Results
91-14-18: 91-14-18 BRITISH AEROSPACE: Amendment 39-7059. Docket No. 91-NM-42-AD. Applicability: All Viscount Model 744, 745D, and 810 series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent reduced structural integrity of the wings, accomplish the following: A. Within 180 days after the effective date of this AD, and thereafter at intervals not to exceed 180 days, perform an eddy current inspection to detect corrosion along the total length of the top surface of the left and right wing spar upper boom in accordance with British Aerospace Preliminary Technical Leaflet (PTL) No. 321, Issue 1, dated January 13, 1989, or PTL No. 190, Issue 1, dated January 13, 1989, as applicable. B. If corrosion is found, prior to further flight, repair in accordance with PTL No. 321, Issue 1, dated January 13, 1989, or PTL No. 190, Issue 1, dated January 13, 1989, as appropriate; or in a manner approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. C. An alternative method 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, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment 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 in order to comply with the requirements of this AD. E. The inspections and repair requirements shall be done in accordance with British Aerospace Preliminary Technical Leaflet (PTL) No. 321, Issue 1, dated January 13, 1989, or PTL No. 190, Issue 1, dated January 13, 1989, as applicable. 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-7059, AD 91-14-18) becomes effective on August 6, 1991.
96-04-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A340 series airplanes. This action requires repetitive inspections to detect corrosion and cracking of the shortening mechanism pins in the main landing gear (MLG), and repair or replacement, if necessary. It also requires replacement of the shortening mechanism pin and retraction mechanism pins; the forward pintle pin; the shortening mechanism bellcrank pin; and the bellcrank subassembly of the shortening mechanism. This amendment is prompted by reports of failure of the shortening mechanism pins due to improper grinding of the chrome plating during manufacture, which led to the initiation of stress corrosion and cracks. The actions specified in this AD are intended to prevent failure of these pins, which could lead to a significant reduction of the shock absorber capability or damage to various components of the MLG, and eventually could lead to the failure of the MLG.
77-18-04: 77-18-04 BEECH: Amendment 39-3028. Applies to Models C90 (serial numbers LJ-689 through LJ-713), E90 (serial numbers LW-167 through LW-216), A100 (serial numbers B-213 through B-232), B100 (serial numbers BE-10 through BE-22), Model 200 (serial numbers BB- 175 through BB-185, BB-187 through BB-193, BB-195 through BB-198, BB-200 through BB- 202, and BB-204 through BB-223), and Model 200T (serial numbers BT-1 and BT-2) airplane certificated in all categories. Compliance: Required as indicated unless already accomplished. To assure proper locking of the cabin door and to prevent possible incidents involving loss of cabin pressurization, within the next 25 hours time-in-service after the effective date of this AD, accomplish the following in accordance with Beechcraft Service Instructions No. 0914 or later approved revisions and instructions contained in the applicable airplane maintenance manual: A. Visually inspect the cabin door latch hooks and pins for proper engagement and adjustment, check the door latch mechanism for over center position when locked, and adjust latch hooks, pins and/or door latch mechanism if necessary. B. Any equivalent means of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective on September 12, 1977.
51-24-01: 51-24-01 LYCOMING: Applies to All Lycoming GO-435-C2 Engines Serially Numbered 1815-11 and Below Not Having the Letter "P" Stamped on the Upper Right-hand Corner of Each Cylinder Rocker Box and Installed in Navion Model B Aircraft. To be accomplished by June 1, 1952 or next overhaul, whichever occurs first. To prevent loosening of the exhaust valve seats in Lycoming GO-435-C2 engines installed in Navion Model B airplanes, the exhaust valve seats are to be peened in the cylinder head. The cylinders must be removed from the engine for this operation. When peening is accomplished, stamp "P" 1/8 inch high on upper right-hand corner of cylinder rocker box flange face near exhaust push rod. (Lycoming Service Bulletin No. 145 also covers this subject.
56-17-02: 56-17-02 de HAVILLAND: Applies to All Model DHC-3 Otter Aircraft. Compliance required as indicated. The Canadian Department of Transport has issued the following directive with which the FAA concurs and considers mandatory. "Any malfunctioning of the flap hydraulic circuit check valve would fail to lock the flaps in any flaps-extended position. When the valve operates properly, the flaps remain stationary when the flap selector lever is moved to the 'up' position and until the flap pump is operated. However, if the valve sticks open for any reason, selecting 'flaps up' results in a rapid flap retraction without use of the pump. Such retraction at high speeds will produce large stick forces and out-of-trim condition which flight tests have shown to be very dangerous when the aeroplane is trimmed for a high flaps-extended speed (full aircraft nosedown trim). "Therefore, until modifications now under development are incorporated, the following restriction is mandatory: "(a) The flap selector must not be placed in the up position until it is desired to retract the flaps, nor at speeds in excess of the following: Flap Setting Maximum Trim Speed Landing (35 degrees) 65 mph IAS Takeoff (30 degrees) 75 mph IAS Climb (15 degrees) 85 mph IAS "The previous '15 degrees climb flap limitation' on the use of flaps is canceled and the special 50- hour inspection of the flap controls, etc., may be discontinued." Incorporation of de Havilland Modifications 3/731, 3/744 and 3/745 is required by November 1, 1957, as outlined in D.H. Engineering Bulletin Series "O" No. 34, dated January 21, 1957. When these Modifications are incorporated, the operating restrictions in section (a) above will no longer be required.
