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95-10-07: 95-10-07 BELL HELICOPTER TEXTRON, INC. : Amendment 39-9224. Docket No. 95-SW-02-AD. Applicability: Model 205A, 205A-1, and 204B helicopters, with a 42-degree tail rotor drive gearbox assembly (42-degree gearbox), part number (P/N) 204-040-003-023, or -037, installed, certificated in any category. NOTE 1: This AD applies to each helicopter identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For helicopters that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must use the authority provided in paragraph (e) to request approval from the FAA. This approval may address either no action, if the current configuration eliminates the unsafe condition, or different actions necessary to address the unsafe condition described in this AD. Such a request should include an assessment of the effect of the changed configuration on the unsafe condition addressed by this AD. In no case does the presence of any modification, alteration, or repair remove any helicopter from the applicability of this AD. Compliance: Required as indicated, unless accomplished previously. To prevent failure of the 42-degree gearbox, loss of tail rotor control, and subsequent loss of control of the helicopter, accomplish the following: (a) Before further flight, after the effective date of this AD, verify that the tail rotor control system is rigged in accordance with the applicable maintenance manual. (b) Before further flight, and thereafter at intervals not to exceed 400 torque events, disassemble the affected 42-degree gearbox and inspect for cracks at the roots of the gear teeth on the pinion, P/N 204-040-500-007 or -009, and gear, P/N 204-040-500-008 or -010, using a fluorescent penetrant inspection method in accordance with the applicable maintenance manual. Only post emulsified fluorescent penetrant inspection materials (Type I, Method B or D, Sensitivity Level 3 or greater) are approved for use. A torque event is defined as a takeoff or a lift (internal or external). (c) If any crack is found at the roots of the gear teeth on the pinion or gear, replace the pinion or gear with an airworthy pinion or gear in accordance with the applicable maintenance manual. (d) Create a component history card for the 42-degree gearbox. Record the number of torque events on a daily basis. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used when approved by the Manager, Rotorcraft Certification Office, FAA, Rotorcraft Directorate. Operators shall submit their requests through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager,Rotorcraft Certification Office. NOTE 2: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Rotorcraft Certification Office. (f) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the helicopter to a location where the requirements of this AD can be accomplished. (g) This amendment becomes effective on May 26, 1995.
96-04-09: This amendment adopts a new airworthiness directive (AD), applicable to all Fokker Model F28 Mark 0100 series airplanes, that requires an inspection to detect the presence of a drain hole in certain mounting frames of the auxiliary power unit (APU). If a drain hole is present, the AD requires an inspection to detect corrosion of the mounting frame, and eventual replacement of the mounting frame. This amendment is prompted by a report indicating that corrosion was found on a number of mounting frames of the APU. The actions specified by this AD are intended to prevent such corrosion, which could lead to failure of the frame and consequently render the APU inoperative and/or create a potential fire hazard.
2013-19-15: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 747-100, -100B, -100B SUD, -200B, -200C, - 200F, -300, -400, -400D, -400F, and 747SR series airplanes. This AD was prompted by reports of cracking at the aft upper corner of the main entry door (MED) 5 cutout. This AD requires inspecting for the presence of repairs and measuring the edge margin at certain fastener locations around the upper aft corner of the door cutout, inspecting for any cracking of the fuselage skin assembly and bear strap in the aft upper corner area of the door cutout, and repairing or modifying the fuselage skin assembly and bear strap if necessary. We are issuing this AD to detect and correct cracking of the skin and bear straps at the aft upper corner of the MED 5 cutout, which could result in in-flight depressurization.
99-08-18: This amendment adopts a new airworthiness directive (AD), that requires revisions to the Time Limits Section of the manufacturer's Instructions for Continued Airworthiness (ICA) for General Electric Company (GE) CF6-6, CF6-45, and CF6-50 series turbofan engines to include required enhanced inspection of selected critical life-limited parts at each piece-part exposure. This amendment also requires an air carrier's approved continuous airworthiness maintenance program to incorporate these inspection procedures. This amendment is prompted by a Federal Aviation Administration (FAA) study of in-service events involving uncontained failures of critical rotating engine parts which indicated the need for improved inspections. The improved inspections are needed to identify those critical rotating parts with conditions, that if allowed to continue in service, could result in uncontained failures. The actions specified by this proposed AD are intended to prevent critical life-limited rotating engine part failure, which could result in an uncontained engine failure and damage to the airplane.
