Results
2005-19-03: The FAA is superseding an existing airworthiness directive (AD), which applies to all BAe Systems (Operations) Limited Model ATP airplanes. That AD currently requires revising the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness to incorporate life limits for certain items and inspections to detect fatigue cracking in certain structures. This new AD requires a revision to the ALS of the Instructions for Continued Airworthiness to incorporate new inspections to detect fatigue cracking of certain significant structural items (SSIs) and to revise life limits for certain equipment and various components. This AD is prompted by a determination that existing inspection techniques are not adequate for certain SSIs and by the revision of certain life limits. We are issuing this AD to detect and correct fatigue cracking of certain structural elements, which could adversely affect the structural integrity of these airplanes.
2010-24-06: We are superseding an existing airworthiness directive (AD) that applies to the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an airworthiness authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as Subsequent to accidents involving Fuel Tank System explosions in flight * * * and on ground, * * * Special Federal Aviation Regulation 88 (SFAR88) * * * required a safety review of the aircraft Fuel Tank System * * *. * * * * * Fuel Airworthiness Limitations are items arising from a systems safety analysis that have been shown to have failure mode(s) associated with an `unsafe condition' * * *. These are identified in Failure Conditions for which an unacceptable probability of ignition risk could exist if specific tasks and/or practices are not performed in accordance with the manufacturers' requirements. * * * * * Weare issuing this AD to require actions to correct the unsafe condition on these products.
98-07-08: This amendment adopts a new airworthiness directive (AD) that is applicable to all IAI, Ltd., Model 1125 Westwind Astra and Astra SPX series airplanes. This action requires disabling of the baggage compartment electrical heating blankets. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified in this AD are intended to prevent overheating of the electrical heating blankets, and consequent increased risk of fire in the baggage compartment.
98-07-05: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires replacement of the main landing gear (MLG) trunnion fittings with reinforced trunnion fittings. This amendment is prompted by the issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent collapse of the MLG due to fatigue cracking of the MLG trunnion fittings.
79-22-01: 79-22-01 BELLANCA: Amendment 39-3596. Compliance is required within the next 30 days or 10 hours of aircraft time in service, whichever occurs first, after the effective date of this AD, unless already accomplished. To prevent exhaust system cracking, accomplish the following on Bellanca Model 7ECA (S/N 985-74 thru 1319-79), 8KCAB (S/N 120-74 thru 550-79 equipped with Lycoming AEIO-360 series engine), and 8GCBC(S/N 1-74 thru 323-79) aircraft: 1. Remove the upper and lower engine cowling. 2. Inspect exhaust system with particular attention to the welded area between the riser tube and the exhaust flange, for cracks, fractures or evidence of exhaust leakage. Remove the heater shroud and inspect the muffler body for cracks, fractures or evidence of exhaust leakage. If any exhaust system component is cracked or otherwise damaged, remove the exhaust system and repair/replace damaged parts in accordance with FAA Advisory Circular 43.13-1A. 3. Loosen exhaust port studnuts several turns; check bead clamps for tightness such that the clamps cannot rotate on the exhaust system with hand pressure. The riser flanges (1) must have equal spacing to the exhaust port pad at both studs (a small amount of flange bow is acceptable), (2) must be free to move up and down on the exhaust port studs without binding and (3) must all contact the exhaust port pads together. 4. If any of the alignment checks are unsatisfactory, determine the cause for the misalignment and repair or replace the part as required. 5. Assemble exhaust system and install on engine with loose exhaust port stud nuts and bead clamp bolts. Torque exhaust port stud nuts to the correct value. Tighten bead clamp bolts until clamps secure risers to exhaust system but allow clamps to rotate with hand pressure; the bead clamps should not be rigidly clamped to the tubes but should be able to rotate on the tubes with moderate hand pressure on the clamp assembly. NOTE: Torque all exhaust port stud nuts evenly and tighten bead clamp bolts evenly to insure uniform loads within the exhaust system parts; torquing bolts individually can cause very large stresses. 6. Inspect exhaust system for proper clearance between ducts, wiring, controls, etc. before reinstallation of the cowling. Install lower cowling and inspect for proper clearance between exhaust outlet and cowl. 7. Reinstall the lower and upper engine cowling. Bellanca Service Letter Number C-138 covers this same subject. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Great Lakes Engine.. This amendment becomes effective October 29, 1979.
