PRIMUS 1000

Integrated Avionics System

Pilot's Manual for the
Embraer 145

AD-52817@
Business and Commuter Aviation Systems
Honeywell Inc.
Box 29000
Phoenix, Arizona 85038--9000

PRIMUSr1000 Integrated
Avionics System

for the
Embraer Model 145

Pilot’s Manual

Printed in U.S.A. Pub. No. A28--1146--112--00 August 1997

 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
(18 USC 1905)

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

information disclosed herein falls within exemption (b) (4) of 5 USC 552
and the prohibitions of 18 USC 1905.

S97

ASSOCIATE
MEMBER

Member of GAMA

General Aviation
Manufacturer’s Association

PRIMUS is a trademark of Honeywell Inc.

E 1997 Honeywell Inc.

 PRIMUSr1000 Integrated Avionics System

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Upon receipt of a revision, insert the latest revised pages and dispose
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 PRIMUSr1000 Integrated Avionics System

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4--55 0 System (EICAS)
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Honeywell Product Support C--11 0
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 PRIMUSr1000 Integrated Avionics System

Table of Contents

Section Page

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

2. SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . . . . . 2-1

Flight Guidance System (FGS) . . . . . . . . . . . . . . . . . 2-3
Electronic Display System (EDS) . . . . . . . . . . . . . . . 2-4
Air Data System (ADS) . . . . . . . . . . . . . . . . . . . . . . . 2-5
PRIMUSâ 660 Weather Radar System . . . . . . . . . . 2-6
Attitude and Heading Reference System (AHRS) . 2-6
PRIMUSâ II Integrated Radio System . . . . . . . . . . . 2-7
Radio Altimeter System . . . . . . . . . . . . . . . . . . . . . . . 2-7
Traffic Alert and Collision Avoidance System
(TCAS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-8
PRIMUSâ 880 Weather Radar System (Option) . . 2-8
Flight Management System (FMS) (Option) . . . . . . 2-9
Other Switches and Controls . . . . . . . . . . . . . . . . . . . 2-9

3. ELECTRONIC DISPLAY SYSTEM (EDS) . . . . . . . 3-1

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-1
MFD Bezel Controller Basic Operation . . . . . . . . . . 3-5
Map Range Control . . . . . . . . . . . . . . . . . . . . . . . . 3-6
Controller Conventions . . . . . . . . . . . . . . . . . . . . . 3-6
Reversionary Controllers and Dimming Panel . . . . 3-7
System Reversionary Panel . . . . . . . . . . . . . . . . 3-7
Data Acquisition Unit (DAU) Reversionary Panel 3-8
EDS Dimming Panel . . . . . . . . . . . . . . . . . . . . . . . 3-8

4. ELECTRONIC FLIGHT INSTRUMENT SYSTEM

(EFIS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-1
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-5
Primary Flight Display (PFD) Bezel Controller . . . 4-5
Multifunction Display (MFD) Bezel Controller . . . . 4-6
Display Controller (DC) . . . . . . . . . . . . . . . . . . . . . 4-20
Guidance Controller Course and Heading
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-22
EFIS Preflight Self--Test . . . . . . . . . . . . . . . . . . . . 4-23
Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-25
Primary Flight Display (PFD) . . . . . . . . . . . . . . . . 4-25
Typical PFD Presentations . . . . . . . . . . . . . . . . . . 4-53
PFD Caution and Failure Display . . . . . . . . . . . . 4-61

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Table of Contents (cont)

Section Page

4. ELECTRONIC FLIGHT INSTRUMENT SYSTEM

(EFIS) (CONT)
Multifunction Display (MFD) . . . . . . . . . . . . . . . . . 4-68
MFD System Displays . . . . . . . . . . . . . . . . . . . . . 4-86
System Display Failure Annunciation . . . . . . . . . 4-95
Standby Navigation Display . . . . . . . . . . . . . . . . . 4-96

5. ENGINE INSTRUMENT AND CREW ALERTING

SYSTEM (EICAS) . . . . . . . . . . . . . . . . . . . . . . . . . . 5-1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-1
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-5
EICAS Bezel Controller . . . . . . . . . . . . . . . . . . . . 5-5
Master Warning and Caution Annunciation
Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-5
Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-6
EICAS Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-6
Primary Engine Displays . . . . . . . . . . . . . . . . . . . 5-7
Crew Alerting System (CAS) . . . . . . . . . . . . . . . . 5-22
CAS Messages . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-23
CAS Display Failure Indications . . . . . . . . . . . . . 5-29
EICAS Reversionary Modes . . . . . . . . . . . . . . . . 5-30
Backup EICAS Display . . . . . . . . . . . . . . . . . . . . . 5-31
Backup EICAS Display Failure Indications . . . . 5-34

6. FLIGHT GUIDANCE SYSTEM (FGS) . . . . . . . . . . . 6-1

Guidance Controller (GC) . . . . . . . . . . . . . . . . . . . . . 6-1
Autopilot (AP)/Yaw Damper (YD)/Couple (CPL)
Buttons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-1
Flight Director (FD) Mode Selection . . . . . . . . . . 6-3
Remote Autopilot Switches and Annunciators . . . 6-5
Autopilot Power--Up Test . . . . . . . . . . . . . . . . . . . 6-5
FGS Caution and Advisory Messages . . . . . . . . 6-5
Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-6
Pitch Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-6
Turn Knob . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-7

7. SYSTEM LIMITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-1

Glossary of Terms . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-1
Attitude Director Indicator (ADI) Command
Cue . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-1
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Section Page

7. SYSTEM LIMITS (CONT)

Glideslope (GS) Gain Programming as a Radio
Altitude (RA) Function . . . . . . . . . . . . . . . . . . . . 7-2
GS Gain Programming as a RA and Vertical
Speed (VS) Function . . . . . . . . . . . . . . . . . . . . . 7-2
GS Capture (CAP) . . . . . . . . . . . . . . . . . . . . . . . . 7-2
Lateral Beam Sensor (LBS) . . . . . . . . . . . . . . . . . 7-2
Lateral Gain Programming . . . . . . . . . . . . . . . . . . 7-3
Localizer/Back Course (BC) Track . . . . . . . . . . . 7-3
True Airspeed (TAS) Gain Programming . . . . . . 7-3
Vertical Beam Sensor (VBS) . . . . . . . . . . . . . . . . 7-4
VOR Capture (CAP) . . . . . . . . . . . . . . . . . . . . . . . 7-4
VOR Over Station Sensor (OSS) . . . . . . . . . . . . 7-4
VOR After Over Station Sensor (AOSS) . . . . . . 7-4
System Performance and Operating Limits . . . . . . 7-5

8. MODES OF OPERATION . . . . . . . . . . . . . . . . . . . . . 8-1

Heading (HDG) Hold Mode, Wings Level . . . . . . . . 8-1
Roll Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-2
Heading Select Mode . . . . . . . . . . . . . . . . . . . . . . . . . 8-3
VOR Navigation (NAV) Mode . . . . . . . . . . . . . . . . . . 8-4
VOR Approach (VAPP) Mode . . . . . . . . . . . . . . . . . . 8-9
Flight Management System (FMS) Navigation
Mode (Option) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-10
Localizer (LOC) Mode . . . . . . . . . . . . . . . . . . . . . . . . 8-12
Back Course (BC) Mode . . . . . . . . . . . . . . . . . . . . . . 8-17
Instrument Landing System (ILS) Approach Mode . . 8-20
Pitch Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-25
Pitch Hold Mode Operation, AP Engaged . . . . . 8-25
Pitch Hold Mode Operation, AP Not Engaged . . . 8-25
TakeOff Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-26
Vertical Speed Hold Mode . . . . . . . . . . . . . . . . . . . . . 8-27
Speed Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-28
Flight Level Change (FLC) Mode . . . . . . . . . . . . . . . 8-31
Altitude Preselect (ASEL) Mode . . . . . . . . . . . . . . . . 8-32
Altitude Hold (ALT HOLD) Mode . . . . . . . . . . . . . . . . 8-37
Windshear (WDSHEAR) Mode . . . . . . . . . . . . . . . . . 8-38
Go--Around (GA) Mode (Wings Level) . . . . . . . . . . . 8-40

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Table of Contents (cont)

Section Page

9. TROUBLESHOOTING . . . . . . . . . . . . . . . . . . . . . . . 9-1
Pilot Writeup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-1
Commonly Used Terms . . . . . . . . . . . . . . . . . . . . 9-3
Typical Problems . . . . . . . . . . . . . . . . . . . . . . . . . . 9-4
Ground Maintenance Test . . . . . . . . . . . . . . . . . . . . . 9-9
Multifunction Display (MFD) Checklist Upload
Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-11

10. HONEYWELL PRODUCT SUPPORT . . . . . . . . . . 10-1

Publication Ordering Information . . . . . . . . . . . . . 10-4

11. ABBREVIATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . 11-1

APPENDICES

A--1 AHZ--800 ATTITUDE AND HEADING REFERENCE

SYSTEM (AHRS) . . . . . . . . . . . . . . . . . . . . . . . . . . . A--1
Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--1
Attitude and Heading Reference Unit . . . . . . . . . . . . A--5
AHRS Control and EICAS Annunciation . . . . . . . . . A--5
AHRS Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--5
EICAS Annunciations . . . . . . . . . . . . . . . . . . . . . . A--6
Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--7
Standard Operation . . . . . . . . . . . . . . . . . . . . . . . . A--7
Reduced Performance Operation . . . . . . . . . . . . A--8
Power--Up Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--8
Pilot Self--Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--9
Ground Initialization . . . . . . . . . . . . . . . . . . . . . . . . . . A--9
Abnormal Operations . . . . . . . . . . . . . . . . . . . . . . A--10

B--1 PRIMUSâ 660 WEATHER RADAR SYSTEM . . . . B--1

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . B--1
Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . B--1
Weather Radar Controller . . . . . . . . . . . . . . . . . . . . . B--2
Controls and Indicators . . . . . . . . . . . . . . . . . . . . . B--2
Normal Operation . . . . . . . . . . . . . . . . . . . . . . . . . B--8
Maximum Permissible Exposure Level (MPEL) . . . . B--12

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Table of Contents (cont)

Section Page

C--1 PRIMUSâ II RADIO SYSTEM . . . . . . . . . . . . . . . . . . C--1

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--1
System Description . . . . . . . . . . . . . . . . . . . . . . . . . . . C--5
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--5
Integrated Navigation Unit . . . . . . . . . . . . . . . . . . C--5
Integrated Communication Unit . . . . . . . . . . . . . . C--6
Radio Management Unit (RMU) . . . . . . . . . . . . . C--6
Audio Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--6
Tuning Backup Control Head . . . . . . . . . . . . . . . . C--7
Radio Management Unit (RMU) Operation . . . . . . . C--7
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--7
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--8
Memory Pages . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--11
System (SYS) On--Off Page . . . . . . . . . . . . . . . . C--11
VHF Communications Transceiver Operation . . C--13
Navigation Receiver Operation . . . . . . . . . . . . . . C--17
ATC Transponder and TCAS Operations . . . . . C--20
ADF Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--24
Cross--Side Operation . . . . . . . . . . . . . . . . . . . . . C--27
Built--In--Test (BIT) . . . . . . . . . . . . . . . . . . . . . . . . . C--28
Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--32
RMU Display Dimming . . . . . . . . . . . . . . . . . . . . . C--32
Audio Panel Operation . . . . . . . . . . . . . . . . . . . . . . . . C--33
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--33
Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--33
Tuning Backup Control Head Operation . . . . . . . . . C--35
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--35
Normal Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--36
Emergency Mode . . . . . . . . . . . . . . . . . . . . . . . . . C--36
Controls and Annunciations . . . . . . . . . . . . . . . . . C--37

INDEX . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Index--1

A28--1146--112--00 Table of Contents

TC--5
PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Illustrations
Figure Page
1--1 PRIMUSâ 1000 in the Embraer 145 Cockpit . . . . . 1-5

2--1 PRIMUSâ 1000 Block Diagram . . . . . . . . . . . . . . . . . 2-11

3--1 Cockpit Layout of EFIS . . . . . . . . . . . . . . . . . . . . . . . 3-3

3--2 MFD Bezel Controls . . . . . . . . . . . . . . . . . . . . . . . . . . 3-5
3--3 System Reversionary Panel . . . . . . . . . . . . . . . . . . . 3-7
3--4 DAU Reversionary Panel . . . . . . . . . . . . . . . . . . . . . . 3-8

4--1 EFIS Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . 4-3

4--2 PFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . 4-5
4--3 MFD Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . . . 4-6
4--4 Menu Inop Display . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-6
4--5 MFD Menu Tree . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7
4--6 Main Menu . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-8
4--7 System Menu Selections . . . . . . . . . . . . . . . . . . . . . . 4-8
4--8 Main Menu -- MFD Button . . . . . . . . . . . . . . . . . . . . . 4-10
4--9 MFD Main Menu (Without FMS) . . . . . . . . . . . . . . . . 4-10
4--10 MFD Main Menu (With FMS) . . . . . . . . . . . . . . . . . . 4-11
4--11 VSPEEDS Submenu . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-11
4--12 Joystick Submenu . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-13
4--13 Checklist Menu . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-15
4--14 TCAS Selection from Main Menu . . . . . . . . . . . . . . . 4-18
4--15 WX Selection from Main Menu . . . . . . . . . . . . . . . . . 4-18
4--16 MAP Plan Selection From Main Menu . . . . . . . . . . . 4-19
4--17 Display Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-20
4--18 Guidance Controller . . . . . . . . . . . . . . . . . . . . . . . . . . 4-22
4--19 PFD Functional Divisions . . . . . . . . . . . . . . . . . . . . . . 4-25
4--20 ADI and Annunciation Display . . . . . . . . . . . . . . . . . . 4-27
4--21 PFD With CAT II Annunciations . . . . . . . . . . . . . . . . 4-35
4--22 HSI Compass Display On PFD . . . . . . . . . . . . . . . . . 4-37
4--23 HSI Arc Display With FMS Map and Weather
Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-41
4--24 Airspeed Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-45
4--25 Altimeter Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-48
4--26 Vertical Speed Display . . . . . . . . . . . . . . . . . . . . . . . . 4-51
4--27 Takeoff Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-53
4--28 Climb to Initial Altitude . . . . . . . . . . . . . . . . . . . . . . . . 4-54
4--29 Enroute Cruise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-55
4--30 Setup for Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-56
Table of Contents A28--1146--112--00
TC--6
 PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Illustrations (cont)
Figure Page
4--31 TCAS II Resolution Advisory Approaching the
Glideslope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-57
4--32 Approach Capture Tracking at DH . . . . . . . . . . . . . . 4-58
4--33 Comparison Monitor Annunciators . . . . . . . . . . . . . . 4-59
4--34 PFD With ARC HSI Display and Caution/Failure
Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-61
4--35 PFD Caution Indications for Vertical and Course
Deviation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-63
4--36 Failure Indications for ATT, MADC, HDG, CRS
Select and FD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-65
4--37 Excessive Attitude Declutter . . . . . . . . . . . . . . . . . . . 4-66
4--38 PFD Test Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-68
4--39 MFD Common Symbols . . . . . . . . . . . . . . . . . . . . . . . 4-69
4--40 MFD Map Display Symbols . . . . . . . . . . . . . . . . . . . . 4-72
4--41 Map Mode With Weather . . . . . . . . . . . . . . . . . . . . . . 4-74
4--42 Map Failure/Warning . . . . . . . . . . . . . . . . . . . . . . . . . . 4-75
4--43 Map Mode Without Weather . . . . . . . . . . . . . . . . . . . 4-76
4--44 Map Mode With Weather . . . . . . . . . . . . . . . . . . . . . . 4-77
4--45 MFD Plan View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-78
4--46 Plan Mode Caution and Warning Displays . . . . . . . 4-79
4--47 Typical Plan Mode With Navaid Components and
Flight Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-80
4--48 Typical Checklist Display . . . . . . . . . . . . . . . . . . . . . . 4-83
4--49 MFD With Map Mode Display and Optional TCAS 4-84
4--50 Takeoff Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-87
4--51 Environmental Control System (ECS) System
Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-88
4--52 Fuel Page . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-90
4--53 Hydraulic (HYD) System Page . . . . . . . . . . . . . . . . . 4-92
4--54 Electrical (ELEC) System Display . . . . . . . . . . . . . . 4-93
4--55 Standby Navigation Display (Typical) . . . . . . . . . . . . 4-97
4--56 RMU Backup Navigation Display . . . . . . . . . . . . . . . 4-98
4--57 RMU Backup Navigation Display With Failures . . . 4-101

5--1 EICAS Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . 5-3

5--2 EICAS Bezel Controller . . . . . . . . . . . . . . . . . . . . . . . 5-5
5--3 Master Warning and Caution Annunciation Switches . . 5-5
5--4 EICAS Display Functional Divisions . . . . . . . . . . . . . 5-6
5--5 EICAS Engine Instruments Display . . . . . . . . . . . . . 5-8
5--6 EICAS in the Takeoff Condition (Typical) . . . . . . . . 5-20
5--7 Engine Instrument Display With Failures . . . . . . . . 5-21

A28--1146--112--00 Table of Contents

TC--7
PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Illustrations (cont)
Figure Page
5--8 Crew Alerting System on EICAS Display . . . . . . . . 5-22
5--9 EICAS Invalid Display Formats . . . . . . . . . . . . . . . . . 5-30
5--10 Backup EICAS Display on RMU -- Page 1 . . . . . . . 5-32
5--11 Backup EICAS Display on RMU -- Page 2 . . . . . . . 5-33
5--12 RMU Backup Engine Display Failure Indication . . . . 5-34

6--1 Guidance Controller . . . . . . . . . . . . . . . . . . . . . . . . . . 6-1

6--2 Autopilot Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-6

8--1 Heading Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . 8-1

8--2 Roll Hold Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-2
8--3 VOR Navigation Mode Radial Intercept View . . . . . 8-5
8--4 VOR Navigation Mode Intercept PFD . . . . . . . . . . . 8-5
8--5 VOR Navigation Mode Capture Plan View . . . . . . . 8-6
8--6 VOR Navigation Mode Capture . . . . . . . . . . . . . . . . 8-7
8--7 VOR Navigation Mode Tracking . . . . . . . . . . . . . . . 8-8
8--8 VOR Navigation Mode -- Tracking Over Station . . . . 8-9
8--9 Long Range Navigation Mode Display . . . . . . . . . . . 8-11
8--10 Localizer Mode Arm Display . . . . . . . . . . . . . . . . . . . 8-13
8--11 Localizer Mode Capture Plan View . . . . . . . . . . . . . 8-14
8--12 Localizer Capture Display . . . . . . . . . . . . . . . . . . . . . 8-15
8--13 Localizer Mode Tracking Profile View . . . . . . . . . . . 8-16
8--14 Localizer Tracking Display . . . . . . . . . . . . . . . . . . . . . 8-16
8--15 Back Course Mode Intercept Plan View . . . . . . . . . 8-17
8--16 Back Course Tracking Display . . . . . . . . . . . . . . . . . 8-19
8--17 ILS Approach Localizer Intercept . . . . . . . . . . . . . . . 8-20
8--18 ILS Approach Mode Glideslope Capture, Profile
View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-23
8--19 ILS Approach Mode Tracking PFD . . . . . . . . . . . . . . 8-24
8--20 ILS Approach Mode Tracking, Profile View . . . . . . . 8-25
8--21 Takeoff Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-26
8--22 Vertical Speed Hold Mode PFD . . . . . . . . . . . . . . . . 8-28
8--23 Speed Hold Mode PFD . . . . . . . . . . . . . . . . . . . . . . . 8-29
8--24 Altitude Preselect Plan View . . . . . . . . . . . . . . . . . . . 8-34
8--26 Windshear Mode PFD . . . . . . . . . . . . . . . . . . . . . . . . 8-39
8--27 Go--Around Mode (Wings Level) PFD . . . . . . . . . . . 8-41

9--1 Lateral Mode Conditions and Problems . . . . . . . . . . 9-5

9--2 Vertical Mode Conditions and Problems . . . . . . . . . 9-7
9--3 Ground Maintenance Test Display (Typical) . . . . . . 9-10

Table of Contents A28--1146--112--00

TC--8
 PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Illustrations (cont)
Figure Page
A--1 AHRU Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . A--3
A--2 AHRS Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . A--5

B--1 Weather Radar Controller . . . . . . . . . . . . . . . . . . . . . B--2

B--2 PFD Display Weather Radar Test Pattern . . . . . . . . B--9
B--3 MFD Display Weather Radar Test Pattern . . . . . . . B--10
B--4 Radar Beam Illumination High Altitude
12--Inch Radiator . . . . . . . . . . . . . . . . . . . . . . . . . . . B--11
B--5 Radar Beam Illumination Low Altitude
12--Inch Radiator . . . . . . . . . . . . . . . . . . . . . . . . . . . B--11
B--6 MPEL Boundary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . B--12

C--1 PRIMUSâ II Radio System Interface Block Diagram C--3

C--2 Radio Management Unit Controls . . . . . . . . . . . . . . . C--8
C--3 Page Menu . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--10
C--4 COM Frequency Storage Page . . . . . . . . . . . . . . . . . C--11
C--5 System On--Off Page . . . . . . . . . . . . . . . . . . . . . . . . . C--12
C--6 RMU in COM Tuning Mode . . . . . . . . . . . . . . . . . . . . C--14
C--7 RMU Tuning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--14
C--8 Exchanging the Preset and Active Frequencies . . . . C--15
C--9 COM Message (Squelch Open) . . . . . . . . . . . . . . . . C--16
C--10 NAV Frequency Select Window . . . . . . . . . . . . . . . . C--17
C--11 Toggling Active and Preset Frequencies . . . . . . . . . C--18
C--12 DME Hold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--19
C--13 RMU ATC/TCAS Controls . . . . . . . . . . . . . . . . . . . . . C--21
C--14 ADF Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--25
C--15 Audio Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--33
C--16 Tuning Backup Control Head . . . . . . . . . . . . . . . . . . C--36
C--17 Tuning Backup Control Head Controls and
Annunciations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--37

A28--1146--112--00 Table of Contents

TC--9
PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Tables
Table Page
1--1 Supplemental Pilot’s Manuals . . . . . . . . . . . . . . . . . 1-1
1--2 Equipment List . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-2

4--1 Cursor Movement/Action . . . . . . . . . . . . . . . . . . . . . 4-16

4--2 Navigation Bearing Pointers . . . . . . . . . . . . . . . . . . 4-21
4--3 Yaw Damper Status Annunciations . . . . . . . . . . . . 4-26
4--4 Pitch Limit Indicator . . . . . . . . . . . . . . . . . . . . . . . . . 4-29
4--5 Pitch Scale Tape Markings . . . . . . . . . . . . . . . . . . . 4-29
4--6 On--Side Sensor Sources . . . . . . . . . . . . . . . . . . . . 4-31
4--7 Autopilot Status Annunciations . . . . . . . . . . . . . . . . 4-32
4--8 Marker Beacon Annunciations . . . . . . . . . . . . . . . . 4-33
4--9 Bearing Pointer Selections . . . . . . . . . . . . . . . . . . . 4-38
4--10 GSPD, TTG, ET Ranges . . . . . . . . . . . . . . . . . . . . . 4-39
4--11 Navigation Symbol Colors . . . . . . . . . . . . . . . . . . . . 4-40
4--12 Approach and Terminal Annunciators . . . . . . . . . . 4-42
4--13 Weather Radar Annunciations . . . . . . . . . . . . . . . . 4-43
4--14 Selected Source Distance Range . . . . . . . . . . . . . 4-44
4--15 VSPEED Bug Identification . . . . . . . . . . . . . . . . . . . . . 4-46
4--16 Low Speed Awareness Color Logic . . . . . . . . . . . . 4-47
4--17 Comparison Monitor Parameters . . . . . . . . . . . . . . 4-60
4--18 Yaw Damper Status Annunciations . . . . . . . . . . . . 4-64
4--19 TCAS Failure Annunciations . . . . . . . . . . . . . . . . . . 4-64
4--20 Excessive Attitude Declutter Parameters . . . . . . . 4-67
4--21 Weather Radar Mode Annunciations on MFD . . . 4-70
4--22 Designating a New Waypoint Procedure . . . . . . . . 4-81
4--23 Engine T/O Mode Logic . . . . . . . . . . . . . . . . . . . . . . 4-88
4--24 Oxygen Pressure Limits in PSI . . . . . . . . . . . . . . . . 4-89
4--25 Bleed Air Temperature Color Limits . . . . . . . . . . . . 4-89
4--26 Fuel Quantity Color Coding . . . . . . . . . . . . . . . . . . . 4-90
4--27 Hydraulic Pressure and Fluid Quantity Limits . . . 4-92
4--28 Brake Temperature Limits . . . . . . . . . . . . . . . . . . . . 4-93
4--29 Marker Beacon Colors . . . . . . . . . . . . . . . . . . . . . . . 4-99

5--1 N1 Analog Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-9

5--2 FADEC Modes of Operation . . . . . . . . . . . . . . . . . . . 5-10
5--3 ITT Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-11
5--4 Ignitor Annunciation Color Definition . . . . . . . . . . . . 5-11
5--5 N2 Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5-12
5--6 Oil Pressure Limits for N2 <88% . . . . . . . . . . . . . . . . 5-12
5--7 Oil Pressure Limits for N2 >88% . . . . . . . . . . . . . . . . 5-13

Table of Contents A28--1146--112--00

TC--10
 PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Tables (cont)
Table Page
5--8 Oil Temperature Limits . . . . . . . . . . . . . . . . . . . . . . . . 5-14
5--9 Engine Vibration Color Limits . . . . . . . . . . . . . . . . . . 5-14
5--10 Landing Gear Position Annunciations . . . . . . . . . . . 5-15
5--11 Flaps Position Symbol Color . . . . . . . . . . . . . . . . . . . 5-15
5--12 Spoiler Color Coding . . . . . . . . . . . . . . . . . . . . . . . . . . 5-16
5--13 Pitch Trim Color Description . . . . . . . . . . . . . . . . . . . 5-17
5--14 Color Differential Pressure Display . . . . . . . . . . . . . . 5-18
5--15 APU Turbine Speed as a Percentage of RPM . . . . . 5-19
5--16 APU Exhaust Temperature . . . . . . . . . . . . . . . . . . . . 5-19
5--17 Message Color Conventions . . . . . . . . . . . . . . . . . . . 5-23
5--18 Red Warning Messages . . . . . . . . . . . . . . . . . . . . . . . 5-24
5--19 Amber Caution Messages . . . . . . . . . . . . . . . . . . . . . 5-25
5--20 Cyan Advisory Messages . . . . . . . . . . . . . . . . . . . . . 5-28
5--21 Conditions for Inhibit Function . . . . . . . . . . . . . . . . . . 5-29

6--1 FD1/FD2 Switch Operation . . . . . . . . . . . . . . . . . . . . 6-4

6--2 Red (Warning) Flight Guidance System
Messages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-5
6--3 Amber (Caution) Flight Guidance System
Messages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6-6

7--1 System Performance/Operating Limits . . . . . . . . . . 7-5

7--2 Air Data Display Parameters and Ranges . . . . . . . . 7-8

8--1VOR Navigation Mode Engagement Procedure . . . . 8-4

8--2VOR Approach Mode Engagement Procedure . . . . . 8-9
8--3FMS Mode Engagement Procedure . . . . . . . . . . . . . 8-10
8--4Localizer Mode Engagement Procedure . . . . . . . . . 8-12
8--5Back Course Flight Procedure . . . . . . . . . . . . . . . . . 8-17
8--6ILS Approach Mode Procedure . . . . . . . . . . . . . . . . . 8-21
8--7Speed Mode Engagement Procedure . . . . . . . . . . . 8-30
8--8Flight Level Change Descend and Climb
Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8-31
8--9 Altitude Preselect Mode Engagement Procedure . . . 8-32
8--10 Altitude Preselect Mode Procedure . . . . . . . . . . . . . 8-34
8--11 Altitude Hold Mode Engagement Procedure . . . . . . 8-37

9--1 Definitions of Terms . . . . . . . . . . . . . . . . . . . . . . . . . . 9-3

9--2 Lateral Mode Problems . . . . . . . . . . . . . . . . . . . . . . . 9-4
9--3 Vertical Mode Problems . . . . . . . . . . . . . . . . . . . . . . . 9-7

A28--1146--112--00 Table of Contents

TC--11
PRIMUSr1000 Integrated Avionics System

Table of Contents (cont)

List of Tables (cont)
Table Page
9--4 Problems Common to Both Vertical and Lateral
Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9-9
9--5 Ground Maintenance Test Procedure . . . . . . . . . . . 9-9
9--6 MFD Checklist Upload Procedure . . . . . . . . . . . . . . 9-11
9--7 Troubleshooting Procedure . . . . . . . . . . . . . . . . . . . . 9-13

A--2 Pilot Self--Test Annunciations . . . . . . . . . . . . . . . . . . A--8

A--3 Pilot Initiated Self--Test Annunciations . . . . . . . . . . . A--9

B--1 Target Alert Characteristics . . . . . . . . . . . . . . . . . . . . B--4

B--2 Rainfall Rate Color Cross Reference . . . . . . . . . . . . B--6
B--3 PRIMUSâ 660 Weather Radar System Precautions . . B--8

C--1 COM Messages . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--16

C--2 DME Cycling Sequence . . . . . . . . . . . . . . . . . . . . . . . C--19
C--3 DME Function Key Operating Procedure . . . . . . . . C--20
C--4 Transponder Modes . . . . . . . . . . . . . . . . . . . . . . . . . . C--22
C--5 Extended Altitude Modes . . . . . . . . . . . . . . . . . . . . . . C--23
C--6 ADF Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . C--25
C--7 PAST Self--Test Approximate Durations . . . . . . . . . C--29
C--8 VHF COM PAST Procedure . . . . . . . . . . . . . . . . . . . C--29
C--9 ATC Transponder and TCAS PAST Procedure . . . . C--30
C--10 VOR/ILS/Marker/DME PAST Procedure . . . . . . . . . C--31
C--11 ADF Past Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . C--32

Table of Contents A28--1146--112--00

TC--12
 PRIMUSr1000 Integrated Avionics System

1. Introduction

This document describes the operation, components, typical flight

applications, and operating procedures for the PRIMUSR 1000
Integrated Avionics System shown as part of the cockpit layout in figure
1--1, in the Embraer 145 aircraft. Subsystems described in this manual
include:
D PRIMUSR 1000 Flight Guidance System (FGS)
D Electronic Display System (EDS)
D ADZ--850 Air Data System (ADS)
D AHZ--800 Attitude and Heading Reference System (AHRS)
D SRZ--850 PRIMUSR II Integrated Radio System
D PRIMUSR 660 Weather Radar System
D AA--300 Altimeter System.

The traffic alert and collision avoidance system (TCAS) is covered in

another manual. Optional systems, including the PRIMUSR 880
Weather Radar Systems and the NZ--2000 Flight Management System
(FMS) are described in separate pilot’s manuals. These manuals are
listed in table 1--1. The equipment covered in this manual is listed in
table 1--2.

System Publication No.

Traffic Alert and Collision A28--1146--070

Avoidance System (TCAS)
PRIMUSR 880 Weather Radar A28--1146--102
System
NZ--2000 Series Flight Refer to aircraft supplemental
Management System (FMS) flight manual for current
software versions.

Supplemental Pilot’s Manuals

Table 1--1

A28--1146--112--00 Introduction
1-1
PRIMUSr1000 Integrated Avionics System

Model Unit Qty Part No.

Cockpit Mounted Equipment

DU--870 Electronic Display Unit 5 7014300--901
BL--870 Primary Flight Display (PFD) 2 7014331--921
Bezel Controller
BL--871 Multifunction/Engine 2 (MFD)
Instrument and Crew 7014332--931
Alerting System Displays 1 (EICAS)
(MFD/ EICAS) Bezel 7014332--951
Controllers
PC--400 Autopilot Controller 1 7003897--925
DC--550 Display Controller (DC) 2 7016986--401
(without FMS)
7016986--501
(with FMS
installed)
WC--660 Weather Radar Controller 1 7008471--VAR
GC--550 Guidance Controller 1 7021170--951
AV--850A Audio Panel 3 7510100--939
RM--855 Radio Management Unit 2 7013270--VAR
(RMU)
CD--850 Tuning Backup Control Head 1 7513000--VAR
Remote Mounted Equipment
IC--600 No. 1 Integrated Avionics 1 7017000--82404
Computer (IAC)
IC--600 No. 2 Integrated Avionics 1 7017000--83404
Computer
SM--200 Servo Drives 3 4006719--910/906
SB--201B Servo Brackets 3 4005842
RT--300 Radio Altimeter 1 7001840--937

Equipment List
Table 1--2 (cont)

Introduction A28--1146--112--00
1-2
 PRIMUSr1000 Integrated Avionics System

Model Unit Qty Part No.

RNZ--851 Integrated NAV Unit 2 7510100--VAR

RCZ--851 Communications Unit with 1/2 7510700--806
Diversity Transponder
RCZ--851 Communications Unit without 1 7510700--809
Diversity Transponder
RT--910 TCAS II Computer 1 4066010--904
AT--910 TCAS Antenna 1 7514060--902
AZ--850 Micro Air Data Computer 2 7014700--918
(MADC)
DA--800 Data Acquisition Unit (DAU) 2 7013348--911
AH--800 Attitude and Heading 2 HG2010AC02
Reference Unit
WU--660 Radar Receiver Transmitter 1 7021450--601
Antenna (RTA)
Optional Equipment
ECP--800 Programmable Checklist 1 7021060--901

Equipment List
Table 1--2

A28--1146--112--00 Introduction
1-3/(1-4 blank)
 PRIMUSr1000 Integrated Avionics System

PRIMUSâ 1000 in the Embraer 145 Cockpit

Figure 1--1
A28--1146--112--00 Introduction
1-5/(1-6 blank)
 PRIMUSr1000 Integrated Avionics System

2. System Description

The PRIMUSâ 1000 Integrated Avionics System is a system that

includes display, flight director guidance, autopilot, and yaw damper
and elevator trim functions. The integrated avionics system block
diagram is shown in figure 2--1, and consists of the following units and
subsystems:

STANDARD
D Integrated Avionics Computer (IAC) that includes:
- Flight Guidance System (FGS)
- Electronic Display System (EDS)
D Air Data System (ADS)
D Attitude and Heading Reference System (AHRS)
D PRIMUSâ II Integrated Radio System
D Traffic Alert and Collision Avoidance System (TCAS)
D Radio Altimeter System
D PRIMUSâ 660 Weather Radar System

OPTIONAL
D PRIMUSâ 880 Weather Radar System
D Flight Management System (FMS)

The PRIMUSâ 1000 Integrated Avionics System is a fail--passive

autopilot/flight director and display system that has a full complement
of horizontal and vertical flight guidance modes. These include all radio
guidance modes, optional long range navigation system tracking, and
air data vertical modes. Either the pilot’s or copilot’s electronic flight
instrument system (EFIS) can be coupled to the FGS for control of the
aircraft.

The IAC flight guidance function digitally processes heading,

navigation, and air data information to satisfy pilot requirements. The
data is presented to each pilot on EFIS displays.

The EDS displays attitude, heading, air data, navigation, and engine
instrument and crew alerting (EICAS) information.

A28--1146--112--00 System Description

2-1
PRIMUSr1000 Integrated Avionics System

EFIS annunciators indicate the selected autopilot/flight director mode,

altitude alert, radio altitude (RA) reference and go--around mode
engagement. Pitch and roll steering commands, calculated by the IAC
flight director function in conjunction with the guidance controller, are
displayed by command bars that direct the pilot to reach and/or
maintain the required flightpath or altitude.

The IAC is the focal point of information flow in this system. It converts
input data and information to the pilot--selected formats and displays
them on the attitude director indicator (ADI) and the horizontal situation
indicator (HSI) on the primary flight display (PFD). The IAC generates
information for display on the multifunction display (MFD). It also
computes the necessary flight director command bar steering
information for the autopilot function. The IAC also collects information
for the EICAS displays.

The two IACs are interconnected with high level data link control
(HDLC) lines. This and other interconnects allow the flight guidance and
symbol generators to share, compare, and communicate information.
The PRIMUSâ II Radio System is connected to the FGS by the radio
system bus (RSB).

When engaged and coupled to the flight director commands, the

PRIMUSâ 1000 IACs control the aircraft using the same commands
that are displayed on the ADI. When the autopilot is engaged and
uncoupled from the flight director commands, manual pitch commands
can be entered using the touch control steering (TCS) button or the
autopilot PITCH wheel.

The TCAS receives air data information from the ADS by way of the
diversity transponder. Heading data is supplied by the appropriate
AHRS. TCAS supplies the electronic display system with selected
aircraft traffic and conflict avoidance information.

The IAC has built--in multilevel test capability. The self--test includes an
automatic power--up self--check and initiated testing. It also includes
and on--ground maintenance test and fault storage. Refer to Section 9,
Troubleshooting, for pilot--initiated test information.

System Description A28--1146--112--00

2-2
 PRIMUSr1000 Integrated Avionics System

FLIGHT GUIDANCE SYSTEM (FGS)

The flight guidance system includes the following components:

D Inside the IAC

- Flight guidance computer (No.1 IAC only)
- System monitors
D Guidance Controller
D Autopilot Controller
D Servo Motor (pitch/roll/yaw)

The No. 1 IAC flight guidance computer is the fail--passive driver for the
flight director, autopilot, and pitch trim. Comparator monitor circuits on
the pilot and copilot AHRS and servo command outputs, are used to
operate the system fail passively. Servo command outputs from a
computed servo model in all three axes are compared to actual
command. If the difference between commands exceeds certain
threshold values, the FGS is disconnected from the servos. Normal
flight guidance functions are computed based on the No. 1 AHRS.

Flight director operation also comes from the No. 2 IAC.

The guidance controller contains the autopilot (AP) engage, yaw

damper (YD) engage, and low bank switch. Either side’s PFD and flight
director can be coupled to the autopilot.

Mode annunciations are displayed on the guidance controller and on

the PFD. The flight director command bars on the PFD, follow flight
director commands to display visual guidance for the selected mode.
The navigation sensor used for the chosen mode is selected on the
display controller (DC), and is annunciated on the PFD.

The yaw damper executes basic yaw damper functions with or without
the autopilot. The yaw damper is active when either the YD or AP
engage buttons on the autopilot controller is pushed.

The autopilot controller is used to manually control other autopilot,

using the PITCH wheel and TURN knob.

A28--1146--112--00 System Description

2-3
PRIMUSr1000 Integrated Avionics System

ELECTRONIC DISPLAY SYSTEM (EDS)

The EDS includes the following components:
D Display Symbol Generator (Inside each IAC)
D Display Unit (DU) -- PFD, MFD, and EICAS
D Bezel Controller
D MFD/EICAS Bezel Controller
D Display Controller
D Data Acquisition Unit (DAU)
D Reversionary Control Panel
D EICAS Reversionary Panel
D Remote Mounted Joystick (option).

In addition, the system outputs backup displays for the PRIMUSâ II

RMUs as follows:
D Navigation Display
D Select engine instruments and data.

The EDS displays pitch and roll attitude, heading, course orientation,
flightpath commands, weather presentations, checklists, mode and
source annunciations, air data parameters, engine data, and fault
warning information. The displays are organized as follows:

D Primary Flight Display -- The PFD integrates attitude, heading, air

data information, flight director modes and command bars, TCAS
resolution advisory (RA) commands, weather radar, and navigation
information.

D Multifunction Display -- The MFD displays heading, navigation

maps, weather radar information, and checklist.

D Engine Instrument and Crew Alerting System -- The EICAS

displays engine data, aircraft systems information, and warning/
caution/advisory messages.

The EDS displays essential display information from sensor systems,

automatic flight control, navigation performance, and caution--warning
systems in the pilot’s prime viewing area.

System Description A28--1146--112--00

2-4
 PRIMUSr1000 Integrated Avionics System

PFD formats are selected using the DC. MFD navigation and aircraft
performance displays are selected using the MFD mounted bezel
controllers. If an FMS is installed, the remote--mounted MFD joystick
menu is used for waypoint designation. The guidance controller contains
the course, heading, selected speed, and altitude preselect controls.

Each symbol generator (SG) can drive all five displays. When a PFD
or EICAS display fails, the on--side reversionary control panel is used
to replace the onside MFD with either a PFD or EICAS display.

The display dual processor symbol generator in the IAC functions as

the data processor for the display system. It receives digital and
discrete inputs, organizes the information into the correct formats as
defined by the DC, and transmits these formats to the display units.

When the display system is in its normal (no failure) configuration, IAC No.
1 drives the pilot’s PFD, MFD and the EICAS. IAC No. 2 drives the copilot’s
PFD and the MFD display. Wraparound signals are used for critical
parameters (such as pitch or roll data) to ensure information accuracy.

The dual channel DAUs transmit all the aircraft engine information to the
symbol generator for display on the EICAS display. They also transmit
analog data for display of some aircraft subsystem information. DAU 1
inputs left engine data and DAU 2 inputs right engine data.

AIR DATA SYSTEM (ADS)

The ADS includes dual micro air data computers (MADC). The aircraft’s
air data system also includes standby airspeed and altimeter indicators
that are driven directly from the aircraft pitot--static system.

The MADC is a microprocessor based digital computer that performs

digital computations, and supplies digital outputs. The MADC receives
pitot--static pressures and total air temperature inputs for computing the
standard air data functions. The MADC outputs, drives the altimeter
and Mach/airspeed displays, transponder, flight recorder, flight director,
and autopilot, as well as other elements of the flight control system. The
guidance controller is used to select the desired altitude reference for
the altitude alerting and altitude preselect functions. The computations
for each of these functions is performed by the IAC’s flight director
computer.

A28--1146--112--00 System Description

2-5
PRIMUSr1000 Integrated Avionics System

PRIMUSâ 660 WEATHER RADAR SYSTEM

The PRIMUSâ 660 Weather Radar System includes the following
elements:
D Weather Radar Receiver/Transmitter/Antenna
D Weather Radar Controller.

The PRIMUSâ 660 Weather Radar System is a lightweight, X--band

radar with alphanumerics designed designed for weather depiction and
GMAP. In the weather depiction mode, storm intensity levels are
displayed in four bright colors contrasted against a deep black
background. Areas of very heavy rainfall are displayed in magenta,
heavy rainfall in red, less severe rainfall in yellow, moderate rainfall in
green, and little or no rainfall in black (background). Range marks and
identifying numerics are used to facilitate the evaluation of storm cells.

NOTE: The PRIMUSâ 660 Weather Radar System transmitter

power is much stronger than previous Honeywell weather
radar systems. Refer to the PRIMUSâ 660 Weather Radar
System pilot’s manual for ground safety information.

In the GMAP, the system parameters are optimized to improve

resolution and to better identify small targets at short ranges. The
reflected signals from ground surfaces are displayed as magenta,
yellow, or cyan (most to least reflective).

A brief description is included in Appendix B, PRIMUSâ 660 Weather

Radar System, of this manual. A complete description is given in
Honeywell Pub. No. 28--1146--111.

ATTITUDE AND HEADING REFERENCE SYSTEM

(AHRS)
The AHRS includes the following components:
D Panel--Mounted Control Switches
D Attitude and Heading Reference Unit (AHRU)
D Thin Flux Valve.

The AHRU uses fiber optic gyros, accelerometers, and flux valve inputs
to derive attitude, heading, angular rates, and linear acceleration
information. This information is supplied to the PFD, MFD, flight
guidance system, weather radar, and other aircraft systems and
instruments. The AHRU differs from conventional vertical and
directional gyro systems, in that the optical fiber gyros output rate of
change data from the initialized reference attitude.

System Description A28--1146--112--00

2-6
 PRIMUSr1000 Integrated Avionics System

A digital computer in the AHRU mathematically integrates the rate data

to obtain magnetic heading and pitch and roll attitude. The thin flux
valve and accelerometers in the AHRU provide long--term references
for the system.

A memory module, attached to the AHRU assembly, retains flux valve

compensation and configuration information. The panel--mounted
heading/free switches are used to select the heading or DG submodes.

An operational description is given in Appendix A, Attitude and Heading

Reference System (AHRS), of this manual.

PRIMUSâ II INTEGRATED RADIO SYSTEM

The PRIMUSâ II Integrated Radio System (also known as the
PRIMUSâ II Radio System) consists of the following components:
D Integrated Navigation Unit
D Integrated Communication Unit
D Radio Management Unit (RMU)
D Audio Panel
D Tuning Backup Control Head.

The PRIMUSâ II Radio System is the source for VHF communication,

VOR/ILS/DME/marker beacon/ADF navigation data and air traffic
control (ATC) transponder functions. The units are connected through
the RSB that is a high speed, two--way data exchanger. The audio
signals are transmitted from the remote units to the audio panel through
a dedicated data bus.

The system is briefly described in Appendix C, PRIMUSâ II Radio

System, of this manual. For a more complete description of the
PRIMUSâ II Radio System, refer to Honeywell Pub. No. 28--1146--050.

RADIO ALTIMETER SYSTEM

The radio altimeter system includes the following components:
D One Radio Altimeter Receiver Transmitter
D Two Embraer--Supplied Antennas (Sensor Systems).

The radio altimeter system gives an absolute altitude display from 0 to

2500 feet, failure annunciation, and internal self--test. It also provides
inputs for gain programming for ILS approaches and other mode logic
in the FGS.

A28--1146--112--00 System Description

2-7
PRIMUSr1000 Integrated Avionics System

TRAFFIC ALERT AND COLLISION AVOIDANCE

SYSTEM (TCAS)
This TCAS II system is described in Honeywell Pub. No. 28--1146--070,
and includes the following elements:
D Computer
D Antenna.

TCAS is designed to act as a backup to the ATC and the ”see and avoid”
concepts. TCAS continuously surveys the airspace around the aircraft
seeking ATC transponder replies from other aircraft in the vicinity.
Flightpaths of these aircraft are predicted based on their tracks. When
TCAS is activated through the radio management unit and selected for
display on the MFD, nearby traffic is displayed. When traffic poses a
collision threat, the PFD displays a climb or descent vertical speed
command.

PRIMUSâ 880 WEATHER RADAR SYSTEM (OPTION)

The PRIMUSâ 880 Weather Radar System, which is fully described in
Honeywell Pub. No. A28--1146--102, includes the following components:
D Weather Radar Receiver Transmitter Antenna
D Weather Radar Controller.

The PRIMUSâ 880 Weather Radar System is an X--band radar

designed for weather detection and analysis. The PFD and MFD show
storm intensity levels in bright colors contrasted against a deep black
background. Areas of heaviest rainfall appear in magenta, significant
rainfall is in red, rainfall of medium intensity appears in yellow, and areas
of weakest rainfall appear in green. After proper evaluation, the pilot can
chart a course around these storm areas. The radar can also be used
for ground mapping. In the map mode, prominent landmarks are
displayed that help the pilot identify coastline, hilly or mountainous
regions, cities, or even large structures. In ground mapping mode,
video levels of increasing reflectivity are displayed as black, cyan,
yellow, and magenta. The weather radar system has a turbulence
(TURB) detection mode and an autotilt function.

System Description A28--1146--112--00

2-8
 PRIMUSr1000 Integrated Avionics System

FLIGHT MANAGEMENT SYSTEM (FMS) (OPTION)

The single FMS consists of the following:
D Flight Management Computer
D Control Display Unit (CDU)
D Data Loader (DL).
The FMS displays lateral and vertical navigation guidance parameters,
and is coupled laterally to the FGS. The CDU displays selected flight
plan data. The alphanumeric keyboard is the interface between the pilot
and the system.

The FMS uses five long--range sensors that are connected through
ARINC 429 buses. The FMS is also connected to the dual PRIMUSâ
II Radio System’s VOR/DME receivers.

The FMS interfaces to the MADC, EFIS, and FMS. FMS pilot interface
is done with the cockpit mounted CDU. VOR/DME and global
positioning system (GPS) are the sensor inputs. Under normal
conditions, FMS position is always based on the GPS inputs.

The FMS is the source of map data for the MFD. Map displays are
generated from an FMS navigation database. The navigation database
is normally updated every 28 days. The data loader is used to load pilot
defined flight plans, waypoint data, and aircraft performance data.

OTHER SWITCHES AND CONTROLS

D AP (Autopilot) Disconnect Switch -- The AP disconnect switch is
mounted on the control wheel. When pushed, it disconnects the
autopilot and yaw damper.

D TCS (Touch Control Steering) Button -- The TCS button is

mounted on the control wheel. When pushed, the pilot can change
aircraft attitude, altitude, airspeed, and/or vertical speed, without
permanently disengaging the autopilot.

