Results
97-07-10: This amendment adopts a new airworthiness directive (AD) that applies to de Havilland DHC-6 series airplanes that do not have a certain wing strut modification (Modification 6/1581) incorporated. This action requires inspecting the wing struts for cracks or damage (chafing, etc.), replacing wing struts that are found damaged beyond certain limits or are found cracked, and incorporating Modification No. 6/1581 to prevent future chafing damage. This AD results from several reports of wing strut damage caused by the upper fairing rubbing against the wing strut. The actions specified by this AD are intended to prevent failure of the wing struts, which could result in loss of control of the airplane.
74-08-07: 74-08-07 MCDONNELL DOUGLAS: Amendment 39-1812 as amended by Amendment 39-1928, 39-2115, 39-2230, and 39-2798 is further amended by Amendment 39-2819. Applies to all Douglas Aircraft Company DC-10-10, 10F, -30, -30F and -40 airplanes certificated in all categories. \n\n\tCompliance required as indicated, unless previously accomplished. \n\n\t(a)\tWithin 10 hours additional time in service after the effective date of this airworthiness directive, on all DC-10-10/-30/-40 series airplanes incorporating the automatic landing system installed per DC-10 Service Bulletin 22-48 or 22-56, or production equivalent, incorporate the following placard on the cockpit directional guidance control panel: \n\n\t"Do not use autoland." \n\n\t(b)\tWithin 300 hours additional time in service after the effective date of this airworthiness directive accomplish (1) and (2). \n\n\t\t(1)\tIncorporate revisions in the FAA-approved Airplane Flight Manual Flight Guidance Appendices IV, IVA and etc., as applicable, Documents MDC-J1010, MDC-J1030, MDC-J5830 and MDC-J1040, as follows: \n\n\t\t\t(a)\tAdd the following heading in Section I, Limitations, to read: \n\n\t\t\t\t"Automatic Landing System \n\n\t\t\t\tDo not use automatic landing system except for crew training and test flights. Do not use automatic landing system unless reported weather conditions are equal to or better than Category I minimums." \n\n\t\t\t(b)\tRevise the existing heading "Category I and II Automatic Landing," in Section I, Limitations, to read: \n\n\t\t\t\t"Category I Automatic Landing" \n\n\t\t\t(c)\tDelete the existing Category IIIA Automatic Landing system limitations from Section I. (All limitations under the following heading, including the heading, are to be deleted from applicable airplane flight manual appendices.) \n\n\t\t\t\t"Category IIIA Automatic Landing" \n\n\t\t\t(d)\tThe following heading in Section I, Limitations, may be added in lieu of the requirements of paragraph (b)(1)(a) above: \n\n\t\t\t\tAutomatic Landing System \n\n\t\t\t\tDo not use automatic landing system ("LAND" Mode) to touchdown except during crew training and test flights with the reported weather conditions equal to or better than Category I minimums. \n\n\t\t\t\tThe automatic landing system ("LAND" Mode) may be used as an approach coupler for all operations, provided the autopilot is disconnected prior to reaching 100 feet height above touchdown or the published Category II decision height, whichever is higher. \n\n\t\t(2)\tRemove the placard installed per (a), above, and install the following placard: \n\n\t\t\t"Do not use Autoland. See AFM Limitations." \n\n\t(c)\tOperators may elect, at any time, to deactivate the automatic landing system using a method approved by the Chief, Aircraft Engineering Division, FAA Western Region. In addition, when the automatic landing system is deactivated, the "LAND" pushbutton must be placarded "INOP." If this procedure is accomplished, the placards installed per (a) and (b), above, may be removed. \n\n\t(d)\tAn operator may use the DC-10-10, -10F, -30 or -30F automatic landing system for revenue service down to and including Category II meteorological conditions when all of the following are accomplished. \n\n\t\t(1)\tAll DC-10-10 and DC-10-30 airplanes in an individual operator's fleet have been modified per Douglas DC-10 Service Bulletin 22-78, dated February 7, 1975, or later FAA-approved revision. \n\n\t\t(2)\tDouglas DC-10 Service Bulletin 22-80, dated February 7, 1975, or later FAA-approved revision is accomplished. \n\n\t\t(3)\tThe applicable Flight Guidance Appendix to the Airplane Flight Manual must incorporate the applicable revision, as listed below, approved on February 20, 1975 or later FAA-approved revision. \n\n\t\t\tReport No. MDC-J1010 Revision No. 62 \n\n\t\t\tReport No. MDC-J1030 Revision No. 41 \n\n\t\t\tReport No. MDC-J5830 Revision No. 18 \n\n\t\t(4)\tApproval to conduct automatic landings is obtained from the Principal Operations Inspector assigned to the individual operator. \n\n\t\t(5)\tRemove the placard installed by paragraph (b)(2), above, when paragraphs d(1), (2), (3) and (4) are accomplished. \n\n\t(e)\tIf the requirements of paragraph (d) above are met, the following limitations apply to use of the DC-10-10 and DC-10-30 automatic landing systems: \n\n\t\t\tAutomatic Landing System \n\n\t\t\tDo not use Automatic Landing mode until DC-10 Service Bulletin 22-41 and 22-48 and flight functional in accordance with MDC Report Number J6204 or production equivalents are accomplished (DC-10-10 airplanes only). \n\n\t\t\tDo not exceed 235 knots with Single Land or Dual Land modes of the autopilot engaged. \n\n\t\t\tCategory III Automatic Landing \n\n\t\t\tIn addition to the Automatic Landing System Limitations listed above, the following limitation applies: \n\n\t\t\tDo not use Automatic Landing System for Category III operation. \n\n\t(f)\tAn operator may use the DC-10-40 automatic landing system for revenue service down to and including Category II meteorological conditions when all of the following are accomplished.