98-06-25: This amendment adopts a new airworthiness directive (AD) that applies to Fairchild Aircraft Inc. (Fairchild) Models SA226-AT, SA226-TC, SA227-AC, and SA227-AT airplanes. This action would require inspecting the cargo door lower belt frames at the cargo latch receptacles for cracks in the belt frames, repairing any cracks, and reinforcing the cargo door lower belt frames by installing doublers. The AD is the result of a decompression incident during flight caused by fatigue at the bottom of the cargo door on a Fairchild Model SA226-TC. The actions specified by this AD are intended to prevent the failure of the cargo door in flight, which could cause decompression injuries to passengers and substantial structural damage to the airplane.
2017-10-22: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-600, -700, -700C, -800, and -900 series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the web lap splices in the aft pressure bulkhead are subject to widespread fatigue damage (WFD). This AD requires repetitive inspections of the web lap splices in the aft pressure bulkhead for cracking of the fastener holes, and repair if necessary. We are issuing this AD to address the unsafe condition on these products.
2017-09-02: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Deutschland GmbH Helicopters (Airbus Helicopters) Model MBB-BK 117 C-2 and MBB-BK 117 D-2 helicopters. This AD requires installing rivets to the air inlet cover rings (rings). This AD was prompted by reports of rings detaching. The actions of this AD are intended to prevent the unsafe condition on these products.
79-01-01: 79-01-01 BEECH: Amendment 39-3386. Applies to the following models and serial number airplanes, certificated in all categories, except those that have previously complied with AD 74-13-02: MODELS SERIAL NUMBERS ITEM NUMBERS (Refers to Items in Table I) F33A CE-447 through CE-684 (1), (2) F33A CE-372 (2) F33C CJ-31 through CJ-128 (1), (2) G33 CD-1292 through CD-1304 (1), (2) 35-C33 CD-970 (1), (2) 35-C33, 35-C33A CD-1066 & CE-62 (2) 35-B33 CD-700 (2) V35B D-9478 through D-9480, D-9482 through D-9531, D-9533 through D-9536 and D-9539 through D-9968 (1) A36 E-323 (2) A36 E-441 through E-954 (1), (2) D55 & D55A TE-485, TE-615, TE-703 and TE-710 (2) E55 & E55A TE-790, TE-911 (2) E55 & E55A TE-905, TE-923 through TE-937, TE-939, TE-940, TE-942, and TE-943 through TE-1084 (1), (2) 56TC TG-78 and TG-83 (2) 58 & 58A TH-49, TH-77, TH-114, TH-180, TH-217, TH-324 (2) 58 & 58A TH-145, TH-340 and TH-343 through TH-779(1), (2) 58P & 58PA TJ-2 through TJ-97 (1), (2) 58TC & 58TCA TK-1 through TK-37 (1), (2) 65-B80 LD-441 through LD-511 (1), (3), (5), (6), (7) 70 LB-31 through LB-35 (1), (3), (5), (6), (7) B90 & C90 LJ-500 through LJ-706 (1), (3), (5), (6), (7) E90 LW-1 through LW-209 (1), (3), (5), (6), (7) 95-A55 TC-241 (1) 95-B55 & 95-B55A TC-661 and TC-1510 (2) 95-B55 & 95-B55A TC-1570 through TC-2011 (1), (2) 95-C55 & 95-C55A TE-271 (2) 95-C55 & 95-C55A TE-415 (1) 99, 99A, A99A, B99 U-134 through U-164 (1), (3), (4), (5), (6), (7) 100 & A100 B-61 through B-178, B-180 through B-204 and B-206 through B-231 (1), (3), (4), (5), (6), (7) B100 BE-1 through BE-19 (1), (3), (4), (5), (6), (7) 200 BB-2 through BB-185, & BB-187 through BB-199 (3), (4), (8), (9), (10) COMPLIANCE: Required as indicated unless already accomplished. To assure continued structural integrity of certain screws in the airplane control system, in accordance with instructions set forth herein and in Beechcraft