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77-13-11: 77-13-11 HAWKER SIDDELEY AVIATION, LTD: Amendment 39-2936. Applies to Model DH/BH-125 airplanes, all series, certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent possible failure of the knife edges of the emergency brake reducing valve, P/N AC.61516, and consequent complete loss of emergency braking capability without advance warning, accomplish the following: (a) Comply with paragraph (b) or (c) of this AD as follows, and, thereafter, continue to comply with paragraph (b) or (c) of this AD at intervals not to exceed 4500 landings since last compliance. (1) For airplanes having emergency brake control valve knife edges that have accumulated less than 4300 landings, since new, on the effective date of this AD, compliance is required prior to the accumulation of 4500 landings. (2) For airplanes having emergency brake control valve knife edges that have accumulated 4300 or more landings, since new, onthe effective date of this AD, compliance is required prior to the accumulation of an additional 200 landings. (3) For airplanes for which no records exist that indicate the number of landings the emergency brake control valve knife edges have accumulated, compliance is required prior to the accumulation of 200 landings after the effective date of this AD. (b) Replace the emergency brake control valve knife edges with new parts in accordance with Part A of Section 2 entitled "Accomplishment Instructions" of Hawker Siddeley Aviation, Ltd., Service Bulletin 32-167, dated January 27, 1976, or an FAA-approved equivalent. (c) Replace the entire emergency brake reducing valve, P/N AC.61516, with a valve fitted with new knife edges and install the valve in accordance with Part B of the section entitled "Accomplishment Instructions" of Hawker Siddeley Aviation, Ltd., Service Bulletin 32-167, dated January 27, 1976, or an FAA-approved equivalent. This amendment becomes effective July 25, 1977.
2009-25-11: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 747 airplanes. The existing AD currently requires repetitive inspections for cracking, and repair as necessary, of lower lobe body frames (sections 42 and 46) of the fuselage. The existing AD also provides for optional modification of the frames, which terminates the repetitive inspections. This new AD requires additional repetitive inspections for cracking of certain fuselage frames, and corrective actions if necessary. This AD results from a new report of a crack found in a body frame with a tapered side guide bracket at fuselage station 1800, located on the left side between stringers 39 and 40; the frame was severed. We are issuing this AD to detect and correct the loss of structural integrity of the fuselage, which could result in rapid depressurization of the airplane.
90-22-02: 90-22-02 PRATT & WHITNEY: Amendment 39-6783. Docket No. 90-ANE-27. Applicability: Pratt & Whitney (PW) PW4050, PW4052, PW4056, PW4060, PW4060A, PW4152, PW4156, PW4158, and PW4460 turbofan engines installed on, but not limited to, Boeing 747, Boeing 767, Airbus A300, and Airbus A310 type aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent a low pressure turbine (LPT) shaft failure that can cause a total loss of thrust and inflight shutdown, accomplish the following: (a) Conduct a one-time eddy current inspection of the inner diameter bore surface of the LPT shaft, between 23 and 44 inches forward of the aft end, in accordance with the Accomplishment Instructions of PW Alert Service Bulletin (ASB) No. PW4ENG 72-328, dated September 28, 1990, for LPT shafts installed on engines identified by serial number in Table 2 of the ASB, prior to October 31, 1990. (b) Conduct a one-time eddy current inspection of the inner diameter bore surface of the LPT shaft, between 23 and 44 inches forward of the aft end, in accordance with the Accomplishment Instructions of PW ASB No. PW4ENG 72-328, dated September 28, 1990, for LPT shafts installed in engines identified by serial number as follows, excluding those shafts subject to the requirements of (a) above, prior to December 31, 1990. Engine Model Engine Serial Number PW4050 P724201 through P724208, inclusive PW4052 P723701 through P723706, inclusive PW4056 P717501 through P717696, inclusive, except P717692 PW4060 P724101 through P724164, inclusive PW4060A P724501 through P724506, inclusive PW4152 P717701 through P717775, inclusive PW4156 P717201 through P717205, inclusive PW4158 P724001 through P724076, inclusive PW4460 P723801 through P723828, inclusive (c) Remove from service, prior to further flight, LPT shafts found cracked during inspections conducted in accordance with (a) or (b) above. (d) LPT shafts which have been previously inspected in accordance with PW Special Instruction 70F-90, dated September 14, 1990, are in compliance with this AD. (e) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (f) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD may be approved by the Manager, Engine Certification Office, ANE-140, Engine & Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803. The eddy current inspections of the LPT shafts shall be done in accordance with the following PW document: Document Number Page Number Issue/Revision Date PW ASB PW4ENG 72-328 All Original 9/28/90 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 Pratt & Whitney, Publications Department, P. O. Box 611, Middletown, Connecticut 06457. Copies may be inspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, FAA, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the Office of the Federal Register, 1100 L Street, NW, Room 8301, Washington, DC 20591. This amendment (39-6783, AD 90-22-02) becomes effective on October 18, 1990.
