Results
94-02-07: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model MD-11 series airplanes, that requires replacing the anti-skid control unit. This amendment is prompted by three reports of failure of the center landing gear drag link, after which the center landing gear swung aft and struck the fuselage. The actions specified by this AD are intended to prevent failure of the center landing gear drag link, which could result in extensive damage to the fuselage structure.
94-13-05: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 0100 series airplanes. This action requires modification of the electrical connections in a certain relay panel. This amendment is prompted by a report that, as a result of a previous modification, certain electrical wires with a positive voltage were connected to an electrical connector that is also wired to the liftdumper system, which could negatively affect the liftdumper system. The actions specified in this AD are intended to ensure protection against uncommanded deployment of the liftdumper.
95-13-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320 series airplanes, that requires inspections to determine the gap of the seat track joints at frame 64, and correction of discrepancies. This amendment also requires eventual modification of the seat tracks on all affected airplanes, which terminates the requirement of repetitively removing or repositioning the seats. This amendment is prompted by in-service inspection reports, which have revealed that a gap between the forward and aft seat track at frame 64 could exceed the tolerance limit due to a method used on the assembly line to control the position of the seat track. The actions specified by this AD are intended to ensure that the gap of the seat track joints do not exceed the tolerance limit and subsequently lead to separation of the passenger seats from the seat track under emergency landing conditions.
50-06-02: 50-06-02 BOEING: Applies to All Model 75 Series Aircraft. Compliance required at each annual inspection. For military aircraft, compliance also required prior to original certification. Remove the center section gas tank and inspect both front and rear spars for cracks, checks and warping. Defective spars should be replaced or repaired in accordance with CAM 18. Ascertain that all drain holes are open. Repeated removal of the tank at each annual inspection is not necessary, if after accomplishment of the items mentioned above, the gap between the gas tank and the upper surface of the center section is sealed by doping on grade A fabric tape, or equivalent sealing means, to prevent moisture entering the tank compartment. This supersedes AD 45-51-01.
46-44-05: 46-44-05 CESSNA: (Was Mandatory Note 5 of AD-768-4.) Applies Only to 120 and 140 Aircraft Serial Numbers 8001 to 8517, Inclusive. Compliance required prior to January 1, 1947. Replace each of the four internal wrenching bolts which attach the engine to the engine mount with an AN 6 bolt and a special offset washer. AN 6-47 bolts should be used at the upper fittings and AN 6-35 bolts at the lower fittings. The special washer is made of 4130 steel 7/8 inch in diameter and 1/4 inch in thickness with a through hole 0.377 inch in diameter and the O.D. machined to a 0.600-inch diameter a depth of 1/8 inch. The 0.600 inch-diameter offset fits into the aft end of the attachment fitting and the head of the replacement bolt bears directly on the special washer. Also, an AN 960-616 washer should be added between the nut and the AN 970-6 washer at the front face of the rubber bushing. This change is made to prevent the bolts from pulling through the 1 1/2-inch diameter x 0.049-inch plate welded to the front of each fitting. (Cessna Service Letter No. 18 covers this same subject.)
47-07-03: 47-07-03 FAIRCHILD: (Was Service Note 4 of AD-724-2.) Applies to M-62 Series Aircraft. Prior to original certification and at each periodic inspection thereafter, and as otherwise noted, make the following inspections: (1) Inspect the wing center section bottom surface for cracks. This inspection should be made after each severe landing. Cracks extending into the spar flange area indicate cracked spar flanges and should be investigated very thoroughly. (2) Inspect the butt ends of the spars to assure that the butt plates are in place and properly attached. (3) Inspect the strap hinge fittings for looseness. Clearance between the spar webs and hinge plates is not critical as long as the plates are bolted tight to the bushings if the bushings protrude. If bushings are loose, replace. (4) Inspect the plywood spar webs for checks or cracks. This inspection should always be made after any damage to the landing gear. Cracks other than those parallel to the face grain generally indicate serious damage to the spar web. (5) Inspect the trailing edge of the wing center section and outer panel over flap area for deterioration due to accumulated moisture. (6) Inspect the forward face of front spar and belly skin at engine cutout in wing center section for oil soaking and skin separation. (These inspections and methods of repair are covered in greater detail in Fairchild Service Bulletin 47-62-1 dated January 24, 1947. Owners may obtain copies from Fairchild Personal Planes Division of Fairchild Engine and Airplane Corp., Hagerstown, Md.)
