Results
73-01-02: 73-01-02 CESSNA: Amdt. 39-1579. Applies to Model 310N (Serial Numbers 310N0001 thru 310N0198), Model 310P (Serial Numbers 310P0001 thru 310P0240), and Model 310Q (Serial Numbers 310Q0001 thru 310Q0130) airplanes. NOTE: This AD is not applicable to Model T310 airplanes which fall within these serial number ranges. Compliance: Required as indicated, unless already accomplished. To prevent moisture from possibly entering the alternate air system and freezing on powerplant induction system components when operating in below freezing temperatures, accomplish the following: Within 50 hours' time in service after the effective date of this AD, install deflector baffles over the alternate air inlet in the vertical baffle and enlarge the warm air cut out in the induction air canister in accordance with Cessna Service Letter ME70-43 and Service Kit SK310-82D, or any other method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Thisamendment becomes effective January 5, 1973.
71-06-03: 71-06-03 PRATT & WHITNEY: Amdt. 39-1168 as amended by Amendment 39-1211. Applies to all Pratt & Whitney Aircraft JT4A series turbojet engines which incorporate the following serial number 12th stage compressor rotor disc part number 360112. 2G5181 2G6360 2G5182 2G6362 2G5183 2G6363 2G5446 2G6445 2G5807 2G6448 2G5812 2G6449 2G5818 2G6454 2G5820 2G6581 2G5821 2G6583 2G5825 2G6585 2G5826 3G0201 2G6027 2G6030 5F7645 7 5F858 6F8362 5F7646 5F8588 6F9707 5F8544 through 5F8695 6F9739 5F8547 6F7206 6F9751 5F8550 through 6F7207 6F9753 through 5F8553 6F7210 6F9759 5F8558 6F7214 6F9760 through 5F8576 6F7218 6F9765 5F8577 6F7501 6F9846 5F8581 through 6F7502 6F9853 5F8584 6F7504 6F9861 6F7949 6F9866 6F8168 7F0156 6F8198 7F0480 7F0556 Pratt & Whitney Aircraft telegram PSE/RS/1-3-10-1 covers this same subject. Compliance required as indicated. To preclude twelfth stage compressor rotor disc failure as the result of suspected material deficiency, accomplish the following: 1. For all the previously listed serial number twelfth stage compressor rotor discs: A. With 18,000 hours or more in service within the next 50 hours in service after the effective date of this AD, replace the suspect serial number disc. B. With 18,000 hours or less in service prior to the accumulation of 18,050 hours in service, replace the suspect serial number disc. 2. Upon submission of substantiating data by an owner or operator through an FAA Maintenance Inspector, the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, may adjust the compliance time. Amendment 39-1168 was effective March 12, 1971. This amendment 39-1211 is effective May 19, 1971 and was effective for all recipients of the telegram dated March 12, 1971 which contained this amendment.
68-21-03: 68-21-03 PIPER: Amendment 39-667. Applies to Model PA-23-250 and PA-E23-250 Aircraft Serial Nos. 27-2505 through 27-3858 that Incorporate AiResearch Turbosupercharged Lycoming IO-540-J4A5 or IO-540-C4B5 Engines Installed in accordance with Supplemental Type Certificate No. SA909WE or SA978WE, or in accordance with Piper Aircraft Corporation Drawing No. 32016. Compliance required within 10 hours of aircraft operations after the effective date of this Airworthiness Directive, unless previously accomplished. To preclude failures of the exhaust tailpipe assemblies, AiResearch P/N 286-P23-060-5 and P/N 286-P23-060-9: (a) Visually inspect both the left and right hand engine tailpipes for cracks or deformation. If a crack or deformation is found in either tailpipe, retire the affected tailpipe assembly and replace with a new tailpipe assembly of the same part number. (b) Visually inspect both the left and right hand engine tailpipes for sufficient clearance betweenthe tailpipes and the firewalls and between the tailpipes and the cowl flaps in accordance with AiResearch Aviation Service Company Service Bulletin No. 14.1.8 dated April 25, 1968, or later FAA approved revisions. If sufficient clearance does not exist, install AiResearch Kit P/N 301-P23-063 and adjust for sufficient clearance in accordance with the above service bulletin. This amendment becomes effective October 28, 1968.