2009-25-03: 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: The manufacturer has advised of receiving a report that during start up on ground a RH propeller gear box (PGB) on the airship has failed resulting in free rotation of the propeller. Investigation performed by the manufacturer revealed that the bevel gear in the propeller gearbox had cracked near the hub area. During an extensive metallurgical investigation of the cracked bevel gear some different manufacturing deviations outside of the specifications were detected. Deviations in the heat treatment, wall thickness of the bevel gear near the hub area, and score marks caused during the production process have been established as causal factors for this failure. We are issuing this AD to require actions to correct the unsafe condition on these products.
78-05-07: 78-05-07 SIKORSKY: Amendment 39-3147. Applies to Model S-61N rotorcraft prior to and including Sikorsky Serial No. 61805. Compliance required within the next 300 hours time in service after the effective date of this AD, unless already accomplished. To prevent inadvertent landing gear retraction, modify the main landing gear electrical system in accordance with Sikorsky Service Bulletin 61B55-40 or later FAA approved revision. The manufacturer's 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 these documents from the manufacturer may obtain copies upon request to Sikorsky Aircraft, Commercial Customer Service, Stratford, Connecticut 06602. These documents may also be examined at the Office of the Regional Counsel, New England Region, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment becomes effective April 13, 1978.
2009-26-10: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: As a result of the Movable Flap Track Fairing (MFTF) 6 crack findings (ref. AD 2008-0216), a detailed review has been launched for all MFTF 2 to 6. This investigation has revealed some cracking at MFTF 4 pivot support-ring. This condition, if not corrected, could lead to in-flight loss of MFTF 4, potentially resulting in injuries to persons on the ground. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
48-24-01: 48-24-01 MARTIN: Applies to Model 202 Aircraft Serial Numbers 9125 Through 9133, and 9158 Through 9167. Compliance required by September 1, 1948. Reinforce the Nos. 1 and 2 (top and middle) rudder hinge brackets and fairing in accordance with Martin Service Bulletin No. 31, dated March 22, 1948. Other reinforcements shown to be equivalent to those covered in the Service Bulletin will also be acceptable.
60-04-07: 60-04-07 WRIGHT: Amdt. 106 Part 507 Federal Register February 20, 1960. Applies to All 977C9HD1, 2 and 3 Engine Models Installed in Helicopters. Compliance required at first engine overhaul after March 15, 1960, but not later than October 31, 1960. To alleviate failures of the master rod assemblies, strengthened master and articulating rods with associated parts must be installed in accordance with the instructions contained in Wright Aeronautical Division Service Bulletin No. C9-353.
