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79-11-06: 79-11-06 PIPER: Amendment 39-3481 as amended by Amendment 39-3545. Applies to all PA-23 series aircraft, certificated in all categories, which have landing gear selector lever, P/N 752303, installed. To prevent possible failure of the landing gear selector lever, Piper Part No. 752303 accomplish the following: a. Within the next 50 hours in service after the effective date of this AD, unless already accomplished, inspect the landing gear selector lever in accordance with Instructions Section in Piper S/B No. 635 steps (1) through (4) or equivalent inspection. b. If cracks are found, replace the landing gear selector lever with lever Piper P/N 761213 or equivalent before further flight. c. If no cracks are found, repeat the inspection in (a) at intervals not to exceed 100 hours in service. The requirements of this AD may be cancelled upon installation of landing gear selector lever Piper P/N 761213 or equivalent. d. Equivalent inspections and replacementsmust be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. e. Upon submission of substantiating data by an owner or operator through an FAA Maintenance Inspector, the Chief, Engineering and Manufacturing Branch, Eastern Region, may adjust the inspection intervals specified in this AD. Amendment 39-3481 was effective June 5, 1979. This Amendment 39-3545 is effective August 29, 1979.
88-12-10: 88-12-10 GARRETT ENGINE DIVISION, ALLIED-SIGNAL INCORPORATED (FORMERLY GARRETT TURBINE ENGINE COMPANY, FORMERLY AIRESEARCH MANUFACTURING COMPANY OF ARIZONA): Amendment 39-5910. Applies to model TPE331-10, -10R, -10U, -10UA, -10UF, -10UG, -10UGR, -10UR, and -11U turboprop engines equipped with second stage turbine rotors, part numbers 3102106-1, -6, and -8, installed in aircraft certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent an uncontained engine failure, accomplish the following: (a) Remove from service the second stage turbine rotor per the schedule below in accordance with the accomplishment instructions of Garrett Alert SB TPE331-A72-0571, dated March 31, 1988: Second Stage Turbine Rotor Cycles Since New (CSN) Removal Schedule 0 to 4,400 Prior to 4,800 cycles since new CSN. 4,401 to 5,000 Within 400 cycles after the effective date of this AD or 5,200 CSN, whichever occurs first. 5,001 to 5,900 Within 200 cycles after the effective date of this AD or 6,000 CSN, whichever occurs first. 5,901 to 6,800 Within 100 cycles after the effective date of this AD or 6,800 CSN, whichever occurs first. (b) Remove from service prior to January 1, 1989, all second stage turbine rotors with 4,800 or more cycles since new CSN, regardless of the schedule provided in paragraph (a) above. NOTE: Garrett SB TPE331-72-0180 Revision 12, dated March 31, 1988, which defines critical component service life limits includes the required life limit and states the cycle definitions. (c) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (d) Upon request, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Los Angeles Aircraft Certification Office, Federal Aviation Administration, Northwest Mountain Region, 4344 Donald Douglas Drive, Long Beach,California 90808. (e) Upon submission of substantiating data by an owner or operator through an FAA maintenance inspector, the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region, may adjust the compliance schedule specified in this AD. Garrett SB, TPE331-A72-0571, dated March 31, 1988, identified and described in this document, is 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 the manufacturer's SB may obtain copies upon request to Garrett General Aviation Services Division, Distribution Center, 2340 East University, Phoenix, Arizona 85034. This document may also be examined at the Office of the Regional Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, Room 311, Rules Docket Number 88-ANE-18, between the hours of 8:00 a.m. and 4:30 p.m., Monday through Friday, except federal holidays. This amendment, 39-5910, becomes effective June 6, 1988.
