Results
46-41-01: 46-41-01 BELLANCA: (Was Mandatory Note 2 of AD-773-5.) Applies to Models 14-13, 14-13-2 Serial Numbers 1060 to 1111, Inclusive. Compliance required prior to November 15, 1946. Replace rudder bellcrank (Bellanca P/N 9817) located at the left and right ends of the rudder torque tube with parts furnished by the manufacturer which are stamped "heat-treat" in ink. (Bellanca Service Bulletin No. 2 dated August 26, 1946, covers this same subject.)
71-24-09: 71-24-09 BEECH: Amendment 39-1347. Applies to Model 56TC (Serial Numbers TG-1 thru TG-76) Airplanes. Compliance: Required as indicated, unless already accomplished. To provide information reflecting applicable operating limitations and margin between maximum structural cruising speed and never exceed speed, within the next 50 hours' time in service after the effective date of this AD, revise placards and change airspeed indicator marking as follows: 1) Install placard at lefthand cabin side, adjacent to ignition switch panel reading: "This airplane must be operated as a Normal Category airplane in compliance with the operating limitations stated in the form of placards, markings and manuals (Pilot's Check List). Occupied seats must be in upright position during takeoff and landing. Maximum weight 5990 lb. No acrobatic maneuvers including spins approved. Max. speed w/landing gear extended (normal) (TG-1 thru TG-71) - 165 m.p.h. (143 knots) (TG-72 and up) - 175 m.p.h. (152 knots) Max. speed with flaps extended (15 degrees down) - 175 m.p.h. (152 knots) Max. speed with flaps extended (normal) - 144 m.p.h. (125 knots) Max. design maneuver speed - 183 m.p.h. (159 knots) Minimum control speed single engine - 97 m.p.h. (84 knots) Max. structural cruising speed (S.L. to 20,000 ft. alt.) - 233 m.p.h. (202 knots) Max. structural cruising speed (25,000 ft. alt.) - 222 m.p.h. (193 knots) Max. structural cruising speed (30,000 ft. alt.) - 214 m.p.h. (186 knots) Never exceed speed (S.L. to 20,000 ft. alt.) - 262 m.p.h. (227 knots) Never exceed speed (25,000 ft. alt.) - 249 m.p.h. (216 knots) Never exceed speed (30,000 ft. alt.) - 240 m.p.h. (208 knots) 2) Install placard on floating instrument panel near airspeed indicator reading: "See limitations placard for 'max. structural cruise' and 'never exceed limits'." 3) Re-mark airspeed indicator to extend yellow arc from 240 m.p.h. to 233m.p.h. so that green arc does not enter this range. Beechcraft Service Instruction No. 0173-016 considers this subject. This amendment becomes effective November 30, 1971.
63-03-01: 63-03-01 BOEING AND DOUGLAS: Amdt. 532 Part 507 Federal Register February 5, 1963. Applies to Boeing Models 707-100B, 707-300B, and 720-000B Series Aircraft, and to Douglas DC-8-50 Series Aircraft With Pratt & Whitney JT3D Series Engines. \n\n\tCompliance required within the next 4,000 hours' time in service after the effective date of this AD, unless already accomplished. \n\n\tClogging of engine main oil filters by foreign matter has caused lubrication system malfunctions which have resulted in engine mechanical failures affecting safety of flight. To prevent such failures, accomplish the following: \n\n\t(a)\tFor Pratt & Whitney JT3D Series engines with serial numbers listed in Pratt & Whitney Engine Service Bulletin No. 327 dated January 8, 1962: \n\n\t\t(1)\tModify the engine oil filter assembly to provide for the installation of a differential pressure switch between the bypass port and the filter drain port, and provide an additional spring in the bypass valve to increase the pressure at which bypass occurs, in accordance with Service Bulletin No. 327, or FAA approved equivalent. \n\n\t\t(2)\tInstall a pressure switch across the engine main oil system filter, set to be actuated when the differential pressure between the inlet and outlet ports reaches a value of approximately 50 p.s.i. This change shall be accomplished in accordance with Boeing Service Bulletin No. 