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78-18-06: 78-18-06 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-3296 is further amended by amendment 39-3498. Applies to Lockheed-California Company Model L-1011-385 Series airplanes certificated in all categories. Compliance required as indicated, unless previously accomplished. To prevent the transfer of potentially hazardous quantities of vapors from the engine compartment into the afterbody of the airplane, accomplish the following: (a) For airplanes with 7500 or more hours total time in service on the effective date of this AD, within the next 1000 hours time in service, inspect the center engine "S" duct flexible fire seal Part No. 1506933-101 per paragraph (c) below. (b) For airplanes with less than 7500 hours total time in service on the effective date of this AD, prior to the accumulation of 8500 hours total time in service, inspect the center engine "S" duct flexible fire seal per paragraph (c) below. (c) Visually inspect the center engine "S" duct flexible fire seal for depth of groove and/or tears per the procedures set forth in paragraph 2.A of Lockheed California Service Bulletin 093-54-030 dated March 28, 1978 (hereinafter referred to as SB 093-54-030). NOTE: Service Bulletin 093-54-030 dated March 28, 1978 and Service Bulletin 093-54-030, Revision 1, are suitable for demonstrating compliance with paragraphs (c) through (h) of this AD. (d) If there are no tears, and evidence of grooving is observed, or if groove depth is determined to be less than 0.035 inch using the procedures of paragraph 2.A of SB 093-54-030, the seal may be retained in service with performance of repetitive inspections at intervals not to exceed 1500 hours total time in service. (e) If there are no tears and seal is worn to a groove depth greater than 0.035 inch but less than 0.050 inch: (1) Continue worn part in service and reinspect within 1000 hours additional time in service; or (2) Repair seal per paragraph 2.C of Service Bulletin 093-54-030dated March 28, 1978 or per paragraph 2.D of Service Bulletin 093-54-030, Revision 1, and reinspect within 8500 hours additional time in service; or (3) Replace seal with like serviceable part per paragraph 2.B of SB 093-54- 030 and reinspect per paragraph 2.A of SB 093-54-030 within 8500 hours additional time in service. (f) If seal is worn to a groove depth equal to or greater than 0.050 inch but is intact, within 500 hours additional time in service, repair seals per paragraph 2.C of SB 093-54-030 or replace seal per paragraph 2.B of SB 093-54-030 and reinspect per paragraph 2.A of SB 093-54- 030 within 8500 hours additional time in service. (g) If seal is worn through or torn, within 250 hours additional time in service repair seal per paragraph 2.C of SB 093-54-030 or replace seal per paragraph 2.B of SB 093-54-030 and reinspect per paragraph 2.A of SB 093-54-030 within 8500 hours additional time in service. (h) Actions required per this AD may be terminated whenchafing protection (Belt Assembly Part Number 1619429-101) is installed per paragraph 2.D of SB 093-54-030 in conjunction with a new, used, or repaired center engine "S" duct flexible fire seal which meets the criteria of paragraph (d) or (e)(2) of this AD. Revert to routine maintenance/inspection practices. (i) Equvalent inspection procedures and repairs may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. (j) Special flight permits may be issued in accordance with FARs 21.197 and 21.199 to operate aircraft to a base for accomplishment of inspections and/or maintenance required by paragraph (b) of this AD. (k) Upon request of operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region may adjust the initial and 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 that operator. Amendment 39-3296 became effective October 13, 1978. This amendment, 39-3498, becomes effective June 25, 1979.
2000-01-08: This amendment adopts a new airworthiness directive (AD), applicable to all British Aerospace BAe Model ATP airplanes, that requires a one-time inspection of the orientation of certain bolts of the rudder standby control system (SCS), and reinstallation of the bolts, if necessary. 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 uncommanded engagement of the rudder SCS, which could result in reduced controllability of the airplane.
88-16-51 R1: 88-16-51 R1 EMPRESA BRASILEIRA DE AERONAUTICA S.A. (EMBRAER): Amendment 39-6053. Applies to all Model EMB-120 series airplanes, certificated in any category. Compliance required within 24 hours after the effective date of this AD, unless previously accomplished. To prevent an overspeed occurrence, accomplish the following: A. Insert the following into the FAA-approved Airplane Flight Manual (AFM) and alert all flight crews. This may be accomplished by inserting a copy of this airworthiness directive (AD) into the appropriate section of the AFM. 1. LIMITATIONS: POWERPLANT - INFLIGHT OPERATIONS AT POWER SETTINGS BELOW FLT IDLE NOT APPROVED. 2. EMERGENCY PROCEDURES IN CASE OF PROPELLER OVERSPEED: - POWER LEVER (AFFECTED ENGINE) FLT IDLE - CONDITION LEVER (AFFECTED ENGINE) FEATHER - SYNCHROPHASING OFF - POWER LEVER (OPERATIVE ENGINE) ......FLT IDLE (ALTITUDE PERMITTING) - FLAPS..................15 DEGREES (AIRSPEED PERMITTING) - AIRSPEED..........................................125 KIAS MINIMUM NOTE POWER REDUCTION ON THE OPERATIVE ENGINE WILL INCREASE DECEL RATE TO 125 KIAS AND WILL PREVENT ADVERSE AIRCRAFT HANDLING QUALITIES ASSOCIATED WITH Vmc. OPERATIVE ENGINE MAY BE USED AS NECESSARY TO MAINTAIN ALTITUDE AIRSPEED. NOTE WITH Np ABOVE 120 PERCENT, BOTH MECHANICAL AND ELECTRICAL AUXILIARY FEATHER SYSTEMS MAY NOT HAVE SUFFICIENT AUTHORITY TO FEATHER THE PROPELLER. THEREFORE, IT IS NECESSARY TO REDUCE THE Np BELOW 120 PERCENT IN ORDER TO OBTAIN SATISFACTORY FEATHERING ACTION. - IF Np IS BELOW 120 PERCENT: o ELECTRIC FEATHERING SWITCH ON - IF Np IS ABOVE 120 PERCENT: o AIRSPEED/FLAPS REDUCE AIRSPEED EXTEND FLAPS PER TABLE