Service Instructions No. 0629- 150, Revision II or later approved revisions, accomplish the following: A) Within 100 hours time-in-service after the effective date of this AD: For affected models and serial numbers, determine those Item Numbers applicable thereto by referring to the column above entitled "Item Numbers". Next, determine the number, type and location of screws by referring to Table I of this AD (see below). Replace each screw at the specified location except that an existing screw need not be replaced if inspection shows that the head has at least one raised or depressed "X" mark. Replace the screws one at a time at the hinge brackets to retain original alignment of the bracket. B) Any equivalent means of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. TABLE I Item 1. Two AN502-10-10 screws securing each hinge bracket at leading edge of rudder or ruddervator. Item 2. Two AN502-10-10 screws securing each hinge bracket at leading edge of elevator. Item 3. Two AN502-10-12 screws securing elevator control cables to forward bellcrank under pilot's compartment left floorboard. Item 4. Two AN503-8-8 screws securing elevator control cables to aft bellcrank in fuselage tail section. Item 5. Two AN503-8-10 screws securing rudder control cables to forward bellcrank under pilot's compartment floorboard aft of left rudder pedals. Item 6. Two AN503-8-6 screws securing rudder control cables to aft bellcrank in the fuselage tail section. Item 7. Two AN503-8-6 screws securing aileron control cables to bellcrank under floorboard just forward of the rear spar. Item 8. Two AN503-8-8 screws securing rudder control cables to forward bellcrank under pilot's compartment floorboard aft of left rudder pedals. Item 9. Two AN503-8-8 screws securing rudder control cables to the aft quadrant assembly in the fuselage tail section. Item 10. Two AN503-8-8 screws securing aileron control cables to bellcrank under floorboard just forward of the rear spar. This AD supersedes AD 74-13-02, Amendment 39-1874 (39 FR 20784). NOTE: Scheduling accomplishment of this Airworthiness Directive simultaneously with accomplishment of Airworthiness Directive No. 78-22-05 will result in a savings of man-hours. This amendment becomes effective January 8, 1979.
63-22-01: 63-22-01 GRUMMAN: Amdt. 632 Part 507 Federal Register October 25, 1963. Applies to Model G-164 Aircraft Serial Numbers 1 through 220. Compliance required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. To prevent damage to fuel and oil lines from spray pump fan blades, rivet two 0.063 inch thick 2024-T3 aluminum alloy sheets approximately 6 inches wide by 18 and 28 inches long each, to the left and right lower accessory cowl panels using at least 10 and 16 AD 4 type rivets respectively. Center the plate's long dimension opposite the spray pump fan blade arc. (Grumman Aircraft Engineering Corporation Model G-164 Service Bulletin No. 23, dated June 17, 1963, pertains to this same subject and also lists the availability of Grumman armor plates P/N's A1624-111 and -113.) This directive effective October 25, 1963.
2017-09-05: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Model AS332C, AS332C1, AS332L, AS332L1, AS332L2, and EC225LP helicopters. This AD requires repetitively checking screws in the emergency flotation gear. This AD is prompted by a report that a screw ruptured on a Model AS332 helicopter's emergency flotation gear. These actions are intended to correct an unsafe condition on these products.
