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63-08-02: 63-08-02 DOUGLAS: Amendment 39-630. McDonnell Douglas. Applies to McDonnell Douglas Model DC-8 Series Aircraft equipped with P/N 3703218 (no dash number) elevator control tab push rod assembly. \n\n\tAs a result of damage near the midpoint of the elevator control tab push rod assembly due to wear from rubbing against the guide support assembly and the guide support attach rivets, accomplish the following: \n\n\t(a) Unless already accomplished, within the next 300 hours' time in service after the effective date of this AD: \n\n\t\t(1) Remove both left and right-hand elevator control tab push rod assemblies, P/N 3703218, and conduct a close visual inspection of the push rods for evidence of wear due to contact of the push rod with guide support assembly, P/N 5708625-3. \n\n\t\t(2) Push rods showing evidence of wear shall, prior to further flight: \n\n\t\t(i) be replaced either with an undamaged part; or \n\n\t\t(ii) be reworked in accordance with the rework procedures outlined in Figure(1) of Step (7) of Douglas DC-8 Service Bulletin No. 27-51 Reissue No. 1 dated September 25, 1962, or an FAA approved equivalent. Push rods showing evidence of wear which require removal of material in excess of 0.025 inch in depth and one inch in length by 0.375 inch in width on one side of the push tube, or which have dents or sharp gouges, or are found worn or cracked in more than one area may not be reworked and must be replaced. When push rod assemblies are reworked, they must be reinspected using dye penetrant method or equivalent, to insure that no cracks exist after the rework is accomplished. \n\n\t\t(3) Following reinstallation of push rod assemblies, and before further flight, conduct an initial check for clearance per Figure (1), Step (1); and , as necessary, accomplish the adjustment and rework outlined in Figure (1), Steps (2) through (6), of Douglas DC-8 Service Bulletin NO. 27-51, Reissue No. 1, dated September 25, 1962, or FAA approved equivalent. \n\n\t(b) At intervals of not less than 400 nor more than 600 hours' time in service following the initial clearance check prescribed by (a)(3), unless other inspection intervals have been approved for an operator by the Chief, Engineering & Manufacturing Branch, FAA Western Region, remove and again inspect rods replaced or reworked per (a)(2) for any evidence of wear or contact with the guide support assembly. Any rods showing evidence of wear must be reworked or replaced per (a)(2), and the reinstallation clearance check and such adjustment and rework provisions of (a)(3), as found necessary, shall be accomplished. \n\n\t(c) If, subsequent to compliance with (a), an airplane is altered by changing an elevator, elevator control tab, elevator control tab push rod assembly, or any combinations of these, the following new procedure is required: \n\n\t\t(1) Prior to further flight conduct an initial check for clearance and any necessary adjustment and rework as described in (a)(3). \n\n\t\t(2) At intervals of not less than 400 nor more than 600 hours' time in service following the initial clearance check required by (c)(1), unless other inspection intervals have been approved for an operator by the Chief, Engineering & Manufacturing Branch, FAA Western Region, the elevator control tab push rod assembly associated with this change shall be removed and inspected for any evidence of wear or contact with the guide support assembly as described in (a)(1). Any rods showing evidence of wear must be reworked or replaced as indicated in (a)(2) and the reinstallation clearance check and adjustment and rework provisions of (a)(3), as found necessary, shall be accomplished. \n\n\t(d) The periodic reinspection prescribed by (b) and (c)(2) may be discontinued when: \n\n\t\t(1) It is determined that no wear or contact with the guide support assembly has developed during the preceding reinspection interval, or \n\n\t\t(2) The elevator control tab push-rod assembly Douglas P/N 3703218 is replaced with Douglas P/N 3703218-501, in accordance with the procedure outlined in DC-8 Service Bulletin No. 27-150 dated November 5, 1963, or by an FAA approved equivalent part and procedure. \n\n\tNOTE: The P/N 3703218-501 steel push rods presently being installed during production have a smaller diameter than the original P/N 3703218 (no dash number) aluminum push rods. \n\n\t(e) Upon request of the operator an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operation period of the operator if the request contains substantiating data to justify the increase or decrease for such operator. \n\n\t(Douglas DC-8 Service Bulletins No. 27-51, Reissue No. 1, dated September 25, 1962, and No. 27-150, dated November 5, 1963, cover this same subject.) \n\n\tThis directive effective April 18, 1963. \n\n\tRevised November 9, 1963. \n\n\tRevised April 15, 1964. \n\n\tRevised September 16, 1968.