47-21-02: 47-21-02 FAIRCHILD: (Was Mandatory Note 8 of AD-707-2 and Mandatory Note 5 of AD-706-1.) Applies to 24R-46 and -46S, and 24W-46 and -46S Aircraft. Compliance required prior to July 1, 1947. Replace the landing light fuse with one of 20-ampere capacity. (Fairchild Service Bulletin 47-24-1 dated January 8, 1947, covers this same subject.)
47-43-05: 47-43-05 CESSNA: (Was Service Note 3 of AD-768-5.) Applies to 120 and 140 Aircraft Serial Numbers 8001 to 13780, Inclusive. Inspection required upon each 100 hours of operation until reinforcing channels are installed at all hinge fittings. Inspect for fatigue cracks in the elevator spar web at the hinges. These cracks start either at the rivets or at an edge of the fitting and progress around the fitting until the elevator breaks loose from the hinge fitting. If cracks less than 1/2 inch in length are found a reinforcing channel, Cessna P/N 0434151 at the outboard hinge or 0434152 at the inboard hinge, should be installed on the aft side of the spar with the flanges riveted between the spar flanges and the skin with two AN 455AD3 rivets per flange. Four AN 442AD4 rivets should be used to attach each fitting to the spar web and reinforcing channel. If any cracks are longer than 1/2 inch the spar should be replaced and the reinforcing channels added. (Cessna Service Letter No. 46 dated July 31, 1947, covers this same subject.)
47-51-05: 47-51-05 CURTISS-WRIGHT Applies to Model C-46 Series Aircraft Equipped with Aileron Horn Assembly, Curtiss-Wright P/N 20-050-5715. Compliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished. The aileron horns part 20-050-5715 have failed due to cracking of the horn between the attaching bolt holes and the outer edge. Inspection should be made to determine if this part has been replaced by P/N SK-10213. If not, part 20-050-5715 which is a casting should be replaced by a machined horn manufactured from 24ST material in accordance with Curtiss-Wright Drawing No. SK-10213. (Army Technical Order 01-25L-102 also covers this same subject.) Revised December 28, 1964.
50-25-01: 50-25-01\tSTINSON: Applies to All Model 108 Series Aircraft.\n\tCompliance required not later than September 1, 1950.\n\tReports have been received of fuel seepage into the space between the inner cabin trim and the outer fabric covering of the fuselage. This results in soaking of insulating material in the cabin wall. The source of the fuel can be spillage during filling of tanks, thermal expansion of fuel in full tanks, or tank leakage. This fuel runs to the under surface of the wing, adhering to the lower curved surface of the trailing edge of the wing at the flap well, thence inboard to the fuselage and across the rear window. Since the window seal is often not perfectly tight the fuel may then enter the cabin wall.\n\tTo preclude the fire hazard of fuel soaked insulation within the cabin wall due to these causes, a drip strip similar to that shown in Figure 1 should be installed on the underside of each wing. This drip strip will prevent fuel from flowing from the wing to the fuselage.\n\t(Piper Service Bulletin No. 115, dated March 31, 1950, covers this same subject.)
94-23-05: This amendment adopts a new airworthiness directive (AD) that is applicable to AlliedSignal Inc., (formerly Garrett Engine Division) TFE731-3A-200G and -3AR-200G model turbofan engines. This action requires removing from service certain low pressure turbine (LPT) disks, imposing an hourly life limit on the first stage and second stage LPT disks, performing a dimensional inspection of second stage LPT disks at repetitive intervals, and incorporating honeycomb material in the second stage LPT nozzle air seal. This amendment is prompted by reports of LPT disk web separations. The actions specified in this AD are intended to prevent LPT disk web separations, which can result in an uncontained engine failure and damage to the aircraft.
