# AC 20-78: Maintenance Inspection Notes for McDonnell Douglas DC-8 Series Aircraft - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Historical - Approved: July 11, 1972 - Last reviewed: September 23, 2024 - This change: Canceled by memo signed by Steven Douglas (AFS-300), dated 12/20/2011. - Family: AC 20-78 - 14 CFR parts: 39 - Pages: 58 - PDF: https://openairdocs.org/files/sha256/07/80/07801fbf59eadb86d92cf3008114b169a514416bb6356586baf189edb8da11a3 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/CA9E23878F964EC6862569AE006D3C82.0001 - DRS record last changed: November 22, 2025 - Page: https://openairdocs.org/ac/20-78 ## Versions - [AC 20-78](https://openairdocs.org/ac/20-78) (Historical), approved Jul 11, 1972, 58 pages. Canceled by memo signed by Steven Douglas (AFS-300), dated 12/20/2011. ## Text Each page's text from the PDF; open the PDF at a page with #page=N. This is OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. ### Page 1 (OCR) AC NO: 20-78 DATE: 11 July 72 **OF TRANSPO ADVISORY** YES OF AMER CIRCULAR MAINTENANCE INSPECTION NOTES FOR MCDONNELL DOUGLAS DC-8 SERIES AIRCRAFT **DEPARTMENT OF TRANSPORTATIONC** **FEDERAL AVIATION ADMINISTRATION** ### Page 3 (OCR) AL NO: 20~78 NA DE TRANSPORT ADVISORY DATE: 11 July 72 **AND TRES ARE CIRCULAR** **DEPARTMENT OF TRANSPORTATION** FEDERAL AVIATION ADMINISTRATION SUBJECT: MAINTENANCE INSPECTION NOTES FOR McDONNELL DOUGLAS DC-8 SERIES AIRCRAFT 10 PURPOSE. This handbook provides maintenance inspection notes which can be used for the maintenance support program for certain structural parts of the DC-8 series aircraft. **2. REFERENCES.** a. FAA Advisory Circular 20-50, Ultrasonic Testing January 1967. b. FAA Advisory Circular 20-61, Nondestructive Testing for Aircraft, May 1968. c. FAA Advisory Circular 65-9, Airframe and Powerplant Mechanics General Handbook 1970. d. FAA Advisory Circular 65-12, Airframe and Powerplant Mechanics Powerplant Handbook 1971. e. Douglas Service Magazine, Volume XXIII, Issue No. 2 1965. f. Douglas Service Magazine, Volume XXIV, 1966. DESCRIPTION. Maintenance inspection matters on the wing and fuselage are reviewed with a view toward supplementing information currently available. **40 HOW TO GET THIS PUBLICATION.** a. Order additional copies of this publication from: Department of Transportation Distribution Unit. TAD-484.3 Washington, D.C. 20591 Initiated by: FS-303A ### Page 4 (OCR) AC 20-78 11 July 72 Identify this publication as: Advisory Circular 20-78 Maintenance Inspection Notes for McDonneil Douglas DC-8 cameligin Series Aircraft. **C. R. MELUGIN, JR.** Acting Director, Flight Standards Service Page ii ### Page 5 (OCR) 11 July 72 AC 20-78 TABLE OF CONTENTS Page No. **CHAPTER 1. MAINTENANCE INSPECTION NOTES. 1** 1. Introduction. 2. Description. Background. 4. General Information. 5• Type of Construction. 6. General Visual Inspection Tips. Nondestructive Testing (NDT). Aircraft Station Diagrams. 9. Definitions. 10. Abbreviations Used In This Document. **CHAPTER 2. WING.** 11. DC-8 Aircraft Maintenance Information. **CHAPTER 3. FUSELAGE. 11** 12. DC-8 Aircraft Maintenance Information. 11 CHAPTER 4. FLIGHT CONTROLS. 13 13. DC-8 Aircraft Maintenance Information. 13 CHAPTER 5. LANDING GEAR. 15 14. DC-8 Aircraft Maintenance Information. 15 CHAPTER 6. PYLONS/NACELLES. 17 15. DC-8 Aircraft Maintenance Information. 17 CHAPTER 7. STABILIZERS. 19 16. DC-8 Aircraft Maintenance Information. 19 CHAPTER 8. MISCELLANEOUS. 21 17. DC-8 Aircraft Maintenance Information. 21 Page iii ### Page 6 (OCR) AC 20-78 11 July 72 Page No. CHAPTER 9• CORROSION PRONE AREAS. 23 18. General. 19. Exhaust Trail Areas. 20. Engine Frontal Areas and Cooling Air Vents. 21. Battery Areas and Vents. 22. Lavatories, Galleys, and Cabin Floors. 23. Wheel Wells and Landing Gears. 24. Control Surface Recesses. 25. Spot-Welded Skins. 26. 21. Lap Joints Between Aluminum Extruded Sections. 28. Fluid Entrapment Areas. 29. Electronic Package Compartments. 30. Control Cables. 31. Microbiological Fuel Contamination and Corrosion. 26 CHAPTER 10. MAINTENANCE INFORMATION. 27 32. DC-8 Aircraft Maintenance Information. 27 APPENDIX 1. DESCRIPTION OF STATIONS. (3 pages) 1 APPENDIX 2. STATTON CHARTS. (6 pages) 1 Figure 1. Basic Dimensions. Figure 2. Fuselage Station Charts. Figure 3. Wing Station Charts. Figure 4. Pylon/Nacelle Station Charts. Figure 5• Horizontal Stabilizer Station Charts. Figure 6. Vertical Stabilizer Station Charts. APPENDIX STRUCTURAL ITEMS. (8 pages) 1 Figure DC-8 Aircraft Significant Structural Items. Figure Fuselage Aft Pressure Bulkhead. Figure 9• Freon Compressor Exhaust Door Opening. Figure 10. Forward Passenger Door Area Figure 11. Aft Passenger Door Area. Figure 12. Aft Lower Wing Skin at Centerline Splice. Page iv ### Page 7 (OCR) 11 July 72 AC 20-78 CHAPTER 1. MAINTENANCE INSPECTION NOTES INTRODUCTION. This advisory circular provides maintenance information which can be used by mechanics, repair agencies, owners, and operators in developing maintenance programs, making improvements in existing programs, and conducting inspections and repairs on certain structural parts of the McDonnell Douglas DC-8 series of airplanes. The material is based, in part, upon information made available through discussions with personnel who have maintained these types of airplanes for thousands of hours of time in service. The intent of the circular is to impart some of this knowledge to other interested persons so that it is not lost. DESCRIPTION. This circular contains guidance material for performing inspection and maintenance on wings, fuselage, empennage, flight controls, pylons/nacelles, and landing gear structures. The information has been derived from service experience. It does not comprise a full and complete maintenance program for the subject aircraft but should be considered as supplemental maintenance data. Included in the circular are diagrammatic sketches and station identifications of the wing, fuselage, and flight controls. In addition, there is a listing of selected maintenance difficulties which were reported during 1970 and 1971. Appendix 3 contains special structural items which are recommended for rework and/or periodic inspection to assure continued structural integrity of high time DC-8 aircraft. **BACKGROUND.** Aircraft Use. The Administrator has realized that several different types of transport aircraft are being phased out of service by some airlines because of the availability of newer equipment. Such older aircraft are being purchased by other operators who may not be familiar with the scope of required maintenance and the means which have been used to keep the aircraft in a safe and airworthy condition. b. Maintenance "Know How." Since maintenance "know how" is not transferred with the aircraft, the new operator generally goes through a learning cycle before he is able to rapidly pinpoint the critical problem areas of the aircraft. In this respect, identification of known areas where structural problems have been experienced will help in the preparation of an initial maintenance program by a new operator. It can also serve as a guide to other operators who have not accumulated sufficient experience to have knowledge of all the problem areas of the aircraft. Chap 1 Par 1 Page 1 ### Page 8 (OCR) AC 20-78 11 July 72 **4. GENERAL INFORMATION.