Results
76-19-01: 76-19-01 BRITISH AIRCRAFT CORPORATION: Amendment 39-2723. Applies to Viscount Model 744, 745D, and 810 Series airplanes, certificated in all categories. Compliance is required as indicated. To prevent possible failure of the elevator spring servo tab control mechanism, accomplish the following: (a) Replace the spigot bracket, P/N 70120-367, that attaches the elevator spring tab system bellcrank to the left elevator torque tube with a new bracket of the same part number as follows: (1) If neither paragraph (b) or (c) of AD 71-4-2 has been complied with prior to the effective date of this AD, replace the bracket within the next 100 hours time in service after the effective date of this AD or prior to the accumulation of 12,000 hours total time in service on the bracket, whichever occurs later. (2) If paragraph (b) or (c) of AD 71-4-2 has been complied with prior to the effective date of this AD, replace the bracket at the latest of the following:(i) Within the next 50 hours time in service after the effective date of this AD. (ii) Within 1000 hours time in service after complying with AD 71-4-2 if the bracket was not replaced in complying with that AD. (iii) Prior to the accumulation of 12,000 hours total time in service on the bracket. (3) After complying with paragraph (a)(1) or (a)(2) of this AD, as appropriate, continue to replace the brackets prior to the accumulation of 12,000 hours total time in service after installation. (b) Operators who have not kept records of total hours time in service on individual spigot brackets, P/N 70120-367, must substitute in lieu thereof the total hours time in service of the airplane. This supersedes Amendment 39-1154 (36 FR 2562), AD 71-4-2. This amendment becomes effective October 15, 1976.
2005-17-02: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 777-200 and -300 series airplanes. This AD requires inspecting the valve control and indication wire bundles of the fuel system of the wing rear spar for discrepancies, and corrective action if necessary. This AD is prompted by reports of six incidents of the wire bundles chafing against the rear spar stiffeners outside the fuel tank. We are issuing this AD to prevent this chafing, which could result in wire damage leading to a short circuit, subsequent ignition of flammable vapors, and possible uncontrollable fire during fueling or flight.
97-19-12: This amendment adopts a new airworthiness directive (AD), applicable to Pratt & Whitney JT8D-1, -1A, -1B, -7, -7A, -7B, -9, -9A,-11, -15, -17, and -17R series turbofan engines, that requires initial and repetitive fluorescent penetrant and eddy current inspections of 4th stage low pressure turbine (LPT) hubs for cracks, and, if necessary, replacement with serviceable parts. This amendment is prompted by a report of an uncontained 4th stage LPT blade release. The actions specified by this AD are intended to prevent a 4th stage LPT blade release due to hub cracking, which can result in an uncontained engine failure and damage to the aircraft.
97-19-06: This amendment supersedes an existing priority letter airworthiness directive (AD), applicable to Sikorsky Aircraft Corporation Model S-61A, D, E, L, N, NM, R, and V helicopters, that currently requires inspecting certain main rotor blade assemblies (blades) to determine if a blade has a blade trailing edge pocket assembly (pocket assembly) that was anodized by Poly-Metal Company during a specified time period, and if so, replacing it with an airworthy blade. This amendment requires the same actions as the existing AD, but corrects two serial numbers in the list of the applicable blades. This amendment is prompted by the manufacturer's issuance of a service bulletin with a revised list of blade serial numbers. The actions specified by this AD are intended to prevent disbonding and separation of portions of the blade, subsequent excessive vibrations, and loss of control of the helicopter.
78-08-05: 78-08-05 MCDONNELL DOUGLAS: Amendment 39-3184. Applies to DC-10-10, -10F, -30, -30F and -40 Series airplanes certificated in all categories, that are equipped with Litton -51 or Litton -72 Inertial Navigation Systems or Litton -58 Inertial Sensor Systems. Airplanes which are equipped with these Litton inertial systems after the effective date of this AD must be modified as defined herein prior to installation of the inertial systems. \n\n\tCompliance is required as indicated unless already accomplished. \n\n\tTo prevent simultaneous loss of primary attitude and heading information during takeoff roll and subsequent flight accomplish the following: \n\n\t(a)\tOn or before July 31, 1979 unless already accomplished, or unless incorporated in production, modify the airplane's navigation rack in accordance with McDonnell Douglas DC-10 Service Bulletin 25-251 dated August 26, 1977, or later FAA approved revisions. \n\n\t(b)\tEquivalent modifications, procedures, or revisions may be used whenapproved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(c)\tSpecial flight permits may be issued in accordance with FAR's 21.197 and 21.199 to operate airplanes to a base for accomplishment of the modification required by this AD. \n\n\tThis amendment becomes effective May 16, 1978.