BELOW TO REDUCE Np BELOW 120 PERCENT: MINIMUM AIRSPEED FLAPS 125 KIAS 15 DEGREES 120 KIAS 25 DEGREES 115 KIAS 45 DEGREES o POWER LEVER (OPERATIVE ENGINE) AS REQUIRED (SEE TABLE ABOVE) NOTE WITH FLAPS LOWERED PAST 15 DEGREES, THE LANDING GEAR WARNING WILL SOUND. o ELECTRICAL FEATHER SWITCH ON NOTE THE ELECTRICAL AUXILIARY FEATHERING PUMP IS AUTOMATICALLY TURNED OFF AFTER 20 SECONDS. THEREFORE, FOR FURTHER PUMP OPERATION, IT IS NECESSARY TO TURN THE SWITCH OFF, THEN ON. THE PUMP IS CAPABLE OF SIX CONSECUTIVE OPERATIONS. - IF PROPELLER DOES NOT FEATHER: WARNING! DO NOT SHUT DOWN THE AFFECTED ENGINE UNLESS ADDITIONAL FAILURES WARRANT SHUTDOWN. o AIRSPEED 25 KIAS (MIN) o FLAPS 15 DEGREES o LAND AS SOON AS POSSIBLE USE PROCEDURES FOR A ONE ENGINE INOPERATIVE APPROACH AND LANDING. MAINTAIN Vref +10 UNTIL LANDING ASSURED. - WHEN PROPELLER IS FEATHERED: o CONDITION LEVER FUEL CUT OFF o COMPLETE PRECAUTIONARY ENGINE SHUTDOWN IF PROPELLER IS FEATHERED PRIOR TO LANDING. B. An alternate means of compliance which provides an acceptable level of safety may be used when approved by the Manager, Atlanta Aircraft Certification Office, FAA, Central Region. NOTE: The request should be forwarded through an FAA Principal Operations Inspector (POI), who may add any comments and then send it to the Manager, Atlanta Aircraft Certification Office. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer, may obtain copies upon request to Embraer, 276 S.W. 34th Street, Fort Lauderdale, Florida 33315. This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at FAA, Central Region, Atlanta Aircraft Certification Office, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia. This amendment, 39-6053, revises Telegraphic AD 88-16-51 issued August 5, 1988. This amendment, 39-6053 (AD 88-16-51 R1), becomes effective November 7, 1988.
56-19-02: 56-19-02 HAMILTON STANDARD: Applies to All 2J17 Steel Propeller Blades Installed on Boeing 377 Aircraft. Compliance required as specified herein. This note consolidates all items of AD notes and revisions issued on the subject propeller blades prior to August 1, 1956. Only those items that are still in effect are included. I. The inspection and maintenance procedures given in Hamilton Standard Service Bulletin No. 302, except as modified by this AD must be accomplished on a continuing basis. II. No external deicers shall be used on the exposed metal portions of the blade. III. Blades of the "stiffener type" shall not be used. IV. The following r.p.m. restrictions shall be included in the aircraft placard: (A) For aircraft having zinc-plated blades: (1) "Avoid ground running under static conditions above 2,600 r.p.m." (2) "Avoid continuous ground operation between 1,400 and 2,000 r.p.m." (3) "Avoid continuous operation in flight below 1,750 r.p.m. except 1,400 r.p.m. may be used for level cruise but not for descent." (B) For aircraft having nickel-plated blades: (1) "Avoid ground running under static conditions above 2,500 r.p.m." This shall be accomplished by increasing the settings of the low pitch stops in the propeller hub. (2) "Avoid continuous ground operation between 1,400 and 1,900 r.p.m." (3) "Avoid continuous operation in flight below 1,750 r.p.m. except 1,400 r.p.m. may be used for level cruise but not for descent." V. Propeller blades in service that have not been renovated or rebuilt by Hamilton Standard. Blade Serial Numbers below 649,400. (A) Conduct hand magnetic inspection of entire blade at 400-500 hour intervals of operation. (B) Blades on which corrosion was found and repaired within acceptable tolerances must not have these reworked areas covered with any material which would preclude discovery of a crack or other defect. Direct inspection, the use of hand magnetic inspection procedures or equivalent, must remain effective. (C) Conduct hand magnetic inspection of garter area daily on blades on which corrosion was found and repaired. On blades on which no corrosion was found, conduct similar inspection at 65-hour intervals of operation. (D) Conduct electrical leakage check in accordance with Section G-3 of Hamilton Standard Service Bulletin No. 302 at 120-hour intervals of operation for blades with Serial Numbers below 619,000 and 500-600 hour intervals for blades with Serial Numbers from 619,000 to 649,400. (E) Conduct electrical resistance check in accordance with Section G-4 of Hamilton Standard Service Bulletin No. 302 as follows: (1) Within 65 hours after each deicer circuit energization. (NOTE: When propeller deicing is used regularly over an extended period of time resistance checks may be conducted at 65-hour intervals of operation rather than after each use.) A means should be provided at the deicer circuit switches that will clearly indicate when a switch has been operated to energize the circuits. Maintenance procedures and instructions should be established to provide for the resistance check within the specified time whenever the indicator shows that the deicer circuits have been energized. (2) At each composite service, not to exceed 200 hours. VI. New propeller blades incorporating special treatment of unplated area and an 8- inch rubber sleeve in place of narrow garter previously used. These new blades can be identified by (1) presence of 8-inch sleeve in place of narrow garter, (2) Serial Numbers above 649,400, (3) model designation 2J17H3-8W change AE or later (zinc plate), 2J17Z3-8W change R or later (nickel plate), 2J17AG3-8W (zinc plate) or 2J17AH3-8W (nickel plate). (A) At a maximum of 1,500 hours operation, remove the sleeve, and inspect the exposed area for corrosion, and the entire blade visually and hand or machine magnetically for other defects. If the blade satisfactorily passes inspection, a new sleeve shall be installed prior to further service. These inspections shall be repeated at intervals of 1,500 hours of operation maximum. (B) Conduct electrical leakage check in accordance with Section G-3 of Hamilton Standard Service Bulletin No. 302 at the time of hand magnetic