Results
68-01-04: 68-01-04 FOUND BROTHERS AVIATION LTD: Amendment 39-561. Applies to FBA-2C aircraft. Compliance required as indicated. To preclude the failure of the forward wing to fuselage attachment either by the failure of the attachment bolt or by the cracking of the wing root rib web around the attachment fitting, accomplish the following: (a) Replace the wing to fuselage forward attachment NAS 146-42 bolt with an unused bolt of the same part number or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, within the next 150 hours' time in service after the effective date of this AD, unless already accomplished within the last 350 hours' time in service, and thereafter at intervals not to exceed 500 hours' time in service from the last replacement. (b) Within the next 150 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter at intervalsnot to exceed 250 hours' time in service from the last inspection, visually inspect for cracks the wing root ribs repaired in accordance with either Found Brothers repair 2C39-18, Issue 2, or 2C39-19, Issue 2, or equivalent repair approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (c) Upon request with substantiating data submitted through an FAA maintenance inspector, the compliance times specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This amendment is effective January 9, 1968.
75-26-02: 75-26-02 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE, (S.N.I.A.S., formerly Sud Aviation). Amendment 39-2454. Applies to Aerospatiale Alouette III Helicopter Models SA-315B, SE-3160, SA-316B, SA-316C, and SA-319B, certificated in all categories, incorporating main rotor heads P/N's 3160S. 12.10.000.11 through .14, P/N's 3160S. 12.20.000.4 through .7, P/N's 3160S.12.10.000.1 through .10 modified in accordance with Modification No. S296-AM 1108 or Alouette Service Bulletin No. 65-52, or P/N's 3160S.12.20.000.1 through .3 modified in accordance with Modification No. S296-AM 1108 or Alouette Service Bulletin No. 65-52. Compliance is required as indicated, unless already accomplished. To prevent failure of the fixed levers of the main rotor head hydraulic drag dampers, accomplish the following: (a) Upon the effective date of this AD, and thereafter once on each day of operation, until accomplishment of paragraph (c) of this AD, visually inspect each of the three hydraulicdamper fixed levers for cracks in the area of the eccentric attachment hole. (b) If cracks are found in any hydraulic damper fixed levers, before further flight, replace the cracked hydraulic damper fixed lever with a serviceable unit of the same part number. (c) Within the next 100 hours' time in service after the effective date of this AD, remove the three hydraulic drag dampers from the main rotor head, inspect, rectify as necessary, and reinstall in accordance with subparagraph 1C of Lama Service Bulletin No. 65.15, dated September 23, 1974, for Model SA-315B, or subparagraph 1C of Alouette Service Bulletin No. 65.101, dated September 23, 1974, for the other designated models, or an FAA-approved equivalent of the applicable Service Bulletin. This amendment becomes effective December 23, 1975.
58-07-03: 58-07-03 VICKERS: Applies to All Viscount 700 Series Aircraft. Compliance required before accumulation of 6,000 flight hours. Investigations have proved it is necessary to replace the 1/2-inch diameter bolts securing the top inboard attachment fittings of the inner nacelles to the leading edge member at Station 96 at 6,000 hours. The P/N's of the bolts, which are to be replaced at 6,000 hours are as follows: 70103-4405 (Mod. D.1031 embodies); 80203-2405 (Mod. D.1327 or D.2025 embodied). Vickers Mod. D.2581 introduces redesigned nuts and bolts as direct replacement for the above bolts. This design ensures that any bending moments present will be taken by the full shank diameter of the bolts. The modified bolt assemblies are split pinned. Vickers-Armstrong has issued PTL 179, Issue 2, and Modification D.2581 covering this same subject. The British Air Registration Board considers this mandatory. The FAA concurs with this action and considers compliance therewith mandatory.
97-17-06: This amendment adopts a new airworthiness directive (AD), applicable to Bell Helicopter Textron, Inc. (BHTI) Model 214ST helicopters, that requires replacement of each emergency float inflation solenoid valve (valve). This amendment is prompted by two inadvertent inflations of emergency float systems that resulted from self-activations of the valves. The actions specified by this AD are intended to prevent self-activation of the valves, and subsequent inadvertent inflation of the emergency float system, which could lead to loss of control of the helicopter.
75-16-22: 75-16-22 DeHAVILLAND DH-114: Amendment 39-2298. Applies to all DeHavilland Model DH-114 airplanes modified in accordance with Supplemental Type Certificate (STC) SA1685WE. Compliance required within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 1500 hours' time in service, and thereafter at intervals not to exceed 1500 hours' time in service from the last inspection. To prevent excessive wear of the counterweight bushings and subsequent ineffectiveness of the counterweight function, accomplish the following: Inspect and replace, if required, crankshaft counterweight pins and bushings in accordance with Teledyne Continental Overhaul Manual X-30039 or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, ASO-210, P.O. Box 20636, Atlanta, Georgia 30320. This amendment becomes effective August 8, 1975.
