Results
2009-19-05: We are adopting a new airworthiness directive (AD) for certain Boeing Model 747 airplanes. This AD requires repetitive inspections for cracking of the fuselage frames in section 41, and corrective actions if necessary. This AD results from reports of cracking in fuselage frames made of 2024 aluminum alloy that were installed during previous modification of the frames in section 41 and during production. We are issuing this AD to detect and correct frame cracks, which could result in cracking of the adjacent fuselage skin and consequent rapid decompression of the airplane.
81-09-07: 81-09-07 SIKORSKY AIRCRAFT: Amendment 39-4099. Applies to S-76A series helicopters certificated in all categories with P/N 76150-09000 series and P/N 76150-09100-041, -042, -043, main rotor blades. For main rotor blades with 340 or more hours time in service, compliance required within the next 25 hours time in service after the effective date of this AD, unless already accomplished. For main rotor blades with less than 340 hours time in service on the effective date of this AD, compliance required before the accumulation of 365 hours time in service. To prevent operation with cracked bolts in the main rotor blade tip plate attachment joint, accomplish the following: 1. In accordance with Sikorsky Alert Service Bulletin No. 76-65-23, dated April 16, 1981, replace the four NAS624H6 bolts which mate the 76150-09030 tip plate assembly with the 76150-09000 or 76150-09100 main rotor blade, per paragraphs D(1) through D(8), and subsequently inspect for torque per paragraph D(9), or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region. 2. Report within 24 hours any discrepancies found during the rework and inspections required herein, including main rotor blade time in service, to the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803. Reporting approved by the Office of Management and Budget under OMB No. 04-401 74. 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) (l). All persons affected by this directive, who have not already received these documents from the manufacturer, may obtain copies upon request to Sikorsky Aircraft, Division of United Technologies Corporation, Stratford, Connecticut 06602. These documents may also beexamined at FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. This AD supersedes AD 80-14-05. This amendment becomes effective May 7, 1981.
2009-20-02: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 767-200 and -300 series airplanes, that requires replacing certain door-mounted escape slides and slide- raft assemblies with new slide-raft assemblies. This AD also requires the following actions, as applicable: replacing certain escape system latches with new latches; modifying or replacing certain counterbalance assemblies with new counterbalance assemblies; and adjusting the door counterbalance system. The actions specified by this AD are intended to prevent the escape slides and slide-rafts of the forward and mid-cabin entry and service doors from being too steep for evacuation in the event that the airplane rotates onto the aft fuselage into the extreme tip-back condition. In the extreme tip-back condition, the forward and mid-cabin exits could result in steeper sliding angles, which could cause injury to passengers and crewmembers during an emergency evacuation. This action is intended to address the identified unsafe condition.
60-07-07: 60-07-07 VICKERS: Amdt. 116 Part 507 Federal Register March 23, 1960. Applies to Viscount Model 745D Serial Numbers 103 to 107 Inclusive, 109 to 134 Inclusive, 136 to 139 Inclusive, 183, 184, 185, 191, 198 to 217 Inclusive, 231, 232, 233, 234, 285, 334. Compliance required as indicated. As a result of instances of corrosion which have been found to occur in the skin to wing spar boom attachment holes, it is necessary that aircraft built to Modification D.953 standard have oversized skin to spar attachment bolts installed in the inner and outer top booms (unbushed) in accordance with part (e) of Vickers Modification Bulletin No. D.2081. The oversize bolts are of S.80 material cadmium plated either 1/32-inch or 1/16-inch oversize as required, depending upon the state of the holes; this is determined by the inspections detailed in the Preliminary Technical Leaflet 197. Thickol is used as sealant when the oversize bolts are fitted. Compliance: (a) Bolt installation in accordance with Mod. D.2081, part (e), is required within 10,000 hours' time in service or five calendar years from the date of aircraft manufacture, whichever occurs first, unless a satisfactory sampling inspection as covered in (b) is accomplished. (b) Bolt installation in accordance with Mod. D.2081, part (e), may be accomplished by an operator within 13,000 hours' time in service provided a satisfactory sampling inspection for corrosion is conducted on five complete aircraft sets of top skin to spar attachment bolts. This sampling inspection must be conducted on the operator's aircraft which have between 9,000 and 10,000 hours' time in service. Corresponding spar bolt holes also must be inspected for corrosion when the bolts are removed. If any corrosion is found on the bolts or in the spar bolt holes, Mod. D.2081, part (e) must be accomplished within 10,000 hour's time in service or five calendar years from the date of aircraft manufacture, whichever occurs first.(Modification Bulletin D.2081 and Preliminary Technical Leaflet No. 197 Issue 5 (700 Series) cover this subject. This supersedes AD 60-01-08. Revised March 13, 1964.
