(19) Aircraft returned to ramp, crew reporting flaps inoperative.
Inspection of the flap system revealed that the flap power unit assembly worm gear teeth were stripped. Flap power unit was replaced, aircraft released.
(20) There is a fleetwide corrosion maintenance program in operation to watch for corrosion and stress corrosion cracking on the upper and lower spar caps (both forward and rear spars) from RSS 450 outboard to about station 499. The upper cap is 7178-T6 material and is especially susceptible to stress corrosion cracking. The lower spar cap is 2024-T3 and is not quite so critical as far as stress corrosion is concerned.
So far, only overseas operators have been severely affected.
* (21) During routine inspection, two cracks were detected in the web of BBL 41.0 lower left beam in the air conditioning bay.
TAT 3117 hours.
(22) During scheduled inspections, the left and right wheel well gussets at station 663.75 were found cracked on four aircraft, gusset P/N's 65-45402-59 and -60. TAT ranged from 4869 hours to 6909 hours. *
c. Empennage •
(1) Routine inspection disclosed a cracked L/H horizontal stabilizer mounted vortex generator. The crack extended forward from the aft edge through four fastener holes.
TAT 5462 hours.
(2) Five cracks were found in the skin at BS 1104. The cracks propagated from the upper fastener hold for R/H #2 vortex generator. TAT 2715 hours.
(3) A crew report of a rumble during letdown resulted in finding a portion of the R/H horizontal stabilizer vortex generator missing. TSO 1269 hours.
(4) During fleet campaign the L/H and R/H horizontal stabilizer vortex generators were found cracked at the 900 bend radius.
TSN 1940 hours.
(5) On arrival, crew reported fuselage vibration of moderate intensity from 800 to 1200 CPS at all speeds above 240 knots.
Inspectors found the elevator tab inboard hinge brackets, left and right, were worn resulting in excessive play. Hinge brackets were replaced and no vibration noted on test flight.
MRR 70-082-11 dated 17 Apri1 1970.
Par 8 Page 7
d. Main Landing Gear Problems.
(1) Extensive research and testing by Boeing have been performed to solve the MLG broken torque link and ensuing shimmy problem.
* The MIG is responsive to the correct oil and air in the shock strut and proper tire pressure. Correct servicing results in improved service life. Service Bulletin 32-1064 describes the installation of a shimmy damper with increased stroke which provides a greater margin of resistance to shimmy. *
(2) During routine replacement of the main landing gear brake assemblies for normal wear, three shear studs (P/N 69-59091-1) were found failed and two cracked at the left inboard brake installation. Further investigation revealed one shear stud failed and one cracked at the left outboard brake installation.
All 18 replaced MRR 70-135-047 dated 7 July 1970). Investigation revealed the studs had been installed with excessive torque.
Affected operators have been advised of the correct torque by Service Bulletin 32-1068 Rl. In addition, Service Bulletin 32- 1094 describes the installation of new improved studs and bolts.* Page 8 Par 8
17 Jan 73 AC 20-83 Appendix 1 APPENDIX 1. CORROSION
1. CAUSES OF CORROSION.
a. Metal corrosion is the deterioration of the metal by chemical or electrochemical attack. Corrosion can take place internally as well as on the surface of the metal. This deterioration may change the smooth surface, weaken the interior or damage or loosen adjacent parts.
Water or water vapor containing salt and combined with oxygen in the atmosphere produces main source of corrosion in aircraft. Thus, aircraft operating in a marine environment or in areas where the atmosphere contains industrial fumes which are corrosive are particularly susceptible to corrosive attacks. If unchecked, corrosion can cause eventual structural failure.
2. TYPES OF CORROSION. There are two general classifications of corrosion, direct chemical attack and electrochemical attack. In both types, the metal is converted into an oxide, hydroxide, or sulfate. The corrosion process involves the anode which is oxidized and cathode (or the corrosive agent) which is reduced.
a. Direct chemical attack. Corrosion by direct chemical attack results from direct exposure to caustic liquids or vapors. The anodic and cathodic change occurs at the same point. Direct chemical attack in aircraft structure deposits are caused by (1) spilled battery acid or fumes, (2) residual flux deposits from welds, and (3) trapped caustic cleaning fluids.
b. Electrochemical attack. An electrochemical attack is similar to the electrolytic reaction in electroplating or in a dry cell battery.