2010-20-11: 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: In completing a review of Engine Manual repair/acceptance limits for titanium compressor shafts, Rolls-Royce has found the specified limits to be incorrect such that the shot peened surface layer at life critical features (the axial dovetail slots) may have been inadvertently removed in-service. Removal of the shot peened layer results in increased vulnerability of the part to tensile stresses, which could reduce the life of the shaft to below the published life limits. We are issuing this AD to prevent failure of the intermediate-pressure (IP) and high-pressure (HP) shaft, which could result in an overspeed condition, possible uncontained disc failure and damage to the airplane.
71-25-04: 71-25-04 NORTH AMERICAN ROCKWELL: Amendment 39-1352. Applies to Models NA-265, NA-265-20, NA-265-30, NA-265-40, NA-265-60, NA-265-70. Compliance required as indicated. To prevent failure of the aileron control cables, accomplish the following: Within the next 100 hours time in service after the effective date of this AD, unless already accomplished within the last 500 hours time in service, and thereafter at intervals not to exceed 600 hours time in service or 12 months, whichever occurs first, inspect the aileron control cables and replace as necessary; provided however that, if as a result of any inspection, more than three wires are found to be broken, the repetitive inspection interval will be decreased, or replacement required, as follows: (a) With four to six wires broken, repeat the inspection at intervals not to exceed 100 hours time in service. (b) With more than six wires broken, or if an equivalent reduction to the cable cross section area is present due to wear, replace the cable with a new or serviceable cable before further flight. Inspect the aileron control cables (P/N's 246-52324, 246-52325, 246-52339, 276-523005- 11, 276-523006-11, 276-523008-11, as applicable) in accordance with the following instructions: 1. Remove aileron control cables from the aircraft and inspect per step 9 or follow steps 2 through 16. 2. Lower wing flaps. 3. Open main wheel well doors or remove both wheel well cover assemblies as applicable. NOTE: Use normal safety precautions such as disconnecting the batteries to prevent inadvertent wing flap or landing gear wheel well door actuation. 4. In the left hand wheel well, disconnect the lower left hand aileron cable turnbuckle. 5. In the right hand wheel well, disconnect the upper left hand cable from the left hand aileron sector (P/N 246-52314). 6. Disconnect the left hand outboard aileron sector (P/N 246-52305-1), accessible through the left hand flap well, by removing the sector pivot bolt. 7. With the aileron sector pivot bolt removed disconnect the upper & lower left hand aileron cables from the sector. 8. Cable slack will now be available to allow pulling the upper left hand cable down into the landing gear strut well for inspection per step 9. 9. Clean the cable for a visual inspection. The cables must be bent in a "U" and inspected with a four power, or greater, magnifying glass in the area of pulley contact. 10. The lower left hand aileron control cable must be pulled inboard into the wheel well for inspection of the cable that passes over the pulley. Inspect per Step 9. 11. If the inspection of the left hand aileron control cables shows that they do not require replacing, reconnect and rig the left hand aileron control cables (See Note, below). 12. Disconnect the lower right had aileron cable turnbuckle located in the right hand wheel well. 13. Disconnect the upper cable at the aileronsector (P/N 246-52364) located in the right hand wheel well. 14. Pull the upper aileron cable down into the right hand main landing gear strut well and inspect per step 9. 15. Pull the lower aileron control cable into the right hand wheel well and inspect per step 9. 16. If the inspection of the two right hand cables reveals that they do not require replacing, reconnect and rig the aileron control system (See Note, below). NOTE: Instructions pertaining to the installation of new or serviceable cables and the rigging of the aileron control system are contained in the applicable maintenance documents. This amendment becomes effective December 4, 1971.