2005-16-14: The FAA is adopting a new airworthiness directive (AD) for all Gulfstream Model G-IV and GV series airplanes, and certain GIV-X and GV-SP series airplanes. This AD requires a one-time inspection to determine if a certain floor heater pad system is installed, and deactivation of the subject floor heater pad system if it is installed. This AD results from an incident of short-circuiting of the floor heater pads, in which no circuit breakers tripped in response to the short-circuiting. We are issuing this AD to prevent short-circuiting of the floor heater pad system, which could result in a fire in the airplane cabin.
97-18-04: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 757 series airplanes, that currently requires repetitive inspections to detect fatigue cracking in the midspar fuse pins; replacement with new or refinished fuse pins, if necessary; and repetitive inspections of newly installed fuse pins. This AD requires earlier initial inspections and replacements; more frequent repetitive inspections of certain fuse pins; and replacement with new fuse pins, if necessary. This amendment is prompted by reports of fatigue cracking of the midspar fuse pins and by fatigue test data indicating that current inspection thresholds and intervals for repetitive inspections are inadequate to detect fatigue cracking in a timely manner. The actions specified in this AD are intended to detect and correct such fatigue cracking, which could lead to separation of the strut and engine from the wing of the airplane.
97-18-05: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 757 series airplanes, that currently requires repetitive inspections to detect cracking in the midspar fuse pins; replacement with new or refinished fuse pins, if necessary; and repetitive inspections of newly installed fuse pins. This AD requires earlier initial inspections and replacements; more frequent repetitive inspections of certain fuse pins; and replacement with new fuse pins, if necessary. This amendment is prompted by reports of fatigue cracking of the midspar fuse pins and by fatigue test data indicating that current inspection thresholds and intervals for repetitive inspections are inadequate to detect fatigue cracking in a timely manner. The actions specified in this AD are intended to detect and correct such fatigue cracking, which could lead to separation of the strut and engine from the wing of the airplane.
92-03-09: 92-03-09 SAAB-SCANIA: Amendment 39-8164. Docket 91-NM-159-AD. Applicability: Model SF-340A and 340B series airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent a fire during the refueling process, accomplish the following: (a) Within 30 days after the effective date of this AD, disconnect the lighting to the refuel/defuel panel lights as follows: (1) Remove the refuel/defuel panel assembly, P/N 7239160-502, -503, or -504, as applicable, in accordance with the Airplane Maintenance Manual (AMM) 28-21-05. (2) With the refuel/defuel panel removed, loosen the four screws securing the lighting panel to the front, and remove the rear cover. (3) Locate the lighting panel Jack 30 QA under the rear cover, and remove the screw securing the wire QA636-20. Remove the wire from Jack 30 QA, and reinstall the screw. (4) Cap and stow wire QA636-20, and reassemble the refuel/defuel panel.(5) Placard the lighting panel with "Lights inop." (6) Reinstall and test the refuel/defuel panel in accordance with AMM 28-21-05. (b) An 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, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. (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) This amendment (39-8164, AD 92-03-09) becomes effective on March 18, 1992.
2001-26-01: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model DC-8 series airplanes that have been converted from a passenger-to a cargo-carrying ("freighter") configuration. This amendment requires, among other actions, modification of the main deck cargo door structure and fuselage structure; modification of a main deck cargo door hinge; modification of the main deck cargo floor; and installation of a main deck cargo 9g crash barrier; as applicable. The actions specified by this AD are intended to prevent opening of the cargo door while the airplane is in flight or collapse of the main deck cargo floor, and consequent rapid decompression of the airplane including possible loss of flight control or severe structural damage. These actions are intended to address the identified unsafe condition.
77-16-10: 77-16-10 EDO-AIRE MITCHELL: Amendment 39-3002. Applies to Model NSD-360 Navigation Systems containing any of the following Navigation Situation Display instruments: 52D136-0020, 52D136-0120, 52D136-0220, 52D136-1020, 52D136-1120, 52D136-1220, 52D137-1020, 52D137-1120, and 52D137-1220, when used with VOR/LOC converters other than the Edo-Aire Mitchell 1C707 and 1C707-1 converters. Compliance with (a) or (c) is required within the next 50 hours in service after the effective date of this AD, unless already accomplished. (a) Install a placard stating "NSD-360 NOT APPROVED AS PRIMARY VOR DISPLAY FOR IFR NAVIGATION PRIOR TO COMPLIANCE WITH EAM MB-13." (b) The placard described in subparagraph (a) is to be installed as close as practical to the affected navigation situation display and remain there until EAM NAV Flag Adapter P/N 1C775 is installed as in (c) below, at which time the placard is to be removed. EAM placard, P/N 13A895 specified by EAM Service Bulletin MB-13, is to be used to comply with (a). (c) Install EAM NAV Flag Adapter, P/N 1C775, in accordance with installation instruction items 1 through 5, pages 2 and 3 of EAM Service Bulletin MB-13 revised May 3, 1977. (d) The manufacturer's instructions identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(l). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Director, Products and Service, Edo-Aire Mitchell, P.O. Box 610, Municipal Airport, Mineral Wells, Texas 76067, Telephone No. 817-325-2517. 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. This amendment becomes effective September 8, 1977.