69-13-01: 69-13-01 BRITTEN NORMAN LTD: Amdt. 39-783. Applies to Britten Norman Models BN-2 and BN-2A Aircraft Serial Numbers 3 through 43 and Serial Numbers 45 and 46. Compliance required as indicated unless already accomplished. To prevent a possible failure of the aileron, rudder, or nose wheel steering control system, accomplish the following: (a) Within the next 25 hours' time in service, inspect the threaded female portion of the fork-ends, P/N NB 45-B-879, of each turnbuckle assembly in the aileron, rudder, and nose wheel steering systems for evidence of thread defects in accordance with Britten-Normal Service Bulletin BN-2/SB.15, dated April 16, 1969, or later ARB-approved issue or later FAA-approved equivalent. (b) If the threaded female portion of any turnbuckle fork-end is found to be defective during the inspection required by paragraph (a), replace each defective fork-end with a serviceable fork-end of the same part number before further flight. This amendment becomes effective June 23, 1969.
54-26-01: 54-26-01 GRUMMAN: Applies to All Models G-44 and G-44A Aircraft. Compliance required by June 15, 1955. There have been reported numerous instances of the landing gear locking mechanism failing because of either hydraulic system leaks or failure of the mechanical locks. These malfunctions have been reported in both the up and down position of the landing gear. To prevent future similar malfunctions, provide a more positive means of holding the gear in its locked position, both in the fully extended and fully retracted positions. Grumman Service Bulletin No. 24, October 18, 1954, accomplishes this by providing a closed center hydraulic system. This arrangement provides hydraulic pressure to hold the gear in the selected position and unwanted extension or retraction is prevented even though the mechanical locks may fail or leaks develop in the hydraulic system. This supersedes AD 48-05-05.
52-11-02: 52-11-02 CONVAIR: Applies to All Model 240 Aircraft. Compliance required not later than the first major engine overhaul after February 1, 1953. To improve further the engine nacelle fire resistance of 240 aircraft, steel facings must be installed over certain aluminum alloy components of the engine cowl panels, the oil cooler duct, and the nacelle structure forward of the firewall. (Convair Service Bulletin No. 240-425, Revision 2, describes these changes in detail. Preliminary information on this modification is contained in Convairogram No. 30, dated April 8, 1952.)
68-06-03: 68-06-03 HAWKER SIDDELEY: Amdt. 39-566. Applies to Model DH. 125 airplanes, Series 1A, 1A/522 and 3A. Compliance required as indicated. To prevent a fully asymmetric flap condition in the lift dump position, within the next 150 hours' time in service after the effective date of this AD, replace the flap center hinge bolt, P/Ns 25CF71, 25CF1837, 25CF2387 and 25CF2357, with a self-retaining bolt, P/N 3110-7681, in accordance with Hawker Siddeley Service Bulletin 27-49-(1894) Revision 2, dated November 27, 1967, or later ARB-approved revision or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region. This amendment becomes effective April 20, 1968.
75-04-07: 75-04-07 PRATT & WHITNEY AIRCRAFT: Amendment 39-2084. Applies to all Pratt & Whitney Aircraft JT3D-3, JT3D-3B, and JT3D-7 turbofan engines containing tenth stage compressor disk, P/N 701810. Compliance required as indicated. To ensure adequate life limit margin for tenth stage compressor disk, P/N 701810, the cyclic life limits on these disks have been reduced below the figures currently approved. Unless already accomplished, remove from service the tenth stage compressor disk prior to exceeding the revised life limit listed below or within the next 25 cycles in service after the effective date of this AD, whichever comes later. Engine Model Previous Life Limit (Cycles) Revised Life Limit (Cycles) JT3D-3 30,000 25,000 JT3D-3B 30,000 25,000 JT3D-7 25,000 23,000 If a disk has been used in more than one engine model, the disk is limited to the lowest cyclic life permitted for the engine models in which it has been exposed. This amendment becomes effective February 19, 1975.
69-10-02: 69-10-02 VICKERS: Amdt. 39-765. Applies to Viscount Models 744, 745D and 810 Series Airplanes. Compliance required within the next 1500 hours' time in service after the effective date of this AD, unless already accomplished. To improve the fire protection of air system ducting adjacent to cabin compressor outlets in the engine nacelles, accomplish the following: (a) For Viscount Models 744, 745D and 810 Series airplanes, replace glasscloth ducts on the outlet side of the cabin compressor in the right inboard, the right outboard and the left inboard engine nacelle with aluminum alloy ducts in accordance with BAC Modification Bulletin No. D3198 Issue 2 (700 Series) or FG.2070 Issue 2 (810 Series) or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East region. (b) For Viscount Model 810 Series airplanes only, replace the right and left inboard nacelle fiber glass saddle brackets with stainless steel saddle brackets in accordance with BAC Modification Bulletin No. FG.2070 Issue 2 (800 Series) or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East region. This amendment is effective 12 June 1969.
