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87-03-11: 87-03-11 SOCATA Groupe AEROSPATIALE: Amendment 39-5575. Applies to Models TB 10, TB 20, and TB 21 (Serial Numbers (S/N) 1 through 709) airplanes certificated in any category. Compliance: Required as indicated in the body of the AD, unless already accomplished. To prevent loose stabilator rod ends that may result in loss of pitch control, accomplish the following: (a) Prior to further flight, and each 100 hours time-in-service (TIS) thereafter until repaired or replaced with a modified stabilator actuator rod assembly, inspect the stabilator actuator rod ends and rivets for relative motion as described in paragraph A of SOCATA TB Aircraft S/B No. 29 dated October 1986. If there is any detectable play in any axis, loose or working rivets, prior to further flight, accomplish the actions in paragraphs (i) or (ii) below: (i) Repair the stabilator actuator rod assembly Part Number (P/N) TB 10.27.000.024 and replace the rivets with threaded fasteners as described in paragraphB of SOCATA TB Aircraft S/B No. 29 dated October 1986; or (ii) Replace the existing riveted stabilator actuator rod assembly P/N TB 10.27.000.024 with actuator rod assembly P/N TB 10.27.000.026, modified with the threaded fasteners as described in paragraph B of SOCATA TB Aircraft S/B No. 29 dated October 1986. (b) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (c) An equivalent means of compliance with this AD may be used if approved by the Manager, Brussels Aircraft Certification Office, FAA, AEU-100, Europe, Africa, and Middle East Office, c/o American Embassy, B-1000 Brussels, Belgium; Telephone 513.38.30. All persons affected by this directive may obtain copies of the document(s) referred to herein upon request to SOCATA Groupe AEROSPATIALE, B.P. 38, 65001-Tarbes, France; Telephone 62.93.97.30; or Product Support Deputy Manager, U.S., Mr. Bernard M. Veyssiere, 12605 Woodygrove Drive, St. Louis, Missouri 63146; Telephone (314) 469-7490; or may examine the document(s) referred to herein at the FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on March 11, 1987, to all persons except those to whom it has already been made effective by priority letter from the FAA dated February 2, 1987, and is identified as AD 87-03-11.
50-34-01: 50-34-01 KOPPERS: Applies to All Aircraft Equipped With Model F200 "Aeromatic" Propellers (Does Not Apply to "Aeromatic Model 220 Propellers"). Compliance required in all cases no later than April 1, 1952. 1. Universal (Stinson) Models 108-2 and 108-3 aircraft: Compliance required no later than first 200 hours of propeller operation. 2. Universal (Stinson) Models 108 and 108-1 aircraft: Compliance required no later than first 400 hours of propeller operation. 3. If the total propeller operation time is unknown, or if a reasonably accurate estimate of total time cannot be made, compliance is required not later than the next 50 hours of operation. (Except for Universal (Stinson) Series aircraft, compliance is required by not later than the next 50 hours of operation if the total operation time as of August 29, 1949, exceeds 500 hours.) Blade retaining flanges, P/N 3277 must be replaced with P/N 3277-1. When this change is accomplished a "-1" (dash one) is tobe suffixed to the propeller assembly number on the nameplate to indicate compliance. Koppers Service Bulletin No. 24 covers this same subject. Universal (Stinson) Models 108-2 and 108-3 only: (Compliance required by May 16, 1949). To avoid the possibility of crankshaft or propeller failures resulting from excessive torsional vibration in the 2,700 to 2,800 r.p.m. range, all engine operation must be restricted to 2,650 r.p.m. maximum and propeller readjusted in accordance with Koppers Service Bulletin No. 22. As a further safety measure it is required that propellers which have accumulated any operating time in the 2,650 to 2,800 r.p.m. range be equipped with new blade retainer flanges P/N 3277-1. (Koppers Service Bulletin No. 23-E covers this same subject.) This supersedes AD 49-42-01, for the purpose of clarifying the date of compliance.
2018-06-04: We are superseding Airworthiness Directive (AD) 2004-03-07, which applied to certain Airbus Model A320-111, -211, -212, and -231 series airplanes. AD 2004-03-07 required repetitive inspections for fatigue cracking around the fasteners attaching the pressure panel to the flexible bracket at a certain frame (FR), adjacent to the longitudinal beams on the left and right sides of the airplane; and repair as necessary. This new AD retains certain requirements of AD 2004-03-07, expands the applicability, and requires an inspection of the fastener holes on the pressure panel and modification or repair as applicable. This AD was prompted by fatigue tests which revealed cracking around the fasteners attaching the pressure panel to the flexible bracket, and by the discovery of additional cracks under the longitudinal beams at locations that are not included in the inspection area required by AD 2004-03-07. We are issuing this AD to address the unsafe condition on these products.
84-19-03 R1: 84-19-03 R1 AVCO LYCOMING: Amendment 39-4907 as amended by Amendment 39-4921. Applies to Avco Lycoming model TIO-540-V2AD engines Serial Numbers L-8620-61A and below and model LTIO-540- V2AD Serial Numbers L-2820-68A and below, having P/N LW-18852 cylinder assemblies. These engines are installed in Piper PA-31P-350 "Mojave" aircraft. Compliance is required as indicated, unless already accomplished. Cracks have occurred in the area of the intake valve seat of the cylinder head in the above engine models. To prevent possible failure of the engine cylinder assembly, P/N LW-18852, accomplish the following: (a) Within the next 50 hours of time-in-service after the effective date of this AD, or prior to accumulating 550 hours time-in-service, whichever occurs later, replace all P/N LW-18852 cylinder assemblies with P/N LW-19281 cylinder assemblies which are included in kit P/N LW-19297-S. Cylinder part number identification is located on the cylinder head casting in the rockerbox area. (b) In accordance with FAR 21.197 and 21.199, the aircraft may be flown to a location where the inspections or alterations required by this AD can be performed. (c) An equivalent means of compliance with the requirements of this AD may be approved by the Manager, New York Aircraft Certification Office. Amendment 39-4907 becomes effective on September 20, 1984. This Amendment 39-4921 becomes effective on September 20, 1984.