77-21-01: 77-21-01 BELL HELICOPTER TEXTRON: Amendment 39-3054. Applies to all Model 47 helicopters with tail rotor gear boxes, part number 47-640-075-1 or 47-640-075-5, installed. NOTE: Retrofit kits incorporating these gear boxes or the associated internal gears and bearings delivered from Bell Helicopter Textron after December 1, 1976, have been verified to be in compliance with this airworthiness directive and will not require inspection and/or further retrofit. Compliance is required within the next 100 hours time in service after the effective date of this Airworthiness Directive, unless already accomplished. To minimize the possibility of loss of directional control of the helicopter resulting from failure of the bearing, part number 47-620-629-3 or 47-620-629-5, or failure of the input pinion shaft, part number 47-645-205-3, located in the tail rotor gear box, part number 47-640-075-1 or 47-640-075-7, conduct the inspection and replacement activities prescribed by paragraphs 1, 2, 2a, 2b, 3, and 4 of Bell Helicopter Textron Service Bulletin No. 47-77-1, dated February 14, 1977, or later FAA approved revision. Comments in paragraph 5 and 6 of this bulletin involving warranties, replacement part sources, or reporting activities, are not a part of this Airworthiness Directive. After accomplishment of paragraphs 1, 2, 3, and 4 of Service Bulletin No. 47-77-1, reassemble and reinstall the tail rotor gear box in accordance with the Maintenance and Overhaul Instruction for the applicable model helicopters. 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 by requesting same from Service Manager, Bell Helicopter Textron, P.O. Box 482, Fort Worth, Texas, 76101. These documents may also be examined at Officeof the Regional Counsel, Southwest Region, Federal Aviation Administration, 4400 Blue Mound Road, Fort Worth, Texas, 76106, and the Federal Aviation Administration Headquarters, 800 Independence Avenue, S.W., Washington, D.C., 20591. Equivalent means of compliance with the modifications prescribed by this Airworthiness Directive may be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration, P.O. Box 1689, Fort Worth, Texas, 76101. This amendment becomes effective November 14, 1977.
2016-21-01: We are adopting a new airworthiness directive (AD) for certain Bell Helicopter Textron (Bell) Model 430 helicopters. This AD requires establishing a life limit for a certain main rotor hub attachment bolt (bolt) and removing from service each bolt that has met or exceeded its life limit. This AD was prompted by a documentation error that omitted the life limit of a certain part-numbered bolt from the Airworthiness Limitations section of the maintenance manual. The actions of this AD are intended to establish a life limit for a certain part-numbered bolt to prevent failure of a bolt, failure of a main rotor hub, and subsequent loss of control of a helicopter.
89-24-05: 89-24-05 BOEING: Amendment 39-6395. Docket No. 89-NM-213-AD. \n\n\tApplicability: Model 767 series airplanes, listed in Boeing Alert Service Bulletin 767- 31A0029, dated March 23, 1989, certificated in any category. \n\n\tCompliance: Required within 45 days after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent binding of the pilot's control wheel, accomplish the following: \n\n\tA.\tDeactivate the control wheel Aircraft Integrated Data System (AIDS) by removing the position transducer, crank assembly, and rod assembly in accordance with Boeing Alert Service Bulletin 767-31A0029, dated March 23, 1989. \n\n\tB.\tAn 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concuror comment and then send it to the Manager, Seattle Aircraft Certification Office. \n\n\tC.\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\tAll persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment (39-6395, AD 89-24-05) becomes effective on December 1, 1989.
2005-16-13: The FAA is adopting a new airworthiness directive (AD) for certain Gulfstream Aerospace LP Model Galaxy and Gulfstream 200 airplanes. This AD requires a one-time general visual inspection for any damaged wiring, splice, connector, and pins for the fuel standby feed pumps and replacement of any damaged wiring, splice, connector, or pin. This AD also requires replacement of the power and ground wires for the fuel standby feed pumps. This AD results from reports of evidence of overheating found on the feeder wires of the left and right fuel standby feed pumps. We are issuing this AD to detect and correct damaged wiring for the fuel standby feed pumps, which could result in an ignition source in an area where fuel vapor may be present, and a consequent fire or explosion.
