Results
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.
2009-13-03: We are adopting a new airworthiness directive (AD) for certain Boeing Model 747-400 and -400F series airplanes. This AD requires modifying certain thrust reverser control system wiring to the flap control unit (FCU). This AD results from a report of automatic retraction of the leading edge flaps during takeoff due to indications transmitted to the FCU from the thrust reverser control system. We are issuing this AD to prevent automatic retraction of the leading edge flaps during takeoff, which could result in reduced climb performance and consequent collision with terrain and obstacles or forced landing of the airplane.
80-07-04: 80-07-04 HILLER HELICOPTERS: Amendment 39-3722. Applies to Hiller Model UH-12D, UH-12E and UH-12 (4 place) helicopters including military models UH-23D, OH-23G, H-23F and turbine-powered models, equipped with main rotor blade assembly, Parsons P/N 2253-1101-03 or 2253-1101-04 certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent fatigue failure of the main rotor blade anti-node bars accomplish the following: a. Before the accumulation of 2500 hours' time in service, or within 10 additional hours' time in service on main rotor blades with 2500 or more hours' time in service on the effective date of this AD, whichever occurs sooner, inspect the threaded ends of the anti-node bar per instructions specified in Part II accomplishment instructions of Hiller Aviation Service Bulletin No. 51-5 dated January 22, 1980, to determine whether the threads are rolled or cut. b. If the anti-node bar thread inspection ofparagraph (a) of this AD reveals that the bar has cut threads, remove the anti-node bar from service and replace with like serviceable part in accordance with paragraph 4.55 of the UH-12E Structural Repair Manual. c. If the inspection of paragraph (a) of this AD does not provide satisfactory evidence that the threads are either cut or rolled, remove the rod from service and replace with a like serviceable part. d. If the inspection of paragraph (a) of this AD reveals that the bar has rolled threads and the total time on the bar is less than 6670 hours, reinstall the bar in accordance with the instructions of paragraph 4.55 of the UH-12E Structural Repair Manual. Note: Caution; Use extreme care in reinstalling the anti-node bar assembly to ensure that the screws attaching the anchor nut to the anti-node bar are not sheared during insertion. Hand pressure is the maximum force allowed in installing the anti-node bar. e. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate rotorcraft to a base for the accomplishment of inspections required by this AD. f. 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 amendment becomes effective March 31, 1980.
80-02-03 R1: 80-02-03 R1 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-3661 as amended by amendment 39-3964. Applies to Lockheed-California Company L-1011-385 series airplanes certificated in all categories. To preclude possible failure of the main landing gear forward and aft trunnion pins P/Ns 1523068(-103), 1535918(-101) and 1523069(-101), respectively, perform the following: Compliance required as indicated. (a) Within the next 48 calendar days after January 17, 1980, unless already accomplished; (1) Visually inspect the main landing gear forward and aft trunnion pins in accordance with the accomplishment instructions of paragraph 2B of Lockheed-California Company Alert Service Bulletin 093-32-A167, Revision 1, March 6, 1980. If a crack(s) or fracture is found, replace the trunnion pin(s) prior to further aircraft operation, or; (2) Visually inspect the main landing gear forward and aft trunnion pins with retainers removed in accordance with paragraph 2A of Lockheed-California Company Service Bulletin 093-32-167, Revision 1, dated September 18, 1980 and install MLG forward and aft trunnion pin retainers per paragraphs 2B and 2C of Lockheed-California Service Bulletin 093-32- 167, Revision 1, dated September 18, 1980. If a crack(s) or fracture is found, replace the trunnion pin(s) prior to further aircraft operation, or; (3) Remove the MLG forward and aft trunnion pins and retainers, if installed and inspect by visual and magnetic particle methods, and reidentify per paragraphs 2A, 2B, 2C, and 2D of Lockheed-California Service Bulletin 093-32-169, Revision 1 dated September 18, 1980. If a crack(s) or fracture is found, replace the trunnion pin(s) prior to further flight. Defects in chrome plating only may be repaired per paragraphs 2E and 2F of Lockheed-California Service Bulletin 093-32-169, Revision 1, dated September 18, 1980. NOTE 1: The repetitive inspection requirements of paragraph (b) of this AD are not applicable to trunnion pins inspected per paragraph (a)(3). (b) Repeat the visual inspections of paragraph (a) of this AD as specified: (1) Within 50 hours' time in service since the last inspection conducted per paragraph (a)(1) of this AD and thereafter at intervals not to exceed 50 hours' additional time in service, repeat the visual inspections required by paragraph (a)(1) of this AD. (2) Within 1500 hours' time in service since the inspection and retainer installation accomplished per paragraph (a)(2) of this AD and thereafter at intervals not to exceed 1500 hours' additional time in service, repeat the visual inspections required by paragraph (a)(2) of this AD. (c) Once per each day in which the aircraft is operated following the accomplishment of the inspections of paragraph (a)(1) above, and excluding the days on which the inspection of paragraph (b), above, is accomplished, conduct visual check of the main landing gear forward and aft trunnion pins in accordance with the accomplishment instructions of paragraph 2A of Lockheed-California Company Alert Service Bulletin 093-32-A167, Revision 1. If an obvious migration of either or both of the pins exists relative to the normal installation configuration, perform the visual inspections of paragraph (a), above. If a crack(s) or fracture is found, replace the pin(s) prior to further aircraft operation. NOTE 2: The daily check requirements are not applicable if paragraph (a)(2) or (a)(3) of this AD is accomplished. (d) Alternate checks, inspections or other actions which provide equivalent level of safety may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. Amendment 39-3661 became effective January 17, 1980. This amendment 39-3964 becomes effective November 6, 1980.