2009-18-11: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Following a red illuminated "DOOR NOT LOCKED" status light indication on the door lock indication panel after lift off, the cabin crew operated the door lock handle. This resulted in inadvertent opening of the downward opening passenger door in flight. * * * After inspection, it was found that the false red light might be the result of an incorrect clearance between lever Part Number (P/N) A26997-003 and the Up-Limit Switch. If the Up-Limit Switch has an incorrect clearance, the combination with cabin differential pressure build-up after lift-off might result in a false steady illuminating red "DOOR NOT LOCKED" indication on the Door Indication Panel.* * * * * * * * The unsafe condition is inadvertent opening of the door lock handle in flight, which could result in rapid decompression of the airplane or ejection of a passenger or crewmember through the door. We are issuing this AD to require actions to correct the unsafe condition on these products.
79-19-03: 79-19-03 SHORT BROTHERS LIMITED: Amendment 39-3553. Applies to Model SD3-30 airplanes, S/Nos. SH 3004, 3005, 3006, 3007, and 3008, certificated in all categories. Compliance is required prior to the accumulation of 10,000 flights, unless already accomplished. To prevent failure of the fuselage skin panel buttstraps, accomplish the following: (a) Replace or reinforce the original fuselage skin panel buttstraps in accordance with Section 2, "Accomplishment Instructions" of Short Brothers Ltd. Service Bulletin SD3-53- 28, dated May 5, 1978, or an equivalent approved by the Chief, Aircraft Certification Staff, AEU- 100, Europe, Africa, and Middle East Region, Federal Aviation Administration, c/o American Embassy, Brussels, Belgium. (b) For purposes of complying with this AD, a flight is defined as one takeoff and one landing. This amendment becomes effective October 4, 1979.
2009-17-04: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: One case of elevator servo-control disconnection has been experienced on an aircraft of the A320 family. Failure occurred at the servo-control rod eye-end. Further to this finding, additional inspections have revealed cracking at the same location on a number of other servo-control rod eye-ends. In one case, both actuators of the same elevator surface were affected. * * * A dual servo-control disconnection on the same elevator could result in an uncontrolled surface, the elevator surface being neither actuated nor damped, which could lead to reduced control of the aircraft. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
64-20-03: 64-20-03 VICKERS: Amdt. 796 Part 507 Federal Register August 20, 1964. Applies to Viscount Model 810 Series Aircraft. Compliance required as indicated. Fatigue cracking has occurred in the rib structure of the inboard rib at Station 96 of the inboard nacelles, illustrated in Figure 1 of Preliminary Technical Leaflet No. 113 (800/810 Series). To preclude further failures, accomplish the following in accordance with the PTL referenced herein or FAA approved equivalent: (a) Within 275 landings after the effective date of this AD unless already accomplished within the last 225 landings, conduct dye penetrant or an FAA approved equivalent inspection for cracks in the rib reinforcing plate and rib web plate in accordance with PTL 113. (b) If no cracks are found, reinspect at intervals not exceeding every 500 landings from the last inspection until Modification FG.1960 Part "B" is accomplished, after which no further inspection for this defect will be required. (c)If cracks are found during the initial inspection described in paragraph (a), accomplish paragraphs (e), (f), or (g), as appropriate, within 275 landings after the effective date of this AD. (d) If cracks are found during a reinspection, accomplish paragraphs (e), (f), or (g), as appropriate, within 275 landings from the time the cracks are found. (e) If a crack is found in the rib reinforcing plate only, incorporate the repair scheme Figures 2 and 3 of PTL 113 or an FAA approved equivalent and reinspect within every 1,500 landings to ensure that there is no progression of damage in the reinforcing plate and no initiation of damage in the web plate. These repetitive inspections may be discontinued after incorporation of Modification FG. 1960 Part "C". (f) If a crack is found in the rib web plate only, incorporate Mod. FG.1960 Part "B" and reinspect the rib web plate within every 3,000 landings to ensure that cracking has not been initiated in the reinforcing plate. Theserepetitive inspections are no longer necessary after the incorporation of Mod. FG.1960 Part "C". (g) If cracks are found in both the reinforcing plate and the rib web plate, incorporate Mod. FG.1960 Part "C". (Vickers-Armstrongs Preliminary Technical Leaflet No. 113 Issue 2 (800/810 Series) and Modification FG.1960 cover this subject.) This directive effective September 21, 1964.
