Results
2009-22-02: We are adopting a new airworthiness directive (AD) for all American Champion Aircraft Corp. Models 7ECA, 7GCAA, 7GCBC, 7KCAB, 8KCAB, and 8GCBC airplanes, manufactured prior to 1989 and equipped with folding rear seat backs. This AD requires inspection of the rear seat back hinge areas for cracking and excessive elongation of the rear seat hinge bolt hole and, if cracking or excessive elongation is found, replacement of the rear seat frame. We are issuing this AD to detect and correct cracking of the rear seat back hinge area and excessive elongation of the rear seat hinge bolt hole, either of which could result in failure of the seat back. This failure could lead to a rear- seated pilot or passenger inadvertently interfering with the control stick while attempting to not roll to the rear of the airplane upon seat back failure. Consequently, this failure could result in loss of control.
2009-22-01: We are superseding an existing airworthiness directive (AD) for the products listed above. That AD currently requires initial and repetitive inspections of the low-pressure (LP) turbine discs stage 2 and stage 3 for corrosion, on certain serial number engines. This AD requires the same actions, but extends the applicability to additional engine serial numbers. 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: Strip results from some of the engines listed in the applicability section of this directive revealed excessively corroded low-pressure turbine disks stage 2 and stage 3. The corrosion is considered to be caused by the environment in which these engines are operated. Following a life assessment based on the strip findings it is concluded that inspections for corrosion attack are required. The action specified by this AD is intended to avoid a failure of a low-pressure turbine disk stage 2 or stage 3 due to potential corrosion problems which could result in uncontained engine failure and damage to the airplane. We are issuing this AD to detect corrosion that could cause the stage 2 or stage 3 disk of the LP turbine to fail and result in an uncontained failure of the engine.
78-24-04: 78-24-04 ISRAEL AIRCRAFT INDUSTRIES (IAI): Amendment 39-3351. Applies to Model 1121, 1121A, and 1121B airplanes, Serial Numbers through 150 except 107 and equipped with Ni-Cad battery overtemperature warning or combination overtemperature warning and indicator systems, which rely on a common single sensor per battery. Compliance required within the next 150 hours time in service after the effective date of this AD, unless already accomplished. To reduce the possibility of damage to the flexible fuel lines and potential fire in the event of an uncontrolled Ni-Cad battery thermal runaway, modify the flexible fuel line installation and after modification perform inspections and leak tests in accordance with Commodore Jet Service Bulletin No. CJ-16, dated May 31, 1977, or equivalent, approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region, c/o American Embassy, APO NY 09667. This amendment becomes effective December 20, 1978.
2009-21-09: The FAA is superseding an existing airworthiness directive (AD) for Rolls-Royce plc RB211 Trent 875-17, Trent 877-17, Trent 884- 17, Trent 892-17, Trent 892B-17, and Trent 895-17 turbofan engines with high-pressure (HP) compressor rotor rear stage 5 and 6 discs and cone shafts, part numbers (P/Ns) FK25230 and FK27899 installed. That AD currently requires removal from service of these HP compressor rotor rear stage 5 and 6 discs and cone shafts before reaching newly reduced life limits. This AD requires removing these parts at new reduced cycle limits. This AD results from Rolls-Royce plc reducing the lives of these parts and changing the life calculating method to use "Standard Duty Cycles'' with "Multiple Flight Profile Monitoring'' and "Flight Cycles'' with "Heavy Flight Profile Monitoring''. We are issuing this AD to prevent stage 5 and 6 disc crack initiation and propagation that might lead to uncontained disc failure and damage to the airplane.
2009-21-01: We are adopting a new airworthiness directive (AD) for certain Boeing Model 737-300 and 737-400 series airplanes. This AD requires repetitive inspections to detect cracking of the aft fuselage skin, and related investigative/corrective actions if necessary. This AD results from reports of cracks in the aft fuselage skin on both sides of the airplane. We are issuing this AD to detect and correct cracking in the aft fuselage skin along the longitudinal edges of the bonded skin doubler, which could result in reduced structural integrity of the airplane.
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.