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93-02-12: 93-02-12 BEECH AIRCRAFT CORPORATION: Amendment 39-8495. Docket 92-NM-183-AD. Applicability: Model 400A airplanes; serial numbers RK-2 through RK-29, inclusive, RK-31, and RK-32; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent the loss of standby power and the possibility of an electrical fire, accomplish the following: (a) Within 100 hours time-in-service after the effective date of this AD, inspect the left-hand interstage turbine temperature (LH ITT) circuit breaker wiring, in accordance with Beechcraft Service Bulletin No. 2458 (ATA Code 39-10), dated August 1992. Prior to further flight, correct any discrepancies found, in accordance with the service bulletin. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Wichita Aircraft Certification Office (ACO), FAA, Small Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Wichita ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Wichita ACO. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The inspection shall be done in accordance with Beechcraft Service Bulletin No. 2458 (ATA Code 39-10), dated August 1992. (NOTE: The issue date of Beechcraft Service Bulletin No. 2458 is indicated only on "page 1 of 5"; no other page is dated.) This 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 Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201-0085.Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Small Airplane Directorate, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on March 26, 1993.
74-25-05: 74-25-05 SIKORSKY: Amendment 39-2030. Applies to S-61L, S-61N, S-61NM and S- 61R helicopters certificated in all categories including Military Type CH-3C helicopters equipped with S6115-20501 series, S6115-20601 series, S6117-20101 series, and S6188-15001 series main rotor blades except for blades with serial numbers subsequent to 61M-6350-6105 and blades which have had all main rotor blade pockets replaced by Sikorsky Aircraft after October 30, 1972. Compliance required within 5 hours time in service after the effective date of this AD. The maximum never exceed speed is hereby reduced to 110 knots. Install placards as close as practicable to the airspeed indicators reading as follows: "Never exceed speed 110 knots." This supersedes telegram dated October 8, 1974 and supersedes the maximum never exceed speed established by Airworthiness Directive 74-20-7, Amendment 39-1989. This amendment becomes effective upon publication in the FEDERAL REGISTER for all persons except those to whom it was made effective immediately by telegram dated November 6, 1974.
2025-06-08: The FAA is adopting a new airworthiness directive (AD) for all Deutsche Aircraft GmbH (Type Certificate previously held by 328 Support Services GmbH; AvCraft Aerospace GmbH; Fairchild Dornier GmbH; Dornier Luftfahrt GmbH) Model 328-100 and Model 328-300 airplanes. This AD was prompted by a report of a nose landing gear (NLG) uplock bracket assembly cracking. This AD requires an inspection of the affected part and applicable on-condition actions, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference (IBR). The FAA is issuing this AD to address the unsafe condition on these products.
2013-07-09: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-700, -700C, -800, and -900ER series airplanes, Model 747-400F series airplanes, and Model 767-200 and -300 series airplanes. This AD was prompted by reports indicating that certain crew oxygen mask stowage box units were possibly delivered with a burr in the inlet fitting. The burr might break loose during test or operation, and might pose an ignition source or cause an inlet valve to jam. This final rule adds a step to identify and label certain crew oxygen mask stowage box units that have already been inspected and reworked by the supplier, and allows operators to install new or serviceable crew oxygen mask stowage box units, and requires a general visual inspection for affected serial numbers of the crew oxygen mask stowage box units, and replacement or re-identification as necessary. We are issuing this AD to prevent an ignition source, which could result in an oxygen-fed fire; or an inlet valve jam in a crew oxygen mask stowage box unit, which could result in restricted flow of oxygen.
2000-05-06: This amendment adopts a new airworthiness directive (AD), applicable to certain Raytheon (Beech) Model 400A and 400T series airplanes, that requires a one-time inspection to detect incorrect wiring of the engine fire extinguisher bottle squibs, and corrective action, if necessary. It also requires a modification to the wiring and the addition of wire harness and bottle labeling for future reference. This amendment is prompted by reports of incorrect wiring of the engine fire extinguisher bottle squibs. The actions specified by this AD are intended to prevent failure of the engine fire extinguisher bottle to discharge, or discharge of the wrong engine fire extinguisher bottle.