2016-19-12: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 747-400, 747-400D, and 747-400F series airplanes. This AD was prompted by a determination that a certain fastener type in the fuel tank walls has insufficient bond to the structure, and an electrical wiring short could cause arcing to occur at the ends of fasteners in the fuel tanks. This AD requires the installation of new clamps and polytetrafluoroethylene (TFE) sleeves on the wire bundles of the front spars and rear spars of the wings. This AD also requires inspecting the existing TFE sleeves under the wire bundle clamps for correct installation, and replacement if necessary. We are issuing this AD to prevent potential ignition sources in the fuel tank in the event of a lightning strike or high-powered short circuit, and consequent fire or explosion.
76-25-05: 76-25-05 BEECH: Amendment 39-2789. Applies to Models A23-19, 19A, M19A and B19 (Serial Numbers MB-1 thru MB-545 and MB-547), Models 23, A23, A23A, B23 and C23 (Serial Numbers M-2 thru M-1415), Models A-23-24 and A24 (Serial Numbers MA-1 thru MA- 368), and Models A24R (Serial Numbers MC-2 thru MC-127) airplanes. Compliance: Required as indicated, unless already accomplished. To prevent interchanging left and right ailerons during reinstallation and resulting reversed aileron control response, accomplish the following: A) Within the next 100 hours' time in service after the effective date of this AD or at the next removal of the ailerons or within one year after the effective date of this AD, whichever occurs first, install one each Beech P/N 169-110000-481 strap in the left and right wing trailing edge in accordance with Beechcraft Service Instruction 0510-032 or later approved revisions. B) Any alternate method of compliance with this AD must be approved by theChief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective January 24, 1977.
2005-16-11: The FAA is adopting a new airworthiness directive (AD) for certain Boeing transport category airplanes listed above. This AD requires repetitive inspections for cracks of the upper surface of the aft lower spar web of the inboard and outboard struts, as applicable; and repetitive inspections for cracks of the upper surface of the intermediate web bay of the aft lower spar. This AD also requires repetitive inspections and torque checks of the bolts common to the aft lower spar chords and the fitting of the rear engine mount bulkhead for missing, loose, or fractured bolts, as applicable; and corrective action, if necessary. This AD is prompted by reports of cracking in the aft lower spar web and reports of missing and fractured bolts. We are issuing this AD to detect and correct cracking of the aft lower spar web, and to prevent missing, loose, or fractured bolts common to the aft lower spar chords and the fitting of the rear engine mount bulkhead, which could result in the lossof the aft lower spar load path and reduced structural capability of the pylon, which may result in the separation of the engine from the airplane.
82-14-02: 82-14-02 PRATT & WHITNEY AIRCRAFT: Amendment 39-4411. Applies to all Pratt & Whitney Aircraft JT3D turbofan engines. Compliance required as indicated, unless already accomplished. To detect cracks in third stage turbine disks, P/Ns 418903, 438903, and 675803, which could result in fracture of the third stage turbine disk and an uncontained third stage turbine disk failure, inspect third stage turbine disks, P/Ns 438903 and 675803, for cracks in the tierod holes, counterweight holes, and spacer snap diameter fillet radii. Inspect in accordance with Pratt & Whitney Aircraft Alert Service Bulletin No. 5358, dated February 5, 1982, or later FAA approved revision, or equivalent means approved by the Chief, Engine Certification Branch, Federal Aviation Administration, New England Region, per the following schedule: a. Disks with less than 4,500 cycles on the effective date of this AD shall be inspected before 5,500 cycles. b. Disks with between 4,500 and 4,999 cycleson the effective date of this AD shall be inspected within 1,000 cycles. c. Disks with between 5,000 and 5,500 cycles on the effective date of this AD shall be inspected before 6,000 cycles. d. Disks with more than 5,500 cycles on the effective date of this AD shall be inspected within 500 cycles. Cracked third stage turbine disks cannot be continued in service. Any third stage turbine disk which has a crack which is within the allowable crack repair limit(s) specified in the above referenced Pratt & Whitney Aircraft Alert Service Bulletin may be returned to service if it is repaired in accordance with this service bulletin. Any third stage turbine disk which has had all of the tierod holes and all of the counterweight holes repaired in accordance with this service bulletin must be reinspected in accordance with the service bulletin prior to the accumulation of 4,000 additional cycles after inspection. All other third stage turbine disks which are suitable for continued usage per this service bulletin must be reinspected in accordance with this service bulletin prior to the accumulation of 2,500 additional cycles after inspection. These repetitive inspections must be repeated throughout the service life of the third stage turbine disk. Third stage turbine disks, P/N 418903, shall be removed from service within 500 cycles from the effective date of this AD or by December 31, 1982, whichever comes first. The inspection and rework provisions of Pratt & Whitney Aircraft Alert Service Bulletin No. 5358 do not apply to P/N 418903 third stage turbine disks. NOTE: The established life limits for the third stage turbine disks shall not be exceeded. Upon request of the operator, an FAA Maintenance Inspector, subject to prior approval of Chief, Engine Certification Branch, FAA, New England Region, may adjust the inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. All persons affected by this directive who have not already received the referenced Alert Service Bulletin from the manufacturer may obtain copies upon request to Pratt & Whitney Aircraft, Division of United Technologies Corporation, 400 Main Street, East Hartford, Connecticut 06108. This document may also be examined at Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803. A historical file on this AD is maintained by the FAA at the New England Region Office. This amendment becomes effective on July 1, 1982.