67-06-02: 67-06-02 BRITISH AIRCRAFT: Amdt. 39-359 Part 39 Federal Register March 1, 1967. Applies to Model BAC 1-11 200 and 400 Series Airplanes. Compliance required as indicated. To prevent the inflight loss of engine stub access panels, accomplish the following: (a) Within the next 150 hours' time in service after the effective date of this AD, accomplish the following: (1) Inspect the access panels Camloc fasteners and sliding bar type fasteners to ensure they are locked in accordance with British Aircraft Corporation One-Eleven Alert Service Bulletin No. 54-A-PM-2649, dated September 22, 1966, or later ARB-approved issue. (2) Inspect the access panels for vertical movement. If found in excess of 0.1 inch, replace the Camloc grommets with steel grommets P/N 4002-0 or an FAA-approved equivalent. (b) Within the next 600 hours' time in service after the effective date of this AD, accomplish the following: (1) Inspect to ensure that each of the clearances betweenthe leading and trailing edges of L.H. panels F3-3 and F3-4 and R.H. panels F3-24 and F3-25 is not greater than 0.080 inch. If any clearance is found to be greater, remove and reposition the two spigots P/N AB 16-549 to the forward face of the channel section and open up and countersink the existing rivet holes to receive 5/32 inch diameter 100 degree rivets to material Specification L86, or FAA- approved equivalent, to attach the spigots. (2) Inspect to ensure that a continuous line has been stenciled across the sliding bar fastener boss and skin in accordance with British Aircraft Corporation One-Eleven Alert Service Bulletin No. 54-A-PM-2649, dated September 22, 1966, or later ARB-approved issue. If this has not been accomplished, aline the locked indicator marks on the sliding bar fastener boss and the skin and stencil, with paint of a contrasting color to the local finish, a continuous line across the boss and the skin in line with the existing locked indicator marks.(3) Operate the sliding bar fasteners to ensure that the snapover locking action is positive and that no looseness exists. Lubricate and operate several times, in accordance with British Aircraft Corporation One-Eleven Alert Service Bulletin No. 54-A-PM-2649, dated September 22, 1966, or later ARB-approved issue, any fastener that fails to meet this requirement. Replace any fastener that fails to respond to this treatment with a new part. (4) Replace Camloc fastener grommets with steel grommets P/N 4002-0 or an FAA-approved equivalent. (c) Within the next 1,800 hours' time in service after the effective date of this AD, remove the access panels and re-lubricate the sliding bar fasteners using grease to Process Specification VP 93/2, or an FAA-approved equivalent. (d) The requirements of this AD do not apply to those airplanes that have been modified in accordance with British Aircraft Corporation One-Eleven Alert Service Bulletin No. 54-A-PM-2649, dated September 22, 1966, or later ARB-approved issue. This directive effective March 6, 1967.
91-11-09: 91-11-09 CESSNA: Amendment 39-7006. Docket No. 90-CE-66-AD. Supersedes AD 88-22-01. Applicability: Model T303 airplanes (all serial numbers), certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent failure of the nose gear actuator attachment fitting, accomplish the following: (a) Within the next 50 hours time-in-service (TIS) after the effective date of this AD or upon the accumulation of 550 hours TIS, whichever occurs later, and thereafter at intervals of 50 hours TIS, inspect the nose gear actuator attachment fitting for cracks in accordance with Cessna Service Bulletin (SB) MEB88-4, dated August 5, 1988. (1) If any of the legs of the fitting are cracked, prior to further flight, replace the cracked fitting with an improved fitting, part number P/N 2543004-5, in accordance with Cessna SB MEB88-4, Revision 2, dated October 26, 1990. (2) If no cracks are found, return the airplane to service and continue the repetitive inspections specified in paragraph (a) of this AD. (b) Within the next 12 calendar months after the effective date of this AD, unless already accomplished in accordance with paragraph (a)(1) of this AD, replace the nose gear actuator attachment fitting, P/N 2543004-1 or 2543004-3, with an improved fitting, P/N 2543004-5, in accordance with Cessna SB MEB88-4, Revision 2, dated October 26, 1990. (c) The inspections required by paragraph (a) of this AD are no longer required when the nose gear actuator attachment fitting has been replaced with P/N 2543004-5 in accordance with paragraphs (a)(1) or (b) of this AD. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a location where the requirements of this AD can be accomplished. (e) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (f) All persons affected by this directive may obtain copies of the documents referred to herein upon request to the Cessna Aircraft Company, Customer Services, P.O. Box 7704, Wichita, Kansas 67277; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This amendment supersedes AD 88-22-01, Amendment 39-6041. This amendment (39-7006, AD 91-11-09) becomes effective on June 28, 1991.