D GA (Go--Around) Button -- The GA button is located on the throttle.

Under normal conditions, if the AP is engaged, it remains engaged
and the automatic flight control system (AFCS) commands coupled
go--around minimum speed hold function for the command bars. If
only the flight director is available, a fixed pitch angle of 12° is
commanded.
D External AHRS Switches -- Refer to Appendix A, Attitude and
Heading Reference System (AHRS), for a description of the
magnetic/ directional gyro (MAG/DG) and DG CW/CCW switches.

A28--1146--112--00 System Description

2-9
PRIMUSr1000 Integrated Avionics System

D Joystick (Option) -- This pedestal--mounted control is used with

MFD JOYSTICK submenu selections to control the pilot designator
(MAP view), to position the route of flight line (PLAN view), and for
alternative control of the auxiliary checklist.

D Master Warning/Master Caution Switches -- Located on the

glareshield, directly in front of each pilot, these switches flash when
a respective warning or caution message is displayed on the EICAS.
Pushing either side (as appropriate) terminates the flashing.

D N2 Override Switch -- When selected, the N2 switch overrides the

EICAS limits on a minimum N2 before certain crew advisory system
(CAS) messages are displayed. The switch is located on the
maintenance panel behind the pilot’s seat.

System Description A28--1146--112--00

2-10
 PRIMUSr1000 Integrated Avionics System

RM--855 RMU

W
E

E
24 24

12

12
21 21

15

15
S S

SCI

ICB IC--600
INTEGRATED
AVIONICS
COMPUTER

429

AH--800
AHRU

RIGHT
FADEC T

AD--39738--R6@

PRIMUSÒ 1000 Block Diagram

Figure 2--1

A28--1146--112--00 System Description

2-11/(2-12 blank)
 PRIMUSr1000 Integrated Avionics System

3. Electronic Display System

(EDS)

INTRODUCTION
The integrated avionics computer (IAC) electronic display system show
the flightcrew flightpath, navigation, engine, and system information.
Each pilot has a dedicated primary flight display (PFD) and a
multifunction display (MFD). The right center display unit (DU) contains
the engine instrument and crew alerting system (EICAS) display. See
figure 3--1 for a cockpit layout of the electronic flight instrument system
(EFIS) and EICAS.

Each pilot has a dedicated display controller (DC) and a reversionary

controller. In general, these dedicated controllers are used to control
PFD display functions.

Display dimming is controlled from the dimming control panel.

The PFD and MFD DUs have a bezel--mounted controller. The PFD
bezel controllers are identical, and the MFD controllers are also
identical.

A detailed description of the EFIS part of the display system is given in

Section 4, Electronic Flight Instrument System, of this manual. That
section covers PFD, MFD, and associated controller functions.

Section 5, Engine Instrument and Crew Alerting System, gives a

detailed description of the EICAS part of the display system.

A28--1146--112--00 Electronic Display System (EDS)

3-1/(3-2 blank)
 PRIMUSr1000 Integrated Avionics System

Cockpit Layout of EFIS

Figure 3--1

A28--1146--112--00 Electronic Display System (EDS)

3-3/(3-4 blank)
 PRIMUSr1000 Integrated Avionics System

MFD BEZEL CONTROLLER BASIC OPERATION

The MFD bezel controller shown in figure 3--2, gives the pilot access to
a variety of supplemental information through the menu item keys and
rotary knob. Menu item selections are sent to the respective display
controller. If one fails, the remaining display controller transmits the
bezel controller commands.

AD--51639@

MFD Bezel Controls

Figure 3--2

When the desired menu is displayed, the menu item keys are used to
do the following functions:

D Submenu Selection -- This selection, sequences through possible

choices with that specific menu item key.

D Toggle Selection -- This selection, sequences through possible

choices with that specific menu item key.

D Momentary Selection -- The menu item is selected as long as the

menu item key is held.

D Parameter Display Selection -- Pushing the menu item key boxes

the menu item, and the parameter is displayed on the MFD.
Sometimes pushing another menu item key boxes the previous
menu items and the previous parameter is displayed on the MFD.

D Settable Parameters -- When this type of menu item is available,

dashes are displayed around the menu item. The rotary knob on the
bezel controller is used to dial in the proper value. When the knob
is turned, the dashes are replaced with the desired value.

A28--1146--112--00 Electronic Display System (EDS)

3-5
PRIMUSr1000 Integrated Avionics System

Map Range Control

When a parameter is not selected or being set, the rotary knob is used
to select the map display range.

NOTES: 1. When weather is displayed on the MFD, range is

controlled from the weather radar controller.
2. The rotary knob on the EICAS bezel controller is only
used to scroll the amber and cyan CAS messages.

Controller Conventions
When a menu item is boxed, that parameter is being displayed on the
PFD or MFD. Selecting a boxed item deselects the item.

D Parameter Selected for Display -- A selected parameter is

displayed only when the item is boxed on the appropriate menu. If
there is no box around the item, it is not displayed.

D Parameter Selected for Setting -- If the parameter is not boxed, the

first selection of the menu key below it, boxes the parameter or
causes dashes to be boxed if the parameter is being set. The rotary
knob is used to change a set value. Once the value is set, it is
displayed by pushing any other toggle key. Pushing the same toggle
key, deselects the parameter and erases entered data.

Electronic Display System (EDS) A28--1146--112--00

3-6
 PRIMUSr1000 Integrated Avionics System

REVERSIONARY CONTROLLERS AND DIMMING

PANEL
System Reversionary Panel
If an attitude and heading reference system (AHRS) or micro air data
computer (MADC) fails, the system can be reconfigured to display the
remaining good sensor on both sides of the cockpit. The remaining
good sensor is selected using the AHRS or air data computer (ADC)
switches on the reversionary controller shown in figure 3--3. A failure of
a PFD unit is controlled by using the on--side display reversionary
panel. Both are described below.

NORM
PFD EICAS

MFD ADC AHRS SG

AD--50892--R1@

System Reversionary Panel

Figure 3--3

D MFD Knob -- This knob selects the on--side MFD display between
PFD, NORM (normal), or EICAS.

NOTES: 1. When the selection is PFD or EICAS, the normal

display unit for the PFD or EICAS is blanked.
2. When the SG reversionary switch is selected
cross--side, the MFD switch operates normally.
However, under this condition, the selection displays
the cross--side PFD or cross--side SG supplied
EICAS on the MFD display tube.

D ADC Button -- This button selects between NORM, or cross--side

MADC to supply the on--side PFD/MFD.

NOTES: 1. If reversionary ADC is selected, the current

airspeed modes are dropped, but they can be
re--engaged.
2. It is possible for both pilots to select ADC
reversionary. In that case, each pilot is receiving air
data information from the cross--side MADC.

A28--1146--112--00 Electronic Display System (EDS)

3-7
PRIMUSr1000 Integrated Avionics System

D SG (Symbol Generator) Button -- This button selects between

NORM or cross--side SG to supply the on--side displays.

NOTE: When the cross side SG is selected, the PFD and MFD on
both sides of the cockpit are identical since they are both
driven by the cross--side SG.

D AHRS Button -- This button selects between NORM or cross--side

AHRS to supply the on--side PFD attitude and heading.

NOTES: 1. When an AHRS fails and, if reversionary AHRS is

selected, the AP is not available.
2. It is possible for both pilots to select AHRS
reversionary. In that case, each pilot receives
AHRS information from the cross--side AHRS.
3. Selecting reversionary AHRS, disengages all flight
director modes and the AP. If both AHRSs are
working normally, the modes (and AP) can be
re--engaged.

Data Acquisition Unit (DAU) Reversionary Panel

The DAU reversionary panel, shown in figure 3--4, selects between
NORM, and reversionary for each DAU to supply data to the SGs. In
NORM, each DAU uses its A channel data for display. In reversion,
each DAU uses its respective B channel for display.

AD--50893@

DAU Reversionary Panel

Figure 3--4

EDS Dimming Panel

The EDS dimming controls are located on the glareshield dimming
panel. They are used to set the dimming reference for DU light sensor
operation.

Electronic Display System (EDS) A28--1146--112--00

3-8
 PRIMUSr1000 Integrated Avionics System

4. Electronic Flight Instrument

System (EFIS)

INTRODUCTION
The EFIS shown in figure 4--1, is a subsystem of the PRIMUSâ 1000
Integrated Avionics System. The EFIS consists of the following
components:
D Display Units (DU)
D Display Symbol Generators
- Part of the integrated avionics computer (IAC)
- Shared with engine instrument and crew alerting system
(EICAS) for reversionary operation
D PFD Bezel Controllers
D MFD Bezel Controllers
D Display Controllers (DC)
D Guidance Controller Course/Heading Controls
D Remote--Mounted Joystick.

The EFIS is an integrated system that displays flight attitudes,

airspeeds, vertical speed, altitude, headings, course orientation,
flightpath commands, weather and mapping presentations, traffic alert
and collision avoidance system (TCAS) data, and source
annunciations.

A28--1146--112--00 Electronic Flight Instrument System (EFIS)

4-1/(4-2 blank)
 PRIMUSr1000 Integrated Avionics System

FD1 HDG NAV AP SPD FLC ALT FD2

CRS 1 HDG ASEL CRS 2

APR CPL SPD VS

BNK YD
PUSH SYNC PUSH SYNC PUSH SYNC

PFD
260 LOC HDG IAS GS 5000 FMS 360 300 TAS FMS 360 300 TAS 260 LOC HDG IAS GS 5000
AP YD KDVT +15 SAT KDVT +15 SAT AP YD
280 280
12.5NM 245 GSPD 12.5NM 245 GSPD
20 20 12 MIN N 12 MIN N 20 20
GS GS

33

33
260 3 3 260
10 10 4500 10 10 4500
PLAB1 PLAB1
43 20 20
1 1

30

30
240
9 00
80
6
PLAB2
*PBD01
LL01
6
PLAB2
* PBD01
LL01
240
9
43
00
80
10 10 KDVT KDVT 10 10
220 220
25 50 50 25 50 50

W
E

E
350 350
200 200
.410 M 200 RA 29.92 IN STAB STAB .410 M 200 RA 29.92 IN
TGT TGT
REVERSIONARY PANEL REVERSIONARY PANEL
24 24

12

12
359CRS TX TX 359 CRS
NORM ILS1 3 --16 --16 ILS1 3 NORM
PFD EICAS PFD EICAS

33

33
3 N 2
21 21
3 N 2
ENGINE DOORS ENGINE DOORS

15

15
S S

30

30
6 13.1 NM 1 6 13.1 NM 1
REF TO TEMP:--99 C REF TO TEMP: --99C
25 REF A--ICE: OFF REF A--ICE: OFF 25

W
0 1000 0 1000

E
VOR1 REF FLX TEMP:--99 C REF FLX TEMP:--99 C VOR1
24
1 DOOR DOOR 24
1

12

12
MFD ADC SG AHRS ADF2 21 4 QT OIL LVL 1 QT OPEN 4 QT OIL LVL 1 QT OPEN ADF2 21 MFD ADC SG AHRS
S 2 S 2

15

15
HDG TTG 3 RESET M/P RESET M/P HDG TTG 3
001 TGT 5MIN RTN T/O ECS FUEL HYD ELEC RNG RTN T/O ECS FUEL HYD ELEC RNG 001 TGT 5MIN

AD--50894--R1@

EFIS Block Diagram

Figure 4--1

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4-3/(4-4 blank)
 PRIMUSr1000 Integrated Avionics System

CONTROLS

Primary Flight Display (PFD) Bezel Controller

The PFD bezel controller shown in figure 4--2, is used to control various
elements on the PFD.

AD--50895@

PFD Bezel Controller

Figure 4--2

D IN/HPA Button -- The IN/HPA button is used to set the barometric

altimeter correction in inches of mercury (inHg), or hectopascals
(hPa). Power up is with IN selected.

D Inclinometer -- The ball in the glass track is used to indicate a slip

or skid flight condition.

D STD (Standard) Button -- Pushing the STD button, returns the

barometric altimeter correction to standard value (29.92 inHg or
1013 hPa).

D BARO (Barometric) SET Knob -- The BARO SET knob is used to

enter the barometric altimeter correction in either inHg or hPa.

NOTES: 1. When the pilots are displaying cross--side micro air

data computer (MADC) data on their PFD, they do
not have control over the displayed BARO setting
from their respective DC.
2. The BARO set function is independent from the DC
and does not require that the DC be functional to
set data.

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PRIMUSr1000 Integrated Avionics System

Multifunction Display (MFD) Bezel Controller

The MFD bezel controller is shown in figure 4--3.

AD--51639@

MFD Bezel Controller

Figure 4--3

The MFD bezel controller accesses several display menus and

submenus. The selections from these menus are described later in this
section.,

NOTE: When the SG reversionary switch selects the cross--side SG,

only the DU bezel buttons on the same side as that symbol
generator, are used to make MFD/EICAS selections.

D M/P RNG (Map/Plan Range) Rotary Knob -- The rotary knob

normally controls the range of the map or plan display in preset
increments. The same menu select buttons change the knob
function to set specific flight parameters. When weather is selected
for display, the weather controller controls range.

D MENU INOP (Inoperative) -- This annunciation shown in figure 4--4,

is displayed when the DC is inoperative.

AD--51640@

Menu Inop Display

Figure 4--4

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 PRIMUSr1000 Integrated Avionics System

An MFD menu operation tree is shown in figure 4--5.

A B

AD--50909--R1@

MFD Menu Tree

Figure 4--5

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PRIMUSr1000 Integrated Avionics System

MFD MAIN MENU

The main menu shown in figure 4--6, is the MFD power--up menu. Push
the SYS MENU key on the main menu to select the MFD system menu
shown in figure 4--7.

AD--50897@

Main Menu
Figure 4--6

SYSTEM MENU

Pushing the SYS button from the main menu shown in figure 4--6,
selects the system menu shown in figure 4--7.

NOTE: Unless weather is selected on the MFD, or unless central

maintenance computer (CMC), TCAS, or checklist are
selected for display, a system display is always shown in the
lower part of the MFD.

AD--50898@

System Menu Selections

Figure 4--7

Electronic Flight Instrument System (EFIS) A28--1146--112--00

4-8
 PRIMUSr1000 Integrated Avionics System

D RTN (Return) Button -- This button returns the system menu to the
main menu. If a system page is being displayed when the RTN
button is pushed, the system page remains displayed.

D T/O (Takeoff) Button -- Pushing this button displays the takeoff

system page. When first pushed, the box around the previously
selected system name and its respective system page are removed,
the menu label T/O is boxed in white, and the takeoff system page
is displayed.

Subsequently pushing the T/O button, while the T/O page is

displayed, has no effect.

D ECS (Environmental Control System) A/I (Anti--Ice) Button --

Pushing this button displays the ECS/A/I page on the MFD. When
first pushed, the box around the previously selected system name
and its respective system page are removed, the menu label ECS
A/I is boxed in white and the environmental control/anti--ice system
page is displayed.

Subsequently pushing the ECS A/I button, while the ECS A/I page
is displayed, has no effect.

D FUEL Button -- This button selects the fuel system page for display
on the MFD. When first pushed, the box around the previously
selected system name and its respective system page are removed,
a RESET label is displayed above the FUEL label, the FUEL label
is boxed in white, and the fuel system page is displayed.

Subsequently pushing the FUEL button, arms the FUEL USED

quantity reset mode. The white box is removed from FUEL, and the
label RESET is boxed in white. Another push of the RESET button
resets the FUEL USED quantity to zero, the box around the RESET
label is removed, and the FUEL label is again boxed in white.

NOTE: If the pilot does not want to reset the fuel used quantity,
the boxed RESET mode can be cancelled by selecting
another system button, then reselecting FUEL.

D HYD (Hydraulic) Button -- This button selects the hydraulic system

page for display on the MFD. When first pushed, the box around the
previously selected system name, and its respective system page
are removed, the menu label HYD is boxed in white, and the
hydraulic system page is displayed.

Subsequently pushing the HYD button, while the HYD page is

displayed, has no effect.

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PRIMUSr1000 Integrated Avionics System

D ELECT (Electrical) Button -- This button selects the electrical

system page for display on the MFD. When first pushed, the box
around the previously selected system name, and its respective
system page are removed, the menu label ELECT is boxed in white,
and the electrical system page is displayed.

Subsequently pushing the ELECT button while the ELECT page is

displayed, has no effect.

MFD MENU

Pushing the MFD button from the main menu shown in figure 4--8,
selects the multifunction display menu shown in figure 4--9 or 4--10.

AD--50899@

Main Menu -- MFD Button

Figure 4--8

AD--51212@

MFD Main Menu (Without FMS)

Figure 4--9

Electronic Flight Instrument System (EFIS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

AD--51211@

MFD Main Menu (With FMS)

Figure 4--10

D RTN Button -- This button returns the menu to the main menu.

D SPDS (Speeds) Submenu -- Pushing the SPDS key on the MFD

menu shown in figure 4--10, selects the VSPEEDS submenu shown in
figure 4--11.

AD--50896@

VSPEEDS Submenu
Figure 4--11

- RTN BUTTON -- This button returns the speeds submenu to the

main menu. If a system page is being displayed at the time the
RTN button is pushed, the system page remains displayed.

- V1 BUTTON -- The V1 button is used to select and set the V1

takeoff reference speed. When the button is pushed for the first
time after power--up, the cyan dashes are replaced with the V1
default value of 89 knots. The selected V1 speed is displayed in
magenta.

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PRIMUSr1000 Integrated Avionics System

- VR BUTTON -- The VR button is used to select and set the

rotation reference speed. When the button is pushed for the first
time after power--up, the cyan dashes are replaced by the
greater of 89 knots or the value of V1, if V1 has been previously
set. The selected VR speed is displayed in cyan.
- V2 BUTTON -- The V2 button is used to select and set the V2
takeoff reference speed. When the button is pushed for the first
time after power--up, the cyan dashes are replaced by the
greater of 89 knots or the value of VR, if VR has been previously
set. The selected V2 speed is displayed in white.
- AP BUTTON -- The AP button is used to select and set the
approach reference speed. When the button is pushed for the
first time after power--up, the cyan dashes are replaced by the
greater of 89 knots or the value of V2, if V2 has been previously
set. The selected AP speed is displayed in green.
NOTE: Typically, for landing, VREF is entered as the value
for AP.
- SET BUTTON -- The pilot uses the SET knob to change the
VSPEEDS when the two white boxes are present.
NOTES: 1. For the VSPEEDS, on power--up, three cyan
dashes are displayed under the VSPEED labels.
2. All labels and boxes are white.
3. When the button is pushed for the first time after
power--up, the cyan dashes are replaced by a
default digital value. Two white boxes are
displayed, an inner box around the default
VSPEED value, and an outer box around the inner
box and the VSPEED label. The dual boxes denote
that the VSPEED is active and can be changed by
the pilot using the SET knob.
4. The next push of the button removes the inner
box, which deselects the SET knob input. The
displayed selected VSPEED remains active.
5. The next push of the button removes the outer
box. The VSPEED remains displayed on the bezel
menu but is removed from the airspeed display
on the PFD.
6. Repeatedly pushing the VSPEED button cycles the
sequence.
7. The takeoff VSPEED order of V1, VR, V2 is always
maintained in increasing order of magnitude.

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 PRIMUSr1000 Integrated Avionics System

D JSTK (Joystick) SUBMENU (option if FMS is installed) --

Pushing the joystick key on the MFD menu shown in figure 4--10,
selects the joystick submenu shown in figure 4--12.

AD--50903@

Joystick Submenu
Figure 4--12

The joystick submenu buttons are used in conjunction with the

remote joystick to control the designator on the MFD map or plan
views.

NOTES: 1. The joystick cannot be used on the copilot’s MFD

with the designator if the checklist is selected for
display.
2. The remote mounted joystick is only installed if an
FMS is installed.

- RTN (RETURN) BUTTON -- This button returns the joystick

submenu to the main menu. If a system page is being displayed
when the RTN button is pushed, the system page remains
displayed.

- SKP (SKIP) BUTTON -- Push the SKP button to skip the

designator’s home position to the next displayed waypoint.
When SKP is pushed, with the designator at the last displayed
waypoint, the designator returns to present position.

When displaying the MAP format, if the designator is co--located

with a connected waypoint, pushing the SKP button positions the
designator box over the next waypoint.

When displaying the PLAN format, if the designator is

co--located with a connected waypoint, pushing the SKP button
positions the the flight plan so the next waypoint is displayed over
the designator.

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PRIMUSr1000 Integrated Avionics System

- RCL (RECALL) BUTTON -- In the MAP mode, when the

designator is not at its referenced waypoint, pushing the RCL
button recalls the designator to that position. If the referenced
waypoint was the aircraft, pushing RCL recalls the designator to
the aircraft (i.e., present) position. In the PLAN mode, pushing
RCL returns the designator to the referenced waypoint.

When displaying the MAP format, if the designator is co--located

with a connected waypoint, pushing the RCL button positions the
designator box at the present position of the aircraft symbol and
removes the designator from the display.

When displaying the PLAN format, if the designator is

co--located with a connected waypoint, pushing the RCL button
positions the the flight plan so the present position of the aircraft
symbol is displayed over the designator.

- ENT (ENTER) BUTTON -- When the designator is offset from the

home position or a waypoint, pushing ENT transmits the
LAT/LON of the designator to the FMS scratchpad, as a
requested waypoint.

NOTE: How the FMS displays the transmitted waypoint

depends on the FMS type and installation.

- M/P RNG KNOB -- The map/plan range knob is used to change

the range scale, when in either the map or plan display mode.

D NAV APT Button -- This button selects the combination of

VOR/DME navigation names, and airport names for display.
Power--up default is with neither selected. Pushing the button
repeatedly, cycles the selection from neither selected, to navaids
only, to airports only, to both navaids and airports, and back to
neither selected. The selection is annunciated by a white box drawn
around the selected name. When no box is present, the selection is
off.

D DATA Button -- Pushing this button, displays the flight plan

waypoint identifiers. Power--up default is with DATA on. Repeatedly
pushing the button cycles data on and off. The DATA on
annunciation is shown with a white box drawn around the word
DATA.

D MAINT (Maintenance) Button -- This button is only active from the

pilot’s side when the aircraft is on the ground and when the CAS
display is not indicating “CMC FAIL.”When pushed, the currently
displayed system page is removed and maintenance data from the
central maintenance computer is displayed. If RTN is pushed, the
maintenance data is removed from display.

Electronic Flight Instrument System (EFIS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

D M/P RNG Knob -- The pilot uses the map/plan range knob to change
the range scale when in the map or plan display mode.

CHECKLIST MENU

Pushing the CKLST button on the main menu (shown in figure 4--8),
displays the checklist menu on the MFD, as shown in figure 4--13.

NOTES: 1. The checklist can be selected for display on either

MFD. However, if it is displayed on one, the menu item
CKLST is removed from the other MFD and, therefore,
cannot be selected on that MFD.
2. When the checklist display is selected for the first time
(after power--up) the checklist disclaimer page is
shown. Push PAG to go to the master index of
procedures.
3. Normal procedures, emergency procedures,
abnormal procedures and waypoint listing indices (as
available) are selectable with the SKP (skip) button
from the checklist menu first displayed. Once the
procedures index is selected, the checklist display lists
all the procedures in that index.
4. To switch from one procedures index to another (for
example, to go from NORMAL to EMERGENCY)
select the checklist menu first displayed, and make the
proper selection.

AD--50905@

Checklist Menu
Figure 4--13

A28--1146--112--00 Electronic Flight Instrument System (EFIS)

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PRIMUSr1000 Integrated Avionics System

D RTN Button -- Pushing this button, returns the bezel menu to the
main menu.

D SKP Button -- Push this button to skip the checklist cursor to the
next incomplete item on the current list. If the cursor is moved
beyond the last item, the cursor moves to the beginning of the list
and goes to the next incomplete item. If all items are complete, the
button has no effect.

D LN BK (Line Back) Button -- Push the LN BK button to move the

cursor to the preceding item.

D PAG (Page) Button -- Push the PAG button to advance to the next
page. The cursor moves to the first incomplete item on the new
page. If an incomplete item is not found on the new page, the cursor
moves to the first item on the page. Pushing the page button while
on the last page, moves the cursor to the first incomplete item or (if
none are incomplete) the first item, on the first checklist page.

D RCL Button -- The first push of the recall button moves the checklist
cursor to the first incomplete item on the current list. If the cursor is
already on the first incomplete item, the first push has no effect. If
all items on the current list are complete, the first push moves the
cursor to the top of the list.

NOTE: Pushing the recall button twice in succession, reselects the

master index so that the other item indexes can be selected.

D Joystick -- The remote--mounted joystick is used for additional

paging and cursor control. Each movement results in the action
described in table 4--1.

Cursor Movement Action

Up arrow Moves the active selection to the next

lower order item
Down arrow Moves the active selection to the next
higher order item
Left arrow Displays the previous page
Right arrow Displays the next page

Cursor Movement/Action
Table 4--1

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 PRIMUSr1000 Integrated Avionics System

D ENT Button -- Push the enter button to either select an item or to

change a checklist item from incomplete to complete. After
selection, the cursor moves to the next incomplete item in the
current checklist.

- ON AN INDEX PAGE -- Pushing the ENT button displays the

checklist corresponding to the active index line selection. The
checklist is displayed at the page containing the lowest order
incomplete item with the active selection at that item. If the
checklist has been completed, the system forces all items in the
checklist to be incomplete and displays the first page of the
checklist with the active selection at the first item.

- ON A CHECKLIST PAGE -- Pushing the ENT button forces the

active selection to be completed and advances the active
selection to the next incomplete item. If ENT is pushed with the
active selection at the last item in a checklist, the operation
depends on the completion 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 completed), the system

presents the index page containing the next higher order
checklist with the active selection at that checklist.

D M/P RNG Button -- The pilot uses the map/plan range knob to
change the range scale when in the map or plan display mode.

TCAS

Pushing the TCAS button on the main menu shown in figure 4--14,
selects the TCAS display in the display window. A second push returns
the window to the previously selected system display. TCAS is
annunciated by a white box drawn around the menu name.

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PRIMUSr1000 Integrated Avionics System

AD--50906@

TCAS Selection from Main Menu

Figure 4--14
WX (WEATHER)
Pushing the WX key shown in figure 4--15, adds weather information
to the MAP display. A second push of the button removes the weather
information. WX on is annunciated by a white box.
NOTES: 1. If the weather radar is off or not transmitting, pushing
the WX button has no effect.
2. If the PLAN mode was being displayed when WX is
pushed, the display changes to the MAP mode.
3. Weather radar selections are controlled from the
weather radar controller. MAP range is also controlled
by the weather radar controller when weather is
selected for display.

AD--50907@

WX Selection from Main Menu

Figure 4--15

Electronic Flight Instrument System (EFIS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

MAP/PLAN

Pushing the MAP PLAN button shown in figure 4--16, selects the MAP
or PLAN MFD formats. Power--up default is MAP. Repeated pushes of
the button cycles from MAP to PLAN and back. The selection is
annunciated by a white box drawn around the selected display.

NOTE: If WX is selected, selecting the PLAN format turns the

weather display off. When MAP is reselected, the weather
display is turned back on if it has not been deselected.

AD--50908@

MAP Plan Selection From Main Menu

Figure 4--16

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PRIMUSr1000 Integrated Avionics System

Display Controller (DC)

The DC is used to select various features on the PFD. These include
horizontal situation indicator (HSI) formats, navigation sources,
weather display, and bearing pointer selection. The location of buttons
and knobs is shown in figure 4--17.

OFF FMS
NAV 1 ADF

ADF NAV 2

FMS OFF
RA TEST

AD--50910@

Display Controller
Figure 4--17

D FULL/WX Button -- This button toggles between the HSI full 360_
compass, and a partial compass with weather.

NOTE: When the HSI weather display is selected, the display

range is controlled by the weather radar controller.

D GSPD/TTG (Groundspeed/Time--To--Go) Button -- Normally, the

HSI displays GSPD that is computed from the selected navigation
source. The GSPD/TTG button toggles between groundspeed and
TTG to the next waypoint or selected navaid.

D ET (Elapsed Time) Button -- When the ET button is pushed,

groundspeed or time to go is replaced with ET. Pushing this button
starts, stops, and resets the digital display of ET.

D NAV (Navigation) Button -- When the NAV button is pushed,

VOR/LOC (localizer) information is selected for display on the HSI.

NOTE: To ensure that the proper source is being used with the
selected flight director modes, when NAV sources are
changed (master/coupled side), the EFIS resets the
lateral FD modes (except for heading).

D FMS (Flight Management System) Button (Optional) -- When

the FMS button is pushed, FMS information is displayed on the HSI.

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 PRIMUSr1000 Integrated Avionics System

D BRG (Bearing) Knobs -- The HSI can display two independent

bearing pointers. The selectable bearing sources for each pointer
are described in table 4--2.

BRG f BRG Z

OFF OFF
NAV 1 NAV 2
ADF ADF
FMS FMS
NOTE: FMS is optional.

Navigation Bearing Pointers

Table 4--2

D RA (Radio Altitude) Knob/TEST Button -- These two concentric

controls are used to set RA minimum altitude setting (outer knob),
and to run the TEST function (inner button).

When the TEST button is pushed, the on--side PFD and MFD enter
the test mode. In the test mode, flags and cautions are displayed
along with a check of the radio altimeter. The EICAS also displays
a test pattern when the integrated avionics computer (IAC) that is
driving the EICAS page is in test. Normally, this is the left IAC.

NOTE: Once airborne, only the radio altimeter is tested but not if
the flight director has captured an ILS glideslope.

A detailed description of the display test results is contained in EFIS

Preflight Checklist, later in this section.

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PRIMUSr1000 Integrated Avionics System

Guidance Controller Course and Heading Controls

The course pointer and heading bug are controlled using the guidance
controller course (CRS) and heading (HDG) knobs, shown in figure
4--18. The heading knob controls the heading bug on both sides of the
cockpit (HSI/MFD). The onside course knob controls the onside course
pointer when the NAV (VOR/LOC) is the selected navigation source.

Additional information on the guidance controller is given in Section 6,

Flight Guidance System (FGS).

AD--50911@

Guidance Controller
Figure 4--18

When the center button in the HDG or CRS knob is pushed and
released, the EFIS performs a synchronization as described below.

D HDG Sync -- The heading bug centers under the compass lubber
line.

D CRS Sync -- The CRS sync only occurs when VOR or VOR/DME
is the selected source. When the center button is pushed, the course
pointer points to the VOR station with the CDI centered.

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EFIS Preflight Self--Test

To run the EFIS self--test, push and hold the TEST button on the DC and
verify the following:

NOTES: 1. The EFIS test is only functional when the aircraft is on

the ground. The test is functional at all times except
during GS CAP/TRK.
2. A localizer frequency must be tuned on both NAV
receivers to annunciate ILS.
3. Self--test is limited to radio altimeter test only in the air
and is inhibited after glideslope capture.
4. FD FAIL annunciation is not displayed during a
pilot--initiated test.
5. If the aircraft is on the ground and the DC TEST button
is pushed longer than 5 or 6 seconds, the system enters
the initiated test mode. Refer to Section 9,
Troubleshooting, for information about initiated testing.

EFIS SELF--TEST

D PFD Display Test Results

- The EFIS test is only functional when the aircraft is on the ground.
The radio altimeter test is functional at all times except during GS
CAP/TRK.
- Course select, heading select, RA set, distance and GSPD/ TTG
digital displays are replaced with amber dashes.
- ATT and HDG displays are flagged.
- All pointers/scales are flagged.
- All heading related bugs/pointers are removed.
- Flight director command cue is biased from view.
- Radio altimeter digital readout displays radio altimeter self--test
value.
- The comparator monitor annunciates ATT, HDG, and ILS (if ILS
sources are selected on both sides).
- The word TEST (in magenta) is annunciated in the lateral capture
location on the top left of the ADI.
- The WDSHEAR (windshear) annunciation is displayed.
- FD mode annunciators are removed.

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PRIMUSr1000 Integrated Avionics System

- MIN is annunciated.
D MFD Display Test Results
- MAP symbols removed or replaced with dashes.
- SAT/TAS/TAT replaced with dashes.
- TCAS (option), checklist, weather displays are removed.
- CHK EICAS message is displayed.
- System pages digital displays and analog scales are invalid.

D EICAS Display Test Results

- Engine displays invalid
- Flaps and pitch trim invalid
- Fuel invalid
- Various engine annunciators invalid.

SWITCHING TEST

Check each switch function on the DC. Switching should operate as

described in the EFIS preflight test.

INSTRUMENT TEST

Perform normal instrument test of navigational sources.

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 PRIMUSr1000 Integrated Avionics System

DISPLAYS
Primary Flight Display (PFD)
The PFD is an integrated display of attitude, navigation, and air data
flight information. The PFD is divided into four functional groups shown
in figure 4--19 and described below.

MODE
ANNUNCIATIONS

AIRSPEED ALTITUDE
DISPLAY ADI DISPLAY
DISPLAY

VERTICAL
HSI SPEED
DISPLAY DISPLAY

AD--50912@

PFD Functional Divisions

Figure 4--19

D Mode Annunciations -- The PFD displays mode annunciations

from the flight guidance system (FGS), HSI source selection, and
comparison monitor functions.
D ADI (Attitude Director Indicator) Display -- The ADI symbols use
a truncated sphere format to display standard attitude information.
The attitude display receives its inputs from the AHRS. When either
the pitch or roll data becomes invalid; all scale markings are
removed, the attitude sphere turns cyan, and a red annunciation of
ATT FAIL is put at the top center portion of the sphere.
D HSI (Horizontal Situation Indicator) Display -- The HSI heading
display receives its inputs from the AHRS. HSI displays include full
compass, arc, weather radar and navigation information.

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D Air Data Displays -- The PFD airspeed, altitude, and vertical speed
displays receive inputs from the micro air data computer (MADC).

D Slip--Skid Display -- Standard slip--skid indications are displayed

by the inclinometer on the PFD bezel control.

PFD ATTITUDE DIRECTOR INDICATOR (ADI)

See figure 4--20 for the location of the symbols and annunciations
described below.

D Flight Director Mode Annunciations -- The flight director mode

annunciations are displayed full--time on both PFDs. Lateral mode
annunciations are located on the left--hand corner of the mode
annunciation area above the ADI. Vertical mode annunciations are
located in the right--hand corner of the mode annunciation area
above the ADI. Armed modes are displayed in white characters, and
captured modes are displayed in green characters. In addition, a
white box surrounds a captured mode annunciation for 7 seconds
after the capture logic is satisfied.

NOTE: The annunciated flight director modes, on both PFDs, are

the result of the flight director mode selections on the
guidance controller and the resulting flight director
computer operation on the coupled side.

D Yaw Damper Annunciations -- These annunciations display the

engage status of the yaw damper. The annunciation is displayed
above and right of center of the attitude sphere. See table 4--3.

Display
Yaw Damper Status Annunciation Type Color

Engaged YD Steady Green

Autopilot test AP TEST Steady Amber
Normal YD disconnect YD 5 second Amber
flash
Abnormal YD disconnect YD 5 second Amber
flash, then
steady

Yaw Damper Status Annunciations

Table 4--3

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IAS HDG
RA ILS
LOC ATT
ALT

ADI and Annunciation Display

Figure 4--20

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 PRIMUSr1000 Integrated Avionics System

D Pitch Limit Indicator -- The pitch limit indicator symbol is displayed

on the PFD attitude indicator sphere as shown in table 4--9.

Margin to Stall Symbol Color

Greater than 11° Symbol removed

Between 11°and 6° Green
Between 3°and 5° Amber
2°or less Red

Pitch Limit Indicator

Table 4--4

D Roll Scale and Roll Pointer -- The roll scale is linear with white
markings at 10°, 20°, 30°, 45°, and 60° of roll. The 30° mark is
highlighted with a double stroke tick mark. A triangle marks 45°of
roll. The sky pointer is filled in and protected. The roll pointer is also
filled in, and it points to the value of roll.

D Pitch Scale -- The pitch scale tape (not shown) consists of the white
scale markings listed in table 4--5.

Up Down

10_ 10_
20_ 20_
30_ 30_
40_ 45_
60_ 60_
90_ 90_

Pitch Scale Tape Markings

Table 4--5

There are reference marks every 5_between 10 and 30_. Red ”fly
down”pitch warning chevrons appear at 45 and 65_pitch up, and
”fly up”warnings appear at 35, 50, and 60_pitch down.

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D Vertical Deviation Scale -- The white vertical deviation scale

appears on the right side of the attitude sphere. The navigation
source for this scale is selected by the DC from any one of the
following sources:
- ILS glideslope (green pointer)
- VNAV from the FMS (magenta pointer).
NOTE: When the cross--side navigation source is selected, or if
both sides are using the same source, the pointer is
shown in yellow.

A white GS is displayed above the vertical deviation scale when the

vertical deviation is supplied by an ILS glideslope, and a white FMS
annunciation is displayed when the vertical deviation is supplied by
an FMS.

When invalid information is present from the ILS glideslope receiver,

the pointer is removed and a red X is drawn through the scale. The
scale, label, and pointer are removed for invalid FMS data.

D Command Bar Display Logic (Normal Mode) -- Normally,

command bars are displayed on each PFD. The on--side PFD
displays the information for the respective IAC to compute steering
for the command bars. Command bar display can be either
single--cue or cross--pointer.

NOTES: 1. The command bar display (single--cue or

cross--pointer) is selected at the time of installation.
2. If the FD1/FD2 switches are selected, command
bars are normally displayed on both PFDs. They
show guidance that satisfies the flight director
modes selected by the guidance controller.
However, there is one exception. If the pilots do not
have the same navigation source (selected with the
DC) the non--coupled command bars are removed
from the ADI.
3. If both pilots have NAV as the selected navigation
source, but have selected different situations (i.e.,
different courses or different frequencies on the NAV
receiver), both command bars are in view. However,
the flight director mode annunciation of course, arm,
and capture, is only a function of the coupled side.
Futher, the non--coupled side command bars could
capture the selected course and not be in agreement
with the displayed flight director modes that are
supplied by the coupled side.

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D Low Altitude Awareness Display -- Low altitude awareness is

displayed on the barometric altitude tape. At 550 feet AGL, a rising
ground brown raster, topped with a yellow horizontal line, that
represents 0 feet AGL, starts rising from the bottom of the altitude
tape. The rising brown raster displaces the normal gray raster field
and altimeter scale data. At a radio altitude of zero feet, the entire
lower half of the altitude tape is shaded brown.

If either radio altitude or barometric altitude is invalid, the low altitude

awareness raster and the yellow horizontal line are removed.

D Aircraft Symbol -- The aircraft symbol is a stationary

representation of the aircraft. Aircraft pitch and roll altitudes are
displayed as the relationship between the aircraft symbol and the
ADI sphere. The aircraft is flown to align the command cue to the
symbol to satisfy FD commands.

D Low Bank Arc -- When BNK is selected on the guidance controller,

a ±14°green roll arc is displayed on the ADI roll scale.

D Flight Director Couple Arrow -- A green arrow is displayed above

the ADI to indicate which PFD is coupled to the flight guidance
system.

D PFD Source Annunciations -- The PFD annunciates the following

sources of display:
- AHRS attitude (ATT)
- Micro air data computer (MADC)
- AHRS and heading (MAG/DG)
When the normal (on--side) source is selected, the source
annunciations are suppressed. The on--side is defined in table 4--6.
If the normal source is not selected, these annunciations are
normally white, but they turn amber on both displays when the pilot
and copilot are on the same source.

Sensor Pilot PFD Copilot PFD

Attitude ATT1 ATT2

MADC ADC1 ADC2
Heading MAG1 MAG2

On--Side Sensor Sources

Table 4--6

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NOTES: 1. When the AHRS is in DG mode, the on--side

heading source annunciations are DG1, DG2,
respectively.
2. When the MAG heading is invalid, the source
annunciation is an amber HDG1 or HDG2.
D Autopilot Status Annunciations -- Displays the engage status of
the autopilot. The annunciation is displayed above and left of center
of the attitude sphere. See table 4--7.

Autopilot Status Annunciation Display Type Color

Engaged AP Steady Green

Autopilot Test AP TEST Steady Amber
Normal YD YD 5 second flash Amber
Disconnect
Abnormal YD YD 5 second steady Amber
Disconnect
Normal AP AP 5 second flash Amber
disconnect
Abnormal AP AP/YD 5 second flash, Red
disconnect then steady
Normal AP AP 5 second flash Red
disconnect in
CAT 2
Touch Control TCS Steady, while White
Steering TCS button is
held
Turn Knob Out of TKNB Steady Amber
Detent

Autopilot Status Annunciations

Table 4--7

D Vertical Track Alert (VTA) Annunciation -- An amber VTA is

displayed above the vertical deviation scale when the FMS VNAV
is selected. The input is sent from the selected FMS.
D Altitude Trend Vector -- The altitude trend vector is displayed when
the aircraft is in either a climb or a descent. The top of the vector
points to the altitude that the aircraft is at, if the current vertical speed
is maintained.

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D Radio Altitude (RA) Decision Height Annunciator -- When actual

radio altitude decreases to within 100 feet of the set value, a box is
placed in the lower right corner of the attitude sphere. When the
actual radio altitude is equal to or less than the set value, and amber
MIN is placed in the box and it flashes for 10 seconds.

D Marker Beacon -- The marker beacons are displayed in the lower

right corner below the attitude sphere. When active, each beacon
annunciation (O, M, or I) is displayed in a white box that flashes at
the proper rate. The beacon symbol colors are listed in table 4--8.

Marker Beacon Symbol Symbol Color

Outer O Blue
Middle M Yellow
Inner I White

Marker Beacon Annunciations

Table 4--8

D Radio Altitude Decision Height Set Data -- The RA set data is

located in the lower right corner of the attitude director indicator
(ADI).

The range for radio altitude set is 0 to 999 feet. The minimum altitude
value can be set within 10 feet above 200 feet above ground level
(AGL), and within 5 feet below that value. Above 999 feet AGL, the
set data is removed.

D Radio Altitude Display -- The radio altitude is displayed as a green

digital display, ranging from --20 feet to +2550 feet, surrounded by
a white box, and located at the bottom of the ADI sphere. For ranges
greater than 2550 feet, the boxed RA digital display is removed. It
is also displayed for low altitude awareness on the barometric
altitude tape as a rising ground brown raster. The input for the radio
altitude display is from the radio altimeter.

For an invalid radio altimeter, the digits are replaced with an amber
boxed --RA--.

D Comparison Monitor Annunciation -- A variety of comparison

monitor annunciations are displayed in various locations on the
PFD. They are described later in this section.

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D Attitude Source Annunciation -- The attitude source, either AHRS

1 or AHRS 2, is selected using the switches on the reversionary
control panel. These switches select the source for the attitude
display on the PFD. This display is always amber and it indicates
that the displayed attitude source is in reversion.

D MAX SPD/MIN SPD or Underspeed Warning -- When the flight

director detects an overspeed condition, a MAX SPD or MIN SPD
warning is displayed in amber to the left of the ADI. The warning
remains annunciated as long as the overspeed conditions exists.

D Airspeed Trend Vector -- The airspeed trend vector indicates the

airspeed in 10 seconds, if the current acceleration is maintained.

D Symbol Generator (SG) Source -- When an SG reversion mode

(SG1 or SG2) is selected on the system reversionary controller, the
source is annunciated in the upper left corner on both PFDs. An
amber SG1 indicates the No. 1 symbol generator (pilot’s side) is
driving all five displays. An amber SG2 indicates No. 2 SG (copilot’s
side) is driving the displays.

NOTE: As a result of selecting SG reversion mode, all the sensor

source annunciations (ADC, ATT, and so forth) are
annunciated, and the annunciation in amber.

D Windshear (WDSHEAR) Annunciation -- When the WDSHEAR

annunciation is displayed, it flashed for 10 seconds and goes on
steadily. An amber annunciation indicates caution, red indicates
warning.

D CAT2 (Category 2) ILS Annunciations -- The symbol generator

displays a green CAT2 category annunciator on the PFD. These
mode annunciations are located above the vertical deviation scale
as shown in figure 4--21.

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CAT 2 ANNUNCIATOR
260 LOC CAT2 50 00
AP YD
280
20 20
GS
260
10 10 4500
1 20 CAT 2
240 43 00 DEVIATION
SCALE AND
9 80 POINTER
10 10
220

350
200
.410 M 200 RA 29.92 IN

359 CRS
ILS1 3
N
2
13.1 NM 1
25 0 1000
W

VOR1
1
ADF2
2
HDG TTG 3
001 TGT 5MIN
AD--50915--R1@

PFD With CAT II Annunciations

Figure 4--21

The green CAT2 mode annunciation is an indication that the

conditions for a CAT2 approach are present and the excessive ILS
deviation monitors are active on the PFD.

The symbol generator activates the CAT2 mode annunciator on the

PFDs whenever APR (approach) mode is selected and the following
criteria are satisfied.

- The DC RA set data must be set to between 200 and 100 feet and
both radio altimeters must be valid, and indicating greater than
1000 feet.

- Independent and valid navigation sources are present on pilot’s

and copilot’s PFDs.

- The on--side radios must be selected for display, both must be

tuned to the ILS, and the localizer and glideslope deviations must
be valid.

- Attitude and heading must be valid on both PFDs.

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- IAS and baro altitude must be valid on both PFDs.

- There are no SG, AHRS, or ADC reversions.

- There are no comparison monitor annunciations

- Two symbol generators must be operational.

If these criterias are not met when the APR mode is selected, the
system defaults out of CAT2 because it is programmed to assume
that the pilot does not care to perform the ILS approach with the
CAT2 monitors active. There is no approach category annunciation
given on the PFD in this case. On the other hand, if CAT2 conditions
were met, but then lost, CAT1 flashes for 5 seconds and goes on
steadily.

D CAT2 Deviation Scale and Pointer -- When the CAT2 mode

annunciation is displayed, as shown in figure 4--21, the ILS
excessive deviation monitors are active. If the localizer deviation
exceeds the CAT2 window requirements with radio altitude less than
500 feet, the deviation pointer and scale are changed from green to
amber. The scale also flashes. The display reverts back to green if
the deviation is brought back within the threshold. The same logic
and symbols apply to the glideslope deviation scale. In addition, the
CAT2 annunciation turns amber and flashes. The monitors are
independent with the thresholds set in accordance with regulatory
guidance for Category 2 ILS operations.

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FULL COMPASS DISPLAY

The HSI annunciations on the PFD are shown in figure 4--22. Refer to
the figure to see the location of the annunciations described.

260 LOC HDG IAS GS 50 00

AP YD
280
20 20
GS
260
10 10 4500

10 10
220
350
200
.410 M 200 RA 29.92 IN
359 CRS
ILS1 3
N
2
13.1 NM 1
25 0 1000
VOR1
1
ADF2
S 2
HDG TTG 3
001 TGT 5MIN

AD--50916--R2@

HSI Compass Display On PFD

Figure 4--22

D Navigation (NAV) Source Annunciations -- The selected NAV

source for display on the course deviation indicator (CDI) is
transmitted from the on--side DC. If the on--side controller is invalid,
the SG reverts to on--side primary NAV (i.e., pilot -- VOR1, copilot
-- VOR2). The annunciation of NAV source is displayed in the upper
left corner of the HSI area.

D Heading Select Bug With Digital Display -- The cyan heading bug
rotates around the compass arc. The heading bug is positioned by
rotating the heading select knob on the GC. The white label HDG is
displayed in the lower left--hand corner of the HSI area. Directly
below the label is a cyan digital display of the heading bug’s current
selected value.

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D Drift Angle Bug -- If available from the FMS, 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 the actual track of the aircraft. The bug is displayed as
a magenta triangle that moves around the outside of the compass
card (in either FULL or ARC modes).

D Bearing Pointers With Annunciation -- The pilot can display

bearing pointers on the compass display. The pointers rotate around
the compass display center in the same way as the course pointer.
Bearing pointer source selections come from the on--side display
controller. If the on--side DC fails, the default sources are VOR1 on
circle (pilot’s) and VOR2 on diamond (copilot’s).

The annunciation for the bearing pointers is displayed in the lower

left--hand corner of the HSI area. The annunciation color matches
the bearing pointer color. The annunciation and bearing pointer are
cyan for the circle pointer and dim white for the diamond pointer. The
bearing pointer selections are listed in table 4--9.

BRG f BRG Z

OFF OFF
NAV 1 NAV 2
ADF 1 ADF 2
FMS FMS
NOTE: The ADF bearing pointer annunciator depends on the number of
installed systems.