83-07-07: 83-07-07 BRITISH AEROSPACE, AIRCRAFT GROUP, SCOTTISH DIVISION: Amendment 39-4605. Applies to Model HP.137 Jetstream MK-1 and Jetstream Series 200 airplanes certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent possible jamming of the rudder, accomplish the following: a) Within 100 hours time-in-service, after the effective date of this AD: 1) Remove the access panel located between the two bottom ribs in the rudder. Using a light and mirror visually inspect the two rudder skin panel stiffeners for damage in accordance with British Aerospace, Aircraft Group, Jetstream Service Bulletin (SB) No. 8/2, Revision 1, dated November 10, 1982, hereinafter referred to as the SB. Ensure that any debris from damaged stiffeners is removed from the rudder. 2) Fit a fabric patch (debris net) on the undersurface of Rib No. B and provide effective water drainage in accordance with the SB. b) Incorporate British Aerospace Modification 5210 to the rudder assembly as specified in SB No. 8/2, by December 31, 1985. c) Aircraft may be flown in accordance with FAR 21.197 to a location where this Airworthiness Directive (AD) can be accomplished. d) An equivalent method of compliance with this AD, if used, must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on April 11, 1983.
2009-14-02: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 747 airplanes. That AD currently requires repetitive inspections to detect evidence of wear damage in the area at the interface between the vertical stabilizer seal and fuselage skin, and corrective actions, if necessary. The existing AD also provides for an optional terminating action for the repetitive inspections. For all Boeing Model 747-100, 747-100B, 747- 100B SUD, 747-200B, 747-200C, 747-200F, 747-300, 747-400, 747-400D, 747-400F, 747SR, and 747SP series airplanes, this new AD requires repetitive inspections for wear damage and cracks of the fuselage skin in the interface area of the vertical stabilizer seal and fuselage skin, a detailed inspection for wear damage and cracks of the surface of any skin repair doubler in the area, and corrective actions if necessary. For airplanes on which the fuselage skin has been blended to remove wear damage, this new AD requires repetitive external detailed inspections or high frequency eddy current inspections for cracks of the blended area of the fuselage skin, and corrective actions if necessary. This AD results from reports of wear damage on airplanes with fewer than 8,000 total flight cycles. In addition, there have been three reports of skin wear damage on airplanes that applied Boeing Material Specifications 10-86 Teflon-filled coating (terminating action per the existing AD). We are issuing this AD to detect and correct wear damage and cracks of the fuselage skin in the interface area of the vertical stabilizer seal and fuselage skin in sections 46 and 48, which could cause in-flight depressurization of the airplane. \n\n\nDATES: This AD becomes effective August 3, 2009. \n\tThe Director of the Federal Register approved the incorporation by reference of a certain publication listed in the AD as of August 3, 2009. \n\tOn February 10, 2003 (68 FR 476, January 6, 2003), the Director of the Federal Register approved the incorporation by reference of Boeing Alert Service Bulletin 747-53A2478, dated February 7, 2002.