63-08-02: 63-08-02 DOUGLAS: Amendment 39-630. McDonnell Douglas. Applies to McDonnell Douglas Model DC-8 Series Aircraft equipped with P/N 3703218 (no dash number) elevator control tab push rod assembly. \n\n\tAs a result of damage near the midpoint of the elevator control tab push rod assembly due to wear from rubbing against the guide support assembly and the guide support attach rivets, accomplish the following: \n\n\t(a) Unless already accomplished, within the next 300 hours' time in service after the effective date of this AD: \n\n\t\t(1) Remove both left and right-hand elevator control tab push rod assemblies, P/N 3703218, and conduct a close visual inspection of the push rods for evidence of wear due to contact of the push rod with guide support assembly, P/N 5708625-3. \n\n\t\t(2) Push rods showing evidence of wear shall, prior to further flight: \n\n\t\t(i) be replaced either with an undamaged part; or \n\n\t\t(ii) be reworked in accordance with the rework procedures outlined in Figure(1) of Step (7) of Douglas DC-8 Service Bulletin No. 27-51 Reissue No. 1 dated September 25, 1962, or an FAA approved equivalent. Push rods showing evidence of wear which require removal of material in excess of 0.025 inch in depth and one inch in length by 0.375 inch in width on one side of the push tube, or which have dents or sharp gouges, or are found worn or cracked in more than one area may not be reworked and must be replaced. When push rod assemblies are reworked, they must be reinspected using dye penetrant method or equivalent, to insure that no cracks exist after the rework is accomplished. \n\n\t\t(3) Following reinstallation of push rod assemblies, and before further flight, conduct an initial check for clearance per Figure (1), Step (1); and , as necessary, accomplish the adjustment and rework outlined in Figure (1), Steps (2) through (6), of Douglas DC-8 Service Bulletin NO. 27-51, Reissue No. 1, dated September 25, 1962, or FAA approved equivalent. \n\n\t(b) At intervals of not less than 400 nor more than 600 hours' time in service following the initial clearance check prescribed by (a)(3), unless other inspection intervals have been approved for an operator by the Chief, Engineering & Manufacturing Branch, FAA Western Region, remove and again inspect rods replaced or reworked per (a)(2) for any evidence of wear or contact with the guide support assembly. Any rods showing evidence of wear must be reworked or replaced per (a)(2), and the reinstallation clearance check and such adjustment and rework provisions of (a)(3), as found necessary, shall be accomplished. \n\n\t(c) If, subsequent to compliance with (a), an airplane is altered by changing an elevator, elevator control tab, elevator control tab push rod assembly, or any combinations of these, the following new procedure is required: \n\n\t\t(1) Prior to further flight conduct an initial check for clearance and any necessary adjustment and rework as described in (a)(3). \n\n\t\t(2) At intervals of not less than 400 nor more than 600 hours' time in service following the initial clearance check required by (c)(1), unless other inspection intervals have been approved for an operator by the Chief, Engineering & Manufacturing Branch, FAA Western Region, the elevator control tab push rod assembly associated with this change shall be removed and inspected for any evidence of wear or contact with the guide support assembly as described in (a)(1). Any rods showing evidence of wear must be reworked or replaced as indicated in (a)(2) and the reinstallation clearance check and adjustment and rework provisions of (a)(3), as found necessary, shall be accomplished. \n\n\t(d) The periodic reinspection prescribed by (b) and (c)(2) may be discontinued when: \n\n\t\t(1) It is determined that no wear or contact with the guide support assembly has developed during the preceding reinspection interval, or \n\n\t\t(2) The elevator control tab push-rod assembly Douglas P/N 3703218 is replaced with Douglas P/N 3703218-501, in accordance with the procedure outlined in DC-8 Service Bulletin No. 27-150 dated November 5, 1963, or by an FAA approved equivalent part and procedure. \n\n\tNOTE: The P/N 3703218-501 steel push rods presently being installed during production have a smaller diameter than the original P/N 3703218 (no dash number) aluminum push rods. \n\n\t(e) Upon request of the operator an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operation period of the operator if the request contains substantiating data to justify the increase or decrease for such operator. \n\n\t(Douglas DC-8 Service Bulletins No. 27-51, Reissue No. 1, dated September 25, 1962, and No. 27-150, dated November 5, 1963, cover this same subject.) \n\n\tThis directive effective April 18, 1963. \n\n\tRevised November 9, 1963. \n\n\tRevised April 15, 1964. \n\n\tRevised September 16, 1968.