(1)\tAll DC-10-40 airplanes in an individual operator's fleet have been modified per Douglas DC-10 Service Bulletin 22-78 dated March 14, 1975, or late FAA-approved revision. \n\n\t\t(2)\tDouglas DC-10 Service Bulletin 22-80 dated May 13, 1975, or later FAA-approved revision is accomplished. \n\n\t\t(3)\tThe Flight Guidance Appendix to the Airplane Flight Manual, Report No. MDC-J1040, must incorporate Revision No. 22 approved on May 22, 1975, or later FAA-approved revision. \n\n\t\t(4)\tApproval to conduct automatic landings is obtained from the FAA Principal Operations Inspector assigned to the individual operator. \n\n\t\t(5)\tRemove the placard installed by paragraph (b)(2), above, when paragraphs f(1), (2), (3) and (4) are accomplished. \n\n\t(g)\tIf the requirements of paragraph (f) above are met, the following limitations apply to use of the DC-10-40 automatic landing system: \n\n\t\tAutomatic Landing System \n\n\t\tDo not use Automatic Landing mode until DC-10 Service Bulletin 22-56 or production equivalent is accomplished. \n\n\t\tDo not exceed 235 knots with Single Land or Dual Land modes of the autopilot engaged. \n\n\t\tCategory III Automatic Landing \n\n\t\tIn addition to the Automatic Landing System Limitations listed above, the following limitation applies: \n\n\t\tDo not use Automatic Landing System for Category III operation. \n\n\t(h)\tAn operator may use the DC-10-10, -10F, -30, -30F, or -40 Automatic Landing System to Category III meteorological conditions for revenue service, as provided in the applicable FAA approved Airplane Flight Manuals when all of the following are accomplished: \n\n\t\t(1)\tParagraph (d) or (f), as applicable, is accomplished. \n\n\t\t(2)\tInstallation of a redundant VOR/localizer antenna and modification of the existing VOR/localizer antenna on all airplanes in an operator's fleet to provide a redundant antenna. Modify per McDonnell Douglas Service Bulletin 34-78, dated December 6, 1976 or later FAA-approved revisions or production equivalent.(3)\tModification of all yaw computer Part Numbers 3757082-7 or 3757091-9, as applicable, in an operator's fleet to provide additional electrical protection of relay terminals, in accordance with McDonnell Douglas Service Bulletin 22-93, dated December 17, 1976 or later FAA-approved revisions or production equivalent. \n\n\t\t(4)\tIncorporation of the applicable pages of the Flight Guidance Appendix to the Airplane Flight Manual, as listed below, approved on December 27, 1976 or later FAA-approved revisions to provide for removal of the Category III limitations. \n\n\t\t\tReport No. MDC-J 1010 Revision No 74 \n\n\t\t\tReport No. MDC-J 1030 Revision No. 54 \n\n\t\t\tReport No. MDC-J 5830 Revision No. 28 \n\n\t\t\tReport No. MDC-J 1040 Revision No. 27 \n\n\t\t\tReport No. MDC-J 2140 Revision No. 06 \n\n\t\t(5)\tApproval to conduct automatic landings is obtained from the FAA Principal Operations Inspector assigned to the individual operator. \n\n\t(i)\tNotwithstanding the requirements of paragraph (e) or (g),after accomplishment of paragraph (h), the Automatic Landing System may be operated in accordance with the limitations defined in the appropriate FAA approved airplane flight manual as provided in paragraph (h)(4). \n\n\tAmendment 39-1812 became effective April 11, 1974. \n\n\tAmendment 39-1928 became effective August 26, 1974. \n\n\tAmendment 39-2115 became effective March 7, 1975. \n\n\tAmendment 39-2230 became effective June 9, 1975. \n\n\tAmendment 39-2798 became effective January 10, 1977. \n\n\tThis Amendment 39-2819 becomes effective February 1, 1977.
92-24-51: 92-24-51 BOEING: Amendment 39-8439. Docket No. 92-NM-212-AD. Supersedes AD 92- 21-51 R1, Amendment 39-8414. \n\n\tApplicability: All Model 747 series airplanes, certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tNOTE: Paragraphs (d) and (f) of this AD require inspections from both ends of the nacelle strut midspar fuse pins, whereas AD 92-21-51 R1, Amendment 39-8414 (57 FR 53546, November 12, 1992), which is superseded by this AD, required inspection from only one end of the fuse pins. As allowed by the phrase, "unless accomplished previously," paragraphs (d) and (f) of this AD do not require that the inspections performed previously from one end of the fuse pins in accordance with AD 92-21-51 R1 be repeated. For those fuse pins, only the end of the fuse pin not inspected previously must be inspected to comply with the initial inspection requirements of this AD. \n\n\tTo prevent failure of the nacelle strut midspar fuse pins, accomplish the following: \n\n\t(a)\tWithin 30 days after the effective date of this AD, remove all old style nacelle strut midspar fuse pins and replace with new style fuse pins, in accordance with Boeing Service Bulletin 747-54-2063, Revision 9, dated April 23, 1992. When an old style fuse pin is removed, the engine must be removed in accordance with the Boeing Model 747 Maintenance Manual, Section 54-10-03; or supported in accordance with the service bulletin; or supported in a manner approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\t(b)\tAs of 30 days after the effective date of this AD, no person shall install an old style nacelle strut midspar fuse pin on any airplane. \n\n\t(c)\tPerform the inspection required by paragraph (d) of this AD at the times specified in paragraph (c)(1), (c)(2), or (c)(3) of this AD, as applicable. \n\n\t\t(1)\tFor airplanes equipped with Pratt and Whitney or Rolls Royce engines on which the new style nacelle strut midspar fuse pins have accumulated 5,000 or more landings as of the effective date of this AD: Inspect inboard engine positions 2 and 3 within 30 days after the effective date of this AD; and inspect outboard engine positions 1 and 4 within 60 days after the effective date of this AD. \n\n\t\t(2)\tFor all other airplanes equipped with Pratt and Whitney, Rolls Royce, or General Electric engines having new style nacelle strut midspar fuse pins, other than those identified in paragraph (c)(1) of this AD: Inspect inboard engine positions 2 and 3 at the later of the times specified in paragraph (c)(2)(i) or (c)(2)(ii) of this AD. \n\n\t\t\t(i)\tPrior to the accumulation of 3,000 landings on the fuse pin or within 3 years since installation of the fuse pin, whichever occurs first; or \n\n\t\t\t(ii)\tWithin 60 days after the effective date of this AD. \n\n\t\t(3)\tFor all other airplanes equipped with Pratt and