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66-14-02: 66-14-02 LEARJET: Amdt. 39-242 Part 39 Federal Register June 1, 1966. Applies to Models 23 and 24 Airplanes. Compliance required as indicated, unless already accomplished. (a) On Model 23 airplanes, before further flight remove windshield deicing alcohol cans. (b) On Model 23 airplanes, the following applies to all Serial Numbers except 003, 011, 016, 020, 024, 026, 033, 035, 039, 043, 044, 047, 050, 051, 062, 065A, 069, 070, 072, 073, 074, 075, 076, 077, 078, 079, 080, 081, 082, 083, 087, 090, 092, 093: further flight is limited to day VFR meteorological conditions and to flight levels below 240 until installation of an attitude indicator (gyro horizon) usable by the pilot and powered by a source separate from the airplane's primary electrical system. (c) Modify the electrical system on Model 23 airplanes, and on Model 24 airplanes S/N 24-100 through 24-129, in accordance with Lear Jet Engineering Change Record No. 340, 227, 230 or 233 (as applicable) or equivalent data approved by the Chief, Engineering and Manufacturing Branch, Central Region within the next 550 hours' time in service after the effective date of this AD. The affected airplanes and applicable data are as follows: (1) Serial Numbers 23-012 and 23-031, Engineering Change Record No. 340. (2) Serial Numbers 23-003 through 23-011, 23-013 through 23-030, and 23- 032 through 23-099, Engineering Change Record No. 340, 227, 230, 233. (3) Serial Numbers 24-100 through 24-129, Engineering Change Record No. 340. This directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated May 21, 1966. Revised August 4, 1966. Revised November 22, 1966.
67-24-04: 67-24-04 RATIER-FIGEAC: Amdt. No. 39-467, Part 39, Federal Register August 22, 1967. Applies to Model FH 76-1-7 Propellers Installed on Pilatus PC-6a Series Aircraft. Compliance required within the next 200 hours' time in service after the effective date of this AD, unless already accomplished. To prevent jamming of the pitch change actuator, replace the bronze actuator socket, P/N FH 76-1-120-02, with a steel actuator socket, P/N FH 76-2-120-02, in accordance with Ratier Figeac Service Bulletin 64-45, dated October 1966, or later SGAC-approved issue, or an FAA-approved equivalent. This amendment effective August 22, 1967.
71-19-05: 71-19-05 BRITISH AIRCRAFT CORPORATION: Amdt. 39-1292. Applies to Model BAC 1- 11 200 series airplanes. Compliance is required as indicated. To prevent failure of the saddle bracket structure located at Station 575 in the main landing gear bay, accomplish the following: (a) For airplanes with saddle bracket assemblies with 9,000 or more landings on the effective date of this AD, within the next 25 landings after the effective date of this AD, unless already accomplished within the last 175 landings, and thereafter at intervals not to exceed 200 landings from the last inspection, inspect the saddle bracket assembly in accordance with paragraph (c). (b) For airplanes with saddle bracket assemblies with less than 9,000 landings on the effective date of this AD, within the next 25 landings after the effective date of this AD, or before the accumulation of 9,000 landings on the saddle bracket assembly, whichever occurs later, unless already accomplished within the last 175 landings, and thereafter at intervals not to exceed 200 landings from the last inspection, inspect the saddle bracket assembly in accordance with paragraph (c). (c) Visually inspect the main landing gear door jack attachment saddle bracket assembly for cracks or damage in accordance with BAC 1-11 Alert Service Bulletin No. 53-A- PM3620, Issue 2, dated March 4, 1971, or an FAA-approved equivalent. (d) If a saddle bracket assembly is found to have cracks only in the top closing plate, P/N AB27-12079, and the cracks do not exceed the acceptable limits defined in BAC 1-11 Alert Service Bulletin No. 53-A- PM3620, Issue 2, dated March 4, 1971, during an inspection required by paragraph (c), before further flight repair the saddle bracket assembly in accordance with paragraph 3.1 of that service bulletin or an FAA-approved equivalent, or comply with paragraph (e). (e) If a saddle bracket assembly is found to have cracks in the top closing plate, P/N AB27-12079, which exceed theacceptable limits defined in BAC 1-11 Alert Service Bulletin No. 53-A-PM3620, Issue 2, dated March 4, 1971, or is found to have cracks or damage to any other part of the assembly during an inspection required by paragraph (c), before further flight either - (1) Replace the affected saddle bracket assembly with a serviceable assembly of the same part number; or - (2) Replace the affected saddle bracket assembly with a serviceable assembly incorporating BAC Modification PM3620. (f) The repetitive inspections required by paragraphs (a) and (b) may be discontinued after compliance with paragraph (e)(2). (g) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Certification Staff, FAA Europe, Africa, and Middle East 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 that operator. This supersedes Amendment 39-687 (33 F.R. 17895), AD 68-25-01, as amended by Amendment 39-910 (35 F.R. 144). This amendment becomes effective September 20, 1971.