83-25-02: 83-25-02 HILLER AVIATION: Amendment 39-4784. Applies to Models UH-12D, UH-12E, UH12E 4 Place, H-23F, and OH-23G series helicopters equipped with main rotor blades P/N 53200-03, serial numbers 038 through 286, certificated in any category. Hiller Model UH-12 series helicopters converted to turbine power by STC SH177WE or SK178WE are affected by this AD. Compliance is required as indicated (unless already accomplished). To prevent possible main rotor blade failure due to the delamination of the main rotor blade root doublers or skin bond separation, accomplish the following: (a) Prior to further flight after the effective date of this AD, and prior to each subsequent flight, visually check the main rotor blade root doubler area for evidence of delamination. This check includes all doublers on the leading edge and trailing edge surfaces, upper and lower, from the fork attachment plate to the tip of the outboard doubler. Delamination will be evidenced by a lifting of the doubler edges and/or cracks in the painted surface. (b) If the check indicates delamination, prior to further flight, replace the blade with a like serviceable part or consult the Manager, Western Aircraft Certification Office, for instructions and disposition of the blade. Descriptive data regarding the extent of the delamination must be made available to the FAA. (c) The check required by paragraph (a) of this AD may be performed by the pilot, provided his logbook is endorsed by a properly rated mechanic stating that the pilot has been trained to conduct the check. (d) Within the next 50 hours time in service after the effective date of this AD, and thereafter at intervals not to exceed 50 hours time in service from the last inspection, inspect the main rotor blade root doubler area for evidence of delamination in accordance with paragraph 2.B of Hiller Aviation Service Letter 51-3A, Revision 1, dated October 17, 1983, or an equivalent as noted in paragraph (i).If delamination is found, comply with paragraph (b). (e) Within the next 100 hours time in service after the effective date of this AD, and thereafter at intervals not to exceed 100 hours time in service from the last inspection, inspect main rotor blades for evidence of skin bond separation in accordance with paragraph 2.A of Hiller Aviation Service Bulletin UH-12-51-7, Revision 1, dated October 3, 1983, or an equivalent as noted in paragraph (i). NOTE: The main rotor blade bond separation occurred on the blade tip leading edge and extended a distance of approximately 2 feet. (f) If the main rotor blade inspection indicates skin bond separation, prior to further flight, replace blade with a like serviceable part or consult the Manager, Western Aircraft Certification Office, for instructions and disposition of the blade. Descriptive data regarding the extent of the skin bond separation must be made available to the FAA. (g) Mark all rejected blades "UNAIRWORTHY"on the blade upper and lower surface at the approximate mid span with lettering at least 2 inches high, and using a metal stamp, impress "UNAIRWORTHY" on the data plate. (h) All checks and inspections are to be recorded in aircraft maintenance records. (i) Alternative inspections, modifications, or other actions which provide an equivalent level of safety may be used when approved by the Manager, Western Aircraft Certification Office, FAA, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009. This supersedes Amendment 39-3897 (45 FR 56331), as amended by Amendment 39- 3935 (45 FR 65997), AD 80-18-01 R1. This amendment becomes effective December 29, 1983.
2017-09-07: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Deutschland GmbH (Airbus Helicopters) Model MBB-BK117 C-2 helicopters. This AD requires inspecting the pilot collective wiring harness and was prompted by a report that the heat-shrinkable sleeve prevented the twist grip on the collective from being fully engaged during a flight test. The actions of this AD are intended to prevent an unsafe condition on these products.
2017-08-11: We are superseding airworthiness directive (AD) 2012-04-01 for all Rolls-Royce plc (RR) RB211-Trent 800 model turbofan engines. AD 2012-04-01 required removal from service of certain critical engine rotating parts based on reduced life limits. This AD makes additional revisions to the life limits of certain critical engine rotating parts. This AD was prompted by RR further revising the life limits of certain critical engine rotating parts. We are issuing this AD to correct the unsafe condition on these products.
96-13-13: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 0100 series airplanes, that currently requires certain maximum brake wear limits to be incorporated into the FAA-approved maintenance inspection program. That AD also currently requires that the Airplane Flight Manual (AFM) be revised to include certain procedures concerning operations in the event of a rejected takeoff (RTO). This amendment requires the incorporation of new maximum brake wear limits for additional brake units into the FAA-approved maintenance program. This action also deletes the previous requirement for the AFM revision. This amendment is prompted by the determination of the maximum allowable brake wear limits for additional brake unit part numbers. The actions specified by the AD are intended to prevent the loss of brake effectiveness during a high energy RTO.