Whitney, Rolls Royce, or General Electric engines having new style nacelle strut midspar fuse pins, other than those identified in paragraph (c)(1) of this AD: Inspect outboard engine positions 1 and 4 at the later of the times specified in paragraph (c)(3)(i) or (c)(3)(ii) of this AD: \n\n\t\t\t(i)\tPrior to the accumulation of 3,000 landings on the fuse pin or within 3 years since installation of the fuse pin, whichever occurs first; or \n\n\t\t\t(ii)\tWithin 90 days after the effective date of this AD. \n\n\t(d)\tIn accordance with the compliance times specified in paragraph (c) of this AD, perform a detailed visual inspection to detect corrosion of the new style nacelle strut midspar fuse pins from each end of the fuse pin with the insert removed, in accordance with Boeing Alert Service Bulletin 747-54A2150, Revision 1, dated November 13, 1992. When a new style fuse pin is removed, the engine must be removed in accordance with the Boeing Model 747 Maintenance Manual, Section 54-10-03; or supported in accordance with the service bulletin; or supported in a manner approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\t(e)\tIf corrosion is detected as a result of the inspection required by paragraph (d) of this AD, prior to further flight, accomplish the following: \n\n\t\t(1)\tIf the amount of corroded material that must be removed exceeds the 0.010-inch limit on the fuse pin inner diameter specified in the service bulletin, replace the fuse pin with a new style fuse pin. Thereafter, accomplish the actions required by this AD on the newly-installed fuse pins. \n\n\t\t(2)\tIf the amount of corroded material that must be removed is more than light, and equal to or less than the 0.010-inch limit on the fuse pin inner diameter specified in the service bulletin, rework the fuse pin in accordance with the service bulletin instructions, or replace the pin with a new style fuse pin. "Light" corrosion is characterized by discoloration or pitting to a depth of not more than 0.001-inch maximum. Thistype of corrosion can be removed normally by light hand sanding. A fuse pin that has been reworked in accordance with Boeing Alert Service Bulletin 747-54A2150, dated October 5, 1992; or Revision 1, dated November 13, 1992, must be replaced with a new fuse pin prior to the accumulation of 3,000 landings on the fuse pin, or 3 years since the pin was reworked and reinstalled, whichever occurs first. \n\n\t\t(3)\tIf the corrosion is light, remove the corroded material in accordance with the service bulletin. Thereafter, repeat the inspections required by paragraph (h) of this AD. \n\n\t(f)\tFollowing accomplishment of the actions required by paragraphs (d) and (e) of this AD, if the fuse pin has been found to be corrosion free, or if the pin has been reworked on the airplane to remove light corrosion, prior to further flight, perform an ultrasonic inspection to detect cracks in the fuse pin from each end of the fuse pin with the insert removed, in accordance with Boeing Alert Service Bulletin 747-54A2150, Revision 1, dated November 13, 1992. When a new style fuse pin is removed, the engine must be removed in accordance with the Boeing Model 747 Maintenance Manual, Section 54-10-03; or supported in accordance with the service bulletin; or supported in a manner approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\t(g)\tIf any crack is found as a result of the inspections required by paragraph (d) or (f) of this AD, prior to further flight, replace the pin with a new style fuse pin in accordance with Boeing Alert Service Bulletin 747-54A2150, Revision 1, dated November 13, 1992. Thereafter, accomplish the actions required by this AD on the newly-installed fuse pins. \n\n\t(h)\tThereafter, repeat the actions required by paragraphs (d), (e), (f), and (g) of this AD at intervals not to exceed 500 landings. \n\n\t(i)\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport 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, Seattle ACO. \n\n\tNOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Seattle ACO. \n\n\t(j)\tSpecial 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. \n\n\t(k)\tThe inspections, replacement, and rework shall be done in accordance with Boeing Alert Service Bulletin 747-54A2150, Revision 1, dated November 13, 1992; and Boeing Service Bulletin 747-54-2063, Revision 9, dated April 23, 1992; as applicable. 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. Incorporation by reference of Boeing Service Bulletin 747-54-2063, Revision 9, dated April 23, 1992, was approved previously by the Director of the Federal Register as of November 27, 1992 (57 FR 53546, November 12, 1992). Copies may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124- 2207. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. \n\n\tThis AD 92-24-51 supersedes AD 92-21-51 R1, Amendment 39-8414, which superseded AD 86-22-01, Amendment 39-5437, and AD 91-09-01, Amendment 39-6970. \n\n\t(l)\tThis amendment becomes effective on January 4, 1993, to all persons except those persons to whom it was made immediately effective by telegraphic AD T92-24-51, issued on November 13, 1992, which contained the requirements of this amendment.
2011-22-08: This amendment supersedes an existing airworthiness directive (AD) that applies to MD Helicopters, Inc. (MDHI) Model MD900 helicopters. That AD currently requires turning ON both Vertical Stabilizer Control System (VSCS) switches and turning OFF the autopilot (AP/SAS) switch; pulling certain AP/SAS circuit breakers; installing a placard near the AP/SAS master switch; installing an airspeed limitation placard on the instrument panel; and making changes to the Rotorcraft Flight Manual (RFM). This amendment retains those requirements and provides an option of replacing each affected tube adapter with a newly-designed tube adapter, which provides terminating action for the unsafe condition. This amendment is prompted by the manufacturer introducing an improved, newly-designed tube adapter. The actions specified by this AD are intended to prevent loss of yaw control and subsequent loss of control of the helicopter.
