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91-16-08: 91-16-08 BOEING: Amendment 39-7098. Docket No. 90-NM-236-AD.\n \n\tApplicability: Model 737 series airplanes, equipped with a Boeing aft cargo bay auxiliary fuel tank; listed in Boeing Service Bulletin 737-28-1088, dated September 6, 1990; certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo reduce the potential for a fire in the aft cargo compartment due to fuel leaking from the auxiliary fuel tank, accomplish the following: \n\n\tA.\tWithin 90 days after the effective date of this AD, accomplish one of the following: \n\n\t\t1.\tDeactivate the auxiliary fuel system and attach a placard in the flight compartment to indicate that the auxiliary fuel tank is inoperative, in accordance with Boeing Alert Service Bulletin 737-28A1034, Revision 2, dated December 6, 1990, or Boeing Service Bulletin 737-28-1088, dated September 6, 1990; or \n\n\t\t2.\tInstall a check valve and a pressure activated shutoff valve in the auxiliary fuel systemnear the center wing tank in accordance with Boeing Alert Service Bulletin 737-28A1034, Revision 2, dated December 6, 1990; and perform the following inspections of the auxiliary fuel tank support structure in accordance with the inspection procedures below: \n\n\t\t\ta.\tWithin 500 flight cycles after the effective date of this AD, to detect a disbonded or cracked side panel in the auxiliary fuel tank, accomplish one of the following: \n\n\t\t\t\t(1)\tConduct an inspection of the lower sidewall (curved) panels of the auxiliary fuel tanks for disbonding, in accordance with Part I of the Accomplishment Instructions in Boeing Service Bulletin 737-28-1088, dated September 6, 1990. \n\n\t\t\t\t(2)\tPerform a leak check of the auxiliary fuel tanks in accordance with Part III of Boeing Service Bulletin 737-28-1088, dated September 6, 1990. If any fuel leakage is detected, repair prior to further flight in accordance with Part III of the service bulletin. Repeat leak check prior to each flight. \n\n\t\t\tb.\tWithin 12,000 flight cycles after the effective date of this AD, conduct an inspection of the auxiliary fuel tank and support structure in accordance with Part II of Boeing Service Bulletin 737- 28-1088, dated September 6, 1990. Repeat this inspection at intervals not to exceed 12,000 flight cycles. Accomplishment of this inspection constitutes terminating action for the inspection requirements of paragraph A.2.a. of this AD. \n\n\t\t\tc.\tIf a disbonded or cracked panel is detected during the inspections required by paragraphs A.2.a. or A.2.b. of this AD, accomplish one of the following prior to further flight: \n\n\t\t\t\t(1)\tReplace the panel in accordance with Part IV of Boeing Service Bulletin 737-28-1088, dated September 6, 1990; or \n\n\t\t\t\t(2)\tDeactivate the auxiliary fuel tank in accordance with Boeing Alert Service Bulletin 737-28A1034, Revision 2, dated December 6, 1990; or Boeing Service Bulletin 737-28-1088, dated September 6, 1990; or \n\n\t\t\t\t(3)\tRemove the auxiliary fuel tankin accordance with the Boeing 737 Maintenance Manual Subject 28-14-0. \n\n\tNOTE: A deactivated auxiliary fuel tank will require inspections per paragraph A.2. of this AD when reactivated. Auxiliary fuel tanks that are deactivated but remain in an airplane accumulate the same number of flight cycles as the airplane. \n\n\tB.\tAuxiliary fuel tanks currently not installed in an airplane must be inspected in accordance with Boeing Service Bulletin 737-28-1088, dated September 6, 1990, prior to installation in an airplane if they have accumulated more than 4,000 flight cycles. If any cracking or delamination is detected, repair prior to installation in an airplane. \n\n\tC.\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, Seattle 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, Seattle ACO. \n\n\tD.\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 Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tThis amendment (39-7098, AD 91-16-08) becomes effective on September 6, 1991.
2011-23-12: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Several instances of fan blade cracking have been reported. The results of the subsequent technical investigation concluded that the cracking was caused by fan blade flutter at certain engine settings during prolonged ground running. This condition, if not corrected, could affect the integrity of the fan blades, leading to cracking of multiple fan blades and could possibly result in engine failure and release of uncontained high energy debris. We are issuing this AD to prevent fan blade flutter, which could result in an uncontained engine failure and damage to the airplane.