The reaction requires a medium, like moisture, capable of conducting electricity. When a metal comes in contact with a corrosive agent (dissimilar metal) and is connected by a liquid path, the metal decays or corrodes. The electrochemical attack is responsible for most forms of corrosion on aircraft structure.
3. FORMS OF CORROSION. There are many forms of corrosion which depend on the metal involved, size, shape, atmospheric conditions and corrosion producing agents.
Surface corrosion. This may be caused by either direct chemical or electrochemical attack. Surface corrosion appears as a general roughening or pitting of the surface of a metal accompanied by a powdery deposit of corrosion products.
b. Dissimilar metal corrosion. Extensive pitting damage may result from contact between dissimilar metal parts in the presence of a conductor.
A galvanic action like electroplating occurs at points of contact when insulation has broken down or was omitted.
Par l Page 1
AC 20-83 17 Jan 73 Appendix 1
c. Intergranular corrosion. The grain boundaries of an alloy are attacked by this type of corrosion. Intergranular corrosion may This is a flaking of the metal at the surface caused by pressure of corrosion residual product buildup.
d. Stress corrosion. This type corrosion occurs as the result of the combined effect of tensile stresses and corrosive environment. Stress corrosion is found in most metals. However, it is a particular characteristic of aluminum, certain stainless steels, and high strength steels.
e. Fretting corrosion. This occurs when two mating surfaces are subject to relative motion although normally at rest with respect to each other. It is characterized by surface pitting and generation of finely divided debris.
4. FACTORS AFFECTING CORROSION. Many factors affect the speed, cause, type, and seriousness of metal corrosion which include:
a. Climate.
b. Size and type of metal.
c. Foreign material.
5. CORROSION PREVENTION. Corrosion-preventive maintenance includes:
a. An adequate cleaning.
b. Thorough periodic lubrication.
c. Detailed inspection for corrosion and failure of protective systems.
d. Prompt treatment of corrosion and touchup of damaged paint areas.
e. Keeping drainholes free of obstruction.
f. Daily wipedown of exposed critical areas.
g. Sealing of aircraft against water during foul weather and proper ventilation on warm, sunny days.
h. Making maximum use of protective covers on parked aircraft.
Page 2 Par 3
AC 20-83 CHG 1 8/8/74 APPENDIX 2, STATION CHARTS Appendix 2 (941 FOR -100) 25° 25% 37 F 00000000000000000000000000 96 FT 11 IN. (90'-7" FOR -100) 17 FT 2 IN. 93 FT 13 FT 2 IN. 37 FT 4 IN. -100) (34'-4" FOR * 13 FT Figure l. Principal Dimensions (-200) Page 1
AC 20-83 CHG 1 Appendix 2 8/8/74 € - WINDOWS & . UPPER RADII WL 208.1 TOP OF FLOOR BEAM WL 148.5 @STABLIZER PIVOT BWL 274 B STA 1154 EXHAUST SHIELD TAILCONE ACCESS DOOR APU ACCESS DOOR NOTE: 1. WL OF GROUND AT Q MAIN LANDING GEAR TIRE 1060 (NOMINAL STRUT AND TIRE CONDITION SYM ABOVE WL 208,1 UPPER LOBE TYPICAL FROM STA 480 TO STA 905 BELOW WL 208.1 LOWER LOBE TYPICAL FROM STA 390 TO STA 540 REAR VIEW PRESSURE BULKHEAD ACCESS DOOR ENTRY DOOR (LH) 30 X 72 GALLEY DOOR (RH) 30 X 65 - SECTION 48 -