74-08-10: 74-08-10 BOEING: Amendment 39-1817. Applies to forward electronic hatches on all model 747 series airplanes certificated in all categories. Compliance required as indicated. \n\tTo detect cracks in the hatch, accomplish the following: \n\t(a)\tWithin the next 100 hours time in service from the effective date of this AD, visually inspect the forward electronic hatch on airplanes having 5,000 or more hours time in service, unless inspected within the last 100 hours. If no cracks are found, or if repaired in accordance with paragraph (e), or if the hatch is replaced with a hatch which has not been modified in accordance with Boeing Service Bulletin 747-52-2088, repeat the inspection every 500 hours time in service until modified in accordance with the termination action specified in Boeing Service Bulletin 747-52-2088, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Northwest Region. \n\t(b)\tIf cracks are found in the frame or outer skin, repair the cracks in accordance with paragraph (e), or replace the hatch before further flight. \n\t(c)\tIf cracks are found in the channel stiffeners or in the horizontal angles on the lower side of the hatch, replace or repair in accordance with paragraph (e), or reinspect every 100 hours and repair within 1,000 hours time in service from effective date of this AD or from discovery of the cracks, whichever is later. The hatch must be replaced or repaired prior to further flight, if cracks in any one horizontal angle P/N 65B11898-2 or -4 exceed 5 inches in length or a combined crack length of 8 inches in both angles. \n\t(d)\tAny broken latch pin must be replaced prior to further flight. \n\t(e)\tHatch repairs shall be accomplished in accordance with the FAA approved Boeing 747 Structural Repair Manual, or Boeing Service Bulletin 747-52-2088 or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Northwest Region. \n\t(f)\tModification of the forwardelectronic hatch in accordance with Boeing Service Bulletin 747-52-2088 or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Northwest Region, constitutes terminating action under the provisions of this AD. \n\t(g)\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 specified in this AD to permit compliance at an established inspection period of the operator, if the request contains substantiating data to justify the adjustment. \n\tAircraft may be ferried to a base for maintenance in accordance with Sections 21.197 and 21.199 of the Federal Aviation Regulations.\n \tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herewith and made a part hereof, pursuant to 5 U.S.C. 552(a)(i). All persons affected by this directive who have not already received these documents may obtain copies upon request to The Boeing Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA Northwest Region, Boeing Field, Seattle, Washington. \n\n\tThis amendment becomes effective upon publication in the Federal Register.
2009-10-09 R2: We are revising an existing airworthiness directive (AD) for certain Cessna Aircraft Company (Cessna) Models 150F, 150G, 150H, 150J, 150K, 150L, 150M, A150K, A150L, A150M, F150F, F150G, F150H, F150J, F150K, F150L, F150M, FA150K, FA150L, FRA150L, FA150M, FRA150M, 152, A152, F152, and FA152 airplanes. That AD currently requires either installing a placard prohibiting spins and other acrobatic maneuvers in the airplane or replacing the rudder stop, the rudder stop bumper, and the attachment hardware with a rudder stop modification kit. This new AD requires a change to the modification kit and removal of a small amount of material from the rudder horn assembly for those that have not yet complied with the existing AD or for those who can not comply with the existing AD (because they were unable to obtain full rudder travel with the existing kits). This AD was prompted by operators who have reported difficulty in obtaining full rudder travel with the existing modification kit. We are issuing this AD to revise the kits to use longer rivets and allow a small amount of material to be removed from the rudder horn assembly, which allows operators to obtain full rudder travel.
97-06-14: This amendment adopts a new airworthiness directive (AD), applicable to General Electric Company CF34 series turbofan engines, that reduces the allowable operating cyclic life limit for affected fan disks. This amendment is prompted by an updated stress and life analysis. The actions specified by this AD are intended to prevent fan disk rupture, engine failure, and damage to the aircraft.
91-25-04: 91-25-04 GARRETT AUXILIARY POWER DIVISION: Amendment 39-8105. Docket No. 91-NM-151-AD. \n\n\tApplicability: Model TSCP700-4B auxiliary power units (APU) prior to serial number 90697, as installed in, but not limited to, McDonnell Douglas Model DC-10 and KC-10 (military) series airplanes; and Model TSCP700-5 APU's prior to serial number 80443, as installed in, but not limited to, Airbus Industrie Model A300 series airplanes; certificated in any category. \n\n\tCompliance: Required within 24 months after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent uncontained high pressure turbine (HPT) disc failures, accomplish the following: \n\n\t(a)\tReplace the HPT containment ring, part number (P/N) 976850-1, with P/N 3614975-1; and replace the HPT containment support, P/N 3604274-1, with P/N 3614934-1; in accordance with the accomplishment instructions in Garrett Service Bulletin TSCP700-49-5892, Revision 2, dated October 10, 1990. \n\n\t(b)\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(c)\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(d)\tThe replacement requirements shall be done in accordance with Garrett Service Bulletin TSCP700-49-5892, Revision 2, dated October 10, 1990, which contains the following list of effective pages: \n \n\nPage Number\nRevision Level\nDate\n1, 7/8\n2\nOctober 10, 1990 \n2, 3/4, 5, 6\nOriginal)\nMay 14, 1990 \n9/10\n1\nJuly 3, 1990\n\n\tThis 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 Garrett Airlines Services Division, Technical Publications, Department 65-70, P. O. Box 52170, Phoenix, Arizona 85072-2170. 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\tThis amendment (39-8105, AD 91-25-04) becomes effective on January 7, 1992.
2011-23-02: This amendment supersedes an existing airworthiness directive (AD) for Bell Model 205B and 212 helicopters with certain main rotor blade (blade) assemblies installed. That AD currently requires washing the upper and lower surfaces of each blade and visually inspecting the grip plates, doublers, and the remaining upper and lower surfaces of the blades in the area between blade stations 24.5 to 40 for an edge void, corrosion, or a crack. This amendment retains the requirements of that AD for the affected part-numbered blades but increases the scope and frequency of the inspections and expands the applicability to include the Model 205A-1 and 210 helicopters, additional blade part numbers, and all helicopter serial numbers for the affected helicopter models. This amendment also requires applying a light coat of preservative oil (C-125) to all surfaces of the blade in addition to the inspection areas as required in the existing AD. This amendment is prompted by an additional report of a fatigue crack on a blade installed on a Model 212 helicopter. The actions specified by this AD are intended to detect an edge void, corrosion, or a crack on a blade, and to prevent loss of a blade and subsequent loss of control of the helicopter.
