Results
96-23-10: This amendment adopts a new airworthiness directive (AD), applicable to Pratt & Whitney (PW) JT3D series turbofan engines, that requires inspection of steel high pressure compressor (HPC) disks for corrosion, recoating or replating those disks, or replacing those disks as necessary. This amendment is prompted by reports of a failure of a PW JT8D steel HPC disk, which is similar in design to the PW JT3D steel HPC disks. The actions specified by this AD are intended to prevent steel HPC disk failure due to corrosion, which could result in an uncontained engine failure and damage to the aircraft.
90-08-14: 90-08-14 BEECH: Amendment 39-6563. Docket No. 89-CE-26-AD. Applicability: The following airplanes certificated in any category. MODELS SERIAL NUMBERS 95, B95, B95A, D95A, E95 TD-1 through TD-721 95-55, 95-A55, 95-B55 and 95-B55A TC-1 through TC-2456, except TC-350 95-C55, 95-C55A, D55, D55A, E55 and E55A TC-350 and TE-1 through TE-1201 95-B55B (T42A) TF-1 through TF-70 56TC, A56TC TG-1 through TG-94 58, 58A TH-1 through TH-1475 Compliance: Required as indicated in the body of the AD, unless already accomplished. To prevent cracks in the wing forward spar carry-through web structure from propagating to lengths that could compromise the integrity of the wing attachment to the fuselage, accomplish the following: (a) Within the next 100 hours time-in-service (TIS), after the effective date of this AD, or upon the accumulation of 1,500 hours total TIS, whichever occurs later, and thereafter at the intervals specified below,inspect the wing forward spar carry-through web structure in accordance with the instructions in Beech Service Bulletin (SB) No. 2269, Revision 1, dated March 1990. (1) If no cracks are found, repeat the inspection at 500 hour TIS intervals thereafter. (2) For cracks in the bend radius: (i) If the crack length is less than 2.25 inches, prior to further flight stop drill the crack in accordance with the instructions in Beech SB No. 2269, Revision 1, and reinspect for crack progression every 200 hours TIS thereafter. Only one stop drilled crack for the left side and one stop drilled crack for the right side of the web structure are permissible. (ii) If the crack length is greater than 2.25 inches but less than 4.0 inches, prior to further flight stop drill the crack in accordance with the instructions in Beech SB No. 2269, Revision 1, and within the next 100 hours TIS, repair the web structure with the applicable Beech Part Number (P/N) 58-4008 kit as specifiedin the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition. (iii) If the crack length is greater than 4.0 inches, prior to further flight repair the web structure with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft CertificationOffice at the address below for disposition. (3) For cracks in the web face, in the area of the huckbolt fasteners: (i) If the crack length is less than 1.0 inch, reinspect for crack progression every 100 hours TIS thereafter. Only one crack for the left side and one crack for the right side are permissible, provided neither crack exceeds 1.0 inch in length. NOTE 1: Do not stop drill these cracks due to the possibility of damaging the structure behind the web face. (ii) If any crack length is greater than 1.0 inch, or a crack is connecting two fastener holes, within the next 25 hours TIS, repair the web face with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in thesesubsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition. (iii) If any crack passes through two fastener holes and extends beyond the holes for more than 0.5 inch, prior to further flight repair the web face with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition. (4) If cracks are found on the same side of the airplane in both the forward and aft web face, or the bend radii, and any of the cracks are more than 1.0 inch long, prior to further flight repair the webstructure with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition. NOTE 2: If a fuselage skin crack is discovered around the opening for the lower forward carry-through fitting, an external doubler may be required. (b) Airplanes may be flown in accordance with FAR 21.197 to a location where the AD may be accomplished. (c) An alternate method of compliance or adjustment of the initial or repetitive compliance times, which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, Room100, 1801 Airport Road, Wichita, Kansas 67209. NOTE 3: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the document referred to herein upon request to Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or may examine this document at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment (39-6563, AD 90-08-14) becomes effective on May 7, 1990.
2022-19-11: The FAA is adopting a new airworthiness directive (AD) for certain Costruzioni Aeronautiche Tecnam S.P.A. (Tecnam) Model P2006T airplanes. 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. This AD requires performing a detailed visual inspection (DVI) of the aileron control assembly, repairing the aileron control assembly if any crack or damage (including missing paint, nicks, or scrapes) is found, measuring the length of the screws installed on the ceiling cover panel, and replacing the screws if found to be of excessive length. The FAA is issuing this AD to address the unsafe condition on these products.
99-24-17: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires modification of the airplane by coldworking fastener holes at the front and rear wing spars and by installing modified support angles for the lower trailing edge panel of the wing. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent fatigue cracking in the lower spar cap of the wing rear spar and in the lower skin at the wing front spar, just outside the nacelle, on the left-hand and right-hand side of the airplane, which could result in fuel leakage and consequent fire in or around the wing.
2022-17-03: The FAA is adopting a new airworthiness directive (AD) for all The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747-200C, 747-200F, 747-300, 747SR, and 747SP series airplanes. This AD was prompted by significant changes, including new or more restrictive requirements, made to the airworthiness limitations (AWLs) related to fuel tank ignition prevention. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate the latest revision of the AWLs. The FAA is issuing this AD to address the unsafe condition on these products.
