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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.