71-13-03: 71-13-03 HAWKER SIDDELEY AVIATION: Amdt. 39-1230 as amended by Amendment 39-1259. Applies to Model DH-114 "Heron" airplanes. Compliance is required as indicated. To prevent possible failure of the Dunlop compressed air bottles used in the emergency landing gear extension system and emergency braking system, accomplish the following on or before August 31, 1971. (a) For all airplanes, inspect the air bottle (P/N AH.7360 or AC.11038) used in either of the emergency landing gear extension system air bottle assemblies (P/N AC.11768) located under the pilot's seat: If the air bottle was manufactured before January 1, 1959, replace the air bottle assembly with a serviceable assembly of the same part number which incorporates an air bottle (P/N AH.7360 or AC.11038) manufactured on or after January 1, 1959. The date of manufacture is etched on the collar of the bottle. (b) For airplanes which have incorporated Modification 281 (Emergency Braking System), inspect the airbottle (P/N AC.10685 or AC.11038) used in the emergency braking system air bottle assembly (P/N ACM.16784) located on the left forward face of the crew cabin sloping bulkhead. If the air bottle was manufactured before January 1, 1959, replace the air bottle assembly with a serviceable assembly of the same part number which incorporates an air bottle (P/N AC.10685 or AC.11038) manufactured on or after January 1, 1959. The date of manufacture is etched on the collar of the air bottle. (Hawker Siddeley Technical News Sheet, Series: Heron (114), No. S.6, Issue 2, covers this subject.) Amendment 39-1230 became effective July 30, 1971. This Amendment 39-1259 becomes effective July 31, 1971.
82-12-02 R1: 82-12-02 R1 BRITISH AEROSPACE (HAWKER SIDDELEY): Amendment 39-4392 as amended by Amendment 39-5058. Applies to British Aerospace Model HS/BH/DH 125 up to and including series 700 airplanes certificated in all categories except those airplanes incorporating Modification 252772. To prevent structural failure of the flap, accomplish the following within the next 100 hours time in service after the effective date of this AD or before the accumulation of 600 hours time in service on the airplane, whichever is later, unless already accomplished. 1. Visually inspect the flap outboard hinge nose ribs for cracks in accordance with the instructions in paragraph 2A of British Aerospace, Aircraft Group, 125 Service Bulletin (SB) No. 57-58, Revision 3, dated September 1, 1983. a. If no cracks are found, no further action is necessary. b. If cracks are found during the inspection required by paragraph 1, above, which do not exceed the criteria in paragraph 2A(5) of the service bulletin, reinspect at intervals not exceeding 40 hours time in service from the last inspection until a permanent repair is incorporated. c. If cracks are found during the inspection required by paragraph 1 or during the repetitive inspection required by paragraph 1.b, above, which exceed the criteria in paragraph 2A(5) of the service bulletin, replace the flap with a serviceable flap or contact the manufacturer for instructions to make a permanent repair before further flight. 2. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. 3. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA Northwest Mountain Region. The manufacturer's specifications and procedures identified and described in this directive are incorporated hereinand made a part hereof pursuant to 5 U.S.C. 552(a)(1). Amendment 39-4392 became effective June 8, 1982. This amendment becomes effective June 17, 1985.