55-03-01: 55-03-01 CONVAIR: Applies to All Model 340 Aircraft. Compliance required as indicated. A 340 was involved in a flight accident due to the AN 310-4 nut coming off of the AN 174-33 bolt that attaches the left-hand elevator servo tab idler to the structure resulting in flutter of the elevator servo tab, with subsequent failures at the elevator control system and loss of elevator control. As a precautionary measure, the following is required: As soon as practical but not later then the next 15 hours service, inspect the control systems for the servo tab on the left elevator and on the right elevator trim tab on all 340 aircraft. Inspection is to cover all nuts, bolts, idlers, bellcranks, pushrods and general security check of all nuts and bolts.
2009-26-03: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 737-300, -400, -500, -600, -700, -700C, -800, and -900, and 747-400 series airplanes; and Model 757, 767, and 777 airplanes. This AD requires modifying the static inverter by replacing resistor R170 with a new resistor and relocating the new resistor. This AD results from evaluation of the carbon resistor, which revealed a failure mode that can cause the resistor to ignite, involving adjacent capacitors as well. We are issuing this AD to prevent a standby static inverter from overheating, which could result in smoke in the flight deck and cabin and loss of the electrical standby power system.
2009-24-16: 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: Zinc-coated starter ring gears installed on Solo 2625 01 and 2625 02 engines have shown to be prone to cracking. For that reason, AD 2009-0169-E has been published in July 2009. From that date, collected in-service data have been revealed that painted starter ring gears with lightening holes are also subject to cracks. The reason for these cracks is still unknown at the present time. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
64-21-04: 64-21-04 FAIRCHILD: Amdt. 814 Part 507 Federal Register September 11, 1964. Applies to Model F-27 Aircraft, Serial Numbers 1 through 97, and 99. Compliance required as indicated. There have been cracks found in the rib caps and stringers in the vertical stabilizer where the rib at water line 201.575 attaches to the hat and zee section stringers. To preclude the condition from developing in other such aircraft, accomplish the following: (a) On aircraft with less than 1,000 hours' time in service on the effective date of this AD, comply with (c) prior to the accumulation 1,050 hours' time in service, unless already accomplished within the last 200 hours' time in service, and thereafter at intervals not to exceed 200 hours' time in service from the last inspection. (b) On aircraft with 1,000 or more hours' time in service on the effective date of this AD, comply with (c) within the next 50 hours' time in service after the effective date of this AD, unless already accomplished within the last 150 hours' time in service, and at intervals thereafter not to exceed 200 hours' time in service from the last inspection. (c) Inspect for cracks in the skin, rib and stringers on both sides of the vertical stabilizer along and directly adjacent to water line 201.575, from the trailing edge of the stabilizer, forward, to the auxiliary spar (located directly aft of the stabilizer leading edges). Use X-ray or dye penetrant in conjunction with at least a 10-power glass, or FAA-approved equivalent inspection. If cracks are found accomplish (d). (d) Repair cracked parts in accordance with an FAA-approved repair or replace cracked parts with a part of the same part number or an FAA approved equivalent before further flight except that a ferry flight may be made in accordance with the provisions of CAR 1.76. (e) The repetitive inspection intervals specified in (a) and (b) may be increased from 200 hours' time in service to 800 hours' time in service from the last inspection on aircraft that are modified in accordance with Fairchild Service Bulletin No. 55-4 dated March 1, 1963, 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. (Fairchild Service Bulletins 55-4 dated March 1, 1963, and 55-5 dated March 1, 1963, cover this subject.) This supersedes AD 63-10-2. This directive effective September 12, 1964.