2009-14-01: The FAA is superseding an existing airworthiness directive (AD) for Turbomeca S.A. Arrius 2F turboshaft engines that have not incorporated Turbomeca Modification Tf75. That AD currently requires replacing the O-ring on the check valve piston in the lubrication unit at repetitive intervals. This AD requires the same repetitive replacements and would require incorporating Modification Tf75 as terminating action to the repetitive O-ring replacements. Modification Tf75 replaces the check valve piston with a piston design not requiring an O-ring. This AD results from the European Aviation Safety Agency (EASA) and Turbomeca S.A. mandating the incorporation of Modification Tf75. We are issuing this AD to prevent an uncommanded in-flight shutdown of the engine, which could result in a forced autorotation landing and damage to the helicopter.
63-13-04: 63-13-04 MARTIN: Amdt. 575 Part 507 Federal Register June 13, 1963. Applies to All Model 404 Aircraft With Main Landing Gear Piston/Axle Assembly P/N 202SD82087 (Menasco P/N 511017). Compliance required as indicated. There have been cracks found in the upper portion of the 0.250 inch radius at the junction of the axle and the piston terminal (approximately 2.4 inches from the centerline of the main gear) of main landing gear piston/axle assemblies. As this condition is likely to develop in other such aircraft, accomplish the following: (a) Thoroughly clean the radius area around the periphery of the section on the left and right sides of both main landing gear in the region of the junction of the axle to the piston terminal, removing all grease and dirt. (b) Visually inspect for cracks all main landing gear assemblies with 5,000 or more hours' time in service on the effective date of this AD, within the next 150 hours' time in service and thereafter within each 150 hours' time in service from the last inspection. (c) Visually inspect for cracks all main landing gear assemblies with less than 5,000 hours' time in service on the effective date of this AD, prior to 5,150 hours' time in service and thereafter within each 150 hours' time in service from the last inspection. (d) At each fifth visual inspection i.e., each 750 hours' time in service after the initial inspection, inspect for cracks all main landing gear assemblies using dye penetrant, Zyglo or FAA approved equivalent, in conjunction with at least a 10-power magnifying glass. (e) If a crack is found, the following rework or replacement is required: (1) Parts with a crack that may be eliminated by removal of not more than 0.050 square inch of material without exceeding a material depth of 0.020 inch shall be reworked prior to further flight except that a ferry flight may be made in accordance with the provisions of CAR 1.76 prior to rework. In eliminating the crack, blend the reworked section into the radius, removing all nicks and dents. Polish the affected area to an RMS 32 finish. (2) Parts with a crack that cannot be removed as set forth in (e)(1) shall be replaced prior to further flight with parts of the same part number (Menasco P/N 511017) or an FAA approved equivalent. (f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. This directive effective July 15, 1963.
84-06-05: 84-06-05 PARACHUTE HARNESSES MANUFACTURED UNDER TSO-C23B: Amendment 39-4839. Applies to all of the following manufacturer's parachute harnesses using the 3-Ring Inc. release system: Adventure Loft, Inc. Aerotech Inc. Altitude Shop The Annex Bill Coe Bureau of Land Management The Chute Shop Rec Whse Flying High Embury Sky Systems LLB Enterprise G.Q. Security Parachutes, Inc. National Parachute Supply, Inc McLaughlin Para Center North American Aerodynamics Nytech Inc. Parachute Associates, Inc. Para Flite, Inc. Para Gear Equipment Co. Para Phernalia Para Wing SA Pioneer Parachute Co., Inc. Ray Hara C/O Cazer Para Loft Reltny Parachute Service Relative Workshop, Inc Rogersport Sky Supplies SSK Industries Inc. Strong Enterprise, Inc. Thomas Sports Equipment LTD Weckbecker Westgaard Parachute Enterprise Westway Parachuting Enterprise Compliance: Required prior to next jump, unless already accomplished. To prevent possible fouling of a reserve parachute canopy by a main canopy which cannot be separated from the harness, accomplish the following: (a) Visually inspect the harness to determine whether or not it incorporates a 3-Ring Inc. release system, and if incorporated, whether or not the large ring of this assembly is identified by either Part Number RW-1-82 or RW-1-83. (b) If either finding of the inspection in paragraph (a) is negative, no further action is required. (c) If both findings of the inspection in paragraph (a) are positive, replace or test the large rings in accordance with 3-Ring Safety Bulletin No. 3 dated February 15, 1984. (1) Replace defective rings or identify acceptable rings in accordance with the instructions in this bulletin. (d) Replacement and testing of the large rings must be accomplished by an FAA certified Parachute Rigger, an FAA certificated Parachute Loft or the manufacturer of the parachute harness involved. (e) An equivalent means of compliancewith this AD may be used if approved by the Manager, Atlanta Aircraft Certification Office, 1075 Inner Loop Road, College Park, Georgia 30337. This amendment becomes effective on April 9, 1984.