92-13-09: 92-13-09 CANADAIR, LTD.: Amendment 39-8279. Docket No. 92-NM-33-AD. Supersedes AD 92-03-06, Amendment 39-8161. Applicability: Model CL-600-1A11 series airplanes, serial numbers 1004 to 1085, except serial number 1037; Model CL-600-2A12 series airplanes, serial numbers 3001 to 3066; and Model CL-600-2B16 series airplanes, serial numbers 5001 to 5049; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent severe damage to an airplane in the event of an engine fire, accomplish the following: (a) Within 30 days after February 11, 1992 (the effective date of AD 92-03-06, Amendment 39-8161), accomplish the following: (1) For Model CL-600-1A11 series airplanes: Perform an inspection for potential crossed wiring in the engine fire extinguishing system, and inspect the electrical connectors for unlocked or inoperative pins, in accordance with Canadair Alert Service Bulletin A600-0581, dated September 8, 1989.(2) For Model CL-600-2A12 and CL-600-2B16 series airplanes: Perform an inspection for potential crossed wiring in both the engine fire detection and warning system and the engine fire extinguishing system, and inspect the electrical connectors for unlocked or inoperative pins, in accordance with Canadair Alert Service Bulletin A601-0309, dated September 8, 1989. (b) If any wiring discrepancies are detected as a result of the inspections required by paragraph (a) of this AD, prior to further flight, correct the discrepancies and replace any discrepant electrical connectors found, in accordance with Canadair Alert Service Bulletin A600-0581, dated September 8, 1989 (for Model CL-600-1A11 series airplanes); or Canadair Alert Service Bulletin A601-0309, dated September 8, 1989 (for Model CL-600-2A12 and CL-600-2B16 series airplanes); as applicable. (c) Within 120 days after the effective date of this AD, or the next time the fire bottles are removed from the airplane, whichever occurs first, modify the engine fire extinguishing warning harnesses and perform a functional test, in accordance with Canadair Alert Service Bulletin A600-0581, dated September 8, 1989 (for Model CL-600-1A11 series airplanes); or Canadair Alert Service Bulletin A601-0309, dated September 8, 1989 (for Model CL-600-2A12 and CL-600-2B16 series airplanes); as applicable. (d) 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, New York Aircraft Certification Office (ACO), ANE-170, FAA, Engine and Propeller Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, New York ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York ACO. (e) 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. (f) The inspection and modification shall be done in accordance with Canadair Alert Service Bulletin A600-0581, dated September 8, 1991 (for Model CL-600-1A11 series airplanes); or Canadair Alert Service Bulletin A601-0309, dated September 8, 1989 (for Model CL-600-2A12 and CL-600-2B16 series airplanes); as applicable. This incorporation by reference was approved previously by the Director of the Federal Register as of February 11, 1992 (57 FR 3006, January 27, 1992). Copies may be obtained from Bombardier, Inc., Canadair, Aerospace Group, P.O. Box 6087, Station A, Montreal, Quebec H3C 3G9, Canada. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; at the FAA, Engine and Propeller Directorate, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream,New York; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. (g) This amendment becomes effective on July 23, 1992.
76-21-04: 76-21-04 GENERAL ELECTRIC: Amendment 39-2747 as amended by Amendment 39-3085. Applies to Models CF6-6D and CF6-6D1 Turbofan Engines. Compliance required by June 30, 1978, unless previously accomplished. To prevent excessive overpressure in the high pressure compressor, remove the abradable material from the inside diameter of the Fan Stator Shroud Mid Ring (Booster Stage) in accordance with General Electric Service Bulletin (CF6-6) 72-647 or subsequent FAA Approved Revision thereto. The manufacturer's service bulletins identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to General Electric Company, Cincinnati, Ohio 45215. These documents may also be examined at the FAA Great Lakes Region, 2300 E. Devon Avenue, Des Plaines, Illinois 60018 and at FAA headquarters,800 Independence Avenue, S.W., Washington, D.C A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Great Lakes Region. Amendment 39-2747 became effective October 20, 1976. This amendment 39-3085 becomes effective November 30, 1977.
59-01-02: 59-01-02 de HAVILLAND: Applies to All Model 104 "Dove" Aircraft. Compliance required as follows: Aircraft prior to Serial Number 04463-April 1, 1959; Aircraft Serial Numbers 04463 through 04477 September 30, 1959. Cases have occurred where the main undercarriage locking levers P/N 4U 139A (Pre Modification 231) and P/N 4U 461A (Post Modification 231) have cracked in service. Dove Modification 868 has therefore been introduced which provides for locking lever assemblies in a material to a revised specification. On aircraft Pre Modification 231 standard, Dove Modification 187-repositioning the undercarriage warning lamp and microswitches-and Dove Modification 308-improving the operation of the main undercarriage mechanical position indicator-must be embodied at the same time as Modification 868. The British Air Registration Board considers this mandatory. The FAA concurs with this action and considers compliance therewith mandatory. (de Havilland TNS CT(104) No. 155 covers the same subject.)