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AIR TRANSPORT INTERNATIONAL

DC-8 MAINTENANCE MANUAL VOLUME 2 ATA 27 THRU 34

This manual is the property of Air Transport International, Limited Liability Company. It is issued for the use of Air Transport International, Limited Liability Company authorized personnel.

Revision Number:

Issued:

36

04/06/11

All rights reserved, including the right to reproduce this publication. No part may be reproduced: stored in any retrieval system: or transmitted in any part or form by any means electronic, photocopying, microfilm, microfiche, mechanical, or otherwise, without prior written permission of Air Transport International, Limited Liability Company.

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27- FLIGHT CONTROLS

TABLE OF CONTENTS Page 27.01.1

27.01

ELEVATOR, GUST LOCK OFF WHILE PARKED-GUST LOCK CHECK ---------------------------------------------------------------------------

27.02

SEAL AILERON AEROBALANCE REPAIR-----------------------------------

27.02.1

27.03

SEAL, CONTROL CABLE PRESSURE-WEAR LIMITS/EXTENSION/ LIMITS/CLEANING/REPAIR-----------------------------------------------------

27.03.1

27.04

LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK --------------------------------------------------------------------------------

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27- FLIGHT CONTROLS

27.01 – ELEVATOR, GUST LOCK OFF WHILE PARKED – GUST LOCK CHECK TABLE OF CONTENTS Job Card 27.01-1

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ELEVATOR GUST LOCK OFF WHILE PARKED CHECK------------

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JOB CARD NUMBER

27.01-1 JOB TITLE

ELEVATOR GUST LOCK OFF WHILE PARKED CHECK J/C Reviewed By:

Date:

Page 1 of 3 Mechanic

Inspector

WARNING: DAMAGE CAN OCCUR TO THE ELEVATOR TABS/LINKAGES IF THE AIRCRAFT GUST LOCK IS NOT ON WHILE THE AIRCRAFT IS PARKED. General Information The following procedures must be adhered to any time DC-8 elevators and/or geared tabs are found to be in a position other than faired with the horizontal stabilizer while the airplane is parked. NOTE: If both of the following conditions are met, it will be necessary only to re-engage the gust lock and dispatch the airplane. If any one condition is not met, the entire check MUST be accomplished. CONDITIONS (To be checked in order given): (1) Engagement of the gust lock fairs both elevators and geared (outboard) tabs. (2) The airplane is known to have been parked in winds not exceeding 15 knots for the entire time that the gust lock was disengaged. Tools and Materials Two stands capable of reaching DC-8 elevators Procedure 1. Gust Lock Check a. In tail cone compartment, locate crank. (See Figure 2.) b. Check visually that crank and adjacent parts have no cracks or breaks. c. Replace any cracked or broken parts found. 2. Check elevator geared tab crank assemblies. NOTE: Cranks to be checked are located between leading edge of elevator fairings, lower elevator surface, two locations each side. (Ref. Figure 1.) One person will be required to move the elevator trailing edge by hand, full throw both directions, while the other person is performing the check. Check should be performed in little or no wind. a. Take necessary precautions to ensure safety of personnel and equipment, then release gust lock.

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12-05-03

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

27.01-1 JOB TITLE

ELEVATOR GUST LOCK OFF WHILE PARKED CHECK J/C Reviewed By:

Date:

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Inspector

b. With small mirror, visually check geared tab cranks (two each side) from below for cracks, breaks, and any interference with adjacent structure. This will require looking at cranks while elevator trailing edge is moved from one extreme to the other. c. Replace any cracked or broken cranks found. This will probably not require elevator removal/replacement. (Test flight is required if elevator is removed and replaced.) d. If crank interferes with elevator leading edge cutout, make a minimum of ¼ inch clearance by removing material from leading edge cut-out. e. If cranks have been nicked by previous contact with elevator leading edge cut-out and no cracks are found, blend out nicks with fine emery cloth. Maximum depth of bleed area is .060 inch. Deeper nicks will require crank replacement. Contact LIT Maintenance Control regarding cranks which require blending in a previously blended area. FR prime areas reworked. f. Engage gust lock. 3. Make aircraft logbook entry stating compliance with this job card and any corrective action taken. Comply with Type 1 Airworthiness Release Procedures per GMM 3.101. Attach this completed job card to the log page and forward to the Aircraft Records Section.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS RECURRING NON-ROUTINE

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27.01-1 JOB TITLE

ELEVATOR GUST LOCK OFF WHILE PARKED CHECK J/C Reviewed By:

Date:

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Inspector

GEARED TAB DRIVE LINKAGE Figure 1

GUST LOCK MECHANISM Figure 2 Revision 19

12-05-03

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27- FLIGHT CONTROLS

27.02 – SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016) TABLE OF CONTENTS Paragraph 1.

GENERAL-----------------------------------------------------------------------

Page 27.02.1

2.

SPECIAL TOOLS AND MATERIALS -------------------------------------

27.02.1

3.

REPAIR LIMITATIONS ------------------------------------------------------

27.02.1

4.

REPAIR PROCEDURE--------------------------------------------------------

27.02.3

FIGURES 27.02-1

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SEAL AILERON AEROBALANCE REPAIR MATERIALS -----------------

12-03-01

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.02 SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016) 1. General a. Aerobalance Seals (draft curtains) installed on ailerons prevent flutter at high mach numbers and help effect aerodynamic balancing, thereby reducing control forces. b. Draft holes are provided in the seals and are fully accounted for in the design, but additional venting as might occur with cuts and tears cannot be tolerated. Immediate repair of torn seals is mandatory. Refer to paragraph 3.a. for interim repair limits and paragraph 4.a for interim repair procedure. 2. Special tools and materials a. Tools – None required b. Materials (1) Fabric cap Part Number CRP2386 (aileron curtain). Recommend line stations stock two lineal yards. Note: Available from: Chemprene, Inc. 1360 Logan No. 103 Costa Mesa, California 92623 (213) 229-0709 or (714) 979-0709 (2) Adhesive EC776 (Butadiene Acrlonitrile BUNA-M oil resident elasttomer base.) (3) Adhesive PR10051 Line STA Only (Fuel tank topcoat sealant BUNA-N type MILS- 4383B.) (4) Solvent E145S4 (Ethyl Acetate FED spec TT-E-751A.) (5) Solvent AE0179 (Methyl Ethyl Ketone [MEK] FED Speck TT-M-261B S5) WARNING: ABOVE ITEMS 2 THRU 5 ARE CLASS 1 MATERIAL – EXTREMELY FLAMMABLE (6) Cloth 309147 3. Repair Limitations Notes:

(1) Permanent Repairs may be done at line stations per paragraph 4.b.(following). (2) The general intent of these limitations is to be certain that no aircraft flies for more than 25 hours with more than one permanent tear repair and or 53 coating repair.

Revision 4

12-01-97

27.02.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.02 SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016)(cont’d) a. Interim repair limits – Line Stations only (1) Holes or tears of any size may be repaired per Paragraph 4.a.(following) if the patch extends a minimum of one inch in every direction from the edge of the hole or from the tear. (2) No repairs need to be made to curtains with deteriorated rubber coating IF the exposed fabric is intact and in good condition. (3) If any interim repair is made the one of (a) or (b) following must be done with 25 hours of this time of the repair. (a) If the repair was made to a single hole or tear of five inch or less in length, a permanent repair must be made. (b) If the repair was made to two or more holes or tears, of if the single hole or tear was more than five inches in length, then the curtain must be replaced. (4) Obtain Carry Over Item (Limited to 25 hours). b. Permanent repair limits Note:

Permanent repairs must be done per paragraph 4.b.(following).

(1) Any curtain which has been repaired in any way and in addition has a tear of any size must be replaced. (2) Any curtain which has a two or more tears of any size(whether or not it has had interim repairs) must be replaced. (3) Any curtain which has a single tear more than five inches long (whether or not it has had an interim repair) must be replaced. (4) Any curtain with more than 5% of its area with the rubber coating dissolved by hydraulic fluid or removed by abrasion must be replaced. (5) Any curtain with a single tear five inches or less in length may be repaired so long as it is in otherwise good condition. (6) Any curtain with less than 5% of its area having the rubber coating dissolved by hydraulic fluid or removed by abrasion may be repaired provided the underlying fabric is not damaged.

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Revision 4

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.02 SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016) 4. Repair Procedure a. Interim Repair (1) Secure aileron in the full trailing edge down position. (2) Cut a single patch CRP 2386 to extend at least one inch in all directions beyond the edges of the defect. (3) Clean the fraying surfaces of the patch and seal using a clean cloth 309147 dampened with E14594. Patch will be installed on the lower surface of seal. (4) Apply a liberal coat of PR1005L to patch and seal. Then allow 15 minutes for adhesive to become tacky. (5) Firmly press patch into place on seal ensuring the edges of the patch are sealed. A second application of adhesive around the edges may be desirable. (6) Allow as much time as possible before flexing the seal after installation of the patch. b. Permanent Repairs (1) Patch dimensions and requirements shall be as follows: (a) For cuts or tears up to 3 inches: a single patch which extends ½ inch in all directions from the cut or tear. (b) For cuts or tears 3 to 5 inches: a single patch which extends ¾ inch in all directions from the cut or tear. (c) Holes up to 1 ½ inches in diameter: (double patch on each side of the curtain) which extends ¾ inch in all directions from the borders of the hole. (2) Clean the fraying surfaces of the patch and seal, using a clean cloth 309147 dampened with E14554. (a) Disconnect 2600 series camloc fasteners attaching the seal to wing structure to gain access to upper surface of seal. On airplanes with removable seals it may be easier to remove seal for bench repair. (b) Be sure that run-off fluids on adjacent areas are cleaned up, thus obviating contamination of freshly cleaned areas.

Revision 4

12-01-97

27.02.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.02 SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016) (c) Swab freshly cleaned areas with a cloth 309147 to prevent re-depositing of oils and grime. Note: An absolutely clean surface is the only surface to which the adhesive will stick, if in doubt re-clean the fraying surfaces. (d) apply two uniform brush coats of EC776 to the fraying surfaces of patch and curtain. (e) Assemble repair 30 minutes after application of second coat of adhesive. The adhesive will exhibit pressure sensitive characteristics at end of this period. 1) If the second coat of adhesive becomes inadvertently tack free, it can be activated by swabbing with a clean white cotton cloth dampened with AE0179. (f)

In line maintenance where time is limited, the following recommended time for curing may not be possible, less curing time is acceptable in service. 1) Reconnect camloc fasteners 4 hours after patch is applied. Do not flex ailerons for 12 hours after patch is applied.

(g) Abrasion damage 1) Abrasion damage which is characterized by removal of coating from back fabric but which has not resulted in damage to the fabric may be repaired by applying one heavy coat of EC776. 2) Abrasion damage which is characterized by damage to the fabric and/or holes in the fabric shall be repaired as per previous paragraphs on holes installing a doubler patch.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.02 SEAL AILERON AEROBALANCE REPAIR (AOL 8-1016)

ID 3 5 7 9 11 13

L/H PART NUMBERS P/N DIM 5766710-3 1.5 ft 5766710-5 3 ft 5766710-7 1.5 ft 5766710-9 2 ft 5766710-11 2 ft 5766710-13 7 ft 17 ft

ID 4 6 8 10 12 14

R/H PART NUMBERS P/N 5766710-4 5766710-6 5766710-8 5766710-10 5766710-12 5766710-14

DIM 1.5 ft 3 ft 1.5 ft 2 ft 2 ft 7 ft 17 ft

SEAL AILERON AEROBALANCE REPAIR MATERIALS Figure 27.02-1

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27- FLIGHT CONTROLS

27.03 – SEAL, CONTROL CABIN PRESSURE – WEAR LIMITS / EXTENSION LIMITS / CLEANING / REPAIR TABLE OF CONTENTS Paragraph 1.

GENERAL-----------------------------------------------------------------------

Page 27.03.1

2.

SPECIAL TOOLS AND MATERIALS -------------------------------------

27.03.1

3.

WEAR LIMITS -----------------------------------------------------------------

27.03.2

4.

EXTENSION LIMITS ---------------------------------------------------------

27.03.2

5.

CLEANING OLD TYPE SEALS (WITH CABLE SLEEVES) ----------

27.03.3

6.

CLEANING NEW TYPE SEALS (BARE CABLE)-----------------------

27.03.4

7.

REPAIR – GENERAL ---------------------------------------------------------

27.03.6

8.

REPAIR – SPECIAL TOOLS AND MATERIALS------------------------

27.03.6

9.

REPAIR --------------------------------------------------------------------------

27.03.6

FIGURES 27.03-1

PRESSURE SEAL SLEEVE EXTENSION --------------------------------------

27.03.5

27.03-2

RUBBER DIAPHRAGM ------------------------------------------------------------

27.03.8

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AIR TRANSPORT INTERNATIONAL 27.03 SEAL, CONTROL CABLE LIMITS/CLEANING/REPAIR

DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS PRESSURE – WEAR LIMITS/EXTENSION

1. General a. There are two types of cable pressure seals. Old type uses a swaged sleeve on the cable which slides through a bras bushing. New type uses a two piece plastic bushing through which the bare cable slides. b. The two types of cable are interchangeable. c. Wear limits for old (swaged) seals are as specified herein. Wear limits for new (bare cable) seals are as specified in UAMM, Vol. 1 Bk. 3, GN/MM 1-0-1-3. d. If, on old type seals, the swaged sleeve moves on the cable, the cable must be replaced. e. Cable seals will be cleaned periodically to prevent control friction and damage to the seal. f. Repair damaged seals as per paragraphs 7 thru 9. 2. Special Tools and Materials a. Special Tools – None Required b. Materials (1) Solvent, #840-0010-SOL3004-1 CAUTION:

COMBUSTIBLE.

(2) Fluid #843-0233-FLU3003-2. (3) Grease, #843-0026 GRE4503-6 (MIL-C-16173 Grade 2) (4) Pad, #912-0042-PA134 (5) Oil, #843-0041-OIL4501-10SC (6) Coating (Corrosion Protection), #843-0042-OIL4501-11SC (MIL-C-16173)

Revision 10

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27.03.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d) c. Referenced Procedures (1) Required – None (2) Optional (a) Paragraph 7 thru 9/27.03-2, Seal Repair 3. Wear Limits (sleeved cables) a. SFOOV: (1) Cable sleeves – minimum acceptable outside diameter. CABLE DIAMETER MINIMUM TUBE OD 1/16 INCH .140 INCH 3/32 INCH .140 INCH 1/8 INCH .205 INCH (2) Seal bushings – no limit. Cable tube can be expected to wear out before bushings. b. Line Stations (1) Cable assembly will be replaced if tube swaged on cable is worn through to cable wires and following condition is present at worn area: (a) Any roughness or turned up edges at the worn area in tube. (b) Any broken wires in cable at worn area. 4. Extension Limits (Aft Pressure Dome Only) a. To enable measurement of minimum extension of a seal sleeve on both sides of a given pressure seal, while working on one side, the following maximum extension table is given. A seal sleeve is within limits if on one side of a seal the minimum extension is .500 inch and the maximum is as specified. Table values are conservative so that any sleeve which fails to meet the maximum limit should be checked on the other side of the seal for a minimum of .500 inch extension, prior to repositioning cable. (Ref. Figure 27.03-2).

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d) CABLE FUNCTION

NO. OF MAXIMUM EXTENSION CABLES LIMIT (INCHES) Elevator Control 4 10.37 Rudder Control 2 14.53 Rudder Trim 2 73.66 Gust Lock 2 11.66 Rudder Throw Limiter 2 14.53 Rudder Power Shutoff 2 4.16 Horizontal Stabilizer Position Indicator 2 14.66 Horizontal Stabilizer Motor Control 2 5.16 Horizontal Stabilizer Fluid Supply 2 5.16

5. Cleaning Old Type Seals (With Cable Sleeves) a. Using cloth and SOL3004-1, (1) Use a lint-free cloth dampened in SOL3004-1 to clean the cable sleeve. Do not allow SOL3004-1 to get into the cable strands. Wring out any excess fluid before using cloth. If SOL3004-1 soaks into the cable strands, apply OIL4501-11SC (MIL-C-16173) or GRE4503-6 (MIL-C-16173 Grade 2). CAUTION:

THESE OILS BECOME EXCEEDINGLY STIFF AT LOW TEMPERATURES. AVOID SPRAYING ON MOVING PARTS. REMOVE ANY OVERSPRAY OR BUILD-UP OF THESE MATERIALS FROM MOVINGPARTS.

(2) Get proper clearance to move cables being cleaned and have the cable moved to expose as much of the sleeve as possible on one side of the seal. Clean the exposed part. (3) Reverse the cable travel and clean the opposite end of the sleeve. It will be necessary to work from both sides of the pressure seal to completely clean the sleeves. (4) After cleaning, apply a thin coat of FLU3003-2 to the entire length of the cable sleeve. Move the cable to expose the sleeve inside the seal and coat that part of the sleeve also.

Revision 10

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27.03.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d) b. Using special cleaning pad PA134. (1) Apply FLU3003-2 to the cleaning material. Clean the cable sleeves as in Paragraph 5.A.(2), (3) and (4). NOTE:

Special solvent called out in Paragraph 5.A.(1) need not be used if the cleaning pad, PA134, and fluid, FLU3003-2, are used.

(2) Discard cleaning pad after use. 6. Cleaning New Type Seals (Bare Cable) a. Clean bare cable by spraying with OIL4501-10SC and wiping off with a dry cloth. Do not rub with cleaning pad PA134 as the pad will remove cable plating. NOTE:

If OIL4501-10SC is not available, use SOL3004-1.

b. After cleaning, coat cables by spraying with OIL4501-11SC.

27.03.4

NOTE:

If OIL4501-11SC (MIL-C-16173) is not available, use GRE4503-6 (MILC-16173 Grade 2).

CAUTION:

THESE OILS BECOME EXCEEDINGLY STIFF AT LOW TEMPERATURES. AVOID SPRAYING ON MOVING PARTS. REMOVE ANY OVERSPRAY OR BUILD-UP OF THESE MATERIALS FROM MOVING PARTS.

12-03-01

Revision 14

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d)

PRESSURE SEAL SLEEVE EXTENSION Figure 27.03-1

Revision 10

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27.03.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d) 7. Repair – General a. Pressure seals, consisting of a rubber diaphragm and brass brushing, through which the aluminum tube swaged on the control cable passes, are inspected and cleaned periodically. Cleaning , in accordance with specifications in Paragraphs 1 thru 6, is done to remove tar deposits from cable tube and brass bushing in order to reduce friction and to provide a lubricant at these parts. If cleaning and lubrication is neglected, binding of cable tube within the brass bushing pulls bushing from the rubber diaphragm. With bushing loose on the cable, it is possible for bushing to move along cable to a fair-lead or pulley and cause high friction loads in cable system. b. During routine inspection or cleaning, the following action will be taken if a brass bushing has been pulled from rubber diaphragm portion of the pressure seal. 8. Repair – Special Tools and Materials a. Special Tools – None Required b. Materials (1) Solvent, #840-0151-SOL3000-2 (2) Cloth, #912-0041-CL511 (3) Sealant, #841-0450-SEA3750-5 9. Repair NOTE: REF TO UAMM Vol. 1 Bk. 3 GN/MM-6-0-2-9 for SEA3750-5 mixing and curing. a. If rubber diaphragm is in good condition, bushing may be reinstalled as follows: (1) Clean bushing surface that contacts rubber with a rag dampened in SOL300-2 and wipe dry with a clean dry CL511. CAUTIONS: (A)

(B)

DO NOT ALLOW SOLVENT TO CONTACT THE CABLE SURFACE BEYOND THE CABLE TUBE. DO NOT ALLOW SOLVENT TO PUDDLE ON SEALANT OR RUBBER AS THESE MATERIALS WILL SWELL AND DEGRADE.

(2) Check rubber diaphragm hole for loose particles of rubber and general condition. (3) Coat bushing with a brush coat of SEA3750-5 around center part and push back into rubber diaphragm.

27.03.6

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Revision 14

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d) b. If rubber diaphragm is torn or split it may be repaired with SEA3750-5 as long as there is sufficient rubber left to form a good base. (1) Clean away fragments or loose rubber parts. (2) Clean and coat bushing with sealant as noted in 9. a. (3). (3) Place bushing in rubber diaphragm and build up as necessary with SEA3750-5. Insure that no sealant is allowed to contact or adhere to cable. (4) Be sure no sealant forms on aluminum tube in cable. c. If rubber diaphragm has deteriorated so as to require replacement and cables need not be loosened or removed for another reason, proceed as follows: NOTE:

This procedure to be used only on diaphragms with four or fewer holes AND support (by plate, skin, etc.) over their entire area. (Ref. Figure 27.03-2). If diaphragm has more than four holes and/or is not supported over the entire area, contact SFOEG.

(1) Carefully cut new diaphragm from the closest edge to each bushing hole. (Ref. Figure 27.03-2 for a typical diaphragm showing cuts.) (2) Cut out old diaphragm and remove fasteners. (3) Install new diaphragm in place and coat cut edges with SEA3750-5 so that they will adhere when cured. Add sealant over seam as necessary. Coat bushing and slip in place. (4) Be sure no sealant forms on aluminum tube on cable. d. Clean and lubricate cable tube and bushing ID per Paragraphs 1 thru 6.

Revision 10

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27.03.7

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.03 SEAL, CONTROL CABLE PRESSURE – WEAR LIMITS/EXTENSION LIMITS/ CLEANING/REPAIR (cont’d)

RUBBER DIAPHRAGM Figure 27.03-2

27.03.8

12-03-01

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS

27.04 – LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK TABLE OF CONTENTS Paragraph A.

GENERAL--------------------------------------------------------------------------

Page 27.04.1

RECURRING NON-ROUTINE JOB CARDS Page 27.04.2

Job Card 27.04-1

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION ----------

27.04-2

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK ----------- 27.04.10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK 1. General a. Non-routine Job Cards DC8-27.04-1 and DC8-27.04-2 are applicable for the DC-8 longitudinal trim jackscrew. b. Neither job card checks for a failure. Instead, JC DC8-27.04-1 provides a lubrication procedure for the jackscrew. JC DC8-27.04-2 checks for wear to make certain that longitudinal trim jackscrew wear has not exceeded the service limits. c. Accomplishment Intervals: (1)

Accomplish Non-routine Job Card DC8-27.04-1 at intervals not to exceed 1,200 flight hours.

(2)

Accomplish Non-routine Job Card DC8-27.04-2 at intervals not to exceed 8,500 flight hours.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

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DATE

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ATI-001

JOB CARD NUMBER

DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

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References Boeing Report MDC-02K9030, SCI 27-1 “DC-8 Longitudinal Trim Jackscrew Lubrication” Mechanic Inspector Procedure 1. General: a. The grease removal and lubrication procedures must be done together and in the sequence listed. b. The use of clean, dry rags instead of solvent to remove the grease from the jackscrews and jackscrew drive nuts is recommended to prevent solvent contamination in the jackscrew drive nuts. c. When lubricating vented joints (grease can freely and visibly flow out), apply grease to the fittings until only new grease comes out of the joints. d. Do not intermix different grease specifications or different grease brands within the same specification. e. When changing from one grease specification or brand to another, follow the procedure “Lubrication Grease Change – Maintenance Practices” in DC-8-60 (or 70) MM 12-80-0, as applicable. f. Refer to applicable DC-8-60 (or 70) MM, Chapter 12 for general lubrication instructions. g. The longitudinal trim jackscrews are located in the aft fuselage section, forward of the horizontal stabilizer center section. Access to the jackscrews is through the aft fuselage access door. 2. Remove the used grease from the longitudinal trim jackscrews: a. Move the horizontal stabilizer to the full aircraft nose down (AND) position. WARNING: BEFORE PRESSURIZING THE MAIN HYDRAULIC POWER SYSTEM, MAKE SURE THAT THE LANDING GEAR GROUND LOCKPINS ARE INSTALLED AND THAT APPLICABLE CONTROLS ARE IN THE CORRECT POSITION TO PREVENT INADVERTENT OPERATION FO THE LANDING GEAR AND FLIGHT CONTROL SYSTEMS. (1) Pressurize the main hydraulic power system (see DC8-60 (or 70) MM, Chapter 29).

27.04.2

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

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ATI-001

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DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

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Inspector

WARNING:

BEFORE MOVING THE HORIZONTAL STABILIZER, MAKE SURE THAT THE AREAS AROUND THE STABILIZER ARE CLEAR OF PERSONNEL AND EQUIPMENT. (2) Using the longitudinal stabilizer trim slide buttons on either aileron control wheel, move the stabilizer to the full AND position. (3) Depressurize the main hydraulic power system and relieve the pressure from the system reservoir (see DC-8-60 (or 70) MM, Chapter 29). WARNING: TAG AND SAFETY CIRCUIT BREAKERS. (4)

Open and tag the following circuit breakers on the EPC circuit breaker panel: (a) Three Autopilot and Alternate Longitudinal Trim CBs located in the radio AC bus section. (b) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (c) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. (5) Put a “DO NOT OPERATE” tag on the longitudinal trim control handles on the pilot’s control pedestal. b. Clean the exposed threads on the left and right jackscrews above the jackscrew drive nuts and the tops of the jackscrew drive nuts with clean, dry rags to remove as much grease as possible (see Figure 27.04-1). c. Move the horizontal stabilizer to the full aircraft nose up (ANU) position. (1) Remove the safety tags and close the following circuit breakers, as required, on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (b) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system.

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27.04.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

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DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

Page 3 of 8 Mechanic

Inspector

WARNING:

BEFORE PRESSURIZING THE MAIN HYDRAULIC POWER SYSTEM, MAKE SURE THAT THE LANDING GEAR GROUND LOCKPINS ARE INSTALLED AND THAT APPLICABLE CONTROLS ARE IN THE CORRECT POSITION TO PREVENT INADVERTENT OPERATION OF THE LANDING GEAR AND FLIGHT CONTROL SYSTEMS. (2) Pressurize the main hydraulic system (see DC-8-60 (or 70) MM, Chapter 29. WARNING: BEFORE MOVING THE HORIZONTAL STABILIZER, MAKE SURE THAT THE AREAS AROUND THE STABILIZER ARE CLEAR OF PERSONNEL AND EQUIPMENT. (3) Using the longitudinal stabilizer trim slide buttons on either aileron control wheel, move the stabilizer to the full ANU position. (4) Depressurize the main hydraulic power system and relieve the pressure from the system reservoir (see DC-8-60 (or 70) MM, Chapter 29). WARNING: TAG AND SAFETY CIRCUIT BREAKERS. (5) Open and tag the following circuit breakers on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (b) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system. d. Clean the exposed threads on the left and right jackscrews below the jackscrew drive nuts and the bottoms of the jackscrew drive nuts with clean, dry drags to remove as much grease as possible. 3. Lubricate the longitudinal trim jackscrews (see Figure 27.04-1): a. Put a layer of grease around the complete circumference of the exposed threads on the left and right jackscrews. Make sure grease is liberally applied to the threads, including the flanks and bottom of the thread valley.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

Page 4 of 8 Mechanic

Inspector

b. Move the horizontal stabilizer to the full AND position. (1)

Remove the safety tags and close the following circuit breakers, as required, on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (b) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system. WARNING: BEFORE PRESSURIZING THE MAIN HYDRAULIC POWER SYSTEM, MAKE SURE THAT THE LANDING GEAR GROUND LOCKPINS ARE INSTALLED AND THAT APPLICABLE CONTROLS ARE IN THE CORRECT POSITION TO PREVENT INADVERTENT OPERATION OF THE LANDING GEAR AND FLIGHT CONTROL SYSTEMS. (2) Pressurized the main hydraulic power system (see DC-8-60 (or 70) MM, chapter 29. WARNING: BEFORE MOVING THE HORIZONTAL STABILIZER, MAKE SURE THAT THE AREAS AROUND THE STABILIZER ARE CLEAR OF PERSONNEL AND EQUIPMENT. (3) Using the longitudinal stabilizer trim slide buttons on either aileron control wheel, move the stabilizer to the full AND position. (4) Depressurize the main hydraulic power system and relieve the pressure from the system reservoir (see DC-8-60 (or 70) MM, Chapter 29). WARNING: TAG AND SAFETY CIRCUIT BREAKERS. (5) Open and tag the following circuit breakers on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections.

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27.04.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

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DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

Page 5 of 8 Mechanic

Inspector

(b)

One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system. c. Put a layer of grease around the complete circumference of the exposed threads on the left and right jackscrews. Make sure grease is liberally applied to the threads, including the flanks and bottom of the thread valley. d. Apply grease to the left and right bearing housings. e. Move the horizontal stabilizer to distribute the grease equally in the jackscrews. (1) Remove the safety tags and close the following circuit breakers, as required, on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (b) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system. WARNING: BEFORE PRESSURIZING THE MAIN HYDRAULIC POWER SYSTEM, MAKE SURE THAT THE LANDING GEAR GROUND LOCKPINS ARE INSTALLED AND THAT APPLICABLE CONTROLS ARE IN THE CORRECT POSITION TO PREVENT INADVERTENT OPERATION OF THE LANDING GEAR AND FLIGHT CONTROL SYSTEMS. (2) Pressurize the main hydraulic power system (see DC8-60 (or 70) MM, Chapter 29). WARNING: BEFORE MOVING THE HORIZONTAL STABILIZER, MAKE SURE THAT THE AREAS AROUND THE STABILIZER ARE CLEAR OF PERSONNEL AND EQUIPMENT. (3) Using the longitudinal stabilizer trim slide buttons on either aileron control wheel, move the stabilizer to the full ANU and then the full AND positions.

27.04.6

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

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DC8-27.04-1 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW LUBRICATION Job Card Reviewed By:

Date:

Page 6 of 8 Mechanic

(4)

Inspector

Move the horizontal stabilizer to the neutral position.

(5)

Depressurize the main hydraulic power system and relieve the pressure from the system reservoir (see DC-8-60 (or 70) MM, Chapter 29). WARNING: TAG AND SAFETY CIRCUIT BREAKERS. (6) Open and tag the following circuit breakers on the EPC circuit breaker panel: (a) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (b) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. Do not close this circuit breaker unless the auxiliary pump is used to pressurize the hydraulic system. f. Remove the grease that is not necessary from the left and right jackscrews. Do not remove the grease from the working surfaces. g. Clean the structural areas around the jackscrews. Remove any grease that is not necessary and any loose materials (sand, dirt, drill shavings, etc.) that could contaminate the jackscrews. h. Remove the safety tags and close the following circuit breakers on the EPC circuit breaker panel: (1) Three Autopilot and Alternate Longitudinal Trim CBs located in the radio AC bus section. (2) Two Wheel Longitudinal Trim CBs located in the radio AC and DC bus sections. (3) One Auxiliary Hydraulic Pump Control CB located in the cabin bus section. i. Remove the “DO NOT OPERATE” tag from the longitudinal trim control handles on the pilot’s control pedestal. 4. Ensure all job card sign offs are complete.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d)

LONGITUDINAL TRIM JACKSCREWS Figure 27.04-1 (1 of 2) 27.04.8

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) ITEM NO. 1 2

ITEM DESCRIPTION (NO. PER AIRCRAFT) Jackscrew (2) Bearing Housing (2)

LUBE TYPE

APPLICATION

GMD GMD

Brush Gun

NUMBER OF FITTINGS OR AREAS 2

LONGITUDINAL TRIM JACKSCREWS Figure 27.04-1 (2 of 2)

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

JOB CARD NUMBER

DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 1 of 8

References Boeing Report MDC-02K9030, SCI 27-2 “DC-8 Longitudinal Trim Jackscrew Wear Check” Tools Needed ITEM NAME NUMBER MANUFACTURER USE A Horizontal stabilizer 3965691-1 Boeing Apply load to jackscrew jackscrew endplay during wear check. tool B Dial indicator and Local Measure end play. clamps or magnetic base C Torque wrench Local Measure load applied to 75-100 ft-lb range jackscrew during wear check. D FR Primer or similar Local Mark roller and jackscrew paint sprocket for reinstallation Mechanic Inspector Procedure 1. General: a. The longitudinal trim jackscrews are located in the aft fuselage section, forward of the horizontal stabilizer center section. Access to the jackscrews is through the aft fuselage access door. b. The item numbers in parentheses in the following text refer to items in Figures 27.04-2 and 27.04-3. 2. Inspection/Check Longituindal Trim Jackscrew Wear: a. Lubricate the longitudinal trim jackscrews (see JC DC827.04-1). WARNING: BEFORE MOVING THE HORIZONTAL STABILIZER, MAKE SURE THAT THE AREAS AROUND THE STABILIZER ARE CLEAR OR PERSONNEL AND EQUIPMENT. b. Using the alternate longitudinal trim control levers, move the horizontal stabilizer to the 4-degree nose up (ANU) position. NOTE: The stabilizer is at the 4-degree ANU position when the distance between the serrated mounting flanges for the left jackscrew stops (1 and 2) is 7-7/32 (± 1/16) inches (see Figure 27.04-2, View B-B, Dimension “A”).

27.04.10

04-28-10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

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ATI-001

JOB CARD NUMBER

DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 2 of 8 Mechanic

Inspector

c. Depressurize the main hydraulic power system and relieve the pressure from the system reservoir (see DC-8-60 (or 70) MM, Chapter 29). WARNING: TAG AND SAFETY CIRCUIT BREAKERS. d. Open and tag the following circuit breakers on the EPC circuit breaker panel: (1) Three Autopilot and Alternate Longitudinal Trim circuit breakers located in the Radio AC bus section. (2) Two Wheel Longitudinal Trim circuit breakers located in the radio AC and DC bus sections. (3) One Auxiliary Hydraulic Pump Control circuit breaker located in the cabin bus section. e. Put a “DO NOT OPERATE” tag on the longitudinal trim control handles on the pilot’s control pedestal. f. Loosen the bolts that attach the sprocket housing sufficiently to allow the housing to move laterally (see Figure 27.04-2). g. Check the first jackscrew; either the left or right jackscrew may be checked first: (1) Loosen the bolts in the clamp that attached the roller chain support to the turnbuckle, or remove the support from the turnbuckle. (2) Clean small, adjacent areas on the roller chain and the jackscrew sprocket. Using FR primer or a similar paint, put index marks on the chain and sprocket. NOTE: The index marks will be used to reinstall the chain at its original location on the jackscrew sprocket.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

JOB CARD NUMBER

DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 3 of 8 Mechanic

Inspector

CAUTION: DO NOT LET THE ROLLER CHAIN DISENGAGE FROM THE DRIVE SPROCKET IN THE SPROCKET HOUSING, WHILE THE CHAIN TENSION IS RELIEVED AND THE CHAIN IS REMOVED FROM THE JACKSCREW SPROCKET. THE CHAIN CAN BE MISPLACED WHEN IT IS REINSTALLED ON THE DRIVE SPROCKET, CAUSING THE LEFT AND RIGHT JACKSCREWS TO BECOME UNSYNCHRONIZED. THIS CONDITION COULD CAUSE INCORRECT JACKSCREW STOP CONTACT AND/OR HORIZONTAL STABILIZER BINDING.

(3) (4)

IF THE ROLLER CHAIN BECOMES DISENGAGED FROM THE DRIVE SPROCKET, DO NOT MOVE THE HORIZONTAL STABILIZER UNTIL THE JACKSCREWS AND JACKSCREW STOPS HAVE BEEN ADJUSTED, THE CHAINS ARE INSTALLED, AND THE CHAIN TENSION (SLACK) IS ADJUSTED PER THE JACKSCREW INSTALLATION PROCEDURE (SEE DC-8-60 (OR 70) MM, CHAPTER 27-41-1, AS APPLICABLE. Adjust the turnbuckle to relieve the roller chain tension. Remove the roller chain from the jackscrew sprocket.

(5)

Remove the cotter pin (3), nut (4), and washer (5) from the jackscrew attach bolt (6). (6) Manually rotate the jackscrew sprocket to relieve the load on the jackscrew attach bolt (6) and then remove the bolt. (7) Install the jackscrew endplay tool (see Figure 27.043). CAUTION: TIGHTEN THE ENDPLAY TOOL SCREW ASSEMBLY BY HAND ONLY TO PREVENT DAMAGE TO THE AIRCRAFT STRUCTURE. (8) Tighten the endplay tool screw assembly by hand to hold the tool against the horizontal stabilizer center section structure.

27.04.12

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

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ATI-001

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DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 4 of 8 Mechanic

Inspector

NOTE:

The screw assembly is sufficiently tightened if the tool remains against the stabilizer structure during the wear check. (9) Clean any grease or loose material from the area of the drive nut upper stop (2) where the dial indicator probe will touch. CAUTION: BE CAREFUL NOT TO DAMAGE THE SURFACE OF THE JACKSCREW WHEN ATTACHING AND REMOVING THE DIAL INDICATOR. (10) Attach the dial indicator to the jackscrew using clamps or a magnetic base. (a) Place the indicator probe against the upper surface of the drive nut upper stop (2). (b) Make sure the dial indicator probe is placed against the level (not sloped) portion of the stop and that the probe is parallel to the jackscrew. (11) Move the dial indicator toward the stop until the probe is compressed approximately ½ travel and the indicator needle is pointed up. (12) Connect the torque wrench to the endplay tool. Align the torque wrench parallel to the section of the tool arm that points down, as shown in Figure 27.04-3. (13) Pull up on the torque wrench to obtain a reading of 75-100 ft-lbs and then hold this load on the jackscrew. (14) While holding the load on the jackscrew, set the dial indicator to indicate zero. NOTE: When the load on the jackscrew is released, the dial indicator indication may move away from zero. Do not reset the indicator to zero when the load is released. (15) Push down on the torque wrench to obtain a reading of 75-100 ft-lbs and then hold this load on the jackscrew. (16) Record the dimension shown on the dial indicator. (17) Pull up on the torque wrench to obtain a reading of 75-100 ft-lbs and then hold this load on the jackscrew.

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27.04.13

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

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ATI-001

JOB CARD NUMBER

DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 5 of 8 Mechanic

Inspector

(18) While holding the load on the jackscrew, check that the dial indicator returned to zero. If it did not, reset the indicator to zero before releasing the load. (19) Repeat steps (15) through (18) until three consecutive dial indicator dimensions are recorded that are no more than 0.001 inch apart, with a return to zero each time. Retain these three consecutive dimensions. (20) Review the three consecutive dimensions retained in step (19). If any of the dimensions are less than 0.004 inch or more than 0.045 inch, replace the jackscrew. (21) Remove the dial indicator and attachment clamps or magnetic base and the jackscrew endplay tool. (22) Install the jackscrew attach bolt (6), washer (5), and nut (4). Safty the nut with a new cotter pin (3). (23) Install the roller chain on the jackscrew sprocket with the painted index marks on the chain and sprocket aligned. If necessary, manually rotate the jackscrew sprocket to align the marks. (24) Adjust the turnbuckle to provide enough tension in the roller chain to prevent the chain from falling off of the drive and jackscrew sprockets. This is a preliminary chain adjustment. h. Perform steps g(1) thru g(24) on the second jackscrew being checked. i. Alternately adjust the left and right turnbuckles until the slack at the midspan position of each chain is ¼ to ½ inch when a force of 8 to 10 pounds is applied. j. Tighten the bolts that attach the sprocket housing. k. Tighten the bolts in the clamps that attach the left and right roller chain supports to the turnbuckles or, if removed, install the supports on the turnbuckles. l. Lubricate the left and right roller chains and the left and right jackscrew attach bolts (6) and clevises (see DC-8-60 (or 70) MM, Chapter 12-82-1, as applicable). m. Remove the safety tags and close the following circuit breakers on the EPC circuit breaker panel: (1) Three Autopilot and Alternate Longitudinal Trim circuit breakers located in the Radio AC bus section.

27.04.14

04-28-10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

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POS

DATE

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ATI-001

JOB CARD NUMBER

DC8-27.04-2 JOB TITLE

DC-8 LONGITUDINAL TRIM JACKSCREW WEAR CHECK Job Card Reviewed By:

Date:

Page 6 of 8 Mechanic

Inspector

(2)

Two Wheel Longitudinal Trim circuit breakers located in the radio AC and DC bus sections. (3) One Auxiliary Hydraulic Pump Control circuit breaker located in the cabin bus section. n. Remove the “DO NOT OPERATE” tag from the longitudinal trim control handles on the pilot’s control pedestal. o. Test the jackscrews (see DC-8-60 (or 70) MM Chapter 2741-1, paragraph 4.A. “Test Jackscrew,” as applicable). 3. Ensure all job card sign offs are complete.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d)

DC-8 LONGITUDINAL TRIM JACKSCREW & HOUSING Figure 27.04-2 27.04.16

04-28-10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 27 – FLIGHT CONTROLS 27.04 LONGITUDINAL TRIM JACKSCREW LUBRICATION & WEAR CHECK (cont’d)

DC-8 LONGITUDINAL TRIM JACKSCREW & ENDPLAY TOOL Figure 27.04-3

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INTENTIONAL BLANK

27.04.18

04-28-10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL

FUEL TABLE OF CONTENTS 28.01

SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS -----------

Page 28.01.1

28.02

ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK-------------------------------------------------

28.02.1

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INTENTIONAL BLANK

28-ii

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 - FUEL

28.01 – SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS TABLE OF CONTENTS Paragraph A.

