Avionic Emb 145-3

Business and Commuter Aviation Systems Honeywell Inc. Box 29000 Phoenix, Arizona 85038

PRIMUS® 1000 Integrated Avionics System

Embraer 145

System Maintenance Manual Volume III — System Test and Fault Isolation

22-05-14 TITLE PAGE T-1 PRINTED IN U.S.A.

PUB. NO. A15-1146-065

1 NOVEMBER 1996

SYSTEM MAINTENANCE MANUAL EMBRAER 145

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.

S96

22-05-21 TITLE PAGE T-2 Copyright 1996 Honeywell Inc. All Rights Reserved

1 SEPTEMBER 1996


RECORD OF REVISIONS – VOLUME III Upon receipt of a revision, insert the latest revised pages and dispose of superseded pages. Enter revision number and date, insertion date, and the incorporator’s initials on the Record of Revisions. The typed initials HI are used when Honeywell Inc. is the incorporator. Revision Number

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By

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Revision Date

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


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SECTION 7

SYSTEM TEST AND FAULT ISOLATION This section provides a Ground Maintenance Test (GMT) in Table 7-1 for checking the units of the PRIMUS® 1000 system for correct installation and proper operation. Table 7-1 also contains fault verification and corrective action procedures. NOTE:

The procedures provided in Table 7-1, Ground Maintenance Test (GMT) Procedures, are based on Engineering Bulletin EB7023389, Rev -.

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Table 7-1. Ground Maintenance Test (GMT) Procedures

TABLE OF CONTENTS Para

Title

Page

SECTION I IC-600 SELF TEST ............................................................................................... 7-9 1. INTRODUCTION ............................................................................................................................ 7-9 1.1 SCOPE .............................................................................................................................................. 7-9 1.2 REFERENCE DOCUMENTS .................................................................................................................... 7-9 1.3 TERMS AND DEFINITIONS ..................................................................................................................... 7-9

2. SYSTEM BENCH REQUIREMENTS ................................................................................................. 7-10 3. T EST M ODES ............................................................................................................................. 7-10 3.1 IN-AIR-INITIATED TEST ...................................................................................................................... 7-10 3.2 ON-GROUND INITIATED TEST ............................................................................................................. 7-12 3.2.1 FAILURE M ODE ANNUNCIATION AND FAMILIARIZATION .................................................................. 7-12 3.2.1.1 EICAS TEST ............................................................................................................. 7-16 3.2.2 INTERNAL BUILT-IN TEST ......................................................................................................... 7-18 3.3 M AINTENANCE TEST M ODE ................................................................................................................ 7-20 3.3.1 INDEX PAGES ......................................................................................................................... 7-23 3.3.1.1 INDEX PAGE 1 OF 4 (RA SETTING 600) ........................................................................ 7-23 3.3.1.2 INDEX PAGE 2 OF 4 (RA SETTING 610) ........................................................................ 7-24 3.3.1.3 INDEX PAGE 3 OF 4 (RA SETTING 620) ........................................................................ 7-25 3.3.1.4 INDEX PAGE 4 OF 4 (RA SETTING 630) ........................................................................ 7-26 3.3.2 HARDWARE PAGES ................................................................................................................. 7-27 3.3.2.1 HW ID 1 (RA SETTING 640)....................................................................................... 7-27 3.3.2.2 HW 2 (RA SETTING 650) ........................................................................................... 7-28 3.3.3 DISPLAY UNIT PAGES .............................................................................................................. 7-29 3.3.3.1 DISPLAY UNITS - PILOT’S SIDE (RA SETTING 660) ......................................................... 7-29 3.3.3.2 DISPLAY UNITS - COPILOT’S SIDE (RA SETTING 660C) .................................................. 7-30 3.3.4 EVENT CODES (RA SETTING 670) ............................................................................................ 7-31 3.3.5 POWER AND INTERNAL TEMPERATURE PAGE (RA SETTING 680) ................................................... 7-33 3.3.6 AHRS DATA (RA SETTING 690) .............................................................................................. 7-34 3.3.7 DC VARIABLES (RA SETTING 700) ........................................................................................... 7-35 3.3.8 M ICRO AIR DATA COMPUTER (RA SETTING 710) ........................................................................ 7-36 3.3.9 DC-550 PORT (RA SETTING 720) ............................................................................................ 7-38 3.3.10 WEATHER RADAR (RA SETTING 730) ...................................................................................... 7-41 3.3.11 DISCRETES 1 (RA SETTING 740) ............................................................................................ 7-42 3.3.12 DISCRETES 2 (RA SETTING 750) ............................................................................................ 7-43 3.3.13 DISCRETES 3 (RA SETTING 760) ............................................................................................ 7-44 3.3.14 DISCRETES 4 (RA SETTING 770) ............................................................................................ 7-45 3.3.15 DISCRETES 5 (RA SETTING 780) ............................................................................................ 7-46 3.3.16 DISCRETES 6 (RA SETTING 790) ............................................................................................ 7-47 3.3.17 DISCRETES 7 (RA SETTING 800) ............................................................................................ 7-48 3.3.18 DISCRETES 8 (RA SETTING 810) ............................................................................................ 7-49 3.3.19 DISCRETES 9 (RA SETTING 820) ............................................................................................ 7-50 3.3.20 LAMP TEST 9 (RA SETTING 830)............................................................................................ 7-51

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Table 7-1. Ground Maintenance Test (GMT) Procedures 3.3.21 3.3.22 3.3.23 3.3.24 3.3.25 3.3.26 3.3.27 3.3.28

ARINC 1 (RA SETTING 840) ................................................................................................. 7-53 ARINC 2 (RA SETTING 850) ................................................................................................. 7-54 ARINC 3 (RA SETTING 860) ................................................................................................. 7-55 ARINC 4 (RA SETTING 870) ................................................................................................. 7-56 RADIO SYSTEM BUS (RA SETTING 880) .................................................................................. 7-57 CHECKLIST LOADING (RA SETTING 890) ................................................................................. 7-58 SYSTEM CONFIGURATION (RA SETTING 900)............................................................................ 7-59 BLANK PAGE (RA SETTING 910) ............................................................................................ 7-59

4. EVENT CODES ........................................................................................................................... 7-60

SECTION II STANDARD GROUND TEST PROCEDURE .............................................................. 7-73 1. INTRODUCTION .......................................................................................................................... 7-73 2. DISCRETE INPUT VERIFICATIONS ................................................................................................. 7-73 3. DU-870 (DISPLAY UNIT) ............................................................................................................ 7-74 4. DC-550 DISPLAY CONTROLLER .................................................................................................. 7-74 5. IC-600 INTEGRATED COMPUTER ................................................................................................. 7-75 6. AZ-850 M ICRO AIR DATA COMPUTER ......................................................................................... 7-75 7. AH-800 ATTITUDE AND HEADING REFERENCE UNIT ...................................................................... 7-75 8. PRIMUS II RADIOS ................................................................................................................... 7-75 9. AA-300 RADIO ALTIMETER ......................................................................................................... 7-76 10. DAU DATA ACQUISITION UNIT .................................................................................................. 7-76 11. TCAS T RAFFIC COLLISION AVOIDANCE SYSTEM ........................................................................ 7-76 11.1 PROCEDURES ................................................................................................................................ 7-76

12. PRIMUS 650/660 W EATHER RADAR ........................................................................................ 7-78 12.1 PROCEDURES ................................................................................................................................ 7-78

13. GC-550 F LIGHT GUIDANCE CONTROLLER .................................................................................. 7-80 14. PC-400 AUTOPILOT CONTROLLER ............................................................................................ 7-80 15. SM-200 .................................................................................................................................. 7-80 16. AUTOPILOT AND ELECTRIC T RIM ............................................................................................... 7-81

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SECTION III EXPANDED GROUND TEST PROCEDURE ............................................................. 7-83 1. INTRODUCTION .......................................................................................................................... 7-83 2. PRIMUS 650/660 W EATHER RADAR SYSTEM CHECK .................................................................. 7-83 2.1 PRELIMINARY CONTROL SETTINGS ...................................................................................................... 7-83 2.2 PRECAUTIONS .................................................................................................................................. 7-83 2.3 TEST PATTERN ............................................................................................................................... 7-83 2.4 STABILIZATION INPUT ........................................................................................................................ 7-84 2.4.1 PRELIMINARY CHECKS ............................................................................................................. 7-84 2.4.2 ELEVATION FEEDBACK CHECK/ADJUSTMENT .............................................................................. 7-84 2.4.3 PITCH GAIN ADJUSTMENT/CHECK ............................................................................................. 7-84 2.4.4 ROLL GAIN ADJUSTMENT/CHECK .............................................................................................. 7-84 2.5 FAULT M ONITORS ............................................................................................................................. 7-85

3. EFIS CHECK ............................................................................................................................. 7-85 3.1 PRELIMINARY CONTROL SETTINGS ...................................................................................................... 7-85 3.2 DU-870 DISPLAY UNITS ................................................................................................................... 7-85 3.3 DC-550 DISPLAY CONTROLLER ......................................................................................................... 7-86 3.4 DIMMING CONTROLLERS .................................................................................................................... 7-87 3.5 REVERSIONARY PANEL ..................................................................................................................... 7-87 3.6 MFD BEZEL BUTTONS AND ROTARY SET KNOB ................................................................................... 7-87

4. MADC CHECK .......................................................................................................................... 7-89 4.1 NORMAL M ODE ................................................................................................................................ 7-89 4.2 SELF-TEST M ODE ............................................................................................................................ 7-90

5. RADIO ALTIMETER T EST ............................................................................................................. 7-91 6. AHRS CHECK ........................................................................................................................... 7-91 6.1 AHRS POWER-UP CHECKOUT ........................................................................................................... 7-92 6.2 AHRS TEST M ODE .......................................................................................................................... 7-92 6.3 AHRS ATTITUDE OUTPUTS CHECK ..................................................................................................... 7-92 6.4 AHRS FLUX VALVE CHECK ............................................................................................................... 7-92 6.5 VERIFICATION OF M EMORY M ODULE PROGRAMMING .............................................................................. 7-93

7. PRIMUS II RADIOS CHECK ........................................................................................................ 7-95 7.1 BUS OPERATION CHECKOUT .............................................................................................................. 7-95 7.2 UNIT POWER ON SELF TEST (POST).................................................................................................. 7-97 7.3 AIRCRAFT INTERFACE/RAMP TESTS .................................................................................................... 7-98 7.3.1 COM 1 TESTS ....................................................................................................................... 7-98 7.3.2 TRANSPONDER 1 TESTS .......................................................................................................... 7-99 7.3.3 NAV 1 TESTS ...................................................................................................................... 7-100 7.3.3.1 VOR T ESTS ............................................................................................................ 7-100 7.3.3.2 LOC 1 TESTS ......................................................................................................... 7-101 7.3.3.3 GLIDE SLOPE 1 TESTS ............................................................................................. 7-101 7.3.3.4 M ARKER BEACON 1 TESTS ....................................................................................... 7-101 7.3.4 DME 1 TESTS ...................................................................................................................... 7-102 7.3.5 ADF 1 TESTS ...................................................................................................................... 7-102

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Table 7-1. Ground Maintenance Test (GMT) Procedures 7.3.6 RMU TESTS ........................................................................................................................ 7-103 7.3.6.1 DIMMING CHECK ...................................................................................................... 7-103 7.3.6.2 POWER ON/OFF CHECK .......................................................................................... 7-103 7.3.7 TUNING BACKUP CONTROL HEAD (CLEARANCE DELIVERY HEAD - CDH) TESTS ............................ 7-104 7.4 SYSTEM STATUS M ESSAGE CHECKS ................................................................................................. 7-104 7.5 IN-FLIGHT TESTS ........................................................................................................................... 7-105 7.5.1 AUDIO PANEL FLIGHT TEST PROCEDURES ................................................................................ 7-105

8. T RAFFIC COLLISION AVOIDANCE SYSTEM (TCAS) ...................................................................... 7-106 8.1 TCAS CHECK ............................................................................................................................... 7-106

9. EICAS CHECK ........................................................................................................................ 7-111 9.1 REVERSIONARY CONTROL ............................................................................................................... 7-111 9.2 CREW ALERTING SYSTEM ................................................................................................................ 7-112 9.2.1 TONE GENERATOR AND M ESSAGE ACKNOWLEDGMENT............................................................... 7-112 9.2.2 TAKE OFF INHIBIT FUNCTION ................................................................................................... 7-113 9.2.3 RMU ENGINE BACKUP .......................................................................................................... 7-113 9.2.4 RMU NAVIGATION BACKUP .................................................................................................... 7-113

10. FLIGHT DIRECTOR T EST ......................................................................................................... 7-114 11. AUTOPILOT T EST ................................................................................................................... 7-122

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List of Figures (Section I) Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure

3.1.1. DC-550 Display Controller .................................................................................. 7-10 3.1.2. PFD In-Air Initiated Test Mode ........................................................................... 7-11 3.2.1. PFD On-Ground Initiated Test Mode ................................................................... 7-13 3.2.2. MFD On-Ground Intitiated Test Mode ................................................................. 7-15 3.2.3. EICAS On-Ground Initiated Test Mode ............................................................... 7-17 3.2.4. Interface Self-Test Display ................................................................................. 7-18 3.3.1. Maintenance Display Formats on the DU-870...................................................... 7-22 3.3.2. Index Page 1 of 4 .............................................................................................. 7-23 3.3.3. Index Page 2 of 4 .............................................................................................. 7-24 3.3.4. Index Page 3 of 4 .............................................................................................. 7-25 3.3.5. Index Page 4 of 4 .............................................................................................. 7-26 3.3.6. Hardware Identification Page 1........................................................................... 7-27 3.3.7. Hardware Identification Page 2........................................................................... 7-28 3.3.8. Display Unit Page (Pilot’s Side) .......................................................................... 7-29 3.3.9. Display Unit Page (Copilot’s Side) ...................................................................... 7-30 3.3.10. Event Code Page ............................................................................................. 7-31 3.3.11. Power/Internal Temperature Page .................................................................... 7-33 3.3.12. AHRS Data Page ............................................................................................. 7-34 3.3.13. DC Variables Page........................................................................................... 7-35 3.3.14. Air Data Page .................................................................................................. 7-36 3.3.15. DC-550 Port Page ............................................................................................ 7-38 3.3.16. W eather Radar Page ....................................................................................... 7-41 3.3.17. Discretes 1 ...................................................................................................... 7-42 3.3.18. Discretes 2 ...................................................................................................... 7-43 3.3.19. Discretes 3 ...................................................................................................... 7-44 3.3.20. Discretes 4 ...................................................................................................... 7-45 3.3.21. Discretes 5 ...................................................................................................... 7-46 3.3.22. Discretes 6 ...................................................................................................... 7-47 3.3.23. Discretes 7 ...................................................................................................... 7-48 3.3.24. Discretes 8 ...................................................................................................... 7-49 3.3.25. Discretes 9 ...................................................................................................... 7-50 3.3.26. Lamp Test ....................................................................................................... 7-51 3.3.27. ARINC #1 ........................................................................................................ 7-53 3.3.28. ARINC #2 ........................................................................................................ 7-54 3.3.29. ARINC #3 ........................................................................................................ 7-55 3.3.30. ARINC #4 ........................................................................................................ 7-56 3.3.31. Radio System Bus............................................................................................ 7-57 3.3.32. Check Loading Page ........................................................................................ 7-58 3.3.33. System Configuration ....................................................................................... 7-59 3.3.34. Blank Page ...................................................................................................... 7-59

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List of Figures (cont) (Section II) Figure 11.1.1. TCAS Test Display........................................................................................... 7-77 Figure 12.1.1. PRIMUS® 650 MFD W eather Radar Test Mode ................................................ 7-78 Figure 12.1.2. PRIMUS® 660 MFD W eather Radar Test Mode ................................................ 7-79 (Section III) Figure 3.6.1. MFD Bezel Button Menu Tree............................................................................. 7-88 Figure 8.1.1. Antenna Test Points (Rack Side) ...................................................................... 7-107

List of Tables Table 4.1. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes That Prevent Autopilot Engagement)....................................................... 7-60 Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Continuous Event Codes That Disengage the AP and YD Autopilot Engagement) ........................... 7-66 Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes) ............................................................................................................. 7-68 Table 4.4. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Event Codes That Record An Action) ............................................................................................ 7-70 Table 4.5. Autopilot and Interloop Monitors Event Codes (Power Up Event Codes That Prevent Autopilot Engagement)...................................................................................... 7-71 Table 4.6. Autopilot and Interloop Monitors Event Codes (Continuous Event Codes That Prevent Autopilot Engagement)...................................................................................... 7-72 Table 4.7. Autopilot and Interloop Monitors Event Codes Execution Failure.............................. 7-72 Table 8.1. CU Fault Reporting and Corrective Actions................................................................... 7-108 Table 8.2. Antenna Wiring Resistance....................................................................................... 7-109 Table 8.3. CU Self-Test Execution............................................................................................ 7-110 Table 10.1. Flight Director Test ............................................................................................ 7-114 Table 11.1. Autopilot Test .................................................................................................... 7-122

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SECTION I IC-600 SELF TEST 1. 1.1

INTRODUCTION SCOPE This specification outlines the self-test capabilities of the IC-600. It decribes in detail how to execute each test from an aircraft and/or system bench. The available self-tests within the IC-600 include the In-Air-Initiated Test (Weight Off Wheels) and the On-Ground-Initiated Test (Weight On Wheels).

1.2

REFERENCE DOCUMENTS Embraer, EMB 145 System and Software Requirements Document, PS7021214

1.3

TERMS AND DEFINITIONS T ERM

D EFINITION

AC

Aircraft

ADC

Air Data Computer

ADI

Attitude Direction Indicator

BIT

Built In Test

DU

Display Unit

HSI

Horizontal Situation Indicator

ICB

Integrated Computer Bus

LRN

Long Range Nav

MFD

Multi-Function Display

RSB

Radio System Bus

RX

Receiver

TX

Transmitter

W OW

W eight On W heels

W SP

W ord Sequence Position

WX

W eather Radar

LRU

Line Replaceable Unit

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2.

SYSTEM BENCH REQUIREMENTS In order to perform the IC-600 test on the system bench, the following requirements must be met: • The 115 V ac 400 Hz bench power is ON. • The 28 V dc bench power is ON. • Factory test identification discretes are all open. (These discretes into the IC-600 should not be wired in the aircraft or system bench. They are intended for use in a factory repair environment only.) • The IC-600, displays, and remaining avionics must be powered on to optimize self test results.

3. 3.1

TEST MODES IN-AIR-INITIATED TEST An in flight (Weight Off Wheels) test may be initiated by pushing and holding the TEST button on the DC-550 (Figure 3.1.1). NOTE:

This test is inhibited when Glide Slope is captured.

The test results in the active FD mode annunciations on the PFD being replaced by a magenta “TEST” annunciation (Figure 3.1.2). The Radio Altimeter will also be commanded into the test mode resulting in the display of a fixed radio altitude of 100 ft. The MFD and EICAS will not change during this test. No other visible testing of the IC-600 will be executed during in this “in-air” test. However, background testing of the IC-600 is continually in process verifying such parameters as power supply levels, analog to digital functions, software/hardware validity, and other parameters.

FULL WX

GSPD TTG

ET

NAV

FMS

OFF NAV 1

ADF NAV 2

ADF

OFF

FMS

BRG

FMS

RA

TEST

BRG AD-50629-R1@

Figure 3.1.1. DC-550 Display Controller

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260

TEST AP

YD

20

20

10

10

145 00

280

260

M A X

251

240 24 9

2 R 1

220

200

A O A

14500 20

14300 80

S P D

10

ATT2

20

.410 M

10 350

N

359 CRS

14000 M

3000 3

3

2

E

W

6

S

21 TGT

0 1

15

HDG 001

1

12

24

25 VOR1 ADF2

29.92 IN

200 RA

30

33

MIN

20

MAG2

13.1 NM

FMS

2

TTG 5MIN

3

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS AD-51487@

Figure 3.1.2. PFD In-Air Initiated Test Mode

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3.2

ON-GROUND INITIATED TEST The IC-600 On-Ground (Weight On Wheels) initiated test has two modes of operation. These modes are: • Failure Mode Annunciation and Familiarization • Initiated Test Mode

3.2.1

F AILURE M ODE ANNUNCIATION AND FAMILIARIZATION To initiate the test, push and hold the TEST button on the DC-550 (Figure 3.1.1). Failure mode annunciation and familiarization will occur for the first 5 to 6 seconds. NOTE:

This test is only functional on the ground. The radio altimeter test is functional at all times except during GS CAP.