79-05-05: 79-05-05 LOCKHEED-GEORGIA COMPANY: Amendment 39-3427 as amended by Amendment 39-3641. Applies to Model 382 series aircraft certificated in all categories, S/N 3946 and 4101 through 4655, which have not incorporated Lockheed Service Bulletin 382-76-4. Compliance is required as indicated, unless otherwise accomplished: (a) For throttle cables with 3000 hours or more time in service, accomplish the following: (1) Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished within the past 1800 hours' time in service, perform an inspection and replace, as necessary, the throttle cables located at FS-245 (upper and lower) and at the engine horse collar in accordance with Lockheed Maintenance Manual SMP 515A, Work Card No. SP-77 dated 1/31/75. (2) If as a result of an inspection required by this AD it is found that the throttle cables are not within the criteria of paragraph (a)(1), replace cable before further flight. Cablereplacement may be made in accordance with Lockheed Maintenance Manual SMP 581, Chapter 76. (3) Repeat the above inspection and cable replacement, as necessary, at each 1800 hours' time in service after the initial inspection. An alternate compliance time must be substantiated by the operator and recommended by the assigned FAA Principal Maintenance Inspector and approved by the Chief, Engineering and Manufacturing Branch, Southern Region. (b) Lockheed Service Bulletin 382-76-4 provides for the installation of larger diameter pulleys and replacement of cables. If compliance with this Service Bulletin is shown or later FAA approved revision or an equivalent modification approved by the Chief, Engineering and Manufacturing Branch, Southern Region, the above inspections will no longer be required. (c) Special flight permits may be issued in accordance with FAR's 21.197 and 21.199 to authorize operation of an airplane to a base for the accomplishment of the modifications required by this AD. Amendment 39-3427 became effective March 12, 1979. This Amendment 39-3641 becomes effective January 2, 1980.
64-17-01: 64-17-01\tBOEING: Amdt. 766 Part 507 Federal Register July 17, 1964. Applies to Models 707 and 720 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tFailures and malfunctions occurring in the fuel dump chute limit switches of the retractable dump chute system have caused switch overheating sufficient to scorch the connecting wires. To prevent this, accomplish the following: \n\n\t(a)\tWithin 500 hours' time in service after the effective date of this AD, and thereafter at periods not to exceed 500 hours' time in service, perform a resistance check on the fuel dump chute extend limit switch, P/N's 5EN11-6B or H11-186, and on the fuel dump chute up limit switch, P/N's 2EN15-6 or H11-162. If any switch is found defective, replace before further flight. (Boeing telegraphic messages 6-7161-1-9101 and 6-7161-1-9159 sent to all operators on February 3, and March 28, 1963, respectively, cover this resistance check.) Upon compliance with paragraph (b) the provisions of this paragraph may be discontinued. \n\n\t(b)\tWithin the next 2,500 hours' time in service after August 17, 1964, replace limit switches on each fuel dump chute that do not incorporate a sealed receptacle with improved switches incorporating a sealed receptacle and replace the switch pigtail type wires with new wires and a mating sealed connector, in accordance with Boeing Service Bulletin No. 1877, or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(c)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, 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(Boeing Service Bulletin 1877 dated January 23, 1964, and Boeing telegraphic messages 6- 7161-1-9101 and 6-7161-1-9159 dated February 3, and March 28, 1963, respectively, cover this subject.)\n \n\tThis directive effective August 17, 1964. \n\n\tRevised June 25, 1966.
2005-20-06: The FAA is adopting a new airworthiness directive (AD) for all the Airbus models identified above. This AD requires modifying the electrical power supply logic for the integral lighting of the standby horizon indicator in the cockpit, accomplishing repetitive operational tests of the integral lighting logic system, and performing corrective action if necessary. This AD is prompted by a report of temporary loss of six cathode ray tube (CRT) flight displays and the integral lighting of the standby horizon indicator backlight in the cockpit during takeoff, due to failure of the normal electrical power circuit. That power circuit supplies power to both the CRTs and the standby horizon indicator backlight. We are issuing this AD to prevent loss of the integral lighting due to failure of the normal electrical power circuit, which could result in inability of the pilot to read the backup attitude information during takeoff, and possible deviation from the intended flight path.