91-05-18: 91-05-18 SCHWEIZER AIRCRAFT CORPORATION (HUGHES HELICOPTERS, INC.): Amendment 39-6857. Docket No. 90-ASW-46. Applicability: All Model 269C series helicopters, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent fatigue failure of the magnesium tailboom center attachment (saddle) fittings which could result in loss of the tailboom of the helicopter, accomplish the following: (a) For helicopters with a magnesium tailboom center attachment fitting, P/N 269A2324-7, installed with 400 hours' or less total time in service, perform the following in accordance with the 269 Series Basic Handbook of Maintenance Information (HMI), as revised by Temporary Revision No. R-42, dated October 8, 1990 (HMI): (1) Prior to further flight after the effective date of this AD, and thereafter at intervals not to exceed 50 hours' time in service from the last inspection, inspect the magnesium tailboom center attachment fitting and tailboom visually using a 10-power or higher magnifying glass, in accordance with the HMI. (2) Within 25 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 100 hours' time in service from the last inspection, inspect the magnesium tailboom center attachment fitting using a dye penetrant inspection in accordance with the HMI. (3) Remove and replace the magnesium tailboom center attachment fitting, P/N 269A2324-7, with an aluminum tailboom center attachment fitting, P/N 269A2324-11 or P/N 269A2324-11T, as prescribed by the HMI prior to attaining 500 hours' total time in service. (b) For helicopters with magnesium tailboom center attachment fitting, P/N 269A2324-7, installed with more than 400 hours' total time in service, perform the following in accordance with the HMI: (1) Prior to further flight and thereafter at an interval not to exceed 50 hours' time in service from the last inspection, inspect thetailboom center attachment fitting using a dye penetrant inspection as prescribed in paragraph (a)(2) of this AD. (2) Before the first flight of each day inspect the tailboom center attachment fitting and tailboom visually using a 10-power or higher magnifying glass as prescribed in paragraph (a)(1) of this AD. (3) Remove and replace the magnesium tailboom center attachment fitting, P/N 269A2324-7, with an aluminum tailboom center attachment fitting, P/N 269A2324-11 or P/N 269A2324-11T, as prescribed in the HMI, within the next 100 hours time in service after the effective date of this AD. (c) Remove and replace the magnesium tailboom center attachment fitting, P/N 269A2324-7, with an aluminum tailboom center attachment fitting, P/N 269A2324-11 or P/N 269A2324-11T, before further flight if it is cracked or otherwise unserviceable. NOTE: Schweizer Service Bulletin B-239, dated October 8, 1990, pertains to this AD. The tailboom assembly service life is 2,100 hours.(d) Aircraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the requirements of this AD can be accomplished. (e) Alternate inspections, modification, or adjustments of the compliance times, which provide an equivalent level of safety may be used when approved by the Manager, New York Aircraft Certification Office, FAA, 181 South Franklin Avenue, Room 202, Valley Stream, NY 11581. This amendment (39-6857, AD 91-05-18) becomes effective on March 29, 1991.
2013-19-16: We are adopting a new airworthiness directive (AD) for Sikorsky Aircraft Corporation (Sikorsky) Model S-92A helicopters to require modifying the No. 1 engine forward firewall center fire extinguisher discharge tube (No. 1 engine tube) and inspecting the outboard discharge tube to determine if it is correctly positioned. This AD was prompted by the discovery that the No. 1 engine tube installed on the helicopters is too long to ensure that a fire could be effectively extinguished in the helicopter. The actions are intended to ensure the No. 1 engine tube allows for complete coverage of an extinguishing agent in the No. 1 engine compartment area, ensure that a fire would be extinguished and prevent the loss of helicopter control.
