Results
82-19-03: 82-19-03 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE (SNIAS): Amendment 39-4460. Applies to Model SA330J series helicopters certificated in all categories that are equipped with main rotor spindles P/N 330A31.1122.03, 330A31.1122.06 to 330A31.1122.08. Compliance required as indicated. To detect possible cracks and to prevent fretting corrosion and possible cracking of the main rotor head outboard spindle lugs, accomplish the following: (a) Within the next 10 hours' spindle time in service after the effective date of this AD, unless already accomplished, remove the spindles from the helicopter and inspect each main rotor spindle lug bore for cracks using a magnetic particle inspection method. (b) Within 750 hours' spindle time in service after inspection per paragraph (a) of this AD, inspect each lug bore for cracks using a dye penetrant or equivalent inspection method. (c) Within 750 hours' spindle time in service after inspection per paragraph (b) of thisAD, remove the spindle from the helicopter and inspect each lug bore for cracks using a magnetic particle inspection method. (d) Within 1,500 hours' spindle time in service after the effective date of this AD or at next overhaul of the main rotor head, whichever comes first, install spindles that have bushings installed in accordance with Aerospatiale Modification Instructions, AMS 07.43.078, or FAA approved equivalent. The inspections, paragraphs (b) and (c) of this AD, are not required after these bushings are installed. (e) Replace spindles having a cracked bore prior to further flight. Install serviceable spindles. (f) Alternative modifications or other actions which provide an equivalent level of safety may be used when approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, Brussels, Belgium. (Aerospatiale Service Bulletin No. 01.35 pertains to visual inspections and protective finish of the spindle lug bore.) This amendment becomes effective September 27, 1982.
2009-14-03: 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: There have been several Stick Pusher Capstan Shaft failures causing the dormant loss or severe degradation of the stick pusher function. * * * Dormant loss or severe degradation of the stick pusher function could result in reduced controllability of the airplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
2009-14-09: 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: An internal review of design data has shown that the web of the left hand side (LH) stringer 13 near frame 8 might have been improperly trimmed on a few aircraft. If not corrected, possible crack initiations could occur in the upper stringer web, and therefore could impair the structural strength of the adjacent door stop. This latent failure could ultimately lead to the loss of redundancy of the door stops, thereby affecting the structural integrity of the fuselage. * * * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
2008-16-10: 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: The 3 supporting blocks [installed on hydraulic tubes] were made of Teflon, which is unsuitable material for this application. Excessive wear of the blocks was discovered on numerous aircraft, as well as several cases of chafing between the loosely supported tubes. In one case, hydraulic fluid was lost due to fatigue failure of an inadequately supported tube. Loss of hydraulic fluid causes subsequent multiple failures of hydraulically operated systems. Multiple failures of hydraulically operated systems (for the flight air brake actuators, brake system, right thrust reverser, etc.) could result in reduced controllability of the airplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
79-24-06: 79-24-06 MESSERSCHMITT-BOLKOW-BLOHM GmbH: Amendment 39-3622. Applies to all Model BO-105 helicopters, certificated in any category, incorporating tail rotor blades P/N 105-31742 or P/N 105-87161 that have not been modified in accordance with the Accomplishment Instructions, paragraph 2.B., of Messerschmitt-Bolkow-Blohm Alert Service Bulletin No. AB-16, Revision 1, dated December 22, 1978, hereinafter referred to as the Service Bulletin, or an FAA-approved equivalent. Compliance required as indicated. To prevent in-flight loss of tail rotor balance trim weights, and consequent imbalance of the tail rotor blades, accomplished the following: (a) Within the next 10 hours time in service after the effective date of this AD, and thereafter following the last flight of each day upon which the accumulated time in service since the preceding inspection reaches 10 hours, until the modifications required by paragraph (c) of this AD are accomplished, inspect the tail rotor blades for condition in accordance with paragraph 2.A. of the Accomplishment Instructions of the Service Bulletin, or an FAA-approved equivalent. (b) If the inspection required by paragraph (a) of this AD reveals any cracks, or bonding separation that is unacceptable in accordance with