Results
2010-26-01: We are adopting a new airworthiness directive (AD) for certain Model 777-200 series airplanes. This AD requires installing a new insulation blanket on the latch beam firewall of each thrust reverser (T/R) half. This AD results from an in-flight shutdown due to an engine fire indication; an under-cowl engine fire was extinguished after landing. The cause of the fire was uncontained failure of the starter in the engine core compartment; the fire progressed into the latch beam cavity and was fueled by oil from a damaged integrated drive generator oil line. We are issuing this AD to prevent a fire from entering the cowl or strut area, which could weaken T/R parts and result in reduced structural integrity of the T/R, possible separation of T/R parts during flight, and consequent damage to the airplane and injury to people or damage to property on the ground.
96-03-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB SF340A and SAAB 340B series airplanes, that requires repetitive operational tests of the valve limit switch of the propeller brake. This amendment also provides for an optional terminating action for the repetitive tests. This amendment is prompted by a report that when the propeller brake was not properly engaged the crew did not receive a "PROP BRAKE" warning due to a faulty valve limit switch. The actions specified by this AD are intended to prevent a valve limit switch from failing to send input to the warning system; absence of a "PROP BRAKE" warning could result in the crew being unaware that the propeller brake is not properly engaged and the propeller may turn without warning.
52-14-02: 52-14-02 CONVAIR: Applies to All 240 Airplanes With Hamilton Standard Propellers Except as Otherwise Indicated. Item IV Also Applies to All 340 Airplanes With Hamilton Standard Propellers. Items I through IV are to be accomplished by means of a progressive modification program to be submitted to and approved by the FAA. This program shall begin no later than August 1, 1952, and shall be completed no later than August 1, 1953. I. The following changes to the electrical circuits are to be accomplished: (NOTE: These changes apply to airplanes which have not been modified since they were manufactured. (See item II for modified airplanes): A. Insulate exposed terminals at unfeathering relay, install insulating separator between throttle microswitches, and secure wires as specified in Convairogram No. 4, dated March 7, 1951. II. To prevent inadvertent actuation of the reversing solenoid valves, the following changes to the electrical circuits are to be accomplished toprotect the reversing solenoid circuits from all other circuits and to protect the reversing solenoid circuits from each other: (NOTE: These changes are known to apply to some aircraft which were modified by operators so that they differ from the originally manufactured configuration. Other airplanes which have design features in the reversing solenoid circuits which are similar to those outlined below, but which are not specifically referred to in this list, should have these points protected in a manner equivalent to that described herein.) A. Modify the following multiple pin connector assemblies as specified in item 2 of attachment A (see 52-13 for attachment A): 1. Connector at wing-fuselage disconnect. 2. Connector at Hamilton Standard Reversing box No. 80340 (covered by change specified in item C.) B. Modify the following terminal strips as specified in item 1 of attachment A: 1. Terminal strip at firewall junction box. 2. Terminal strip in junction box at fuselage Station No. 109. C. Hamilton Standard reversing relay box: Reversing solenoid circuit relay contacts, etc., to be shielded from all other circuits which are energized at any time except when reversing is desired. If reversing relay boxes are used which have separate pin connectors for the reversing solenoid wire and the remaining circuits, it shall not be possible inadvertently to interchange any connectors in the two relay boxes. D. Reversing solenoid circuit wiring: Modify in accordance with item 4 of attachment A. E. Protect the exposed terminals of the secondary throttle lock relays, (if used), as specified in item 3 of attachment A. F. Install insulating separator between throttle microswitches, and secure wires as specified in Convairogram No. 4, dated March 7, 1951. III. Other circuit modification: All airplanes are to be modified to comply with Hamilton Standard Service Bulletin No. 221. IV. Reverse solenoid lock assembly on all airplanes which do not have "lift up" throttles, either (a) Install a warning light system as described in Convair Service Bulletin No. 240-381 except that the system shall be so arranged that it will indicate to the crew when the solenoid lock has just started to move to the open position, or (b) adjust the lock actuating handles so that not less than 1 inch of movement is required before the lock opens. V. Maintenance