2009-18-11: 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: Following a red illuminated "DOOR NOT LOCKED" status light indication on the door lock indication panel after lift off, the cabin crew operated the door lock handle. This resulted in inadvertent opening of the downward opening passenger door in flight. * * * After inspection, it was found that the false red light might be the result of an incorrect clearance between lever Part Number (P/N) A26997-003 and the Up-Limit Switch. If the Up-Limit Switch has an incorrect clearance, the combination with cabin differential pressure build-up after lift-off might result in a false steady illuminating red "DOOR NOT LOCKED" indication on the Door Indication Panel.* * * * * * * * The unsafe condition is inadvertent opening of the door lock handle in flight, which could result in rapid decompression of the airplane or ejection of a passenger or crewmember through the door. We are issuing this AD to require actions to correct the unsafe condition on these products.
79-19-03: 79-19-03 SHORT BROTHERS LIMITED: Amendment 39-3553. Applies to Model SD3-30 airplanes, S/Nos. SH 3004, 3005, 3006, 3007, and 3008, certificated in all categories. Compliance is required prior to the accumulation of 10,000 flights, unless already accomplished. To prevent failure of the fuselage skin panel buttstraps, accomplish the following: (a) Replace or reinforce the original fuselage skin panel buttstraps in accordance with Section 2, "Accomplishment Instructions" of Short Brothers Ltd. Service Bulletin SD3-53- 28, dated May 5, 1978, or an equivalent approved by the Chief, Aircraft Certification Staff, AEU- 100, Europe, Africa, and Middle East Region, Federal Aviation Administration, c/o American Embassy, Brussels, Belgium. (b) For purposes of complying with this AD, a flight is defined as one takeoff and one landing. This amendment becomes effective October 4, 1979.
2009-17-04: 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: One case of elevator servo-control disconnection has been experienced on an aircraft of the A320 family. Failure occurred at the servo-control rod eye-end. Further to this finding, additional inspections have revealed cracking at the same location on a number of other servo-control rod eye-ends. In one case, both actuators of the same elevator surface were affected. * * * A dual servo-control disconnection on the same elevator could result in an uncontrolled surface, the elevator surface being neither actuated nor damped, which could lead to reduced control of the aircraft. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
64-20-03: 64-20-03 VICKERS: Amdt. 796 Part 507 Federal Register August 20, 1964. Applies to Viscount Model 810 Series Aircraft. Compliance required as indicated. Fatigue cracking has occurred in the rib structure of the inboard rib at Station 96 of the inboard nacelles, illustrated in Figure 1 of Preliminary Technical Leaflet No. 113 (800/810 Series). To preclude further failures, accomplish the following in accordance with the PTL referenced herein or FAA approved equivalent: (a) Within 275 landings after the effective date of this AD unless already accomplished within the last 225 landings, conduct dye penetrant or an FAA approved equivalent inspection for cracks in the rib reinforcing plate and rib web plate in accordance with PTL 113. (b) If no cracks are found, reinspect at intervals not exceeding every 500 landings from the last inspection until Modification FG.1960 Part "B" is accomplished, after which no further inspection for this defect will be required. (c)If cracks are found during the initial inspection described in paragraph (a), accomplish paragraphs (e), (f), or (g), as appropriate, within 275 landings after the effective date of this AD. (d) If cracks are found during a reinspection, accomplish paragraphs (e), (f), or (g), as appropriate, within 275 landings from the time the cracks are found. (e) If a crack is found in the rib reinforcing plate only, incorporate the repair scheme Figures 2 and 3 of PTL 113 or an FAA approved equivalent and reinspect within every 1,500 landings to ensure that there is no progression of damage in the reinforcing plate and no initiation of damage in the web plate. These repetitive inspections may be discontinued after incorporation of Modification FG. 1960 Part "C". (f) If a crack is found in the rib web plate only, incorporate Mod. FG.1960 Part "B" and reinspect the rib web plate within every 3,000 landings to ensure that cracking has not been initiated in the reinforcing plate. Theserepetitive inspections are no longer necessary after the incorporation of Mod. FG.1960 Part "C". (g) If cracks are found in both the reinforcing plate and the rib web plate, incorporate Mod. FG.1960 Part "C". (Vickers-Armstrongs Preliminary Technical Leaflet No. 113 Issue 2 (800/810 Series) and Modification FG.1960 cover this subject.) This directive effective September 21, 1964.