2021-23-15: The FAA is adopting a new airworthiness directive (AD) for certain Airbus SAS Model A319-111, -112, -113, -114, -115, -131, -132, and -133 airplanes; Model A320-211, -212, -214, -216, -231, -232, and - 233 airplanes; and Model A321-111, -112, -131, -211, -212, -213, -231, and -232 airplanes. This AD was prompted by a report that during re- engineering of galley G5, a 9G forward full scale qualification test was performed, and the door of the waste compartment opened before the required load was reached. This AD requires modifying the waste compartment door of each affected galley, as specified in a European Aviation Safety Agency (EASA) AD, which is incorporated by reference The FAA is issuing this AD to address the unsafe condition on these products.
88-11-10 R1: 88-11-10 R1 AIRBUS INDUSTRIE: Amendment 39-5931 as amended by Amendment 39- 6005. Applies to Model A300 series airplanes, equipped with General Electric CF6-50 engines, without a secondary latching system on core cowl doors, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent structural damage to the airplane due to engine core cowl door separation, accomplish the following: A. Within 10 days after the effective date of this AD, check the core cowl door latches of each engine once each day, and re-check after each core cowl door is opened and subsequently closed. 1. If the latch is open, before further flight, properly close the latch. 2. If the latch will not engage, adjust the latch, in accordance with the A300 maintenance manual. 3. If the latch cannot be properly adjusted, replace the latch prior to further flight. B. The checks required by paragraph A., above, may be discontinued after a secondary latching system is installed, in accordance with Airbus Industrie Service Bulletin A300- 71-053, Revision 2, dated January 6,1987. C. 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. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Standardization Branch, ANM-113. D. 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 the modification required by this AD. All persons affected by this directive who have not already received the appropriate service information from the manufacturer, may obtain copies upon request to Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. This information may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This AD revises AD 88-11-10, Amendment 39-5931. This amendment, 39-6005, becomes effective September 6, 1988.
60-23-04: 60-23-04 PRATT & WHITNEY: Amdt. 215 Part 507 Federal Register November 4, 1960. Applies to Turbo Wasp JT3C-7 Turbojet Engines Prior to Serial Number P632425B and all JT3C-6 Turbojet Engines. Compliance required at first engine overhaul after December 15, 1960. To prevent failure of the bevel accessory drive gearshaft and resultant loss of engine power, replace P/N 350460 gearshaft with P/N 401493 gearshaft.
92-15-05: 92-15-05 FOKKER: Amendment 39-8299. Docket No. 92-NM-39-AD. Applicability: Model F28 Mark 0100 series airplanes, serial numbers 11340 through 11348 airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent wire burnthrough or meltdown, that could render the fire detection system ineffective in detecting fire in the engine, accomplish the following: (a) Within 30 days after the effective date of this AD, remove existing engine fire detection wires and replace them with fire resistant ones, in accordance with Fokker Service Bulletin SBF100-26-004, dated August 23, 1991. (b) 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, Standardization Branch, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch. (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) The removal and replacement shall be done in accordance with Fokker Service Bulletin SBF100-26-004, dated August 23, 1991. 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 Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. (e)This amendment becomes effective on September 8, 1992.