practices (to be instituted not later than August 1, 1952): A. At each nearest scheduled service to 350 hours: 1. Inspect all points specified in items I and IIB. The inspections of item IIB may be discontinued if the modifications made to the system are of the type described in item 1(a) or 1(b) of attachment A. B. At any time that an electrical fault occurs in a circuit which is carried in the same bundles or the same conduits as the reversing solenoid circuit, representative terminal points in the faulty circuit are to be inspected to determine whether any damage may have occurred within the bundles or conduit. If there is evidence of possible damage, all the wiring involved is to be removed and inspected. Damaged wiring is to be replaced as necessary. C. At each nearest scheduled service to 350 hours, perform an electrical check of the reverse safety switches in the pedestal assembly to assure that the switch is open when the throttles are moved forward out of the reverse position, unless it is shown that failure of any of the reverse safety switches to open will be clearly apparent to the flight crew by reason of improper operation of the propeller control system. Because of the many technical considerations involved, analyses showing that the objective of this revision has been accomplished should be referred to the FAA for engineering evaluation and approval. D. At any time that operations are performed which may affect the relative position of the solenoid lock and throttle switches, but in anyevent at intervals not to exceed 1,500 hours: Check the relationship between the position of the pedestal strikers when they are: (a) In contact with the solenoid latch; (b) at the point where the detent roller contacts the first detent cam, and (c) when the reversing microswitches are actuated. It shall not be possible for the switches to be actuated before the latch and the detent engage the striker and the cam. This determination shall be made by positive measurements rather than observation of engine r.p.m. at which these actions take place. VI. Operating instructions: Comply with item 5 of Attachment A, AD 52-13-02 Lockheed. VII. (NOTE: Propeller governor design changes which are under development and whose purpose is to provide a high pressure hydraulic circuit bypass to safeguard against inadvertent reversing and ability to feather even when the reversing solenoid is energized, are still under consideration and may be the subject of a further directive.)
2022-02-05: The FAA is adopting a new airworthiness directive (AD) for certain Pratt & Whitney (P&W) PW1500G and PW1900G model turbofan engines. This AD was prompted by an analysis of an event involving an International Aero Engines AG (IAE) V2533-A5 model turbofan engine, which experienced an uncontained failure of a high-pressure turbine (HPT) 1st-stage disk that resulted in high-energy debris penetrating the engine cowling. This AD requires removing certain HPT 1st-stage and HPT 2nd-stage disks from service and replacing with parts eligible for installation. The FAA is issuing this AD to address the unsafe condition on these products.
2010-24-07: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: Damage to the lower lateral fittings of the 80VU rack, typically elongated holes, migrated bushes [bushings], and/or missing bolts have been reported in-service. In addition damage to the lower central support fitting (including cracking) has been reported. In the worst case scenario a complete failure of the 80VU fittings in combination with a high load factor or strong vibration could lead to failure of the rack structure and/or computers or rupture/disconnection of the cable harnesses to one or more computers located in the 80VU. This rack contains computers for Flight Controls, Communication and Radio-navigation. These functions are duplicated across other racks but during critical phases of flight the multiple system failures/re-configuration may constitute an unsafe condition. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
96-03-03: This amendment adopts a new airworthiness directive (AD) that applies to certain Fairchild Aircraft SA226 and SA227 series airplanes. This action requires replacing the nuts that attach the power control cable to the lever attach point clevis with nuts that have safety wire holes, safety-wiring the power control cable to the lever attach point clevis, inspecting to assure that the power cable is securely attached to the power control cable bracket, and correcting any attachment problems. Reports of power control cable attaching hardware failure on two of the affected airplanes prompted this action. In one of these instances, the power control cable disconnected from the lever attach point clevis, resulting in engine shutdown. The actions specified by this AD are intended to prevent such power control cable disconnection, which could result in engine shutdown and subsequent loss of control of the airplane.
