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60-24-03: 60-24-03 PIPER: Amdt. 225 Part 507 Federal Register November 17, 1960. Applies to PA-24 and PA-24 "250" Aircraft Serial Numbers 24-1 to 24-2161 Inclusive. Compliance required within the next 10 hours of operation or at the next periodic inspection, whichever occurs first, after the effective date of this directive. To prevent any interruption in fuel flow should the vent tubes become obstructed, the two fuel cell vent tubes which are located under the wings shall be modified in the following manner: Drill an 0.098-inch diameter hole (#40 drill) in the aft side of each tube three-fourths of an inch from the end. (Piper Service Bulletin No. 193 covers this subject.) This directive effective December 19, 1960.
75-22-11: 75-22-11 GENERAL ELECTRIC: Amendment 39-2398. Applies to Models CF6-6D and CF6-6D1 Turbofan Engines. Compliance required as indicated. To prevent possible disintegration of the high pressure turbine thermal shield accomplish the following: (a) Within the next 400 operating cycles after the effective date of this Airworthiness Directive, unless already accomplished, and every 400 cycles thereafter, borescope inspect the high pressure turbine thermal shield on all CF6-6D and CF6-6D1 engines except those noted in paragraph (b) below in accordance with the instructions of General Electric Service Bulletin (CF6-6) 72-603 dated October 1, 1975 or subsequent FAA Approved revision. (b) Inspection is not required: (1) On engines containing thermal shields replaced or modified in accordance with General Electric Service Bulletin (CF6-6) 72-442, dated October 8, 1973, Revision 1 or subsequent FAA Approved revision, or General Electric Service Bulletin (CF6-6) 72-502 dated October 23, 1974 or subsequent FAA Approved revision. (2) On engines incorporating thermal shields, General Electric part number 9687M67P08, assembly number 9687M67G12; part number 9687M67P09, assembly number 9687M67G13; part number 9687M67P12, assembly number 9687M67G16; or part number 9687M67P17, assembly number 9687M67G21. (c) For the purposes of this Airworthiness Directive, the definition of a "cycle" is the definition appearing in the General Electric CF6-6 Shop Manual GEK 9266, Section 72-00-00, Page 301 dated August 1, 1975 or subsequent FAA Approved revision. 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)(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. This amendment becomes effective October 28, 1975.
60-16-03: 60-16-03 DOUGLAS: Amdt. 188 Part 507 Federal Register August 6, 1960. Applies to All DC3 Series Aircraft With Geared Rudder Tab Installations Based On Data Approved Prior to the Effective Date of This Airworthiness Directive. \n\n\tCompliance is required as indicated. \n\n\t(a)\tIn order to correct rudder force reversal tendencies on existing installations, the following shall be accomplished: \n\n\t\t(1)\tWithin two weeks after the effective date of this directive and until the aircraft has been flight tested or modified in accordance with this directive, a placard shall be placed in the aircraft in full view of the pilot which reads as follows: \n\n\t\t\t"Possible rudder force reversal and/or rudder lock may be experienced in this aircraft if rudder application is not coordinated with lateral control. Avoid yawed flight." \n\n\t\t\tThis placard shall be retained in the airplane and complied with until either of the applicable procedures described in (2) have been accomplished. \n\n\t\t(2)\tToremove the placard, either of the following procedures must be accomplished: \n\n\t\t(i)\tINSPECTION AND TEST OF THE GEARED TAB INSTALLATION. \n\n\t\t\t(a)\tCheck the rigging of the geared rudder tab installation in accordance with the manufacturer's approved installation data to prove conformity of this installation prior to the required flight test below. The results of the rigging check must be recorded in the aircraft logbook and signed by the individual making the check. \n\n\t\t\t(b)\tContact the nearest FAA Regional Office and make arrangements through the Flight Test Branch for having the aircraft tested. The results of this flight test must be recorded in the aircraft logbook and signed by the individual conducting the flight test. \n\n\t\t\t(c)\tIf the rudder control characteristics in the flight test are found to meet the requirements of Civil Air Regulations, Part 4a, Sections 4a.758-T (or Civil Air Regulations, Part 4b, Section 4b.157), the placard in paragraph (1) may be removed.(d)\tIf the rudder control characteristics in the flight test are found not to meet the requirements of Civil Air Regulations, Part 4a, Section 4a.758-T (or Civil Air Regulations, Part 4B, Section 4b.157), the placard may not be removed until a corrective design modification has been made, officially inspected and flight tested, and found to comply with the above regulations. \n\n\t\t(ii)\tREPLACEMENT WITH AN APPROVED NEW OR MODIFIED GEARED TAB INSTALLATION. \n\tAt such time as a "fix" or a new design installation has been developed, officially inspected and flight tested, and found to comply with the regulations, such an FAA approved modification or design may be installed in accordance with the manufacturer's specifications, a rigging and installation check made and recorded in the aircraft logbook by the individual who made the check. No mandatory flight tests will be necessary for such installations and the above-mentioned placard may be removed at this time. \n\n\t\t(b)\tTo precludethe installation on other aircraft of geared tabs of the same design which may have rudder force reversal tendencies, the following shall be accomplished prior to each approval: \n\n\t\t(1)\tAn official flight test shall be arranged with the nearest FAA Regional Office to determine that the installation complies with the regulations. The results of this flight test, as well as the prior inspection for conformity with approved installation data, must be recorded in the aircraft logbook and signed by the individuals conducting the installation inspection and flight test. \n\n\tThis directive shall become effective 30 days after the date of its publication in the Federal Register.