** Manufacturer's Bulletins. It must be emphasized that the manufacturer has published several service bulletins concerning the inspection, repair, and modification of McDonnell Douglas DC-8 series aircraft. Service bulletins highlight the importance of maintaining structural integrity on aircraft with particular reference to areas known to have experienced crack and corrosion damage. Operators are urged to become conversant with the manufacturer's recommendations and make certain that responsible. maintenance personnel are knowledgeable on this subject. Airworthiness Directives. It is emphasized that the material in this circular does not supersede any of the requirements of airworthiness directives issued under Part 39 of the Federal Aviation Regulations. Certification Basis. The United States type certificate holder is McDonnell Douglas Corporation, Long Beach, California. The basic design and first production airplane was the DC-8-ll Model approved for transport category by the FAA on August 8, 1959. The Type Certificate Data Sheet No. 4A25 prescribes conditions and limitations for which the type certificate was issued. However, the data sheet may be revised from time to time as conditions and limitations change. Thus, it is important that the latest data sheet revision be available whenever a problem arises that may be covered in the data sheet. Twenty-one models of the DC-8 have been manufactured, the basic being the DC-8-11 and the latest being the DC-8-63F. WIPE OF CONSTRUCTION. The major structural components of the aircraft are the wing group, the fuselage group, and the tail group. The primary structure, consists of the fuselage, pylons, wing, and the vertical and horizontal stabilizers. These components have been designed in accordance with the "fail safe" policy. This policy provides miltiple load paths and reduced stress levels, such that a single failure of any structural element will not result in the loss of the complete structural component. Several flap system components meet the "safe life" requirements and the life limits for these parts are listed in Note 3 of Type Certificate Data Sheet No. 4A25. The wing is an all-metal, full cantilever-type structure and con- Lime to the dihedrel and sweepback at station Key69, 5, 440d, and sists of a center section which extends from the airplane center- board section extending from the sweepback to station the outboard section from station Xw408 to the wing tip. The primary structure consists of three spars, ribs, bulkheads, and paneling of multiple stringers and stressed plating. The areas between the front and rear spars are used as integral fuel tanks. The wing has four high lift leading edge slots and provisions are Chap i Page 2 Par 4 ### Page 9 (OCR) 11 July 72 AC 20-78 incorporated in the structure for attachment of flaps, spoilers, and ailerons. The leading edge and tips are the only portions not permanently joined to the wing structure; all other portions are not disassembled except for major repair of primary structure. The fuselage is an all metal structure composed of stressed aluminum alloy skin attached to transverse frames and longitudinal stiffeners. It is divided into three main sections: The nose section (station 40 to 310), the center section (station 310 to 1490), and the tail section (station 1490 to 1795). Where plating of .071 inch thickness or less is attached ment and beaded doublers are used in the window belt area forward of the front spar. The frames are generally located at 20-inch intervals throughout and longitudinal fuselage splices are reinforced with beaded inner doublers and pi-section longerons. (3) A keel is used to carry structural loads in the main gear wheel-well area and extends from station 680 to 1090 with a splice at station 920. The fuselage is fully pressurized above the floor from station 109 to the aft pressure bulkhead at station 1490, and below the floor in the cargo and accessories compartments. The cargo compartments are Class D compartments, and by design, will not support combustion which eliminates the need of having fire extinguishing equipment in the cargo compartments; they are also heated to prevent freezing. The gylons are multiple spar and bulkhead structures of high strength aluminum alloy, steel, and titanium. The engine mounts are connected to the pylon bulkheads and loads are transferred to the wing through fittings, angles, and clips. d. The stabilizers, vertical and horizontal, are full cantilever portions of the empennage section. The vertical stabilizer is a fixed portion of the fuselage with a removable leading edge and tip. The horizontal stabilizer is a removable three-section unit consisting of a center section and two outboard sections with operating bulkheads between the center and outboard sections. The leading edge and tips of the outboard sections are removable. The vertical and horizontal stabilizers contain fittings for rudder and elevator attachment. Chap 1 Par 5 Page 3 ### Page 10 (OCR) AC 20-78 11 July 72 GENERAL VISUAL INSPECTION The primary structure of the aircraft is designed to provide resistance to variable forces imposed while in operation by dispensing the forces through a structural pattern of "force flow" to the primary structural members of the wing and fuselage. the obviously failed area often indicate secondary damage caused by transmission of stress from the failed area. Misalignment of doors and panels may indicate distortion of internal structure. Internal structural damage, although not always apparent, may be found by closely examining the exterior surface. For example: a. Buckled skin between rivets at the end of a stiffener or stringer could mean that the last attaching rivet has failed, or that the stiffener or stringer is buckled in the area of the skin buckle. when a detailed inspection of the failed area is to be performed, functional parts should be actuated to determine if the failure has caused binding. b. Deep diagonal skin buckles, located over a frame, former, or rib, could mean the member is distorted. when doubt exists concerning internal condition, the area in question should be opened and carefully inspected. 7• NONDESTRUCTIVE TESTING (NDI). Nondestructive testing is a maintenance inspection procedure which includes utilization of such maintenance tools as X-ray, ultrasonic, magnetic particles, eddy current, dye penetrant, and others. a. Maintenance Inspection. NDT permits maintenance inspections without removing components from aircraft or tearing down complex assemblies. Defects in various aircraft systems which would escape detection through normal visual inspection can be identified by NDI. Training Required. Special NDT training is desirable to make sure that the operator is capable of operating the equipment and interpreting the results. Also, many states require that an X-ray operator have an approved certificate for use of X-ray in industrial applications. This is to minimize improper use with attendant health hazard of X-ray equipment. 8. AIRCRAFT STATION DIAGRAMS. The wing, fuselage, and empennage station diagrams included in this document were developed for the DC-8 aircraft and are used as a general reference only. Several models of these aircraft were manufactured and have different station locator numbers based on the particular configuration. Since the defective areas generally apply to all models of aircraft, the referenced area can be compared with a similar area and locator on the appropriate station diagram for the particular model of aircraft. Chap 1 Page 4 Par 5 ### Page 11 (OCR) 11 July 72 AC 20-78 9• DEFINITIONS. Fatigue. The progressive fracture of a metal by means of a fault which develops and spreads under repeated cycles of stress. b. Stress. The force per unit area of a body that tends to produce a deformation. c. Stress raiser. A scratch, groove, rivet hole, forging defect, or other structural discontinuity giving rise to & focal point for a local concentration of stress. d. Corrosion. Gradual destruction of a material by chemical action. Often evidenced by oxide buildup on the surface. (See Chapter 9) **10. ABBREVIATIONS USED IN THIS DOCUMENT.