Bearing Pointer Selections

Table 4--9

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D Groundspeed, Time--To--Go, Elapsed Time -- Depending on the

relation with the DC, this display can be GSPD, TTG to the active
waypoint, or ET. The range of values for each parameter are shown
in table 4--10.

Parameter Range

GSPD 0--999 kts (NAV)

0--4095 kts (FMS)
TTG 0--399 min. (NAV)
0--512 min. (FMS)
ET 0--9 hr, 59 min.

GSPD, TTG, ET Ranges

Table 4--10

D Course Pointer With Digital Display -- The course pointer rotates

around the center of the arc heading display.

With a short range NAV selected, the course pointer is positioned

by rotating the course knob located on the GC.

When the flight management system (FMS) is selected, the course

select data is generated by the FMS desired track input.

The course pointer data is located in the upper left--hand corner of

the HSI region. If short range NAV has been selected, the label is
CRS. If long range NAV has been selected, the label is desired track
(DTK). Directly beside the label is a digital display of the current
course pointer value.

NOTE: When HSI ARC display is selected, the course pointer can
be rotated to nearly off--scale (not visible). The
course/track digital readout is still available.

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D Navigation Symbol Colors -- The color of the course pointer,

distance display, groundspeed, lateral deviation, and navigation
source annunciation is listed in table 4--11.

On--Side Navigation
Source Selected Color

Short range navigation Green

FMS Magenta
Same navigation source on both sides Yellow
or secondary NAV source

Navigation Symbol Colors

Table 4--11

D NAV Course Deviation Indicator (CDI) With TO/FROM Indication

-- The center of the HSI contains an aircraft symbol. The lateral
deviation scale is displayed in the form of two dots on either side of
the aircraft symbol. This represents NAV deviation from the selected
source. The lateral deviation dots rotate around the center of the
course pointer.

NOTE: If the optional FMS is installed and selected, and the FMS
is operating in the approach mode, full scale deflection is
0.3 nautical miles (NM).

The TO/FROM indicator is displayed as a white arrow in front of the

aircraft (TO) or behind the aircraft (FROM). A TO indication is
displayed as long as the selected course pointer is within ±90_of the
selected NAV source bearing.

D Wind Vector (Available if FMS is Installed) -- Wind vector

information is displayed to the left of the HSI. The wind is shown in
magenta with velocity and direction. Wind information is displayed
as a vector arrow showing the direction of the wind relative to the
airplane symbol. The digital quantity indicates wind velocity.

D NAV Source Distance -- The NAV source distance is displayed on

the left side of the HSI. The distance represents distance to the
station for a short range NAV and the distance to the TO waypoint
for the FMS. If DME hold is selected when the VOR is displayed, an
amber H is displayed next to the DME distance.

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ARC DISPLAY

Most of the arc display annunciators, shown in figure 4--23 , are the
same as that for full compass HSI display. The differences are
described below.

260 LNAV VS VNAV 14500

I 280
AP YD ASEL A
A ADC2 L
S SG2 20 20 VTA T
FMS
260
M ATT 10 10 14500
1 A
20
240 X
143 00
9 2
S 80
P
R 10 10 MIN
220 1 D
FMS HEADING 14000
(FHDG)
ATT2 350 M
FMS STATUS 200 RA
LOC 200 RA
ANNUNCIATORS .410 M CAS 29.92 IN
AMBER CYAN
MAG2 HDG HEADING SOURCE
359 DTK FHDGDR 360 MSG TCAS 3000 ANNUNCIATION
INTEG TERM
WPT APP
FMS APP TEST 3
DR SXTK KDVT N 2 FMS MSG
DGR ANNUNCIATION
ALERT 25.9 H 1
”TO”WAYPOINT 25 0 1000 WEATHER RETURN
DME HOLD VOR1
WX 50 1
ADF2 --3.5
2
OFFSCALE HDG GSPD
ARROW
270 TGT 212KTS 3

WEATHER RADAR WEATHER RADAR RANGE RING

TGT/VAR MODES AND TILT VALUE
ANNUNCIATIONS ANGLE ANNUNCIATION AD--50917@

HSI Arc Display With FMS Map and Weather Radar

Figure 4--23

D Heading Source Annunciations -- When the cross--side heading

source is selected, or when the AHRS is in DG, the heading source
is shown above the compass card.

D FMS Status Annunciation -- Some critical FMS status displays are

annunciated to the left of the compass. These are: waypoint (WPT),
integrity (INTEG), offset (SXTK), terminal (TERM), approach
(APP), degrade (DGR) and dead reckoning (DR). TERM, SXTK and
APP are displayed in cyan; DR, DGR, and WPT are displayed in
amber. Message (MSG) is displayed to the right of the compass in
amber.

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NOTES: 1. The APP and TERM annunciators are related to

FMS supplied approaches as described in table
4--12.
2. The INTEG annunciation indicates that the GPS
sensor does not meet the required integrity
calculations for the current phase of flight.
3. The MSG annunciator flashes until the FMS
condition is cleared.

FMS Annunciation Action

Full accuracy approach APP Flashes for 10

(typically with stored seconds.
approaches)
Normal RNAV accuracy APP Steady
approach annunciator
Enter the terminal area TERM Steady
annunciator

Approach and Terminal Annunciators

Table 4--12

D Weather Returns -- Weather radar returns can be displayed to

show the location of weather cells in relation to the range and
direction from the aircraft.

D Weather Radar Range Ring Value -- The half range ring and value
are controlled by the weather radar controller. Normal range can be
selected between 5--300 NM. FPLN (flight plan) range can be
selected up to 1000 NM.

D Weather -- When weather is selected on the DC, weather

information is displayed on the arc format. Each PFD can
independently display selected weather information. Depending on
the installed weather radar system, the mode annunciations
described in table 4--13 are displayed to the left of the compass arc.

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Mode Annunciation
R/TMode Annunciation Color

R/T in warm--up WAIT Green

REACT mode RCT Green
Ground Clutter Reduction GCR P880 Green (Option)
RCT and GCR Modes GR/R P880 Green (Option)
Active
Forced Standby FSBY Green
Standby STBY Green
Test Mode TEST Green

Weather Mode WX Green

(Note 1)
Variable VAR Amber
Weather and Turbulence WX/T P880 Green (Option)
RCT and Turbulence R/T P880 Green (Option)
Groundmap Mode GMAP Green
Flight Plan Mode FPLN Green

R/T Fail FAIL Amber

(Note 2)
R/T Off OFF Green
Transmitting, but not TX (Note 3) Amber
selected for display
NOTES: 1. When weather radar is invalid, WX is displayed in amber.
2. When on the ground and the weather test display is selected, failures
of the weather radar are indicated by fault codes in the tilt angle field.
3. Early versions of the P1000 software annunciates an amber TX while
the radar is in the warmup mode. This warmup is annunciated by a
green WAIT, in later P1000 software updates.

Weather Radar Annunciations

Table 4--13

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D Weather Radar TGT/VAR Annunciations -- When the weather

radar supplies a target, a cyan TGT is displayed below the HSI.
When the target is triggered, TGT turns amber. When variable radar
gain is selected, TGT is replaced with VAR.

D Off Scale Arrows -- In the arc mode, the heading bug can be rotated
off the compass scale. When the HDG bug is off--scale, a cyan arrow
is displayed on the outer compass ring to indicate the shortest
direction to its location.

D DME Hold -- The distance display is shown to the left of the compass
display. The values for distance, shown in table 4--14, depend on the
selected navigation source.

Distance Range
Selected Source (NM)

DME 0--409.5
FMS 0--4095

Selected Source Distance Range

Table 4--14

NOTE: When DME distance is selected to HOLD, an amber H is

placed after the distance display.

D FMS to Waypoint -- The active FMS waypoint identifier is shown in

magenta to the left of the HSI compass.

D FMS Heading (FHDG) Annunciation (If Available) -- When

heading guidance is supplied from the FMS, a magenta FHDG
annunciation is displayed on the upper left of the HSI.

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AIRSPEED DISPLAY
The airspeed section of the PFD is shown in figure 4--24. A description
of airspeed annunciations is given below.
AIRSPEED BUG
DIGITAL DISPLAY 260
280 Vmo/Mmo
INDICATOR BAR

AIRSPEED SET
260 BUG

AIRSPEED TREND
ROLLING AIRSPEED 1 VECTOR
DIGITAL DISPLAY 240
AIRSPEED REFERENECE
9 LINE

220 IAS ANALOG SCALE

200 2 VSPEED
MACH DISPLAY .410 M SET BUGS

1 V1
R VR
2 V2
160 AP VA PP
80
AIRSPEED I
COMPARISON A
MONITOR
ANNUNCIATOR
S 60

1
240
9
135 2
125 R TAKEOFF VSPEED
SET DISPLAY
120 1

AD--50918--R1@

Airspeed Display
Figure 4--24
D VMO Overspeed Bar -- The Vmo is a fixed red bar originating at VMO
and extending to the scale’s end. This bar is located along the right
side of the airspeed scale.
D Airspeed/Mach Reference Bug and Display -- Using the guidance
controller SPD set knob, and engaging the flight director SPD mode,
the pilot--adjustable airspeed or Mach reference is displayed above the
airspeed tape. The corresponding airspeed bug is shown on the right
side of the airspeed tape. The digital display and bug are both cyan.

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When FLC is selected on the guidance controller, the FGC selects the
preprogrammed values for a climb or descent. The preprogrammed
value for climb is an airspeed, and for descent, a rate of descent. When
in a climb, the cyan airspeed bug is placed at the pre--programmed
airspeed value. When in a FLC descent,the cyan airspeed bug is
removed, and the vertical speed bug is placed at its preprogrammed
value.
NOTE: When the flight director is not in the SPD or FLC mode, the
flight director airspeed bug, airspeed target, and target
box on that side remain at the last selected value.
D Airspeed Trend Vector -- The airspeed trend vector is positioned
along the outer right side of the airspeed tape. It is referenced to the
airspeed reference line. The magenta vector indicates what the
projected value of indicated airspeed should be in approximately 10
seconds, if the present trend is maintained.
D Indicated Airspeed (IAS) Analog Scale -- The airspeed tape is a
moving scale display with fixed pointer and calibrated airspeed
marks. The white scale markings on the tape are in 10--knot
increments. The scale digits move such that larger numbers come
from the top of the display. The scale and its markings are white.
D VSPEED Set Bugs -- In addition to the airspeed bug, VSPEED bugs,
corresponding to speeds for various phases of flight, can be displayed
on the PFD’s airspeed tape. This data comes from pilot inputs using
the MFD bezel keys. The VSPEEDS travel along the right side of the
airspeed tape. The VSPEEDS bug colors are described in table 4--15.

VSPEED Label Speed Definition Color

V1 1 Takeoff decision Magenta

VR R Takeoff rotation Cyan
V2 2 Takeoff safety White
VAPP AP Landing Speed Green
NOTES: 1. VAPP can also be used to set VREF for landing speed reference.
2. V1, VR, and V2, can be set equal to each other.
3. VSPEED bugs are removed when IAS exceeds V2+42 knots.
4. If either PFD is in the reversionary mode (i.e., displayed on the
MFD) VSPEEDS are set by the cross--side MFD bezel menu.

VSPEED Bug Identification

Table 4--15

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When the FGS enters the MAX SPEED mode, the annunciator MAX
SPEED is displayed in amber to the left of the ADI.

D Takeoff VSPEED Display -- If the aircraft is on the ground V1, VR, and
V2 are displayed in digital tabular form inside the lower portion of the
airspeed tape. As the airspeed tape increases and the values come
into view on the airspeed tape, the tabular display is removed and
subsequently replaced by the appropriately positioned VSPEED bug.

D Indicated Airspeed Comparison Monitor -- If the on--side and

cross--side calibrated airspeed differ by 5 knots or more for greater
than 2 seconds, an amber IAS is displayed vertically near the top of
the airspeed tape. When initially activiated, the display flashes for
ten seconds, then goes on steadily.

D Mach Display -- The digital Mach speed display is of the present

Mach number and is shown below the airspeed tape. The color of
the digits always agrees with the digital airspeed display. Mach
speed is displayed when the airspeed increases beyond 0.45M and
is removed when airspeed drops below 0.40M.

D Airspeed Rolling Digit Display -- A rolling digit display of the actual

current IAS value is contained within the display window. This data
is a magnification of the digits on the scale and they are readable to
a one--knot resolution. The digits within the pointer are green. When
the current airspeed value is equal to or exceeds the maximum
allowable airspeed (VMO) the digits turn red. When the airspeed trend
vector exceeds Vmo by one knot, the rolling digits turn amber unless
red is required.

D Low Speed Awareness (LSA) -- LSA (not shown) is indicated by

the three segment bar on the lower right side of the airspeed tape.
The bar indicates angle of attack (AOA) relative to Vs. The reference
for the bar is the airspeed reference line. The top of each segment
is defined in table 4--16.

Segment VS Reference (VSO)

White 1.23
Amber 1.13
Red 1.0

Low Speed Awareness Color Logic

Table 4--16

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ALTIMETER DISPLAY

The altimeter display is shown in figure 4--25. Refer to the figure for the
location of annunciations described in the paragraphs that follow.

ALTITUDE ALERT
14500 PRESELECT DISPLAY

ALTITUDE SELECT
BUG

14500
ALTITUDE TREND
VECTOR 20 CURRENT ALTITUDE
14300 DIGITAL DISPLAY
80
ALTITUDE REFERENCE
LINE

14000
ALTITUDE ANALOG
SCALE
BAROMETRIC
29.92 IN ALTIMETER DISPLAY

IN INCHES Hg
HPA HECTOR PASCALS

5000

4500

20
4300
80
LOW ALTITUDE
AWARENESS
SYMBOL

29.92 IN
AD--50919--R2@

Altimeter Display
Figure 4--25

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D Altitude Alert Preselect Display -- This data is displayed at the top

of the altitude scale. The data is set by the guidance controller
altitude select knob. The set data is cyan. When the aircraft is within
the altitude alert operating region, the box around the set data turns
from white to amber. The set data also turns amber. When a
departure from the selected altitude occurs, the select display box
also changes from white to amber.

The altitude alert operating region is defined as that time when the
aircraft enters the region where it is within 1000 feet of the
preselected altitude during a capture maneuver. At this time the box
around the set data turns to amber. Once the aircraft is within 250
feet of the preselected altitude, the box turns back to cyan. After
capture, the aircraft re--enters the altitude alert operating region if it
departs more than 250 feet from the selected altitude. A momentary
audio alert is sounded when the aircraft is 1000 feet from the
preselected altitude or has departed 250 feet from the select altitude
after capture.

D Altitude Digital Display -- A digital display of the actual altitude

value is contained in the display window. This data is a magnification
of the digits on the scale and is readable to within a 20--foot
resolution. The digits within the pointer are green. Below 10,000
feet, boxed hash marks are used to show that the
ten--thousand--foot digit is missing.

D Barometric Altimeter Display -- The baro set display is located

directly below the altitude tape. The pilot can set the altimeter in
either inches of mercury (inHg) or hecto Pascals (hPa) as selected
with the PFD bezel controller. If the on--side DC is invalid, the SG
defaults to the last selection (In or hPa). The baro set data is always
cyan.

D Low Altitude Awareness Symbol -- At radio altitudes of 550 feet

or less, the lower part of the altitude tape linearly changes from a
gray raster to brown and the altimeter scale markings are removed.
At zero radio altitude, the brown raster touches the altimeter
reference line.

D Altitude Analog Scale -- The altitude tape is a moving scale display

with fixed pointer. The scale markings on the tape are labeled in
500--foot increments. The scale digits are arranged with the larger
numbers at the top. The scale and its markings are white. The
500--foot tick marks are enhanced below 10,000 feet.

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Each 1000-- and 500--foot altitude increment is enhanced,

respectively, with a double or single chevron shape. This shape,
along with the shape of the altimeter reference line and altitude
select bug, are designed to align when the selected and current
altitudes are on an increment of 1000 or 500 feet.

D Altitude Reference Line -- The reference line is at the center point

of the altitude display. The rolling digit display readout to the right of
the reference line is the current altitude. The reference line is also
the 0 point for the altitude trend vector.

D Altitude Trend Vector -- The altitude trend vector originates at the

altitude reference line. The trend vector is a magenta thermometer
shape that corresponds to altitude increase or decrease rate and
moves along the left side of the altitude tape. The tip of the vector
indicates the altitude that will be reached, at the present rate, in six
seconds. Altitude rate is output from the MADC.

D Altitude Select Bug -- The cyan altitude select bug travels along the
left side of the altitude tape. The altitude select bug is notched to fit
the 1000 or 500 foot altitude tape chevron format. The bug is
displayed on the scale across from the altitude value set in the
altitude alert select display. If the bug is moved off the current scale
range, half of the bug remains on the scale to indicate the direction
to the set bug.

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VERTICAL SPEED (VS) DISPLAY

The VS display is shown in figure 4--26. Refer to the figure for the
location of annunciations described below.

TCAS STATUS

WHITE AMBER

TCAS TEST TCAS TEST

TCAS OFF TCAS OFF
TA ONLY TA ONLY

VERTICAL
SPEED ANALOG
SCALE

FLIGHT DIRECTOR FLIGHT DIRECTOR

VS VISUAL TARGET VS TARGET
DISPLAY BUG

TCAS ”FLY--TO”
ZONE

CURRENT VERTICAL
SPEED DISPLAY

TCAS ”NO--FLY”
ZONE

AD--50920--R1@

Vertical Speed Display

Figure 4--26

D Flight Director VS Target Display and Bug -- Engaging the

vertical speed mode brings the VS target bug into view. The VS
target bug moves along the right side of the VS scale. The bug lines
up with the value on the VS scale that is set with the guidance
controller. The bug is always cyan. The digital readout of the target
is displayed on top of this vertical speed scale. The target comes
from the flight guidance system.

NOTE: When the flight level change (FLC) FD mode is used in a

descent, the vertical speed bug and target display are
positioned at the preprogrammed rate of descent value.

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D TCAS II Resolution Advisory Display -- The TCAS II system

displays a green “fly to target”band and a red “do--not--fly”band on
the vertical speed display that commands the pilot to comply with a
resolution advisory (RA) to avoid a potential aircraft conflict.

D Vertical Speed Digital Display -- A green digital display of the

actual VS value is located on the zero reference line. This data is a
magnification of the digits on the scale and is readable to a 50 feet
per minute resolution. The maximum value is 9900 feet per minute.
For values between ±500 feet per minute, the digital display is
removed. At values beyond ±500 feet per minute, the digital value
of vertical speed is displayed.

For vertical speeds greater than ±3500 feet per minute, the pointer
is positioned in the appropriate direction near the end of the scale.
The digital display shows the actual vertical speed value.

D Vertical Speed Analog Scale -- The VS scale is a fixed scale with

a moving pointer. The scale on the display ranges from +3500 to
--3500 feet per minute. Display scale markings are 0, 1, 2 and 3. The
scale and its markings are white. The scale is expanded between
the +1000 to --1000 feet per minute range.

D TCAS Status Message -- The TCAS status messages are

displayed to the top left of the vertical speed display. When a TCAS
II RA is displayed, the vertical speed digital display matches the
color of the red or green band where the pointer is located.

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Typical PFD Presentations

This section shows PFD presentations that the pilot can see during
various flight phases. These examples do not show all display
possibilities, they are intended to represent common presentations.

TAKEOFF (TO) MODE

The aircraft has started the takeoff roll, as shown in figure 4--27. TO and
HDG modes have been selected. The VSPEEDS are set to V1 -- 119, VR
-- 124, V2 -- 132, the preselected altitude is 7000 feet mean sea level
(MSL), and the radio altitude is 5. The roll attitude is level while the flight
director cue is displaying the wings level, 12_ nose up takeoff
command.

169M 30 30

30 30

Takeoff Mode
Figure 4--27

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CLIMB TO INITIAL ALTITUDE

The aircraft is on climb, approaching 250 knots in HDG select and IAS
hold modes, as shown in figure 4--28. The AP and YD are engaged.
Altitude select is armed and the amber altitude preselect digits indicate
that the aircraft is between 1000 and 250 feet from the 7000--foot
selected altitude. Vertical speed is 1500 feet per minute. The altitude
trend vector indicates climb. Flight director commands are satisfied.

30 30

30 30

AD--50922--R1@

Climb to Initial Altitude

Figure 4--28

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ENROUTE CRUISE

The aircraft is in straight and level flight on a preselected altitude of

31,000 feet, flying HDG select and ALT hold with VOR armed for
capture, as shown in figure 4--29. AP and YD are engaged. Airspeed
is 220 knots, which is 0.70 Mach.

M
30 30

30 30

AD--50923--R2@

Enroute Cruise
Figure 4--29

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SETUP FOR APPROACH

The aircraft is flying HDG select and vertical speed (VS) while armed
for localizer (LOC) and glideslope (GS) capture, as shown in figure
4--30. The glideslope and localizer displays are in view. The amber
ASEL digits and box indicate the aircraft is in excess of 1000 feet above
the approach fix of 4000 feet. Currently the aircraft is flying at 180 knots.

30 30

5000

30 30

Setup for Approach

Figure 4--30

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TERMINAL AREA

The aircraft is in the terminal area, flying the heading select mode with
the LOC and GS armed. Altitude (ALT) hold is on, as shown in figure
4--31. A TCAS descent resolution advisory is displayed on the vertical
speed scale. The aircraft is in a descent to satisfy the command. AP is
disengaged.

30 30

30 30

TCAS II Resolution Advisory Approaching

the Glideslope
Figure 4--31

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APPROACH CAPTURE TRACKING AT RA MINIMUMS

The aircraft is setup on final approach with LOC and GS captured, as

shown in figure 4--32. The radio altitude of 100 feet is below the
200--foot minimum altitude, therefore the amber MIN is boxed. The
rising brown in the altimeter is 100 feet away from the altitude reference
line. VAPP is set for 130 knots.

M
30 30

1500

30 30

Approach Capture Tracking

at RA Minimums
Figure 4--32

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COMPARISON MONITORING

The amber comparison monitor annunciators shown in figure 4--33, are

in various locations on the PFD. Active messages are cleared when the
miscompare situation has been corrected. Comparison monitor
parameters are given in table 4--17.

I
A
S

AD--50914--R1@

Comparison Monitor Annunciators

Figure 4--33

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Parameter Monitored Tolerances

Pitch (PIT) Attitude ±5_

Roll (ROL) Attitude ±6_
Attitude (ATT) Active only when both pitch and roll
comparators are already tripped.
Heading (HDG) ±6_ (level flight)
Altitude (ALT) ±200 feet
Airspeed (IAS) ±5 knots
Localizer (LOC) deviation ±40 µA (»1/2 dot below 1200 feet
AGL)
Glideslope (GS) deviation ±50 µA (»1/2 dot below 1200 feet
AGL)
Instrument Landing System Active when both LOC and GS
(ILS) comparators are already tripped
Radio Altimeter (RA) 1/8 x (RA1+RA2) + 10 feet
2
CAS MSG Annunciated when there is a
miscompare between the symbol
generators and the processing of
the EICAS CAS messages.

Comparison Monitor Parameters

Table 4--17

NOTE: The comparison is made when the pilot and copilot have the
same type but different sources selected for display. If, for
example, the pilot and copilot both have ILS1 selected
(amber annunciation of the source), no comparison monitor
is active on that data (LOC, GS).

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PFD Caution and Failure Display

CAUTION (AMBER) ANNUNCIATIONS

The PFD with caution and failure annunciations is shown in figure 4--34.

AD--50928@

PFD With ARC HSI Display and

Caution/Failure Annunciations
Figure 4--34

D Same Attitude Source -- If the pilot and copilot are using their
normal attitude sources, there is no attitude source annunciated. If
the pilot and copilot have selected the same attitude source, that
attitude source is annunciated in amber on both PFDs.

D Common Symbol Generator -- When the reversionary mode and

one symbol generator is driving both pilot’s and copilot’s displays,
a reversionary warning is given in amber that indicates the
information source. This annunciation is displayed next to the upper
left corner next to the sphere and reads SG1 or SG2 depending on
whether the pilot’s or copilot’s symbol generator is the source.
D Same Air Data Source -- Under normal conditions, the on--side source
annunciator is not shown. When both pilots have selected the same
source, an amber source annunciation is displayed in the upper left
corner of the PFD, above the attitude director indicator (ADI).

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D Course Select Failure -- Failure of the course select signals,

replaces the display with amber dashes and removes the course
pointer from the display. This indication is also given in the event of
an invalid heading display or FMS source.
D Flight Director (FD) Failure -- In the event of a flight director failure,
an amber FD FAIL is displayed in the lateral mode annunciation box,
and the flight director mode annunciations and cue are removed.
D Radio Altimeter Failure -- In the event of a failure of the radio
altimeter, an amber RA replaces the digital radio altitude value.
D Vertical Deviation Failure -- An invalid condition or failure of the
radio source driving the vertical navigation scale is shown by
removing the deviation pointer and displaying a red X through the
scale deviation dots, as shown in figure 4--35. The scale and pointer
are removed for invalid FMS data.
D Distance Display Failures -- Failure of either the DME or FMS
distance signals is indicated by replacing the digital distance value
with amber dashes.

D Course Deviation Failure -- An invalid condition or failure of the

course deviation data is shown by removing the deviation bar and
displaying a red X through the scale deviation dots, shown in figure
4--35.

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M 30 30 00

30 30

AD--50929--R1@

PFD Caution Indications for Vertical and

Course Deviation
Figure 4--35

D Autopilot Status Annunciations -- These annunciations display

the engage status of the autopilot. The annunciation is displayed
above and left of center of the attitude sphere. Refer to table 4--3 for
annunciations and colors.

D Yaw Damper Status Annunciations -- Displays the engage status of

the yaw damper. The annunciation is displayed above and right of
center of the attitude sphere. Refer to table 4--18 for annunciations and
colors.

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Yaw Damper Status Annunciation Display Type Color

Engaged YD Steady Green

Autopilot test AP TEST Steady Amber
Normal YD disconnect YD 5 second flash Amber
Abnormal YD YD 5 second Amber
disconnect flash, then
steady

Yaw Damper Status Annunciations

Table 4--18

D TCAS Messages -- The TCAS status display is located to the left

of the vertical speed scale. Messages for TCAS are described in
table 4--19.

Annunciation Color Status

TCAS TEST White Pilot activated test

TCAS OFF White TCAS is off
TA ONLY White Traffic advisory only
TCAS FAIL Amber TCAS system failed
RA FAIL Amber Resolution advisory failure

TCAS Failure Annunciations

Table 4--19

D Maximum/Minimum Speed (MAX/MIN SPD) -- MAX/MIN SPD is

displayed to the left of the ADI. MIN SPD is displayed when in the
VS/IAS modes and the indicated airspeed (IAS) drops below 80
knots. MAX SPD is displayed anytime IAS exceeds Vmo/Mmo.

D Windshear Annunciation (WDSHEAR) -- When the windshear

detection system detects windshear, this annunciator flashes for 10
seconds and then goes on steadily. Depending on the severity of the
windshear detected, the annunciator can be amber or red. If the GA
button is pushed during a windshear caution or warning, the FD
vertical mode becomes windshear (WSHR) and vertical FD
guidance directs the aircraft.

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FAILURE (RED) ANNUNCIATIONS

PFD failure annunciations are shown in figure 4--36. Refer to the figure
while reading the descriptions that follow.

-- --

33

Failure Indications for ATT, MADC,

HDG, CRS Select and FD
Figure 4--36

D Attitude Reference System Failure -- Failure of either the pitch or

roll data is indicated by removing the pitch scale markings, turning
the entire attitude sphere to cyan and displaying a red ATT FAIL in
the top center of the attitude sphere.

D Micro Air Data Computer (MADC) Failures -- In the case of the

IAS and altitude scales, data is removed from the current value
pointer, the scale markings are removed and a red X is drawn
through the scale. The digital Mach CAS display failure is shown by
replacing the numerical value with amber dashes.

In the case of the vertical speed, the current value pointer is

removed, and amber dashes replace the digital display.

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D Heading (HDG) Select Failure -- Failure of the heading select

signals replaces the HDG display with amber dashes. The heading
bug is removed from the display. This indication is also given in the
event of an invalid heading display.

EXCESSIVE ATTITUDE DECLUTTER

The excessive attitude declutter display shown in figure 4--37, is

decluttered when an unusual attitude condition is displayed. The
unusual attitude conditions, anyone of which causes declutter, are
defined in table 4--20.

10

20 10

30 20

30

AD--50932--R1@

Excessive Attitude Declutter

Figure 4--37

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Monitored Parameter Tolerance

Bank angle Greater than ±65_

Pitch attitude up Greater than 30_
Pitch attitude down Greater than 20_

Excessive Attitude Declutter Parameters

Table 4--20

The following items are removed from the display when the logic in table
4--20 has been satisfied:
D FD mode annunciations and command bars and couple arrow
D Marker beacons
D Vertical deviation scale, pointer, and annunciator
D ADI localizer scale
D Speed bugs and readout
D Radio altitude display
D Altitude select, airspeed display and vertical speed bug data
D All flags and comparators except ATT and ADC (IAS/ALT)
D Low bank annunciator
D RA set digits.

NOTE: Declutter mode is inhibited when windshear flight director

mode is engaged.

PFD TEST MODE

This test can be initiated with the DC when airspeed is valid and less
than 60 knots, and the weight--on--wheels (WOW) switch is in the
on--ground mode. Selecting the various tests shows the invalid flags for
the following PFD information shown in figure 4--38:
D Micro Air Data Computer
D Flight Director
D Radio Altitude
D Attitude and Heading
D Navigation.

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ATT

AD--50931--R1@

PFD Test Mode

Figure 4--38

Multifunction Display (MFD)

The MFD has two major format areas:
D MFD Map
D MFD Plan.

In either of these formats, a window area can display any of the

following selection:

D System Displays

NOTE: One of several system displays is shown in the lower part

of the MFD, but not when weather, TCAS, or the checklist
display is displayed.
D Checklist
D TCAS
D Central maintenance computer display (only when the aircraft is on
the ground).

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Certain symbols are available with either display format. Some symbols
are only available with specific display formats. The lower portion of the
MFD is always reserved for displaying the display menu and submenu
sections.

NOTE: The symbols related to navigation are available on the MFD

only if the optional FMS is installed.

MFD COMMON SYMBOLS

Figure 4--39 shows MFD common navigation symbols.

VOR/DME

DME

VOR

AIRPORT

TOC/TOD

WAYPOINT

NDB
AD--50933@

MFD Common Symbols

Figure 4--39

D Upper Right Area

- TRUE AIRSPEED (TAS) -- TAS is output from the MADC and is
displayed in green digits.
- STATIC AIR TEMPERATURE (SAT) -- SAT is output from the
MADC and is displayed in green digits.
- TOTAL AIR TEMPERATURE (TAT) -- TAT is output from the
MADC and is displayed in green digits.

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D Upper Left Area

- FMS source annunciator
- Active waypoint in magenta
- Distance and time--to--go to active waypoint.
D Mid--Left Side
- WIND VECTOR -- Wind vector shows the magnitude and
direction of the wind, in green.
- WEATHER (WX) RADAR MODE ANNUNCIATIONS -- All
modes, except weather radar range, are annunciated in this area
as described in table 4--21.

Mode Annunciation
Weather Radar Mode Annunciation Color

R/T in Warm--up WAIT Green

REACT Mode RCT Green
Forced Standby FSBY Green
Standby STBY Green
Test Mode TEST Green
Weather Mode WX Green
Weather and Turbulence WX/T Green
Groundmap Mode GMAP Green
Flight Plan Mode FPLN Green
R/T Fail FAIL Amber
WX Interface Failure WX Amber
Stabilization Off STAB Amber

Weather Radar Mode Annunciations on MFD

Table 4--21

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- AMBER FAIL ANNUNCIATION -- When the amber FAIL

annunciation is displayed (optional PRIMUSâ 880 Weather Radar
System only), selecting the test mode displays the weather radar
system fault code data in place of the tilt information. Refer to the
PRIMUSâ 880 Weather Radar maintenance handbook for specific
information on fault code interpretation.

NOTE: The PRIMUSâ 660 and optional PRIMUSâ 880 radars

can also display plain English fault displays when the
TEST mode is selected.

- RADAR TILT -- Radar tilt is displayed in the form of a cyan number

with + (up) or -- (down) sign. If attitude compensated tilt (part of
the optional PRIMUSâ 880 Weather Radar only) is engaged, an
A is displayed in this location.

- TARGET ALERT (TGT) -- TGT alert is shown above the tilt angle.
It overrides the display of variable gain. When armed, TGT is
displayed in green. When activated, TGT turns amber and
flashes.

- VARIABLE GAIN -- When radar gain is selected as a variable, it

displays a value for radar gain in proportion to the radar controller
gain setting.

D Mid--Right Side

- MAP DESIGNATOR (if FMS is installed) -- This section of the

display, shown in figure 4--39, shows the cyan designator bearing
and distance position data. The designator is always cyan.

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MFD MAP DISPLAY

MFD map displays are shown in figure 4--40

HEADING SOURCE
ANNUNCIATION

MAG 1/2
HEADING DIGITAL
DG 1/2 (NOTE)
REFERENCE HEADING
HDG 1/2
LINE DISPLAY

FMS 360 MAG1 --35 SAT DRIFT BUG

KDVT --25 TAT
HEADING
12.5NM 300 TAS
SELECT BUG 12 MIN

PLAB1
MID--RANGE
NAVIGATION *PBD01 ANNUNCIATION
SYMBOLS LL01
PLAB2
KDVT AIRCRAFT SYMBOL
25 50 50
W

PILOT
DESIGNATOR
TGT 24
12

TX LATERAL DEVIATION
DIGITAL DISPLAY
1.0L
21
15

S
SYSTEM
DISPLAY
AREA

V1 VR V2
RTN 89 110 -- -- -- SET
NOTE: THE NORMAL (MAG), ON SIDE HEADING SOURCE IS
NOT ANNUNCIATED. AD--50935--R2@

MFD Map Display Symbols

Figure 4--40

D Heading Reference Line -- The heading reference line points to the

current heading (digitally displayed above it) on the compass rose.

D Heading Display -- The displayed compass rose consists of an

expanded 120_arc that is divided into 5_increments and is labeled
every 30_. The compass rose rotates around the stationary aircraft
symbol to show actual heading information. The compass indices
and labels are displayed in white. The display is augmented with a
white digital readout of actual heading above the compass lubber
line.

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D Heading Source Annunciations -- The heading source for this

display is the same source that drives the on--side PFD. Cross--side
data is displayed in amber. When the pilot and copilot horizontal
situation indicator (HSI) heading sources are the same, they are
annunciated in amber as is the heading source on the MFD.
NOTE: When the selected heading source has failed, the heading
source is annunciated as HDG 1/2.
D Drift Bug -- A magenta drift or actual track is shown on the map
display. It is supplied by the FMS. The position of the bug, relative
to the compass display, represents the actual track across the
ground. Once established, the position of the bug, relative to the
heading reference line on the desired course, represents the aircraft
drift angle.
D Range Rings -- Range rings are displayed to determine the position
of radar returns and active flight plan parameters. The range ring
boundary is the compass card arc and it displays the MFD controller
selected range. A mid--range ring is displayed and labeled with the
half--range distance. The half--range ring is changed to an arc when
a vertical profile is displayed.
NOTE: When weather radar data is displayed, the radar controller
is used to control the range.
D Aircraft Symbol -- The white aircraft symbol is a visual cue to
aircraft position, relative to actual heading and selected heading.
D FMS Lateral Deviation Digital Display -- Lateral deviation is
shown as a combination digital display with L/R direction. It is
displayed centered above the systems display area.
NOTE: When the TCAS or checklist window is active, lateral
deviation is not displayed.
D Heading Select Bug -- A heading bug is displayed on the compass
arc. Its position follows the pilot’s HSI heading bug. When the bug
is not in view, a cyan arrow is displayed to indicate the shortest
direction to the bug’s position.
D Pilot Designator -- Use the remote--mounted joystick and the MFD
menu, joystick submenu functions from the MFD bezel menu, to
change the position of the square map designator. Distance and
bearing location of the designator, relative to its reference point, is
shown above the system display area. When the designator is
parked in its home position (the aircraft symbol) it is not displayed,
When anchored at a waypoint, it has the same color as the waypoint.
The color of the designator distance/bearing readout matches the
designator.

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NOTES: 1. If checklist is selected for display, the joystick

cannot be used with the MAP designator.

2. The JSTK submenu can only be accessed if the

optional FMS is installed.

D Navigation Symbols -- Both the map and plan MFD displays use
common navigation symbols to describe flight plans.

WEATHER

Weather information from the radar is displayed only on the map format,
as shown in figure 4--41, whenever the radar controller has been turned
on and weather has been selected for display with the MFD bezel menu.

Activating the TEST, WX, or GMAP modes, with the aircraft on the
ground and the plan mode displayed on the MFD, displays an amber
TX annunciator.
W

24
12

21
15

AD--50936--R1@

Map Mode With Weather

Figure 4--41

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MAP FAILURE AND WARNING DISPLAYS

Map failure and warning displays are shown in figure 4--42.

E
24

12
21 15
S

AD--50937--R1@

Map Failure/Warning
Figure 4--42
D EICAS Failure Message -- CHK EICAS is displayed to indicate that
the EICAS wrap--around monitor has detected a difference between
what EICAS is displayed and what the DAUs have sent to the EICAS
for display.
D Heading Failure and Flags -- Failure of the displayed heading from
the AHRS is shown by removing the flight plan from the display. The
digital heading readout is replaced by amber dashes. As shown in
figure 4--42, a red HDG FAIL label is displayed below the lubber line.
NOTE: When the EFIS test is selected, HDG FAIL is replaced
with HDG TEST.
D MADC Failures -- MADC failures are indicated by replacing the
TAS, SAT, and TAT numerical values with amber dashes.
D FMS Failures -- A failure of the FMS removes the active lateral flight
plan waypoints, navaids, and airports from the display. This
indication is also given in the event of an invalid heading display.

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D Weather Radar Failure -- This failure is shown by displaying an

amber FAIL in the weather radar mode annunciation box.

NOTE: If the optional PRIMUSR 880 Digital Weather Radar System

is installed, when the amber FAIL annunciation is displayed,
selecting the test mode causes the weather radar system
fault codes to be displayed (in amber) in place of the digital
tilt data. Refer to the PRIMUSR 880 Digital Weather Radar
System’s System Description and Installation Manual, Pub.
No. A09--3944--001, for interpretation of fault codes.

TYPICAL MAP MODE PRESENTATIONS

Figure 4--43 shows a typical map mode without weather display. Figure
4--44 shows a typical map mode with weather displayed.

NOTE: A system display is shown in each illustration. Unless the

checklist or TCAS is displayed, a selected system display is
always shown.
W

24
12

21
15

AD--50939--R1@

Map Mode Without Weather

Figure 4--43

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21

3
S

6
15
299

E
12
2900

AD--51826--R1@

Map Mode With Weather

Figure 4--44

MFD PLAN MODE

The plan mode shown in figure 4--45, is a north--up presentation of the

active flight plan. The active waypoint is displayed in magenta at the
center of the display. The track lines are oriented to true north. A range
ring, showing the selected display range, is centered around the active
waypoint. In the plan mode, the aircraft symbol is always located at
present position in referenced to present heading. In addition to MFD
symbols described under MFD common symbols, this section details
unique plan mode symbols.

The plan mode consists of the following functions:

D True north--up map presentation
D Heading source annunciation (when cross--side heading source is
selected)
D FMS source annunciation
D FMS waypoint annunciations
D FMS waypoint, airport, and navaid display provisions

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D FMS desired track lines.

NOTE: Weather display is not available in the plan mode.

MFD Plan View

Figure 4--45

D Aircraft Symbol -- The aircraft symbol moves on the display as a

function of aircraft present position. The aircraft symbol is a visual
cue to the actual aircraft position relative to true north and the active
flight plan.

D Range Ring -- A range ring circle is displayed to determine the

position of the active flight plan parameters. The circle radius
corresponds to the selected range.

D North--Up Arrow -- The north--up arrow is permanently displayed

in the plan mode to indicate the mode’s north--up reference.

D Pilot Designator -- The MFD JSTK submenu functions and the

remote--mounted joystick still function with the designator,
However, the primary use of the joystick and designator in the plan
view is to position the circular viewing ring so that either the route
being flown or the maneuvering aircraft can be better observed. This
feature becomes useful in maintaining position orientation in the
terminal area as the aircraft is being vectored in for final approach.

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A bearing and distance display of the designator’s position relative

to its anchor waypoint is also displayed.
PLAN MODE FAILURE AND WARNING DISPLAYS
Plan mode caution and warning annunciations are shown in figure
4--46. Refer to the figure to find the location of failure and warning
annunciations described below.

AD--50943--R3@

Plan Mode Caution and Warning Displays

Figure 4--46
D MADC Failures -- When the MADC fails, the SAT, TAS, and TAT
numerical values are replaced with amber dashes.
D Heading Failures and Flags -- Failure of the displayed heading is
shown by removing the active flight plan and wind display, and
placing a red HDG FAIL at the top center of the plan range ring.
NOTE: When EFIS TEST is selected, HDG FAIL is replaced with
HDG TEST.
D WX (Weather Radar) Failure -- Weather radar failures are
annunciated with an amber FAIL in the middle left side of the display.
D FMS Failure -- When the FMS fails, the active flight plan, navaids,
and airports are removed from the display. This indication is also
given in the event of an invalid heading display.

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TYPICAL PLAN MODE PRESENTATION

Figure 4--47 shows a typical plan mode display.

TAS

21

3
S

6
12

AD--50941--R1@

Typical Plan Mode With Navaid Components

and Flight Plan
Figure 4--47

CHANGING A WAYPOINT

Adding a waypoint is a routine operation on the MFD. Table 4--22

describes the procedure to change a waypoint. The procedure inserts
a new waypoint between DRK1 and LL03. The MAP display is normally
used to define a new waypoint.

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Step Procedure

1 Use the JSTK submenu from the MFD’s MFD menu (and
the external joystick) to move the designator to the location
of the new waypoint, as shown in the display below.

21

3
S

6
15

E
2 Use the ENT bezel button from the JSTK submenu to send
the new waypoint designated position to the FMS control
display unit. When entered, the display appears as shown
in the display below.
21

S
6

12

3 Enter a new waypoint into the FMS.

NOTE: The degree to which the new waypoint can be combined with the
existing FMS flight plan depends on the installed FMS.

Designating a New Waypoint Procedure

Table 4--22

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MFD DISPLAY WINDOW

The checklist and TCAS displays are shown on an MFD display

window. When either the map or plan function is selected, the MFD’s
map or plan navigation displays are slightly repositioned toward the top
of the MFD. The display window is shown in the bottom portion of the
MFD between the lower left and right display windows.

NOTE: When checklist or TCAS is displayed, a portion of the MFD’s

map or plan mode are removed to make more space.

CHECKLIST (OPTION)

The checklist window can be displayed with either the MFD map or plan
mode, as shown in figure 4--48. The checklist is selected for display in the
display window, using the MFD display controller. The integrated avionics
computer (IAC) is capable of storing and displaying 400 pages of text.
These pages are stored in electronically programmable memory 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.

NOTES: 1. The checklist is stored in both IACs. Cockpit--mounted

plugs are used to access each IAC.
2. The checklist is generated by the operator using the
operator’s PC. The checklist is loaded onto the IAC
through an aircraft mounted plug and a PC. Refer to
Section 9, Troubleshooting, for checklist uploading
procedures.
3. A checklist can be displayed on only one MFD at a
time. When selected on one side, it is not available for
selection on the other side.
4. The checklist stored in each IAC are not synchronized
to each other. If, during operations, the pilot on the side
displaying the checklist rotates the MFD reversionary
knob to PFD, the checklist is restarted. If the checklist
is subsequently selected to the cross--side MFD, the
checklist reinitializes to the checklist power--up
sequence (disclaimer page).

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AD--50945--R2@

Typical Checklist Display

Figure 4--48

Pushing the CKLST MFD bezel menu button, displays the master
index. Normal, abnormal, or emergency indexes are selected from the
master index. The normal checklists are arranged in the order of
standard flight operations. Pushing the button displays the normal
checklist index page that contains the lowest order incomplete and
unskipped checklist with the active selection of that checklist.

The SKP, LNBK, RCL, PAG, and ENT menu selections control the
checklists.

The emergency index on the master index enters the abnormal and
emergency checklist display function. When a selection is made, an
index (arranged by aircraft system) is displayed, The crew can then
select the listing for the malfunctioning system area that in turn
accesses the specific malfunction checklist. Therefore, the format of
the electronic checklists closely follows the approved abbreviated
checklist for the aircraft.

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The SKP, LNBK, RCL, PAG, and ENT menu selections control the
emergency checklist. The emergency checklist is used the same way
as the NORM except for the action taken at the completion of a
checklist. When the emergency checklist is completed, all checklist
items are removed from the page and EMERGENCY PROCEDURE
COMPLETE is written below the amber checklist title. This is cleared
when the master checklist is selected.

NOTE: If the optional FMS is installed, the remote mounted joystick

can also be used to select checklist line entries and to move
between pages.

TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS)

The TCAS display is selected for display on the MFD map mode using
the TCAS menu button on the MFD bezel controller. The TCAS display
shown in figure 4--49, has the following specific symbols:
W

24
12

12
21
6
15

S --10

--05

00

AD--50946--R1@

MFD With Map Mode Display and Optional TCAS

Figure 4--49

D TCAS Mode and Failure Annunciations -- When the TCAS

system fails, the various traffic symbols are removed and a mode
annunciation is displayed.

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D No Bearing Target Readout -- No bearing data when system has

range but not bearing information.

D TCAS Range Ring -- The TCAS range ring is a white 60°arc with
the aircraft symbol as its centerpoint. The digital value at the right
end of the arc is the range in NM. The range is set through the RMU.

D Proximity Advisory Arrow -- This symbol is a target relative

altitude arrow that indicates targits with greater than a 500--foot per
minute climb/descent arrow.

D 2 NM Range Ring -- The 2 NM TCAS range ring is a circle of dots

that surround the aircraft

NOTES: 1. When the selected TCAS range is less than 20 NM,

the 2 NM range ring is displayed. At ranges greater
than 20 NM the range ring is removed. The TCAS
display shows a mid--range ring.
2. TCAS range is controlled with the RMU controller.

D Traffic Symbols
- Other traffic diamond (hollow blue)
- Proximate traffic diamond (blue)
- Caution area/traffic advisory circle (amber)
- Warning area/resolution advisory square (red)

D TCAS Altitude Display Submodes -- The following are TCAS

submodes:

- ABV/BLW (ABOVE/BELOW) -- TCAS relative altitude is selected

to look above or below the normal TCAS altitude band.

- NRM (NORMAL) (not annunciated) -- TCAS relative altitude is

selected to look at the normal TCAS altitude band.

- FL (FLIGHT LEVEL) -- The actual altitude of the traffic is

displayed.

D TCAS Auto -- When selected with the RMU, the TCAS window
automatically switches to TCAS when a TCAS TA or RA is detected.

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MFD System Displays

The data acquisition units (DAU) collect all the information required for
the system displays. This information is displayed in the lower third of
the MFD and it is selected using the MFD bezel menu SYS button.

NOTE: The last selected system display is shown on the MFD unless
the TCAS or checklist display is selected.

The system display area can show the following type of system
information:

D Takeoff (T/O) -- The takeoff digital display shows engine takeoff

reference temperatures, anti--ice system status, and engine oil quantity
levels. In addition, a synoptic of aircraft door status is displayed.

D Environmental Control/Anti--Ice System (ECS A/I) -- The ECS A/I

display is a synoptic display of cabin and cockpit temperatures, oxygen
pressure, and engine bleed air temperatures (for anti--ice purposes).

D FUEL -- The FUEL display is a synoptic display of aircraft fuel

quantity, fuel used, fuel temperature, and fuel tank pump status.

D Hydraulic (HYD) -- The HYD display is a synoptic display of

hydraulic fluid quantity and pressure, pump status, and brake
temperatures.

D Electrical (ELEC) -- The ELEC display is a synoptic display of the

electrical bus connections, generators, APU, GPU, and battery
status.

TAKEOFF (T/O)

The takeoff display shown in figure 4--50, displays the following

information.

