Results
96-03-05: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model MD-11 series airplanes and Model DC-10-30, DC-10-40, and KC-10A (military) airplanes. For Model MD-11 series airplanes, the AD requires an inspection to determine the serial number of the forward trunnion bolts on the main landing gear (MLG), and rework or replacement of the bolts, if necessary. For Model DC-10-30, DC-10-40, and KC-10A (military) airplanes, the AD requires an inspection for evidence of missing chrome and for corrosion on the chrome surfaces, or verification that the forward trunnion bolts have been chrome plated in a specific manner; and rework or replacement of the bolts, if necessary. This amendment is prompted by reports of chrome flaking on the bearing surface of the trunnion bolts due to improper cleaning of the base material prior to chrome plating. The actions specified by this AD are intended to prevent premature failure of the trunnion bolts and subsequent collapse of the MLG as a result of severe corrosion on the bearing surface and in the mechanical fuse due to chrome flaking.
80-19-14: 80-19-14 PIPER (Ted Smith): Amendment 39-3915. Applies to Aerostar Model 600, 601 and 601P airplanes Serial Nos. 0001 through 0807 certificated in all categories. To prevent possible loss of control due to structural failure of the vertical and horizontal stabilizer aft attach fittings, accomplish the following. (a) Within 100 hours' time in service since the last dye penetrant or visual inspection or within 5 hours' time in service from the effective date of this AD, if no dye penetrant or visual inspection previously accomplished, visually inspect the aft horizontal and vertical stabilizer attach fittings for cracks per part I.A. of Piper Service Bulletin 600-88A dated July 29, 1980. (b) If any crack extends into the web area of the fitting, replace cracked part with like serviceable part and replace fasteners using fasteners specified in Piper Kit 59 600-88 or Kit 59 600-88A. (c) If more than 3 cracks are found in the flange area, replace cracked part withlike serviceable part and replace attach fitting flange fasteners using fasteners specified in Piper Kit 59 600-88 or Kit SB 600-88A. (d) If three (3) or less cracks are found in the flanges in any one fitting and do not extend into the web area of the fitting, within 200 hours' time in service or 180 days whichever occurs sooner from discovery of cracks, replace the fitting and fitting fasteners using fasteners specified in Piper Kit 59600-88 or Kit 59600-88A. (e) Alternative 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. This supersedes Amend. 39-3835 (45 FR 46346), AD 80-12-08. This amendment becomes effective September 25, 1980.
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.