74-23-06: 74-23-06 MCDONNELL DOUGLAS: Amendment 39-2005 as amended by Amendment 39-2999 is further amended by Amendment 39-3031. Applies to McDonnell Douglas DC-8 series airplanes, certificated in all categories, incorporating Pratt and Whitney JT3D engines. \n\n\tWithin 24 hours after receipt of this telegram, incorporate the operating limitations and procedures, set forth in paragraphs (1), (2), and (3), below, into the Douglas DC-8 FAA approved Airplane Flight Manual. Make appropriate notations on the log of pages. Operators shall promptly implement these limitations and procedures. \n\n\t(1)\tRevise Section I, Limitations, to include a new item relative to fuel management. \n\n\t\tFUEL BOOST AND/OR FEED PUMP OPERATION \n\n\tPrior to descent, except during landing, the main tank pumps must be in the boost and feed position. \n\n\t(2)\tRevise Section I, Limitations, to include a new item relative to engine operation at idle power. \n\n\t\tFLIGHT IDLE OPERATION \n\n\tInflight at altitudes above 6,000 feet MSL and at indicated airspeeds below 200 knots, a minimum N2 engine rotor speed of 62 percent must be maintained except during landing. \n\n\t(3)\tRevise Section II, Emergency Procedures, to include a new item relative to recovery from a condition where an engine(s) fail to accelerate in flight after the throttle levers are advanced from the idle settings: \n\n\t\tENGINE RESPONSE TO THROTTLE LEVER(S) \n\n\tIf the engine(s) fail to accelerate in flight after the throttle lever(s) are advanced, the following procedure should be used after it has been determined that a flameout has not occurred. This procedure need not be continued and affected systems may be reactivated if engine operation has been restored to normal. \n\t\n\nPHASE I - II\n\nThrottle Lever(s) affected engine(s)\nAdvance\nEngine Anti-Ice \nOff \nAirspeed\nIncrease (as practicable)\n\t\t \t \nPHASE III \n\nIgnition \nOverride/Both \nMain Tank Pumps \nBoost and Feed\nPneumatic Bleed (Affected Engine(s))\nReduce \nElectrical Load\nRemove Non-required \n\t\n\tNote: One reservoir feed pump may be inoperative provided: \n\n\t(a)\tThe affected reservoir feed pump is placarded inoperative at the feed pump switch position. \n\n\t(b)\tEstablished maintenance procedures for this item are followed. \n\n\t(c)\tSufficient fuel is carried in the associated tank to provide a minimum of 2000 pounds of additional fuel in excess of the fuel (including reserves) needed for the flight. \n\n\t(d)\t'FUEL LOADING AND MANAGEMENT' is in accordance with the FAA Approved Airplane Flight Manual. \n\n\tAmendment 39-2005 became effective November 14, 1974, for all persons except those to whom it was made effective immediately by telegram dated October 11, 1974. \n\n\tAmendment 39-2999 became effective August 10, 1977. \n\n\tThis amendment 39-3031 becomes effective September 15, 1977.
2011-20-08: This amendment adopts a new airworthiness directive (AD) for the specified Agusta model helicopters. This action requires inspecting certain modules and related connectors for corrosion. If there is corrosion on the connectors, this AD requires cleaning the connectors before further flight. If there is corrosion on a module, before further flight, this AD requires replacing the module with an airworthy module. This AD also requires modifying the Number 2 Modular Avionic Unit (MAU) ventilation duct. This amendment is prompted by some in- flight emergencies due to internal corrosion of the MAU circuit card assemblies. The actions specified in this AD are intended to detect corrosion of certain modules to prevent the display of misleading data to the flight crew, disengagement of the flight director modes of the autopilot or other alert system, increased workload of the flight crew, and subsequent loss of control of the helicopter.
97-06-12: This amendment supersedes two existing airworthiness directives (AD), applicable to British Aerospace Model BAe 146 and Avro 146-RJ series airplanes, that currently require inspections to detect cracking of the upper main fitting of the nose landing gear (NLG), and replacement or repair of cracked parts, if necessary. Those actions were prompted by reports of cracking in the main fittings of the NLG. This amendment requires that, for certain airplanes, the inspections be accomplished at reduced intervals. This amendment is prompted by the results of new analyses of the cracking that were conducted by the manufacturer of the NLG. The actions specified by this AD are intended to prevent failure of the main fitting, which could lead to collapse of the NLG during landing.
2022-02-12: The FAA is adopting a new airworthiness directive (AD) for all Leonardo S.p.a. Model AB139 and AW139 helicopters. This AD was prompted by the identification of certain parts needing maintenance actions, including life limits and maintenance tasks. This AD requires incorporating into maintenance records requirements (airworthiness limitations), as specified in a European Aviation Safety Agency (now 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.
2005-13-33: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Model A300 B2 and B4 series airplanes. This AD requires modifying the wiring of the autopilot pitch torque limiter switch. This AD is prompted by several reports of pitch trim disconnect caused by insufficient length in the wiring to the pitch torque limiter lever. We are issuing this AD to prevent possible trim loss when the flightcrew tries to override the autopilot pitch control, which could result in uncontrolled flight of the airplane.
95-06-03: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 95-06-03 which was sent previously to all known U.S. owners and operators of Robinson Helicopter Company (Robinson) Model R22 helicopters by individual letters. This AD requires an inspection and modification of the main rotor (M/R) gearbox. This amendment is prompted by a report of an incident involving a Model R22 helicopter in which the two M/R mast spanner nuts (nuts) became loose, resulting in failure of the M/R mast support structure. The actions specified by this AD are intended to prevent M/R separation and subsequent loss of control of the helicopter.
95-12-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Aerospatiale Model ATR72 series airplanes. This action requires repetitive inspections to detect displacement of the rear hinge bush, and to detect cracking or rupture of the rear hinge pin on the main landing gear (MLG) leg; and the correction of any discrepancies. This amendment is prompted by a report of the failure of this hinge pin on an in-service airplane. The actions specified in this AD are intended to prevent failure of the hinge pin, which can lead to failure of the MLG leg or MLG attachment assembly.
2011-21-17: We are adopting a new airworthiness directive (AD) for all General Electric Company (GE) CT7-8A, CT7-8A1, CT7-8E, and CT7-8F5 turboshaft engines with a fuel filter differential pressure switch, part number (P/N) TD028VF0H7Y5 (part of the fuel filter assembly, P/N 4110T53P06) installed. This AD requires daily visual inspections of the fuel filter differential pressure switch for fuel leaks and for excessive cracking of the switch mounting flanges due to stress- corrosion. This AD also requires the installation of a collar kit over the fuel filter differential pressure switch as terminating action to the daily inspections. This AD was prompted by reports of 47 fuel filter differential pressure switches found with stress-corrosion cracking of the mounting flanges. We are issuing this AD to prevent unrecoverable in-flight engine shutdown, engine bay fire due to fuel leakage, and forced landing or accident.