2002-08-03: This amendment supersedes an existing airworthiness directive (AD) for Enstrom Helicopter Corporation (EHC) Model F-28, F-28A, F-28C, F-28F, 280, 280C, 280F, and 280FX helicopters. That AD currently requires determining the radius of the shaft fillet, performing certain visual and dye-penetrant inspections before further flight, and replacing certain main rotor transmissions. This amendment requires the same actions as the previous AD, adds additional main rotor gear box part numbers, and corrects various errors contained in the current AD. This amendment is prompted by a commenter who noted that two additional main rotor gear box part numbers should have been included in the AD. The actions specified by this AD are intended to prevent shaft failure and subsequent loss of control of the helicopter.
91-08-04: 91-08-04 AVIATECH INC.: Amendment 39-6925. Docket No. 90-ANE-14. Applicability: Aviatech Inc., TSO-C39a Pilot and Co-Pilot seats Model 394 series as follows: Affected Pilot and Co-Pilot seats Part Numbers (P/N), 394 (000) 01,-02, -03, -04, -05, -06, and 394-(001) -01, -02, -03, -04,-05, -06, installed on, but not limited to, Boeing of Canada, Ltd., deHavilland Division, Model DHC-8. Compliance: Required within the next 90 days after the effective date of this AD, unless already accomplished. To prevent the pilot and co-pilot shoulder harness from becoming excessively worn and frayed, which could result in failure of the shoulder harness, accomplish the following: (a) Inspect the pilot and co-pilot seats to determine if any of the above P/Ns are inscribed on the FAA - TSO-C39a tag. (b) Modify all seats with any of the above P/Ns by incorporating the Aviatech Inc. Kit No. 394-25-002, which requires installing new backrest cushion-covers and a roller-guide for the shoulder harness, in accordance with Aviatech Service Bulletin No. 2, Model 394, Revision A, dated March 1, 1990 (Paragraph 2, Accomplishment Instructions). (c) Special flight permits may be issued in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (d) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance specified in this AD may be approved by the Manager, New York Aircraft Certification Office, Engine & Propeller Directorate, Aircraft Certification Service, FAA, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581. The modifications and installation procedures shall be done in accordance with the following Aviatech Service Bulletin (SB): DOCUMENT NO. PAGE NO. ISSUE/REVISION DATE Aviatech Inc. 1, 2, 4, 8, 10 Rev. A March 1, 1990 SB No. 2 3, 5, 6, 7, 9, 11 Original June 29, 1989 The incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Aviatech Inc., 2400 Guenette Street, St. Laurent, Quebec, Canada H4R 2H2. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-6925, AD 91-08-04) becomes effective on May 6, 1991. AVIATECH 1
77-08-04: 77-08-04 BEECH: Amendment 39-2875. Applies to Models 58P and 58PA (Serial Numbers TJ-46 thru TJ-74, TJ-76, TJ-78 thru TJ-80, TJ-82 thru TJ-85, TJ-88 and TJ-89) and Models 58TC and 58TCA (Serial Numbers TK-7 thru TK-18, TK-20, TK-21, TK-25, TK-27, TK-30, TK-32, TK-34, TK-37 and TK-38) airplanes certified in all categories. Compliance: Required as indicated, unless already accomplished. To assure that the correct bolts are installed in the engine mount fittings, within the next 100 hours' time in service after the effective date of this AD, accomplish the following: A. Inspect, and if necessary, replace the engine mount fitting bolts in accordance with the following procedures: 1. Remove the access covers just aft of the engine firewall on the upper and lower inboard and outboard sides of both right and left engine nacelles to gain access to the engine mount fittings (four per engine). 2. Visually inspect the two attach bolts in each fitting (eight bolts for each engine) to determine that NAS bolts are installed. NAS bolts may be identified by a concave indentation on the bolt head and NAS stamped on the bolt head. 3. If NAS bolts are installed, reinstall the access covers and proceed to Paragraph A.6. below. If NAS bolts are not installed, proceed with Paragraphs A.4. thru A.6. below. 4. Using engine hoist and sling as noted in the engine removal section of the applicable Beech Maintenance Manual, lift the engine weight off of the engine mount fittings. 5. Remove any of the engine mount fitting bolts that are not NAS bolts (remove and replace one bolt at a time). Discard the bolts and self-locking nuts removed and install new NAS1104-7 or NAS1104-7M bolts using new AN960-416L washers and new MS21042-4 self-locking nuts. Torque bolts to 50 to 70 inch pounds. Reinstall access covers and proceed to Paragraph A.6. 6. Make proper entry in the aircraft maintenance records that are to be transferred with the aircraft showing compliance with this AD. B. Compliance time for this AD may be extended up to 10 hours to a maximum of 110 hours to allow compliance with this AD at previously scheduled maintenance periods. C. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. D. Beechcraft Service Instructions No. 0875-038 covers the subject matter of this AD. This amendment becomes effective May 26, 1977.