- ENTRY DOOR (LH) CARGO DOOR (LH) 34 × 72 86 X 134 -200 ONLY UPPER LOBE 100000000000 WL 232.5 LOWER LOBE CARGO DOOR 35 × 48 ELECTRONIC ACCESS DOOR 30.5 X 24.5 SECTION 43 LEFT SIDE VIEW -200 ONLY 000000000000000 000 CARGO DOOR 33 x 48 (RH) SECTION 46 GALLEY DOOR (RH) 30 × 65 OCI FW D ACCESS DOOR 14 × 27 NOSE GEAR WHEEL WELL FORWARD AIR STAIRS DOOR (LH) 16 X 44 SECTION 41 ESCAPE HATCH 20 × 38 (LH & RH) SL899 689 WING CENTER WHEEL A iL CREASE LINE WL 208.1 PRESSURE LEVEL SECTION Figure 2. Body Station Diagram Page 2
Appendix 2 =63.29391 - WBL557.5 1. BASIC LEADING EDGE ANGLE TAN = 52230862
35.9059* - WBL 522.00
19.32- WBL 546 WBL522.0 SURGE TANK WBL 583.0 SLAT W STA 557.0 AILERON 2. SEMI SPAN 46.50 FEET W STA WBL 417.00 COUTBOARD FRONT SPAR 531.0 W STA 505.0 - OUTBOARD FLAP ACCESS DOORW STA 479.0 WBL438.77 - W STA 447.47 TYP W STA 453.0 SLAT -SPOILER -W STA 395.47 -OUTBOARD REAR SPAR WBL 314.76 WBL 329.75 WBL 304.5- WBL 355.0- • SPOILER W STA 342.77 • SPOILER W STA 290.07 - WBL 175.5€ NACELLE & FRONT SPAR WBL 279.25 - - WBL 204.4 25% BASIC CHORD WBL 254.0 25% SWEEPBACK WBL 204.4 SLAT WBL 227.0 - INBOARD FLAP - WBL 180.75 - SPOILER - & INBOARD REAR SPAR WBL 198.6 WBL 87.75 WBL 191 WBL183.4 LE STA 222.7 LE STA 213.0 INBOARD FRONT SPAR WBL 135.5 LE STA 169.0 FLAP R8 1332380 - & AIRPLANE (SYM) LE STA 119.0- FLAP- WBL 114.0-2 WBL 92.5 • 133.1701 • 124.8268 W STA 55. 6 989 60.9735 WBL 71.247 21.8721 41.5233 185.60 -288.0968 W STA
0.00 Figure 3. Wing Station Diagram Page 3
AC 20-83 CHG 2 1/31/75 Appendix 2
COWL SPLIT COWL SPLIT
LINE LINE
• & FWD MOUNT RING NAC STA
03.47 NAC E INLET. BL 0.8706 ENG AND NAC (BL O.O) WING BL 191,00 & AFT MOUNT RING
PLAN VIEW AS SHOWN ENGINE NO. I
NAC NAC TR TR A THRUST REVERSER & WING CHORD STA C ENGINE STA STA STA AND TAIL PIPE PLANE - 139.037 218.65 92.92 96.47 NAC 4 FRONT SPAR • REAR SPAI TR STA AT WBL 191.00 AT WBL 191.0 STA
63.41 154.66
35.867 1° ANGLE OF 38* 4.00
INCIDENCE 32,25 COWL SPLIT
COWL SPLIT LINE LINE- & ENGINE CAFT @ FWD MOUNT ENG INTERSECTION OF ENGINE C
MOUNT RING STA AND THRUST REVERSER & RING I NACELLE 224.67 NAC ENG (NAC WL 100.0) ENG 3° SPACER
STA SIA STA
99.25 136.00 218.65
INTERSECTION OF ENGINE C
AND NACELLE & - Figure 4. Engine And Nacelle Station Diagrams Page 4
" POINT "E" -JACK SCREW PIVOT POINT NOTE: 1/31/75 POINT "A" (STAB NEUTRAL) BODY HORIZONTAL STABILIZER RIBS FROM STA 83.50 TO 249.10 HAVE C STABILIZER 35.50 BL 9.20 INCH CENTERS CLOSURE RIB POINT "G" Figure 5. Horizontal Stabilizer and Elevator Station Diagrams GREAR BODY STA 1093.417 POINT IF". POINT "B" C HINGE (CNTR) SPAR & FRONT SPAR (CNTR) POINT "H"
249.10- 3.00 - VIEW 1 AIRPLANE (SYM) 21.75 LBODY BL BODY STA 1154,000 BODY STA 1142.495 STAB STA LE (SYM) AIRPLANE LE STA STAR TRUSS STAB STA 68.14 STAB STA OI STAB STA 25% CHORD LINE.