A magenta “TEST” will be displayed in the lateral arm location (top left center) as shown in Figure 3.2.1. The PFD test page format displays the following quantities as invalid: • Pitch and Roll Attitude • Vertical Deviation (Pointer and Scale) • Barometric Correction • Vertical Speed Set Digital Readout • Altitude Preselect Digital Indicator • Heading (Scale and digital readout) • Lateral deviation (Pointer and Scale) • Distance Digital Readout • GSPD or TTG or Elapsed Time • Selected course or desired track (pointer and digital readout) • Selected heading (digital) • Airspeed Indicator • Mach Digital Readout • Airspeed Set Digital Readout • Altitude Indicator (Tape and Digits) • Vertical Speed Indicator

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--I A S

TEST

WDSHEAR AP

A L T

ATT FAIL

ATT

A O A RA .- - - M ILS CAS MSG

1610 O

I

M

800 RA HDG

3

33

6 1

24

12

V S

0 1

15 S

21

HDG ---

E

W

- - . - NM

VOR1 VOR2

3

2

HDG FAIL

30

- -.- - IN TCAS FAIL

N

- - - CRS ILS 1

-----

YD

2

TTG - - -MIN

3

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS AD-51488@

Figure 3.2.1. PFD On-Ground Initiated Test Mode

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Table 7-1. Ground Maintenance Test (GMT) Procedures The following will be removed from the test page display: • All bugs (airspeed, altitude, heading, drift, vertical speed) • Flight Director information (command bars/cue, modes, FPA symbology) • All pointers (bearing, To/From) • Low Speed Awareness • Vspeed bugs and digital readouts • Vmo/Mmo • Airspeed trend vector • Altitude trend vector The following display items will be forced on: • All comparison monitors • All three marker beacons • Windshear warning annunciation If heading is valid upon entering this test mode, the Heading source annunciation shall remain a valid annunciation - (DG1or DG2) or (MAG1 or MAG2) - and will NOT be forced to (HDG1or HDG2) as it normally would be when heading is invalid. MFD Test Mode will always be displayed in MAP mode (Figure 3.2.2). The MFD test page format displays the following quantities as invalid: • Heading • Weather Radar Tilt • True Airspeed • Static Air Temp • Total Air Temperature • Nav Source distance to “TO” Waypoint • Time to “TO” Waypoint • Bezel Menu

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----

FMS

- - - SAT - - - TAT

- -. - NM ---

- - - TAS

MIN

N

33

3

6

30

HDG FAIL

12

W 24

50

E

50

WX

21

S

- - - QT

DOORS

--------

REF TO TEMP: REF A-ICE:

15

ENGINE T/O MODE:

---

C

--OIL LVL

- - - QT

MENU INOP

AD-51489@

Figure 3.2.2. MFD On-Ground Intitiated Test Mode

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Table 7-1. Ground Maintenance Test (GMT) Procedures The following symbology will be removed from the display: • Heading Source Annunciation • TCAS, Checklist or System Pages • Weather Patch/Boundary Line • Drift Bug • Wind Vector • Heading Select Bug • Flight Plan Data • Airports • Navaids • Designator Information 3.2.1.1

EICAS Test A magenta “TEST” will be displayed on the PFD to signify that the On-Ground Test has been initiated (Figure 3.2.1). The EICAS test page format displays the following quantities as invalid as shown in Figure 3.2.3: • Landing Gear Status • N1 • ITT • N2 • Fuel Flow and Quantity • Oil Pressure and Temperature. • Vibration for Low and High Pressure Rotors • Flaps • Spoilers • All Cabin and APU Parameters • All trim values The CAS message field will be filled with an X as shown in Figure 3.2.3. The following symbology will be removed from the display: • Reversion Annunciation • Ignition Annunciation • FADEC in Control Annunciation • All Engine Bugs • All Trim Bugs

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------

------

ATTCS

---.-

---.-

N1

A A

IGN A

IGN

----

A

----

ITT

%

N2

----

PPH

FF

---.----

----

LB

FQ

----

% PPH LB

VIB

OIL

LDG GEAR

FLAPS

--- ---

--- --TEMP

PRESS

CAB ALT CAB

P

CAB RATE APU

---

%

LP

------.-----

PSI

---

C

FT

FPM

SPLRS

HP

CLD

ROLL

PITCH

YAW

--

AD-51490@

Figure 3.2.3. EICAS On-Ground Initiated Test Mode

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Table 7-1. Ground Maintenance Test (GMT) Procedures 3.2.2

INTERNAL BUILT-IN T EST The internal built-in test (IBIT) executes 5 - 7 seconds after the Failure Mode Annunciation and Familiarization if the TEST button remains pushed. The test exercises the IC-600s internal interfaces using internal wraparounds. The test remains active as long as the TEST button is pushed and the PFD display format will cycle (every 10 seconds) between the IBIT1 and IBIT2 test results. A short description of the hardware tested, followed by the status annunciation, will be displayed as shown in Figure 3.2.4. Releasing the DC-550 test button terminates the test and resumes normal operation.

I C - 6 0 0

I B I T

W A I T

A U T O P I L O T

P A S S

F L I G H T

P A S S

D I R E C T O R

S Y M B O L

P A S S

G E N E R A T O R

D I S P L A Y

U N I T

P A S S

A T T / H D G

I N T E R F A C E

P A S S

R A D

I N T E R F A C E

P A S S

A L T

R A D I O

D I S P L A Y L A M P A D C

P A S S

I N T E R F A C E C O N T R O L L E R

D R I V E R S

P A S S P A S S P A S S

D A T A

AD-51492@

Figure 3.2.4. Interface Self-Test Display As each segment of hardware is tested the status is annunciated: ANNUNCIATION

C OLOR

D ESCRIPTION

“VALD”

Green

Valid Data

“INVD”

Red

Invalid data is detected

“PASS”

Green

Successful end-around hardware test

“FAIL”

Red

Unsuccessful end-around hardware test

“N/A”

W hite

Not applicable/Not installed. Determined through the configuration discretes

A PASS/FAIL status indicates that the test is performed utilizing the manipulation of “end-around” hardware signals. These tests multiply the internal signals into the appropriate hardware. In the case of analog signals, the internal test voltages are used. Each test is run again after the first failure to avoid erroneous FAIL determination. A VALD/INVD (Valid/Invalid) status implies that the test is performed on incoming signals and NOT on internally wrapped signals.

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Table 7-1. Ground Maintenance Test (GMT) Procedures As each segment of hardware is tested the red FAIL/INVD annunciation will be replaced by a green PASS or valid (VALD) if the hardware test is successful. It may be necessary to run through the tests more than once to isolate intermittents. IBIT determines the status of the following items within the IC-600 as directed by the configuration: • Autopilot status (if installed) • Flight Director status • Symbol Generator status IBIT also tests the external or “blind sensors” are checked for valid data (a “blind sensor” is one whose failure is NOT apparent): • Display Unit (DU-870) • ATT/HDG interface • Radio Alt Interface • Radio Interface • Display Controller • Lamp Drivers • ADC data (MADC). The results of the Initiated Test are displayed on the PFD without further operator action. When the FAIL/INVD annunciation remains after performing the tests, the following actions should be accomplished: • Verify the appropriate LRU providing the tested interface is installed and power is applied. • Verify the wiring interface between the tested IC and LRU is correct. Replace the tested LRU with the cross-side unit to verify if condition remains. If not, replace suspect LRU and document the failed condition. Details of specific actions may be found in the manufacturers aircraft fault isolation manual.

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3.3

MAINTENANCE TEST MODE The maintenance displays provide the capability for the end user to display maintenance pages on the PFD. To enter the Maintenance Test Mode: • Use the decision height (RA) knob on the DC-550 (Figure 3.1.1) to set a RA value between 600 and 990 feet. Please note that four (4) INDEX pages at RA settings of 600, 610, 620, and 630 are provided as a reference to the available maintenance pages. • Select decesion height knob on DC-550 for 600 or above on PFD. • Push and hold TEST button/knob on the DC-550 for approximately 5 to 7 seconds . • While holding TEST button, push ET button on the DC-550. • Release the TEST button. • The IC-600 should now be in the maintenance test mode and will remain in test (via software) until it is cancelled by another press of the TEST button OR selecting a RA setting below 600. In Maintenance Test mode, the MFD will be exactly the same as in the On-Ground-Initiated Test mode. The PFD will be replaced with the results of the various maintenance tests. These results are displayed on several pages which are selected by the RA set knob on the DC-550. In addition to test results, there are also pages of sensor status and other miscellaneous information. There will be a distinct page (and possible subpages) for each setting of RA, beginning with 600 ft. and then for every 10 ft. above (i.e., 610, 620, etc). There are two sets of maintenance pages which are selectable based on the pushbutton selections. One set of maintenance pages, customer maintenance pages, are displayed for use in determining the valid operation of the IC-600. These pages are activated by pressing the TEST pushbutton approximately 5 to 7 seconds and then pressing the ET pushbutton. The second set of maintenance pages, engineering maintenance pages, are displayed to assist in troubleshooting the IC-600 internal operation. These engineering pages are activated by setting RA to 990 or greater and pressing the FULL/WX pushbutton while in the customer maintenance pages. Once the engineering pages have been entered, to display the customer mainenance pages upon re-entering the Maintenance Test Mode the ET pushbutton must be selected again or the engineering pages will be displayed.

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Table 7-1. Ground Maintenance Test (GMT) Procedures NOTE:

A description of the engineering maintenance pages are not provided in this document. If you observe maintenance pages not described in the following text you may have inadvertently entered engineering maintenance pages. These are useful to engineering and testing personnel only.

The following maintenance pages are displayed based on the RA set value: RA S ETTING

P AGE N AME

600 610 620 630 640 650 660 660C 670 680 690 700 710 720 730 740 750 760 770 780 790 800 810 820 830 840 850 860 870 880 890 900 910

Index Index Index Index Hardware 1 Identification Hardware 2 Identification Pilot’s Display Unit Page Copilot’s Display Unit Page Event Codes Power/Internal Temperature AHRS Data DC Variables Air Data DC-550 Port W eather Radar Discretes 1 Discretes 2 Discretes 3 Discretes 4 Discretes 5 Discretes 6 Discretes 7 Discretes 8 Discretes 9 Lamps Test ARINC 1 ARINC 2 ARINC 3 ARINC 4 Radio Sys Bus Checklist Loading Sys Config Blank Page

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Table 7-1. Ground Maintenance Test (GMT) Procedures The approximate location and size of Maintenance Display Formats on the DU-870 is shown in Figure 3.3.1. Each page is identified by a title (in green) on the first line. A textual description of the test result or data identification is displayed in white with the actual test results or data in cyan. Operator instructions are displayed in white with key words displayed in amber.

HELPING YOU TO CONTROL YOUR WORLD

MAINTENANCE PAGE AREA

180 RA

AD-51493@

Figure 3.3.1. Maintenance Display Formats on the DU-870

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INDEX PAGES Four Index pages presenting the Customer Index pages are provided.

3.3.1.1

Index Page 1 of 4 (RA Setting 600) I N D E X

P A G E S

1

O F

4

R A

I N D E X

6 0 0

I N D E X

6 1 0

I N D E X

6 2 0

I N D E X

6 3 0

H W

I D

1

6 4 0

H W

I D

2

6 5 0

D I S P L A Y E V E N T

A H R S

6 6 0

U N I T S

6 7 0

C O D E S

P O W E R / I N T E R N A L D A T A

S E T

T E M P

6 8 0 6 9 0

- N E X T -

600 RA AD-51494@

Figure 3.3.2. Index Page 1 of 4

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Table 7-1. Ground Maintenance Test (GMT) Procedures 3.3.1.2

Index Page 2 of 4 (RA Setting 610) I N D E X D C A I R

P A G E S

2

V A R I A B L E S

P O R T

W E A T H E R

4

R A

S E T

7 0 0 7 1 0

D A T A

D C - 5 5 0

O F

R A D A R

7 2 0 7 3 0

D I S C R E T E S

1

7 4 0

D I S T R E T E S

2

7 5 0

D I S C R E T E S

3

7 6 0

D I S C R E T E S

4

7 7 0

D I S C R E T E S

5

7 8 0

D I S C R E T E S

6

7 9 0

- N E X T -

610 RA AD-51495@


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Index Page 3 of 4 (RA Setting 620) I N D E X

P A G E S

3

O F

4

R A

D I S C R E T E S

7

8 0 0

D I S C R E T E S

8

8 1 0

D I S C R E T E S

9

8 2 0

L A M P

8 3 0

T E S T

A R I N C

1

8 4 0

A R N I C

2

8 5 0

A R N I C

3

8 6 0

A R N I C

4

8 7 0

R A D I O

S Y S

C H E C K L I S T

S E T

B U S

8 8 0

L O A D I N G

8 9 0

- N E X T -

620 RA AD-51496@


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Page 7-25 Nov 1/1996




Index Page 4 of 4 (RA Setting 630) I N D E X S Y S

P A G E S

C O N F I G

4

O F

4

R A

S E T

9 0 0

B L A N K

9 1 0

B L A N K

9 2 0

B L A N K

9 3 0

B L A N K

9 4 0

B L A N K

9 5 0

B L A N K

9 6 0

B L A N K

9 7 0

B L A N K

9 8 0

B L A N K

9 9 0

630 RA AD-51497@


22-05-14

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HARDWARE PAGES The Hardware Identification HW ID 1 page displays the following information: • Base part number of IC-600 installed. • Dash number of IC-600 installed. • Serial number of IC-600 installed. • Software mod status of the IC-600.

3.3.2.1

HW ID 1 (RA Setting 640) H W

I D

1

B A S E

P A R T

D A S H

N O .

S E R I A L S W

M O D

N O .

N O . S T A T U S

7 0 1 7 0 0 0 - 5 0 1 0 X X X X X X X X X X X

640 RA AD-51498@

Figure 3.3.6. Hardware Identification Page 1

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HW 2 (RA Setting 650) H W

2

H W

M O D

S T A T U S

A B C D E F G H J K L M N P Q R S T U V W X Y Z

H W

M O D

S T A T U S

A _


H W

M O D

S T A T U S

B _


650 RA AD-51499@

Figure 3.3.7. Hardware Identification Page 2

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DISPLAY UNIT PAGES This page displays the status of the Display Units. Data is wrapped back from display units, the MFD and EICAS on the HDLC bus and the PFD on a ARINC 429 channel. If any of the wrapped parameters do not agree, the DU WRAP will be set to FAIL. When this occurs, the following acctions should be taken: • Verify the appropriate DU is installed with power applied to the unit. • Verify the appropriate Display Wrap Around interface between the IC-600 and the suspect DU. • Replace the suspect DU with a DU from another position and if the FAIL condition continues to be observed, document the condition and remove the unit for repair. PFD, MFD, and EICAS VALIDs remain true if the labels being sent from the DU to IC-600 are fresh. Each PFD (DUs #1 and #5) provide data via Arinc 429, while each MFD and the EICAS (DUs #2, 3, and 4) provide data via HDLC. The DU Cooling Fan status is monitored by the Dispaly Units and the status is sent to the IC-600 via the Wrap Arounds. If the FAIL annunciation is displayed, the following actions should be taken: • Verify that nothing is obstructing the fan inlet or outlet for the appropriate DU. • Replace the suspect DU with a DU from another position and if the FAIL condition continues to be observed, document the condition and remove the unit for repair.

3.3.3.1

Display Units - Pilot’s Side (RA Setting 660) D I S P L A Y

U N I T S

( P )

S T A T U S

D U

# 1

W R A P

P A S S

D U

# 1

V A L I D

P A S S

D U

#

D U

# 2

W R A P

N / A

D U

# 2

V A L I D

N / A

D U

# 2

C O O L I N G

D U

# 3

W R A P

P A S S

D U

# 3

V A L I D

P A S S

D U

# 3

C O O L I N G

1

C O O L I N G

F A N S

F A N S

F A N S

P A S S

N / A

P A S S

660 RA AD-51500@

Figure 3.3.8. Display Unit Page (Pilot’s Side)

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Display Units - Copilot’s Side (RA Setting 660C)

TO BE SUPPLIED

660C RA AD-51501@

Figure 3.3.9. Display Unit Page (Copilot’s Side) L ABEL D ISPLAYED AS

ANNUNCIATION

P ARAMETER

L ABEL

“DU #1 or 5 W RAP” “DU #2,3,or 4 W RAP” “DU#1 or 5 VALID”

“PASS” or “FAIL” “PASS” or “FAIL” “PASS” or “FAIL”

PFD DU W rap-Around MFD/EICAS DU W rap-Around DU Data on HDLC

See Note 1 356, Bits 16-30 See Note 2

DU#1 or 5 COOLING FANS”

“PASS” or “FAIL1” or “FAIL2” or “FAIL 1-2”

Fan/airflow 0 test failure Fan/airflow 1 test failure Fan/airflow 0 & 1 test fail

350, Bit 22 350, Bit 21 350, Bit 22 & 21

“PASS” or “FAIL” or “N/A” or “FAIL 1” or “FAIL 2” or “FAIL 1-2”

Fan/airflow 0 test failure Fan/airflow 1 test failure Fan/airflow 0 & 1 test fail

See Note 2 350, Bit 22 350, Bit 21 350, Bit 22 & 21

“DU# 2,3, or 4 VALID” “DU# 2,3, or 4 COOLING FANS

NOTE 1: PFD wrapback data is set via an ARINC 429 bus. NOTE 2: PFD, MFD, and EICAS VALIDs are true if the DU data received over the HDLC bus is fresh.

22-05-14

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EVENT CODES (RA SETTING 670) This page displays non-volatile flight data, referred to as Event Codes (EC). A total of 64 flights (each flight is defined as one weight on wheels cycle) can be stored. The 65th flight will automatically overwrite flight number one’s data. Flights will only be stored when an EC has been recorded. This maintenance page powers up displaying the EC addresses for the current flight, i.e., ECs for the upcoming flight will be stored in the displayed addresses. If no recordable event codes have occurred during the most recent flight, the last flight with recorded event codes will be displayed. The maximum number of ECs stored per flight is 48. Note that only 8 ECs can be displayed on the screen at once. To view the previous 8 ECs, (“FULL/WX”) and to view the next 8 ECs, use DC-550 “ET,” or press the NAV PB on. E V E N T

F L I G H T

C O D E S

1

X X X X : X X

E C

X X

C C C C

W O W

Y Y Y Y : Y Y

E C

X X

C C C C

W O W

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

E C

X X

C C C C

A I R

Y Y Y Y : Y Y

670 RA

DC550 PUSHBUTTONS GPSD/TTG ET NAV

INCR ECODE NEXT FLT PREV FLT

AD-51502@

Figure 3.3.10. Event Code Page

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Table 7-1. Ground Maintenance Test (GMT) Procedures The flights are numbered from 1 - 64 (Flight 1 is shown in the example) and the most recent flight number will be displayed in green. If a ‘weight on wheels’ cycle has occurred, the start of flight timestamp will be displayed (as indicated by ‘XXXXXXX’), otherwise 00000:00 will be displayed for the timestamp. See Table 4.1 and 4.2 for Event Code Descriptons. The Event Codes are recorded for Autopilot Monitors, Flight Director, EFIS, and EICAS (Table 4.1) and Autopilot Inner Loop Monitors (Table 4.2). The information displayed on the event code page includes: P ARAMETER

S YMBOL

C OLOR

Event Code Prefix

EC

W hite

Event Code Number

xx

Magenta

Event Code

cccc

Cyan

W here EC Occurred

W OW /AIR

Cyan

Time of EC

HHHHH:MM

Cyan

Start of Flight

XXXXX:XX

Cyan

Display Controller (DC-550) Push Buttons: DC-550 P USHBUTTON

F UNCTION

“NAV”

View previous 8 event codes (current flight)

“ET”

View next 8 event codes (current flight)

GSPD/TTG

Increments 1 flight (48 event codes)

22-05-14

Page 7-32 Nov 1/1996




POWER AND INTERNAL T EMPERATURE PAGE (RA SETTING 680) This page displays the power measured by the IC-600, main power and servo power. The average power while the aircraft was in flight for the main and servo power is displayed to provide power degradation monitoring (approximately 3 hours of flight). P O W E R / I N T E R N A L

T E M P

M A I N

P W R

2 8 V D C

2 8 . 0

M A I N

P W R

A V G

2 7 . 5

S E R V O

P W R

2 8 V D C

2 8 . 0

S E R V O

P W R

A V G

2 6 . 0

T E M P

0 3 8 . 0

E N C O U N T

0 0 1 . 0

I N T E R N A L L O W E S T H I G H E S T

E N C O U N T

V

V

1 1 0 . 0

680 RA AD-51503@

Figure 3.3.11. Power/Internal Temperature Page CAUTION:

DO NOT INSTALL ANOTHER UNIT UNTIL THE DISCREPANCIES WITH THE POWER BUS ARE RESOLVED.

The power bus in question should be verifed per the Installation EB, and any discrepancies should be resolved. The Installation Bulletin for the EMB-145 provides information for the IC-600 power requirements . If a value is found to be outside the limits (32.0
S TATUS D ISPLAYED

C OLOR

Value within limits

“V”

Green

Value outside limits

“I”

Red

Not installed

“N/A”

W hite

The INTERNAL TEMP of the IC-600 is displayed in Celsius. The lowest and highest temperatures since the unit was initialized by factory personnel is displayed.

22-05-14

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AHRS DATA (RA SETTING 690) This page shows the pitch, roll, and heading data from the primary and secondary AHRS ARINC 429 data bus inputs. The data is displayed in degrees from 0 to +/- 180.0. The bus status is displayed to the right of the digital readout as follows: L ABEL

S TATUS DISPLAYED

S TATUS C OLOR

Valid

V

Green

Test

V

Green

NCD

I

Red

Fail

I

Red

Whenever the staus is displayed as “I”, the following actions should be taken: • Enter GMT on the cross-side IC-600 and observe the parameter in question. For example, if the parameter is question is the Primary Pitch, then the cross-side IC-600 Secondary pitch should be observed. • If the two separate GMTs agree, then replace the AHRS with the cross-side AHRS and repeat the GMT. If the status remains displayed as “I”, then the AHRS should be replaced and the discrepant condition documented. Note that the primary pitch for IC-600 No.1 is AHRS No. 1 and the Secondary pitch for IC-600 No. 2 is AHRS No. 2, while the Primary Pitch for IC-600 No. 2 is AHRS No. 2 and the Secondary Pitch for IC-600 No. 2 is AHRS No. 1. • If the two separate GMTs do not agree, verify the ARINC 429 interface to the IC-600s and the AHRS. A H R S

P R I S E C

P R I S E C

P R I S E C

D A T A

P I T C H P I T C H

R O L L R O L L

H D G H D G

X X X . X

V

X X X . X

V

X X X . X

V

X X X . X

V

X X X . X

V

X X X . X

V

690 RA AD-51504@

Figure 3.3.12. AHRS Data Page

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DC VARIABLES (RA SETTING 700) D C

V A R I A B L E S

R A D

A L T

B E L

V

X X . X

V

R A D

A L T

A B V

V

X X . X

V

V

X X . X

V

V

X X X . X

V

T U R N

K N O B

P I T C H

W H E E L

700 RA AD-51505@

Figure 3.3.13. DC Variables Page Line 3 contains the Radio Altitude voltage (J2A-50,51) Line 4 contains the PC-400 Turn knob voltage (J2A-73) Line 5 contains the PC-400 Pitch Wheel voltage (J2B-68,69) The Turn Knob and Pitch Wheel on the PC-400 should be rotated in both directions to verify their interface with the IC-600. When an invalid condition is displayed, the wiring interface between the IC-600 and the PC-400 should be verified. Note that the No. 2 IC-600 will normally display N/A for the Turn Knob and Pitch Wheel. C ONDITION

S TATUS D ISPLAYED

VALID

“V”

Green

INVALID

“I”

Red

Not installed

“N/A”

C OLOR

W hite

22-05-14

Page 7-35 Nov 1/1996




M ICRO AIR DATA COMPUTER (RA SETTING 710) A I R

D A T A

P R I

M A D C

S E C

P R I S E C P R I S E C P R I S E C

V A L I D

M A D C

B A R O

A L T

B A R O A L T

V A L I D

A L T

P A S S F A I L

X X X X X X X X X X X X X X X X

R A T E

A L T C A L C A L

R A T E

X X X X

A I R S P E E D

X X X X X X X X

A I R S P E E D

710 RA AD-51506@

Figure 3.3.14. Air Data Page The Barometric Altitude can be changed by rotating the Baro Set knob on the PFD Bezel Controllers. This will verify the interface from the PFD Bezel Controller to the MADC, as well as the ARINC 429 interface from the MADC to the IC-600.