77-23-08: 77-23-08 HUGHES HELICOPTERS: Amendment 39-3081. Applies to Hughes Models 269A, 269A-1, 269A-2, and 269B helicopters equipped with main rotor blade P/N 269A1125, all blade Serial Numbers; P/N 269A1131, all blade Serial Numbers; P/N 269B1145, all blade Serial Numbers; P/N 269A1131-1, all blade Serial Numbers; P/N 269B1145-1, all blade Serial Numbers; and P/N 269B1145-25, all blade serial numbers. Compliance required as indicated. Small cracks have been discovered on several main rotor blades under the blade root fittings, radiating from the outboard bolt hole on the upper and lower side of the blade. To prevent main rotor blade failure due to extension of these cracks under the fittings, unless already accomplished, accomplish the following: (a) For main rotor blade P/N 269A1125, all Serial Numbers; P/N 269A1131, all Serial Numbers; and P/N 269B1145, Serial Nos. 0001 through 1313 - (1) Prior to the accumulation of 210 hours' time in service for main rotor bladeshaving less than 200 hours' time in service on January 16, 1967, and within the next 10 hours' time in service for main rotor blades having between 200 and 1000 hours' time in service on January 16, 1967, unless already accomplished within the last 390 hours' time in service, and thereafter at periods not to exceed 400 hours' time in service from the date of the last inspection, until a total of 1000 hours' time in service is reached, inspect in accordance with Hughes Service Information Notice No. N-9.2, dated October 3, 1977, or later FAA approved revision. (2) For main rotor blades accumulating a total of 1000 hours' time in service, subsequent to January 16, 1967, and for main rotor blades having 1000 or more hours' time in service on January 16, 1967, within the next 10 hours time in service, unless already accomplished within the last 90 hours' time in service, and thereafter at periods not to exceed 100 hours' time in service from the date of the last inspection until themain rotor blade is retired from service, inspect in accordance with Hughes Service Information Notice No. N-9.2, dated October 3, 1977, or later FAA approved revision. (b) For main rotor blade P/N 269A1131-1, all Serial Numbers; P/N 269B1145-1, all Serial Numbers; P/N 269B1145, Serial Nos. 1314 and subsequent; and P/N 269B1145-25, all Serial Numbers. Prior to the accumulation of 1025 hours' time in service for main rotor blades having less than 1000 hours' time in service on the effective date of this AD, and within the next 25 hours' time in service for main rotor blades having 1,000 or more hours' time in service on the effective date of this AD, unless already accomplished within the last 75 hours' time in service from the date of the last inspection until the main rotor blade is retired from service, inspect in accordance with Hughes Service Information Notice No. N-9.2, dated October 3, 1977, or later FAA approved revision. (c) Cracked blades must be removed before further flight, marked conspicuously to avoid inadvertent return to service, and replaced with new or serviceable used blades in accordance with (1), (2), or (3) below. Blades listed in (1), (2), and (3) are different types. Do not intermix types (1), or (2) or (3) blades. Main rotor blades, either those originally installed or replacement blades, must meet the requirements of this AD and must be retired from service before they exceed their maximum service life of 1366 hours' time in service. (1) Main rotor blade P/N 269B1145 and/or P/N 269B1145-1 and/or P/N 269B1145-25. (2) Main rotor blade P/N 269A1131 and/or P/N 269A1131-1. (3) Main rotor blade P/N 269A1125. This supersedes Amendment 39-642 (33 FR 12085), AD 68-17-07. This amendment becomes effective December 27, 1977.
2010-25-01: We are adopting a new airworthiness directive (AD) for the products listed above. This AD requires changing the emergency open doors procedure by incorporation of a temporary revision into the FAA- approved airplane flight manual (AFM) for all airplanes. This AD also requires replacement of the passenger door retaining bracket with an improved design retaining bracket for certain airplanes. This AD was prompted by several reports of the rear passenger door departing the airplane in flight. We are issuing this AD to change the emergency open doors procedure and retrofit the rear passenger door retaining bracket, which if not corrected could result in the rear passenger door departing the airplane in flight.
2010-06-06: This amendment supersedes an existing airworthiness directive (AD) for MD Helicopters, Inc. (MDHI) model MD-900 helicopters that currently requires applying serial numbers to certain parts, increasing the life limit for various parts, maintaining a previously established life limit for a certain vertical stabilizer control system (VSCS) bellcrank assembly and bellcrank arm, and correcting the part number for the VSCS bellcrank arm. This amendment requires the same actions as the existing AD, except it reduces the life limit of the swashplate spherical slider bearing (slider bearing). It further corrects what was described as a ''bellcrank arm'' life limit in the current AD and correctly describes it as another ''bellcrank assembly'' life limit. This amendment is prompted by two reports of cracks in the slider bearing that occurred well before the previously increased retirement life of 2,030 hours time-in-service (TIS) was reached. The actions specified by this AD are intendedto establish appropriate life limits for various parts, and to prevent fatigue failure of those parts and subsequent loss of control of the helicopter.