2009-17-05: The FAA is adopting a new airworthiness directive (AD) for Honeywell International Inc. TPE331-10 and TPE331-11 series turboprop engines. This AD requires removing certain first stage turbine disks from service. This AD results from a report of an uncontained failure of a first stage turbine disk that had a metallurgical defect. We are issuing this AD to prevent uncontained failure of the first stage turbine disk and damage to the airplane.
62-26-03: 62-26-03 LOCKHEED: Amdt. 512 Part 507 Federal Register December 6, 1962. Applies to All Models 49, 149, 649, 649A, 749, 749A, and 1049-54 Series Aircraft Incorporating Main Landing Gear Crossheads, P/N 307866 or P/N 288982 Which Have Accumulated 10,000 or More Hours' Time in Service. Compliance required as indicated. To detect fatigue cracking in the 0.25-inch radii adjacent to the one-inch diameter bearing surfaces on main landing gear crossheads, the failure of which would prevent normal extension and retraction of the main landing gear, the following shall be accomplished: (a) Within the next 700 hours' time in service after the effective date of this AD, unless already accomplished within the last 1,800 hours' time in service prior to the effective date of this AD, and thereafter at intervals not to exceed 2,500 hours' time in service from the last inspection, inspect all crossheads as follows: The crosshead shall be removed from the aircraft and inspected by themagnetic particle method or FAA approved equivalent for cracks in the 0.25-inch radii adjacent to the one-inch diameter bearing surfaces. All cracked crossheads shall be replaced with sound ones before the aircraft is returned to service. Crosshead replacement for 1049-54 aircraft shall be P/N 307866 only. Crosshead replacement for 49, 149, 649, 649A, 749, and 749A aircraft shall be either P/N 307866 or P/N 288982. (b) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. (Lockheed Field Service Letter FS/251565L, dated March 31, 1961, covers this same subject.) This directive effective January 7, 1963.
81-01-02 R1: 81-01-02 R1 GOVERNMENT AIRCRAFT FACTORIES (GAF): Amendment 39-3999 as amended by Amendment 39-4147. Applies to Models N22B (Serial Nos. N22B-5 and up) and N24A (Serial Nos. N24A and up), certificated in all categories, which are equipped with fuel selector cables U2000 LAS-2-( ) and U2000 LA-2-( ); i.e., the U2000L series. Compliance required as indicated. To prevent failure of the fuel tank selector valves or the fuel shut-off valves to operate, accomplish the following: (a) Within the next 25 hours time in service after the effective date of this AD, unless already accomplished, inspect the cable sleeves on the fuel tank selector valve and fuel shut-off valve. If they are incorrectly crimped, cracked, or loose, before further flight, repair No. ANMD- 28-11 (hereinafter referred to as the Service Bulletin) dated August 21, 1980, or an FAA-approved equivalent. (b) Cable sleeves on the fuel tank selector cable which have been repaired in accordance with GAF Nomad Alert Service Bulletin No. ANMD-28-11, dated August 21, 1980, or an FAA-approved equivalent, must be visually inspected prior to the first flight of each day in accordance with paragraph 4 of the service bulletin, and replaced prior to the accumulation of 200 hours time in service from the time of repair. (c) If an equivalent is used in complying with paragraph (a) of this AD, that equivalent must be approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. NOTE: All persons affected by this directive who have not already received the Service Bulletin from the manufacturer may obtain copies upon request to the Government Aircraft Factories, 226 Lorimer Street, Port Melbourne 3207 Vic., Australia. These documents may be examined at the FAA, Engineering and Manufacturing District Office, 300 Ala Moana Blvd., Room 7321, Honolulu, Hawaii 96850, or Rules Docket, Room 916, FAA, 800 Independence Ave., S.W., Washington, DC20591. Amendment 39-3999 became effective January 5, 1981. This amendment 39-4147 becomes effective June 22, 1981, as to all persons except those persons to whom it was made immediately effective by telegraphic AD T81-01-02 R1, issued January 16, 1981, which contained this amendment.