71-13-03: 71-13-03 HAWKER SIDDELEY AVIATION: Amdt. 39-1230 as amended by Amendment 39-1259. Applies to Model DH-114 "Heron" airplanes. Compliance is required as indicated. To prevent possible failure of the Dunlop compressed air bottles used in the emergency landing gear extension system and emergency braking system, accomplish the following on or before August 31, 1971. (a) For all airplanes, inspect the air bottle (P/N AH.7360 or AC.11038) used in either of the emergency landing gear extension system air bottle assemblies (P/N AC.11768) located under the pilot's seat: If the air bottle was manufactured before January 1, 1959, replace the air bottle assembly with a serviceable assembly of the same part number which incorporates an air bottle (P/N AH.7360 or AC.11038) manufactured on or after January 1, 1959. The date of manufacture is etched on the collar of the bottle. (b) For airplanes which have incorporated Modification 281 (Emergency Braking System), inspect the airbottle (P/N AC.10685 or AC.11038) used in the emergency braking system air bottle assembly (P/N ACM.16784) located on the left forward face of the crew cabin sloping bulkhead. If the air bottle was manufactured before January 1, 1959, replace the air bottle assembly with a serviceable assembly of the same part number which incorporates an air bottle (P/N AC.10685 or AC.11038) manufactured on or after January 1, 1959. The date of manufacture is etched on the collar of the air bottle. (Hawker Siddeley Technical News Sheet, Series: Heron (114), No. S.6, Issue 2, covers this subject.) Amendment 39-1230 became effective July 30, 1971. This Amendment 39-1259 becomes effective July 31, 1971.
82-12-02 R1: 82-12-02 R1 BRITISH AEROSPACE (HAWKER SIDDELEY): Amendment 39-4392 as amended by Amendment 39-5058. Applies to British Aerospace Model HS/BH/DH 125 up to and including series 700 airplanes certificated in all categories except those airplanes incorporating Modification 252772. To prevent structural failure of the flap, accomplish the following within the next 100 hours time in service after the effective date of this AD or before the accumulation of 600 hours time in service on the airplane, whichever is later, unless already accomplished. 1. Visually inspect the flap outboard hinge nose ribs for cracks in accordance with the instructions in paragraph 2A of British Aerospace, Aircraft Group, 125 Service Bulletin (SB) No. 57-58, Revision 3, dated September 1, 1983. a. If no cracks are found, no further action is necessary. b. If cracks are found during the inspection required by paragraph 1, above, which do not exceed the criteria in paragraph 2A(5) of the service bulletin, reinspect at intervals not exceeding 40 hours time in service from the last inspection until a permanent repair is incorporated. c. If cracks are found during the inspection required by paragraph 1 or during the repetitive inspection required by paragraph 1.b, above, which exceed the criteria in paragraph 2A(5) of the service bulletin, replace the flap with a serviceable flap or contact the manufacturer for instructions to make a permanent repair before further flight. 2. 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. 3. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA Northwest Mountain Region. The manufacturer's specifications and procedures identified and described in this directive are incorporated hereinand made a part hereof pursuant to 5 U.S.C. 552(a)(1). Amendment 39-4392 became effective June 8, 1982. This amendment becomes effective June 17, 1985.