GENERAL--------------------------------------------------------------------------

Page 28.01.1

B.

DC8 AIRCRAFT-------------------------------------------------------------------

28.01.1

RECURRING NON-ROUTINE JOB CARDS Job Card 28.01-1 28.01-2 28.01-3 28.01-4 28.01-5 28.01-6 28.01-7 28.01-8 28.01-9 28.01-10 28.01-11 28.01-12 28.01-13 28.01-14 28.01-15 28.01-16

Revision 36

Page DC-8 REPAIR WIRE GAUGE, TYPE, & INSTALLATION INSPECTION ---------------------------------------------------------------------DC-8 FUEL QUANTITY INDICATION SYSTEM (FQIS) WIRE ISOLATION INSPECTION-----------------------------------------------------DC-8 IN-TANK WIRE SPLICE INSPECTION------------------------------DC-8 PROHIBITION OF POWER FEEDER SPLICES IN AREAS ABOVE FUEL TANKS INSPECTION ---------------------------------------DC-8 FUEL TANK ACCESS DOOR BONDING INSPECTION ---------DC-8 FUEL TANK STRUCTURAL DESIGN INSPECTION-------------DC-8 WIRING NEAR FUEL TANKS & FUEL PIPES INSPECTION---DC-8 PROHIBITION OF SAFETY WIRE, LOCKWIRE, SAFETY CABLE, OR COTTER PINS INSIDE FUEL TANKS INSPECTION ----DC-8 FUEL PUMP (BOOST OR TRANSFER) AND HOUSING ASSEMBLIES INSPECTION --------------------------------------------------DC-8 FUEL PUMP (BOOST OR TRANSFER) AND HOUSING ASSEMBLIES INSPECTION (WELDON FUEL PUMP)------------------DC-8 FUEL PUMP ELECTRICAL CONNECTOR WIRE/CONDUIT ASSEMBLY INSPECTION-----------------------------------------------------DC-8 OUTER WING SURFACE PREPARATION INSPECTION -------DC-8 FUEL QUANTITY MASTER INDICATORS INSPECTION ------DC-8 FUEL PUMP CONNECTOR INSPECTION --------------------------DC-8 FUEL PUMP CIRCUIT BREAKER RESET INSPECTION--------DC-8 FLOAT SWITCH CIRCUIT INSPECTION----------------------------

04-06-11

28.01.3 28.01.4 28.01.5 28.01.6 28.01.7 28.01.8 28.01.9 28.01.10 28.01.11 28.01.13 28.01.14 28.01.15 28.01.17 28.01.18 28.01.19 28.01.20

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INTENTIONAL BLANK

28.01-ii

04-06-11

Revision 36

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS [FAR 121.1113, 25.981(b), 25.1309(b)(d), 25.1529, 91.403(c), SFAR 88] A. General – SFAR 88 requirements have been developed by the manufacturer to enhance the inspection program and to define the process of maintenance being performed to ensure the highest level of safety. The following guidelines will be followed for each aircraft within the ATI fleet. B. DC8 Aircraft - Report number MDC-02K9030 was written to provide guidance of Special Compliance Items as required by FAR 121.1113 and 25.1529 and SFAR 88 Airworthiness Limitation Instructions and Critical Configuration Control Limitations required by FAR 25.981 and 25.1529 and has been included in the ATI Continuous Airworthiness Maintenance Program. 1. Critical Design Configuration Control Limitations a.

b. c. d.

CDCCL’s are features that Boeing has identified that are not always obvious or intuitive to the mechanic, that a change could impact safety. By identifying certain design configuration features, Boeing developed these CDCCL’s that are intended to preclude a fuel tank ignition source for the operational life of the aircraft. The terms of the appropriate DC-8 Type Certificate are only valid when the design configuration aspects specified by these CDCCL’s are maintained as stated. CDCCL’s are mandatory and cannot be changed or deleted without the approval of the FAA Engineering at the LA-ACO. Recurring Non-Routines have been developed for each CDCCL and have been incorporated in the ATI DC8 Maintenance Manual, Chapter 28 and must be accomplished whenever repairs are accomplished to the fuel system.

2. Airworthiness Limitation Instructions a. b. c. d.

ALI’s are a means of detecting safety significant conditions potentially making the aircraft non-compliant with SFAR 88 and FAR 25.981. When an ALI detects such a condition, maintenance action to correct it is required. ALI’s are mandatory and cannot be changed or deleted without the approval of the FAA Engineering at the LA-ACO. Recurring Non-Routines have been developed for each ALI and have been incorporated in the ATI DC8 Maintenance Manual, Chapter 28.

NOTE: When accomplishing repairs relating to fuel system, each mechanic is responsible for the adherence to and the accomplishment of the Recurrent Non-Routine for the type work being performed. 3. Short Term Extensions a.

Revision 31

Short Term Extensions may be authorized in certain cases. Chapter 9 of the ATI Reliability Control Program Manual (RCPM) details the limitations of such an escalation. All escalations must have the concurrence of the FAA cognizant inspector prior to entering the extension. 11-20-08

28.01.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) 4. Paragraph J, “ATI Accomplishment of Boeing SCIs, CDCCLs, & ALIs,” in the Introduction to this Manual (Section 00) lists the recurring non-routines that will be used when performing repairs on the fuel system. After completion of the job card, attach the job card to the aircraft logbook page and forward to LIT Q.C.

28.01.2

04-28-10

Revision 34

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-1 JOB TITLE

DC-8 REPAIR WIRE GAUGE, TYPE, & INSTALLATION INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-1 “DC-8 Repair Wire Gauges” Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-4 “DC-8 Repair Wire Types” Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-5 “DC-8 Wire Installation” Applicability Mechanic Procedure 1. Verify that all repair wiring used inside or within 6” of the fuel tanks and fuel pipes uses the exact same Applicable Not Applicable wire gauge as the wire being replaced. [Reference CDCCL 20-1.] 2. Verify that any repair wiring used inside or within 6” of the fuel tanks and fuel pipes is not one of the following wire types: (a) (b) (c) (d)

DES-SM2006 (PVC/Nylon) 7484444 (“Quad 4,” PVC/Fiberglass/Nylon) 7616964 (PVC/Nylon & PVC/Fiberglass/Nylon) 7706560 (MIL-W-81044/9, XL-Polyolefin, XLKynar) (e) 7891145 (was Poly-X, then later was Stilan) (f) BXS7007 (Aromatic Polyimide)

Inspector

Applicable Not Applicable

[Reference CDCCL 20-4.] NOTE: Aircraft as originally built with previously approved wire types are acceptable. New repairs should only be made with OEM approved wire types. 3. Verify that all repair of wiring used inside or within 6” of the fuel tanks and fuel pipes uses the existing wire routing and clipping established per OEMApplicable approved wire installation or maintains wire Not Applicable clearances as prescribed in the applicable sections of SWPM 20-00-06 and/or SWPM 10-10-01. [Reference CDCCL 20-5.] NOTES:

Revision 36

04-06-11

28.01.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-2 JOB TITLE

DC-8 FUEL QUANTITY INDICATION SYSTEM (FQIS) WIRE ISOLATION INSPECTION Job Card Reviewed By: Date: Page 1 of 1 References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-2 “DC-8 FQIS Wire Isolation” Applicability Mechanic Procedure 1. Verify that all FQIS wires are of wire type DA12266 Applicable DA14286, 358-XE or equivalent, from the Master Not Applicable Indicator to the Tank entry point(s). 2. Verify that all FQIS wires maintain existing wire routing established per OEM-approved wire Applicable installation drawings or maintain wire clearances as Not Applicable prescribed in the applicable sections of SWPM 2000-06 and/or SWPM 20-10-01. 3. Verity that all FQIS wires are isolated or separated from other aircraft wiring in order to prevent chafing Applicable Not Applicable that can lead to shorting to an incompatible power source. NOTES:

28.01.4

04-06-11

Inspector

Revision 36

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-3 JOB TITLE

DC-8 IN-TANK WIRE SPLICE INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-3 “DC-8 In-Tank Wire Splices” Applicability Mechanic Procedure 1. Verify that no wire splices of any type, including crimp or solder type, are present on wires in the fuel Applicable Not Applicable tank or on wires inside conduits that are in the fuel tank. NOTE: This includes (1) wiring going to AC and DC fuel pumps, (2) wiring belonging to the Fuel Quantity Gauging System (FQGS), and (3) wiring going to fuel high-level and low-level float switches, as applicable; but this does not include wiring inside a LRU. NOTES:

Revision 31

11-20-08

Inspector

28.01.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-4 JOB TITLE

DC-8 PROHIBITION OF POWER FEEDER SPLICES IN AREAS ABOVE FUEL TANKS INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-6 “DC-8 Prohibition of Power Feeder Splices in Areas Above Fuel Tanks” Applicability Mechanic Inspector Procedure 1. Verify that Power Feeders above any fuel tank are maintained without splices. Any needed repair Applicable splice needs to be implemented at a location at least Not Applicable 3 feet or 2 wire supports (whichever is further) from the fuel tank boundaries. NOTE: Power feeders are defined as any wire 10 gauge and larger (10, 8, 6, 4, 2, 0, 00, 000 and 0000 wire gauges). NOTES:

28.01.6

11-20-08

Revision 31

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-5 JOB TITLE

DC-8 FUEL TANK ACCESS DOOR BONDING INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-7 “DC-8 Fuel Tank Access Door Bonding” Applicability Mechanic Inspector Procedure CAUTION: DEPARTING FROM THE OEM DESIGNED ELECTRICAL CURRENT PATH ACROSS DIFFERENCE INTERFACES COULD CREATE HAZARDOUS IGNITION SOURCES INSIDE THE TANK. 1. Verify that the following are maintained as defined in the applicable OEM design data which includes Boeing design drawings or the Boeing SRM or the Boeing AMM or Boeing approved repairs or any combination of these: (a) Conductive path design (b) Surface treatments, sealants, sealing procedures, seals, and O-rings (c) Fastener types, materials, conductive rubstrips, interference fit fasteners, and coatings (d) Non-conductive features of the door installation, including anodizing, priming and sealing, as well as the seals on the door and the dome nuts NOTES:

Revision 36

04-06-11

Applicable Not Applicable

28.01.7

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-6 JOB TITLE

DC-8 FUEL TANK STRUCTURAL DESIGN INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-8 “DC-8 Fuel Tank Structural Design” Applicability Mechanic Inspector Procedure 1. Verify that the following are maintained as defined in the applicable OEM design data which includes Boeing design drawings or the Boeing SRM or the Boeing AMM or Boeing approved repairs or any combination of these: Applicable Not Applicable

(a) Conductive and non-conductive path design (b) Surface treatments, sealants, and sealing procedures (c) Fastener types, and coatings NOTES:

28.01.8

04-06-11

Revision 36

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-7 JOB TITLE

DC-8 WIRING NEAR FUEL TANKS & FUEL PIPES INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-9 “DC-8 Wiring Near Fuel Tanks and Fuel Pipes” Applicability Mechanic Inspector Procedure 1. Verify that all wires/wire assemblies, when installed with 6 inches of any fuel tank or fuel pipe, conform to Standard Wiring Practice Manual (Boeing Applicable Not Applicable Document #D6-82481), Section 20-00-06, “Inspection Instructions for Standard Wiring Installations”. NOTES:

Revision 31

11-20-08

28.01.9

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-8 JOB TITLE

DC-8 PROHIBITION OF SAFETY WIRE, LOCKWIRE, SAFETY CABLE, OR COTTER PINS INSIDE FUEL TANKS INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-1 “DC-8 Prohibit Safety Wire, Lockwire, Safety Cable, or Cotter Pins Inside Fuel Tanks” Applicability Mechanic Inspector Procedure 1. Verify that no safety wire, lockwire, safety cable, or cotter pins are installed inside the fuel tanks. Applicable Components or equipment installed inside fuel tanks Not Applicable shall not use safety wire or cotter pins for hardware retention. NOTE: Safety wire, lockwire, safety cable, and cotter pins may be used if they are contained within the housing of an explosion proof, tank mounted LRU and must be installed per applicable OEM design/repair/maintenance documentation. In-shop LRU repairs must adhere to OEM CMM documented procedures and sole use of parts specified in the OEM’s CMM. NOTES:

28.01.10

11-20-08

Revision 31

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-9 JOB TITLE

DC-8 FUEL PUMP (BOOST OR TRANSFER) AND HOUSING ASSEMBLIES INSPECTION Job Card Reviewed By:

Date:

Page 1 of 2

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-2 “DC-8 Fuel Pump (Boost or Transfer) and Housing Assemblies” Applicability Mechanic Inspector Procedure NOTE: This card is applicable only to the following: (1) Fuel Pump, Impeller, and Motor Assembly P/N 112-303-1, 112-303-11, 1711460, 1711461 (2) Fuel Pump, Scroll Assembly P/N 112-304, 112-332, 112-631 (3) Pump Inlet System Components Any repair to the part numbers listed above must be made in accordance with the OEM’s Component Maintenance Manuals (CMMs), 28-1-1, Revision 8, or an FAA approved service bulletin. Any deviation from the OEM’s CMM must be justified by the modifier and approved by the FAA LA-ACO. CAUTION: EXPOSURE OF PUMP(S) AND VOLUTE/SCROLL ALUMINUM COMPONENTS TO MERCURY MAY RESULT IN MATERIAL EMBRITTLEMENT, WHICH COULD COMPROMISE CRITICAL DESIGN FEATURES OF THE PUMP OR VOLUTE/SCROLL. EXPOSURE OF ANY PUMP OR HOUSING ASSEMBLY, OR THEIR SUBCOMPONENTS, TO MERCURY IS PROHIBITED. NOTE: Faulty pump stators must be replaced in strict adherence with OEM documentation to protect the pump’s dry running and overheat performance as demonstrated in qualification tests. 1. Verify that the fuel pump(s) retain all features designed by the OEM to provide explosion proof properties. These include but are not limited to Applicable bearing materials and clearances, housing materials Not Applicable and thickness, size of vent and cooling holes and installation of flame-arresting plates, and metal-tometal faying surfaces. 2. Verify that the fuel pump(s) retain all OEM features of the inlet housing shape, clearances, and materials, as well as the materials of all other rotating Applicable Not Applicable components and stationary interfacing surfaces (reprime impellers, liners, plates, bearings, bearing assemblies, shaft, etc.). 3. Verify that pump volute/scroll and inlet screens are Applicable maintained as designed by Boeing and/or OEM Not Applicable where applicable in order to prevent FOD ingestion.

Revision 34

04-28-10

28.01.11

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-9 JOB TITLE

DC-8 FUEL PUMP (BOOST OR TRANSFER) AND HOUSING ASSEMBLIES INSPECTION Job Card Reviewed By:

Date:

Page 2 of 2

Applicability Procedure Verify that electrical connectors are repaired or replaced in strict accordance with OEM documentation to protect against electrical failures and maintain explosion proof characteristics. 5. Verify that the pump stator lead wire routing inside the pump is installed in strict accordance with OEM documentation so that it cannot contact the rotor and charge the rotor, which could lead to an electrical arc outside of the pump’s explosion proof volume. The protector, which provides a barrier between the lead wires and rotor assembly and shaft, must be installed in strict accordance with OEM documentation. 6. Verify that the electrical bonding paths from the motor assembly to the volute/scroll and to ground are maintained per OEM, as changes to the material coatings at motor housing joints, motor-tovolute/scroll interface, and volute/scroll-to-structure interface could defeat fault current bonding paths. NOTES:

Mechanic

Inspector

4.

28.01.12

04-28-10

Applicable Not Applicable

Applicable Not Applicable

Applicable Not Applicable

Revision 34

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-10 JOB TITLE

DC-8 FUEL PUMP (BOOST OR TRANSFER) AND HOUSING ASSEMBLIES INSPECTION (WELDON FUEL PUMP) Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-3 “DC-8 Fuel Pump (Boost or Transfer) and Housing Assemblies” Applicability Mechanic Inspector Procedure NOTE: This card is applicable only to the following: (1) Weldon Fuel Pump Assembly, P/N 10045-A 1.

Verify that the installation location of the fuel pump Applicable is as defined by the original aircraft OEM design Not Applicable data. NOTE: Any repair to components that affect location of the fuel pump relative to distance from the fuel tank wall or components mounted between the pump and fuel tank wall must be made in accordance with the airframer’s design documentation (e.g., installation drawings and/or Illustrated Parts Catalog) or an FAA approved service bulletin. Any changes or deviations to the features described above must be justified by the modifier and approved by the FAA LA-ACO. NOTES:

Revision 31

11-20-08

28.01.13

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-11 JOB TITLE

DC-8 FUEL PUMP ELECTRICAL CONNECTOR WIRE/CONDUIT ASSEMBLY INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-4 “DC-8 Fuel Pump Electrical Connector Wire/Conduit Assembly” Applicability Mechanic Inspector Procedure 1. Verify that only connector P/N’s 14158-1 or 14158Applicable 2 are used in a rebuilt or repaired pump conduit Not Applicable assembly. NOTE: Repairs to the fuel pump conduit/wire assembly must be done with the assembly removed from the aircraft and in accordance with Douglas DC-8 Overhaul Manual, Chapter 28-10-2, Temporary Revision 28-4 (dated July 15, 2000) or later. Assembly procedures for the wire harness/conduit are to be used to maintain the design intent to prevent connector, wiring, and conduit damage and/or failure inside the tank. NOTE: Not all DC-8 aircraft were in compliance with CDCCL 28-4 at delivery. Boeing Service Bulletin DC8-28-83, Revision 2, dated 8/30/1975 or later is required to be completed on any aircraft not conforming to CDCCL 28-4. NOTES:

28.01.14

11-20-08

Revision 31

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-12 JOB TITLE

DC-8 OUTER WING SURFACE PREPARATION INSPECTION Job Card Reviewed By:

Date:

Page 1 of 2

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-5 “DC-8 Outer Wing Surface Preparation” Applicability Mechanic Inspector Procedure NOTE: The outer outboard portion of the DC-8 wing, with skin thickness less than 0.080 inch, is not capable of withstanding a severe lightning channel attachment without melting through when it is painted. Since all of the skin panels of the wing box are thicker than 0.064 inch, then can safely handle a potential lightning strike when the skin is bare, anodized, alodined, or has aluminum clad finish. 1. Verify that all paint has been removed from the exterior portion of the upper and lower skin panels Applicable Not Applicable of the fuel tank box in the areas identified in the following list: APPLICABLE AIRPLANE FSNs 45410 45935 45600-45606 45610-45624, 45636-45812, 4581745881 45813, 45886, 45888, 45891, 45898, 45900-45902, 45907, 45914-45915, 45938-45939, 45944-45950, 45952, 45973-45978, 45979-45981, 4599245994, 46011-46012, 46015-46018, 46029-46032, 46037, 46040, 46048, 46055-46056, 46064-46065, 46072 45570 45905, 45920-45921, 45925, 4595345956, 46023-46024

Revision 34

PAINT REMOVAL OUTBOARD OF STATION NO. Upper Panels: XRS = 589 Lower Panels: XRS = 589 Upper Panels: XRS = 589 Lower Panels: XRS = 693 Upper Panels: XRS = 624 Lower Panels: XRS = 589 Upper Panels: XRS = 624 Lower Panels: XRS = 624 Upper Panels: XRS = 624 Lower Panels: XRS = 659

Upper Panels: XRS = 624 Lower Panels: XRS = 727 Upper Panels: XRS = 659 Lower Panels: XRS = 693

04-28-10

28.01.15

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-12 JOB TITLE

DC-8 OUTER WING SURFACE PREPARATION INSPECTION Job Card Reviewed By:

Date:

Page 2 of 2

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-5 “DC-8 Outer Wing Surface Preparation” Applicability Mechanic Inspector Procedure APPLICABLE AIRPLANE FSNs 45903-45904, 45909, 45922-45924, 45926-45928, 45936, 45960-45969, 45984, 45986, 45988-45991, 4599946008, 46013, 46019-46022, 46027, 46033-46036, 46041-46044, 4606746071, 46073-46082, 46085, 4608646115, 46117-46163

PAINT REMOVAL OUTBOARD OF STATION NO. Upper Panels: XRS = 727 Lower Panels: XRS = 727

NOTE: These unpainted panels still require corrosion protection per AMM structural processes. NOTES:

28.01.16

04-28-10

Revision 34

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-13 JOB TITLE

DC-8 FUEL QUANTITY MASTER INDICATORS INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-6 “DC-8 Fuel Quantity Master Indicators” Applicability Mechanic Inspector Procedure NOTE: This card is applicable to all DC-8 airplanes with utilize the following fuel quantity gauging system (FQGS) components:

1.

(1) Master indicator P/N B118-xxx series for Goodrich/Simmonds/Liquidometer FQGS (2) Master indicator P/N 206-001-001 thru 206-001-004 for Parker FQGS Verify that any repair to a FQGS master indicator was performed in accordance with the OEM’s Applicable Not Applicable Component Maintenance Manual (CMM) listed below or an FAA approved service bulletin: PART NUMBER(S)

CMM

REVISION

206-001-001 thru 206-001-004

28-40-01

1

B118-xxx Series

28-40-007

C

NOTE: Any deviation from the OEM’s CMM must be justified by the modifier and approved by the FAA LA-ACO. NOTES:

Revision 34

04-28-10

28.01.17

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-14 JOB TITLE

DC-8 FUEL PUMP CONNECTOR INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 28-7 “DC-8 Fuel Pump Connector” Applicability Mechanic Inspector Procedure NOTE: This card is applicable to the following tank mounted electrically operated fuel pumps: (1) Pesco Fuel Pump P/N 112-303-1, 112-303-11, 1711460, 1711461 NOTE: An under torqued connector may leak fuel into the connector or may back off and open the connector, leading to short circuits of the pump power wires and potentially an ignition source within the fuel tank. 1. Every time a pump connector is installed, a second Applicable individual must witness inspection of pump Not Applicable connector/torquing. NOTES:

28.01.18

04-28-10

Revision 34

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-15 JOB TITLE

DC-8 FUEL PUMP CIRCUIT BREAKER RESET INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), ALI 20-1 “DC-8 Fuel Pump Circuit Breaker Reset” Tools Needed Standard Electrical Tools Applicability Mechanic Inspector Procedure CAUTION: ANY TRIPPED FUEL PUMP CIRCUIT BREAKER(S) IS NOT TO BE RESET UNTIL IT HAS BEEN DETERMINED IT IS SAFE TO DO SO. NOTE: This non-routine is to be accomplished only when the open circuit breaker condition occurs in service. NOTE: Tripped fuel pump circuit breakers shall be reported by the flight crew and will generate maintenance action. 1. Verify that it is safe to reset the circuit breaker(s) by following the applicable OEM troubleshooting procedures. Fault(s) that resulted in circuit breaker Applicable Not Applicable trip must be isolated and corrected prior to reset if the fault(s) occurred inside the fuel tank or adjacent to any fuel tank wall. 2. Check “pass” or “fail”. Pass – It is safe to reset the circuit breaker(s) if the troubleshooting procedures are successfully Pass completed and the problem has been corrected. Fail Fail – It is not safe to reset the circuit breaker(s) if the troubleshooting procedures are not successfully completed. NOTES:

Revision 31

11-20-08

28.01.19

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 – FUEL 28.01 SFAR 88 COMPLIANCE AND NON-ROUTINE JOB CARDS (cont’d) NOTE: This jobcard is applicable after incorporation of Boeing Service Bulletin DC8-28-090. RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.01-16 JOB TITLE

DC-8 FLOAT SWITCH CIRCUIT INSPECTION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030 (SFAR 88), CDCCL 20-10 “DC-8 Float Switch Circuit” Boeing Service Bulletin DC8-28-090 FAA AD 2010-21-03 Applicability Mechanic Procedure 1. Verify that any replacement fuse installed to protect the float switch circuits is the same type as Applicable Not Applicable prescribed in SB DC8-28-090, Revision New or later, or Boeing approved part, per SWPM 20-20-02. 2. Verify that any replacement or repair of wiring or sleeving from the fuse to tank penetration point has Applicable Not Applicable DMS 2109 or DMS 2379 Type 4 sleeving installed per SWPM 20-10-01. NOTE: For fuse installations that incorporate a drip loop near the fuse, sleeving will only extend from tank penetration point to bottom of the drip loop. NOTES:

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GENERAL--------------------------------------------------------------------------

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RECURRING NON-ROUTINE JOB CARDS Job Card 28.02-1

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 - FUEL 28.02 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK 1. General a. Non-routine Job Card DC8-28.02-1 “DC-8 Alternate and Center Auxiliary Tank Fuel Pump Control Systems Check” is applicable to all Alternate and Center Auxiliary Tank Fuel Pump Control System Switches, Control Relays, and wiring. b. This Non-routine inspects for failures in the fuel pump control system that would allow the fuel pump to continue to operate after it has been commanded OFF via the fuel pump control switch on the FE’s panel. Failures could include welded contacts in the FE’s Panel control switches, welded contacts in the Center Auxiliary Tank fuel pump control relay, or wire shorts. Failures are found when the pressure in the engine fuel feed system will not decrease after the fuel tank pump being tested is commanded OFF. c. Accomplishment of Non-routine Job Card DC8-28.02-1 is not to exceed 8,000 flight hours or 24 calendar months, whichever comes first.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.02-1 JOB TITLE

DC-8 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK Job Card Reviewed By:

Date:

Page 1 of 4

References Boeing Report MDC-02K9030, ALI 28-1 “DC-8 Alternate & Center Auxiliary Tank Fuel Pump Control Systems Check” FAA AD 2010-21-03 Mechanic Inspector Procedure Perform an ON/OFF check of the affected fuel pump(s) to verify the pump is not running when it is commanded OFF. This can be monitored from engine fuel feed pressure gauge when the pump is commanded on or off. 1. Apply electrical power to the airplane and verify that all 4 electrical buses are powered. NOTE: Suggest total fuel loaded on the airplane be at least 20,000 Lbs (9075 Kg) to be distributed in each individual fuel tank as noted in the following check steps. A lesser total amount can be used and transferred between tanks as necessary to achieve the minimum amounts noted in the following check steps. 2. Test No. 1 Alternate Tank Fuel Boost Pump: a. Verify that at least 1200 Lbs (545 Kg) of fuel is available in the subject tank. b. Verify that the Fuel Panel is in the NORMAL configuration: • All fuel pumps are OFF, • All Main Tank Fuel Selector Valves are in the MAIN position, • Center Wing Aux and Forward Aux Selector Valves are in the “OFF” position, • All the Crossfeed Valves are in the NORMAL position, • All Fill Valves are in the CLOSED position, • All Fire Handles are in the “NORMAL” position.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.02-1 JOB TITLE

DC-8 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK Job Card Reviewed By:

Date:

Page 2 of 4 Mechanic

Inspector

Procedure c. Move the Number 1 fuel tank Selector Valve to the ALT position. d. Verify that the fuel pressure supplied to the number 1 engine on the fuel pressure gauge is less than 10 PSIG. e. Select the Number 1 Alternate Tank Boost Pump switch to the BOOST position. f.

Verify that the fuel pressure supplied to the number 1 engine on the fuel pressure gauge indicates greater than 15 PSIG. NOTE: It may take up to 2.5 minutes for the fuel pump to produce this pressure. g. Select the Number 1 Alternate Tank Boost Pump switch to the OFF position. h. If the fuel pressure decays to less than 15 PSIG within 10 seconds, the test is satisfactory. If the pressure holds on the number 1 engine fuel pressure gauge for longer than 15 seconds, momentarily place the number 1 Alternate Tank Crossfeed Valve in the CROSSFEED position and open the Number 1 Alternate Tank Fill Valve by placing the switch in the FILL position to relieve the pressure to below 10 PSIG. After the pressure is relieved, place the Number 1 Alternate Tank Fill Valve in the CLOSED position and move the crossfeed lever back to the NORMAL position. Wait for at least 30 seconds and verify that the Number 1 engine fuel pressure does not climb above 15 PSIG. i. If the fuel pressure again climbs above 15 PSIG, the test is failed. Troubleshoot the cause of the uncommanded operation of the fuel pump before further filght or deactivate the defective fuel pump by removing all electrical power from the pump by opening and collaring the applicable fuel pump power circuit breakers. j. For DC-8-71 Series airplanes, repeat steps 2.a through 2.i for Alternate fuel tanks 2, 3 and 4. NOTE: Use the Fuel Pressure Gauge, Crossfeed, Tank Selector and Fill Valves, and Tank Boost Pump associated with the respective engine position.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.02-1 JOB TITLE

DC-8 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK Job Card Reviewed By:

Date:

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Inspector

Procedure k. For DC-8-62, -72, and –73 Series airplanes, repeat steps 2.a through 2.i for Alternate fuel tank 4. NOTE: Use the Fuel Pressure Gauge, Crossfeed, Tank Selector and Fill Valves, and Tank Boost Pump associated with the respective engine position. 3. Test Center Wing Auxiliary Tank Fuel Pumps if installed. a. Verify that at least 3000 Lbs (1360 Kg) of fuel is available in the subject tank (Center Wing Auxiliary Tank). b. Verify that the Fuel Panel is in the NORMAL configuration such as: • All fuel pumps are OFF, • All Fuel tank Selector Valves are in the NORMAL (MAIN) position, • All the Crossfeed Valves are in the NORMAL (CLOSED) position, • All Fill Valves are in the CLOSED position, • All Fire Handles are in the NORMAL (OPEN) position. c. Verify that the fuel pressure supplied to the number 3 engine on the fuel pressure gauge is less than 10 PSIG. d. Select the Center Wing Aux Manual Operated Gate Valve to CENTER WING AUX position. e. Select the Center Wing Aux Tank Fuel Boost Pump switch to the BOOST position. f.

Select the No.3 Crossfeed Valve to the CROSSFEED position.

g. Verify that the fuel pressure supplied to the number 3 engine on the fuel pressure gauge is greater than 15 PSIG. NOTE: It may take up to 2.5 minutes for the fuel pump to produce this pressure. h. Select the Center Wing Aux Tank Boost Pump Switch to the OFF position.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 28 - FUEL 28.02 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

DC8-28.02-1 JOB TITLE

DC-8 ALTERNATE AND CENTER AUXILIARY TANK FUEL PUMP CONTROL SYSTEMS CHECK Job Card Reviewed By:

Date:

Page 4 of 4 Mechanic

Inspector

Procedure i. If the fuel pressure to the number 3 engine decays to less than 15 PSIG within 10 seconds, the test is satisfactory. If the pressure holds on the number 3 engine fuel pressure gauge for longer than 15 seconds, momentarily place the Center Wing Auxiliary Tank fill valve to the FILL position to relieve the pressure to below 10 PSIG. After the pressure is relieved, place the Center Wing Auxiliary Tank fill valve in the CLOSED position. Wait for at least 30 seconds and verify that the Number 3 engine fuel pressure does not climb above 15 PSIG. j. If the fuel pressure again climbs above 15 PSIG, the test is failed. Troubleshoot the cause of the uncommanded operation of the fuel pump(s) in the Center Wing Auxiliary Tank before further flight or deactivate the defective fuel pump by removing all electrical power from the pump by opening and collaring the applicable fuel pump circuit breakers. k. After all test steps are done, redistribute fuel as necessary. 4. Ensure all job card sign offs are completed.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29- HYDRAULICS

TABLE OF CONTENTS Page 29.01

VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP (STC SA4201NM-D, E.O. 29-DC8-82) -------------------------------------------

29.01.1

29.02

FILTER, PUMP CASE DRAIN----------------------------------------------------

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VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP BYPASS SOLENOID REPLACEMENT-----------------------------------------

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FLIGHT CONTROL PRIORITY VALVE TEST -------------------------------

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GENERAL-----------------------------------------------------------------------------

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2.

TOOLS AND EQUIPMENT REQUIRED ----------------------------------------

29.01.1

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REMOVAL / INSTALLATION OF ENGINE DRIVEN HYDRAULIC PUMPS ---------------------------------------------------------------------------------

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FIGURES 29.01-1

PV3-160-4 ENGINE DRIVEN PUMP IPC (DC8-62/-63) ----------------------

29.01.6

29.01-2

HYDRAULIC PUMP MOUNTING CLAMP & ADPATER IPC -------------

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CASE DRAIN FILTER ASSY., P/N 1732733 IPC ------------------------------

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.01 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP 1. General a. Aircraft modified in accordance with STC SA4201NM-D or E.O. 29-DC8-82, which installs the Vickers PV3-160-4 inline engine driven hydraulic pump. The long duct aircraft case drain filter has been moved from the rear spar location to a location in the Q.E.C. plumbing. The short duct aircraft filter remains the same. b. Access to the hydraulic pumps is through the right hand side engine nacelle doors and removal of the right hand side engine bypass duct (see Chapter 71). c. Removal/installation/inspection procedures for the #2 and #3 engine hydraulic pumps are identical, except as noted. 2. Tools and Equipment Required: Grease: Electro-Moly/92 (Electrofilm, Inc.) or equivalent, is used for lubrication of hydraulic pump shaft splines. 3. Removal/Installation of Engine Driven Hydraulic Pumps a. Remove engine driven hydraulic pump. (1)

Open 2 and 3 engine hydraulic pump bypass circuit breaker located on the miscellaneous (dc bus) section of the EPC circuit breaker.

WARNING: TAG AND SAFETY CIRCUIT BREAKER(S). (2)

Place fire control handle(s) in fuel, air, and hydraulic shut-off position.

(3)

Relieve hydraulic head pressure from the system using plunger located in left hand wheel well.

(4)

Disconnect case drain hose from case drain port. CAUTION:

Revision 14

THE FIRE DETECTOR SENSING ELEMENTS WHICH ARE INSTALLED WITHIN THE ENGINE AREA ARE VERY SNESITIVE. PHYSICAL CONTACT WITH THESE ELEMENTS CAN CAUSE HIDDEN DAMAGE. A CONSTANT DIMENSION MUST BE MAINTAINED BETWEEN ELEMENTS AND THEIR SUPPORT STRUCTURE. A MINIMUM CLEARANCE OF 3/4 INCH MUST BE MAINTAINED BETWEEN ELEMENTS AND ENGINE COMPONENTS AND LINES.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.01 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP (CONT’D) (4)

Disconnect pressure line from outlet port.

(5)

Disconnect suction hose from suction port.

(6)

Disconnect electrical connector from bypass solenoid.

(7)

Remove the mounting clamp, which attaches the pump body to the mount adapter flange, while supporting the pump. Pull the pump straight back to disengage the spline and the anti-rotation pins.

(8)

Remove the pump.

(9)

Remove the hydraulic fittings from the pump ports and retain for use on the new pump. Note: In the event the pump is removed due to a malfunction of the pump and physical internal damage is suspected, the following filters and check valves should be removed and cleaned or replaced. (a) Dual filter and relief valve filter element. (b) Hydraulic reservoir return fluid filter element. (c) Engine driven case drain check valve, located in the pylon. (d) Engine driven pump pressure line check valve, located in the pylon. (e) Case drain filter element.

b. Install the Engine Driven Hydraulic Pump (1)

Make certain that engine 2 and 3 hydraulic pump bypass circuit breaker located on the miscellaneous (dc bas) section of the EPC circuit breaker panel is open.

(2)

Install the fittings in the pump pressure, suction, and case drain ports, replacing 0rings as required.

(3)

Drive Spline Treatment (a) PV3-160-4: Lubricate pump shaft spline with Eletro-Moly/92 or equivalent grease.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.01 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP (cont’d) (b) PV3-160-4B: This spline has a plastic wear feature. (4)

Check the pump mount adapter flange for condition and security. Correct any discrepancy prior to installing the pump.

(5)

Fill the pump case with clean hydraulic fluid through the case drain port. CAUTION:

(6)

BEFORE OPERATING THE ENGINE DRIVEN HYDRAULIC PUMP, THE PUMP MUST BE FILLED WITH CLEAN HYDRAULIC FLUID THROUGH THE CASE DRAIN PORT TO PREVENT THE PUMP FROM OPERATING WITHOUT LUBRICATION DURING THE SELF-PRIMING PERIOD. Install the pump on the engine accessory pad and secure. CAUTION:

THE FIRE DETECTION SENSING ELEMENTS WHICH ARE INSTALLED WITHIN THE ENGINE AREA ARE VERY SENSITIVE. PHYSICAL CONTACT WITH THESE ELEMENTS CAN CAUSE HIDDEN DAMAGE. A CONSTANT DIMENSION MUST BE MAINTAINED BETWEEN ELEMENTS AND THEIR SUPPORT STRUCTURE. A MINIMUM CLEARANCE IF 3/4 INCH MUST BE MAINTAINED BETWEEN ELEMENTS AND ENGINE COMPONENTS AND LINES.

(a) Insert the pump spline into its gearbox drive. Align and engage the pump antirotational pins. Confirm that the pump seats fully on its adapter flange without any force being applied. (b) Install the mounting clamp around the mated pump and mount adapter flanges. Secure the “T” bolt, taking up all slack in the clamp. (c) Inspect the pump-to-adapter flange mating. Make certain that the mounting clamp has equally engaged all parts of the pump/adapter flange. (d) Torque the “T” bolt to 45 to 55 inch/pounds and lockwire clamp. (7)

Connect the pressure line hose to the pump outlet port fitting.

(8)

Connect the suction hose to the pump suction port fitting.

(9)

Revision 5

Connect the case drain hose to the pump case drain port fitting.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.01 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP (cont’d) (10) Connect the electrical connector to the pump bypass solenoid and lockwire. (11) Return the fire control handle (s) to the normal position. (12) Bleed the entrapped air from the pump by cracking the case drain line and allowing fluid flow from the reservoir to establish itself. Re-secure the case drain hose at the pump. (13) Reset the generator which was tripped when the fire control handle was pulled. (14) Close engine 2 and 3 hydraulic pump bypass circuit breaker located on the miscellaneous (dc bus) section of the EPC circuit breaker panel. (15) Check the operation of the hydraulic pump in accordance with paragraph 4 below. (16) Service hydraulic head pressure to 35 psi IAW DACO MM Chapter 29. c. Inspection/Check Engine Driven Hydraulic Pump (1) Start engines No. 2 and 3. (Ref. DC-8 ERH.) (2) Place the hydraulic system selector control lever in the general system (normal) position. (3) Check the main hydraulic pressure indicator for 2800 to 3000 psi system pressure. (4) Operate the wing flaps through a complete cycle of operation. WARNING:

BEFORE OPERATING THE FLAPS, MAKE CERTAIN THAT THE FLAP AREA IS CLEAR OF PERSONNEL AND OBSTRUCTIONS.

(5)

Place both engine driven pump control switches in the bypass position. Hydraulic pressure should drop to approximately zero psi.

(6)

Place one control switch in the on position. Hydraulic pressure should build up to 2800 to 3000 psi.

(7)

Return that switch to the bypass position.

(8)

Place the other switch in the on position. Hydraulic pressure should build up to 2800 to 3000 psi.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.01 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP (cont’d) (9)

Place both switches in the on position for normal operation, and observe that the hydraulic pressure returns to the normal range.

(10) Confirm that the generators are operational. (11) Shut down the engine (s). (12) Perform a leak check of the pump and its hoses and fittings. (13) Make a final inspection of the pump, hoses, and electrical connector for security.

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

9.

10.

NOMENCLATURE CASE DRAIN HOSE ASSY. PRESSURE HOSE ASSY. SUCTION HOSE ASSY. CLAMP (ATTACH) HYDRAULI C PUMP CONNECTOR (Bypass Solenoid) BYPASS SOLENOID SUCTION FITTING a. ELBOW b. NUT c. O-RING d. ADAPTER PRESSURE FITTING a. ELBOW b. NUT c. O-RING CASE DRAIN FITTING a. ELBOW b. NUT c. O-RING

PART NUMBER(S) 170F005-6CR-0096 or AE2460201G0100 170F003-10CD-0130 or AE2460101J0134 AE6630M0091-037 or AE6647M0091 374105 PV3-160-4 or PV3-160-4B MS3106E-10SL-4S 411200 AN833-20D AN924-20D NAS1612-20 AN894D20-16 AN833-10 AN924-10 NAS1612-10 AN833-6 AN924-6 NAS1612-6

PV3-160-4 ENGINE DRIVEN PUMP IPC FOR THE DC-8-62 AN DC-8-63 ONLY Figure 29.01-1

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ITEM 1 2 3 4

NOMENCLATURE Adapter and Pin Assembly Clamp, Mounting Nut Bolt, T

PART NUMBER 387999 374105 138116 393101

HYDRAULIC PUMP MOUNTING CLAMP AND ADAPTER IPC Figure 29.01-2

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NOTE: When P/N 7501227 is used the following steps must be performed: A. Ensure Filter Element 7580341 has been installed. B. Add a Union(AN815-6) with a Packing(NAS1612-6) to the head of the Filter Assembly. The head is the larger diameter end of the

ITEM 1 2 3 4 5

NOMENCLATURE Filter Assy Element Spring Packing Packing

PART NUMBER 1732733 or 7501227 7580341 A2808J NAS 1612-20 109661-1-6

CASE DRAIN FILTER ASSEMBLY Figure 29.01-3

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GENERAL-----------------------------------------------------------------------------

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REMOVAL SERVICE / INSTALLATION---------------------------------------

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.02 FILTER, PUMP CASE DRAIN - REMOVAL/SERVICE/INSTALLATION 1. General a. This procedure for DC8-61/-62 and -63 aircraft only. b. The case drain filter assembly is located on the upper end of the Q.E.C. case drain hard line in place of the union which was originally there. On the DC-8-61 series, the filter is located on the upper case drain line just prior to the pylon in the flap well. c. The disposable paper element is to be replaced whenever its related engine driven pump is changed. Contamination found in this element during scheduled replacement shall be used to determine the serviceability of its respective pump. d. Inspection of this paper filter element will be accomplished by freeing the pleated portion of the filter from its end caps and unfolding it in order to reveal the entire contents of the element. e. The spring installed on top of the filter element prevents bypassing, and subsequent downstream contamination. It also prevents blockage of the case drain circuit if the filter element is installed backwards. 2. Removal Service/Installation a. Remove the Filter Assembly. (1)

Relieve hydraulic system pressure.