22-05-14

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Table 7-1. Ground Maintenance Test (GMT) Procedures Whenever a “FAIL” condition is displayed for all the parameters, the following actions should be taken: • Verify the appropriate MADC is installed and power is applied. • Perform GMT on the cross-side IC-600. If all the parameters also show a “FAIL” condition, the suspect MADC should be replaced. • If the parameters show a “VALID” condition for the cross-side GMT test, then the ARINC 429 interface for the suspect MADC should be verified. Whenever a “FAIL” condition is displayed for only one or some of the parameters, the following actions should be taken: • The MADC should be replaced with the cross-side MADC. If the parameters now display a “VALID” condition, the suspect MADC should be replaced. • If the “FAIL” condition remains, verify the pitot/static interface with the suspect MADC. L ABEL

P ARAMETER

ARINC 429

S TATUS

“MADC VALID”

Discrete #1

270, BIT 13

Pass, Fail, or N/A

“BARO ALTITUDE (ft)”

Altitude

204

“ALTITUDE RATE (ft/min)”

Altitude Rate

212

“CAL AIRSPEED (KTS)”

Calibrated Airspeed

206

“BARO SETTING (IN-HG)”

Barometric Setting

235

“TOTAL AIR TEMP (DEG C)”

TAT

211

22-05-14

Page 7-37 Nov 1/1996




DC-550 PORT (RA SETTING 720) D C - 5 5 0

P O R T

C R S

K N O B

X X

H D G

K N O B

X X

S P D

K N O B

X X

A S E L M F D

K N O B K N O B

X X X X

E I C A S

K N O B

X X

P B

1

-

2 5

Y Y Y Y Y Y Y Y

S W

1

-

8

Z Z

720 RA AD-51507@

Figure 3.3.15. DC-550 Port Page

22-05-14

Page 7-38 Nov 1/1996



Table 7-1. Ground Maintenance Test (GMT) Procedures The various knob data is shown is HEX format. The knobs should be rotated and a corresponding change noted in the HEX data to verify the interface. DC-550, GC-550, PFD/MFD bezel controllers, and other controller pushbutton activity are sent to the IC-600 via discretes. The following pushbutton (PB 1-25) discretes are received: P USHBUTTON

L OCATION

FULL/W X

DC-550

GSPD/TTG

DC-550

ET

DC-550

NAV

DC-550

FMS

DC-550

MFD Bezel Pushbuttons 1-6

MFD Bezel

Master Caution

Inst. Panel

Master W arn

Inst. Panel

In/Pa

PFD Bezel

SPD IAS Mach

GC-550

CRS 1 Sync (Note)

GC-550

HDG Sync

GC-550

NOTE:

Copilots side will display CRS 2.

Engaging the pushbutton will change the status from ‘0’ to ‘1’. Disengaging the pushbutton will return the status back to ‘0’. Additional GC-550 functions are sent through the DC-550 controller before being received by the IC-600. • * - AP and YD are only brought in on the pilot side IC-600, and hence will only be displayed if the DU is being driven by the pilot’s IC-600.

22-05-14

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Table 7-1. Ground Maintenance Test (GMT) Procedures The Reversionary Controller switch inputs are sent to the IC-600 via the DC-550. The following switch(SW 1-8) discretes are received: S WITCH

L OCATION

AHRS Reversion

Reversionary Controller

MADC Reversion


DAU Channel A Reversion

DAU Rev. Controller

DAU Channel B Reversion

DAU Rev. Controller

DU #2/#4* PFD Select


DU #2/#4* EICAS Select


NOTE:

* Copilot’s side.

Engaging the switch will change the status from ‘0’ to ‘1’. Disengaging the pushbutton will return the status back to ‘0’. Whenever a pushbutton or switch is activated and the status does not change from ‘0’ to ‘1’, the following actions should be followed. • Verify the pushbutton or switch activity on the cross-side IC-600 GMT. • If the cross-side GMT functions correctly, verify the appropriate wiring interface to the DC-550 or IC-600. If the wiring interface is verified to be correct, replace the DC-550 with the cross-side DC-550 and verify the switch/pushbutton operation. If it now functions replace the suspect DC-550 • If the cross-side GMT also does not correctly function, verify proper operation of the switch or pushbutton in question.

22-05-14

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W EATHER RADAR (RA SETTING 730) WEATHER RADAR

WX STATUS

PASS

730 RA

DC550 PUSHBUTTONS FULL/WX

SEL/NOT SEL

AD-51508@

Figure 3.3.16. Weather Radar Page The Weather Radar status field will display the following: FAIL is displayed if either the WX Status, WX R/T, WX Mode/Range, or WX Tilt, data messages on the WX control bus are invalid. PASS is displayed if the above data messages are all received valid. NOT SEL is displayed if the WX is not selected for display on the PFD. Whenever a FAIL condition is observed, verify the wiring interface between the Radar Controller and the IC-600.

22-05-14

Page 7-41 Nov 1/1996




DISCRETES 1 (RA SETTING 740) D I S C R E T E S

1

J 1 A - 0 7 6

M N

T R M

D N

M

2 8

V

J 1 A - 0 7 6

M N

T R M

D N

A P

2 8

V

J 1 A - 0 7 7

M N

T R M

U P

M

2 8

V

J 1 A - 0 7 7

M N

T R M

U P

A P

2 8

V

J 1 A - 0 7 8

M N

T R M

E N A B L

2 8

V

J 1 A - 0 7 9

R E S E R V E D

O P N

J 1 B - 0 1 4

N 2

O P N

J 1 B - 2 1

O V R R I D E

L A M P

T E S T S E L

I N

J 1 B - 0 2 2

H D G

J 1 B - 0 2 3

R E S E R V E D

O P N O P N

P B

O P N

740 RA AD-51509@

Figure 3.3.17. Discretes 1 IC-600 P IN #

F UNCTON

S OURCE

J1A-076

Manual Trim Down

Manual Pitch Trim Switch

J1A-077

Manual Trim Up

Manual Pitch Trim Switch

J1A-078

Reserved

J1A-079

Reserved

J1B-014

N2 Override

J1B-021

Lamp Test Input

J1B-022

Heading F/D Pushbutton

J1B-023

Reserved

Maintenance Panel

GC-550, 11J1-11

22-05-14

Page 7-42 Nov 1/1996





2

J 1 B - 0 2 4

N A V

S E L

P B

O P N

J 1 B - 0 2 5

A P P

S E L

P B

O P N

J 1 B - 0 2 6

A L T

S E L

P B

O P N

J 1 B - 0 2 7

S P D

S E L

P B

O P N

J 1 B - 0 2 8

F L C

S E L

P B

O P N

J 1 B - 0 2 9

G A

S E L E C T

O P N

J 1 B - 0 3 0

V S

S E L

P B

O P N

J 1 B - 0 4 0

R E S E R V E D

O P N

J 1 B - 0 4 1

R E S E R V E D

2 8

J 1 B - 0 4 4

M S T R

O P N

W A R N

V

750 RA AD-51510@


F UNCTION

S OURCE

J1B-024

NAV F/D Pushbutton

GC-550, 11J1-13

J1B-025

APR F/D Pushbutton

GC-550, 11J1-14

J1B-026

ALT F/D Pushbutton

GC-550, 11J1-15

J1B-027

SPD F/D Pushbutton

GC-550, 11J1-16

J1B-028

FLC F/D Pushbutton

GC-550, 11J1-17

J1B-029

GA F/D Pushbutton

Throttle Levers

J1B-030

VS F/D Pushbutton

GC-550, 11J1-19

J1B-040

Reserved

J1B-041

Reserved

J1B-044

AW C Master W arning

Aural W arning Computer

22-05-14

Page 7-43 Nov 1/1996





3

J 1 B - 0 4 5

J S T K

R I G H T

J 1 B - 0 4 8

F D

J 1 B - 0 4 9

R E S E R V E D

2 8 V

J 1 B - 0 5 2

R E S E R V E D

5 V

J 1 B - 0 7 6

J S T K

U P

O P N

J 1 B - 0 7 7

J S T K

D O W N

O P N

J 1 B - 0 7 8

S T I C K

J 1 B - 0 7 9

J S T K

J 2 A - 0 4 9

R A D

J 2 A - 0 5 4

L

S E L

2 8 V

O N

G N D

S H A K E R L E F T

A L T

S I D E

V A L

S E L

2 8 V O P N 2 8 V O P N

760 RA AD-51511@


F UNCTION

S OURCE

J1B-045

Joystick Input (Right)

Optional Joystick

J1B-048

F/D 1 ON Pushbutton

GC-550, 11J1-7

J1B-049

Reserved

J1B-052

Reserved

J1B-076

Joystick Input (Up)

Optional Joystick

J1B-077

Joystick Input (Down)

Optional Joystick

J1B-078

Stall W arning

Stall Protection Computer

J1B-079

Joystick Input (Left)

Optional Joystick

J2A-049

Radio Alt 28V Valid

RAD ALT, 20J1-Y

J2A-054

Left Side Select

Aircraft W iring

22-05-14

Page 7-44 Nov 1/1996





4

J 2 A - 0 5 9

R E S E R V E D

G N D

J 2 A - 0 6 0

R E S E R V E D

O P N

J 2 A - 0 6 1

G E A R

C M D

O P N

J 2 A - 0 6 2

S P D

B R A K E

O P N

J 2 A - 0 6 3

S G

J 2 A - 0 6 4

W O W

G N D

J 2 A - 0 6 5

R E S E R V E D

O P N

J 2 A - 0 6 6

E N G / M E T R I C

O P N

J 2 A - 0 6 7

R E S E R V E D

O P N

J 2 A - 0 6 8

F M S

G N D

R E V

G N D

X S

C O N F I G

770 RA AD-51512@


F UNCTION

J2A-059

Reserved

J2A-060

Reserved

J2A-061

Ldg Gear Command

Landing Gear Lever

J2A-062

Speed Brakes Deployed

Speed Brakes

J2A-063

Cross-Side SG Rev

SG Reversion Switch

J2A-064

W eight on W heels (W OW )

Squat Switch

J2A-065

Memory Program Enable Switch

Memory Program

J2A-066

System Config Pins, Ref

Aircraft W iring

J2A-067


Aircraft W iring

J2A-068


Aircraft W iring

NOTE:

S OURCE

Changing WOW, weight on wheels, to “OPN” will result in the maintenance test being exited.

22-05-14

Page 7-45 Nov 1/1996





5

J 2 A - 0 6 9

A D F

C O N F I G

J 2 A - 0 7 0

F D

J 2 A - 0 7 1

C A T I I

J 2 A - 0 7 2

A / C

C O N F

1

O P N

J 2 A - 0 7 3

A / C

C O N F

2

O P N

J 2 A - 0 7 4

A / C

C O N F

3

G N D

J 2 A - 0 7 5

A / C

C O N F

4

O P N

J 2 A - 0 7 6

A / C

C O N F

5

G N D

J 2 A - 0 7 7

T E S T

I D

1

O P N

J 2 A - 0 7 8

T E S T

I D

2

O P N

G N D

B A R S

O P N O P N

E N

780 RA AD-51513@


F UNCTION

S OURCE

J2A-069

Sys Config Pins, Ref

Aircraft W iring

J2A-070


Aircraft W iring

J2A-071


Aircraft W iring

J2A-072

Aircraft Configuration Pins

Aircraft W iring

J2A-073


Aircraft W iring

J2A-074


Aircraft W iring

J2A-075


Aircraft W iring

J2A-076


Aircraft W iring

J2A-077

Test Idents Used by Repair and Testing Facilities

Should always read OPN when installed in aircraft

J2A-078


Should always read OPN when installed in aircraft.

22-05-14

Page 7-46 Nov 1/1996





6

J 2 A - 0 7 9

T E S T

I D

3

O P N

J 2 A - 0 8 0

T E S T

I D

4

O P N

J 2 B - 0 2 2

R E S E R V E D

O P N

J 2 B - 0 4 5

R E S E R V E D

2 8

J 2 B - 0 4 8

S G

G N D

J 2 B - 0 4 9

R E S E R V E D

2 8

J 2 B - 0 5 5

C P L

G N D

J 2 B - 0 5 6

L O W B A N K

J 2 B - 0 5 7

R E S E R V E D

G N D

J 2 B - 0 5 8

R E S E R V E D

G N D

R E V

O S

S E L

P B

2 8

P B

V

V

V

790 RA AD-51514@


F UNCTION

S OURCE

J2A-079



J2A-080



J2B-022

Reserved

J2B-045

Secondary Trim Up

Secondary Trim Switch

J2B-048

SG Reversion

SG Reversion Switch

J2B-049

Reserved

J2B-055

Couple Push Button Input

GC-550, 11J1-10

J2B-056

Low Bank PB Input

GC-550, 11J1-12

J2B-057

Reserved

J2B-058

Reserved

22-05-14

Page 7-47 Nov 1/1996




DISCRETES 7 (RA SETTING 800) DISCRETES

7

J2B-066

MN

TRIM

UP

OPN

J2B-067

MN

TRIM

DN

OPN

J2B-070

TCS

MAIN

OPN

J2B-070

TCS

AP

OPN

J2B-071

RESERVED

OPN

J2B-072

RESERVED

28V

J2B-075

TRN

OPN

J2B-076

AP

DISC

M

28V

J2B-076

AP

DISC

AP

28V

J2B-083

YD

ENG

KNOB

PB

GND

800 RA AD-51515@


F UNCTION

S OURCE

J2B-066

Manual Trim Up

Trim Switch

J2B-067

Manual Trim Dn

Trim Switch

J2B-070

TCS 28V Input MAIN CPU

TCS Switch

J2B-070

TCS 28V Input AP CPU

TCS Switch

J2B-071

Reserved

J2B-072

Reserved

J2B-075

Turn Knob Out of Detent

J2B-076

M Disc Input to main CPU

M disc switches; 28V when not pushed, OPEN when pushed

J2B-076

AP Disc Input to main CPU

AP disc switches; 28V when not pushed, OPEN when pushed

J2B-083

Yaw Engage Select PB In

GC-550, 11J1-9

22-05-14

Page 7-48 Nov 1/1996





8

J 2 B - 0 8 5

A P

E N G

J 2 B - 0 9 0

A P

J 2 B - 0 9 0

A P

J 2 B - 0 9 1

M S T R

J 2 B - 0 9 4

R

P B

G N D

D I S C

M

2 8 V

D I S C

A P

2 8 V

C A U T

S I D E

O P N

S E L

O P N

810 RA AD-51516@


F UNCTION

S OURCE

J2B-085

AP Engage Select PB Input

GC-550, 11J1-8

J2B-090

AP Disc 28V In Main CPU

IC-600, J2B-076

J2B-090

AP Disc 28V In AP CPU

IC-600, J2B-076

J2B-091

AW C Master Caution

Aural W arning Computer

J2B-094

Right Side Select

Aircraft W iring

22-05-14

Page 7-49 Nov 1/1996





9

A P

C L U T C H

M A I N

O P N

A P

C L U T C H

A P

O P N

Y D

C L U T C H

M A I N

O P N

Y D

C L U T C H

A P

O P N

A P

E N G

E N *

M A I N

L O

A P

E N G

E N *

A P

L O

Y D

E N G

E N *

M A I N

L O

Y D

E N G

E N *

A P

L O

A P

D I S C

E A

L O

D I R

L O

E L E V

O U T

S E R V O

820 RA AD-51517@

Figure 3.3.25. Discretes 9 This page displays IC-600 internal discretes. The Pitch, Roll, and Elevator Servo Clutches are all activated from a single discrete output. This output is wrapped around to both processors’ input discretes internally. The autopilot processor then passes its discrete back to the main processor for display on the maintenance page.

22-05-14

Page 7-50 Nov 1/1996




L AMP T EST 9 (RA SETTING 830) LAMPS

TEST

ANNUNCIATED

DC550

LAMP

MSTR

830 RA

PUSHBUTTONS

GPSD/TTG ET

WRN

INCR TEST DECR TEST

AD-51518@

Figure 3.3.26. Lamp Test

22-05-14

Page 7-51 Nov 1/1996



Table 7-1. Ground Maintenance Test (GMT) Procedures The lamp driver test is executed by sequencing through the lamp test sequence shown below. L AMP B EING T ESTED

T EST S EQUENCE

Cockpit Lamps

MSTR W RN, MSTR CAU

GC-550

FD1, HDG, NAV, APR, BNK, AP, CPL LEFT, YD, SPD, FLC, VS, ALT.

Cockpit Lamps

MSTR W RN, MSTR CAU.

GC-550

FD2, HDG, NAV, APR, BNK, AP, CPL RIGHT, YD, SPD, FLC, VS, ALT.

P ILOT ’ S S IDE :

C OPILOT ’ S S IDE :

The GSPD/TTG pushbutton increments to the desired lamp driver test, and the ET pushbutton is decrementa to the desired lamp driver test. The “ANNUNCIATED LAMP” line corresponds to the lamp being tested.

22-05-14

Page 7-52 Nov 1/1996




ARINC 1 (RA SETTING 840) A R I N C

0

1

J 2 A - 0 2 1

L R N

P

VAL

DATA

J2A - 02 2 1

J 2 B - 0 1 1

A P

AP

CPU

C M C

UNUSED

J2B - 01 2

2

J 2 A - 0 2 3 J2A - 02 4

3

J 2 B - 0 2 3

A D C

P

N C D

D A T A

J2B - 02 4

840 RA AD-51519@

Figure 3.3.27. ARINC #1 Because equipment installed on the aircraft may vary as identified by the configuration discretes, the ARINC 429 channels shown must be tailored accordingly. For instance, there would be no channel descriptions next to the pin and the status would be “UNUSED” if the equipment were not installed. S TATUS D ISPLAYED

C OLOR

“VAL DATA”

Green

“TST DATA”

Amber

“NCD DATA”

Amber

“INV DATA”

Red

“NO DATA”

Red

“UNUSED”

Dim white, displayed if equipment is not installed

“NOT SEL”

Amber, displayed for channels 0, 2, 3 if the selected source is secondary

N/A

N/A for A/P ARINC is normal for IC-600 #2

22-05-14

Page 7-53 Nov 1/1996





4

2

J 2 A - 0 2 9

A H R S

P

U N U S E D

J 2A - 03 0 5

J 2 B - 0 2 9

T C A S

V A L

D A T A

J 2B - 03 0

6

J 2 A - 0 3 1

RD A U

S

N O

L D A U

S

N C D

D A T A

J 2A - 03 2 7

J 2 B - 0 3 1

D A T A

J 2B - 03 2

850 RA AD-51520@

Figure 3.3.28. ARINC #2 Because equipment installed on the aircraft may vary as identified by the configuration discretes, the ARINC 429 channels shown must be tailored accordingly. For instance, there would be no channel descriptions next to the pin and the status would be “UNUSED” if the equipment were not installed. S TATUS D ISPLAYEED

C OLOR

“VAL DATA”

Green

“TST DATA”

Amber

“INV DATA”

Red

“NCD DATA”

Amber

“NO DATA”

Red

“UNUSED”


“NOT SEL”

Amber, displayed for channels 6 and 7 if the selected source is secondary

22-05-14

Page 7-54 Nov 1/1996





8

3

J 2 A - 0 4 1

RD A U

P

V A L

D A T A

L D A U

P

V A L

D A T A

V A L

D A T A

J 2A - 04 2 9

J 2 B - 0 8 1 J 2B - 08 2

1 0

J 2 A - 0 4 3

A D C

S

J 2A - 04 4 1 1

J 2 A - 0 8 3

A H R S

S

U N U S E D

J 2A - 08 4

860 RA

AD-51521@


C OLOR

“VAL DATA”

Green

“TST DATA”

Amber

“INV DATA”

Red

“NCD DATA”

Amber

“NO DATA”

Red

“UNUSED”


“NOT SEL”


22-05-14

Page 7-55 Nov 1/1996





1 2

4

J 2 A - 0 4 0

S P A R E

U N U S E D

W N D S H R

V A L

S P A R E

U N U S E D

O S

U N U S E D

J2 A - 0 2 6 1 3

J 2 B - 0 1 3

D A T A

J2 B - 0 2 6

1 4

J 2 A - 0 2 7 J2 A - 0 2 8

1 5

J 2 B - 0 2 7

P F D

J2 B - 0 2 8

870 RA AD-51522@


C OLOR

“VAL DATA”

Green

“TST DATA”

Amber

“INV DATA”

Red

“NCD DATA”

Amber

“NO DATA”

Red

“UNUSED”


“NOT SEL”


22-05-14

Page 7-56 Nov 1/1996




RADIO SYSTEM BUS (RA SETTING 880) R A D I O

S Y S

B U S

S R N

P R I

V A L

D A T A

S R N

S E C

V A L

D A T A

D M E

P R I

V A L

D A T A

D M E

S E C

V A L

D A T A

880 RA AD-51523@

Figure 3.3.31. Radio System Bus Equipment installed on the aircraft may vary as identified by the configuration discretes for the SRN and data provided via the RSB for the DME. Therefore, if the equipment is not installed the RADIO SYS BUS page will display “UNUSED” for that equipment in the status column. S TATUS D ISPLAYED

C OLOR

“VAL DATA”

Green

“INV DATA”

Red

“NO DATA”

Red

“UNUSED”

Dim white

D ESCRIPTION

1. For SRN check config discretes 2. For DME check configuration via RSB

22-05-14

Page 7-57 Nov 1/1996




Checklist Loading (RA Setting 890) The IC-600 supports the loading of user defined checklists via RS-232 interface with the User-Friendly Checklist System (UFCS), HPN (TBD). The UFCS will interface with the IC-600 through the BOOT software. The IC-600 will program the checklist via commands from the UFCS. The UFCS will create the data exchange and user interface. CHECKLIST LOADING

PROGRAMMABLE CHECKLIST EQUIPMENT IS REQUIRED

SCREEN WILL BLANK

CYCLE IC-600 CIRCUIT BREAKER TO RECOVER

DC550 PUSHBUTTONS

890 RA

NAV

AD-51524@

ENTER CKLST

Figure 3.3.32. Check Loading Page

22-05-14

Page 7-58 Nov 1/1996




SYSTEM CONFIGURATION (RA SETTING 900) S Y S

A / C

C O N F I G

C O N F I G

E M B - 1 4 5

E N G L I S H / M E T R I C

E N G L I S H

F M S

N O T

A D F

S I N G L E

F D C A T

C O M M A N D I I

A P R

B A R S

S I N G

I N S T

C U E

D I S A B L E D

900 RA

AD-51525@

Figure 3.3.33. System Configuration 3.3.28

BLANK PAGE (RA SETTING 910) BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE BLANK PAGE

910 RA AD-51526@

Figure 3.3.34. Blank Page

22-05-14

Page 7-59 Nov 1/1996




4.