91-26-09: 91-26-09 McDONNELL DOUGLAS: Amendment 39-8122. Docket No. 91-NM-136-AD. Supersedes AD 91-02-13, Amendment 39-6867. \n\n\tApplicability: Model DC-9-10, -20, -30, -40, and -50 series airplanes and C-9 (Military) series airplanes, operating in a passenger or passenger/cargo configuration, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tNOTE: The requirements of this AD become applicable at the time an airplane in an all-cargo configuration is converted to a passenger or passenger/cargo configuration. \n\n\tTo prevent failure of the tailcone release system, accomplish the following: \n\n\t(a)\tWithin 60 days after February 11, 1991 (the effective date of AD 91-02-13, Amendment 39-6867), unless previously accomplished within the last 60 days, inspect the interior and exterior tailcone release handles for cracks, in accordance with the Accomplishment Instructions of McDonnell Douglas Alert Service Bulletin A53-242, dated December20, 1990, or Alert Service Bulletin A53-243, Revision 1, dated February 8, 1991; and accomplish a tailcone release system functional test in accordance with the Accomplishment Instructions of McDonnell Douglas Alert Service Bulletin A53-243, dated January 10, 1991, or Revision 1, dated February 8, 1991. \n\n\t\t(1)\tCracked or broken tailcone release handles must be replaced prior to further flight. \n\n\t\t(2)\tDiscrepancies in the operation of the tailcone release system found as a result of the functional test must be repaired prior to further flight. \n\n\t(b)\tRepeat the inspection of the interior and exterior tailcone release handles and conduct the functional test required by paragraph (a) of this AD at intervals not to exceed 3,000 flight hours or 15 months, whichever occurs first. \n\n\t(c)\tWithin 30 days after discovery, report any cracked or broken tailcone release handles or any discrepancies found during the accomplishment of the inspection and functional tests required by paragraph (a) of this AD to the Manager, Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California 90806-2425. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1990 (Pub. L 96-511) and have been assigned OMB Control Number 2120-0056. \n\n\t(d)\tWithin 90 days after the effective date of this AD, replace or modify the internal and external tailcone release system cable and handle assemblies, in accordance with the Accomplishment Instructions of McDonnell Douglas Service Bulletin 53-245, Revision 1, dated June 12, 1991. Accomplishment of such replacement or modification constitutes terminating action for the repetitive inspections of the interior and exterior tailcone release handles for cracks, as required by paragraph (b) of this AD. However, the repetitive functional tests of the tailcone release system required by paragraph(b) of this AD must continue to be accomplished. \n\n\tNOTE: The following portions of the continuing repetitive functional tests and inspections of the tailcone release system are not necessary to accomplish once the replacement/modification of the cable and handle assembly is completed: Those procedures specified in paragraphs H. and L., and the second paragraph of the Notes of paragraphs F. and J., of the Accomplishment Instructions of McDonnell Douglas Alert Service Bulletin A53-243, dated January 10, 1991, Revision 1, dated February 8, 1991, and all of McDonnell Douglas Alert Service Bulletin A53-242, dated December 20, 1990. \n\n\t(e)\tAn alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles ACO. \n\n\t(f)\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. \n\n\t(g)\tThe inspection requirements of this AD shall be accomplished in accordance with McDonnell Douglas Alert Service Bulletin A53-242, dated December 20, 1990, or McDonnell Douglas Alert Service Bulletin A53-243, Revision 1, dated February 8, 1991. The functional test requirements of this AD shall be accomplished in accordance with McDonnell Douglas Alert Service Bulletin A53-243, dated January 10, 1991, or McDonnell Douglas Alert Service Bulletin A53-243, Revision 1, dated February 8, 1991. The replacement and modification requirements shall be accomplished in accordance with McDonnell Douglas Service Bulletin 53-245, Revision 1, dated June 12, 1991. This 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 McDonnell Douglas Corporation, Post Office Box 1771, Long Beach, California 90801, ATTN: Business unit Manager, Technical Publications, Technical Administration Support, C1-L5B(45-60). Copies may be inspected at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. \n\n\t(h)\tAirworthiness Directive 91-26-09, supersedes AD 91-02-13, Amendment 39-6867. \n\n\t(i)\tThis amendment (39-8122, AD 91-26-09) becomes effective on February 17, 1992.