1586 dated April 11, 1962, for Boeing aircraft, and in accordance with Douglas Service Bulletin No. 79-11 (to be issued later) for DC-8 aircraft, or FAA approved equivalent. Prior or concurrent incorporation of (a)(1) is required with this change. \n\n\t(b)\tFor Boeing Models 707-100B, 707-300B, and 720-000B Series aircraft with serial numbers listed in Boeing Service Bulletin No. 1586 dated April 11, 1962, and for Douglas DC-8-50 Series aircraft listed in Douglas Service Bulletin DC-8 No. 79-11 (to be issued later): \n\n\t\t(1)\tProvide means in the cockpit to give corresponding indication of the actuationof the differential pressure switch on each engine in accordance with Boeing Service Bulletin No. 1586 for Boeing aircraft, and in accordance with Douglas Service Bulletin 79-11 for DC-8 aircraft, or FAA approved equivalent. \n\n\tNOTE: Any person may submit an equivalent means of compliance with the objective of this directive. Such equivalent means shall be submitted to FAA, Western Region, Attention, Chief, Engineering and Manufacturing Branch, for evaluation and approval. Adequate substantiation of equivalency will be required. If approved, the equivalent means, when accomplished, shall be deemed as compliance with (a) and (b). The objective of this directive is to provide means of preventing serious mechanical damage to engines which would affect safety of flight as a result of lubrication failure of engine main bearings. \n\n\t(c)\tWhen the modifications prescribed in (a) and (b) are accomplished or when an equivalent means of compliance is approved and accomplished, the engine oil filter inspections prescribed by AD 61-24-01 are no longer required. \n\n\t(d)\tAppropriate revisions to the FAA Airplane Flight Manual covering procedures required in connection with devices installed shall be prepared and submitted to FAA, Western Region, Attention, Chief, Engineering and Manufacturing Branch, for approval. \n\n\tThis directive effective March 7, 1963.
62-27-04: 62-27-04 DOUGLAS: Amdt. 520 Part 507 Federal Register December 20, 1962. Applies to DC-8 Standard Leading Edge Aircraft Powered With Pratt & Whitney JT3C, JT4A or Conway Engines. \n\n\tNOTE: Does not apply to aircraft with extended leading edge and to JT3D powered aircraft with standard leading edge. \n\n\tCompliance required as indicated. \n\n\tAs a result of failure of the upper inboard spar cap structure of the outboard pylons, accomplish the following: \n\n\t(a) Unless already accomplished within the last 425 hours' time in service, within the next 25 hours' time in service, inspect upper inboard spar cap structure of the outboard pylon for evidence of cracks. Gain access to the area to be inspected by removing the pylon leading edge nose cap between Station YOP 214 and 244 and access doors numbers 110, 113, 411, and 414. Using close visual or dye penetrant methods, inspect the upper inboard cap and adjacent structure for cracks in the area of Station YOP 230 and at the edgesof support fitting P/N 3647306-501. \n\n\t(b) If cracks are found, repair in accordance with Douglas Drawing 5776811 or FAA approved equivalent, prior to further flight. \n\n\t(c) If no cracks are found the inspections outlined in (a) must be repeated at periods not to exceed 500 hours' time in service from the last inspection. \n\n\t(d) The repetitive inspections may be discontinued on aircraft repaired in accordance with Douglas Drawing 5776811 and on aircraft modified to incorporate preventive rework accomplished in accordance with FAA engineering approved technical data. \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 operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Douglas DC-8 Alert Service Bulletin A54-33, Revision No. 2, dated January 24, 1964, covers this same subject.) \n\n\tThis directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated November 21, 1962. \n\n\tRevised June 8, 1963. \n\n\tRevised June 23, 1964.