EVENT CODES NOTE 1: If an Event Code No. is displayed and it is not listed in this document, Event Codes, please contact your local Customer Support Engineer for further instruction. NOTE 2: Please write the Event Code No. on your exchange or repair transaction. Please provide details regarding maintenance actions required to fix system anomalies. There are two categories of IC-600 failure and associated codes: (a) Power-up failures - These result in a latched failure which can only be reset by cycling the 28V dc circuit breaker. Power-up failures are identified as Event Codes 0001-0199 and all 90XX codes. (b) Continuous monitored failures - These result in AP/YD disengage with an associated Event Code, all 02XX and 91XX codes. If one of these monitors trips, such as in overpowering the autopilot, and the AP FAIL annunciation is displayed on the EFIS/EICAS, the failure can be reset by holding the AP DISC button for 2 second. Resetting the IC-600 circuit breaker will also reset the computer, however, the preferred method is the AP disconnect button. Not all of the 02XX or 91XX codes will result in the AP failing. They may result in the AP disconnecting, for example, a code 0207 - Primary Pitch invalid. Event codes 03XX and 04XX are codes which are logged but do not prevent AP engagement. CAUTION:

IT IS IMPORTANT TO NOTE THAT AT NO TIME SHOULD MAINTENANCE PERSONNEL ATTEMPT TO REPLACE A HONEYWELL LRU BASED ON EVENT CODES ALONE. IN MOST CASES, THE RECORDED EVENT CODES SHOULD BE USED AS A TOOL ALONG WITH TRADITIONAL GROUND TESTING AND OTHER STANDARD MAINTENANCE PRACTICES.

Table 4.1. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes That Prevent Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Engtest

0001

TCS engaged during power up test

Engtest

0002

GA engaged during power up test

Engtest

0003

Heartbeat latch is not reset before executing Engtest

Engtest

0004

Software cannot disengage the autopilot

Engtest

0005

Software cannot disengage the yaw damper

Engtest

0006

AP CPU Engtst acknowledge timeout error

Engtest

0007

Heartbeat latch cannot disengage the autopilot

Engtest

0008

Heartbeat latch cannot enable autopilot and yaw damper

Engtest

0009

AP valid discrete cannot disable the autopilot

Engtest

0010

AP valid discrete cannot enable the AP engage enable discrete

Engtest

0011

YD valid discrete cannot disable the yaw damper

Engtest

0012

YD valid discrete cannot enable the YD engage enable discrete

Engtest

0013

AP CPU Engtst acknowledge timeout error

Srvtst

0014

Power supply flipflop is not reset after clock pulse

22-05-14

Page 7-60 Nov 1/1996



Table 7-1. Ground Maintenance Test (GMT) Procedures Table 4.1. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes That Prevent Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Srvtst

0015

AP elevator servo tach moved when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0016

AP aileron servo tach moved when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0017

AP rudder servo tach moved when the AP servo valid was valid and the YD servo valid was invalid

Srvtst

0018

Trim end-arounds enabled when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0019

Elevator servo tach moved when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0020

Aileron servo tach moved when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0021

Rudder servo tach moved when the AP servo valid was valid and the YD servo valid was invalid

Srvtst

0022

Excessive AP elevator servo current when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0023

Excessive AP aileron servo current when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0024

Excessive AP rudder servo current when the AP servo valid was valid and the YD servo valid was invalid

Srvtst

0025

Excessive elevator servo current when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0026

Excessive aileron servo current when the AP servo valid was invalid and the YD servo valid was valid

Srvtst

0027

Excessive rudder servo current when the AP servo valid was valid and the YD servo valid was invalid

Srvtst

0028

Elevator servo direction discrete indicates wrong current direction

Srvtst

0029

Elevator trim end-arounds are not indicating that elevator is trimming up

Srvtst

0030

AP elevator servo tach moved too far after 10 milliseconds

Srvtst

0031

AP aileron servo tach moved too far after 10 milliseconds

Srvtst

0032

AP rudder servo tach moved too far after 10 milliseconds

Srvtst

0033

Elevator servo tach moved too far after 10 milliseconds

Srvtst

0034

Aileron servo tach moved too far after 10 milliseconds

Srvtst

0035

Rudder servo tach moved too far after 10 milliseconds

Srvtst

0036

AP elevator current not high enough after 10 milliseconds NOTE:

Srvtst

0037

Occurs when servos are not connected

AP aileron current not high enough after 10 milliseconds

22-05-14

Page 7-61 Nov 1/1996




E VENT C ODE #

D ESCRIPTION

Srvtst

0038

AP rudder current not high enough after 10 milliseconds

Srvtst

0039

Elevator current not high enough after 10 milliseconds

Srvtst

0040

aileron current not high enough after 10 milliseconds

Srvtst

0041

Rudder current not high enough after 10 milliseconds

Srvtst

0042

AP elevator servo tach did not move enough after 60 milliseconds

Srvtst

0043

AP aileron servo tach did not move enough after 60 milliseconds

Srvtst

0044

AP rudder servo tach did not move enough after 60 milliseconds

Srvtst

0045

Elevator servo tach did not move enough after 60 milliseconds

Srvtst

0046

Aileron servo tach did not move enough after 60 milliseconds

Srvtst

0047

Rudder servo tach did not move enough after 60 milliseconds

Srvtst

0048

Excessive AP elevator current after 60 milliseconds

Srvtst

0049

Excessive AP aileron current after 60 milliseconds

Srvtst

0050

Excessive AP rudder current after 60 milliseconds

Srvtst

0051

Excessive elevator current after 60 milliseconds

Srvtst

0052

Excessive aileron current after 60 milliseconds

Srvtst

0053

Excessive rudder current after 60 milliseconds

Srvtst

0054

Elevator servo direction discrete indicates wrong current direction

Srvtst

0055

Elevator trim end-arounds are not indicating elevator is trimming down

Srvtst

0056

AP elevator current too high after 10 milliseconds of polarity reversal

Srvtst

0057

AP aileron current too high after 10 milliseconds of polarity reversal

Srvtst

0058

AP rudder current too high after 10 milliseconds of polarity reversal

Srvtst

0059

Elevator current too high after 10 milliseconds of polarity reversal

Srvtst

0060

Aileron current too high after 10 milliseconds of polarity reversal

Srvtst

0061

Rudder current too high after 10 milliseconds of polarity reversal

Srvtst

0062

AP elevator servo tach did not return to zero 25 milliseconds after polarity reversal

Srvtst

0063

AP aileron servo tach did not return to zero 25 milliseconds after polarity reversal

Srvtst

0064

AP rudder servo tach did not return to zero 25 milliseconds after polarity reversal

22-05-14

Page 7-62 Nov 1/1996




E VENT C ODE #

D ESCRIPTION

Srvtst

0065

Elevator servo tach did not return to zero 35 milliseconds after polarity reversal

Srvtst

0066

Aileron servo tach did not return to zero 35 milliseconds after polarity reversal

Srvtst

0067

Rudder servo tach did not return to zero 35 milliseconds after polarity reversal

Svtsthb

0068

Power supply flipflop is not reset at power up

Svtsthb

0069

Power supply flipflop is not set after clock pulse

Svtsthb

0070

Heartbeat flipflop is not reset at power up

Svtsthb

0071

AP elevator servo tach moved with the heartbeat latch disengaged

Svtsthb

0072

AP aileron servo tach moved with the heartbeat latch disengaged

Svtsthb

0073

AP rudder servo tach moved with the heartbeat latch disengaged

Svtsthb

0074

Elevator trim end-arounds enabled with the heartbeat latch disengaged

Svtsthb

0075

Elevator servo tach moved with the heartbeat latch disengaged

Svtsthb

0076

Aileron servo tach moved with the heartbeat latch disengaged

Svtsthb

0077

Rudder servo tach moved with the heartbeat latch disengaged

Srvtst

0078

Heartbeat flipflop is not set before executing Srvtst

Svtstps

0079

Heartbeat flipflop is not set before executing Svtstps

Svtstph

0080

Power supply flipflop is not reset before executing Svtstps

Svtstps

0081

AP elevator servo tach moved with the PS flipflop disengaged

Svtstps

0082

AP aileron servo tach moved with the PS flipflop disengaged

Svtstps

0083

AP rudder servo tach moved with the PS flipflop disengaged

Svtstps

0084

Elevator trim end-arounds enabled with the PS flipflop disengaged

Svtstps

0085

Elevator servo tach moved with the ps flipflop disengaged

Svtstps

0086

Aileron servo tach moved with the ps flipflop disengaged

Svtstps

0087

Rudder servo tach moved with the ps flipflop disengaged

Svtstps

0088

AP CPU srvtstact acknowledge timeout error

Svtstps

0089

AP CPU srvtstact acknowledge timeout error, used in two different parts of the module, at the beginning and at the end

Srvtst

0090

AP CPU srvtstact acknowledge timeout error

Pstest

0091

Power supply flipflop was not reset before executing Pstest

Pstest

0092

Undervoltage power supply test did not reset the 80960 CPU

Pstest

0093

Power supply undervoltage test did not reset the power supply flipflop

22-05-14

Page 7-63 Nov 1/1996




E VENT C ODE #

D ESCRIPTION

Pstest

0094


Pstest

0095

Overvoltage power supply test did not reset the 80960 CPU

Pstest

0096

Power supply overvoltage test did not reset the power supply flipflop

Ap_powerup

0097

(W OW ) Arinc 429 timeout error, no communication from the 80188 CPU

Hbmtest

0098

Heartbeat flipflop is not set before executing Hbmtest

Ap_powerup

0099

Arinc 429 timeout error, a_pwruprdy was not set by the 80188 CPU

Ap_powerup

0100

(W OW ) Arinc 429 timeout error, no communication from the 80188 CPU

Ap_powerup

0101


Ap_powerup

0102

Heartbeat flipflop is not reset after clock pulse

Srvtst

0103

Elevator current is too low after 60 milliseconds

Srvtst

0104

Aileron current is too low after 60 milliseconds

Srvtst

0105

Rudder current is too low after 60 milliseconds

Srvtst

0106

AP elevator current too high after 10 milliseconds of polarity reversal

Srvtst

0107

AP aileron current too high after 10 milliseconds of polarity reversal

Srvtst

0108

AP rudder current too high after 10 milliseconds of polarity reversal

Hbmtest

0109

Power supply flipflop was not reset before executing Hbmtest

Hbmtest

0110

Power supply flipflop reset after heartbeat monitor failure

Hbmtest

0111

Heartbeat flipflop did not reset after heartbeat failure

Hbmtest

0112

Heartbeat flipflop did not reset after clock pulse

Hbmtest

0113

Power supply flipflop reset after heartbeat monitor failure

Hbmtest

0114

Heartbeat flipflop did not reset after heartbeat monitor failure

Hbmtest

0115


Hbmtest

0116

Two quick writes did not trip the heartbeat monitor

Hbmtest

0117

Heartbeat monitor did not trip after not updating for 90 milliseconds

Ap_powerup

0118


Ap_powerup

0119


Engtest

0120

AP disconnect active during power up test

Engtest

0121

YD select button is pushed or the AP controller is not installed

Engtest

0122

AP select button is pushed during power up test

22-05-14

Page 7-64 Nov 1/1996




E VENT C ODE #

D ESCRIPTION

Pkdet_tst_pu/c

0123

Primary peak detect initially failed

Pkdet_tst_pu/c

0124

Primary peak detect cannot be failed

Pkdet_tst_pu/c

0125

Primary peak detect failed to go valid

Pkdet_tst_pu/c

0126

Secondary peak detect initially failed

Pkdet_tst_pu/c

0127

Secondary peak detect cannot be failed

Pkdet_tst_pu/c

0128

Secondary peak detect failed to go valid

Errlgc

0129

Aircraft ID mismatch

Pkdet_tst_c

0130

AC reference failed to go valid before the timeout

Pkdet_tst_c

0131

Peak detect test executed an undefined state

Srvtst

0132

AP elevator servo tach moved at the beginning of srvtst

Srvtst

0133

AP aileron servo tach moved at the beginning of srvtst

Srvtst

0134

AP rudder servo tach moved at the beginning of srvtst

Srvtst

0135

Elevator servo tach moved at the beginning of srvtst

Srvtst

0136

Aileron servo tach moved at the beginning of srvtst

Srvtst

0137

Rudder servo tach moved at the beginning of srvtst

Srvtst

0139

Excessive AP pitch current at the beginning of srvtst

Srvtst

0140

Excessive AP roll current at the beginning of srvtst

Srvtst

0141

Excessive AP yaw current at the beginning of srvtst

Srvtst

0142

Excessive pitch current at the beginning of srvtst

Srvtst

0143

Excessive roll current at the beginning of srvtst

Srvtst

0144

Excessive yaw current at the beginning of srvtst

Main

0145

Timer #3 count register failure

Main

0146

Timer #3 interrupt failure

Main

0176

Application Prom CRC test ticket failure

22-05-14

Page 7-65 Nov 1/1996




Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Continuous Event Codes That Disengage the AP and YD Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Errlgc

0200

AP monitor ticket failure when the autopilot/yaw damper was engaged

Errlgc

0201

Ram partition failure when the autopilot/yaw damper was engaged

Errlgc

0202

Primary mux A/D-D/A wrap around failure when the autopilot/yaw damper was engaged

Errlgc

0203

Secondary mux A/D-D/A wrap around failure when the autopilot/yaw damper was engaged

Errlgc

0204

The AP monitor was not finished in the time allotted when the autopilot/yaw damper was engaged

Errlgc

0205

Continuous test detected a CRC checksum failure when the autopilot/yaw damper was engaged

Errlgc

0206

Rate gyro went invalid when the autopilot/yaw damper was engaged

Errlgc

0207

Primary pitch VG went invalid when the autopilot/yaw damper was engaged

Errlgc

0208

Secondary pitch VG went invalid when the autopilot/yaw damper was engaged

Errlgc

0209

Primary roll VG went invalid when the autopilot/yaw damper was engaged

Errlgc

0210

Secondary roll VG went invalid when the autopilot/yaw damper was engaged

Errlgc

0211

The 80188’s AP valid went invalid when the autopilot/yaw damper was engaged

Errlgc

0212

AP Arinc 429 went invalid when the autopilot/yaw damper was engaged

Errlgc

0213

Flight director data went invalid when the autopilot/yaw damper was engaged

Errlgc

0214

Stick shaker went invalid when the autopilot/yaw damper was engaged

Errlgc

0215

Pgyro’s pitch attitude monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0216

Priatt’s pitch attitude monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0217

Rgyro’s roll attitude monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0218

Priatt’s roll attitude monitor tripped when the autopilot/yaw damper was engaged

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Page 7-66 Nov 1/1996



Table 7-1. Ground Maintenance Test (GMT) Procedures Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Continuous Event Codes That Disengage the AP and YD Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Errlgc

0219

Logic’s AP disconnect monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0220

Logic’s YD disconnect monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0221

Pmodel’s pitch servo monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0222

Rmodel’s roll servo monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0223

Ymodel’s yaw servo monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0224

Accmon’s vertical acceleration monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0225

Rolmon’s roll rate monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0226

Tmrun’s trim runaway monitor tripped when the autopilot/yaw damper was engaged

Errlgc

0227

Autopilot monitor input ram corruption

Errlgc

0228

Autopilot monitor output ram corruption

Errlgc

0229

Flight director went invalid when the autopilot/yaw damper was engaged

Errlgc

0230

YD monitor ticket failure when the autopilot/yaw damper was engaged

Errlgc

0231

The 80188’s YD valid went invalid when the autopilot/yaw damper was engaged

Errlgc

0234

Non-AP side flight director went invalid when the autopilot/yaw damper was engaged

Errlgc

0235

Loss of servo power when the autopilot/yaw damper was engaged

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Page 7-67 Nov 1/1996




Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes) M ODULE

E VENT C ODE #

D ESCRIPTION

Main

0300

Arinc ram test failure

Ca_init

0301

Local ram test failure, address test

Ca_init

0302

Local ram test failure, marching 1’s

Ca_init

0303

W arm start ram checksum failure

Ca_init

0304

Local ram test failure AA’s, 55”s, FF’s, 00”s

Hdlc_powerup_w

0305

Power up ICB wraparound test failure, timeout failure

Hdlc_powerup_w

0306

Power up ICB wraparound test failure, checksum failed 3 consecutive times

Hdlc_powerup_w

0307

Power up ICB wraparound test failure, TX not done or RX not done

Ap_powerup

0308

Transfer of AP software ID failure, cold start, W OW

Main

0309

RS-422 channel #1 UART TX data timeout error

Main

0310

RS-422 channel #1 UART DMA transfer failure, either the DMA failed or the UART failed to request DAM

Main

0311

RS-422 channel #1 UART TX or RX parity or framing error

Main

0312

RS-422 channel #1 RX data failure, DMA xfer occurred to wrong address or ram didn’t accept the data

Main

0313

RS-422 channel #1 UART DMA terminal count failure

Main

0314


Main

0315


Main

0316


Main

0317


Main

0318


Main

0319


Main

0320


Main

0321


Main

0322


Main

0323


Main

0324


Main

0325


Main

0326


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Page 7-68 Nov 1/1996



Table 7-1. Ground Maintenance Test (GMT) Procedures Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event Codes) Main

0327


Main

0328


Main

0330


Main

0331


Main

0332


Main

0333


Main

0336

Power up partition failure in supervisor mode

Main

0337

Power up partition failure in user/essential mode

Main

0338

Power up partition failure in user/non-essential mode

Main

0340

Stored Arinc #1 CRC didn’t agree with the CRC stored in eeprom and the calculated CRC

Main

0341

Stored Arinc #1 CRC didn’t agree with the CRC stored in eeprom

Main

0342

Stored Arinc #1 CRC didn’t agree with the calculated CRC

Main

0344

Stored Arinc #2 CRC didn’t agree with the CRC stored in eeprom and the calculated CRC

Main

0345

Stored Arinc #2 CRC didn’t agree with the CRC stored in eeprom

Main

0346

Stored Arinc #2 CRC didn’t agree with the calculated CRC

Main

0350

Stored Application CRC didn’t agree with the CRC stored in eeprom and the calculated CRC

Main

0351

Stored Application CRC didn’t agree with the CRC stored in eeprom

Main

0352

Stored Application CRC didn’t agree with the calculated CRC

22-05-14

Page 7-69 Nov 1/1996




Table 4.4. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Event Codes That Record An Action) M ODULE

E VENT C ODE #

D ESCRIPTION

Errlgc

0400

80960 CPU had to reset the 80188 CPU, Arinc 429 timeout error

Evcode

0401

Event code buffer full

Afcs_pr_sup

0402

Run time reinitialization of ASIC #1 Arinc receivers

Evcode

0403

IC-500 over-temperature has occurred

Evcode

0404

IC-500 Over-temperature Intermittent problem

Xs_input

0405

The non-AP side IC did not acknowledge being the FD master in the time allotted

Rev_proc

0406

Invalid DU reversionary mode selected

Tcas_enabled

0407

TCAS flash failure

Pfdwpr

0408

IAS DU wrap around failure

Pfdwpr

0409

ALT DU wrap around failure

Pfdwpr

0410

Pitch DU wrap around failure

Pfdwpr

0411

Roll DU wrap around failure

Ca_init

0412

W arm start bit stuck in the warm start position, or IC reset due to heartbeat monitor

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Page 7-70 Nov 1/1996




Table 4.5. Autopilot and Interloop Monitors Event Codes (Power Up Event Codes That Prevent Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Boot188a

9000

AP power up ram test failure

Engtst

9001

TCS engaged during power up testing

Powrup

9002

Power up ticket failure

Engtst

9003

AP engaged during power up testing

Engtst

9004

YD engaged during power up testing

Engtst

9005

AP or YD engage cannot be enabled

Engtst

9006

AP engage valid cannot disable the autopilot or the AP engage valid disables the yaw damper

Engtst

9007

YD engage valid cannot re-enable YD engage enable

Promtest

9008

AP rom CRC checksum failure

Engtst

9009

AP engage valid cannot re-enable AP engage enable

W raptest

9010

AP A/D power up wrap around failure (mux 0)

W raptest

9011

AP A/D mux 1 failure

W raptest

9012

AP-A/D mux 2 failure

B429test

9013

Arinc transmitter #0 not empty

B429test

9014

Timeout error, transmitter interrupt (int3) did not occur

B429test

9015

Timeout error, DMA interrupt did not occur

B429test

9016

Timeout error, receive interrupt error, could be a label detection error

B429test

9017

Timeout error, Arinc warp around error

B429test

9018

Arinc 429 DMA wrap around data failure

B429test

9019

Arinc 429 RXO wrap around data failure

B429test

9020

Arinc 429 RX1 wrap around data failure

Btxint

9021

TX interrupt was not cleared by read cycle

Engtst

9022

YD engage valid cannot disable the yaw damper or the YD engage valid disables the autopilot

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Page 7-71 Nov 1/1996




Table 4.6. Autopilot and Interloop Monitors Event Codes (Continuous Event Codes That Prevent Autopilot Engagement) M ODULE

E VENT C ODE #

D ESCRIPTION

Elogic

9100

VG was invalid when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9101

AP Arinc 429 was invalid when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9102

AP ticket failure when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9103

AP continuous A/D wrap around failure when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9104

AP AC reference failure when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9105

AP analog ground failure when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9106

AP detected that the primary processor’s AP ticket was invalid when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9107

Stack runaway error, exceeded lower ram memory boundary

Elogic

9108

Event code buffer runaway error, exceeded upper ram memory boundary

Dynrmt

9109

AP continuous ram test failure

Elogic

9110

AP detected that the primary processor’s YD ticket was invalid when the autopilot/yaw damper was engaged or when the AP button was pressed

Elogic

9111

Background ticket failure

Table 4.7. Autopilot and Interloop Monitors Event Codes Execution Failure M ODULE

E VENT C ODE #

D ESCRIPTION

Undefint

9200

Executed an undefined interrupt
Badint

9201

Executed an undefined interrupt > type 31

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Page 7-72 Nov 1/1996




SECTION II STANDARD GROUND TEST PROCEDURE 1.