89-25-01: 89-25-01 SIKORSKY AIRCRAFT: Amendment 39-6401. Docket No. 89-ASW-33. Applicability: Model S-58 series helicopters equipped with reciprocating engines, certificated in any category except restricted. Compliance: Required within the next 10 hours' time in service after the effective date of this AD, unless already accomplished. To prevent or reduce the hazards of an in-flight fire which could result in loss of the helicopter, accomplish the following: (a) Inspect the helicopter for the installation of an approved engine compartment fire extinguishing system. (b) If an approved engine fire extinguishing system is installed, no further action is necessary. (c) If an approved engine fire extinguishing system is not installed, accomplish the following: (1) Prior to further flight, install a placard, decal, or markings, in full view of and legible to the pilot in daylight which states "Engine Fire Extinguishing System Not Installed." (2) Prior to further flight, revise the helicopter's FAA-approved Rotorcraft Flight Manual (RFM) Operating Limitations Section by attaching or placing in front of the RFM a durable card or paper which contains the following limitations: (i) "No passengers allowed, except required crew and passengers necessary to accomplish a work program when carried on the flights to and from the work site." (ii) "Prior to the first flight of the day, test and determine that the fire detector system is properly functioning in accordance with the normal procedure section of the RFM." NOTE: An engine fire detector system is part of the approved type design and is required by the certification basis. Installation and proper operation are necessary for a determination of airworthiness in accordance with FAR Section 91.29. (3) Prior to November 30, 1990, install an approved engine compartment fire extinguishing system in accordance with technical data approved by the Administrator. When the fire extinguishing system has been installed, the placard, etc., required by paragraph (c)(1) and the RFM limitations required by paragraph (c)(2) may be removed. (d) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (e) An alternate method of compliance or adjustment of the compliance times, which provides an equivalent level of safety, may be used if approved by the Manager, Boston Aircraft Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, 12 New England Executive Park, Burlington, Massachusetts 01803. NOTE: An operator's request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Boston Aircraft Certification Office. This amendment (39-6401, AD 89-25-01) becomes effective on December 26, 1989.
2013-07-04: We are superseding an existing airworthiness directive (AD) for certain Airbus Model A319, A320, and A321 series airplanes. That AD currently requires installing spacer assemblies at the attachment points of the YZ-latches of the cargo loading system (CLS) in the forward and aft cargo compartments, as applicable. This new AD also requires modifying the attachment points of fixed YZ-latches of the CLS lower deck cargo holds on those airplanes on which one or both lower deck cargo holds have not been modified, which terminates the existing requirements. This AD was prompted by results from tests that have shown that the attachment points of the YZ-latches of the cargo loading system (CLS) fail under maximum loads and reports that installation has been applied only on one of the lower deck cargo holds, instead of on both forward and aft cargo holds, and that some airplanes could have installed the affected YZ-latches through the instructions of the cargo conversion manual. We areissuing this AD to prevent failure of the attachment points of the YZ-latches, which could result in unrestrained cargo causing damage to the fire protection system, hydraulic system, electrical wiring, or other equipment located in the forward and aft cargo compartments. This damage could adversely affect the continued safe flight of the airplane.
75-06-03: 75-06-03 BELL: Amendment 39-2122 as amended by Amendment 39-2146 and 39-2350 is further amended by Amendment 39-2386. Applies to Bell Models 206A, 206B, 206A-1, and 206B-1 helicopters, certificated in all categories, equipped with pitch link assemblies, P/N 206- 010-330 or 206-010-342. Compliance required within 10 hours' time in service after March 12, 1975, unless already accomplished. To detect possible fatigue cracks in each main rotor pitch link assembly, upper and lower clevis, accomplish the following. a. Remove each main rotor blade pitch link assembly from the helicopter and measure the distance between the bolt holes. Remove the upper and lower clevis from each pitch link assembly in accordance with Bell Model 206A or 206B maintenance and overhaul instructions. b. Inspect the threaded shank of each clevis using fluorescent penetrant or an equivalent inspection method. c. Replace each clevis that has a cracked shank before further flight. d. Assemble the pitch link assemblies in accordance with the Model 206A or 206B maintenance and overhaul instructions and set the pitch link assembly to the appropriate length measured in paragraph (a) of this AD. e. Determine that each bearing, P/N 206-010-469-1, installed in the swashplate outer ring horns has a breakaway force that does not exceed 10 pounds when measured as specified in the Mailgram dated February 15, 1975, from Bell Helicopter Company to all 206A, 206B, and TH57A operators or as specified in an FAA approved equivalent procedure. f. Replace each swashplate outer ring horn bearing, P/N 206-010-469-1, that exceeds 10 pounds breakaway force measured in paragraph (e) of this AD, prior to further flight, in accordance with procedures specified in Section XIV of the Bell Model 206A or 206B maintenance and overhaul instructions dated November 1, 1972, or later revision, or as specified in an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration. g. Install the pitch link assemblies in accordance with the Bell Model 206A or 206B maintenance and overhaul instructions. h. This AD does not apply to the main rotor pitch link assemblies, P/N 206-010- 355. The 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). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Company, P. O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S. W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. Amendment 39-2122 became effective March 12, 1975. Amendment 39-2146 became effective March 19, 1975. Amendment 39-2350 became effective October 2, 1975. This amendment 39-2386 becomes effective October 20, 1975.