56-02-03: 56-02-03 ROLLS-ROYCE: Applies to All Dart 506 and 510 Engines. Compliance required as indicated. Due to the possibility of low stage compressor impeller failure, all impellers, P/N's RK13782, RK17877, RK19795, RK20156 and RK20181, must be removed from service and not reused after a maximum of 725 hours service time, except that when Rolls-Royce Modifications 335 and 348 have been accomplished the maximum service time may be increased to 750 hours. The British Air registration Board considers this parts replacement program mandatory and the FAA concurs. Operation beyond 750 hours total time is not authorized.
76-04-01: 76-04-01 CESSNA: Amendment 39-2517 as amended by Amendment 39-2556, 39-2686 and 39-2767 is further amended by Amendment 39-2810. Applies to Models 210 through 210D (Serial Numbers 57001 thru 57575 and 21057576 thru 21058510) airplanes on which are installed Electrol manufactured main gear rotary actuator assemblies, Cessna P/Ns 1280102-1 and -2 (Electrol P/Ns EA 1471-1 and -2) or Cessna P/Ns 1280501-1 and -2 (Electrol P/Ns EA 1614-1 and -2). This AD does not apply to those airplanes on which Cessna P/Ns 1280511-3/4 and 1298100-1/2 actuator assemblies are installed. Compliance: Required as indicated, unless already accomplished in accordance with this AD, previous maintenance or AD 71-24-07. To decrease the possibility of main gear extension failure, accomplish the following: On or before April 1, 1977, or within 100 hours' time in service after February 26, 1976, whichever occurs later, install Cessna Kits 1209005-1 R/L in accordance with Cessna Service Letter SE75-21 dated October 3, 1975, or later approved revisions, or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. NOTE 1) The landing gear actuator assemblies having Cessna P/Ns and manufactured by Electrol can be identified by using an inspection mirror through the inspection plate located forward of the strut doors to read the actuator nameplate. NOTE 2) A significant savings of manhours will result if compliance with this AD and the modifications and/or inspections required by AD 76-14-07 (Amendment 39-2670 - landing gear saddle fittings) are accomplished at the same time. NOTE 3) It is imperative that Cessna P/N 1209005-1 R/L Kits be ordered from the manufacturer immediately to assure that a sufficient supply of kits will be available to allow modification of all affected aircraft on or before April 1, 1977. Amendment 39-2517 superseded AD 71-24-07, Amendment 39-1345. Amendment 39-2517 became effective February 26, 1976. Amendment 39-2556 became effective March 26, 1976. Amendment 39-2686 became effective August 12, 1976, and replaces Amendments 39-2517 and 39-2556. Amendment 39-2767 became effective November 2, 1976, and supplements Amendment 39-2686. Amendment 39-2686 in turn replaced Amendments 39-2517 and 39-2556. This amendment, 39-2810, becomes effective January 27, 1977, and supplements Amendments 39-2686 and 39-2767.