90-12-10: 90-12-10 AIRBUS INDUSTRIE: Amendment 39-6624. Docket No. 90-NM-100-AD. Applicability: Model A320 series airplanes, as listed in Airbus Industrie Service Bulletin A320-28-1024, Revision 1, dated February 20, 1990, certificated in any category. Compliance: Required within 100 landings after the effective date of this AD, unless previously accomplished. To ensure proper operation of the fuel quantity indicating system and to prevent the possibility of a spark in the fuel system in the event of a lightning strike, accomplish the following: A. Verify the clearance between fuel probes 23QT1, 23QT2, and 28QT1, and the adjacent structures; install an insulating sleeve on the vent pipe bonding lead; and adjust fuel probe clearance, in accordance with Airbus Industrie Service Bulletin A320-28-1024, Revision 1, dated February 20, 1990, or All Operators' Telex (AOT) 28/89/03, dated November 20, 1989. B. 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. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6624, AD 90-12-10) becomes effective on June 20, 1990.
71-10-02: 71-10-02 AEROSTAR: Amdt. 39-1206. Applies to all Model 600 and 601 airplanes certificated in all categories. Compliance required within the next 10 hours' time in service after the effective date of this AD, unless already accomplished. To assure all over-voltage relays are properly grounded, accomplish the following: A. On S/N 61-0001 and 60-0001 through 60-0005: (1) Inspect each overvoltage relay installation and assure that the overvoltage relay base and the aircraft structure are in direct contact. (2) If the installation appears satisfactory from the inspection, use an ohmmeter to determine that low resistance continuity of less than one (1) ohm exists between the base of each overvoltage relay and the airplane structure. (3) If a satisfactory ohmmeter indication is not achieved, ground the overvoltage relay base to the aircraft structure using good aircraft grounding practice. B. On all others; (1) Inspect each installation of thecrimp wire terminal which connects to the overvoltage relay base and assure that the crimp terminal and relay base are in contact. (2) If the results of the inspection are not satisfactory, rearrange the components as follows: (a) Install the nylon bearing from the bottom side of the overvoltage relay base so that it isolates the overvoltage relay base from the aircraft structure. (b) Install the crimp wire terminal over the nylon bearing so that it is in electrical contact with the overvoltage relay base. (c) Install the AN960 washer over the nylon bearing so it is in contact with the crimp wire terminal. (d) Install the two nylon washers on top of the AN960 washer and insert the NAS221 screw through the two nylon washers and the nylon bearing, and tighten the screw into the nut-plate provided. (e) Assure that the nylon bearing is seated properly to prevent electrical contact of the crimp wire terminal, the AN960 washer, and the overvoltage relay base with the NAS221 screw and the aircraft structure. (f) Repeat the above procedure for the opposite side of each overvoltage relay except for elimination of the crimp wire terminal. (3) Using an ohmmeter, determine that low resistance continuity of one (1) ohm or less exists between the base of each overvoltage relay and the airplane structure with the respective alternator switch in the "on" position. Low resistance continuity should not exist with the respective alternator switch in the "off" position. (4) If the existing installation of either overvoltage relay will not give a satisfactory check and cannot be changed as described above, accomplish an equivalent FAA approved modification. If Aerostar Aircraft Corporation Service Bulletin No. S.B. 600-24 dated November 23, 1970, or later FAA approved revision, has been compiled with and an appropriate entry made in the airplanes' permanent maintenance record, the requirements of this AD will be considered satisfied. This amendment becomes effective on May 14, 1971.
2021-23-07: The FAA is adopting a new airworthiness directive (AD) for certain Saab AB, Support and Services Model SAAB 340B airplanes. This AD was prompted by a report that the circuit breaker for the emergency cabin lighting tripped without fault in the system. This AD requires replacing a certain circuit breaker with a part having a higher rating, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
48-13-03: 48-13-03 PIPER: Applies to PA-12 Airplanes With Battery Holddown of Metal Bracket With Fiber Insulation at Its End. Compliance required by October 1, 1948. To eliminate battery short circuits caused by defective battery holddown brackets, replace brackets by wood blocks, Piper P/N 84682-3 and 84682-9 or equivalent. (Piper Service Bulletin No. 105 dated February 18, 1948, covers this same subject.)