** AOL All operators letter AD Airworthiness Directive F/S Fuselage station MIG Main landing gear NIG Nose landing gear P/N Part Number SB Service Bulletin W/S Wing station I/B Inboard 0/B outboard R/W Right wing I/W Left Wing Chap 1 Page 5 (and 6) Par 9 ### Page 13 (OCR) 11 July 72 AC 20-78 CHAPTER 2. WING 11. DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing oi significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. ixcept where noted, wing structural references apply to both left and right wings. a. McDonnell Douglas Center Wing Section Reference For Details The vertical flange of the inner SB 57-9 wing bulkhead lower chord member 6/28/60 found tracky. Inspotion bhould at station X also include upper chord member. (2) The forward tang of the lower SB 57-37 front spar cap has been found 1/29/65 cracked at W/S Xew 46.000. (3) Cracks have been found in the web SB 57-58 of the wing trailing edge panel 1/19/68 between stations Xw 5.812 and XwI 60.000. Corrosion has been found on the SB 57-31 inner surface of integral fuel 8/23/63 tank access doors and doubler Reissued lands on aircraft with approxi- 11/1/64 mately 9000 hours flying time. The corrosion was found on the topside doors and lands. b. Inboard Wing Section. (1) Leading edge formers have been SB 57-11 Pound cracked at the upper and 5/19/60 lower segments of the aft end. The cracks were between 1/4" to 2 3/4" in length. Inspection of these formers should be from station Xes 138.750 through Xes 245.000. Cracks have been found in the bulk- SB A57-7 head upper and lower caps at station Reissued Ars 139.000 just forward of the rear 9/24/64 SB 57-7 Rev 1 7/25/60 and AD 60-8-3 Chap 2 Far 11 Page 7 ### Page 14 (OCR) AC 20-78 11 July 72 (3) An alert bulletin was issued due to cracks SB AST-17 being found in the lower wing flap support Reissue No. 2 fitting chord member in the bulkhead at 3/20/61 station Ars 358.000. The cracks occurred in the bulkhead chord member between the barrel nut retainer hole and the end of the part. Cracks have been found in the lower lead- SB 57-19 ing edge skin of the pylon stub wing Revision No. 2 adjacent to the I/B side of each 1/B pylon. 1/26/67 The cracks emanate from the inside radius of the skin cutout at the intersection of the edge of the skin and the pylon. This also applies to the outboard wing section leading edge skin inboard of the 0/B pylons. Several operators have found cracks in the SB 57-30 lower front spar cap tang at W/S Xis 67.531. Rev 2 The cracks occurred on the forward tang of 10/31/67 and the spar cap at the splice. Cracks have SB 57-34 also been found in the upper front spar cap Rev 2 tang at the splice at W/S Xps 67.531. 10/27/67 Loose attachments in the top trailing edge SB 57-41 skin have been reported by operators on 9/7/66 aircraft with more than 10,000 flight hours. The manufacturer modified all applicable models prior to delivery of fuselage #285 and sub sequent. The modification installs a doubler from W/S Xw 70.000 to Xw 113.000. A service bulletin was issued to inspect for SB 57-46 cracks in the I/B aft corner of the leading 3/6/67 edge upper slot door assemblies. This bulletin applies to aircraft prior to fuselage #315. c. Outboard Wing Section. (1) An inspection and rework bulletin was issued SB A57-15 as a result of a report of an aileron hinge Reissue #l fitting lug breaking off at station Xa 448. 8/8/60 Electrical junction box support clips on the SB 57-16 rear spar have been found cracked between 11/29/60 station Yrs 470 and Xrs 479. The cracks occurred in the clip lobe's at the rear spar attach point. Page 8 Chap 2 Par 11 ### Page 15 (OCR) AC 20- 78 11 July 72 (3) Operators have reported cracks in the SB 57-25 stub wing leading edge of the 0/B Rev 2 pylons in the area of the pneumatic 8/27/62 duct cutout. (4) Cracks have been found in the wing SB 57-48 leading edge lower structure between 3/3/67 the opening for access door 520J and percent wing leading edge. 526.760 on aircraft with four Cracks have been found in the wing SB 57-59 leading edge lower surface skin on 4/11/68 aircraft with 15,000 to 20,000 flight hours. The cracks developed between access door 520J and W/S X 454.000 on aircraft with standard wing leading Par 11 Chap 2 Page 9 (and 10) ### Page 17 (OCR) 11 July 72 AC 20-78 CHAPTER 3. FUSELAGE 12. DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following 1s & listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. Except where noted, fuselage structural references apply to left and right sides of the fuselage. Tail cone structure has been found cracked. SB 53-17 The cracks are a result of sonic vibration 10/5/60 which is prevalent in the tail cone area SB 53-11 during operation of the aircraft. Rev 2 5/1/61 SB 53-13 Rev 2 3/24/61 SB 53-19 11/1/60 SB 53-25 Rev 1 2/27/62 b. Cracks have been reported in the aft vertical SB 53-21 stiffeners of the 0/B MG wheel well webs. 3/24/61 c. Cracked fuselage to skin shear clips have been found cracked on aircraft with 5,000 hours Rev 2 SB 53-28 flight time. The affected areas are at F/S 6/26/53 Y 640.000 between longerons 22 and 23 and F/S Y 660.000 between the cusp line and longeron 23. d. Cracks have been found in longeron 24 at F/S SB 53-32 980.000 on aircraft with 10,000 or more Rev l flight hours. 9/17/65 e. Cracks in the MG I/B door actuating cylinder support fittings have been reported. SB 53-33 1/21/66 fittings are located on the left and right sides of the keel beam at F/S 920.000. Cracks in the domed aft pressure bulkhead have been reported. Cracks were a result of SB 53-41 10/17/67 fatigue and occurred at approximately 15,000 flight hours. Cracked vertical angles located at the upper SB 53-46 portion of the aft flat pressure bulkhead 3/28/68 have been reported by one operator on an aircraft with 14,000 flight hours. Chap 3 Par 12 Page 11 ### Page 18 (OCR) AC 20-78 11 July 72 h. Fuselage skin corrosion has been found on SB 53-53 an aircraft with approximately 25,000 12/20/68 flight hours. The corrosion was in the scuff plate area of passenger, service, and cargo doors. Fuselage frame fittings have been found AOL 8-169A corroded on aircraft with between 21,000 8/30/68 and 30,000 flight hours. The corroded fittings were located at F/S 920 and 960. Chap Page 12 Par 1 ### Page 19 (OCR) 11 July 72 AC 20-78 CHAPTER 4. FLIGHT CONTROLS 13. DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. Except where noted, ilight control structural references apply to left and right sides. a. Flap support fittings on the aft side of SB A27-5 the rear spar have been found cracked. Reissue #2 Examination indicated stress corrosion 3/1/61 developed and cracks propagated from the corroded area. b. Operators have experienced cracking of the SB 27-48 rudder nose rib webs at the three hinge point locations and in the rib to spar attach angles, (Stations Zy 105.000, Zy 162.906, and Zr 220.813). c. An alert bulletin was issued by the manufacturer SB A27-66 due to a report of a left 0/B elevator geared 6/20/60 tab push rod being broken at the aft rod end fitting, rod P/N 4702890. d. The aileron reversion mechanism forward gripper SB 27-95 has been found cracked at the lower clevis ends Reissue #l which attach to the end of the aileron tab 5/18/62 lockout cylinder piston end fitting. e. Clevis P/N 2619862 on the rudder lockout SB A27-100 cylinder was found broken. This allowed Reissue #2 the rudder to revert to manual with hydraulic 9/24/64 and power still on causing the rudder to lock AD 61-6-4 in its last position. f. Cracks have been found in the aileron tab SB 27-115 lockout mechanism support brackets Rev #1 10/25/01 P/N 4643350. and AD 62-2-4 g. Wing flap actuating cylinder barrels have SB 27-118 been found cracked. Examination of failed Reissue #2 cylinders disclosed corrosion in the internal 12/4/61 portion of the barrel. Wing flap support track assemblies have been SB 27-123 found cracked at W/S Xp 219.498. Analysis 12/27/61 indicated the cracks could have been caused by stress corrosion. Chap 4 Par 13 Page 13 ### Page 20 (OCR) AC 20-78 11 July 72 Wing flap rod end bearings have been found SB 27-127 cracked with more than 5000 hours. Rev. #2 The cracked bearings were located at W/S 1/28/64 and 98 and 219 in the area of the bearing case AD 62-12-4 adjacent to the bearing dust shield. j• Wing flap actuating cylinder barrels have SB 27-134 failed. Both mid-wing and 0/B actuating Reissue #1 cylinders have been found cracked. 10/19/62 SB A27-134 Reissue #2 10/13/64 and ADS 62-16-2 and 62-24-4 Wing flap link assemblies P/N 5644642-501 SB 27-137 at W/S X 97.906 and P/N 5644644-501 at Rev. #1 W/S Xp 219.498 have been found broken on 1/8/64 several occasions. 1. Cracks have been found along the horizontal SB 27-142 centerline of the forward attach lugs of the flap Reissue #l link support assemblies on aircraft with 5/15/63 more than 8000 hours, m. Operators have experienced failure of the SB 27-159 0/B wing slot assembly, P/N 3770143-1, on Reissue #1 aircraft with four percent extended wing 2/23/65 leading edge. n. Cracked flap hinge support fittings have been SB 27-217 found in the 0/B flap center hinge base Rev. #1 assembly on aircraft with between 8000 and 10/9/68 11,000 hours.. o. Rudder link, P/N 2619201, failures have SB 27-218 occurred on aircraft with 23,000 to 26,000 Rev. #1 hours. 