66-20-04: 66-20-04 LYCOMING: Amdt. 39-277, Part 39, Federal Register August 18, 1966. Applies to Model O-320, IO-320, O-340, O-360, IO-360, O-540, and IO-540 Series Engines Equipped With AC Oil Filters Except O-320- A, -E Series, Engine Serial Number 16128-27 and Higher; O-320-B, -C, and -D Series, Engine Serial Number 6217-39 and Higher; IO-320 Series, Engine Serial Numbers 2110-55A, 2113-55A and Higher; O-360 Series, Engine Serial Number 9346-36A and Higher; O-540 Series, Engine Serial Numbers 9770-40, 9800-40, 9803-40 and Higher; and IO-540 Series, Engine Serial Numbers 2831-48, 2835-48, 2840-48 and Higher. Compliance required within the next 50 hours' time in service after the effective date of this AD, unless already accomplished. To prevent further failures of oil filter adapter gasket, P/N 74904, accomplish the following: (a) Replace gasket, P/N 74904, with gasket, P/N 76691. (b) Inspect the stud, cap screws, and tapped holes in the accessory housing mounting padfor proper length or depth, as applicable, in accordance with Lycoming Service Bulletin No. 307 or later FAA-approved revision. Replace studs and cap screws of improper length and retap holes of insufficient depth as necessary in accordance with Bulletin No. 307 or later FAA-approved revision. This directive effective August 27, 1966.
92-26-08: 92-26-08 ALLIED-SIGNAL INC., GARRETT ENGINE DIVISION: Amendment 39-8606. Docket 92-ANE-58. Applicability: Allied-Signal Inc., Garrett Engine Division, Model TPE331-1, -2, -2UA, -3U, -3UW, -5, -5A, -6, and -6A turboprop and Model TSE331-3U turboshaft engines incorporating third stage stator assemblies, Part Number (P/N) 868379-3, repaired at National Flight Services between February 20, 1990, and April 6, 1992, and identified by serial numbers listed in National Flight Services Alert Bulletin No. NF331-A72-11921, dated November 9, 1992. These engines are installed on but not limited to Mitsubishi MU-2B series (MU-2 series); Construcciones Aeronauticas, S.A. (CASA) C-212 series; Fairchild SA226 series (Swearingen Merlin and Metro series); Prop-Jets, Inc. Model 400; Twin Commander 680 and 690 (Jetprop Commander); Rockwell Commander S-2R; Shorts Brothers and Harland, Ltd. SC7 (Skyvan); Dornier 228 series; Beech 18 and 45 series and Models JRB-6, 3N, 3NM, 3TM, and B100; Pilatus PC-6series (Fairchild Porter, Peacemaker); De Havilland Model DH 104 series 7AXC (Dove); and Ayres S-2R series airplanes; and Sikorsky S-55 series helicopters. Compliance: Required as indicated, unless accomplished previously. To prevent an uncontained failure of the third stage turbine wheel, accomplish the following: (a) Replace affected third stage stator assemblies, P/N 868379-3, with serviceable assemblies in accordance with the following schedule: Third Stage Stator Cycles in Service Since Repair by National Flight Services Replacement Schedule 900 or more cycles Within 50 cycles in service after the effective date of this AD 450 to 899 cycles Within 150 cycles in service after the effective date of this AD, but not to exceed 950 cycles Less than 450 cycles Prior to accumulating 600 cycles NOTE: The FAA has determined that cracking of third stage stator assemblies is related to operating cycles, rather than operating hours. (b) If cycles cannotbe determined, calculate cycles by multiplying third stage stator assembly hours time in service by 1.5. (c) An alternative method of compliance or adjustment of the initial compliance time that provides an acceptable level of safety may be used if approved by the Manager, Los Angeles Aircraft Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Los Angeles Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Los Angeles Aircraft Certification Office. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the aircraft to a location where the requirements of this AD can be accomplished. (e) Third stage stator assemblies, P/N 868379-3, repaired at National Flight Services between February 20, 1990, and April 6, 1992, are identified by serial numbers listed in the following alert bulletin: Document No. Pages Date NF331-A72-11921 Total pages: 10. 1-10 November 9, 1992 This 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 National Flight Services, Inc., 10971 E. Airport Service Road, Swanton, Ohio 43558; telephone (419) 865-2311. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (f) This amendment becomes effective July 15, 1993, to all persons except those persons to whom it was made immediately effective by priority letter AD 92-26-08, issued December 16, 1992, which contained the requirements of this amendment.