91-23-16: 91-23-16 AEROSTAR AIRCRAFT, INC.: Amendment 39-8085. Docket No. 91-CE-34-AD. Applicability: Model PA60-700P (formerly Piper) airplanes (serial numbers 60-8423001 through 60-8423025), certificated in any category. Compliance: Required within the next 50 hours time-in-service after the effective date of this AD, unless already accomplished. NOTE 1: The compliance time in this AD takes precedence over that cited in the referenced service information. To prevent engine damage that could result from incorrect manifold pressure operations, accomplish the following: (a) Install placard, part number 87369-77, in accordance with the instructions in Piper Special Advisory No. 60-7, dated January 11, 1991, and operate the airplane accordingly. NOTE 2: This placard (part number 87369-77) is enclosed in Piper Special Advisory No. 60-7, dated January 11, 1991, which may be obtained from the manufacturer at the address in paragraph (e) of this AD. (b) Insert Report VB-1220, Revision 4, dated December 14, 1990, into the limitations section of the PA-60-700P Pilot's Operating Handbook and operate the airplane in accordance with these limitations. NOTE 3: Copies of Report VB-1220, Revision 4, dated December 14, 1990, may be obtained from the manufacturer at the address in paragraph (e) of this AD. (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) An alternative method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager, Seattle Aircraft Certification Office, 1601 Lind Avenue, S.W., Renton, Washington 98055-4056. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Seattle Aircraft Certification Office. (e) The installation required by this AD shall be done in accordance with Piper Special Advisory No. 60-7, dated January 11, 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 Aerostar Aircraft, Inc., 3608 S. Davison Boulevard, Spokane, Washington 99204. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-8085, AD 91-23-16) becomes effective on December 20, 1991.
76-18-02: 76-18-02 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE: (formerly SUD AVIATION): Amendment 39-2710. Applies to Alouette III Models SE-3160, SA-316B, SA- 316C, and SA-319B helicopters, Serial Nos. 2032 and below, certificated in all categories. Compliance is required as indicated. To prevent failure of tail rotor control cables and the consequent loss of directional control, accomplish the following: (a) Within the next 300 hours time in service after the effective date of this AD, unless already accomplished, replace tail rotor control cables, fittings, and guides with improved cables, fittings, and guides in accordance with Alouette Service Bulletin No. 65.72, dated March 29, 1971, and Alouette Service Bulletin No. 65.93, dated November 14, 1972, or their equivalents approved by the Chief, Aircraft Certification Staff, c/o American Embassy, APO New York, N.Y. 09667. (b) Within the next 100 hours time in service after the accomplishment of the modification specified in paragraph (a) of this AD, or within the next 100 hours after the effective date of this AD, whichever occurs later, and thereafter at intervals not to exceed 100 hours time in service from the last inspection, inspect the cables for the proper tension, for exposed or broken wire strands, and for tearing of the wirelon coating. NOTE: During the inspection required by paragraph (b) of this AD, particular attention should be directed to condition of cables in the areas of the fittings, pulleys, and cable guides. (c) If an exposed wire strand or a broken wire strand is found in any cable or tearing of the wirelon coating of any cable is found as a result of any inspection required by paragraph (b) or (c) of this AD, before further flight, replace that cable with a serviceable cable of same part number and continue to inspect the cables in accordance with paragraph (b) of this AD at intervals not to exceed 100 hours time in service from the last inspection. (d) If, asa result of any of the inspections required by this AD, improper tension is found in any cable, adjust that cable for proper tension and continue to inspect the cables in accordance with paragraph (b) of this AD at intervals not to exceed 100 hours time in service from the last inspection. This amendment becomes effective on September 29, 1976.