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ENGINE T/O
STATUS ARROW

ENGINE T/O DATA

AIRCRAFT DOOR
TEMPERATURE STATUS
ENTRY

ANTI--ICE STATUS
ANNUNCIATION

ENGINE OIL LEVEL

AD--50947--R1@

Takeoff Display
Figure 4--50
D Aircraft Door Status -- The display in figure 4--50, also indicates the
status of the aircraft doors. Closed doors are shown in green. An
open door is indicated by a red DOOR OPEN message. Under
certain conditions, opening a door displays the DOOR OPEN status
message boxed and flashing for 5 seconds.
D Engine Oil Level -- The oil quantity for each engine is shown.
Normally, the quantity is green. When the quantity drops to 5 quarts
and below, the digits are boxed and displayed in amber. Once the
aircraft is airborne, the oil level is removed.
D Anti--Ice (REF A--ICE) Status Annunciation -- The REF A--ICE
indicates the current state of the bleed air anti--ice system. The
annunciation indicates either ON or OFF.
D Reference T/O Temperature (REF TO TEMP) -- The pilot can enter
a reference temperature for the engine FADEC to use for takeoff
power calculation. This input can be normal or flexible T/O power.
The pilot can also select engine anti--ice for T/O. Depending on the
combination of entries, the engine T/O status indicator points to the
selected engine T/O condition. The digits are either cyan or amber.
D Engine T/O Mode -- The engine take off mode readout is based on
the logic described in table 4--23.

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Takeoff Mode 1 Takeoff Mode 2 Annunciation

0 0 T/O--1
0 1 ALT T/O--1
1 0 ALT T/O--2
1 1 ALT T/O--3

Engine T/O Mode Logic

Table 4--23

D Engine T/O Status Arrow -- This arrow moves up and down the
three categories (T/O MODE, REF TO TEMP, and REF A--ICE) to
indicate a valid status condition. In the case of REF A--ICE, ON is
the valid condition.

ENVIRONMENTAL CONTROL SYSTEM/ANTI--ICE (ECS A/I)

The ECS display, shown in figure 4--51, displays the following information.

AD--50948@

Environmental Control System (ECS)

System Display
Figure 4--51

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D Cabin and Cockpit (CKPT) Temperature -- Current temperatures

in each section is displayed in °C and ranges from --54°to 54°C.

D Oxygen Pressure -- Oxygen pressure is displayed in pounds per

square inch (PSI) and has color codes that are described in table
4--24. Both the analog scale and digital display are shown.

Color Lower Limit Upper Limit

Red ------ 250

Amber 250 410
Green 410 2080
NOTE: When the pressure range is displayed in amber or red, the
digits are boxed in that color.

Oxygen Pressure Limits in PSI

Table 4--24

D Engine Output Bleed Air Temperature (Anti--Ice) -- An analog

scale display the left and right engine bleed air temperature being
supplied to the ECS. The color of the scale and pointer are defined
in table 4--25.

Color Limits

Amber Less than 260 °C

Green 261--304 °C
Amber Greater than 305 °C

Bleed Air Temperature Color Limits

Table 4--25

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FUEL SYSTEM

The fuel system display shown in figure 4--52, displays summary

information concerning the fuel supply. It gives the following information.

AD--50949@

Fuel Page
Figure 4--52

D Tank Fuel Quantity -- Left and right tank fuel quantity is shown by
an analog scale and pointer with a digital readout. Fuel quantity can
be displayed in pounds (lb) or kilograms (kg). The range of values
and their color coding are described in table 4--26.

Color Lb Range Kg Range

Red Less than 620 lb Less than 280 kg

Amber 630--880 lb 290--400 kg
Green More than 890 lb More than 410 kg

Fuel Quantity Color Coding

Table 4--26

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NOTES: 1. The pointer and digital readout have the same color
as the scale.
2. When the fuel quantity enters the amber or red
range, the digital display is boxed. The box flashes
for the first five seconds.
3. Left wing fuel quantity is supplied by DAU No. 1.
Right wing and fuselage fuel quantity is supplied by
DAU No. 2.
4. The kg or lb designation is set during installation.

D Total Fuel Quantity -- Total fuel quantity is computed from the sum
of the fuel in the left and right fuel tanks.

D Fuel Used -- Fuel used is determined by periodically evaluating fuel

flow rate and calculating the amount of fuel used for each engine and
maintaining a running total. APU fuel consumption is also added to
the total.

D Fuel Tank Temperature -- Wing tank fuel temperature is shown in

green digits, the range is from --64°to +64 °C with 1°resolution.
When the fuel temperature is less than --40 °C, the display is shown
in amber.

D Fuel Boost Pump Status -- The left and right fuel boost pump
status is shown in the fuel quantity display. The A, B, and C identify
pumps, or OFF can be displayed in green.

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HYDRAULIC (HYD) SYSTEM

The hydraulic system display shows the major elements of the

hydraulic system that is shown in figure 4--53. The elements include:

AD--50950@

Hydraulic (HYD) System Page

Figure 4--53

D HYD Pressure and Fluid Quantity -- The hydraulic pressure and

fluid quantity limits for both systems are given in table 4--27.

Color Pressure in PSI Quantity in Qts

Amber (low) Less than 1200 Less than 1

Green 1201 -- 3300 1.1 minimum
Amber (high) Greater than 3300 ------
NOTE: When the status of the hydraulic pressure enters the amber re-
gion, the digital value is boxed, flashes for 5 seconds, and then
goes on steadily.

Hydraulic Pressure and Fluid Quantity Limits

Table 4--27

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D Electric Hydraulic Pump Status -- The electric hydraulic pumps

can be switched ON and OFF. The ON/OFF status is shown on the
display.
D Brake Temperature -- Brake temperature is displayed on vertical
analog scales that act as a monitor reference. The pointer and scale
are green and remains green until the pointer enters the yellow
region. The color changes according to limits in table 4--28.

Color Range

Green Less than 200 °C

Amber More than 200 °C

Brake Temperature Limits

Table 4--28

ELECTRICAL (ELEC) SYSTEM

The electrical system display shown in figure 4--54, shows the dc
electrical block diagram. Bus voltage and amperage loads and battery
temperatures are shown.

AD--50951@

Electrical (ELEC) System Display

Figure 4--54

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D Bus Voltage and Amperage Loads -- A generator volts and amps

display is normally shown in green for the following dc power
sources:
- Generators 1 through 4 (GEN 1/2/3/4)
- Auxiliary power unit (APU) and ground power unit (GPU)

D DC Power Interconnect Buses -- The dc power interconnect buses

are normally displayed in green. The resolution of the voltage
display is 0.1 vdc. The resolution of amperage readouts is 5 amps.

NOTES: 1. When a digital value exceeds the limit, the value turns
amber, is boxed and flashes for 5 seconds.
Depending on the circumstances, this annunciation
changes if there has already been an abnormal
situation.
2. The annunciation of power sources exceedances
only occurs when an electrical power system
message is displayed on the CAS.
3. Failure of an interconnect bus is indicated by a solid
amber line.
4. The GPU power display does not include an
amperage load indication.
5. APU data is displayed under the following
circumstances:
- APU master is ON
- APU turbine speed is greater than 10%
6. GPU data is only displayed with weight--on--wheels
(WOW).

D Battery Voltages and Operating Temperatures -- Battery

voltages and operating temperatures in °C are displayed in a digital
format. Normal indications are displayed in green.

NOTES: 1. Amber (abnormal) annunciations occur when an

electrical power system message is displayed on
CAS.
2. The displayed battery temperature is only transmitted
from the primary sensor. The battery temperature
CAS message can be triggered by either the primary
or secondary battery sensor.
3. Valid/invalid indications for the battery display are
only transmitted from the primary sensor.

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System Display Failure Annunciation

Failures (or invalid data) are shown on system displays by replacing
digital data with dashes and remove the analog scale pointers. These
can be seen when EFIS TEST button on the DC is pushed.

DISPLAY SYSTEM REVERSIONARY MODES

D EFIS 1 (Pilot Normal) Failure -- A failure of SG 1 in DGC 1 (or its

interface) shows itself in one of the following manners:
- Red X on the pilot’s PFD
- Red X on the pilot’s MFD
- Red X on the EICAS
- Red X on both the pilot’s PFD, MFD, and EICAS.

To recover from this failure, push the SG button. The NO. 2 SG (No.
2 DGC) now drives all five DUs.

D EFIS 2 (Copilot Normal) Failure -- A failure of the SG (in DGC 2)

displays a red X on the copilot’s PFD and the MFD.

To recover from this failure, push the SG on the copilot’s

reversionary control panel. When this is done, the No. 1 SG (No. 1
DGC) drives all five DUs.

DISPLAY CONTROLLER FAILURES

D Pilot’s DC Fails -- If the pilot’s DC fails, the SG driving the pilot’s

display automatically reverts to the display source selections given
below:
- AHRS 1
- MADC 1
- NAV 1
- RAD ALT 1
- NO. 1 BEARING POINTER -- NAV 1
- NO. 2 BEARING POINTER -- OFF
- BARO -- Last selected setting
- FULL HSI
- MFD No. 1 menu inoperative.

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D Copilot’s DC Fails -- If the copilot’s DC fails, the SG driving the

copilot’s display automatically reverts to the display source
selections given below:
- AHRS 2
- MADC 2
- NAV 2
- RAD ALT 2
- NO. 2 BEARING POINTER -- NAV 2
- NO. 1 BEARING POINTER -- OFF
- BARO -- Last selected setting
- FULL HSI
- MFD No. 2 menu inoperative.

D Both Pilot’s and Copilot’s DC Fail -- If both the pilot’s and the
copilot’s display controllers fail, the symbol generator driving the
EICAS displays automatically uses DAU 1 channel A and DAU 2
channel B. The pilot/copilot’s PFD and the MFD sources for this
condition are the same as described under the Pilot DC Fails and
Copilot DC Fails paragraphs.

Standby Navigation Display

Either radio management unit (RMU) can be selected to display certain
navigation data in an HSI arc format. The RMU operation is described
in Honeywell Pub. No. A28--1146--050. This section details the use of
the RMU for the standby navigation display role.

The RMU PGE (page) button is used to select the standby navigation
display. Redundant sources of 28 V dc power are supplied to RMU No.
1 NAV, No. 1 AHRS, and both DAUs to ensure display and data
availability. The display shows the following information:
D Heading arc display
D VOR/ILS navigation course information
D DME
D f (VOR) and Z (automatic direction finder (ADF)) bearing pointers
D NAV1/ADF1 frequency display
D Marker beacon.

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The standby navigation display shown in figure 4--55, displays a typical

enroute VOR course with the ADF and VOR bearing pointers in view.

AD--50953@

Standby Navigation Display (Typical)

Figure 4--55

The paragraphs that follow detail specific symbols. Refer to figure 4--56
for an RMU backup navigation display.

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MARKER BEACON

OM
MM
IM

AD--50954--R1@

RMU Backup Navigation Display

Figure 4--56

D Z ADF Bearing Pointer -- The bearing pointer, and its legend and
digital position display are white. When the pointer (or its tail) is not
in view, a double white arrow is displayed above the compass arc,
on either the left or right side, indicating its location.

D VOR/ILS and ADF Frequency Display -- The VOR/ILS tuned

frequency is shown in the upper left and the ADF frequency is shown
in the upper right of the display; both are white. When the RMU line
select key adjacent to the respective window is pushed, that
frequency is enclosed in the RMU yellow tuning box. The tune knob
is used to select the desired frequency.

D Digital Heading -- This is the digital value of the current heading.

D f VOR Bearing Pointer -- The bearing pointer, and its legend and
digital position display are cyan. When the pointer (or its tail) is not
in view, a single cyan arrow is displayed above the compass arc, on
either the left or right side (as appropriate), indicating its location.

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D VOR Pointer Out of View Right/Left -- When the VOR pointer is

out of the range displayed on the RMU ND, an arrow above the
compass to either the right or left, shows the shortest direction to the
VOR pointer.
D ADF Pointer Out of View Right/Left -- When the ADF pointer is out
of the range displayed on the RMU ND, an arrow below the VOR
pointer to either the right or left, shows the shortest direction to the
ADF pointer.
D Heading Compass -- The compass display is a standard compass
card that displays 90°of the compass.
D Glideslope Pointer -- The glideslope scale and pointer indicate the
amout of vertical deviation during an approach.
D Glideslope Scale -- When a localizer frequency is tuned, the white
glideslope is displayed, and the VOR (f ) bearing pointer, VOR
bearing readout, and bearing label are removed. The valid
glideslope pointer is shown in green.
D Lateral Deviation Scale -- The lateral devaition scale is the
standard four dot scale that indicates the amout of left or right
deviation from the aircraft course (see the course pointer below).
D DME Distance -- DME distance to the selected VOR or ILS is shown
in green digits in the lower right corner.
D Marker Beacon -- The beacon display is to the left of the DME
information. The mnemonics flash at the proper rate and are as
should in figure 4--29.

Beacon Color

OM Cyan
MM Amber
IM White

Marker Beacon Colors

Table 4--29

D Aircraft Symbol -- The aircraft symbol indicates the aircraft’s

position as it relates to the rest of the display.
D TO/FROM Indicator -- The center of the HSI contains an aircraft
symbol. Lateral deviation is shown on a scale with two dots on either
side of a centered position. This represents navigation deviation
from the selected course for the specific source. A green TO or FR
(from) indication is shown to the right of the digital course readout.

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D Tune Box -- Items that are displayed in the tune box can be changed
using the RMU TUNE knob.

D Selected Course -- The selected course in the tune box can be set
using the RMU TUNE knob.

D Lateral Deviation Indicator -- The lateral deviation indicator is part

of the course pointer. It slides left and right along the lateral deviation
scale to indicate the amount of lateral devation from the selected
course.

D Digital ADF Bearing -- This readout indicates the number of

degrees to the tuned ADF.

D Digital VOR Bearing -- This readout indicates the number of

degrees to the tuned VOR.

D Heading Arc Compass Rose -- 90_of heading is displayed. The

digits and arc are shown in white.

D Course Pointer With Display -- A green course arrow is displayed.

A digital readout (in green) of the selected course is shown in the
lower left corner. When the RMU line select key next to the course
readout is pushed, it is boxed in yellow. The RMU tune knob is used
to select a new course, The inner knob selects 1_changes, the other
knob selects 10_changes.

NOTE: The yellow tune box returns to the course window after 20
seconds of RMU inactivity.

D Standby Navigation Display Failure Indications -- See figure

4--57 for an illustration of the RMU backup navigation display with
failures.

- COURSE GLIDESLOPE DEVIATION -- Deviation bar or pointer

is removed and replaced with a red X.

- DIGITAL DISPLAY OF DME DISTANCE, NAV OR ADF

FREQUENCY, DIGITAL BEARING -- When these inputs fail, the
data is replaced with amber dashes.

- HEADING FAILURE -- A red HDG FAIL is displayed in the center

of the arc display. The course pointer and its digital readout are
removed. All bearing pointer information is removed.

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AD--50955@

RMU Backup Navigation Display With Failures

Figure 4--57

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5. Engine Instrument and Crew

Alerting System (EICAS)

INTRODUCTION
The EICAS shown in figure 5--1, is a part of the electronic display
system. It has the following components:
D Electronic Display Unit (1)
D Display Symbol Generator
- Part of the Display Guidance Computer
- Shared with EFIS
D Data Acquisition Units (DAU) (2)
D EICAS Bezel Controller (1)
D Electronic Flight Instrument System (EFIS) Reversionary Control
Panels (2)
D EICAS Display Reversionary Panel (1)
D Backup EICAS Display on the Radio Management Units (RMUs)
D External Master Warning (Red) and Master Caution (Amber) Lamps
and Switches

The EICAS is an integrated digital computer/display system that

replaces the majority of the traditional gauges and warning lights
scattered throughout the cockpit.

In the event of a failure of the EICAS display unit, either reversionary

control panel can select the EICAS in place of a multifunction display
(MFD) and that MFD’s bezel controller knob controls the EICAS menu.

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 PRIMUSr1000 Integrated Avionics System

TAS TAS

NM NM

S S

AD--50956--R1@

EICAS Block Diagram

Figure 5--1

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 PRIMUSr1000 Integrated Avionics System

CONTROLS
EICAS Bezel Controller
The EICAS bezel controller is shown in figure 5--2. The rotary knob is
used to scroll the amber and cyan CAS messages.

AD--50957--R2@

EICAS Bezel Controller

Figure 5--2

Master Warning and Caution Annunciation Switches

The master caution and warning annunciation switches are located on the
glare shield adjacent to the display controllers (DC). These switches,
shown in figure 5--3, are used to acknowledge the display of a warning or
caution message. The messages are shown on the EICAS display.

AD--50958@

Master Warning and Caution

Annunciation Switches
Figure 5--3
The display of a warning crew advisory system (CAS) message causes
WARN to flash in red. A caution CAS message causes CAUT to flash
in amber.

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DISPLAYS

EICAS Display
The engine instrument and caution advisory display is divided into five
functional areas, as shown in figure 5--4. The areas include the
following:
D Primary Engine Display
D Crew Alerting System
D Gear/Flaps/Spoiler
D Cabin Pressurization and Auxiliary Power Unit (APU) Status
D Pitch/Roll/Yaw Trim Display.

CREW
ALERTING
SYSTEM

PRIMARY
ENGINE
DISPLAYS

GEAR/
FLAPS/
SPOILER

CABIN PITCH/ROLL
PRESSURIZATION YAW
AND TRIM
APU STATUS DISPLAY
AD--50959--R1@

EICAS Display Functional Divisions

Figure 5--4

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Primary Engine Displays

The engine displays shown in figure 5--5, receive their information from
two DAUs. The two DAUs are each dual channel and receive various
aircraft engine and sensor data. One channel is primary, and both
channels process the data in parallel.

The symbol generator (SG) supplying the EICAS display (normally the
NO.1 SG) also supplies display processing logic for some CAS
messages.

The primary engine displays are shown in the left half of the EICAS
display. The display is divided into seven areas:
D N1 fan (FAN)
D Interstage turbine temperature (ITT)
D N2 turbine
D Fuel flow (FF) in pounds per hour (PPH) or kilograms per hour (KPH)
D Fuel quantity (FUEL lbs, or kg L/R)
D Oil pressure/temperature (OIL PRESS/OIL TEMP)
D Engine vibration for low pressure (LP) and high pressure (HP).

Some engine warning and status messages are also displayed.

N1, ITT, OIL PRESS/TEMP, and VIB LP/HP are gauges. N2, OIL, fuel
flow, and fuel quantity are only digital readouts.

The round engine gauges are a fixed arc with moving pointer and digital
current value window located above the pointer rotation point. Each
engine gauge type is labeled. Vertical gauges have a pointer that moves
on a fixed scale. Some vertical gauges also have a digital display.

The following is a list of the parameters shown on the engine display.

D Analog fan (N1) scales with digital readout.
D Analog ITT scales with digital readout.
D Digital turbine (N2) readout.
D Digital FF readout.
D Digital fuel quantity readout.
D Digital and analog readouts of oil pressure.
D Digital and analog readouts of oil temperature.
D LP and HP vibration analog displays

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D Annunciations are as follows:

- Ignition (IGN)
- Thrust reverser display status
- Thrust rating mode
- Full authority digital engine control (FADEC) status
- Engine auxiliary power reserve (ATTCS).

PSI

AD--50960@

EICAS Engine Instruments Display

Figure 5--5

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FAN (N1) SPEED DISPLAY

D N1 Analog Scale -- The fan arc is white. The ranges are listed in
table 5--1.

Lower Limit Upper Limit

Color (%RPM) (%RPM)

White 0 100.0
Red 100.1 110.0

N1 Analog Scale
Table 5--1

When the analog pointer enters the red range, the normally white
pointer and digital fan readout turn red.

D Fan Digital Display -- Displays digital readout of fan (N1) RPM as

a percent of rotation speed. Display values range from 0 to 110% in
0.1% increments.

NOTE: The digital readout can indicate as high as 199.9% N1.

D N1 Request Bug -- The N1 request bug is displayed as a solid green

triangle. The bug indicates the throttle position as set by the pilot.

D FADEC N1 Bug and Digital Target -- The FADEC N1 bug and digital
target are shown as a cyan T--shaped symbol that indicates the
thrust target for the phase of flight, as determined by the FADEC.

D FADEC Mode Annunciations -- The FADEC modes, that are

determined by the throttle position, are annunciated in cyan.
Normally, a single annunciation is supplied at the top of the engine
display. The modes are described in table 5--2.

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Annunciation Meaning

CON Continous
CLB Climb
CRZ Cruise
PWT (To be determined)
T/O--1 Takeoff Mode 1
T/O--2 Takeoff Mode 2
T/O--3 Takeoff Mode 3
ALT T/O--1 Alternate Takeoff Mode 1
ALT T/O--2 Alternate Takeoff Mode 2
ALT T/O--3 Alternate Takeoff Mode 3
R--MODE Reversion Mode
NOTE: At the time of publication of this manual, not all FADEC modes were
activated.

FADEC Modes of Operation

Table 5--2

D FADEC in Control -- The FADEC in control is annunciated as an A

or B next to the N1 display. It indicates which FADEC is controlling
the engine.

D Automatic Takeoff Thrust Control System (ATTCS) -- When the

reserve takeoff power function is selected, ATTCS is annunciated
in green.

D Reverse Thrust Annunciation -- When reverse thrust is armed,

and the thrust reverser is normal, a green REV is displayed near the
N1 display.

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INTERSTAGE TURBINE TEMPERATURE (ITT) DISPLAY

D ITT Analog Scale -- The scale listed in table 5--3, is color banded,
as to indicate normal and parameter exceedance ranges.

Lower Upper
Operating Condition Color Limit Limit
All modes except White 0 _C 921 _C
ALT--T/O
Red 922 _C 1000 _C
NOTES: 1. Color ranges change for engine start and CON, the CLB, and
CRZ conditions.
2. ALT--T/O modes have different ITT limits.

ITT Limits
Table 5--3

The scale ranges from 0 to 1000 _C. The scale from 800 to 1000 _C
is twice the rest of the range.

In the red range, the pointer and digital display change color.

A red tick mark, indicates the beginning of the red range.

D ITT Digital Display -- Ranges from 0 to 1000_C in 1.0 _C

increments.

ENGINE IGNITION (IGN) ANNUNCIATION

Each engine’s dual ignitors are monitored for activation. The

annunciation color relates to ignitor status as listed in table 5--4.

Number of Operating IGN Annunciation

Ignitors Color

2 Green
0 Amber

Ignitor Annunciation Color Definition

Table 5--4

The annunciation can be A, B or AB, depending on the selection by the

FADEC. The annunciation can also be OFF.

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DIGITAL TURBINE (N2) DISPLAY

This shows the speed of the engine turbine as a percentage of rotation

speed. Refer to table 5--5 for the range of N2.

Lower Limit Upper Limit

Color (%RPM) (%RPM)

Green 0 102.4
Red 102.5 120.0
NOTE: When APR is selected, the white region is only to 100.0%.

N2 Limits
Table 5--5

OIL PRESSURE (OIL PSI) DISPLAY

The analog and digital displays for oil pressure are shown in PSI.
Resolution is 1 PSI. Color coding of the digits is listed in table 5--6 for
N2 <88% and table 5--7 for N2 >88%.

Lower Limit Upper Limit

Color (PSI) (PSI)

Red 0 31
Green 32 90
Amber 91 99

Oil Pressure Limits for N2 <88%

Table 5--6

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Lower Limit Upper Limit

Color (PSI) (PSI)

Red 0 31
Amber 32 49
Green 50 90
Amber 91 99

Oil Pressure Limits for N2 >88%

Table 5--7

FUEL FLOW (FF PPH FLOW) DIGITAL DISPLAY

The white digits display a range of 0 to 4400 PPH with a resolution of

10 PPH.

NOTE: At the time of installation, the fuel flow can be configured for
KPH. If KPH is selected, the value range is 0 to 1820 KPH.
The display is labeled FF KPH.

FUEL QUANTITY (FUEL LB) DIGITAL DISPLAY

Aircraft fuel quantity, in pounds, is shown in digital format for the left and
right wing.

The displays read in 10--pound increments. When each wing indication

reaches 880 pounds or less, the digits turns amber. When each wing
indicates 620 pounds or less, the digits turn red.

Fuel imbalance between the left and right wing tanks is shown by a CAS
message.

NOTE: If selected at installation, fuel quantity can be shown in

kilograms (KG). The label LBS is replaced with KG.

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OIL TEMPERATURE (OIL _C) DISPLAY

The digital display for oil temperature is displayed in _C. Resolution is

1 _C. Color coding of the digits is listed in table 5--8.

Lower Limit Upper Limit

Color (_C) (_C)

Amber ------ 20
Green 21 126
Red 127 126

Oil Temperature Limits

Table 5--8

ENGINE VIBRATION DISPLAY

The engine vibration display is an analog scale that shows both the LP
and HP turbine sections. The scale and pointer are described in table
5--9.

Condition Limit (IPS)

Color LP HP
Green Less than 1.8 1.1
Amber 1.9 and above 1.2 and above

Engine Vibration Color Limits

Table 5--9

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GEAR/FLAPS/SPOILER POSITIONS

D Landing Gear Position -- The landing gear position is indicated by

annunciations in three boxes; left, nose and right. Table 5--10,
defines the annunciations.

Gear Position Annunciation Color

Up UP White (note 1)
Down DN Green (note 1)
Transit Diagonal lines Amber
NOTES: 1. If the position of the landing gear handle does not agree with the
landing gear itself, the annunciation is amber.
2. If the MFD reversionary switch is selected to EICAS, the leading
gear display is on the selected MFD display.

Landing Gear Position Annunciations

Table 5--10

D Flap Position -- A digital display of the degrees of flap extension is

shown. The digits range from 0°to 47°(full flaps). Table 5--11 defines
the color logic for this display.

Condition Color Logic

Condition A Red
--Weight--on--wheels (WOW)
--Flaps are at 9°or 22°
--Thrust lever angle greater than 60°
Condition B Amber
--Flap failure or fault
Condition C White
--Flaps position 0

Flaps Position Symbol Color

Table 5--11 (cont)

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Condition Color Logic

Condition D Green
--Flaps position is as selected
NOTES: 1. When flap condition A or B exists, the flap position digital display
is boxed in the same color.
2. Only 0°, 9°,22°, 33°, 45°are displayed if flap control units are
valid (other positions are denoted by a dash (--)).
3. If either flap control unit fails, the actual flap position from the DAU
is displayed (0°to 47°).

Flaps Position Symbol Color

Table 5--11

D Spoiler Position -- The ground spoiler position is shown by a boxed

annunciation. Table 5--12 defines the color coding for the spoiler
position.

Spoiler Position Annunciation Color

Closed CLD White (see note)

Open OPN Green (see note)
NOTE: When the following conditions exist, the box and annunciation turn red:
- Any spoilers switch sensed to be open
- Weight on wheels (WOW)
- Thrust lever angle > 60°.

Spoiler Color Coding

Table 5--12

PITCH/ROLL/YAW TRIM DISPLAY

The aircraft surface trim positions are shown in the lower right corner
of the EICAS. The symbols are described below.

D Pitch Trim -- The pitch trim is indicated by digital and analog

displays. The digital display ranges from --10°(full nose down) to +4°
(full nose up). A green band is displayed to mark the takeoff trim
range. The color of the pointer and digital display are described in
table 5--13.

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NOTES: 1. When the pitch trim is > +1°, UP is annunciated

above the digital display.
2. When the pitch trim is < --1°, DN is annunciated
above the digital display.
3. When pitch trim is < --1°, the minus (--) sign is
removed.

Digits and Pointer

Condition Color

Condition A
--Weight--on--wheels (WOW)
Red
--Trim out of takeoff range
--Trust lever angle >60°
Condition B
--Weight--on--wheels (WOW)
Amber
--Trim out of takeoff range
--Trust lever angle <60°
Condition C
Green
--Airborne
NOTE: Digits are also enclosed in a box of the same color.

Pitch Trim Color Description

Table 5--13

D Roll Trim -- An analog roll trim scale displays the current roll trim.
The green trim arrow indicates the trim position. An amber mistrim
arrow is displayed when the autopilot is engaged and the ailerons
must be retrimmed. The mistrim arrow indicates the direction of the
trim required to retrim the aircraft.

D Yaw Trim -- Rudder trim is displayed with an analog scale and a

green pointer. The pointer indicates the current trim position.

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CABIN PRESSURE AND AUXILIARY POWER UNIT (APU)

STATUS

Cabin pressurization and the APU status are displayed in the lower left
corner of the EICAS. The annunciations are described below.

D Cabin Pressurization -- Cabin climb/descent rate is displayed in

feet per minute (fpm). Cabin pressure differential is displayed in PSI.
Cabin altitude is displayed in feet. All displays are cyan. The
resolution of the cabin settings are as follows:
- Climb/descent rate: 50 fpm
- Pressure differential: 0.1 PSI
- Altitude: 100 ft.
When cabin pressure differential exceeds its limits, the pressure
differential digits turn amber in an amber box or red in a red box.

When cabin altitude exceeds 8100 feet, the cabin altitude digits turn
amber in an amber box. WHen cabin altitude exceeds 10,000 feet,
the digits and box turn red.

The cabin altitude rate has a range of from --2000 feet per minute
(fpm) to +2000 fpm in 50--foot increments. It is displayed in green.

Cabin pressure differential (PD) ranges from --0.4 PSID to + 10.00

PSID. The colors of the digital display are described in table 5--14.

Value (PSID) Color

--0.5 to --0.4 Red

--0.3 to --0.1 Amber
0.0 to +7.9 Green
8.0 to 8.3 Amber
> 8.4 Red

Color Differential Pressure Display

Table 5--14

D APU Status -- The APU status consists of the APU turbine speed
and the APU exhaust temperature. The APU annunciation and
labels are white. The digits follow the logic given in tables 5--15 and
5--16 .

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Color Percentage of RPM

Amber < 95
Green 96 -- 104
Amber 104 -- 110
Red 111 <

APU Turbine Speed as a Percentage of RPM

Table 5--15

Color Percentage of RPM

Green < 680 °C

Amber 681°C -- 717°C
Red 718°C <

APU Exhaust Temperature

Table 5--16

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Figure 5--6 shows a typical EICAS display.

AD--50961--R1@

EICAS in the Takeoff Condition (Typical)

Figure 5--6

PRIMARY ENGINE DISPLAYS FAILURE INDICATIONS

D Data Acquisition Unit (DAU) Failure -- A DAU failure is indicated

on the primary engine display by removing the data pointers from all
the arc displays on the failed side. The current value digits are
replaced with amber dashes.

NOTE: When a DAU fails, all the CAS messages supplied by the
DAU are removed until DAU REV is selected.

D Engine Sensor/Interface Failure -- The failure indication of an

individual sensor or interface is a subset of the DAU failure. The
difference is that the failure indication is shown only on the affected
parameter.

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D Engine Instrument Test -- If on the ground, when the No. 1 EFIS

DC test button is pushed, the failure flags are shown. The test
display is shown only as long as the TEST button is pushed. Figure
5--7 shows the gear/flaps/spoiler display, cabin pressurization and
APU status display, and pitch/roll/yaw trim displays. The following
green messages are shown (master caution is not activated when
the messages are caused by the TEST sequence).
- REV
- IGN AB
- ATTCS

AD--50962@

Engine Instrument Display With Failures

Figure 5--7

NOTES: 1. The CAS display goes to its test format during TEST.
2. If metric units are displayed for certain parameters,
metric labels are shown during TEST.

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Crew Alerting System (CAS)

The crew alerting system has the following characteristics:

D 15 message lines maximum
D 18 characters per line maximum
D One message status line
D CAS display declutter through the EICAS bezel rotary knob
D Master warning/caution button acknowledgement of messages
D CAS computed system messages.

The display shown in figure 5--8, is controlled from the No. 1 symbol
generator (normally) that receives warning, caution, and advisory signals
from various subsystems in the aircraft. The symbol generator processes
these aircraft inputs, including multiple combinations, to determine which
CAS message to display. Some CAS messages are the result of several
aircraft inputs, these inputs are computed messages.

0 2

Crew Alerting System on EICAS Display

Figure 5--8

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The CAS display is in the upper right half of the EICAS display. The color
convention and displayed priority for messages on the CAS display are
defined in table 5--17.

Color Message

Red Warnings (top of the message stack)

Amber Cautions (middle of the message stack)

Cyan Advisories and status (bottom of the message

stack)

Message Color Conventions

Table 5--17

Amber and cyan messages can be scrolled off the display with the
EICAS bezel controller knob. The quantity of scrolled messages is
displayed at the bottom of the CAS area.

NOTES: 1. Red (warning) messages cannot be scrolled.

2. No messages can be scrolled if any amber messages
that are displayed have not been acknowledged
through the caution annunciator.
3. The END message cannot be scrolled off the display.

CAS Messages
WARNING MESSAGES (RED)

When a red message becomes active, the following indications are

given:
D Red message displayed flashing on the CAS display
D Red WARN annunciator on the warning/caution annunciator is lit.

When the warning/caution annunciator/pushbutton is pushed on the

warning panel, the red annunciator is extinguished and the red
message is displayed steadily on the CAS display.

If there are red messages already on the CAS display when a new red
message becomes active, the new message is put at the top of the red
message stack. Messages are removed from the CAS display when the
condition causing the message is remedied. Table 5--18 lists the
warning messages.

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APU FIRE D ENG 1 FIRE D

AUTOPILOT FAIL ENG 2 FIRE D
BAGG SMOKE ENG ATTCS FAIL
BATT 1 OVTEMP D FUEL 1 LO LEVEL
BATT 2 OVTEMP D FUEL 2 LO LEVEL
BLD 1 LEAK GPWS D
BLD 2 LEAK ICE COND -- A/I INOP D
BLD 1 OVTEMP LAV SMOKE
BLD 2 OVTEMP LG/LEVER DISAGREE D
BLD APU LEAK MAIN DOOR OPN
E1 ATTCS NO MRGN NO TAKEOFF CONFIG
E2 ATTCS NO MRGN PIT TRIM 1 INOP
E1 LOW N1 PIT TRIM 2 INOP
E2 LOW N1 SPS 1 INOP D
E1 OIL LOW PRESS SPS 2 INOP D
E2 OIL LOW PRESS SERVICE DOOR OPN D
ELEC ESS XFR FAIL D
NOTE: Red messages with a D are not inhibited during takeoff and landing.

Red Warning Messages

Table 5--18

D NO TAKEOFF CONFIG -- This message is displayed when the

aircraft is on the ground, either engine thrust lever is at greater than
60°, and any of the following conditions exist:
- The parking brake is on
- Ground spoilers are deployed
- Pitch trim is greater than --4°or less than --8°
- Flaps are not at either 9°or 22°.
D ELEC ESS XFR FAIL -- The following combination triggers this
message:
- The L air/ground switch indicates airborne and
- Either L or K essential bus contact indicates open or L or R dc
bus indicates off and
- Generators 1 through 4 line contacts indicate open and
- APU generator line contact indicates open.

CAUTION MESSAGES (AMBER)

When an amber message becomes active, the following indications are
given:
D Amber message displayed flashing on the CAS display
D Amber caution annunciator on the warning/caution annunciator is lit.

When the warning/caution annunciator/pushbutton is selected on the

master warning panel, the amber annunciator is extinguished and the
amber message is displayed steadily on the CAS display.

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 PRIMUSr1000 Integrated Avionics System

If there are amber messages already on the CAS display when a new
amber message becomes active, the new message is put at the top of
the amber message stack. If the existing amber or cyan messages are
scrolled of the display, the amber and cyan message stacks are brought
back on the CAS display with the new message displayed at the top of
the amber stack. Messages are removed from the CAS display when
the condition that caused the message is remedied. Table 5--19 list the
caution messages.

Amber messages with a D are not inhibited during takeoff and landing.

115 VAC BUS OFF BLD 2 TEMP

ACCESS DOORS OPEN BLD 2 VLV FAIL
AHRS 1 OVERHEAT BRAKE OVERHEAT D
AHRS 2 OVERHEAT BRAKE DEGRADED
A/ICE SWITCH OFF BRK INBD INOP
AIL SYS 1 INOP BRK OUTBD INOP
AIL SYS 2 INOP CHECK PFD 1
AP AIL MISTRIM CHECK PFD 2
AP ELEV MISTRIM CROSS BLD FAIL
APU BLD VLV FAIL CROSS BLD SW OFF
APU CNTOR CLSD DAU1 A FAIL
APU EXTBTL INOP DAU2 A FAIL
APU FAIL DAU 1 ENG MISCMP
APU FIREDET FAIL DAU 1 SYS MISCMP
APU FUEL LO PRESS DAU 1 WRN MISCMP
APU FUEL SOV INOP DAU 2 ENG MISCMP
APU GEN OFF BUS DAU 2 SYS MISCMP
APU GEN OVLD DAU 2 WRN MISCMP
APU OIL HI TEMP DC BUS 1 OFF
APU OIL LO PRESS DC BUS 2 OFF
AOA 1 HEAT INOP DFDR FAIL
AOA 2 HEAT INOP E1 A/ICE FAIL
AURAL WARN FAIL ELEKBAY OVTEMP
AUTO TRIM FAIL EMERG LT NOT ARMD
AUTOPILOT FAIL EMERG EXIT OPN
BAGGAGE DOOR OPN EMRG BRK LO PRES
BATT 1 OFF BUS ENG 1 REV FAIL
BATT 2 OFF BUS ENG 1 TLA FAIL
BKUP BATT OFF BUS ENG 2 REV FAIL
BLD 1 TEMP ENG 2 TLA FAIL
BLD 1 VLV FAIL ENG NO TO DATA

Amber Caution Messages

Table 5--19 (cont)

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PRIMUSr1000 Integrated Avionics System

E1 ATS SOV OPN IC 2 OVERHEAT

E1 CTL A FAIL IC 2 WOW INOP
E1 CTL B FAIL IC BUS FAIL
E1 EXTBTL A INOP ICE DET 1 FAIL
E1 EXTBTL B INOP ICE DET 2 FAIL
E1 FIREDET FAIL ICE DETECTORS FAIL
E1 FUEL LO PRESS ITT 1 LIMIT FAIL
E1 FUEL LO TEMP ITT 2 LIMIT FAIL
E1 FUEL SOV INOP LG AIR/GND FAIL D
E2 A/I FAIL NO ICE A/ICE ON
E2 ATS SOV OPN OXYGEN LO PRESS
E2 CTL A FAIL OVER WG ICE DET
E2 CTL B FAIL OVER WG ICE INOP
E2 EXTBTL A INOP PACK 1 OVHT
E2 EXTBTL B INOP PACK 1 OVLD
E2 FIREDET FAIL PACK 1 VLV FAIL
E2 FUEL LO PRESS PACK 2 OVHT
E2 FUEL LO TEMP PACK 2 OVLD
E2 FUEL SOV INOP PACK 2 VLV FAIL
ELEC EMERG ABNORM PITOT 1 INOP
ENG REF A/I DISAG PITOT 2 INOP
ESS BUS 1 OFF PITOT 3 INOP
ESS BUS 2 OFF PRESN AUTO FAIL
FLAP FAIL RAM AIR VLV FAIL
FUELING DOOR OPN RUD HDOV PROT FAIL
FUEL IMBALANCE RUDDER OVERBOOST
FUEL TANK LO TEMP RUDDER SYS1 INOP
FUEL XFEED FAIL RUDDER SYS2 INOP
GEN 1 OFF BUS SHED BUS 1 OFF
GEN 1 OVLD SHED BUS 2 OFF
GEN 2 OFF BUS SPBK LVR DISAGREE
GEN 2 OVLD SPOILER FAIL
GEN 3 OFF BUS SPS ADVANCED D
GEN 3 OVLD STAB A/ICE FAIL
GEN 4 OFF BUS STEER INOP
GEN 4 OVLD STICK PUSHER FAIL D
GPWS INOP D TAT 1 HEAT INOP
HS VLV 1 FAIL TAT 2 HEAT INOP
HS VLV 2 FAIL WG 1 A/ICE FAIL
HYD SYS 1 FAIL WG 2 A/ICE FAIL
HYD SYS 1 OVHT WG A/ICE ASYMETRY
HYD SYS 2 FAIL WINDSHEAR INOP D
HYD SYS 2 OVHT W/S 1 HEAT FAIL
IC 1 OVERHEAT W/S 2 HEAT FAIL
IC 1 WOW INOP YAW DAMPER FAIL

Amber Caution Messages

Table 5--19

Engine Instrument and Crew Alerting System (EICAS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

COMPUTED CAUTION (AMBER) MESSAGES

D BLD 1/2 TEMP -- This message results from a combination of

engine bleed air valve positions and high cooler outlet temperatures
on No. 1 and No. 2 engines.

D ELEC EMER ABNORM -- This message is annunciated when the

contacts for any of the four generators, the APU generator, or the
GPU generator indicate they are closed but both the L and R
essential contacts also indicate they are closed.

D FUEL IMBALANCE -- Fuel imbalance is displayed when both the

L and R fuel tank quantities indicate that they are valid and the
difference between the two tank quantities is 800 lbs (363 kg). The
message latches until the difference drops to 100 lbs (45 kg).

D PITOT 1--2--3 INOP -- This message is displayed when the aircraft

is on the ground, either engine’s N2 is above 60%, and the
respective pitot status sensor indicates a failure, or, when the pitot
status sensor is valid and either N2 is less than 50%.

D ENG 1/2 A/ICE FAIL -- This message is displayed when the aircraft
is on the ground, either power lever is advanced, and the anti--ice
switch, engine bleed air pressure sensor, or engine anti--ice bleed
valve switch are not properly set. It is also displayed at altitudes less
than 23,000 feet, at thrust lever angles less than 43°, and when the
engine anti--ice indicates low pressure.

D ENG 1--2 REV DISAGREE -- This message is displayed with the

aircraft on the ground. It indicates a disagreement between the
sensor or the thrust reverser being deployed and the hydraulic
pressure monitors for the reverse thrust.

D BRAKE DEGRADED -- This message is displayed when either the

inboard or outboard brake indicates a degraded condition and either
main gear indicates down and locked.

A28--1146--112--00 Engine Instrument and Crew Alerting System (EICAS)

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PRIMUSr1000 Integrated Avionics System

ADVISORY MESSAGES (CYAN)

The cyan message flashes on the CAS display for 5 seconds, and then
goes steady. No pilot interaction is required to acknowledge this type
of message.

If there are cyan messages already on the CAS display when a new
cyan message becomes active, the new message is displayed at the
top of the cyan message stack. If the existing amber or cyan messages
are scrolled off the display, the cyan message stack is brought back on
the CAS display with the new message displayed a the top of the stack.
Cyan messages are removed from the CAS display when the condition
that caused the message is remedied. The cyan advisory messages
are listed in table 5--20.

All cyan messages are inhibited for takeoff and landing.

AHRS1 BASIC MODE E1 OIL IMP BYP

AHRS2 BASIC MODE E2 ADC DATA FAIL
APU FUEL SOV CLSD E2 FADEC FAULT
BLD 1 VLV CLSD E2 FUEL IMP BYP
BLD 2 VLV CLSD E2 FUELSOV CLSD
CHECKLIST MISMATCH E2 HYD PUMP FAIL
CMC FAIL E2 HYD SOV CLSD
CROSS BLD OPEN E2 IDL STP FAIL
DAU 1 B FAIL E2 OIL IMP BYP
DAU 1 REVERSION FLAP LOW SPEED
DAU 2 B FAIL FUEL XFEED OPEN
DAU 2 REVERSION GEN 1 BRG FAIL
DU 1 FAN FAIL GEN 2 BRG FAIL
DU 2 FAN FAIL GEN 3 BRG FAIL
DU 3 FAN FAIL GEN 4 BRG FAIL
DU 4 FAN FAIL HYD1 LO QTY
DU 5 FAN FAIL HYD EPUMP FAIL
E1 ADC DATA FAIL HYD PUMP SELEC OFF
E1 FADEC FAULT IC 1 FAN FAIL
E1 FUEL IMP BYP IC 2 FAN FAIL
E1 FUELSOV CLSD ICE CONDITION
E1 HYD PUMP FAIL PACK 1 VLV CLSD
E1 HYD SOV CLSD PACK 2 VLV CLSD
E1 IDL STP FAIL

Cyan Advisory Messages

Table 5--20

Engine Instrument and Crew Alerting System (EICAS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

COMPUTED ADVISORY MESSAGES

D E 1--2 IDL STP FAIL -- This message indicates that the flight idle
stop has failed regardless of whether the aircraft is on the ground or
in the air.

D Message Inhibit -- During the takeoff and landing phases of flight,

all red, amber, and cyan messages (except as noted in the
respective message lists) are inhibited from being displayed. The
conditions that activate the inhibit function are described in table
5--21 (all messages must be valid for inhibit to activate).

Condition Parameter Description

CAS Greater than V1 -- 15 knots

Takeoff
L, R WOW Valid
Approach Radio Altimeter Less than 200 feet

Conditions for Inhibit Function

Table 5--21

CAS Display Failure Indications

SYMBOL GENERATOR FAILURE

Under normal conditions, SG1 (in the display guidance computer

(DGC) 1)) drives the EICAS display. If it fails, the CAS messages are
replaced with a red X.

EICAS TEST

When the pilot’s (normally) DC TEST button is pushed while weight is

on wheels and calibrated airspeed is less than 50 knots, the EICAS
display invalid formats shown in figure 5--9 are displayed.

A28--1146--112--00 Engine Instrument and Crew Alerting System (EICAS)

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PRIMUSr1000 Integrated Avionics System

AD--50964@

EICAS Invalid Display Formats

Figure 5--9

EICAS Reversionary Modes

LOSS OF THE DISPLAY UNIT

Setting either pilot’s reversionary control panel MFD knob to EICAS

displays the EICAS on that side’s MFD.

LOSS OF DISPLAY GUIDANCE COMPUTER (DGC) SYMBOL

GENERATOR

This is indicated by a red X on the EICAS display unit. Pushing the

pilot’s SG button on the reversionary control panel selects the other
symbol generator in DGC No. 2.

Engine Instrument and Crew Alerting System (EICAS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

Backup EICAS Display

Either radio management unit (RMU) can be selected to display certain
EICAS data in abbreviated format. Under some conditions, the No. 1
RMU can be automatically selected to display engine data. Normal
operation of the RMU is described in Honeywell Pub. No. 28--1146--050.
This section describes the use of the RMU for the backup EICAS
display.

Normally, the RMU page (PGE) button is used to select the EICAS
displays. Redundant sources of 28 Vdc power are connected to the
RMUs and both DAUs to ensure display and data availability. If neither
DGC is outputting the EICAS display format, the No. 1 RMU
automatically displays the first of two EICAS pages.

NOTES: 1. Manual or automatic selection of the RMU EICAS

display allows access to the first of two pages. Even if
the second page is selected, the first page is always
displayed after 20 seconds of no RMU selections.
2. Returning the RMU to the communication and
navigation function is done by pushing the RMU PGE
button. However, if the No. 1 RMU is displaying EICAS
information due to an automatic selection, that RMU
returns to the EICAS display 20 seconds after the last
pilot selection on the RMU. When the automatic
selection conditions are normal, the No. 1 RMU returns
to the main RMU tuning page by itself.

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PRIMUSr1000 Integrated Avionics System

PAGE 1

Figure 5--10 shows page 1 of the EICAS display on the RMU. It displays
this data:
D N1 engine fan speed analog and digital display
D Digital display of:
- ITT -- interstage turbine temperature
- N2 turbine speed
- FF PPH -- fuel flow in pounds/hr
- OIL P -- oil pressure
- OIL T -- oil temperature spoiler and flap positions.

MAX--TO MAX--TO

102.5 N1 99.9

520 ITT 490

95.0 N2 96.7
850 FF PPH 910
250 OIL P 195
145 OIL T 145
PAGE 2 4 MSGS

TUNE

AD--50965--R1@

Backup EICAS Display on RMU -- Page 1

Figure 5--10

Engine Instrument and Crew Alerting System (EICAS) A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

PAGE 2

Figure 5--11 shows page 2 of the EICAS display on the RMU. It displays
the following data:

D Digital display of:

- FQ LB -- fuel quantity in pounds
- FLAPS -- flap position
- Landing gear down and locked
- EICAS messages.

1200 FQ LB 2100
FLAPS 22

LG DOWN LOCKED
MESSAGE LINE #2
MESSAGE LINE #3
MESSAGE LINE #4
MESSAGE LINE #5
MESSAGE LINE #6
MESSAGE LINE #7

PAGE 1

TUNE

AD--50966@

Backup EICAS Display on RMU -- Page 2

Figure 5--11

NOTE: When the installation is set to display weight in kilograms, the

fuel quantity on the RMU page 2 is displayed in kg.

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PRIMUSr1000 Integrated Avionics System

Backup EICAS Display Failure Indications

See figure 5--12 for an illustration of the failure indications of the backup
EICAS display.
D Failure of the RMU to display engine data is annunciated with
ENGINE DATA UNAVAILABLE.
D N1 data failure is shown with a red X and removal of the pointer.
D Digital data failure is shown by amber dashes.