99-08-17: This amendment adopts a new airworthiness directive (AD), that requires revisions to the manufacturer's Life Limits Section of the Instructions for Continued Airworthiness (ICA) for General Electric Company (GE) GE90 series turbofan engines to include required enhanced inspection of selected critical life-limited parts at each piece-part exposure. This amendment will also require an air carrier's approved continuous airworthiness maintenance program to incorporate these inspection procedures. This amendment is prompted by a Federal Aviation Administration (FAA) study of in-service events involving uncontained failures of critical rotating engine parts which indicated the need for improved inspections. The improved inspections are needed to identify those critical rotating parts with conditions, that if allowed to continue in service, could result in uncontained failures. The actions specified by this AD are intended to prevent critical life-limited rotating engine part failure, which could result in an uncontained engine failure and damage to the airplane.
99-07-18: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 99-07-18 which was sent previously to all known U.S. owners and operators of Robinson Helicopter Company (RHC) Model R44 helicopters by individual letters. This AD requires, before further flight, inserting a Special Pilot Caution into the Normal Procedures section of the Rotorcraft Flight Manual (RFM). This amendment is prompted by several reports of sprag clutch assemblies with cracked or fractured sprag ends. The sprag clutch failures, determined to be due to a change in the manufacturing process, could result in loss of main rotor revolutions-per-minute (RPM) during autorotations. The intent of this AD is to alert pilots of the potential for the sprag clutch failing to overrun during autorotation, loss of main rotor RPM, and subsequent loss of control of the helicopter.
2013-19-22: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 717-200 airplanes. This AD was prompted by multiple reports of cracks of overwing frames. This AD requires repetitive inspections for cracking of the overwing frames, and corrective actions if necessary. We are issuing this AD to detect and correct such cracking that could sever a frame, which may increase the loading of adjacent frames, and result in damage to the adjacent structure and consequent loss of structural integrity of the airplane.
56-21-02: 56-21-02 CESSNA: Applies to All Model T-50 (AT-17 Series and UC-78 Series) Aircraft. Compliance required on or before the next periodic inspection but not later than January 1, 1957, and at each periodic inspection thereafter. Receipt of reports of spar deterioration subsequent to the corrective measures of AD 49- 20-01 necessitates repetitive and detailed inspections. It is therefore requested that the rear spar just outboard of the fuselage fittings in the area of the inside corner of the fuel tank compartment be inspected annually for indications of wood deterioration of the spar. To accomplish the inspection, remove a section of plywood wing covering 6 inches x 4 inches adjacent to and outboard of the rear spar fuselage fittings and/or remove the fuel tank. If the inspection reveals any evidence of deterioration or separation of glue lines, removal of the lower cap strip between ribs in the affected area for detail inspection is required. To facilitate periodic inspection a permanent inspection opening may be incorporated provided it is reinforced to maintain strength equivalent to the original panel. If spar deterioration does not exceed 1 inch of spar depth the spar may be repaired by the following method: Remove the damaged wood of the spar and fit with a matched inlay by gluing to the spar with the ends of the inlay tapered in a ratio of no less than 15 to 1. A bottom reinforcing strip or plate of an approximate thickness of 0.4 of the depth of thickness of the inlay with a maximum of 0.4-inch thickness is to be matched and glued to the bottom surface of the spar. This bottom reinforcing strip should extend approximately 3 inches beyond the end of the glue line of the inlay strip. The installation of this reinforcement plate may require a rework of the spacer block at the attachment fitting and a rework of the fore and aft stringers where they attach to the bottom surface of the spar. The following contingencies apply in the repair of the spar: (1) In the "5,700 pound" wing spar, the material for the inlay lamination and bottom reinforcement strip or plate should be fabricated from birch, maple or its equivalent. (2) In the "5,100 pound" wing spar, the material for the inlay lamination and bottom reinforcement strip or plate may be fabricated from spruce. Workmanship, gluing process, quality of materials and other limitations and repair requirements of CAM 18 apply except where otherwise noted. After completion of the spar repair and repair or replacement of deteriorated gussets, stringers, etc., as necessary, drain holes should be incorporated in wings if not already installed. If the drain holes are installed it should be ascertained that they are open. The felt padding under fuel tanks should be inspected for evidence of absorption of moisture, and the padding replaced, if necessary, with neoprene coated felt or equivalent nonabsorbing material. This supersedes AD 49-20-01.