97-06-09: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-300, -400, and -500 series airplanes. This AD requires replacing certain aileron/rudder trim control modules with an improved module that contains an improved rudder trim switch that precludes the problems of sticking associated with the existing switch. This amendment is prompted by reports of sticking conditions in the rudder trim switch. The actions specified by this AD are intended to prevent such sticking, which could result in uncommanded movement of the rudder and consequent deviation of the airplane from its set course.
69-26-06: 69-26-06 MCDONNELL DOUGLAS: Amendment 39-902. Supersedes Amendment 39-738, AD 69-06-03. Applies to all McDonnell Douglas Model DC-9 Series aircraft. \n\n\tCompliance required as indicated, unless already accomplished. \n\n\tTo prevent heat damage to the H.F. (if installed) and V.H.F. coaxial cables and other wiring located in the tail compartment of DC-9 Series aircraft, accomplish the following: \n\n\t(A)\tWithin the next 200 hours' time in service after the effective date of this AD, unless already accomplished, perform the following: \n\n\t\t(1)\tDetermine that the H.F. (if installed) and V.H.F. coaxial cables located in the tail compartment of the aircraft adjacent to the 8th stage bleed duct have not deteriorated due to excessive heat. The determination may be accomplished by the use of electrical tests such as fault finder pulse indications, reflectometer measurements, or X-ray inspections or a satisfactory equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. If the electrical tests indicate any coaxial cable impedance change in the areas of the 8th stage bleed duct or the APU exhaust shroud, or the X-ray inspections show physical change, such as noticeable drift of the center conductor, or any unsatisfactory condition in these areas, replace the damaged coaxial cables or repair the damaged areas of the cables by use of proper connectors and new coaxial sections, in conjunction with (2) and (3), below. In lieu of electrical testing of X-ray inspection an operator may replace the cables within this 200 hour period. \n\n\t\t(2)\tProvide maximum possible clearance (at least one inch) between the H.F. (if installed) and V.H.F. conduits, and the right engine 8th stage bleed duct by rotating the conduit clamps and reworking the spacers as necessary. NOTE: McDonnell Douglas DC-9 Alert Service Bulletin A23-24, dated February 21, 1969 describes this work. \n\n\t\t(3)\t(a)\tVisually inspect the APU exhaust shroud for any indications of overtemperature condition, such as shroud discoloration or exterior airframe paint discoloration around the shroud outlet. If the APU exhaust duct has been deformed or leaks, and continued use of the APU is desired, replace the duct in accordance with McDonnell Douglas S.B. 49-8, dated May 2, 1966, and Service Letters AOL-9 No. 74, dated February 6, 1967, and AOL-9 No. 139, dated September 29, 1967, or later FAA approved revisions, or an equivalent duct replacement approved by the Chief, Aircraft Engineering Division, FAA, Western Region. \n\n\t\t\t(b)\tInspect all wiring adjacent to the APU exhaust shroud for heat damage. Replace or repair to an airworthy condition all wiring found damaged. \n\n\t\t(4)\tPressure test the pneumatic duct installation in the DC-9 tail cone area in accordance with the DC-9 Maintenance Manual Temporary Revision 36-19, dated December 3, 1969, or the subsequent equivalent revision, or an equivalent pressure test approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t\t(5)\tSteps (1) and (3) above, must be repeated prior to any further IFR operation after every report of a pneumatic duct malfunction or an APU exhaust duct failure until (A)(7) or (B), below, has been accomplished. \n\n\t\t(6)\tStep (4), above, must be repeated whenever pneumatic duct maintenance is performed in the tail cone area. \n\n\t\t(7)\tSteps (1) through (5), above, need not be accomplished if the operator accomplishes Step B below, within 200 hours' time in service from the effective date of this AD. \n\n\t(B)\tWithin the next 2000 hours' time in service from the effective date of this AD, unless already accomplished, perform the following in accordance with McDonnell Douglas Service Bulletin 23-24, Rev. 2, dated June 23, 1969; S.B. 23-28, dated December 3, 1969, and S.B. 27-104, Rev. 2, dated April 15, 1969, or later FAA approved revisions, or an equivalent installation and modification approved by the Chief, Aircraft Engineering Division, FAA, Western Region: \n\n\t\t(1)\tInstall an insulation blanket on the 8th stage bleed duct adjacent to the HF and VHF coaxial cable conduits. \n\n\t\t(2)\tReroute the HF (if installed) and VHF coaxial cables, and the other wiring bundle (flight recorder, interphone wiring, etc.) away from the APU exhaust shroud area where they are now located. \n\n\t\t(3)\tReplace the sections of the polyethylene dielectric type VHF and HF (if installed) coaxial cables with sections of electrically equivalent polytetrafluoroethylene (teflon) dielectric type coaxial cables from just forward of the pressure dome feed-through throughout the tail compartment, or from between just aft of the pressure dome feed-through to just aft of the exhaust duct from the air condition pack heat exchangers, and apply PF105-700 glass fiber batt and CRS-102 silicon wrap heat insulation material over all exposed low temperature cable which is not installed in conduit. NOTE: No additional rerouting or repositioning, other than that specified inparagraph (A) (2) and (B)(2), is required. \n\n\t\t(4)\tAdd a metal heat shield between the eighth stage pneumatic duct and electrical wire bundle in the tail cone R.H. side just aft of pressure panel and forward of the eighth stage bleed duct. \n\n\t\t(5)\tReposition the wire harnesses FBC and DDC, containing overheat sensor wiring and APU generator control wiring located in the tailcone L.H. side to a new position more outboard of the wing ice protection duct. \n\n\tNOTE: Compliance with the coaxial cable separation and rerouting modification also provided in AOL No. 9-333, dated August 27, 1969, and Service Bulletin No. 23-28, dated December 3, 1969, is optional. \n\n\tThis AD supersedes amendment 39-738, (34 F.R. 5427) AD 69-6-3. \n\n\tThis amendment becomes effective on December 30, 1969.
