D.(8)(d), the go-around mode is discussed in paragraph

3. D.(8)(e), and the emergency descent mode (EDM) is discussed in paragraph 3.D.(8)(f).

(a)

Lateral Steering Command with Autopilot Disengaged

With the autopilot disengaged, the selected flight director steering command is

routed through the following:

Rate Limiter

Bank Angle Limit

Roll Hoki Switch

Summing Point

AP Engage Switch

Roll Bar Bias Switch

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Flight director valid comprises the following:

CPU “A” heartbeat monitor valid, which provides validity logic to

engage/disengage the autopilot functions. The heartbeat monitor is hardware

independent from the CPU, so that no single fault can disable both the CPU and

the monitor. The heartbeat monitor output is provided as an interrupt to the

nonrnonitored processor.

Flight director flag/annunciator valid, which is a direct discrete output from the

“A” CPU to drive the FD flag on the ADI and enable the annunciator drivers on

the same side guidance controller channel.

FGC power supply valid, which monitors the internal power supply voltages for

proper operating levels.

3. 0. (8) (b)

Lateral Steering Command with Autopilot Engaged

With the autopilot engaged, the selected flight director

through the:

. Rate Limiter

Bank Angle Limiter

. Roll Hold Switch

. First Summation Point

~M Switch

Second Summation Point

steering command is routed

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Second, the rate feedback signal is integrated to obtain position feedback, gain

adjusted and summed with the steering command. When these signals are equal,

the aileron is in the proper position to satisfy the steering command. As the aircraft

responds, roll attitude and roll rate information provided by the AHRS, is summed

with the steering command prior to the t45° bank limiter. This allows the feedback

position signal to drive the aileron back to its neutral position.

The aileron position synchro output adds to the tach generator signal to

compensate for any deadzone in the aileron rigging.

As the

steering command is satisfied and diminishes in size, the roll attitude signal

bermmes dominant and provides a command to move the aileron in the opposite

direction, to return the aircraft to a wings level attitude. The sewo loop 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 is sent to the command rate taker. The signal is changed

to rate, and summed with tach generator rate feedback. The summing of these two

signals is then integrated to obtain position data and summed with the steering

command. This boost helps eliminate flight path or attitude standoffs.

3. D. (8) (C) Wings Level and Heading Hold Mode

If the autopilot is engaged and no flight director mode has been selected, then a

zero roll command becomes the desired steering command. This zero command is

rouhxf through the following:

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3. D. (8) (d)

When we are HDG HOLD (refer to figure 208, sheet 1), heading information from

the AHRS is routed to the heading hold reference synchronizer. When the system

recognizes the heading hold criteria as being true, the HDG HOLD switch opens,

locking the reference heading in the synchronizer. Actual heading information is

now compared to the reference heading. Any difference between the two is TAS

gain programmed and routed to figure 208, sheet 4. The heading hold error signal

is now routed to the aileron servo drive, as previously described in paragraph

3.D.(8)(b).

Roll Hold Mode

The autopilot recognizes the roll hold mode as being operational when:

No lateral flight director mode is selected.

The aircraft bank angle is greater than 6 degrees.

Touch Control Steering (TCS), was used to initiate the bank maneuver.

The roll hold mode can be used by the pilot to maneuver the aircraft into a bank

and utilize the autopilot to hold the bank angle.

With the roll hold criteria being met, roll attitude information from the AHRS is

entered into the roll hold reference block. Wtih the ROLL HOLD switch activated to

the up position, desired roll attitude is compared against actual roll attitude at the

second summation point. Since these signals are equal and opposite, no

command is issued to the aileron sefvo drive, and the autopilot maintains the

8/11/2019 spz-8000 difcs


desired bank angle.

The autopilot generates a pitch down command proportional to the difference

between IAS and VMO. This command signal is limited to 6 degrees,

Simultaneous to the pitch down maneuver, the autopilot commands a 35 degree

bank angle for a perimf of approximately 48 seconds. During the emergency

descent, the system performs a flare computation of the form (h + 20 hs O) where

h is the difference between the aircraft altitude (ft) and the 15,000 ft flare altitude,

and h is the aircraft vertical speed (ft/see). At the point the flare computation is

satisfied, the autopilot switches into the EDM FLARE mode. In this mode, the

autopilot generates a pitch command proportional to the difference between the

aircraft altitude and the flare altitude. The system remains latched into the EDM

FLARE mode until cancelled by disengaging the autopilot.