2009-13-08: We are adopting a new airworthiness directive (AD) for all McDonnell Douglas Model MD-90-30 airplanes. This AD requires repetitive inspections for cracks of the upper aft skin panels on the horizontal stabilizer, and related investigative and corrective actions if necessary. This AD results from a report of cracks found in the aft skin panels on the upper right side of the horizontal stabilizer at the aft inboard corner. We are issuing this AD to detect and correct cracks in the fail-safe structure that may not be able to sustain limit load, which could result in the loss of overall structural integrity of the horizontal stabilizer.
2009-04-18: The FAA is adopting a new airworthiness directive (AD) for PW models JT9D-7, -7A, -7AH, -7H, -7F, and -7J turbofan engines. This AD requires initial and repetitive borescope inspections of the 2nd stage high-pressure turbine (HPT) rotor and stator assembly. This AD results from an uncontained failure of a 2nd stage HPT rotor disk that caused the engine to separate from the airplane. We are issuing this AD to prevent failure of the 2nd stage HPT rotor disk, which could result in uncontained engine failure, damage to the airplane, and the engine separating from the airplane.
61-11-02: 61-11-02 BRANTLY: Amdt. 292 Part 507 Federal Register May 30, 1961. Applies to All Model B-2 Helicopters Prior to Serial No. 95. Compliance required within the next 25 hours' time in service after effective date of this directive. Severe oil leaks have been reported recently from failure of the welded seam of the mixture heater oil jacket, Lycoming P/N 72396, on the Lycoming VO-360 engines in Brantly B-2 helicopters. As this condition introduces a fire hazard as well as the probability of engine failure from oil starvation, the following modifications must be accomplished. (a) Disconnect the oil system hose assemblies from the inlet and outlet ports of the mixture heater oil jacket. Join these assemblies with AN 815-8D union, thus by-passing the mixture heater oil jacket. Support the joined hose assemblies with AN 742-D-14C hose clamps at the right rear starter mount bolt and the left rear transmission mount tube. Remove the mixture heater oil jacket, clean thoroughly, apply zinc chromate to the oil passage cavity, and reinstall with the oil ports capped with AN 928-8 cap assemblies. (b) Modify engine cooling boxes in accordance with Brantly Service Bulletin No. 12. This modification includes adding covers over the rear openings in the right hand cooling box, deleting the cylinder barrel baffle assemblies from around No. 2 and No. 4, cylinders, making a 1 3/4-inch diameter cutout in the bottom of the left hand cooling box under the No. 3 cylinder, and relocating the cylinder head temperature thermocouple from the No. 4 cylinder to the No. 3 cylinder. (Brantly Service Bulletin No. 12 covers these modifications.) This directive effective May 30, 1961.
2009-13-05: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: During a SOCATA flight test, it has been noted some difficulties for the pilot to release oxygen. After investigation it has been found that, due to the design of the oxygen generator release pin, one of the mask's lanyard linked to the pin can be jammed when it is pulled by a pilot or a passenger. This condition, if not corrected, would lead, in case of an emergency procedure due to decompression to a risk of generator fault with subsequent lack of oxygen on crew and/or passenger. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
82-06-01: 82-06-01 CANADAIR: Amendment 39-4337. Applies to Canadair Model CL-600-1A11 airplanes, Serial Numbers (S/N) 1005-1008 and 1010-1018, certificated in all categories. Compliance required as indicated. To ensure structural integrity of the windshield lower sill members, accomplish the following, unless already accomplished: 1. Prior to the accumulation of 600 hours time in service or within 25 hours time in service after the effective date of this AD, whichever occurs later, reinforce the sill members as described in paragraph 2 of Canadair Alert Service Bulletin A600-0035 dated May 28, 1981. 2. Airplanes with Canadair Modification Summary 600-680, Issue NC, installed are exempt from this AD. 3. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108. 4. Airplanes may beflown in accordance with FAR 21.197 to a maintenance base for the accomplishment of the modification required by this AD. The manufacturer's specifications and procedures identified and described in this directive are 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 these documents from the manufacturer may obtain copies upon request to the addresses listed above. These documents may also be examined at the FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective March 16, 1982.