2009-17-05: The FAA is adopting a new airworthiness directive (AD) for Honeywell International Inc. TPE331-10 and TPE331-11 series turboprop engines. This AD requires removing certain first stage turbine disks from service. This AD results from a report of an uncontained failure of a first stage turbine disk that had a metallurgical defect. We are issuing this AD to prevent uncontained failure of the first stage turbine disk and damage to the airplane.
62-26-03: 62-26-03 LOCKHEED: Amdt. 512 Part 507 Federal Register December 6, 1962. Applies to All Models 49, 149, 649, 649A, 749, 749A, and 1049-54 Series Aircraft Incorporating Main Landing Gear Crossheads, P/N 307866 or P/N 288982 Which Have Accumulated 10,000 or More Hours' Time in Service. Compliance required as indicated. To detect fatigue cracking in the 0.25-inch radii adjacent to the one-inch diameter bearing surfaces on main landing gear crossheads, the failure of which would prevent normal extension and retraction of the main landing gear, the following shall be accomplished: (a) Within the next 700 hours' time in service after the effective date of this AD, unless already accomplished within the last 1,800 hours' time in service prior to the effective date of this AD, and thereafter at intervals not to exceed 2,500 hours' time in service from the last inspection, inspect all crossheads as follows: The crosshead shall be removed from the aircraft and inspected by themagnetic particle method or FAA approved equivalent for cracks in the 0.25-inch radii adjacent to the one-inch diameter bearing surfaces. All cracked crossheads shall be replaced with sound ones before the aircraft is returned to service. Crosshead replacement for 1049-54 aircraft shall be P/N 307866 only. Crosshead replacement for 49, 149, 649, 649A, 749, and 749A aircraft shall be either P/N 307866 or P/N 288982. (b) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western 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. (Lockheed Field Service Letter FS/251565L, dated March 31, 1961, covers this same subject.) This directive effective January 7, 1963.
81-01-02 R1: 81-01-02 R1 GOVERNMENT AIRCRAFT FACTORIES (GAF): Amendment 39-3999 as amended by Amendment 39-4147. Applies to Models N22B (Serial Nos. N22B-5 and up) and N24A (Serial Nos. N24A and up), certificated in all categories, which are equipped with fuel selector cables U2000 LAS-2-( ) and U2000 LA-2-( ); i.e., the U2000L series. Compliance required as indicated. To prevent failure of the fuel tank selector valves or the fuel shut-off valves to operate, accomplish the following: (a) Within the next 25 hours time in service after the effective date of this AD, unless already accomplished, inspect the cable sleeves on the fuel tank selector valve and fuel shut-off valve. If they are incorrectly crimped, cracked, or loose, before further flight, repair No. ANMD- 28-11 (hereinafter referred to as the Service Bulletin) dated August 21, 1980, or an FAA-approved equivalent. (b) Cable sleeves on the fuel tank selector cable which have been repaired in accordance with GAF Nomad Alert Service Bulletin No. ANMD-28-11, dated August 21, 1980, or an FAA-approved equivalent, must be visually inspected prior to the first flight of each day in accordance with paragraph 4 of the service bulletin, and replaced prior to the accumulation of 200 hours time in service from the time of repair. (c) If an equivalent is used in complying with paragraph (a) of this AD, that equivalent must be approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. NOTE: All persons affected by this directive who have not already received the Service Bulletin from the manufacturer may obtain copies upon request to the Government Aircraft Factories, 226 Lorimer Street, Port Melbourne 3207 Vic., Australia. These documents may be examined at the FAA, Engineering and Manufacturing District Office, 300 Ala Moana Blvd., Room 7321, Honolulu, Hawaii 96850, or Rules Docket, Room 916, FAA, 800 Independence Ave., S.W., Washington, DC20591. Amendment 39-3999 became effective January 5, 1981. This amendment 39-4147 becomes effective June 22, 1981, as to all persons except those persons to whom it was made immediately effective by telegraphic AD T81-01-02 R1, issued January 16, 1981, which contained this amendment.