the inspection criteria contained in paragraph 2.A. of the Service Bulletin, or an FAA-approved equivalent, before further flight, except that the helicopter may be flown in accordance with FAR 21.197 to a place where the required work can be performed - (1) Replace the affected blade with a serviceable blade of the same part number and continue to comply with the repetitive inspection and modification requirements of paragraph (a) and (c) of this AD; or (2) Replace both blades with blades of improved design, P/N 105-31743 or P/N 105-31744 after which paragraphs (a) and (c) of this AD do not apply. (See paragraph (d) of this AD.) (c) Within the next 100 hours time in service after the effective date ofthis AD, inspect the tail rotor blades in accordance with paragraph 2.A. of the Service Bulletin or an FAA-approved equivalent and before further flight - (1) If the inspection reveals any cracks, or bonding separation that is not acceptable in accordance with the inspection criteria contained in paragraph 2.A. of the Service Bulletin or an FAA-approved equivalent - (i) Replace the affected blade with a serviceable blade of the same part number and modify the tail rotor blade balance provisions of both blades in accordance with paragraph 2.B. of Accomplishment Instructions of the Service Bulletin or an FAA-approved equivalent; or (ii) Replace both blades with blades of improved design, P/N 105-31743 or P/N 105- 31744. (See paragraph (d) of this AD.) (2) If the inspection does not reveal any cracks or bonding separation that is not acceptable in accordance with the inspection criteria contained in paragraph 2.A. of the Service Bulletin or an FAA-approved equivalent, modify the tail rotor blade balance provisions of both blades in accordance with paragraph 2.B. of the Service Bulletin or an FAA-approved equivalent. (d) If tail rotor blades are to be changed in compliance with this AD, both blades must be of the same part number. (e) For the purpose of this AD, an FAA-approved equivalent must be approved by the Chief, Aircraft Certification Staff, AEU-100, FAA, Europe, Africa, and Middle East Region, c/o American Embassy, Brussels, Belgium. This amendment becomes effective December 10, 1979.
2009-14-04: We are adopting a new airworthiness directive (AD) for certain Boeing Model 737-100, -200, -200C, -300, -400, and -500 series airplanes. For certain airplanes, this AD requires deactivating or modifying the wiring to the outboard landing lights, until the wire bundles and electrical connectors have been replaced. For all airplanes, this AD also requires inspecting for any broken, damaged, or missing fairleads, grommets, and wires in the four electrical junction boxes of the main wheel well, and corrective actions if necessary. For certain airplanes, this AD also requires replacing certain wire bundles for the landing lights and fuel shutoff valves, and related investigative, other specified, and corrective actions if necessary. For certain airplanes, this AD also requires replacing of certain electrical connectors and backshell clamps. This AD results from reports of uncommanded engine shutdowns and burned and damaged wire bundles associated with the outboard landing lights and engine fuel shutoff valves. This AD also results from reports of damaged and missing grommets and broken and damaged fairleads in the electrical junction boxes of the main wheel well. We are issuing this AD to prevent a hot short between the outboard landing light and fuel shutoff valve circuits, which could result in an uncommanded engine shutdown. We are also issuing this AD to prevent corrosion of the electrical connectors of the wing rear spars, which could result in short circuits and consequent incorrect functioning of airplane systems needed for safe flight and landing.
62-02-03: 62-02-03\tBOEING: Amdt 389 Part 507 Federal Register January 19, 1962. Applies to Models 707 and 720 Series Airplanes Which Have Not Previously Been Modified In Accordance With Boeing Service Bulletin No. 1359, Dated June 30, 1961, (Service Bulletin No. 1359 Contains a List of Such Airplanes), and to Model 707 Airplanes on Which Retractable Dump Chutes Have Been Installed Per Boeing Service Bulletin No. 1200. \n\tCompliance required as indicated. \n\tIn order to prevent leakage through the secondary seal of the fuel dump chute when fuel is allowed to enter the manifold for any reason, the following modification shall be accomplished within 3,250 hours' time in service after the effective date of this directive: \n\tRemove the secondary fuel seal assembly, Boeing P/N 66-2538, and rebuild using new parts from Boeing kit, P/N 65-9566-1. Upon completion of the rebuilding, change the part number of secondary seal assembly to 69-16258-1. Use new "O" ring seal P/N MS29513-238 when installingsecondary seal assembly, P/N 69-16258-1. \n\t(Boeing Service Bulletin No. 1359, dated June 30, 1961, covers this modification.) \n\tThis directive effective February 20, 1962.