68-26-07: 68-26-07 MCDONNELL DOUGLAS: Amdt. 39-700. Publication of this correction necessary to show correct rework tolerance in Part II (b)(2)(ii). Note: correction is underlined. Applies to Model DC-8 and DC-8F airplanes having certain identified main landing gear forward bogie beam assemblies installed. \n\n\tCompliance required as indicated. \n\n\tSeveral failures in the "critical area" on the bottom of certain main landing gear forward bogie beams have been attributed to fatigue resulting from localized high loads during ground operation. The "critical area" is defined in McDonnell Douglas Letter C1-78-987/WBM. Failures have also occurred in the main landing gear forward bogie beam swivel pin lower lugs due to cracking as a result of corrosion. To preclude further failures of this nature, accomplish the following or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\tPART I ("keyway area") \n\n\t(a)\tPart I of this AD applies to airplanes having a main landing gear forward bogie beam assembly, P/N's 5760590, 5760602, or 5719123, installed (hereinafter referred to as the Bogie Beam Assembly). \n\n\tNOTE: AD 63-27-01, which is superseded, lists parts other than these. These parts may be found either as a component of the main landing gear forward bogie beam and axle assemblies, P/N's 5760631, 5760633, or 5719124, or as a component of main landing gear aft and forward bogie beam assemblies, P/N's 5760630, 5760635, or 5716469. \n\n\t(b)\tRework each Bogie Beam Assembly per that portion of Appendix II, McDonnell Douglas Letter C1-78-228/WBM (or later FAA approved revision) dated February 16, 1965, entitle "IV. Rework Instructions," except that, in lieu of Notes 1, 2, and 3 under 5.A, remove .040 inches of material in a manner prescribed in 5.B following removal of crack indications. Rework may also be accomplished in accordance with procedures approved by the Chief, Aircraft Engineering Division, FAA Western Region. The applicable times are: \n\n\t\t(1)\tFor bogie beams that have never been reworked: \n\n\tWithin the next 400 landings after the effective date of this AD or before the accumulation of a total of 4,000 landings, whichever occurs later; and again within 6,000 landings (for Bogie Beam Assembly P/N's 5760602 and 571923) or 10,000 landings (for Bogie Beam Assembly P/N 5760590) after the initial rework. \n\n\t\t(2)\tFor bogie beams that have undergone one instance of rework in which .040 inches of material have been removed from an uncracked Bogie Beam Assembly or in which .040 inches of material have been removed following removal of crack indications: \n\n\tWithin 6,000 landings (for Bogie Beam Assembly P/N's 5760602 and 5719123) or 10,000 landings (for Bogie Beam Assembly P/N 5760590) after the initial rework, or within the next 400 landings after the effective date of this AD, whichever occurs later. \n\n\t\t(3)\tFor bogie beams that have undergone one or more instances of rework in which less than the specified .040 inches of material have been removed from an uncracked Bogie Beam Assembly or in which less than .040 inches of material have been removed following removal of crack indications: \n\n\tThe applicable times will be as established by the Chief, Aircraft Engineering Division, FAA Western Region, for each specific case. \n\n\t(c)\tIn accomplishing the rework per (b) of Part I of this AD, the total depth of material removed at the centerline of the keyway, including the material removed during all phases of any rework previously performed, shall not exceed .120 inches without prior approval by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(d)\tWithin the next 400 landings after the effective date of this AD or before the accumulation of a total of 4,000 landings, whichever occurs later, install "Kit A" or "Kit E" per that portion of McDonnell Douglas Service Bulletin No. 32-79, Reissue No. 1, Revision No. 3 (or later FAA approved revision) dated January 26, 1968, entitled "2. Accomplish Instructions. Kit A and Kit E, " or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(e)\tLubricate each Bogie Beam Assembly at the bogie beam swivel joint in accordance with the lubrication instructions contained in Figure 2, Sheets 1 through 3, of McDonnell Douglas Service Bulletin No. 32-79, Reissue No. 1, Revision No. 3, dated January 26, 1968, or later FAA approved revision, or by a method approved by the