78-22-02: 78-22-02 HILLER HELICOPTER: Amendment 39-3320. Applies to Hiller Model UH- 12D, UH-12E, and UH-12E(4 Place) certificated in all categories including Military Models UH- 23D, UH-23G, H-23F and turbine powered models, equipped with main rotor blades assembly, Parson Part No. 2253-1101-03 or 2253-1101-04. Compliance required as indicated. To prevent failure of the main rotor blade Parson Part No. 2253-1101-03 and 2253-1101- 04 antinode outboard retention system, which could result in a loss of the rotorcraft, accomplish the following: (a) Before further flight after receipt of this AD, unless already accomplished, inspect the threaded end of the antinode bar outboard of the antinode weight anchor nut according to the following instructions: (1) Remove the antinode bar assembly per paragraph 4.53 of UH-12E/UH- 12L Series Structural Repair Manual. (2) Clean thread outboard of the antinode weight anchor nut with a suitable wire brush. (3) Using a micrometer, measure the thread major diameter. (i) If the rod thread major diameter is between .3678 and .3750 inches, no further action is required for this AD and the rod may be reinstalled per instructions contained in the Structural Repair Manual. NOTE: CAUTION. Use extreme care in reinstalling the antinode bar assembly to ensure that the screws attaching the anchor nut to the antinode bar are not sheared. Hand pressure is the maximum force allowed. (ii) If the rod thread major diameter measures .3677 inch or less, before returning to service: (A) Return the rotor blade(s) to Hiller Aviation or an FAA approved repair station to replace the antinode bar assembly, or (B) Replace the antinode bar assembly per Hiller Service Bulletin 51-4 dated September 26, 1978. NOTE: Hiller Service Bulletin 51-4 dated September 26, 1978 is the only version of the service bulletin suitable for compliance with this AD. (b) Special flight permits may beissued in accordance with FARs 21.197 and 21.199 to operate aircraft to a base for accomplishment of the repair required by paragraph (a) of this AD. (c) Equivalent inspection procedures and repairs may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. (d) Record the rod thread major diameter on the appropriate blade record card. (e) Report the helicopter serial number, blade serial number, blade total time and the antinode rod thread major diameter to the Chief, Aircraft Engineering Division, FAA Western Region, P.O. Box 92007, World Way Postal Center, Los Angeles, Ca 90009. (Reporting approved by the Bureau of the Budget under BOB No. 04-R0174.) This amendment is effective October 27, 1978 and was effective upon receipt for all recipients of the priority mail letter dated September 22, 1978 which contained this amendment.
77-14-11: 77-14-11 BRITISH AIRCRAFT CORPORATION: Amendment 39-2963. Applies to BAC 1- 11 200 and 400 series airplanes certificated in all categories, and equipped with any of the following flap assemblies: Inboard Flap Assy. Center Flap Assy. Outboard Flap Assy. AB09 A1/2 or AB09 A3/4 or AB09 A5/6 or AK09 A121/2 or AK09 A123/4 or AK09 A125/6 or AK09 A1579/80 or AP09 A3/4 or AP09 A5/6 or AP09 A1/2 or EN09 A3/4 EN09 A5/6 AP09 A501/2 or EN09 A1/2 Compliance is required as indicated. To detect cracks and prevent possible failure of the flap structure, accomplish the following: (a) For airplanes equipped with flap assemblies (R/H and L/H) which do not incorporate Modification 57-PM2961, the following inspections apply: 1) Upon accumulating 20,000 landings on the flap assemblies or within 50 landings after the effective date of this AD, whichever occurs later, unless accomplished within the preceding 200 landings for inboard flap assemblies or 950 landings for center and outboard flap assemblies, visually inspect the top surface skin panels of all flaps for cracks, paying particular attention to the rivet attachment areas. 2) Repeat the inspection of paragraph (a)(1) of this AD at intervals from the last inspection not to exceed 250 landings for the inboard flaps and 1000 landings for the center and outboard flaps. 3) If, during an inspection required by paragraph (a)(1) or (a)(2) of this AD, signs of movement or distress are found at a group of rivets, within 500 landings, unless already accomplished within the preceding 500 landings, conduct an X-ray or visual inspection of the flap internal structure in accordance with paragraph 2.1.1 of the section titled "Accomplishment instructions" of BAC Alert Service Bulletin 57-A-PM5381, Issue 2, dated July 5, 1976, (hereinafter referred to as BAC ASB 57-A-PM5381), or an FAA-approved equivalent. 4) Upon accumulating 20,000 landings on the inboard flapsor within 100 landings after the effective date of this AD, whichever occurs later, unless already accomplished within the preceding 3,400 landings, conduct an X-ray inspection of the flap structure in accordance with paragraph 2.1.3 of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. 5) Repeat the X-ray inspection of the flap structure of the inboard flaps required by paragraph (a)(4) of this AD at intervals not to exceed 3,500 landings from the last inspection. 6) If an inspection required by paragraph (a)(4) or (a)(5) of this AD confirms that the internal structure is free of cracks, and if the skin panels are free of cracks, the intervals of 250 landings for the repetitive visual inspection of the skin panels of the inboard flaps required by paragraph (a)(2) of this AD may be increased to 1,200 landings. 7) Upon accumulating 23,000 landings on the center and outboard flaps, or within 500 landings after the effective date of this AD, whichever occurs later, unless already accomplished, conduct an X-ray inspection of the flap structure in accordance with the instructions contained in paragraph 2.1.6 of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. (b) For airplanes equipped with flap assemblies (R/H and L/H) that incorporate Modifications 57-PM-1931 and 57-PM-2961, the following inspections apply: 1) Upon accumulating 25,000 landings on the flap assemblies or within 1,000 landings after the effective date of this AD, whichever occurs later, unless already accomplished, conduct an X-ray inspection of the flap structure in accordance with paragraph 2.2.1 of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. 