AD--50967@

RMU Backup Engine Display Failure Indication

Figure 5--12

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 PRIMUSr1000 Integrated Avionics System

6. Flight Guidance System (FGS)

GUIDANCE CONTROLLER (GC)

The GC shown in figure 6--1, is used to operate the functions listed
below.

FD1 HDG NAV AP SPD FLC ALT FD2

CRS 1 HDG ASEL CRS 2

SPD
APR CPL VS

BNK YD
PUSH SYNC PUSH SYNC PUSH SYNC

AD--50968@

Guidance Controller
Figure 6--1
D Engages and disengages the autopilot and yaw damper.
D Couples the left or right flight director to the single autopilot.
D Selects lateral or vertical flight modes for left and right side flight
directors
D Is used to enter the following:
- Altitude preselect value
- Airspeed or vertical speed targets
- Heading select value
- Selected course value.

When the autopilot or yaw damper are selected, they are annunciated
on the GC and the primary flight display (PFD). When the flight director
mode is selected, they are also displayed on the GC and PFD. In most
cases, when a button is pushed it is annunciated with a lighted green
bar on the selected button.

Autopilot (AP)/Yaw Damper (YD)/Couple (CPL)

Buttons
D AP Button -- The AP button is used to engage the autopilot and yaw
damper simultaneously. When engaged, the button lights. The
autopilot can be engaged with the aircraft in any reasonable attitude.
With no flight director modes selected, engaging the autopilot rolls
the aircraft to a wings level attitude and holds the existing pitch
attitude. If any flight director modes are selected before the autopilot
is engaged, the autopilot automatically couples to the selected flight
director mode.

A28--1146--112--00 Flight Guidance System (FGS)

6-1
PRIMUSr1000 Integrated Avionics System

D AP Disengage -- Normally, the autopilot is disengaged by pushing

the AP disconnect button located on the outboard side of each
control wheel. The autopilot can also be disengaged by any one of
the following:
- Pushing the AP or YD button on the GC
- Operating the pilot’s or copilot’s manual elevator trim switch or
backup elevator trim.

D CPL Button -- The CPL button is used to couple the left or right side
HSI data, and that side’s flight director computations to the autopilot.
The system powers up coupled to the left side.

NOTES: 1. If the selected navigation source is not the same on

both PFDs, and the FD NAV mode is selected, the
command bar on the uncoupled side is removed.
2. Under some conditions, particularly when the FD
NAV mode is selected, the command bars on each
side may not show the capture of a navigation
course at the same time. This causes a split in the
right and left command bars.

When the system is transferred to the cross--side HSI, all flight director
modes are cancelled. Operating modes must be reselected. The
pointer on the right or left side of the CPL button lights to indicate which
HSI are selected. A couple arrow is also displayed on both PFDs.

NOTE: The arm or capture status of the flight director modes are
only annunciated on the PFD.

D YD Button -- The YD button is used to select only yaw damper.

D YD Disengage -- The yaw damper can be disengaged by any one

of the following:
- Pushing the YD button on the GC
- Pushing the AP disconnect button on the outboard side of each
of the control wheels.

NOTE: The autopilot and yaw damper are automatically

disengaged when either aircraft stall protection system
detects an actual or stick shaker stall condition.

Flight Guidance System (FGS) A28--1146--112--00

6-2
 PRIMUSr1000 Integrated Avionics System

Flight Director (FD) Mode Selection

D HDG (Heading) Button -- When the HDG button is pushed, the
flight director computer is commanded to follow the inputs from the
heading bug on the selected HSI. The command bars on the PFD
are driven to follow the position of the heading bug. While in the
heading mode, a lower bank limit can be selected with the BNK
(Bank) button on the GC.

D NAV (Navigation) Button -- Pushing the NAV button directs the

flight guidance computer to arm, capture and track the selected
navigation signal sources (VOR, LOC, BC, LNAV, or FMS).

D APR (Approach) Button -- Pushing the APR button selects the

gains to arm and capture the lateral deviation signal for VOR, LOC,
and BC sources, and selects both lateral and vertical navigation
signals for the ILS to meet approach criteria. Deselection of APR
with LOC source selected removes the vertical navigation signals.

D ALT (Altitude) Button -- Pushing the ALT button commands the

system to hold the current altitude.

D SPD (Airspeed) Button -- When the SPD button is pushed, the pilot
can use the SPD knob or touch control steering (TCS) function to
select an airspeed value for the aircraft to hold during a climb or
descent.

Pushing the SPD knob sets the speed reference in knots or Mach.
The value and airspeed bug is displayed on the airspeed tape on the
PFD.

NOTE: The SPD function is inoperative when FLC is selected.

D FLC (Flight Level Change) Button -- Pushing the FLC button

commands the system to maintain a predetermined airspeed profile
as the aircraft climbs or a predetermined vertical speed profile as the
aircraft descends.

NOTE: The pilot cannot set a speed target, or a vertical airspeed

target with the guidance controller SPD knob when FLC
is selected.

D VS (Vertical Speed) Button -- This button selects the system to

maintain the current vertical speed and to select and maintain a new
vertical speed using either the SPD knob or by using the TCS button.
The vertical speed target is displayed on the PFD.

A28--1146--112--00 Flight Guidance System (FGS)

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PRIMUSr1000 Integrated Avionics System

D FD1/FD2 Buttons -- The FD1/FD2 buttons are used to select and

deselect the display of on--side flight director command bars as
described in table 6--1.

Condition
Effect
Autopilot FD Command Bars
AP disengaged Out of view (no FD
modes selected)
In view (both sides) Removes only the
on--side command bars
On--side only Disengages all flight
director modes
AP engaged Out of view -- Cross--side N/A
In view -- Coupled side
(coupled side always in
view)
In view -- Cross--side Can be selected OFF
NOTES: 1. If CAT II is to be enabled, both PFD’s command bars must be on.
2. When the FD1/2 switches are selected, they are annunciated by the
display of the command bars, lights on the GC switches, and in some
cases, by the flight director mode annunciations.

FD1/FD2 Switch Operation

Table 6--1

D CRS (Course) Knob -- When the navigation mode is VOR, LOC, or

BC, the CRS knob is used to set the course on the PFD HSI. When
in VOR, pushing the PUSH SYNC button on the CRS knob points
the HSI’s course pointer to the VOR station with the course deviation
centered.

D HDG Knob -- The HDG knob is used to set the heading bug on both
HSIs, and on the multifunction display (MFD) when the MAP mode
is displayed. Pushing the PUSH SYNC button on the HDG knob
synchronizes the selected heading to the current aircraft heading.

D YD OFF -- This annunciator lights when the flight guidance system

yaw damper is disconnected by the FGS.

D AP OFF -- This annunciator lights when the autopilot has been

manually disengaged or it has disengaged itself because of an
internal malfunction, or by the AHRS, or when the DG slew has been
engaged.

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 PRIMUSr1000 Integrated Avionics System

Remote Autopilot Switches and Annunciators

D Pitch Wheel and Turn Knob Controls -- Refer to the autopilot
controller section (page 6--6), for a description of the PITCH wheel
and TURN knob controls.

D AP Disconnect Switch -- The AP disconnect switch is mounted on

the control wheel. When pushed, it disconnects the autopilot and
yaw damper.

D TCS Button -- The TCS button is mounted on the control wheel.

When pushed, the pilot can change aircraft attitude, altitude,
airspeed, and/or vertical speed, without permanently disengaging
the autopilot.

D GA (Go--Around) Button -- The go--around button is located on the

throttle. Under normal conditions, if the autopilot is engaged, it
remains engaged and the AFCS commands coupled go--around
minimum speed hold function for the command bars. If only the flight
director is available, a fixed pitch of 8°nose--up is commanded.

Autopilot Power--Up Test

When the autopilot is powered up, it performs a self--test on the servo
circuits, and the autopilot control and disconnect paths.

FGS Caution and Advisory Messages

The engine instrument and crew alerting system (EICAS) displays
messages that relate to the FGS. Table 6--2 contains the red warning
FGS messages that are shown on the EICAS. Table 6--3 contains all
the FGS amber caution messages.

Message Description

AUTOPILOT FAIL Autopilot has disengaged and

indicates failed.
PIT TRIM 1 INOP Main pitch trim is inoperative.
PIT TRIM 2 INOP Backup pitch trim is inoperative.

Red (Warning) Flight Guidance System Messages

Table 6--2

A28--1146--112--00 Flight Guidance System (FGS)

6-5
PRIMUSr1000 Integrated Avionics System

Message Description

AP AIL MISTRIM Autopilot has detected the aileron out

of trim.
AP ELEV MISTRIM Autopilot has detected the elevator out
of trim.
AUTO TRIM FAIL Autopilot elevator trim has failed.
AUTOPILOT FAIL The autopilot has failed.
YAW DAMPER FAIL Autopilot has detected that the yaw
damper has failed.

Amber (Caution) Flight Guidance System Messages

Table 6--3

AUTOPILOT CONTROLLER
The autopilot controller shown in figure 6--2, is used to manually control
the autopilot through the PITCH wheel and TURN knob.

TURN
DESCEND

P
I
T
C
H
CLIMB

AD--50485@

Autopilot Controller
Figure 6--2

PITCH Wheel
Rotating the PITCH wheel with the autopilot engaged changes the pitch
attitude proportional to the rotation of the wheel and in the direction of
wheel movement. Movement of the PITCH wheel cancels only altitude
hold or altitude preselect capture.

Flight Guidance System (FGS) A28--1146--112--00

6-6
 PRIMUSr1000 Integrated Avionics System

TURN Knob
The TURN knob must be in detent (center position) before the autopilot
can be engaged. Rotating the TURN knob out of detent commands a
roll. The roll angle is proportional to and in the direction of TURN knob
rotation. Rotating the TURN knob cancels any other previously
selected flight director lateral mode. When the knob is returned to the
detent, the heading hold mode is entered.

A28--1146--112--00 Flight Guidance System (FGS)

6-7/(6-8 blank)
 PRIMUSr1000 Integrated Avionics System

7. System Limits

GLOSSARY OF TERMS
This section describes the flight guidance system (FGS) major
functions that are referred to throughout Section VIII, Modes of
Operation. These functions are listed alphabetically below:
D Attitude Director Indicator (ADI) Command Cue
D Glideslope (GS) Gain Programming as a Radio Altitude (RA) Function
D GS Gain Programming as a RA and Vertical Speed (VS) Function
D GS Capture (CAP)
D Lateral Beam Sensor (LBS)
D Lateral Gain Programming
D Localizer/Back Course (BC) Track
D Navigation On Course (NOC)
D True Airspeed (TAS) Gain Programming
D Vertical Beam Sensor (VBS)
D VOR Capture (CAP)
D VOR Over Station Sensor (OSS)
D VOR After Over Station Sensor (AOSS).

Attitude Director Indicator (ADI) Command Cue

When a command signal is applied to the cue input, the cue (either
crosspointer or single cue) moves left or right (roll), or up or down
(pitch). This gives the required visual command so the pilot can
maneuver the aircraft in the proper direction to reach the flightpath.

If the information required to fly the flightpath becomes invalid, the

appropriate flight director mode will be cancelled.

A28--1146--112--00 System Limits

7-1
PRIMUSr1000 Integrated Avionics System

Glideslope (GS) Gain Programming as a Radio

Altitude (RA) Function

Gain programming starts after the VBS trips. The gain is programmed
as a function of RA. If the RA altimeter is invalid, gain programming
occurs at GS capture and is controlled by a runway height estimator.
The value estimated is a function of GS capture, GS track, and middle
marker. At GS capture, the height is estimated at 1200 feet. At GS track,
and the middle marker is passed, the height is 250 feet. If the micro air
data computer (MADC) is not valid, or the true air temperature (TAT)
probe fails, airspeed is preprogrammed at 120 knots.

GS Gain Programming as a RA and Vertical Speed

(VS) Function

Gain programming starts after the VBS trips. The gain is programmed
as a function of RA and vertical speed VS. If the radio altimeter is
invalid, gain programming occurs at GS CAP and is controlled as a
function of VS, TAS, and time.

GS Capture (CAP)

The following conditions are required for GS CAP:

D Glideslope is valid
D Glideslope is armed
D The localizer mode is captured
D Glideslope deviation is less than one dot.
D Either of the following are satisfied:
- The VBS is tripped
- GS deviation is less than 20 mV.

Lateral Beam Sensor (LBS)

When flying to intercept the VOR, LOC, or BC beam, the LBS is tripped
as a function of beam deviation, course error, TAS, and distance
measuring equipment (DME), if DME is available and not on HOLD. In
the VOR or LOC mode, the course error is compared with the beam
deviation signal and rate of crossing the beam to determine the LBS trip
point.

System Limits A28--1146--112--00

7-2
 PRIMUSr1000 Integrated Avionics System

When the LBS trips, the flight director commands a turn toward the
desired VOR radial or runway at the optimum point to capture the beam.
If the intercept angle to the beam center is very shallow, the LBS does
not trip until the aircraft is near beam center. For this reason, an override
on the LBS occurs when the beam deviation reaches the specified
minimum. The minimum beam sensor trip point for the VOR mode is
±30 mV (»1/2 dot). In the LOC mode, the minimum trip point is ±60 mV.
(»2/3 dot) The maximum LBS trip point is ±175 mV (»2 1/3 dots) for
VOR and ±195 mV (»2 1/2 dots) for LOC.

Lateral Gain Programming

Lateral gain programming is used in both the VOR and localizer modes
of operation. In the VOR mode, gain programming is computed as a
function of DME distance to the station (if available) and TAS. This
adjusts for the aircraft moving toward or away from the station.

In the localizer mode, gain programming is required to adjust for the

aircraft closing on the localizer antenna and beam convergence caused
by the directional properties of the localizer antenna. In the localizer
mode, lateral gain programming is controlled by the change in radio
altitude (RA) with the aircraft below 2400 feet RA when the RA is valid.
If the RA is not valid, then gain programming starts as a function of the
localizer beam capture, TAS, and time.

Localizer/Back Course (BC) Track

Localizer and BC track signifies that the aircraft is on beam center and
crosswind correction can take place. The track phase occurs when the
following conditions are satisfied:
D LOC or BC is captured plus 30 seconds
D Localizer beam rate is less than 30 feet/second
D Localizer beam deviation is less than 20 mV
D Aircraft bank angle is less than 6°.

True Airspeed (TAS) Gain Programming

TAS gain programming is used to program heading select/track error,
course select error, PITCH wheel command, air data commands, and
glideslope deviation to maintain the same aircraft response regardless
of the aircraft’s airspeed and altitude. The TAS computation is derived
from airspeed and altitude.

A28--1146--112--00 System Limits

7-3
PRIMUSr1000 Integrated Avionics System

Vertical Beam Sensor (VBS)

The VBS determines the point of glideslope capture using a number of
inputs. The VBS is armed when the NAV radio is tuned to a LOC
frequency, the LOC receiver is valid, and the LBS is tripped. The VBS
trips as a function of VS, TAS, and glideslope deviation. The VBS trips
when deviation is less than 150 mV and a capture sensor is satisfied.
The capture sensor combines airspeed, rate of change of beam
deviation, and acceleration to determine the best capture point. In the
event the aircraft is paralleling the beam (i.e., no beam closure rate),
the VBS trips at a vertical deviation less than 36 µA. This resets the
previously selected pitch mode and changes aircraft attitude to capture
the glideslope beam.

VOR Capture (CAP)

VOR CAP occurs when the following conditions are satisfied:
D The VOR mode has been armed plus 3 seconds of elapsed time (ET)
D The LBS has tripped.

VOR Over Station Sensor (OSS)

The OSS is used to detect the erratic radio signal encountered in the
area above the VOR transmitter. When these radio signals reach a
certain level of deviation, they are no longer useful and the OSS
eliminates them from the control signal.
OSS is computed when, with DME valid, a 30_VOR station zone of
confusion is assumed and calculated, using barometric altitude for
VOR navigation. During VOR approach, the formula computes the
zone with radio altitude.
NOTE: If the radio altimeter is invalid, the computer assumes a 2500
foot altitude for VOR approach computations.
With DME invalid:
D VOR track has occurred plus 3 seconds of elapsed time
D Lateral deviation is greater than 75 mV and the rate of deviation is
greater than 8 mV per second and the DME is not present.

VOR After Over Station Sensor (AOSS)

During the period when neither the OSS control function, nor the NOC
control function are notified, the system operates in AOSS. AOSS
permits large bank angles so that prior to track controlling the aircraft,
a larger course correction can be made.

System Limits A28--1146--112--00

7-4
 PRIMUSr1000 Integrated Avionics System

SYSTEM PERFORMANCE AND OPERATING LIMITS

Table 7--1 lists the system performance limits for the aircraft. Table 7--2
defines the units of measure and their data range for this aircraft.

Control or
Mode Sensor Parameter Value

Yaw Damper Yaw Engage Engage Limit No limit

Autopilot A/P Engage Engage Limit Roll: Up to ±45_

Engage Pitch: Up to ±30_

Basic Touch Control Roll Control Limit Roll: Up to ±45_

Autopilot Steering (TCS) Pitch Control Limit Pitch: Up to ±30_

TURN Knob Roll Angle Limit ±30_

Roll Rate Limit 3_/sec.
PITCH Wheel Pitch Angle Limit ±15_Pitch

Heading Hold Roll Angle Limit ±30°/sec

Roll Rate Limit 5.5_/sec.

Heading Heading Knob Roll Angle Limit ±27_

Select Low Bank ±14_
Roll Rate Limit ±4.0_/sec.

Capture
VOR, Course Knob, Beam Angle Up to ±90_
VORAPR, or NAV Receiver, Intercept (HDG
FMS DME Receiver, SEL)
FMS
Roll Angle Limit ±27.5_

Roll Rate Limit 5.5_/sec VORAPR

3_/sec VOR
Course Cut Limit 3_/sec VOR or FMS
at Capture ±45_Course

Capture Point Function of beam, beam

rate, course error, DME
distance. Maximum trip
point 175 µA
Minimum trip point
30 µA

System Performance/Operating Limits

Table 7--1 (cont)

A28--1146--112--00 System Limits

7-5
PRIMUSr1000 Integrated Avionics System

Control or
Mode Sensor Parameter Value

On Course
VOR, Roll Angle Limit ±12_
VORAPR, or
FMS (cont) Roll Rate Limit 1_/sec
Crosswind Correction Up to ±45_Course Error
in VOR, ±30_in VAPP
Over Station
Course Change Up to ±90_
Roll Angle Limit ±7_of Roll

Roll Rate Limit 3_/sec

LOC Capture
LOC, APR, or Course Knob Beam Intercept Up to±90_
BC and Nav
Receiver and Roll Angle Limit ±27.5_
Radio Altimeter
Roll Rate Limit 5.5_/sec

Capture Point Function of Beam, Beam

Rate and Course Error
Max. trip point 175 µA
Min. trip point 60 µA
LOC Track
Roll Angle Limit ±17_roll

Crosswind Correction ±30_of course error

Limit

Gain Programming Starts at 1200 ft. radio

altitude gain reduction =
1 to 0.5

GS Capture
GS GS Receiver, Air Beam Capture Function of beam and
Data Computer beam rate
and Radio
Altimeter or MLS Pitch Command Limit ±10_
Receiver
Path Damping Vertical Acceleration
f (TAS)
Pitch Rate Limit Gain Starts at 1200 ft radio
Programming (First altitude, gain reduction
Stage) = 1 to 0.33

Gain Programming Starts at 250 ft. radio

(Second Stage) altitude, gain reduction =
0.33 at 250 ft.= 0 at 0 ft.

GA Control Switch Fixed Pitch--up 12_Pitch up

on Throttle Command,
Wings Level

System Performance/Operating Limits

Table 7--1 (cont)

System Limits A28--1146--112--00

7-6
 PRIMUSr1000 Integrated Avionics System

Control or
Mode Sensor Parameter Value

Pitch Sync TCS Switch on Pitch Attitude ±20_max.

Wheel Command

ALT Hold Micro Air Data ALT Hold Engage 0 to 37,000 ft

Computer Range

ALT Hold Engage ±20 ft

Error

Pitch Limit ±20_

Pitch Rate Limit f (TAS)

VS Hold Micro Air Data VERT Speed Engage 0 to ±6,000 ft/min

Computer, Range
Accelerometer
VERT Speed Hold ±30 ft/min
Engage Error
Pitch Limit (see note) ±20_

Pitch Rate Limit (see f (TAS) 0.3 g’s max

note)

SPD Hold, Micro Air Data SPD Engage Range 80 to 330 kts
FLC Computer
(DGC for FLC) SPD Hold Engage ±2 knots
Error
Pitch Limit ±20_

Pitch Rate Limit f (TAS)

ALT Preselect Micro Air Data Preselect Capture 0 to 37,000 ft

Computer/ Range
Altitude
Preselect on Max. VS for Capture ±6000 ft/min.
MFD
Max. Gravitational ±0.275g
Force During Capture
Maneuver
Pitch Limit ±20_

Pitch Rate Limit f (TAS)

NOTE: FLC, on climb (airspeed) and descent (vertical speed) is controlled by the
respective IAS and V/S error and pitch/pitch rate limit values.

System Performance/Operating Limits

Table 7--1

A28--1146--112--00 System Limits

7-7
PRIMUSr1000 Integrated Avionics System

Display Parameter Units Data Range

Altitude Pressure Altitude feet --1000 to 60,000

Baro Corrected feet --1000 to 60,000

Altitude

Altitude Rate ft/min. --20,000 to +20,000

Preselect Altitude feet 0 to 60,000

Baro Set inHg 22 to 32

Baro Set hPa 745 to 1083.5

Airspeed/ Indicated Airspeed Knots 40 to 400

Mach
Mach Mach 0.2 -- 1.0

Vertical Vertical Airspeed ft/min. --6600 to +6600

Airspeed
Vertical Speed ft/min. --00 to +00
Target

Air Data Display Parameters and Ranges

Table 7--2

System Limits A28--1146--112--00

7-8
 PRIMUSr1000 Integrated Avionics System

8. Modes of Operation

HEADING (HDG) HOLD MODE, WINGS LEVEL

The basic lateral mode of the autopilot is heading hold. The heading
hold mode is defined as:
D Autopilot is engaged
D No lateral flight director mode is selected
D Bank angle is less than 6_.

When the three conditions are met, the autopilot rolls the aircraft to a
wings--level attitude. One second after the bank angle is less than 6_,
the flight guidance system (FGS) uses the current heading as the
heading hold reference. The primary flight display (PFD) annunciates
ROL in the lateral flight director mode annunciation position. Figure 8--1
shows the autopilot (AP) and yaw damper (YD) on and the autopilot
altitude hold (ALT) and roll hold (ROL) modes engaged.

30 30

30 30

AD--50971--R1@

Heading Hold Mode

Figure 8--1

A28--1146--112--00 Modes of Operation

8-1
PRIMUSr1000 Integrated Avionics System

ROLL HOLD MODE

The autopilot recognizes the roll hold mode as being operational when:
D No lateral flight director mode is selected, and
D The aircraft bank angle is greater than 6_, but less than 35_, and
D The autopilot is engaged.

When the three conditions are met, the autopilot maintains the
established bank angle. The touch control steering (TCS) button can be
used while the bank angle is being adjusted. A typical roll hold mode
PFD is shown in figure 8--2. The ROL annunciation is shown in the PFD
lateral flight director mode annunciation area during roll hold.

30

30

30

30

280

AD--50972--R1@

Roll Hold Mode

Figure 8--2

Modes of Operation A28--1146--112--00

8-2
 PRIMUSr1000 Integrated Avionics System

HEADING SELECT MODE

The heading select mode is used to maintain a magnetic heading. The
desired heading is set using the HDG knob on the guidance controller
(GC). The mode is selected with the HDG button on the GC and is
annunciated on the HDG button and also by a green HDG on the top
left side of the PFD. In the HDG mode, the integrated avionics computer
(IAC) uses the command bars to indicate a turn to the heading selected
with the heading bug on the horizontal situation indicator (HSI). The
heading select signal is gain programmed as a function of airspeed.
Selecting low bank (BNK) on the GC limits bank angles to 14_.

When turning the heading bug greater than 180°, the aircraft continues
to turn in the same direction as the bug was originally moving.
Therefore, if the bug is turned to the right 270°, the aircraft turns to the
right in a 270°turn instead of turning left 90°.

The the heading select mode is cancelled by any one of the following
actions:
D Capture of any other lateral steering mode
D Selection of symbol generator (SG) reversionary mode
D Selection of go--around (GA)
D Placing TURN knob out of detent with autopilot engaged
D Selection of attitude and heading reference system (AHRS)
reversionary mode
D Pushing the couple (CPL) button on the GC
D Pushing the heading (HDG) button.

NOTES: 1. Above 25,000 feet mean sea level (MSL), the bank
angle in heading select mode is automatically reduced
to 14_. When descending below 24,750 MSL, the bank
angle is returned to the standard bank value of 27_.
2. Each of the flight director modes can be cancelled by
a variety of pilot (or systems) options. Only the most
typical methods are shown here, and for each flight
director mode in this section.

A28--1146--112--00 Modes of Operation

8-3
PRIMUSr1000 Integrated Avionics System

VOR NAVIGATION (NAV) MODE

The VOR mode automatically intercepts, captures, and tracks a
selected VOR radial, using the selected navigation source displayed on
the HSI. The navigation source displayed on the HSI is a function of the
NAV source buttons located on the display controller (DC). Table 8--1
is the procedure used to engage the VOR NAV mode.

Step Procedure

1 Couple either the pilot’s or copilot’s FD to the autopilot with

the CPL switch on the GC.
2 Tune the coupled side navigation receiver to the frequency
of the VOR in use.
3 Select NAV as the navigation source on the coupled side
DC.
4 Set the pointer on the coupled side HSI for the course to
be flown.
5 Set the heading bug on the HSI to the intercept heading for
the selected course.
6 Push the NAV button on the GC.

VOR Navigation Mode Engagement Procedure

Table 8--1

When the aircraft is outside the normal capture range of the VOR signal,
as shown in figure 8--3, the pilot engages the NAV button on the GC.
The HDG and NAV buttons annunciate. On the PFD, HDG annunciates
in green and VOR ARM in white, as shown in figure 8--4. The IAC is
armed to capture the VOR signal and generates a roll command to fly
the heading select mode.

Modes of Operation A28--1146--112--00

8-4
 PRIMUSr1000 Integrated Avionics System

VOR Navigation Mode Radial Intercept View

Figure 8--3

30 30

9500
30 30

NM

AD--50970--R1@

VOR Navigation Mode Intercept PFD

Figure 8--4

A28--1146--112--00 Modes of Operation

8-5
PRIMUSr1000 Integrated Avionics System

When reaching the lateral beam sensor (LBS) trip point, as shown in
figure 8--5, the system automatically drops the heading select mode
and switches to the VOR capture phase. The following is displayed on
the PFD (see figure 8--6):
D The white VOR annunciator extinguishes
D The green HDG annunciator extinguishes
D A green VOR is annunciated and is enclosed in a white box for 8
seconds to emphasize the capture phase of operation.

The IAC generates the proper roll command to bank the aircraft to
capture and track the selected VOR radial.

When the course select pointer is set on the PFD using the course
(CRS) knob on the GC, the course select error signal is 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 function to the flight control function in the IAC.

The radio signal is routed from the navigation receiver to the IAC. The
IAC lateral gain programs the signal in its flight control function.

The lateral gain programming is performed as a function of distance

measuring equipment (DME), distance to the station (if available), and
true airspeed (TAS). Gain programming adjusts for the aircraft either
coming toward or moving away from the VOR station.

NOTE: Avoid operating in DME hold during VOR capture and

tracking operation if possible. When in DME hold, the
computer cannot use DME distance for gain programming.

VOR Navigation Mode Capture Plan View

Figure 8--5

When flying a VOR intercept, the optimum intercept angle is 45_or less.

Modes of Operation A28--1146--112--00

8-6
 PRIMUSr1000 Integrated Avionics System

30
30

30
30

AD--50973--R1@

VOR Navigation Mode Capture

Figure 8--6

When the aircraft meets the VOR track conditions the course error
signal is removed from the lateral steering command. This leaves radio
deviation and DME gain programming (if available) to track the VOR
signal and to compensate for beam standoff in the presence of a
crosswind. The system automatically compensates for a crosswind of
up to 45_course error. Figure 8--7 shows the PFD display in the tracking
mode.

Cancel the VOR navigation mode with any one of the following actions:
D Push the NAV button on the GC
D Select go--around
D Select another lateral mode
D Change the selected navigation or heading source.

A28--1146--112--00 Modes of Operation

8-7
PRIMUSr1000 Integrated Avionics System

30 30

30 30

AD--50975--R1@

VOR Navigation Mode Tracking

Figure 8--7

When the aircraft approaches a VOR station, it enters a zone of

unstable radio signal, shown in figure 8--8. 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 best to remove
it from the roll command. The over--station sensors (OSS) monitor
when the aircraft enters the zone of confusion and removes radio
deviation from the roll command. The system also uses the co--located
DME signal (if available), to adjust tracking gains.

When over the VOR station, the system can accept and follow a course
change of up to 90_.

Modes of Operation A28--1146--112--00

8-8
 PRIMUSr1000 Integrated Avionics System

ZONE OF CONFUSION

OSS A0SS1 A0SS2

AD--50976@

VOR Navigation Mode -- Tracking Over Station

Figure 8--8

VOR APPROACH (VAPP) MODE

A VOR approach procedure is described in table 8--2. The APR button

annunciator on the GC is lit and VAPP is displayed in white on the PFD.
The IAC applies gains appropriate for an approach. When the selected
course is captured, the PFD displays VAPP in green.

NOTE: To get the best results, establish the aircraft on the selected
course at least 3 miles before reaching the VOR station.

Step Procedure

1 Couple the pilot’s or copilot’s flight director to the autopilot

with the CPL switch on the GC.
2 Tune the coupled side navigation receiver to the frequency
of the VOR to be used.
3 Select NAV as the navigation source on the coupled side DC.
4 Set the course pointer on the coupled side HSI for the
course to be flown.

VOR Approach Mode Engagement Procedure

Table 8--2 (cont)

A28--1146--112--00 Modes of Operation

8-9
PRIMUSr1000 Integrated Avionics System

Step Procedure

5 Set the heading bug on the HSI to the intercept heading for
the selected course.
6 Push the APR button on the GC.

VOR Approach Mode Engagement Procedure

Table 8--2

FLIGHT MANAGEMENT SYSTEM (FMS) NAVIGATION

MODE (OPTION)

The FMS mode is flown when the pilot selects FMS on the DC. The
FMS mode is flown as described in table 8--3.

Step Procedure

1 Couple either the pilot’s or copilot’s flight director to the

autopilot with the CPL switch on the GC.
2 Enter the data required for navigation into the FMS.
3 Select FMS as the navigation source on the coupled side
DC.
4 Push the NAV button on the GC.

FMS Mode Engagement Procedure

Table 8--3

Characteristics common to FMS are described below.

D Instead of using course error and radio deviation from the SG, a
composite lateral steering command is used from the FMS
navigation computer through the IAC symbol generator function.

D The IAC supplies the flight control function with required steering
commands.

D This lateral steering command is lateral gain programmed in the

FMS navigation computer, and therefore, is not gain programmed
again in the IAC.

Modes of Operation A28--1146--112--00

8-10
 PRIMUSr1000 Integrated Avionics System

Pushing the flight director NAV mode select button annunciates the
FMS mode in green on the PFD as shown in figure 8--9. The flight
control function of the IAC directs the aircraft to a desired track
intercept.

The FMS mode is cancelled by any one of the following:

D Pushing the NAV button on the GC
D Selecting the HDG mode
D Selecting go--around
D Selecting another navigation or heading source.

30 30

30 30

AD--50977--R1@

Long Range Navigation Mode Display

Figure 8--9

A28--1146--112--00 Modes of Operation

8-11
PRIMUSr1000 Integrated Avionics System

LOCALIZER (LOC) MODE

The localizer mode automatically intercepts, captures, and tracks the

front course localizer beam to line up on the centerline of the runway
in use. Before mode engagement, the pilot performs the procedure
described in table 8--4.

Step Procedure

1 Couple either the pilot’s or copilot’s flight director to the

autopilot using the CPL button the the GC.
2 Tune the coupled side navigation receiver to the published
front course localizer frequency for the runway in use.
3 Set the course pointer on the coupled side HSI for the
inbound published localizer course.
4 Set the heading bug on the HSI for the heading to perform
a course intercept, using the GC HDG knob.
5 Push the NAV button on the coupled side DC to select
NAV as the navigation source.
6 Push the NAV button on the GC. The system annunciates
HDG in green and LOC in white.
NOTE: The APR button can also be in this step. However, when pushing
the APR button, and assuming an invalid glideslope, the NAV
button remains the only button lit. If APR is used, and the glides-
lope is valid, the FGS captures the GS at the proper time.

Localizer Mode Engagement Procedure

Table 8--4

The HSI displays the relative position of the aircraft to the center of the
localizer beam and the inbound course, as shown in figure 8--10. With
the heading bug set for course intercept, the heading select mode is
used to perform the intercept. Outside the normal capture range of the
localizer signal, pushing the NAV button on the GC annunciates LOC
on the PFD, as shown in figure 8--10.
D LOC is in white, HDG is green.

D The NAV button on the mode selector also annunciates.

D The aircraft flies the selected heading to the intercept point and the
system is armed for automatic localizer beam capture.

Modes of Operation A28--1146--112--00

8-12
 PRIMUSr1000 Integrated Avionics System

M 30 30

20 20

30 30
--2000

AD--50980--R1@

Localizer Mode Arm Display

Figure 8--10

With the aircraft approaching the selected course intercept, as shown

in figure 8--11, the LBS monitors localizer beam deviation, beam rate,
and TAS. At the computed time, the lateral beam sensor (LBS) trips and
captures localizer signal. The flight director drops the heading select
mode and generates the proper roll command to bank the aircraft
toward the localizer beam center. When the LBS trips, the PFD displays
LOC in green, as shown in figure 8--12. LOC is enclosed in a white box
for 5 seconds.

A28--1146--112--00 Modes of Operation

8-13
PRIMUSr1000 Integrated Avionics System

Localizer Mode Capture Plan View

Figure 8--11

NOTES: 1. When flying a localizer intercept, the optimum

intercept angle is 45_.
2. The pilot should not intercept the localizer at angles
greater than 90_. At angles greater than 90_, the
system automatically captures the back course.

When the course select pointer is set on the coupled side HSI using the
CRS knob on the GC, the course select error signal is established. This
signal represents the difference between actual aircraft heading and
the selected aircraft course.

The course error signal is routed from the IAC’s SG to the IAC’s flight
control function.

The radio deviation signal is routed from the navigation receiver to the
electronic flight instrument system (EFIS), and then to the flight control
function where the signal is lateral gain programmed.

Modes of Operation A28--1146--112--00

8-14
 PRIMUSr1000 Integrated Avionics System

30
30

30
30
CRS

Localizer Capture Display

Figure 8--12

Lateral gain programming is required to adjust the gain applied to the

localizer signal due to the aircraft approaching the localizer antenna,
and beam convergence caused by the directional qualities of the
antenna.

The localizer mode is cancelled with any one of the following actions:
D Pushing the NAV button on the GC
D Changing the displaced heading source or NAV source
D Selecting go--around
D Selecting the heading select mode
D Pushing the CPL button on the GC
D Deselecting an SG reversionary selection.

A28--1146--112--00 Modes of Operation

8-15
PRIMUSr1000 Integrated Avionics System

When the system acquires the localizer capture conditions, the course
error signal is removed from the lateral calculations. This leaves
localizer gain programming to track the localizer signal, as shown in
figure 8--13, and to compensate for localizer beam standoff in the
presence of a crosswind. The system automatically compensates for
a crosswind of up to 45_course error. Figure 8--14 shows the PFD in
the localizer tracking mode.

Localizer Mode Tracking Profile View

Figure 8--13

30 30

20 20

60
40

20 20

30 30 IN

AD--50983--R1@

Localizer Tracking Display

Figure 8--14

Modes of Operation A28--1146--112--00

8-16
 PRIMUSr1000 Integrated Avionics System

BACK COURSE (BC) MODE

The BC mode automatically intercepts, captures, and tracks the back

course localizer signal, as shown in figure 8--15. When flying a back
course localizer approach, glideslope capture is automatically inhibited.
The BC mode is set up and flown similar to a front course localizer
approach. The BC procedure is described in table 8--5.

Back Course Mode Intercept Plan View

Figure 8--15

Step Procedure

1 Couple either the pilot’s or copilot’s flight director to the

autopilot using the CPL button on the GC.
2 Tune the coupled side navigation receiver to the published
front course localizer frequency for the runway in use.
3 Set the course pointer on the coupled side HSI for the
inbound published localizer course.
4 Set the heading bug to the course intercept heading.
5 Push the NAV on the coupled side DC.

Back Course Flight Procedure

Table 8--5 (cont)

A28--1146--112--00 Modes of Operation

8-17
PRIMUSr1000 Integrated Avionics System

Step Procedure

6 Push the NAV button on the GC. With the aircraft outside
the normal back course localizer capture limits, and more
than 105_off the localizer front course heading, the PFD
annunciates BC in white and HDG in green, as shown in
the display below.

M 30 30

20 20

30 30
--2000

AD--50984--R1@

7 At back course capture, the PFD annunciates BC in green,

as shown in the display below. BC is enclosed in a white
box for 8 seconds.

M 30

30

30

30
--2000
3
2
1

1
2
3

AD--50986--R1@

Back Course Flight Procedure

Table 8--5

Modes of Operation A28--1146--112--00

8-18
 PRIMUSr1000 Integrated Avionics System

When the back course mode is selected by pushing the NAV button on
the mode selector with ILS displayed on EFIS, and the selected course
being more than 105_off the aircraft heading, logic in the IAC’s flight
control function establishes internal reverse polarity of the course error
and localizer signals. A gain change takes place in the computer when
BC is selected because the aircraft is closer to the localizer antenna by
the length of the runway plus 1000 feet.

At back course capture, the IAC flight control function generates a roll
command to capture and track the back course localizer signal, as
shown in figure 8--16.

The BC mode is cancelled by any one of the following:

D Pushing NAV on the GC
D Selecting the heading or go--around modes
D Changing the displayed navigation or heading source.

30 30

AP

30 30

3
2
1

1
2
3

AD--50987--R1@

Back Course Tracking Display

Figure 8--16

A28--1146--112--00 Modes of Operation

8-19
PRIMUSr1000 Integrated Avionics System

INSTRUMENT LANDING SYSTEM (ILS) APPROACH

MODE
The ILS approach mode automatically intercepts, captures, and tracks
the front course localizer and glideslope signals, as shown in figure
8--17, so the pilot can fly a fully coupled ILS approach. The mode is set
up and flown similar to the localizer mode. The mode is interlocked so
that the glideslope capture is inhibited until the localizer is captured. As
with the localizer mode, heading select is used to initiate the localizer
approach intercept. The procedure for ILS approach is described in
table 8--6.

ILS Approach Localizer Intercept

Figure 8--17

Modes of Operation A28--1146--112--00

8-20
 PRIMUSr1000 Integrated Avionics System

Step Procedure

1 Couple either the pilot’s or copilot’s flight director to the

autopilot using the CPL switch on the GC.
2 Tune the coupled side navigation receiver to the published
ILS frequency for the runway in use.
3 Push the NAV button on the coupled side DC to select ILS
as the navigation source.
4 Set the radio altitude (RA) minimum altitude (MIN) on the
PFD with the RA knob on the DC.
5 Set the course pointer and heading bug on the coupled
side PFD for localizer intercept, using the GC HDG and
CRS knobs.
6 Push the APR button on the GC. The PFD annunciates, as
shown in the display below.
D LOC in white
D GS in white
D HDG in green.

M
30 30

30 30

AD--50990--R1@

ILS Approach Mode Procedure

Table 8--6 (cont)

A28--1146--112--00 Modes of Operation

8-21
PRIMUSr1000 Integrated Avionics System

Step Procedure
7 With the localizer captured, and outside the normal
glideslope capture limits, the PFD annunciates, as shown
in the display below.
D LOC in green
D GS in white

30
M

30

30
30
--2000

AD--50991--R1@

ILS Approach Mode Procedure

Table 8--6

The glideslope portion of the approach mode is used to automatically

intercept, capture, and track the glideslope beam. The beam is used to
guide the aircraft to the runway in a linear descent. Typical glideslope
beam angles vary between 2_and 3_, dependent on local terrain. The
mode is interlocked so that glideslope capture is inhibited until localizer
capture has occurred.

With the localizer captured and outside the normal glideslope capture
limits, the PFD annunciates the following modes:
D LOC in green
D GS in white
D Any other vertical mode in use at this time is also displayed.

Modes of Operation A28--1146--112--00

8-22
 PRIMUSr1000 Integrated Avionics System

As the aircraft approaches the glideslope (GS) beam, as shown in

figure 8--18, the vertical beam sensor (VBS) monitors TAS, vertical
speed, and GS deviation to determine the correct capture point. At
glideslope capture, the computer drops any other vertical mode that
was in use, and automatically generates a pitch command to track the
glideslope beam.

AD--50989@

ILS Approach Mode Glideslope Capture, Profile View

Figure 8--18

The PFD shown in figure 8--19, annunciates:

D LOC in green
D GS in green. The annunciation is enclosed in a white box for 8
seconds.

The NAV and APR mode selector buttons are also annunciated on the GC.

NOTE: The autopilot must be disengaged before reaching 200 feet

(CAT I) above ground level (AGL).

A28--1146--112--00 Modes of Operation

8-23
PRIMUSr1000 Integrated Avionics System

30 30

1500
20 20

30 30

AD--50993--R1@

ILS Approach Mode Tracking PFD

Figure 8--19

Gain programming is performed on the GS signal to compensate for the

aircraft closing on the GS antenna and beam convergence caused by
the directional properties of the antenna, as shown in figure 8--20. GS
programming is normally a computed function of radio altitude (if
available).

If radio altitude is not available, then GS gain programming is a function

of vertical speed (VS), true air speed (TAS), and time.

The approach mode is cancelled by any one of the following actions:

D Pushing the APR or NAV buttons on the GC
D Loss of ILS GS data
D Selecting go--around
D Selecting HDG on the mode selector
D Changing navigation or heading sources.

Modes of Operation A28--1146--112--00

8-24
 PRIMUSr1000 Integrated Avionics System

RADIO ALTITUDE
BETWEEN 1200
AND 1500 FEET
RADIO ALTITUDE
BETWEEN 200
AND 300 FEET

OUTER MIDDLE RUNWAY

MARKER MARKER
4.0--7.0 MILES 3500 FEET AD--50992@

ILS Approach Mode Tracking, Profile View

Figure 8--20

PITCH HOLD MODE

The pitch hold mode is the basic autopilot and flight director vertical
mode. This mode is best described by comparing the operation of the
pitch hold mode with the AP engaged and with the AP not engaged.

Pitch Hold Mode Operation, AP Engaged

Pitch hold mode with AP engaged operates as follows:

D If no vertical FD modes are active, the AP holds the pitch attitude

that existed at the time the AP was engaged. The vertical flight
director mode is annunciated as PIT.

D The pitch reference can be changed by pushing and holding the

TCS (touch control steering) button on the pilot’s or copilot’s control
wheel and manually flying the aircraft to a new pitch attitude
reference. When the TCS button is released, this new reference is
held by the autopilot. The pitch reference can also be changed by
moving the PITCH wheel on the autopilot controller.

Pitch Hold Mode Operation, AP Not Engaged

Pitch hold mode with AP not engaged operates as follows:
D Selecting an FD lateral mode with no active FD vertical mode
displays the command bar. The command represents the aircraft
pitch attitude at the moment the FD lateral mode was selected. The
vertical flight director mode is annunciated as PIT.

D The pitch hold mode is cancelled by manual or automatic selection

of any FD vertical mode.

A28--1146--112--00 Modes of Operation

8-25
PRIMUSr1000 Integrated Avionics System

TAKEOFF MODE
When the aircraft is on the ground, pushing the throttle--mounted
go--around switch sets the PFD command bars to a fixed pitch up
attitude of 12°. The vertical mode annunciation is a green TO, as shown
in figure 8--21.

HDG TO
169M 30 30
80
1500
20 20
60
10 10
40
40 11 20
00

10 10
132 2
124 R 20 5 20
110 1
30 30 30.17 IN

180 CRS
3
VOR1
2

1.8 NM 1

1
2
HDG GSPD 3
180 00 KTS

AD--50921--R3@

Takeoff Mode
Figure 8--21

Modes of Operation A28--1146--112--00

8-26
 PRIMUSr1000 Integrated Avionics System

VERTICAL SPEED HOLD MODE

The vertical speed (VS) hold mode is used to automatically maintain the
aircraft at a pilot--selected vertical speed reference. To initiate the VS
mode, the pilot can maneuver the aircraft to any climb or descent
attitude, establish the vertical speed reference, and engage the VS
mode. Or, with the autopilot engaged, the reference VS can be changed
by using the SPD knob on the GC. With the VS mode engaged, a new
VS value can be engaged by pushing the TCS button on the control
wheel and maneuvering the aircraft to a new VS reference and
releasing the TCS button.

When the VS mode is engaged, as shown in figure 8--22, the following

occurs:
D VS is annunciated in green on the PFD
D The PFD displays the vertical speed target value above the vertical
speed scale in plus or minus feet per minute.

When the VS reference is changed by using the SPD knob on the GC,
the target value changes and the vertical speed reference bug is
repositioned. Actual aircraft VS is displayed on the VS indicator. When
VS is selected, it resets all previously selected vertical modes.

The VS mode is cancelled by any one of the following actions or event:

D Pushing the VS button on the GC
D Selecting another vertical mode
D Selecting go--around
D Coupling to the cross--side FD
D If on the active side, selecting SG reversionary selection.
D Failure of the on--side micro air data computer (MADC).

NOTES: 1. When the VS mode is selected, the vertical speed

reference bug is displayed on the vertical speed display.
2. Aircraft airspeed protection is active in this mode. Refer
to Speed Hold Mode description on the next page.

A28--1146--112--00 Modes of Operation

8-27
PRIMUSr1000 Integrated Avionics System

HDG VS ASEL
200M
240 30AP YD30
8500

20 20
220
10 10
20
200 80 00
80

180 10 10

20 20
160 7500
29.92 IN
30 30
1000
090 CRS
3
VOR2
2

15.0 NM 1

0 1000

1
VOR2
2
HDG GSPD 3
360 225 KTS

AD--50994--R3@

Vertical Speed Hold Mode PFD

Figure 8--22

SPEED HOLD MODE

The speed (SPD) hold mode is engaged by pushing the SPD button on
the GC. The indicated airspeed (IAS)/Mach reference is synchronized
to the IAS/Mach being flown when the mode is engaged. When a new
reference is manually selected using the GC SPD knob, the system
automatically flies the new reference.

The IAS/Mach speed target comes from the coupled side PFD.
Depending on whether the reference is identified as IAS or Mach (input
from the controller), the system flies IAS or the Mach reference.
Switching from IAS to Mach (or Mach to IAS) does not move the
reference, but simply changes the units used on the digital readout on
the PFD. No aircraft maneuvering changes occur when units of
measure are switched. Push the SPD knob on the GC to switch
between the IAS and Mach modes.

Modes of Operation A28--1146--112--00

8-28
 PRIMUSr1000 Integrated Avionics System

The SPD mode is basically an airspeed mode, however, it differs from

a standard IAS or Mach mode in the following ways:

D Although SPD mode (in the long term), tracks the reference airspeed,
short--term emphasis is on vertical speed. This minimizes vertical
speed excursions due to disturbances or large airspeed changes.

D Vertical operations are monitored to prevent the aircraft from flying too
slowly. Below 120 knots, the vertical mode is overridden and is
replaced by SPD hold mode. This is underspeed (MIN SPD)
protection.

D All vertical operations must be monitored to prevent the aircraft from

exceeding Vmo/Mmo. If, however, Vmo/Mmo is exceeded, the flight
guidance system switches to the SPD hold mode. This is overspeed
(MAX SPD) protection.

NOTE: The only annunciation of MIN SPD or MAX SPD

submodes is displayed on the left side of the attitude
director indicator (ADI).

The SPD mode is annunciated on the PFD by a green IAS or MACH at

the vertical capture location, as shown in figure 8--23.