Chief, Aircraft Engineering Division, FAA Western Region, within the next 100 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 100 hours time in service from the last lubrication. Lubrication under this paragraph may be discontinued upon the installation of "Kit A" or "Kit E" in accordance with (d) of Part I or this AD. \n\n\tNOTE: While not part of this AD, Bogie Beam Assemblies reworked per "Kit A" or "Kit E" should still be lubricated between the swivelpin and the lower swivel lug bushing in accordance with good maintenance practice. \n\n\t(f)\tFor the purpose of complying with Part I of this AD: \n\n\t\t(1)\tOperators who have not kept records of the number of landings accumulated by individual Bogie Assemblies shall, in lieu thereof, substitute the total number of landings of the airplanes on which the Bogie Beam Assembly has been installed. \n\n\t\t(2)\tSubject to acceptance by the assigned FAA Maintenance Inspector, the operator may determine the number of landings by dividing each airplane's hours time in service by the operator's fleet average time from take-off to landing for the airplane type. \n\n\t\t(3)\tUpon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may adjust the repetitive intervals for lubrication in (e) of Part I of this AD if the request contains substantiating data to justify the increase for such operator. \n\n\tPART II ("critical area") \n\n\t(a)\tPart II of this AD applies to airplanes having a main landing gear forward bogie beam assembly, P/N's 5760602, 5719123, 5773032, 5773033, or 5777346, installed (hereinafter referred to as the Bogie Beam Assembly). \n\n\tNOTE: AD 64-17-7, which is superseded, does not list all of these parts. These parts may also be found either as a component of main landing gear forward bogie beam and axle assemblies, P/N's 5760631 or 5719124, or as a component of main landing gear aft and forward bogie beam assemblies, P/N's 5760630, 5716469, or 5774700. \n\n\t(b)\tInspect each Bogie Beam Assembly "critical area" in accordance with McDonnell Douglas Letter C1-78-987/WBM, Revision 1, dated July 17, 1964, or later FAA approved revision, or in accordance with an inspection procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region, within the next 400 landings after the effective date of this AD or before the accumulation of a total of 4000 landings after the effective date of this AD or before the accumulation of a total of 4000 landings, whichever occurs later, unless an initial .006 inches of material has been removed (per McDonnell Douglas Letter C1-78-987/WBM) with the last 800 landings prior to the effective date of this AD or unless .040 inches of material has been removed (per McDonnell Douglas Letter C1-78-228/WBM) within the last 2600 landings prior to the effective date of this AD. \n\n\t\t(1)\tIf no cracks are found and the bogie beam has not previously been reworked in the "critical area," perform (i) or (ii) below. \n\n\tIf the bogie beam has previously been reworked in the "critical area," perform (ii) below. \n\n\t\t \t(i)\tRemove .006 inches of material in accordance with McDonnell Douglas Letter \nC1-78-987/WBM, Revision 1, dated July 17, 1964, or later FAA approved revision, or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. (For subsequent inspection and rework see (c) below).(ii)\tRemove .040 inches of material in accordance with Section IV of Appendix I of McDonnell Douglas Letter C1-78-228/WBM dated February 16, 1965, or later FAA approved revision, or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. (For subsequent inspection and rework see (d) below). \n\n\t\t(2)\tIf cracks are found, remove all crack indications in accordance with the rework instructions in Section IV of Appendix I of McDonnell Douglas Letter C1-78-228/WBM, dated February 16, 1965, or later FAA approved revision or a method approved by the Chief, Aircraft Engineering Division, FAA Western Region. After removal of crack indications, perform one of the following: \n\n\t\t\t(i)\tIf no crack indications were evident after removal .003 inches of material and the bogie beam has not previously been reworked in accordance with McDonnell Douglas Letter C1-78-987/WBM or undergone any other rework, complete the rework to a total depth of .006 inches as specified in McDonnell Douglas Letter C1-78-987/WBM, Revision 1, dated July 17, 1964, or later FAA approved