2) Repeat the inspection required by paragraph (b)(1) of this AD, for the inboard flaps only, at intervals not to exceed 5,000 landings. (c) For airplanes equipped with flap assemblies (R/H and L/H) that were originally manufactured to the standard of Modification 57-PM-2961, upon accumulating 25,000 landingson the flap assemblies, or within 1000 landings after the effective date of this AD, whichever occurs later, unless already accomplished, conduct an X-ray inspection of the flap structure in accordance with paragraph 2.3.1 of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. (d) If one or more cracks are found in the top skin panels during an inspection required by this AD, comply with the following as applicable: 1) If skin panel cracks are found that do not exceed two inches in length (measured from the skin panel edge) and no more than three such cracks exist over the span of any one spar with no more than one crack at any rib station, the flap may continue in service provided that each crack is measured for propagation at intervals not to exceed 250 landings. 2) If more than three skin panel cracks are found over the span of any one spar or more than one crack is found at any rib station, all of which measure less than two inches in length (measured from the skin panel edge), before further flight, accomplish temporary repairs in accordance with Figure 2A or 2B of BAC ASB 57-A-PM5381, or accomplish a repair in accordance with the BAC 1-11 Structural Repair Manual, Chapter 57-02-4, or an FAA-approved equivalent. 3) If a skin panel crack is found that exceeds two inches in length (measured from the skin panel edge), before further flight, accomplish a repair in accordance with the BAC 1-11 Structural Repair Manual, Chapter 57-02-4, or an FAA-approved equivalent. 4) If one or more skin cracks that do not exceed six inches in length are found along a rib flange rivet line, and providing not more than one crack exists at any one rib station and not more than two cracks are found on any one flap surface, before further flight, accomplish a temporary repair of the cracked area in accordance with Figure 3 of BAC ASB 57-A- PM5381, or accomplish a repair in accordance with BAC 1-11 Structural Repair Manual, Chapter 57-02-4, or an FAA-approved equivalent. 5) If a skin crack that exceeds six inches in length is found along a rib flange rivet line or if more than two cracks of any length are found on any one flap surface, before further flight, accomplish a repair of the cracked areas in accordance with the BAC 1-11 Structural Repair Manual, Chapter 57-02-4, or an FAA-approved equivalent. 6) If skin cracks are found emanating from beneath rubbing strips or external patches, before further flight, remove the rubbing strip or external patch and accomplish a repair in accordance with the criteria established in paragraph (d)(4) or (d)(5) of this AD as applicable. 7) If temporary repairs are accomplished for cracked areas in the skin panels in compliance with paragraph (d)(2), (d)(4), or (d)(6) of this AD, within 15 landings after accomplishing the temporary repair, conduct an X-ray inspection of the internal structure of the flap in the area of the repair in accordance with paragraph 2.1.3 of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. If the X-ray inspection confirms that the internal structure is free of cracks, the flap with temporary repairs incorporated may remain in service for a period not to exceed 1,200 landings at which time the temporary repairs must be replaced by repairs in accordance with the BAC 1-11 Structural Repair Manual, Chapter 57-02-4, or an FAA-approved equivalent. (e) If one or more cracks are found in the flap spars as a result of an inspection required by this AD, the following apply: 1) Except as provided in paragraph (e)(2) of this AD, before further flight, repair the cracked area of the spar in accordance with Figure 4 or Figure 5, or both, as applicable, of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. 2) If one or more cracks are found in the flanges of the flap spars but they have not progressed into the web of the spar, the flap assembly may remain in service for an additional 600 landings provided that, before further flight, a temporary repair of the affected area is accomplished in accordance with Figure 2A or 2B of BAC Alert Service Bulletin 57-A-PM5381, Issue 2, dated July 5, 1976, or an FAA-approved equivalent. Upon accumulating the 600 additional landings, repair the cracked area of the spar in accordance with Figure 4 or Figure 5, or both, as applicable, of BAC ASB 57-A-PM5381, or an FAA-approved equivalent. (f) For the purpose of complying with this AD, subject to acceptance by the assigned Airworthiness Inspector, the number of landings, if not recorded, may be determined by dividing the actual number of flight hours for the particular flap assembly by the operator's fleet average time from takeoff to landing for the airplane type. (g) Upon the request of an operator, the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region, c/o American Embassy, APO New York, N.Y. 09667, may adjust crack length limitations, approve alternate repetitive inspection intervals and procedures, and approve alternate rework procedures to facilitate continued operations by the operator, if data substantiating such action are submitted by the operator. (h) The repetitive inspections required by paragraphs (a)(2) and (a)(5) of this AD may be discontinued upon the incorporation of Modifications 57-PM-1931, 57-PM-2961, and 57- PM5381. (i) The repetitive inspections required by paragraph (b)(2) of this AD may be discontinued upon the incorporation of Modification 57-PM5381. This amendment becomes effective July 21, 1977.