2009-13-04: 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: Excessive wear on a guide pin of a power lever has been detected during inspections. The total loss of the pin could cause loss of the flight idle stop and lead to inadvertent activation of the beta mode in flight. The inadvertent activation of beta mode in flight can result in loss of control of the airplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
88-12-01: 88-12-01 GENERAL ELECTRIC: Amendment 39-5933. Applies to General Electric (GE) CF6-6 turbofan engines. Compliance is required as indicated, unless already accomplished. To prevent failure of high pressure compressor (HPC) rear shafts, and/or compressor discharge pressure (CDP) seal supports, accomplish the following: (a) Remove from service, HPC rear shafts, Part Numbers 9021M68G02 through 9021M68G04 inclusive; 9021M68G07 through 9021M68G09 inclusive; 9021M68G12; and 9021M68G13; and 1380M55G01 through 1380M55G06 inclusive, as follows: (1) Remove from service HPC rear shafts which have accumulated 17,600 or more cycles since new on the effective date of this AD, within the next 400 cycles in service. (2) Remove from service HPC rear shafts which have accumulated less than 17,600 cycles since new on the effective date of this AD, at or prior to accumulating 18,000 cycles since new. (b) Remove from service, CDP seal supports, Part Numbers 9686M62P02 through 9686M62P08 inclusive; and 9686M62P10, as follows: (1) Remove from service CDP seal supports which have accumulated 17,600 or more cycles since new on the effective date of this AD, within the next 400 cycles in service. (2) Remove from service CDP seal supports which have accumulated less than 17,600 cycles since new on the effective date of this AD at or prior to accumulating 18,000 cycles since new. NOTE: This action establishes new life limits of 18,000 service cycles for the parts noted in (a) and (b) above. The new limits are published in Chapter 5 of the CF6-6 Maintenance and Shop manuals. (c) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD may be accomplished. (d) Upon submission of substantiating data by an owner or operator through an FAA airworthiness inspector, the Manager, Engine Certification Office, New England Region, may adjust the compliance schedules specified in this AD.This amendment, 39-5933, becomes effective June 6, 1988.
69-21-02: 69-21-02 GODFREY: Amdt. 39-858. Applies to Godfrey Cabin Superchargers Type 15, Marks 6, 9 and 14, installed on, but not necessarily limited to British Aircraft Corporation, Viscount Models 744, 745D, and 810; Armstrong Whitworth Argosy AW-650; Fokker F-27, Marks 100 and 300; Fairchild Hiller F-27 and FH-227 all series; Grumman Model G-159 and Nihon YS-11 all series airplanes. Compliance required as indicated. To prevent loss of oil from Godfrey cabin compressor due to distortion of the banjo adapter in the oil filter assembly, accomplish the following, unless already accomplished: (a) For British Aircraft Corporation Viscount Models 744, 745D, and 810 airplanes, at the next overhaul of the Supercharger or within the next 750 hours' time in service, whichever occurs first, after the effective date of this AD, comply with either paragraph (d) or (e) of this AD. (b) For Armstrong Whitworth Argosy AW-650, Grumman Model G-159, and Nihon YS-11 series airplanes, at the next overhaul of the supercharger or within the next 1,500 hours' time in service, whichever occurs first, after the effective date of this AD, comply with either paragraph (d) or (e) of this AD. (c) For Fokker Model F-27, Marks 100 and 300 series airplanes, and Fairchild Hiller Models F-27 and FH-227 series airplanes, at the next overhaul of the supercharger or within the next 1,500 hours' time in service, whichever occurs first, after the effective date of this AD, comply with either paragraph (d) or (f) of this AD. (d) Replace Godfrey banjo adapter, P/N V8653, with Godfrey banjo adapter, P/N 139313, in accordance with Godfrey Precision Products Ltd., Service Bulletin Nos. 21-120, dated July 1968 or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East. (e) Replace the Godfrey oil feeder pipe, P/N V9249, and banjo adapter, P/N 8653, with a Godfrey oil feeder pipe, P/N 139328, and banjo adapter, P/N 139322, in accordance with Godfrey Precision Products Ltd., Service Bulletin No. 21-121, dated 5 April 1968 or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East. (f) Replace the Godfrey oil feeder pipe, P/N V9979, and banjo adapter, P/N V8653, with a Godfrey oil feeder pipe, P/N 139330, and banjo adapter, P/N 139322, in accordance with Godfrey Precision Products Ltd., Service Bulletin No. 21-129, Revision 1, dated 9 October 1968 or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East. This amendment becomes effective October 9, 1969.