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22- 05=07

Pages 263/264

Jun 1/93

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3. E. Pitch Channel Functional Operation

(1) Flight Director Pitch Attitude Hold Mode (See figure 209, sheet 1.)

The pitch attitude hold mode is the basic vertical fl ight director mode. It is activated when

a fl ight director roll mode is selected without an accompanying pitch mode and is not

annunciated on the EADI. The pitch command on the EADI provides the pilot with a pitch

reference corresponding to the pitch attitude existing at the moment the rol l mode was

selected. This pitch reference may be changed with the TCS button located on the pilot’s

and copilot’s control wheel.

The reference pitch attitude may also be changed as a function of the pitch wheel on the

GC-81O when the autopilot is engaged. [Refer to discussion in paragraph 3. E.(8)(c).]

Prior to the mode being operative, AHRS pitch attitude information is applied to a

summation point and then routed through a closed [PITCH HOLD + (AP ENG . T=s

NO VERT F/D MODE)] switch to the input of a synchronizer. The output of the

synchronizer is of opposite polarity to the pitch attitude signal, and therefore the two

signals cancel each other. This results in a zero signal out of the summation point.

When only a lateral flight director mode is selected, the [PITCH HOLD + (AP ENG o=S

“NO VERT F/D MODE)] switch opens. This clamps the synchronizer output as a

reference for the pitch hold mode. As long as the pitch attitude of the aircraft remains

unchanged, there will be no command to drive the pitch cue on the EADI. If the aircraft

deviates from the reference atti tude established at mode engagement, an error signal

corresponding to the difference between the actual aircraft attitude and the reference

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atti tude will drive the pitch fl ight director cue in the proper direction to fly the aircraft back

The output of the integrator is rate limited and fed to a summing point where it is

compared to actual aircraft vertical speed. Should a difference between these signals

occur, the difference is TAS gain programmed and routed as a VS command signal to

figure 209, sheet 6. TAS gain programming accurately adjusts the VS command signal as

a function of the aircraft’s current speed and barometric altitude. The VS command signal

will driie the flight director command cue on the EADI in the proper direction to fly the

aircraft back to the pilot selected vertical speed. As the aircraft returns to the reference

vertical speed, the VS command will decrease towards zero. The aircraft has now

returned to the vertical speed reference. For discussion of the vertical speed command

signal as it is processed on figure 209, sheet 6, refer to paragraph 3. E.(8).

3. E.

(3)

Flight Level Change (FLC) Mode (See figure 209, sheet 2.)

Activation of the FLC pushbutton on the GC-81 O Flight Guidance Controller selects the

Flight Level Change mode and overrides all active pitch F/D modes except VNAV. The

FLC mode will fly to the airspeedhnach reference which is displayed on the EADI. The

speed target is selectable by the pilot to be either IAS or MACH as a function of the

change over (C/0) pushbutton on the GC-81 O Flight Guidance Controller.

The FLC mode is set up to change level from present altitude to the preselected altitude.

It will ty to maintain the speed reference over the long term and allow vertical speed to

change, as a function of power setting. For example, throttle retard in a climb will cause

the system to track the speed reference, while bleeding off vedical speed.

In the FLC mode, the AFCS should fly to the new preselect altitude at the target speed

from EFIS when aircraft thrust is set appropriately for climb or descent. When the power

is not set appropriately, then the AFCS should maintain zero vertical speed in order not to

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In the second instance, with the mode engaged and the speed synchronized to existing

aircraft speed, the pilot advances the throttles to maintain the speed reference during the

FLC maneuver. Initially, the aircraft starts to accelerate. The increase in TAS and

longitudinal acceleration is changed to a potential speed rate, with normal acceleration

added as a damping term. This potential speed rate is changed to an altitude rate signal;

and the commanded vertical speed signal is processed, as previously discussed, is routed

to figure 209, sheet 6, and is processed as discussed in paragraph 3.E.(8).