47-42-17: 47-42-17\tDOUGLAS: (Was Service Note 3 of AD-781-1.) Applies to the Following DC-6 Aircraft Serial Numbers: Douglas 43061; AAL 42854 to 42865, Inclusive; 42879 to 42880, Inclusive; 42882 to 42896, Inclusive; and 43035 to 43044, Inclusive; UAL 42866 to 42875, Inclusive; and 43000 to 43024, Inclusive; Panagra 42876 to 42878, Inclusive; National 43055 to 43058, Inclusive; Sabena 43062 to 43064, Inclusive; Braniff 43105, 43106; KLM 43111 to 43112, Inclusive; and AAF 42881. \n\n\tInspection required at each 300 hours (or at each 150 hours for non-air-carrier operations).\n \n\tInspect the center spar web between Stations 167 and 184 for cracks in the web along the lower row of rivets which attach the spar web to the leg of the upper spar cap. For aircraft with the 10-tank fuel system this inspection can be properly made only by removing the fuel tank inspection opening near the affected area, since the spar web attaches to the forward side of the spar cap leg and small cracks in the web cannot be detected without close examination. If cracks are found during this inspection or, if between the inspections, leaks occur which are caused by cracks in the center spar web between Stations 167 and 184, the spar web must be reinforced by installing a doubler in accordance with Douglas Drawing 5356664.\n\n\tWhen the spar web reinforcement has been incorporated the special inspection required by this Note may be eliminated. All DC-6 aircraft not mentioned above will be reinforced at the factory.\n \n\t(Douglas Service Bulletin DC-6 No. 29, "Rework Center Spar Web, Stations 167-184, Integral Wing Fuel Tank DC-6 Airplane", covers the same reinforcement as described on Drawing 5356664.)
95-12-21: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A340-211 and 311 series airplanes. This action requires the installation of doublers on certain stringers located in the center fuselage. This amendment is prompted by the results of the manufacturer's full-scale fatigue test which indicate that fatigue cracking can occur at these stringer locations. The actions specified in this AD are intended to prevent reduced structural integrity of the fuselage due to the problems associated with fatigue cracks in the subject stringers.
71-09-06: 71-09-06 MARTIN AIRCRAFT: Amdt. 39-1200. Applies to Models 202, 202A and 404 airplanes certificated in all categories. Compliance required as indicated. To detect cracks or corrosion in the main landing gear aluminum gland nut, P/N 202SD81548, (Menasco Drawing No. 511034), accomplish the following: (a) Within the next 25 hours in service, after the effective date of this AD, unless already accomplished within the last 475 hours in service, and thereafter at intervals not to exceed 500 hours in service from the last inspection, comply with (b). (b) Unscrew the left and right main landing gear aluminum gland nut, P/N 202SD81548, so that all threads are visible. Inspect the nut for evidence of cracks or corrosion, using dye penetrant in conjunction with at least a 10-power glass or an equivalent inspection. Cracked or corroded parts must be replaced with an unused part or an equivalent before further flight, except that the airplane may be flown in accordance withFAR 21.197 to a base where the repair can be performed. (c) The repetitive inspection specified in (a) may be discontinued when the aluminum gland nut is replaced by a part made of steel, as per ECD 32590 to Menasco Drawing No. 511034, change L, dated 8 March 1961, or replaced by a steel gland nut in accordance with Eastern Air Lines Drawing No. 204-8125, or with an equivalent part. (d) Upon submission of substantiating data by an owner or operator through an FAA Maintenance Inspector, the Chief, Engineering & Manufacturing Branch, FAA, Eastern Region, may adjust the initial inspection and the repetitive inspection interval specified in this AD. Equivalent parts and inspections must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective May 4, 1971.
49-09-02: 49-09-02 NAVION: Applies to Airplanes Equipped With Adel Electric Booster Pumps. The Following Adel Pumps Do Not Require Modification in Accordance With This Directive: (1) Pumps With Serial Numbers above 2451, (2) Pumps Having a Red Painted Band on the Pump Housing, (3) Pumps Having the Letters "G" or "S" Suffixed to the Pump Serial Number. To be accomplished as soon as possible but not later than April 1, 1949. Several instances of air leakage into the fuel system have been reported on Navions equipped with Adel electric booster pumps. It has been determined that air can enter the fuel system through the 0.062-inch diameter hole in the plate at the rear of the Adel pump inlet chamber. This hole was originally provided to prevent overboard drainage of fuel through a faulty pump shaft seal while the pump was running. All of the pumps affected require blocking of the hole at the rear of the pump inlet chamber. This is accomplished in the field by means of an Adel manufactured wire plug which is inserted into the hole through the pump inlet port. Pumps with Serial Numbers below 1600 which do not have the letter "R" suffixed to the serial number also require replacement of the pumpshaft running seal spring. Adel Accessories Service Bulletin No. 147-49 describes these changes. The required plug and spring and copies of the Adel Bulletin and Ryan's covering Service Letter No. 57, may be obtained from the Ryan Aeronautical Co., San Diego, California.