2005-20-08: The FAA is adopting a new airworthiness directive (AD) for certain Airbus transport category airplanes, identified above. This AD requires an inspection to determine if a certain lower pin (p-pin) of the retraction actuator of the main landing gear (MLG) is installed. If the affected p-pin is installed, this AD requires a one-time inspection of the p-pin for correct grease hole position and cracking; repetitive daily inspections for pin migration; and eventual replacement of all p- pins with new p-pins. For any p-pin that is cracked or shows pin migration, this AD requires immediate replacement with a new p-pin. Replacing the p-pin with one that is correctly manufactured (i.e., that has the correct grease hole position) is terminating action for the repetitive inspections. This AD results from a report that a cracked p- pin was found when the MLG was removed for overhaul. We are issuing this AD to prevent failure of the p-pin, which could result in degradation of the MLG structural integrity and possible hazardous landing.
2002-07-11: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-200 and -200C series airplanes, that requires repetitive inspections to find cracking of certain fuselage lap joint areas, and repair of any cracking found. This amendment also requires eventual modification of those areas, which constitutes terminating action for the repetitive inspections. This action is necessary to find and fix cracking of certain fuselage lap joint areas, which could result in rapid decompression of the airplane. This action is intended to address the identified unsafe condition.
98-06-22: This amendment supersedes an existing airworthiness directive (AD), applicable to all Airbus Model A310 series airplanes, that currently requires measurement of the force required to move the interior control handle of the emergency exit doors, and various follow-on corrective actions, if necessary. This amendment adds repetitive functional tests to measure the force necessary to move the interior control handle of the emergency exit doors; and requires adjusting an emergency exit door or replacing the bearing of the door lifting mechanism, if necessary. This amendment is prompted by reports of seizure of a bearing and increased door handle forces that were outside the limits of the required hand forces due to seizure of two teflon line bearings on the lifting shaft. The actions specified by this AD are intended to prevent impeding passenger evacuation during an emergency due to difficulty in lifting the interior control handle that is used to open the emergency exit door.
59-09-04: 59-09-04 VERTOL: Applies to all Models 44A, B, and C Helicopters. Compliance required as indicated, but no later than May 15, 1959. (1) A failure of the low velocity fan disc assembly, P/N 42P6422, has recently been experienced in a Vertol 44 helicopter. The failure originated as a crack at the blade retaining bolt hole in the disc flange forward face, progressing radially into the flange and disc radius and continuing circumferentially in the radius area for 14 to 16 inches. This portion of the fan disc, containing six fan blades, then separated from the fan and was thrown through the left-hand side of the fuselage. (2) To preclude recurrence of said failures, fan disc assemblies, P/N 42P6422, must be retired at 400 hours' time. (3) It is permissible to operate the fan to 500 hours, providing the following requirements are complied with at, or before, 400 hours' time: a. Remove spinner, P/N 42P6420. b. Visually inspect the disc flange in the blade retaining bolt hole area for evidence of cracks. c. Reassemble, adding AN 960-416 washers under the 32 nuts in place of removed spinner. (The spinner must not be reinstalled, to facilitate future inspections.) d. Inspect the above-mentioned areas for evidence of cracks every four (4) flight-hours. e. Any evidence of cracks will be cause for immediate retirement of the fan disc assembly. Fans removed from service because of cracks will be destroyed or mutilated so as to preclude the possibility of their being returned to service. (Vertol Service Department Technical Memorandum No. 1901 covers this same subject.)