48-51-01: 48-51-01 BEECH: Applies to AT-11, C18S, D17S Airplanes Equipped With Beech Half-Circle Type Control Wheels. To be accomplished as soon as possible, but not later than March 1, 1949. To preclude the possibility of failure of the control column wheel shaft due to too great a counterbore depth in the shaft, the following inspection should be accomplished: (Bow-tie type control wheels attached to the control shaft by three bolts through a flange on the shaft are satisfactory and need not be inspected). (1) Drill a 3/8-inch hole in the center of the hub of the control wheels. Do not allow the drill to go more than 1 inch into the hub or the pin securing the wheel will be damaged. (2) Place a narrow scale or straight wire through the 3/8-inch hole and obtain the distance from the bottom of the counterbore (not the peak of the counterbore cone) to the face of the hub. (3) Insert a small hook scale or bent wire in the hole and obtain the distance from the end of the shaft to the face of the hub. (4) Subtract the distance obtained in step No. 3 from the distance obtained in step No. 2 to obtain the depth of the counterbore in the end of the shaft. If this distance is over 1 7/16 inches, the shaft must be replaced. All shafts having a counterbore less than 1 7/16 inches deep are satisfactory. (Beech Service Bulletin No. C-18-9 dated November 22, 1948, covers this same subject.)
62-08-01: 62-08-01 AERO COMMANDER: Amdt. 415 Part 507 Federal Register April 10, 1962. Applies to All Models With Serial Numbers 1 Through 1170, Except Serial Numbers 1133, 1153, 1156, 1162, Models 720, and Pressurized Versions of Model 680F. Compliance required within the next 25 hours' time in service after the effective date of this directive. It has been found that fuel and fuel fumes are released within the cockpit area behind the instrument panel, through the fuel pressure gage case vent as a result of gage diaphragm rupture. Accordingly, a fuel drain line shall be installed leading from the fuel pressure gage vent connection and routed overboard through the belly of the aircraft in accordance with Aero Commander Service Bulletin No. 78 dated March 16, 1962, or an FAA approved equivalent. This directive effective April 10, 1962.
2018-06-03: We are superseding Airworthiness Directive (AD) 2009-18-16, which applied to certain Airbus Model A310-203, -204, -221, -222, -304, -322, -324, and -325 airplanes. AD 2009-18-16 required an inspection for cracking of certain fastener holes on certain frames, and related investigative and corrective actions if necessary; and modification of certain fastener holes. This new AD reduces the compliance times. This AD was prompted by the identification of a structural modification that falls within the scope of the work related to the extension of the service life of the affected airplanes and widespread fatigue damage evaluations. We are issuing this AD to address the unsafe condition on these products.
87-15-09: 87-15-09 AIRBUS INDUSTRIE: Amendment 39-5685. Applies to all Model A300 B2 series airplanes, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent failure of the wing front spar, accomplish the following: A. Within the next 100 landings after the effective date of this AD, or prior to the accumulation of 8,000 landings, whichever occurs later, visually inspect the wing front spar webs in accordance with paragraph 2B of Airbus Industrie All Operators Telex AOT/57/87/02, Issue 2, dated April 22, 1987. B. Within the next 300 landings after the effective date of this AD, or prior to the accumulation of 8,000 landings, whichever occurs later, perform an ultrasonic inspection in accordance with paragraph 2C of Airbus Industrie All Operators Telex AOT/57/87/02, Issue 2, dated April 22, 1987. C. If cracks are found as a result of the inspections required by paragraph A. or B., above, repair prior to further flight in accordance with Airbus Industrie All Operators Telex AOT/57/87/02, Issue 2, dated April 22, 1987. D. 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. E. 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 inspections 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, 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 amendment, 39-5685, becomeseffective on August 10, 1987.
98-23-17: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A300 series airplanes, that requires modification of the emergency evacuation slide/raft system. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent the container release cable of the emergency evacuation slide/raft system from jamming, which could result in the inability to open the emergency exit doors or to correctly deploy the emergency evacuation slide/rafts, and consequent delay or impedance of passengers exiting the airplane during an emergency.
62-02-06: 62-02-06 PIPER: Amdt. 387 Part 507 Federal Register January 18, 1962. Applies to Model PA-25 Aircraft, Serial Numbers 25-1 to 25-619 Inclusive, 25-622, 25-623, 25-625, and 25-626. Compliance required as indicated. As a result of inspections performed on the aluminum front spar wing-to-fuselage attachment fittings, Piper P/N's 61111-00 and 61111-01, it has been determined that the following action is necessary. Within the next 25 hours' time in service after the effective date of this AD, but not later than February 15, 1962, unless already accomplished, replace the aluminum front spar wing-to-fuselage attachment fittings, Piper P/N's 61111-00 and 61111-01, with steel forgings, Piper P/N's 60113-00 and 60113-01, or equivalent parts approved by the FAA Eastern Region, Engineering and Manufacturing Branch. The steel forgings have 1/8 inch raised digits "60112" or "60112-1" on the parts. The replacement shall be accomplished in accordance with Piper Service BulletinNo. 206, dated August 24, 1961, or FAA approved equivalent. (Piper Service Bulletin No. 206, dated August 24, 1961, pertains to this subject.) This supersedes AD 61-13-02. This directive effective January 18, 1962.