47-49-04: 47-49-04 LOCKHEED: (Was Mandatory Note 37 of AD-763-3.) Applies to Model 49 Serials 2068 to 2088, Inclusive. Compliance required within next 50 hours of operation unless the 1 3/16-inch headless drive pin has been installed. Inspect attachments of rudder pedal lever arms to the clip assembly in the 284587 rudder pedal slot cover guide assemblies to determine whether it is possible for the flat head pin to cause jamming of the system. If any possibility of jamming exists, the flat head pin should be replaced with a headless drive pin 1 3/16 inches long. (LAC Service Bulletin 49/SB-260 covers this same subject.)
47-06-01: 47-06-01 GLOBE: (Was Service Note 1 of AD 766-5.) Applies to Models GC-1A and GC-1B Aircraft. To be accomplished prior to April 1, 1947, and upon each 100 hours operation thereafter. Inspect main landing gear retraction system to determine that adjustments are as follows: (1) When the side brace is against the down stop the middle joint should be 1/8 inch to 1/4 inch above dead center (3/16 inch to 5/16 inch if measured from edges of links in accordance with Globe Customer Service Maintenance Bulletin No. 7). (2) When the side brace is against the down stop and the down lock plunger is fully extended, covering at least 1/2 of the adjustment screw head, the clearance between the plunger and the screw head should be from 0.001 inch to 0.005 inch. (3) When the side brace is against the down stop the limit switch plunger should be depressed approximately 1/32 inch beyond the cutoff point. (4) The turnbuckle in the emergency extension cable should be adjusted so that on manual extension of the gear both down locks operate before the handcrank has been wound to the full down position. After it has been determined that the turnbuckle adjustment is satisfactory in this respect it should be determined also that with the handcrank wound to the full up position the cable length is sufficient to permit the up limit switches to cut off. (Globe Customer Service Maintenance Bulletin No. 7 covers this same subject.)
47-42-09: 47-42-09 DOUGLAS: (Was Mandatory Note 6 of AD-781-1.) Applies to DC-6 Serial Numbers 42854, 42855, 42858 Through 42865, 42869 Through 42880, 42882 Through 42891, 43000 Through 43009, 43055 and 43056. \n\nTo be accomplished not later than next No. 3 inspection (or not later than next 150 hours for noncarrier operations). \n\nCertain cases have been found where the aileron hinge plates at wing Stations 421, 485, 585, and 675 were fabricated from overgage stock resulting in an interference fit between the plate and the clevis fitting. The following hinge plates and fittings should be inspected to determine whether or not they conform with the tolerances listed below. If plates are found which exceed the widths noted below, they should be reworked with emery cloth to specified limits and touched up with zinc chromate primer. Fittings which have been installed over an oversize plate should be anodized and carefully inspected before being reinstalled. \n\n\n\nHinge\n\n\n\n\nStation\nNo.\nPlate\nThickness\nFitting\nWidth\n421\n1\n3320118\n0.249 \n4334619\n0.249\n\n\n\n0.237\n\n0.254\n485\n2\n3323460\n0.238*\n4345756\n0.334\n\n\n\n0.243\n\n0.350\n585\n3\n3323461\n0.311\n4345755\n0.311\n\n\n\n0.297\n\n0.316\n575\n4\n3323462\n0.249\n4345754\n0.249\n\n\n\n0.237\n\n0.254\n\n*Thickness of 0.093 angle not included.\nTotal thickness should not exceed 0.334. \n\n(Douglas Service Letter A-214-529.004/RLT dated July 21, 1947, and attached sketches cover this same subject.)
47-33-02: 47-33-02\tDOUGLAS: (Was Mandatory Note 14 of AD-618-3, Supplement 2; and Mandatory Note 15 of AD-669-3, Supplement 2.) Applies to All DC3 Series Aircraft. \n\nTo be accomplished not later than the first engine change after September 1, 1947, but in any event not later than December 1, 1947. \n\nIn order to preclude cowl flap hydraulic line failures and possible subsequent fires, replace grommets and lines forward of the firewall with AN 833-4 elbows and AN 924-4 nuts, or equivalent, and new fire resistant flexible hose assemblies of proper length.