77-19-03: 77-19-03 HUGHES HELICOPTERS: Amendment 39-3038. Applies to Hughes Model 369D helicopters, certificated in all categories, which have P/N 369D21200 main rotor hub serial numbers 0001 through 0178 installed. Compliance required as indicated. To prevent failure of the main rotor retention straps which can result in loss of a main rotor blade accomplish the following: A. Before further flight after the effective date of the AD, visually inspect the lead leg of each main rotor retention strap pack, P/N 369D21210, at the inboard ends to determine if two adjacent metallic anti-fretting laminates P/N 369D21271-5, -51, -53 or -55, have been installed on the upper side of the lead leg of any strap pack. (1) If two adjacent metallic anti-fretting laminates have been installed, visually inspect to determine if one or both are cracked or have been broken off where they extend from the hub. (i) If no anti-fretting laminate is found cracked or broken, repeat the inspections in (1), above, at times in service not to exceed 25 hours. This repetitive inspection can be discontinued when the main rotor hub, P/N 369D21200, has been replaced by a serviceable hub having only one anti-fretting laminate on the upper side of the lead leg of each main rotor strap pack. (ii) If any anti-fretting laminate is found cracked or broken, further visually inspect the main rotor retention strap laminates adjacent to the anti-fretting laminates. (a) If any strap laminate is broken or cracked, replace the main rotor hub, P/N 369D21200, before further flight. (b) If no strap laminate is cracked or broken, replace the main rotor hub, P/N 369D21200, within ten hours time in service. (2) If the inspection reveals that only one metallic anti-fretting laminate has been installed on the upper lead leg of each strap pack no further action is required by this AD. This amendment becomes effective September 27, 1977, for all persons except those to whom it was made effective by telegram dated August 12, 1977, which contained this amendment.
76-24-02: 76-24-02 LAKE AIRCRAFT, DIVISION OF CONSOLIDATED AERONAUTICS, INC: Amendment 39-2774. Applies to all Lake Model LA-4-200 airplanes certified in all categories equipped with Stewart Warner Model 8406J engine oil coolers. Compliance required as indicated unless already accomplished. To preclude possible oil cooler failures allowing rapid loss of engine oil which could result in engine stoppage, accomplish the following: A. Before next flight of the affected airplanes, inspect the engine oil cooler (fluid fitting side) to determine whether it is a Model 8406J S/N 101 through 1500. 1. If the oil cooler is a Model 8406J S/N 101 through 1500, prior to further flight, replace this cooler with an FAA approved oil cooler not of the above model and serial number. 2. If the oil cooler is not of the model and serial number listed above, make an entry in the aircraft maintenance records indicating that this airworthiness directive has been accomplished and the airplane may be returned to service. B. The inspection and maintenance record entry required by paragraph A2 may be accomplished by holder of a pilot's certificate issued under Part 61 of the Federal Aviation Regulations on any aircraft owned or operated by him. C. Equivalent methods of compliance with this AD may be approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. NOTE: A ferry permit to accomplish a needed oil cooler replacement may be issued under the provisions of FAR 21.197 by FAA District Offices, with appropriate limitations. This amendment becomes effective immediately upon publication in the Federal Register for all persons except those to whom it was made effective immediately upon receipt of airmail letter dated October 15, 1976.
2009-26-08: We are superseding an existing 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: Late in 2002 the manufacturer advised CASA of another Nomad accident which was possibly caused by aileron flutter with the flaps at 38 degrees. This, along with the other flutter incidents, has resulted in the manufacturer issuing ANMD-57-18 Issue 1 as a precautionary measure while they further investigate the issue. The manufacturer has now completed their investigation and issued Alert Service Bulletin ANMD-27-53 to modify flap actuation linkages to restore the necessary rigidity to the outboard flap, and hence the aileron. The unacceptable flexibility of the outboard flap mechanism allows flutter to occur in extreme circumstances. We areissuing this AD to require actions to correct the unsafe condition on these products.