67-01-03: 67-01-03 PRATT & WHITNEY: Amdt. 39-336 Part 39 Federal Register January 4, 1967. Applies to Model JT4A Series Turbojet Engines. Compliance required as indicated unless already accomplished. To prevent failure of the fuel manifold assembly, accomplish the following: (a) Within the next 3,400 hours' time in service after the effective date of this AD, inspect all P/N's 378155, 378156, 391957, 391959, 447330 and 447339 fuel manifold assemblies for cracks using the fluorescent penetrant inspection procedures outlined in Pratt & Whitney Aircraft JT4A Overhaul Manual. (1) If cracks are found, replace before further flight the fuel manifold assembly with one of the fuel manifold assemblies listed above or with a P/N 572766 or 572767 fuel manifold assembly. (2) If no cracks are found, glass bead peen all cluster tee fillets (eight places per each fuel manifold assembly) prior to return to service and thereafter at each fuel mainfold assembly overhaul in accordance with Pratt& Whitney Aircraft JT4A Overhaul Manual Temporary Revision No. 73-9 dated October 20, 1966, or Pratt & Whitney Aircraft JT4A Overhaul Manual Revision No. 39 which includes Temporary Revision No. 73-9. (b) At each fuel manifold assembly overhaul, inspect and glass bead peen in accordance with (a) all fuel manifold assemblies P/N's 572766 and 572767. NOTE: Fuel Manifold Assemblies P/N's 572766 and 572767 were glass bead peened during initial fabrication. (c) Prior to use, inspect and glass bead peen in accordance with (a) all spare fuel manifold assemblies P/N's 378155, 378156, 391957, 391959, 447330, and 447339. (Pratt & Whitney Aircraft letter dated March 23, 1966, to all operators of JT4A turbojet engines covers this subject.) This directive effective January 5, 1967.
62-27-02: 62-27-02 BELL: Amdt. 522 Part 507 Federal Register December 28, 1962. Applies to All Model 47 Series Helicopters Equipped With P/N 47-642-020-1 Wood Tail Rotor Blades. Compliance required as indicated. There have been several failures of wood tail rotor blades resulting from wood deterioration. To preclude further wood blade failures the following must be accomplished: (a) Within 50 hours' time in service after the effective date of this AD: (1) Remove wood tail rotor blades in accordance with the applicable Bell Maintenance and Overhaul (M&O) Manual. (2) Remove the fiberglass wrapping from the root end area of blades and remove the fiberglass blade covering from areas underneath the wrapping in accordance with the applicable Bell M&O instructions for repair of wood main rotor blades. Cut blade covering by lightly sanding cover as a knife or other sharp instrument can cause damage. (3) Inspect root end of blades from root end of blade to 6 inchesoutboard for: (i) Elongated bolt holes. Maximum allowable diameter 0.260 inch. (ii) Decay of wood. Detection of decay can be made visually by noting discoloration of the basic material. (Generally decay will start as a grayish discoloration and deepens to a brown color during the later stages.) (iii) Cracks in the stainless steel leading edge strip and grip plates using at least a 5-power magnifying glass. (4) Blades found with bolt hole diameters exceeding 0.260 inch, with decay, or with any cracks, shall be removed from service prior to further flight. (5) Blades without defects may be returned to service after: (i) Recovering the blade root area in accordance with patching procedures given in the applicable Bell M&O Manual; and (ii) Rewrapping the root area with two pieces of MIL-P-8013 No. 181 fiberglass cloth 2 x 27 inches in accordance with Bell Service Bulletin No. 75 dated September 17, 1951. (b) Blades returned to serviceafter compliance with (a) shall be retired from service prior to the accumulation of 200 hours' time in service since reinstallation in accordance with (a). This directive effective January 29, 1963.