A28--1146--112--00 Modes of Operation

8-29
PRIMUSr1000 Integrated Avionics System

HDG IAS ASEL

200
240 30AP YD30
8500

20 20
220
10 10
20
200 8000
80

180 10 10

20 20
160 7500
29.85 IN
30 30
090 CRS
3
VOR1
2

12.5 NM 1

25 0 1500

1
VOR2
2
HDG GSPD 3
001 225 KTS

AD--50995--R3@

Speed Hold Mode PFD

Figure 8--23

The pilot can maneuver the aircraft without disengaging the mode, by
pushing and holding the TCS button on the control wheel. When TCS
is released, the airspeed target is the current airspeed.

To fly the SPD mode in a climb to a preselected altitude from a straight

and level condition, follow the procedure described in table 8--7.

Step Procedure

1 Set alert altitude more than 250 feet from the current
altitude.
2 Push the SPD button on the GC.

Speed Mode Engagement Procedure

Table 8--7 (cont)

Modes of Operation A28--1146--112--00

8-30
 PRIMUSr1000 Integrated Avionics System

Step Procedure

3 With the autopilot engaged, set the speed reference on the

coupled side PFD to the reference IAS/Mach number, using
the SPD knob.
4 Advance throttle position to attain climb power.

Speed Mode Engagement Procedure

Table 8--7

The aircraft climbs toward the preselected altitude, maintaining the

speed reference. The amount of throttle position change varies the
aircraft rate of climb.
In the SPD mode, all armed pitch flight director modes are valid, but
capture of any armed pitch mode overrides the SPD mode.
NOTE: When the SPD mode is selected, the speed command bug
is displayed on the airspeed display.
The SPD mode is cancelled by any one of the following actions or event:
D Pushing the SPD button on the GC
D Selecting any other vertical mode on, or capture
D Coupling to the cross--side FD.
D Coupled side MADC becomes invalid
D If on the active side, selecting SG reversionary selection.

NOTES: 1. In a climb above 25,100 feet MSL, the SPD reference

automatically switches from IAS to MACH.
2. In a normal descent below 24,900 feet, the SPD
reference automatically switches from MACH to IAS.
3. The FGS cannot fly to an airspeed reference outside
the normal aircraft flight envelope. The FGS limits the
commanded airspeed between 120 knots and the
maximum speed of the aircraft. This feature is armed
in the SPD, FLC, VS, VNAV, and pitch hold modes.

FLIGHT LEVEL CHANGE (FLC) MODE

Pushing the FLC button on the GC selects the preprogrammed
climb/descent values mode. In a climb, FLC is annunciated as CLB and
the preprogrammed airspeed target is maintained. In a descent, FLC
is annunciated as DES and the preprogrammed rate of descent is
maintained. Descend and climb parameters are listed in table 8--8.

A28--1146--112--00 Modes of Operation

8-31
PRIMUSr1000 Integrated Avionics System

Preprogrammed
Direction Altitude Values

Descend All altitudes --2000 fpm

0--10,000 feet 240 knots
10,000 feet to 240 knots to 270
12,000 feet knots linearly
Climb 12,000 feet to
270 knots
17,377 feet
17,377 feet to
0.56 Mach
37,000 feet

Flight Level Change Descend and Climb Parameters

Table 8--8

When the FLC button is pushed, the SPD or V/S modes are not
available. In CLB, the SPD hold mode control laws are used. In DES,
the VS hold control laws are used. When the TCS button is pushed and
held, the pitch and roll autopilot servos are disengaged and the pilot can
maneuver the aircraft. When the TCS button is released, the FGS
returns the aircraft to the preset FLC target.

In FLC, all armed pitch flight director modes are allowed, but if any
armed pitch mode is captured, the FLC mode is overridden.

NOTE: During an FLC descent, the normally available airspeed/Mach

reference bug and digital display are removed.

The FLC mode is cancelled by any one of the following actions or event:
D Pushing the FLC button on the GC
D Selecting any other vertical mode on, or captured
D Switching to SG reversionary mode
D Selecting go--around
D Coupling to cross--side FD
D Coupled side MADC or AHRU goes invalid.

ALTITUDE PRESELECT (ASEL) MODE

The ASEL mode is used in conjunction with another vertical mode to
automatically capture, level off, and hold the altitude that was set with
the ASEL knob on the GC. The altitude preselect mode captures and

Modes of Operation A28--1146--112--00

8-32
 PRIMUSr1000 Integrated Avionics System

levels off on the correct altitude, while the other vertical mode is used
to fly to the desired altitude. The procedure to fly an ASEL mode is
described in table 8--9.

Step Procedure

1 Set the desired altitude in the PFD altitude preselect

window.
2 Initiate the ascent or descent toward the new altitude. This
process arms the ASEL mode.
3 Engage another vertical mode such as VS, SPD, or FLC on
the GC. The PFD displays the following messages:
D ASEL in white
D The other vertical mode in green.

Altitude Preselect Mode Engagement Procedure

Table 8--9

The aircraft flies toward the preset altitude using one vertical mode,
while ASEL is armed to automatically capture the preset altitude.

A28--1146--112--00 Modes of Operation

8-33
PRIMUSr1000 Integrated Avionics System

The ASEL mode automatically arms when the aircraft climbs or

descends towards the alert altitude and the following conditions are
met:
D ASEL altitude is more than 250 feet from the current altitude
D Computed vertical speed is greater that actual vertical speed
D Vertical speed is greater than 100 fpm for 3 seconds
D Glideslope is not captured.

The ASEL mode is cancelled in altitude hold, or at glideslope capture.

When the altitude select capture detector trips, the ASEL mode is
captured and the other active vertical mode is dropped. The PFD
displays ASEL in green.

At ASEL capture, a command is generated to flare the aircraft onto the

selected altitude.

The aircraft remains in the ASEL capture modes until the following
conditions exist simultaneously:
D Altitude error less that 25 feet
D Altitude rate less than 5 feet/second.

When these three conditions exist simultaneously, the system

automatically switches to altitude (ALT) hold.

NOTE: Changing the selected altitude on the GC while in the capture

phase, cancels the active ASEL mode and the armed ASEL
annunciation is displayed.

Table 8--10 is an illustrated step through of the ASEL procedure. The

numbers in figure 8--24 correspond to the procedural step numbers in
table 8--10. Figures associated with each step number show the PFD
annunciations after the step has been completed.

Modes of Operation A28--1146--112--00

8-34
 PRIMUSr1000 Integrated Avionics System

1 18,000 FEET PRESENT ALTITUDE

2 16,000 FEET

3 FLARE

4
SELECTED ALTITUDE 15,000 FEET
AD--50996@

Altitude Preselect Plan View

Figure 8--24

Step Procedure

1 Set the selected altitude on the PFD. Engage the VS mode

to descend toward the selected altitude. See display
below.

M 30 30

20 20

10 10

10 10

20 20

.390 M 30 30

001

Altitude Preselect Mode Procedure

Table 8--10 (cont)

A28--1146--112--00 Modes of Operation

8-35
PRIMUSr1000 Integrated Avionics System

Step Procedure

2 The mode is armed, as shown in the display below.

30 30

30 30

3 The altitude flare point (ASEL capture) is dependent on

vertical speed, as shown in the display below.

30 30 15500
M

40
153 20
00

30 30

3
2
1

1
2
3

Altitude Preselect Mode Procedure

Table 8--10 (cont)

Modes of Operation A28--1146--112--00

8-36
 PRIMUSr1000 Integrated Avionics System

Step Procedure

4 ALT SEL capture is dropped and ALT HOLD is

automatically engaged, as shown in the display below.

M 30 30

20 20

10 10

10 10

20 20

.470 M 30 30

3
2
1

1
2
3

Altitude Preselect Mode Procedure

Table 8--10

A28--1146--112--00 Modes of Operation

8-37
PRIMUSr1000 Integrated Avionics System

ALTITUDE HOLD (ALT HOLD) MODE

The altitude hold (ALT HOLD) mode is a vertical flight director mode
used to maintain a barometric altitude reference. To fly ALT HOLD
mode, follow the procedure in table 8--11.

Step Procedure

1 Put the aircraft into any lateral flight director mode.

2 Push the ALT button on the GC. The PFD annunciates, the
ALT annunciator is green while ALT HOLD is active, as
shown in the display below.

M
30 30

30 30

Altitude Hold Mode Engagement Procedure

Table 8--11

Altitude hold maintains the barometric altitude that exists when the
mode is engaged. The reference altitude can be changed by pushing
the touch control steering (TCS) button on the control wheel,
maneuvering the aircraft to a new altitude, and releasing the TCS
button. Selecting the ALT HOLD mode cancels any other previously
selected vertical mode.

The ALT HOLD mode is cancelled by any one of the following actions
or event:
D Pushing the ALT button on the GC
D Selecting any other vertical mode to on, or captured
D Selecting go--around

Modes of Operation A28--1146--112--00

8-38
 PRIMUSr1000 Integrated Avionics System

D Selecting the SG reversionary on the active side

D Coupling to the cross--side FD.
D The on--side AHRU or MADC become invalid.

WINDSHEAR (WDSHEAR) MODE

The windshear system uses a combination of detection and guidance.
The windshear detection computer annunciates an amber or red
WDSHEAR on the PFD for increasing performance or decreasing
performance windshears, respectively.

Any time an amber or red WDSHEAR is annunciated , as shown in

figure 8--25, the pilot is expected to apply full throttle, activate the
windshear mode, and execute a missed approach while following the
windshear guidance.

The windshear mode is activated manually by pushing the GA button

while an amber or red WDSHEAR annunciation is displayed. It is
activated automatically from the TO or GA modes if an amber or red
WDSHEAR occurs. It is also activated automatically from any mode if
a red WDSHEAR occurs and the throttles are greater than 80% of the
throttle lever angular travel.

When the windshear mode is active, the autopilot is disconnected and

the command bars are displayed. The command bars guide the aircraft
to gain energy or conserve energy depending on the conditions of the
airmass in which the aircraft is flying. This energy management is done
first by maintaining a slightly negative (climb) flight path angle. This
continues until the aircraft has either reached excess energy and
begins to control to a speed target for climb out or until the aircraft has
pitched up near stickshaker while trying to maintain the flightpath.

If the aircraft has reached excess energy, the aircraft switches from
flightpath control to speed control so the aircraft can reach whatever
climb rate it can while maintaining airspeed. Excess energy is defined
as the latched speed at time of windshear engagement, which is the
greater of 1.3Vs + 30 kts or current IAS. This speed latch is increased
by 20 knots if an increasing performance (amber) windshear is last
detected.

If the aircraft has pitched up while trying to maintain flightpath angle, the
windshear guidance controls to stickshaker until a climb can again
begin.

A28--1146--112--00 Modes of Operation

8-39
PRIMUSr1000 Integrated Avionics System

The windshear mode is cancelled by one of the following events:

D Selecting any vertical mode on the GC
D Selecting GA below 2500ft without a WDSHEAR annunciation present
D Coupling to cross side PFD
D Selecting the SG reversionary on the active side.

NOTES: 1. Automatic vertical modes are inhibited when in

windshear mode.
2. PFD declutter mode is inhibited when the windshear
mode is active.

AP

10 10

20 20

30 30

Windshear Mode PFD

Figure 8--25

Modes of Operation A28--1146--112--00

8-40
 PRIMUSr1000 Integrated Avionics System

GO--AROUND (GA) MODE (WINGS LEVEL)

The GA mode is normally used to transition from a landing approach to
a climb out condition in the event of a missed approach. The pilot
selects GA mode by pushing the GA button located on either outboard
throttle handle. With GA mode selected, all flight director modes are
cancelled. Laterally, the pilot sees a heading hold command on the
PFD, as shown in figure 8--26.

If the autopilot was engaged before go--around was selected, the

autopilot remains engaged. The command bars show a wings level and
pitch command to hold a minimum speed of 1.3 Vs.

CAUTION

DURING A COUPLED (AUTOPILOT ENGAGED) GO--AROUND UN-

DER HEAVY AND FORWARD CG CONDITIONS, AN ALTITUDE
LOSS OF APPROXIMATELY 100 FEET CAN OCCUR.

NOTES: 1. For an invalid AOA, minimum speed defaults to 120

kts.
2. The go--around mode ensures that minimal altitude is
lost during the maneuver.
3. The maximum pitch angle is 12°.
4. Under low power conditions, the mode first maintains
altitude (see CAUTION above) until 1.3 Vs is reached.
5. MIN SPD protection mode is active during go--around.
6. If a windshear condition is encountered and
go--around is activated, the command bars are driven
by windshear recovery commands. The vertical flight
director mode is WS.

The GA mode is cancelled by selecting another pitch mode or changing

the couple to cross--side.

A28--1146--112--00 Modes of Operation

8-41
PRIMUSr1000 Integrated Avionics System

151
30 30

1500

20 20

30 30

Go--Around Mode (Wings Level) PFD

Figure 8--26

Modes of Operation A28--1146--112--00

8-42
 PRIMUSr1000 Integrated Avionics System

9. Troubleshooting

This section describes conditions associated with suspected

malfunctions in the flight control system. Proper awareness helps the
pilot make write--ups that convey the necessary information to
maintenance personnel. As a result, unnecessary and expensive flight
time to verify pilot squawks can be greatly reduced.

This section does not imply that the pilot must troubleshoot the system
down to the box level. Rather, the paragraphs that follow contain
information required to give the pilot a complete understanding of
system problems.

CAUTION

THE FLIGHT CONTROL SYSTEM HAS BEEN DESIGNED TO EX-

HIBIT A HIGH DEGREE OF FUNCTIONAL INTEGRITY. HOWEVER,
THE USER MUST RECOGNIZE THAT IT IS NOT PRACTICAL TO
MONITOR AND/OR SELF--TEST FOR ALL CONCEIVABLE SYS-
TEM FAILURES. IT IS POSSIBLE THAT THE SYSTEM CAN OPER-
ATE ERRONEOUSLY WITHOUT A FAULT INDICATION. IT IS THE
RESPONSIBILITY OF THE PILOT TO DETECT SUCH AN OCCUR-
RENCE USING CROSS--CHECKS WITH REDUNDANT OR CORRE-
LATED INFORMATION AVAILABLE IN THE COCKPIT.

PILOT WRITEUP
The following paragraphs present general guidelines for making pilot
writeups in the logbook. As a guideline, the following two rules should
always be adhered to.

Rule 1 -- Before making an entry, determine conditions under which the

problem exists.

Rule 2 -- When making an entry, always write a complete description

of the problem.

NOTE: It is important that all system anomalies be correctly and

completely logged in. Effective communication significantly
improves the troubleshooting process.

A28--1146--112--00 Troubleshooting
9-1
PRIMUSr1000 Integrated Avionics System

Rule 1 states that system conditions must be considered before the

writeup is made. To aid in determining these conditions, the following
type of questions must be answered.
D Are any flags in view or fault annunciators lit (obvious problems)?
D Is the problem in pitch, roll, or yaw axis or a combination thereof?
D Is the problem present in all modes or only under specific conditions
such as:
- Flaps or gear up or down, or speed brakes in or out
- Certain aircraft power configuration
- Certain speed
- Certain altitude
- Two or more modes
- Certain sequence in mode selection
- Specific radio frequencies (NAV or COM)
- When keying a transmitter
- When weather radar is operating
- Certain electrical configurations (are all circuit breakers in)?

D Does the autopilot follow the commands as shown by the flight

director command cue and horizontal situation indicator lateral
deviation bar?

D Can the pilot fly the flight director commands with the autopilot
disconnected?

D Does some problem exist with autopilot engaged in a heading hold

and pitch hold mode?

D In radio modes, are certain conditions, such as another aircraft in

front of LOC or GS transmitter (overflight disturbances), VOR beam
scallops, etc., present?

Rule 2 states that an adequate description must be made when making

a writeup. The description should define the problem and should always
include the specific conditions under which the problem exists, such as:
D Flags showing (which ones, if any)
D Mode or modes selected.
D The IAS present when the problem occurs
D Period and magnitude of any oscillations
D Do any inputs fail to work (such as heading bug when in HDG
mode)?

NOTE: The preceding paragraphs are not intended to cover every

possible system problem. They should be used as guides in
preparing a pilot writeup.

Troubleshooting A28--1146--112--00
9-2
 PRIMUSr1000 Integrated Avionics System

Commonly Used Terms

Frequently, phrases or terms used in pilot writeups are not clearly
understood by all persons concerned. Table 9--1 lists some of the most
common terms and their definitions.

Term Definition

Autopilot Active Controls continually move in still air with small

command errors.
Autopilot Loose Autopilot does not null command bars
satisfactorily in most modes.
Porpoising Low frequency oscillation in the pitch axis,
typically 10--second period or longer.
Pumping The control wheel moves back and forth, usually
with a low frequency, and typically a 1-- to
10--second period.
Stick Bump Controls give a quick moderate movement,
usually with virtually no aircraft movement,
mostly associated with autopilot engagement or
during mode changes.
Stick Buzz With autopilot engaged, a high frequency, small
movement of control wheel can be felt without
aircraft movement.

Definitions of Terms
Table 9--1

A28--1146--112--00 Troubleshooting
9-3
PRIMUSr1000 Integrated Avionics System

Typical Problems
Some of the typical problems associated with flight control systems are
listed below. The list assumes the autopilot is engaged and is broken
down into lateral mode problems, vertical mode problems, and
problems that are common to both vertical and lateral modes.
Illustrations that show the most common lateral and vertical mode
problems are included. The list of problems and the illustrations are not
all inclusive, but are typical of the problems most often encountered.

LATERAL MODE PROBLEMS

Refer to table 9--2 and see figure 9--1 for an in--flight graphic representation
of lateral mode problems.

Mode Problems

HDG Mode -- Slow capture

-- Oscillates
-- Does not hold
NAV, BC or VOR APR mode; -- Undershoots capture
also Localizer portion of APR -- Overshoots capture
mode -- Missed capture
-- Standoff
-- Oscillates

Lateral Mode Problems

Table 9--2

Troubleshooting A28--1146--112--00
9-4
 PRIMUSr1000 Integrated Avionics System

AD--51002@

Lateral Mode Conditions and Problems

Figure 9--1 (cont)

A28--1146--112--00 Troubleshooting
9-5
PRIMUSr1000 Integrated Avionics System

AD--51003@

Lateral Mode Conditions and Problems

Figure 9--1

Troubleshooting A28--1146--112--00
9-6
 PRIMUSr1000 Integrated Avionics System

VERTICAL MODE PROBLEMS

Refer to table 9--3 and see figure 9--2 for an in--flight pattern graphic
representation of vertical mode problems.

Mode Problem

Air data hold modes (ALT, VS, -- Oscillates

IAS, MACH) -- Porpoising
-- Does not hold reference
Altitude preselect (ASEL) -- Misses capture
-- Undershoots capture
-- Overshoots capture
-- Standoff
GS mode (vertical portion of -- Captures early
APR mode) -- Standoff
-- Oscillates

Vertical Mode Problems

Table 9--3

AD--51004@

Vertical Mode Conditions and Problems

Figure 9--2 (cont)

A28--1146--112--00 Troubleshooting
9-7
PRIMUSr1000 Integrated Avionics System

AD--51005--R1@

Vertical Mode Conditions and Problems

Figure 9--2

Troubleshooting A28--1146--112--00
9-8
 PRIMUSr1000 Integrated Avionics System

COMBINED VERTICAL AND LATERAL MODE PROBLEMS

Refer to table 9--4 for combined vertical and lateral mode problems.

Mode Problems

Mode Logic Problems -- Modes do not engage

-- Modes do not clear
Autopilot Problems -- Autopilot does not engage
-- Autopilot does not follow commands
-- Stick bump
-- Stick buzz

Problems Common to Both Vertical and Lateral Modes

Table 9--4

GROUND MAINTENANCE TEST

On the ground, the crew can use the PRIMUSR 1000 Integrated
Avionics System to access to the status of several key internal
functions in the IC--600 Integrated Avionics Computer (IAC). The
ground maintenance test procedure is described in table 9--5, and the
ground maintenance display is shown in figure 9--3.

Step Procedure

1 Initiate system test by powering up the aircraft on the ground.

Steps 2--5 are all done on the display controller (DC).
2 Push and hold the TEST button. For the first 5 to 6
seconds, the EFIS displays the standard preflight test.
3 Continue to hold the TEST button. The PFD displays a listing
of key IAC internal functions.

Ground Maintenance Test Procedure

Table 9--5 (cont)

A28--1146--112--00 Troubleshooting
9-9
PRIMUSr1000 Integrated Avionics System

Step Procedure

4 As each function is satisfactorily tested, the FAIL/INVD

(fail/invalid) annunciator changes to PASS/VALD (pass/valid).
NOTE: Air data sensor, accelerometer, and rate or turn sensor inputs are
the only inputs tested during this test. All other tests are strictly
internal IAC processing tests.

5 When the test is complete, release the TEST button. More

detailed tests are available for maintenance personnel.
Access to those tests are made by pushing the TEST and
WX buttons and operating the RA set between 800 and 990
feet.

Ground Maintenance Test Procedure

Table 9--5

HELPING YOU CONTROL YOUR WORLD

HW ID 1

BASE PART NO 7017000

DASH NO --80152
SERIAL NO 93050155
SW MOD STATUS

640 RA

AD--51006--R2@

Ground Maintenance Test Display (Typical)

Figure 9--3

Troubleshooting A28--1146--112--00
9-10
 PRIMUSr1000 Integrated Avionics System

MULTIFUNCTION DISPLAY (MFD) CHECKLIST

UPLOAD PROCEDURE
The MFD checklist is stored in each display guidance computer. Table
9--6 is the procedure used to set up the aircraft to access checklists.

Step Procedure

1 The aircraft must be on the ground and powered up in

standby.
2 Locate the interface plug (located behind the pilot’s seat)
and connect a personal computer (PC) using a RS232
interconnect cable to the plug.
NOTES: 1. The PC must have the checklist programming software and
the correct checklist available for use. Refer to the ECP--800
Programmable Checklist equipment for details.
2. If the PC has WindowsTM, do not access the ECP--800
software from the WindowsTM prompt. Instead, use the DOS
prompt to start the checklist software.

3 Apply power to the avionics. Wait for the ADI to become

valid before processing.
4 Using the on--side DC, use the RA (DH) knob to set 890
RA (DH) on the PFD.
5 Push and hold the display controller TEST knob.
6 While pushing in the TEST knob, momentarily push the
display controller’s fourth selection button from the left.
The EFIS momentarily blanks and the following display
appears on the PFD:
NOTE: Do not follow any instructions on the display. These instructions
are not required actions.

CHECKLIST LOADING

PROGRAMMABLE CHECKLIST
EQUIPMENT IS REQUIRED

SCREEN WILL BLANK

CYCLE IC--600 CIRCUIT
BREAKER TO RECOVER

WHEN READY TO BEGIN:

PRESS PB#1 ON DC--500

MFD Checklist Upload Procedure

Table 9--6 (cont)

A28--1146--112--00 Troubleshooting
9-11
PRIMUSr1000 Integrated Avionics System

Step Procedure

7 Continue to hold the TEST knob for about 10 seconds and

then release the knob. The page in step 6 remains on the
PFD.
8 Momentarily push the first (#1) button from the left on the
EFIS DC. The PFD blanks and a large red X is displayed.
The red X remains until step 11.
9 Use the electronic programmable checklist software on the
PC to output the checklist and upload it to the pilot’s IAC.
NOTE: If a checklist is already in the IAC, error code 5100 is displayed on
the PC. Follow the instructions on the PC.

10 When the upload is complete, follow the instructions on the

PC to finish.
NOTE: If the No. 2 IAC needs a checklist upload, pull the No. 2 IAC cir-
cuit breaker, move the PC’s RS232 cable to the other plug by the
pilot’s right rudder pedal. Reset the No. 2 IAC circuit breaker and
repeat steps 3 through 10.

11 Pull the appropriate IAC circuit breaker to power down the

IAC.
12 Remove the RS232 cable from the airplane connector.
13 Push in the appropriate IAC circuit breaker to power up the
IAC.
14 When the IAC is powered up, verify that the checklist can
be selected by pushing either the NORM or EMER buttons
on the MFD controller located in the center pedestal.
15 Review the checklist for accuracy.
16 The procedure is complete.

MFD Checklist Upload Procedure

Table 9--6

ERROR CODE 500

If error code 500 is displayed, the PC cannot communicate with the IAC.
Use table 9--7 to clear the problem.

Troubleshooting A28--1146--112--00
9-12
 PRIMUSr1000 Integrated Avionics System

Step Procedure

1 Cycle power to the IAC by pulling the circuit breaker. When

the IAC has powered up, try programming it again. Do not
push R for retry, instead choose A for abort to start again.
2 Check the RS232 cable connection on the PC and the IAC
diagnostic connector.
3 Verify that the serial port (COM port) that the cable is
plugged into is the same as the one defined in the loading
configuration (F5 in the checklist loading program).
4 Check that the cable is plugged into the correct IAC
diagnostic connector. To program the pilot’s (primary
checklist) use the connector marked J1330. To program
the copilot’s IAC (backup checklist) use connector J1324.
5 Try moving the connector to another serial port. The PC
may have a bad serial port. Set the configuration in the
checklist loading program to correspond to the serial port
on the PC.
6 The extended memory manager (EMM) to network drivers
on the PC may interfere with the operation of the checklist
loading program. If network drivers are installed in the PC,
try to disable them and reprogram. If the PC has an EMM,
try disabling the EMM. Either of these functions can
usually be disabled by editing the PC’s CONFIG.SYS
and/or AUTOEXEC.BAT files.
7 Try reprogramming the other IAC. If the other IAC takes
the program, the problem is likely to be the IAC or aircraft
wiring.
8 Contact Honeywell Checklist Product Support that is given
in the Electronic Programmable Checklist manual.

Troubleshooting Procedure
Table 9--7

ERROR CODE 5005

An error code 5005 usually occurs when the IAC has never been
programmed with a checklist. Usually selecting RETRY allows the
checklist loading program to continue.

A28--1146--112--00 Troubleshooting
9-13/(9-14 blank)
 PRIMUSr1000 Integrated Avionics System

10. Honeywell Product Support

Honeywell SPEX program for corporate operators provides an

extensive Exchange and Rental service that complements a worldwide
network of support centers. An inventory of more than 9000 spare
components assures that your Honeywell equipped aircraft will be
returned to service promptly and economically. This service is available
both during and after warranty.

The aircraft owner/operator is required to ensure that units provided

through this program have been approved in accordance with their
specific maintenance requirements.

All articles are returned to Reconditioned Specifications limits when

they are processed through a Honeywell repair facility. All articles are
inspected by quality control personnel to verify proper workmanship
and conformity to Type Design and to certify that the article meets all
controlling documentation. Reconditioned Specification criteria are on
file at Honeywell facilities and are available for review. All Exchange
units are updated with the latest performance reliability MODs on an
attrition basis while in the repair cycle.

When contacting a Honeywell Dealer or Customer Support Center for

service under the SPEX program, the following information regarding
the unit and the aircraft are required:
D Complete part number with dash number of faulty unit
D Complete serial number of faulty unit
D Aircraft type, serial number and registration number
D Aircraft Owner
D Reported complaint with faulty unit
D Service requested (Exchange or Rental)
D Is faulty unit IN WARRANTY
- Type of warranty (NEW PRODUCT, MAINTENANCE
CONTRACT, or SPEX)
- Date warranty started
- Warranty ID number (if applicable)
D If faulty unit is NOT IN WARRANTY, provide billing address
D Purchase order number.

A28--1146--112--00 Honeywell Product Support

10-1
PRIMUSr1000 Integrated Avionics System

The Honeywell Support Centers listed below will assist with processing
exchange/rental orders.

24--HOUR EXCHANGE/RENTAL SUPPORT CENTERS

U.S.A. -- DALLAS CANADA -- OTTAWA

800--USA--SPEX 800--267--9947
(800--872--7739)

ENGLAND -- BASINGSTOKE AUSTRALIA -- TULLAMARINE

44--256--51111 61--3--330--1411

FRANCE -- TOULOUSE FRANCE -- DFS

33--6171--9662 33--1--4934--5050

SINGAPORE GERMANY -- AOA GAUTING

65--542--1313 49--89--89317--0
49--89--850--3695(After Hours AOG)

CUSTOMER SUPPORT CENTERS -- NORTH AMERICA

Dallas Support Center Canada Support Center

Honeywell Inc. Honeywell Inc.
Commercial Aviation Systems Commercial Aviation Systems
7825 Ridge Point 3 Hamilton Avenue North
IRVING, TX 75063 OTTAWA, ONTARIO, K1Y 4J4
Telephone: 214--402--4300 Telephone: 613--728--4681
Telex: 795539 Telex: 0533637
Fax: 214--402--4999 Fax: 613--728--7084

Central Support Center Ohio Support Center

Honeywell Inc. Honeywell Inc.
Commercial Aviation Systems Commercial Aviation Systems
1830 Industrial Avenue 8370 Dow Circle
WICHITA, KS 67216 STRONGVILLE, OH 44136
Telephone: 316--522--8172 Telephone: 216--243--8877
Telex: 417444 Telex: 985441
Fax: 316--522--2693 Fax: 216--243--1954

Honeywell Product Support A28--1146--112--00

10-2
 PRIMUSr1000 Integrated Avionics System

United Kingdom Support France Support Center

Center Honeywell Aerospace Ltd
Honeywell Avionics Systems Ltd 1, Rue Marcel--Doret
Edison Road, Ringway North 31700 TOULOUSE--BLAQNAC
BASINGSTOKE, HANTS, FRANCE
RG 21 2QD Telephone: 33--62--121500
ENGLAND After Hours -- AOG:
Telephone: 44--256--51111 33--61--71--9662
Telex: 85--8067 Telex: 52--1635
Fax: 44--256--474932 Fax: 33--61--300258

Singapore Support Center Australia Support Center

Honeywell Pty Ltd Honeywell Ltd
2 Loyang Crescent Trade Park Drive
SINGAPORE 1750 TULLAMARINE, VICTORIA 3043
Telephone: 65--542--1313 AUSTRALIA
Telex: RS 56969 Telephone: 61--3--330--1411
Fax: 65--542--1212 Telex: 35789 HWLTUL
Fax: 61--3--330--3042
Germany Support Center
Apparatebau Gauting Gmbh
Ammerseestrasse 45--49
D 8035 Gauting
GERMANY
Telephone: 49--89--89317--0
Telex: 0521702
Fax: 49--89--89317--183

A28--1146--112--00 Honeywell Product Support

10-3
PRIMUSr1000 Integrated Avionics System

Publication Ordering Information

Additional copies of this manual can be obtained by contacting:
Honeywell Inc.
P.O. Box 29000
Business and Commuter Aviation Systems
Phoenix, Arizona 85038--9000
Attention: Publication Distribution, Dept. M/S V19A1

Telephone No.: (602) 436--5553

FAX: (602) 436--1588

Honeywell Product Support A28--1146--112--00

10-4
 PRIMUSr1000 Integrated Avionics System

11. Abbreviations

Abbreviations used in this manual are defined as follows:

ABBREVIATION EQUIVALENT
ABNORM Abnormal
ABV Above
ADC Air Data Computer
ADF Automatic Direction Finder
ADI Attitude Director Indicator
ADS Air Data System
AFCS Automatic Flight Control System
AGL Above Ground Level
AHRS Attitude and Heading Reference System
AHRU Attitude and Heading Reference Unit
A/I Anti--Ice
ALT Altitude
AOA Angle of Attack
AOSS After Over Station Sensor
AP Autopilot
APP, APR Approach
APT Airport
APU Auxiliary Power Unit
ASEL Altitude Select
ATC Air Traffic Control
ATT Attitude
ATTCS Automatic Takeoff Thrust Control System
AUX Auxiliary

BARO Barometer
BFO Beat Frequency Oscillator
BLW Below
BNK Bank
BRG Bearing

CAP Capture
CAS Crew Advisory System
CAT Category
CAUT Caution
ccw Counterclockwise

A28--1146--112--00 Abbreviations
11-1
PRIMUSr1000 Integrated Avionics System

ABBREVIATION EQUIVALENT
CDI Course Deviation Indicator
CDU Control Display Unit
CHK Check
CKLST Checklist
CKPT Cockpit
CLB Climb
CLD Closed
CMC Central Maintenance Computer
COM Communication
CON Continues
CRS Course
CRZ Cruise
cw Clockwise

DAU Data Acquisition Unit

DC Display Controller
DES Descend
DG Directional Gyro
DGC Display Guidance Computer
DGR Degrade
DIM Dimming
DL Data Loader
DME Distance Measuring Equipment
DN Down
DR Dead Reckoning
DTK Desired Track
DU Display Unit

ECS Environmental Control System

EDS Electronic Display System
EFIS Electronic Flight Instrument System
EICAS Engine Instrument and Crew Alerting System
ELEC, ELECT Electrical
EMER Emergency
ENG Engine
ENT Enter
ERR Error
ET Elapsed Time

Abbreviations A28--1146--112--00
11-2
 PRIMUSr1000 Integrated Avionics System

ABBREVIATION EQUIVALENT
FADEC Full Authority Digital Engine Control
FD Flight Director
FF Fuel Flow
FGS Flight Guidance System
FHDG FMS Heading
FL Flight Level
FLC, FLCH Flight Level Change
FMS Flight Management System
FP, FPLN Flight Plan
FPM Feet Per Minute
FR From
FSBY, FSTBY Forced Standby

GA Go--Around
GCR Ground Clutter Reduction
GEN Generator
GMAP Ground Mapping
GPS Global Positioning System
GPU Ground Power Unit
GR/R React and Ground Clutter Reduction Modes
GS Glideslope
GSPD Groundspeed

HDG Heading
HDLC High Level Data Link Control
HDPH Headphone
HF High Frequency
HP High Pressure
hPa Hectopascals
HSI Heading Situation Indicator
HYD Hydraulic

IAC integrated Avionics Computer

IAS Indicated Airspeed
ID Identification
IDL Idle
IGN Ignition
ILS Instrument Landing System
IM Inner Marker

A28--1146--112--00 Abbreviations
11-3
PRIMUSr1000 Integrated Avionics System

ABBREVIATION EQUIVALENT
inHg Inches of Mercury
INPH Interphone
INTEG Integrity
ITT Interstage Turbine Temperature

JSTK Joystick

kg Kilogram
KPH Kilograms Per Hour

LAT/LON Latitude/Longitude
LBS Lateral Beam Sensor
LN BK Line Back
LOC Localizer
LP Low Pressure
L/R Left/Right
LSA Low Speed Awareness

MADC Micro Air Data Computer

MAG Magnetic
MAINT Maintenance
MAX Maximum
MFD Multifunction Display
MIC Microphone
MIN Minimum
MKR Marker
MM Middle Marker
M/P Map/Plan
MPEL Maximum Permissible Exposure Level
MSG Message
MSL Mean Sea Level

NAV Navigation
NB Narrow Bandwidth
NDB Non--Directional Beacons
NM Nautical Mile
NOC Navigation On Course
NORM, NRM Normal

Abbreviations A28--1146--112--00
11-4
 PRIMUSr1000 Integrated Avionics System

ABBREVIATION EQUIVALENT
OM Outer Marker
OPN Open
OSS Over Station Sensor

PAG Page
PAST Pilot--Activated Self--Test
PAX Passenger Address
PFD Primary Flight Display
POST Power--On Self--Test
PPH Pounds Per Hour
PRESS Pressure
PSI Pounds per Square Inch

RA Radio Altitude
RCL Recall
RCT Rain Echo Attenuation Compensation Technique
REF Reference
RMU Radio Management Unit
RNAV Area Navigation
RNG Range
ROL Roll
RSB Radio System Bus
RT, R/T Receiver/Transmitter
RTA Receiver Transmitter Antenna

SAT Static Air Temperature

SECT Sector
SG Symbol Generator
SKP Skip
SLV Slaved
SPD Speed
SPKR Speaker
SQ Squelch
SQUITTR Squitter
S.T. Sidetone
STAB Stabilization
STBY Standby
STC Sensitivity Time Control
STK Stuck

A28--1146--112--00 Abbreviations
11-5
PRIMUSr1000 Integrated Avionics System

ABBREVIATION EQUIVALENT
STO Storage, Store
STP Stop
SXTK Offset
Sync Synchronize

TA Traffic Advisory
TAS True Airspeed
TAT Total Air Temperature
TCAS Traffic Alert and Collision Avoidance System
TCS Touch Control Steering
TEMP Temperature
TERM Terminal
TGT Target
TKNB Turn Knob
TO, T/O Takeoff
TRK Track
TST Test
TTG Time--To--Go
TURB Turbulence
TX Transmitter

VAPP VOR Approach

VAR Variable
VBS Vertical Beam Sensor
VHF Very high frequency
VNAV Vertical Navigation
VOR VHF Omnidirectional Range
VS Vertical Speed
VTA Vertical Track Alert

WB Wide Bandwidth
WDSHEAR Windshear
WOW Weight on Wheels
WPT Waypoint
WS Windshear
WX Weather
WX/T Weather and Turbulence

YD Yaw Damper

Abbreviations A28--1146--112--00
11-6
 PRIMUSr1000 Integrated Avionics System

Appendix A
AHZ--800 Attitude and Heading
Reference System (AHRS)

INTRODUCTION
The AHRS is a self--contained strapdown system that computes
velocity, heading, and attitude data. It does not compute position
information.

The AHRS digital outputs the following:

D Primary attitude
D Magnetic heading
D Body linear accelerations
D Body angular rates.

Each AHRS contains the attitude and heading reference unit (AHRU),
configuration memory module, and flux valve. Both primary and
auxiliary 28 V dc are supplied to each AHRU. The system is controlled
using the panel--mounted HEADING/FREE and L/R switches.

The AHRS interfaces with the following equipment:

D Flight guidance system (FGS)
D Flight management system (FMS) (option)
D Micro air data computer (MADC)
D Electronic display system (EDS)
D Weather radar
D Backup radio management unit (RMU) navigation display.

The AHRU and the equipment it interfaces with are shown in figure A--1.

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--1/(A--2 blank)
 PRIMUSr1000 Integrated Avionics System

MADC TRUE AIRSPEED AHRU INTEGRATED

1 AVIONICS
PITCH, ROLL, HEADING, COMPUTER
FUNTIONS CONTROL NO. 1

MADC EDS
2 FMS
FGS
MAGNETIC NORTH
FLUX LONG TERM STABILITY
VALVE INTEGRATED
AVIONICS
COMPUTER
FLUX VALVE COMPENSATION, NO. 2
MEMORY INSTALLATION ORIENTATION SERIAL
MODULE BUS EDS

WEIGHT--ON--WHEELS
PITCH, ROLL P650
SHOP MAINTENANCE INPUT
RADAR
DISCRETES
MEMORY ACCESS MODE SEL

MAINTENANCE RMU
ACCESS OUTPUT
DISCRETE
AHRS FAIL ANNUNCIATOR
+28 VDC
INPUT
AUX +28 VDC POWER
FLIGHT
RECORDER
AD--51007@

AHRU Block Diagram

Figure A--1

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--3/(A--4 blank)
 PRIMUSr1000 Integrated Avionics System

ATTITUDE AND HEADING REFERENCE UNIT

The AHRU is the main electronics assembly of the AHRS. It contains
an inertial sensor assembly that includes the fiber optic gyros and
accelerometers. These measure accelerations and angular rates of the
aircraft.

The AHRU microprocessors compute the data required for flight

control, and display functions through digital bus interconnect. In
addition, an externally mounted memory module stores the flux valve
compensation data and aircraft AHRU orientation data.

The power supplies receive dual input dc power from the aircraft. The
AHRU supplies 26 Vac for the flux valve. True airspeed (TAS)
information is supplied by the MADCs.

The system can be tested and has troubleshooting functions that are
maintenance interconnect.

AHRS CONTROL AND EICAS ANNUNCIATION

AHRS Controls
AHRS controls shown in figure A--2, are located on the instrument
panel, adjacent to the reversionary panel.

DG AHRS CW

SLVD CCW

AD--51726@
AHRS Controls
Figure A--2

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--5
PRIMUSr1000 Integrated Avionics System

D MAG/DG (Magnetic/Directional Gyro) Switch -- When the pilot

inputs the MAG value, either a flux valve compensated heading, or
DG only heading is displayed.

NOTES: 1. If MAG is selected while in level flight, the AHRS

instantly returns to flux valve slaving. However, if
MAG is selected while in a turn, slaving starts only
after the bank is reduced to a small value and
occurs at 1_per minute.
2. While in MAG, no display of flux valve slaving
occurs.

D CW/CCW (Clockwise/Counterclockwise) Switch -- When a DG

heading is displayed, the CW/CCW switches let the pilot enter the
desired heading.

EICAS Annunciations
Table A--1 lists AHRS annunciations that are displayed on the EICAS.

Annunciation Color Definition

AHRS 1--2 OVHT Amber The AHRU is operating outside

normal temperature limits.
Damage can result.
AHRS 1--2 BASIC Cyan AHRS has lost all true airspeed
MODE information from both MADCs. It
is operating in the basic mode.

EICAS AHRS Annunciations

Table A--1

AHZ--800 Attitude and Heading Reference System (AHRS) A28--1146--112--00

A--6
 PRIMUSr1000 Integrated Avionics System

OPERATION

Standard Operation
The standard AHRS operating modes are the normal mode for
attitude, and the MAG mode for heading. After the system is turned on,
AHRS enters these modes automatically, if all system components
and signals are valid. In the normal mode, true airspeed from the MADC
is used to compensate for acceleration--induced error sources inherent
to the vertical gyro system. In the MAG heading mode, the flux valve
is used to establish the magnetic heading reference. System operation
in this mode is similar to that of a conventional magnetic compass, that
is stabilized by a gyroscope.

In the MAG mode, when valid flux valve data is lost, a HDG flag is
displayed on the heading instruments. The HDG flag clears when the
DG mode is entered again.

when exiting the DG mode, the AHRS automatically synchronizes the

heading outputs to the present flux valve magnetic heading. This
feature can also be used if a heading error develops 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 done by selecting
DG mode with the MAG/DG switch.

NOTES: 1. In normal operation, no flux valve slaving annunciation

is displayed.
2. The pilot cannot fast--erect the AHRS. The system
constantly updates the vertical calculations.

The AHRS has two power source inputs. On the pilot’s side, the primary
and auxiliary power is from the left essential bus. On the copilot’s side,
the power is from the essential bus; auxiliary power is from the
emergency battery bus. Separate circuit breakers are used for each
power source. In flight power loads or bus switching, transients do not
shut down the AHRS because power is automatically transferred within
the AHRU to the auxiliary power input. When the primary power input
is restored, the AHRS switches back to the primary power source.

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--7
PRIMUSr1000 Integrated Avionics System

Reduced Performance Operation

Two reduced performance modes are available: basic mode

for attitude, and DG mode for heading. The basic mode is annunciated
and entered automatically when true airspeed from the MADC
becomes invalid. The basic mode is annunciated by lighting a white
AHRS BASIC MODE message on the EICAS. AHRS basic mode
operation results in an attitude system that behaves like a conventional
vertical gyro (VG) with pitch and roll erection cutoffs. It is subject to drift
and acceleration errors. For this reason, AHRS in the basic mode
operates with reduced attitude accuracy.

The DG mode disables heading output automatic slaving. This mode

is entered by settomg the MAG/DG switch to DG. After the mode is
entered, the MAG annunciation on the PFD is replaced with a DG above
the compass display. AHRS DG mode operation acts like a free
directional gyro. It is subject to drift and turn error. For this reason, the
AHRS DG mode operates with reduced heading accuracy.

While in the DG mode, the heading card can be manually set to any
heading using the CW/CCW switch to enter an initial heading value.
The control can not be used in the slaved mode.

Moving the CW/CCW switch to CW increases the heading value. The

HDG flag on the HSI is displayed when a slew command is initiated, and
clears when the compass card reaches the selected heading. In the
DG mode, the system can be manually slaved using the cross--side
heading or standby compass as reference.

POWER--UP TEST

When AHRS power is applied, the system completes the power--up

test. On the ground, this test lasts two minutes. During this time, the
annunciations listed in table A--2 are displayed.

Annunciation Location

ATT FAIL In the PFD attitude sphere

HDG FAIL On the PFD and MFD

Pilot Self--Test Annunciations

Table A--2

AHZ--800 Attitude and Heading Reference System (AHRS) A28--1146--112--00

A--8
 PRIMUSr1000 Integrated Avionics System

PILOT SELF--TEST
After the AHRS is valid, the pilot can manually initiate AHRS test with
the display controller (DC) TEST button. When the AHRS test is
activated, the annunciations in table A--3 are displayed. The following
parameters are also displayed:
D 5°pitch up attitude
D 45°roll attitude right wing down
D 15°heading.

Annunciation Location

ATT TEST In the PFD attitude sphere

HDG TEST In the PFD heading circle
HDG TEST On the MFD

Pilot Initiated Self--Test Annunciations

Table A--3

NOTE: AHRS cannot be tested once airborne.

The system returns the correct values, clears the flags, and
extinguishes the annunciators after the 10--second test sequence is
complete. System modes (e.g., basic or DG) are not affected by the test
sequence.

The test function can extended beyond 10 seconds by continuing to

hold the TEST button. In this case, the test ends when the button is
released. If the test is valid, the TEST annunciation clears at the end
of the test.

GROUND INITIALIZATION
The AHRS system requires approximately 2 minutes to initialize after
the system is turned on. The initialization is complete when the ATT
and HDG flags on the ADI and HSI clear. During the initialization, the
aircraft must be stationary. Wind gusts and aircraft buffeting are not
limiting in this respect. All normal preflight operations, including engine
starts and passenger loading, can be carried out while the AHRS is
initializing. If the initialization requires more than 2 minutes, the AHRS
probably detected excessive aircraft motion. If the aircraft moves
during initialization, the AHRS starts a 5--second alignment process.
The 2--minute cycle can be monitored by watching the HSI heading

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--9
PRIMUSr1000 Integrated Avionics System

move from 115°to 0°. If the alignment is successful, the failure flags
are removed and the AHRS is valid. If not, the 5--second cycle starts
again.

If the heading card stops and does not step to an indication of 0°, the
initialization of the AHRS has not been completed correctly. The main
and auxiliary dc power to that AHRS must be removed by opening and
closing the circuit breakers, to restart the initialization.

NOTE: It is necessary that both breakers (primary and auxiliary

(AUX)) are pulled. Resetting each breaker individually does
not reset the AHRS.

In order to increase satisfactory ground initialization, the following must

be considered:

D The aircraft must remain stationary on the ground until the attitude
and heading flags are pulled out of view. Normal passenger and
cargo loading, engine start, and engine run--up procedures can be
performed during the initialization. Wind buffeting is not limiting in
this respect. The aircraft cannot be towed, nor can it taxi during
AHRS initialization.

D Two (2) minutes after power is connected to the dc buses, verify that
the attitude and heading flags are not displayed. If the timer has
stopped, the AHRS is not serviceable, and it should be reinitialized.
With the aircraft stationary, push and release the TEST button on
the AHRS controller. If the flags do not pull out of view after 10
seconds, reinitialize that AHRS.

D Normal preflight taxi checks of pitch, roll heading, and rate--of--turn

must be made on each system.

D If the aircraft is moved during AHRS initialization, both AHRS must

be reinitialized. This is done by pulling all four AHRS circuit
breakers. The four circuit breakers are then reset to their normal
position.

Abnormal Operations
TAKEOFF IN BASIC MODE

Takeoff with one AHRS in basic mode is not recommended. Verify that
both MADCs are operating.

NOTE: If a third attitude source is available, it can be used as the

primary data in place of the AHRS in the basic mode.

AHZ--800 Attitude and Heading Reference System (AHRS) A28--1146--112--00

A--10
 PRIMUSr1000 Integrated Avionics System

TAKEOFF IN DG MODE

Takeoff with one AHRS in the DG mode is not recommended. Return

the system to the normal slaved mode by setting the MAG/DG to MAG.

NOTE: If a third heading source is available, it can be used as the

primary data in place of the AHRS ,if slaved operation cannot
be achieved.

FLIGHT OPERATIONS IN BASIC MODE

If a malfunction causes the AHRS to revert to the basic mode, it is

annunciated on the EICAS. Normal flight operations can be continued
in the basic mode subject to the limitations of the aircraft flight manual.

In the basic mode, the pilot must avoid sustained, shallow banked turns
of less than 6°(e.g., a constant turn to hold DME arc). Keep the aircraft
properly trimmed.

NOTE: If an attitude error develops while in the basic mode, it cannot

be removed by the pilot. The system attempts to maintain the
error within bounds.