INTRODUCTION The purpose of the following tests is to verify functionality of each component and its associated wiring. Inputs and outputs for each LRU will be tested. IC-600 internal functions such as gain programming and monitors will not be exercised as they are verified at power up by software checks. Prior to testing ensure all sensors and radios are valid. In many instances the reader will be referenced to Section I, IC-600 Self Test, where component and interface functionality can be verified. All values stated in these tests have a ± 10% tolerance.

2.

DISCRETE INPUT VERIFICATIONS The following table lists various discrete inputs to the IC-600. Reference Section I, IC-600 Self Test, RA Settings, 740 through 810, to observe the input change as the respective switch is activated or power source (CB, circuit breaker) is pulled. F UNCTION

S WITCH /S OURCE

IC-600 RA S ETTING

LAMPS TEST

LAMP TEST INPUT

830

MASTER W ARN

AW C MASTER W ARNING

830

STICK SHAKER

STALL W ARNING CB

760

RAD ALT VALID

RAD ALT CB

760

W OW

SQUAT SW ITCH

770

LOW BANK PB

BANK SW ON GC-550

790

SG REV

REVERSIONARY CONTROLLER

790

TCS

TCS SW ON CONT W HEEL

800

GO AROUND

GA SW ON THROTTLE

750

TURN KNOB

TK ON PC-400

800

AP DISC

AP DISC SW ON CONT W HEEL

800 AND 810

YD ENG PB

YD BUTTON ON GC-550

800

AP ENG PB

AP BUTTON ON GC-550

810

MASTER CAUTION

AW C MASTER CAUTION

810

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3.

DU-870 (DISPLAY UNIT) The PFD, MFD, and EICAS DUs are identical except for the Bezel Controller installed on the front of the units. The displays may be swapped from one position to another to verify a suspected defective display or wiring problem.

4.

Step 1:

Ensure MFD switch on Reversionary Controller is in “NORM” (normal) position.

Step 2:

After applying aircraft power, ensure all displays light up.

Step 3:

Check each display for color, clarity, and general overall appearance.

Step 4:

Shine a light into each of the ambient light sensors (photo cells) on each display and observe the display brightness level increases.

Step 5:

Refer to Section I, paragraph 3.3.3.1, RA Setting 660. The STATUS column should indicate PASS for each item.

DC-550 DISPLAY CONTROLLER The display controller communicates with the IC-600 over a two wire digital bus. This digital bus can be verified by selecting any function on the controller and observing the appropriate response on the displays. (a) Refer to Section I, paragraph 3.3.9, RA Setting 720. DC-550 VALID should indicate PASS. (b) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate each button/knob/switch on the DC-550. A message corresponding to each activation should be displayed. (c) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate the IAS/M, HDG SYNC, and CRS SYNC pushbuttons on the GC-550. A message corresponding to each activation should be displayed. (d) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate the Master Warning and Master Caution acknowledge pushbuttons. A message corresponding to each activation should be displayed. (e) Refer to Section I, paragraph 3.3.9, RA Setting 720. Rotate the CRS, HDG, ALT, and CAS knobs on the GC-550 and the M/P RNG knob on the MFD and observe the values change for each respective knob.

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5.

IC-600 INTEGRATED COMPUTER The IC-600 Integrated Computer provides symbol generator, fault warning, flight director, and autopilot (if installed) functions. Refer to Section I, paragraph 3.2.2, regarding IC-600 self testing.

6.

AZ-850 MICRO AIR DATA COMPUTER (a) Reference Section I, paragraph 3.3.8, RA Setting 710. MADC VALID should display PASS indicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600. (b) Reference Section I, paragraph 3.3.21, RA Setting 840. MADC 1 should display VAL DATA indicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600. (c) Reference Section I, paragraph 3.3.23, RA Setting 860. MADC 2 should display VAL DATA indicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600. (d) Exit the IC-600 self test by pressing the test button on the DC-550. (e) Using a pitot static tester, simulate 200 kts of airspeed and 1000 ft/minute rate of climb. (f)

Select the IAS mode on the GC-550 and observe the display of approximately 200 kts on the PFD in the air data command location.

(g) Select the VS mode on the GC-550 and observe the display of approximately 1000 ft/minute on the PFD in the air data command location. (h) Set simulated airspeed and vertical speed to zero.

7.

AH-800 ATTITUDE AND HEADING REFERENCE UNIT (a) Reference Section I, paragraph 3.3.22, RA Setting 850. AHRS 1 should display VAL DATA indicating a valid AH-800 and its associated wiring to the IC-600 (b) Reference Section I, paragraph 3.3.23, RA Setting 860. AHRS 1 should display VAL DATA indicating a valid AH-800 and its associated wiring to the IC-600

8.

PRIMUS II RADIOS (a) Reference Section I, paragraph 3.3.25, RA Setting 880. VAL DATA should be displayed for each item, if applicable, indicating a valid radio and the Radio System Bus wiring to the IC-600.

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9.

AA-300 RADIO ALTIMETER (a) Reference Section I, paragraph 3.3.13, RA Setting 760. J2A-049 RAD ALT VALID should display 28V. (b) Reference Section I, paragraph 3.3.7, RA Setting 700. RAD ALT BEL and RAD ALT ABV should indicate a “V” along with the voltage display.

10. DAU DATA ACQUISITION UNIT (a) Reference Section I, paragraph 3.3.22, RA Setting 850. J2A-031 R DAU S and J2B-031 L DAU S should both indicate VAL DATA, indicating a valid DAU channel and associated ARINC 429 wiring to the IC-600. (b) Reference Section I, paragraph 3.3.23, RA Setting 860. J2A-041 R DAU P and J2B-081 L DAU P should both indicate VAL DATA, indicating a valid DAU channel and associated ARINC 429 wiring to the IC-600.

11. TCAS TRAFFIC COLLISION AVOIDANCE SYSTEM 11.1

PROCEDURES (a) Ensure that the 115 VAC breaker is in for the TCAS computer. (b) With the PRIMUS II radio system energized, position the cursor in the TCAS window of the RMU. (c) Press and hold the TST button. The word test will appear in the TCAS window. (d) Continue holding the TST button. The words TCAS TEST will be displayed indicating that the TCAS computer is performing self-test. If the Aural Warning Unit is enabled, a TCAS test advisory will be heard. Additionally, a TCAS Test Pattern will be displayed on the MFD, as shown in Figure 11.1.1. (e) Continue holding the TST button. Upon completion of the test, the words TCAS PASS in green, or ATC ERR in red will appear in the TCAS window. In addition, the TCAS aural warning will annunciate TCAS PASS or TCAS FAIL separately, as will the MFD annunciate TCAS PASS or TCAS FAIL.

22-05-14

Page 7-76 Nov 1/1996




FMS CHK EICAS KDVT 12.5 NM 12 MIN

33

360 MAG2 N

+15 SAT +25 TAT 300 TAS

3

PLAB1

30 PLAB2

LL01 KDVT

50

STAB TCAS TEST TGT ABV FL TX -16 +20

12

24

RA NO BRG TA NO BRG

-11

15

W

50

E

25 TCAS AUTO

6

*PBD01

315 46.0 6

21

S

+02 -03

BEZEL MENU DISPLAY AREA

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS. AD-51728@

Figure 11.1.1. TCAS Test Display

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12. PRIMUS 650/660 W EATHER RADAR PROCEDURES (a) Install a dummy load on the P-650/660 WX Receiver-Transmitter. (b) Rotate the RADAR knob on the WC-650 WX Controller to TST (Test) position. NOTE:

The radar transmits in all modes except SBY (Standby Mode).

(c) Observe radar test pattern displayed on MFD. It should appear as shown in Figure 12.1.1 or Figure 12.1.2. A broken or missing noise band usually indicates a problem with the AFC or magnetron.

FMS KDVT 12.5 NM 12 MIN

+15 SAT +25 TAT 300 TAS

360 N

3

33

TEST PATTERN

6

30

12.1

25 50

TEST MODE FAIL ANNUNCIATOR

50

315 47.0

FAIL 25

FAULT CODE

SYSTEM PAGE DISPLAY AREA BEZEL MENU DISPLAY AREA

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS. AD-51259-R1@

Figure 12.1.1. PRIMUS® 650 MFD Weather Radar Test Mode

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+15 SAT +25 TAT 300 TAS

360

XM

ST

A

3

CHK L CA 0 N O S A TT UN 7 P O C T 2 B 1F1B B AB E: N! :S T D O R IT

25 50

TEST MODE ANNUNCIATOR

6

33

CO

N

C SR C SR P S A

TEST PATTERN WITH TEXT FAULT SHOWN

30

FMS KDVT 12.5 NM 12 MIN

50

315 47.0

TEST -10

SYSTEM PAGE DISPLAY AREA BEZEL MENU DISPLAY AREA

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS. AD-51619@

Figure 12.1.2. PRIMUS® 660 MFD Weather Radar Test Mode

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13. GC-550 FLIGHT GUIDANCE CONTROLLER (a) Reference Section I, paragraph 3.3.9, RA Setting 720. Activate each pushbutton on the GC-550 controller. Engaging a pushbutton will change the status from ‘0’ to ‘1’. Disengaging the pushbutton will return the status back to ‘0’. Note that AP and YD are only brought in on the pilot side IC-600, and therefore will only be displayed if the DU is being driven by the pilot’s IC-600. (b) Additional GC-550 functions are sent through the DC-550 controller before being received by the IC-600. See paragraph 4.0

14. PC-400 AUTOPILOT CONTROLLER These tests verify the Turn Knob and Pitch Wheel functions and their associated wiring between the PC-400 and the IC-600. (a) Reference Section I, paragraph 3.3.7, RA Setting 700. The TURN KNOB and PITCH WHEEL signals should display a green “V”, valid, and a signal level of approximately 0.00 V dc. (b) TURN KNOB test: Engage the Autopilot by pressing the AP ENG button on the GC-550. Rotate the Turn Knob fully clockwise. The TURN KNOB readout should be -15 V dc. Rotate the Turn Knob fully counterclockwise. The TURN KNOB readout should be approximately 15 V dc. (c) PITCH WHEEL test: Engage the Autopilot by pressing the AP ENG button on the GC-550. Rapidly rotate the pitch wheel first in the climb and then in the descend directions. Observe a voltage reading change for the PITCH WHEEL signal.

15. SM-200 During system power up the IC-600 tests the servo motor and tach feedback functions by commanding the servo to move and then measuring the appropriate servo movement via the tach feedback signal. Proper aircraft cable tensions are required for proper autopilot and yaw damper operation. Loose or overly tight cables may result in abnormal autopilot/yaw damper operation. (a) Engage the autopilot by pressing the AP engage button on the GC-550. (b) Servo Clutch Test: Reference Section I, paragraph 3.3.19, RA Setting 820. The AP and YD clutch states should change from OPN to HI when the Autopilot is engaged. This checks the IC-600 clutch output status, which is also verified during software power up tests. (c) With Autopilot engaged ensure servos are engaged by moving the Turn Knob and Pitch Wheel on the PC-400 and observing the appropriate roll and pitch responses of the aircraft control wheel.

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16. AUTOPILOT AND ELECTRIC TRIM Pitch trim operation may be compromised by high friction control cables, high breakout forces in the trim control system, faulty horizontal stabilizer system, or improper trim output signals from the IC-600. (a) With the autopilot not engaged, manually rotate the aircraft trim wheel to the NOSE DOWN limit. (b) Reference Section I, paragraph 3.3.11, RA Setting 740. Move the control wheel trim switches to the UP position. Pitch trim should rotate 3 complete revolutions in approximately 30 seconds and J1A-077 TRIM UP should display 28 V. (c) Reference Section I, paragraph 3.3.11, RA Setting 740. Move the control wheel trim switches to the DOWN position. Pitch trim should rotate 3 complete revolutions in the opposite direction in approximately 30 seconds and J1A-076 TRIM DOWN should display 28 V. (d) With 0 kts of airspeed, manually rotate aircraft trim wheel to NOSE DOWN limit. Center control column and engage autopilot. Over power the control column in the nose down direction. After a 1-second delay, the pitch trim wheel should rotate in the NOSE UP direction, one revolution in approximately 12 seconds. Disengage autopilot. (e) Simulate greater than 200 kts (190 kts is the switching point for trim speed change). Manually rotate the aircraft trim wheel to the NOSE UP limit. Center the control column and engage autopilot. Over power the control column in the nose up direction. After a 1-second delay, the pitch trim wheel should rotate in the NOSE DOWN direction, one revolution in approximately 17 seconds.

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(Blank Page)

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SECTION III EXPANDED GROUND TEST PROCEDURE 1.

INTRODUCTION The format of the tests is to identify the required response without elaboration. The Work Steps define the preliminary setup for each test but it may be necessary to reset and reengage the mode during the course of the test. For these reasons these tests should only be used by technicians who understand the system theory and operation. All values stated in these tests should have a ± 10% tolerance.

2.

PRIMUS 650/660 WEATHER RADAR SYSTEM CHECK The following paragraphs provide the procedures for operation and checkout of the radar.

2.1

PRELIMINARY CONTROL SETTINGS Place the system controls on the WC-650 Controller in the following positions before applying power from the aircraft electrical system: MODE OFF GAIN Fully counterclockwise TILT Fully clockwise to +15 position Function WX

2.2

PRECAUTIONS If radar is to be operated in any mode other than STBY , included TEST mode, while the aircraft is on the ground: (a) Direct the nose of the aircraft so that the antenna scan sector is free of large metallic objects such as hangars or other aircraft for a distance of 100 feet and tilt antenna fully upwards. (b) Avoid operations during refueling of aircraft or other refueling operations within 100 feet. (c) Avoid operation if personnel are standing too close in the 270 degree forward section of the aircraft. Referring to the Maximum Permissible Exposure Level, this distance would be a 5–foot radius from the antenna.

2.3

TEST PATTERN After warm up, select TEST mode and verify that the test pattern appears as shown in Section II Figure 12.1.1 or Figure 12.1.2. If the noise band is missing or broken it is indicative of problems with the system.

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2.4 2.4.1

STABILIZATION INPUT PRELIMINARY CHECKS (a) Level the pitch and roll reference axes of the aircraft relative to the earth’s surface. (b) Verify that the mounting surface of the RTA is aligned to the pitch and roll reference axes of the aircraft within ± ¼ degree.

2.4.2

ELEVATION FEEDBACK CHECK/ADJUSTMENT (a) Set the XMTR switch on the RTA housing to the OFF position. (b) On the radar controller, select the STANDBY mode. (c) Set the SCAN switch on the RTA housing to the OFF position. On the radar controller, adjust the tilt control for 0.0 degree, as displayed on the EFIS/MFD display. NOTE:

2.4.3

Make sure the flat-plate antenna is at 0.0 ± 0.5 degree with respect to the RTA mounting surface.

PITCH GAIN ADJUSTMENT/CHECK (a) Set SCAN and XMTR switches, on RTA housing, to OFF (toward antenna) and select Standby on the Controller. The RTA is now in pitch and roll calibration mode. (b) Select preset GAIN. This forces a 0-degree condition in azimuth by the processor. Azimuth position of the antenna does not matter in calibration mode. (c) Adjust TILT control on Controller for 0-degree on the face of the flat-plate radiator with an inclinometer. NOTE:

Since the No. 2 AHRS is used to provide the stabilization input to the RTA, press the No. 2 AHRS test button to simulate 10-degree nose up pitch and 20 degrees of right roll.

(d) Verify that the antenna tilts down 25 ± 1.5 degrees. 2.4.4

ROLL GAIN ADJUSTMENT/CHECK (a) Verify that the GAIN control is out of the preset position. This will force a 90-degree right azimuth condition by the processor. NOTE:

Since the No. 2 AHRS is used to provide the stabilization input to the RTA, press the No. 2 AHRS test button to simulate 10-degree nose up pitch and 20 degrees of right roll.

(b) Verify that the antenna tilts up 25 ± 1.5 (c) Return the GAIN control to preset, turn the system off, and set the SCAN and XMTR switches to ON. (d) Re-install the vertical reference in its aircraft mounting location.

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2.5

FAULT MONITORS Critical functions in the RTA are continuously monitored. The RTA utilizes the EFIS as the display mechanism by transmitting detected faults to it via the control bus. The EFIS will typically display amber WX on the MFD if no SCI data is present or when an RTA failure has occurred. This may be due to an actual bus failure or if the RTA is not powered up. When the RTA serial bus is functioning normally, the EFIS will indicate the presence of any failures detected by the RTA fault monitoring system and by displaying an amber fault code number in the tilt angle location of the display when in Test mode. NOTE:

Past faults are stored in the internal memory of the radar. These are accessible when the unit is returned to a Honeywell repair facility. Only the current faults are displayed when the unit is placed in Test mode.

The Fault Codes and a brief description of each are listed in Volume I, Section 2.5, Tables 2-5-5 and 2-5-6 of this system maintenance manual.

3.

EFIS CHECK The following paragraphs provide the procedures for checkout of the EFIS.

3.1

PRELIMINARY CONTROL SETTINGS (a) Reversionary Controller: The MFD Selector should be set to NORM. The ADC, SG, and AHRS pushbuttons should not be selected to REV. (b) Display Controller: BRG 0 and BRG <> knobs selected to OFF.

3.2

DU-870 DISPLAY UNITS (a) After applying aircraft power, ensure that the displays light up. Use the Dimming Controller to adjust each display to the desired brightness. (b) Check each display for color, clarity, and general overall appearance. (c) Shine a light into each of the photocells on the displays and observe that the display brightness level increases. NOTE:

It is normal for the display to overdrive since the photocell is calibrated for changes in ambient light and not for a direct light source such as a flashlight.

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3.3

DC-550 DISPLAY CONTROLLER The Display Controller communicates with the Symbol Generator (located in the IC-600) over a two-wire digital bus. This digital bus can be verified by selecting any function on the controller and observing the appropriate response on the EFIS displays. (a) Decision Height (RA) knob: A Decision Height Set Control allows the operator to adjust the decision height. Rotation of this knob allows the decision height on the PFD to be adjusted between 5 and 200 feet in 5-foot increments, and 200 and 990 feet in 10-foot increments. Rotating the RA knob completely counterclockwise removes the decision height from the PFD display. (b) TEST pushbutton: A system test is provided by a momentary action switch that is actuated by depressing the decision height (RA) set knob. The TEST pushbutton will place the IC-600 in the test mode. Reference Section 1, paragraph 3.2.2, for expected results when the TEST button is activated. (c) BRG <> and BRG 0 knob test: Rotate the BRG <> and BRG 0 knobs and on the PFD and MFD (Compass and Map modes) observe the bearing pointers and annunciators change sources accordingly: BRG <>OFF NAV 2 ADF (single installation, ADF 2 (dual installation) FMS (option) BRG 0OFF NAV 1 ADF (single Installation, ADF 1 (dual installation) FMS (option) If the DC-550 is invalid, the onside NAV BRG is displayed by default. (d) FULL/WX pushbutton: Successive pushes of the FULL/WX button will alternately change the PFD HSI format from a full compass and ARC mode where radar will be displayed if it is operating. (e) GSPD/TTG pushbutton: Actuation of this button will change the display on the PFD between GSPD and TTG. If ET is currently displayed, pressing the GSPD/TTG button will select whichever parameter was previously displayed. Power up state is GSPD. (f)

ET pushbutton: Pushbutton control of an elapsed timer that appears on the PFD in the HSI location dedicated to GSPD/TTG. The initial actuation enters the mode at the previous sequence position. Subsequent actuations follow this toggle sequence: RESET ELAPSED TIME - STOP - REPEAT.

(g) NAV pushbutton: This pushbutton allows selection and toggling of short range navigation sources. The sources to be toggled are: Pri VORLOC - Sec VORLOC. (h) FMS pushbutton: Allows selection of LRN as the NAV source.

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3.4

DIMMING CONTROLLERS The Dimming Controllers consist of five rotary knobs which interface with each of the DUs such that they can be independently dimmed. Verify that each knob allows brightness control of its respective DU, and that with all the knobs in the minimum brightness position that all the DUs appear of equal intensity.