79-24-05: 79-24-05 EMPRESA BRASILEIRA de AERONAUTICA, S.A. (EMBRAER): Amendment 39-3619. Models EMB-110P1 and EMB-110P2, certificated in all categories. Compliance is required within the next 50 hours time in service, unless already accomplished, and thereafter at intervals not to exceed 250 hours time in service. To prevent failure of the flap supports and possible loss of the flaps, accomplish the following: A. With the wing flaps extended, using a 10-power magnifying glass or dye- penetrant method, conduct an inspection of all the flap supports, part numbers listed below, installed on the wing and on the flaps, for cracks in the components near the attachment bolts. Flap Support Part Numbers 4A-2611.46.01 4A-2621.46.01 4A-2611.47.01 4A-2611.48.01 4A-2621.48.01 4A-2116.01.01 4A-2116.02.01 or 4A-2116.02.01N 4A-2216.02.01 or 4A-2216.02.01N 4A-2116.03.01 4A-2216.03.01 If any cracks are found, replace the component before further flight. B. Uponrequest of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, Southern Region, may adjust the inspection interval if the request contains substantiating data to justify the increase for that operator. C. Compliance with the provisions of this AD may be accomplished in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, Southern Region. This amendment is effective November 21, 1979.
78-23-12: 78-23-12 PRATT & WHITNEY AIRCRAFT: Amendment 39-3315. Applies to Pratt & Whitney Aircraft JT8D-1, -1A, -1B, -7, -7A, -7B, -9 and -9A turbofan engine models not incorporating third stage turbine blade retention rivets, P/N 759351 or P/N 618749. Compliance required as indicated unless already accomplished. To prevent third stage turbine blade rivet failure which could result in failure of the low turbine shaft and/or noncontainment of turbine blade and vane debris, accomplish the following: 1. Inspect engines for proper position of the third stage turbine blade in the disk blade slot in accordance with the procedures in the Pratt & Whitney Aircraft JT8D Maintenance Manual, P/N 481671, Section 72-00, Borescope or Radioisotope Inspection of Third Stage Turbine Blade, or equivalent means approved by the Chief, Engineering and Manufacturing Branch, New England Region, prior to the accumulation of 3,000 hours time in service since installed in disk or within the next 600 hours time in service after the effective date of this AD, whichever is later. Engines with no measurable third stage turbine blade mismatch or displacement must be re-inspected every 1,000 hours time in service thereafter. Engines with third stage turbine blade roots displaced axially more than .032 inch relative to the disk rear surface or engines with third stage turbine blade root platform rear surface displaced axially more than .032 inch relative to an adjacent third stage turbine blade root platform rear surface must be removed prior to further flight. Engines with third stage turbine blade mismatch or axial displacement .032 inch or less shall be subject to 300 hour repetitive displacement inspections. Engines with third stage turbine blade mismatch confirmed by an initial displacement inspection and two 300 hour repetitive displacement inspections during which there is no change in blade position, indicating blades were mismatched at last assembly, may then revert to the 1,000 hour repetitive inspection interval. NOTE: a. Mismatch of the blade relative to the disk or the blade root platform rear surface relative to an adjacent blade root platform rear surface is the result of manufacturing tolerance build-up. b. Displacement of the blade is axial movement of the blade relative to its position when originally installed. c. A piece of .032 inch safety wire may be used with the borescope technique as a guide to determine the position of the blade relative to the disk rear face or the blade root platform rear surface relative to an adjacent blade root platform rear surface. d. The rigid borescope and radioisotope inspection methods provide blade displacement information by comparing one blade root platform rear surface to an adjacent blade root platform rear surface. The flexible borescope inspection method provides blade displacement information by comparing the blade root with the disk rear surface. 2. Install by June 30, 1980, improved third stage turbine blade retention rivets, P/N 759351, in accordance with Pratt & Whitney Aircraft Service Bulletin No. 4592, Revision 1, dated August 20, 1976, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA, New England Region, may adjust the 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 increase for that operator. The manufacturer's alert service bulletin identified and described in this directive is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received this document from the manufacturer may obtain copies upon request to Pratt & Whitney Aircraft, Division of United Technologies Corporation, 400 Main Street, East Hartford, Connecticut 06108. This document may also be examined at the Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. 20591. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its Headquarters in Washington, D.C., and at New England Region. This amendment becomes effective December 20, 1978.
2009-14-08: The FAA is adopting a new airworthiness directive (AD) for GE CF6-80C2B5F turbofan engines. This AD requires removing certain part number (P/N) high-pressure compressor rotor (HPCR) stages 11-14 spool/ shafts before they exceed a new, reduced life limit. This AD results from an internal GE audit that compared the life limited parts certification documentation to the airworthiness limitations section (ALS) of the instructions for continuing airworthiness (ICA). We are issuing this AD to prevent HPCR stages 11-14 spool/shaft fatigue cracks caused by exceeding the life limit, which could result in a possible uncontained failure of the HPCR spool/shaft and damage to the airplane.