10/18/68 Cracked aileron tab rod end assemblies have SB 27-223 been found on aircraft with 3,000 to 29,000 12/6/68 hours. Chap 4 Page 14 Par 13 ### Page 21 (OCR) 11 July 72 AC 20-78 CHAPTER 5. LANDING GEAR DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. Except where noted, landing gear structural references apply to left and right landing gear. Intergranular corrosion and cracking have SB 32-28 been reported on main and nose gear Rev. #1 aluminum 'B" nuts on landing gear pipe 11/13/61 assemblies. Cracks have been found in the MIG retract SB 32-51 strut barrels. The cracks occurred in a Reissue #1 flash weld area of the barrel. Rev. #1 6/28/61 c. Operators have reported failure of the bolt SB 32-53 assembly attaching the aft brake compensating Reissue #1 links to the MIG piston. 5/4/61 Cracked and broken MIG bogie beam pivot bolts SB. 32-60 have been found which appeared to be caused Reissue #l by lack of lubrication. 8/3/61 Rev. #1 2/27/62 MIG bogie beam trim cylinder forward attach SB 32-61 bolts have failed allowing the bogie beam Reissue #l to rock freely without snubbing. 2/2/62 f. MIG actuating cylinder upper attach brackets SB A32-76 P/N 5641950-1, -2, -501, or -502 have been Reissue #l found broken. Rev. #l 4/5/62 and AD 62-23-3 MLG bungee cylinder rod-end eyebolts have been SB A32-87 found cracked in the threaded area. Byebolt Reissue #1 P/N 4772490-1. 5/4/62 h. Fatigue cracks have been found in the support SB 32-89 flange and O-ring groove of the MIG retract Reissue #l cylinder gland assembly on aircraft with 7/16/63 5,000 or more hours flight time. Cracks have been found in the MIG forward axle SB 32-91 beam pivot arm upper lug. 11/12/62 Chap S Par 14 Page 15 ### Page 22 (OCR) AC 20- 78 11 July 72 J. MIG forward bogie beam swivel pin lower lugs McDaco letter have been found cracked and failures of bogie C1-78-987/WBM beams in "critical area" on the bottom of Rev. #1 7/17/64 certain forward bogie beams have occurred. and C1-78-228 The failures were determined to be a result 2/16/65 and of corrosion. AD 68-26-7 Failures of the nose gear forward door operating SB 32-104 mechanism bellcrank have occurred on aircraft 9/17/64 having more than 7,000 hours flight time. 1. Several operators reported nose gear uplatch SB 32-112 upper link assemblies broken at the lower 11/8/65 attach point. m. Malfunctions of the MIG uplatch hook have been SB 32-113 due to inadequate lubrication of the pivot bolt Rev. #1 bearings. Aircraft fuselage #348 and subsequent 8/10/67 were modified prior to delivery. The modification replaced the pivot bolt with one that permits lubrication without removing the bolt. Some operators have reported binding of the SB 32-114 bearing in the lower support bracket of the MIG 3/23/66 I/B door latch due to corrosion. 0. There have been numerous instances of minute SB 32-140 cracks in the machine threaded root of the 11/6/68 main gear bogie trim cylinder eyebolt. These cracks were detected on aircraft with approximately 15,000 flight hours. Manufacturer recommends replacing bolt, with a bolt incorporating rolled threads. This applies to series 60 aircraft; fuselage #445 and subsequent were modified prior to delivery. P. Failures of landing gear bogie beams have AD 64-5-2 occurred due to cracks initiated by corrosion in areas where the bogie beam surface had been damaged. Page 16 Chap 5 Par 14 ### Page 23 (OCR) 11 July 72 AC 20-78 CHAPTER 6. PYLONS/NACELLES DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. Except where noted, pylon/nacelle structural references apply to left and right sides of the aircraft. Access doors have been found cracked on SB 54-7 the R/H side of nacelles in the area of 5/18/60 cooling air exit ducts. Cracks have been found in the tunnel and SB 54-17 junction box of each pylon. The cracks Rev. #1 occurred due to the welded seam which 9/26/61 joins the tunnel to the junction box. Cracks have been found in the zee angles SB 54-20 located on the left side nacelle door in Rev. #1 the area of the upper aft pylon drain 8/7/61 fitting support. d. Lower 0/B pylon-to-wing attach angles and SB 54-24 flexible ribs below pylon spars have been Reissued T/25/62 found cracked. and Revised 8/20/62 e. The vertical leg of the 0/B spar cap on SB 54-27 the O/B pylon upper spar was found cracked 1/26/62 at station Yop 246.000. 1. Failure of the upper I/B spar cap structure SB A54-33 has occurred in the O/B pylons in the area Rev. #2 1/24/64 oi station I 230 at the edges of support and AD 62-27-4 fitting P/N 3647306-501. g. Failure of the upper 0/B spar cap structure SB 54-34 has occurred in the I/B pylons in the area of Reissue #l stations Yip 225 to Yip 233. 7/6/64 and AD 65-5-2 h. Cracks have been found in the I/B and 0/B Daco Service pylon stub wing structure. The cracks have Engineering occurred in skins, bulkheads, and spar caps. Letter C1-78-2016/DBA 10/18/66 SB 54-57 Rev. #1 12/9/69 and AD 67-5-3 Chap 6 Par 15 Page 17 ### Page 24 (OCR) AC 20- 78 11 July 72 Skin and doubler cracks have been found on SB 54-41 the 0/B side of the O/B pylons between Rev. #1 aircraft vith four percent extended wing <30.000 and Top 240.000 on 9/27/68 leading edge. Aircraft fuselage #148 and and subsequent were modified prior to delivery. Cracks have been found on 0/B pylon stub SB 54-42 wings on aircraft with more than 6,000 1/6/67 flight hours. The cracks occurred at station Xw 454.000 and the I/B intersection adjacent to the leading edge lower slot door. Operators have reported cracks in the SB 54-43 Pylon cant bulkhead fitting on aircraft Rev. #1 with 17,000 and 18,000 flight hours. 5/12/67 1. Cracks have been found in the D-duct spars, SB 54-44 adjacent structure, and pylon skin of the Rev. #1 I/B pylons. The reported cracks were on 2/10/67 aircraft with 14,000 to 21,000 flight hours between station Lip 204.000 and Iip 220.000. Cracks have been found in the I/B pylon stub SB 54-49 wing skin and internal doubler at the cutout 8/25/67 for the wing leading edge lower slot door on The cracks were detected in the area aft of Thrust reverser stow latch pivot bolt failures SB 54-52 have occurred. This applies to series 60 3/28/68 aircraft; fuselage #332 and subsequent were modified prior to delivery. Five failures of forward thrust reverser stow SB 54-53 latch on series 60 aircraft at approximately 8/20/68 2,000 flight hours have occurred. An alert bulletin was issued due to reported SB A54-58 cracks in P/N 5753016-1 and -2 wing skin Rev #1 pylon support doubler. The cracks originated 12/16/68 at the forward edge of the doubler in the intersection of the I/B side of the I/B pylons and the wing front spar lower cap at W/S Xw 234.000. Aircraft fuselage #451 and subsequent were modified prior to delivery. Although cracks were not reported in 0/B pylons, the structure is similar and should also be inspected. Chap 6 Page 18 Par 15 ### Page 25 (OCR) 11 July 72 AC 20-78 CHAPTER 7. STABILIZERS 16. DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. ixcept where noted, stabilizer structural references apply to leit and right sides. Cracks have been reported in the formers SB 55-7 of the vertical stabilizer leading edge at 10/18/61 stations Zes 117.250 to Zfs 179.750 and Zfs SB 55-11 186.000 to Zfs 350.500. 5/2/63 b. Cracks have been detected in the radius of SB 55-8 the vertical leg of the horizontal stabilizer Rev. 1 closing bulkhead caps, P/N 5654549-1, -2 8/13/62 -501, and -502. The cracks have been up to 4" in length in the aft end of the caps. Operators have reported cracked formers in SB 55-10 the horizontal stabilizer leading edges 5/2/63 originating at the aft end of the formers. Flying time on aircraft with reported cracks was in excess of 5,000 hours. Fatigue cracking has been reported in the SB 53-45 vertical stabilizer front spar to fuselage Rev. #1 attach fitting at F/S 1500.000 on aircraft 3/12/68 with 15,000 to 20,000 fight hours. Chap ar 16 Page 19 (and 20) ### Page 27 (OCR) 11 July 72 AC 20-78 CHAPTER 8. MISCELLANEOUS 17. DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is a listing of significant maintenance difficulties that have been reported by air carriers. This information may be useful in identifying structural inspection areas. a. Cracks have been found in the vertical SB 51-2 flange of the keel beam fittings between Reissue #1 the center and rear spar. 1/3/61 Operators have reported cracks in fuselage SB 51-7 frame at refrigerant condenser forward 9/3/63 exhaust flap hinges. C. Forward and aft passenger entrance door SB 52-14 mechanism housings have been found cracked. 