3.5

REVERSIONARY PANEL (a) Press the SG REV pushbutton on the Pilot’s Reversionary Panel and verify that an amber “SG2” is displayed in the upper left-hand corner of each PFD. (b) Press the SG REV pushbutton on the Copilot’s Reversionary Panel and verify that an amber “SG1” is displayed in the upper left-hand corner of each PFD. (c) Press the ADC REV pushbutton on the Pilot’s Reversionary Panel and verify that an amber “ADC2” is displayed in the upper left-hand corner of each PFD. (d) Press the ADC REV pushbutton on the Copilot’s Reversionary Panel and verify that an amber “ADC1” is displayed in the upper left-hand corner of each PFD. (e) Press the AHRS REV pushbutton on the Pilot’s Reversionary Panel and verify that an amber “ATT2” is displayed in the middle left-hand corner of each PFD. (f)

3.6

Press the AHRS REV pushbutton on the Copilot’s Reversionary Panel and verify that an amber “ATT1” is displayed in the middle left-hand corner of each PFD.

MFD BEZEL BUTTONS AND ROTARY SET KNOB Six bezel buttons are located on the bottom of the MFD display along with a rotary set knob. The menu for the bezel buttons is shown above the buttons on the bottom of the MFD CRT. These menus will change depending on the mode that is selected on the MFD. The bezel button tree menu is shown in Figure 3.6.1. (a) Main Bezel Button Menu: The main bezel button menu is the default power up. (b) Press the SYS button and verify that the menu changes to the SYS Submenu. Press the RTN button and verify that the menu returns to the main bezel menu. Repeat this procedure for each of the remaining buttons on the main bezel button menu. (c) With the main bezel button menu displayed, press the MAP/PLAN button and then rotate the M/P RNG knob and observe the range change on the MFD.

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MENU SYS

MFD

CKLST

TCAS

WX

MAP PLAN

M/P RNG

CHECK LIST RTN

T/O

ECS A/I

FUEL

ELEC

M/P RNG

RTN

SKP

RTN

SPDS

LNBK

PAG

RCL

ENT

M/P RNG

MENU

MENU RTN

HYD

SPDS

NAV APT

A

C

DATA

MAINT

M/P RNG

JSTK

NAV APT

DATA

MAINT

M/P RNG

B

JOYSTICK RTN

V1 89

VR ---

V2 ---

AP ---

SET

RTN

SKP

RCL

M/P RNG

ENT

AD-51208@

Figure 3.6.1. MFD Bezel Button Menu Tree

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4.

MADC CHECK The MADC has two operating modes: normal and self-test. The normal mode is for normal aircraft operation. The self-test mode is initiated by applying a DC ground to the self-test pin on the rear connector. This would normally be interlocked with a WOW switch, and is inhibited when more than 50 knots of airspeed is present. In the self-test mode, the outputs of the MADC are driven to preset values to check the operation of the MADC, interconnects, and the indicators.

4.1

NORMAL MODE (a) After applying power to the system, verify on the PFD that the Airspeed, Altitude, and Vertical Speed displays show valid information from the MADCs. (b) Press the ADC REV pushbutton on the Pilot’s Reversionary Panel and verify that the Pilot’s PFD shows “ADC 2” in the upper left-hand corner and that the Airspeed, Altitude, and Vertical Speed displays show valid information. (c) Press the ADC REV pushbutton on the Copilot’s Reversionary Panel and verify that the Copilot’s PFD shows “ADC 1” in the upper left-hand corner and that the Airspeed, Altitude, and Vertical Speed displays show valid information. (d) Press the ADC REV pushbuttons again to return the displays to their normal configuration. (e) Press the IN/HPA pushbutton on the PFD bezel controllers and verify that the Barometric Correction display on the respective PFDs toggle between IN (inHg) or HPA (HectoPascals). (f)

Rotate the BARO knobs on the PFD controllers and verify that the Barometric Correction Displays on the respective PFDs change value either by 0.01 inHg per click or 1 HPA per click, depending on which is selected.

(g) Press the STD pushbutton on the PFD bezel controllers and verify that the Barometric Correction Displays on the respective PFDs display standard baro correction values of 29.92 in inHg is selected or 1013 HPA if HectoPascals is selected. (h) Using a pitot static tester, simulate 200 knots of airspeed and 1000 ft/min rate of climb. Verify that the Airspeed and Vertical Speed Displays on the PFD show the simulated values from the pitot static tester. Press the GC-550 CPL pushbutton to select the side to be tested (left or right). The MADC selected for display on the side selected will be the MADC tested in steps (i) and (j) (i)

Press the SPD pushbutton on the GC-550 Flight Guidance Panel and verify that the current airspeed is displayed as the Airspeed Target above the Airspeed Tape on the PFD (for altitudes less than 34,000 feet). The target switches automatically from indicated airspeed target to mach number target as the system climbs through 34,000 feet. The target automatically switches from mach number target to indicated airspeed target as the aircraft descends through 33,750 feet.

(j)

Press the VS pushbutton on the GC-550 and verify that the Vertical Speed Command Display and Command Bug on the Vertical Speed Display on the PFD displays the current vertical speed.

(k) Return the simulated airspeed and vertical speed to zero. (l)

Press the CPL pushbutton on the GC-550 to indicate a left arrow. This will select the left MADC as the master for altitude preselect operation. Also verify that MADC 1 is selected for display on the left PFD and MADC 2 is selected for display on the right PFD.

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Table 7-1. Ground Maintenance Test (GMT) Procedures (m) Rotate the ASEL knob on the GC-550 and verify that the Altitude Preselect Display value on both PFDs changes correspondingly. (n) Press the CPL pushbutton on the GC-550 to indicate a right arrow. This will select the right MADC as the master for altitude preselect operation. Also verify that MADC 1 is selected for display on the left PFD and MADC 2 is selected for display on the right PFD. (o) Rotate the ASEL knob on the GC-550 and verify that the Altitude Preselect Display value on both PFDs changes corresponding.

4.2

SELF-TEST MODE (a) Place system in WOW mode (Weight-On-Wheels) and apply a DC ground to MADC pin J1-70 (Functional Test Discrete). Ensure that airspeed is less than 50 knots. This places the MADC in Self-Test mode. (b) The MADC being tested will output the following on the respective PFD and MFD. O UTPUTS

D ATA V ALUE

Pressure Altitude

4000 Feet

Baro-Corrected Altitude

1000 Feet

Altitude Rate

5000 Ft./Min

Calibrated Airspeed (CAS)

325 knots

True Airspeed (TAS)

325 knots

Mach

0.77 M

Static Air Temperature (SAT)

-45 deg C

Total Air Temperature (TAT)

-16 deg C

Baro-Correction inHg

29.92 inHg

Baro Correction hPa

1013 hPa

Static Pressure

29.92 inHg

Total Pressure

29.92 inHg

Dynamic Pressure

0 inHg

Max Operating Speed (VMO)

0.76 M

Vmo W arning

Active (GND) for 2 sec

MADC Valid

Inactive (OPEN)

Programmable Output Discretes

Active (GND)

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5.

RADIO ALTIMETER TEST CAUTION:

UNDER NO CIRCUMSTANCES SHOULD POWER BE TURNED ON WITH THE TRANSMIT ANTENNA DISCONNECTED FROM THE TRANSMITTER.

The following paragraph provide the procedures for the checkout of the Radio Altimeter System.

6.

Step 1:

Depress and hold the TEST button on the DC-550 Display Controller to place the Radio Altimeter system in TEST.

Step 2:

Set the Decision Height display on the PFD to 0 feet, using the Decision Height (RA) knob and place the system in the “air” mode.

Step 3:

Press and hold the TEST button for a minimum of 5 seconds. The Radio Altimeter display on the PFD should increase to 100 ± 10 feet.

Step 4:

Observe that a box is drawn.

Step 5:

Release the Test pushbutton and observe that “MIN” is drawn inside the box and flashes for 10 seconds.

AHRS CHECK The AHRS has six fundamental operating modes. They are Initialization, Full Performance, DG, Basic, Test, and Maintenance modes. The initialization mode is entered upon power up of the system. During this mode the system performs self tests to determine the condition of its components (sensors, CPU, power supply, etc). At the end of the initialization process, the system enters full performance mode unless a system input (lack of input) or operator command has placed it in one of the reversionary modes (basic or DG). The test mode can be activated by the operator at any time while the aircraft is on the ground. During the test the system outputs are driven to preset test values to verify proper operation of the data channels, interconnects, and indicators. The maintenance mode is designed for use in installing and maintaining the AHRS. An installer uses the mode for setting discretes (orientation, SDI, and ARINC 29 Hi/Low Select), determining the flux valve compensation coefficients, and determining the mounting tray alignment coefficients. Honeywell field support uses the mode for special data retrieval on the aircraft. A special function in the maintenance mode is the automatic compass swing. This mode is accessed by grounding the Maintenance Test pin on the AHRU and communicating via RS-232 with a laptop and Honeywell supplied software.

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6.1

AHRS POWER-UP CHECKOUT Verify the AHRS Power-up by performing the following: (a) Make sure that all AHRS circuit breakers are set to OFF. (b) Set the AHRS 28 V dc circuit breaker to ON. (c) Verify that the AHRS cooling air fan has turned on by checking for airflow at the air exit holes in the AHRS cover. (d) Verify that the PFD is displaying valid Attitude and Heading information. Allow 2 minutes for the AHRS to become valid.

6.2

AHRS TEST MODE (a) Place the system in WOW (Weight-On-Wheels) mode. (b) Place the AHRS in Test mode by pressing the remote mounted AHRS TEST pushbutton on the maintenance panel. (c) Verify the AHRS outputs the correct test values on the respective PFD and MFD.

6.3

AHRS ATTITUDE OUTPUTS CHECK Verify the AHRS attitude by performing the following: (a) Place the aircraft on jacks and level the aircraft per the manufacturer’s applicable aircraft leveling procedure. The aircraft should be leveled to within ± 0.1 degree. (b) Raise or lower the nose of the aircraft until it is 2 degrees away from the level position. (c) Verify that the ADIs on the PFDs display a minimal change in roll angle. (d) Verify that the ADIs on the PFDs indicate a pitch with the correct polarity.

6.4

AHRS FLUX VALVE CHECK Verify the AHRS flux valve by performing the following: (a) Move the aircraft to a level site that is free of buried metal objects and is at least 200 feet from the nearest metal structure. (b) Move the aircraft around a 360-degree compass swing. During the compass swing verify that the heading output of the AHRS on the HSI of the PFDs correctly indicates the heading of the aircraft to within ± 1 degree throughout the compass swing.

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6.5

VERIFICATION OF MEMORY MODULE PROGRAMMING If the previous AHRS CHECK procedures have been performed then it has been verified that the memory module on the mounting tray has been programmed correctly. The programming of each location in the memory module was verified during the AHRS checkout procedure as described below. (a) Memory Module Programmed Flag. The AHRS will not enter the flux valve calibration procedure unless the Memory Module Programmed Flag in the memory module has been programmed correctly. If the flux valve calibration was completed successfully then it has been verified that the Memory Module Programmed Flag in the memory module has been programmed correctly. (b) Orientation Select. If this location is not programmed correctly, the AHRS pitch and roll outputs will be incorrect. If the orientation stored in the memory module is 90 degrees off from the actual orientation of the AHRS, then the pitch of the aircraft will appear in the AHRS roll output, and the roll of the aircraft will appear in the AHRS pitch output. If the orientation stored in the memory module is 180 degrees off from the actual orientation of the AHRS, then the pitch and roll outputs of the AHRS will have the wrong sign. (i.e., they will be positive when they should be negative and vice versa.) If the AHRS Attitude Outputs Checkout procedure described in Section 6.3 was completed successfully, then the orientation select in the memory module has been programmed correctly. (c) Source/Destination Identifier. If this location is not programmed correctly, then the AHRS will not respond when it is addressed on the ASCB or ARINC bus. If the AHRS ARINC Checkout (Section I paragraph 3.3.22 and 3.3.23, RA 850 and 860) procedure was completed successfully, then the source/destination identifier in the memory module has been programmed correctly. (d) ARINC Configuration. If this location is not programmed correctly, then the AHRS will not respond when it is addressed on the ARINC bus. If the AHRS ARINC Checkout (Section I paragraph 3.3.22 and 3.3.23, RA 850 and 860) procedure was completed successfully, then the ARINC configuration in the memory module has been programmed correctly. (e) Controller Configuration. If this location is not programmed correctly, then the AHRS will not respond to controller inputs. If the AHRS was successfully placed in the output test mode via a controller command during the AHRS Test Mode (Section 6.2) procedures then the controller configuration in the memory module has been programmed correctly. (f)

Mounting Tray Misalignment Angle. If these locations are not programmed correctly, the AHRS pitch and roll outputs will be incorrect. If the AHRS Attitude Outputs Checkout procedure described in Section 6.3 was completed successfully, then the mounting tray misalignment angle in the memory module has been programmed correctly.

(g) Discrete Data Sumcheck. If this location is not programmed correctly the memory module sumcheck power-on BITE test in the AHRS will fail, and the AHRS outputs shall indicate an AHRS fault. If the AHRS was successfully powered up without the outputs indicating an AHRS fault, then the discrete data sumcheck in the memory module is correct.

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Table 7-1. Ground Maintenance Test (GMT) Procedures (h) Flux Valve Calibration Coefficients. If these locations are not programmed correctly, the output of the AHRS flux valve will not be properly compensated for errors caused by sources of magnetism other than the earth’s magnetic field. If the AHRS Flux Valve Checkout (Section 6.4) procedure was completed successfully, then the flux valve calibration coefficients must have been correctly stored in the memory module. (i)

Flux Valve Calibration Date. These locations are not used by the AHRS, therefore there will be no faults or errors if these locations are not programmed correctly, as long as the flux valve calibration data sumcheck in the memory module accurately reflects the flux valve calibration date data.

(j)

Flux Valve Calibration Coefficients Sumcheck. If this location is not programmed correctly the memory module sumcheck power-on BITE test in the AHRS will fail, and the AHRS outputs shall indicate an AHRS fault. If the AHRS was successfully powered up without the outputs indicating an AHRS fault, then the flux valve calibration coefficients sumcheck in the memory module is correct.

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7.

PRIMUS II RADIOS CHECK This section provides procedures for checking the PRIMUS® II, SRZ-85X Series Integrated Radio System for the correct installation and proper operation of all components in the system.

7.1

BUS OPERATION CHECKOUT Step 1:

With all power off, verify that the remote radio units are installed in their respective mounting trays. Install all other radio units and connect them to their respective mating connector(s).

Step 2:

Apply 28 V dc power to System 1 RMU.

Step 3:

Verify that upon initial application of power, the RMU page comes up for approximately 8 seconds during which time the Power On Self Tests (POST) are being performed.

Step 4:

If no errors are detected, the main tuning page will be displayed after completion of the POST. If an error is detected, it will be displayed on either an RMU/SYSTEM TEST RESULTS page or a RADIO TEST RESULTS page. If an error is displayed, ignore the test results at this time and go to the main page by momentarily pressing the TST button on the RMU.

Step 5:

Select the Maintenance Menu page on the RMU by simultaneously pressing the COM and NAV transfer buttons (top button on each side of the RMU).

Step 6:

Select the RMU SETUP menu. The RMU may indicate “MLS ENABLED.” The ML-850 MLS Receiver is not part of the EMB-145 radio system, therefore momentarily press the line select button adjacent to the “MLS ENABLED” display. The display should change to “MLS DISABLED.”

Step 7:

Return to the main radio page as indicated by the instructions on the RMU. If the MLS window was turned off, the MLS window on the RMU should be blank. If the MLS window was left on, the MLS display window should be on.

Step 8:

Apply 28 V dc power to System 1 Com Units by closing the COM 1 and ATC 1 circuit breakers.

Step 9:

Momentarily press the COM line select button on the RMU and verify that the yellow cursor box moves to line 2 of the COM window.

Step 10: Rotate the RMU tuning knobs and verify that the COM frequency does change. Press the COM transfer button and verify that the active and the temporary frequencies switch locations. Step 11: Press and hold the COM line select button on the RMU and verify that the yellow cursor box moves to line 1 of the COM window for direct tuning. Press and hold the line select button to return to normal tuning. Step 12: Momentarily press the transponder code line select button on the RMU and verify that the yellow cursor box moves to line 1 of the ATC window. Cursor automatically returns to the COM window after 20 seconds of inactivity. Step 13: Rotate the RMU tuning knobs and verify that the ATC codes change. Step 14: Press and hold the transponder line select button on the RMU and verify that the active ATC code is replaced by the stored ATC code.

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Table 7-1. Ground Maintenance Test (GMT) Procedures Step 15: Momentarily press the transponder mode line select button on the RMU and verify that the yellow cursor box moves to line 2 of the ATC window. Step 16: To change the memory, change the code using the RMU tuning knobs, press store (STO),change the code to another value, press and hold the transponder mode line select button again and the stored value should be should be displayed. Step 17: Rotate the RMU tuning knobs and verify that the ATC operating modes change. If the transponder is in STBY mode, rotating the tuning knobs will have no effect. Step 18: Momentarily press the transponder mode line select button on the RMU and verify that the mode changes from the previously selected System 1 mode (ATC ON, ATC ALT, TA/RA, TA ONLY) to STANDBY. Step 19: Apply 28 V dc power to System 1 NAV Unit by closing the VOR/ILS 1, DME 1, and ADF 1 circuit breakers. Step 20: Momentarily press the NAV line select button on the RMU and verify that the yellow cursor box moves to line 2 of the NAV window. Cursor automatically returns to the COM window after 20 seconds of inactivity. Step 21: Rotate the RMU tuning knobs and verify that the NAV frequency does change. Press the NAV transfer button on the RMU and verify that the active and temporary frequencies switch locations. Step 22: Press and hold the NAV line select button on the RMU and verify that the yellow cursor box moves to line 1 of the NAV window for direct tuning. Press and hold the line select button to return to normal tuning. Step 23: Momentarily press the DME button on the RMU and verify that the NAV window goes into a split mode with DME on the bottom (line 2) Step 24: Rotate the RMU tuning knobs and verify that the DME frequency changes. Step 25: Momentarily press the DME button on the RMU again and verify that the DME frequency switches to a Tacan channel number. Step 26: Rotate the RMU tuning knobs and verify that the Tacan channel numbers change. Step 27: Momentarily press the DME button on the RMU again and verify that the NAV window is no longer in the split mode. Step 28: Momentarily press the ADF frequency line select button on the RMU and verify that the yellow cursor box moves to line 1 of the ADF window. Press and hold the line select button to view the memory value. Step 29: Rotate the RMU tuning knobs and verify that the ADF frequency changes. Step 30: Momentarily press the ADF frequency mode select button on the RMU and verify that the yellow cursor box moves to line 2 of the ADF window. Cursor automatically returns to the COM window after 20 seconds of inactivity.

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Table 7-1. Ground Maintenance Test (GMT) Procedures Step 31: Rotate the RMU tuning knobs and verify that the various ADF operating modes (ADF, BFO, VOICE, ANT) cycle through the window. Step 32: Momentarily press the ADF mode line select button on the RMU and verify that the mode changes from the previously selected System 1 mode (ADF, BFO, VOICE, ANT) cycle through the window. Cursor automatically returns to the COM window after 20 seconds of inactivity. Step 33: If any of the frequency, code, or mode functions fail (dashes in cursor box), the most probable cause will be a miswired bus. Remove all power and check bus wiring. Step 34: Select the Maintenance Menu page on the RMU by simultaneously pressing the COM and NAV transfer buttons (top button on each side of the RMU). Verify that the RMU leaves its normal operating mode and displays a menu allowing the operation to enter the Aircraft Maintenance Mode. Step 35: Select the STRAP menu and check the strap display of each unit (COM UNIT, NAV UNIT, and RMU) to ensure that the proper strap options have been incorporated per the installation diagrams. Step 36: Repeat steps (a) through (hh) for System 2 radio components. Step 37: Remove all radio system power.

7.2

UNIT POWER ON SELF TEST (POST) Step 1:

Ensure that all radio system power is OFF.

Step 2:

To ensure proper POST sequencing, apply 28 V dc power to all System 1 components, ensuring that the power is applied to all components simultaneously.

Step 3:

Upon initial application of power to the RMU, the system will automatically perform a Power On Self Test (POST). If no errors are detected, the main tuning page will then be displayed. If an error is detected, it will be displayed on either an RMU/SYSTEM TEST RESULTS page or a RADIO TEST RESULTS page.

Step 4:

Annotate all errors for later troubleshooting. Momentarily press the TST button on the RMU to proceed to the main tuning page.

Step 5:

Repeat steps (a) through (d) for System 2 radio components.

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7.3 7.3.1

AIRCRAFT INTERFACE/RAMP TESTS COM 1 T ESTS Step 1:

Ensure that the RMU cursor box is in the COM window. Press and hold the TST button on the RMU and verify that COM self test is being performed. If the COM self test does not pass, annotate failure for later troubleshooting.

Step 2:

Tune COM 1 to any COM frequency and momentarily press the SQ button on the RMU. Verify that SQ is displayed in the banner of the RMU COM window and that noise can be heard in the audio system.

Step 3:

Momentarily press the SQ button on the RMU again and verify that audio noise disappears.

Step 4:

Using the RMU tune COM 1 and 2 to a local ground station or to a hand held COM. Choose the frequency to avoid blocking a busy channel.

Step 5:

On the System 1 audio panel, select the COM 1 microphone switch. Set the COM 1 volume control to mid range.

Step 6:

Momentarily key a System 1 microphone and verify that a TX is displayed in the banner of the RMU COM 1 window.

Step 7:

Transmit to the local ground station or hand held requesting a radio check.

Step 8:

Verify proper transmission and reception of audio.

Step 9:

Verify sidetone is working (hear yourself talk in headphones and optionally in speaker).

Step 10: Select the COM 2 MIC select switch on the System 1 audio panel. Step 11: Repeat steps (f) through (i). Step 12: Select the EMER switch on the System 1 audio panel. Step 13: Repeat Steps (f) through (I). Step 14: Repeat COM Tests for System 2 equipment if applicable.