76-17-08: 76-17-08 ENSTROM: Amendment 39-2700 as amended by Amendment 39-3043. Applies to Models F-28, F-28A, and 280 helicopters certificated in all categories, except for those helicopters which have the following main rotor shaft gear box serial numbers. 032FS 07-012-74PS 39-002-13PS 041PS 08-001-72S 44-011-76PS 067PS 10-002-72PS 44-019-76PS 070PS 19-024-74PS 53-003-73PS 091 34-029-75PS To detect tool marks, surface irregularities, and cracks which may develop into failure of the Main Rotor Shaft (Enstrom Part Number 28-13104) accomplish the following: A. Before further flight visually inspect the main rotor shaft in the area of the radius beneath the rotor hub shoulder, using an eight power or greater magnifying glass and report to the Enstrom Helicopter Corporation, Menominee County Airport, P.O. Box 277, Menominee, Michigan, Phone 906-863-9971, any evidence of circumferentially disposed tool marks or surface irregularities. B. Before further flight, unless previously accomplished within the last 20 hours time in service or one month, and every 20 hours or one month thereafter, whichever occurs first, inspect the main rotor shaft for cracks in the area of the radius beneath the rotor hub shoulder using a 3-step dye-penetrant method. The inspections are to be performed by maintenance personnel familiar with the dye-penetrant inspection method. C. Within 10 hours time in service or 15 days, whichever occurs first after each dye- penetrant inspection, visually check the main rotor shaft for cracks in the area of the radius beneath the rotor hub shoulder using an eight power or greater magnifying glass. D. Before further flight remove from service any main rotor shaft found to contain cracks or other evidence of damage and replace with an airworthy shaft of the same part number or later FAA approved part number. E. Immediately check in accordance with Paragraph B above any helicopter which develops unusual once-per-rotor-revolution vibration. Such vibrations serve as warning of imminent failure. F. Within the next 200 hours time in service or six months after the effective date of this amendment, whichever occurs first, remove main rotor shaft from service or return gearbox and shaft to Enstrom Helicopter Corporation for modification in accordance with Enstrom Service Directed Bulletin No. 0036. G. Repetitive inspections of paragraphs B and C above may be discontinued after main rotor shaft modification in accordance with Enstrom Service Directed Bulletin No. 0036. 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 the Enstrom Helicopter Corporation, Menominee County Airport, P.O. Box 277, Menominee, Michigan 49858.These documents may also be examined at the FAA Great Lakes Region, Engineering and Manufacturing Branch, AGL-210, 2300 East Devon Avenue, Des Plaines, Illinois. A historical file on this AD which includes the incorporated material in full is also maintained at that office. Amendment 39-2700 superseded Amendment 39-2472 (40 F.R. 59197), AD 75-26-19. Amendment 39-2700 became effective September 1, 1976 and was effective immediately for all recipients of the airmail letters dated August 6, 1976 which contained this amendment. This amendment 39-3043 becomes effective October 4, 1977.
98-12-33: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320 series airplanes, that requires repetitive inspections to detect fatigue cracking on the connecting angle between frame 56 and the right-hand frame support at stringer 38; and replacement of the connecting angle, if necessary. This amendment also provides for an optional terminating action for the repetitive inspections. 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 detect and correct fatigue cracking on the connecting angle, which could result in reduced structural integrity of the airplane.
2020-07-02: The FAA is adopting a new airworthiness directive (AD) for all Pratt & Whitney (PW) PW1519G, PW1521G, PW1521G-3, PW1521GA, PW1524G, PW1524G-3, PW1525G, and PW1525G-3 model turbofan engines. This AD requires the removal from service of certain electronic engine control (EEC) full authority digital electronic control (FADEC) software and the installation of a software version eligible for installation. This AD was prompted by reports of four in-flight shutdowns (IFSDs) due to failure of the low-pressure compressor (LPC) rotor 1 (R1) and by subsequent findings of cracked LPC R1s during inspections. The FAA is issuing this AD to address the unsafe condition on these products.
2020-04-22: The FAA is superseding Airworthiness Directive (AD) 2018-19-27 and AD 2014-16-12, which applied to certain Dassault Aviation Model FALCON 2000EX airplanes. AD 2018-19-27 and AD 2014-16-12 required revising the existing maintenance or inspection program, as applicable, to incorporate new maintenance requirements and airworthiness limitations. This AD retains those actions and requires revising the existing maintenance or inspection program, as applicable, to incorporate additional new or more restrictive airworthiness limitations. This AD was prompted by the FAA's determination that new or more restrictive airworthiness limitations are necessary. The FAA is issuing this AD to address the unsafe condition on these products.
79-17-05: 79-17-05 PIPER: Amendment 39-3531. Applies to Model PA-38-112 serial numbers 38-78A0002 thru 38-78A0027, 38-78A0029 thru 38-78A0037, 38-78A0039 thru 38-78A0041, 38-78A0043, 38-78A0045 thru 38-78A0064, 38-78A0066 thru 38-78A0072, and 38-78A0074 thru 38-78A0104. Compliance required as indicated unless already accomplished. a. To prevent the occurrence of insufficient or intermittent grounding of the fuel gauges and engine instrument cluster, accomplish the following: Within the next 25 hours in service after the effective date of this AD, install ground wire harness assembly P/N 77960-400 and connect leads to instruments and terminal ground strip in accordance with Piper Service Bulletin No. 603, or an approved equivalent alteration. Equivalent alterations must be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration (FAA), Eastern Region. b. Aircraft may be flown in accordance with 21.197 to a base where the modification required by this AD can be accomplished. This amendment is effective August 22, 1979.