9/26/60 The cracks were a result of maladjustment of the latching mechanism. d. The forward upper cargo door mechanism apex SB 52-47 pin assembly has been found cracked on aircraft rev. #1 with approximately 8,000 hours. 8/16/68 Chap 8 Par 17 Page 21 (and 22) ### Page 29 (OCR) 11 July 72 AC 20-78 CHAPTER 9. CORROSION PRONE AREAS 18. GENERAL• Corrosion is most likely to strike at the joints or attachment points in metallic structure because they provide entrapment areas for corrosive agents and moisture to accumulate. However, the problem can be compounded if metallic surfaces are exposed to an extremely corrosive environment, or if the surfaces are difficult to inspect and clean. The most important corrosion preventive measure is to keep them clean and dry. 19. EXHAUST TRAIL AREAS. Both jet and reciprocating engine exhaust deposits contain compounds that are very corrosive, and the structure in the path of exhaust gas is more likely to suffer from corrosion than is any other structure. It is customary to coat structure exposed to exhaust gases with protective finishes to prevent these gases from coming in direct contact with the metal. However, the deposits which collect on top of the protective finishes must be removed before they permeate the film and attack the metal. Most troublesome are those areas where exhaust gas deposits may become entrapped and cannot be reached by normal cleaning methods. Typical of these areas are seams, gaps, hinges, or fairings located in the exhaust gas path. 20. ENGINE FRONTAL AREAS AND COOLING AIR VENTS. These areas are subject to erosion by airborne contaminants, rain, and from foreign objects on the runways. Erosion will remove the protective finishes or oxide film from the metal surfaces, leaving them vulnerable to corrosive attack. In addition, much of the equipment installed within these recesses is also vulnerable to corrosion. when an aircraft is operated in a marine environment, salt deposits may accumulate in these areas and the ensuing corrosive attack can be rapid and destructive. It is imperative that these areas be frequently inspected and cleaned, and protective finishes be maintained. 21. BATTERY AREAS AND VENTS. Due to the highly corrosive nature of battery acid and its fumes, battery areas and their vents are protected by special acid-resistant paints. Generally, this is enougn to stem the tide of corrosion if scrupulous attention is devoted to keeping these areas clean. 22. LAVATORIES, GALLEYS, AND CABIN FLOORS. The usual spillage, condensation, and other contamination of these areas are extremely corrosive to aircraft metals. The most corrosive agents are acidic foods and beverages, and human excreta. Also, most chlorinated disinfectants are acidic and corrosive, and cannot be recommended for use in aircraft. It is impractical to assume that these areas can be kept clean and dry at all times, but it is important to inspect the structure carefully at suitable intervals, cleaning and renewing the finish as necessary. Areas behind lavatories, sinks, or ranges where waste may collect are potential trouble spots, as are personnel reliei and waste disposal vents or openings on the exterior of the aircraft. Chap 9 Par 18 Page 23 ### Page 30 (OCR) AC 20- 78 11 July 72 **23. WHEEL WELLS AND LANDING GEARS.** Equipment installations in the wheel-well areas probably absorb more punishment than any other portion of an aircraft because they are exposed to moisture and flying debris during takeoffs and landings, and when the aircraft is perked, they are exposed to atmospheric moisture. After an airplane has entered service, it is difficult to maintain protective paint film on landing gears, equipment installations, and wheel well surfaces because the many complicated shapes, assemblies, and fittings in these areas obscure other surfaces. Items that should receive special attention during wheel-well inspections are: (1) Magnesium wheels (especially the areas around the bolt heads, lugs, and wheel webs). (2) Portions of rigid tubing obscured by clamps and identification tapes. (3) Exposed electrical equipment. (4) Crevices between ribs, stiffeners, and lower skin surfaces which can serve as water or debris entrapment areas. b. Corrosion control in wheel well areas may best be attained by frequent cleaning, lubrication, paint touch-up and judicious use of wheel covers. CONTROL SURFACE RECESSES. Control surface recesses are potential problem areas because normally they are difficult to inspect. If corrosive agents gain entry and accumulate in these areas, they may go unnoticed for some time. Frequent inspection and cleaning of the surfaces and installations located in these recesses will prevent corrosion from gaining a foothold. SPOT-WELDED SKINS. Corrosive agents may become entrapped between the layers of metal adjacent to the spot-weld beads. If moisture enters the area, an electrolytic cell can be set up between the dissimilar metal phases in the spot-weld area, and one or more of these phases will be subject to preferential corrosive attack. Whenever practical, structural areas with spot-welded assemblies should be sealed to prevent the ingress of contaminants and moisture. HINGES. Hinges are natural traps for corrosive agents. Often they are susceptible to galvanic corrosion when the hinges and pins are made of dissimilar metals. The most practical means of corrosion control is to inspect and lubricate hinges frequently. After lubrication, actuation of the door through several cycles is necessary to ensure complete penetration of the lubricant. Chap 9 Page 24 Par 23 ### Page 31 (OCR) 1l July 72 AC 20-78 **27. LAP JOINTS BETWEEN ALUMINUM ALLOY EXTRUDED SECTIONS.** Extruded sections of high-strength materials such as 7075 and 7178 aluminum alloys are more susceptible to intergranular corrosion than are other aluminum alloys. Application and maintenance of chemical processes, sealants, paints, or combinations of these protective measures have proven most effective in controlling corrosive attack on these materials. -I intergranular corrosion occurs, it is usually founa around fasteners or in lap joints. It is evidenced by raised areas or lumps under the paint film, and there have been instances where the attack has progressed to such an extent that there were actually bulges in the faying surfaces. These raised areas, lumps, and bulges are due to the accumulation of corrosion products, which take up more volume than did the affected material before it corroded. 28. FLUID ENTRA MEVI AREAS. Design specifications require that aircraft have drains installed in areas where water may collect. However, if drains are rendered ineffective because they are clogged by debris, sealant, etc., or because the level of the aircraft is changed from that of a normal ground attitude, corrosive agents can collect in a localized area. Low-point areas and drains should be inspected frequently to prevent the inception of corrosive attack. Low points of integral fuel tanks are areas where water condensate can collect. This water and condensate is of doubtful purity, and if permitted to stand, it can permeate the protective coating of the tank and bring many corrosive agents held in suspension in contact with the vulnerable metallic surfaces. Water condensate should be drained regularly from fuel tanks and the integrity of the tank sealant maintained to prevent corrosion of integral tank surfaces. b. When considering fluid entrapment areas, one inevitably thinks of aircraft drinking water and wash water systems. Corrosion in these systems ia rare because they are fabricated from nonmetallic materials and/or stainless steel. Water stagnation should pose no problems so long as sanitation regulations are heeded. 