2005-10-13: The FAA is adopting a new airworthiness directive (AD) for Rolls-Royce Corporation (RRC) (formerly Allison Engine Company) 250-B17B, - B17C, -B17D, -B17E, -C20, -C20B, -C20F, -C20J, -C20S, and -C20W turboprop and turboshaft engines that do not have turbine energy absorbing ring, RRC part number (P/N) 23035175, or an equivalent FAA- approved serviceable turbine energy absorbing ring, installed. This AD requires installation of a turbine energy absorbing ring in the plane of the 1st stage turbine wheel. This AD results from an unacceptable rate of uncontained 1st stage turbine wheel failures. We are issuing this AD to minimize the risk of uncontained 1st stage turbine wheel fragments from causing damage to the aircraft or damage to the second engine on twin-engine installations, which could lead to loss of control and loss of the aircraft.
2015-25-06: We are superseding Airworthiness Directive (AD) 2010-06-04, for certain Airbus Model A300 B2-1C, B2-203, B2K-3C, B4-103, B4-203, B4-2C airplanes; Model A310 series airplanes; Model A300 B4-600 series airplanes; and Model A300 B4-600R series airplanes. AD 2010-06-04 required repetitive inspections to detect cracks of the pylon side panels (upper section) at rib 8; and corrective actions if necessary. This new AD continues to require repetitive inspections for cracking of the pylons 1 and 2 side panels (upper section) at rib 8 with reduced compliance times, and corrective actions if necessary. This AD also requires repetitive post-repair and post-modification inspections and repair if necessary. This AD also removes certain airplanes having a certain modification from the applicability. This AD was prompted by reports of cracks found on pylon side panels at rib 8 and a fleet survey and updated fatigue and damage tolerance analyses. We are issuing this AD to detect and correct cracking of pylon side panels (upper section) at rib 8, which could lead to reduced structural integrity of the pylon primary structure, which could cause detachment of the engine from the fuselage.
73-24-01: 73-24-01 ROCKWELL INTERNATIONAL: Amdt. 39-1743. Applies to Rockwell Commander Model 112 airplanes, Serial Numbers 3 through 120, certificated in all categories. Compliance required before further flight unless already accomplished. To prevent failure of the aileron hinges and/or the elevator trim tab hinges, accomplish the following: (a) Inspect all hinge halves of both ailerons and both elevator trim tabs from underside of aircraft to determine if the hinges are of the formed type made by rolling the edge of a 0.040 inches thick flat sheet or of the extruded type 0.060 inches thick. (b) If extruded hinge halves are found in all locations, no further action is required. (c) If a formed hinge piece is found, that complete hinge must be replaced with Rockwell Commander Part No. 42251-1 for the aileron hinges or Rockwell Commander Part No. 44020-5 for the elevator trim tab hinges before further flight. If no cracks are visually evident in any formed hinges,the airplane may be flown in accordance with FAR 21.197 to a base where the replacement can be performed. Rockwell International Service Bulletin No. SB-112-6 pertains to this same subject. This amendment becomes effective November 19, 1973.