254.02 ề. 28 LEADING EDGE- FRONT SPARE 83.50 STAB 75.00 STA 57.93 STA LE STA 56.01 YBODY STA 1049.602 10% CHORD 15.10 T
60.00 45.00 33.00 · STABILIZER CONNECT POINT "A" SEE VIEW STAB 27.75 STABILIZER CONNECT POINT "B" SEE VIEW 1 83.05 30° SWEEPBACK ELEV FRONT SPAR C 123.10 113.25 BODY STA 1096.428 TRAILING EDGE 151.00 REAR SPAR € ELEV STA 213.32 ELEV 55% CHORD ELEV. HINGE RIB C STA ELEV NOSE RADIUS& & HINGE 149.12- ELEV STA ELEV HINGE RIB C ELEV - BODY STA 1146.50฿
174.64 ELEV BALANCE PANEL STA Page 5 ELEV STA 121.59 ELEV HINGE RIB C ELEVATOR REAR SPAR C ELEV STA 66.54 ELEV HINGE RIB C TAB HINGE RIB Q (TYP) ELEV STA 94.27 108.16- 89,52 19.63 ELEV STA 24.90 ELEV STA 23.40 37.75 ELEV HINGE C 75 CHORD Appendix AC 20-83
CONTROL TAB HINGE &
Page 6 NOTE: -62.00. Appendix 2 AC 20-83 CHG 2
46.50 →
1. FIN STATIONS ARE PERPENDICULAR TO 34.10 11.00 $ 8.00 55% LINE 6.20- LE STA 312,35
2. LEADING EDGE STATIONS ARE PERPENDICULAR TO LEADING EDGE FIN STA 299.15 FIN STA 292.7 + FWL 230.25 Figure 6. Vertical Stabilizer and Rudder Station Diagrams 3. RUDDER STATIONS ARE PERPENDICULAR TO 75% LINE TYP PIVOT POINT - RUD
4. VERTICAL STABLIZER RIBS FROM FIN STATION 81.10 TO FIN STATION 292.70 ON 9,2 INCH CENTERS 214.77 STA RUD -RUD STA 233.09 STA 237.64
RUDE RUDDER SPAR
5. TO CONVERT FIN STATION AT FRONT SPAR (10%) LE STA=FIN STA 1.01756521 + 3.59150051 E REAR SPAR 25% CHORD LINE FWL 196.46 STA TANGENT A = .52293481 55% CHORD 161.0 TANGENT B = 78884390 - RUD TANGENT C = 084793481 STA
174.12 242.00 FIN STA 81. 44 FIN STA 69. 10% CHORD I FRONT SPAR FWL RUDDER L RUD STA TANGENT 112.356 141.52 RADIUS 220.0 POINT •RUD BWL 378,218 STA
121.64 FIN STA 57.77.
FIN STA 46.1 & RUDDER HINGE FIN STA 45.5. RUD LE 55.36 86.59 STA O
870.913 B STA TONGENT STA O FIN STA 9.5 FIN 350 6.50 RUD BALANCE WEIGHT - STA 31.65 60.15 RUD STA FIN WL O (BWL 300.50)
•20,5 - 51.25 101.897 - CONNECTION -B STA 1016.00 B STA 1088.20 POINTS - 112,75 FIN 52.248 STA THRUST B STA STA 56.57 FIN - RUD STA 0.00 1176.5167.70 B STA FIN STA RUD STA 13.33 RUD STA 23.27 - BODY CROWN LINE 1/31/75 RIB 1126.56 153,75 122,397 205.00
Appendix 3 APPENDIX 3. NONDESTRUCTIVE TESTING
1. MAINTENANCE TOOL. Simply stated nondestructive testing (NDI) is a tool for performing maintenance inspections such as X-ray, ultrasonic, magnetic particles, eddy current, dye penetrant, and visual with or without magnification.