(2)

Safety tag cockpit hydraulic and engine controls.

(3)

Relieve the main hydraulic reservoir head pressure.

(4)

Disconnect the hoses from both ends of the filter assembly. Cap/plug Q.E.C. plumbing.

b. Service the Filter Assembly (1)

Unscrew the cap from the filter assembly.

(2)

Carefully remove the filter element, and look for contamination in the case drain filter assembly housing.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 27.02 FILTER, PUMP CASE DRAIN - REMOVAL/SERVICE/INSTALLATION(cont’d) (3) (4)

Clean the filter assembly housing and cap with Sroddard Solvent or equivalent, and dry. Install a new packing (O-ring) on the housing parting surface.

(5)

Install a new filter element and reassemble the filter assembly. (a) Check that the spring is installed first, in the male half of the housing. The large loop of the spring may be spread outward slightly to prevent the spring from falling out of the housing. (b) Insert the element with its open end away from the spring. (c) Screw on the female half of the housing and tighten it. (d) Re-safety the filter housing to the cap.

c. Installation (1)

Reconnect the filter assembly into its previous position in the Q.E.C. case drain plumbing. CAUTION:

THE FILTER MUST BE INSTALLED SO THAT THE FLOW WILL BE FROM THE OUTSIDE OF THE ELEMENT TO THE INSIDE. THIS IS DONE BY INSTALLING THE ‘OUT’ OR CAP END SO THAT IT IS EITHER UP OR AFT.

(2)

Remove the cockpit safety tags.

(3)

Check and service the main hydraulic reservoir fluid level and head pressure as required.

(4)

Pressure check the filter installation for no leakage by motoring or running the engine.

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29.03 – VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP – BYPASS SOLENOID REPLACEMENT TABLE OF CONTENTS Paragraph 1. 2. 3.

GENERAL-----------------------------------------------------------------------------

Page 29.03.1

REMOVAL / INSTALLATION OF ENGINE HYDRAULIC PUMP BYPASS SOLENOID----------------------------------------------------------------

29.03.1

INSPECTION / CHECK OF ENGINE DRIVEN HYDRAULIC PUMP BYPASS SOLENOID----------------------------------------------------------------

29.03.2

FIGURES 29.03-1

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BYPASS SOLENOID IPC ----------------------------------------------------------

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29-HYDRAULICS 29.03 VICKERS PV3-160-4 ENGINE DRIVEN HYDRAULIC PUMP - BYPASS SOLENOID REPLACEMENT 1. General a. This procedure covers the removal and replacement of the engine driven hydraulic pump bypass solenoid valve without removing the pump. b. Extreme care must be taken during this action to prevent damage to the packings and backup rings. c. Access to the bypass solenoid is through the engine nacelle doors. (See chapter 71.) d. Removal/installation/inspection/check procedures for the #2 and #3 engine driven hydraulic pump solenoid valves are identical. 2. Removal /Installation of Engine Hydraulic Pump Bypass Solenoid (Refer Figure 29-5) a. Remove engine driven hydraulic pump bypass solenoid. (1)

Open 2 and 3 engine hydraulic pump bypass circuit breaker located on the miscellaneous (dc bus) section of the EPC circuit breaker panel. WARNING:

TAG AND SAFETY CIRCUIT BREAKER (S) AND ENGINE CONTROLS

(2)

Place fire control handle (s) in fuel, air, hydraulic shutoff position.

(3)

Remove the bypass solenoid connector from the bypass valve.

(4)

Remove and retain the three screws that attach the solenoid to the pump.

(5)

Carefully withdraw the solenoid and value assembly from the pump housing port. CAUTION:

THE SOLENOID AND VALVE ARE A MATCHED SET AND CANNOT BE INTERCHANGED WITH PARTS OF A DIFFERENT ASSEMBLY.

b. Install engine driven hydraulic pump bypass solenoid assembly. (1)

Revision 4

Inspect the hydraulic pump bypass solenoid port interior for presence of foreign objects or o-ring debris.

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Install packings and backup rings on the solenoid valve (ref. Figure 1).

(3)

Lubricate the packings and backup rings with clean hydraulic fluid.

(4)

Carefully withdraw the solenoid and valve assembly from the pump housing port. CAUTION:

THE PACKING OF THE SOLENOID VALVE ASSEMBLY ARE EXTREMELY SUSCEPTIBLE TO DAMAGE DURING ASSEMBLY.

(5)

Install the three retaining screws and safety as required.

(6)

Install bypass solenoid connector and safety.

(7)

Place fire control handle in the cockpit to the normal position.

(8)

Reset the bypass control circuit breaker for the 2 and 3 engine driven pumps previously opened.

(9)

Remove safety tags from controls and circuit breaker (s).

(10) Reset generator breaker on Flight Engineer’s panel. 3. Inspection/Check of Engine Driven Hydraulic Pump Bypass Solenoid a. Motor the effected engine IAW JRH and cycle the effected pump switch on and off to bleed entrapped air from the solenoid valve. Let the pump build up pressure to 3000 PSI and turn the switch OFF. Let the pressure decrease to approximately 300 PSI and turn the pump ON. WARNING:

DO NOT EXCEED STARTER DUTY LIMITS DURING MOTORING.(Reference Jet Run Handbook)

b. Check the bypass solenoid for external leakage and security. c. Check the bypass solenoid connector for security. d. Close and secure engine nacelle doors.

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ITEM 1 2 3 4

NOMENCLATURE Bypass Solenoid Valve Retaining Screw Packing Ring-Backup

PART NUMBER 411200 MS24678-10 271133 MS28774-013

BYPASS SOLENOID IPC Figure 29.03-1

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29.04 – FLIGHT CONTROL PRIORITY VALVE TEST TABLE OF CONTENTS Paragraph 1.

GENERAL-----------------------------------------------------------------------------

Page 29.04.1

2.

TEST -----------------------------------------------------------------------------------

29.04.2

FIGURES 29.04-1

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HYDRAULIC SYSTEM PRIORITY VALVE INSTALLATION-------------

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29- HYDRAULICS 29.04 FLIGHT CONTROL PRIORITY VALVE TEST 1. General a. The priority valve is installed in the hydraulic system between the left and right hand manifolds and connected so as to provide hydraulic power to only the ailerons, rudder and stabilizer in case the system pressure drops to a certain level. The purpose is to prevent flight control reversion from power to manual configuration in case other systems are operated in flight. b. The valve operation may be tested by using the auxiliary pump or by using a hydraulic bench. If the bench is used, the flow rate should be limited to approximately 5 gpm. 2. Test – Clear the airplane for hydraulic system pressure and for operation of the wing flaps, ailerons, and/or the rudder. Operation of the wing flaps and ailerons should be sufficient. a. Pressurize the hydraulic system. b. Turn on power to the ailerons and/or rudder. c. Operate the wing flaps, and while the flaps are moving, operate the selected control (ailerons and/or rudder) at a rate to cause the test stand pressure or auxiliary pump pressure to drop to below 1700 psi, as read on the main system accumulator pressure gauge in the left wheel well. d. Note that as the pressure drops below 1700 psi the flaps stop moving. e. The ailerons or rudder should continue to move on power and not revert to manual control. NOTE:

If controls are operated too fast, ailerons or rudder will revert to manual operation when system pressure drops below 1100 (+100, -50) psi due to the aileron or rudder low pressure shutoff valves.

f. Decrease or stop the operating rate of the controls to allow the test stand or auxiliary pump pressure to build up. g. At a maximum of 2250 psi, the priority valve should operate and allow the flaps to continue moving. Read pressure on the accumulator gauge in the left wheel well.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 29- HYDRAULICS 29.04 FLIGHT CONTROL PRIORITY VALVE TEST (cont’d)

HYDRAULIC SYSTEM PRIORITY VALVE INSTALLATION Figure 29-04-1 29.04.2

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS

TABLE OF CONTENTS 31.01

DIGITAL FLIGHT DATA RECORDS - DATA RETENTION POLICY ----

Page 31.01.1

31.02

ARINC 542A DFDR EXPANDED PARAMETERS ----------------------------

31.02.1

31.03

UNDERWATER ACOUSTIC BEACON TESTERS----------------------------

31.03.1

31.04

UNDERWATER ACOUSTIC BEACONS ---------------------------------------

31.04.1

31.05

EXHAUST GAS TEMPERATURE INDICATOR UPGRADE----------------

31.05.1

31.06

ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT)--------------

31.06.1

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.01 DIGITAL FLIGHT DATA RECORDS - DATA RETENTION POLICY 1. As digital flight data recorder units are changed either on a routine or non-routine basis, the 25 hour data retention requirement mandated by the F.A.R.’s can be interrupted. In order to ensure that there is continuity of recorded flight data for any individual aircraft which experiences the replacement of a DFDR unit, that unit will be retained by Stores in its unserviceable condition for the period of time required for that aircraft to accumulate 25 flight hours. Only then can the unserviceable DFDR be released to the vendor for repair.

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31.02 – ARINC 542A DFDR EXPANDED PARAMETERS TABLE OF CONTENTS Paragraph 1.

Page DIGITAL FLIGHT DATA RECORDER------------------------------------------ 31.02.01

2.

GENERAL----------------------------------------------------------------------------- 31.02.02

3.

EQUIPMENT INSTALLED -------------------------------------------------------- 31.02.02

4.

EQUIPMENT / AIRCRAFT SYSTEM INTERFACE--------------------------- 31.02.03

5.

ELECTRICAL ------------------------------------------------------------------------ 31.02.03

6.

EQUIPMENT LOCATION --------------------------------------------------------- 31.02.04

7.

PART LIST ---------------------------------------------------------------------------- 31.02.10

8.

DFDR CALIBRATION AND CORRELATION TEST ------------------------- 31.02.12

9.

SENSOR ALIGNMENT------------------------------------------------------------- 31.02.12

FIGURES 31.02-1

WIRE RUNS AND SENSOR LOCATIONS ------------------------------------- 31.02.04

31-02-2

DC-8 RUDDER SURFACE --------------------------------------------------------- 31.02.05

31-02-3

DC-8 AILERON ---------------------------------------------------------------------- 31.02.06

31.02-4

DC-8 CONTROL WHEEL ---------------------------------------------------------- 31.02.07

31.02-5

DC-8 CONTROL COLUMN ------------------------------------------------------- 31.02.08

31.02-6

DC-8 RUDDER PEDAL------------------------------------------------------------- 31.02.09

TABLES 31.02-1

Revision 21

PARTS LIST --------------------------------------------------------------------------

05-20-04

31.02.10

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JOB CARDS 31.02-1

F800 DRDR GROUND TEST PROCEDURE – 17 PARAMETERS ---------

31.02.13

31.02-2

F1000 DRDR GROUND TEST PROCEDURE – 17 PARAMETERS --------

31.02.21

31.02-3

SUNSTRAND 980-4120-RXUS – DFDR GROUND TEST PROCEDURE (71CX)----------------------------------------------------------------------------------

31.02.31

ALLIED-SIGNAL 980-4100 GQUS/RQUS – DFDR GROUND TEST PROCEDURE (-73 SERIES ONLY)-----------------------------------------------

31.02.43

31.02-4

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS 1. Digital Flight Data Recorder a. Description. The Digital Flight Data Recorder is an expandable twenty-three parameter unit which is currently operated by ATI with 17 parameters. It is essentially a “plug-in” digital flight data recorder system contained in one unit. The data is stored digitally on an endless loop of magnetic tape, very similar to the design of the CVR. All data transducers are in the unit itself. When installed in an aircraft, the digital flight data recorder records aircraft operating information consecutively on six tracks for a maximum of 25 hours. All data is read out 3 seconds after recording and is fed into Built-In Test Equipment (BITE), which continuously monitors such things as the incoming data and power, internal operation of the DFDR, faulty track changes, and other critical DFDR operating parameters. The cockpit warning light indicates the operating status of the DFDR as determined by the BITE. This test circuitry also checks to ensure that the DFDR is installed in the proper aircraft configuration. Because of the endless loop design, the DFDR never stops, reverses, or changes speeds from the time it is turned on until it is turned off, continuously recording all data without interruption. The 25 hour tape loop meets the FAR requirement for flight data retention. b. Removal and Installation. Because of the ARINC 542 interchangeability configuration of this unit, it is externally identical to the foil medium recorder it replaced. Therefore, all removal and replacement will be accomplished in accordance with the Douglas DC-860 Series Maintenance Manual. c. Operational Test. After installation, the BITE circuitry will be used to operationally test this unit. With power on the aircraft, turn the DFDR switch to the “OFF” position. The fail warning light should be on. Move the DFDR switch to the “ON” position. Observe the fail warning light for a period of one minute, it should remain extinguished. This means that the DFDR’s internal BITE test has determined that the unit is fully operational. d. System Operational Test. On special occasions such as heavy maintenance or problem with a chronic system, it may be necessary to perform a test of the entire DFDR system. This can be accomplished through the use of the Avionica RSU, P/N 802-0108, with applicable cable.

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31.02.01

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d) 2. General a. This section describes the modifications made to acquire 17 or 24 parameters from the Digital Flight Recorder. b. Prior to any flight control maintenance, i.e. “rigging” or movement of Flight Controls beyond its normal operating condition, first retract the applicable sensor cable ends from the appropriate flight control attachment ends.

3. Equipment Installed – The following information regarding the equipment installed is presented as part of the installation description: a. DFDR -The DFDR is a 25 hour (typically 30 hours) solid state unit that utilizes a digital method of recording and storing data. It receives its data in an ARINC 542A format. The DFDR acquires inputs from sensors and signals within the aircraft. The DFDR contains an internal self-test circuitry (BITE), which invalidates the need of additional pre-flight test procedure 31.02. b. Part Numbers for Allied Signal or Loral-Fairchild DFDR (DC-8-62/-71/-72): (1) PATRIOT P/N: 3001-01-101-2 (3001-01-111 799AL only), DUAL AXIS ACCELEROMETER LONGITUDINAL, VERTICAL, AND LATERAL SENSORS (2) FSA 342418A CONTROL WHEEL SENSOR (3) FSA 342419A RUDDER PEDAL POSITION SENSOR (4) FSA 342446A LATERAL CONTROL SURFACE (AILERON) SENSOR (5) FSA 342436A CONTROL COLUMN POSITION SENSOR (6) FSA 342444A RUDDER SURFACE POSITION SENSOR c. Part Numbers for ASM Installed ARINC 542A DFDR (DC-8-73) (1) PATRIOT 3001-01-111-2 DUAL AXIS ACCELEROMETER LONGITUDINAL, VERTICAL, AND LATERAL SENSORS (2) ASM T0409-001, Designator T2-9205 CONTROL COLUMN POSITION SENSOR

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d) (3) ASM T0409-001, Designator T2-9203 RUDDER PEDAL POSITION SENSOR (4) ASM T0406-001, Designator T2-9202 LATERAL CONTROL INPUT SENSOR (5) ASM T0406-001, Designator T2-9201 AILERON POSITION SENSOR (6) ASM T0406-001, Designator T2-9204 RUDDER POSITION SENSOR 4. Equipment/Aircraft System Interface a. The DFDR increases existing 11 parameters to 17 parameters acquiring data from sensors installed in 3.b. through 3.g. and existing aircraft signal for the elevator surface. b. The DFDR receives its parameters with the following interfaces with the aircraft. (1) Time intervals are recorded from an internal clock in the DFDR. (2) Altitude and airspeed are taken from the air data computer. Air temperature is recorded from existing temperature sensor. (3) Time of each radio transmission is obtained in the same manner as original design. (4) Heading information is obtained in the same manner as the original design. Pitch and roll attitude signals are derived from the first officer’s attitude director indicator. (5) Vertical, longitudinal and, if applicable, lateral acceleration are provided by the triple axis accelerometer identified in 3.b. (6) The five position sensors, rudder surface, aileron surface, rudder pedal, control column, and control wheel identified in 3.c. through 3.g. provide the appropriate signals. Elevator position is obtained from existing signal. (7) Engine thrust is obtained from existing signals within the aircraft. (8) Autopilot engagement status is an existing discrete signal made available.

5. Electrical – Wiring diagrams vary depending on manufacturer and model of DFDR. Refer to individual Aircraft Wiring Diagram Manuals for particular electrical pin-out information.

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WIRE RUNS AND SENSOR LOCATIONS Figure 31.02-1 31.02.04

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DC-8 RUDDER SURFACE Figure 31.02-2

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DC-8 AILERON Figure 31.02-3

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DC-8 CONTROL WHEEL Figure 31.02-4 Revision 20

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d)

DC-8 CONTROL COLUMN Figure 31.02-5 31.02.08

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DC-8 RUDDER PEDAL Figure 31.02-6 Revision 20

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31.02.09

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d) 7. Parts List QTY.

DESCRIPTION

PART NUMBERS

1 ea

Lat. Cont. Surface (Aileron) Sensor

FSA 342446A

1 ea

Rudder Pedal Sensor

FSA 342419A

1 ea

Cont. Column Sensor

FSA 342436A

1 ea

Cont. Wheel Sensor

FSA 342418A

1 ea

Rudder Surface Sensor

FSA 342444A

10 ea

Screw (For Potentiometer)

NAS 1096-2-7

10 ea

Nut (For Potentiometer)

NAS 679-A08

10 ea

Washer (For Potentiometer)

NAS 960 8L

3 ea

Cable Bracket

FSA342437

1 ea

Rudder Pedal Bracket

FSA342420

760 ft

Shielded Triple Conductor

M27500-22-TG3T14

300 ft

Shielded Double Conductor

M27500-22-TG2T14

30 ft

Single Wire 22 AWG Conductor

M22759/16-22

1 ea

Clamp (Rudder Pedal)

MS21919WDG25

1 ea

Bracket (Control Column)

FSA342425

1 ea

Torque tube (For Control Column & Control Wheel)

FSA 342421

1 ea

Aileron Attachment

FSA342447

1 ea

Rudder Surface Attachment

FSA342448

1 ea

Aileron Surface Bracket

FSA342421

5 ea

Spacer

FSA3424123

PARTS LIST Table 31.02-1

31.02.10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d) a. The following list reflects the alternate part numbers to the existing part numbers. Existing P/N FSA342418 FSA342419 FSA342436 FSA342444 FSA342445 FSA342446 FSA342452 FSA342454 FSA342455 FSA342458

Alternate P/N FSA342418A FSA342419A FSA342436A FSA342444A FSA342445A FSA342446A FSA342452A FSA342454A FSA342455A FSA342458A

NOTE: It should be affirmed that the new alternate part number does not disqualify or invalidate the in-service or to-be-in-service sensors, only to qualify the new part number as an improved direct replacement of the existing part numbers.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS 31.02 ARINC 542A DFDR EXPANDED PARAMETERS (cont’d) 8. DFDR Calibration and Correlation Test a. Equipment and Materials: As listed on Job Card 31.02-X. b. ARINC 542A DFDR on DC-8, 17 Parameters - Targeted values were derived from ideal conditions, tolerance tighter than airframe manufacturers tolerance for “rigging.” As such, readings may not be exact. However, general trend should match. c. Parameters Checkouts (1) To complete a full operational checkout, comply with Recurring Non-routine Job Card 31.02-1 (F800), 31.02-2 (F1000), 31.02-3 (RXUS) or 31.02-4 (GQUS). (2) However when replacing a DFDR itself, only a download and verification of values in each of the 18 Parameter Words (in OCTAL) is required; not a full operational checkout. Use only the applicable portions of the applicable Job Card. (3) When replacing only a sensory unit, i.e. a stab position transmitter, only a check of that parameter is required. Use only the applicable portions of Job Card 31.02-1 (F800), 31.02-2 (F1000), 31.02-3 (RXUS) or 31.02-4 (GQUS). (4) Attach the completed Job Card to the aircraft log page and forward to LIT Q/C Records. 9. Sensor Alignment – If a sensor output does not fall within the expected octal value during test, adjust per the following procedures. If a sensor does not adjust to within expected octal range, isolate the cause. a. To adjust linear slide potentiometers: (1) Secure the item being sensed in the neutral position. (2) Loosen the clamp securing the potentiometer. (3) Slide the potentiometer in the clamp until the desired readout is displayed on the portable tester. (4) Secure the clamp around the potentiometer. (5) Verify that the readout displayed on the portable tester is within tolerance. b. To adjust synchros: (1) Secure the item being sensed in the neutral position. (2) Loosen one of the screws that secure the synchro in the clamp. (3) Rotate the synchro in the clamp until the desired readout is displayed on the portable tester. (4) Secure the potentiometer in the clamp. (5) Verify that the readout displayed on the portable tester is within tolerance. 31.02.12

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-1 JOB TITLE

F800 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

Date:

Page 1 of 8 Mechanic

Inspector

A. GENERAL This ground test verifies F800 DFDR inputs and the recording capability of the DFDR unit. Tools and Material Avionica, Inc. RSU DFDR Tester/Transcriber P/N 802-0108, or Equivalent Digital Protractor Tilt Table and DG/VG Extender Harnesses (Optional) Pitot-static Test Equipment Signal Generator HP204D or equivalent Work-stand capable of reaching the aft side of the horizontal stabilizer The DFDR monitors and records 17 parameters (Checkout Location): 1. TIME (D) 10. HEADING/COMPASS (I) 2. VERT. ACCEL (M) 11. ALTITUDE (G) 3. LONG. ACCEL. (M) 12. AUTOPILOT (F) 4. ELEVATOR SURF. (Q) 13. CONTROL WHEEL POS (N) 5. MIC. KEY (E) 14. RUDDER PEDAL POS. (R) 6. AIRSPEED (H) 15. AILERON SURFACE (O) 7. PITCH ATTITUDE (K) 16. CONTROL COLUMN POS. 8. ROLL ATTITUDE (J) (P) 9. ENGINE ACCELERATION (L) 17. RUDDER SURFACE (S) B. TEST SETUP 1. Record the following: a. Aircraft Serial Number/Tail Number: __________________________ b. DFDR Part Number: __________________________ c. DFDR Serial Number: __________________________ d. Biaxial Accelerometer Part Number (Left Wheel Well): ____________ e. Biaxial Accelerometer Serial Number: __________________________ 2. Check that the following aircraft systems are powered: DFDR power, 115 VAC Left Emergency Buss, 28 VAC Instrument Buss 2 & 3, 28 VDC Left Emergency Buss, Pilots & Co-pilots Horizon & Heading, VHF Communication Equipment 3. Connect front connector J1 on the DFDR to the DFDR test equipment. WARNING: WHEN POWER TO THE DFDR IS ON, THE DFDR WILL RECORD OVER PREVIOUSLY RECORDED DATA (25 HOURS EARLIER). TO MINIMIZE DESTRUCTION OF PREVIOUSLY RECORDED DATA, ONLY HAVE DFDR ON DURING TESTING. 4. Turn “ON” DFDR Switch and verify that the FLIGHT RECORDER OFF light extinguishes.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-1 JOB TITLE

F800 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

Date:

Page 2 of 8 Mechanic

Inspector

5.

Verify FDR ON/OFF switch panel lighting works from “BRIGHT” to “OFF” using the FE panel light control. C. SYNC WORD 1. Select Word 1. 2. Verify the following SYNC Octal reading by Frame: FRAME # WORD #1 1 1107 2 2670 3 5107 4 6670 D. TIME 1. Select Word 17. 2. Verify that Word 17 counts upward. E. MICROPHONE KEY 1. Select Word 16. 2. Set power to Captain’s and First Officer’s VHF and HF Radios. NOTE: Select unused VHF and HF frequencies for keying tests. 3. Select and key Captain’s Audio Panel on VHF 1 and verify that Bit 1 toggles to “0” or the OCTAL number changes down by 1. 4. Repeat step 3 for the Captain’s Audio Panel with HF 1 selected. 5. Verify that Bit 3 (counting from the right of the 12 bit word) is “1”. Key F.O.’s Audio Panel on VHF 2 and verify that Bit 3 toggles to “0” or the OCTAL number changes down by 4. 6. Key F.O.’s Audio Panel on HF 2 and verify that Bit 3 toggles to “0”. F. AUTOPILOT ENGAGE STATUS 1. Select Word 16. 2. Verify that Bit 2 (second from right of 12 bit Word) is “0”. 3. Engage Autopilot and verify that Bit 2 toggles to “1” (or the OCTAL number changes up by 2). 4. AUTOPILOT OCTAL POSITION READING ENGAGED DISENGAGED G. ALTITUDE 1. Select Word 13 for Course and Word 29 for Fine. 2. Connect Pitot-static test equipment to the Pitot tube and Static port of the Pitot-static system that is connected to CADC-1 and to the Captain’s Pitotstatic system for reference. Reference the Pitot-static section of the DC-8 Maintenance Manual for test procedure standards.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-1 JOB TITLE

F800 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

Date:

Page 3 of 8 Mechanic

3.

4.

Inspector

Record Pitot and static system source used for CADC-1 air data. Static: ____________________________________________ Pitot: ____________________________________________ Using the Pitot-static test box, step through the following altitudes, verifying that the altitude displayed on the Captain’s Altimeter is the expected value: ELECTRIC ALTITUDE Word 13 Course Altitude Frames 1 thru 4 ACTUAL ALTITUDE CAPTAIN’S EXPECTED OCTAL ALTITUDE OCTAL ±100 Feet VALUE 4000 0235-0250 8000 0430-0445 22000 1165-1177 Word 29 Fine Altitude Frames 1 thru 4 ALTITUDE ±100 Feet

CAPTAIN’S ALTITUDE

EXPECTED OCTAL

ACTUAL OCTAL VALUE

4000 6244-6507 8000 4553-4757 22000 3020-3244 H. AIRSPEED 1. Select Word 19. 2. Verify that Pitot-static test equipment is still connected to CADC-1 and Captain’s pitot-static system. 3. Using the Pitot-static test box, step through the following airspeeds. Ensure that the airspeed displayed on the Captain's Airspeed Indicator is the expected value as selected with the Pitot-Static Test Set. ELECTRIC AIRSPEED Word 19 Frames 1 thru 4 ACTUAL AIRSPEED ACTUAL EXPECTED OCTAL AIRSPEED OCTAL ±5 Knots VALUE 100 7110-0075 200 1250-1361 300 2721-3020 350 3417-3546 I.

HEADING/COMPASS 1. Select Word 3.

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-1 JOB TITLE

F800 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

Date:

Page 4 of 8 Mechanic

2.

J.

Inspector

Apply electrical power to the directional gyro compass system. Manually position the First Officer’s RMI and verify reading on CAPT MHR Indicator to the following: HEADING Word 3 Frames 1 thru 4 ACTUAL HEADING ACTUAL EXPECTED OCTAL HEADING OCTAL ±1 ° VALUE 0 7755-0022 90 1755-2022 180 3755-4064 270 5755-6064

ROLL ATTITUDE NOTE: If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #2. 2. Remove Vertical Gyro #2 and set on tilt table or employ extender harness to keep Vertical Gyro on mounting bracket. Reset power to Vertical Gyro #2. 3. Using a digital protractor, tilt Vertical Gyro #2 to the following angles. Verify that roll is shown on the F.O.’s Horizon Indicator. 4. Select Word 20. ROLL ATTITUDE Word 20 Frames 1 thru 4 ACTUAL ROLL ACTUAL EXPECTED OCTAL ROLL OCTAL ±1 ° VALUE 7745-0022 0° 0545-0600 30° right 1157-1222 60° right 7167-7224 30° left 6535-6600 60° left

K. PITCH ATTITUDE NOTE: If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Ensure that Vertical Gyro #2 is still sitting on tilt table or employ extender harness to keep the Vertical Gyro on the mounting bracket. 2. Using a digital protractor, tilt Vertical Gyro #2 to the angles in the following table. Verify that pitch is indicated on the F.O.’s Horizon Indicator. 3. Select Word 4.

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PITCH ATTITUDE Word 4 Frames 1 thru 4 PITCH ±1 °

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

Neutral 7745-0020 0545-0600 30° Up 1157-1222 60° Up 7167-7222 30° Down 6535-6600 60° Down L. ENGINE ACCELERATION N1 (-70 SERIES ONLY) 1. Using the signal generator, inject the signal at N1 Speed Probe (Tach Generator) pins 1 and 2 (or by running engines). Record N1 for the following Frame and Word. 2. Monitor the N1 indicator for the engine under test for N1 settings. N1 Frame 1 N1% Target

ENGINE #1 FRAME 1 WORD 11 OCTAL

ENGINE #2 FRAME 1 WORD 27 OCTAL

N1 Frame 2 ENGINE #3 FRAME 2 WORD 11 OCTAL

ENGINE #4 FRAME 2 WORD 27 OCTAL

30 60 90

N1% Target 30 60 90

EPR (60 SERIES A/C) 1. Using approved EPR test equipment per the maintenance manual or by running engines, record EPR for the following Frame and Word.

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Monitor the EPR indicator for the engine under test for EPR settings. EPR Frame 1 ENGINE #1 ENGINE #2 EPR FRAME 1 FRAME 1 WORD 11 WORD 27 Target OCTAL OCTAL 1.0 1.2 1.4

EPR Target

EPR Frame 2 ENGINE #3 FRAME 2 WORD 11 OCTAL

ENGINE #4 FRAME 2 WORD 27 OCTAL

1.0 1.2 1.4 M. VERTICAL & LONGITUDINAL ACCELERATION 1. With the biaxial accelerometer (left-hand wheel well) in the normal position, verify the following. Note expected values assume that the aircraft is approximately level. ACCELERATION ACCEL. Vert. Long.

WORD

EXPECTED OCTAL

2,6,10, 14,18,22,26,30 7,15,23,31

3022-3715 3640-4277

ACTUAL OCTAL VALUE

2.

Verify values change when the accelerometer is tilted vertically and longitudinally respectively. N. CONTROL WHEEL POSITION 1. Provide main system hydraulic power following General Precautions in Douglas DC-8 MM, Chapter 27-00-00 “Maintenance Practices”. 2. Select Word 24.

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Enter OCTAL values for various control wheel positions indicated below. CONTROL WHEEL POSITION NEUTRAL FULL LEFT FULL RIGHT

EXPECTED OCTAL

OCTAL READING WORD 24

4061-4163 2570-6261 5360-6200

O. AILERON SURFACE 1. Select Word 8. 2. Enter OCTAL values for the aileron surface positions indicated below.

AILERON SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

EXPECTED OCTAL

OCTAL READING WORD 8

2735-3555 0014-0634 5656-6476

P. CONTROL COLUMN POSITION 1. Select Word 12. 2. Enter OCTAL values for various control column positions indicated below. CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

EXPECTED OCTAL

OCTAL READING WORD 12

1236-1362 1676-2020 1014-1136

Q. ELEVATOR SURFACE 1. Select Words 9 and 25. 2. Use Control Column Neutral pin position for the Neutral elevator position. (Ref. MM 27-30, page 506.)

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ELEVATOR POSITION Word 9 and 25 Frames 1 thru 4 PITCH NEUTRAL FULL UP FULL DOWN

EXPECTED OCTAL

ACTUAL OCTAL VALUE

4061-4163 6157-6261 3001-3103

R. RUDDER PEDAL POSITION 1. Provide main system hydraulic power following General Precautions in Douglas DC-8 MM, Chapter 27-00-00 “Maintenance Practices”. 2. Select Word 28. 3. Enter OCTAL values for various rudder pedal positions indicated below: RUDDER PEDAL POSITION

EXPECTED OCTAL

OCTAL READING WORD 28

NEUTRAL 2450-2760 FULL LEFT 1616-2160 FULL RIGHT 3542-4052 S. RUDDER SURFACE 1. Select Word 32. 2. Enter OCTAL values for the rudder surface positions indicated below: RUDDER SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

EXPECTED OCTAL

OCTAL READING WORD 32

3146-3656 0244-0753 6050-6560

T. TEST CLOSEOUT 1. If actual “OCTAL VALUE/READING” fall outside “EXPECTED OCTAL” range, generate a Non-Routine or Logbook entry and refer to L3 F800 Component Maintenance Manual for acceptable limitations. 2. Remove all test equipment, secure all LRU’s, and reinstall any panels that were removed to accomplish this test. 3. Review pages of the Checkout for omissions, clarity, or errors, and return to the Quality Control Department for retention in aircraft records file.

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F1000 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

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Page 1 of 10

Mechanic Inspector A. GENERAL This ground test verifies F1000 DFDR inputs and the recording capability of the DFDR unit. This test is applicable only to those aircraft with the Fairchild F1000 installed. Tools and Material Avionica, Inc. RSU DFDR Tester/Transcriber, Fairchild/Loral Portable Display Unit (PDDU), or Equivalent Digital Protractor Tilt Table and DG/VG Extender Harnesses (Optional) Pitot/Static Test Equipment Signal Generator HP204D or equivalent Work-stand capable of reaching the aft side of the horizontal stabilizer The DFDR monitors and records 17 parameters (Checkout Location): 1. TIME (D) 10. ENGINE THRUST (M, N) 2. RADIO TRANSMISSIONS (E) 11. VERTICAL ACCELERATION (O) 3. AUTOPILOT ENGAGE (F) 12. LONGITUDINAL ACCELERATION (O) 4. ALTITUDE (G) 13. CONTROL WHEEL POSITION (P) 5. AIRSPEED (H) 14. AILERON SURFACE (Q) 6. HEADING/COMPASS (I) 15. CONTROL COLUMN POS. (R) 7. ROLL ATTITUDE (J) 16. RUDDER PEDAL POSITION (S) 8. PITCH ATTITUDE (K) 17. RUDDER SURFACE (T) 9. ELEVATOR SURFACE (L) B. TEST SETUP 1. Record the following: a. Aircraft Serial Number/Tail Number: __________________________ b. DFDR Part Number: __________________________ c. DFDR Serial Number: __________________________ d. Biaxial Accelerometer Part Number (Left Wheel Well): ____________ e. Biaxial Accelerometer Serial Number: __________________________ 2. Check that the following aircraft systems are powered: DFDR power, 115 VAC Right & Left Emergency Buss, 28 VAC Instrument Buss 2 & 3, 28 VDC Right & Left Emergency Buss, Pilots & Co-pilots Horizon & Heading, VHF Communication Equipment 3. Connect front connector J1 on the DFDR to the DFDR test equipment. WARNING: WHEN POWER TO THE DFDR IS ON, THE DFDR WILL RECORD OVER PREVIOUSLY RECORDED DATA (25 HOURS EARLIER). TO MINIMIZE DESTRUCTION OF PREVIOUSLY RECORDED DATA, ONLY HAVE DFDR ON DURING TESTING. 4. Turn “ON” DFDR Switch and verify that the FLIGHT RECORDER OFF light extinguishes.

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5.

Verify FDR ON/OFF switch panel lighting works from “BRIGHT” to “OFF” using the FE panel light control. C. SYNC WORD 1. Select Word 1. 2. Verify the following SYNC Octal reading by Frame: FRAME # WORD #1 1 1107 2 2670 3 5107 4 6670 D. TIME 1. Select Word 57. 2. Verify that Word 57 counts upward. E. MICROPHONE KEY 1. Select Word 11. 2. Set power to Captain’s and First Officer’s VHF and HF Radios. NOTES: Skip HF Testing if HF Systems are deactivated. 3. Verify that Bit 1 (counting from the right) is “1”. NOTES: Select unused VHF and HF frequencies for keying tests. 4. Select and key Captain’s Audio Panel on VHF 1 or 2 and verify that Bit 1 toggles to “0” or the OCTAL number changes down by 1. 5. Repeat steps 3 and 4 for the Captain’s Audio Panel with HF 1 selected. 6. Verify that Bit 3 (counting from the right of the 12 bit word) is “1”. Key F.O.’s Audio Panel on VHF 2 and verify that Bit 3 toggles to “0” or the OCTAL number changes down by 4. 7. Repeat step 6 for F.O.’s Audio Panel with HF 2 selected. F. AUTOPILOT ENGAGE STATUS 1. Select Word 11. 2. Verify that Bit 2 (counting from the right) is “0”. 3. Engage Autopilot and verify that Bit 2 toggles to “1” (or the OCTAL number changes up by 2). 4. Note the Octal Reading for each status: AUTOPILOT OCTAL POSITION READING ENGAGED DISENGAGED G. ALTITUDE 1. Select Word 19 for Course Altitude and Word 58 for Fine Altitude. 2. Connect Pitot-static test equipment to the Pitot tube and Static port of the Pitot-static system that is connected to DADC-1 and to the Captain’s Pitotstatic system for reference. Reference the Pitot-static section of the DACO DC-8 Maintenance Manual for test procedure standards.

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Record Pitot and static system source used for DADC-1 air data. Static: ____________________________________________ Pitot: ____________________________________________ Using the Pitot-static test box, step through the following altitudes, verifying that the altitude displayed on the Captain’s Altimeter is the expected value. Ensure Captain’s Altimeter is in self-sensing mode. ELECTRIC ALTITUDE Word 19 Coarse Altitude Frames 1 thru 4 ACTUAL ALTITUDE CAPTAIN’S EXPECTED OCTAL ALTITUDE OCTAL ±100 Feet VALUE 4000 0235-0250 8000 0430-0445 22000 1165-1177 Word 58 Fine Altitude Frames 1 thru 4 ALTITUDE ±100 Feet

CAPTAIN’S ALTITUDE

EXPECTED OCTAL

ACTUAL OCTAL VALUE

4000 6244-6507 8000 4553-4757 22000 3020-3244 H. AIRSPEED 1. Select Word 10. 2. Connect Pitot-static test equipment to the Pitot and Static Systems operating the #1 DADC. Reference the Pitot-static section of the DACO DC-8 Maintenance Manual for test procedure standards. 3. Using the Pitot-static test box, step through the following airspeeds. Verify indicated airspeed on Captain’s airspeed indicator is the expected value. Ensure that the Captain’s ASI is in the remote sensing mode of operation. ELECTRIC AIRSPEED Word 10 Frames 1 thru 4 ACTUAL AIRSPEED ACTUAL EXPECTED OCTAL AIRSPEED OCTAL ±5 Knots VALUE 100 7110-0075 200 1300-1420 300 2760-3160 400 3665-4670 I.

HEADING/COMPASS 1. Select Word 9.

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Apply electrical power to the directional gyro compass system. Either rotate the aircraft on a compass rose or manually position the F.O.’s Compass Indicator (typically Captain’s RMI controls the F.O.’s Compass Indicator) to the following: HEADING Word 9 Frames 1 thru 4 ACTUAL HEADING ACTUAL EXPECTED OCTAL HEADING OCTAL ±1 ° VALUE 0 7755-0022 90 1755-2022 180 3755-4064 270 5755-6064

J. ROLL ATTITUDE NOTE: If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #2. 2. Remove Vertical Gyro #2 and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. Reset power to Vertical Gyro #2. 3. Using a digital protractor, tilt Vertical Gyro #2 to the following angles. Verify that roll is shown on the First Officer’s Horizon Indicator. 4. Select Word 14 and 46 and note values in the tables below. ROLL ATTITUDE Word 14, Frames 1 thru 4 ACTUAL ROLL ACTUAL EXPECTED OCTAL ROLL OCTAL ±1 ° VALUE 7754-0022 0° 0545-0600 30° right 1157-1222 60° right 7167-7224 30° left 6535-6600 60° left ROLL ATTITUDE Word 46, Frames 1 thru 4 ROLL ±1 ° 0° 30° right 60° right 30° left 60° left

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ACTUAL ROLL

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7754-0022 0545-0600 1157-1222 7167-7224 6535-6600

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K. PITCH ATTITUDE NOTE: If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #2. 2. Remove Vertical Gyro and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. 3. Reset power to Vertical Gyro #2. 4. Using a digital protractor, tilt Vertical Gyro #2 to the angles in the following table. Verify that pitch is indicated on the First Officer’s Horizon Indicator. 5. Select Words, 13, 29, 45, and then 61 and note values in the tables below. PITCH ATTITUDE Word 13, Frames 1 thru 4 ACTUAL PITCH ACTUAL EXPECTED OCTAL PITCH OCTAL ±1 ° VALUE Neutral 7745-0022 0545-0600 30° Up 1157-1222 60° Up 7167-7222 30° Down 6535-6600 60° Down PITCH ATTITUDE Word 29, Frames 1 thru 4 PITCH ±1 °

ACTUAL PITCH

Neutral 30° Up 60° Up 30° Down 60° Down

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0545-0600 1157-1222 7167-7222 6535-6600 PITCH ATTITUDE Word 45, Frames 1 thru 4

PITCH ±1 ° Neutral 30° Up 60° Up 30° Down 60° Down

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ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0545-0600 1157-1222 7167-7222 6535-6600

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F1000 DFDR GROUND TEST PROCEDURE – 17 PARAMETERS J/C Reviewed By:

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PITCH ATTITUDE Word 61 Frames 1 thru 4 PITCH ±1 °

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

Neutral 7745-0020 0545-0600 30° Up 1157-1222 60° Up 7167-7222 30° Down 6535-6600 60° Down L. ELEVATOR SURFACE 1. Select Word 23 and 55. 2. Using a digital protractor, move the Elevator to the following positions. 3. Use Control Column Neutral pin position for the neutral elevator position. ELEVATOR POSITION Word 23 Frames 1 thru 4 PITCH ACTUAL VALUE +1° NEUTRAL FULL UP FULL DOWN ELEVATOR POSITION Word 55 Frames 1 thru 4 ACTUAL PITCH VALUE +1° NEUTRAL FULL UP FULL DOWN M. ENGINE THRUST N1 TACH (DC-8-70 Series) 1. Remove the N1 engine TACH transmitter connector on each engine. 2. Using the signal generator, connect High to Pin 1, Low to Pin 2 to the TACH transmitter shipside connector.