96-03-09: This amendment adopts a new airworthiness directive (AD), applicable to certain de Havilland Model DHC-8 series airplanes, that requires modification of a certain battery temperature monitor. This amendment is prompted by reports of failure of the battery temperature monitor, which resulted in smoke in the flight compartment. The actions specified by this AD are intended to prevent failure of the battery monitor, which could result in smoke in the flight compartment.
96-03-10: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-400 series airplanes. This action requires an inspection for damage of the fuel tube located in the forward engine strut, and repair, if necessary; installation of a new support bracket and clamps on the power feeder conduit; and an inspection for proper positioning of the power feeder conduit in each engine strut, and adjustment, if necessary. This amendment is prompted by reports of worn fuel tubes that were caused by the power feeder conduit moving and touching the fuel tube. The actions specified in this AD are intended to prevent wear of the fuel tube, which could result in a fuel leak in the engine strut and a consequent fire hazard.
2021-26-05: The FAA is superseding Airworthiness Directive (AD) 2020-07- 17, which applied to all Saab AB, Support and Services Model SAAB 2000 airplanes. AD 2020-07-17 required revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations. Since the FAA issued AD 2020-07- 17, it has determined that new or more restrictive airworthiness limitations are necessary. This AD retains the requirements of AD 2020- 07-17 and requires revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations, as specified in a European Union Aviation Safety Agency (EASA) AD. The FAA is issuing this AD to address the unsafe condition on these products.
65-22-02: 65-22-02 LOCKHEED: Amdt. 39-139 Part 39 Federal Register September 29, 1965. Applies to Model 1329 Airplanes, Serial Numbers 5001 through 5063. Compliance required within the next 10 hours' time in service after the effective date of this AD, unless already accomplished. To prevent uncontrolled pitch trim actuation, accomplish the inspection and repair as described in Lockheed Alert Service Bulletin No. 329-218, "Horizontal Stabilizer Trim System- Inspection of A-723L, A-723LA, A-723LD, and A-723N Stabilizer Trim Contactors," or perform an equivalent FAA-approved inspection and repair. This directive effective September 29, 1965.
87-09-03: 87-09-03 MESSERSCHMITT-BOLKOW-BLOHM GMBH: Amendment 39-5995. Final copy of priority letter AD issued May 7, 1987. Applies to Model BO-105, BO-105LS, and BK-117 series helicopters certificated in any category with main rotor blade secondary bolt P/N's 105-14101.35, 105-14101.79, 105-14101.11, and 105-141021.05 installed. Compliance is required as indicated, unless already accomplished. To prevent failure of the secondary main rotor blade retention bolt and possible loss of control of the main rotor blade, accomplish the following: (a) Inspect all bolts having more than 100 hours' time in service for cracks at the junction of the head and shank using a magnetic particle inspection or a dye penetrant inspection method prior to further flight. If the dye penetrant inspection method is used, bolts with solid film lubricant coating shall have the coating removed in the affected area (to approximately 5 MM below the bolt head) using Scotch Brite or equivalent. Use MEK or equivalent to clean the surface prior to dye checking. Coated crack-free bolts need not be recoated prior to reinstallation. (b) Inspect bolts with less than 100 hours' time in service in accordance with paragraph (a) prior to accumulating 100 hours' time in service. (c) If cracks are found, remove bolt from service and replace with an airworthy bolt prior to further flight. (d) Conduct the inspection of paragraph (a) at intervals not to exceed 100 hours' time in service since last inspection. (e) Reinstall bolts in accordance with the following instructions. Lubricate crack-free bolts (per MBB Maintenance Manual) with Molykte BR2 (MBB Maintenance Manual item CM- 102) or Aeroshell grease #22 (MBB Maintenance Manual item CM-101) and install per the applicable MBB Maintenance Manual with the following exceptions: (1) Make certain that no lubricant is on the threaded area of the bolt; (2) Before installing the nut, lightly tap the head of the bolt with aplastic or rubber mallet to seat the bolt in the hole; (3) Install nut, and torque to 70-90 inch pounds (8-11 Newton meters); and (4) If required to install cotter pin, back off the nut to next available cotter pin hole, and install cotter pin. (f) Upon request, an alternate means of compliance which provides an equivalent level of safety with the requirements of this AD may be used when approved by the Manager, Aircraft Certification Division, Department of Transportation, Federal Aviation Administration, Fort Worth, Texas 76193-0100. (g) Special flight permits may be issued in accordance with Sections 21.197 and 21.199 to operate helicopters to a base for the accomplishment of inspections required by this AD. This amendment, 39-5995, becomes effective August 26, 1988, as to all persons except those persons to whom it was made immediately effective by priority letter AD No. 87-09-03, issued May 7, 1987, which contained this amendment.