FLIGHT OPERATIONS IN DG MODE

If a heading flag is observed during a flight, set the MAG/DG switch to

DG.

The AHRS heading must be checked every 5 minutes with reference

to a known accurate heading source. Any errors can be removed by
using the INC/DEC switch to set the heading card to agree with the
known reference.

IN--AIR INITIALIZATION

In--air initialization is not recommended. If it is necessary to perform an

in--air initialization, the aircraft must be maintained in wings--level
unaccelerated flight during the alignment. In the air, alignment time is
normally 15 seconds. Both power source circuit breakers must be
pulled to restart the system.

NOTE: Since the AHRS continuously attempts to realign itself, if it

does not, pulling the circuit breakers to force a restart may not
be successful.

A28--1146--112--00 AHZ--800 Attitude and Heading Reference System (AHRS)

A--11/(A--12 blank)
 PRIMUSr1000 Integrated Avionics System

Appendix B
â
PRIMUS 660 Weather Radar
System

INTRODUCTION
The PRIMUSR 660 Weather Radar System is a lightweight, X--band
digital radar that is designed for weather detection and ground
mapping.

The purpose of the system is to detect storms along the flightpath and
give the pilot a visual color indication of rainfall intensity and turbulence
content. After proper evaluation, the pilot can chart a course to avoid
storm areas.

This appendix is an abbreviated operational description of the

PRIMUSR 660 Weather Radar System that is installed in the Embraer
145. For complete operating instructions on the PRIMUSR 660 Weather
Radar System, refer to Honeywell Pub. No. A28--1146--111.

WARNING

THE SYSTEM PERFORMS ONLY THE FUNCTIONS OF WEATHER

DETECTION OR GROUND MAPPING. IT IS NOT INTENDED THAT
THIS SYSTEM EITHER BE USED OR RELIED UPON FOR PROX-
IMITY WARNING OR ANTICOLLISION PROTECTION.

DESCRIPTION
The system consists of a receiver transmitter antenna (RTA) and a
single controller (dual controllers are optional). Radar information is
normally displayed on the multifunction display (MFD) in the MAP
mode.

In the weather detection mode, storm intensity levels are displayed in

four bright colors, contrasted against a deep black background. Areas
of very heavy rainfall are displayed in magenta, heavy rainfall in red,
less severe rainfall in yellow, moderate rainfall in green, and little or no
rainfall in black (background). Areas of detected turbulence are
displayed in soft white.

Range marks and identifying numerics, displayed in contrasting colors,

are used to evaluate the location of storm cells relative to the aircraft.

A28--1146--112--00 PRIMUS â 660 Weather Radar System

B--1
PRIMUSr1000 Integrated Avionics System

The ground mapping (GMAP) function is used to improve resolution

and identification of small ground targets at short ranges. The reflected
signals from ground surfaces are displayed as magenta, yellow, or cyan
(most to least reflective).

WEATHER RADAR CONTROLLER

Controls and Indicators

Controls and display features described below are numbered to match
the numbered callouts in figure B--1. All legend and controls on the
indicator that light are controlled by the dimming bus that controls the
aircraft panel.

1 2 3 4 5

OFF

RCT STAB TGT SECT

PULL WX GMAP
VAR SBY FP +
OFF TST

MIN MAX
--
GAIN RADAR SLV TILT

9 8 7 6

AD--52931--R1@

Weather Radar Controller

Figure B--1

1 RANGE SWITCHES

The range switches are two momentary contact buttons that are used
to set the operating radar range. WX ranges can be set from 5 to 300
NM full scale. In the flight plan (FPLN) mode, ranges of 500 and 1000
miles can be set. The up arrow selects increasing ranges, and the down
arrow selects decreasing ranges. One--half the selected range is
annunciated at the one--half scale range mark on the PFD or MFD.

NOTE: For dual controller installations, the weather radar range is

controlled by the on--side weather radar controller.

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--2
 PRIMUSr1000 Integrated Avionics System

2 RCT (RAIN ECHO ATTENUATION COMPENSATION

TECHNIQUE) SWITCH

The RCT switch is a toggle--contact switch that is used to select the

RCT mode. The RCT circuitry compensates for attenuation of the radar
signal as it passes through rainfall. The cyan field indicates areas where
further compensation is not possible. Any target detected within the
cyan field cannot be calibrated and should be considered severe
weather. All targets in the cyan field are displayed as fourth level
precipitation (magenta)

RCT is a submode of the WX mode and selecting RCT forces the

system to preset gain. When RCT is selected, the RCT legend is
displayed on the PFD or MFD.

3 STAB (STABILIZATION)

Pushing the STAB button selects or deselects the stabilization function

that automatically compensates for for aircraft roll and pitch
maneuvers.

4 TGT (TARGET) BUTTON

The TGT button is used to enable and disable the radar target alert
feature. Target alert is selectable in all but the 300 mile range. When
selected, target alert monitors beyond the selected range and 7.5_on
each side of the aircraft heading. If a return with certain characteristics
is detected in the monitored area, the target alert changes from the cyan
armed condition to the amber TGT warning condition (refer to table B--1
for target alert characteristics). The target advises the pilot of a
potentially hazardous condition directly in front of and outside the
selected range. When the amber warning is received, the pilot must
select longer ranges to view the target. Note that the target alert is
inactive within the selected range.

Selecting target alert forces the system into preset gain. Target alert
can be selected in the WX RCT and FPLN modes.

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PRIMUSr1000 Integrated Avionics System

Selected Range Target Depth Target Range

(NM) (NM) (NM)

5 2 5--155
10 2 10--160
25 4 25--150
50 4 50--150
100 6 100--175
200 6 200--250
300 6 300--350
FP (Flight Plan) 2 5--155

Target Alert Characteristics

Table B--1

5 SECT (SECTOR) SWITCH

The SECT switch is a toggle switch that selects either the normal 14
looks/minute 120° scan, or the faster update 20 looks/minute 60°sector
scan.
6 TILT KNOB
The rotary TILT knob is used to set the tilt angle of the antenna beam
with relation to the aircrafts longitudinal axis. Clockwise (cw) rotation
tilts the beam upward to +15_; counterclockwise (ccw) rotation tilts the
beam downward to --15_.

A digital readout of the antenna tilt angle is displayed on the MFD.

WARNING
TO AVOID FLYING UNDER OR OVER STORMS, FREQUENTLY SE-
LECT MANUAL TILT TO SCAN BOTH ABOVE AND BELOW THE
AIRCRAFT’S FLIGHT LEVEL. ALWAYS USE MANUAL TILT FOR
WEATHER ANALYSIS.
7 SLV (SLAVED) ANNUNCIATOR
The SLV dead--front annunciator is only used in dual controller
installations. With dual controllers, one controller can be slaved to the
other by selecting OFF with the radar MODE switch. This slaved
condition is shown with the SLV annunciator.

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--4
 PRIMUSr1000 Integrated Avionics System

In the slaved conditions, both controllers must be off before the radar
system turns off.

8 MODE Switch

The MODE switch is a rotary switch that is used to select one of the
following functions:

D OFF -- Turns the radar system off. An amber WX is displayed in the

MFD mode field.

D STBY (Standby) -- The radar system is placed in standby, a ready

state, with the antenna scan stopped. The transmitter is inhibited,
and the display memory is erased. A green STBY is displayed in the
MFD mode field.

If STBY is selected before the warm--up period is over

(approximately 45 seconds), a green WAIT legend is displayed in
the MFD mode field. When the warm--up period is over, the system
automatically switches to the STBY mode.

D FSBY (Forced Standby) -- Forced standby is an automatic,

nonselectable radar mode. FSBY mode is a safety feature that
inhibits the transmitter on the ground to eliminate the X--band
microwave radiation hazard. The controller is wired to the
weight--on--wheels (WOW) switch. The RTA is in forced standby
mode when the aircraft is on the ground. In the forced standby
mode, the transmitter and antenna scan are both inhibited, the
memory is erased, and the cyan FSBY legend is displayed in the
MFD mode field. When in FSBY, pushing the STAB button 4 times
within 3 seconds overrides the FSBY mode.

NOTE: When weather radar is displayed on the MFD, forced

standby is dropped once the aircraft is airborne.

D WX (Weather) -- Selecting WX places the radar system in the

weather detection mode. The system is fully operational and all
internal parameters are set for enroute weather detection.

If WX is selected before the initial RTA warm--up period is over

(approximately 45 seconds), a green WAIT legend is displayed. In
the WAIT mode, the transmitter and antenna scan are inhibited and
the memory is erased. When the warm--up period is over, the
system automatically switches to WX mode and a green WX is
displayed in the MFD mode field.

In preset gain, the system is calibrated as described in table B--2.

A28--1146--112--00 PRIMUS â 660 Weather Radar System

B--5
PRIMUSr1000 Integrated Avionics System

Rainfall Rate (mm/hr)* Color

1--4 Green
4--12 Yellow
12--50 Red
Greater than 50 Magenta
* Millimeters per hour

Rainfall Rate Color Cross Reference

Table B--2

D GMAP -- Selecting the GMAP position places the radar system in

the ground mapping mode. The system is fully operational and all
internal parameters are set to enhance returns from ground targets.
RCT compensation is inactive.

CAUTION

WEATHER TYPE TARGETS ARE NOT CALIBRATED WHEN THE

RADAR IS IN THE GMAP MODE. BECAUSE OF THIS, THE PILOT
SHOULD NOT USE THE GMAP MODE FOR WEATHER
DETECTION.

As a constant reminder that GMAP is selected, the green GMAP

legend is displayed and the color scheme is changed to cyan,
yellow, magenta. Cyan represents the least reflective return, yellow
is a moderate return, and magenta is a strong return.

If GMAP is selected before the initial RTA warm--up period is over

(approximately 45 seconds), a green WAIT legend is displayed. In
the WAIT mode, the transmitter and antenna scan are inhibited and
the memory is erased. When the warm--up period is over, the
system automatically switches to GMAP mode and a green GMAP
is displayed in the MFD mode field.

WARNING

THE SYSTEM ONLY PERFORMS THE FUNCTIONS OF WEATHER

DETECTION OR GROUND MAPPING. IT IS NOT INTENDED TO
BE USED OR RELIED UPON FOR PROXIMITY WARNING OR
ANTI--COLLISION PROTECTION.

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--6
 PRIMUSr1000 Integrated Avionics System

D FP (Flight Plan) -- In the FP position, the WX transmitter is placed

in standby and the PFD or MFD map range has been selected up
to 1000 NM. There is no radar data displayed in this mode.

NOTE: When weather is not selected for display, the MFD has its
own range control. The PFD does not require range
control.

D TEST -- The TEST position selects the radar test mode. A special
test pattern is displayed to verify system operation. The green TEST
legend is displayed in the MFD mode field.

WARNING

IF THE AIRCRAFT IS ON THE GROUND AND FORCED STANDBY

(FSBY) IS OVERRIDDEN, THE TRANSMITTER IS ON AND
RADIATING X--BAND MICROWAVE ENERGY IN THE TEST MODE.
REFER TO MAXIMUM PERMISSIBLE EXPOSURE LEVEL (MPEL)
IN THIS APPENDIX.
9 GAIN KNOB
The GAIN knob is a rotary control and a push/pull switch that controls
the receiver gain. When the GAIN switch is pushed, the system enters
the preset, calibrated gain mode. Calibrated gain is the normal mode
and is used for weather avoidance. In calibrated gain, the rotary
function of the GAIN control is disabled.

When the GAIN switch is pulled, the system enters the variable gain
mode. Variable gain is used for additional weather analysis and for
ground mapping. In the WX mode, variable gain can increase receiver
sensitivity over the calibrated level to show very weak targets, or it can
be reduced below the calibrated level to eliminate weak returns.

WARNING

HAZARDOUS TARGETS ARE ELIMINATED FROM THE DISPLAY

WITH LOW SETTINGS OF VARIABLE GAIN.
In the GMAP mode, variable gain is used to reduce the level of strong
returns from ground targets.

Minimum gain is set with the control at its fully ccw position. Gain
increases as the control is rotated in a cw direction from full ccw to the
12 o’clock position. At the 12 o’clock position, both the gain and the
sensitivity time control (STC) are at their maximum values. Additional
cw rotation removes STC. At the full cw position, the gain is at maximum
and the STC is at minimum.

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PRIMUSr1000 Integrated Avionics System

STC reduces the receiver gain at the start of the trace, and then
increases it as the more distant returns are received. With STC, a
uniform display of cell strength is displayed for both near and distant
cells.

The variable (VAR) legend annunciates variable gain. Selecting RCT,

or TGT forces the system into preset gain. Preset gain is not
annunciated.

Normal Operation
PRELIMINARY CONTROL SETTINGS

Place the MODE control, GAIN control, and TILT control, as shown
below, before powering up the aircraft electrical system.

MODE Control: Off

GAIN Control: Preset Position
TILT Control: +15

PRECAUTIONS

If the radar system is operated in any mode other than standby, while
the aircraft is on the ground, follow the precautions given in table B--3.

Step Precautions

1 Direct nose of aircraft so that antenna scan sector is free

of large metallic objects (such as hangars or other
aircraft), for a minimum distance of 100 feet, and tilt
antenna fully upwards.
2 Do not operate during aircraft refueling or during refueling
operations within 100 feet.
3 Do not operate if personnel are standing too close to the
270_forward sector of aircraft. (Refer to Maximum
Permissible Exposure Level (MPEL) in this appendix.)
4 Operating personnel should be familiar with FAA AC
20--68B, that is referenced in Honeywell Pub. No.
28--1146--120.

PRIMUSR 660 Weather Radar System Precautions

Table B--3

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--8
 PRIMUSr1000 Integrated Avionics System

POWERUP

On powerup, select either the standby or test mode. When power is first
applied, the radar is in WAIT mode for 45 seconds to let the magnetron
warm up. Power sequences ON--OFF--ON lasting less than the initial
45--second wait result in a 6--second wait period.

After warm--up, select the TEST mode and verify that the weather radar
test pattern shown in figure B--2 for the PFD, and figure B--3 for the
MFD, is displayed. Check the function of the TGT control.

169M 30 30

30 30

PFD Display Weather Radar Test Pattern

Figure B--2

A28--1146--112--00 PRIMUS â 660 Weather Radar System

B--9
PRIMUSr1000 Integrated Avionics System

12
21
S

AD--53423--R2@

MFD Display Weather Radar Test Pattern

Figure B--3

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--10
 PRIMUSr1000 Integrated Avionics System

TILT MANAGEMENT

Figures B--4 and B--5 show the relationship between tilt angle, flight
altitude, and selected range. The figures show the distance above and
below aircraft altitude that is illuminated by the flat--plate radiator during
level flight with 0_tilt and shows a representative low altitude situation,
with antenna adjusted for 3.95_up--tilt.

80,000
ELEVATION IN FEET

70,000
60,000 41,800 FT
ZERO TILT
50,000 20,000 FT
10,500 FT
CENTER OF RADAR BEAM
10,500 FT
30,000 20,000 FT
7.9
20,000 41,800 FT
10,000
0
0 25 50
100
RANGE NAUTICAL MILES
AD--35693@

Radar Beam Illumination High Altitude

12--Inch Radiator
Figure B--4

Radar Beam Illumination Low Altitude

12--Inch Radiator
Figure B--5

A28--1146--112--00 PRIMUS â 660 Weather Radar System

B--11
PRIMUSr1000 Integrated Avionics System

MAXIMUM PERMISSIBLE EXPOSURE LEVEL (MPEL)

Heating and radiation effects of weather radar can be hazardous to life.
Personnel should remain at a distance greater than R (shown in figure
B--6) from the radiating antenna in order to be outside the envelope in
which radiation exposure levels equal or exceed 10 mW/cm2, the limit
recommended in FAA Advisory Circular AC No. 20--68B, August 8,
1980, Subject: Recommended Radiation Safety Precautions for
Ground Operation of Airborne Weather Radar. The radius, R,
distance to the maximum permissible exposure level boundary is
calculated for the radar system on the basis of radiator diameter, rated
peak--power output, and duty cycle. The greater of the distances
calculated for either the far--field or near--field is based on the
recommendations outlined in AC No. 20--68B.

The American National Standards Institute, 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 B--6 shows the MPEL for the 12--inch antenna and
PRIMUSR 660 Weather Radar power.

MPEL Boundary
Figure B--6

PRIMUS â 660 Weather Radar System A28--1146--112--00

B--12
 PRIMUSr1000 Integrated Avionics System

Appendix C
â
PRIMUS II Radio System

INTRODUCTION
The PRIMUSR II Radio System, also known as the integrated radio
system, displays communication and navigation information to the
flightcrew. The navigation data is fed to the integrated avionics
computers (IAC) for display on the electronic display system (EDS),
and for use by the automatic flight control system (AFCS), both of which
are resident in the IACs. Navigation data also goes to the flight
management system (FMS) computer.

It is a dual system, with system No. 1 being the left (pilot’s) side and
system No. 2 being the right (copilot’s) side.

The cockpit controls consist of a radio management unit (RMU), an

audio panel, and a tuning backup control head.

A functional block diagram of the integrated radio system is shown in

figure C--1.

This appendix is a summary of the systems operation. Refer to

Honeywell Pub. No. 28--1140--50 for a complete description of the
PRIMUSR II Radio System.

A28--1146--112--00 PRIMUS â II Radio System

C--1/(C--2 blank)
 PRIMUSr1000 Integrated Avionics System

12

2. 3.

PRIMUSâ II Radio System Interface

Block Diagram
Figure C--1

A28--1146--112--00 PRIMUS â II Radio System

C--3/(C--4 blank)
 PRIMUSr1000 Integrated Avionics System

SYSTEM DESCRIPTION
General
The PRIMUSÒ II Radio System is comprised of the following cockpit
controls and remote radio units.
D Integrated Navigation Unit
D Integrated Communications Unit
D Audio Panel
D Radio Management Unit (RMU)
D Tuning Backup Control Head.
Except for the audio panel, all data for these control functions are
transmitted through a radio system bus. Audio switching is controlled
by the controls on the audio panel. The audio signals are transmitted
from the remote units to the audio panel through a dedicated digital
audio bus.
All the navigation data is displayed on the EDS. Data is also displayed
on either RMU display when they are operating as a standby navigation
display.
There are three levels of built--in test. First, at power up, the RMU
checks itself and the buses, and then commands each of the remote
functions to run through a full self--test. Any error is annunciated on the
RMU. Second, the pilot can activate a self--test of any individual
function through the RMU. Third, key parameters in the radios are
monitored at all times. If a discrepancy occurs, it is recorded in
nonvolatile memory located in the remote module. The discrepancy is
recorded in the memory and remains there, even after power down.
Discrepancy information includes: the discrepancy itself, the time into
flight it took place, and the temperature. The memory can be read in the
aircraft on the RMU, or the remote unit can be taken into the shop where
a technician can read out the results on a digital interface unit.

Integrated Navigation Unit

The integrated navigation unit, also known as the NAV unit, is a
complete self--contained navigation system. It contains the VOR,
localizer, glideslope and marker beacon receiver module, automatic
direction finder (ADF) module, and a six--channel scanning distance
measuring equipment (DME) module.
In the NAV unit, each module has its own independent housing and
power supply. The DME and ADF are mounted on the outside edges
of the unit and the VOR and cluster module are in the center. The DME
and ADF modules are both part of the NAV unit structure. They reduce
the RMUs installed volume and they heat sink to the outside surfaces.

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PRIMUSr1000 Integrated Avionics System

Integrated Communication Unit

The integrated communication unit, also known as the COM unit, is
identical in concept to the integrated NAV unit. It contains internal
modules that interface through a cluster module to the radio system bus
for operation. The modules in the COM unit are the very high frequency
(VHF) communication transceiver and the air traffic control (ATC)
transponder. The traffic alert and collision avoidance system (TCAS) is
controlled by this unit.

Radio Management Unit (RMU)

The RMU is the central control unit for the radio system. It controls the
operating mode, frequencies, and codes within all the units of the radio
system. The RMU can switch operation from its primary radio system
to the cross--side system. The RMU is an electronic controller with a
color and flat--panel display. It is used to select functions by pushing a
line select key adjacent to the parameter to be controlled. Any
selectable parameter, such as a VOR station frequency, can be
changed by pushing the line key next to the parameter and then rotating
the dual concentric controller tuning knobs.

The RMU display is divided into five dedicated windows. Each window
groups the data associated with a particular function of the radio
system. The windows (COM, NAV, transponder, ATC/TCAS, and
ADF) each control of the frequency, code, and operating mode of the
associated function. The RMU has other display modes, called pages,
that perform additional features and functions to control the radio
system. A menu of the pages access these additional features and
functions. In addition to standard RMU functions, the RMU can also
display navigation and engine data.

Audio Panel
The audio panel receives digitized audio from the remote navigation
(NAV) and communication (COM) units through a high speed digital
audio bus from each side of the airplane. Each audio panel selects the
proper channels from the digital audio bus and reconstitutes
headphone and speaker signals. The system can individually select the
radio function the crew member wants to hear using the audio panel.

The audio panel also has inputs for intercom, crew annunciation, crew
communication, hot microphone, etc., and full time emergency warning
inputs from aircraft systems.

PRIMUS â II Radio System A28--1146--112--00

C--6
 PRIMUSr1000 Integrated Avionics System

Tuning Backup Control Head

The tuning backup control head is an additional backup system that
controls the copilot’s VOR receiver and communication transceiver.

RADIO MANAGEMENT UNIT (RMU) OPERATION

General
The RMU is the central control unit for the radio system. Two RMUs are
mounted on the center pedestal, forward of the engine throttle
quadrants.

The RMU is used to control operating modes, frequencies, and codes

in all the units of the PRIMUSâ II radio system. Frequencies can also
be controlled by a radio control input from an external flight
management system (FMS). When the FMS tunes the radios, the
digital signals from the FMS come into the integrated radio system
where they act as if the RMU tuning knob were being operated. For
detailed FMS operation, refer to the applicable FMS operating
procedures. A third frequency control method uses the tuning backup
control head.

The RMU is also used to control the TCAS, a TCAS II system, that
displays conflict resolution advisories in the form of vertical guidance.

The RMU screen is divided into several dedicated windows. Each

window groups the data associated with a function. The windows,
COM, NAV, transponder, ATC/TCAS, and ADF, each control both the
frequency and operating mode of the associated function. The RMU
has other display modes, called pages, that control additional features
of the radio system.

The RMU is accomplished through the use of line select keys, function
keys, and concentric tuning knobs. Any selectable parameter, such as
a radio frequency, can be changed by pushing the line key next to the
displayed parameter. This places a cursor box around that parameter.
The values for that parameter can be set by rotating the tuning knobs.

On the ground, the RMU initiates a system power--on self--test when

power is first applied, or with weight--on--wheels (WOW) when power
has been off more than 10 seconds.

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C--7
PRIMUSr1000 Integrated Avionics System

Controls
See figure C--2 for the location of the controls and indicators described
below.

Radio Management Unit Controls

Figure C--2

D Transfer Key -- When pushed, the transfer keytoggles between the

active frequency (top line) and the preset frequency (second line) of
the window.

D Line Select Keys -- The first push of the line select key moves the
yellow cursor to surround the data field associated with the line select
key. This electronically connects that data field to the tuning knobs so
that the frequency or mode can be changed. For some functions,
additional pushes of the line select key toggles modes or recalls stored
frequencies. The line select key, if held for certain functions, recalls
automatic direction finder (ADF) and ATC memories, and enters and
exits the direct tune mode for COM and NAV.

D Tuning Knobs -- The tuning knobs are used to change the

frequency values or mode in the data field enclosed by the cursor.

D Cursor -- The yellow cursor encloses the data field selected by the
line select key. The cursor in the COM or NAV window can enclose
either the preset frequency or the memory mnemonics. Selection is
made by cycling the preset frequency line select key. The yellow
cursor ”homes”to its last position in the COM window 20 seconds
after the last tuning operation on the RMU. Home can be either the
preset COM frequency or the memory mnemonics.

PRIMUS â II Radio System A28--1146--112--00

C--8
 PRIMUSr1000 Integrated Avionics System

D Function Keys

- SQ (SQUELCH) KEY -- Pushing the SQ key opens the COM

radio squelch and any noise or signal present in the radio can be
heard in the audio system. The squelch button is strictly a toggle.
The letters SQ are annunciated along the top line of the COM
window when the squelch is opened.

- DIM (DIMMING) KEY -- Pushing the DIM key connects the RMU
brightness control to the tuning knob and the display can then be
adjusted to match overall cockpit brightness.

- 1/2 (CROSS--SIDE) KEY -- With the cursor in any window except

the ATC or TCAS displays, pushing the 1/2 key transfers the entire
RMU operation and display to the cross--side system. If the cursor
is in the ATC or TCAS display window, pushing this key selects
transponder for operation. With enhanced TCAS, the pilot pushes
the key to control TCAS data in the cross--side display.

- STO (STORE) KEY -- Pushing the STO key stores a temporary

(TEMP) COM/NAV preselect frequency stored in memory and
assigns a numbered location, if the cursor has first been placed
around that frequency. The ADF and ATC each have one memory
location. When the cursor is placed around the current ADF
frequency or ATC code, push the STO key to store it in memory.

- ID (IDENTIFICATION) KEY -- Pushing the ID key places the

transponder in the identification response mode. The
identification squawk terminates after 18 seconds.

- PGE (PAGE) KEY -- The PGE key accesses the page menu,
shown in figure C--3. From this menu the following functions can
be accessed (depending on the installation):
RADIO PAGE -- Returns the display to the normal tuning page
COM MEMORY -- Preset COM frequencies
NAV MEMORY -- Preset VOR NAV frequencies
ATC/TCAS -- A variety of ATC and TCAS display features
SYS ON/OFF -- Discrete radios ON/OFF switches
NOTE: The RADIO ON--OFF page can be disabled at
installation.
RETURN -- Returns the display to normal tuning page (same
as RADIO PAGE)
MAINTENANCE -- Ground only maintenance menu access.

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PRIMUSr1000 Integrated Avionics System

Page Menu
Figure C--3

- TST (TEST) KEY -- Pushing TST activates the internal self--test

circuits for the component associated with the yellow cursors
present position. The self--test is a complete end--to--end test
of the function. Hold the TST button down for the duration of the
test, about 2 seconds for COM transceiver, 5--to--7 seconds for
DME, ATC, ADF, and about 20 seconds for NAV (VOR/ILS).
Releasing the TST button at any time immediately returns the
function to normal operation. If the TST button is held for 30
seconds or more, the radios automatically return to normal
operation.

- DME (DISTANCE MEASURING EQUIPMENT) KEY -- This key

disassociates the DME from the active VOR frequency, so a
different DME channel can be tuned without changing active VOR
(DME Hold). Subsequently pushing the DME key switches the
display and DME channel formats between VHF and TACAN.

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Memory Pages

There are 12 memory locations associated with both the NAV and COM
windows. In all cases, these two windows preset frequency memory
functions work identically. Changing a preset NAV or COM frequency on
the main tuning page, shown in figure C--4, is done by putting the yellow
cursor over the frequency using the line select button, and changing the
frequency with the tuning knobs. Each memory location is numbered, and
is displayed on one of the two memory pages. Each page stores up to six
frequencies. Pushing the COM or NAV line select key a second time,
moves the cursor around the third line of the COM or NAV window.
Frequencies stored in memory and their corresponding memory labels
can be scrolled in the preset window by using the tuning knob.

AD--51017@

COM Frequency Storage Page

Figure C--4

System (SYS) On--Off Page

The SYS ON/OFF page is selected from the page menu shown in figure
C--5. This page can be inhibited by a strap option at installation.
This page lists all integrated COM and NAV radio modules. Using this
page, the pilot can turn off any one or more functions independently.
The line select keys opposite the listed modules act as ON--OFF
toggles. The ON--OFF page effects only on--side radios; the 1/2 switch
for cross--control is disabled.

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PRIMUSr1000 Integrated Avionics System

CAUTION

ALL SWITCHES POSITIONED ON THIS PAGE REMAIN IN

THE SAME STATUS AFTER POWER DOWN. THE RADIO PRE-
FLIGHT SHOULD INCLUDE THE RADIO ON--OFF PAGE TO EN-
SURE THAT ALL MODULES ARE ON. A SWITCH LEFT OFF IS INDI-
CATED BY DASHES IN THE CORRESPONDING FREQUENCY
WINDOW ON THE MAIN PAGE, AND BY AN ERROR MESSAGE FOR
THE TURNED--OFF MODULE DURING POWER--ON SELF--TEST.

AD--51018@

System On--Off Page

Figure C--5

NOTES: 1. This feature is dependent upon installation, but

typically only one side’s COM and NAV is wired for
ON--OFF controls. The COM and NAV (VOR) that are
the emergency radios (and, therefore, cannot be
turned ON or OFF) have an N/A annunciation in place
of ON or OFF. In some installations, the RMU can be
inhibited from calling up the RADIO ON--OFF page.

2. Single DME and ADF installations have their

respective ON--OFF functions removed from the No.
2 RMUs radio ON--OFF page.

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VHF Communications Transceiver Operation

The COM frequency select is in the upper left corner of the screen. The
COM window has three lines, all associated with frequency. The top line
displays the active frequency of the COM, while the second line
displays either the memory frequency or a scratchpad frequency,
depending on the last operation by the pilot.

Pushing the line select key associated with the lower (preset) COM
frequency display encloses the present frequency with the yellow
cursor box. Rotating the tuning knobs changes the preset frequency. If
it is necessary to store the preset frequency, push the STO key. The
COM storage mnemonics change to show the memory location.
Pushing the same line select key a second time encloses the storage
mnemonics with the yellow cursor box. Rotating the tuning knob
progresses through the stored COM frequencies by memory location
number. As each memory location (channel) is selected, the stored
frequency is shown in the COM preset area.

Storing overwrites anything previously in the memory location. When

overwriting occurs, the nomenclature beneath the bottom COM
frequency changes back to MEMORY.

Pushing the upper left transfer key on the RMU bezel (Y y B), swaps the
preset frequency and the active frequency location and function. This
switches the COM so it channels the new active frequency (previously,
the preset frequency). At the same time, the frequency in the active
portion of the window drops down to the second line and becomes the
preset frequency, labelled TEMP (as if the pilot had changed the preset
frequency with the tuning knobs). The new frequency can be stored,
using the STO button, or it can be modified to a new preset frequency
using the tuning knobs. If required, the new frequency can also be left
in the scratchpad for instant recall.

BASIC COM TUNING

After completing power on self--test, the RMU displays the main tuning
page, as shown in figure C--6. The data configuration is the same as it
was before the last power down.

Normally, the cursor is parked at the COM preset window. If not,

pushing the line select key alongside the COM preselect window moves
the cursor to that window. This action connects the tuning knobs to the
COM preset frequency.

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PRIMUSr1000 Integrated Avionics System

AD--51019--R2@

RMU in COM Tuning Mode

Figure C--6

Use the tuning knobs to set the frequency (i.e., 123.55, as shown in
figure C--7). The TEMP--1 label indicates that the temporary frequency
can be stored in MEMORY--1.

AD--51021--R2@

RMU Tuning
Figure C--7

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Push the top left COM transfer key, shown in figure C--8, to exchange
the preset and active frequencies. If required, the 118.00 preset
frequency can be changed to the next frequency, using the tuning knobs
or by pushing the line select key to bring up the next frequency from
memory.

COM 1 NAV 1

AD--51020--R2@

Exchanging the Preset and Active Frequencies

Figure C--8

Stored frequencies are selected by cycling the COM preset key until the
tuning box encloses the memory mnemonic. Rotating the tuning knob
scrolls through the stored frequencies and displays them in the preset
area.

NOTE: The frequency that is displayed in the top section of the

window, whether on the main tuning page or a memory page,
is ACTIVE.

COM MESSAGES

The COM messages and their conditions are listed in table C--1. The
following messages are displayed in the message area on top of the
COM tuning window, as shown in figure C--9.

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Message Description

MIC STK When the microphone is stuck on for approximately

two (2) minutes, a beep sounds on the audio and an
MIC STK is displayed until the MIC button is released.
Ten seconds after the MIC STK annunciation is
displayed, the selected transmitter automatically turns
off.
AUX ON AUX ON indicates that the auxiliary COM control
head is turned on or clearance delivery control is in
emergency mode. The respective COM is being
channeled by the AUX control head or the clearance
delivery control. The RMU is locked out from control
of that COM.
TX TX indicates that the transmitter is ON.
SQ SQ indicates that the squelch has been opened with
the SQ button.
NB NB indicates that a narrow bandwidth has been
selected.
WB WB indicates that a wide bandwidth has been selected.

COM Messages
Table C--1

AD--51022--R2@

COM Message (Squelch Open)

Figure C--9

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Navigation Receiver Operation

The NAV frequency select window is in the upper right corner of the
main tuning page shown in figure C--10. The NAV window is identical
to the COM in that it has two lines, each associated with frequencies.

AD--51024--R2@

NAV Frequency Select Window

Figure C--10

BASIC NAV TUNING

NAV radio tuning is done using the same procedure as COM radio
tuning. The NAV tuning window can also accommodate split DME
tuning.

Normal NAV tuning is as follows:

D The top line of the NAV box displays active VOR or ILS frequency
D The bottom line displays the memory frequency or a scratchpad
frequency.

The cursor encloses either the preset frequency and mnemonics or the
memory mnemonics.

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The message AUX ON is displayed at the top of the NAV window when
the NAV is being controlled by the tuning backup control head. The
RMU is then locked out from control of that NAV.

NOTE: Under certain circumstances, some integrated automatic

flight control/flight director systems have a VOR
tuning--inhibit signal to the navigation receiver that prevents
retuning the radio after approach course capture. If this
capability exists, the message INHIBIT is displayed in the
NAV window.

The upper right transfer key on the RMU bezel, shown in figure C--11,
toggles the active and preset frequencies to channel the NAV receiver
to what had been the preset frequency. The old active frequency moves
onto the line below and becomes the preset frequency.

NOTE: Whatever frequency is displayed in the top section of the

window is ACTIVE, whether it is on the main tuning page or a
memory page.

AD--51023--R2@

Toggling Active and Preset Frequencies

Figure C--11

SINGLE DME INSTALLATION

Single DME installations operate the same as dual configurations.

However, the No. 2 NAV has an unused DME channel in the No. 1 DME
as its own DME. There is no No. 2 maintenance event log.

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DME HOLD

DME hold is annunciated by splitting the NAV window, as shown in

figure C--12. While in hold, the actual DME can be tuned independent
of the active VOR or ILS.

AD--51025--R2@

DME Hold
Figure C--12

In normal VOR/ILS/DME operations, one of the six DME channels is

paired with the active frequency and another with the preset frequency.
Pushing the DME function key splits the NAV box on the main tuning
page and the active DME channel can be selected separately from the
active VOR/ILS frequency.

Cycling the DME select button sequences the NAV window, as

described in table C--2.

Push # Mode

1 Normal
2 VOR/ILS and DME split tuning
3 VOR/ILS and TACAN channel split tuning
4 Normal
DME Cycling Sequence
Table C--2

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When the NAV window is split, an amber H (hold) is displayed in the

lower (DME) window. The H indicates that the distance display (DME
or TACAN) is not paired with the VOR/ILS navigation data. When the
H is displayed, the other aircraft systems HOLD annunciators also light.

Display of the DME (or TACAN) channel being held positively identifies
the navigation channel. The DME station identifier is also displayed.

NOTE: It can take up to 2 or 3 minutes for the station identifier to be

displayed.

The DME function key operation is described in table C--3.

Step Procedure

1 Select the DME function key. The NAV window splits with
the upper half VOR/ILS tuning and the lower half DME
channel tuning.
2 Move the cursor to the DME window (lower) and tune
115.60, as shown in figure C--13. PXR is the identifier for
that channel. An amber H is shown.
3 Move the cursor to the upper VOR/ILS window with the top
line select key and tune 108.30.
4 When it is no longer necessary to hold the DME channel,
cycle the DME function button until the split is removed
from the NAV window.

DME Function Key Operating Procedure

Table C--3

ATC Transponder and TCAS Operations

GENERAL

ATC transponder and TCAS operations are controlled by code, mode,

range, and surveillance line select keys, transponder (1/2) button, and
code/mode select (TUNE) knobs on the RMU.

CONTROLS AND ANNUNCIATIONS

The following controls and annunciations are used to operate the ATC
transponder and TCAS. Refer to figure C--13 for locations.

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 PRIMUSr1000 Integrated Avionics System

RMU ATC/TCAS Controls

Figure C--13

The line select key opposite the ATC legend places the cursor over the
transponder code to be changed. The large outer tuning knob controls
the left two digits; the smaller inner knob controls the two right digits.

D Line Select Keys -- The first push of a line select key moves the
yellow cursor to surround the data line associated with that line
select key. This electronically connects that data field to the TUNE
knobs so that the ATC/TCAS mode or transponder code can be
changed.

D Code Line Select Key -- Push this key to place a cursor around the
four--digit transponder code. The code can be changed using the
large outer TUNE knob to control the left two digits, and the smaller
inner TUNE knob to control the right two digits. The code line select
key can also be used to retrieve a transponder code that has been
stored in memory.

D Transponder Code Storage (STO) Button -- A selected

transponder code can be stored in memory by pushing the STO
button while the cursor is on the transponder code. To retrieve the
code from memory, push and hold the code line select key for three
seconds.

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D Transponder Select (1/2) Button -- When the cursor is located in

the ATC window, this key is used to select the cross--side
transponder. The transponder in operation is indicated by a 1 or 2
located in front of the selected mode. Since only one transponder
can be operated at a time, the selected ATC transponder code is
duplicated on each RMU. For the RMU being operated from
cross--side, the data is displayed in yellow.

D Mode Line Select Key -- Pushing this key moves the cursor to the
mode line and enables the following functions:

- Repeated pushes of the mode key toggles the transponder

between standby and the last active mode.

- With the cursor on the mode line and an active mode selected,
push the 1/2 key to toggle between transponder 1 active and
transponder 2 active.

- Turn either TUNE knob to change the active mode. The available
selections are listed in table C--4.

Annunciation Mode

ATC ON Replies on modes A and S, no altitude reporting

Replies on modes A, C, and S, with altitude
ACT ALT
reporting
TA ONLY The TCAS traffic advisory mode is selected
The TCAS traffic advisory/resolution advisory
TA/RA
mode is selected

Transponder Modes
Table C--4

D Range Line Select Key -- Pushing this key moves the cursor to the
range select line. The TCAS range can be selected by turning either
TUNE knob or by repeated pushes of the range select key. The
available ranges are 6, 12, 20, and 40 NM.

D TCAS Extended Altitude Select Key -- Pushing this key moves the
cursor to the extended altitude select line. The extended altitude
selections can be selected by turning either TUNE knob or by
repeated pushes of the range select key repeatedly. The available
selections are listed in table C--5.

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Annunciation Extended Altitude Window Size

2700 feet above own aircraft and 2700 feet

NORMAL
below own aircraft
7000 feet above own aircraft and 2700 feet
ABOVE
below own aircraft
2700 feet above own aircraft and 7000 feet
BELOW
below own aircraft

Extended Altitude Modes

Table C--5

D Code/Mode Select (TUNE) Knobs -- The TUNE knobs are used to

change the transponder codes and the TCAS modes.

D TCAS Operational Select (PGE) Button -- This button is used to

select various TCAS operational features. Pushing the PGE button
accesses a page menu. This menu access is the TCAS operational
function that are selected using the ATC/TCAS button. The RMU
main tuning page selects and displays the primary TCAS selection
in the lower left window. The selections are as follows:

- TCAS DSPY 1/2 -- This annunciation shows whether the pilots (1)
or copilots (2) TCAS display features are being controlled. When
the cursor is in the window, the 1/2 button is used for the
selection. The selections are stored when the aircraft is powered
down.

- FLIGHT LEVEL 1/2 -- This is a display of the transponder’s

encoded altitude and the air data source for that altitude (for
example, DADC 1 and DADC 2).

D ATC Ident (ID) Button -- Pushing the ID button activates the

transponder ident mode for approximately 18 seconds. The ATC
reply annunciator (yellow rectangular box) is displayed along the top
edge of the transponder window, indicating that the transponder is
in the identification mode.

D ATC Reply Annunciator -- A reply annunciator (rectangular box),

located in the upper right corner of the ATC/TCAS window, turns
yellow when the transponder is replying to an interrogation from a
ground station or TCAS.

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ATC FAILURE WARNINGS

If the squitter function or the transmitter of the active Mode S

Transponder fails in flight, a SQUITTR INOP warning message is
displayed in red at the bottom of the ATC window. This message
indicates that the selected TCAS compatible Mode S Transponder has
lost some of its ability to operate as part of a collision avoidance system.

If the SQUITTR INOP message is displayed, do a pilot--activated

self--test (PAST) on the ATC. If the test fails (ATC ERR in red in the ATC
window), the transponder is most likely not transmitting properly. The
independent, other side transponder should be selected. If the
transponder passes PAST, the squitter function only is not operating
properly; this only slightly reduces the TCAS reply time. It is preferable
to use the other side transponder.

If a transponder fails (i.e., it does not accept a command from the RMU),
the RMU displays an ATC1 INOP or ATC2 INOP warning message in
red at the bottom of the ATC window. If ATC INOP is displayed, the
affected transponder should not be used. The independent, other side
transponder should be selected.

ADF Operation
GENERAL

The ADF receiver can drive a variety of bearing pointers. Its receiving
frequency range is 100 to 2183 KHz. The ADF frequency is turned like
for the COM and NAV modules, using the frequency select key next to
the ADF frequency window to position the yellow cursor box, and tuning
with the TUNE knobs.

The extended frequency range of the ADF tracks the new low frequency
non--directional beacons (NDBs) below 200 kHz as well as receiving the
marine emergency band at 2181 to 2183 kHz (installation dependent).

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 PRIMUSr1000 Integrated Avionics System

CONTROLS
The following controls are used to operate the ADF. Refer to figure C--14
for control locations.

ADF Controls
Figure C--14

D Frequency Select Key -- Push this key to place a cursor around the
ADF frequency. The frequency can be changed using the large outer
TUNE knob to control the left two digits, and the smaller inner TUNE
knob to control the right two digits. The frequency select key may
also be used to retrieve an ADF frequency that has been stored in
memory. To retrieve the stored frequency from memory, push and
hold the frequency select key for three seconds.
D ADF Mode Select Key -- Pushing this key moves the cursor to the
ADF mode line. Repeated pushes sequences modes that are
described in table C--6. The modes can also be selected by turning
the frequency/mode select (TUNE) knob.

Annunciation Mode

Maximum sensitivity and range. Aural beacon

ANT
ident and voice only. No bearing determination.
Bearing determination. Some loss of received
ADF
audio range and sensitivity.

ADF Modes
Table C--6 (cont)

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Annunciation Mode

ADF adds a beat frequency oscillator for

BFO
detection of cw signals.
Maximum audio clarity and fidelity. No bearing
VOICE
determination.

ADF Modes
Table C--6

Part of a dual bandwidth design, the VOICE mode transmits high

quality audio in the pilot headsets or cabin speakers. The high
quality audio comes from the wider bandwidth that is activated by
selecting VOICE. Direction finding requires narrow band reception
for bearing resolution.

NOTE: Because the VOICE mode has a wider audio bandwidth,

it is possible to have some beat frequency interference
when listening to an ADF station within 3 kHz of another
ADF station. This interference can be eliminated by
selecting the ANT mode.

D Frequency/Mode Select (TUNE) Knobs -- The TUNE knobs are

used to select the ADF operating frequency (when the yellow cursor
surrounds ADF frequency line) and can also be used to select the
ADF operating mode (when the yellow cursor is surrounding the
ADF mode line.)

Turning the knob clockwise raises the frequency, and

counterclockwise decreases the frequency. The large outer knob
functions in 10 kHz increments; the small inner knob in .5 kHz steps.
Slow knob movement produces slow frequency change. Rapid
turning, causes a several step jump, proportional to twisting speed.
Large frequency changes can be tuned quickly using a quick
twisting motion.

D ADF Frequency STO Button -- The RMU can store an ADF

frequency. This is done by selecting the desired frequency and then
pushing the STO button on the RMU while the cursor is still on the
ADF frequency. To retrieve the stored frequency from memory, push
and hold the frequency select key for three seconds.

ADF OPERATION WITH HF TRANSMITTERS

When an HF COM is keyed, the ADF receiver freezes the ADF bearing
needles. After the HF transmits continuously for 5 seconds, the receiver
”parks”the ADF bearing needles of electromechanical indicators at a

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 PRIMUSr1000 Integrated Avionics System

90°relative bearing position and removes ADF bearing needles from

electronic flight instrument displays until HF transmission ceases. This
avoids possible interference caused by bearing inaccuracy.

SINGLE ADF INSTALLATION

Single ADF installations operate the same as dual configurations

except that the ADF window annunciator does not identify a side
number. The main tuning page can be used to turn the ADF.

Cross--Side Operation
The RMU normally controls tuning of the radio system on its side of the
aircraft, but it can be switched into a mode called cross--side operation
to display and control the other radio system in the airplane.

If the pilot whose RMU normally tunes system No. 1, wants to tune the
copilot’s set of radios, the pilot pushes the 1/2 function key and replicate
the co--pilot’s No. 2 system (active frequencies, modes, and codes, but
not memory or preset frequencies) on the No. 1 RMU. Both RMU
displays are identical (except for memory and preset frequencies) but
the pilot’s RMU shows the function legends on the main tuning and
memory pages in magenta to indicate that cross--control is being
exercised. In addition to having access to the No. 2 system, the pilot still
has the NAV and COM memory frequencies in the No. 1 RMU available
to recall for use with the No. 2 system. Both pilots have this control
transfer function. The crew can coordinate tuning as well as have a
back--up mode in the event one RMU becomes inoperative.

NOTES: 1. Cross--side operation shifts between complete

systems. It is not possible to display/control on one
RMU some functions from the pilot’s side and some
from the copilot’s side at the same time.
2. The 1/2 function key does not affect the ATC window.
(Refer to note 3, below.)
3. Moving the cursor to the ATC window and pushing the
1/2 function key activates the opposite side
transponder.

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Built--In--Test (BIT)
GENERAL

The radio system BIT has three major functions:

D Performs a high confidence level pre--flight test
D Checks for latent failures that affect redundancy
D Assists in troubleshooting failures.

A power--on self--test (POST) is initiated when power is applied to the

RMU. POST causes the RMU to test itself, the primary digital bus that
connects both sides, each side secondary or back--up bus and the
interface to the remote units. Then it will command each remote unit to
perform its self--test. For more information on POST, refer to Honeywell
Pub. No. 28--1140--50.

In addition, each remote radio function has a pilot--activated self--test

(PAST) that performs a very comprehensive check of the individual
functions.

The BIT performs a thorough check of numerous system and internal

unit parameters. The entire system is designed to function properly
when all parameters are within their tolerance range.

NOTE: The errors identified by the self--test indicate that one or more
internal parameters were measured to be outside their
self--test limit but may not necessarily indicate non--operation
of the function. The pilot should verify operation of the
function. If any question remains, an authorized service
agency should be consulted.

The BIT also monitors important parameters and temperatures on a

continuous basis. If a discrepancy occurs, the code for the discrepancy
is recorded in nonvolatile memory along with the power--up count, the
elapsed time since power--up, the module temperature, and the
measured value of the discrepant parameter.

To aid in troubleshooting, the maintenance memory can be read in the

aircraft on the RMU when the aircraft is on the ground. Or the remote
unit can be taken into the shop where a digital interface unit can be used
to determine the failures.

PILOT--ACTIVATED SELF--TEST (PAST)

The pilot can initiate and conduct a self--test sequence of any radio
function. This pilot--activated self--test requires pushing and holding
down the TST button for the duration of the test, after the cursor is
positioned inside the window for the function to be tested.

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Sub--System Test Duration in Seconds

COM 2
DME 5 to 7
ATC 5 to 7
ADF 5 to 7
NAV (VOR/ILS) 20

PAST Self--Test Approximate Durations

Table C--7

After the test is complete, the green PASS or red ERR (Error) legend
is displayed in the window.

Releasing the test button at any time, immediately returns the function
to normal operation.

NOTES: 1. The errors identified by the self--test indicate one or more

internal parameters were outside their self--test limit but
do not necessarily indicate non--operation of the
function. The pilot should verify operation of the function.
If any question remains, an authorized service agency
should be consulted.
2. Later release RMUs may inhibit PAST during flight.

The following tables gives the procedures for the individual modules.

VHF COM PAST

Refer to table C--8 for the VHF COM PAST procedures.