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3.

Using the following Frame, Word, and Frequencies, monitor the respective N1 indicator while taking data. N1 FRAME 1 N1% ENGINE #1 ENGINE #2 TARGET FRAME 1-4 FRAME 1-4 WORD 8 WORD 40 30 60 90 N1 FRAME 2 ENGINE #3 ENGINE #4 N1% TARGET FRAME 1-4 FRAME 1-4 WORD 24 WORD 56 30 60 90 4. Remove wiring from signal generator to the N1 transmitter and reconnect the TACH Generator connector. N. ENGINE THRUST EPR (DC-8-60 Series) 1. Locate and set-up an EPR Test Set or prepare the aircraft for an engine run if the test set is not available. 2. Using the test set or the engine indicators set the EPR as shown in the table below. 3. Using the following Frame, Word, and Frequencies record the octal values in the space provided: EPR FRAME 1 ENGINE #1 ENGINE #2 EPR FRAME 1-4 FRAME 1-4 Target WORD 16 WORD 48 1.0 1.2 1.4 EPR FRAME 1 ENGINE #3 ENGINE #4 EPR FRAME 1-4 FRAME 1-4 Target WORD 32 WORD 64 1.0 1.2 1.4

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O. ACCELERATION 1. With the Accelerometer in the normal position, verify the following: ACCELERATION EXPECTED OCTAL

ACCEL

WORD

Vert. Long.

4,12,20,28,36,44,52,60 5,21,37,44,53

ACTUAL VALUE

3022-3715 3640-4277

2. Verify values change when Accelerometer is tilted. P. CONTROL WHEEL POSITION 1. Select Word 31. 2. Enter OCTAL values for various control wheel positions indicated below. REFERENCE ONLY: Null control wheel sensor to an approximate OCTAL reading of 3777. CONTROL WHEEL POSITION

OCTAL READING WORD 31

NEUTRAL FULL LEFT FULL RIGHT Q. AILERON SURFACE POSITION 1. Select Word 18. 2. Enter OCTAL values for the aileron surface positions indicated below. AILERON SURFACE POSITION NEUTRAL FULL LEFT (Aileron Up) FULL RIGHT (Aileron Down)

OCTAL READING WORD 18

R. CONTROL COLUMN POSITION 1. Select Word 30.

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2. Enter OCTAL values for various control column positions indicated below. NOTE: Use control column neutral rig-pin or put the gust lock in the “ON” position to obtain neutral position. CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

OCTAL READING WORD 30

S. RUDDER PEDAL POSITION 1. Provide main system hydraulic power following General Precautions in Douglas DC-8 MM, Chapter 27-00-00 “Maintenance Practices”. 2. Select Words 6, 22, 38, and 54. 3. Enter OCTAL values for various rudder pedal positions indicated below: RUDDER PEDAL POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 6

RUDDER PEDAL POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 22

RUDDER PEDAL POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 38

RUDDER PEDAL POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 54

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T. RUDDER SURFACE POSITION 1. Select Word 51. 2. Enter OCTAL values for the rudder surface positions indicated below: RUDDER SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 51

U. TEST CLOSEOUT 1. If actual “OCTAL VALUE/READING” falls outside “EXPECTED OCTAL” range, generate a Non-Routine or Logbook entry and refer to L3 F1000 Component Maintenance Manual for acceptable limitations. 2. Remove all test equipment, secure all LRU’s, and reinstall any panels that were removed to accomplish this test. 3. Review pages of the Checkout for omissions, clarity, or errors, and return to the Quality Control Department for retention in aircraft records file.

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SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 1 of 12

Mechanic Inspector A. GENERAL This ground test verifies 980-4120-RXUS inputs and the recording capability of the unit. This test is applicable to 71CX only. Tools and Material Avionica, Inc. RSU DFDR Tester/Transcriber, or Equivalent Digital Protractor, Kell-Strom Pro 3600 or equivalent. Rigging Protractor. Tilt Table and DG/VG Extender Harnesses (Optional) Pitot-static Test Equipment Work-stand capable of reaching the aft side of the horizontal stabilizer The DFDR monitors and records 17 parameters (Checkout Location): 1. RUDDER PEDAL (C) 10. TIME (K) 2. CONTROL COLUMN (D) 11. RADIO TRANSMISSION (L) 3. LATERAL CONTROL INPUT (E) 12. AIRSPEED (M) 4. AILERON POSITION (F & S) 13. PITCH ATTITUDE (N) 5. ELEVATOR POSITION (G) 14. ROLL ATTITUDE (O) 6. RUDDER POSITION (H) 15. ENGINE THRUST (P) 7. VERTICAL ACCELERATION (I) 16. HEADING/COMPASS (Q) 8. LONGITUDINAL ACCELERATION (I) 17. ALTITUDE (R) 9. AUTOPILOT ENGAGE (J) NOTE: If a sensor output does not fall within the expected octal value range during this test, refer to Section U for adjustment procedures. If a sensor does not adjust to within expected octal value range, initiate a non-routine discrepancy to isolate the cause. Mechanic Inspector B. TEST SETUP 1. Record the following: a. Aircraft Serial Number/Tail Number: __________________________ b. DFDR Part Number: __________________________ c. DFDR Serial Number: __________________________ d. Biaxial Accelerometer Part Number (Left Wheel Well):____________ e. Biaxial Accelerometer Serial Number: __________________________ 2. Check that the following aircraft systems are powered: DFDR power; 115 VAC Right & Left Emergency Buss; 28 VAC Instrument Buss 2 & 3; 28 VDC Right & Left Emergency Buss; Pilots & Co-pilots Horizon & Heading; VHF Communication Equipment. 3. Connect the RSU to the DFDR front test connector using the Adapter Cable. WARNING: WHEN POWER TO THE DFDR IS ON, THE DFDR WILL RECORD OVER PREVIOUSLY RECORDED DATA(25 HOURS EARLIER). TO MINIMIZE DESTRUCTION OF PREVIOUSLY RECORDED DATA, ONLY HAVE DFDR ON DURING TESTING. 4. Apply power to all the Allied Signal flight recorder system by setting the flight recorder and DFDR Sensor Ref. circuit breakers. 5.

Apply power to all of the appropriate systems that the DFDR monitors.

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SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

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Page 2 of 12 Mechanic

Inspector

C. RUDDER PEDAL POSITION TEST 1.

Ensure the rudder pedals are in the neutral position.

2.

On the portable tester select word 29, subframe 0.

3.

Verify that the position of the rudder pedals with reference to each other is the same. Verify the portable tester readout is between 1210 and 1260 octal.

4.

If the readout on the portable tester is not correct see Section U for adjustment procedures. 5. Move the rudder pedals to the positions listed below and record the portable tester readout as shown: SUBFRAME TEST CONDITION ACTUAL VALUE WORD 29

1 (2,3,4)

LEVEL (pedals even)

29

1 (2,3,4)

FULL RIGHT

29

1 (2,3,4)

FULL LEFT

6.

Repeat step “5” and confirm that readings are within +/- 62 octal of the value previously recorded.

D. CONTROL COLUMN POSITION TEST 1. Ensure the control column is in the neutral position. 2.

On the portable tester select word 17, subframe 0.

3.

Set the control column to it’s full forward position and verify that the readout on the portable tester is between 7760 and 0030 octal. 4. If readout on the portable tester is not between 7760 and 0030 see Section U for adjustment procedures. 5. Move the control column to the positions listed below and record the portable tester readout as shown: SUBFRAME TEST CONDITION ACTUAL VALUE WORD 17

1 (2,3,4)

FULL FORWARD

17

1 (2,3,4)

10° AFT

17 1 (2,3,4) FULL AFT 6. Repeat step “5” and confirm that readings are within +/- 62 octal of the value previously recorded.

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ATI001

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31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 3 of 12 Mechanic

Inspector

E. LATERAL CONTROL INPUT POSITION TEST 1. Ensure the lateral movement of the yoke is in the neutral position. 2.

On the portable tester select word 11, subframe 0.

3.

Using the digital protractor to measure the yoke (0 degrees is referenced as vertical) verify the yoke position is at 0 degrees. Verify the portable tester readout is between 3100 and 3160 octal. 4. If the readout on the portable tester is not between 3100 and 3160 see Section U for adjustment procedures. 5. Move the yoke to the right and left to the angles listed below and record the portable tester readout as shown: SUBFRAME TEST CONDITION ACTUAL VALUE WORD 11

1 (2,3,4)

LEVEL (0 degrees)

11

1 (2,3,4)

RIGHT (+30 degrees)

11

1 (2,3,4)

FULL RIGHT

11

1 (2,3,4)

LEFT (+30 degrees)

11

1 (2,3,4)

FULL LEFT

6. F.

Repeat step “5” and confirm that readings are within +/- 42 octal of the value previously recorded.

AILERON POSITION TEST (LEFT INBOARD) 1.

Move the aileron to the trail position. With the digital protractor, measure the angle of the aileron. Note this angle and use it as the reference angle for the following steps.

2.

On the portable tester select word 8, subframe 0.

3.

Ensure the aileron is at its reference position per step 1. Verify the portable tester readout is between 2720 and 3000 octal.

4.

If readout on the portable tester is not between 2720 and 3000 octal, refer to Section U for adjustment procedures.

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 4 of 12 Mechanic

Inspector

5. Manually move the aileron to the positions listed below and record the portable tester readout as shown: ACTUAL VALUE WORD SUBFRAME TEST CONDITION 8

1 (2,3,4)

TRAIL (Ref. Pos.)

8

1 (2,3,4)

3” DOWN (from Ref. Pos.)

8

1 (2,3,4)

FULL DOWN (from Ref. Pos.)

8

1 (2,3,4)

3” UP (from Ref. Pos.)

8

1 (2,3,4)

FULL UP (from Ref. Pos.)

6.

Repeat step “4” and confirm that readings are within +/- 42 octal of the value previously recorded.

G. ELEVATOR POSITION TEST 1. Move the elevator to the trail position. With the digital protractor, measure the angle of the elevator. Note this angle and use it as the reference angle for the following steps. 2.

On the portable tester select word 30, subframe 0.

3.

Ensure the elevator is at its reference position per step 1. Verify the portable tester readout is between 2300 and 2360.

4.

If the readout on the portable tester is not between 2300 and 2360 see Section U for adjustment procedures.

5.

Manually move the elevator to the positions listed below and record the portable tester readout as shown: ACTUAL VALUE WORD SUBFRAME TEST CONDITION 30

1 (2,3,4)

LEVEL (Ref. Pos.)

30

1 (2,3,4)

5” DOWN (from Ref. Pos.)

30

1 (2,3,4)

FULL DOWN (from Ref. Pos.)

30

1 (2,3,4)

10” UP (from Ref. Pos.)

30

1 (2,3,4)

FULL UP (from Ref. Pos.)

6.

Repeat step “5” and confirm that readings are within +/- 42 octal of the value previously recorded.

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 5 of 12 Mechanic

Inspector

H. RUDDER POSITION TEST 1. Move the rudder to the trail (reference) position. Note this position and use it as the reference for the following steps. With a tape measure, measure the travel of the trailing edge of the rudder. 2.

Position the flaps at 10 degrees.

3.

On the portable tester select word 22, subframe 0.

4.

Ensure the rudder is at its reference position per step 1.

5.

Verify the portable tester is between 5330 and 5377

6.

If the readout on the portable tester is not between 5330 and 5377 see Section U for adjustment procedures.

7. Move the rudder to the positions listed below and record the portable tester readout is as shown: WORD SUBFRAME TEST CONDITION

ACTUAL VALUE

22

1 (2,3,4)

TRAIL (Ref. Pos.)

22

1 (2,3,4)

13” RIGHT (from Ref. Pos.)

22

1 (2,3,4)

FULL RIGHT (from Ref. Pos.)

22

1 (2,3,4)

13” LEFT (from Ref. Pos.)

22 1 (2,3,4) FULL LEFT (from Ref. Pos.) 8. Repeat step “7” and confirm that readings are within +/- 42 octal of the value previously recorded.

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 6 of 12 Mechanic

I.

ACCELERATION TEST 1. Disconnect the accelerometer from the aircraft and free the end of the wire harness so that it can be turned in all directions. 2. With the accelerometer disconnected from the aircraft, turn it to the positions indicated and record the readings on the portable tester for Vertical Acceleration (Word 2), and Longitudinal Acceleration (Word 4) vary as the position changes as follows: TEST CONDITION a.

b.

J.

Inspector

SPEC.

EXPECTED

Vertical

1.000g +/- .2g

3520 to 3636

Longitudinal

0.000g +/- .2g

3614 to 4425

ACTUAL VALUE

Up +1g pointed up:

Forward +1g pointed up: Vertical

0.000g +/- .2g

2633 to 2752

Longitudinal

1.000g +/- .2g

7704 to 7777

AUTOPILOT ENGAGE STATUS TEST 1. Ensure the autopilot is disengaged. 2.

On the portable tester select word 9, subframe 0.

3.

Record bit 1 of word 9 and verify that it is set to “0”.

4.

Engage the autopilot. Verify that bit 1 of word 9 changes to “1” and record.

K. TIME TEST 1. Select word 14 or 37. 2.

Verify that word 14 or 37 counts upwards.

L. RADIO TRANSMISSION TEST 1. On the portable tester select word 25, subframe 0. 2. Select and key VHF 1 on the Captain’s audio control panel and verify word 25 bit 2 toggles to 0. 3. If HF systems are activated: Select and key HF 1 on the Captain’s audio control panel and verify that word 25 bit 2 toggles to 0. 4. Select and key VHF 2 on the First Officer’s audio control panel and verify that word 25 bit 2 toggles to 0. 5. If HF systems are activated: Select and key HF 2 on the First Officer’s audio control panel and verify that word 25 bit 2 toggles to 0.

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 7 of 12 Mechanic

Inspector

M. AIRSPEED TEST 1. On the portable tester select word 19 or 51. 2. 3.

Connect Pitot-static test equipment to the Pitot and Static Systems operating the #1 DADC. Using the Pitot-static test box, step through the following airspeeds. Verify indicated airspeed on Captain’s airspeed indicator is the expected value. ELECTRIC AIRSPEED Word 19 or 51 Frames 1 thru 4 ACTUAL AIRSPEED ACTUAL EXPECTED OCTAL AIRSPEED OCTAL ±5 Knots VALUE 100 7750-0020 200 1350-1420 300 2750-3020 400 3760-4020

N. PITCH ATTITUDE NOTE If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #1. 2. Remove Vertical Gyro and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. 3. Reset power to Vertical Gyro #1. 4. Using a digital protractor, tilt Vertical Gyro #1 to the angles in the following table. Verify that pitch is indicated on the Captain’s Horizon Indicator. 5. Select Words 9, 25, 41 and then 57 and note values in the tables below. PITCH ATTITUDE Word 9 Frames 1 thru 4 ACTUAL PITCH ACTUAL EXPECTED OCTAL PITCH OCTAL ±1 ° VALUE Neutral 7745-0022 0550-0620 30° Up 1160-1230 60° Up 7160-7230 30° Down 6540-6620 60° Down

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 8 of 12 Mechanic

Inspector

PITCH ATTITUDE Word 25 Frames 1 thru 4 PITCH ±1 ° Neutral 30° Up 60° Up 30° Down 60° Down

PITCH ±1 °

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0550-0620 1160-1230 7160-7230 6540-6620 PITCH ATTITUDE Word 41 Frames 1 thru 4 ACTUAL PITCH

Neutral 30° Up 60° Up 30° Down 60° Down

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0550-0620 1160-1230 7160-7230 6540-6620 PITCH ATTITUDE Word 57 Frames 1 thru 4

PITCH ±1 ° Neutral 30° Up 60° Up 30° Down 60° Down

31.02.38

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0550-0620 1160-1230 7160-7230 6540-6620

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 9 of 12 Mechanic

Inspector

O. ROLL ATTITUDE NOTE: If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #1. 2. Remove Vertical Gyro #1 and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. Reset power to Vertical Gyro #1. 3. Using a digital protractor, tilt Vertical Gyro #1 to the following angles. Verify that roll is shown on the Captain’s Horizon Indicator. 4. Select Word 13 and 45 and note values in the tables below. ROLL ATTITUDE Word 13 Frames 1 thru 4 ACTUAL ROLL ACTUAL EXPECTED OCTAL ROLL OCTAL ±1 ° VALUE 7754-0022 0° 0550-0620 30° right 1160-1230 60° right 7150-7220 30° left 6530-6610 60° left ROLL ATTITUDE Word 45 Frames 1 thru 4 ROLL ±1 ° 0° 30° right 60° right 30° left 60° left

Revision 20

ACTUAL ROLL

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7754-0022 0550-0620 1160-1230 7150-7220 6530-6610

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS RECURRING NON-ROUTINE

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 10 of 12 Mechanic

P.

Inspector

ENGINE THRUST TEST 1. Prepare each engine in turn for simulated EPR testing as per DC-8 MM chapter 77. 2. Pressurize the PT2 and PT7 line of each EPR transmitter to cause the cockpit indicator to read EPR settings listed below. 3. Record the octal value for the words listed below. EPR FRAME 1 EPR ENGINE #1 ENGINE #2 TARGET FRAME 1-4 FRAME 1-4 WORD 33 WORD 35 1.0 1.4 2.0 EPR FRAME 2 EPR ENGINE #3 ENGINE #4 TARGET FRAME 1-4 FRAME 1-4 WORD 7 WORD 15 1.0 1.4 2.0 4. Return the EPR systems to normal conditions.

Q. HEADING/COMPASS 1. Select Word 3. 2. Apply electrical power to the directional gyro compass system. Either rotate the aircraft on a compass rose or manually position the F.O.’s RMI to the following: HEADING Word 3 Frames 1 thru 4 ACTUAL HEADING ACTUAL EXPECTED OCTAL HEADING OCTAL ±1 ° VALUE 0 7755-0022 90 1755-2022 180 3755-4064 270 5755-6064

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ATI-001

JOB CARD NUMBER

31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 11 of 12 Mechanic

Inspector

R. ALTITUDE 1. Select Word 23 for Course Altitude and Word 5 for Fine Altitude. 2. Connect Pitot-static test equipment to the Pitot tube and Static port of the Pitot-static system that is connected to DADC-1 and to the Captain’s Pitot-static system for reference. Reference the Pitot-static section of the DACO Maintenance Manual for test procedure standards. 3. Record Pitot and static system source used for DADC-1 air data. Static: ____________________________________________ Pitot: ____________________________________________ 4.

Using the Pitot-static test box, step through the following altitudes, verifying that the altitude displayed on the Captain’s Altimeter is the expected value. Ensure Captain’s Altimeter is in self-sensing mode. ELECTRIC ALTITUDE Word 23 Course Altitude Frames 1 thru 4 ACTUAL ALTITUDE CAPTAIN’S EXPECTED OCTAL ALTITUDE OCTAL ±100 Feet VALUE 4000 0110-0150 15000 0620-0660 30000 1640-1670 Word 5 Fine Altitude Frames 1 thru 4 ALTITUDE ±100 Feet 4000 15000 35000

S.

CAPTAIN’S ALTITUDE

EXPECTED OCTAL

ACTUAL OCTAL VALUE

6220-6260 7770-0040 7730-7770

AILERON POSITION TEST (RIGHT INBOARD) 1.

Move the aileron to the trail position. With the digital protractor, measure the angle of the aileron. Note: this angle and use it as the reference angle for the following steps.

2.

On the portable tester select word 16, subframe 0.

3.

Ensure the aileron is at its reference position per step 1. Verify the portable tester readout is between 2430 and 2470 octal.

4.

If readout on the portable tester is not between 2430 and 2470 octal, refer to Section U for adjustment procedures.

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ATI-001

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31.02-3 JOB TITLE

SUNSTRAND 980-4120-RXUS - DFDR GROUND TEST PROCEDURE (71CX) J/C Reviewed By:

Date:

Page 12 of 12 Mechanic

Inspector

5. Manually move the aileron to the positions listed below and record the portable tester readout as shown: ACTUAL VALUE WORD SUBFRAME TEST CONDITION 16

1 (2,3,4)

TRAIL (Ref. Pos.)

16

1 (2,3,4)

3” DOWN (from Ref. Pos.)

16

1 (2,3,4)

FULL DOWN (from Ref. Pos.)

16

1 (2,3,4)

3” UP (from Ref. Pos.)

16

1 (2,3,4)

FULL UP (from Ref. Pos.)

T. RETURN AIRCRAFT TO SERVICEABLE CONDITION 1. Remove all test equipment, secure all LRU’s, and reinstall any panels that were removed to accomplish this test. 2. Review pages of this Job Card for omissions, clarity, or errors and return to the Quality Control Department for retention in aircraft records file. U. SENSOR ADJUSTMENT PROCEDURES 1. To adjust linear slide potentiometers; a) b) c) d) e) 2.

Secure the item being sensed in the neutral position. Loosen the clamp securing the potentiometer. Slide the potentiometer in the clamp until the desired readout is displayed on the portable tester. Secure the clamp around the potentiometer. Verify that the readout displayed on the portable tester is within tolerance.

To adjust synchros; a) b) c) d) e)

31.02.42

Secure the item being sensed in the neutral position. Loosen one of the screws that secure the synchro in the clamp. Rotate the synchro in the clamp until the desired readout is displayed on the portable tester. Secure the potentiometer in the clamp Verify that the readout displayed on the portable tester is within tolerance.

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31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS - DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 1 of 12

Mechanic Inspector A. GENERAL This ground test verifies 980-4100-GQUS DFDR inputs and the recording capability of the unit. This test is applicable to DC-8-73 series only. Tools and Material Avionica, Inc. RSU DFDR Tester/Transcriber, or Equivalent Digital Protractor, Kell-Strom Pro 3600 or equivalent. Rigging Protractor. Tilt Table and DG/VG Extender Harnesses (Optional) Pitot-static Test Equipment Signal Generator HP204D or equivalent Work-stand capable of reaching the aft side of the horizontal stabilizer The DFDR monitors and records 17 parameters (Checkout Location): 1. TIME (D) 10. ENGINE THRUST (M) 2. RADIO TRANSMISSIONS (E) 11. VERTICAL ACCELERATION (N) 3. AUTOPILOT (F) 12. LONGITUDINAL ACCELERATION (N) 4. ALTITUDE (G) 13. CONTROL WHEEL POS. (O) 5. AIRSPEED (H) 14. AILERON SURFACE (P) 6. HEADING/COMPASS (I) 15. CONTROL COLUMN POS. (Q) 7. ROLL ATTITUDE (J) 16. RUDDER PEDAL POS. (R) 8. PITCH ATTITUDE (K) 17. RUDDER SURFACE POS. (S) 9. ELEVATOR SURF. (L) NOTE: If a sensor output does not fall within the expected octal value range during this test, refer to Section U for adjustment procedures. If a sensor does not adjust to within expected octal value range, initiate a non-routine discrepancy to isolate the cause. Mechanic Inspector B. TEST SETUP 1. Record the following: a. Aircraft Serial Number/Tail Number: __________________________ b. DFDR Part Number: __________________________ c. DFDR Serial Number: __________________________ d. Biaxial Accelerometer Part Number (Left Wheel Well): ____________ e. Biaxial Accelerometer Serial Number: __________________________ 2. Check that the following aircraft systems are powered: DFDR power, 115 VAC Right & Left Emergency Buss, 28 VAC Instrument Buss 2 & 3, 28 VDC Right & Left Emergency Buss, Pilots & Co-pilots Horizon & Heading, VHF Communication Equipment 3. Connect front connector J1 on the DFDR to the DFDR test equipment. WARNING: WHEN POWER TO THE DFDR IS ON, THE DFDR WILL RECORD OVER PREVIOUSLY RECORDED DATA (25 HOURS EARLIER). TO MINIMIZE DESTRUCTION OF PREVIOUSLY RECORDED DATA, ONLY HAVE DFDR ON DURING TESTING. 4. Turn “ON” DFDR Switch and verify that the FLIGHT RECORDER OFF light extinguishes.

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ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 2 of 12 Mechanic

5.

Inspector

Verify FDR ON/OFF switch panel lighting works from “BRIGHT” to “OFF” using the FE panel light control.

C. SYNC WORD 1. Select Word 1. 2. Verify the following SYNC Octal reading by Frame: FRAME # WORD #1 1 1107 2 2670 3 5107 4 6670 D. TIME 1. Select Word 57. 2. Verify that Word 57 counts upward. E. MICROPHONE KEY 1. Select Word 9. 2. Set power to Captain’s and First Officer’s VHF and HF Radios. NOTES: Skip HF Testing if HF Systems are deactivated. 3. Verify that Bit 1 (counting from the right) is “1”. NOTES: Select unused VHF and HF frequencies for keying tests. 4. Select and key Captain’s Audio Panel on VHF 1 or 2 and verify that Bit 1 toggles to “0” or the OCTAL number changes down by 1. 5. Repeat steps 3 and 4 for the Captain’s Audio Panel with HF 1 selected. 6. Key F.O.’s Audio Panel on VHF 2 and verify that Bit 1 toggles to “0” or the OCTAL number changes down by 1. 7. Repeat step 6 for F.O.’s Audio Panel with HF 2 selected. F.

AUTOPILOT ENGAGE STATUS 1. Select Word 9. 2. Verify that Bit 2 (counting from the right) is “0”. 3. Engage Autopilot and verify that Bit 2 toggles to “1” (or the OCTAL number changes up by 2). 4. AUTOPILOT OCTAL POSITION READING ENGAGED DISENGAGED

G. ALTITUDE 1. Select Word 6 for Altitude. 2. Reference the Pitot-static section of the DACO DC-8 Maintenance Manual for test procedure standards.

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ATI-001

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31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 3 of 12 Mechanic

3.

Inspector

Using the Pitot-static text box, step through the following altitudes, verifying that the altitude displayed on the Captain’s Altimeter is the expected value. Ensure Captain’s Altimeter is in self-sensing mode. NOTE: Set Captain’s & F/O’s altimeter baro correction to 29.92 In.Hg. ALTITUDE Word 6 Frames 1 thru 4 ACTUAL ALTITUDE CAPTAIN’S EXPECTED OCTAL ALTITUDE OCTAL ±100 Feet VALUE 0 2204-2231 4,000 2715-2750 8,000 3365-3421 22,000 5020-5063

H. AIRSPEED 1. Select Word 21. 2. Reference the Pitot-static section of the DACO DC-8 Maintenance Manual for test procedure standards. 3. Using the Pitot-static test box, step through the following airspeeds. Verify indicated airspeed on Captain’s airspeed indicator is the expected value. Ensure that the Captain’s ASI is in the remote sensing mode of operation. AIRSPEED Word 21 Frames 1 thru 4 ACTUAL AIRSPEED ACTUAL EXPECTED OCTAL AIRSPEED OCTAL ±5 Knots VALUE 100 1410-1444 200 2134-2162 300 3250-3311 400 5010-5044

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ATI-001

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31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 4 of 12 Mechanic

I.

Inspector

HEADING/COMPASS 1. Select Word 3. 2. Apply electrical power to the directional gyro compass system. Position the F.O.’s Compass Indicator. HEADING Word 3 Frames 1 thru 4 ACTUAL HEADING ACTUAL EXPECTED OCTAL HEADING OCTAL ±1 ° VALUE 0 7755-0022 90 1755-2022 180 3755-4064 270 5755-6064

J. ROLL ATTITUDE NOTE If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #1. 2. Remove Vertical Gyro #1 and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. Reset power to Vertical Gyro #1. 3. Using a digital protractor, tilt Vertical Gyro #1 to the following angles. Verify that roll is shown on the Captain’s Horizon Indicator. 4. Select Word 13, 29,45, and 61 and note values in the tables below. ROLL ATTITUDE Word 13 Frames 1 thru 4 ACTUAL ROLL ACTUAL EXPECTED OCTAL ROLL OCTAL ±1 ° VALUE 7754-0022 0° 0420-0500 30° right 1260-1340 60° right 7300-7360 30° left 6430-6520 60° left

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31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 5 of 12 Mechanic

Inspector

ROLL ATTITUDE Word 29 Frames 1 thru 4 ROLL ±1 °

ACTUAL ROLL

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7754-0022 0420-0500 1260-1340 7300-7360 6430-6520

0° 30° right 60° right 30° left 60° left

ROLL ATTITUDE Word 45 Frames 1 thru 4 ROLL ±1 °

ACTUAL ROLL

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7754-0022 0420-0500 1260-1340 7300-7360 6430-6520

0° 30° right 60° right 30° left 60° left

ROLL ATTITUDE Word 61 Frames 1 thru 4 ROLL ±1 ° 0° 30° right 60° right 30° left 60° left

Revision 22

ACTUAL ROLL

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7754-0022 0420-0500 1260-1340 7300-7360 6430-6520

05-20-05

31.02.47

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31 - INSTRUMENTS RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 6 of 12 Mechanic Inspector

K. PITCH ATTITUDE NOTE If the aircraft has FOG gyros installed, substitute FOG gyro for Vertical Gyro in the following steps. 1. Remove power to Vertical Gyro #1. 2. Remove Vertical Gyro and set on mounting bracket or employ extender harness to sit Vertical Gyro on tilt table. 3. Reset power to Vertical Gyro #1. 4. Using a digital protractor, tilt Vertical Gyro #1 to the angles in the following table. Verify that pitch is indicated on the Captain’s Horizon Indicator. 5. Select Words, 9, 25, 41, and then 57 and note values in the tables below. PITCH ATTITUDE Word 9 Frames 1 thru 4 ACTUAL PITCH ACTUAL EXPECTED OCTAL PITCH OCTAL ±1 ° VALUE Neutral 7745-0022 0420-0500 30° Up 1300-1377 60° Up 7300-7377 30° Down 6410-6510 60° Down PITCH ATTITUDE Word 25 Frames 1 thru 4 PITCH ±1 °

ACTUAL PITCH

Neutral 30° Up 60° Up 30° Down 60° Down

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0420-0500 1300-1377 7300-7377 6410-6510 PITCH ATTITUDE Word 41 Frames 1 thru 4

PITCH ±1 ° Neutral 30° Up 60° Up 30° Down 60° Down

31.02.48

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0420-0500 1300-1377 7300-7377 6410-6510

05-20-05

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 7 of 12 Mechanic

Inspector

PITCH ATTITUDE Word 57 Frames 1 thru 4 PITCH ±1 ° Neutral 30° Up 60° Up 30° Down 60° Down

ACTUAL PITCH

EXPECTED OCTAL

ACTUAL OCTAL VALUE

7745-0022 0420-0500 1300-1377 7300-7377 6410-6510

L. ELEVATOR POSITION 1. Select Word 15 and 47. 2. Using a digital protractor, move the Elevator to the following positions. 3. Use Control Column Neutral pin position for the neutral elevator position. ELEVATOR POSITION Word 15 Frames 1 thru 4 ACTUAL PITCH VALUE +1° NEUTRAL FULL UP FULL DOWN ELEVATOR POSITION Word 47 Frames 1 thru 4 PITCH ACTUAL VALUE +1° NEUTRAL FULL UP FULL DOWN M. ENGINE THRUST N1 TACH 1. Remove the N1 engine TACH transmitter connector on each engine. 2. Using the signal generator, connect High to Pin 1, Low to Pin 2 to the TACH transmitter shipside connector. 3. Using the following Frame, Word, and Frequencies, monitor the respective N1 indicator while taking data. N1 N1% ENGINE #1 ENGINE #2 TARGET FRAME 1-4 FRAME 1-4 WORD 33 WORD 35 30 60 90

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31.02.49

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 8 of 12 Mechanic

N1 ENGINE #3 FRAME 1-4 WORD 38

N1% TARGET

Inspector

ENGINE #4 FRAME 1-4 WORD 53

30 60 90 4.

Remove wiring from signal generator to the N1 transmitter and reconnect the TACH Generator connector.

N. ACCELERATION 1. With the Accelerometer in the normal position, verify the following: ACCELERATION

2.

EXPECTED OCTAL

ACCEL

WORD

Vert. Long.

2,10,18.26,34,42,50,58 4,20,36,52

ACTUAL VALUE

3022-3715 3640-4277

Verify values change when Accelerometer is tilted.

O. CONTROL WHEEL POSITION 1. Select Word 19, 51. 2. Verify the tester readout is between 7734 and 7777 or 0000 and 0042 in neutral, if not reference section U for alignment procedures. 3. Enter OCTAL values for various control wheel positions indicated below. CONTROL WHEEL POSITION

OCTAL READING WORD 19

OCTAL READING WORD 51

NEUTRAL FULL LEFT FULL RIGHT

31.02.50

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 9 of 12 Mechanic

P.

Inspector

AILERON SURFACE POSITION 1. Select Word 7, 23, 39, 55. 2. Verify the tester readout is between 7734 and 7777 or 0000 and 0042 in neutral, if not reference Section U for alignment procedures. 3. Enter OCTAL values for the aileron surface positions indicated below. AILERON SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 7

AILERON SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 23

AILERON SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 39

AILERON SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 55

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TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 10 of 12 Mechanic

Inspector

Q. CONTROL COLUMN POSITION 1. Select Word 8,16,24,32,40,48,56, and 64. 2. Verify the tester readout is between 3605 and 3667 in neutral, if not reference Section U for alignment procedures. 3. Enter OCTAL values for various control column positions indicated below.

31.02.52

CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

OCTAL READING WORD 8

OCTAL READING WORD 16

CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

OCTAL READING WORD 24

OCTAL READING WORD 32

CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

OCTAL READING WORD 40

OCTAL READING WORD 48

CONTROL COLUMN POSITION NEUTRAL FULL FORWARD FULL AFT

OCTAL READING WORD 56

OCTAL READING WORD 64

05-20-05

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 11 of 12 Mechanic

Inspector

R. RUDDER PEDAL POSITION 1. Provide main system hydraulic power following General Precautions in Douglas DC-8 MM, Chapter 27-00-00 “Maintenance Practices”. 2.

Position the flaps at 10 degrees.

3. 4.

Select Words 17, and 49. Verify the tester readout is between 3605 and 3667 in neutral, if not reference Section U for alignment procedures. Enter OCTAL values for various rudder pedal positions indicated below:

5.

RUDDER PEDAL POSITION

OCTAL READING WORD 17

OCTAL READING WORD 49

NEUTRAL FULL LEFT FULL RIGHT S.

RUDDER SURFACE POSITION 1. Select Word 5. 2. Verify the tester readout is between 7734 and 7777 or 0000 and 0042 in neutral, if not reference Section U for alignment procedures. 3. Position the flaps at 10 degrees. 4. Enter OCTAL values for the rudder surface positions indicated below: RUDDER SURFACE POSITION NEUTRAL FULL LEFT FULL RIGHT

OCTAL READING WORD 5

T. TEST CLOSEOUT 1. Remove all test equipment, secure all LRU’s, and reinstall any panels that were removed to accomplish this test. 2. Review pages of this Job Card for omissions, clarity, or errors, and return to the Quality Control Department for retention in aircraft records file.

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31.02.53

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ATI A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

31.02-4 JOB TITLE

ALLIED-SIGNAL 980-4100-GQUS DFDR GROUND TEST PROCEDURE (-73 Series Only) J/C Reviewed By:

Date:

Page 12 of 12 Mechanic

Inspector

U. SENSOR ALIGNMENT 1. To adjust linear slide potentiometers: a) b) c) d) e) 2.

Secure the item being sensed in the neutral position. Loosen the clamp securing the potentiometer. Slide the potentiometer in the clamp until the desired readout is displayed on the portable tester. Secure the clamp around the potentiometer. Verify that the readout displayed on the portable tester is within tolerance.

To adjust synchros: a) b) c) d) e)

31.02.54

Secure the item being sensed in the neutral position. Loosen one of the screws that secure the synchro in the clamp. Rotate the synchro in the clamp until the desired readout is displayed on the portable tester. Secure the potentiometer in the clamp Verify that the readout displayed on the portable tester is within tolerance.

05-20-05

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS

31.03 – UNDERWATER ACOUSTIC BEACON TESTERS TABLE OF CONTENTS Paragraph 1.

DUKANE MODEL PL-1 ------------------------------------------------------------

Page 31.03.1

2.

BENTHOS ATS-260 ACOUSTIC TEST SET -----------------------------------

31.03.1

3.

ALTERNATIVE BATTERY CHECK---------------------------------------------

31.03.2

FIGURES 31.03-1

PINGLITE MODEL PL-1 TESTER -----------------------------------------------

31.03.1

31.03-2

BENTHOS ATS-260 TESTER -----------------------------------------------------

31.03.1

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INTENTIONAL BLANK

31.03-ii

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Revision 14

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.03 UNDERWATER ACOUSTIC BEACON TESTERS A. This chapter covers three methods of testing underwater acoustic beacons/batteries using either the Dukane Model PL-1, the Benthos ATS-260, or a high impedance digital voltmeter. 1. Dukane Model PL-1. The Dukane Model PL-1 Pinglite is designed to check operational status of Underwater Acoustic Beacons (pingers) in the 37.5 kHz range. a. It includes a shorting tab on an attached wire to prevent foreign object damage due to loose articles dropper(i.e., paper clips, braid, wire) within the aircraft. b. Using the shorting tab, short water switch on beacon from center pin to case. Press end of PL-1 to beacon body approximately 1” from switch end. Depress and hold operation button. Light will flash with each output pulse of beacon indicating it is operating.

Pinglite PL-l Figure 31.03-1

2. Benthos ATS-260 Acoustic Test Set. To verify battery condition and to test the operation of the beacon of the Benthos ELP-362 Emergency Locator Beacon, use the following procedures: a. attach the cable clip directly to the beacon housing as shown in Figure 31.03-2. It is not necessary to remove the beacon from the mount to conduct the test. Although the cable clip can touch the beacon label, it is important that at least part of the clip touch the bare aluminum surface of the beacon.

Benthos ATS-260 Figure 31.03-2

b. apply the test set probe to the water activation switch as shown.

Revision 12

06-01-01

31.03.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.03 UNDERWATER ACOUSTIC BEACON TESTERS(cont’d) c. to test the battery, hold down the button labeled PUSH TO TEST on the test set handle. Check the beacon battery condition by observing the green and red indicators on the test set. If green, battery is good and if red battery should be replaced. d. to test the beacon operating condition, listen for an audible tone from the test set and/or observe the amber indicator. If there is an audible tone and/or the amber indicator is flashing, the beacon is operating properly. If there is no tone or flashing amber indicator, the beacon is not operating properly. 3. Alternative Battery Check. In addition to using the above test sets to check the beacon’s battery condition, a high-impedance (minimum input impedance of 10 megohms) digital voltmeter can be used to measure the battery voltage. Use the following procedure to measure the battery voltage: a. Place the negative meter lead on the water activation switch. b. Place the positive meter lead on the bare aluminum surface of the beacon housing. Do not place the meter lead on the label. c. Measure the battery voltage. The beacon is operable at the given minimum acceptable voltage as shown: Manufacturer P/N Benthos ELP-362D Dukane DK100/DK120 Dukane DK100/DK120 Dukane DK100/DK120 Dukane DK100/DK120 * Located on beacon label.

31.03.2

Battery Code* NA A B C No Code

05-20-04

Minimum Acceptable Voltage 6.0 Volts 3.55 Volts 2.97 Volts 2.97 Volts 7.10 Volts

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS

31.04 – UNDERWATER ACOUSTIC BEACONS TABLE OF CONTENTS Paragraph 1.

DUKANE MODELS DK100 AND DK120---------------------------------------

Page 31.04.1

2.