2010-23-26: We are superseding an existing airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: Following the occurrence of cracks on the MLG [main landing gear] rib 5 RH [right-hand] and LH [left-hand] attachment fitting lower flanges, DGAC [Direction Generale de l'Aviation Civile] France AD 2003-318(B) [parallel to part of FAA AD 2006-12- 13] was issued to require repetitive inspections and, as terminating action * * *[.] Subsequently, new cases of cracks were discovered during scheduled maintenance checks by operators of A300B4 and A300-600 type aeroplanes on which the terminating action * * * [was] embodied. This condition, if not corrected, could affect the structural integrity of those aeroplanes. * * * * * We are issuingthis AD to require actions to correct the unsafe condition on these products.
2010-06-03: Drive Shaft
95-26-15: This document publishes in the Federal Register an amendment adopting an airworthiness directive that was sent previously by individual letters to all known U.S. owners and operators of various transport category airplanes equipped with Allied Signal Commercial Avionics Systems CAS-81 TCAS. This amendment is prompted by reports of failure of the audio output of the CAS-81 TCAS. This AD requires a revision to the Airplane Flight Manual to provide the flightcrew with procedures to cycle power to the TCAS processor via the circuit breaker or power bus, and to perform a TCAS functional test to verify proper operation of the TCAS. The actions specified by this AD are intended to ensure that the flightcrew is advised of the potential hazard associated with failure of the audio output of the CAS-81 TCAS, and of the procedures necessary to address it.
72-03-07: 72-03-07 CESSNA: Amdt. 39-1391. Applies to Models 310 (Serial Numbers 310Q0074 through 310Q0425, except 310Q0261, 0271, 0279, 0402, 0409, 0411, 0416, 0420, 0424); Models 340 (Serial Numbers 340-0001 through 340-0009, except 340-0005); Models 401 (Serial Numbers 401B0029 through 401B0204); Models 402 (Serial Numbers 402B0007 through 402B0201); Models 414 (Serial Numbers 414-0052 through 414-0173, except 414-0152, 0153, 0168, 0172); and Models 421 (Serial Numbers 421B0001 through 421B0209, except 421B0144, 0201, 0205, 0208); and all other 300 and 400 series airplanes which may have the original barrel assemblies replaced with new or used assemblies shipped by the manufacturer between January 1, 1970, and December 15, 1971. Compliance: Required as indicated, unless already accomplished. To prevent landing gear failure accomplish the following in accordance with Cessna Service Letter ME71-28 dated December 24, 1971, and Supplement No. 1 dated January 28, 1972, or any equivalent methods approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region: A) Within the next 5 hours' time in service after the effective date of this AD, visually inspect barrel assemblies of the main landing gear upper struts for hydraulic leaks and surface cracks. If such discrepancies are noted, prior to further flight, replace the appropriate barrel assemblies. B) To accomplish the inspection required by Paragraph A the airplane may be flown in accordance with FAR 21.197 to a base where the inspection may be performed. C) Within the next 25 hours' time in service after the effective date of this AD, conduct chemical tests to determine if barrel assemblies of the main landing gear upper struts are composed of dissimilar or incorrect metals. If such discrepancies are noted, prior to further flight, replace with appropriate and correct barrel assemblies. D) Report all defects found in complying with this AD. Such reports must be made in writing and sent to Chief, Engineering and Manufacturing Branch, FAA, Central Region, and should include such items as aircraft serial number, total time in service and nature of defect. (Report approved by the Bureau of the Budget under BOB No. 04-R0174.) This amendment becomes effective February 9, 1972.