2025-05-12: The FAA is superseding Airworthiness Directive (AD) 2008-10-01 and AD 2010-05-51, which applied to certain Eurocopter France (now Airbus Helicopters) Model EC120B helicopters. AD 2008-10-01 required replacing certain part-numbered and serial-numbered spherical thrust bearings. AD 2010-05-51 required repetitively inspecting the main rotor (M/R) head rotor hub (rotor hub) and, depending on the results, taking corrective action. Since the FAA issued those ADs, the manufacturer revised the airworthiness limitations section (ALS) to incorporate various airworthiness limitations, tasks, and associated thresholds and intervals that were previously contained in service bulletins, as well as incorporate a new task. This AD requires revising the ALS of the existing maintenance manual (MM) or instructions for continued airworthiness (ICAs) and the existing approved maintenance or inspection program, as applicable, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
88-03-51 R1: 88-03-51 R1 BOEING OF CANADA, LTD., DE HAVILLAND DIVISION: Amendment 39- 5868 as amended by Amendment 39-6505. Docket No. 89-NM-132-AD. Applicability: DeHavilland Model DHC-8-100 series airplanes, Serial Numbers 3 through 119, inclusive, with Modification No. 8/0467 incorporated, equipped with Eldec Proximity Switch Electronic Control Unit (PSEU), P/N 8-410-03, 8-410-04, or 8-410-05, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To preclude the possibility of the nose gear cockpit indication system indicating erroneous nose gear position, accomplish the following: A. Within 24 hours after March 25, 1988 (the effective date of AD 88-03-51) add the following to the Limitations Section of the Airplane Flight Manual (AFM) and notify all crew members. This may be accomplished by inserting a copy of this AD in the AFM: "1. Perform the following check prior to each flight. This check is to be performed even when the airplane is being operated with the anti-skid inoperative under the minimum equipment list: a. Anti-skid switch - "OFF" b. Anti-skid switch - "ON" c. Check that inboard and outboard anti-skid caution lights illuminate, and then extinguish within 6 seconds. Should the lights fail to function as noted above, dispatch is prohibited until maintenance action clears the fault." "2. While performing the 'After Take Off' and 'Approach' procedures: a. Monitor the landing gear indication system during landing gear retraction and extension. b. If there is any irregularity in gear indication or operation at any time throughout the flight, the gear must be confirmed DOWN AND LOCKED using the alternate down-lock verification system, irrespective of gear DOWN AND LOCKED (green) indication on the normal landing gear indicating system." B. Any in-flight landing gear irregularity must be corrected prior to further flight. C. Within 60 days after the effective dateof this amendment, modify the landing gear control system, in accordance with deHavilland Service Bulletin 8-32-70, Revision B, dated December 2, 1988. This modification constitutes terminating action for the requirements of paragraphs A. and B. of this AD, and the revised operating procedures may be removed from the AFM. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received theappropriate service documents from the manufacturer may obtain copies upon request to Boeing of Canada, Ltd., deHavilland Division, Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This AD revises AD 88-03-51, Amendment 39-5868. This amendment (39-6505, AD 88-03-51 R1) becomes effective on March 19, 1990.
69-11-07: 69-11-07 SLINGSBY: Amdt. 39-771. Applies to all Slingsby Model T.53.B Gliders. Before further flight after the effective date of this AD, modify the fuselage center section structure in accordance with Slingsby Technical Instruction No. 39, dated March 1969, or later ARB-approved issue or FAA-approved equivalent. This amendment becomes effective June 2, 1969.
70-21-02: 70-21-02 HAWKER SIDDELEY AVIATION, LTD: Amdt. 39-1087. Applies to de Havilland Model DH.104 "Dove" airplanes. To prevent failure of the flap datum hinge assemblies, unless already accomplished, accomplish the following within the next 3,000 hours' time in service after the effective date of this AD, or by March 31, 1971, whichever occurs first: (a) Inspect the wall thickness of the bearing housing recess of both the right wing and left wing flap datum hinge links in accordance with Hawker Siddeley Aviation, Limited, Technical News Sheet CT(104) No. 216 Issue 1, June 8, 1970, or later ARB-approved issue or an FAA-approved equivalent. If the wall thickness is found to be less than 0.17 inches, replace the flap datum hinge link with a serviceable link of Modification 982 standard. (b) Incorporate Modification 982 by replacing the flap datum hinge assemblies P/N 4WF.16A(R.H.) and P/N 4WF.15A(L.H.) with assemblies P/N 14WF.456A(R.H.) and P/N 14WF.455A (L.H.) in accordance with de Havilland Aircraft Company, Limited, Modification No. Dove 982 dated August 20, 1956, or later ARB-approved issue or an FAA-approved equivalent. This amendment becomes effective November 5, 1970.