BENTHOS / DATASONICS MODEL ELP-362D-------------------------------

31.04.6

FIGURES 31.04-1

DK100 / DK120 UAB’s AND MOUNTING KIT--------------------------------

31.04.1

31.04-2

BEACON BATTERY END COVER REMOVAL-------------------------------

31.04.3

31.04-3

BEACON EXPLODED VIEW -----------------------------------------------------

31.04.5

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06-03-02

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INTENTIONAL BLANK

31.04-ii

12-03-01

Revision 14

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS A. This chapter contains description, theory and maintenance for Underwater Acoustic Beacons manufactured by Dukane Corporation and Benthos, Inc. (formerly Datasonics). 1. Dukane Models DK100 and DK120

DK100/DK120 UAB’s AND MOUNTING KIT Figure 31.04-1 a. The DK100/DK120 Beacon consists of a self-contained battery, an electronic module and a transducer. It is housed in a cylindrical watertight aluminum case capable of withstanding high-G impact shock and deep water immersion. b. The Beacon is mounted to a flight recorder or the cockpit voice recorder by means of a mounting kit. This mount consists of an extruded channel, which in conjunction with a securing plate, retains the beacon. c. DK-100 Maintenance. This section contains DK100 Beacon cleaning, battery testing, disposal and storage procedures. Initially beacons must be tested at every installation or beacon change. The recommended schedule for beacon cleaning and testing is every 24 months, when the beacon is installed on a recorder and the recorder is installed in accordance with its TSO. Otherwise, the recommended schedule is every six months. The required scheduled for battery replacement, available only at Dukane Corp. is every 6 years. (1) Cleaning. Clean the switch end of the beacon with a soft cloth and mild detergent, then dry thoroughly with a clean cloth. Clean the switch end insulator to prevent leakage currents from occurring across the switch. This will affect battery life. The water switch should be cleaned at any time if dirt or dust becomes apparent.

Revision 15

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31.04.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) (2) Precautions. (a)

The beacon must not be exposed to heat in excess of 160°F (71°C).

(b)

The beacon must not be disassembled in any way. Although the beacon may be removed from the mount, the water switch cap and battery cap must remain in place at all times.

(c)

Any situation that could possibly crush or penetrate the case of the beacon should be avoided.

(3) Battery. Replace the beacon as an assembly. DK100 or DK120 may be used as a replacement. The DK100 beacon requires NO battery maintenance. DO NOT remove the battery at any time. The beacon must be taken out of service according to the expiration date printed on the beacon case and returned to TOL Stores for disposition. WARNING:

FAILURE TO OBSERVE THE ABOVE PRECAUTIONS COULD RESULT IN THE RELEASE OF HAZARDOUS CHEMICALS.

(4) Disposal. Should it be necessary to dispose of the beacon for any reason DO NOT THROW IT AWAY. Ship the beacon, in accordance with current IATA Dangerous Goods Regulations, to TOL Stores for proper disposal. (5) Storage. When long term storage of the DK100 beacon is required it should be stored in the original shipping container (or equivalent) in a cool dry environment. d. DK120 Maintenance. This section contains DK120 Beacon cleaning, battery testing, battery replacement and testing, disposal and storage procedures. Initially beacons must be tested at every installation. The required scheduled for battery replacement is every 6 years. (1) Cleaning. Clean the switch end of the beacon with a soft cloth and mild detergent, then dry thoroughly with a clean cloth. Clean the switch end insulator to prevent leakage currents from occurring across the switch. This will affect battery life. The water switch should be cleaned at any time if dirt or dust becomes apparent. (2) Testing. Make sure the beacon case is clean and dry prior to testing. Test in accordance with procedures in Section 31.03.

31.04.2

06-01-01

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) (3)

Precautions. (a) The beacon must not be exposed to heat in excess of 160°F (71°C). (b) Any situation that could possibly crush or penetrate the case of the beacon should be avoided.

(4)

Disassembly. Disassembly of the beacon is limited to battery replacement, as outlined in paragraph below.

(5)

Battery Replacement and Testing. Battery replacement should be done under clean conditions to prevent dust from contaminating O-ring and lubricant. Because the old O-ring may have developed a set with age, O-ring replacement is mandatory at the time of battery change. O-ring lubrication should be applied to new O-ring and threads before installation.

WARNING:

INCORRECT BATTERY INSTALLATION WILL CAUSE PERMANENT DAMAGE TO THE BEACON.

CAUTION:

TO AVOID INTERNAL DAMAGE, DO NOT CLAMP THE BEACON IN A VISE, UNLESS A VISE-CLAMP (P/N 810-546) OR EQUIVALENT IS USED.

Battery End Cover Removal with Vise Clamp and Spanner Wrench Figure 31.04-2

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31.04.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) (a) Secure the beacon with Vise-Clamp (P/N 810-546) or equivalent as shown in figure 31.04-2. (b) Use Spanner Wrench (P/N 810-325) to remove end cover marked “BATTERY ACCESS” by unscrewing counterclockwise. Break away torque is usually high so spanner wrench should be held firmly in contact with battery end cover in order to prevent damage to wrench holes. (c) Remove shock cushion (P/N 810-419) from battery end of beacon if not removed with cover. See Figure 31.04-3. (d) Remove the old O-ring from the cover. Do not use steel screwdriver or sharp tool because of danger of damaging O-ring groove. (e) Remove the old battery. NOTE: Ensure shock cushion 810-371A remains in beacon. (f) Clean the threads, O-ring groove in the body and the threads on the cover by wiping them thoroughly with solvent. CAUTION:

FOREIGN SUBSTANCES IN LUBRICANT ON SEALING SURFACES MAY DAMAGE THREADS AND/OR ALLOW WATER LEAKAGE THROUGH THE O-RING SEAL. SCRATCHES OR GOUGES ON SEALING SURFACES WILL ALSO CAUSE WATER LEAKAGE.

(g) Carefully install new O-ring (P/N 810-342) on battery cover. Apply a thin coating of O-ring lubricant (P/N 810-346, 810-500 or 810-1168) to O-ring, Oring groove and threads. Note: O-ring, lubricant and battery are provided in battery replacement kit 810-2007/K used with Battery Code B or 810-2008/K used with Battery Code C. (h) Install new battery. Be sure the end marked “INSERT THIS END FIRST” goes in first. See Figure 31.04-4. Contact Dukane Corporation, Seacom Division at (630) 584-2300, Ext. 860 for replacement 810-2007/K and 810-2008/K battery kits. WARNING:

31.04.4

REPLACE BATTERY WITH 810-2007 USED WITH BATTERY CODE B OR 810-2008 USED WITH BATTERY CODE C BATTERY ONLY. USE OF ANOTHER BATTERY MAY PRESENT A RISK OF FIRE OR EXPLOSION. USE OF AN UNAUTHORIZED BATTERY WILL VOID THE WARRANTY AND MAY CAUSE AN INOPERATIVE OR DANGEROUS CONDITION

05-20-04

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) CAUTION:

RISK OF FIRE, EXPLOSION, AND BURNS. DO NOT RECHARGE, DISASSEMBLE, HEAT ABOVE 160°F (71°C) OR INCINERATE. DISPOSE OF BATTERIES PROMPTLY, KEEP AWAY FROM CHILDREN.

Beacon Exploded View Showing Relative Location of Battery and Related Parts Figure 31.04-3 (i) Position rubber shock cushion (P/N 810-419) around contact spring and on the inside of the battery cover. (j) Replace the end cover and tighten until the cover flange contacts the body or leaves less than a 0.003 gap. Use hand force only on the wrench. Hold the beacon in a vise clamp as shown in Figure 31.04-2. Clean beacon exterior of excess O-ring grease. (k) Perform operational test of beacon as outlined in section 31.03. (l) Replace the battery replacement date label on the outside of the beacon case with the new label supplied with the battery. Locate the label on the beacon body in the space provided. Never put label on the water switch end of the beacon. (6)

Battery Disposal. Dispose of battery in accordance with all local, state and federal regulations. Dispose of batteries promptly, keep away from children.

(7)

Beacon Storage. When long term storage of a beacon is required, it should be stored in the original shipping container(or equivalent). Make sure it is stored in a cool dry environment. Beacon may be stored without battery.

Revision 21

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31.04.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) 2. BENTHOS/Datasonics Model ELP-362D. This section covers the cleaning, recommended testing interval, and disposal and storage procedures for the ELP-362D Emergency Locator Beacon. NOTE:

The Benthos and Datasonics model ELP-362D beacons and batteries (listed below) are interchangeable.

a. Cleaning. The end cap with the water switch should remain free and clear of dirt, grease and dust. If necessary, should be cleaned with a detergent and dried thoroughly with a clean cloth. b. Maintenance Precautions. The following precautions should be exercised when handling or storing the beacon: (1)

The beacon should not be exposed to temperatures in excess of 160 degrees F, as the battery life can be reduces by storage in a high temperature environment. When long-term storage is required, the beacon should be stored in a cool, dry environment in its original shipping container.

(2)

Any situation that could possibly crush or penetrate the case of the beacon should be avoided.

c. Battery Maintenance. (1)

To determine the battery maintenance procedure, first check the serial number of the beacon. If the unit does not have an “S” before the serial number: (a) Replace the battery by the date stamped on the beacon label. (b) Use the battery replacement procedure provided with the lithium battery replacement kit P/N B362-06192-2(standard lithium battery) or P/N B36206192-1(optional lithium battery). The kit includes the battery and a lubricated O-ring.

(2)

If the unit does have an “S” before the serial number it must be returned to Benthos for battery replacement.

WARNING: DO NOT REMOVE THE BATTERY WITH AN “S” BEFORE THE SERIAL NUMBER UNDER ANY CIRCUMSTANCES. IF THE BEACON IS DISASSEMBLED IN ANY WAY, THE WARRANTY IS VOID. WARNING: FAILURE TO OBSERVE THESE PRECAUTIONS COULD RESULT IN THE RELEASE OF HAZARDOUS CHEMICALS.

31.04.6

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.04 UNDERWATER ACOUSTIC BEACONS(cont’d) d. Return Procedures. Do not throw the beacon away. If it is necessary to return the beacon for service, battery replacement or disposal, contact RJE International or Benthos for a Return Material Authorization number and shipping instructions: RJE International, Inc. Tel: 949-727-9399 Fax: 949-727-0070 e-mail: [email protected]

Revision 12

Benthos, Inc. Tel: 508-563-1000 Fax: 508-563-6444 e-mail: [email protected]

06-01-01

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31.05 – EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (B&D 65016-003) TABLE OF CONTENTS Paragraph 1.

GENERAL-----------------------------------------------------------------------------

Page 31.05.1

2.

OPERATING CHARACTERISTICS----------------------------------------------

31.05.1

3.

GROUND CHECK -------------------------------------------------------------------

31.05.3

4.

MAINTENANCE---------------------------------------------------------------------

31.05.5

FIGURES 31.05-1

EGT INDICATOR--------------------------------------------------------------------

31.05.2

31.05-2

LIGHT TRAY REPLACEMENT --------------------------------------------------

31.05.5

TABLES 31.05-1

TAKE-OFF MAXIMUM TEMPERATURE--------------------------------------

31.05.2

31.05-2

TEMPERATURE ADJUSTMENT CONTROL KNOB ROTATION---------

31.05.4

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (B&D 65016-003) 1. General a. The ATI Engineering Order (77-DC8-920) installed a new technology digital/analog EGT indicator (P/N 65016-003 made by B&D Instruments and Avionics), which improved EGT system accuracy and reliability. In addition, it modified the aircraft wiring in accordance with Douglas Service Bulletin 77-26, which removed the balance resistor and the need to calibrate aircraft line resistance with the balance resistor. Furthermore, installation of these indicators allowed an increase in EGT limitations to the engine certification limit of 905° C. 2. Operating Characteristics a. Start-Up (1) Immediately after power is applied to the indicator a self-test shall occur. Four dashes “----” on the digital display and no pointer movement indicates a failure. Successful completion of the self-test results in a pointer reset and a digital display test. Normal operation begins within 2 seconds for the analog display and the 0 to 5 VDC output. The digital display begins normal operation after one cycle of the digit test, which has a duration of 5 ± 1 seconds. b. EGT Process (1) The data is sufficiently filtered to minimize pointer jittering and digital flickering. The inputs are converted to outputs in accordance with the National Bureau of Standards monograph 125, table A7.1.1. All digital indications of 906 ± 3° C and above flash at a rate of 2 ± 1 Hz. c. Take-Off Maximum Temperature (1) The EGT Indicator continues to record the maximum temperature until the Weight on Wheels switch (Pin A) is open for 50 (-1/+3) seconds. (Reference Table 31.051)

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (cont’d) CONDITION Starting (ground and in-flight)

Maximum Continuous

EGT LIMIT 725°C 750°C 800°C 835° C 870° C*

TIME LIMIT Continuous 30 Seconds 10 Seconds Continuous Continuous

Takeoff

870° C 5 Minutes 905° C* 5 Minutes * For aircraft incorporating Douglas Service Bulletin Number 77-26 or equivalent and using digital/analog instruments, the Maximum Continuous EGT LIMIT is 870° C; the Takeoff EGT LIMIT is 905° C. TAKE-OFF MAXIMUM TEMPERATURE Table 31.05-1 d. EGT Indicator Operation

EGT INDICATOR Figure 31.05-1 (1) EGT Pointer: When the airplane is electrically powered, the EGT Pointer indicates below 300° for approximately two seconds, then indicates static engine temperature. No pointer movement indicates a failure.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (cont’d) (2) EGT Digital Display: When the airplane is electrically powered, the EGT Digital Display cycles “0000,” “1111,” “2222,” “3333,” “4444,” “5555,” “6666,” “7777,” “8888,” and “9999” for approximately 0.5 seconds of each number for a total of 5 ± 1 seconds. Four dashes “----” on the digital display indicates a failure. (3) Recall Button: The EGT Indicator records the maximum temperature from the time power is applied until 50 (-1/+3) seconds after the Weight on Wheels Switch (pin A) is open. The maximum temperature is displayed on the digital display when the recall button in the dial face is depressed. The digital display returns to normal operation within 2 (± 0.5) seconds after the button is released. The pointer continues to indicate the actual engine temperature during this period. 3. Ground Check a. Equipment Required: EGT Test Set: Barfield TT 1000A, or equivalent b. Materials: None NOTE:

Check each installation for secure mounting and proper electrical connections.

c. Procedure NOTE:

Check each instrument.

(1) Apply 28 VDC main power to the indicator(s) and 5 VAC/VDC lamp power to the indicator(s). (2) Self test: Power-up starts a self-test. (a) Pass: The digital display cycles “0000,” “1111,” “2222,” “3333,” “4444,” “5555,” “6666,” “7777,” “8888,” and “9999” for approximately 0.5 seconds of each number for a total of 5 ± 1 seconds and the pointer resets to lowest position. (b) Fail: No pointer movement, four dashes “----” on digital display. (3) Normal operation: Within two (2 ± 1) seconds for analog display; within six (6) seconds for digital display. Indicator(s) will then display actual engine temperature(s).

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (cont’d) d. Calibration Test (1) Disconnect the pylon interface connector. NOTE: The indicator display may blank with an open in the thermocouple circuit. (2) Connect the adapter cable (connector 10SL3-P) to the aircraft side of the pylon connector. (3) Turn the EGT Test Set on and set the function selector to “RESISTANCE MEASURE” and the resistance range selector to 200 ohms (the tester display shows 1.) (4) Short the test clips together and press the “PUSH TO SET” button. Adjust the system resistance knob all the way counterclockwise until 2 ohms or less is displayed. Release the “PUSH TO SET” button. (5) Turn the function selector to Indicator Test; the display shows a 3-digit number and a degree symbol. (6) Connect the red clip to the chromel (+) (gold colored) wire and the black clip to the alumel (-) (magnetic) wire. (7) Rotate the temperature adjustment control knob for the values shown in Table 31.05-2. Record the values for the Number 1 Engine. INPUT

INDICATOR

TOLERANCE

0

± 15

500

± 15

700

± 5

900

± 5

906*

± 5

50

± 30

* Note: At 906° C the digital indication should start flashing. TEMPERATURE ADJUSTMENT CONTROL KNOB ROTATION Table 31.05-2

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (cont’d) (8) Return the system to normal. (9) Repeat the calibration test procedure for each new indicator. 4. Maintenance a. Light Tray Replacement

LIGHT TRAY REPLACEMENT Figure 31.05-2

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31-INSTRUMENTS 31.05 EXHAUST GAS TEMPERATURE INDICATOR UPGRADE (cont’d) (1) Remove access door retaining screws (1). (2) Remove access door (2) and gasket (3). (3) Remove defective light tray assembly (4). (4) Install a new light tray assembly (4). (5) Align access door (2) and gasket (3) with screw holes. (6) Install access door retaining screws (1) using Loctite #222 (or equivalent thread locker) on the threads and secure. (7) Light tray assembly replacement complete.

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31.06 – ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) TABLE OF CONTENTS Paragraph A.

GENERAL-----------------------------------------------------------------------------

Page 31.06.1

B.

DESCRIPTION -----------------------------------------------------------------------

31.06.2

C.

LABELS -------------------------------------------------------------------------------

31.06.4

D.

INSTALLATION AND CONFIGURATION OF THE BEACON IN AUTOMATIC SURVIVAL MODE (WATER ACTIVATION) ---------------

31.06.6

E.

UTILIZATION------------------------------------------------------------------------

31.06.6

F.

MAINTENANCE---------------------------------------------------------------------

31.06.9

G.

PARTS LIST -------------------------------------------------------------------------- 31.06.11

FIGURES 31.06-1

ELT ADT 406 AP DETAILED DESCRIPTION ---------------------------------

31.06.3

31.06-2

PRODUCT NAME IDENTIFICATION LABEL --------------------------------

31.06.4

31.06-3

SERIAL NUMBER IDENTIFICATION LABEL--------------------------------

31.06.4

31.06-4

BEACON CONTROL PANEL LABEL -------------------------------------------

31.06.5

31.06-5

IDENTIFICATION LABEL --------------------------------------------------------

31.06.5

31.06-6

BATTERIES LABEL ----------------------------------------------------------------

31.06.5

31.06-7

BEACON REMOVAL #1 -----------------------------------------------------------

31.06.7

31.06-8

BEACON REMOVAL #2 -----------------------------------------------------------

31.06.8

31.06-9

BEACON REMOVAL #3 -----------------------------------------------------------

31.06.8

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) CAUTION 1: THIS TRANSMITTER IS AUTHORIZED FOR USE ONLY DURING SITUATIONS OF GRAVE AND IMMINENT DANGER. CAUTION 2: IT IS IMPERATIVE THAT EACH BEACON OWNER REGISTERS THEIR BEACON. CONTACT ELTA OR APPROVED AGENT TO OBTAIN THE INFORMATION RELATIVE TO THIS REGISTRATION. CAUTION 3: THE BEACON MUST BE IMPERATIVELY PROGRAMMED WITH THE RELEVANT IDENTIFICATION AUTHORIZED BY THE LOCAL AIRWORTHINESS. CAUTION 4: ONCE THE BEACON HAS BEEN PROGRAMMED, IT IS GENERALLY ASSIGNED TO AN AIRCRAFT (NAME AND IDENTIFICATION). IF THIS BEACON IS USED ON ANOTHER AIRCRAFT, IT WOULD HAVE TO BE PROGRAMMED (NEW NAME AND IDENTIFICATION). CAUTION 5: BEFORE INSTALLING OR USING THIS EQUIPMENT, THE VALIDITY OF THE INFORMATION ON THE BEACON LABEL MUST BE IMPERATIVELY CHECKED. A. General – The ELT ADT 406S consists of a beacon that incorporates means of attachment and is designed to transmit a digital distress signal to satellites that are part of the COSPAS/SARSAT System. These satellites transmit the captured signal to the reception stations on the ground. This signal is transmitted on the 406.028 MHz frequency and is used to precisely locate and identify the ELT ADT 406S. It also transmits a 121.5 MHz and 243 MHz signal to facilitate the final approach of the distress scene (homing process). It can be manually or automatically triggered by means of a water sensor. This ELT meets the latest JAR OPS and ICAO recommendation. Any encoding protocol defined by COSPAS/SARSAT can be used with ELT ADT 406S, including country code assignment. NOTE: ELT non-water activated batteries must be replaced after 1 hour of operation or 50 percent of the life or charge of the battery as per FAR 121.339.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) B. Description - This beacon from the ADT 406 family is based from the ADT 406 AF/AP model. Consequently, this beacon features the functional capabilities of the AF/AP. In addition to its flotation gear, the attachment interface and its flexible antenna are entirely compatible with COSPAS/SARSAT. The ELT ADT 406 essentially consists of: •

ADT 406S beacon transmitter consisting of: • • • • • • •

31.06.2

Power supply module Electronic assembly Mechanical assembly Float Whip antenna Labels Water sensor/identification module

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1 2 3 4 5 6 7

Whip Antenna TX LED (indicates beacon activation) ARMED/OFF/ON Toggle Switch Front Face Float Velcro Strap Support

8 9 10 11 12 13

Back Face Metallic Strap Water Sensor/Identification Assy. Hoop Switch Guard Test Push Button Polyester Strap

ELT ADT 406 AP DETAILED DESCRIPTION Figure 31.06-1

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) C. Labels 1. Identification Label – This label indicates: • • •

Beacon name, manufacturer details Condition in which the beacon may be used (warning) Approval number, identification, part number, service bulletin, serial number, and conditions on the use of the beacon

PRODUCT NAME IDENTIFICATION LABEL Figure 31.06-2

SERIAL NUMBER IDENTIFICATION LABEL Figure 31.06-3

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) 2. Beacon Control Panel Label – This label indicates the switch use and the location of the connector on the front face of the electronic assembly.

BEACON CONTROL PANEL LABEL Figure 31.06-4 3. Identification Label – This label is fixed to the water sensor/identification module.

IDENTIFICATION LABEL Figure 31.06-5 4. Batteries Label – This label indicates the batteries’ expiration date (next battery servicing).

BATTERIES LABEL Figure 31.06-6

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31.06.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) D. Installation and Configuration of the Beacon in Automatic Survival Mode (Water Activation) CAUTION: TO BE OPERATIONAL, THE BEACON SHALL BE IN “ARMED” POSITION ONCE IT HAS BEEN CORRECTLY INSTALLED. 1. Make sure that the antenna is correctly connected to the TNC connector. 2. Switch the Beacon ARMED/OFF/ON lockable toggle switch to ARMED (pull and slide) so that the system is ready for use. The water sensor is then activated. CAUTION:

AS SOON AS THE BEACON IS ARMED, IT CAN AUTOMATICALLY ACTIVATED BY THE WATER SENSOR.

BE

E. Utilization CAUTION: IN THE EVENT OF AN UNTIMELY BEACON ACTIVATION, SHUT DOWN THE BEACON AND INFORM THE CLOSEST SEARCH AND RESCUE (SAR) OR AIRPORT CONTROL TOWER IMMEDIATELY. 1. The ELT ADT 406 can be activated automatically when the water sensor is triggered or manually by placing the beacon switch to the “ON” position. a.

Automatic Activation (1) The aircraft has to be evacuated. The beacon is removed from its support or bag and dipped into the water with the survivors. (2) The water sensor has detected sufficient water to trigger it. The indicator light (Figure 31.06-1, item 2) and the aural indicator indicate beacon activation. (3) Do not do anything, and leave the beacon in operation until the rescue team arrives.

b.

Manual Activation (1) There are two cases in which a distress signal may be triggered manually: • •

31.06.6

The water sensor has not been triggered but a distress signal must be sent (injured passengers, aircraft out of operation). The aircraft is on the ground and must be evacuated.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) (2) From the beacon: (a) Place the beacon switch (Figure 31.06-1, item 3) in the ON position (pull and slide) and do not do anything. (b) An automatic self-test sequence is performed. (c) Then the beacon is positioned in a waiting condition for about 30 seconds. This state is displayed by flashing on the indicator light (Figure 31.06-1, item 2) 1.75 seconds ON, 0.25 seconds OFF. This delay will avoid unwanted activation (false maneuver). (d) Then the indicator light and the aural indicator detect beacon activation. The actual distress signal is transmitted. This state is displayed by flashing on the indicator light: 0.5 seconds ON, 0.5 seconds OFF. CAUTION:

IN THE TWO CASES OF ACTIVATION, MAKE SURE THAT THE ANTENNA HAS DEPLOYED AND IS IN VERTICAL POSITION. FOR MAXIMUM POWER TRANSMISSION, THE ANTENNA MUST BE IN AN UNOBSTRUCTED AREA FAR FROM ANY METALLIC PIECES.

2. How to Remove the Beacon from the Bracket a.

If the aircraft has to be evacuated, follow these instructions: (1) Unfasten the retaining Velcro strap (not illustrated).

BEACON REMOVAL #1 Figure 31.06-7

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) (2) Remove the beacon from its bracket.

BEACON REMOVAL #2 Figure 31.06-8 (3) Pull firmly to break the retaining metallic strap. (4) Deploy the antenna, take the beacon with you, switch it ON, and evacuate aircraft with the beacon.

BEACON REMOVAL #3 Figure 31.06-9 31.06.8

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) b.

The best data transmission of the beacon is obtained in water floating conditions. (Don’t forget to anchor the beacon to the survivors or to the life raft with the orange lanyard attached to the beacon front panel.)

c.

It is recommended to place the beacon in an unobstructed area (avoiding transmission masks), and check regularly that the antenna is vertical. If necessary, shim the beacon vertically (e.g., in case of windswept area).

3. Beacon Shutdown – In the event of a false maneuver or untimely operation, shut down the beacon by placing the ARMED/OFF/ON toggle switch in OFF position. CAUTION: THE BEACON MUST BE SHUTDOWN BEFORE ANY MAINTENANCE OPERATIONS ARE PERFORMED. THE BEACON MUST BE COMPLETELY SHUTDOWN BEFORE IT IS REMOVED. ESTABLISH THAT ITS ARMED/OFF/ON TOGGLE SWITCH IS IN OFF POSITION.

F. Maintenance – The following test has been added to the applicable “B” and “C” Check Inspection Jobcards and should not be performed otherwise unless directed by Maintenance Control. 1. Beacon Self-test CAUTION 1:

THE SELF-TEST PROCEDURE MUST BE RESPECTED TO AVOID UNTIMELY BEACON ACTIVATION.

CAUTION 2:

THE SELF-TEST SHALL BE PERFORMED WITHIN THE FIRST FIVE MINUTES OF ANY HOUR BECAUSE THE ELT IS SENDING A 121.5 MHZ SIGNAL FOR THOSE FIVE SECONDS DURING THIS PROCESS (LIMITATION OF FALSE ALERT).

a.

The ELT ADT 406S is designed to perform a self-test by using the TEST push button when the ON/OFF/ARMED toggle switch is in ARMED position. Actual test transmission on 121.5 MHz for five seconds can be accessed on any VHF receiver.

b.

Remove the beacon from its bracket and extend the antenna before performing the self-test. Hold the beacon at arm’s length in the aisle to free the antenna from the wall.

c.

Place the beacon’s ARMED/OFF/ON toggle switch (Figure 31.06-1, item 3) in ARMED position to authorize the self-test process.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) d.

Press the TEST pushbutton (Figure 31.06-1, item 12) until the two acknowledgments blink on LED.

e.

After the two short blinks (acknowledgment) and a delay of three seconds, check that the indicator light (Figure 31.06-1, item 2) comes ON and the beacon’s buzzer sounds during six seconds.

f.

After a brief blink, the self-test report is displayed: (1) 10 seconds permanent illumination of the indicator light for correct operation (2) 20 seconds blinking condition of the indicator light for failure detection. Blinking rate indicates the failure source detection as follows: (a) 125 ms ON, 125 ms OFF (frequency 4 Hz), ELT Check Sum Failure (software problem) (b) 250 ms ON, 250 ms OFF (frequency 2 Hz), ELT power failure (UHF and/or VHF) (c) 500 ms ON, 500 ms OFF (frequency 1 Hz), signal identification missing

g.

The self-test is finished and the beacon returns to ARMED mode.

h.

Fold up the antenna, return the beacon to its bracket, and grip the strap.

2. Water Sensor Test a.

Place the ARMED/OFF/ON toggle switch (Figure 31.06-1, item 3) to ARMED, and make a short circuit between the two hexagon socket screws of the water sensor assy. This short circuit must last two to three seconds to activate the beacon.

b.

When the beacon is activated, switch OFF the beacon and return to ARMED if necessary. The water sensor works correctly.

3. Following completion of tests, fold up the antenna, return beacon to its bracket, and secure in aircraft.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 31- INSTRUMENTS 31.06 ADT 406S EMERGENCY LOCATOR TRANSMITTER (ELT) (cont’d) 4. Unit Failure or Battery Replacement – The battery has a five year life as indicated on its label (figure 31.06-6). When the battery reaches the end of its service life or if the ELT fails a self-test, the entire ADT 406S ELT shall be returned to the manufacturer for repair.

G. Parts List Part Name ADT 406S ELT Mounting Bracket Clear Protective ELT Cover

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Part Number 01N65920 04N64250 04N64265

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 32- LANDING GEAR TABLE OF CONTENTS 32.01

STRUT SEALER PRACTICES ----------------------------------------------------

Page 32.01.1

32.02

DC-8 LANDING GEAR BOGIE LUBRICATION AND INSPECTION ----

32.02.1

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GENERAL --------------------------------------------------------------------------

Page 32.01.1

2.

PROCEDURE-----------------------------------------------------------------------

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 32-LANDING GEAR 32.01 STRUT SEALER PRACTICES 1. General a. Granville aircraft strut sealer is FAA approved for use on all aircraft. The fluid swells the existing O-Rings, including any spares, installed in the strut. The amount of swelling is controlled by the amount of sealant added. b. This sealant will not stop leaks due to damaged O-Rings, piston chrome, or excessive wear of internal components of the strut. Do not expect immediate results, it will take a few days for the O-rings to swell. Leakage inspection will occur during normal routine maintenance checks. c. Do not proceed with this procedure if structural integrity of the landing gear strut is in question. Excessive wear or damage to the strut may be indicated by large amounts of fluid being expelled from the strut during push back or turns. d. Initiate a Carry-Over item in accordance with the G.M.M. referencing this procedure. Due at - next heavy maintenance visit. The affected strut should be repacked at the earliest opportunity. The repack will also include replacing the spare seals. 2. Procedure. All normal maintenance functions to be accomplished in accordance with the Douglas DC-8 maintenance manual. a. Inspect the leaking strut for structural defects in and around the leak area. If any defects are noted the strut must be repaired or replaced prior to flight. b. Deflate the leaking strut completely. c. Drain strut hydraulic fluid (MIL-H-5606) in order to replenish the fluid with the Granville fluid. d. Add Granville fluid in the following proportions: (1) (2)

DC-8 Main Strut DC-8 Nose Strut

-

2 quarts 1 quart

e. Service the strut to full with hydraulic fluid (MIL-H-5606). f. Inflate strut with dry nitrogen. g. Advise Maintenance Control and initiate Carry-Over item.

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32.02 – DC-8 LANDING GEAR BOGIE LUBRICATION AND INSPECTION TABLE OF CONTENTS Paragraph 1.

GENERAL--------------------------------------------------------------------------

Page Page 32.01.1

RECURRING NON-ROUTINE JOB CARDS Job Card 32.02-1

DC-8 LANDING GEAR BOGIE ATTACH POINT LUBRICATION-----

Page 32.02.2

32.02-2

DC-8 MAIN LANDING GEAR BOGIE INSPECTION ---------------------

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 32 – LANDING GEAR 32.02 DC-8 LANDING GEAR BOGIE LUBRICATION & INSPECTION 1. General a. Non-routine Job Cards DC8-32.02-1 “DC-8 Landing Gear Bogie Attach Point Lubrication” and DC8-32.02-2 “DC-8 Main Landing Gear Bogie Inspection” are applicable to DC-8 series aircraft with visits to Short Wavelength Roughness (SWR) Level 2 Runways, as defined in the latest revisions of Service Letters DC-8-SL-32-112. b. Accomplishment Interval (1) Accomplish Job Card DC8-32.02-1 to lubricate the pivot bolts on both main landing gear truck beams within 10 cycles before and within 10 cycles after each visit to a SWR Level 2 runway. After completion of the lubrication within 10 cycles before and after each SWR Level 2 visit, return to the standard lubrication interval. (2) Accomplish Job Card DC8-32.02-2 to inspect the pivot bolts on or before 500 SWR Level 2 Runway landings or 24 months of first SWR Lever 2 runway landing, whichever comes first, after MLG overhaul.

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ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-32.02-1 JOB TITLE

DC-8 LANDING GEAR BOGIE ATTACH POINT LUBRICATION Job Card Reviewed By:

Date:

Page 1 of 1

References Boeing Report MDC-02K9030, SCI 32-1 “DC-8 Landing Gear Bogie Attach Point Lubrication” SL DC-8-SL-32-112 Douglas DC-8-60 (or –70) MM, 12-81-0 Mechanic Inspector Procedure 1. Lubricate the Truck Beam Pivot Pin Attach Points on the Left and Right Main Landing Gear Strut and Main Landing Gear Truck per the Douglas DC-8-60 (or -70) MM, 12-81-0, as applicable. 2. Ensure all job card sign offs are complete.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 32 – LANDING GEAR 32.02 DC-8 LANDING GEAR BOGIE LUBRICATION & INSPECTION (cont’d) RECURRING NON-ROUTINE

ATI A/C NUMBER

TAT

TAC

DATE

POS

STATION

ATI-001

JOB CARD NUMBER

DC8-32.02-2 JOB TITLE

DC-8 MAIN LANDING GEAR BOGIE INSPECTION Job Card Reviewed By:

Date:

References Boeing Report MDC-02K9030, SCI 32-2 “DC-8 Main Landing Gear Bogie Inspection” SL DC-8-SL-32-112 Douglas DC-8-60 (or –70) MM, 32-11-5 Mechanic Procedure 1. Inspect the Bogie Pivot Pin and Pivot Pin Bushings:

Page 1 of 1

Inspector

a. Remove the left or right main gear bogie pivot pin for inspection per DC-8-60 (or –70) MM, 32-11-5 “Removal/Installation Main Gear bogie Beam and Axle” b. Do a detailed inspection of the bogie pivot pin and bushings: (1)

Do a detailed inspection of the outer diameter of the pivot pin. If cracks are found in the pivot pin, or if the pivot pin has bronze transfer, corrosion, or damage to the chrome plate, replace the pivot pin with a new, overhauled, or serviceable pivot pin before further flight. (2) Do a detailed inspection of the pivot pin bushings in the main landing gear truck assembly and the strut assembly. If the bushing is out of limits, replace the bushing by using Standard Overhaul Procedures Manual 20-20-07. c. Return landing gear to serviceable condition per DC-8-60 (or –70) MM, 32-11-5 “Removal/Installation Main Gear Bogie Beam and Axle”. d. Repeat steps a thru c for the other main landing gear. 2. Ensure if there is failure of inspection (cracks are found in the pivot bolt, the pivot bolt exhibits bronze transfer, corrosion, or damage to the chrome plate, or bushings out of tolerance), the failure is corrected. 3. Ensure all job card sign offs are complete.

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PASSENGER EMERGENCY ESCAPE PATH LIGHTING SYSTEM (PEEPLS)------------------------------------------------------------------------------

33.01.1

COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) -------------------------------------------------------------------------

33.02.1

GRIMES STROBE UPPER & LOWER ANTI-COLLISION LIGHTS (STC ST127CH) -----------------------------------------------------------------------------

33.03.1

MAIN CARGO DECK LIGHTS MODIFICATION (LUFTHANSA E.O. 33012/1)

33.04.1

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33.01 – PASSENGER EMERGENCY ESCAPE PATH LIGHTING SYSTEM (PEEPLS) TABLE OF CONTENTS Paragraph 1.

DESCRIPTION ---------------------------------------------------------------------

Page 33.01.1

2.

TESTING ----------------------------------------------------------------------------

33.01.1

3.

TROUBLESHOOTING -----------------------------------------------------------

33.01.1

4.

ILLUSTRATED PARTS LIST ---------------------------------------------------

33.01.3

TABLES 33.01-1

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PEEPLS TROUBLESHOOTING CHART---------------------------------------- 33.01.2

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AIR TRANSPORT INTERNATIONAL L.L.C. DC-8 MAINTENANCE MANUAL CHAPTER 33-LIGHTING 33.01 PASSENGER EMERGENCY ESCAPE PATH LIGHTING SYSTEM (PEEPLS) 1. Description - The Bruce PEEPLS consists of Exit Indicators to assist passengers in quickly locating the nearest exit during an emergency evacuation. Input power is provided by 6.0 Volt emergency power supplies charged and controlled by the existing aircraft emergency system. The floor path lighting portion of the originally installed PEEPLS has been replaced by a Photoluminescent System (reference Chapter 33.02). 2. Testing - To verify the system is in good working condition proceed as follows: a. Perform continuity test between existing emergency power supply and PEEPLS Emergency Supply per the respective wiring diagrams. b. Turn on the PEEPLS input power by activating the aircraft emergency lighting system in accordance with existing operating procedures for the particular aircraft involved. c. All Exit indicators should be fully “ON”. With fully charged batteries they will remain illuminated for a period in excess of 10 minutes. Daily testing should be of minimum duration, one minute or less, to avoid depleting batteries. NOTE:

Discharged batteries will take a minimum of 16 hours of charging by the aircraft electrical system to fully recover.

3. Troubleshooting. Refer to the following table for the troubleshooting procedures. A system that fails to perform as indicated in the testing procedure should be visually inspected. Check all electrical component and connections for possible damage. Fuse, diodes and lamps should be checked with an ohmmeter.

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AIR TRANSPORT INTERNATIONAL L.L.C. DC-8 MAINTENANCE MANUAL CHAPTER 33-LIGHTING 33.01 PASSENGER EMERGENCY ESCAPE PATH LIGHTING SYSTEM (PEEPLS) (cont’d) TROUBLE No Power (One or more outputs not working)

No Output on New Unit or Get Output for a short time (less than 10 minutes) One/Several Lamps Inoperative

CAUSE

ISOLATION PROCEDURE

REMEDY

Blown fuses

Check Battery Pack fuses with Ohmmeter (5A, 125V, P/N 276005)

Replace Fuses (Make Sure they are the correct rating)

Loose Wiring Connections at the Mailing Connector/ poor Crimp Joints

Check Crimping Joints at the Mating Sockets

Replace Sockets

Bad Diode

Check Diode With Ohmmeter

Replace Diode

Poor Crimping Joint

Check Continuity with Ohmmeter

Replace the splice with Butt Connector

Battery not fully charged

Charge Battery for minimum of 16 hours and retest. See Battery Pack C.M.M. RPS1-B for Charging Procedure

Bad Lamp

Check Lamp with Ohmmeter

Replace Lamp

Poor Connections at lamp holder contacts

Check continuity with Ohmmeter

Clean the contacts

Check continuity with Ohmmeter Exit Identifiers Inoperative

Poor Crimp Joints Poor Connections at the connector Crimping Joints (Pin/Socket)

Check continuity with Ohmmeter

Replace the splice Replace the Pin/Socket

Check Lamp with Ohmmeter

Replace Lamp

Check continuity with Ohmmeter

Replace P.C. Board

Bad Lamp Damaged Printed Circuit Board/Or Solder Joints

PEEPLS TROUBLESHOOTING CHART Table 33.01-1

33.01.2

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AIR TRANSPORT INTERNATIONAL L.L.C. DC-8 MAINTENANCE MANUAL CHAPTER 33-LIGHTING 33.01 PASSENGER EMERGENCY ESCAPE PATH LIGHTING SYSTEM (PEEPLS) (cont’d) 4. Illustrated Parts List a. Components of Exit Identifiers & Floodlights: (1)

BASE (Exit Identifier & Flood Light) 07950-() (-1 for BR9264 & -3 for BR9279) NOTE: 07950-1 Superseded by 07950-3

(2)

PRINTED CIRCUIT BOARD (Exit Identifier & Flood Light) 08030-() (-3 Pigtails & -4 with Connectors)

(3)

CONNECTOR PLUG & CONTACT SOCKET AMP P/N 35077-1 350689-2

(4)

CONNECTOR CAP & CONTACT PIN AMP P/N 350778-1 350690-2

Note: Pins & Sockets purchased separately, Require AMP tool: (1) 455830-1 Insertion Tool (2) 458994-1 Extraction Tool (3) 90300-1 Crimping Tool Revision 15

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LENS COVER (without screw hole) 08031-Series

(6)

EXIT IDENTIFIER / FLOODLIGHT BR9264 / BR9279 Series Note: BR9264 Superseded by BR9279-Series

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GUARD, EXIT IDENTIFIER 08543-01

(8) Screws, Exit Identifier Mounting MA51957-28 (2 each required per Exit Identifier) (Used to mount Exit Identifier and Flood Lights to Brackets or Structure)

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POWER SUPPLY 100865 (Radiant Illumination P/N RPS1-B)

(2)

33.01.6

Reference Chapter 24.01 for battery pack maintenance requirements.

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AIR TRANSPORT INTERNATIONAL L. L. C. DC-8 MAINTENANCE MANUAL CHAPTER 33- LIGHTING

33.02 – COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) TABLE OF CONTENTS Paragraph 1.

INTRODUCTION ------------------------------------------------------------------

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2.

MINIMUM EQUIPMENT FOR INTERIOR LIGHTING --------------------

33.02.3

3.

MAINTENANCE REQUIREMENTS-------------------------------------------

33.02.3

4.

MAINTENANCE LIMITATIONS AND INTERVALS ----------------------

33.02.4

5.

REQUIRED EQUIPMENT -------------------------------------------------------

33.02.7

6.

TRACK REMOVAL AND MAINTENANCE ---------------------------------

33.02.7

7.