47-27-04: 47-27-04 DOUGLAS: (Was Mandatory Note 18 of AD-762-7.) Applies to All C-54 and DC-4 Series Airplanes Having Exhaust Collector Rings Made Up of Top Segments P/N 5174842-56 L.H. and 5174529-56 R.H. \n\n\tTo be accomplished not later than the first engine change subsequent to July 15, 1947, but in any event not later than October 15, 1947. \n\n\tSeveral reports have been received of cracking failure of the collector ring "Y" outlet assembly due to breathing of the exhaust stack. This induces failure which creates a fire hazard. This type of exhaust collector "Y" is not reinforced with a flange and is shown on page 4, Douglas Service Bulletin No. DC-4 No. 31. To correct this condition weld a scalloped stiffener flange on the exhaust collector aft of the "Y" outlet assembly. \n\n\t(Douglas Service Bulletin No. DC-4 No. 68 covers this same subject.) \n\n\tUntil this repair is accomplished, inspection for cracks should be made immediately and at periods not to exceed 50 hours of operation.
47-07-04: 47-07-04 FAIRCHILD: (Was Mandatory Note 6 of AD-724-2.) Applies to M-62 Series Aircraft. Compliance required at next periodic inspection. In order to eliminate the possibility of the control sticks becoming disengaged, both front and rear control sticks should be reworked by drilling through stick and socket, and installing and safetying an AN393-51 clevis pin. Washer AN 960-10L and cotter AN 380-2-2. (Fairchild Service Bulletin 44-62-2 dated October 31, 1944, covers this same subject.)
47-51-16: 47-51-16 DOUGLAS: Applies to DC-6 Serial Numbers 42854 Through 42880, Inclusive; 42882 Through 42891, Inclusive; and 43000 Through 43009, Inclusive; 43055; and 43056. \n\nTo be accomplished not later than August 1, 1948. \n\nIn order to increase the strength of the flap support assembly at wing Station 378 and to replace the temporary rework outlined in Douglas Co. Service Letter of May 12, 1947, which was necessitated by failure of the flap hinge support assembly on an airplane in flight, the following must be accomplished: \n\n(a)\tRemove the two Shafer bearings, P/N AB-4A from flap link assembly, P/N 4325008 and press in new Shafer bearings, P/N AB-5A and stake in place. \n\n(b)\tRemove outboard flap link support assembly, P/N 5107188, and line ream (0.312-0.313) diameter through to permit use of 5/16-inch bolt for attachment of upper end of link assembly, P/N 4325008. Assembly, P/N 5107188 becomes P/N 5107188-500 after rework. \n\n(c)\tPress out old bushing, P/N 1338719, two places in flap hinge bracket assembly, P/N 3320998, and press in new bushings, P/N 1338719-500. \n\n(d)\tAfter replacing P/N 5107188-500 replace P/N 4325008, using bolts, P/N 2356375-22; washers, P/N 124682-5-12-6 and P/N AN 960-516; nut, P/N AN 310-5 and cotter pin, P/N AN 380-2-3. \n\n(Douglas Service Bulletin DC-6 No. 66 covers this same rework.)