83-20-02: 83-20-02 BRITISH AEROSPACE: Amendment 39-4735. Applies to Model BAC 1-11 200 and 400 series airplanes, certificated in all categories. To prevent failure of the No. 1 left and right side fuselage mounted flap beam on aircraft that have not incorporated Modification No. PM5805, accomplish the following unless previously accomplished: A. Inspect and repair or replace, as necessary, the forward and aft spigots of the No. 1 flap beam in accordance with paragraph 2, Accomplishment Instructions, of British Aerospace Alert Service Bulletin No. 53-A-PM5805, Issue 2, dated May 4, 1982, per the following schedule: 1. For aircraft which have accumulated 27,000 or more landings on the effective date of the AD, compliance is required prior to the accumulation of 30,000 landings, or within the next 1,000 landings, whichever occurs later. 2. For aircraft which have accumulated less than 27,000 landings on the effective date of the AD, compliance is required prior to the accumulation of 18,000 landings, or within the next 3,000 landings, whichever occurs later. B. Repeat the actions of paragraph A., above, at intervals not to exceed 6,000 landings. C. Incorporation of modification PM5805 terminates the repetitive inspection requirement of paragraph B., above. D. For the purpose of this AD, and when approved by an FAA maintenance inspector, the number of landings may be computed by dividing each airplane's time in service by the operator's fleet average time from takeoff to landing for the aircraft type. E. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. F. 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. This amendment becomes effective November 3,1983.
2007-03-17 R1: We are revising 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: This Airworthiness Directive (AD) was prompted by reports of loose rivets on frames C18 BIS and C19, which could result in a reduced structural integrity of the tail area. We are issuing this AD to require actions to correct the unsafe condition on these products. DATES: This AD becomes effective September 9, 2009. On September 9, 2009, the Director of the Federal Register approved the incorporation by reference of SOCATA TBM Aircraft Mandatory Service Bulletin SB 70-129, AMENDMENT 1, dated February 2009, listed in this AD. As of March 15, 2007 (72 FR 5923, February 8, 2007), the Director of the Federal Register approved the incorporationby reference of SOCATA TBM Aircraft Mandatory Service Bulletin SB 70-129, dated June 2005, listed in this AD.
83-06-08: 83-06-08 SHORT BROTHERS AND HARLAND LTD.: Amendment 39-4595. Applies to Model SC 7, Series 3 airplanes certificated in any category. COMPLIANCE: Required as indicated, unless already accomplished. To prevent failure of a main landing gear brake flange and subsequent loss of both main and emergency brake systems, accomplish the following: a) Within the next 200 landings after the effective date of this AD for airplanes equipped with brake flange Part No. EH183134 and having 5,000 or more landings in service or for airplanes equipped with brake flange Part No. EH183712 and having 1,000 or more landings in service, perform a dye penetrant inspection of the main landing gear brake flanges as instructed in paragraph 2 of Dowty Rotol Service Bulletin No. 32-10M, Rev. 1, dated February 17, 1982. 1) If no cracks are found, return the airplane to service. 2) If cracks are found, prior to further flight, replace the brake flange with a serviceable unit of the same part number.b) Repeat the inspection of paragraph a) above at 1,000 landing intervals. When the brake flange is replaced, these repeated inspections may be discontinued until the newly installed brake flange has accumulated the total number of landings as prescribed in paragraph a) above. c) Operators who have not kept records of total landings in service may convert airplane hours time-in-service to total landings in service at the rate of two landings per hour. d) The intervals between the repetitive inspections required by this AD may be adjusted up to 10 percent of the specified interval to allow accomplishing these inspections concurrent with other scheduled maintenance of the airplane. e) Aircraft may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished. f) An equivalent method of compliance with this AD if used must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on April 4, 1983.