77-17-06: 77-17-06 AIRESEARCH MANUFACTURING COMPANY of ARIZONA: Amendment 39-3016. Applies to AiResearch Model TSCP700-4B and -5 Auxiliary Power Units (APUs) which have first stage compressor discs P/Ns 969600-1 or -2 installed. Compliance required before accumulating a total of more than 3000 cycles on the first stage compressor discs, or within the next 300 cycles after the effective date of this AD, whichever occurs later, unless already accomplished within the last 1700 cycles, and thereafter at intervals not to exceed 2000 cycles since the last inspection. To prevent a high energy release of first stage compressor blades and disc parts due to the possible fatigue failure of the disc, accomplish the following: (a) Remove the first stage disc from the compressor section of the APU and inspect the blade dove tail slots of the disc in accordance with either of the following methods or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. (1) Use an eddy current probe made in accordance with instructions contained in paragraphs 2.B.(4) of AiResearch Service Bulletin TSCP700-49-A3912, dated May 16, 1977, and a visual display instrument Nortec Model No. NDT-6D. In addition to this test apparatus, A P/N 969600-1 or -2 compressor disc, or segment thereof, known to be free of cracks and a corresponding disc, or segment thereof, which has been verified to contain a detectable crack must be available to calibrate the inspection test apparatus. These sample disc(s) may be obtained from the AiResearch Manufacturing Company or, subject to the acceptance of the assigned FAA maintenance inspector, furnished by the operator. An acceptable procedure for calibrating and using this test equipment is provided in paragraph 2.A., 2.B.(1), 2.B(2), 2.B(3), 2.B.(5) and 2.B.(6) of AiResearch Service Bulletin TSCP700-A3912, dated May 16, 1977, supplemented by AiResearch Gage Operating Instructions MSC-4769, MSC-4770 or MSC-4771, as appropriate. (2) Use tooling described in paragraphs 1.G and procedures prescribed in paragraph 2. of AiResearch Service Bulletin TSCP700-49-A3912, Revision 2, dated June 27, 1977, or later FAA approved revisions. (b) All discs found to not meet the inspection criteria covered by the procedure described in paragraphs (a)(1) or (a)(2) must be removed from service. Except as provided in paragraph (c)(4), replacement discs of P/Ns 969600-1 or -2 must be inspected in accordance with paragraph (a) before being installed in the APU for service, unless already inspected within the last 3000 cycles in service, and thereafter at intervals not to exceed 2000 cycles since the last inspection. (c) For purposes of this AD: (1) A cycle is defined as a start and acceleration to at least 95% high pressure spool (N2) rpm followed by a shutdown, during which low pressure spool (N1) rpm reaches, or exceeds, 97% rpm nominal. If, in any start, operating and shutdown sequence, the low pressure spool (N1) is prevented from exceeding 91% rpm nominal, only one half of a cycle must be recorded. (2) Operators who have not kept a record of operating starts on individual discs may assume two starts have occurred for each recorded APU operating hour of service, or any other cycle per hour ratio approved by the operators' assigned FAA maintenance inspector, provided the request contains substantiating data to justify the alternative ratio. (3) Operators who have not kept a record of APU operating hours of service shall estimate hours of APU operation by equating APU operation to airplane hours time in service using a ratio approved by the operators' assigned FAA maintenance inspector and justified by substantiating data. (4) Unused replacement discs installed per (b) above may be assumed to have zero cycles and need not be inspected prior to installation. NOTE: AiResearch Model TSCP 700-4B and -5 APU are known to be installed in McDonnell Douglas Model DC-10 series aircraft and Aerospatiale Model A-300B aircraft. This amendment becomes effective August 24, 1977.