47-51-02: 47-51-02 CURTISS-WRIGHT Applies to Model C-46 Series Aircraft Equipped with Horizontal Stabilizer and Elevator Assemblies, Curtiss-Wright P/N's 20-110-5001 and 20-130- 5701, Respectively. Compliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished. The attachment bolts in the elevator hinges and the spring and trim tab bellcranks, located in the stabilizer, tend to loosen with resultant elongation of the holes and grooving of the bolts. 1. Replace the eight AN 5-14 hinge bolts on the 20-130-5701 elevator installation with NAS 55-14 or AN 175-14 bolts. 2. Replace the eight AN 4-26 bolts on the 20-130-5700 elevator installation with NAS 54-26 or AN 174-26 bolts. 3. Replace four AN 23-11A and two AN 23-12A bolts through 20-110-5020 or 20- 110-5112 blocks on each of the outboard stabilizer ribs with four NAS 53A-7 or AN 173-7A and two NAS 53A-10 or AN 173-10A bolts. Replace six existing bolts on eachof the inboard hinges with NAS 53A-7 or AN 173-7A bolts. 4. Fabricate spacers from 24ST material having an O.D. of 0.590 to 0.594 inch, 1.562 plus 0.005, minus 0.000 inches in length and drill concentric hole lengthwise 0.250 inch in diameter. Install spacer between the two hub bearings on "Idler Assem-Elev Trim Tab" P/N 20- 530-5722 and "Idler Assem-Elev Trim Tab (L.H.)" P/N 20-530-5775 so that the AN 24 bolts attaching the idlers to their mating bracket may be so tightened as to prevent rotation of the bolt in the inner bearing race or in the holes of the bracket. 5. Fabricate 0.75 diameter X1.012 inches plus 0.005, minus 0.000 spacers (Curtiss P/N 20-530-5709-1201) from 24ST alclad and drill 0.250. Install these spacers between the two hub bearings in the 20-530-5709 spring tab bellcranks. 6. P/N 1007-D-4-250 shoulder bushings should be installed in each 20-130-5775-2 bracket. Revised December 28, 1964.
73-08-01: 73-08-01 BOEING VERTOL: Amendment 39-1618 as amended by Amendment 39-2120. Applies to all 107-II type helicopters. Compliance required as indicated unless already accomplished. In order to detect surface defects which can cause complete failure of the P/N 107D2067 aft transmission quill shaft, accomplish the following: 1. Within 25 flight hours after the effective date of this AD, inspect and code mark, in accordance with paragraph 3, all P/N 107D2067-1 or -3 quill shafts having 600 hours total time except those quill shafts specified in Service Bulletin A107-320, Revision C. 2. Quill shafts with less than 600 hours total time, except those quill shafts specified in Service Bulletin A107-320, Revision C, must comply with paragraph 3 prior to accumulating 600 hours total time in service. 3. Strip quill shaft at the .375 inch diameter pin hole; inspect under microscope magnification, code mark acceptable units with suffix letter "R" and apply surface treatments in accordance with Boeing-Vertol Telex 8-1420-1-291 dated 16 November 1972, or Service Bulletin No. A107-320 dated April 30, 1973, and subsequent approved revisions, or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. Inspect for surface pits, arc burns and cracks. Replace any quill shaft which has surface pits, arc burns or cracks with a shaft which has been inspected in accordance with this paragraph or a shaft referenced in Service Bulletin No. A107-320, Revisions A and C. (Boeing-Vertol Service Bulletin Nos. A107-320A, A107-320B and A107-320C, and Boeing Vertol Telex Nos. 8-1420-1- 268 and 8-1420-1-254 and available specimen photographs pertain to this subject.) Amendment 39-1618 was effective April 9, 1973. This amendment 39-2120 is effective March 11, 1975.
2008-26-07: We are adopting a new airworthiness directive (AD) for all McDonnell Douglas airplanes identified above. This AD requires repetitive inspections of the lower skin and stringers at stations Xw=408 and Xw=-408, and corrective actions if necessary. This AD results from reports of cracks in the skins and stringers at the end fasteners common to the stringer end fittings at stations Xw= 408 and Xw=-408 wing splice joints. We are issuing this AD to detect and correct fatigue cracking in the skins and stringers at the end fasteners common to the stringer end fittings at certain station and wing splice joints, which could result in wing structure that might not sustain limit load, and consequent loss of structural integrity of the wing.
71-16-05: 71-16-05 BRITISH AIRCRAFT CORPORATION: Amdt. 39-1249. Applies to Model BAC 1- 11 200 and 400 series airplanes. To prevent a hazardous drift in the Voltage Sensing Unit, within the next 800 hours' time in service after the effective date of this AD, incorporate Rotax Modification SP 7174 by replacing the transistor T.1 in the Rotax Voltage Sensing Unit Type U.3619 or U.3619/1 with a new transistor, P/N.197235/1, in accordance with Rotax Service Bulletin No. 24-368 dated May 4, 1970, or later ARB-approved issue or an FAA-approved equivalent. (British Aircraft Corporation Model BAC 1-11 Service Bulletin 24-PM 4641, Revision 2, dated June 15, 1970, refers to this subject.) This amendment becomes effective July 29, 1971.