inspection specified in VI. (A). (C) Conduct electrical resistance check in accordance with Section G-4 of Hamilton Standard Service Bulletin No. 302 as follows: (1) Within 65 hours after each deicer circuit energization. (NOTE: When propeller deicing is used regularly over an extended period of time, resistance checks may be conducted at 65-hour intervals of operation rather than after each use.) A means should be provided at the deicer circuit switches that will clearly indicate when a switch has been operated to energize the circuits. Maintenance procedures and instructions should be established to provide for resistance check within the specified time whenever the indicator shows that the deicer circuits have been energized. (2) At each composite service, not to exceed 200 hours. VII. Factory renovated propeller blades incorporating special treatment of unplated area and an 8-inch rubber sleeve in place of narrow garter previously used. These blades can be identified by (1) presence of 8-inch sleeve in place of narrow garter, (2) Serial Numbers between 645,300 and 649,400, (3) model designation 2J17H3-8W prior to change AE (zinc plate), or 2J27Z3-8W prior to change R (nickel plate). (A) At a maximum of 1,500 hours of operation remove the sleeve, and inspect the exposed area for corrosion, and the entire blade visually and hand or machine magnetically for other defects. If the blade satisfactorily passes inspection a new sleeve shall be installed prior to further service. These inspections shall be repeated at intervals of 1,500 hours of operation maximum. (B) Conduct electrical leakage check in accordance with Section G-3 of Hamilton Standard Service Bulletin No. 302 at the time of hand magnetic inspection specified in VII.(A). (C) Conduct electrical resistance check in accordance with Section G-4 of Hamilton Standard Service Bulletin No. 302 as follows: (1) Within 65 hours after each deicer circuit energization. (NOTE: When propeller deicing is used regularly over an extended period of time, resistance checks may be conducted at 65-hour intervals of operation rather than after each use.) A means should be provided at the deicer circuit switches that will clearly indicate when a switch has been operated to energize the circuits. Maintenance procedures and instructions should be established to provide for the resistance check within the specified time whenever the indicator shows that the deicer circuits have been energized. (2) At each composite service, not to exceed 200 hours. VIII. Check strength of hand magnets for conformance withHamilton Standard Service Bulletin No. 302 or equivalent before each use, or daily when in continued use. IX. To supplement the above precautions, while in flight continue to monitor the vibration indicator at least hourly for any indication of progressive unbalance. X. While in flight continue to monitor the propeller deicer current load meter for any indication of deicer heater resistance change. This supersedes AD 55-21-01.
75-07-03: 75-07-03 LOCKHEED: Amendment 39-2133 as amended by Amendment 39-2744. Applies to L-1011-385-1 series airplanes certificated in all categories. Compliance required as indicated. To prevent the possibility of an out-of-trim Autopilot upon manual disconnect from occurring at close proximity to the ground, and to prevent operation at speeds at which the airplane may not meet stability requirements with inoperative MACH TRIM and MACH FEEL systems, and to advise the flight crew of possible incorrect failure annunciations, accomplish the following: (a) Within 100 hours time in service after the effective date of this AD, unless already accomplished, (1) Install the following placard in full view of the flight crew, "DUAL PITCH TRIM REQ'D TO 100 FT. FOR A/P USE BELOW 100 FT." (2) Revise L-1011-385-1, FAA Approved Airplane Flight Manual (AFM) Limitations sections as follows: LR 25925 by incorporating pages 1-6 and 1-7 dated March 6, 1975, or later FAA-approved revisions; and LR 25225 by incorporating pages 1-6.2 and 1-6.3 dated March 6, 1975, or later FAA-approved revisions. Also revise appropriate operations manuals to incorporate the TRIM/FEEL system failure limitations included in the above AFM pages. (b) An operator may remove the placards and discontinue the instructions imposed by this AD on his fleet of airplanes after the following actions have been accomplished. (1) All Trim Augmentation Computers, Lockheed P/N 672 443-(105, -107 or -109) in service and in spares inventory are modified per Lockheed Service Bulletin 093-22-069, dated November 26, 1974, or later FAA-approved revisions; and (2) A system of parts pooling is established to insure that only spares, modified as defined in (b)(1), above, are installed. (c) The Lockheed Company may operate and deliver an airplane to an operator without the placard required by this AD after the individual operator provides written notification to the Lockheed Company that his fleetno longer requires the placard as provided in (b), above. (d) All Trim Augmentation Computers, Lockheed P/N 672 443- (105, -107, or -109), must be modified in accordance with Lockheed Service Bulletin 093-22-069, dated November 26, 1974, or later FAA - Approved revisions, by July 1, 1977. (e) Equivalent procedures and modifications may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (f) An airplane may be flown to a base for the performance of the work required by this AD, per FAR's 21.197 and 21.199. (g) For those airplanes not provided for in paragraph (c), above, the Lockheed Company may operate these airplanes without the placard required by this AD if the following actions are accomplished: (1) A Lockheed P/N 672 443-113 Trim Augmentation Computer or later FAA-approved version is installed; and (2) a sign-off indicating compliance with paragraph (g) of the AD is made on the aircraft release form prior to each flight. Amendment 39-2133 became effective March 27, 1975. This Amendment 39-2744 becomes effective October 20, 1976.