revision, or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. (For subsequent inspection and rework, see (c) below). \n\n\t\t\t(ii)\tIf no crack indications were evident after removal of .003 inches of material, continue to rework by removing an additional .040 inches of material below the bottom of the crack in accordance with Section IV of Appendix I of McDonnell Douglas Letter C1-78-228/WBM, or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. (For subsequent inspection and rework, see (d) below). \n\n\t\t\t(iii)\tIf crack indications were evident after removal of .003 inches of material but no crack indications were evident after removal of up to .030 inches of material, continue to rework by removing an additional .040 inches of material below the bottom of the crack in accordance with Section IV of Appendix I of McDonnell Douglas Letter C1-78-228/WBM or perform a rework approved by the Chief, Aircraft Engineering Division, FAA Western Region. (For subsequent inspection and rework, see (d) below). \n\n\t(c)\tReinspect each Bogie Beam Assembly "critical area" per McDonnell Douglas Letter C1-78-987/WBM, Revision 1, dated July 17, 1964, or later FAA Approved revision, or in accordance with a method approved by the Chief, Aircraft Engineering Division, FAA Western Region, within the next 400 landings after the effective date of this AD, or within the next 1,200 landings following removal of .006 inches of material per McDonnell Douglas Letter C1-78-987/WBM, whichever occurs later. If cracks are found, remove the cracks as specified per (b)(2)(iii). If no cracks are found, accomplish rework per (b)(2)(ii). \n\n\t(d)\tReinspect each Bogie Beam Assembly "critical area" per McDonnell Douglas Letter C1-78-987/WBM, Revision 1, dated July 17, 1964, or later FAA approved revision, or in accordancewith a method approved by the Chief, Aircraft Engineering Division, FAA Western Region, within the next 400 landings after the effective date of this AD, or within the next 3,000 landings following rework per McDonnell Douglas Letter C1-78-228/WBM, whichever occurs later, and reinspect at intervals not to exceed 3,000 landings following any rework performed in accordance with this paragraph. If cracks are found, remove the cracks per (b)(2)(ii) or (b)(2)(iii). If no cracks are found, remove .040 inches of material per (b)(2)(ii). \n\n\t(See paragraph (e) of Part II of this AD for bogie beam limitations.) \n\n\t(e)\tThe following limitations apply to bogie beams reworked in accordance with (b), (c) and (d) above: \n\n\t\t(1)\tThe bogie beam must not be reworked more than 3 times in accordance with Section IV or Appendix I of McDonnell Douglas Letter C1-78-228/WBM or an FAA approved equivalent rework. \n\n\t\t(2)\tThe bogie beam must not be used beyond a total of 3,000 landings after the thirdrework in accordance with Section IV of Appendix I of McDonnell Douglas letter C1-78-228/WBM or an FAA approved equivalent rework. \n\n\t(f)\tA bogie beam assembly specified in this AD may be replaced with one of the following: \n\n\t\t(1)\tA new bogie beam approved for installation on an airplane with an approved takeoff weight equal to or in excess of the takeoff weight of the airplane receiving the replacement bogie beam. The following table lists the airplane approved takeoff weights with the appropriate bogie beam assembly part numbers. \n\n\n\nT.O. Weight\n(1000 lbs.)\nBogie Beam \nAssembly P/N\n\n\n276\n5760590\n\n5773031\n300\n5760602\n\n5773032\n315\n5719123\n\n5773033\n325\n5777346\n\n5778946\n\n\t\t(2)\tA bogie beam assembly inspected and reworked in accordance with this AD that has not exceeded the landing life limits of this AD. \n\n\t\t(3)\tA bogie beam assembly approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(g)\tFor the purpose of complying with PartII of this AD: \n\n\t\t(1)\tOperators who have not kept records of the number of landings accumulated by individual Bogie Beam Assemblies shall, in lieu thereof, substitute the total number of landings of the airplanes on which the Bogie Beam Assembly has been installed. \n\n\t\t(2)\tSubject to acceptance by the assigned FAA Maintenance Inspector, the operator may determine the number of landings by dividing each airplane's hours time in service by the operator's fleet average time from take-off to landing for the airplane type. \n\nThis amendment becomes effective February 1, 1969.