29. ELECTRONIC PACKAGE COMPARIMENTS. Often the safety of those on board an aircraft rests on the proper function of a little black box. The environment of electronic package compartments is carefully controlled to provide the most ideal conditions that can be achieved. The degree to which such sensitive equipment is exposed to corrosive agents is very small, but even small quantities of moisture and contaminants can adversely affect equipment reliability. Components in these areas should be inspected for corrosive attack as thoroughly as possible during routine checks, and advantage should also be taken of nonscheduled component removals for further inspection. Chap 9 Par 27 Page 25 ### Page 32 (OCR) AC 20- 78 11 July 72 30. CONTROL CABLES. Control cables have preservative coatings which, when intact, prevent corrosive attack. Due to their vital function, it is necessary to inspect these cables frequently to ensure that they are adequately protected. During these inspections, incipient failures due to other causes also may be detected. It is recommended that control cables be inspected periodically, then cleaned and treated as necessary before reapplication of the preservative. 31. MICROBIOLOGICAL FUEL CONTAMINATTON AND CORROSION. Microbiological contamination in turbine fuel is caused by bacteria and fungi which feed on the constituents of the turbine fuel. The result is a sludge, or mat deposit, which has been found in some aircraft fuel tanks and is often loosely referred to as green slime. However, some deposits have been found in various shades of grey, brown, red, and white. If allowed to develop and grow in the aircraft fuel tanks, microbiological contamination can cause a myriad of problems not the least of which is severe aluminum alloy corrosion in the aircraft integral fuel tanks. Since the subject is large and complex, we recommend Douglas Service Magazine, Vol. KAlll, Issue #2, 1905, which contains information on microbiological contamination and other types of corrosion, their causes, and their control. Vol. XXIV, 1966 issue, also contains information on fuel tank microbiological contamination. Chap 9 Page 26 Par 30 ### Page 33 (OCR) 11 July 72 AC 20-78 CHAPTER 10. MAINTENANCE INFORMATION DC-8 AIRCRAFT MAINTENANCE INFORMATION. The following is & listing of selected significant maintenance difficulties that have been reported by air carriers as mechanical reliability reports from 1970 and 1971 records. This information may be useful in identifying additional structural inspection problem areas. In all cases, check the corresponding trouble area, left right, top bottom, forward aft,etc. Wing: During routine inspection a l2" crack was found in the upper R/W skin at W/S Xrs 367. The total time on the aircraft was 30,358 hours. The operators report referred to this as being a corrosion type crack. During overhaul a 6-inch crack was found in the L/W bulkhead upper cap fitting P/N 5615314-1 at W/S Xrs 508. The crack was adjacent to the center spar along the bulkhead cap fitting radius. The aircraft total time 38,513 hours. (3) Routine inspection disclosed a 2-inch crack in the L/W leading edge lower skin P/N 5754700-3 above #2 pylon and a one-inch crack above #3 pylon. A 4-inch crack was also found in the R/W leading edge skin P/N 5754701-3 above #4 pylon. (Reference marufacturer's All Operator Letter (AOL) 8-490 dated 10/28/70.) (4) Routine inspection disclosed a 2-inch crack in L/W front spar upper splice angle at W/S Xps 63, and an 11-inch crack in the R/w front spar forward cap progressing 1/B from W/S Afs 67.5. Aircrait total time 31,992 hours. (5) During overhaul found a 4-inch crack in L/W stringer #8 The crack was located along the radius formed by the vertical and 'Y" legs of the splice. Aircraft total time was 34,888 hours. During overhaul a 2 3/4-inch crack was detected in the L/W lower front spar cap at W/S Xrs 67.5. The crack originated at an attach hole and progressed through two other attach holes. Aircraft total time 17,902 hours. During approach the aircraft had a tendency to roll with wing flaps extended beyond 25° accompanied by a split flap indication. Inspection after landing disclosed a broken compensator link and shaft, P/N 2647101. Aircraft time since overhaul 1,048 hours. Chap 10 Par 32 Page 27 ### Page 34 (OCR) AC 20-78 11 July 72 (8) During inspection an 8-inch crack was found in the R/W flap support fitting, P/N 5614374-2, at W/S Xys 219. The aircraft total time was 34,261 hours. During overhaul a 3-inch crack was found in the L/W support fitting, P/N 5614374-1 at W/S Xrs 219. Total aircraft time was 17,700 hours. (10) During overhaul a 5-inch crack was found in the radius of attach bracket, P/N 3653812-3 located between the I/B support bearing and the R/H 0/B aileron. Total aircraft time 34,345 hours, time since overhaul 10,613 hours. (11) Routine inspection disclosed broken mount lugs on the I/H I/B flap actuating cylinder, P/N 5642074-5015. Time since overhaul was 10,915 hours. Operators report attributed breakage to stress corrosion. (12) One aircraft was taken out of scheduled service and ferried to a maintenance base due to finding broken I/H O/B flap hinge base fittings P/N's 5653754-1 and 5653755-1 at station 219. Total aircraft time 37,788 hours, time since last service check was 329 hours. Another operator reported a 4-inch crack in the R/H 0/B flap hinge base at station 219. Aircrait total time was 37,640 hours. Routine inspection revealed partial separation of the L/W flap follow-up cable at W/S Xps 22.750. Cable P/N 5652981-529, total aircraft time l1,346 hours. Operators report attributed separation to be a result of oxidation. (14) Routine inspection disclosed a cracked R/H O/B flap hinge base fitting at station 339. Total aircraft time was 37,867 hours. (15) During scheduled inspection an Il-inch crack was found in the crank arm of the R/H wing I/B slot torque tube. (16) During routine inspection the L/H wing flap drive link, P/N 5655655-501, was found cracked at station 219. Total time since new was 23,145 hours. During scheduled inspection the R/H flap support bracket, P/N 5614376-2, was found to have a crack approximately 10 inches in length. Total aircraft time was 35,170 hours. (18) A flight returned due to loss of hydraulic fluid. Inspection after landing disclosed a crack in the R/H aileron actuating cylinder, P/N 5613697-5. The crack was along the circumference of the cylinder barrel. Chap 10 Page 28 Par 32 ### Page 35 (OCR) 11 July 72 AC 20-78 b. Lylon. (1) Routine inspection disclosed a 12-inch crack in #4 pylon upper 1/B skin at station Yop 230-240. Total aircraft time was 10,419 hours. During overhaul, angle P/N 5769515-36 and doubler P/N 5769515-155 were found cracked on the I/B side of #3 pylon. Total aircraft time 10,524 hours. During routine inspection a chordwise crack was found in # pylon intercostal P/N 5769523-46. The crack was in the horizontal leg and located 1-inch 0/B of the pylon upper 0/B spar plane. Total aircraft time was 8,091 hours. Routine inspection disclosed the lower spar I/B cap broken in two in #l pylon. The break occurred 12 inches from the found inh crane or spar web at top 248 and 248 and those creast ofted silar ascerpentes 41 pou two skin cracks at I 32,156 hours. at Yop 243 with a total aircraft time of 30,141 hours. (5) Inspection found a crack in the vertical leg of the o/B spar cap at station Yop 220 of #1 pylon. (6) Routine inspection disclosed a cracked top 0/B spar rail at shit in 44 Bylon. Total aircraft time was 34,510 hours. and a 2-inch crack in the web at station Yor A 5-inch crack was found in the 0/B skin of #1 pylon at station Yop 235. Total aircraft time was 8,505 hours. C. Landing Gear. (1) During taxi after landing #3 wheel, P/N 152059, failed at the I/B rim, operator reported a total of 2,297 landings. Another operator reported #3 wheel failure during takeoff, total landings since overhaul was 63. Failure was attributed to cracks which originated at corrosion pits. A flight was interrupted due to inability to retract the MIG after takeoff. Inspection after landing disclosed the L/H MIG down lock bungee cylinder piston rod broken, P/N 4641900. During walkaround inspection found R/H MG bogie beam broken at the lover swivel lug, P/N 5774700. Operator attributed failure to corrosion due to lack of grease penetration and initiated a program of inspection and modification to reduce possibility of additional failures. A