Step Procedure
1 Position the cursor in the COM window. Verify that AUX
ON is not present.
2 Push and hold the TST button. The word TEST is
displayed in the COM window.

VHF COM PAST Procedure

Table C--8 (cont)

A28--1146--112--00 PRIMUS â II Radio System

C--29
PRIMUSr1000 Integrated Avionics System

Step Procedure
3 Continue holding the TST button. The words COM TEST are
displayed, indicating that the COM is performing a self--test.
4 Continue holding the TST button. Upon completion of the
self--test, the words COM PASS in green or COM ERR in
red are displayed in the COM window.
NOTE: If the COM squelch was open prior to running the PAST, the squelch closes
during the test.

VHF COM PAST Procedure

Table C--8

ATC TRANSPONDER AND TCAS PAST

Refer to table C--9 for the ATC transponder and TCAS PAST procedures.

Step Procedure
1 Position the cursor in the ATC/TCAS window and select
transponder 1.
2 Push and hold the TST button. SYS TEST will appear in
the ATC/TCAS window.
3 Continue holding the TST button. The words ATC TEST are
displayed indicating that the transponder is performing
self--test.
4 Continue holding the TST button. Upon completion of the
test, the words ATC PASS in green, or ATC ERR in red are
displayed in the ATC/TCAS window. In addition, the TCAS
aural warning sounds TCAS TEST, TCAS PASS, or TCAS
FAIL separately.
5 Repeat for transponder 2.
NOTES: 1. If a Honeywell TCAS is installed, it does not require that the TST button
to be held to activate its test sequence. A momentary push starts
TCAS (but only TCAS) testing.
2. With a Honeywell TCAS installed, it is possible to hold the TST button long
enough (8 seconds), with other conditions in the right sequence, to force
TCAS and the TCAS traffic display into the maintenance mode. Normal
operation can be re--established by entering a transponder code.
3. TCAS itself can also be tested with the cursor in the TCAS display
features window.

ATC Transponder and TCAS PAST Procedure

Table C--9

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 PRIMUSr1000 Integrated Avionics System

VOR/ILS/MARKER/DME PAST

Refer to table C--10 for the VOR/ILS/marker/DME PAST procedures.

NOTE: While in flight, this test may be inhibited by other aircraft

systems (e.g., by autopilot while on approach tracking).

Step Procedure
1 Adjust the selected course for zero degree.
2 Position cursor in NAV window. Verify that AUX ON is not
present. Push and hold the TST button. VOR TEST and
DME TEST are displayed in the NAV window.
3 VOR/ILS/MARKER PAST Results:
D After 2 seconds, marker lamps and tones annunciate in
the order of 400 Hz (outer); 1300 Hz (middle); and 3000
Hz (inner).
D Localizer and glideslope deviation bars indicate a
centered course for approximately 2 seconds with flags
out of view.
D Localizer and glideslope deviation bars deflect left
(localizer) and up (glideslope) one dot deflection, for
approximately 2 seconds with flags out of view.
D If course selector is on zero, VOR deviation bar centers
on a course of 0°and RMI indicates zero degree north
for approximately 5 seconds with flags out of view.
D The RMU annunciates VOR PASS in green, or VOR
ERR in red.
4 DME PAST Results:
D TEST is displayed as Ident on DME indicator and EFIS.
D 10.0 NM on DME indicator and EFIS
120 KTS
5 MIN
D RMU annunciates DME PASS in green or DME ERR in
red.

VOR/ILS/Marker/DME PAST Procedure

Table C--10

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PRIMUSr1000 Integrated Avionics System

ADF PAST

Refer to table C--11 for the ADF PAST procedures.

Step Procedure

1 Position cursor in ADF window.

2 Push and hold the TST button.
3 The results are:
D ADF TEST message will appear on the RMU.
D ADF bearing pointer slews to 135_± 10_ .
D Audio tone is heard through the audio system.
D RMU annunciates ADF PASS in green or ADF ERR in
red.

ADF Past Procedure

Table C--11

Maintenance
Several pages of maintenance associated data can be accessed
through the maintenance menu page. By simultaneously pushing the
COM and NAV transfer buttons (top button on each side of the RMU)
the RMU will leave its normal operating mode and displays a menu that
the operator can use to enter a maintenance mode of operation. This
menu can be called up only when the aircraft is on the ground. During
this mode, various pages are used to give maintenance personnel
access to software versions, strap option settings, maintenance log
data and operating conditions of the radio system. Pushing the bottom
left line select key returns the RMU to the main tuning page.

Refer to the radio systems maintenance manual for additional information.

RMU Display Dimming

RMU display brightness can be adjusted using the tuning knob after the
DIM function button is pushed. The tuning knob returns to normal
operation when any function or line select key is selected. Setting
display brightness sets the bezel mounted light sensors with a
reference about which they adjust the brightness for varying lighting
conditions.

PRIMUS â II Radio System A28--1146--112--00

C--32
 PRIMUSr1000 Integrated Avionics System

AUDIO PANEL OPERATION

General
Two audio panels (figure C--15) are mounted on the aircraft instrument
panel, outboard of the electronic display units. A third audio panel is
mounted at the observers station.

MASK

Audio Panel
Figure C--15

Controls
The audio panel receives digitized audio from remote radio units,
decodes the audio, controls the gain (volume) and routing of the various
channels, filters audio signals, and outputs the audio to various
speakers and headphones. It controls microphone inputs to various
radios, interphone, and passenger address systems. Amplifiers are
included for driving headphones and speakers.

The audio panel also has inputs for intercom, crew annunciation, crew
communication, hot microphone, etc., and full time emergency warning
inputs from aircraft systems.

Refer to figure C--15 for the locations of the following controls.

A28--1146--112--00 PRIMUS â II Radio System

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PRIMUSr1000 Integrated Avionics System

D COM1, COM2, COM3, and HF Microphone Switches -- These

switches are used to connect the on--side cockpit microphone to the
selected transceiver. The received audio from the selected
transceiver is routed to the on--side speaker or headphone at an
internally preset level, regardless of the setting of the COM or HF
audio source (on/off/volume) control located below the switches.
The received audio can then be adjusted as desired, using the audio
source control.

D Audio Source (COM, HF, NAV, ADF, and DME) Controls -- Each
control functions as an audio connect/disconnect switch and a
volume control. The audio is connected and the volume can be
adjusted when the control is in the unlatched (out) position. The
audio is disconnected from the on--side speaker or headphone when
the control is in the latched (in) position.

D Passenger Address (PAX) Microphone Switch -- When the PAX

button is pushed, the on--side microphone is connected directly to
the passenger address audio amplifier, independent of power being
applied to the audio panel. Passenger address sidetone is internally
generated within the audio panel. Headphone sidetone volume is
controlled by the sidetone (S.T.) and headphone (HDPH) volume
controls. The speaker sidetone volume is controlled by the S.T) and
speaker (SPKR) volume controls. During audio panel power loss, no
sidetones are available. During passenger address operations, all
other audio outputs, except warning audios, are deselected.

D Emergency (EMER) Microphone Switch -- When any audio panel

EMER button is pushed, the on--side microphone is connected
directly to the No. 2 VHF COM and the No. 2 VHF COM received
audio is connected directly to the on--side headphone. The No. 2
VOR/ILS audio is also connected directly to the on--side headphone
if it has been selected by the NAV AUDIO button on the tuning
backup control head. When EMER is selected, headphone volume
is controlled by the on--side HDPH volume control. The EMER
button disables all other audio panel controls.

D BOOM/MASK Switch -- When this switch is in the latched (in)

position, the headphone microphone is selected. When the switch
is in the unlatched (out) position, the oxygen mask microphone is
selected.

D ID/VOICE Control -- The ID/VOICE control is used to filter the VOR

and ADF audio signals. In the ID latched (in) position, the control
enhances Morse code identification. In the VOICE unlatched (out)
position, the audio is filtered to reduce the Morse code signal for
received VOR/ILS audio. ADF audio is unaffected when the VOICE
mode is selected.

PRIMUS â II Radio System A28--1146--112--00

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 PRIMUSr1000 Integrated Avionics System

D SPKR and HDPH Controls -- These controls are used to adjust the
overall on--side speaker and headphone volume. They function in
series with the individual audio source controls.

D S.T. Level Control -- The sidetone level control is used to prevent

undesirable feedback of speaker sidetone audio into the
transmitting microphone. When transmitting, both the on--side and
off--side speaker sidetone audio levels are lowered. In addition, the
off--side headphone and sidetone levels are lowered. When the S.T.
control is in the latched (in) position, the on--side cockpit speaker is
turned off and when the control is in the unlatched (out) position, the
on--side cockpit speaker is on.

D Marker (MKR) Beacon Volume Control -- This control is used to

adjust the marker beacon volume to the speaker and headphone. The
volume cannot be adjusted below an internally preset level. When the
MKR control is in the latched (in) position, the on--side cockpit
speaker/headphone is turned off and when the control is in the
unlatched (out) position, the on--side cockpit speaker/headphone is on.

D Marker Beacon MUTE and HI SENS/LO SENS Sensitivity

Controls -- The MUTE control is used to temporarily silence the
marker beacon audio. When the MUTE button is momentarily
pushed, the audio is muted. The audio remains muted for the time
it takes the aircraft to leave the influence of the marker beacon, then
it resets (unmutes) in preparation for reception of the next marker
beacon. The marker beacon receiver sensitivity is controlled by
rotating the MUTE control to the HI SENS or LO SENS positions.

D Interphone (INPH) Volume Control -- The INPH (intercom)

volume adjusts the on--side headphone audio level when the
interphone function is used.

TUNING BACKUP CONTROL HEAD OPERATION

General

The tuning backup control head shown in figure C--16, is located on the
center console, between the RMUs.

The tuning backup control head can be used before engine start for
initial communications with low power drain. It can act as a stand--alone
control unit or a back--up third control. It has two operating modes,
normal and emergency, which are selected by the MODES rotary knob
on the tuning backup control head front panel.

A28--1146--112--00 PRIMUS â II Radio System

C--35
PRIMUSr1000 Integrated Avionics System

AD--51036--R2@

Tuning Backup Control Head

Figure C--16

Normal Mode
In the normal mode, the tuning backup control head can be used to
control the co--pilot’s COM 2 and NAV 2 frequencies. The frequencies
selected by a crew member using the tuning backup control head are
displayed on both the tuning backup control head and the No. 2 RMU.
Frequencies selected by the co--pilot on the No. 2 RMU, or by the pilot
using the No. 1 RMU in the cross--side operating mode, are also
displayed on the tuning backup control head. The normal mode is not
annunciated.

Emergency Mode
In the emergency mode, the tuning backup control head becomes the
exclusive source for frequency control of the COM 2 and NAV 2 units.
In this mode, No. 2 RMU inputs are inhibited, the tuning backup control
head annunciates EMRG, and the No. 2 RMU annunciates AUX ON
(this annunciation is adjacent to the COM 2 and NAV 2 annunciations).
If the pilot is operating the No. 1 RMU in the cross--side mode, AUX ON
is displayed on the No. 1 RMU as well.

PRIMUS â II Radio System A28--1146--112--00

C--36
 PRIMUSr1000 Integrated Avionics System

Controls and Annunciations

The tuning backup control head controls and indicators shown in figure
C--17, are described below.

Tuning Backup Control Head Controls and Annunciations

Figure C--17

D Remote Tune Annunciator -- The remote tune annunciator is only

active when the system is strapped for NAV only or COM only
tuning. It is on when the radio is tuned from some source other than
what the unit is tuned to.

D Tuning Cursor (" ) -- The tuning cursor is a lighted triangle that is

controlled by the transfer key. It indicates which frequency can be
changed by the tuning knobs.

D Navigation Audio (NAV AUDIO) On Annunciator -- When

displayed, this annunciator indicates that NAV audio has been selected
on.

D Emergency (EMRG) Mode Annunciator -- The EMER

annunciator indicates when the tuning backup control head has
been placed in the emergency mode. This annunciation is not
related to the emergency COM frequency of 121.5.

D Squelch (SQ) Annunciator -- When displayed, this annunciator

indicates the squelch is opened by the SQ switch.

A28--1146--112--00 PRIMUS â II Radio System

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PRIMUSr1000 Integrated Avionics System

D Transmit (TX) Annunciator -- This annunciator indicates when the

copilot’s COM transmitter is transmitting.

D NAV AUDIO On/Off Switch -- This alternate action button is used

to toggle NAV audio on or off to the audio panel in in the EMER
mode.

D Squelch On/Off Switch -- This alternate action button is used to

toggle COM squelch on or off.

D Norm/Emergency MODES Switch -- This rotary switch knob

selects the normal and emergency modes.

D Transfer Key (Y y B) -- The transfer key alternately selects either the

COM frequency (top) or the NAV frequency (bottom) to be
connected to the tuning knobs.

D Radio Tuning Annunciators -- These four annunciations are lit

individually to identify source of the frequencies on the top and
bottom lines.

D System Installation Annunciator (2) -- Annunciates the radio

system to which the tuning backup control head is connected. For
the Embraer 145, aircraft wiring selects the No. 2 radio system.

PRIMUS â II Radio System A28--1146--112--00

C--38
 PRIMUSr1000 Integrated Avionics System

Index

A
Abbreviations, 11-1, 11-2
ADF operation, C--24
ADF operation with HF transmitters, C--26
Controls, C--25
ADF frequency STO button, C--26
ADF mode select key, C--25
Frequency select key, C--25
Frequency/mode select (TUNE) knobs, C--26
General, C--24
Single ADF installation, C--27
Air data system (ADS), 2-5
Altitude hold (ALT HOLD) mode, 8-37
Engagement procedure, 8-37
Altitude preselect (ASEL) mode, 8-32
Mode procedure, 8-34
Engagement procedure, 8-32
ATC transponder and TCAS operations, C--20
ATC failure warnings, C--24
Controls and annunciations, C--20
ATC ident (ID) button, C--23
ATC reply annunciator, C--23
Code line select key, C--21
Code/mode select (TUNE) knobs, C--23
Line select keys, C--21
Mode line select key, C--22
Range line select key, C--22
TCAS extended altitude select key, C--22
TCAS operational select (PGE) button, C--23
Transponder code storage (STO) button, C--21
Transponder select (1/2) button, C--22
General, C--20
Attitude and heading reference system (AHRS), 2--6, A--1, A--5, A--6
AHRS control and EICAS annunciation, A--5
AHRS controls, A--5
EICAS annunciations, A--6
Attitude and heading reference unit, A--5
Ground initialization, A--9
Abnormal operations, A--10
Introduction, A--1

A28--1146--112--00 Index
Index--1
PRIMUSr1000 Integrated Avionics System

Index (cont)
Operation, A--7
Reduced performance operation, A--8
Standard operation, A--7
Pilot self--test, A--9
Pilot initiated self--test annunciations, A--9
Power--up test, A--8
Pilot self--test annunciations, A--8
Audio panel, C--6
Audio panel operation, C--33
Controls, C--33
Audio source (COM, HF, NAV, ADF, and DME) controls,
C--34
BOOM/MASK switch, C--34
COM1, COM2, COM3, and HF microphone switches, C--34
Emergency (EMER) microphone switch, C--34
ID/VOICE control, C--34
Marker (MKR) beacon volume control, C--35
Marker beacon MUTE and HI SENS/LO SENS sensitivity
controls, C--35
Passenger address (PAX) microphone switch, C--34
S.T. level control, C--35
SPKR and HDPH controls, C--35
General, C--33
Autopilot controller, 6-6
PITCH wheel, 6-6
TURN knob, 6-7
Autopilot power--up test, 6-5

Back course (BC) mode, 8-17

Flight procedure, 8-17
Backup EICAS display, 5-31
Backup EICAS display failure indications, 5-34
Built--in--test (BIT), C--28
ADF PAST, C--32
Procedure, C--32
ATC transponder and TCAS PAST, C--30
Procedure, C--30
General, C--28
Pilot--activated self--test, C--28

Index A28--1146--112--00
Index--2
 PRIMUSr1000 Integrated Avionics System

Index (cont)
VHF COM PAST, C--29
Procedure, C--29
VOR/ILS/marker/DME PAST, C--31
VOR/ILS/marker/DME PAST procedure, C--31

Cockpit mounted equipment, 1-2

Controls, 4-5, 5-5
Display controller (DC), 4-20
EFIS preflight self--test, 4-23
Self--test, 4-23
Instrument test, 4-24
Switching test, 4-24
EICAS bezel controller, 5-5
Guidance controller course and heading controls, 4-22
Multifunction display (MFD) bezel controller, 4-6
Checklist menu, 4-15
Map/plan, 4-19
MFD main menu, 4-8
MFD menu, 4-10
System menu, 4-8
TCAS, 4-17
WX (weather), 4-18
Primary flight display (PFD) bezel controller, 4-5
Crew alerting system (CAS), 5-22
CAS display failure indications, 5-29
EICAS test, 5-29
Symbol generator failure, 5-29
CAS messages, 5-23
Advisory messages (Cyan), 5-28
Caution messages (Amber), 5-24
Computed advisory messages, 5-29
Computed caution (AMBER) messages, 5-27
Warning messages (RED), 5-23
Cross--side operation, C--27
Customer support centers -- North America, 10-2

A28--1146--112--00 Index
Index--3
PRIMUSr1000 Integrated Avionics System

Index (cont)

D
Display controller (DC), 4-20
BRG (bearing) knobs, 4-21
ET (elapsed time) button, 4-20
FMS (flight management system) button (optional), 4-20
FULL/WX button, 4-20
GSPD/TTG (groundspeed/time--to--go) button, 4-20
NAV (navigation) button, 4-20
RA (radio altitude) knob/TEST button, 4-21
Displays, 4-25, 5-6
Backup EICAS display, 5-31
Backup EICAS display failure indications, 5-34
Crew alerting system (CAS), 5-22
CAS display failure indications, 5-29
CAS messages, 5-23
EICAS display, 5-6
EICAS reversionary modes, 5-30
Loss of display guidance computer (DGC) symbol
generator, 5-30
Loss of the display unit, 5-30
MFD system displays, 4-86
Electrical (ELEC) system, 4-93
Environmental control system/anti--ice, 4-88
Fuel system, 4-90
Hydraulic (HYD) system, 4-92
Takeoff (T/O), 4-86
Multifunction display (MFD), 4-68
Changing a waypoint, 4-80
Checklist (option), 4-82
Map display, 4-72
Map failure and warning displays, 4-75
MFD common symbols, 4-69
MFD display window, 4-82
MFD plan mode, 4-77
Plan mode failure and warning displays, 4-79
Traffic alert and collision avoidance system (TCAS), 4-84
Typical map mode presentations, 4-76
Typical plan mode presentation, 4-80
Weather, 4-74

Index A28--1146--112--00
Index--4
 PRIMUSr1000 Integrated Avionics System

Index (cont)
PFD caution and failure display, 4-61
Caution (AMBER) annunciations, 4-61
Excessive attitude declutter, 4-66
Failure (RED) annunciations, 4-65
PFD test mode, 4-67
Primary engine displays, 5-7
Cabin pressure and auxiliary power unit (APU) status, 5-18
Digital turbine (N2) display, 5-12
Engine ignition (IGN) annunciation, 5-11
Engine vibration display, 5-14
Fan (N1) speed display, 5-9
Fuel flow (FF PPH FLOW) digital display, 5-13
Fuel quantity (FUEL LB) digital display, 5-13
Gear/flaps/spoiler positions, 5-15
Interstage turbine temperature (ITT) display, 5-11
Oil pressure (OIL PSI) display, 5-12
Oil temperature (OIL °C) display, 5-14
Pitch/roll/yaw trim display, 5-16
Primary engine displays failure indications, 5-20
Primary flight display (PFD), 4-25
Airspeed display, 4-45
Altimeter display, 4-48
ARC display, 4-41
Full compass display, 4-37
Functional divisions, 4-25
PFD attitude director indicator (ADI), 4-26
Vertical speed (VS) display, 4-51
Standby navigation display, 4-96
RMU backup navigation display, 4-98
System display failure annunciation, 4-95
Display controller failures, 4-95
Display system reversionary modes, 4-95
Typical PFD presentations, 4-53
Approach capture tracking at DH, 4-58
Climb to initial altitude, 4-54
Comparison monitoring, 4-59
Enroute cruise, 4-55
Setup for approach, 4-56
Takeoff (TO) mode, 4-53
Terminal area, 4-57

A28--1146--112--00 Index
Index--5
PRIMUSr1000 Integrated Avionics System

Index (cont)

E
Electronic display system (EDS), 2--4, 3-1, 3-5, 3-6
Introduction, 3-1
MFD bezel controller basic operation, 3-5
Bezel controls, 3-5
Controller conventions, 3-6
Map range control, 3-6
Reversionary controllers and dimming panel, 3-7
Data acquisition unit (DAU) reversionary panel, 3-8
Dimming panel, 3-8
System reversionary panel, 3-7
Electronic flight instrument system (EFIS), 4-1, 4-5, 4-6
Controls, 4-5
Display controller (DC), 4-20
EFIS preflight self--test, 4-23
Guidance controller course and heading controls, 4-22
Multifunction display (MFD) bezel controller, 4-6
Primary flight display (PFD) bezel controller, 4-5
Displays, 4-25
MFD system displays, 4-86
Multifunction display (MFD), 4-68
PFD caution and failure display, 4-61
Primary flight display (PFD), 4-25
Standby navigation display, 4-96
System display failure annunciation, 4-95
Typical PFD presentations, 4-53
Introduction, 4-1
Engine instrument and crew alerting system (EICAS), 5-1
Controls, 5-5
EICAS bezel controller, 5-5
Master warning and caution annunciation switches, 5-5
Displays, 5-6
Backup EICAS display, 5-31
Backup EICAS display failure indications, 5-34
Crew alerting system (CAS), 5-22
EICAS display, 5-6
EICAS reversionary modes, 5-30
Primary engine displays, 5-7
Introduction, 5-1

Index A28--1146--112--00
Index--6
 PRIMUSr1000 Integrated Avionics System

Index (cont)
Equipment
Air data system, 2-5
Attitude and heading reference system, 2-6, A--1
Cockpit mounted, 1-2
Electronic display system, 2-4, 3--1
Flight guidance system (FGS), 2-3, 6--1
Flight management system, 2-9
Integrated radio system, 2-7
List, 1-2
Optional, 1-3
Optional list, 2-1
Other switches and controls, 2-9
Radio altimeter system, 2-7
Remote mounted, 1-2
Standard list, 2-1
Traffic alert and collision avoidance system, 2-8
Weather Radar System, 2-6, 2-8

F
Flight guidance system (FGS), 2--3, 6-1, 6-2
Autopilot controller, 6-6
PITCH wheel, 6-6
TURN knob, 6-7
Guidance controller (GC), 6-1
Autopilot (AP)/yaw damper (YD)/couple (CPL) buttons, 6-1
Autopilot power--up test, 6-5
FGS caution and advisory messages, 6-5
Flight director (FD) mode selection, 6-3
Remote autopilot switches and annunciators, 6-5
Flight level change (FLC) mode, 8-31
Descend and climb parameters, 8-31
Flight management system (FMS) (Option), 2-9
Flight management system (FMS) navigation mode (option), 8-10
Engagement procedure, 8-10

G
Glossary of terms, 7-1
Attitude director indicator (ADI) command cue, 7-1
Glideslope (GS) gain programming as a radio altitude (RA)
function, 7-2

A28--1146--112--00 Index
Index--7
PRIMUSr1000 Integrated Avionics System

Index (cont)
GS capture (CAP), 7-2
GS gain programming as a RA and vertical speed (VS) function,
7-2
Lateral beam sensor (LBS), 7-2
Lateral gain programming, 7-3
Localizer/back course (BC) track, 7-3
True airspeed (TAS) gain programming, 7-3
Vertical beam sensor (VBS), 7-4
VOR after over station sensor (AOSS), 7-4
VOR capture (CAP), 7-4
VOR over station sensor (OSS), 7-4
Go--around (GA) mode (wings level), 8-40
Ground initialization, A--9
Abnormal operations, A--10
Flight operations in basic mode, A--11
Flight operations in DG mode, A--11
In--air initialization, A--11
Takeoff in basic mode, A--10
Takeoff in DG mode, A--10
Ground maintenance test, 9-9
Test procedure, 9-9
Guidance controller (GC), 6-1
Autopilot (AP)/yaw damper (YD)/couple (CPL) buttons, 6-1
AP button, 6-1
AP disengage, 6-2
CPL button, 6-2
YD button, 6-2
YD disengage, 6-2
Autopilot power--up test, 6-5
FGS caution and advisory messages, 6-5
Amber (caution) flight guidance system messages, 6-6
Red (warning) flight guidance system messages, 6-5
Flight director (FD) mode selection, 6-3
ALT (altitude) button, 6-3
AP OFF, 6-4
APR (approach) button, 6-3
CRS (course) knob, 6-4
FD1/FD2 buttons, 6-4
FD1/FD2 switch operation, 6-4
FLC (flight level change) button, 6-3
HDG (heading) button, 6-3

Index A28--1146--112--00
Index--8
 PRIMUSr1000 Integrated Avionics System

Index (cont)
HDG knob, 6-4
NAV (navigation) button, 6-3
SPD (airspeed) button, 6-3
VS (vertical speed) button, 6-3
YD OFF, 6-4
Remote autopilot switches and annunciators, 6-5
AP disconnect switch, 6-5
Go--around button, 6-5
Pitch wheel and turn knob controls, 6-5
TCS button, 6-5
Guidance controller course and heading controls, 4-22
CRS sync, 4-22
HDG sync, 4-22

Heading hold mode, wings level, 8-1

Heading select mode, 8-3
Honeywell product support, 10-1, 10-2
Customer support centers -- North America, 10-2
Publication ordering information, 10-4
Support centers, 10-2, 10-3
24--hour exchange/rental support centers, 10-2

Instrument landing system (ILS) approach mode, 8-20

Procedure, 8-21
Integrated communication unit, C--6
Integrated navigation unit, C--5
Introduction, 1-1, 1-2
Equipment list, 1-2
Supplemental pilot’s manuals, 1-1

Localizer (LOC) mode, 8-12

Engagement procedure, 8-12

A28--1146--112--00 Index
Index--9
PRIMUSr1000 Integrated Avionics System

Index (cont)

Maximum permissible exposure level (MPEL), B--12

Memory pages, C--11
MFD bezel controller basic operation, 3-5
Bezel controls, 3-5
Momentary selection, 3-5
Parameter display selection, 3-5
Settable parameters, 3-5
Submenu selection, 3-5
Toggle selection, 3-5
Controller conventions, 3-6
Parameter selected for display, 3-6
Parameter selected for setting, 3-6
Map range control, 3-6
MFD system displays, 4-86
Electrical (ELEC) system, 4-93
Battery voltages and operating temperatures, 4-94
Bus voltage and amperage loads, 4-94
DC power interconnect buses, 4-94
Environmental control system/anti--ice, 4-88
Cabin and cockpit (CKPT) temperature, 4-89
Engine output bleed air temperature, 4-89
Oxygen pressure, 4-89
Fuel system, 4-90
Fuel boost pump status, 4-91
Fuel tank temperature, 4-91
Fuel used, 4-91
Tank fuel quantity, 4-90
Total fuel quantity, 4-91
Hydraulic (HYD) system, 4-92
Brake temperature, 4-93
Electric hydraulic pump status, 4-93
HYD pressure and fluid quantity, 4-92
Takeoff (T/O), 4-86
Aircraft door status, 4-87
Anti--ice (REF A--ICE) status annunciation, 4-87
Engine oil level, 4-87
Engine T/O mode, 4-87
Engine T/O status arrow, 4-88
Reference T/O temperature (REF TO TEMP), 4-87
Index A28--1146--112--00
Index--10
 PRIMUSr1000 Integrated Avionics System

Index (cont)
Modes of operation, 8-1, 8-2
Altitude hold (ALT HOLD) mode, 8-37
Engagement procedure, 8-37
Altitude preselect (ASEL) mode, 8-32
Engagement procedure, 8-32
Mode procedure, 8-34
Back course (BC) mode, 8-17
Flight procedure, 8-17
Flight level change (FLC) mode, 8-31
Descend and climb parameters, 8-31
Flight management system (FMS) navigation mode (option), 8-10
Engagement procedure, 8-10
Go--around (GA) mode (wings level), 8-40
Heading hold mode, wings level, 8-1
Heading select mode, 8-3
Instrument landing system (ILS) approach mode, 8-20
Procedure, 8-21
Localizer (LOC) mode, 8-12
Engagement procedure, 8-12
Pitch hold mode, 8-25
AP engaged, 8-25
AP not engaged, 8-25
Roll hold mode, 8-2
Speed hold mode, 8-28
Engagement procedure, 8-30
TakeOff mode, 8-26
VOR approach (VAPP) mode, 8-9
Engagement procedure, 8-9
VOR navigation (NAV) mode, 8-4
Engagement procedure, 8-4
Windshear (WDSHEAR) mode, 8-38
Multifunction display (MFD), 4-68
Changing a waypoint, 4-80
Designating a new waypoint procedure, 4-81
Checklist (option), 4-82
Map display, 4-72
Aircraft symbol, 4-73
Drift bug, 4-73
FMS lateral deviation digital display, 4-73
Heading display, 4-72
Heading reference line, 4-72

A28--1146--112--00 Index
Index--11
PRIMUSr1000 Integrated Avionics System

Index (cont)
Heading select bug, 4-73
Heading source annunciations, 4-73
Navigation symbols, 4-74
Pilot designator, 4-73
Range rings, 4-73
Map failure and warning displays, 4-75
EICAS failure message, 4-75
FMS failures, 4-76
Heading failure and flags, 4-75
MADC failures, 4-75
Weather radar failure, 4-76
MFD common symbols, 4-69
Mid--left side, 4-70
Mid--right side, 4-71
Upper left area, 4-70
Upper right area, 4-69
MFD display window, 4-82
MFD plan mode, 4-77
Plan mode failure and warning displays, 4-79
FMS failure, 4-79
Heading failures and flags, 4-79
MADC failures, 4-79
WX (weather radar) failure, 4-79
Traffic alert and collision avoidance system (TCAS), 4-84
2 NM range ring, 4-85
No bearing target readout, 4-85
Proximity advisory arrow, 4-85
TCAS altitude display submodes, 4-85
TCAS auto, 4-85
TCAS mode and failure annunciations, 4-84
TCAS range ring, 4-85
Traffic symbols, 4-85
Typical map mode presentations, 4-76
Typical plan mode presentation, 4-80
Weather, 4-74
Multifunction display (MFD) bezel controller, 4-6
Checklist menu, 4-15
Cursor movement/action, 4-16
ENT button, 4-17
Joystick, 4-16
LN BK (line back) button, 4-16

Index A28--1146--112--00
Index--12
 PRIMUSr1000 Integrated Avionics System

Index (cont)
M/P RNG button, 4-17
PAG (page) button, 4-16
RCL button, 4-16
RTN button, 4-16
SKP button, 4-16
M/P RNG (map/plan range) rotary knob, 4-6
Map/plan, 4-19
MENU INOP (inoperative), 4-6
Menu inop display, 4-6
MFD main menu, 4-8
MFD menu, 4-10
Joystick submenu, 4-13
Main menu (with FMS), 4-11
Main menu (without FMS), 4-10
Main menu -- MFD button, 4-10
Vspeeds submenu, 4-11
MFD menu tree, 4-7
System menu, 4-8
ECS (environmental control system) A/I (anti--ice) button, 4-9
ELECT (electrical) button, 4-10
FUEL button, 4-9
HYD (hydraulic) button, 4-9
RTN (return) button, 4-9
T/O (takeoff) button, 4-9
TCAS, 4-17
WX (weather), 4-18
Multifunction display (MFD) checklist upload procedure, 9-11
Error code 500, 9-12
Troubleshooting procedure, 9-13
Error code 5005, 9-13

N
Navigation receiver operation, C--17
Basic NAV tuning, C--17
DME hold, C--18
DME function key operating procedure, C--20
Single DME installation, C--18

O
Optional Equipment, 1-3

A28--1146--112--00 Index
Index--13
PRIMUSr1000 Integrated Avionics System

Index (cont)

PFD caution and failure display, 4-61

Caution (amber) annunciations, 4-61
Autopilot status annunciations, 4-63
Common symbol generator, 4-61
Course deviation failure, 4-62
Course select failure, 4-62
Distance display failures, 4-62
Flight director (FD) failure, 4-62
Maximum/minimum speed (MAX/MIN SPD), 4-64
Radio altimeter failure, 4-62
Same air data source, 4-61
Same attitude source, 4-61
TCAS messages, 4-64
Vertical deviation failure, 4-62
Windshear annunciation (WDSHEAR), 4-64
Yaw damper status annunciations, 4-63
Excessive attitude declutter, 4-66
Parameters, 4-67
Failure (RED) annunciations, 4-65
Attitude reference system failure, 4-65
Heading (HDG) select failure, 4-66
Micro air data computer (MADC) failures, 4-65
PFD test mode, 4-67
Pilot self--test, A--9
Pilot initiated self--test annunciations, A--9
Pilot writeup, 9-1
Commonly used terms, 9-3
Definitions of terms, 9-3
Typical problems, 9-4
Combined vertical and lateral mode problems, 9-9
Lateral mode problems, 9-4
Vertical mode problems, 9-7
Pitch hold mode, 8-25
AP engaged, 8-25
AP not engaged, 8-25
Power--up test, A--8
Pilot self--test annunciations, A--8

Index A28--1146--112--00
Index--14
 PRIMUSr1000 Integrated Avionics System

Index (cont)
Primary engine displays, 5-7
Cabin pressure and auxiliary power unit (APU) status, 5-18
APU exhaust temperature, 5-19
APU status, 5-18
APU turbine speed as a percentage of RPM, 5-19
Cabin pressurization, 5-18
Color differential pressure display, 5-18
Digital turbine (N2) display, 5-12
Engine ignition (IGN) annunciation, 5-11
Ignitor annunciation color definition, 5-11
Engine vibration display, 5-14
Fan (N1) speed display, 5-9
Automatic takeoff thrust control system (ATTCS), 5-10
FADEC in control, 5-10
FADEC mode annunciations, 5-9
FADEC N1 bug and digital target, 5-9
Fan digital display, 5-9
N1 analog scale, 5-9
N1 request bug , 5-9
Reverse thrust annunciation, 5-10
Fuel flow (FF PPH FLOW) digital display, 5-13
Fuel quantity (FUEL LB) digital display, 5-13
Gear/flaps/spoiler positions, 5-15
Flap position, 5-15
Landing gear position, 5-15
Spoiler position, 5-16
Interstage turbine temperature (ITT) display, 5-11
ITT analog scale, 5-11
ITT digital display, 5-11
Oil pressure (OIL PSI) display, 5-12
Oil temperature (OIL °C) display, 5-14
Pitch/roll/yaw trim display, 5-16
Pitch trim, 5-16
Roll trim, 5-17
Yaw trim, 5-17
Primary engine displays failure indications, 5-20
Data acquisition unit (DAU) failure, 5-20
Engine instrument test, 5-21
Engine sensor/interface failure, 5-20

A28--1146--112--00 Index
Index--15
PRIMUSr1000 Integrated Avionics System

Index (cont)
Primary flight display (PFD) bezel controller, 4-5
Airspeed display, 4-45
Airspeed rolling digit display, 4-47
Airspeed trend vector, 4-46
Airspeed/mach reference bug and display, 4-45
Indicated Airspeed (IAS) analog scale, 4-46
Indicated airspeed comparison monitor, 4-47
Low speed awareness (LSA), 4-47
Mach display, 4-47
Takeoff Vspeed display, 4-47
Vmo overspeed bar, 4-45
Altimeter display, 4-48
Altitude alert preselect display, 4-49
Altitude analog scale, 4-49
Altitude digital display, 4-49
Altitude reference line, 4-50
Altitude select bug, 4-50
Altitude trend vector, 4-50
Barometric altimeter display, 4-49
Low altitude awareness symbol, 4-49
ARC display, 4-41
Approach and terminal annunciators, 4-42
DME hold, 4-44
FMS heading (FHDG) annunciation (if available), 4-44
FMS status annunciation, 4-41
FMS to waypoint, 4-44
Heading source annunciations, 4-41
Off scale arrows, 4-44
Weather, 4-42
Weather radar annunciations, 4-43
Weather radar range ring value, 4-42
Weather radar TGT/VAR annunciations, 4-44
Weather returns, 4-42
BARO (barometric) SET knob, 4-5
Full compass display, 4-37
Bearing Pointers with annunciation, 4-38
Course pointer with digital display, 4-39
Drift angle bug, 4-38
Groundspeed, time--to--go, elapsed time, 4-39
Heading select bug with digital display, 4-37

Index A28--1146--112--00
Index--16
 PRIMUSr1000 Integrated Avionics System

Index (cont)
NAV course deviation indicator (CDI) with TO/FROM
indication, 4-40
NAV source distance, 4-40
Navigation (NAV) source annunciations, 4-37
Navigation symbol colors, 4-40
Wind vector (available if FMS is installed), 4-40
Functional divisions, 4-25
ADI (attitude director indicator) display, 4-25
Air Data displays, 4-26
HSI (horizontal situation indicator) display, 4-25
Mode annunciations, 4-25
Slip--skid display, 4-26
IN/HPA button, 4-5
Inclinometer, 4-5
PFD attitude director indicator (ADI), 4-26
Aircraft symbol, 4-31
Airspeed trend vector, 4-34
Altitude trend vector, 4-32
Attitude source annunciation, 4-34
Autopilot status annunciations, 4-32
CAT2 (category 2) ILS annunciations, 4-34
CAT2 deviation scale and pointer, 4-36
Command bar display logic (normal mode), 4-30
Comparison monitor annunciation, 4-33
Flight director couple arrow, 4-31
Flight director mode annunciations, 4-26
Low altitude awareness display, 4-31
Low bank arc, 4-31
Marker beacon, 4-33
MAX SPD/MIN SPD or underspeed warning, 4-34
Pitch limit indicator, 4-29
Pitch scale, 4-29
Radio altitude (RA) decision height annunciator, 4-33
Radio altitude decision height set data, 4-33
Radio altitude display, 4-33
Roll scale and roll pointer, 4-29
Source annunciations, 4-31
Symbol generator (SG) source, 4-34
Vertical deviation scale, 4-30
Vertical track alert (VTA) annunciation, 4-32
Windshear (WDSHEAR) annunciation, 4-34

A28--1146--112--00 Index
Index--17
PRIMUSr1000 Integrated Avionics System

Index (cont)
Yaw damper annunciations, 4-26
STD (standard) button, 4-5
Vertical speed (VS) display, 4-51
Flight director VS target display and bug, 4-51
TCAS II resolution advisory display, 4-52
TCAS status message, 4-52
Vertical speed analog scale, 4-52
Vertical speed digital display, 4-52
PRIMUSâ 660 Weather Radar System, 2-6, B--1, B--2
Description, B--1
Introduction, B--1
Maximum permissible exposure level (MPEL), B--12
Tilt management
Radar beam illumination high altitude -- 12--inch radiator,
B--11
Radar beam illumination low altitude --12--inch radiator, B--11
Weather radar controller, B--2
Controls and indicators, B--2
Normal operation, B--8
PRIMUSâ 880 Weather Radar System (Option), 2-8
PRIMUSâ II Integrated Radio System, 2-7
PRIMUSâ II Radio System, C--1, C--5, C--6
Audio panel operation, C--33
Controls, C--33
General, C--33
Introduction, C--1
Radio management unit (RMU) operation, C--7
ADF operation, C--24
ATC transponder and TCAS operations, C--20
Built--in--test (BIT), C--28
Controls, C--8
Cross--side operation, C--27
General, C--7
Memory pages, C--11
Navigation receiver operation, C--17
System (SYS) on--off page, C--11
VHF communications transceiver operation, C--13
System description, C--5
Audio Panel, C--6
General, C--5
Integrated communication unit, C--6

Index A28--1146--112--00
Index--18
 PRIMUSr1000 Integrated Avionics System

Index (cont)
Integrated navigation unit, C--5
Radio management unit (RMU), C--6
Tuning backup control head, C--7
Tuning backup control head operation, C--35
Controls and annunciations, C--37
Emergency mode, C--36
General, C--35
Normal mode, C--36
Procedures
ADF PAST procedure, C--32
Altitude hold mode engagement procedure, 8-37
Altitude preselect mode engagement procedure, 8-32
Altitude preselect mode procedure, 8-34
ATC transponder and TCAS PAST procedure, C--30
Back course flight procedure, 8-17
Designating a new waypoint procedure, 4-81
DME function key operating procedure, C--20
FMS mode engagement procedure, 8-10
Ground maintenance test procedure, 9-9
ILS approach mode procedure, 8-21
Localizer mode engagement procedure, 8-12
MFD checklist upload procedure, 9-11
Speed mode engagement procedure, 8-30
Troubleshooting procedure, 9-13
VHF COM PAST procedure, C--29
VOR approach mode engagement procedure, 8-9
VOR navigation mode engagement procedure, 8-4
VOR/ILS/marker/DME PAST procedure, C--31
Publication ordering information, 10-4

Radio altimeter system, 2-7

Radio management unit (RMU), C--6
ADF operation, C--24
ADF operation with HF transmitters, C--26
Controls, C--25
General, C--24
Single ADF installation, C--27
ATC transponder and TCAS operations, C--20
ATC failure warnings, C--24

A28--1146--112--00 Index
Index--19
PRIMUSr1000 Integrated Avionics System

Index (cont)
Controls and annunciations, C--20
General, C--20
Built--in--test (BIT), C--28
ADF PAST, C--32
ATC transponder and TCAS PAST, C--30
General, C--28
Pilot--activated self--test (PAST), C--28
VHF COM past, C--29
VOR/ILS/marker/DME PAST, C--31
Cross--side operation, C--27
General, C--7
Maintenance, C--32
Memory pages, C--11
Navigation receiver operation, C--17
Basic NAV tuning, C--17
DME hold, C--18
Single DME installation, C--18
VHF communications transceiver operation, C--13
Basic COM tuning, C--13
COM messages, C--15
Radio management unit (RMU) operation, C--7
Controls, C--8
Cursor, C--8
Function keys, C--9
Line select keys, C--8
Transfer key, C--8
Tuning knobs, C--8
System (SYS) on--off page, C--11
Remote autopilot switches and annunciators, 6-5
AP disconnect switch, 6-5
Go--around button, 6-5
Pitch wheel and Turn knob controls, 6-5
TCS button, 6-5
Remote mounted equipment, 1-2
Reversionary controllers and dimming panel, 3-7
Data acquisition unit (DAU) reversionary panel, 3-8
EDS dimming panel, 3-8
System reversionary panel, 3-7
ADC button, 3-7
AHRS button, 3-8
MDF knob, 3-7

Index A28--1146--112--00
Index--20
 PRIMUSr1000 Integrated Avionics System

Index (cont)
SG (symbol generator) button, 3-8
Reversionary modes, EICAS, 5-30
Loss of display guidance computer (DGC) symbol generator, 5-30
Loss of the display unit, 5-30
Roll hold mode, 8-2

Speed hold mode, 8-28

Engagement procedure, 8-30
Standby navigation display, 4-96
RMU backup navigation display, 4-98
ADF bearing pointer, 4-98
ADF pointer out of view right/left, 4-99
Aircraft symbol, 4-99
Course pointer with display, 4-100
Digital ADF bearing, 4-100
Digital heading, 4-98
Digital VOR bearing, 4-100
DME distance, 4-99
Glideslope pointer, 4-99
Glideslope scale, 4-99
Heading arc compass rose, 4-100
Heading compass, 4-99
Lateral deviation indicator, 4-100
Lateral deviation scale, 4-99
Marker beacon, 4-99
Selected course, 4-100
Standby navigation display failure indications, 4-100
TO/FROM indicator, 4-99
Tune box, 4-100
VOR bearing pointer, 4-98
VOR pointer out of view right/left, 4-99
VOR/ILS and ADF frequency display, 4-98
Supplemental pilot’s manuals, 1-1
System (SYS) on--off page, C--11
System description
Air data system (ADS), 2-5
Attitude heading and reference system (AHRS), 2-6
Electronic display system (EDS), 2-4
Flight guidance system (FGS), 2-3

A28--1146--112--00 Index
Index--21
PRIMUSr1000 Integrated Avionics System

Index (cont)
Flight management system (FMS) (Option), 2-9
Other switches and controls, 2-9
AP (autopilot) disconnect switch, 2-9
External AHRS switches, 2-9
GA (go--around) button, 2-9
Joystick (option), 2-10
Master warning/master caution switches, 2-10
N2 override switch, 2-10
TCS (touch control steering) button, 2-9
PRIMUSâ 660 Weather Radar System, 2-6, B--1
PRIMUSâ 880 Weather Radar System (Option), 2-8
PRIMUSâ II Integrated Radio System, 2-7, C--1
Radio altimeter system, 2-7
Traffic alert and collision avoidance system (TCAS), 2-8
System description, 2-1, 2-2
System display failure annunciation, 4-95
Display controller failures, 4-95
Both pilot’s and copilot’s DC fail, 4-96
Copilot’s DC fails, 4-96
Pilot’s DC fails, 4-95
Display system reversionary modes, 4-95
EFIS 1 (pilot normal) failure, 4-95
EFIS 2 (copilot normal) failure, 4-95
System file, 1-1, 1-2
System limits, 7-1, 7-2
Glossary of terms, 7-1
Attitude director indicator (ADI) command cue, 7-1
Glideslope (GS) gain programming as a radio altitude (RA)
function, 7-2
GS capture (CAP), 7-2
GS gain programming as a RA and vertical speed (VS)
function, 7-2
Lateral beam sensor (LBS), 7-2
Lateral gain programming, 7-3
Localizer/back course (BC) track, 7-3
True airspeed (TAS) gain programming, 7-3
Vertical beam sensor (VBS), 7-4
VOR after over station sensor (AOSS), 7-4
VOR capture (CAP), 7-4
VOR over station sensor (OSS), 7-4
System performance and operating limits, 7-5

Index A28--1146--112--00
Index--22
 PRIMUSr1000 Integrated Avionics System

Index (cont)

TakeOff mode, 8-26

Vertical speed hold mode, 8-27
Tilt management
Radar beam illumination high altitude -- 12--inch radiator, B--11
Radar beam illumination low altitude --12--inch radiator, B--11
Traffic alert and collision avoidance system (TCAS), 2-8
Troubleshooting, 9-1, 9-2
Ground maintenance test, 9-9
Test procedure, 9-9
Multifunction display (MFD) checklist upload procedure, 9-11
Error code 500, 9-12
Error code 5005, 9-13
Troubleshooting procedure, 9-13
Pilot writeup, 9-1
Commonly used terms, 9-3
Typical problems, 9-4
Tuning backup control head, C--7
Controls and annunciations, C--37
Emergency mode annunciator, C--37
NAV AUDIO On/Off switch, C--38
Navigation audio (NAV AUDIO) on annunciator, C--37
Norm/emergency modes switch, C--38
Radio tuning annunciators, C--38
Remote tune annunciator, C--37
Squelch (SQ) annunciator, C--37
Squelch on/off switch, C--38
System installation annunciator, C--38
Transfer key, C--38
Transmit (TX) annunciator, C--37
Tuning cursor, C--37
Emergency mode, C--36
General, C--35
Normal mode, C--36
Tuning backup control head operation, C--35
Typical PFD presentations, 4-53
Approach capture tracking at DH, 4-58
Climb to initial altitude, 4-54
Comparison monitoring, 4-59
Enroute cruise, 4-55
A28--1146--112--00 Index
Index--23
PRIMUSr1000 Integrated Avionics System

Index (cont)
Setup for approach, 4-56
Takeoff (TO) mode, 4-53
Terminal area, 4-57

Vertical speed hold mode, 8-27

VHF communications transceiver operation, C--13
Basic COM tuning, C--13
COM messages, C--15
VOR approach (VAPP) mode, 8-9
Engagement procedure, 8-9
VOR navigation (NAV) mode, 8-4
Engagement Procedure, 8-4

Weather radar controller, B--2

Controls and indicators, B--2
GAIN switch, B--7
MODE switch, B--5
Range switches, B--2
RCT (rain echo attenuation compensation technique)
switch, B--3
SECT (sector) switch, B--4
SLV (slaved) annunciator, B--4
STAB (stabilization), B--3
Target alert characteristics, B--4
TGT (target) button, B--3
TILT knob, B--4
Normal operation, B--8
Powerup, B--9
Precautions, B--8
Preliminary control settings, B--8
Tilt management, B--11
Windshear (WDSHEAR) mode, 8-38

Index A28--1146--112--00
Index--24