89-10-04: 89-10-04 McDONNELL DOUGLAS: Amendment 39-6204. \n\n\tApplicability: Model DC-8 series airplanes, equipped with left (LH) or right (RH) main landing gear (MLG) attach fitting, P/N(s) 5611425-1 through -508, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent severe structural damage to the airplane during takeoff or landing due to stress corrosion failure of the MLG attach fittings, accomplish the following: \n\n\tA.\tWithin the next 12 calendar months after the effective date of this AD, unless already accomplished within the last 12 calendar months, and thereafter at intervals not to exceed 24 calendar months, except as provided below, perform a visual inspection of the MLG attach fittings for cracks at locations in accordance with Figure 1. of McDonnell Douglas DC-8 Service Bulletin 57-94, Revision 1, dated June 23, 1987 (hereafter referred to as the Service Bulletin). \n\n\tB.\tIf no cracks are found, apply LPS-3 corrosion inhibiting oil to the fitting in accordance with the Service Bulletin, and repeat inspections for cracks in accordance with paragraph A. of this AD. \n\n\tC.\tIf cracks are found, accomplish the following: \n\n\t\t1.\tIf cracks are located in area 1 or 3, as defined in Figure 1. of the Service Bulletin, before further flight, replace the fitting, P/N(s) 5611425-1, -2, -501, -502, -503, -504, -505, -506, -507, or -508, with respective P/N(s) 5893930-1, -2, -1, -2, -509, -510, -507, -508, -505, or -506. \n\n\t\t2.\tIf cracks are located in area 2, as defined in Figure 1. of the Service Bulletin, accomplish the following: \n\n\t\t\ta.\tIf cracks are within limits as prescribed by Figure 1. of the Service Bulletin, apply LPS-3 corrosion inhibiting oil to the fitting in accordance with the Service Bulletin, and repeat visual inspections for crack development at intervals not to exceed 7 calendar days, in accordance with the Service Bulletin. \n\n\t\t\tb.\tIf cracks exceed limits as prescribed by Figure 1. of the Service Bulletin, replace the fitting in accordance with paragraph C.1. of this AD before further flight. \n\n\tD.\tReplacement of both the LH and RH MLG attach fittings in accordance with paragraph C.1. of this AD constitutes terminating action for the inspection requirements of this AD. \n\n\tE.\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, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspection (PMI), who may add any comments and then send it to the Manager, Los Angeles Aircraft Certification Office. \n\n\tF.\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 documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director of Publications, C1-LOO (54-60). These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or at 3229 East Spring, Long Beach, California. \n\n\tThis amendment (39-6204, AD 89-10-04) becomes effective on May 29, 1989.
72-04-01: 72-04-01 BOEING: Amendment 39-1392 as amended by Amendment 39-1407. Applies to 747-100 and 747-200B Series airplanes. \n\tCompliance required as indicated. \n\tTo prevent unscheduled stabilizer trim and to maintain stabilizer control capability, accomplish the following: \n\t(a)\tFor airplanes incorporating stabilizer trim modules Boeing P/N 60B80027-2 and 60B80027-3 and/or stabilizer trim drive motors P/N 60B00250-1, within 100 hours time in service after effective date of this A.D., and, thereafter at intervals not to exceed 100 hours time in service from the last inspection, test the stabilizer trim system components per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent tests approved by the Chief, Aircraft Engineering Division, FAA Western Region, until modified in accordance with paragraph (c), below. \n\t(b)\tReplace, or modify, prior to further flight, stabilizer trim system components which are found defective bythe inspections per paragraph (a), above, in accordance with Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent replacements or modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(c)\tThe inspections required per (a), above, may be discontinued after accomplishment of the following: \n\t\t(1)\tReplace stabilizer trim modules, Boeing P/N 60B80027-2 and 60B80027-3, with stabilizer trim modules modified with improved arming and control valve seals and steel retainer caps, per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\tNOTE: Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, incorporates LTV Electrosystems Service Bulletins 27-6, 27-8, and 27-9 in "Part II, Terminating Action." \n\t\t(2)\tReplace stabilizer trimmotor, Boeing P/N 60B00250-1, without suffix "D" identification following unit serial number identification, with stabilizer trim motors modified with solid locking pins, per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\tNOTE: Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, incorporates Vickers Service Bulletin 910274-4, dated February 10, 1972. \n\t(d)\tWithin 100 hours time in service after the effective date of this A.D., unless already accomplished, incorporate in the FAA approved Airplane Flight Manual, "Emergency Procedures" (Section 2), the following procedures: \n\t"UNSCHEDULED STABILIZER TRIM \n\tRecall \n\tStabilizer Trim Hydraulic Switches - \n\tCUTOUT \n\tReference \n\tAutopilot - DISENGAGE \n\tControl column movement in opposition of trim will stop unscheduled trim caused by electrical fault. \n\tMoving stabilizer trim hydraulic switches to CUTOUT will stop any unscheduled trim. Allow sufficient time for valves to operate. \n\tA portion of the system may be determined to be usable by moving one stabilizer trim hydraulic switch at a time to NORM. If trim is normal that system may be used to adjust trim as required." \n\tAmendment 39-1392 became effective February 11, 1972. \n\tThis amendment 39-1407 becomes effective March 14, 1972.
2011-23-05: We are superseding an existing airworthiness directive (AD) for certain Model 737-300, -400, and -500 series airplanes. That AD currently requires repetitive inspections for cracking of the 1.04-inch nominal diameter wire penetration hole, and applicable related investigative and corrective actions. This AD reduces the compliance times for those actions. This AD was prompted by reports of cracking in the frame, or in the frame and frame reinforcement, common to the 1.04- inch nominal diameter wire penetration hole intended for wire routing; and recent reports of multiple adjacent frame cracking found before the compliance time required by the existing AD. Such cracking could reduce the structural capability of the frames to sustain limit loads, and result in cracking in the fuselage skin and subsequent rapid depressurization of the airplane. We are issuing this AD to correct the unsafe condition on these products.
97-06-05: This amendment adopts a new airworthiness directive (AD) that applies to Avions Pierre Robin Model R2160 airplanes. This action requires repetitively inspecting the weld area between the strut and the lower plate of the nose landing gear leg for cracks, and replacing the strut when cracks are found. The AD is the result of several reports of cracks in the weld securing the nose wheel steering bottom bracket to the nose landing gear leg on the affected airplanes. The actions specified by this AD are intended to prevent nose landing gear failure caused by cracks in the weld area between the strut and the lower plate of the nose landing gear leg, which could result in loss of control of the airplane during landing operations.