similar condition was also found on the L/H bogie beam of another aircraft. chap 10 Par 32 Page 29 ### Page 36 (OCR) AC 20- 78 11 July 72 (4) A flight returned due to NIG not retracting, inspection found the NIG side brace broken at the retract strut attach lugs, side brace P/N 5771405-5. Total aircraft time 21,030 hours. Several operators have reported flight interruptions due to left and right MIG retract cylinder attach pins being broken. These reports showed that the first evidence of malfunction usually an unsafe gear position indication. (6) Routine inspection disclosed a cracked L/H MIG bungee cylinder attach casting, P/N 5719176-5511. Flight crew reported a severe jolt when the landing gear was extended and did not get a dom and locked indication. The aircraft was stopped on the runway and inspection showed the L/H MIG was not locked due to broken retract and bungee cylinders. A flight was interrupted due to a loud roar and vibration when gear was retracted after takeoff, fuel was jettisoned and the aircraft returned. Inspection revealed a broken NIG door link, ₽/N 5646507-501. Total aircraft time was 11,830 (9) One operator reported finding a broken O/B brake link rod on the R/H MG, rod P/N 576390. Total time on rod 10,359 hours. During routine brake change, #7 brake carrier lug was found broken off. A total of 225 landings had been made since last brake overhaul. (11) After takeoff,. the left MIG could not be retracted. Inspection after landing revealed a broken I/B bungee unlock bracket, P/N 6F-8574-501, which prevented the gear from being inlocked from the down position. Total aircrait time was 38,107 hours. (12) The #4 wheel broke on landing on one aircrait, wheel f/N 1>42>1-1. There had been 100 landings accumulated since the last overhaul. A flight returned due to the landing-gear warning light remaining on after gear was retracted. Inspection found a broken pivot lock arm, P/N 5769871-1, on the left MIG. Total aircraft time was 21,220 hours. (14) During scheduled inspection, the R/H MIG forward bogie beam was found cracked at the upper attach lug for the pivot unlock arm. Total aircraft time was 28,269 hours. Chap 10 Page 30 Par 32 ### Page 37 (OCR) 11 July 72 AC 20-78 d. Fuselage. During inspection found severe corrosion on the inner surface of the aft center section lower skin panel P/N 5615372-3, F/S 1150-1290. Also found four cracks 1/2" to 4" in length and light corrosion on longerons and doublers. Total aircraft time was 30,963 hours, time since overhaul 1,801 hours. Routine inspection disclosed a 4-inch crack in the upper aft corner of the rear passenger door jamb skin. Total aircraft time 37,221 hours. (3) During scheduled inspection cracks were found in the fittings at the top of the pressure bulkhead at F/S 857 in the left and right wheel wells, fitting P/N's 3646874-1 and -2. Total aircraft time 7,771 hours. Several operators reporting finding cracks in fuselage former #24, L/H side at F/S 920 and 980, cracks ranged from 2 to 4 inches in length. Total aircraft time ranged from 7,771 hours to 33,091 hours. During overhaul, a 22-inch crack in the 0/B flange radius of the L/H MIG support fitting web, support P/N 5750092-501. Repairs were made in accordance with SB 63-61. (6) Several occurrences of excessive corrosion have been reported on metal floorboards in the forward gally, center aisle opposite forward galley, and in the mid-lavatory areas. Total aircraft time ranged from 8,027 hours to 11,075 hours. (7) Moderate corrosion was found during routine inspection at the following locations on the fuselage skin between stringers 34L and 34R; F/S 640 to 650, F/S 1075 to 1260, F/S 1300 to 1300, and F/S 1430 to 1480. Total aircraft time was 32,135 hours. Routine inspection revealed bulged skin and rivets popped at the following locations: F/S 1100, stringer #36 F/S 1300 at stringer #29, F/S 1420, 1440, 1460, and 1240 at 1400, and 1490 at stringers 27 through 31 right. Total aircraft time was 37,399 hours. (9) Scheduled inspection disclosed corrosion of lower fuselage skin adjacent to fasteners and doublers at F/S 420 to 500 and F/S 1180 to 1300 between stringers 33R and 351. Total aircraft time was 34,510 hours. Chap 10 Par 32 Page 31 ### Page 38 (OCR) AC 20- 78 11 July 72 (10) During routine inspection corrosion was found in the lower aft corner of the forward service door jamb and skin laps. Total aircraft time was 34,510 hours. e. Miscellaneous. (1) An attempt to dump fuel in flight failed. Inspection after landing disclosed a bent I/H dump chute link rod, P/N 3765140. The first officer's windshield outer panel shattered during flight on one aircraft. The aircraft was at 28,000 feet with windshield heat in warm position. (3) During routine inspection, two 5-inch cracks were found in the rudder fiberglas section at station Zy 253.0. Chap 10 Page 32 Par 32 ### Page 39 (OCR) 11 July 72 AC 20-78 Appendix 1 STATIONS DESCRIPTION. The airplane is divided into reference points in inches, designated as station numbers. These station numbers provide a means for quickly identifying the location of components, the center of gravity, and the distribution of weight. There are three axes used in the plotting of station numbers: The lateral or X-axis, the longitudinal or I-axis, and the vertical or Z-axis. a. Glossary of Terms. (1) FRP The fuselage reference plane is used as a basic reference. It extends along the fuselage cusps which are formed by the intersection of the upper and lower sections of the fuselage. The reference plane has an angle of one degree nose down with reference to the ground and is the origin of Z-stations. (2) WRP The wing reference plane contains the trailing edge of the wing and has an angle of dihedral of 6-1/2 degrees to the FRP. (3) VRP The vertical reference plane (plane of symmetry) bisects the airplane through its centerline at an angle of 90 degrees to the FRP. (4) X-axis Divided into segments in inches along the lateral plane. (5) Y-axis Divided into segments in inches along the longitudinal plane. There is a I-axis subsystem designation for all subsystems. (6) Z-axis Divided into segments in inches along the vertical plane. b. Wing Subsystem Designations. MOX (I) B62: This subsystem designation refers to the center, or constant, section of the wing. It consists of stations in inches, parallel to the centerline of the airplane and extending outboard 69.5 inches from its origin at the airplane centerline to its 7x (8) This wing subsystem designation denotes stations in inches, parallel to the centerline of the airplane and perpendicular to the WRP, beginning at the dihedral break and extending outboard indefinitely. ### Page 40 (OCR) AC 20-78 11 July 72 Appendix l (3) Xp This designation represents stations in inches, located perpendicular to the wing flap hinge line. (4) Xes This designation is both a wing and a horizontal stabilizer station numbering subsystem, representing stations in inches located perpendicular to their front spars. In the wing, the point of C• Horizontal Stabilizer Subsystem Designations. (1) Xe This designation refers to stations in inches, located perpendicular to the elevator hinge line. (2) Xet This subsystem designation represents stations in inches, located perpendicular to the elevator tab hinge line. (3) XIS This designation represents stations in inches, located perpendicular to the front spar. The point of origin is the centerline of the fuselage. (4) X1 This subsystem designation represents the stations in the horizontal stabilizer system. The subsystem's point of origin is at the fuselage centerline. d. Vertical Stabilizer Subsystem Designations. (1) Zoy This subsystem designation denotes the stations in the vertical stabilizer system. Its origin is located 47 inches above the FRP at fuselage station 1464.2. The stations in this subsystem are segments divided into inches, parallel to the FRP. (2) Zes This designation represents stations in inches along the center spar of the vertical stabilizer. The stations are located perpendicular to the center spar and originate at Zy. (3) Lfs This subsystem designation represents the front spar system of the vertical stabilizer. It consists of stations divided into inches, located perpendicular to the front spar and originating (4) ZIe This subsystem designation represents stations in inches, along and parallel to the leading edge of the vertical stabilizer, originating at Zy, Page 2 Par 1 ### Page 41 (OCR) 11 July 72 AC 20-78 Appendix 1 (5) Ly This subsystem designation refers to the rudder system of the vertical stabilizer. It consists of stations divided into inches, located perpendicular to the rudder hinge line and originating at Zy. (6) Irs This subsystem designation refers to the rear spar system of the vertical stabilizer. It consists of stations divided into inches, located perpendicular to the rear spar and originating at Zy. (7) Zat This designation represents the rudder tab hinge system of the vertical stabilizer. It consists of stations divided into inches, located perpendicular to the hinge line of the rudder tab, and originating at Zy. Par 1 Page 3 ### Page 43 (OCR) 11 July 72 AC 20- 78 Appendix 2 142'4" FUSELAGE REFERENCE PLANE 12'3" 10° 30° 2'6". 