93-24-12: 93-24-12 LOCKHEED: Amendment 39-8761. Docket 92-NM-199-AD. Applicability: Model L-1011-385-1, -385-1-14, -385-1-15, and -385-3 series airplanes; as listed in Lockheed Service Bulletin 093-52-155, Revision 1, dated October 23, 1989, and Lockheed Service Bulletin 093-53-252, Revision 2, dated April 25, 1989; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent failure of the cargo door latching mechanism, which could lead to the cargo door opening during flight and resulting in rapid decompression of the airplane, accomplish the following: (a) For airplanes listed in Lockheed Service Bulletin 093-52-155, Revision 1, dated October 23, 1989: Prior to the accumulation of 8,000 total landings, or within 1,800 landings after the effective date of this AD, whichever occurs later, perform a one-time visual inspection to detect excessive thickness of the shims installed under the hinges of the C-1A cargo door in accordance with Lockheed Service Bulletin 093-52-155, Revision 1, dated October 23, 1989. Accomplishment of this inspection is considered to constitute compliance with the similar inspection requirements of AD 91-05-05, Amendment 39-6878, for this same area. (1) If any shim is found that exceeds 0.125 inch for single leg hinges, or 0.140 inch for double leg hinges, prior to further flight, install a structural doubler on the hinge and a new shim, in accordance with the procedures described in the service bulletin. (2) If any shim is found that equals or is less than 0.125 inch for single leg hinges, or 0.140 inch for double leg hinges, no further action is required for that shim. (b) For airplanes listed in Lockheed Service Bulletin 093-53-252, Revision 2, dated April 25, 1989: Prior to the accumulation of 8,000 total landings or within 1,800 landings after the effective date of this AD, whichever occurs later, inspect the lower sill latch fittings and serrated plates ofthe C-1A cargo door for cracks and corrosion, in accordance with Lockheed Service Bulletin 093-53-252, Revision 2, dated April 25, 1989. Additionally, perform a hardness test to determine the condition of the heat treatment of the serrated plates. NOTE 1: Lockheed Service Bulletin 093-53-252, Revision 2, dated April 25, 1989, refers to Lockheed Service Modification/Kit Drawing 1646587, Revision C, dated August 14, 1987, for additional information concerning the inspection procedures, corrosion limit specifications, crack limit specifications, and modifications relative to the requirements of this paragraph. (1) If any cracked latch fitting is found, prior to further flight, replace the latch fitting with a serviceable part. (2) If any corroded latch fitting is found, prior to further flight, replace the latch fitting with a serviceable part. However, if the latch fitting is of a condition suitable for refurbishment, as referred to in the service bulletin, it may berefurbished and reused. (3) If any cracked serrated plate is found, prior to further flight, replace it with a serviceable part. However, if the cracked serrated plate is determined to be suitable for reuse, as referred to in the service bulletin, it may be reinstalled for an additional 1,000 landings only, at which time it then must be replaced. (4) If no crack or corrosion is found in any serrated plate, prior to further flight, apply cadmium plating to the plate in accordance with the service bulletin. (5) If any serrated plate is found with improper heat treatment, prior to further flight, reprocess the plate or replace the plate with a serviceable part in accordance with the service bulletin. (c) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Atlanta Aircraft Certification Office (ACO), FAA, Small Airplane Directorate. Operators shall submit theirrequests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Atlanta ACO. NOTE 2: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Atlanta ACO. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (e) The inspection and installation shall be done in accordance with Lockheed Service Bulletin 093-52-155, Revision 1, dated October 23, 1989. The inspection, hardness test, replacement, refurbishment, reinstallation, application of cadmium plating, and reprocessing shall be done in accordance with Lockheed Service Bulletin 093-53-252, Revision 2, dated April 25, 1989, which contains the following list of effective pages: Page Number Revision Level Shown on Page Date Shown on Page 1, 4-5 2 April 25,1989 2-3 Original March 15, 1988 The incorporation by reference of these documents was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Lockheed Western Export Company (LWEC), Dept. 693, Zone 0755, 86 South Cobb Drive, Marietta, Georgia 30063. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Small Airplane Directorate, Atlanta Aircraft Certification Office (ACO), 1669 Phoenix Parkway, Atlanta, Georgia; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (f) This amendment becomes effective on January 10, 1994.
89-16-04: 89-16-04 BRITISH AEROSPACE (BAe) PLC: Amendment 39-6268. Applicability: Jetstream Model 3101 (all serial numbers) airplanes equipped with Sperry SPZ-200B or SPZ-500 autopilot installations. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To prevent the possible loss of elevator trim control, accomplish the following: (a) For all airplanes on which the autopilot elevator trim servo cable has accumulated 900 or more hours time-in-service (TIS) on the effective date of this AD, within the next 100 hours TIS and thereafter at intervals not to exceed 1,000 hours TIS, replace the elevator trim servo cable, ET1, BAe Part Number 137187E472, in accordance with BAe Mandatory Alert Service Bulletin Jetstream 22-A-JA861023, Rev 1, dated January 6, 1989. (b) For all airplanes on which the autopilot elevator trim servo cable has accumulated less than 900 hours TIS on the effective date of this AD, at or before reaching1,000 hours TIS and thereafter at intervals not exceeding 1,000 hours TIS, replace the autopilot elevator trim servo cable, ET1, BAe Part Number 137187E472, in accordance with BAe Mandatory Alert Service Bulletin Jetstream 22-A-JA861023, Rev 1, dated January 6, 1989. (c) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (d) An equivalent means of compliance with this AD may be used if approved by the Manager, Aircraft Certification Office, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, B-1000 Brussels, Belgium. All persons affected by this directive may obtain copies of the documents referred to herein upon request to British Aerospace PLC, Manager, Product Support, Civil Aircraft Division, Prestwick Airport, Ayrshire, KA9 2RW, Scotland; or British Aerospace, Inc., Librarian, Box 17414, Dulles International Airport, Washington, D.C. 20041; or may examine these documents at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD supersedes AD 87-07-02, Amendment 39-5579, which became effective on April 17, 1987. This amendment (39-6268, AD 89-16-04) becomes effective on August 24, 1989.
2014-15-18: We are adopting a new airworthiness directive (AD) for certain Mooney International Corporation Models M20C, M20E, M20M, M20R, and M20TN airplanes. This AD requires inspection of the outer empennage attach fittings for correct thickness with replacement as necessary. This AD was prompted by discovery of empennage attach fittings (Lugs) that do not meet the approved design dimensional requirements, which could result in possible reduction in fatigue or static strength and/or corrosion. We are issuing this AD to correct the unsafe condition on these products.