78-16-08: 78-16-08 PIPER AIRCRAFT CORPORATION: Amendment 39-3223 as amended by amendment 39-3497. Applies to Model PA-32R-300 serial numbers 32R-7680001 through 32R- 7880068 inclusive; Model PA-32RT-300 serial numbers 32R-7885001 through 32R-7885207 inclusive, airplanes, certificated in all categories. Compliance is required as indicated. To prevent engine failure due to engine oil loss, accomplish the following: (a) On or before September 30, 1979, unless already accomplished, accomplish either (1) or (2): (1) Replace both engine oil coolers, Piper Part Number 16599-00 (Harrison Part number AP09AU06-01), if installed, with Piper Oil Cooler replacement kit 763-859V. (2) Modify the oil cooler system in accordance with Aircraft Metal Products STC SA3736WE dated September 27, 1978. (b) Prior to the first flight of each day and immediately after the last flight of each day, unless (a) has been accomplished: (1) Check below the nose gear wheel well for signsof fresh engine oil deposits. (2) Check the following areas of the airplane for engine oil deposits. (i) Inside surface on nose gear doors. (ii) Inside lower edge of lower cowl. (iii) Lower forward face of firewall. (iv) Aft face of nose gear strut, linkage, and tire. (v) Forward belly of fuselage. (3) If no oil leakage is evident, make appropriate maintenance record entry. (4) If oil leakage is evident and an oil cooler is determined to be leaking or contains a crack, have it replaced with a serviceable oil cooler. Refer to the appropriate Piper Service Manual for replacement instructions. Special caution is necessary during installation of the oil hoses to prevent damage to the replacement oil cooler. (c) Within the next 10 hours time in service after the effective date of this A.D. and thereafter at intervals not to exceed 10 hours time in service from the last check, unless (a) has been accomplished: (1) Insure thatthe engine oil is at normal operating temperature by conducting this check immediately after a flight or a ground run. (2) Gain access to the engine by removing the top engine cowl. (3) Check both the right and left engine oil coolers in the area of the oil hose assembly end fittings and in the area of the fluted portion at the oil cooler end tanks for cracks and oil leakage. (i) If no oil leakage or cracks are found, make appropriate maintenance record entry. (ii) If an oil cooler is determined to be leaking or contains a crack, have it replaced with a serviceable oil cooler. Refer to the appropriate Piper Service Manual for replacement instructions. Special caution is necessary during installation of the oil hoses to prevent damage to the replacement oil cooler. (d) An alternate method of compliance must be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southern Region. The checks in this A.D. may be accomplished by the pilot and appropriate maintenance record entries made in accordance with FAR 91.173. Inspections and component replacements must be accomplished by a person authorized by FAR 43.3. Piper Service Bulletin 586B, dated July 24, 1978, pertains to this same subject. Amendment 39-3223 superseded Amendment 39-3106, 43 FR 4, A.D. 78-01-03. Amendment 39-3223 became effective August 11, 1978. This amendment 39-3497 becomes effective June 25, 1979.
50-52-01: 50-52-01\tHAMILTON STANDARD: Applies to All Convair Model 240 Series, Douglas DC-6 and Martin Models 202 and 202A Aircraft Equipped With Hamilton Standard 2H17 Series Blades.\n \n\tCompliance required daily until further notice.\n \n\tThere have been several cases in which a crack has been detected in Hamilton Standard 2H17 blades during routine ground inspections and recently there was a case in which a section of the blade tip shell was lost in flight requiring an unscheduled landing. In order to eliminate the possibility of other blade failures, the following precautionary measures must be taken:\n \n\t(1)\tThoroughly clean the entire surface of each blade to remove oil, grease, dirt, etc., so that an adequate inspection of the entire blade surface can be made.\n \n\t(2)\tCarefully examine visually at close range (12 inches-14 inches) and in detail, the entire surface of each blade for cracks and surface defects in accordance with Hamilton Standard Service Bulletins Nos. 177 and 193. Any suspected areas should then be more closely examined by using a suitable magnifying glass, permanent magnet or any other suitable means as required.\n \n\t(3)\tIf any cracks are found in the blade surface, it must be retired immediately from service. All doubtful cases should be referred to the propeller manufacturer.\n \n\tThis directive supplements previous Hamilton Standard information on the same subject to all affected operators.
80-13-05 R1: 80-13-05 R1 ROLLS-ROYCE, LTD: Amendment 39-3804 as amended by Amendment 39-4742. Applies to Rolls-Royce RB211-22B turbofan engines. Compliance required as indicated, unless already accomplished. To preclude possible high pressure compressor (HPC) rotor stage 1 and 2 disk assembly failure, remove from service all HPC rotor stage 1 and 2 disk assemblies, listed by part number and serial number in Appendices 1 and 2 of Rolls-Royce Alert Service Bulletin No. RB211-72-A5722, Revision 3, dated April 17, 1981, in accordance with the following compliance schedule: 1. For disk assemblies listed in Appendix 1: a. After July 31, 1980, no disk assemblies may exceed 7,000 flight cycles. b. After December 31, 1980, no disk assemblies may exceed 6,000 flight cycles. c. All remaining disk assemblies by April 30, 1981. 2. For disk assemblies listed in Appendix 2: a. After the effective date of this amendment, no disk assemblies may exceed 9,000 flight cycles. NOTE: For the purpose of this AD, a flight cycle is considered to be an engine operating sequence from takeoff to landing. Replace with an FAA-approved, serviceable HPC rotor stage 1 and 2 disk assembly. The manufacturer's Alert Service Bulletin identified and described in this directive is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a). All persons affected by this directive who have not already received the service bulletin from the manufacturer may obtain copies upon request to Technical Publications Department, Rolls-Royce, Ltd., Derby, England DE2 8BJ. The service bulletin may also be examined at Federal Aviation Administration, New England Region, Office of the Regional Counsel, 12 New England Executive Park, Burlington, Massachusetts. Amendment 39-3804 became effective upon publication in the Federal Register. This Amendment 39-4742 becomes effective November 28, 1983.