99-25-01: This amendment adopts a new airworthiness directive (AD), applicable to all Raytheon Model BAe.125 series 1000A and 1000B, and Model Hawker 1000 series airplanes, that requires inspection of P1 pitot pipes for chafing or damage, and various follow-on actions. This amendment is prompted by reports of P1 pitot pipes chafing against adjacent flight control cables. The actions specified by this AD are intended to prevent a hole in the P1 pitot pipes, which would lead to erroneous input to the instrumentation and warning systems associated with the pilot's instruments.
57-16-03: 57-16-03 DOUGLAS: Applies to All Model DC-7 Series Aircraft Prior to Fuselage No. 697. \n\n\tCompliance required as indicated. \n\n\tNumerous cases have been reported wherein fatigue failures have occurred in the elevator control tab pushrod assembly, P/N 4499117. Investigation reveals that tubes reamed to accommodate the end fittings are unsatisfactory. Also, it has been determined that P/N 4499117 as well as other assemblies, P/N's 2357984, 3593467, and 4335618, have been reamed to permit installation of the end fittings. \n\n\tUnless already accomplished, inspect all the above assemblies within the next 300 hours for fatigue cracks or reaming. Unless disassembled x-ray is believed to be the only reliable means of verifying whether the tubes have been reamed in excess of the limits specified by Douglas. \n\n\tIf cracks are found or reaming exceeds the limits specified in Douglas Service Bulletin DC-7 No. 181 reissued February 8, 1957, the part must be replaced. \n\n\tAll parts reamed within tolerances specified in the Douglas Service Bulletin must be visually reinspected for fatigue cracks at intervals not to exceed 300 hours. When P/N's 4335618 and 4499117 as well as any P/N 3593467, which have been reamed, are replaced with P/N 359347 now manufactured by controlling the swaging of the tube instead of reaming to permit installation of the end fittings, the repeated inspection may be discontinued. \n\n\t(Douglas Service Bulletin DC-7 No. 132 revised November 13, 1956, covers the installation of P/N 3593467 and the associated changes required to make the installation on aircraft originally incorporating P/N 4335618 or P/N 4499117.) \n\n\tThis supersedes AD 56-24-04.
76-10-02: 76-10-02 HUGHES HELICOPTERS: Amendment 39-2611. Applies to Hughes Model 269 series helicopters equipped with P/N 269A7316-3, -5, -7, -9, or -11 cyclic trim control assembly, certificated in all categories, including military TH-55A. Compliance required as indicated. To prevent the restriction of right lateral cyclic control which will result in the loss of right lateral control of the helicopter, accomplish the following: (a) Within the next 25 hours time in service after the effective date of this AD, unless already accomplished; (1) Remove the lateral cyclic trim control assembly from the helicopter in accordance with Hughes 269 Helicopter Basic Handbook of Maintenance Instructions and inspect the structural bond for any slippage between the 269A7142 tube and the 269A7318-1 housing of the 269A7316-3, -5, -7, -9, or -11 Lateral Cyclic Trim Control Assembly. The inspection of the bond consist of applying a 75 to 80 pound load to the 269A7142 tube in accordancewith Hughes Service Information Notice No. N-138, dated April 30, 1976 or later FAA-approved revision. (2) Those assemblies with no slippage may be reinstalled per Hughes 269 Helicopter Basic Handbook of Maintenance Instructions. (3) Those assemblies with slippage, repair per Hughes Service Information Notice No. N-138, dated April 30, 1976 or later FAA-approved revision or replace with a serviceable part prior to further flight. (b) After the effective date of this AD, inspect and repair, if necessary, the lateral cyclic trim control assembly in spare inventory per Hughes Service Information Notice No. N-138, dated April 30, 1976 or later FAA-approved revision on or prior to installation on a helicopter. (c) Equivalent inspections and repair may be approved by the Chief, Aircraft Engineering Division, Western Region. NOTE: For the requirements regarding the listing of compliance and method of compliance with this AD in the rotorcraft maintenance record, see FAR 91.173. (d) Rotorcraft may be flown to a base for accomplishment of the inspections required by this AD per FAR's 21.197 and 21.199. This amendment becomes effective May 21, 1976.
96-23-05: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747 series airplanes, that requires repetitive inspections to detect cracks and/or corrosion of the girt bar support fitting at certain main entry doors (MED); and repair or replacement of the support fitting. This amendment also provides for various terminating actions for the repetitive inspections. This amendment is prompted by reports that, during scheduled deployment tests of main entry door slides, corrosion was found on the floor structure supports for the escape slides of the main deck entry doors on these airplanes. The actions specified by this AD are intended to prevent such corrosion, which could result in separation of the escape slide from the lower door sill during deployment, and subsequently prevent proper operation of the escape slides at the main entry doors during an emergency.
2002-16-22: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 727 series airplanes that have been converted from a passenger-to a cargo-carrying ("freighter") configuration, that requires, among other actions, installation of a fail-safe hinge, redesigned main deck cargo door warning and power control systems, and 9g crash barrier. This amendment is prompted by the FAA's determination that the main deck cargo door hinge is not fail-safe; that certain main deck cargo door control systems do not provide an adequate level of safety; and that the main deck cargo barrier is not structurally adequate during an emergency landing. The actions specified by this AD are intended to prevent structural failure of the main deck cargo door hinge or failure of the cargo door system, which could result in the loss or opening of the cargo door while the airplane is in flight, and consequent rapid decompression of the airplane, including possible loss of flight control or severe structural damage; and to prevent failure of the main deck cargo barrier during an emergency landing, which could injure occupants.