CLEANING OF SECTIONS------------------------------------------------------

33.02.8

8.

OPERATING PROCEDURES ---------------------------------------------------

33.02.8

FIGURES 33.02-1

TYPICAL PHOTO LUMINESCENT INSTALLATION ----------------------- 33.02.2

33.02-2

PHOTO LUMINESCENT TRACK REMOVAL--------------------------------- 33.02.7

TABLES 33.02-1

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) NOTE: Refer to Page xxi of this manual for aircraft effectivity. 1. INTRODUCTION a. To provide a means of identifying the emergency escape path in the event of an emergency evacuation, photo luminescent strips are installed along the cabin aisles. The photo luminescent material is a strontium aluminate compound, the particles of which, are excited by light and gradually dissipate this energy by glowing. This glow provides the aisle path guidance. Upon reaching an exit, electrically operated lights and markers provide general illumination to allow for identification of the exit. The electrically operated markers are used in conjunction with the floor mounted photo luminescent strips to create a hybrid conjunction with the floor mounted photo luminescent strips to create a hybrid emergency escape path marking system. b. The strips provide a continuous floor level cue to assist the occupants in locating the exits. At each exit, an additional cue is used to guide the occupant to the proper direction of the exit opening. These additional markings are in the form of contrasting arrows or dots to indicate direction change along with an extension of the photo luminescent strip and track system at the aft exit. Several segments are butted together to create a complete system. Each photo luminescent strip is secured in a mounting and covering system. At each aft exit, lateral extensions are fitted and an indicating sign of “EXIT” with an arrow and the international sign of two dots is indicated in each track end. c. The photo luminescent PATHFINDER Emergency Aisle Path Lighting System consists of rigid photo luminescent strips (SG9422) enclosed in transparent protective casing/top track. The protective covers snap onto a bottom track with clips extruded into the cover material. The bottom tracks (SG9602) are adhered to the aircraft floor panels and are extruded with a lip that is designed to accept the internal clips of the upper covers. Application of a light pressure secures the two pieces together and sandwiches the photo luminescent strip and any markings. Figure 33.02-1 shows a typical cross-section of the installation onto the aircraft floor and general layout of aisle path marking. d. The information on the PATHFINDER Emergency Aisle Path Lighting System described in this section was compiled from the Installation Instructions and Maintenance Manual, H72-10, dated April 6, 2000; the Interior Lighting Compliance Inspection Procedures & Checklist Report, H72-12, dated April 6, 2000; the Dispatch Operating Procedures Manual, H72-13, dated April 6, 2000; and the Airplane Flight Manual Supplement H7211, dated April 6, 2000, as published by: STG Aerospace (USA) 6043 NW 167th Street, Suite A-14 Miami, FL 33015 Phone: (305) 828-9811 Fax: (305) 828-2939

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TYPICAL PHOTO LUMINESCENT INSTALLATION Figure 33.02-1

33.02.2

02-02-00

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) (cont’d) 2. MINIMUM EQUIPMENT FOR INTERIOR LIGHTING a. The effectiveness of the system requires the cabin lighting to meet prerequisite levels. The equipment required to meet these values is dependent upon the actual aircraft configuration and available light. The interior lighting requirements are established from typical interior configurations, but areas missing significant lamps, bulbs, ballasts, etc. may render the system ineffective. b. Typical interior lighting in the DC-8-62/-72 COMBI allows for the following conditions to exist: (1) Cabin and work lights should be set in the “BRIGHT“ or “ON“ position for charging prior to and during passenger boarding. (2) The sidewall lighting may be inoperative. lighting is sufficient.

Lighting provided by the overhead

(3) The overhead ceiling lighting may contain one inoperative light. The inoperative light must be clear of galley, cabinets, and life-raft storage areas. (4) Entry way lighting (overhead lighting at aft cabin doors) must operate in the “BRIGHT” setting. c. The Electrical Exit Markers may have two of the five bulbs missing or inoperative. All markers are required to be operational with at least 3 bulbs in each marker assembly functioning. 3. MAINTENANCE REQUIREMENTS a. The photo luminescent escape path system consists of two systems; a floor mounted photo luminescent track system to guide evacuating passengers to the exit and an electrically operated exit marker system at the exit to allow for evacuees to identify the exit. b. Maintenance of the floor track system includes routine inspection and simple cleaning. The floor level system requires no electrical circuits to operate and this virtually eliminates maintenance of the track system. The track system requires periodic cleaning and daily inspection to ensure it is intact. The overhead lighting system is to be maintained per the applicable maintenance manual. The cabin lighting is critical to the performance of the system and it MUST BE PROPERLY MAINTAINED. Refer to the DACO maintenance manual for procedures specific to the equipment. This system is activated prior to each flight for system charging as noted below:

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) (cont’d) (1)

Prior to the FIRST DAILY flight, the system requires an initial charging. The charging is provided by the interior cabin lighting. The charging requires all overhead ceiling, sidewall window, entry, and work lights to be in the “ON”/”BRIGHT” setting. Bright setting only applicable to switches having a “DIM” or “NIGHT” setting. The PSU reading lights are not required to be on, and the overhead bin doors should be closed. Once activated, the lights must remain on for a minimum of 45 minutes prior to extinguishing.

NOTE: During initial charging, cabin activity is limited to minor aisle traffic of crew and personnel. Passenger boarding may shadow the system during charging and IS NOT ALLOWED DURING THE REQUIRED CHARGE TIME (45 MINUTES). The cabin aisle must be clear of obstructions and overhead bin doors in the closed position for initial charge. (2)

At EACH FLIGHT, the overhead ceiling lighting is to be activated while passengers are being boarded or deplaning. Light setting is to be “ON” or “BRIGHT”. Bright setting only applicable to switches having a “DIM” or “NIGHT” setting.

c. The electrically operated exit markers and all components are to be maintained per the DACO maintenance manual. All required checks and intervals for the electrically operated exit markers specified by the DACO maintenance manual take precedent over those noted herein.

4. MAINTENANCE LIMITATIONS AND INTERVALS a. All inspections are general visual (unless otherwise noted) and limitations may not be combined to increase damage limits. However, multiple damage and types may be considered acceptable if within the constraints of the least acceptable limit. The intervals and inspections are listed below. (1) The photo luminescent elements (SG9422) are identified on the underside with the part number and a batch number. The batch number (98/2344, for example) is used to track the system elements and is used in conjunction with the SG9444 test coupon to monitor the coating performance during its service life. The test coupon is a piece of the track located in the forward cabin and measures 7.87”. A filler section is to be installed while the test coupon is submitted to STG for evaluation. (2) Current life expectancy is 10 years and is periodically monitored in accordance with the maintenance program using the SG9444 Test Coupon.

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MAINTENANCE LIMITATIONS AND INTERVALS (continued)

ITEM

INSPECTION & INTERVALS TYPE INSPECTION INTERVAL Preflight A B C 2 Years

LIMITS/ACTIONS

Missing sections or components Top Cover Photoluminenscent Elements

X

X

X

X

X

X

X

X

Test Coupon

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

Bottom Track Exit Indicators (arrows, etc.) End Indicators (dot markers)

End Caps or Ramps Missing Exit Marker Damaged Sections (cracks, stains or scrapes) All Sections Clean System Components

X

X

Missing edge flanges acceptable Less than 10% of any 48” length may be missing. Return to STG every 2 years for evaluation of elements. Less than 6” missing provided remainder is intact & top cover is secured. Replace missing segment at next “C” check. All items required. All items required in extension segments. Mid cabin aisle markers may be missing. Damage is acceptable if inner elements/inserts remain secured. All Marker assemblies required. Acceptable if less than 10% over any 48” length. Clean or replace item if greater than 10% of any 48” length. Clean per Paragraph 7 following.

INSPECTION AND INTERVALS Table 33.02-1 (1 OF 2)

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MAINTENANCE LIMITATIONS AND INTERVALS (continued)

ITEM

INSPECTION & INTERVALS TYPE INSPECTION INTERVAL Preflight A B C 2 Years

LIMITS/ACTIONS

Interior Lighting

Output level of Interior Lighting Missing Ceiling Lamps Missing Ceiling Lamps at Life Raft area Missing Sidewall Lamps

Missing bulbs in Exit Marker

X X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

Measure & record using survey procedures in STG maintenance manual to ensure 15 and 20 LUX limitations Per limitations in Paragraph 2 above. All lights must operate. Replace or repair inoperative lamps. Per limitations in Paragraph 2 above. Acceptable to have 2 of 5 bulbs missing per lamp assembly(minimum of 3 operative bulbs required.

INSPECTION AND INTERVALS Table 33.02-1 (2 OF 2)

33.02.6

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) (cont’d) 5. REQUIRED EQUIPMENT. The photo luminescent aisle path system requires the following minimum equipment to be installed. NO MORE THAN 4.8” OVER ANY 48” LENGTH TRACK SECTION MAY BE STAINED OR OTHERWISE OBSCURED. NO MARKINGS OR EXIT EXTENSIONS MAY BE OBSCURED. a. b. c. d. e. f. g. h.

SG9601 Top Cover along the main cabin aisle and exit extensions. SG9602 Bottom Track along cabin aisle and exit extensions. SG9422 Photo luminescent Insert along cabin aisle and all exit extensions. Track Extensions at aft exits. SG9342 “EXIT” marker at all extensions. End Cap at all exit extensions. Dot Markers at aft end of cabin aisle. Electrically operated exit markers at each exit with a minimum of three operating bulbs per assembly.

6. TRACK REMOVAL AND MAINTENANCE. The track system may be disassembled by removing the top cover, which is secured with clips extruded into the section. The clips are secured to mating flanges in the bottom track. A downward pressure secures the cover, and an upward pressure releases the clips. A smooth non-marring release tool (plastic) can be inserted into the end of a given top section to apply the initial upward force. Once released, a smooth tool can be slid longitudinally under the cover to remove the top cover along with the photo luminescent insert as shown in the figure below. AVOID LIFTING FROM ONE END. This could cause damage to the semi-rigid insert rendering it unserviceable.

TRACK REMOVAL Figure 33.02-2

Revision 9

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33.02.7

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) (cont’d) 7. CLEANING OF SECTIONS a. Cleaning of photo luminescent inserts (SG9422) is accomplished using mild detergent and water. AVOID the used of solvents or petroleum based products. Chemicals that may have a corrosive effect on the base metal (aluminum) should be avoided. b. Cleaning of polycarbonate elements (covers, bottom track, exit markers, and indicators) is accomplished using mild detergents and water. The use of naphtha or 1-1-1 trichloroethane is acceptable if prolonged exposure or soaking is avoided. AVOID the use of abrasive material that will scratch the polycarbonate elements. AVOID using the following elements: Alkais Amines Aromatic Hydrocarbons

Esters Ketones Halogenated Hydrocarbons

8. OPERATING PROCEDURES a. Preflight Inspection (1) Prior to commencement of flight, maintenance shall inspect the installation and ensure the required equipment is in place. The installed components are described in paragraph 5 above. Missing or inoperative equipment limitations are listed in paragraph 4 above. (2) Ensure there is no damage to the system that would allow a floor track to be displaced during normal passenger operations. This type of damage would include floor track that is not securely attached to the base and may be inadvertently removed. (3) Daily performance check for the emergency lighting system should be performed in accordance with the electrical system’s operating instructions and maintenance procedures. (4) Perform preflight charging or activate cabin lights as discussed in Paragraph ‘b’ following.

33.02.8

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 - LIGHTING 33.02 COMBI PHOTO LUMINESCENT AISLE PATH MARKING (STC STO2001AT) (cont’d) b. Preflight Charging (1) The temperature of the cabin must be stabilized between +37 & +120º Fahrenheit (+3 to +49ºC) at time of charging. Aircraft storage may be at any temperature. (2) Prior to the FIRST DAILY flight, the system requires an initial charging. The charging is provided by the interior cabin lighting. The charging requires ALL overhead ceiling, sidewall window, entry, and work lights to be in the “ON”/”BRIGHT” setting. The PSU reading lights are not required to be on, and the overhead bin doors are to be closed for the initial daily charging. Once activated, the lights must remain on for a minimum of 45 minutes prior to extinguishing. Note:

During initial charging, cabin activity is limited to minor aisle traffic of crew and personnel. Passenger boarding may shadow the system during charging and IS NOT ALLOWED DURING THE REQUIRED CHARGE TIME. The cabin aisle must be clear of obstructions for the initial charge. (3) At EACH FLIGHT, the overhead ceiling lighting is to be activated while passengers are being boarded or deplaning. Light setting is to be “ON”/”BRIGHT”. Bright setting only applicable to switches having a “DIM’ or NIGHT” setting.

c. Flight Duration (1) The flight duration is limited by the effectiveness of the system charge. This duration may be increased by providing additional charging to the system during the flight. If no in flight lighting is utilized and the cabin is operated in complete darkness (cabin lighting setting in “NIGHT” or “OFF”), the following limits apply: Night Flight Duration

6.5 Hours (Without cabin lighting being activated in flight)

(2) After 6.5 hours of night flight with complete cabin darkness, CEILING CABIN LIGHTING MUST BE ACTIVATED. Additional charging can be accomplished by allowing the interior cabin lights to remain on during flight. Activation for 15 minutes will increase flight duration an additional hour. (3) Flight duration is not limited if overhead ceiling lighting is operated in the “ON” and “BRIGHT” setting. Bright setting only applicable to switches having a “DIM” or “NIGHT” setting.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33 – LIGHTING 33.03 GRIMES STROBE UPPER AND LOWER ANTI-COLLISION LIGHTS (STC ST127CH) 1. Purpose - To provide replacement part numbers for the upper and lower anti-collision lights P/N 43325-(X) and P/N 43335-(X) respectively with the new Grimes P/N 30-1585-1 and 301586-5 upper and lower positions respectively as detailed in ATI E.O. 33-CD8-786.

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33.04 – MAIN CARGO DECK LIGHTS MODIFICATION (LUFTHANSA E.O. 33012/1) TABLE OF CONTENTS Paragraph 1.

PURPOSE ---------------------------------------------------------------------------

Page 33.04.1

2.

PART NUMBERS------------------------------------------------------------------

33.04.1

FIGURES 33.01-1

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MODIFIED LIGHT ASSEMBLY

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33-LIGHTING 33.04 MAIN CARGO DECK LIGHTS MODIFICATION (LUFTHANSA E.O. 33012/1) 1. Purpose – To provide replacement part numbers for the main deck cargo lights for Lufthansa modified DC-8-73F’s. The modified light assemblies contain two independently powered miniature halogen lamps. The lights are connected by four power circuits. If power is lost on a single bus, every second lamp on the affected side (L/H or R/H) is still lighted. (Reference WD 33-08-28.) 2. Part Numbers: a. Light Assembly: b. Lamp Holder: c. Bulb:

Revision 15

2LA007317-30 9FH408642-01 8GH005448-28

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33.04.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 33-LIGHTING 33.04 MAIN CARGO DECK LIGHTS MODIFICATION (cont’d)

MODIFIED LIGHT ASSEMBLY Figure 33.04-1 33.04.2

06-03-02

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION

TABLE OF CONTENTS 34.01

ELECTRONIC FLAT PANEL DISPLAY----------------------------------------

Page 34.01.1

34.02

ATC TRANSPONDER SYSTEM - RAMP TEST PROCEDURE ------------

34.02.1

34.03

PULSE DME SYSTEM RAMP TEST PROCEDURE--------------------------

34.03.1

34.04

RESERVED FOR FUTURE USE -------------------------------------------------

34.04.1

34.05

TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM - CAS-81

34.05.1

34.06

SUNSTRAND WINDSHEAR WARNING SYSTEM --------------------------

34.06.1

34.07

SAFE FLIGHT WINDSHEAR WARNING SYSTEM -------------------------

34.07.1

34.08

UNS-1D FLIGHT MANAGEMENT SYSTEM ---------------------------------

34.08.1

34.09

RESERVED FOR FUTURE USE -------------------------------------------------

34.09.1

34.10

STANDBY MAGNETIC COMPASS ---------------------------------------------

34.10.1

34.11

LTR-97 FOG VERTICAL GYRO/DIRECTIONAL GYRO SYSTEM-------

34.11.1

34.12

REDUCED VERTICAL SEPARATION MINIMUMS (RVSM) COMPLIANT DIGITAL AIR DATA SYSTEM --------------------------------

34.12.1

COLLINS TCAS II TRAFFIC ALERT & COLLISION AVOIDANCE SYSTEM ------------------------------------------------------------------------------

34.13.1

34.14

HONEYWELL RDR-4A WEATHER RADAR ---------------------------------

34.14.1

34.15

BFGOODRICH MODEL AI-330 STANDBY ATTITUDE INDICATOR---

34.15.1

34.16

TERRAIN AWARENESS & WARNING SYSTEM (TAWS) ----------------

34.16.1

34.17

ELECTRONIC FLIGHT INSTRUMENT (EFI) 600----------------------------

34.17.1

34.13

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34.01 – ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION TABLE OF CONTENTS Paragraph 1.

INTRODUCTION -----------------------------------------------------------------

Page 34.01.1

2.

GENERAL--------------------------------------------------------------------------

34.01.1

3.

DESCRIPTION --------------------------------------------------------------------

34.01.1

4.

SYSTEM FUNCTION ------------------------------------------------------------

34.01.2

5.

LRU LOCATIONS ----------------------------------------------------------------

34.01.3

6.

LRU FUNCTIONS ----------------------------------------------------------------

34.01.4

7.

TROUBLESHOOTING – MAINTENANCE PAGE PROCEDURES -----

34.01.9

8.

TROUBLESHOOTING PROCEDURE ----------------------------------------

34.01.11

9.

OPERATIONAL TEST PROCEDURES – LCD EADI/EHSI INSTRUMENTS -------------------------------------------------------------------

34.01.28

10.

SYSTEM TEST PROCEDURES – LCD EADI/EHSI INSTRUMENTS --

34.01.32

11.

EHSI & EADI REMOVAL-------------------------------------------------------

34.01.47

12.

EHSI & EADI INSTALLATION------------------------------------------------

34.01.48

13.

DISPLAY CONTROL PANEL REMOVAL ----------------------------------

34.01.50

14.

DISPLAY CONTROL PANEL INSTALLATION----------------------------

34.01.50

15.

DISPLAY CONTROL PANEL TEST ------------------------------------------

34.01.51

16.

PREFLIGHT DIFFERENCES ---------------------------------------------------

34.01.51

17.

ELECTRONIC FLAT PANEL INSTRUMENTS IPC------------------------

34.01.52

FIGURES 34.01-1

ELECTRONIC FLAT PANEL INSTRUMENT SYSTEM BLOCK DIAGRAM ----------------------------------------------------------------------------

34.01.3

34.01-2

EHSI DISPLAY MODES (TYPICAL) --------------------------------------------

34.01.5

34.01-3

EADI DISPLAY MODES (TYPICAL)--------------------------------------------

34.01.6

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34.01-i

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION

FIGURES (cont’d) 34.01-4

DCP-955 DISPLAY CONTROL PANEL -----------------------------------------

34.01.7

34.01-5

LRU LOCATIONS -------------------------------------------------------------------

34.01.8

34.01-6

DU STATUS PAGE ------------------------------------------------------------------ 34.01.10

34.01-7

EHSI FAILURE FLAGS & INDICATIONS-------------------------------------- 34.01.11

34.01-8

EADI FAILURE FLAGS & INDICATIONS ------------------------------------- 34.01.12

34.01-9

INDICATOR INSTALLATION ---------------------------------------------------- 34.01.49

34.01-10

ILLUSTRATED PARTS CATALOG---------------------------------------------- 34.01.53

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AIR TRANSPORT INTERNATIONAL 34.01 ELECTRONIC FLAT (S.T.C. ST00987LA-D)

PANEL

DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION INSTRUMENT INSTALLATION

1. Introduction - The Electronic Flat Panel Instruments are electronic displays that replace the existing electromechanical Attitude Direction Indicator (ADI) and Horizontal Situation Indicator (HSI) flight instruments. 2. General - The Configurable FPI-920 (Electronic Flat Panel Instruments) Flight Instrument System is primarily a replacement for existing electromechanical Attitude Direction indicator (ADI) and Horizontal Situation Indicator (HSI) flight instruments. The FPI 920 Flight Instrument System provides the flight crew with primary flight instrument displays, including pitch, roll, attitude, heading, flight director guidance, and radio navigation data (VOR/ILS), FMS, and Flight Plan Leg Display. Depending on other avionics installed in the aircraft, the FPI 920 system can also provide EFIS capabilities. 3. Description a. The Electronic Flat Panel Instruments consists of the following equipment: four Configurable Displays (FPI-920) and two Display Control Panels (DCP-955). The FPI920 is a flight deck mounted, Active Matrix/Liquid Crystal Display (AM/LCD) for both daytime and nighttime readability. b. The display unit is housed in an ARINC configurable Electronic Flat Panel Indicator size 9 inch deep case. Electrical connections are made through a circular 128 connector (J1) and a circular 55-pin (J2) connector located at the rear of the unit. The unit has an assembled aluminum chassis supporting a dot matrix liquid crystal display (LCD), a fluorescent lamp reflector, a heat exchanger and a back assembly c. A bezel surrounds the front of the flat panel LCD. The FPI-920 provides multicolor (blue, red, green, yellow, cyan, white and magenta) LCD presentation of ADI and HSI data as well as EFIS functionality. d. The Display Control Panel (DCP-955) is used to control the modes used with the FPI-920 flat panel indicator. Electrical connections are made through an ARINC 55-pin connector located at the rear of the unit. The DCP-955 is the primary control interface between the pilot and the displays. The panel controls display modes (ADI, Compacted ADI, HSI), navigation source selections, etc. through front panel switches, pushbutton switches and rotary knobs. The DCP-955 has a mix of discrete, analog DC, and ARINC 429 interfaces. The DCP interfaces to the on side displays through a low speed ARINC 429 data bus. e. In addition to the above listed components, ATI -73 aircraft also have two ARINC 429 to ARINC 407 Converters. The two ARINC 429 to ARINC 407 (X, Y, Z Syncro) converters are installed to convert pitch data for both Flight Directors. ARINC 429 HS cross-talk data is taken from both EADI’s and converted to ARINC 407 (X, Y, Z Syncro) data. The converters are programmed to convert only the pitch data, label 324. The X, Y, Z Syncro signal is fed to both Flight Director Systems, Capt EADI pitch to the no. 1 FD and F/O EADI pitch to the no. 2 FD. This pitch information is fed to the Flight Director Control panel and is used for Manual Pitch command.

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34.01.1

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 4. System Function a. The FPI-920 Configurable Displays and the DCP-955 Display Control Panel operate together to provide the required displays and controls for the flight crew. The FPI-920 supports a wide range of interfaces via the two rear connectors. The 55 pin connector (J2) generally models ADI I/O functions and the 128 pin connector (JI) models the HSI and digital port requirements. Due to the flight critical nature of the instruments, segregation rules are applied in the I/O channels. Refer to Figure 34.01-1 for a system block diagram. b. Analog synchro interfaces are provided to interface with existing attitude (pitch and roll), compass, and navigation systems (VOR, ILS). For replacement HSI functions, heading, azimuth (compass) and bearing source inputs are provided. c. Analog AC heading and course datum outputs to the existing flight director system and/or roll steering computer are provided. These datum outputs represent the error between the selected heading or course and the current heading as determined by the compass. d. Analog DC signals are provided for localizer deviation, glideslope deviation, To/From, and flags consistent with electromechanical instruments. Flag discretes are designed as differential inputs to the flexibility to adapt to existing external systems. e. The FPI-920 provides six ARINC 429 input channels and two ARINC 429 outputs. The ARINC 429 inputs are used for FMS data, TCAS data (future growth), EGPWS control data (future growth), cross-talk between instruments and DCP-955 control data. The ARINC 429 outputs are used for instrument cross-talk and EGPWS data (future growth). f. An ARINC 453 color weather radar is also provided (future growth). This input may also be used to provide EGPWS terrain data (future growth). g. The FPI-920 provides an ARINC 568 input for distance data from either DME or FMS depending on the aircraft navigation mode. h. The FPI-920 and DCP-955 are powered by 28 Vdc. Brightness of the displays is controlled through the pilot DCP-955. Brightness of the EADI Turn/Slip indicator is controlled by the individual instrument control panel. i. The FPI-920 Flight Instrument System uses built-in test (BIT) to test the operational integrity of the system. Monitoring occurs at power-up, is continuous during operation, and may be initiated manually. BIT can be manually initiated only when the Aircraft is on the ground by pressing the DU Test button on the DCP-955 control panel. Pressing the DU Test button places both onside displays into diagnostics mode.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 5. Line Replaceable Unit (LRU) – Locations - The Electronic Flat Panel Indicator Flight Instrument System is composed of LRU’s which are installed in the following locations and analytically shown in Figure 34.01-3. Line Replaceable Unit Description Captain’s EHSI & EADI First Officer’s EHSI & EADI Display Control Panels 429/407 Pitch Converter

Location Captain’s Instrument Panel First Officer’s Instrument Panel Center Pedestal Radio Rack Shelf 5 Outbd.

ELECTRONIC FLAT PANEL INSTRUMENT SYSTEM BLOCK DIAGRAM (CAPT SIDE SHOWN, F/O SAME) Figure 34.01-1

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34.01.3

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 6. Line Replaceable Unit (LRU) – Functions - The Electronic Flat Panel Flight Instrument System is composed of LRU’s which perform the following functions: NOTE: When on the ground, the screen(s) should be dimmed to prevent overheating. a. EHSI (FPI-920) (1) The EHSI presents a selectable, dynamic color display of flight progress and plan view orientation. Various display modes presented include HSI, ARC, and Flight Plan Leg Display (a function of FMS Map Mode). (Refer to Figure 34.01-2.) (2) The HSI mode presents a standard compass rose with heading orientation. Heading is supplied by the respective Compass or Flight Management System (FMS). Selected range and system source annunciation are provided with conventional VOR/ILS navigation information. FMS information may also be displayed in a standard CDI format. (3) The ARC mode presents an expanded compass rose with heading orientation. Selected range and system source annunciation is provided with conventional VOR/ILS navigation information. FMS information may also be displayed in a standard CDI format. (4) The EHSI displays Enhanced Ground Proximity Warning System (EGPWS) (future growth) terrain data in three basic colors: green, yellow, and red. Red is used for terrain threat area warnings with yellow used for caution. Reductions in threat are indicated by a change in color from red to yellow to green. Terrain information can only be displayed when the respective EHSI is in ARC or MAP mode. (5) The bearing pointers supply bearing to station or bearing to way-point information. The available bearing sources for each bearing pointer include (blank), VOR I, VOR 2, FMS 1, and FMS 2. The bearing information includes associated distances, if applicable. (6) DU TEST activates the EHSI status and self-test capability. The DU Test capability is available only on the ground. To cancel or exit the DU Test function and return to normal navigation display, press DATA SEL. The DU TEST capability and the related status and fault pages are for ground maintenance proposes only.

34.01.4

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

HSI MODE

FLIGHT PLAN LEG DISPLAY FMS MAP MODE

ARC MODE

EHSI DISPLAY MODES (TYPICAL) Figure 34.01-2 b. EADI (FPI-920) (1) The EADI presents a dynamic color display of the aircraft attitude. Attitude is shown by the cyan-colored sky and brown-colored ground moving about the aircraft symbol as the aircraft pitches and rolls. (Refer to Figure 34.01-3.) (2) The artificial horizon has scaling for pitch and roll and also has a roll attitude pointer. Flight director commands are presented as horizontal and vertical lines. (3) A vertical deviation pointer and scale show vertical deviation from the ILS glideslope or FMS generated vertical navigation. A lateral deviation pointer, scale and rising runway show lateral course deviation guidance localize, VOR or FMSgenerated paths.

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34.01.5

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (4) A Fast/Slow speed pointer shows speed targets generated by the DADC. (5) The EADI provides Navigation source annunciations and Navigation Comparator Warning Monitor annunciations. Instrument Failure, Attitude Failure, Go-around, Nav Source Failure, Radio Altitude Failure, Speed Command Failure, Glideslope Failure and Autopilot Failure flags are displayed on the EADI. Windshear Caution and Windshear Warning annunciations are displayed on the EADI. Minimum Decision Altitude (MDA) annunciations are provided as well.

COMPACT MODE

NORMAL MODE

EADI DISPLAY MODES (TYPICAL) Figure 34.01-3 c. DCP-955 – There are two DCP’s, one for each flight instrument system. Each pilot (Capt and F/O) has separate control of his/her system. The functions of the individual controls are as follows. (Refer to Figure 34.01-4.) (1) BRIGHT/DM Knob - Controls the brightness adjustment for both HSI and ADI. Clockwise rotation increases the brightness. For aircraft having both the EADI and EHSI, separate knobs adjust the brightness of each display. (2) COMPACT Switch - Toggles the ADI full-view display and the ADI/HSI mode. The ADI/HSI mode includes a combined ADI and HSI presentation on the ADI display. (3) TERRAIN Switch (Future Growth) - Selects the ground proximity terrain overlay when the ARC-HSI display mode is selected. Currently inactive.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (4) DU (Display Unit) TEST Switch - Selects the Display Unit (DU) status page on both the ADI and HSI. The DU status page shows software part numbers and results of internal built-in tests (IBIT). After the DU status page is displayed, reselection for two seconds or more will initiate the built-in test (BIT) function for both the ADI and HSI. Use NAV DATA Switch to exit DU Test. (5) CRS (Course) Knob [Inner knob] - Adjusts the selected course numeric readout and corresponding pointer on the HSI when a radio is the navigation source. (6) MODE Knob [Outer knob] - Selects the HSI display between HSI (full 360º compass) and ARC-HSI (partial 80º compass) mode. (7) HDG (Heading) Knob [Inner knob] - Selects the position of the heading bug on the compass card and the heading in the selected heading numeric readout. This function is available for both full and partial compass displays. (8) RNG (Range) Knob [Outer knob] - Changes the full-scale range on the ARC-MAP (80 degree compass with map overlay) display mode. The MAP ranges are 5, 10, 20, 40, 80, 160 and 320 nautical miles. (Visible range ring on display is ½ scale or 2.5, 5, 10, 20, 40, 80, and 160 nautical miles.) The map range is annunciated on the outer full-scale ring of the plan map. Half-scale annunciations are on the inner ring of the plan map and along side the range arc (or present track line) on the map display. (9) NAV DATA Switch – Selects different FMS navigation data for display on the HSI. The data available depends on the FMS. The type of information displayed toggles between (blank), GSPD, TAS, CRS (VHF Nav) or DTK (FMS), FMS (Distance) and TTG.

DCP-955 DISPLAY CONTROL PANEL Figure 34.01-4

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34.01.7

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

LRU LOCATIONS Figure 34.01-5

34.01.8

12-02-02

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 7. Troubleshooting – Maintenance Page Procedures - The Electronic Flat Panel instrument incorporates self-monitoring and self-test capabilities. Access to the DU maintenance capabilities is defined by this section. a. The DU maintenance pages are organized as shown in the following table. The pages are viewed in forward sequence only (next page). Once the last page is viewed, the paging sequence begins again and may be repeated as many times as is required: DU Maint. Pages STATUS & FAULT DU Status (**)

DU BIT Fault Tally (*) DU BIT Fault (*) DETAIL Discrete Outputs

Page(s)

Maint Tests

2

IBIT

1-2 1

(*) (*)

1

CMD

Description & Function DU hardware and software status. DU status displayed. IBIT self-test run from this page. DU BIT fault log. Pass/fail log. (*) I/O Synchro board temperatures displayed. (*)

Discrete output details. Discrete output command self-test run from this page. Discrete Inputs 1-2 Discrete input details. AC Outputs 1-2 CMD AC output details. AC output command self-test run from this page. Io2 AC Inputs 1-3 AC input details. DC Outputs 1 CMD DC output details. DC output command self test run from this page. DC Inputs 1-2 DC input details. ARINC 429 Outputs 1 CMD ARINC 429 output details. ARINC 429 output command self-test run from this page. ARINC 429 Inputs 1 ARINC 429 input details. ARINC 419 Inputs 1 ARINC 419 input details. ARINC 453 Inputs 1 ARINC 453 input details. NOTE(*): DU Fault Talley and Fault pages contain a fault log for the specific unit and are for Factory or Service Center use only. The pass/fail status on these pages may only be reset by Factory or Service Center procedures. NOTE (**):

Revision 15

Once the page sequence has been completed, the DU status will show the following items in yellow: SYNCHROS N/A, AC OUT N/A, DC OUT N/A, and DISCRETE OUT N/A. To reset the above items, return or exit to the normal display mode by pressing NAV DATA on the DCP and re-enter the test by pressing DU TEST. The DU STATUS page will be displayed with the above items reset.

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34.01.9

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) b. The DU TEST button provides access to the DU STATUS page. The DU STATUS page shows software part numbers and results of internal built-in tests (IBIT). (The DU STATUS page can only be accessed on the ground. Pushing the DU TEST button when the aircraft is airborne has no effect on the display.)

DU STATUS PAGE Figure 34.01-6 c. Access to the DU maintenance pages is accomplished from the DCP as follows: (1) Press the DU TEST button to access the DU STATUS page. (2) Press and hold the DU TEST button for at least two seconds. The “PRESS TEST TO CONTINUE IBIT” prompt comes on for five seconds. (3) While the prompt is being displayed, press and hold the DU TEST button for at least two seconds. This starts the internal built-in test (IBIT). (4) When the internal build-in test finishes, the DU STATUS page shows the test results, which are color-coded. Items that passed are white and items that failed are red. Items that show in yellow indicate the system is not monitoring signals from those items. To restore monitoring, exit the DU STATUS page, and then enter the DU STATUS page again. (5) To exit the DU STATUS page, press and release NAV DATA switch. The button to exit the DU STATUS page is identified on the upper right part of the DU STATUS page. (a) Exits maintenance pages and returns to normal flight display at any point in the sequence. Upon exiting the DU STATUS page, it takes 7 to 8 seconds for the display to return to the normal operating mode. 34.01.10

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) NOTE: If any fault table directs reference to data on a specific detail page(s), access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detailed tests may be accomplished to function specific check items. Follow the instructions at the bottom of each applicable detail page(s). 8. Troubleshooting Procedure - If any failure condition, message or flag is displayed during the Operational or System Test(s) accomplish the following troubleshooting procedure. a. EHSI Failure Flags and Indications

EHSI FAILURE FLAGS AND INDICATIONS Figure 34.01-7 Refer to the Rockwell/Collins Air Transport Division, FPI-900/910/920, Pilot Guide, Document No. 523-0780537-002117, dated February 12, 2001 (or later appropriate revision) for additional description and operating details.

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34.01.11

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) b. EADI Failure Flags & Indications

EADI FAILURE FLAGS AND INDICATIONS Figure 34.01-8 Refer to the Rockwell/Collins Air Transport Division, FPI-900/910/920, Pilot Guide, Document No. 523-0780537-002117, dated February 12, 2001 (or later appropriate revision) for additional description and operating details.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) c. If flags are displayed, refer to the following table:

DU Flag (Red) DU/DCP DU DCP

INTERFACE HDG

Revision 15

DU Flag Table Cause of Malfunction

Recommended Corrective Action

Display Internal Failure Display Control Panel Communications Failure

Replace display Check the applicable DCP associated wiring, or replace DCP or replace display

Loss of valid Heading Information

Check the applicable interface equipment, AC references and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment, AC references and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Not Used

LATERAL NAV SOURCE (VOR, ILS, FMS) VERTICAL NAV SOURCE (G/S, FMS) ATTITUDE

Loss of valid Lateral Navigation and Distance Information Loss of valid Vertical Navigation Information

RADIO ALTITUDE

Loss of valid Radio Altitude Information

FLIGHT DIRECTOR

Loss of valid Flight Director Information

SPEED COMMAND

Loss of valid Speed Command Information

FLAG

Fail flag displayed on EADI/EHSI

AUTPILOT

Loss of valid Autopilot Information

AUTOTHROTTLE INSTRUMENT 1 INSTRUMENT 2

Loss of valid Autothrottle Information Instrument failure on Captain side (1) or First Officer (2) side

Loss of valid Attitude Reference Information

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Check the applicable interface equipment and associated wiring or replace display

34.01.13

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) DU Flag Table (cont’d) Cause of Malfunction Recommended Corrective Action Go-around Instrument Check the applicable interface failure on Captain side equipment and associated wiring or (1) or First Officer (2) replace display side NOTE: The IBIT and Detail Maintenance information contained in the DU Test pages may be used to further identify the cause of display interface fault(s). DU Flag (Red) GO-AROUND 1 GO-AROUND 2

d. Do these steps on the applicable DCP: (1) Push the DU TEST button on DCP. (2) Make sure the DU STATUS page shows on the EADI/EHSI. (3) Push and hold the DU TEST button on the DCP for at least two seconds. When the “PRESS TEST TO CONTINUE IBIT” prompt appears, release the DU TEST button. (4) While the “PRESS TEST TO CONTINUE IBIT” prompt is being displayed, push and hold the DU TEST button for at least two seconds. (This starts the built-in test. The display will blank momentarily and then display the results. When the BIT is finished, the DU STATUS page is again displayed. Items displayed in white passed and those in red failed. Items that show in yellow indicate the system is not monitoring signals from those items. To restore monitoring, exit the DU STATUS page and then enter the DU STATUS page again.) (5) Make sure the DU STATUS page on the EADI/EHSI shows the following BIT results: (white text = passed, red text = failed) (a) All PROCESSOR status items displayed in white (Pass). (b) All DISPLAY status items displayed in white (Pass). (c) All INTERFACE status items displayed in white (Pass). (6) Refer to the following DU fault table for corrective action of failed items. NOTE: If any fault table directs reference to data on a specific detail page(s), access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detail tests may be accomplished to function specific detail items. Follow the instructions at the bottom of each applicable detail page(s). DU Status Fault Table DU Status Message (Red) PROCESSOR CPU RAM MEM

34.01.14

Cause of Malfunction

Recommended Corrective Action

Processor Failure RAM checksum or read/write failure

Replace display Replace display

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) DU Status Fault Table (cont’d) DU Status Message (Red) EPROM EEPROM DISPLAY GRAPHIC LCD DRIVER LAMP INTERFACE AC REFERENCE

SYNCHROS

A/D CONVERTER AC OUT

Cause of Malfunction

Recommended Corrective Action

EPROM Failure EEPROM Failure

Replace display Replace display

Graphics Card LCD Driver Failure Lamp Failure

Replace display failure Replace display Replace display

26 VAC reference Failure (AC REF 1, 2 & 3)

Refer to the AC INPUTS detail page(s), identify the failed 26 VAC input(s) (AC REF 1, 2 & 3), check the applicable input source and associated wiring or replace display Refer to the AC INPUTS detail page(s), identify the failed Synchro input(s), check the applicable Synchro interface equipment and associated wiring or replace display Replace display

Synchro Failure a) Synchro Failure b) 26 VAC reference failure (AC REF 1, 2 & 3) Analog to Digital Converter Failure a) 26 VAC ref Failure (AC REF 4) b) AC Output Failure

Refer to the AC INPUTS detail page(s), identify the failed 26 VAC input(s) (AC REF 4), check the applicable input source and associated wiring or replace display DC OUT DC Output failure Refer to the DC OUTPUTS detail page(s), identify the failed DC output(s), check the applicable DC output interface equipment and associated wiring or replace display DISCRETE OUT Discrete Output Failure Refer to the DISCRETE OUTPUTS detail page(s), identify the failed discrete output(s), check the applicable discrete output inter-face equipment and associated wiring or replace display NOTE: Once the page sequence has been completed, the DU status will show the following items in yellow: SYNCHROS N/A, AC OUT N/A, DC OUT N/A, and DISCRETE OUT N/A. To reset the above items, return or exit to the normal display mode by pressing NAV DATA on the DCP and re-enter the test by pressing DU TEST. The DU STATUS page will be displayed with the above items reset. Revision 15

06-03-02

34.01.15

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (7) Make sure the DU STATUS page on each EHSI shows the following hardware and software data: (a) (b) (c) (d) (e) (f)

TYPE: OFP: FLXD: CHAR: SEL OPC: #1: CFG:

FPI-920 091-2006-003 091-1932-002 091-2l’) 5-003 EMERY DC-8 091-2007-Ofl

(8) Make sure no configuration messages show on the lower part of the DU STATUS page of the EHSI. (9) Refer to the following DU Configuration fault table for corrective action of configuration messages. DU Configuration Fault Table DU Configuration Fault Message CONFIGURATION NO CFG LOADED (Black Text on Red) CFG INVALID (Black Text on Red) UNUSED OPC SEL (White Text) TYPE INVALID TYPE CODE (Black Text on Red) INVALID TYPE (Black Text on Red) X COMP UNABLE CONFIG X COMP (Black Text on Yellow)

Cause of Malfunction

Recommended Corrective Action

No OPC Loaded

Replace display

OPC Checksum invalid

Replace display

Incorrect strapping

Check associated indicator strap wiring or replace display

Incorrect type code file loaded

Replace display

No type code file loaded

Replace display

Refer to the ARINC 429 INPUTS detail page(s), check for labels and activity on Cross-side ARINC input, check the applicable cross-side equipment, cross-side EADI, EHSI, DCP, C/B and associated wiring or replace the appropriate display CONFIG MISCOMP X-side configuration Check associated and x-side (Black Text on Red) checksum indicator strap wiring or replace the miscomparison appropriate display NOTE: The DU does not contain field loadable software. If software errors are suspected, replace the display.