3. E. (4)

Altitude Hold Mode (See figure 209, sheet 3.)

The altitude hold mode is a vertical axis flight director mode used to maintain a barometric

altitude reference. The vertical command for altitude hold is displayed on the flight

director command cue on the EADI. To fly utilizing altitude hold, the pilot would:

. Be in any lateral flight director mode

Press the ALT button on the GC-81 O

At this time, the green ALT annunciator is displayed on the EADI while altitude hold is

active. The vettical axis of the flight director will maintain the barometric altitude at the

time of mode engagement. The reference altitude may be changed by using TCS to

maneuver to a new altitude and then releasing the TCS button. Using the pitch wheel

cancels the ALT mode.

Prior to mode engagement, barometric altitude information txovided by the selected

DADC is routed through a summing junction and a closed ALT HOLD + TCS switch to the

input of the altitude hold reference synchronizer. The synchronizer develops an output

equal in amplitude but opposite in polarity to its input. The synchronizer output will sum

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After the .25 G limiter, the altitude error signal is summed with washed out pitch atti tude.

Long term pitch atti tude is washed out so that the aircraft will maintain the pilot desired

attitude. From the summing junction, the ALT SEL CMD signal is filtered, rate limited and

summed with IVV (Instantaneous Vettical Velocity).

IVV is a combination of vertical (normal) acceleration and altitude rate. This signal is

used as a damping term and is summed with the ALT SEL CMD signal to enhance the

smoothness of the flare maneuver. The aircraft will remain in the ALT SEL capture mode

until the following conditions exist simultaneously:

ALT SEL CAP

ALT error is less than 25 feet

ALT rate is less than 5 fthec

At this time, the ALT SEL mode is dropped and the aircraft is automatically placed in the

altitude hold mode.

After being summed with IW, the ALT SEL CMD signal is TAS gain programmed and

routed to figure 209, sheet 6, and is processed as discussed in paragraph 3. E.(8).

3. E. (6) Glideslope (APP) Mode (See figure 209, sheet 5.)

The glideslope mode is used for the automatic intercept, capture and tracking of the

glideslope beam. The beam is used to guide the aircraft down to the runway in a linear

descent. Typical glideslope beam angles vary between two and three degrees,

dependent on local terrain, When the glideslope mode is used as the vertical portion of

the Iocalizer approach mode, it allows the pilot to fly a fully coupled ILS approach. The

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At this time, the EADI will annunciate:

LOC in green

GS* in green

The asterisk denotes the capture phase of mode operation.

The glideslope deviation signal is routed to the SG-816 Symbol Generator from the

navigation receiver on the Collins Proline II bus. From the symbol generator, the signal is

routed to the FZ-800 Flight Guidance Computer through the ASCB.

Gain programming is performed on the glideslope signal to compensate for the aircraft

closing on the glideslope transmitter, and beam convergence caused by the directional

properties of the glideslope antenna. Glideslope programming is normally accomplished

as a function of radio altitude and veftical speed. The radio altitude signal is rate limited,

summed with vertical speed and limited again before gain programming the glideslope

signal. [f the radio altimeter is not valid, then GS gain programming is accomplished as a

function of preset height above runway estimates and run down as a function of true

airspeed. From the GS gain programming block, the glideslope signal is filtered, rate

limited and summed with estimated vertical deviation rate.

Estimated vertical deviation rate is used as a damping term to help maintain a truer track

of the glideslope beam. The estimator util izes vertical acceleration provided from the

AHRS, along with glideslope deviation, to provide an inertially derived vertical rate, with

long term glideslope deviation correction.

The summation of glideslope deviation and vertical deviation rate is then TAS gain

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3.

E.

(6.1) (b)

Vertical Altitude Select (VASL) (See figure 209, sheet 5.2.)

VASL operates identically to ASEL. VASL will arm as soon as VFLC or VPTH is

engaged. When the mode captures, VALT in green will be displayed on the EADI.