20'10" PIVOT AXIS x = 125.0 STATIC GROUND PLANE Y = 921.0 Z =-41.1 150'8" RUDDER →6'3" RIGHT 32 1/2° LEFT 32 1/2° 29'10" 42'4" 137 4'5" 1Ґ з" 415"11" 57'6" PIVOT AXIS Х =Ô FUSELAGE REFERENCE PLANE Y=240.0 Z=-50.I STATIC GROUND PLANE HORIZONTAL GROUND REFERENCE PLANE Fig. l Basic Dimensions Page 1 ### Page 44 (OCR) AC 20-78 11 July 72 Appendix 2 265 1517 1400 123 718 798 0000 1 1661 40 285 LOWER FUSELAGE CONSTANT 1420| 1601 1795 360 SECTION 1180 1490 1310 360 CABIN CONSTANT SECTION oooo oo..aoo.....a. • 0 1220 1100 855.5 683.0 540 374.6 1355 (CENT ERWING (CENTERWING 412.7 **REAR SPAR) FRONT SPAR)** **781.7 ( CENTERWING** CENTER SPAR) Fig. 2. Fuselage Station Charts Page 2 ### Page 45 (OCR) 11 July 72 AC 20- 78 Appendix 2 Хсто Xcw 69.5 OR ₩ 239.8 **467.0 FRONT SPAR (FUSELAGE STATION** **463.0 SPOILERS 663.0)** CENTER SPAR X1$ 508.9 FUSELAGE STATION 781.7) *78 519.0 REAR SPAR Xrs589.0 355.3) FUSELAGE STATION INNER AILERON P$ 624.0 X*- 3.1 iNNER FLAP X18693.0 × 37.9 442.9 X 129.3 X+s 237.7 x 437. TRIM × ,238.8 „761.7 554.5 TAB FLAP EXHAUST GATE X, 303.7 X,697.8 X₩564.0 × 408.0- LX,$ 358.0 **OUTER AILERON - OUTER FLAP** - Х р8763.0 -Х,$428.0 WING STATIONS Fig 3. Wing Station Charts Page 3 ### Page 46 (OCR) AC 20- 78 11 July 72 appendix 2 **ENGINE AFT MOUNT** **56.2 141.4 216.9 NACELLE 100.0 178.0 ORIGIN (YNO)** • EJECTOR STROKE = 72.0 FIREWALL 118.7 186.0 258.0 **FIRESE AL** **ENGINE FORWARD MOUNT** INBOARD PYLON AND NACELLE → _ 266.7 NACELLE ORIGIN 255.2 (YNO) OUTBOARD PYLON AND NACELLE (SAME AS INBOARD EXCEPT AS NOTED) Fig 4. Pylon/Nacelle Station Charts Page 4 ### Page 47 (OCR) 11 July 72 AC 20- 78 Appendix 2 FUSELAGE X£28.8 XE 77.0 **FRONT SPAR** ACTUATING POINT × E 149:0 REAR SPAR XE221.0- HINGE POINT XE 272.0 XET103.2 XET 1852 XE721.5 Xg312.5 X4285.9 HORIZONTAL STABILIZER STATIONS HORIZONTAL STABILIZER ORIGIN (FUSELAGE STATION 1541.4) THIS POINT DESIGNATED XpSO PIVOT AXIS (FRONT SPAR STATION O) (FUSELAGE STATION 1665) FRONT SPAR PLANE (22. 7%) 25% PLANE THIS POINT DESIGNATED XEO (ELEVATOR STATION O) REAR SPAR PLANE (44.7%) THIS POINT DESIGNATES TAB ORIGIN LINE *×E *XET TAB HINGE (93.7%) ELEVATOR HINGE (75.0%) X 2839 ELEVATOR LEADING EDGE (66.2%) **METHOD OF DESIGNATING HORIZONTAL** STABILIZER STATIONS Fig 5. Horizontal Stabilizer Station Charts Page 5 ### Page 48 (OCR) AC 20-78 11 July 72 Appendix 2 FRONT SPAR - Ip 329.2 ZA 283,0 CENTER SPAR- Zp 219:0 REAR SPAR ZR 161.O ZRT 132.7 ZR 103.0 ZRT 56.6 ZA 54.5 VERTICAL STABILIZER STATIONS Zv 307.2 REAR SPAR PLANE (51.0%) CENTER SPAR PLANE (35.5%) - FRONT SPAR PLANE (20.0%) RUDDER LEADING EDGE 25% (52,0%) + Zv CHORD FUSELAGE RUDDER HINGE STATION 1464.2 Zv 103.0 35 ° (65.0%) TAB HINGE (92.9%) VERTICAL STABILIZER +ZAS + ZCS +ZAS+ZR RUDDER TAB ORIGIN ORIGIN -Yv + YV YR FYFS +YCS +YRT 50.1 **89.0 FUSELAGE 127.8 130.5 RUDDER ORIGIN REFERERENCE 162.9 PLANE 235.7** • 250.7 **METHOD OF DESIGNATION** VERTICAL STABILIZER STATIONS Fig. 6. Vertical Stabilizer Station Charts Page 6 ### Page 49 (OCR) 11 July 72 AC 20- 78 Appendix 3 STRUCTURAL ITEMS 1° DESCRIPTION. Five significant structural items are discussed in this appendix. If in a high time DC-8 aircraft a crack occurs in one of the continued airworthiness of the aircraft may be adversely affected a. Fuselage Aft Pressure Bulkhead (1) Location. Aft flat pressure bulkhead vertical angles. (See Figure 8, Appendix 3). (2) Applicable Airworthiness Directive. None (3) Other Reference Documents. (a) Service Bulletin 53-46. (4) Recommendations (a) Accomplish Service Bulletin 53-46. (b) Perform subsequent periodic inspections starting at a later date to insure continued structural integrity. b. Freon Compressor Exhaust Door Opening. (I) Location. Below forward passenger door at longeron 29 (See Figure 9, Appendix 3). (2) Applicable Airworthiness Directive. None (3) Other Reference Data. None (4) Probable Consequences. Extensive cracking around the freon compressor exhaust door opening may lead to more serious structural damage. (5) Recommendations. Periodically inspect to assure continued structural integrity. c. Forward Passenger Door Area. (1) Location. Forward fuselage, left hand side (See Figure 10, Appendix 3). (2) Applicable Airworthiness Directive. None. (3) Other Reference Data. None Par 1 Page 1 ### Page 50 (OCR) (1) FUSELAGE AFT PRESSURE **BULKHEAD** (4) AFT PASSENGER DOOR AREA Fig. 7 DC-8 Structural Inspection Locations STA. 857 STA. 1566 (2) FREON COMPRESSOR EXHAUST **DOOR OPENING WING REAR SPAR - (5) AFT LOWER WING SKIN** **AT CENTERLINE SPLICE** (3) FORWARD PASSENGER **DOOR AREA DC-8 FUSELAGE/EMPENNAGE/WING STRUCTURAL** **INSPECTION LOCATIONS** ### Page 51 (OCR) 11 July 72 Fig. 8 Fuselage Aft Pressure Bulkhead VIEW B (PAPICAL) AT STATION 1566 AFT PRESSURE BULKHEAD VIEW A-A cce OF AIRCRAFT CENTERLINE BULKHEAD PRESSURE AFT VERTICAL ANGLE BULKHEAD PRESSURE INSPECT Appendix 3 AC 20-78 Page 3 ### Page 52 (OCR) Page 4 Appendix 3 AC 20-78 Fig. 9 Freon Compressor Exhaust Door Opening VIEW A TYP. SKIN CRACK LOCATIONS 11 July 72 ### Page 53 (OCR) 11 July 72 AC 20- 78 Appendix 3 **INSPECT ALL CORNERS FOR SKIN** **AND DOUBLER CRACKING** **TYPICAL DOUCLER** **CRACK LOCATION** TYPICAL SKIN CRACK LOCATION VIEW A Fig. 10 Forward Passenger Door Areas Page 5 ### Page 54 (OCR) AC 20-78 11 July 72 Appendix 3 (4) Probable Consequences. Extensive fuselage cracks in this area may lead to more serious structural damage. (5) Recommendations. Periodically inspect to assure continued structural integrity. d. Aft Passenger Door Area. (1) Location. Aft fuselage, left hand side(see Figure 1l, Appendix 3) (2) Applicable Airworthiness Directive. None (3) Other Reference Data. None (4) Pro 20 Come serious damage. Extensive fuselage cracks in this (5) Recommendation. Periodically inspect fuselage skin around door cutout to assure continued structural integrity. e. Aft Lower Wing Skin at Centerline Splice. (1) Location. Aft lower wing skin at the centerline splice plate attach to the rear sper cap. (See Figure 12, Appendix 3) (2) Applicable Airworthiness Directive. None (3) Other Reference Documents. None (4) Probable Consequences. Crack extension in other area can significantly reduce the structural integrity of the wing. (5) Recommendation. (a) Periodically inspect to assure continued structural integrity. (b) Perform preventive rework in this area equivalent to Service Bulletin 57-70 at an appropriate later date depending on the airplane model. Page 6 Par 1 ### Page 55 (OCR) 11 July 72 CRACKING FOR SKIN AND DOUBLER INSPECT ALL CORNERS Fig. ll Aft Passenger Door Area VIEW A TYPICAL CRACK LOCATION Appendix 3 AC 20-78 Page 7 ### Page 56 (OCR) 11 July 72 AC 20-78 Appendix 3 **WING LOWER REAR** **SPAR CAP** STGR STGR #52 #51 LOWER **WING SKIN SKIN CRACKS (TYPICAL)** KEEL (REF) AIRPLANE = 0 AFT **EXTERNAL** **SPLICE PLATE VIEW A** LOOKING UP Fig. 12. Aft Lower Skin at Centerline Splice Page 8 ### Page 58 (OCR) **DEPARTMENT OF TRANSPORTATION** FEDERAL AVIATION ADMINISTRATION Washington, D.C. 20591 POSTAGE AND FEES PAID FEDERAL AVIATION Official Business ADMINISTRATION U.S.MAII DOT 515 'PENALTY FOR PRIVATE USE, $300 --- OpenAirDocs.org is an independent research project of [Contrail Labs](https://contraillabs.com/). It is not affiliated with the FAA or the NTSB. Source documents are U.S. government works in the public domain.