73-23-05: 73-23-05 GENERAL ELECTRIC: Amdt. 39-1742 as amended by Amendment 39-1775. Applies to Models CJ610-1, -4, -5, -6 and J85-GE-17B turbojet and CF700-2C turbofan engines. Compliance required as indicated. 1. Inspect first stage turbine discs, P/N 634E583P4 and P/N 634E583P4Y, (Wheel Assembly P/N 841B690P6) for cracks and a minimum radius of .015 inch in the forward rabbet radius and for cracks and a minimum radius of .020 inch in the aft rabbet radius in accordance with the following schedule. Use the procedures outlined in General Electric Alert Service Bulletin No. (CJ610) A72-70, or (CF700) A72-70, or later FAA approved revision or equivalent inspection method approved by the Chief, Engineering and Manufacturing Branch, New England Region, Federal Aviation Administration. a. Inspect turbine discs with 2300 or more total cycles on the effective date of this AD as follows: (1) Within next 10 cycles if not previously inspected in accordance with G.E. Alert Service Bulletins No. A72-70, or later FAA approved revisions. (2) Prior to accumulation of 800 cycles since last inspection if previously inspected in accordance with G.E. Alert Service Bulletins No. A72-70, or later FAA approved revisions. Discs in excess of both 2300 total cycles, and 800 cycles since last inspection must be inspected within the next 50 cycles. b. Inspect turbine discs with less than 2300 total cycles on the effective date of this AD, in accordance with G.E. Alert Service Bulletin No. (CJ610) A72-70 or (CF700) A72-70, or later FAA approved revisions prior to the accumulation of 2310 cycles or every 800 cycles since last inspection, whichever comes later, it discs have been previously inspected in accordance with G.E. Alert Service Bulletin A72-70. 2. Remove from service all first stage turbine discs in accordance with the following schedule. a. Remove turbine discs with 3090 or less total cycles on the effective date of this AD, from service priorto the accumulation of 3100 cycles. b. Remove turbine discs with more than 3090 total cycles on the effective date of the AD, from service prior to the accumulation of 10 additional cycles. 3. Inspect second stage turbine discs, P/N 646C596P1, for cracks and a minimum radius of .020 inch in the forward rabbet radius in accordance with the following schedule. Use the procedures outlined in General Electric Alert Service Bulletin No. (CJ610) A72-80, or (CF700) A72-80, or later FAA approved revision or equivalent inspection method approved by the Chief, Engineering and Manufacturing Branch, New England Region, Federal Aviation Administration. a. Inspect turbine discs with 2800 or more cycles on the effective date of this AD, in accordance with General Electric Alert Service Bulletin No. (CJ610) A72-80, or (CF700) A72-80, or later FAA approved revisions within the next 10 cycles. b. Inspect turbine discs with 2201 or more cycles on the effective date of this AD, inaccordance with General Electric Alert Service Bulletin No. (CJ610) A72-80, or (CF700) A72-80, or later FAA approved revisions within the next 100 cycles or at 2810 cycles whichever occurs first. c. Inspect turbine discs with 2200 or less cycles on the effective date of this AD, in accordance with General Electric Alert Service Bulletin No. (CJ610) A72-80, or (CF700) A72-80, or later FAA approved revisions at first overhaul or at 2300 cycles whichever occurs first. 4. For the purposes of this AD, a cycle is defined as that set forth in the subject Alert Service Bulletins. 5. Discs with less than the specified radius or which exhibit the specified point-type flourescent indications or cracks are to be replaced with like parts which meet the specified minima for radii and criteria for size and location of point-type flourescent indications and cracks. This supersedes Amendment 39-855 (34 F.R. 15467), AD 69-20-08. Amendment 39-1742 became effective November 20,1973. This Amendment 39-1775 becomes effective upon publication in the Federal Register.