77-16-05: 77-16-05 CESSNA: Amendment 39-2998. Applies to Models 210-5(205), 206, P206/TP206, U206/TU206, 207/T207, and 210/T210 series airplanes of the serial numbers specified below. \n\n\tCompliance: Required within the next 25 hours time-in-service after the effective date of this AD, unless already accomplished. \n\n\tTo prevent malfunction of the P/N C291503-0101 or P/N 1216100-1 fuel selector valve, accomplish the following: \n\n\tA.\tOn Model 210-5(205) series (serial numbers 205-0481 thru 205-0577), Model 206 series (serial numbers 206-0001 thru 206-0275), Model P206/TP206 series (serial numbers P206-0001 thru P20600647), Model U206/TU206 series (serial numbers U206-0276 thru U20603123), Model 207/T207 series (serial numbers 20700001 thru 20700322), Model 210/T210 series (serial numbers 21058221 thru 21061154, and T210-0001 thru T210-0454) airplanes, examine the aircraft maintenance records to determine whether the fuel selector valve has been changed subsequent to December 19, 1975.If the valve has not been changed, make an entry in the maintenance record indicating this AD is not applicable to the airplane and no further action is necessary. Examination of the records and the record entry may be accomplished by the owner/operator. \n\n\tIf the valve has been changed subsequent to December 19, 1975, accomplish a fuel valve inspection and, if indicated, replacement in accordance with paragraph C. below. \n\n\tB.\tOn Model U206/TU206 series (serial numbers U20603124 thru U20603712, U20603714 thru U20603791, U20603793 thru U20603797, U20603799 thru U20603803, U20603805, U20603808 thru U20603812, U20603814 thru U20603846, U20603848, U20603850, U20603851, U20603853 thru U20603857, U20603861, U20603862, U20603867 thru U20603871, U20603875, U20603876, U20603882 and U20603886), Model 207/T207 series (serial numbers 20700323 thru 20700373, 20700375 thru 20700394), and Model 210/T210 series (serial numbers 21061155 thru 21061731, 21061733 thru 21061766, 21061768 thru 21061860, 21061862 thru 21061881, 21061883 thru 21061987, 21061990 thru 21061993, 21061995 thru 21062005, 21062007 thru 21062009, 21062012 thru 21062017, 21062020 thru 21062022, 21062024 thru 21062027, 21062029, 21062033 thru 21062037, 21062043) airplanes, accomplish a fuel selector valve inspection and, if indicated, replacement in accordance with paragraph C. below. \n\n\tC.\t1.\tPlace fuel selector valve in OFF position. \n\n\t\t2.\tRemove selector valve handle and associated parts. Obtain access to the valve by removal of the control valve pedestal and selector valve access plate from the floor. \n\n\t\t3.\tCheck the selector valve serial number. If the serial number is 1421 thru 3269 inclusive, accomplish the pull test inspection described in paragraph C.4. below. If the serial number is not within this block, the valve is not affected, and the aircraft may be returned to service after reassembly. \n\n\t\t4.\tUsing tools fabricated in accordance with Figure 1, or an equivalent test arrangement, accomplish a pull test on the selector valve in accordance with the following procedures: \n\n\t\t\ta.\tRemove safety wire, roll pin, and valve handle shaft from valve control yoke. \n\n\t\t\tb.\tFeed cable through hole in yoke and crimp cable securely. Place bar in loop in other end of cable. \n\n\t\t\tc.\tRemove four screws securing selector valve cover. Break the seal between the cover and body and rotate the cover to positively assure it is free. \n\n\t\t\td.\tAttach a tensiometer to the cable in the section between loops and from a seated or squatting position, with one foot on each side of the valve, using the legs and arms to lift, apply an upward force of 130 pounds minimum to 150 pounds maximum directly in line with the shaft while observing the shaft for movement. Exert upward force gradually. If any upward movement of valve shaft is noted, release force immediately. Drain fuel from airplane and install a new valve in accordance with the aircraft's Service Manual Instructions. \n\nCAUTIONIf the shaft pulls out of the valve body, fuel spillage may result. Therefore, all precautions applicable to working with or around open fuel should be observed. \n\n\t\t\te.\tIf no upward movement of the valve shaft is noted or evident, reassemble the valve and aircraft and return to service. \n\n\n\n\nAD 77-16-05 \n\n\tD.\tAny equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. \n\n\tCessna Service Letter SE77-22, dated June 27, 1977, or later approved revisions pertain to the subject matter of this AD. \n\n\tThis amendment becomes effective August 11, 1977.
79-09-01: 79-09-01 BOEING: Amendment 39-3452. Applies to Model 727-100 series airplanes certificated in all categories listed in Boeing 727 Service Bulletin No. 24-30, revised May 18, 1966, that complied with FAA AD 66-30-02 by having been modified to Option No. 2 in said service bulletin. Compliance required as indicated. To prevent generator electrical lead damage and possible hydraulic fire from cable clamp failure allowing the electrical leads to chafe against hydraulic lines or components, accomplish the following: \n\t1.\tWithin the next 100 hours time-in-service after the effective date of this AD, and every 2,000 hours time-in-service thereafter, inspect the No. 1 generator electrical leads from the pressure feed through fittings (below the floor level) to the engine strut feed through, all generator electrical lead clamps, hydraulic systems, and airframe for routing separation and insulation chafing. Repair or replace any electrical lead, hydraulic line, airframe part or clamp damage as required in accordance with approved maintenance procedures. \n\tThe repetitive 2,000 hour inspection interval may be adjusted by FAA air carrier maintenance inspectors to the nearest scheduled maintenance inspection period. \n\t2.\tIf the generator electrical leads are not routed with at least three (3) inches separation from hydraulic lines or components, provide additional physical protection with aircraft quality Skydrol resistant insulation wrap and clamp as required. \n\t3.\tIf generator electrical leads are routed with at least three (3) inches separation from hydraulic lines or components, no further action is required under this AD. \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). All 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. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis amendment becomes effective May 8, 1979.