80-13-11 R2: 80-13-11 R2 DeHAVILLAND: Amendment 39-3814 as amended by Amendment 39-4703. Applies to all DHC-6 model airplanes, certificated in all categories. To prevent possible loss of control due to cracking of the elevator, flap and aileron control rods, accomplish the following: (a) On aircraft Serial Numbers 1 thru 430 and on those aircraft having as replacement control rods those with part numbers listed in Column 2 of Table 2 in DeHavilland Service Bulletin 6/390, within the next 50 hours in service or 30 days, whichever occurs first, after the effective date of this AD, unless previously accomplished within the last 350 hours in service or 150 days, whichever occurred last, visually inspect tube ends of the rod assemblies in accordance with the dye penetrant method using at least a ten power glass, in the above Bulletin's paragraph 8, 9 and 10 of ACCOMPLISHMENT INSTRUCTIONS, or approved equivalent. (b) If cracks are not or have not been found, repeat inspection in paragraph (a) within 400 hours in service or 180 days, whichever occurs first after the last inspection. Following inspection, install sleeves on rods in accordance with the above Bulletin's ACCOMPLISHMENT INSTRUCTIONS or approved equivalent and the applicable drawings and Mod numbers listed in Column 4 of Table 1. (c) On aircraft Serial Numbers 431 thru 685 and on other aircraft having as replacement control rods those with part numbers listed as "original" in Column 5 of Table 2 in the above Bulletin, within 800 hours in service or one year, whichever occurs first, from the effective date of this AD, unless previously accomplished, inspect in accordance with paragraph (a). Following inspection, install sleeves on rods in accordance with the Bulletin's ACCOMPLISHMENT INSTRUCTIONS or approved equivalent and the applicable drawings and Mod numbers listed in Column 5 of Table 1. (d) On aircraft Serial Numbers 686 and subsequent and all other aircraft on which paragraphs (b) or (c) have been accomplished: (1) Remove external paint and visually inspect with at least a ten power glass in accordance with the above Service Bulletin those areas shown in Figure 1 on all tube ends of the rod assemblies listed in Column 4 or 5 of Table 2 of the Bulletin. (2) The inspections specified in paragraph (d)(1) of this AD must be accomplished at intervals not to exceed 800 hours time-in-service, or 1 year, whichever occurs first, from the last inspection. (e) If cracks are found, the rod assembly must be replaced before further flight with rods of the same part number or equivalent inspected and found serviceable in accordance with paragraph (a); or with new rods of the same part number or equivalent; or with new Post-Mod rods whose parts are listed in Column 5 or 6 of Table 1 in the above Bulletin. (f) Report positive findings, including crack length, from any of the above inspections to the Manager, Aircraft Certification Office, FAA, New England Region, within 10 days of inspection. (Reporting approved by Office of Management and Budget under OMB No. 04-R0174.) (g) Equivalent parts and procedures must be approved by the Manager, Aircraft Certification Office, FAA, New England Region. (h) Compliance times may be increased by the Manager, Aircraft Certification Office, FAA, New England Region, upon receipt of substantiating data submitted through an FAA Maintenance Inspector. Amendment 39-3814 was effective July 1, 1980. This Amendment 39-4703 becomes effective August 10, 1983.
2014-12-11: We are adopting a new airworthiness directive (AD) for the Sikorsky Aircraft Corporation (Sikorsky) Model S-92A helicopter. This AD requires revising the Rotorcraft Flight Manual (RFM) to include the appropriate operating limitations for performing Class D external load- combination operations. This AD was prompted by an inaccurate RFM provision, which was approved without appropriate limitations for this model helicopter for carrying Class D external rotorcraft-load combinations, including human external cargo (HEC). The actions are intended to require appropriate operating limitations to allow operators to perform Class D external load-combination operations, including HEC, in this model helicopter that now meets the Category A performance standard.
2006-04-12: The FAA is superseding an existing airworthiness directive (AD) for General Electric Company (GE) CF34-3A1 and -3B1 series turbofan engines. That AD requires initial and repetitive visual inspections and eddy current inspections (ECIs) of certain stage 5 low pressure turbine (LPT) disks and stage 6 LPT disks, installed in GE CF34-3A1 and -3B1 series turbofan engines. Those engines are installed in certain Bombardier Canadair Regional Jet (RJ) airplanes. This AD requires the same initial and repetitive visual inspections and ECIs, but adds SNs to the affected disk population for RJ airplanes. This AD also adds GE CF34-1 and -3 series turbofan engines with certain stage 5 and stage 6 LPT disks, to the applicability section. Those engines are installed in certain Bombardier Canadair Business Jet (BJ) airplanes. Also, this AD requires eventual replacement of the affected disks as terminating action to the repetitive inspections. This AD results from the discovery of an additionalpopulation of suspect stage 5 LPT disks and stage 6 LPT disks that could fail due to low-cycle fatigue cracking that may start at the site of an electrical arc-out on the disk. We are issuing this AD to prevent low-cycle-fatigue (LCF) failure of stage 5 LPT disks and stage 6 LPT disks, which could lead to uncontained engine failure.
95-17-05: This amendment adopts a new airworthiness directive (AD), applicable to all Airbus Model A310 series airplanes, that requires inspections to detect loose self-locking nuts and damaged cotter pins on the actuating cylinder to drag strut attachment of the left- and right-hand main landing gear (MLG), and correction of discrepancies. This amendment also provides an optional terminating action for the repetitive inspections. This amendment is prompted by reports of loose nuts and sheared cotter pins found on in-service airplanes. The actions specified by this AD are intended to prevent an undampened free fall of the left- and right-hand MLG, which subsequently could lead to the inability to retract the MLG and damage to other airplane systems.
2022-16-09: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-8 airplanes. This AD was prompted by a report that, during production, a small number of fasteners in certain locations of the center fuel tank were cap sealed on top of a black stripe of ink with a clear overcoat. This clear overcoat is not an approved surface for sealing and can potentially compromise sealant adhesion. Compromised sealant adhesion can, over time, affect the lightning-protection properties of the airplane. This AD requires preparation of the affected surface areas to ensure that there is adequate sealant adhesion, and complete encapsulation of the discrepant fastener locations with the approved production sealant. The FAA is issuing this AD to address the unsafe condition on these products.