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T RANSPONDER 1 T ESTS Step 1:

Ensure that the RMU cursor box is in the ATC window. Press and hold the TST button on the RMU and verify that the ATC self test is being performed. If the ATC self test does not pass, annotate failure for later troubleshooting.

Step 2:

Using the RMU, set the transponder code to a local VFR and mode to 1 ON.

NOTE:

When selecting a transponder code, be careful not to radiate any of the following codes unintentionally. Code Code Code Code

0001 7500 7600 7700

(Military intercept code) (Hijack code) (VHF COM receiver failure code) (Emergency code)

Step 3:

Using an ATC ramp tester, verify that the local VFR code is being transmitted.

Step 4:

Verify that the reply annunciator, located in the RMU upper right corner of the ATC window, comes on.

Step 5:

Momentarily press the ID button on the RMU and verify that ID is displayed in the banner of the RMU ATC window. Verify that the ATC ramp tester ID light comes on.

Step 6:

Using the RMU, set the transponder mode to 1 ATC ALT.

Step 7:

Set the aircraft’s cockpit barometric altimeter to indicate 29.92 inHg (1013 mb) and verify that the altitude code segment of the ATC ramp tester indicates the approximate field elevation.

Step 8:

Verify that the Transponder automatic reporting output displayed on the ramp tester and the altitude displayed on the barometric altimeter does not exceed 125 feet.

Step 9:

Check transponder sensitivity by placing the ramp tester antenna approximately 10 feet from the aircraft transponder antenna. Verify that reply rate is grater than 90 percent.

NOTE:

Step 10 needs to be accomplished only at initial transponder installation and as part of the FAA mandated 24 calendar month static system maintenance requirements.

Step 10: Using an air data test fixture, slowly run up the altitude to the aircraft’s maximum allowable ceiling. Verify that the baro altimeter indicator indicates the same altitude as the ATC ramp tester. Step 11: Repeat Transponder Tests for System 2 equipment if applicable.

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NAV 1 T ESTS Place the RMU cursor box in the NAV window. Press and hold the TST button on the RMU and verify that the VOR/DME self test is being performed. Also verify that the navigation displays indicate as follows: Marker

Marker Beacon annunciators flash on in the following sequence: Outer Marker - Blue Middle Marker - Amber Inner Marker - W hite

LOC/GS

Glide Slope deviation needle/bar moves up 1 standard deviation (dot).

VOR

W ith the OBS set to zero degree, the course deviation bar should center and the TO/FROM annunciator should indicate TO.

DME

DME displays 10.0 NMI.

If the VOR/DME self test does not pass, annotate the failure for later troubleshooting. 7.3.3.1

VOR Tests Step 1:

Using a VOR ramp tester, set up a frequency of 108.00 Mhz with a bearing of 0 degree TO the station.

Step 2:

Using the RMU, tune the NAV to 108.00 Mhz. Verify that the aircraft bearing indicator indicates 0 degree TO the station with the VOR Flag out of view.

Step 3:

Rotate OBS 10 degrees and verify full scale deflection of left/right needle.

Step 4:

Switch ramp test TO-FROM switch to FROM. Verify that the bearing indicator TO-FROM annunciator displays 0 degree FROM the station.

Step 5:

Set up ramp tester to transmit an audio Ident signal. Verify presence of ident tone on aircraft audio system. (Verify ident filter is off on Audio System.)

Step 6:

Repeat above steps using the 90, 180, and 270-degree radials.

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7.3.3.3

7.3.3.4

LOC 1 Tests Step 1:

Using a LOC ramp tester, set up to transmit a 108.10 Mhz on course signal.

Step 2:

Using the RMU, tune the NAV to 108.10 Mhz. Verify that the aircraft localizer deviation indicator indicates on course with its flag out of view.

Step 3:

Set up LOC ramp tester to transmit left (0.093 DDM) of course deviation. Verify that the aircraft deviation indicator indicates left of course (One dot).

Step 4:

Set up LOC ramp tester to transmit a right (0.093 DDM) of course deviation signal. Verify that the aircraft deviation indicator indicates right of course (One dot).

Step 5:

Set up ramp tester to transmit an audio signal. Verify the presence of ident tone on aircraft audio system.

Step 6:

Set up ramp tester to transmit 108.00 Mhz. Verify that the aircraft LOC indicator Flag comes into view.

Glide Slope 1 Tests Step 1:

Using a Glide Slope ramp tester, set up to transmit a 108.10 Mhz On Course signal.

Step 2:

Using the RMU, tune the NAV to 108.10 Mhz. Verify that the aircraft GS deviation indicator indicates on course with its flag out of view.

Step 3:

Adjust Ramp Generator Modulation for 0.091 DDM (Up needle). Verify that aircraft GS course needle is a half scale (one dot) up.

Step 4:

Adjust Ramp Generator Modulation for 0.091 DDM (Down needle). Verify that aircraft GS course needle is at half scale (one dot) down.

Step 5:

Set up ramp tester to transmit a frequency of 108.00. Verify that the GS indicator flag comes into view.

Marker Beacon 1 Tests Step 1:

Using a Marker Beacon ramp tester, switch on the outer marker. Verify that the aircraft’s blue marker light is activated and that a 400 Hz audio tone is present on the aircraft audio system.

Step 2:

Switch on the ramp tester middle marker. Verify that the aircraft’s amber marker light is activated and that a 1300 Hz audio tone is present on the aircraft audio system.

Step 3:

Switch on the ramp tester inner marker. Verify that the aircraft’s white marker light is activated and that a 3000 Hz audio tone is present on the aircraft audio system.

Step 4:

Repeat NAV Tests for System 2 NAV equipment if applicable.

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7.3.5

DME 1 T ESTS Step 1:

Using a DME ramp test, set up to transmit on 108.00 Mhz and set in 35 miles and 200 knots to the station.

Step 2:

Using the RMU, select DME split mode (DME Hold) and tune DME to a frequency of 108.00 Mhz. Verify on the aircraft’s DME display a distance decreasing from 35 miles. Verify a DME Hold annunciation.

Step 3:

Activate ident tone on ramp tester. Verify tone on aircraft audio system on both the headphone and speaker outputs.

Step 4:

Repeat DME Tests for System 2 equipment if applicable.

ADF 1 T ESTS NOTE:

Reliable ADF reception and bearing indications cannot be assured unless the aircraft is removed from noise sources, such as power carts and fluorescent lighting. The aircraft should also be removed from large metallic structures such as hangars and power lines that may distort or attenuate the radio fields.

Step 1:

Using the RMU, tune in a local non-directional beacon or AM radio station frequency if a beacon is unavailable using the ADF frequency window.

Step 2:

Select BFO mode in the RMU ADF mode select window. Verify the reception of an audio signal with a beat frequency on the aircraft audio system. Verify that the aircraft’s ADF bearing pointer points to the station.

Step 3:

Select ADF mode in the RMU ADF mode select window. Verify the reception of an audio signal on the aircraft audio system. The ADF needle should point to station.

Step 4:

Select ANT mode in the RMU ADF mode select window. Verify the reception of an audio signal on the aircraft audio system. Verify that the aircraft ADF bearing pointer is removed from the EFIS display.

Step 5:

Tune ADF to a strong local broadcast station frequency. Select VOICE mode and verify extended audio frequency range.

Step 6:

Repeat ADF Tests for System 2 equipment if applicable.

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Table 7-1. Ground Maintenance Test (GMT) Procedures 7.3.6 7.3.6.1

7.3.6.2

RMU T ESTS Dimming Check Step 1:

Momentarily press the DIM button on the RMU and verify that the dimming pattern is displayed. Rotate the tuning knobs and verify that dimming occurs.

Step 2:

Adjust the dimming either up or down to suit operator’s preference. To leave the DIM mode, press any line select button or momentarily press the DIM button again.

Power ON/OFF Check Step 1:

Pull the circuit breakers for each radio unit. Return to the main tuning page and verify that the radios frequencies, codes, and channel windows display dashes indicating that power to the remote unit is off.

Step 2:

Return power to all units. Return to the main tuning page and verify that all frequencies, codes, and channels are displayed as before.

NOTE 1: Repeat RMU Tests for System 2 equipment if applicable. NOTE 2: If dual radio systems are installed, momentarily press the 1/2 button on the RMU. Check cross-side operation by verifying that RMU No. 1 (pilot’s) can control System No. 2 (copilot’s) radios. NOTE 3: Repeat above step by verifying that RMU No. 2 (copilot’s) can control System No. 1 (pilot’s) radios.

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7.4

T UNING BACKUP CONTROL HEAD (CLEARANCE DELIVERY HEAD - CDH) T ESTS NOTE:

The following tests apply only if the CD-850 Clearance Delivery CDU is configuration strapped for clearance delivery mode.

Step 1:

Check Clearance Delivery CDU display for any Power On Self Test (POST) error codes. Annotate any failure codes and then press any switch or rotate any knob to exit POST mode.

Step 2:

Set the Clearance Delivery CDU mode select switch to normal. The COM and NAV frequencies displayed on the Clearance Delivery CDU should be the same as the active frequencies shown in the RMU COM and NAV windows.

Step 3:

Set the Clearance Delivery CDU mode select switch to emergency. Verify that EMRG is displayed vertically in the upper-right corner of the CDU display. Verify that AUX ON is displayed in the RMU COM and NAV windows.

Step 4:

Press the SQ button on the Clearance Delivery CDU. Verify that the SQ annunciator on the CDU display turns on and that noise can be heard in the audio system. Press SQ button again; the SQ annunciator should turn off.

Step 5:

Tune the Clearance Delivery CDU to a local VHF communications frequency and verify clear and understandable reception and transmission.

Step 6:

Tune the Clearance Delivery CDU to a local VOR station and verify that the proper VOR signals are being received.

Step 7:

Press the NAV AUDIO button on the Clearance Delivery CDU. Verify that the NAV AUDIO annunciator on the CDU display turns on and that normal receiver audio can be heard in the audio system.

Step 8:

Press the NAV AUDIO button again; the NAV AUDIO annunciator should turn off.

Step 9:

Set the Clearance Delivery CDU mode select switch to normal. The EMRG annunciator should turn off.

SYSTEM STATUS MESSAGE CHECKS The procedures in this section are to be used to check the maintenance log to determine if any errors have been entered into the log. Step 1:

Simultaneously press the COM and NAV transfer buttons (top button on each side of RMU). Verify that the RMU display leaves its normal operating mode and displays a menu which allows the operator to enter the Aircraft Maintenance Mode (AMM).

Step 2:

Select the MAINTENANCE LOG menu and check the log of each component for any errors that may have been entered during the Power On Self Test (POST) and Pilot Activated Self Test (POST) performed in Sections 7.2 and 7.3. Refer to Section 5 of the PRIMUS II Maintenance Manual, Pub. No. A15-3800-01, Fault Isolation, for troubleshooting procedures.

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7.5

IN-FLIGHT TESTS The in-flight test procedures are not mandatory and it is recommended that they be performed only to check for a suspected fault that cannot be checked on the ground. If in-flight checks are necessary they can be performed in any sequence desired. In addition only selected tests need be performed.

7.5.1

AUDIO PANEL FLIGHT T EST PROCEDURES • Electromagnetic Compatibility - With all systems operating in-flight, verify by observation, that operation of the Audio Panel has no adverse effects on other aircraft systems. • Audio Quality Check - Monitor all audio signals during flight, testing all individual radios including but not limited to: − Intelligibility of VHF communications − Localizer ident − Marker beacon audio − VOR ident − ADF ident − DME ident • Emergency Com Check - Depress “Emer Com” button and verify operation of audio panel in a simulated failure mode.

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8. 8.1

TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS) TCAS CHECK The TCAS II is programmed to perform automatic self-test at power-on and during operation. Resident test programs are executed independently in the computer unit. The test performed include power-on self-test, continuous self-test, and commanded self-test. The computer unit detests system faults and reports them on its front panel lamp display. Its flight leg memory stores status and fault information for 10 consecutive flight legs. A flight leg is the interval between weight-off-wheels and weight-on-wheels during which the TCAS is operative. By recalling the stored data, ground maintenance personnel can evaluate in-flight performance on the ground and fault isolate a current or previous failure to a specific LRU or LRU interface. Table 8.1 lists the functions of the computer unit display lamps and the corresponding troubleshooting actions. If the lamps indicate an antenna problem, the antenna connections should be checked by measuring wiring resistance at the tray per Figure 8.1.1 and Table 8.2. The resistance values in Table 8.2 are between center conductor and shield on each LTP and LMP connector. Table 8.3 summarizes how computer unit self-test is executed at power-on, during TCAS operation, and during commanded self-test. The computer unit can execute commanded self-test only when the aircraft is on the ground.

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OHM METER LEADS

2 1

RTP

BLANK

4

LTP

3

2 1

RMP

SIGNAL

4

LMP

3

RBP

SIGNAL

PWR/GND

LBP

10

AD-45492@

Figure 8.1.1. Antenna Test Points (Rack Side)

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Table 7-1. Ground Maintenance Test (GMT) Procedures Table 8.1. CU Fault Reporting and Corrective Actions L AMP

F AILURE I DENTIFICATION

P OSSIBLE ACTION

TCAS PASS

The TCAS II was operational in the flight displayed.

N/A

TCAS FAIL

TCAS failed in the flight displayed.

Remove and replace TCAS CU, if there is no other FAIL indication.

TOP ANT

Top antenna, or cabling connectors to top antenna are failed or miswired.

Verify antenna cabling and connections by verifying resistance at TCAS CU tray is per Figure 8.1.1 and Table 8.2. Antenna Wiring Resistance(after removing TCAS CU). Repair cabling, or replace antenna, as required.

BOT ANT

Bottom antenna, or cabling and connectors to bottom antenna, are failed or miswired.

Verify antenna cabling and connections by verifying resistance at TCAS CU tray is per Figure 8.1.1 and Table 8.2. Antenna Wiring Resistance(after removing TCAS CU). Repair cabling, or replace antenna, as required.

Heading input function is not used.

N/A

RA LOG

Function is not used.

N/A

TA DISP

The LRU providing the traffic advisory display function (on one or both sides) is not available to TCAS.

Verify wiring and power to TA display (on both sides of the cockpit, if two are installed). TA display may be provided on VSI/TRA, W X radar indicator, or EFIS. Check that TA valid discretes 1 and 2 (RMP-7E and RMP-7J) are 10 ohms to ground maximum. Repair wiring, or replace TA display, as required.

RA DISP

The LRU providing the resolution advisory display function (normally coupled to the VSI display on the VSI or EFIS system) is not responding to TCAS inputs.

Verify wiring and power to the RA displays. Check that RA valid discretes 1 and 2 (RMP14C and RMP-13E) are 10 ohms to ground maximum. Repair wiring, or replace RA display, as required.

RAD ALT

One or both of the radio altimeter inputs to TCAS have failed.

Verify wiring and power to the radio altimeter. Check that RAD ALT 1 and 2 are valid: RMP-2K and RBP-3C should be at 28 V dc, or RMP-13H/J and RBP-3D/E should be valid ARINC 429. Repair wiring, or replace radio altimeter, as required.

Both of the transponder interfaces to TCAS have failed.

Verify wiring and power to the transponder. Check for data on TX 429 buses 1 and 2 (RMP-14F/G and RMP-14H/J). Repair wiring, or replace transponder(s), as required.

Attitude input function is not used.

N/A

HDG

XPDR BUS

ATT

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Table 8.2. Antenna Wiring Resistance ANTENNA

C ONNECTOR S ECTION

P IN

DC R ESISTANCE

Top Directional

LTP

1

1000 ± 100 ohms

2

8060 ± 800 ohms

3

4020 ± 400 ohms

4

2000 ± 200 ohms

1

0 (50 ohms max)

2

Infinite (>50k ohms)

3


4


1

1000 ± 100 ohms

2

8060 ± 800 ohms

3

4020 ± 400 ohms

4

2000 ± 200 ohms

Bottom Omnidirectional

Bottom Directional

LM?P

LMP

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Table 8.3. CU Self-Test Execution T EST S EQUENCE

ACTIVATION

T EST I NDICATIONS

Power-on self-test

Self activated with each application of system power.

a) All front-panel lamps come on during the threesecond lamp test. b) If the TCAS is operational, the TCAS PASS green lamp comes on for a 10-second display period, and then goes off. c) If the TCAS is not operational, one or more fault lamps come on for the 10-second display period. Refer to table 8.1 for corrective action.

Continuous selftest

Executed automatically as part of normal TCAS inflight operation.

No indication, unless a fault is detected. System status/fault data is stored in memory for 10 consecutive flight legs. Data may be recalled by performing commanded self-test on the ground.

Commanded selftest

1) Press front panel PUSH TO TEST button.

Same as power-on test sequence

2) Press PUSH TO TEST button again before the previous 10-second display period has elapsed.

a) Previous fault display is aborted.

3) Press PUSH TO TEST button before the end of each succeeding display period.

a) Status/fault data recorded during a total of 10 flight legs (maximum) is displayed.

b) All lamps come on during a one-second lamp test, c) Status/fault data recorded during the preceding flight leg is displayed for 10 seconds.

b) (b)W hen test data for the earliest recorded flight leg has been displayed, all indicators will flash at a 2.5-Hz rate for three seconds if PUSH TO TEST is again pressed. This indicates the end of recorded test data.

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9. 9.1

EICAS CHECK REVERSIONARY CONTROL Step 1:

With power applied to the system, observe that the EICAS display is displaying valid data for both the No. 1 and No. 2 engines and systems.

Step 2:

Open the circuit breaker for DAU 1A and observe that the data displayed on the EICAS for the left-hand systems becomes invalid.

Step 3:

Press the DAU 1 pushbutton on the EICAS REV panel and observe that the EICAS display is now showing valid data for the No. 1 systems from DAU 1B.

Step 4:

Close the circuit breaker for DAU 1A and open the circuit breaker for DAU 1B. Observe that the data displayed on the EICAS for left-hand systems becomes invalid.

Step 5:

Close the circuit breaker for DAU 1B and press the DAU 1 pushbutton on the EICAS REV panel to select DAU 1A for display and verify that valid data is displayed for the No. 1 Systems on the EICAS display.

Step 6:

Open the circuit breaker for DAU 2A and observe that the data displayed on the EICAS for the right-hand systems becomes invalid.

Step 7:

Press the DAU 2 pushbutton on the EICAS REV panel and observe that the EICAS display is now showing valid data for the No. 2 Systems from DAU 2B.

Step 8:

Close the circuit breaker for DAU 2A and open the circuit breaker for DAU 2B. Observe that the data displayed on the EICAS for right-hand systems becomes invalid.

Step 9:

Close the circuit breaker for DAU 2B and press the DAU 2 pushbutton on the EICAS REV panel to select DAU 2A for display and verify that valid data is displayed for the No. 2 Systems on the EICAS display.

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9.2 9.2.1

CREW ALERTING SYSTEM T ONE GENERATOR AND M ESSAGE ACKNOWLEDGMENT Step 1:

Simulate a Red Warning Message. Verify that the appropriate message is displayed as blinking red in the CAS message field on the EICAS display. Verify also that the warning tones are present from the aural warning system.

Step 2:

Acknowledge the message by pressing the pilot’s Master Warning Pushbutton. Verify that the message goes steady (stops blinking) and that the warning tone stops.

Step 3:

Simulate the Red Warning Message again. Verify that the appropriate message is displayed as blinking red in the CAS message field on the EICAS display. Verify also that the warning tones are present from the aural warning system.

Step 4:

Acknowledge the message by pressing the copilot’s Master Warning Pushbutton. Verify that the message goes steady (stops blinking) and that the warning tone stops.

Step 5:

Simulate an Amber Caution Message. Verify that the appropriate message is displayed as blinking amber in the CAS message field on the EICAS display. Verify also that the caution tones are present from the aural warning system

Step 6:

Acknowledge the message by pressing the pilot’s Master Caution Pushbutton. Verify that the message goes steady (stops blinking) and that the caution tone stops.

Step 7:

Scroll the amber message off screen by using the knob on the EICAS bezel.

Step 8:

Simulate another Amber Caution Message. Verify that the appropriate message is displayed as blinking amber in the CAS message field on the EICAS display. Verify also that the caution tones are present from the aural warning system. Also verify that the first Amber message from step 5 above returned to the CAS field.

Step 9:

Acknowledge the message by pressing the copilot’s Master Caution Pushbutton. Verify that the message goes steady (stops blinking) and that the caution tone stops.

Step 10: Simulate a Blue Advisory Message. Verify that the appropriate message is automatically acknowledged after 5 seconds. Step 11: Scroll the blue message off screen by using the knob on the EICAS bezel. Step 12: Simulate another Blue Advisory Message by. Verify that the appropriate message is automatically acknowledged after 5 seconds. Also verify that the first Blue message from step 10 above returned to the CAS field.

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T AKE OFF INHIBIT FUNCTION Step 1:

Place the system in Takeoff Inhibit Mode, which is: V = V1 < 15KTS and RA < 400 FT and Elapsed time of TO Inhibit < 60 SEC

9.2.3

Step 2:

Simulate a Red or Amber message that is inhibited during Takeoff.

Step 3:

Verify that the message is inhibited, i.e., not displayed in the CAS message field for 60 seconds.

Step 4:

Verify that after approximately 60 seconds the message is displayed in the CAS field and the aural warning is active.

RMU ENGINE BACKUP Step 1:

Press the PAGE button on the RMU.

Step 2:

Select the line select key next to ENGINE.

Step 3:

The RMU Backup Engine Page No. 1 should now be displayed.

Step 4:

Press the Page 2 button.

Step 5:

The RMU Backup Engine Page No. 2 should now be displayed.

Step 6:

Press the PGE button on the RMU and then select the Return to Radios line select key.

Step 7:

The Main Radio Page should now be displayed.

Step 8:

Pull the C.B. for the left IC-600.

Step 9:

RMU1 should automatically display the Engine Backup Page No. 1.