92-16-02: 92-16-02 BOEING: Amendment 39-8311. Docket No. 92-NM-119-AD. Supersedes AD 92-06-13, Amendment 39-8193. \n\n\tApplicability: Model 767 series airplanes equipped with General Electric CF6-80C2 engines, certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent inadvertent deployment of a thrust reverser during flight, accomplish the following: \n\n\t(a)\tFor airplanes listed in Boeing Alert Service Bulletin 767-78A0052, Revision 1, dated February 14, 1992: Within 60 days after March 18, 1992 (the effective date of AD 92-06-13, Amendment 39-8193), revise the wiring in certain panels, the wing-body disconnects, and the wing-strut disconnects, in accordance with Boeing Alert Service Bulletin 767-78A0052, Revision 1, dated February 14, 1992. \n\n\t(b)\tFor airplanes listed in Boeing Alert Service Bulletin 767-78A0052, Revision 2, dated May 28, 1992: Within 60 days after the effective date of this AD, revise the wiring in certain panels, the wing-body disconnects, and the wing-strut disconnects, in accordance with Boeing Alert Service Bulletin 767-78A0052, Revision 2, dated May 28, 1992. Procedures that were accomplished previously in accordance with Revision 1 of the service bulletin, and that have not changed in Revision 2 of the service bulletin, need not be repeated. \n\n\t(c)\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 sent 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(d)\tSpecial flight permits may be issued in accordance with FAR 21.197 and21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.\n \n\t(e)\tThe modification shall be done in accordance with Boeing Service Bulletin 767-78A0052, Revision 2, dated May 28, 1992, which includes the following list of effective pages: \n\n\nPage Number\nRevision Level\nDate \n1, 3-4, 7-8, 12-14\n2\nMay 28, 1992 \n2, 5, 10\n1\nFebruary 14, 1992 \n6, 9, 11\nOriginal\t\nDecember 10, 1991 \n\t\t\t\t\t\t\t\nThis 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. The incorporation by reference of Boeing Service Bulletin 767-78A0052, Revision 1, dated February 14, 1992, was approved previously by the Director of the Federal Register as of March 18, 1992 (57 FR 9381, March 18, 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, 1100 L Street NW., Room 8401, Washington, DC. \n\n\t(f)\tThis amendment becomes effective on July 27, 1992.
92-08-10: 92-08-10 AIRBUS: Amendment 39-8221. Docket No. 91-NM-201-AD. Applicability: Model A310 and A300-600 series airplanes; equipped with pilot and copilot seats manufactured by Sogerma-Socea, as listed in Sogerma-Socea Service Bulletin 25- 188, Revision 1, dated July 2, 1991; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced ability of the flight crew to control the airplane, accomplish the following: (a) Within 30 days after the effective date of this AD, accomplish paragraphs (a)(1) and (a)(2) in accordance with Sogerma-Socea Service Bulletin 25-188, Revision 1, dated July 2, 1991: (1) Perform a visual inspection to detect damage to the aft electrical stop switch (switch reference 3 in Figure 1 of the service bulletin). Prior to further flight, replace any damaged switches found, in accordance with the service bulletin. (2) Determine the manufacturer's serial number on the pilot's andcopilot's seats. If the seats have serial numbers that are less than number 261, or if the horizontal actuator has been replaced, accomplish the following: (i) Measure the amount of clearance between the electrical stop and the mechanical stop of the horizontal actuator. (ii) If the clearance is less than 4mm, prior to further flight, adjust the clearance to more than 4mm in accordance with the service bulletin. (iii) If there is no clearance, prior to further flight, replace the horizontal actuator and adjust the clearance to the proper dimension when fitting the new horizontal actuator, in accordance with the service bulletin. (b) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or commentand then send it to the Manager, Standardization Branch, ANM-113. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The inspection shall be done in accordance with Sogerma-Socea Service Bulletin 25-188, Revision 1, dated July 2, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. (e) This amendment becomes effective on May 29, 1992.