87-21-06: 87-21-06 BRITISH AEROSPACE: Amendment 39-5744. Applies to Model BAC 1-11 - 200 and -400 series airplanes, listed in British Aerospace Campaign Wire 52-CW-PM 5448, Issue 2, dated September 18, 1987, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent sudden decompression of the fuselage as a result of failure of the rear passenger door, accomplish the following: A. Within the next 100 landings, or within 30 days after the effective date of this AD, or prior to the accumulation of 20,000 landings, whichever occurs later, inspect the rear passenger door structure for cracks, in accordance with British Aerospace Campaign Wire 52-CW-PM5448, Issue 2, dated September 18, 1987. B. Repair or replace cracked structure, before further flight, in accordance with campaign wire. C. Repeat the inspection required by paragraph 1., above, at intervals not to exceed: 1. 800 landings for airplanes with doors that have the horizontal members repaired in accordance with the structural repair manual. 2. 1,200 landings for airplanes with doors in which the horizontal members were found undamaged at the last inspection or were replaced. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes unpressurized to a base for the accomplishment of the inspection required by this AD. All persons affected by this directive who have not already received the appropriate service information, may obtain copies upon request to British Aerospace, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. This information may be examined at FAA, Northwest Mountain Region, 17900Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective October 21, 1987.
2000-01-07: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Bombardier Model DHC-8-100, -200, and -300 series airplanes. This action requires the removal and testing of sections of bonded skin from the upper and lower skin panels of the horizontal stabilizer, repair of those areas, and follow-on corrective actions, if necessary. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified in this AD are intended to prevent reduced strength capability and consequent failure of the horizontal stabilizer, which could result in loss of controllability of the airplane.
92-01-02: 92-01-02 FAIRCHILD AIRCRAFT (formerly Swearingen): Amendment 39-8125. Docket No. 91-CE-26-AD. Applicability: Model SA226-T airplanes (Serial numbers (S/N) T201 through T275, and T277 through T291), Model SA226-T(B) airplanes (S/N T(B)276, and T(B)292 through T(B)417), Model SA226-AT airplanes (S/N AT001 through AT074), Model SA226-TC airplanes (S/N TC201 through TC419), Model SA227-TT airplanes (S/N TT421 through TT555), Model SA227-AT airplanes (S/N AT423 through AT599), and Model SA227-AC airplanes (S/N AC406, AC415, AC416, and AC420 through AC599), certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent brake system malfunctions that could result in a fire in the brake area or possible airplane collision during landing, accomplish the following: (a) Within the next 90 calendar days after the effective date of this AD, modify the parking brake system in accordance with the instructions in Fairchild Aircraft Service Bulletin (SB) No. 227-32-017 and SB No. 226-32-049, both dated November 14, 1984, as applicable. (b) On airplanes equipped with B.F. Goodrich brakes, part number 2-1203-3, within the next 100 hours TIS after the effective date of this AD, and thereafter at intervals not to exceed 250 hours TIS, inspect and conduct measurements in accordance with the instructions in B.F. Goodrich Service Letter No. 1498, dated October 26, 1989. If wear measure exceeds the maximum allowed in accordance with the criteria in B.F. Goodrich Service Letter No. 1498, dated October 26, 1989, prior to further flight, overhaul or replace the brakes in accordance with the instructions in the applicable maintenance manual. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) An alternative method of compliance or adjustment of the initial or repetitive compliance timesthat provides an equivalent level of safety may be approved by the Manager, Fort Worth Airplane Certification Office, Fort Worth, Texas 76193-0150. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Fort Worth Airplane Certification Office. (e) The inspections and modifications required by this AD shall be done in accordance with Fairchild Aircraft Service Bulletin (SB) No. 227-32-017 and SB No. 226-32-049, both dated November 14, 1984, and B.F. Goodrich Service Letter (SL) No. 1498, dated October 26, 1989. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Fairchild Aircraft Corporation, P.O. Box 790490, San Antonio, Texas 78279-0490 and B.F. Goodrich Aircraft Wheels and Brakes, P.O. Box 340, Troy, Ohio 45373. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. (f) This amendment (39-8125, AD 92-01-02) becomes effective on January 16, 1992.
88-07-03: 88-07-03 BRITISH AEROSPACE: Amendment 39-5877. Applies to Model BAe-125 series airplanes listed in BAe-125 Service Bulletin 57-66, Revision 2, dated October 25, 1986, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent displacement of the aileron mass balance side plate, and possible control surface interference, accomplish the following: A. Within 6 months after the effective date of this AD, perform a dye penetrant inspection of the side plate of the aileron mass balance assembly for cracking of the attachment lugs, in accordance with BAe Service Bulletin 57-66, Revision 2, dated October 25, 1986. Repeat the inspection at intervals not to exceed 2 years. B. If cracks are detected, repair before further flight, in accordance with Repair Scheme 25WG/R143, issued with Service Bulletin 57-66, Revision 2, dated October 25, 1986. C. Accomplishment of Repair Scheme 25WG/R143 constitutes terminating action for the repetitive inspections required by paragraph A., above. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to British Aerospace, Inc., Service Bulletin Librarian, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective April 25, 1988.
2014-15-16: We are adopting a new airworthiness directive (AD) for certain Airbus Model A319-111, -112, -115, -132, and -133 airplanes; Model A320-214, -232, and -233 airplanes; and Model A321-211, -231, and -232 airplanes. This AD requires a detailed inspection for missing fasteners on the frame between certain stringers; for certain airplanes, a rototest inspection of the fastener holes for cracking; and corrective actions if necessary. This AD was prompted by a report that when the cabin lining was removed during a cabin conversion it was discovered that fasteners were missing on the frame. We are issuing this AD to detect and correct missing fasteners which, if not corrected, could affect the structural integrity of the airframe and could result in rapid decompression.
2000-01-10: This amendment supersedes Airworthiness Directive (AD) 98-08-07, which currently requires replacing the rudder and elevator pivot arms with parts of improved design on certain Pilatus Aircraft Ltd. (Pilatus) Model PC-7 airplanes. This AD requires replacing the rudder and elevator pivot arms with parts that have been improved since issuance of AD 98-08-07. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Switzerland. The actions specified by this AD are intended to prevent failure of the elevator and rudder caused by fatigue cracking of the pivot arms, which could result in reduced airplane controllability and possible loss of control of the airplane.
