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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.