2004-13-05: This amendment adopts a new airworthiness directive (AD) for the specified Eurocopter Deutschland (ECD) model helicopters that requires inspecting the vertical fin skin paneling to determine if it was manufactured with the correct wall thickness. This amendment is prompted by a report from the manufacturer that some vertical fins may have been produced with the wrong vertical fin skin thickness. The actions specified by this AD are intended to prevent failure of the vertical fin and subsequent loss of control of the helicopter.
2010-24-10: 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: Damages to the rudder bar locking adjustment tube of a non- reinforced version have been reported to Soci[eacute]t[eacute] Nouvelle (SN) Centrair. This tube had been reinforced in 1984 with a modification. Gliders produced before the introduction of this modification have not been systematically retrofitted. In case of rudder bar locking adjustment tube breaking in flight when adjusting the rudder pedals position, it might interfere with the rudder pedals which could lead to rudder jam or a restricted rudder movement and consequently, to reduced control of the sailplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
93-17-51: 93-17-51 BOEING: Amendment 39-8699. Docket 93-NM-152-AD. \n\n\tApplicability: Model 737-300, -400, and -500 series airplanes, line positions 2288 through 2515 inclusive, certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent loss of the horizontal stabilizer, which could severely affect controllability of the airplane, accomplish the following: \n\n\t(a)\tFor airplanes having line positions 2288 through 2347 inclusive: Within 24 hours after the effective date of this AD, perform a detailed visual inspection of the left and right horizontal stabilizer hinge fitting to verify installation of the retention devices on the inner and outer hinge pins. For the purposes of this AD, retention devices are cotter pins, nuts, washers, and bushing retainers. \n\n\t(b)\tFor airplanes having line positions 2348 through 2515 inclusive: Within 15 days after the effective date of this AD, perform a detailed visual inspection of the leftand right horizontal stabilizer hinge fitting to verify installation of the retention devices on the inner and outer hinge pins. \n\n\t(c)\tIf no part is missing, no further action is required by this AD. \n\n\t(d)\tIf any part is missing, prior to further flight, replace the outer pin, inner pin, and all associated retention devices on the affected side of the horizontal stabilizer, in accordance with the procedures described in Boeing 737 Maintenance Manual. \n\n\t(e)\tWithin 24 hours after completion of the inspection required by this AD, submit a report of any finding(s) of discrepancies to the Manager, Seattle Manufacturing Inspection District Office, ANM-108S, FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington 98055-4056; fax (206) 227-1181. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (44 U.S.C. 3501 et seq.) andhave been assigned OMB Control Number 2120-0056. \n\n\t(f)\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(g)\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(h)\tThis amendment becomes effective on October 14, 1993, to all persons except those persons to whom it was made immediately effective by telegraphic AD T93-17-51, issued on August 27, 1993, whichcontained the requirements of this amendment.