34.01.16

No x-side data received

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (10) Press the DU test switch, as required, to show the desired detail page(s) for specific input/output functions. NOTE: DU BIT Fault Tally and DU BIT Fault pages contain a fault log for the specific unit and are for Factory or Service Center use only. The pass/fail status on these pages may only be reset by Factory or Service Center procedures. (11) If Discrete Output information suspected as faulty or additional interface details required, refer to the DISCRETE OUTPUTS fault table and make sure each applicable discrete output operates correctly. Discrete Output Fault Table Cause of Function Malfunction

Discrete Output (Pin-out) NOT USED DISCOUT 01 Open = N/A (JI-108) Gnd = N/A NOT USED DISCOUT 02 Open = N/A (J1-119) Gnd = N/A VERT DEV VALID OUTPUT DISCOUT 03 Open = Invalid (J1-118) 28VDC = Valid

LAT DEV VALID OUTPUT DISCOUT 04 Open = Invalid (J1-117) 28VDC = Valid

Recommended Corrective Action

Not used

Not used

Not used

Not used

No or incorrect VERT DEV Valid out

Check the applicable interface equipment and associated wiring or replace display

No or incorrect LAT DEV Valid out

Check the applicable interface equipment and associated wiring or replace display

LOC TUNED DISCOUT 05 Open = VOR Not used Not used (J1-116) 28VDC = ILS NOTE: Access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detail tests may be accomplished to function specific detail items. Follow the instructions at the bottom of each applicable detail page(s). The outputs may be toggled between “Open” and “Active”. Active is “Gnd” or “28VDC” as applicable with initial state being “Open”. (12) If Discrete Input information suspected as faulty or additional interface details required, refer to the DISCRETE INPUTS fault table and make sure each applicable discrete input operates correctly. Revision 15

06-03-02

34.01.17

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

Discrete Input (Pin-out) lo1: (Page 1) AUX ATT Select DISCIN 01 (J2-024) (ADI Only) WNDSHR WARN DISCIN 02 (J2-033) (ADI Only) FD VALID DISCIN 03 (J2-023) (ADI Only) HDG BOTH ON 1 DISCIN 04 (J2-029)

Discrete Input Fault Table Cause of Function Malfunction

Recommended Corrective Action

Open = Norm Gnd = Aux

No or incorrect Aux Attitude Selected input

Open = Norm Gnd = Warn

No or incorrect Windshear Warning input

Open = Invalid Gnd = Valid

No or incorrect Flight Director input

Open = Normal Gnd = Both on 1

No or incorrect heading select input

Not used DISCIN 05 (J2-022) Not used DISCIN 06 (J2-028) WNDSHR CAUT DISCIN 07 (J2-026) (ADI Only) DME HOLD DISCIN 08 (J2-027) HDG BOTH ON 2 DISCIN 09 (J2-021)

Open = N/A Gnd = N/A Open = N/A Gnd = N/A Open = Normal Gnd = Caution

Not used

Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display Not used

Not used

Not used

No or incorrect Windshear Caution input

Open = Normal Gnd = Hold

Not used, Deactivated

Check the applicable interface equipment and associated wiring or replace display Not used, Deactivated

Open = Norm Gnd = Both on 2

No or incorrect heading select input

AT FAIL DISCIN 10 (J2-016) (ADI Only) MDA DISCIN 11 (J2-020) (ADI Only)

Open = Norm Gnd = Fail

Not used, Deactivated

Open = Off Gnd = On

No or incorrect Minimum Decision Alt. Input

34.01.18

12-02-02

Check the applicable interface equipment and associated wiring or replace display Not used, Deactivated

Check the applicable interface equipment and associated wiring or replace display

Revision 16

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) Discrete Input Fault Table (cont’d) Discrete Input Cause of Recommended Corrective Function (Pin-out) Malfunction Action AP FAIL DISCIN Open = Normal No or incorrect Check the applicable interface 12 (J2-015) Gnd = Fail Autopilot fail equipment and associated (ADI Only) input wiring or replace display CWM RESET Open = Normal No or incorrect Check the applicable interface DISCIN 13 Gnd = Reset Comparator equipment and associated (J2-019) Warn. Reset wiring or replace display (ADI Only) input VOR/LOC VALID Open = Invalid No or incorrect Check the applicable interface DISCIN 14 28VDC = Valid VOR/LOC equipment and associated (J2-014) Valid input wiring or replace display FD OFF DISCIN Open = On No or incorrect Check the applicable interface 15 (J2-009) 28VDC = Off Flight Dir off equipment and associated (ADI Only) input wiring or replace display lo1: (Page 2) SPD COM VALID Open = Invalid No or incorrect Check the applicable interface DISCIN 16 28VDC = Valid Speed equipment and associated (J2-013) Command input wiring or replace display (ADI Only) ATT VALID Open = Invalid No or incorrect Check the applicable interface DISCIN 17 28VDC = Valid Attitude Valid equipment and associated (J2-008) input wiring or replace display (ADI only) RAD/FMS SEL Open = Radio No or incorrect Check the applicable interface DISCIN 18 28VDC = FMS Radio Nav/FMS equipment and associated (J2-012) Nav input wiring or replace display DME VALID Open = Invalid No or incorrect Check the applicable interface DISCIN 19 28VDC = Valid DME Valid equipment and associated (J2-007) input wiring or replace display RAD ALT VALID Open = Invalid No or incorrect Check the applicable interface DISCIN 20 28VDC = Valid Radio Alt. equipment and associated (J2-011) Valid input wiring or replace display (ADI only) CMD BAR OUT Open = In view No or incorrect Check the applicable interface OF VIEW DISCIN 28VDC = Out Command Bar equipment and associated 21 (J2-010) of view Bias input wiring or replace display lo2: (ADI only)

Revision 15

06-03-02

34.01.19

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) Discrete Input (Pin-out) lo2: (Page 1) OPC LOAD ENABLE DISCIN 01 (J1-087)

Discrete Input Fault Table (cont’d) Cause of Function Malfunction Open = Normal Gnd = Load Enable (*) Open = Air Gnd = Gnd (*)

Recommended Corrective Action

No or incorrect Data Load Enable

Check the applicable interface equipment and associated wiring or replace display Weight on Wheels No or incorrect Check the applicable DISCIN 02 Air/Gnd input interface equipment and (J1-092) associated wiring or replace display GS VALID Open = Invalid No or incorrect Check the applicable DISCIN 03 28VDC = Valid Glideslope Valid interface equipment and (J1-098) input associated wiring or replace display HDG VALID Open = Invalid No or incorrect Check the applicable DISCIN 04 28VDC = Valid Heading Valid interface equipment and (J1-103) input associated wiring or replace display the title on the detail page(s). The associated input pin and the state of the discrete input may be monitored on these page(s). NOTE: Access the applicable page(s) using the DU test switch. (See instructions above.) The associated input pin and the state of the discrete input may be monitored on these page(s). NOTE (*): These discrete inputs are permanently assigned. (13) If AC Output information suspected as faulty or additional interface details required, refer to the AC OUTPUTS fault table and make sure each applicable AC output operates correctly. AC Output (Pin-out) lo1: (Page 1) NOT USED AC_OUT 1 (P: J2-042) AC_OUT 1 (N: J2-043) NOT USED WARNING 1 (J1-109) WARNING 2 (J1-111) 34.01.20

AC Output Fault Table Cause of Function Malfunction

Recommended Corrective Action

N/A N/A

Not Used

Not Used

N/A N/A

Not Used

Not Used

12-02-02

Revision 16

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) AC Output Fault Table (cont’d) AC Output Cause of Recommended Function (Pin-out) Malfunction Corrective Action lo2: (Page 2) PITCH BOOTSTRAP X AC-OUT 1 N/A Not Used Not Used (J1-075) (ADI Only) PITCH BOOTSTRAP Y AC-OUT 2 N/A Not Used Not Used (J1-076) (ADI Only) CRS DATUM AC_OUT 3 Course Deviation Incorrect Deviation Check the applicable (P: J1-077) output Output interface equipment and AC_OUT 3 associated wiring or (N: J1-078) replace display HDG DATUM AC_OUT 4 Heading Incorrect Deviation Check the applicable (P: J1-079) Deviation output Output interface equipment and AC_OUT 4 associated wiring or (N: J1-080) replace display NOTE: Access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detail tests may be accomplished to function specific detail items. Follow the instructions at the bottom of each applicable detail page(s). The outputs may be toggled between “Minimum,” “Middle,” and “Maximum”. The initial state is “Minimum”. (14) If AC Input information suspected as faulty or additional interface details required, refer to the AC INPUTS fault table and make sure each applicable AC input operates correctly.

AC Input (Pin-out) HEADING AC_IN_12_01 (J1-048) AC_IN_12_02 (J1-049) AC_IN_12_03 (J1-050)

Revision 16

AC Input Fault Table Cause of Function Malfunction SYNCHRO XYZ No or incorrect heading input

12-02-02

Recommended Corrective Action Check 26 VAC Ref 1, check the applicable interface equipment and associated wiring or replace display

34.01.21

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) AC Input (Pin-out) PITCH ATT AC_IN_12_04 (J1-051) AC_IN_12_05 (J1-052) AC_IN_12_06 (J1-053) (ADI Only) VOR BEARING AC_IN_12_07 (J1-054) AC_IN_12_08 (J1-044) AC_IN_12_09 (J1-043) (ADI Only) ROLL ATT AC_IN_12_10 (J1-045) AC_IN_12_11 (J1-046) AC_IN_12_12 (J1-056) (ADI Only) NOT USED AC_IN_12_13 (J1-055) AC_IN_12_14 (J1-057) AC_IN_12_15 (J1-063) CAPT ON NAV 2 AC_IW_12_16 (J1-067)

GA 2 FAIL AC_IN_12_17 (J1-059) (ADI Only)

INST 1 FAIL AC_IN_12_18 (J1-060) (ADI Only)

GA 1 FAIL AC_IN_12_19 (J1-061) (ADI Only)

34.01.22

AC Input Fault Table (cont’d) Cause of Function Malfunction SYNCHRO XYZ No or incorrect Pitch Attitude input SYNCHRO XYZ

SYNCHRO XYZ

SYNCHRO XYZ

DISCRETE Open = Normal 28VDC = Capt Nav 2 DISCRETE Open = Normal 28VDC = Fail DISCRETE Open = Normal 28VDC = Fail DISCRETE Open = Normal 28VDC = Fail

Recommended Corrective Action Check 26 VAC Ref 2, check the applicable interface equipment and associated wiring or replace display

No or incorrect VOR Bearing input

Check 26 VAC Ref 3, check the applicable interface equipment and associated wiring or replace display

No or incorrect Roll Attitude input

Check 26 VAC Ref 2, check the applicable interface equipment and associated wiring or replace display

Not Used

Not Used

No or incorrect Capt On Nav 2 input

Check the applicable interface equipment and associated wiring or replace display

No or incorrect Go-around fail input

Check the applicable interface equipment and associated wiring or replace display

No or incorrect instrument fail input

Check the applicable interface equipment and associated wiring or replace display

No or incorrect Go-around fail input

Check the applicable interface equipment and associated wiring or replace display

12-02-02

Revision 16

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

AC Input (Pin-out) INST 2 FAIL AC_IN_12_20 (J1-064)

LOC TUNED AC_IN_12_21 (J1-065)

AC Input Fault Table (cont’d) Cause of Function Malfunction DISCRETE Open = Normal No or incorrect 28VDC = Fail instrument fail input DISCRETE Open = Normal 28VDC = Loc Tuned

No or incorrect Localizer energized input

Recommended Corrective Action Check the applicable interface equipment and associated wiring or replace display Check the applicable interface equipment and associated wiring or replace display

NOT USED DISCRETE AC_IN_12AC_IN_22 Not Used Not Used Not Used (J1-069) AC_IN_12_23 (J1-071) Not Used Not Used Not Used AC_IN_12_24 (J1-073) Not Used Not Used Not Used AC_IN_12_25 (J1-068) Not Used Not Used Not Used NOTE: A/C inputs ACIN 16 through 25 may be used as Open/28VDC discrete inputs. (The value displayed will be 0.0 VDC for Open discrete inputs and 12.0 VDC for 28 VDC discrete inputs.) AC REF HDG REF AC_REF 1 26 VAC Ref 1 Fail = No or Check the applicable (J1-126) incorrect 26 source and associated VAC Reference wiring or replace display ATT REF (EADI) 26 VAC Ref 2 Fail = No or Check the applicable AC_REF 2 (J1-062) incorrect 26 source and associated (EFTSI) Jump to Ref 3 VAC Reference wiring or replace display BEARING REF 26 VAC Ref 3 Fail = No or Check the applicable AC_REF 3 (J1-066) incorrect 26 source and associated VAC Reference wiring or replace display ATT REF (EADI) 26 VAC Ref 4 Fail = No or Check the applicable AC_REF 4 (J1-074) incorrect 26 source and associated (EHSI) Jump to Ref 3 VAC Reference wiring or replace display NOTE: Access the applicable page(s) using the DU test switch. (See instructions above the title on the detail page(s).) The associated input pin(s) and the AC Synchro and reference voltages may be monitored on these page(s).

Revision 16

12-02-02

34.01.23

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (15) If DC Output information suspected as faulty or additional interface details required, refer to the DC OUTPUTS fault table and make sure each applicable DC output operates correctly. DC Output Fault Table Cause of Function Malfunction

DC Output (Pin-out) lo1: VERT DEV DC-OUT 1 (J1-121)

lo2: LATDEV DC-OUT 1 (J1-082)

Vertical Deviation output

No or incorrect Vertical Deviation output

Recommended Corrective Action

Check the applicable interface equipment and associated wiring or replace display

Lateral Deviation output

No or incorrect Check the applicable Lateral interface equipment Deviation and associated wiring output or replace display NOTE: Access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detail tests may be accomplished to function specific detail items. Follow the instructions at the bottom of each applicable detail page(s). The outputs may be toggled between “Minimum,” “Middle,” and “Maximum”. The initial state is “Minimum”. (16) If DC Input information suspected as faulty or additional interface details required, refer to the DC INPUTS fault table and make sure each applicable DC input operates correctly.

DC Input (Pin-out) lo1: (Page 1) NOT USED DC_IN 1 (P: J2-035) DC_IN 1 (N: J2-034) FD ROLL STRG DC_IN 2 (P: J2-039) DC_IN 2 (N: J2-038) (ADI Only) NOT USED DC_IN 3 (P: J2-041) DC_IN 3 (N: J2-040) DC_IN 4 (P: J1-124) DC_IN 4 (N: J1-125) 34.01.24

DC Input Fault Table Cause of Function Malfunction

Recommended Corrective Action

N/A N/A

Not Used

Not Used

+ Right + Left

No or incorrect Roll Steering input

Check the applicable interface equipment and associated wiring or replace display

N/A N/A N/A N/A

Not Used

Not Used

Not Used

Not Used

12-02-02

Revision 16

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) DC Input Fault Table (cont’d) DC Input (Pin-out) DC_IN 5 (P: J1-122) DC_IN 5 (N: J1-123) TO/FROM DC_IN 6 (P: J2-036) DC_IN 6 (N: J2-037) FD PITCH STRG DC_IN 7 (P: J2-030) DC_IN 7 (N: J2-031) (ADI Only) RAD ALT RAD ALT LOW (J2-018) RAD ALT HIGH (J2-017) (ADI Only) lo2: (Page 2) Glideslope Dev (LL) DC_IN 1 (P: J1-091) DC_IN 1 (N: J1-090) (ADI Only) SPD DEV DC_IN 2 (P: J1-091) DC_IN 2 (N: J1-090) (ADI Only) NOT USED DC_IN 3 (P: J1-097) DC_IN 3 (N: J1-096) DC_IN 4 (P: J1-102) DC_IN 4 (N: J1-101) Lateral Dev (LL) DC_IN 5 (P: J1-084) DC_IN 5 (N: J1-083)

Cause of Malfunction Not Used

Recommended Corrective Action Not Used

+ To + From

No or incorrect To/From input

Check the applicable interface equipment and associated wiring or replace display

+ Up + Down

No or incorrect Pitch Steering input

Check the applicable interface equipment and associated wiring or replace display

Radio Altitude

No or incorrect Radio Altimeter input

Check the applicable interface equipment and associated wiring or replace display

+ Up + Dn

No or incorrect Glideslope Deviation input

Check the applicable interface equipment and associated wiring or replace display

+ Fast + Slow

No or incorrect Speed Command Deviation input

Check the applicable interface equipment and associated wiring or replace display

N/A N/A N/A N/A

Not Used

Not Used

Not Used

Not Used

+Rt +Lt

No or incorrect Lateral Deviation input

Check the applicable interface equipment and associated wiring or replace display

Function N/A N/A

NOT USED DC_IN 6 (P: J1-089) N/A Not Used Not Used DC_IN 6 (N: J1-088) N/A DC_IN 7 (P: J1-095) N/A Not Used Not Used DC_IN 7 (N: J1-094) N/A DC_IN 8 (P: J1-100) N/A Not Used Not Used DC_IN 8 (N: J1-099) N/A NOTE: Access the applicable page(s) using the DU test switch. (See instructions above the title on the detail page(s).) The associated input pin(s) and the DC input voltages may be monitored on these page(s).

Revision 16

12-02-02

34.01.25

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (17) If ARINC 429 Output information suspected as faulty or additional interface details required, refer to the ARINC429 OUTPUTS fault table and make sure each applicable data-bus has labels and activity. ARINC429 Output (Pin-out) On-side/Cross-side Data Bus 1A (J1-039) Data Bus 1B (J1-038)

ARINC429 Output Fault Table Cause of Function Malfunction ARINC 429 data out (Bus 1)

Recommended Corrective Action

No or incorrect Check the applicable ARINC429 data interface equipment output and associated wiring or replace display

NOT USED Data Bus 2A (J1-041) ARINC429 data out Not Used Not Used Data Bus 2B (J1-040) (Bus 2) NOTE: Access the applicable page by pressing the DU test switch. (See instructions above the title on the DU status, fault and detail page(s).) Detail tests may be accomplished to function specific detail items. Follow the instructions at the bottom of each applicable detail page(s). (18) If ARINC429 Input information suspected as faulty or additional interface details required, refer to the ARINC429 INPUTS fault table and make sure each applicable data-bus has labels and activity. ARINC429 Input Fault Table Cause of Function Malfunction

Recommended Corrective Action

ARINC429 Input (Pin-out) FMS Data Bus 1A (J1-026) Data Bus 1B (J1-025)

ARINC429 data in (Bus 1)

No or incorrect Check the applicable ARINC429 data interface equipment input and associated wiring or replace display

CROSS-SIDE Data Bus 2A (J1-028) Data Bus 2B (J1-027)

ARINC429 data in (Bus 2)

No or incorrect Check the applicable ARINC429 data interface equipment input and associated wiring or replace display

ONSIDE Data Bus 3A (J1-030) Data Bus 3B (J1-029)

ARINC429 data in (Bus 3)

No or incorrect Check the applicable ARINC429 data interface equipment input and associated wiring or replace display

34.01.26

06-03-02

Revision 15

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) ARINC429 Input Fault Table (cont’d) ARINC429 Input Cause of Recommended Corrective Function (Pin-out) Malfunction Action TCAS Data Bus 4A (J1-032) ARINC429 data Not Used Not Used Data Bus 4B (J1-031) in (Bus 4) DCP Data Bus 5A (J1-034) ARINC429 data No or incorrect Check the applicable interface Data Bus 5B (J1-033) in ARINC429 data equipment and associated (Bus 5) input wiring or replace display EGPWS ALERT Data Bus 6A (J1-037) ARINC429 data Not Used Not Used Data Bus 6B (J1-036) in (Bus 6) NOTE: Access the applicable page(s) using the DU test switch. (See instructions above the title on the data page(s).) The associated input pin(s) and the ARINC429 input data activity may be monitored on these page(s). (19) If ARINC419 Input information suspected as faulty or additional interface details required, refer to the ARINC419 INPUTS fault table and make sure each applicable data-bus has labels and activity. ARINC419 Input Fault Table Cause of Function Malfunction

ARINC419 Input Recommended (Pin-out) Corrective Action DME Clock 1 Hi (J2-045) ARINC419 data in No or incorrect Check the applicable Clock 1 Lo (J2-044) (Bus 1) ARINC419 data interface equipment Data 1 Hi (J2-049) input and associated wiring Data 1 Lo (J2-048) or replace display Sync 1 Hi (J2-051) Sync 1 Lo (J2-050) FMS Clock 2 Hi (J2-046) ARINC419 data in Not Used Not Used Clock 2 Lo (J2-052) (Bus 2) Data 2 Hi (J2-053) Data 2 Lo (J2-047) Sync 2 Hi (J2-055) Sync 2 Lo (J2-054) NOTE: Access the applicable page(s) using the DU test switch. (See instructions above the title on the detail page(s).) The associated input pin(s) and the ARINC419 input data activity may be monitored on these page(s). (20) If ARINC-453) Input information suspected as faulty or additional interface details required, refer to the ARINC 453 INPUTS fault table and make sure each applicable data-bus has labels and activity. Revision 16

12-02-02

34.01.27

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) ARINC453) Input Fault Table Cause of Function Malfunction

ARINC453 Input Recommended (Pin-out) Corrective Action TERR/WXR IMAGE Data Bus A (J1-021) ARINC453 data in Not Used Not Used Data Bus B (J1-020) NOTE: Access the applicable page(s) using the DU test switch. (See instructions above the title on the detail page(s).) The associated input pin(s) and the ARINC453 input data activity may be monitored on these page(s). This information is enabled by DCP button or controls. (21) Push the NAV DATA button on the DCP. (22) Make sure the EADI/EHSI returns to normal mode with no DU or DCP flag in view. (23) Repeat Operational or System Test(s), as applicable, after troubleshooting and fault correction are completed. e. End of Procedure. 9. Operational Test Procedures – LCD EADI/EHSI Instruments a. Apply electrical power to the aircraft in accordance with the AMM. b. Set all related circuit breakers in accordance with the AMM. (1) Set the following circuit breakers: (a) (b) (c) (d) (e) (f)

CAPT EADI (EPC LH Emergency DC) CAPT EHSI (EPC LH Emergency DC) CAPT DCP (EPC LH Emergency DC) F/O EADI (EPC RH Emergency DC) F/O EHSI (EPC RH Emergency DC) F/O DCP (EPC RH Emergency DC)

(2) Set any other related system circuit breakers. c. Do these steps on both Captain’s and F/O’s DCP: (1) Push the DU TEST button on both DCP’s. (2) Make sure the DU STATUS page shows on Captain and F/O’s EADI and EHSI. (3) Push and hold the DU TEST button on the Captain’s DCP for at least two seconds. When the “PRESS TEST TO CONTINUE IBIT” prompt appears, release the DU TEST button.

34.01.28

06-03-02

Revision 15

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (4) While the “PRESS TEST TO CONTINUE IBIT” prompt is being displayed, push and hold the DU TEST button for at least two seconds. (This starts the built-in test. The display will blank momentarily and then display the results. When the BIT is finished, the DU STATUS page is again displayed. Items displayed in white passed and those in red failed.) (5) Repeat steps 3 and 4 for the F/O’s DU IBIT test. (6) Make sure the DU STATUS page on each EADI/EHSI shows the following BIT results: white text = passed, red text = failed. (a) All PROCESSOR status items displayed in white (Pass). (b) All DISPLAY status items displayed in white (Pass). (c) All INTERFACE status items displayed in white (Pass). (7) Make sure the DU STATUS page on each EADI/EHSI shows the following hardware and software data: (a) (b) (c) (d) (e) (f)

TYPE: OFP: FIXD: CHAR: SEL OPC: CFG:

FPI-920 091-2006-003 091-1932-002 091-2135-003 #1: EMERY DC-8 091-2007-013

(8) Ensure none of the following configuration messages show on the lower part of the DU STATUS page of either EHSI/EADI: (a) (b) (c) (d) (e) (f) (g)

NO CFG LOADED (black text on red) CFG INVALID (black text on red) UNUSED OPC SEL (white text) INVALID TYPE CODE (black text on red) INVALID TYPE (black text on red) UNABLE CONFIG X COMP (black text on yellow) CONFIG MISCOMP (black text on red)

(9) (10) (11) (12)

Push the NAV DATA button on both DCP’s. Ensure both EHSI’s return to HSI mode with no DU or DCP flags in view. Ensure both EADI’s return to ADI mode with no DU or DCP flags in view. Push the NAV DATA button on both DCP’s until the GSPD label is displayed on both EHSI’s. (13) Select a frequency of 108.0 on the Captain’s (F/O) VOR/DME control panel and verify G/S flag or G/S pointer is not displayed. (14) Set both HSI course pointers to 000 degrees using the course select knobs on the DCP.

Revision 16

12-02-02

34.01.29

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (15) Set Captain’s (F/O) EHSI and RMI compass cards to zero degrees using the SET HDG control on Capt’s (F/O’s) compass controller. (16) Press and hold the VOR 1 (2) test switch on the Captain’s (F/O’s) VOR/DME control panel and observe the following: (a) After approximately 5 to 9 seconds, the VOR 1 (2) Flag (black text on red) will be replaced with VOR 1 (2) in white text. (b) After VOR 1 (2) in white text is displayed: 1) 2) 3) 4)

VOR 1 (2) pointers in RMI’s indicate 180 (+ 4) degrees. Captain’s (F/O’s) course pointer indicates 000 degrees. Captain’s (F/O’s) deviation bar is centered (+ ½ dot). Captain’s (F/O’s) TO/FROM pointer indicates FROM.

(c) Deleted. (17) Release VOR 1 (2) test switch. (18) Press and hold the DME test switch on the Captain’s VOR/DME Control Panel and observe the following: (a) EHSI’s DME 1 readouts (distance) must display three red dashes for one second, followed by blanks for one second, followed by a display of 0.0 (+ 0.1) nautical miles. (b) Release DME 1 test switch. After 10 seconds, verify blanks on EHSI DME 1 readouts. (c) Press and hold the DME test switch on the F/O’s VOR/DME Control Panel and observe the following: 1) EHSI’s DME 2 readouts (distance) must display three red dashes for one second, followed by blanks for one second, followed by a display of 0.0 (+ 0.1) nautical miles. (d) Release DME 2 test switch. After 10 seconds, verify blanks on EHSI DME 2 readouts. (19) Select a frequency of 108.1 on the Captain’s (F/O’s) VOR/DME control panel and check the following: (a) EHSI LOC 1 (2) and GS flags displayed (black text on red). (b) EADI GS flag displayed (black text on red) and RWY not displayed. (20) Press and hold the DN/RT ILS test switch on the VOR/DME control panel and observe the following on Captain’s (F/O’s) ADI and HSI: (a) After approximately five seconds, the LOC 1 (2) Flag (black text on red) will be replaced with LOC 1 (2) in white text. After LOC 1 (2) in white text is displayed:

34.01.30

12-02-02

Revision 16

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 1) EADI GS flag not displayed. 2) EHSI GS flag not displayed. 3) EADI runway displayed. (b) EADI and EHSI GS pointers indicate one dot down + 1/8 inch. (c) EHSI course bar indicates one dot right + 1/8 inch, EADI expanded LOC is deflected full scale right and rising runway is fully extended. (d) Deleted. (e) Release the DN/RT test switch. (21) Press and hold the UP/LT ILS test switch on the VOR/DME control panel and observe the following on Captain’s (F/O’s) ADI and HSI: (a) After approximately five seconds, the LOC 1 (2) Flag (black text on red) will be replaced with LOC 1 (2) in white text. (b) After LOC 1 (2) in white text is displayed: 1) EADI GS flag not displayed. 2) EHSI GS flag not displayed. 3) EADI runway displayed. (c) EADI and EHSI GS pointers indicate one dot up + 1/8 inch. (d) HSI course bar indicates one dot left + 1/8 inch, EADI expanded LOC is deflected full scale left and rising runway is fully extended. (e) Deleted. (f) Release the UP/LT switch. d. Select ARC on Capt’s (F/O’s) DCP and repeat steps (13) through (21). e. Select MAP on both DCP’s and after 10 seconds, verify the MAP FAIL message is displayed on each EHSI. f. Secure the aircraft in accordance with the AMM. g. End of Test.

Revision 16

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34.01.31

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 10. System Test Procedures – LCD EADI/EHSI Instruments a. Apply electrical power to the aircraft in accordance with the AMM. b. Set all related circuit breakers in accordance with the AMM: (1)

Set the following circuit breakers. (a) (b) (c) (d) (e) (f)

CAPT EADI (EPC LH Emergency DC) CAPT EHSI (EPC LH Emergency DC) CAPT DCP (EPC LH Emergency DC) F/O EADI (EPC RH Emergency DC) F/O EHSI (EPC RH Emergency DC) F/O DCP (EPC RH Emergency DC)

(2) Set any other system related circuit breakers. c. Set both Captain and F/O VHF/FMS annunciator switches to VHF. NOTE:

The VHF/FMS switch on the NAV Select Panel switches the navigation source between radio (VOR or LOC) and FMS for both the EADI and EHSI.

d. Perform the following VHF Test Procedures: (1) Initial Power On. (a) Apply power to all related aircraft system equipment (VHF, COMM, NAV, and DME). NOTE:

During Power-up. Control panel will perform a built-in test (BIT) procedure. If a control panel function failure occurs, the displays will annunciate “FAIL”. The panel must be removed from aircraft for further testing.

(b) Turn on both COMM and NAV sides of Gables G7418-01 COMM/NAV/DME control panel by rotating clockwise both volume control knobs (outer small knob found on both triple-stacked control knobs). 1) Functions tested by BIT on power-up are: a) Controller functions. b) Code space (Flash memory).

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 2) Functions tested by BIT on power-up and continuously tested during normal operation (a minimum of once every second): a) Non-volatile storage (EEPROM). b) Field Programmable Gate Array (FPGA). c) +15 volt monitor. (c) Verify both LCD backlighting and front panel illuminate. Front plastic panel dimming will be controlled by cockpit dimming control (external to panel); verify dimming functions properly. (d) Verify both volume controls by adjusting the volume control (for both COMM and NAV) as heard through the cockpit speakers. (2) Frequency Selection Verification (a) Verify using left knob stack (COMM side), full VHF COMM frequency range. 1) Normal range mode with 8.33 KHz spacing mode. a) Place front panel 8.33 KHz switch to 8.33 position. b) Rotate large frequency selector knob and verify VHF COMM radio is being tuned to upper limit of 135 MHz and a lower limit of 118 MHz. c) Rotate middle frequency selector knob and verify VHF COMM radio is being tuned to upper limit of .990 MHz and a lower limit of .000 MHz. NOTE:

When control panel is powered-up in 25 KHz spacing mode, only two digits are displayed to the right of the decimal point, number in parenthesis is not displayed. In 8.33 KHz spacing mode, three digits to the right of the decimal point will be displayed.

2) Rotate left VHF COMM volume control knob fully counterclockwise (to the off position) and verify VHF COMM receiver as well as the display lighting are powered off. (b) Verify using right knob stack (NAV side), full NAV frequency range. 1) Rotate large frequency selector knob and verify VHF NAV radio is being tuned to upper limit of 117 MHz and a lower limit of 108 MHz.

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34.01.33

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 2) Rotate middle frequency selector knob and verify VHF NAV radio is being tuned to upper limit of .95(0) MHz and a lower limit of .00(0) MHz. NOTE:

NAV frequency selection is set at 50 KHz spacing. Two digits are displayed to the right of the decimal point. Shown in procedures as .95(0).

3) Rotate right NAV volume control knob fully counterclockwise (to the off position), and verify NAV receivers as well as the display lighting are powered off. a) Verify COMM TEST switch, when pressed, disables squelch in the VHF COMM transceiver. b) Verify DME TEST switch, when pressed, initiates a functional test of the DME interrogator. Front panel displays are not affected by this action. c) Verify VOR TEST switch, when pressed, initiates a VOR functional test. d) Verify 4-position DME function switch as follows: DME HOLD Position (This Function placarded “DEACT”.):

DME ORIDE Position:

(1) (2) (3) (1) (2)

(1) (2) DME NORM Position:

(1) DME STBY Position:

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(2)

Deleted Deleted Deleted Verify NAV side of control panel is turned on. With 4-poisition DME function switch set to ORIDE position, 200 mile limit of the DME instrument is overridden. Verify NAV side of control panel is turned on. With 4-position DME function switch set to NORM position, DME range changes as the frequency being dialed into the control panel changes. Verify NAV side of control panel is turned on. With 4-position DME function switch set to STBY position, DME interrogator is powered but does not provide range.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) e. Captain’s DCP Display Control Function Test: (1) (2) (3) (4) (5) (6) (7) (8) (9)

(10) (11) (12) (13) (14) (15) (16) (17) (18) (19)

(20) (21) (22) (23) (24)

Check the EADI turn/slip indicator and DCP integral lighting. Adjust the ADI - BRT control. Make sure the EADI display responds correctly to the control input. Adjust the HSI - BRT control. Make sure the EHSI display responds correctly to the control input. Rotate the MODE Selector Knob (outer) to HSI (full 360 degree compass). Make sure the EHSI display responds correctly to the control input. Rotate the CRS Selector Knob (inner). Make sure the Course Pointer and the digital CRS value on the EHSI respond correctly to the control input. The CRS data will be displayed in the lower R/H corner of the EHSI. Pressing the NAV DATA button on the associated DCP will scroll through the information available. Rotate the HDG Selector Knob (inner). Make sure the Heading Bug on the EHSI responds correctly to the control input. Push the NAV DATA button (six times). Make sure the navigation data on the EHSI toggles between (blank), GSPD, TAS, TTG, DME 1 (Distance), and CRS. Rotate the MODE Selector Knob (outer) to ARC-HSI (partial 80 degree compass). Make sure the EHSI display responds correctly to the control input. Rotate the MODE Selector Knob (outer) to ARC MAP (partial 80 degree compass with map overlay). (This will bring up the Flight Plan Leg Display.) Make sure the EHSI display responds correctly to the control input. Rotate the RNG Selector Knob (outer). Make sure the Range value on the EHSI responds correctly to the control input (5, 10, 20, 40, 80, 160 and 320 nautical miles). Visible range ring on display is ½ scale or 2.5, 5, 10, 20, 40, 80, and 160 nautical miles. Push the COMPACT mode button. Make sure the EADI changes to compact display mode (ADI/HSI data). Push the COMPACT mode button. Make sure the EADI returns to the normal ADI display. Rotate the MODE Selector Knob (outer) to HSI (full 360 degree compass).

f. F/O’s DCP Display Control Function Test: (1) (2) (3) (4) (5) (6) (7) (8) Revision 16

Check the EADI turn/slip indicator and DCP integral lighting. Adjust the ADI - BRT control. Make sure the EADI display responds correctly to the control input. Adjust the HSI - BRT control. Make sure the EHSI display responds correctly to the control input. Rotate the MODE Selector Knob (outer) to HSI (full 360 degree compass). Make sure the EHSI display responds correctly to the control input. Rotate the CRS Selector Knob (inner). 12-02-02

34.01.35

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (9) Make sure the Course Pointer and the digital CRS value on the EHSI response correctly to the control input. The CRS data will be displayed in the lower R/H corner of the EHSI. Pressing the “Nav Data” button on the associated DCP will scroll through the information available. (10) Rotate the HDG Selector Knob (inner). (11) Make sure the Heading Bug on the EHSI responds correctly to the control input. (12) Push the NAV DATA button six times. (13) Make sure the navigation data on the EHSI toggles between (blank), GSPD, TAS, TTG, FMS 1 (Distance), and CRS. (14) Rotate the MODE Selector Knob (outer) to ARC-HSI (partial 80 degree compass). (15) Make sure the EHSI display responds correctly to the control input. (16) Rotate the MODE Selector Knob (outer) to ARC MAP (partial 80 degree compass with map overlay). (This will bring up the Flight Plan Leg Display.) (17) Make sure the EHSI display responds correctly to the control input. (18) Rotate the RNG Selector Knob (outer). (19) Make sure the Range value on the EHSI responds correctly to the control input (5, 10, 20, 40, 80, 160 and 320 nautical miles). (20) Push the COMPACT mode button. (21) Make sure the EADI changes to compact display mode (ADI/HSI data). (22) Push the COMPACT mode button. (23) Make sure the EADI returns to the normal ADI display. (24) Rotate the MODE Selector Knob (outer) to HSI (full 360 degree compass). g. Compass System Test – Accomplish the Compass System test in accordance with the AMM (Chapter 34-42-0 or equivalent). NOTE: (1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12) (13) (14) (15)

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“COMP” lights are deleted.

Rotate both DCP MODE Selector Knobs to the HSI position. Verify proper operation of Compass #1 and the Captain’s EADI and EHSI. Verify proper operation of Compass #2 and the F/O’s EADI and EHSI. Trip the circuit breaker for Compass #1. Verify “HDG” displayed in red on the Captain’s EADI and EHSI. Select Compass Selector switch to “Both on Compass 2”. Verify “Mag 2” displayed in white on both EADI and EHSI. Verify “HDG” not displayed in red on the Captain’s EADI and EHSI. Verify proper operation of Compass #2 and the Captain’s EADI and EHSI. Verify proper operation of Compass #2 and the F/O’s EADI and EHSI. Set the circuit breaker for Compass #1. Select Compass Selector switch to “Normal”. Trip the circuit breaker for Compass #2. Verify “HDG” displayed in red on the F/O’s EADI and EHSI. Select Compass Selector switch to “Both on Compass 1”.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (16) (17) (18) (19) (20) (21) (22)

Verify “Mag 1” displayed in white on both EADI and EHSI. Verify “HDG” not displayed in white on both EADI and EHSI. Verify proper operation of Compass #1 and the Captain’s EADI and EHSI. Verify proper operation of Compass #1 and the F/O’s EADI and EHSI. Set the circuit breaker for Compass #2. Select Compass Selector switch to “Normal”. Verify “Mag 1” or “Mag 2” is not displayed on both the Captain and F/O’s EADI and EHSI. (23) Verify “HDG” is not displayed in red on both the Captain and F/O’s EADI and EHSI. NOTE:

When Compass is manually set using the associated RMI, the EHSI may respond as follows: the HDG flag may be displayed and the compass card may freeze and may be removed from the display. This does not indicate a fault or failure. The display should return to normal after the desired heading is manually set.

(24) On #1 compass controller (F/O’s RMI), use SET HDG knob to rotate the Captain’s RMI compass card to 90 degrees. (25) On #2 compass controller (Capt’s RMI), use SET HDG knob to rotate the F/O’s RMI compass card to 90 degrees. (26) Make sure the HDG values on both RMI’s, EADI’s, and EHSI’s match within + 2 degrees. Verify rate of turn indicator responds correctly on both EADI’s. (27) Repeat steps (24) through (26) for 180 degrees, 270 degrees, and 0 degrees. (28) Set display mode to normal (non-compact). (29) On #1 compass controller, use SET HDG knob to rotate the Captain’s RMI compass card to 10 degrees. (30) After approximately 10 seconds, verify the text message “HDG” is displayed in yellow at the top of both EADI’s. (31) Using the #1 compass controller, set the Captain’s RMI to 0 degrees. (32) After approximately 10 seconds, verify “HDG” not displayed on either EADI. h. LOC/GS (ILS) Test – Accomplish the test in accordance with the AMM (Chapter 34-310 or equivalent). NOTE:

Accomplish this test in compact mode and normal mode, as applicable.