68-02-04: 68-02-04 FAIRCHILD-HILLER: Amendment 39-544. Applies to Type FH-1100 Helicopters, Serial Numbers 9 through 49. Compliance required as indicated. To prevent fatigue failures of the Cyclic Input Swashplate Ring, P/N 24-34205-3, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, unless already accomplished, and thereafter at intervals not to exceed 25 hours' time in service from the last inspection, visually inspect the cyclic input swashplate ring, P/N 24-34205-3, in accordance with Part A (excluding paragraph 5) of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967, or later revisions approved by the Chief, Engineering & Manufacturing Branch, Federal Aviation Administration, Eastern Region. Equivalent inspections may be approved by an FAA maintenance inspector. (b) If a crack is found, remove the ring from service prior to further flight. (c) Accomplish the following on rings that have not been reworked in accordance with Part B of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967: (1) Remove from service or rework in accordance with Part B of the aforementioned Letter rings with 75 or more hours time in service on the effective date of this AD within the next 25 hours' time in service. (2) Remove from service or rework in accordance with Part B of the aforementioned Letter all other rings before the accumulation of 100 hours' time in service. (d) Rings which have been modified in accordance with Part B of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967, or in accordance with any other method approved by the Chief, Engineering & Manufacturing Branch, Federal Aviation Administration, Eastern Region may be continued in service until the accumulation of 750 hours' time in service. The 25-hour repetitive inspection of (a) may be discontinued on modified rings when a satisfactory inspection for crackshas been accomplished on the ring after it has been modified, by the dye penetrant method or an equivalent approved by an FAA maintenance inspector. This AD is effective January 27, 1968.
62-25-02: 62-25-02 DOUGLAS: Amdt. 510 Part 507 Federal Register November 28, 1962. Applies to All Models DC-6 and DC-7 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tDue to failure of a main gear shock strut cylinder and numerous cases of cracks in the 0.125-inch radii next to the torque link lugs on the cylinders, and on the piston tube axle fittings, the following shall be accomplished. \n\n\t(a) DC-6 Series Aircraft. \n\n\t\t(1) With 30,000 or more hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter within each 300 hours' time in service from the last inspection. \n\n\t\t(2) With less than 30,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) priorto the accumulation of 30,200 hours' time in service, and thereafter within each 300 hours' time in service. \n\n\t(b) DC-7 Series aircraft. \n\n\t\t(1) With more than 15,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter within each 300 hours' time in service from the last inspection. \n\n\t\t(2) With less than 15,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) prior to the accumulation of 15,200 hours' time in service, and thereafter within each 300 hours' time in service. \n\n\t(c) Inspect, using dye penetrant, or magnetic particle, or FAA approved equivalent, for cracks in the 0.125-inch radii at the edges of the torque link lugs in the main landing gear shock strut cylinder and the piston tube axle fitting. \n\n\t(d) If cracks are found, they may be removed by reworking the 0.125-inch radius in accordance with the instructions contained in Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto or Douglas Service Engineering Letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D and 534E attached thereto. If cracks cannot be removed without exceeding limits specified in the Douglas sketches, the gear must be replaced prior to further flight. Parts that can be reworked, and those in which no cracks are found, must be repainted with zinc chromate primer and aluminized lacquer before they are returned to service. \n\n\t(e) When the 0.125-inch radii at the edges of the torque link lugs on the strut cylinders and axle fittings have been enlarged to 0.250-inch radii, holding the tolerances described in Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto, or Douglas Service Engineering letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D and 534E attached thereto, and the parts are refinished as described in (d), the repetitive inspections required herein may be discontinued. \n\n\t(f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering & 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. \n\n\t(Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto, or Douglas Service Engineering letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D, and 534E attached thereto, covers thissame subject.) \n\n\tThis directive effective November 28, 1962. \n\n\tRevised April 4, 1963.
2008-17-13: We are adopting a new airworthiness directive (AD) for certain Boeing Model 737-100, -200, -200C, -300, -400, and -500 series airplanes. This AD requires replacing the existing straight-to-90- degree hose assembly for the Lavatory "A'' water supply. The replacement is a new straight hose assembly and a separate 90-degree elbow fitting. This AD results from a report of a separated hose assembly for the passenger water system. We are issuing this AD to prevent a water leak into the flight deck ceiling, which could result in an electrical short and possible loss of several functions essential to safe flight.
2008-17-19: 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 ATR 42-300 experienced a collapse of the Right (RH) Main Landing Gear (MLG) when taxiing, caused by failure of the side brace assembly. Investigations revealed a crack propagation that occurred from a corrosion pit, in a very high stressed area of the upper arm. * * * * * * * * The unsafe condition is cracking of the upper arms of the secondary side brace assemblies of the MLG, which could result in collapse of the MLG during takeoff or landing, damage to the airplane, and possible injury to the flightcrew and passengers. We are issuing this AD to require actions to correct the unsafe condition on these products.