2009-13-07: 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: During receipt of spare parts at the final assembly line, it was discovered that lugs of the assembly nut * * * had been inverted (wrong orientation of the braking pin) during manufacturing process at the supplier. * * * This lug inversion could give the illusion of correct torque whereas the affected parts are not properly connected. Loose connection could lead to loss of the fire extinguishing system integrity and therefore inability to ensure the adequate agent concentration. In combination with an engine fire event, it could result in a temporary uncontrolled engine fire, which constitutes an unsafe condition. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
2009-13-01: This amendment supersedes an existing airworthiness directive (AD) for Sikorsky Aircraft Corporation (Sikorsky) Model S-92A helicopters. That AD currently requires removing all main gearbox (MGB) filter bowl assembly mounting titanium studs (titanium studs) and replacing them with steel studs. This amendment requires the same actions as the existing AD as well as changes to the Rotorcraft Flight Manual (RFM). This amendment is prompted by an accident, by recent RFM changes made by the manufacturer that were not available when we issued the existing AD, and by our determination that certain MGB Normal and Emergency procedures in the RFM are unclear, may cause confusion, and may mislead the crew regarding MGB malfunctions, in particular the urgency to land immediately after warning indications of loss of MGB oil pressure and oil pressure below 5 pounds per square inch (psi). Replacing the titanium studs is intended to prevent their failure, which could result in rapid loss ofoil, failure of the MGB, and subsequent loss of control of the helicopter. Changing the RFM procedures is intended to clarify and emphasize certain Normal and Emergency procedures to give the crew the best available information in the event of certain MGB malfunctions.
84-19-06: 84-19-06 GATES LEARJET: Amendment 39-4919. Applies to the following model/series airplanes certificated in all categories except airplanes modified in accordance with Supplemental Type Certificate SA944NW (Dee Howard XR Modification). Compliance required as indicated unless already accomplished. MODEL SERIAL NUMBERS 24 100 thru 357 25 003 thru 369 28 001 thru 005 29 001 thru 004 35 001 thru 514 36 001 thru 053 55 001 thru 105 To prevent aileron/trim tab flutter due to a failure or disconnect of the tab control system, accomplish the following: A. Within the next 600 hours time in service, or the next aileron/trim tab removal or rebalance, whichever occurs first, replace the trim tab balance weight and rebalance the left aileron in accordance with the instructions in Gates Learjet Corporation Airplane Modification Kit Number AMK 83-3 for Models 24, 25, 28, 29, 35, and 36; and AMK 55-83-3 for Model 55. B. Special flight permits may be issued in accordance with FAR 21.197 to operate airplanes to a base in order to comply with the modification requirements of this AD. C. Alternate means of compliance with this AD which provide an equivalent level of safety may be used when approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Gates Learjet Corporation, P.O. Box 7707, Wichita, Kansas 67277. These documents may also be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington. This amendment becomes effective October 22, 1984.
62-27-05: 62-27-05 FAIRCHILD: Amdt. 524 Part 507 Federal Register December 28, 1962. Applies to Model F-27 Aircraft Serial Numbers 2 to 95 Inclusive. Compliance required within the next 1,000 hours' time in service after the effective date of this AD. In order to preclude the loss of DC power to the primary DC bus in the event of failure of the forward primary bus relay, and the loss of AC power to certain engine instruments in the event of loss of the 115/26V instrument transformer, the following, or an equivalent approved by Chief, Engineering and Manufacturing Branch, FAA Eastern Region, New York, shall be accomplished: (a) Modify the DC electric distribution system to provide an alternate means of connecting the primary bus to the No. 2 electrical panel battery-generator feeder bus by installing one auxiliary forward primary bus relay, Cutler Hammer P/N 6041H172 in parallel with the existing forward primary bus relay. The control circuit for this auxiliary relay shall be made independent of the control circuit for the existing primary bus relay by connecting it to the flight emergency bus through an added 5-ampere circuit breaker and through a separate pair of contacts in the load monitor switch, P/N MS35059-22. (b) Modify the AC electric distribution system to provide an alternate means of supplying 26V electric power to the AC engine instruments by installing one Eclipse-Pioneer instrument transformer type DW-73-A1 and one selector switch P/N MS 35059-23 which will provide for selection of either one of the two instrument transformers. (Fairchild Service Bulletin 24-9 Revision No. 1, dated June 11, 1962, covers this same subject.) This directive effective January 29, 1963.