79-25-01: 79-25-01 MCDONNELL DOUGLAS: Amendment 39-3627 as amended by Amendment 39-3702. Applies to McDonnell Douglas Model DC-10-10, -10F, -30, -30F, and -40 series airplanes certificated in all categories. \n\n\tCompliance required as indicated, unless already accomplished. \n\n\tTo preclude contamination from preventing proper operation of the AiResearch Positive Pressure Relief Valves, accomplish the following: \n\n\t(1)\tWithin the next 300 hours time in service after the effective date of this AD, \n\n\t\t(a)\tModify and reidentify the AiResearch P/N 103506-2 cabin pressure relief valves by the addition of an improved filter in accordance with Douglas Service Bulletin 21-87 dated December 3, 1975 (AiResearch Service Bulletin 103506-21-2271, Revision #1, dated May 15, 1979) or AiResearch Service Bulletin 103506-21-2307, dated June 15, 1978 immediately following cleaning of the valve metering system in accordance with Douglas Service Bulletin A21-103, Revision 1, dated August 7, 1978; or(b)\tUnless already accomplished within the preceding 3,000 flight-hours prior to the effective date of this AD, \n\n\t\t\t(i)\tDisassemble, clean and reassemble the AiResearch P/N 103506-2 cabin pressure relief valves metering systems and adjust/test valves in accordance with Douglas Service Bulletin A21-103, Revision 1, dated August 7, 1978; or \n\n\t\t\t(ii)\tReplace the AiResearch P/N 103506-2 cabin positive pressure relief valves with P/N 103506-2 valves having clean metering systems and test valves in accordance with Douglas Service Bulletin A21-103, Revision 1, dated August 7, 1978. \n\n\t(2)\tIf paragraph (1)(a) above is accomplished, at intervals not to exceed 8,000 hours' time in service thereafter, change the relief valve filter elements and functionally check per the maintenance manual. For airplanes with relief valve filters installed per paragraph (1)(a) of this AD, with total filter time in service of 8,000 or more hours on January 7, 1980; within the next 1,000 hours' time in service after January 7, 1980, and thereafter, at intervals not to exceed 8,000 hours' time in service change the relief valve filter elements and functionally check the valve per the maintenance manual. \n\n\t(3)\tUnless paragraph (1)(a) is accomplished, paragraph (1)(b)(i) or (1)(b)(ii) must be accomplished within 3,000 hour intervals since previous accomplishment, and the original delivery design positive pressure relief valve filter elements must be changed per the Maintenance Manual within 1,500 hour intervals since previous accomplishment. \n\n\t(4)\tIf AiResearch P/N 103624-1 or -2 cabin positive pressure relief valves are installed, prior to the accumulation of 8,000 hours' time in service on the relief valve filter elements, and at intervals not to exceed 8,000 hours' time in service thereafter, change the relief valve filter elements and functionally check valve per the maintenance manual. \n\n\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections required by this AD. \n\n\tAlternative inspections, modifications or other actions which provide an equivalent level of safety may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\tAmendment 39-3627 became effective January 7, 1980. \n\n\tThis Amendment 39-3702 becomes effective March 3, 1980, and was effective earlier to all recipients of telegraphic AD T80WE-1 dated January 4, 1980.
89-08-12 R1: 89-08-12 R1 BOEING: Amendment 39-6187 as revised by Amendment 39-6462. Docket No. 89-NM-128-AD. \n\n\tApplicability: Model 737-200, -300, and -400 series airplanes, as listed in Boeing Service Bulletin 737-26-1063, dated May 18, 1989, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo reduce the potential for dispatching an airplane with an inoperative fire/overheat system, accomplish the following: \n\n\tA.\tWithin 10 days after April 24, 1989 (effective date of Amendment 39-6187), inspect the engine fire/overheat detection module to determine the part number. \n\n\t\t1.\tIf part number 10-61096-41, -71, -81, -91, -92, or 10-62061-1, -2, -3, - 11, or -12 is installed, add the following Engine Fire Detection System Test Procedure to the Limitations Section of the FAA-approved Airplane Flight Manual (AFM). This may be accomplished by inserting a copy of this AD in the AFM: \n\n\t\t\ta.\tPrior to engine start, accomplish fire/overheat warning system test. \n\n\t\t\tb.\tAfter engine start, and with the electrical power supply system in the flight configuration, accomplish the fire/overheat warning system test. \n\n\t\t\tc.\tIn the event of an electrical power supply configuration change in flight (e.g., generator failure), perform the fire/overheat warning system test. In the event that this test is unsuccessful, land at the nearest suitable airport. \n\n\t\t2.\tIf part numbers other than those listed in paragraph A.1., above, are installed, no further action is required. \n\n\tB.\tWithin 120 days after the effective date of this amendment, modify the engine fire/overheat detection module, in accordance with Boeing Service Bulletin 737-26-1063 dated May 18, 1989. Once this modification is accomplished, the limitation required by paragraph A.1., above, may be removed from the AFM. \n\n\tC.\tAn alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment, and then send it to the Manager, Seattle Aircraft Certification Office. \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 Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis AD revises AD 89-08-12, Amendment 39-6187. \n\tThis amendment (39-6462, AD 89-08-12 R1) becomes effective on February 12, 1990.