Step 10: Reset the EICAS display C.B. 9.2.4

RMU NAVIGATION BACKUP Step 1:

Press the PAGE button on the RMU.

Step 2:

Select the line select key next to NAVIGATION.

Step 3:

The RMU Backup Navigation Page should now be displayed.

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10. FLIGHT DIRECTOR TEST Table 10.1. Flight Director Test T ITLE

T EST N O .

W ORK S TEPS

Power Check

10.1.A. Check all plugs for voltages on correct pins

Go-Around & W NDSHR Modes

10.2.A. Push & release pilot’s Go Around button with W OW

FD CMD

C ONTROL W HEEL

R ESPONSE Use appropriate interconnect to determine pins & voltages

Up

B. Engage AP with W OW .

TO annunciation on PFD will light. Command Bars will show 12 deg nose up. TO mode is reset.

C. Disengage AP. D. Push & release copilot’s Go Around button with W OW .

Up

TO annunciation on PFD shall light. Command Bars will show 12 deg nose up.

E. Engage AP with W OW . F. Simulate air mode, and greater than 400 ft radio altitude. G. Push & release pilot’s Go Around button with A/C in Air Mode. H. Engage AP. The internal monitors may disengage the AP shortly after engagement.

TO mode is reset. Up GA annunciation on PFD will light. GA mode remains engaged.

I. Push & release pilot’s Go Around button. Press AP DISC switch on control wheel to reset AP. J. Push & release copilot’s Go Around button with A/C in Air Mode.

GA mode extinguishes

Engage AP. The internal monitors may disengage the AP shortly after engagement.

GA annunciation on PFD will light..

K. Disengage AP. Note that internal monitors may disengage the AP soon after engagement. If this occurs, press the AP DISC switch on the control wheel to reset the system.

GA mode remains engaged. AP OFF is annunciated on EICAS. If monitors disengaged AP then AP FAIL is annunciated until AP DISC is pressed.

L. Select GA with W OW . M. Engage Test Mode for GPW S/W indshear with switch located on Maintenance Panel. Verify that windshear warning is annunciated on the PFD

TO annunciation on PFD will light. Command Bars will show 12 deg nose up

N. Press GA again.

W DSHR will be annunciated in place of TO Mode. W DSHR is deselected.

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Table 7-1. Ground Maintenance Test (GMT) Procedures Table 10.1. Flight Director Test T ITLE

T EST N O .

Annunciator Lights Test HDG Select & FD OFF Modes & CPL Select

W ORK S TEPS

FD CMD

C ONTROL W HEEL

R ESPONSE

10.3.A Reference maintenance page 1-35 of IC-600 self-test and perform lamp test. 10.4.A. Connect pitot-static tester to pilot’s pitot input. Set Hdg Bug on PFD to fore Lubber Line. Select HDG on GC-550 FD mode selector.

Left

B. Decrease Hdg Bug reading 10 deg less than aircraft heading. C. Increase Hdg Bug reading to 10 deg greater than aircraft heading.

HDG annunciators on GC-550 FD mode selector & PFD will light. FD pitch command in pitch sync mode.

FD Command bar follows Hdg Bug. Right

D. Increase Hdg Bug reading to 90 deg greater than aircraft heading.

FD Command bar follows Hdg Bug. FD Command bar follows Hdg Bug.

E. Push & release BNK button on autopillot GC-550 controller.

Right

Autopilot GC-550 controller LOW bank Limit Light turns on. PFD LOW Bank Limits are displayed.

F. Push and release BNK button on GC-550 autopilot controller,

Right

Autopilot GC-550 controller LOW BANK Limit Light turns off. PFD LOW Bank Limits are removed.

G. Press Hdg SYNC to synch Hdg Bug reading with lubber line.

Null

FD command bar follows Hdg Bug.

H. Increase altitude to 25,000 ft.

Null

At approximately 25,000 ft. the LOW BANK light on the autopilot GC-550 controller will light and the Low Bank will be annunciated on the PFD.

I. Decrease altitude to 24,500 ft

Null

At approximately 24,750 ft. the LOW BANK on the autopilot GC-550 controller and the Low Bank annunciation on the PFD will extinguish.

J. Set simulated airspeed to zero. Push and release FD Off button on the GC-550.

Null

FD command bar is removed on appropriate PFD.

K. Reset HDG mode.

HDG annunciators on GC-550 and PFD are removed.

L. Sel FD ON and push and Select HDG on GC-550 FD mode selector.

HDG annunciators on GC-550 FD mode selector & PFD will light. FD command bars appear on PFD.

M. Push & release CPL button on GC-550.

CPL arrow on GC-550 and PFDs changes direction. HDG mode will extinguish.

N. Select CPL on GC-550 to return CPL arrow to pilot side.

CPL arrow on GC-550 and PFDs points left.

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T EST N O .

NAV Select Mode, VOR Course

NAV Select Mode, VOR Radio Deviation

W ORK S TEPS

FD CMD

C ONTROL W HEEL

R ESPONSE

Left

Green NAV annun on FD MODE SELECTOR GC-550 shall light. Green VOR Annunciator on PFD shall appear.

B. Decrease Course Pointer reading 10 deg.

Right

FD Command bar follows Course Pointer.

C. Increase Course Pointer reading 10 deg from original position.

Left

FD Command bar follows Course Pointer

D. Set simulated airspeed to zero. Reset all FD modes.

Null

10.5.A. Simulate 100 kts airspeed. Tune pilot’s nav receiver to a VOR frequency. Set course Pointer to zero left-right (CDI) needle. Select NAV on FD MODE SELECTOR GC-550.

10.6.A. Simulate 100 kts airspeed. Set Course Pointer 90 deg right of aircraft heading. Simulate greater than 2 dots dev fly left, 25 DME miles & a “TO” the station indication. Set Hdg Bug on PFD to fore lubber line. Select NAV on FD mode selector GC-550. B. Reduce radio deviation slowly.

Null

HDG annunciators on FD model selector GC-550 & PFD shall light. Green NAV annunciator on FD mode selector shall light. W hite VOR annunciator on PFD shall appear. FD pitch command in pitch sync mode.

Right

NAV mode captures at approx. 1/2 dot. HDG annunciators turn off. Green VOR annunciator on PFD shall appear. Nav mode annunciators on PFD & mode selector GC-550 shall go out. FD Command bar is out of view.

C. Reset NAV mode. (Push NAV button on GC-550) Simulate 250 kts airspeed. Simulate greater than 2 dots dev fly left. D. Select NAV on FD mode selector

.E. Reduce radio deviation slowly.

F. Reset all FD modes.

Null

HDG annunciators on FD mode selector GC-550 & PFD shall light. NAV annunciator on FD mode GC-550 selector shall light. W hite VOR annunciator on EADI PFD shall appear. FD pitch command in pitch sync mode.

Right

NAV mode captures at about 1 dot. HDG annunciators turn off. Green VOR annunciator on PFD shall appear. FD Command bar is out of view.

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T EST N O .

NAV Select Mode, VOR APP & OSS

W ORK S TEPS

10.7A. Simulate 100 kts airspeed. Set Hdg Bug & Course Pointer on PFD to fore lubber line. Simulate greater than 2 dots dev fly right, 10 DME miles & a “TO” the station indication. Select APR on FD mode selector GC-550. B. Reduce radio deviation to zero. W ait 20 seconds before starting test no. 7C.

FD CMD Null

Nulled

C ONTROL W HEEL CTR

R ESPONSE HDG annunciator on FD mode selector GC-550 & EADI PFD shall light. W hite VAPP annunciator on PFD shall appear. FD pitch command in pitch sync mode.

HDG annunciators turn off. NAV annunciator on FD mode selector GC-550 shall light. Green VAPP annunciator on PFD shall appear. Allows OSS circuit to time out.

C. Pull C.B. for AZ-850. OSS circuit trips. FD command bar follows course pointer change, not radio dev.

D. Rapidly rotate course pointer plus & minus 30 deg from aircraft heading. E. Set course pointer 5 deg greater than aircraft heading & 1 dot radio dev fly left

NAV Select Mode, Localizer

Right

For about 30 seconds FD command responds only to course error then moves left in response to radio dev signal. Checks OSS circuit time out.

F. Reset APR on FD mode selector GC-550. Set course pointer 10 deg less than aircraft heading & radio dev to zero.

Nav mode annunciators on PFD & mode selector GC-550 shall go out. FD Command bar is out of view.

G. Reset all FD modes.

FD Command bar is out of view. HDG annunciators on FD mode selector GC-550 & PFD shall light. NAV annunciator on FD mode selector GC-550 shall light. W hite LOC annunciator on PFD shall appear. FD pitch command in pitch sync mode.

10.8.A. Tune pilot’s nav receiver to a LOC frequency. Set Course Pointer 90 deg right of aircraft heading. Simulate greater than 2 dots dev fly left. Set Hdg Bug on PFD to fore lubber line. Select NAV on FD mode selector GC-550. B. Reduce radio deviation to zero.

Right

C. Set Course Pointer to aircraft heading.

Left

D. One minute after test no. 8.C. rotate Course Pointer 20 deg left of aircraft heading.

Left

E. Rotate Course Pointer 20 deg right of heading. F. Return Course Pointer to aircraft heading.

Right

HDG annunciators turn off. Green LOC annunciator on PFD shall appear.

FD command initially goes left then nulls. Checks phasing & crosswind washout for LOC. FD command initially goes right then nulls. Checks phasing & crosswind washout for LOC. FD Command bar is out of view.

G. Reset all FD Modes

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T EST N O .

APR Select Mode, Glide Slope

W ORK S TEPS

FD CMD

B. Reduce GS deviation to 1/4 dot fly up.

Dn then Up

C. Reduce GS deviation to zero.

Nulls

D. Simulate 1/4 dot fly down GS deviation.

Dn

Dn then Up

G. Reset ALL FD modes. Set simulated airspeed to zero 10.10.A Tune pilot’s nav receiver to a LOC frequency. Set Course Pointer on PFD to aft lubber line (tail of course pointer is forward). Simulate greater than 2 dots dev fly left. Set Hdg Bug on PFD to fore lubber line. Select APR & ALT on FD mode selector GC-550.

FD command goes down.

Green GS annunciator on PFD shall appear. Initial command is fly down. After 15 seconds command is fly up. FD Command bar is out of view

Null

B. Reduce radio deviation to zero.

C. Rotate tail of Course Pointer 10 deg right of aircraft heading.

Green GS annunciator on PFD shall light. Initial command is fly down. After 15 seconds command is fly up.

NAV CAP, APR ARM & ALT annunciators on FD mode selector GC-550 shall light. Green LOC, ALT & white GS annunciators on PFD shall appear.

E. Open radio altimeter valid or pull radio altimeter circuit breaker. Reestablish original conditions in test no. 9.A.

BC Select Mode, Back Course

R ESPONSE NAV, APR & Alt annunciators on FD mode selector GC-550 shall light. Green LOC, ALT & white GS annunciators on PFD shall appear. AP & YD ENGAGE lights turn on.

10.9.A. Tune pilot’s nav receiver to a LOC frequency. Simulate 150 kts airspeed. Simulate 1500 ft on radio altimeter by putting piece of Echo Sorbs on either Radio Altimeter Antennas. Set Hdg Bug & Course Pointer on PFD to fore lubber line. Simulate zero LOC deviation. Simulate 1 dot fly up GS deviation. Select APR & ALT on FD mode selector GC-500.

F. Reduce GS deviation to zero then return it to 1/4 dot fly up.

C ONTROL W HEEL

HDG, BC & ALT annunciators on FD mode selector GC-550 shall light. Green HDG, ALT & white BC annunciator on PFD shall appear.

HDG annunciators turn off. Green BC annunciator on PFD shall appear. Right

FD command initially goes right then nulls. Checks phasing & crosswind washout for BC.

D. Return tail of Course Pointer to aircraft heading. E. Reset all FD modes.

FD Command bar is out of view.

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T EST N O .

ALT Select Mode, Altitude Hold

W ORK S TEPS

10.11.A. Set Hdg Bug on PFD to fore lubber line. Select HDG & ALT on FD mode selector GC-550. B. Simulate 200 ft greater than present altitude.

FD CMD

R ESPONSE

Null

HDG & ALT annunciators on FD mode selector GC-550 & EADI PFD shall light.

Down

Checks phasing of ALT error signal.

C. Deselect & reselt ALT on GC-550.

Null

System resynchronizes to new altitude at ALT engage.

D. Simulate 200 ft less than original altitude.

Up

Checks phasing of ALT error signal.

E. Reset all FD modes. Disengage AP. ALT SEL Mode Altitude Preselect & Departure W arning

C ONTROL W HEEL

FD Command bar is out of view. HDG & VS annunciators on FD mode selector GC-550 light. Green HDG, VS and white ASEL annunciators appear on PFD. 2000 fpm air data command appears on PFD.

10.12.A. Preselect 10,000 ft using ALTSEL knob. Simulate 2000 fpm climb with ADC tester. Set Hdg bug on PFD to fore lubber line. Select HDG and VS on FD mode selector GC-550. B. Allow altitude to increase towards 10,000 ft.

Null

At approximately 9000 ft the ALT PRESELECT W INDOW on the PFD changes from blue to amber, and the altitude alert horn sounds for 1 second.

Down

At approximately 9600 ft the VS annunciator on the FD mode selector GC-550 turns off and the ASEL arm annunciator on the PFD changes from white to green. A white box around the green ASEL annunciator is displayed for 7 seconds indicating mode capture. At approximately 9750 ft the amber altitude select window changes from amber to blue.

C. Reduce Climb Rate to zero at 10,000 ft.

Null

The green ASEL annunciator on PFD turns off and the green ALT annunciator on the FD mode selector GC-550 and the PFD lights. NOTE: System will not go into ALT hold mode if rate of climb is not less than 300 FPM and within 25 ft of selected altitude.

D. W hile in ALT HLD decrease altitude below 9750 ft

Up

At approximately 9750 ft the Altitude select window changes from blue to amber and a horn sounds for one second for the altitude departure warning function.

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T EST N O .

VS Select Mode, Vertical Speed Hold

W ORK S TEPS

FD CMD

C ONTROL W HEEL

R ESPONSE

Null

HDG & VS annunciators on FD mode selector GC-550 & PFD turn on. 2000 fpm air data command appears on PFD.

B. Increase VS climb rate to 2500 fpm.

Down

Checks phasing of VS error signal.

C. Push & hold TCS button until signal nulls

Nulls

Synchronizes to new VS signal. Air data command displays dashes while TCS button pushed, new VS when button released.

Up

Checks phasing of VS error signal.

Nulls

Synchronizes to new VS signal. Air data command displays dashes while TCS button pushed, new VS when button released.

10.13.A. Simulate 2000 fpm climb with Pitot/ Static tester. Set Hdg Bug on EHSI to fore lubber line. Select HDG & VS on FD mode selector GC-550.

D. Decrease VS climb rate to 1000 fpm. E. Push & hold TCS button until signal nulls.

F. Move pitch wheel on AP controller fore and aft

Resets VS hold mode. VS annunciators on FD mode selector GC-550 & PFD turn off. FD command in pitch sync mode.

G. Reset all FD modes. Set simulated VS rate to zero.


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T EST N O .

SPD Mode

W ORK S TEPS

FD CMD

C ONTROL W HEEL

R ESPONSE

Null

HDG & IAS annunciators on FD mode selector GC-550 turn on. HDG & IAS annunciators on PFD appear. 200 kts air data command appears on PFD.

Down

Checks phasing of IAS error signal.

C. Increase airspeed to 210 kts.

Up

Checks phasing of IAS error signal.

D. Push & hold TCS button until signal nulls.

Nulls

Synchronizes to new IAS signal. Air data command displays dashes while TCS button pushed, new IAS when button released.

E. Move pitch wheel on AP controller in climb or descend.

Sync

Resets IAS hold mode. IAS annunciators on FD mode selector GC-550 & EADI PFD turn off. FD command in pitch sync mode.

F. Select SPD mode on GC-550.

Null

IAS annunciator on FD mode selector GC-550 turns on. IAS annunciator on PFD appears. 200 kts air data command appears on PFD.

10.14.A. Simulate 15,000 ft altitude & 200 kts airspeed with ADC tester. Set Hdg Bug on PFD to fore lubber line. Select HDG & SPD on FD mode selector GC-550. B. Decrease airspeed to 190 kts.

G. Increase altitude slowly to 26,000 Ft.

As altitude passes thru 25,100, MACH is annunciated on the GC-550 and the PFD. The air data command displays a Mach number target.

H. Decrease altitude slowly to 24,000 ft

As altitude passes thru 24,900 ft, IAS is annunciated on the GC-550 and the PFD. The air data command displays an airspeed target.

I. Reset all FD modes. Set simulated altitude to ambient & airspeed to zero.


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11. AUTOPILOT TEST Table 11.1. Autopilot Test T ITLE

T EST N O .

NOTE:

FD CMD

C ONTROL W HEEL

R ESPONSE

Use appropriate interconnect to detremine pins & voltages

AP and YD Engage Disengage Logic

NOTE:

W ORK S TEPS

11.1.A. Push AP Engage button on GC-550 Guidance Controller

AP and YD Annunciators on GC-550 are ON and AP and YD indications are on the PFD..

B. Push AP Engage button on GC-550 Guidance Controller again

AP button annunciator turns OFF on the GC-550 and the YD annunciator remains ON. The AP indication of the PFD flashes for 5 sec and turns OFF. The YD indication remains ON.

C. Push the YD button on GC-550 Guidance Controller

The YD annunciator on the GC-550 turns OFF and the YD indication on the PFD turns OFF.

Engage the AP and YD with the AP button on the GC-550 before each of the following work steps D thru J. D. Activate pilots AP/TRIM/PUSHER DISC switch

AP and YD annunciators on the CG-550 are ON AP and YD indications are on the PFD.

E. Activate copilots AP/TRIM/PUSHER DISC switch

Aural warning system sounds. AP and YD flash for 5 seconds on the PFD, then extinguish. AP and YD annunciators on the GC-550 turn OFF.

F. Activate the pilots UP/Down Trim Switch

Aural warning system sounds. AP flashes for 3 seconds on the PFD, then extinguish. YD remains ON. AP annunciators on the GC-550 turn OFF and YD remains ON.

G. Activate the copilots UP/Down Trim Switch


H. Activate the backup test switch on the pedestal in UP/Down or Down position


I. Pull the IC 1 Circuit Breaker

Aural warning system will sound. On the copilots PFD, AP and YD flash for 5 seconds then extinguish. FD failure will be ON. YD remains ON. The pilots PFD, MFD, and EICAS will show Red X..

J. Pull AP 1 Circuit Breaker

Aural warning system will sound. On the PFD AP (red) and YD (amber) will flash for 5 YD annunciators are OFF. seconds and then remain ON until the failure is cleared. ON the GC-550 AP and

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Table 7-1. Ground Maintenance Test (GMT) Procedures Table 11.1. Autopilot Test T ITLE

T EST N O .

W ORK S TEPS

FD CMD

C ONTROL W HEEL

R ESPONSE

Refer to Note 1 in the introduction section prior to performing the following tests. Autopilot Roll Axis

HDG annunciators on FD mode selector GC-550 and PFD shall light. AP & YD ENGAGE lights turn on.

11.2.A Set Hdg Bug on PFD to fore lubber line. Select HDG on FD mode selector GC-550. Center control wheel & engage AP. B. Rotate TURN knob on autopilot controller full clockwise.

CW

HDG annunciators turn off. Autopilot responds to turn knob command.

C. Rotate TURN knob full counterclockwise.

CCW

Autopilot responds to turn knob command.

D. Disengage autopilot with AP TRIM DISC switch.

AP & YD ENGAGE lights turn off.

E. Push & release AP ENGAGE button.

AP & YD ENGAGE lights remain off.

F. Rotate TURN knob to center detent position. Engage AP.

CTR

AP & YD ENGAGE lights turn on.

G. Depress and hold A/P TCS button. Turn control wheel right and then left and then center control wheel.

CTR

AP & YD remains engaged.

H. Release A/P TCS button. I. Disengage AP & YD. Autopilot Pitch Axis

AP & YD ENGAGE lights turn off. CTR

ALT annunciator on GV-550 FD mode selector & PFD shall light. AP & YD ENGAGE annunciate on PFD.

B. Rotate PITCH wheel on autopilot controller towards DESCEND.

FW D

ALT annunciators turn off. Autopilot responds to pitch wheel command.

C. Rotate PITCH wheel on autopilot controller towards CLIMB.

AFT

Autopilot responds to pitch wheel command.

D. Depress A/P TCS and recenter control column. Release A/P TCS.

CTR

Control column free with TCS depressed. Upon releasing TCS autopilot in pitch old mode.

11.3.A Select ALT on FD mode selector GC-550. Center control column & engage AP.

E. Disengage AP.


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Table 7-1. Ground Maintenance Test (GMT) Procedures Table 11.1. Autopilot Test T ITLE Autopilot Trim

T EST N O .

W ORK S TEPS

FD CMD

C ONTROL W HEEL

AP & YD ENGAGE lights turn on

11.4.A Using Aircraft UP/Down Trim Switches located on Yoke, run trim to NOSE UP limit. Move control column slightly forward & engage AP. B. Rotate PITCH wheel on autopilot controller 2 turns towards DESCEND.

R ESPONSE

FW D

After 3 to 4 seconds autopilot trim drives nose down. Approximately 5 seconds later Pitch Trim Indicator on EICAS will move is a pitch down direction. . Verify trim is running in the proper direction.

C. Disengage AP.


D. Using Aircraft UP/Down Trim Switches located on Yoke, run trim to NOSE DOW N limit. Move control column slightly aft & engage AP. Rotate Pitch W heel on Autopilot Controller 2 turns toward Climb.

After 3 to 4 seconds autopilot trim drives nose down. Approximately 5 seconds later Pitch Trim Indicator on EICAS will move is a pitch down direction. . Verify trim is running in the proper direction.

E. Disengage AP & reduce airspeed to zero


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Avionic Emb 145-3

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