79-25-08: 79-25-08 SIKORSKY AIRCRAFT DIVISION: Amendment 39-3635. Applies to all Model S-76A helicopters equipped with Sikorsky Models 76650-09802-101 and 76650-09802-102 hydraulic pumps. Compliance required as indicated, unless previously accomplished. To check the operation of existing installed pumps and to replace all defective pumps, accomplish the following: 1. If a hydraulic pump is known to have operated after fluid had been lost from the system, it is likely that the pump has run dry and it must be replaced prior to further flight. 2. If hydraulic fluid has been lost from a hydraulic reservoir, but the system has not run dry, perform servicing and bleeding procedures in accordance with the S-76A maintenance manual prior to further flight. 3. During each rotor startup, check for hydraulic pressure indication in the green arc range at 60% NR. The check required by the aforementioned sentence may be performed by the pilot. If hydraulic pressure is not in thisrange, conduct trouble shooting procedures and replace any malfunctioning component prior to further flight. 4 (A). Replace all hydraulic pumps, Sikorsky P/Ns 76650-09802-101 and 76650-09802-102, with Sikorsky P/N 76650-09802-103, in accordance with Paragraphs 4(B) and 4(C) below. These approved replacement pumps may also be identified by the suffix "C" after the serial number. (B). For pumps with 250 or more hours time in service on the effective date of this AD, compliance with paragraph 4(A) is required within the next 100 hours time in service. (C). For pumps with less than 250 hours time in service on the effective date of this AD, compliance with Paragraph 4(A) is required before the accumulation of 350 hours time in service. NOTE: Sikorsky Commercial Customer Service Notice 76-16, dated November 14, 1979, applies to this AD. The manufacturer's Customer Service Notice and Maintenance Manual identified and described in this directive are incorporatedherein 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 S-76A Product Manager, Commercial Customer Service Department, Sikorsky Aircraft Division, North Main Street, Stratford, Connecticut 06602. These documents may also be examined at Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803 and FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. 20591. This amendment becomes effective upon publication in the Federal Register, except for recipients of the Emergency AD, dated November 16, 1979, for whom it became effective upon receipt.
2009-15-11: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: During the landing roll a Corvette aircraft inclined to the Left Hand (LH) side as a result of the uncoupling of the left main landing gear shock absorber upper and lower cylinders, leading the left wheel tire to rub against the left wing under surface and to deflate, and the left wing tip fuel tank to strike the runway surface. The investigation showed that this uncoupling resulted from the loosening of the shock absorber locking system nut and its associated lock washer. * * * * * The unsafe condition is reduced structural integrity of the main landing gear, which could cause the wing tip fuel tank to strike the runway surface and potentially result in a fire. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
78-23-08: 78-23-08 AVCO LYCOMING: Amendment 39-3334. Applies to all IO-540-G1B5, -K1F5, -P1A5 and - S1A5 series engines with serial numbers up to and including L-17835-48 or -45A and all IO-540-G1B5, -K1F5, - P1A5, -S1A5 series engines overhauled (also known as remanufactured) by Lycoming prior to March 30, 1978. Compliance required within the next 50 hours' in service after the effective date of this AD, unless already accomplished. To prevent possible fuel leakage due to failure of the fuel pump to fuel injector tube elbow replace the fuel pump to fuel injector tube part number LW-10445 and the fuel pump to fuel injector tube elbow, part number LW-10446 with flexible hose assembly, part number LW-12877-6-051. Upon submission of substantiating data through an FAA Maintenance Inspector, the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration (FAA), Eastern Region, may adjust the compliance time specified in this AD. (NOTE: AVCO Lycoming Service Bulletin No. 421 covers this subject.) This amendment is effective November 7, 1978.
81-24-02: 81-24-02 GULFSTREAM AMERICAN CORPORATION (GRUMMAN AMERICAN AVIATION CORPORATION): Amendment 39-4259. Applies to Model G-1159 Serial Numbers 1 through 248, 250, 251, 253 through 258 and 775; Model G-1159A Serial Numbers 249, 252, 300 through 338 and 875. Compliance required within the next 75 hours time in service after the effective date of this AD, unless already accomplished. To prevent the loss of fluid flow in the auxiliary hydraulic system following a complete loss of fluid in the combined hydraulic system, install a check valve in the reservoir system in accordance with Gulfstream American Aircraft Service Change (ASC) No. 295 (for the Model G-1159) and ASC No. 21 (for the Model G-1159A), both dated September 1, 1981. An equivalent method of compliance may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southern Region. This amendment becomes effective November 23, 1981.