The mode annunciation wil l flash for 5 seconds to indicate the transition from arm

to capture. VASL is cancelled whenever VALT mode engages. For the altitude

preselect mode description, refer to paragraph 3. E.(5).

(c)

Vertical Attitude Hold (VALT)

VALT operates identically to ALT. VALT will engage automatically after VASL has

captured the target altitude. VALT will also engage whenever the VNAV

pushbutton is activated and the aircraft is within 250 feet of the FMS target altitude.

The FMS ALT mode is annunciated on the EADI by a green VALT. For the Altitude

Hold mode description, refer to paragraph 3.E.(4).

(d) Vertical Path (VPTH) Mode (See figure 209, sheet 5.3.)

VPTH mode is used to fly a fixed flight path angle to a vertical waypoint during

descent. VPTH mode will engage whenever the FMS initiates a path descent

which may occur while in VFLC or VALT modes. When the mode captures, VPTH

in green will be displayed on the EADI. The mode annunciation will flash for 5

seconds to indicate the transition from arm to capture. VPTH mode will be

cancel led by VASL mode capture.

On the FMS CD-800/81 O Control Display Unit (CDU), the pilot has entered the

following parameters to fly a VPTH descent:

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3. E. (7) Dual HSI Approach Mode

During the tracking phase of an lLS/MLS approach, the system will util ize landing aid

vertical fl ight path information from both the pilot’s and copilot’s HSI. This dual phase

shall provide for sensor fail-operational performance through sensor redundancy

management for the safety critical segment of the approach. Initiation of this flight

segment of the approach phase is automatic.

When both the Iocalizer and glideslope signals are on track, radio altitude is below 1200

feet and both navigation receivers are valid, the system will transition to the dual HSI

mode of operation. When this mode is active, both HSI SEL arrows on the GC-81 O will

light. In this mode, both fl ight guidance computers are using information from both

navigation receivers. This allows the approach to be continued in the event of a failure of

one navigation receiver. Should one receiver fai l, the arrow associated with that receiver

on the GC-81 O wil l extinguish and the approach mode will remain active.

(8)

Pitch Autopilot Mode Flow (See figure 209, sheet 6.)

The pitch autopilot diagram shows two signal paths for the vertical command, The first

path is with the autopilot disengaged and routes the vetiical command to the SG-816

Symbol Generator for the EADI onty. This path is discussed in paragraph 3.E.(8)(a). The

second path is with the autopilot engaged, and routes the vertical command to both

symbol generators for the EADI display and to the elevator servo drive motor, This path

is discussed in paragraph 3. E.(8)(b).

When the autopilot is engaged and no vertical flight director mode is selected, the system

will automatically revert to the basic autopilot mode of pitch attitude hold. This is

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Second, the rate feedback signal is integrated to obtain position feedback, gain

adjusted and summed with the steering command. When these signals are equal,

the elevator is in the proper position to satisfy the vettical command. As the aircraft

responds, the flight director command diminishes and the position feedback signal

drives the elevator servo back to its original position.

Should there be a mismatch between vertical command and elevator servo

position, a flight path standoff could occur. To prevent this standoff, any command

at the output of the second pitch limiter is routed through a command rate taker

and limiter.

The signal is changed to rate, and summed with tach generator rate feedback. The

summing of these two signals is then integrated to obtain position data and

summed with the vertical command.

3. E. (8) (C)

Autopilot Pitch Attitude Hold (See figure 209, sheet 1.)

Pitch attitude hold is the basic vertical autopilot mode. It is automatically active if

the autopilot is engaged and no vertical flight director mode has been selected.

Prior to autopilot engagement, pitch attitude is routed thrgh a summing point and

through the normally closed PITCH HOLD + (AP ENG .TCS oNO VERT F/D

MODE) switch to a synchronizer. The output of the synchronizer is inverted and

summed with pitch attitude to give a zero output from the summing point.

If the autopilot is engaged and no vettical flight director mode is selected, the

synchronizer switch opens, clamping the synchronizer with the reference pitch

attitude at the time of autopilot engagement. Should the aircraft deviate from the

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Honeywell

3. E.

(9) Autopilot Pitch Trim (See figure 209,

MAINTENANCE

MANUAL

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sheet 8.)