2000-01-02: This amendment adopts a new airworthiness directive (AD), applicable to all Raytheon Model BAe.125 series 1000A and 1000B airplanes and Model Hawker 1000 series airplanes, that requires an inspection to determine the integrity of the duct connection on both ends of the turbine air discharge duct in the air conditioning system; an inspection to measure the bead height on the ends of the turbine air discharge duct; and corrective actions, if necessary. This amendment is prompted by reports indicating that the turbine air discharge duct disconnected from the cold air unit (CAU) or water separator due to insufficient bead height on the ends of the turbine air discharge duct. The actions specified by this AD are intended to prevent such disconnection from the CAU or water separator, which could result in cabin depressurization.
48-51-02: 48-51-02 CONVAIR: Applies to All Model 240 Aircraft. Recently reported difficulties concerning the horizontal tail surfaces for Convair Model 240 aircraft have indicated the necessity of conducting thorough and frequent inspections of all critical items to detect any fatigue cracks and to minimize the development of hazardous conditions. The several reported difficulties appear to result from tail vibrations induced by the engines and/or propellers at certain r.p.m.'s. Pending completion of the necessary investigations and determination of adequate remedial measures, the following must be accomplished: A. To be accomplished daily on the left horizontal surfaces. (NOTE: Time limit may be extended to each No. 3 operation upon incorporation of AD 49-44-02 and CVAC Service Bulletins 240-219 and 240-247). Utilizing available inspection openings but without removing any surfaces, conduct a close, visual inspection of the following: 1. Horizontal stabilizer skin, rearspar and hinge brackets. 2. Elevator skin, leading edge ribs, spars, hinge brackets and balance weight installations. 3. Tab skin, spar, hinge brackets and balance weight installations. Particular care should be taken to detect any evidence of loose balance weights, sheared rivets or cracked hinge brackets. Any failed parts should be adequately repaired or replaced prior to the next flight. NOTE: Inspection procedures which have been satisfactorily demonstrated to the CAA Agent to provide equivalent safety may be accepted in lieu of the inspection procedures outlined above. NOTE: The following static balance tolerances about the respective hinge lines must be retained after rework of any surface: 1. Left elevator, including flight tab; 72-87 inch-pounds tail heavy with the seal curtain removed. 2. Right elevator, including trim tab: 0-15 inch-pounds nose heavy with the seal curtain removed. 3. Elevator flight tab: 6-8 inch-pounds nose heavy. B. To be accomplished at periods not to exceed 50 hours until close tolerance bolts and bushings have been installed in elevator tab hinges in accordance with CVAC Service Bulletin 240-205. Remove all elevator flight tab hinge pins and inspect the pins, bushings and bearings for sign of wear. Worn parts should be replaced. This supersedes AD 48-47-01.
2014-15-09: We are adopting a new airworthiness directive (AD) for all Airbus Model A330-200 Freighter, A330-200 and -300, and A340-200, -300, -500, and -600 series airplanes. This AD was prompted by reassessment of an unsafe condition related to MZ-type spoiler servo-controls (SSCs) that did not remain locked in the retracted position (hydraulic locking function) after manual depressurization of the corresponding hydraulic circuit. This reassessment resulted in the determination that performing repetitive operational tests of all SSC types is necessary. This AD requires repetitive operational tests of the hydraulic locking function on each SSC installed on the blue or yellow hydraulic circuits, and replacing any affected SSC with a serviceable SSC. We are issuing this AD to detect and correct loss of the hydraulic locking function during take-off, which, in combination with one inoperative engine, could result in reduced controllability of the airplane.
2000-01-05: This amendment supersedes an existing airworthiness directive (AD) applicable to certain Boeing Model 747 series airplanes. That AD currently requires repetitive inspections and tests of the thrust reverser control and indication system on each engine, and corrective actions, if necessary; installation of a terminating modification; and repetitive operational checks of that installation, and repair, if necessary. This amendment is prompted by the results of a safety review, which revealed that in-flight deployment of a thrust reverser could result in significant reduction in airplane controllability. The actions specified in this AD are intended to ensure the integrity of the fail-safe features of the thrust reverser system by preventing possible failure modes, which could result in inadvertent deployment of a thrust reverser during flight, and consequent reduced controllability of the airplane. This action identifies certain repetitive operational checks that were inadvertently omitted from the existing AD, and revises certain procedures for accomplishment of the operational checks and certain follow-on corrective actions.
82-19-02: 82-19-02 GENERAL DYNAMICS: Amendment 39-4458. Applies to Model 340, 440, and military models eligible or to be made eligible for civil use under Type Certificate 6A6, and all such model airplanes converted to turbopropeller power, certificated in all categories. Compliance required as indicated unless previously accomplished. To prevent possible loss of a horizontal stabilizer due to failure of the stabilizer attachment fittings (P/Ns 340-8510150 and 340-8510151) caused by fatigue cracks, accomplish the following: A. Within 250 hours time in service or within 90 days from the effective date of this AD, whichever occurs first, unless previously accomplished within the last 450 hours of time in service, conduct a visual inspection of the upper and lower, forward and aft, horizontal stabilizer attachment fitting lugs in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions," General Dynamics Convair Division Service Bulletin 640 (340D)55-3A, Revision 1, dated June 12, 1981. If cracks are found, replace with a new part before further flight. B. Within 250 hours time in service or within 90 days from the effective date of this AD, whichever occurs first, unless previously accomplished within the last 1150 hours of time in service, conduct an ultrasonic inspection of the upper and lower, forward and aft, horizontal stabilizer attachment fitting lugs in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of SB 640 (340D) 55-3A. If cracks are found, replace with a new part before further flight. C. Repeat the visual inspection required by paragraph A of this AD at intervals not to exceed 700 hours time in service from the last such inspection and repeat the ultrasonic inspection required by paragraph B of this AD at intervals not to exceed 1400 hours time in service from the last such inspection. D. Within 10,000 hours time in service from the effective dateof this AD, unless already accomplished, install bushings in the horizontal stabilizer attachment fitting lug holes in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of SB 640 (340D) 55-3A. Continue to visually inspect at 700 hour intervals per paragraph A and to inspect by ultrasonic procedures at 1400 hour intervals