60-10-01: 60-10-01 BELL: Amdt. 146 Part 507 Federal Register May 10, 1960. Applies to All Helicopter Models: 47B, 47B3, 47D, 47D1, 47G, and 47H1, all Serial Numbers; 47G2 Serial Numbers 1327 Through 2467, 2469, 2470, 2472 Through 2477, 2556 Through 2558; 47J Serial Numbers 1420 Through 1776 (Except For Helicopters On Which Kit No. 47-3410-1 (333SI) Has Been Installed); 47E, and 47K. Compliance required as indicated except Model 47G2, Serial Numbers 2451, 2452, 2457, 2459 through 2467, 2469, 2470, 2472 through 2477, 2556 through 2558, for which compliance date is September 2, 1960. As the result of a number of recent failures of the scissor lever pivot bolts due to excessive wear, the following is required unless already accomplished. (a) Prior to June 30, 1960, except 47E and 47K as to which compliance is required prior to August 15, 1960, inspect the scissor lever pivot bolts, AN 174-31, and bolt holes in the brackets of the collective pitch sleeve weld assembly, P/N 47-150-117-5 for wear. Wear limits and reinspection intervals are specified in the following items (1), (2), (3), and (4). (1) If the diameter of the two AN 174-31 bolts is less than 0.2465 inch in any area, bolts must be replaced prior to next flight. (2) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is 0.2550 inch or more, install four bushings, P/N 47-150-260-3 or equivalent, and new AN 174-31 bolts within the next 25 hours' time in service. (3) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is between 0.2500 and 0.2550 inch, the bolts and bolt holes must be reinspected dimensionally every 25 hours' time in service until bushings P/N 47-150-260-3 are installed. (4) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is 0.2500 inch or less, the bolts and bolt holes must be reinspected dimensionally every 100 hours' time in service until bushings P/N 47-150-260-3 are installed. (b) Upon installation of the bushings P/N 47-150-260-3, the bolts and bushing holes must be inspected every 300 hours' time in service thereafter. (c) Upon installation of the four bushings, P/N 47-150-260-3, designate the reworked collective pitch sleeve weld assembly as P/N 47-150-117-21. (Bell Service Bulletin No. 129SB, dated March 18, 1960, covers this same subject.) Revised July 15, 1960. Revised August 19, 1960.
2005-10-10: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier Model CL-600-2B19 (Regional Jet Series 100 & 440) airplanes. This AD requires revising the Airworthiness Limitations section of the Instructions for Continued Airworthiness of the Canadair Regional Jet Maintenance Requirements Manual by incorporating new repetitive detailed inspections of the secondary load path indicator for the horizontal stabilizer trim actuator (HSTA). This AD is prompted by a report of a potential failure of the horizontal stabilizer trim actuator (HSTA) secondary nut in conjunction with a latent failure of the HSTA primary load path discovered during sampling program activities. We are issuing this AD to detect and correct latent failure of the primary load path of the HSTA, which, in conjunction with a failure of the HSTA secondary nut, could result in loss of horizontal trim control and consequent reduced controllability of the airplane.
70-25-09: 70-25-09 AEROSTAR: Amdt. 39-1127. Applies to Model 601, S/N's 61-001 through 61- 0070, and to all Model 600 airplanes equipped with the Aerostar Model 601 oxygen system. Compliance required within the next 10 hours time in service after the effective date of this AD unless already accomplished. To prevent possible short circuiting of the microphone jack connection by contact with the oxygen outlet receptacle due to their proximity to each other, accomplish the relocation specified in Aerostar Aircraft Corporation Service Bulletin No. S.B. 600-26 dated November 6, 1970, or later FAA-approved revisions, or other equivalent modification approved by the Chief, Engineering and Manufacturing Branch, Southwest Region, FAA. This amendment becomes effective to all known owners of Aerostar Model 601 airplanes and Model 600 airplanes equipped with the Aerostar Model 601 oxygen system upon receipt of individual copies mailed December 8, 1970 and to all other persons on December18, 1970.
96-12-07: This amendment supersedes an existing airworthiness directive (AD), applicable to Teledyne Continental Motors (TCM) (formerly Bendix) S-20, S-1200, D-2000, and D-3000 series magnetos equipped with impulse couplings, that currently requires inspections for wear, and replacement, if necessary, of the impulse coupling assemblies. This amendment requires replacement, if necessary, of worn riveted impulse coupling assemblies with serviceable riveted impulse couplings or snap ring impulse couplings. This amendment is prompted by the availability of an improved design for the impulse coupling assembly. The actions specified by this AD are intended to prevent magneto failure and subsequent engine failure.