93-19-01: 93-19-01 Dowty Rotol (now Dowty Aerospace Gloucester Ltd.): Amendment 39-8694. Docket 93-ANE-10. Supersedes AD 87-21-51, Amendment 39-5929. Applicability: Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Propeller Models (c)R354/4-123-F/13, (c)R354/4-123-F/20, and (c)R375/4-123-F/21 installed on SAAB-SF340A, and SAAB-SF340B series aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent possible loss of the propeller, accomplish the following: (a) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/13 propellers, perform Dowty Aerospace Gloucester Ltd. Modification (C) VP3336 by installing interface shim Part Number (P/N) 660712669 in accordance with Dowty Aerospace Gloucester Ltd. Service Bulletin (SB) No. SF340-61-57, dated February 15, 1991, within 500 hours time in service (TIS) since the last torque check and inspections accomplished in accordance with AD 87-21-51. (b) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model R354/4- 123-F/13 propellers, perform a torque check of the propeller retention bolts for low torque and a magnetic particle inspection of the propeller retention bolts for cracks; perform dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. If propeller retention bolts or hubs are found to have cracks, remove from service prior to further flight, and replace with serviceable propeller retention bolts and hubs, within 500 hours TIS since the last torque check and cracking inspections accomplished in accordance with AD 87-21-51. These actions must be accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991. (c) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/20 and (c)R375/4-123-F/21 propellers, perform Dowty Aerospace Gloucester Ltd. Modification (C) VP3336 by installing interface shim P/N 660712669 in accordancewith Dowty Aerospace Gloucester Ltd. SB No. SF340-61-57, dated February 15, 1991, within 100 hours TIS after the effective date of this AD; or 500 hours TIS since the last torque check and inspections accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991, or Dowty Aerospace Gloucester Ltd. SB No. SF340-61-21, Revision 4, dated October 1, 1987; or 500 hours TIS since new, whichever occurs latest. (d) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/20 and (c)R375/4-123-F/21 propellers, perform a torque check of the propeller retention bolts for low torque; a magnetic particle inspection of the propeller retention bolts for cracks; and dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. Remove from service prior to further flight cracked propeller retention bolts and hubs, and replace with serviceable propeller retention bolts or hubs, within 100 hours TIS after the effective date of this AD; or 500 hours TIS since the last torque check and inspections accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991, or Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) SB No. SF340-61-21, Revision 4, dated October 1, 1987; or 500 hours TIS since new, whichever occurs latest. These actions must be accomplished in accordance with Dowty Aerospace Gloucester Ltd, SB No. SF340-61-58, Revision 1, dated July 18, 1991. (e) Thereafter, for propeller models identified in paragraphs (a) and (c) of this AD, perform a torque check of the propeller retention bolts for low torque; a magnetic particle inspection of the propeller retention bolts for cracks; and dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. Remove from service prior to further flight cracked propeller retention bolts and hubs, and replace with serviceable propeller retention bolts or hubs at intervals not to exceed 1500 hours TIS since the last torque check and inspections performed in accordance with paragraphs (b) or (d), as applicable, of this AD. These actions must be accomplished in accordance with Dowty Gloucester Aerospace Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991. (f) Installation of newly designed Dowty Aerospace Gloucester Ltd. propeller hub assembly P/N 660714241 and remarking of the propeller in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-61, Revision 1, dated October 19, 1992, constitutes terminating action to the inspection requirements of this AD. (g) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Boston Aircraft Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, BostonAircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Boston Aircraft Certification Office. (h) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (i) The inspections shall be done in accordance with the following Dowty Aerospace Gloucester service bulletins: Document No. Pages Revision Date SF340-61-57 1-5 Original February 15, 1991 Total pages: 5. SF340-61-58 1-2 1 July 18, 1991 3-5 Original February 15, 1991 Appendix A 1-3 Original February 15, 1991 Appendix B 1-3 Original February 15, 1991 Appendix C 1-2 1 July 18, 1991 3-4 Original February 15, 1991 Total pages: 15. SF340-61-61 1-3 1 October 19, 1992 4-8 Original September 9, 1992Total pages: 8. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from Dowty Aerospace Gloucester Ltd., Anson Business Park, Cheltenham Road East, Gloucester GL2 9QN England. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (j) This amendment becomes effective on October 29, 1993.
94-17-05: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Airbus Model A340-211 and -311 series airplanes. This action requires replacement of certain circuit breakers for the toilet system vacuum generator. This amendment is prompted by reports of excessive vacuum generator overloads or blocked rotor conditions, in which the thermal protection device and/or related aircraft circuit breakers do not interrupt the power supply. The actions specified in this AD are intended to prevent a fire in the event of excessive current in the electrical circuit due to overheating of the vacuum generator.
95-12-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Model A320-231 series airplanes, that requires repetitive functional checks to detect leakage of the distribution piping of the engine fire extinguishing system, and repair, if necessary; and modification of the piping, which would terminate the inspection requirements. This amendment is prompted by reports of cracking of the engine fire extinguisher pipe, which resulted in leakage of the fire extinguisher agent. The actions specified by this AD are intended to prevent leakage of the fire extinguishing agent, which could prevent the proper distribution of the agent within the nacelle in the event of a fire.