per paragraph B of this AD. E. Within 250 hours time in service or within 90 days from the effective date of this AD, whichever occurs first, unless previously accomplished within the last 450 hours time in service, conduct an internal visual inspection for cracks in the horizontal stabilizer attachment fittings, upper, lower, forward and aft, and associated structure outboard of the lugs and butt rib in the area surrounding the fasteners, in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of General Dynamics Convair Division Service Bulletin 640 (340D) 55-4, Revision 1, dated March 24, 1982. Inspect for evidence of loose rivets or fasteners in the stabilizer attachment fittings with particular attention given to those just outboard of the butt rib which fasten the upper forward fittings to the spar web and rail. If cracks are found in the stabilizer attachment fittings or associated structure, and/or if loose fasteners are detected in the stabilizer attachment fitting: replace fitting with a new part, repair structure, and/or replace the fasteners, in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of SB 640 (340D) 55-4 before further flight. F. Within 250 hours time in service or within 90 days from the effective date of this AD, whichever occurs first, unless previously accomplished within the last 1150 hours time in service, conduct an x-ray inspection for cracks in the upper forward and upper aft horizontal stabilizer attachment fittings in the area outboard of the lugs in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of SB 640 (340D) 55-4. Only the four (4) upper fittings are required to be x-ray inspected as they are the most highly loaded in tension and therefore most susceptible to fatigue cracking. If cracks are found, replace with a new part before further flight. G. Repeat the visual inspection required by paragraph E of this AD at intervals not to exceed 700 hours time in service from the last such inspection and repeat the x-ray inspection required by paragraph F of this AD at intervals not to exceed 1400 hours time in service from the last such inspection. H. Within 10,000 hours time in service from the effective date of this AD, unless already accomplished, replace the first three (3) steel rivets outboard of the butt rib through the two (2) forward upper horizontal stabilizer attachment fittings in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions" of SB 640 (340D) 55-4. Continue to visually inspect at 700 hour intervals per paragraph E and x-ray inspect at 1400 hour intervals per paragraph F of this AD. I. Prior to issuance of a Certificate of Airworthiness for military aircraft being converted for civil certification, and prior to further flight for any aircraft that has been out of service for one (1) year or more, the airplane must be inspected in accordance with paragraphs A, B, E, and F of this AD. J. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections or modifications required by this AD. K. Alternative inspections, modifications, or other actions which provide an equivalent level of safety may be used when approved by the Chief, Western Aircraft Certification Field Office, FAA Northwest Mountain Region, Hawthorne, California. 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 General Dynamics, Convair Division, P.O. Box 80877, San Diego, California 92138. These documents also may be examined at FAA Northwest Mountain Region, 17900 Pacific Highway South, C-68966, Seattle, Washington 98168; or Western Aircraft Certification Field Office, 15000 Aviation Boulevard, Hawthorne, California. This supersedes AD 81-16-07, Amendment 39-4180 (46 FR 39431), issued July 24, 1981, and AD 80-11-01, Amendment 39-3775 (45 FR 35309), issued May 13, 1980. This Amendment becomes effective September 16, 1982.
74-10-10: 74-10-10 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-1842 as amended by Amendment 39-1868 is further amended by Amendment 39-2253. Applies to Model TPE331-1-151B, -2-201C, -3U-307G, -5-251C, -5-251K and -6-251M engines. Compliance required as indicated. To detect, correct and prevent loosening of the fuel control assembly mounting and support bracket fasteners, accomplish the following: (a) Within the next 25 hours time in service after the effective date of this AD, unless previously accomplished, replace the fuel control assembly support bracket with a new bracket in accordance with instructions contained in paragraph 2.B. of AiResearch Service Bulletin TPE331-73-0028, dated April 29, 1974, or later FAA-approved revisions. (b) Within 100 hours time in service after accomplishment of paragraph (a), above, or 25 hours additional time in service after the effective date of this AD, whichever occurs later, and intervals not to exceed 100 hours time in service thereafter, inspect and retorque the fuel control assembly and support bracket mounting fasteners in accordance with the procedures contained in paragraph 2.C. of AiResearch Service Bulletin TPE331- 73-0028, dated April 29, 1974, or later FAA-approved revisions. (c) Equivalent procedures may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiation data. Note: AiResearch Operating Information Letter OI-331-3, Revision A, dated April 19, 1974, provides pertinent operational information. Flight crews are urged to obtain a copy of this letter and acquaint themselves with the information presented therein. (d) Unless previously accomplished, within the next 1000 hours time in service, or within one year after the effective date of this AD, whichever occurs earlier, the fuel control assembly support bracket attachment lug and fastener must be modified in accordance with AiResearch Service Bulletin TPE331-73-0029, dated May 27, 1974, or later FAA-approved revisions. This change has been incorporated in production engines by model and serial number as defined in paragraph 1.A. of the above referenced service bulletin. (e) The recurring inspection required in (b), above, may be discontinued when the modification prescribed in (d), above, have been accomplished. (f) If the replacement of the support bracket required by paragraph (a), above, has not already been accomplished on the effective date of this amendment to the AD, this modification, when accomplished, must be done in accordance with the instructions contained in paragraph 2.B. of AiResearch Service Bulletin TPE331-73- 0028, Revision 2, dated June 20, 1975, or later FAA-approved revisions. (g) If the replacement of the support bracket required by paragraph (a), above, has already been accomplished, at the next inspection required by paragraph (b), above, install the three (3) washers, P/N 960C10, and bolts, P/N MS21379-14, as required, per paragraph 2.B. of AiResearch Service Bulletin TPE331-73-0028, Revision 2, dated June 20, 1975, or later FAA-approved revisions. (h) If the modification described in paragraph (d), above, has already been accomplished, before exceeding an additional 100 hours time in service after the effective date of this AD, as amended, unless already accomplished; or, when accomplishing paragraph (d), install the three (3) washers, P/N 960C10 and bolts, P/N MS21279-14, as required, per paragraph 2.B. of AiResearch Service Bulletin TPE331-73-0028, Revision 2, dated June 20, 1975, or later FAA-approved revisions. Amendment 39-1842 became effective May 13, 1974. Amendment 39-1868 became effective June 14, 1974. This amendment 39-2253 becomes effective July 11, 1975.