74-18-02: 74-18-02 PIPER: Amendment 39-1929. Applies to Models PA-25-235 and PA-25-260 airplanes, Serial numbers 25-7405573 to 25-7405673 inclusive, certificated in all categories. Compliance required within the next 50 hours' time in service after the effective date of this AD, unless already accomplished. To prevent possible fuel unbalance and fuel line chafing, accomplish the following: (a) (1) Install the vent line clamp assembly Piper Part Number 60383-02 using four stainless steel 6 x 3/8 L type A Truss Phillips Head screws, one MS35206-228 screw and one MS20365-632C nut on both right and left wings at the fuel tank vents. (2) Position both right and left fuel tank vents so that they project 3/4 inch below the wing fabric and secure with the vent line clamps. (b) Inspect all fuel line grommets (15 per wing) in the left and right wings for proper position and security as follows: (1) Remove inspection covers from bottom of wing and inspect the grommetson each fuel line (using light and mirror) to insure that they are correctly installed in the holes of the fuel line support plates. (2) Replace all lines that have been damaged by chafing wear with serviceable lines of the same part numbers or equivalent lines approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southern Region. (3) Correct installation of any grommet not properly installed. (4) If any of the grommets are not properly installed at a support plate that cannot be reached through the existing inspection hole, it will be necessary to cut additional access openings in bottom of wing as required, install cover plate grommet part number 85012-78 and inspection cover plate part number 12761-02 over any new opening in fabric and install existing inspection covers. Piper Service letter No. 721 Parts II and III, or other approved later revision, pertains to this same subject. This amendment becomes effective August 27, 1974.
57-16-04: 57-16-04 HAMILTON STANDARD: Applies to all Hamilton Standard Two-Flyweight Models 4U18 and 5U18 Governors. Compliance required as indicated. To prevent the possible occurrence of propeller reversal resulting from oil leakage caused by the mounting holes in the governor body being drilled beyond tolerances, the following must be accomplished: A. Prior to the installation of new or overhauled governors of the above models, perform the following, except that it need be accomplished only once for each governor affected, and need not be accomplished if paragraph C is complied with: 1. Remove reverse solenoid valve assembly. 2. Thoroughly clean the solenoid valve mounting hole in the governor body as described in Hamilton Standard Service Bulletin No. 518. 3. Measure the depths of the hole to the deepest point. 4. Governor bodies having a hole exceeding 0.490 inch deep shall not be used until inspected as specified by paragraph C. B. If the reverse solenoid is loosened or removed while in service, comply with paragraph A unless already accomplished. C. As soon as practicable, but not later than next overhaul of all governors of the above models, comply with the inspection outlined in Hamilton Standard Service Bulletin No. 518. Governor bodies having a wall thickness between the solenoid attaching stud hole and the low pressure relief valve passage of less than 0.035 inch should not be returned to service. If the provisions of Service Bulletin No. 518 have been complied with, it will not be necessary to repeat. (Hamilton Standard Service Bulletin No. 518 covers this same subject.) This supersedes telegraphic instructions dated July 12, 1957.
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.
53-15-02: 53-15-02 LOCKHEED: Applies to All Models 049, 149, 649 and 749 Series Aircraft. Compliance required as indicated. At the first arrival at the main base, unless already accomplished, inspect for cracks in the forward flange of the lower front spar cap at wing Station 326, left and right, with particular reference to spar cap joggle areas using dye penetrant inspection method or equivalent. 1. If crack is found in the forward flange and does not extend into the vertical leg, stop drill the crack unless it terminates in a rivet hole and make permanent repair or install the serviceable repair in accordance with LAC Drawing 325800. When serviceable repair is used, a visual inspection must be conducted at periodic intervals not to exceed 50 hours with dye penetrant inspection or equivalent method to be used at periods not to exceed every 200 hours until incorporation of the reinforcement per LAC Drawing 325667, change A or equivalent. 2. If crack in the front spar flange extends into the vertical leg, remove tank sealant as necessary for skin and web inspection using dye penetrant inspection method or equivalent. Reinforcement per LAC Drawing 325667, change A or equivalent is necessary before resuming commercial operation and normal inspection procedures. 3. If no cracks are found, reinspect using dye penetrant inspection method or equivalent on all aircraft with 10,000 hours or more total flight time, at intervals not to exceed 200 hours, and on all other aircraft at each major airframe inspection period until such time as reinforcement per LAC Drawing 325667, change A or equivalent is accomplished.
73-25-03: 73-25-03 HILLER AVIATION: Amdt. 39-1752. Applies to Hiller UH-12D helicopters certificated in all categories. Compliance required prior to further flight for all UH-12D helicopters which have been converted from the military version (H-23D) before the effective date of this AD, and at the time of conversion for those helicopters which are converted to the UH-12D after the effective date of this AD. To insure safe service life for the finite life components of the Hiller Model UH-12D Helicopters, accomplish the following: Replace the finite life components listed in Hiller Aviation's UH-12D Inspection Guide, Airworthiness Limitations Section, dated November 5, 1973, at the times specified therein with new or serviceable parts. NOTE: A copy of the finite life components list can be obtained from Hiller Aviation, 2075 West Scranton Avenue, Porterville, California, 93257, or from the FAA, Aircraft Engineering Division, P. O. Box 92007, World Way Postal Center, Los Angeles, California 90009. This amendment becomes effective January 10, 1974.