76-18-11: 76-18-11 BOEING: Amendment 39-2720 as amended by Amendment 39-2736 and 39-2771 is further amended by Amendment 39-2808. Applies to all Model 727-100 and 727-100C series airplanes, certificated in all categories, with P/N 65-19670-2 and -3 aft cargo door lowest side stop fittings which have accumulated 15,000 or more pressurization cycles. Compliance required as indicated. \n\tTo detect cracks in the aft cargo door lower stop fittings, accomplish the following: \n\tA.\tWithin the next 500 flights from the effective date of this amendment, unless accomplished within the last 500 flights, visually inspect the aft cargo door for cracks in the four (4) lowest side stop fittings (two forward and two aft) and the attaching door frame structure in accordance with Boeing Alert Service Bulletin No. 727-52-A102, Revision 2, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. Repeat visual inspections of the lowestforward and aft side fittings (total of two) and adjacent door frame structure at intervals not to exceed 1000 flights from the last inspection. \n\tB.\tIf a cracked fitting(s) and/or frame(s) are detected: \n\t\t1.\treplace the fitting and repair or replace the door frame section, as necessary, prior to further flight in accordance with Boeing Alert Service Bulletin No. 727-52-A102, Revision 1, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region; or \n\t\t2.\tcontinue service with no more than one cracked fitting and/or frame for up to 50 flights under the following conditions: \n\t\t\ta.\tframe cracks may not extend along the frame more than 4.5 inches, including the length of the stop fitting cutout, and no cracks may be present in the frame flange radius. Stop drill all frame cracks, which do not terminate in an existing fastener hole. \n\t\t\tb.\tassure that the stop fittings adjacent to the cracked fitting, includingthe lower sill stop fitting, are crack-free by the visual inspection specified in the service bulletin. \n\t\t\tc.\twithin 25 flights repeat the visual inspection of the adjacent fittings and frame per a and b above. \n\t\t\td.\twithin 50 flights replace the fitting with a crack-free 7079-T6 aluminum or steel fitting and repair or replace the door frame, as necessary, in accordance with the service bulletin. \n\tC.\tReplacement of a lowest side stop fitting with a new steel fitting in accordance with Boeing Alert Service Bulletin No. 727-52-A102, Revision 2, or later FAA approved revisions, or equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region, constitutes terminating action for this AD at that fitting, provided the adjacent fitting and attaching door frame were/are inspected and found to be crack-free in accordance with Revision 2 to the service bulletin or equivalent. \n\tD.\tFor the purpose of this AD, when conclusive records are not available to show the number of flights accumulated by a particular fitting, the number of flights may be computed by dividing the airplane time-in-service since the fitting was installed in the airplane by the operator's fleet average time per flight for his Model 727 airplanes. \n\tE.\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Northwest Region, may adjust the repetitive inspection intervals in this AD, if the request contains substantiating data to justify the increase for that operator. \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 these documents from the manufacturer, may obtain copies upon request to Boeing Commercial Airplane Company, P. O. Box 3707, Seattle, Washington 98124. The documents may also beexamined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tAmendment 39-2720 became effective September 21, 1976. \n\tAmendment 39-2736 became effective October 15, 1976. \n\tAmendment 39-2771 became effective November 29, 1976. \n\tThis amendment 39-2808 becomes effective February 18, 1977.
2013-07-11: We are superseding an existing airworthiness directive (AD) for certain The Boeing Company Model 777-200, -200LR, -300, and -300ER series airplanes. That AD currently requires inspecting for scribe lines in the skin along lap joints, butt joints, certain external doublers, and the large cargo door hinges, and doing related investigative and corrective actions if necessary. This new AD adds an inspection for scribe lines where external decals have been applied or removed across lap joints, large cargo door hinges, and external doublers, and related investigative and corrective actions if necessary. This AD was prompted by a determination that scribe lines could occur where external decals are installed or removed across lap joints, large cargo door hinges, or external doublers. We are issuing this AD to detect and correct scribe lines, which can develop into fatigue cracks in the skin. Undetected fatigue cracks can grow and cause sudden decompression of the airplane.
75-17-24: 75-17-24 SIAI-MARCHETTI: Amendment 39-2329. Applies to Model S.205-18/R, - 20/R, and -22/R airplanes, serial numbers 001 thru 003, 101 thru 399, 4-101 thru 4-282, 4-285, 4- 292, 5-302, 5-303 and 5-406, certificated in all categories, and Model S.208 airplanes, serial numbers 001 thru 003, 1-03 thru 1-15, 2-16 thru 2-27, 2-47 thru 2-50, 4-51, 4-60, 4-62, 369, 3- 100, 4-231, 4-233, 4-256, 4-257, 4-258, certificated in all categories. Compliance is required as indicated, unless already accomplished. To detect cracks in the landing gear actuator brackets and prevent bracket failure, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, unless already accomplished within the last 90 hours' time in service, and thereafter at intervals not to exceed 100 hours' time in service from the last inspection, until modified in accordance with paragraph (d) of this AD, visually inspect the landing gear actuator bracket P/N 205-6-209-03 installed on the rear side of the No. 5 frame for cracks, in accordance with Instruction a) of the SIAI Marchetti Service Bulletin No. 205B39, dated September 19, 1973, or an FAA-approved equivalent. (b) If cracks are found, before further flight, comply with paragraph (d) of this AD. (c) If no cracks are found, within 300 hours' time in service after the effective date of this AD, comply with paragraph (d) of this AD. (d) Install landing gear actuator bracket reinforcements, P/Ns U.T. 5270-11 and -13, to inside of the landing gear actuator bracket P/N 205-6-209-03 in accordance with Instruction (b) of the SIAI Marchetti Service Bulletin No. 205B39, dated September 19, 1973, or an FAA- approved equivalent. This amendment becomes effective August 20, 1975.