In the process of mntrolling the pitch axis of the aircraft, the elevator must maintain

certain required surface positions. Maintaining these various surface positions may

require the elevator servo to hold the elevator against a constant air load. The elevator

servo must exert a torque sufficient to hold the surface. Torque is proportional to the

amount of electrical current needed by the servo motor to hold that surface position.

When the current reaches a specified threshold, a signal is generated to operate the

elevator trim actuator, which in turn deflects the horizontal stabilizer. The horizontal

stabilizer is driven in the proper direction and amount to relieve the aerodynamic loading

on the elevator. This reduces the current level in the servo motor and allows the elevator

to return to its neutral position. When elevator setvo current drops below the trim

threshold

limit, the elevator trim actuator stops, and the horizontal stabilizer is in a new

position to hold the aircraft’s desired pitch attitude.

When the trim threshold current has been exceeded as shown in figure 209, sheet 8, a

trim threshold sensor will apply the trim drive signal to a trim up and a trim down sensor.

These sensors are polariiy detectors and determine which direction the trim must run. At

the same time, the trim drive signal is properly gained according to the flap position and

transition status. The trim gain will be increased when the flaps are in motion to

compensate for the change in lift. The trim drive signal is then applied to a time delay

circuit. If the flaps are in motion, a time delay of one second will

occur before the trim

begins to move. Any other time the delay will be 3 seconds.

Dependent on which direction the trim must run, the appropriate logic AND gate will

provide a signal to the pitch trim interface circuitry. The trim drive output will drive the trim

actuator in the proper direction which in turn,

moves

the horizontal stabilizer. Limit

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switches are employed to protect against extreme trim demands.

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Pages 279/280

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Pages 281/282

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Pages 283/284

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MAINTENANCE

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22- 05=07

Pages 291/292

Jun 1193

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pages 298.3/298.4

8/11/2019 spz-8000 difcs


3. F. Yaw Channel Functional Operation (See figure 210, sheets 1 and 2.)

The yaw axis of the autopilot provides directional stability (yaw damping) and directional control

for turn coordination. The yaw axis of the autopilot receives sensor information from the AHRS

and the DADC. AHRS supplies the following information through the ASCB.

Yaw rate

. Roll rate

. Pitch attitude

. Roll attitude

Normal acceleration

Longitudinal acceleration

Lateral acceleration

The DADC supplies the following information through the ASCB.

. Indicated airspeed (IAS)

True airspeed (TAS)

Altitude rate (VS)

The above inputs from AHRS and the DADC are all combined in the FZ-800’S rudder command

processor. (See figure 210, sheet 2.) The rudder command processor will determine the proper

rudder deflection to maintain directional stability and control.

Yaw rate, true airspeed, roll attitude, and lateral acceleration are the primary controlling inputs

for the yaw axis. The rudder command processor looks at yaw rate and computes the control

response necessary to bring the yaw rate of the aircraft to zero, True airspeed, roll attitude, and

lateral acceleration combine to provide turn coordination. The remaining inputs to the processor

8/11/2019 spz-8000 difcs


The output of the servo amplifier is sent to the SM-200 rudder servo drive motor and the “A”

processor current monitor to check for servo runaway current. The SM-200 is a permanent

magnet dc motor that utilizes a dc tachometer for rate feedback. It does not use a position

feedback transducer. As the servo motor drives the rudder, it also drives the dc tach generator

through mechanical coupling (represented by a dotted line). The tach generator provides a rate

feedback signal that serves two functions. First, it acts as a damping term when summed with

the yaw command input to the pulse width command limiter. This helps to stabilize rudder

position and minimize excessive rudder travel.

Second, the rate feedback signal is integrated to obtain position feedback and summed with the

yaw command. When these signals are equal, the rudder is in the proper position to satisfy the

yaw command. As the aircraft responds, the yaw command diminishes and the position

feedback signal drives the rudder servo back to its original position.

MAINTENANCE

8/11/2019 spz-8000 difcs


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Yaw Channel Mode Flow Diagram

Figure 210 (Sheet 1 of 2 )

Pages 298,7/298.8

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