Maintenance Inspection. NDI permits maintenance inspection without removing components from aircraft or tearing down complex assemblies.
Defects in various aircraft systems which would escape detection through normal visual inspection will be identified by NDT.
Training Required. Special NDT training is desirable to make sure that the technician is capable of operating the equipment and interpreting the results. Also, many States require that an X-ray industrial aprica tone. technician 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.
2. METHODS AND APPROACHES FOR DAMAGE DETECTION. Information on the application of NDI methods for the detection of fatigue damage is scattered through literature on metal fatigue and literature on NDT.
Few publications deal with this problem specifically. Some publications are limited to laboratory investigations, and the NDT application is conducted under controlled laboratory conditions. Such investigations contribute considerably to available knowledge and provide new or improved NDI methods for field use. However, there is a great difference between what can be done in the laboratory and what can be used in practical applications under field service conditions.
3. NDI METHOD IN FIELD AND SERVICE USE. Most of the NDT methods that are used under field and service conditions are those capable of detecting fatigue cracks of various sizes. These methods are described in books on the subject of nondestructive testing and other publications such as reports, technical papers, and magazine articles. The NDT methods most commonly used in the field and service for fatigue crack detection follow.
4. VISUAL INSPECTION. Visual inspection is the oldest, simplest, cheapest, and most widely used of all NDT methods. The basic principal used in visual inspection is to illuminate the object and examine the surface with the eye.
The surface should be adequately cleaned before being inspected.
Visual inspection for detection of fatigue cracks can be improved by aids such as mirrors, lenses, microscopes, periscopes, and telescopes. These devices compensate for limitations of the human eye. Boroscopes permit direct visual inspection of the interior of hollow tubes, chambers, and other internal surfaces.
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AC 20-83 17 Jan 73 Appendix 3
10. The capability of visual inspection to detect a fatigue crack depends on many factors such as the size and location of the crack, the illumination used, optical aids employed, and skill of the inspector.
It is often difficult to detect even a relatively large fatigue crack that is located, for example, at the corner of a groove or that coincides with a machining mark. There are, of course, also limitations on the size of cracks that can be detected by visual inspection, depending on optical aids employed.
LIQUID PENETRANT. The liquid penetrant is one of the oldest methods of nondestructive testing and is capable of detecting cracks that may be impossible to find with the most careful visual inspection because either they are too small or because they are difficult to detect due to their location. The principle involves applying to the part surface a liquid penetrant having a low-surface tension and low viscosity. When used on a clean surface that the liquid will wet, the liquid is drawn into the cracks by capillary action. The presence of the liquid in the cracks is revealed when, after wiping the excess liquid from the surface, a developer is applied that acts like a blotter and draws the liquid out.
a. There are two types of liquid penetrants in general use. One contains a dye which usually gives a good color contrast against the selected developer; the other contains dissolved flourescent material, which makes it readily visible when viewed under a "black" (ultraviolet)
b. Liquid penetrant inspection is inexpensive and readily applicable to field use. The surface must be cleaned before inspection and also afterwards to remove the developer.
6. MAGNETIC METHODS. Magnetic inspections are used to detect surface or near surface discontinuities in ferromagnetic materials, and they are well suited for the detection of fatigue cracks.
a. The principle employed here is that once a magnetic field is induced in a naterial, any cracks and flaws that are present, will rerturb or distort that magnetic field. These methods are most sensitive when the crack orientation and the magnetic field direction are perpindicular to each other. When they are parallel, the crack will not be detected.
The magnetic particle method is the most frequently used. It consists of three basic steps:
(1) Establishment of the magnetic field in the part to be inspected.
(2) Application of magnetic particles to the surfaces of the part.
(3) Visual examination of the surfaces for indications of fatigue cracks. These indications are provided by the particles being Page 2 Par 4