75-14-06: 75-14-06 MCDONNELL DOUGLAS: Amendment 39-2250. Applies to all DC-9-10, -20, -30, -40 Series, and Military C-9A, C-9B, and VC-9C Series airplanes, manufacturer's fuselage numbers 1 through 765. Fuselage numbers subsequent to 765 are exempt, provided the thrust reverser driver links and rigging have not been altered subsequent to delivery by the manufacturer. \n\n\tCompliance required as indicated. \n\n\tTo prevent unwanted deployment of the engine reverser doors, accomplish the following inspections, checks, repairs and replacements on aircraft whose thrust reversers have in excess of 8000 reverser cycles, on or after the effective date of this AD. \n\n\tNOTE: Thrust reversers whose cycles cannot be determined must be considered to have in excess of 8000 cycles. \n\n\t(a)\tWithin 72 hours after the effective date of this AD, inspect or check the upper and lower thrust reverser door fairings on each day on which the airplane is operated, to verify the alignment of the door fairing.(1)\tIf the upper door extends more than one-quarter inch higher than the fixed fairing, deploy the reverser and visually inspect the driver links for separation. Replace failed upper driver links prior to further flight. \n\n\t\t(2)\tIf a driver link is broken on the lower door, it will gap similar to the upper door while hanging on the lock latch and a reverse unlock indication will be present in the cockpit. Inspect the lower door driver link for separation. If the link is failed, replace failed lower link, or deactivate the reverser door per DC-9 Maintenance Manual, Chapter 78, prior to further flight. \n\n\tNOTE: McDonnell Douglas Telegram DC-9-COM-10-JER, dated April 26, 1975, covers this subject. \n\n\t(b)\tWithin 1600 additional hours of flight operation or six months after the effective date of this airworthiness directive, whichever occurs first, inspect by dye penetrant all thrust reverser driver links, P/N 5958782-1 and/or 5958782-501, in accordance with the procedures described in McDonnell Douglas Service Bulletin 78-36, Revision 1, dated June 19, 1975, or later FAA-approved revisions. All links found acceptable for proper flange thickness (undercutting) and exhibiting no evidence of corrosion or cracks as described in paragraph 2C of S.B. 78-36, Revision 1, must be checked for proper rigging in accordance with procedures outlined in paragraphs 2D through paragraph 2J of S.B. 78-36, Revision 1, prior to returning the aircraft to service. \n\n\t(c)\tLinks may be returned to service, if within the limits specified herein, provided that, at intervals not to exceed 800 cycles in service thereafter from the last inspection, a dye penetrant inspection is performed per the procedures of paragraph (b), above. Serviceable links shall include: \n\n\t\t(1)\tFlange thickness of .085 - .095 inches, with or without evidence of corrosion, pits or cracks as determined by paragraph 2C(3) and 2C(4) of McDonnell Douglas Service Bulletin, Revision 1, dated June 19, 1975, or laterapproved revisions. \n\n\t\t(2)\tFlange thickness of .090 inches or more with properly blended out cracks and with or without corrosion pits as described per paragraph 2C(4) and 2C(5) of the above referenced Service Bulletin. \n\n\t\t(3)\tFlange thickness of .095 or more with corrosion pits as described in paragraph 2C(4) of the above referenced Service Bulletin. \n\n\t\t(4)\tLinks checked for crown height and found to be over .243 inches deflection per paragraph 2F(1) of the above referenced Service Bulletin. \n\n\tLinks determined to be unserviceable by paragraph 2C(2) must be scrapped. \n\n\tNOTE: Reverser links determined to be in categories (1) through (4), above, must be permanently identified as described in McDonnell Douglas Service Bulletin 78-36, Revision 1, or later FAA-approved revisions to preclude reinstallation as a normal replacement. \n\n\t(d)\tFurther action under this AD may be discontinued as to that airplane when the following conditions are met. \n\n\t\t(1)\tAll links have beeninspected and no evidence of cracks, corrosion pits or undercutting exists and crown height is found to be in limits when properly rigged. \n\n\t\t(2)\tLinks have been replaced, if necessary, and rigged in accordance with the procedures outlined in the McDonnell Douglas Service Bulletin 78-36, Revision 1, or later FAA-approved revisions. \n\n\t(e)\tEquivalent inspections and installations may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiating data. \n\n\tThis supersedes the telegraphic AD adopted June 4, 1975. \n\n\tThis amendment becomes effective July 7, 1975.
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.