60-07-01: 60-07-01 ALLISON: Amdt. 124 Part 507 Federal Register March 29, 1960. Applies to Models 501-D13D and 501-D13E Engines. Compliance required within the next 25 hours of operation. A few cases of third stage turbine blade failures have occurred due to a resonance condition at low speed ground idle. All of these failures to date have resulted in visible damage to fourth stage blades as well as fourth stage vanes. In one case continued operation of an engine with a failed blade resulted in failure of the turbine inlet case-vane case split line bolts. (a) Aircraft not having operating engine vibration detection equipment must observe the following engine operating restriction and inspection. (1) Low speed ground idle operation from time all engines are started to stopping all engines at end of flight not to exceed 4 minutes total time. (2) Conduct inspection of fourth stage turbine blades at intervals not to exceed 25 hours of operation for indications of damage using adequate light and optical aid. (b) Aircraft having operating engine vibration detection equipment shall use this equipment to detect any indications above normal and if found, the above inspection of fourth stage turbine blades shall be conducted upon arrival at the next maintenance base. If any damage is discovered as a result of (a) or (b) it is cause for more detailed inspection and/or engine removal. (c) This restriction will not apply to engines modified in accordance with Allison Commercial Engine Bulletin No. 72-77 by installation of third stage turbine blades P/N 6794773 identified by a stripe of heat and corrosion resistant aluminum polytherm paint 1/2-inch wide and 4 inches long around contour of the inlet casing clockwise starting at the 1:00 position forward of the terminal block mounting flange. (Allison Commercial Engine Bulletin No. 72-77 covers the same subject.)
2009-14-05: The FAA is adopting a new airworthiness directive (AD) for Pratt & Whitney models PW2037, PW2037(M), and PW2040 turbofan engines. This AD requires 12th stage disks of certain high-pressure compressor (HPC) drum rotor disk assemblies, to be inspected for cracks by Pratt & Whitney using a special eddy current inspection procedure. This AD results from six HPC 12th stage disks found cracked during HPC module disassembly at overhaul. We are issuing this AD to prevent uncontained failure of the HPC 12th stage disk and airplane damage.
78-23-02: 78-23-02 ROLLADEN SCHNEIDER: Amendment 39-3336. Applies to Model LS1-f and LS3 gliders, all serial numbers, certificated in all categories. Compliance is required within the next 25 hours time in service after the effective date of this AD, unless already accomplished, and thereafter at intervals not to exceed one year from the last inspection until the tow release mechanism cable is replaced in accordance with the method specified in paragraph (b) of this AD. To prevent a tow release mechanism cable failure due to corrosion or cable strand damage, accomplish the following: (a) Remove the seat pan from the aircraft and using care not to damage the CG tow release mechanism cable, cut out approximately four (4) inches of the plastic fairlead (tube) covering the tow release mechanism cable, preferably at a low point in the cable run and just forward of the landing gear box. Move the tow release mechanism cable forwards and backwards through the cutout in the plastic tube and inspect the cable for corrosion or cable strand damage. (b) If during an inspection required by paragraph (a) of this AD, corrosion or cable strand damage is found, before further flight, replace the two release mechanism cable with a new stainless steel cable of 2.4 mm diameter, LN9389 or U.S. equivalent MIL-W-5424B, in accordance with Rolladen Schneider Technical Bulletins No. 35 dated May 12, 1978, or No. 3007 dated May 12, 1978, as applicable or in accordance with criteria provided in Chapter 4 of FAA Advisory Circular 43.13-1A "Acceptable Methods, Techniques, and Practices - Aircraft Inspection and Repair" or equivalent procedure approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region. NOTE: Rolladen Schneider Technical Bulletins No. 35 dated 5/12/78 applicable to the Model LS1-f glider and No. 3007 dated 5/12/78 applicable to the Model LS3 glider pertain to the same subject as this AD. This amendment becomes effective November 16, 1978.
2009-13-02: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Fokker Model F.28 Mark 0100 airplanes. That AD currently requires revisions to the airplane flight manual (AFM) to include procedures to prohibit use of reverse engine thrust power settings between idle and emergency maximum and to prohibit stabilized engine operation in a certain engine speed range on the ground. This new AD continues to require revising the AFM to include certain procedures. This AD also requires removing the normal maximum (second) detent for the reverse-thrust control. In addition, this AD requires revising the AFM to prohibit use of reverse thrust in flight and to limit operation of Max Reverse thrust. This AD results from issuance of mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. We are issuing this AD to prevent inadvertent operation inthe prohibited stabilized engine speed range on the ground, which could result in uncontained engine fan blade failure due to high cycle fatigue cracking.