(1) Set the MODE Selector on both DCP’s to HSI or ARC-HSI, as desired. (2) Adjust the CRS Selector Knob on both DCP’s to aircraft heading or as required during this test. The CRS data will be displayed in the lower R/H corner of the EADI/EHSI. Pressing the “Nav Data” button on the DCP will scroll through the information available. (3) Perform the following self-test steps on the Capt and F/O’s ILS systems:

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34.01.37

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (a) Set both VOR/DME Control Panels to 108.1 or an inactive LOC frequency and check the following: 1) EHSI LOC 1 (2) and GS flags displayed (black text on red). 2) EADI LOC 1 (2) and GS flag displayed (black text on red). 3) EADI runway not displayed. (b) Press and hold the DN/RT ILS test switch on the Captain’s VOR/DME Control Panel and observe the following on Captain’s EADI and EHSI: 1) After approximately five seconds, the LOC 1 Flag (black text on red) will be replaced with LOC 1 in white text. 2) After LOC 1 in white text is displayed: a) EADI GS flag not displayed. b) EHSI GS flag not displayed. c) EADI runway displayed. 3) EADI/EHSI GS pointers indicate one dot down + 1/8 inch. 4) EADI/EHSI course bar indicates one dot right + 1/8 inch, EADI expanded LOC is deflected full scale right and rising runway is fully extended. 5) Release the DN/RT test switch. (c) Press and hold the UP/LT ILS test switch on the Captain’s VOR/DME Control Panel and observe the following on Captain’s EADI and EHSI: 1) After approximately five seconds, the LOC 1 Flag (black text on red) will be replaced with LOC 1 in white text. 2) After LOC 1 in white text is displayed: a) EADI GS flag not displayed. b) EHSI GS flag not displayed. c) EADI runway displayed. 3) EADI/EHSI GS pointers indicate one dot up + 1/8 inch. 4) EADI/EHSI course bar indicates one dot left + 1/8 inch, ADI expanded LOC is deflected full scale left and rising runway is fully extended. 5) Release the UP/LT test switch. (d) Press and hold the DN/RT ILS test switch on the F/O’s VOR/DME Control Panel and observe the following on F/O’s EADI and EHSI:

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 1) After approximately five seconds, the LOC 2 Flag (black text on red) will be replaced with LOC 2 in white text. 2) After LOC 2 in white text is displayed: a) EADI GS flag not displayed. b) EHSI GS flag not displayed. c) EADI runway displayed. 3) EADI/EHSI GS pointers indicate one dot down + 1/8 inch. 4) EADI/EHSI course bar indicates one dot right + 1/8 inch, EADI expanded LOC is deflected full scale right and rising runway is fully extended. 5) Release the DN/RT test switch. (e) Press and hold the UP/LT ILS test switch on the F/O’s VOR/DME Control Panel and observe the following on F/O’s ADI and HSI: 1) After approximately five seconds, the LOC 2 Flag (black text on red) will be replaced with LOC 2 in white text. 2) After LOC 2 in white text is displayed: a) EADI GS flag not displayed. b) EHSI GS flag not displayed. c) EADI runway displayed. 3) EADI/EHSI GS pointers indicate one dot up + 1/8 inch. 4) EADI/EHSI course bar indicates one dot left + 1/8 inch, EADI expanded LOC is deflected full scale left and rising runway is fully extended. 5) Release the UP/LT test switch. (4) Perform the following ramp-test steps on the Captain and F/O’s ILS systems: (a) Set both VOR/DME Control Panels to 108.1 or other equivalent ramp-test LOC frequency. (b) Set the ramp-test set to LOC and adjust the LOC switch to provide a centered LOC signal and check the following: 1) LOC 1 (2) displayed in white text on EADI/EHSI. 2) EHSI course bars are centered + 1/8 inch. 3) EADI localizer pointers are centered + 1/8 inch. (c) Adjust the ramp-test set to provide a two dots left LOC signal and check the following:

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34.01.39

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 1) EHSI course bars indicate two dots left + 1/8 inch. 2) EADI localizer pointers indicate full scale left. (d) Adjust the ramp-test set to provide a one dot left LOC signal and check the following: 1) EHSI course bars indicate one dot left + 1/8 inch. 2) EADI localizer pointers indicate full scale left. (e) Adjust the ramp-test set to provide two dots right LOC signal and check the following: 1) EHSI course bars indicate two dots right + 1/8 inch. 2) EADI localizer pointers indicate full scale right. (f)

Adjust the ramp-test set to provide a one dot right LOC signal and check the following: 1) EHSI course bars indicate one dot right + 1/8 inch. 2) EADI localizer pointers indicate full scale right.

(g) Adjust the ramp-test set to provide a centered LOC signal and check the following: 1) EHSI course bars are centered + 1/8 inch. 2) EADI localizer pointers are center + 1/8 inch. (h) Adjust the GS switch to provide a centered GS signal and check the following: 1) GS flags not displayed on EADI/EHSI. 2) EADI and EHSI GS pointers are centered + 1/8 inch. (i)

Adjust the ramp-test set to provide a two dots up GS signal and check the following: EADI and EHSI GS pointers indicate two dots up + 1/8 inch. (j) Adjust the ramp-test set to provide a one dot up GS signal and check the following: EADI and EHSI GS pointers indicate one dot up + 1/8 inch. (k) Adjust the ramp-test set to provide a two dots down GS signal and check the following: EADI and EHSI GS pointers indicate two dots down + 1/8 inch. (l) Adjust the ramp-test set to provide a one dot down GS signal and check the following: EADI and EHSI GS pointers indicate one dot down + 1/8 inch. (m) Adjust the GS switch to provide a centered GS signal and check the following: EADI and EHSI GS pointers are center + 1/8 inch.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (5) Set the ramp-test set to OFF and check the following: EADI and EHSI LOC 1 (2) and GS flags displayed (black text on red). (6) Select Radio Navigation switch to “Nav-2”. (7) Set the ramp-test set to LOC, adjust the LOC switch to provide a centered LOC signal, and check the following: (a) LOC 2 displayed in white on Captain’s EADI/EHSI. (b) EHSI course bars are centered + 1/8 inch. (c) EADI localizer pointers are centered + 1/8 inch. (8) Verify proper operation of LOC #2 and the Captain’s EHSI. (9) Select Radio Navigation switch to “Nav-1”. (10) Verify that “LOC 1” and “LOC 2” are displayed in white and not yellow or red on the respective EADI/EHSI’s. i. VOR Test – Accomplish the test in accordance with the AMM (Chapter 34-52-0 or equivalent). NOTE:

Accomplish this test in compact mode.

(1) Set the MODE selector on both DCP’s to HSI or ARC-HSI, as desired. (2) Adjust the CRS Selector Knob on both DCP’s as required during this test. The CRS data will be displayed in the lower R/H corner of the EADI/EHSI. By pressing the “Nav Data” button, the DCP will scroll through the information available. NOTE:

Compass magnetic heading information must match during set (3). Select both compasses to DG and slew to match the Standby Compass value. (Single compass reference may also be used during this test.)

(3) Set HSI course pointers to 000 degrees using the course select knobs on both DCP’s. (4) Rotate single pointer knob to VOR on both RMI’s. (5) Rotate double pointer knob to VOR on both RMI’s. (6) Perform the following self-test steps on the Captain and F/O’s VOR systems: (a) Set both VOR/DME Control Panels to 108.0 or an inactive VOR frequency and check the following: 1) EHSI and EADI GS flags and pointers not displayed. 2) VOR 1 (2) flag displayed (black text on red) on EADI/EHSI. (b) Press and hold the VOR test switch on the Captain’s VOR/DME Control Panel and observe the following:

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34.01.41

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 1) After approximately 5 to 9 seconds, the VOR 1 Flag (black text on red) will be replaced with VOR 1 in white text. 2) After VOR 1 in white text is displayed: a) b) c) d)

VOR 1 pointers in RMI’s indicate 180 (+ 4) degrees. Captain’s EADI/EHSI course pointer indicates 000 degrees. Captain’s EADI/EHSI deviation bar is centered (+ ½ dot). Captain’s EADI/EHSI TO/FROM pointer indicates FROM.

(c) Release VOR test switch. (d) Press and hold the VOR test switch on the F/O’s VOR/DME Control Panel and observe the following: 1) After approximately 5 to 9 seconds, the VOR 2 Flag (black text on red) will be replaced with VOR 2 in white text. 2) After VOR 2 in white text is displayed: a) b) c) d)

VOR 2 pointers in RMI’s indicate 180 (+ 4) degrees. F/O’s EADI/EHSI course pointer indicates 000 degrees. F/O’s EADI/EHSI deviation bar is centered (+ ½ dot). F/O’s EADI/EHSI TO/FROM pointer indicates FROM.

(e) Release VOR test switch. (7) Perform the following ramp-test steps on the Captain and F/O’s VOR systems: (a) Set both VOR/DME Control Panels to 108.0 or other equivalent ramp-test VOR frequency. (b) Set both EHSI course pointers to 000 degrees using the course select knobs on the DCP’s. (c) Set the ramp-test set to VOR and adjust the VOR switches to provide a centered VOR signal at 000 degrees/TO the station and check the following: 1) 2) 3) 4)

VOR 1 (2) displayed in white text on both EADI/EHSI’s. Course deviation bars are centered. TO/FROM pointers indicate TO. RMI single and double pointers indicate 000 (+ 4) degrees.

(d) Rotate Captain’s course selector until EADI/EHSI course deviation bar indicates two dots right. Verify EADI/EHSI CRS values read 350 + 2 degrees. (e) Rotate Captain’s course selector until EADI/EHSI course deviation bar indicates two dots left. Verify EADI/EHSI CRS values read 010 + 2 degrees.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (f) (g) (h) (i) (j) (k) (l) (m) (n)

Rotate F/O’s course selector until EADI/EHSI course deviation bar indicates two dots right. Verify EADI/EHSI CRS values read 350 + 2 degrees. Rotate F/O’s course selector until EADI/EHSI course deviation bar indicates two dots left. Verify EADI/EHSI CRS values read 010 + 2 degrees. Repeat steps (b) and (c) for 030 degrees. Repeat steps (b) and (c) for 060 degrees. Repeat steps (b) and (c) for 090 degrees. Repeat steps (b) and (c) for 120 degrees. Repeat steps (b) and (c) for 150 degrees. Repeat steps (b) and (c) for 180 degrees. Adjust the VOR switches to provide a centered VOR signal at 180 degrees/FROM the station and check the following: 1) 2) 3) 4)

(o) (p) (q) (r) (s)

VOR 1 (2) displayed in white text on both HSI’s. Course deviation bars are centered. TO/FROM pointers indicate FROM. RMI single and double pointers indicate 000 (+ 4) degrees.

Repeat steps (b) and (c) for 210 degrees. Repeat steps (b) and (c) for 240 degrees. Repeat steps (b) and (c) for 270 degrees. Repeat steps (b) and (c) for 300 degrees. Repeat steps (b) and (c) for 330 degrees.

(8) Set the ramp-test set to OFF and check the following: VOR 1 and VOR 2 flags are displayed (black text on red) on the respective EADI/EHSI’s. k. DME Test – Accomplish the test in accordance with the AMM (Chapter 34-54-0 or equivalent). NOTE:

Accomplish this test in compact mode.

(1) Set the MODE Selector on both DCP’s to HSI or ARC-HSI, as desired. (2) DME 1 distance is displayed on the upper left corner on both EHSI’s and on the middle left of both EADI’s. DME 2 distance is displayed on the upper right corner of both EHSI’s and on the middle left of both EADI’s. (3) Perform the following self-test steps on the Captain and F/O’s DME systems: (a) Set both VHF NAV selector panels to 108.0 or an inactive VOR frequency and check the following: VOR 1 (2) flag displayed (black text on red) on EADI/EHSI. (b) Press and hold the DME test switch on the Captain’s VOR/DME Control Panel and observe the following: EADI/EHSI’s DME 1 readouts (distance) must display three red dashes for one second, followed by blanks for one second, followed by a display of 0.0 (+ 0.1) nautical miles. Revision 16

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34.01.43

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (c) Release DME 1 test switch. After 10 seconds, verify blanks on EADI/EHSI DME 1 readouts. (d) Press and hold the DME test switch on the F/O’s VOR/DME Control Panel and observe the following: EADI/EHSI’s DME 2 readouts (distance) must display three red dashes for one second, followed by blanks for one second, followed by a display of 0.0 (+ 0.1) nautical miles. (e) Release DME 2 test switch. After 10 seconds, verify blanks on EADI/EHSI DME 2 readouts. (4) Tune both VOR/DME radios to a local station and verify the following: (a) DME 1 results displayed in DME 1 readouts on both EADI/EHSI’s must agree with know distance within + 1 NM. (b) DME 2 results displayed in DME 2 readouts on both EADI/EHSI’s must agree with know distance within + 1 NM. (5) If no local stations available, accomplish the DME System Test in accordance with AMM Chp. 34-54-0 (or equivalent). l. WINDSHEAR System Test – Reference Honeywell or Safe-Flight AMM Supplement (or equivalent). (1) Accomplish the Windshear system test in accordance with the Honeywell or Safe Flight AMM Supplement (or equivalent). (2) When the Windshear Caution alert is triggered, “WINDSHEAR” will be displayed in yellow on both Captain’s and F/O’s EADI’s. (Windshear caution annunciators are deleted). (3) When the Windshear Warning alert is triggered, “WINDSHEAR” will be displayed in red on both Captain’s and F/O’s EDI’s. (Windshear warning annunciators are deleted.) m. NAV INSTRUMENT FAILURE System Test – Accomplish the test in accordance with the AMM (Chapter 34-24-0 or equivalent). (1) Verify both EADI’s are set to “Normal” mode (not “Compact” mode). (2) Push the INST WARN TEST switch. (3) Verify “INST 1”, “INST 2”, “GA 1”, and “GA 2” display in red on the Captain and F/O’s EADI’s. (4) Release the INST WARN TEST switch. (5) Accomplish the Navigation Instrument Monitor system tests in accordance with the AMM (Chp. 34-24-0 or equivalent). NOTE:

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Bow-tie annunciators are deleted.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (6) Verify proper operation of Comparator. (7) Verify “PITCH”, “ROLL”, “HDG”, “LOC”, and “GS”, as applicable, are displayed in yellow at the top of both EADI’s during test. (“INSTR 1”, “INSTR 2”, “GA 1”, and “GA 2” green annunciators are deleted. “SPD COMD”, “RAD ALT”, “MONITOR”, and “GP EXT” amber annunciators are deleted.) (8) Trip the circuit breakers for the Captain’s EADI/EHSI. (9) Verify “COMP” displayed in red at the top of the F/O’s EADI. (10) Set the circuit breaker for the Captain’s EADI/EHSI. (11) Trip the circuit breakers for the F/O’s EADI/EHSI. (12) Verify “COMP” displayed in red at the top of the Captain’s EADI. (13) Set the circuit breaker for the F/O’s EADI/EHSI. (14) Trip the circuit breaker for VG #1. (15) Verify “INST 1” and “GA 1” are displayed in red on both EADI’s. (16) Set the circuit breaker for VG #1. (17) Verify “INST 1” and “GA 1” are extinguished on both EADI’s. (18) Trip the circuit breaker for VG #2. (19) Verify “INST 2” and “GA 2” are displayed in red on both EADI’s. (20) Set the circuit breaker for VG #2. (21) Verify “INST 2” and “GA 2” are extinguished on both EADI’s. n. FLIGHT DIRECTOR System Test – Accomplish the test in accordance with AMM Chapter 34-23-0 (or equivalent). (1) Set the MODE Selector on both DCP’s to HSI or ARC-HSI, as desired. (2) Adjust the CRS Selector Knob on both DCP’s as required during this test. The CRS data will be displayed in the lower R/H corner of the EHSI. Pressing the NAV DATA button on the associated DCP will scroll through the information available. (3) Adjust the HDG Selector Knob on both DCP’s as required during this test. (4) Accomplish the Flight Director system tests in accordance with the AMM (Ch. 3423-0 or equivalent). (5) Verify proper operation of Flight Director System #1 and the Captain and F/O’s EADI/EHSI. (6) Select Captain’s FD Selector to “HDG”. (7) Verify director needles are displayed on the Captain’s EADI. (8) Trip the AC Circuit breaker for FD #1. (9) Verify director needles are not displayed on the Captain’s EADI. (10) Verify “FD” displayed in red on the Captain’s EADI. (11) Set the AC circuit breaker for FD #1. (12) Verify “FD” is extinguished on the Captain’s EADI. (13) Repeat steps (8) through (11) for DC Circuit Breaker FD #1. (14) Select F/O’s FD Selector to “HDG”. (15) Verify director needles are displayed on the F/O’s EADI. (16) Trip the AC circuit breaker for FD #2.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (17) (18) (19) (20) (21)

Verify director needles are not displayed on the F/O’s EADI. Verify “FD” displayed in red on the F/O’s EADI. Set the AC circuit breaker for FD #2. Verify “FD” is extinguished on the F/O’s EADI. Repeat steps (14) through (20) for DC Circuit Breaker FD #2.

o. RADIO ALTIMETER Test – Accomplish the test in accordance with AMM Chapter 3442-0 (or equivalent). (1) (2) (3) (4) (5) (6) (7) (8) (9) (10) (11) (12)

Set the Captain and F/O’s MDA Bugs to 100 feet. Set both VOR/DME Control Panels to 108.1 or other unused LOC frequency. Set aircraft to “AIR” mode. (Trip weight on wheel circuit breakers.) Push “Press to Test” button on the Captain’s Radio Altimeter Indicator. Verify “Rising Runway” operation on the Captain’s EADI. Verify MDA alert illuminates on the Captain’s Radio Altimeter Indicator. Verify “MDA” is displayed momentarily in yellow on the Captain’s EADI. Push “Press to Test” button on the F/O’s Radio Altimeter Inidcator. Verify “Rising Runway” operation on the F/O’s EADI. Verify MDA alert illuminates on the F/O’s Radio Altimeter Indicator. Verify “MDA” is displayed momentarily in yellow on the F/O’s EADI. Return aircraft to “Ground” mode. (Set weight on wheels circuit breakers.)

p. AUTOPILOT Test – Accomplish the test in accordance with AMM Chapter 22 (or equivalent). (1) Set the MODE Selector on the Captain’s DCP to HSI or ARC-HSI, as desired. (2) Adjust the CRS Selector Knob on the Captain’s DCP as required during this test. The CRS data will be displayed in the lower R/H corner of the EHSI. Pressing the NAV DATA button on the associated DCP will scroll through the information available. (3) Adjust the HDG Selector Knob on the Captain’s DCP as required during this test. (4) Accomplish the Autopilot system tests in accordance with the AMM (Ch. 22 or equivalent) and FMS autopilot test per AMM Supplement (or equivalent). (5) Disconnect the autopilot and verify “AP” is displayed in red on both EADI’s. (6) Push the Autopilot disconnect switch and verify that the red “AP” annunciation is extinguished. q. FMS/GPS Test – Reference UNS-1D FMS/GPS AMM Supplement and/or UNS-1D Installation Manual (or equivalent). (Power-up EHSI display deviation test is not applicable.) NOTE:

34.01.46

Accomplish the following test in compact mode.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (1) Verify that both FMS 1 and FMS 2 are not on. (FMS systems must be off for the following steps.) (2) Set the MODE Selector on both DCP’s to HSI or ARC-HSI, as desired. (3) Set both Captain and F/O’s VHF/FMS annunciator switches to FMS. (4) Make sure “FMS1” distance and “DTK” data are displayed on the Captain’s EHSI lower R/H corner. Pressing the NAV DATA on the Captain’s DCP will scroll through the information available. The “FMS1” source label and VNAV pointers are displayed on the Captain’s EADI and EHSI. (5) Make sure “FMS2” distance and “DTK” data are displayed on the F/O’s EHSI lower R/H corner. Pressing the NAV DATA button on the F/O’s DCP will scroll through the information available. The “FMS2” source label and VNAV pointers are displayed on the F/O’s EADI and EHSI. (6) Energize the FMS systems and accomplish the FMS/GPS system tests in accordance with the UNS-1D FMS/GPS AMM Supplement and the UNS-1D Installation Manual (or equivalent). (7) Verify proper operation of FMS #1 and the Captain’s EADI/EHSI. (8) Verify proper operation of FMS #2 and the F/O’s EADI/EHSI. (9) Set the MODE Selector on both DCP’s to ARC-MAP. (10) Adjust the CRS Selector Knob on both DCP’s. (11) Verify both EHSI displays do not respond to the control input. (12) Verify “FMS 1” source label, “FMS 1” distance,” “DTK”, and map overlay are displayed on the Captain’s EHSI. (13) Verify “FMS 2” source label, “FMS 2” distance, “DTK”, and map overlay are displayed on the F/O’s EHSI. (14) Adjust the RNG Selector Knob on the DCP’s as required during this test. (15) Accomplish the FMS/GPS system tests in accordance with the UNS-1D FMS/GPS AMM Supplement and the UNS-1D Installation Manual (or equivalent). (16) Verify proper map display operation of FMS #1 on the Captain’s EHSI. (17) Verify proper map display operations of FMS #2 on the F/O’s EHSI. r. Secure the aircraft in accordance with the AMM. s. End of Test. 11. EHSI and EADI - Removal a. There are two electronic horizontal situation indicators (EHSI) and two Electronic Attitude Direction Indicators (EADI) located in the flight compartment: one EHSI and EADI on the captain’s instrument panel, and one EHSI and EADI on the first officer’s instrument panel. b. Removal/installation for all indicators is identical except for circuit breakers, which must be opened. If necessary, the captain and first officer’s instrument panels can be opened for access to the indicator wiring.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) (1) Open following circuit breakers as applicable to EHSI/EADI being removed: (a) (b) (c) (d) (e)

CAPT EHSI and CAPT EADI (Left Emergency 28V DC Bus) F/O EHSI and F/O EADI (Right Emergency 28V DC Bus) #1 & 2 Vertical Gyro on RT/LT Emergency A/C. #1 & 2 Directional Gyro on RT/LT Emergency A/C. Any other related system circuit breakers.

WARNING:

TAG AND SAFETY CIRCUIT BREAKERS.

(2) Loosen clamp adjustment (indicator retaining) screws. (Refer to Figure 34.01-9.) (3) Press loosened adjustment screws back flush against panel face. (4) Loosen clamp attachment screws to relieve pressure of clamp on indicator. NOTE: Airplanes with retaining clip on lower right screw rotate clip to clear instrument when removing instrument. CAUTION:

THE EHSI AND EADI ARE ELECTROSTATIC SENSITIVE DEVICES (ESD) AND PRECAUTIONS NECESSARY FOR HANDLING ESD’S MUST BE ADDRESSED. WHEN REMOVING AND INSTALLING AN ESD, THE TECHNICIAN MUST BE PROPERLY GROUNDED TO AIRFRAME USING APPROVED BONDING WRIST STRAP. EQUIPMENT ELECTRICAL CONNECTORS MUST HAVE CONDUCTIVE CAPS INSTALLED WHEN REMOVED AND EQUIPMENT MUST BE PLACED IN CONDUCTIVE BAGS WHEN STORING OR SHIPPING

CAUTION:

DO NOT PRY INDICATOR FROM PANEL WITH SCREWDRIVER OR OTHER TOOL AS INDICATOR OR PANEL CAN BE DAMAGED. REMOVE INDICATOR CAREFULLY TO AVOID STRIKING CONTROL COLUMN AND DAMAGING GLASS.

(5) Pull indicator out of panel face, disconnect, and cap electrical connector. 12. EHSI and EADI - Installation a. Ensure following applicable circuit breakers are opened, tagged and safetied: (1) (2) (3) (4) (5)

CAPT EHSI and CAPT EADI (Left Emergency 28V DC Bus) F/O EHSI and EADI (Right Emergency 28V DC Bus) #1 & 2 Vertical Gyro on RT/LT Emergency A/C. #1 & 2 Directional Gyro on RT/LT Emergency A/C. Any other related system circuit breakers.

NOTE: Observe special handling precautions for ESD devices described in removal instructions.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) b. Remove cap and connect electrical connector to back of indicator. c. Insert indicator through panel face and mounting clamp. NOTE: Instrument clamp should be checked for clamp rivet failure at eyelets before instrument installation. d. Secure clamp adjustment (indicator retaining) screws and clamp attachment screws. Tighten to torque of 6-8 inch lbs. (Refer to Figure 34.01-9.) CAUTION: ENSURE CLAMP ATTACHMENT SCREWS ARE TIGHT OR INDICATOR MAY SLIDE FROM PANEL DURING AIRPLANE ACCELERATION. CAUTION: VERIFY THAT INSTRUMENT COOLING HOLES ARE NOT OBSTRUCTED BY INSTRUMENT MOUNTING CLAMP. NOTE: Airplanes with retaining clip on lower right screw, position clip to engage corner of instrument property before tightening screw. e.

Close following applicable circuit breakers and remove warning signs: (1) CAPT EHSI and CAPT EADI (Left Emergency 28V DC Bus) (2) F/O EHSI and EADI (Right Emergency 28V DC Bus) (3) Any other related system circuit breakers.

f. Check EADI/EHSI integral lighting. g. Set the Captain and F/O’s panel lighting control to the full clockwise position. h. Do an operational test of the EADI/EHSI. (Refer to 34-20-00 Operational Test, pages 501 – 502.)

INDICATOR INSTALLATION Figure 34.01-9

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 13. Display Control Panel – Removal – There are two Display Control Panels (DCP) located on the pedestal in the cockpit. Removal/installation procedures for both DCP’s are identical. a. Open and tag the following circuit breakers: (1) (2) (3) (4) (5)

CAPT DCP (Left Emergency 28V DC Bus) F/O DCP (Right Emergency 28V DC Bus) #1 & 2 Vertical Gyro on RT/LT Emergency A/C. #1 & 2 Directional Gyro on RT/LT Emergency A/C. Any other related system circuit breakers

WARNING: TAG AND SAFETY CIRCUIT BREAKERS CAUTION: THE DCP IS AN ELECTROSTATIC SENSITIVE DEVICE (ESD) AND PRECAUTIONS NECESSARY FOR HANDLING ESD’S SHOULD BE ADDRESSED. WHEN REMOVING AND INSTALLING AN ESD THE TECHNICIAN MUST BE PROPERLY GROUNDED TO AIRFRAME USING APPROVED BONDING WRIST STRAP. EQUIPMENT ELECTRICAL CONNECTORS MUST HAVE CONDUCTIVE CAPS INSTALLED WHEN REMOVED, AND EQUIPMENT MUST BE PLACED IN CONDUCTIVE BAGS WHEN STORING OR SHIPPING. b. Loosen four turn-lock fasteners. c. Carefully remove the DCP from panel. d. Rotate connector-locking ring on aircraft wiring connector counterclockwise and remove connector from DCP. e. Install cap plug on DCP connector. 14. Display Control Panel - Installation NOTE: Observe special handling precautions for ESD devices described in removal instructions. a. Remove cap plug from DCP connector. b. Connect aircraft wiring connector to DCP connector, then secure by rotating locking ring clockwise. c. Install DCP in panel and secure the four turn-lock fasteners. d. Close the following circuit breakers: (1) (2) (3) (4) (5)

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CAPT DCP (Left Emergency 28V DC Bus) F/O DCP (Right Emergency 28V DC Bus) #1 & 2 Vertical Gyro on RT/LT Emergency A/C. #1 & 2 Directional Gyro on RT/LT Emergency A/C. Any other related system circuit breakers

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) 15. Display Control Panel – Test a. Check EHSI, EADI and DCP integral lighting. b. Set the Captain and F/O’s panel lighting and center pedestal lighting control to the full clockwise position. c. Do an operational test of the EADI/EHSI. (Refer to Paragraph 9 – Operational Test Procedures.) 16. Preflight Differences a. On arrival at the aircraft, the displays should appear blank. Confirm EADI, EHSI, and DCP circuit breaks on the left and right DC emergency busses are in. These circuit breakers should be in at all times. Turn the bright control knob on the CAPT and F/O DCP Panel and adjust the EADI and EHSI display to desired intensity. b. Instrument Annunciator Indicator (1) Instrument Annunciator Test Button – Press to test. (a) Observe instrument red warning flags, INST 1, GA 1, INST 2, and GA 2 appear in both CAPT and FO’s EADI’s. Yellow comparator warning flags, ROLL, HDG, G/S, and LOC will not appear. These are tested by the self-test. (2) Annunciator Reset Button (Pilot’s Instrument Panel) – Press to reset. (a) The reset button will reset the red warning flags. c. Nav Receiver Checks – Receiver checks remain the same. d. Flight Director Checks – Flight director checks remain the same, except the heading and course select knobs are located on the DCP installed in the center pedestal. e. Autopilot Checks – Autopilot checks remain the same, except the heading and course select knobs are located on the DCP installed in the center pedestal. The EADI also has a repeating red AP fail flag that will appear anytime that the autopilot fail light located on the FO instrument panel is illuminated. f. Windshear Warning System Checks – Windshear annunciation appears in the EADI; discrete annunciation panel has been removed.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) g. Display and Control Panel (1) Rotate the BRT knobs clockwise to increase display brightness and counterclockwise to decrease brightness. (2) Press the COMPACT button to toggle the display between the basic EADI and the compact flight display. (3) Press the NAV DATA button to display different data available in the lower left corner of the EHSI. (4) Rotate the MODE selector switch between the different EHSI modes. Two available displays are Basic EHSI and ARC EHSI. (5) Rotate the HDG selector to adjust the position of the heading bug on the compass card and the heading in the selected heading readout. (6) Rotate the CRS selector to adjust the desired course readout and pointer. A radio navigation source must be available. (7) Rotate the RNG selector to change the range display of the FMS map display. (8) Press the DU TEST button. This will activate the internal built-in test (IBIT). Press the NAV DATA button to return to normal display. The DU TEST button is deactivated in flight.

17. Electronic Flat Panel Instruments Illustrated Parts Catalog a. This section contains information pertinent to the installation of Rockwell Collins Electronic Flat Panel Instruments in DC-8 aircraft per Hollingsead MDL 2515-01. b. The IPC data contained herein supplements or modifies the information contained in McDonnell Douglas IPC. The information contained in this supplement will supersede or modify the information in the McDonnell Douglas IPC only where specifically stated in this document.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

IPC: LRU INSTALLATION Figure 34.01-10 (1 of 3) Revision 16

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) Fig Item 10-1

Part Number 2515-3420-103

10 822-1624-012

20 64624B 30 MS35214-48 40 MS35214-49 50 822-1255-401

1234567 Nomenclature Installation – Electronic Flat Panel Line Replaceable Unit Panel, Flat, Indicator (FPI-920) (Rockwell-Collins) Clamp, Electronic Flat Panel (MSP Inc., Nashville, TN) Alt 64301B (MSP Inc., Nashville, TN) Screw, Claw Jack Screw Display Control Panel (Rockwell-Collins)

Effect From To

RF A

4

A

4

A A

8 8

A

2

2515-3915-501Blanking Plate A 101 NOTE A: Effectivity: N602AL, N603AL, N604BX, N605AL, N606AL 60

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Unit Per Assy

2

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

IPC: TERMINAL BLOCK INSTALLATION Figure 34.01-10 (2 of 3)

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d) Fig Item 10-2 10 -15 20 21 30 40 50 51 60 70 80 90 NOTE A:

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Part Number

1234567 Nomenclature

2515-3444-100 Installation – Terminal Block 2515-3420-101 Assembly 2515-3915-401Mount Plate (Capt.) 101 (Hollingsead) 2515-3915-401Mount Plate (F/O) 103 (Hollingsead) YHLZG-16-1 Ground Block (Burndy) MS20470AD3-5 Rivet, Universal Head YHLZE8-4 Terminal Block (Burndy) AAMC-6 End Clamp (Burndy) AMY720OT5 Rail 20” (Burndy) MS20426AD4-5 Rivet, Flush Head NAS1801-8-8 Screw NAS1 Washer 149DO332J DAT-18-502 Ref Designator Label (Brady) 2515-1100400Placard (Hollingsead) 105 Effectivity: N602AL, N603AL, N605AL, N606AL

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Effect From To

Unit Per Assy RF

A

1

A

1

A A A A A A A

12 48 112 4 2 18 8

A

8

A

18

A

2

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AIR TRANSPORT INTERNATIONAL GENERAL MAINTENANCE MANUAL MAINTENANCE FORMS & REPORTS 34.01 ELECTRONIC FLAT PANEL INSTRUMENT INSTALLATION (cont’d)

IPC: CIRCUIT BREAKER INSTALLATION Figure 34.01-10 (3 of 3) Fig Item

Part Number

2515-3420-100 2515-3420-101 2515-3420-102 10 BACC18Z1R 20 BACC18Z7R

10-3

Revision 16

1234567 Nomenclature

Effect From To

Installation – Circuit Breakers Circuit Breaker 1 Amp Circuit Breaker 7.5 Amp

12-02-02

Unit Per Assy RF

A A

2 4

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION

INTENTIONAL BLANK

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION

34.02 – ATC TRANSPONDER SYSTEM TABLE OF CONTENTS Paragraph 1.

TEST REQUIREMENTS ---------------------------------------------------------

Page 34.02.01

2.

GENERAL--------------------------------------------------------------------------

34.02.01

3.

FRONT PANEL CONTROLS (ALL UNITS) ---------------------------------

34.02.02

RECURRING NON-ROUTINE JOB CARDS 34.02-1

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FAR 43 ATC TRANSPONDER SYSTEM RAMP TEST RECORD---------- 34.02.05

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34 - NAVIGATION

INTENTIONAL BLANK

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34-NAVIGATION 34.02 ATC TRANSPONDER SYSTEM 1. Test Requirements a. FAR Part 91 requires that every 24 months each aircraft transponder system be checked in accordance with Appendix F of Part 43. b. In addition to the FAR requirements, a transponder system check must be performed when: (1) replacing a component which provides data to the transponder system, such as: (a) (b) (c) (e) (f) (g)

Digital Air Data Computer #1 #2 FMS Digital Air Data Computer (-73 Series Aircraft Only) Altimeter (All Non-73 Series Aircraft) Control Head Transponder Antenna

NOTE: (2)

RVSM compliance requires two separate functioning altitude sources.

repairing wiring for the above listed components and systems.

c. Test points of the component replacement test procedures will be used. d. To verify ATC transponder operation, difference between ATC transponder and altitude reporting equipment shall not exceed 125 feet (FAR 43 Appendix E, Para. (c)). e. Documentation. Form 34.02-1 provides the required steps and space to record readings of a system check. This form must be completed and attached to the action document whenever a transponder/altitude reporting system check required above is performed. f. Test Equipment. (1) Use the T-48D/T-49/T-49C/T-49CF or equivalent. If equivalent is used, follow the instructions contained with that equipment. (2) For ELS equipped aircraft, use the TR-220 Multifunction Test Set or equivalent. Refer to Section 2 of the TR-220 Operating/Maintenance Manual for operating instructions. If the TR-220 Test Set or equivalent is not available, perform the test using the T-48D/T-49/T-49C/T-49CF or equivalent test set, and a carryover must be issued to show that the aircraft is not ELS compliant.

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34-NAVIGATION 34.02 ATC TRANSPONDER SYSTEM (CONT’D) 2. General a. The T-48D XPDR/DME Ramp Test Set is a battery powered, portable unit used to test operation of transponder and DME systems antenna-to-antenna when installed in an aircraft. A direct connection coupler is also provided so that the test set can be directly connected to the antenna connector of the unit under test (UUT). The T-48D test set may be operated in one of three modes of operation; (1) via the use of an Omni antenna to the antenna connector, (2) via the implementation of the TAP-122 in a “direct-connection” configuration between the test set and the UUT RF output connectors, and (3) by placing the TAP-126 antenna fixture to the blade antenna of the UUT. b. The T-49 series of TCAS/XPDR test equipment tests airborne ATCRBS (Mode A/C) and Mode S transponders and TCAS I/II systems. They are self-contained battery-operated test sets that require no direct connection to the equipment under test. The test set receives and radiates signals to the UUT from an antenna. For Mode-S and ATCRBS, an antenna coupler unit is provided to measure transponder transmitter power, frequency, receiver sensitivity, and diversity operation. c. Both the T-48D & T-49 series of test sets are equipped with an Accessories Connector capable of connecting to a standard 9 pin serial communications cable for the purpose of downloading stored data to an IBM PC or compatible. 3. Front Panel Controls (All Units) a. DC Power Switch. All the T-48D/T-49 series units can be operated via AC ground power or DC power supplied via an onboard battery pack. To engage battery power on the T-48D, place the switch marked “DC” to the “ON” position. On the T-49 series test sets, simply initiate the tests with the “AC Power” switch in the off position. This will begin the test/interrogate sequence in the DC power mode of operation. b. Interrogate Switch. Select the appropriate mode of operation with the mode selector knob, and press the “Interrogate Switch” to initiate an automatic sequence of tests. The test sequence continues until a failure occurs (error message displayed). After successful completion of interrogate mode, the test data is stored in internal memory. This data can later be sent to a computer for a hard copy. c. Test Switch. Select the appropriate mode of operation with the mode selector knob, and press the “Test Switch” once to start the first test of the sequence and display relevant data after the completion of the test. The test will repeat continuously and the results displayed until the test switch is pressed again to move onto the next test in the sequence. d. Light Switch. When pressed to MOM (momentary) or ON, the display backlighting is illuminated. Note: the use of the backlight feature significantly decreases battery life. e. AC Power Switch. Line voltage On-Off switch for battery charging. On the T-48D, the yellow AC Power lamp will light when the AC is turned on. On the T-49 series equipment, a red LED will illuminate when the unit is charging. 34.02.02

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34-NAVIGATION 34.02 ATC TRANSPONDER SYSTEM (CONT’D) f. AC Input Connector. AC plug for ground operation or charging of the internal battery. g. Mode Switch. This rotary switch selects the time delay between P1 and P3, and thus the interrogation Modes 1, 2, 3/A, B, C, and D. h. Accessories Connector. Connect a standard 9-pin serial communications cable to an IBM PC or compatible. i. Data Display Window. An alpha-numeric display (two line, twenty character) which provides operational instructions, error messages, scenario progress, and test data. j. Antenna Connector. Connection point for the omni-directional dipole antenna, directional (TAP) antenna, and antenna coupler (TAP) units. k. T-48D Front Panel Features. (1)

Function Rotary Switch. (a) Mode 3a. Interrogates Mode 3a Transponder (b) Mode C. Interrogates Mode C Transponder (c) Mode S. Interrogates Mode S Transponder (d) DME 978MHz. Performs DME tests on channel 17X, 108.00 MHz (e) DME 1104 MHz. Performs DME tests on channel 17Y, 108.05 MHz (f) DME 979 MHz. Performs DME tests on channel 18X, 108.10 MHz

(2)

DME Toggle Switches. (a) IDENT. Adds identification tone (b) To/From. Selects velocity direction

(3)

Range/Velocity Rotary Switch (for use with DME tests only). (a) 0 miles 0 knots (b) 10 miles 10 knots (c) 25 miles 25 knots (d) 90 miles 90 knots (e) 100 miles 100 knots (f) 130 miles 130 knots (g) 180 miles 180 knots (h) 300 miles 300 knots

l. T-49 Series Front Panel Features. (1)

Function Rotary Switch. (a)

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XPDR. Interrogates ATCRBS and/or Mode-S transponder. Performs a preprogrammed test sequence which will continue unless a failure occurs (error message displayed) or until the test is completed successfully (relevant data displayed).

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AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34-NAVIGATION 34.02 ATC TRANSPONDER SYSTEM (CONT’D) (b)

TCAS. Simulates the approach of a TCAS equipped intruder. Works with scenario selector to provide different intruder scenarios. The parameters for each scenario are screened on the front panel. Provides Mode-S replies in response to Mode-S interrogations. Response to uplink format with a Mode-S reply which indicates a simulated intruder is TCAS equipped. Also supplies Mode-S squitter transmissions and simulated replies, not in response to interrogations.

(c)

Mode-S. Simulates the approach of a Mode-S equipped intruder. Works scenario selector as described in the previous “TCAS” description. Also supplies Mode-S squitter transmissions and simulated replies, not in response to interrogations.

(d)

ATCRBS. Simulates the approach of an ATCRBS equipped intruder. Works with the scenario selector as previously noted in the “TCAS” description. Provides ATCRBS/C replies in response to valid ATCRBS/C interrogations.

(2)

Scenario Selector. Simulates the approach of an intruder aircraft scenario. The altitude and closing rate of speed for each scenario is screened on the front panel. The T-49 Fixed Scenario Selector selects a fixed intruder at 4 nmi at ±1000 ft. relative altitude. The T-49C/CF altitude offset selector selects the positive or negative altitude offset for the scenario selected. Also note that the T-49C/CF can stop (fixed intruder mode) and start any scenario in progress by pressing the Test Switch.

(3)

Data Storage Switch. Press after interrogate mode is completed test data in RAM. This data can later be downloaded into a PC. Press after test mode is completed to repeat the current test.

(4)

Setup & Test Procedures. Carefully unpack and inspect the units for damage. Release the latches securing the cover and check inside the cover compartment for accessories such as the line cord, omni-directional dipole antenna, directional antenna, and the appropriate antenna couplers (depends on the test set model) and User Manual. Complete ATI Recurring Non-Routine Job Card 34.02-1 in accordance with the appropriate test set user manual. If no manual is supplied with test equipment, refer to the ATI Test Equipment Reference Manual or contact Maintenance Control or the Director of Quality Control.

34.02.04

05-11-07

Revision 28

AIR TRANSPORT INTERNATIONAL DC-8 MAINTENANCE MANUAL CHAPTER 34-NAVIGATION ATI

RECURRING NONROUTINE

A/C NUMBER

TAT

TAC

POS

DATE

STATION

ATI-001

JOB CARD NUMBER

34.02-1 JOB TITLE

FAR 43 ATC TRANSPONDER SYSTEM RAMP TEST RECORD Job Card Reviewed By

Date

Page 1 of 1 Mechanic

Inspector

A. Record the following transponder information: S/N: Type: Reason for test: Note: Refer to Section 2 of the applicable test equipment user manual for procedural reference. B. Mode C and Mode S: FAR 43 App. F Paragraph (a)

Record Readings

Description

Frequency Check (1090.0 MHz ±1)

freq. =

MHz

(b)

(1)

Suppression: P1 = P2

Pass/Fail

(b) (c)

(2) (1)

Suppression: P2 > P1 Receiver Sensitivity:

Pass/Fail Mode A =

dbm/dbw

(A,C = -73 ±4 dbm/-103±4 dbw)

Mode C =

dbm/dbw

Mode S =

dbm/dbw

(S = -74 ±3 dbm/-104±3 dbw) (c)

(2)

(d)

(1)

Difference between Mode “A” and Mode “C” (Dif