83-06-10: 83-06-10 SIAI-MARCHETTI: Amendment 39-4597. Applies to Models F260, F260B, and F260C airplanes certificated in any category that are equipped with one of the following propeller governor Part Numbers (PN): HARTZELL PN EQUIVALENT SIAI PN F4-4A 260-13-272-01 F4-4A (F4-11) 260-13-272-05 F4-112 260-13-272-03 F6-15A 260-13-272-07 F4-4A 260-13-272-09 Compliance: Required as indicated, unless already accomplished. To preclude failure of the propeller governor relief valve spring and possible propeller overspeed, accomplish the following: a) Within the next 100 hours time-in-service after the effective date of this Airworthiness Directive (AD), and thereafter at each 100 hour time-in-service interval on propeller governors with a relief valve body which has a rectangular port and within the next 500 hours time-in-service from the effective date of this AD and thereafter at each 500 hour time-in- service interval, on propeller governors with a relief valvebody which has a round port, accomplish the following: 1) Visually inspect the governor's relief valve spring PN A-3107 for cracks or breakage in accordance with the "INSTRUCTIONS" section of the manufacturer's Service Bulletin (SB) No. 260B35, dated March 5, 1982, hereinafter referred to as the SB. i) If the relief valve spring is cracked or broken, prior to further flight, repair and bench test the propeller governor assembly at SIAI-Marchetti or at an FAA approved Hartzell propeller governor overhaul repair station. (This requirement is necessary because the relief valve spring will require an adjustment to regulate the oil pressure to a range within factory tolerances.) ii) If the spring is not cracked or broken, reinstall the relief valve spring and return the airplane to service and continue the repetitive inspections as prescribed in paragraph a) of this AD. b) Within the next 500 hours time-in-service after the effective date of this AD or at the time of the next propeller overhaul, whichever occurs first, replace those propeller governor relief valve bodies (Part Number (PN) A-3173) having rectangular ports with relief valve bodies (PN A-3173) having round ports as shown in Fig. No. 2 of the SB and continue the repetitive inspections of paragraph a) of this AD. NOTE: It is not necessary to send the propeller governor to the test bench to change the valve body unless the relief valve spring is changed. SIAI-Marchetti or the FAA approved Hartzell propeller governor certified overhaul repair station provides the inspection and modification identification of all governor assemblies installed on airplanes or supplied as spare parts. Modified governors with the new relief valve body with round ports will be identified by a printed "X" with indelible ink on the data plate. c) The intervals between the repetitive inspections required by this AD may be adjusted up to 10 percent of the specified interval to allow accomplishing these inspections concurrent with other scheduled maintenance of the airplane. d) Operators who have not kept records of hours time-in-service of the propeller governor relief valve body type must substitute airplane hours time-in-service in lieu thereof. e) Aircraft may be flown in accordance with FAR 21.197 to a location where this (AD) can be accomplished. f) An equivalent method of compliance with this AD if used must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on April 7, 1983.
96-11-12: This amendment adopts a new airworthiness directive (AD) that applies to certain Beech Aircraft Corporation (Beech) Model C90A airplanes equipped with an optional Beech electric trim system or a Collins autopilot system. This action requires modifying the elevator electric trim tab actuator assembly. Failure of the elevator electric trim tab system on a Beech Model C90A prompted the proposed AD action. The actions specified by the proposed AD are intended to prevent possible failure of the elevator electric trim tab system, which, if not detected and corrected, could cause loss of airplane maneuverability and possible loss of control of the airplane.