74-10-02: 74-10-02 MCDONNELL DOUGLAS: Amendment 39-1832. Applies to Douglas Model DC-10 series airplanes, certificated in all categories, incorporating Weber Aircraft seats, Part Numbers 818472, 818473, 818474, 818475, 819291, 819812, and 819813, with dash numbers as listed in Weber Aircraft Service Bulletin No. 25-326, dated February 15, 1974. \n\n\tCompliance required within the next 300 hours' time in service after the effective date of this AD, unless already accomplished. \n\n\tTo prevent further inadvertent dislodgement of oxygen generators from seat backs, accomplish the following: \n\n\t(a)\tReplace the oxygen generator torsion retention springs in accordance with Weber Aircraft Service Bulletin No. 25-326, dated February 15, 1974. \n\n\t(b)\tThe Chief, Aircraft Engineering Division, FAA Western Region, may approve equivalent modifications. \n\n\t(c)\tAircraft may be flown to a base for accomplishment of the maintenance required by this AD per FAR's 21.197 and 21.199. \n\n\tThis amendment becomes effective May 6, 1974.
75-04-08: 75-04-08 BOEING: Amendment 39-2089. Applies to Boeing Model 737 airplanes, listed under Group I in Boeing Service Bulletin 29-1004, Revision 1, dated April 2, 1969, or later FAA approved revisions. Compliance required within the next 1,000 hours time in service after the effective date of this AD, unless already accomplished. \n\tTo prevent failure of the "B" hydraulic system electrical wiring and other systems wiring which use a common wire bundle, replace the "B" hydraulic system electrical pump spliced wires in accordance with Boeing Service Bulletin 29-1004, Revision 1, dated April 2, 1969, or later FAA approved revisions, or in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). \n\tAll persons affected by this directive who have not already received thesedocuments from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. The documents may be examined at FAA Northwest Region, 9010 East Marginal Way, Seattle, Washington 98108. \n\tThis amendment becomes effective March 10, 1975.
70-12-03: 70-12-03 FAIRCHILD-HILLER: Amdt. 39-996. Applies to F-27 and FH-227 type airplanes certificated in all categories. To assure that the outboard flaps are contained in the event of over-travel, by the addition of positive stops to the screwjacks, accomplish the following within the next 250 hours in service after the effective date of this AD, unless already accomplished. (a) Comply with the applicable Fairchild Hiller Service Bulletin, No. F-27-27-72 dated January 16, 1970, or No. FH-227-27-30 dated January 16, 1970, or later revision or equivalent method both approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. (b) Upon request with substantiating data submitted through an FAA Maintenance Inspector, the compliance time specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective June 19, 1970.
75-03-04: 75-03-04 FAIRCHILD (HILLER): Amendment 39-2071 as amended by Amendment 39-2251. Applies to model 1100 and FH1100 type helicopters certificated in all categories. To detect cracks in the tail fin spar channel, P/N 24-62030-7 or P/N 24-62030-43 in the area of the tail rotor gear box mount, P/N 24-62006-3, accomplish the following inspection or an equivalent inspection approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, within the next five hours in service after the effective date of this AD unless already accomplished within the last 95 hours in service and at every 100 hours in service thereafter: 1. Remove tail rotor gear box fairing and fin leading edge cover. 2. Clean the tail fin spar in an area one inch in diameter around the left and right forward attachment bolts (two of ten to which the tail rotor gear box mount is attached to the spar) and the spar surface between these two attachments and forward, for a distance of 1 1/2 inches with metachlor or equivalent grease and oil remover by light scrubbing with a stiff bristle brush. 3. Inspect the cleaned area for cracks with at least a ten power magnifying glass by looking through the front end of the tail rotor gear box mount fitting. 4. If a crack is found, replace with an uncracked fin assembly that has been inspected in accordance with the above procedure or alter fin in accordance with an alteration approved by the Chief, Engineering and Manufacturing Branch, Eastern Region before further flight. Amendment 39-2071 was effective January 27, 1975, and was effective for all recipients of the airmail letter of December 9, 1974, upon receipt. This amendment 39-2251 is effective July 8, 1975.