2013-07-10: We are adopting a new airworthiness directive (AD) for certain International Aero Engines AG (IAE), V2525-D5 and V2528-D5 turbofan engines, with a certain No. 4 bearing internal scavenge tube and a certain No. 4 bearing external scavenge tube installed. This AD was prompted by a report of an engine under-cowl fire and commanded in- flight shutdown. This AD would require replacement of certain part number (P/N) No. 4 bearing internal scavenge tubes, and alignment checks of certain P/N No. 4 bearing external scavenge tubes. We are issuing this AD to prevent engine fire and damage to the airplane.
2013-07-08: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 757 airplanes. This AD was prompted by reports that inspections of the wing center section revealed defective, misapplied, or missing secondary fuel vapor barriers on the center fuel tank. This AD requires inspecting for discrepancies and insufficient coverage of the secondary fuel barrier, determining the thickness of the secondary fuel barrier, and corrective actions if necessary. We are issuing this AD to detect and correct defective surfaces and insufficient thickness of the secondary fuel barrier, which could allow fuel leaks or fumes into the pressurized cabin, and allow fuel or fuel vapors to come in contact with an ignition source, which could result in a fire or an explosion.
75-16-11: 75-16-11 PILATUS AIRCRAFT LTD: Amdt. 39-2284. Applies to B4-PC11 gliders, serial numbers 141 and subsequent, certificated in all categories. Compliance is required as indicated. To prevent rudder cable chaffing, jamming, or separation accomplish the following: (a) Within 10 hours' time in service after the effective date of this AD, unless already accomplished, visually inspect the rudder control cables for chaffing in that portion adjacent to the crotch strap attachment brackets and repair, or replace as required in accordance with the Pilatus B4- PC11 Maintenance Manual or an FAA-approved equivalent. (b) Within 25 hours' time in service after the effective date of this AD, unless already accomplished, install a polyurethane foam block P/N 112.50.11.065 on the underside of the pilot seat to provide a grooved channel to cover the crotch strap attachment bracket in accordance with the accomplishment instructions set forth in Pilatus Aircraft Service Bulletin No. 1002 datedMay 1974, or an FAA-approved equivalent. This amendment becomes effective July 29, 1975.
2013-04-04: We are superseding an existing airworthiness directive (AD) for certain The Boeing Company Model 757 airplanes equipped with Rolls- Royce RB211-535E engines. That AD currently requires repetitive inspections for signs of damage of the aft hinge fittings and attachment bolts of the thrust reversers, and related investigative and corrective actions if necessary. The existing AD also provides for an optional terminating modification for the repetitive inspections. For certain airplanes, this new AD adds a one-time inspection of the washers installed under the attachment bolts of the aft hinge fittings for correct installation sequence, and reinstallation if necessary. This new AD also adds an option for installing a redesigned aft hinge fitting with the trim already done, instead of trimming an existing or new hinge fitting, which is included in the existing optional terminating modification. This AD was prompted by reports of incorrectly installed washers under the attachment bolts of the aft hinge fittings of the thrust reversers. We are issuing this AD to prevent failure of the attachment bolts and consequent separation of a thrust reverser from the airplane during flight, which could result in structural damage to the airplane.
2013-07-02: We are adopting a new airworthiness directive (AD) for all Airbus Model A318, A319, and A320 series airplanes. This AD was prompted by fuel system reviews conducted by the manufacturer, which revealed that certain fuel pumps under certain conditions can create an ignition source in the fuel tank. This AD requires modification of the center tank fuel pump control circuit by installation of ground fault interrupters (GFIs). This AD would also require either replacement of the GFI or deactivation of the associated fuel pump following failure of any post-modification operational test of the GFI. We are issuing this AD to prevent the potential of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane.
2025-05-14: The FAA is adopting a new airworthiness directive (AD) for all Airbus SAS Model A350-941 and A350-1041 airplanes. This AD was prompted by a report indicating that the thrust reverser and pylon thermal blankets were found damaged due to air leaking from the pre-cooler exchanger (PCE). This AD requires repetitively testing the PCE for air leaks and reporting the results, and, depending on findings, inspecting the thermal blankets for damage and replacing the PCE, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.