69-03-01: 69-03-01 VICKERS VISCOUNT: Amdt. 39-717. Applies to Model 744 Airplanes. Compliance required as indicated. (a) For airplanes with inner wing lower spar booms which have accumulated 7,000 or more landings on the effective date of this AD, before further flight, except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed, replace inner wing lower spar booms with new booms of the same part number. (b) For airplanes with inner wing lower spar booms which have accumulated less than 7,000 landings on the effective date of this AD, replace inner wing lower spar booms before the accumulation of 7,000 landings with new booms of the same part number. (c) All new inner wing lower spar booms that are installed in accordance with the requirements of paragraph (a) or (b) of this AD must be replaced before the accumulation of 7,000 landings. (d) For the purpose of complying with this AD, subject to acceptance by the assignedFAA maintenance inspectors, the number of landings may be determined by dividing each airplane's hours' time in service by the operator's fleet average time from takeoff to landing for the airplane time. This amendment becomes effective upon publication in the FEDERAL REGISTER and was effective upon receipt for all recipients of the telegram dated January 7, 1969, which contained this amendment.
67-33-02: 67-33-02 de HAVILLAND AIRCRAFT: Amdt. 39-532, Part 39, Federal Register December 30, 1967. Applies to Type DHC-6 Aircraft. Due to flutter of the left elevator trim tab, accomplish the following after the effective date of this AD unless already accomplished. (a) Prior to the next flight, install the following placard in full view of the pilot: "SPEED SHALL NOT EXCEED 125 KNOTS IAS (144 M.P.H. IAS)". (b) Prior to the next flight, accomplish the following one-time inspections on the left elevator-trim tab assembly: (1) Remove the bolt between the push rod and the tab operating lever and operate the trim tab control in the cockpit until the screw jack in the elevator is fully retracted. Lower the tab until the bolt can be inserted through the operating arm and the push rod without exerting pressure on the trim tab. If this can be accomplished, reconnect the tab to the push rod. If the holes do not align, adjust the push rod sufficiently to allow assembly withoutexerting pressure on the trim tab. (2) With the trim tab in the neutral position, assure that the total vertical free play, measured adjacent to the operating lever, does not exceed one-tenth (0.1) inches at the trailing edge of the tab. Comply with this limit prior to further flight. (c) Before each flight, visually inspect the left elevator trim tab top and bottom skins for diagonal buckling or concavity, other than the manufactured flutes. Trim tabs found to contain diagonal buckles or concavities must be replaced before further flight with a part of the same part number that has been inspected in accordance with this paragraph, or with an equivalent part approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. The repetitive inspection required by this paragraph also applies to replacement parts. (d) The placard required by (a) may be removed, and the repetitive inspection required by (c) may be discontinued, when a trim tab modification,approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, is incorporated. (e) Upon request with substantiating data submitted through an FAA maintenance inspector, the compliance time specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This amendment effective December 30, 1967, for all persons except those to whom it was made effective immediately by telegram dated December 1, 1967.
2008-26-02: The FAA is superseding an existing airworthiness directive (AD) for certain GE CT7-8A turboshaft engines. That AD currently requires initial and repetitive inspections of the electrical chip detectors for the No. 3 bearing. This AD requires removing from service certain GE CT7-8A turboshaft engines within 6,200 cycles-since-new. This AD results from investigation for the root causes of two failures of the No. 3 bearing. We are issuing this AD to prevent failure of the No. 3 bearing due to contamination by aluminum oxide, which could result in a possible in-flight shutdown of the engines and loss of control or forced landing of the aircraft.
67-07-02: 67-07-02 DASSAULT: Amdt. 39-354, Part 39, Federal Register February 25, 1967. Applies to Fan Jet Falcon Airplanes, Serial Numbers 1 thru 69, except, Numbers 59, 62, 63, and 67. Compliance required as indicated, unless already accomplished. To prevent inflight jamming of the horizontal stabilizer actuator, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 50 hours' time in service from the date of the last inspection, inspect the horizontal stabilizer actuator jack assembly, P/N's 21-59-0 or 21-59-1-A, for the presence of moisture in accordance with AMD Service Bulletin No. 270 (27-34), dated January 13, 1967, or later SGAC-approved revision. (b) If, during the inspections required by (a), the jack body and electric motors show no trace of moisture, before further flight, drill a drain port in the covers of the actuator motors, P/N SEB993A, in accordance with AMDService Bulletin No. 88, Revision 2, dated January 10, 1967, or a later SGAC-approved revision. If the jack body is dry but the motors show traces of moisture, before further flight, dry the motors and comply with the requirements of the first sentence of this paragraph. If the jack body shows traces of moisture, before further flight, dry the jack and replace both the normal and emergency motors, P/N SEB993A, with motors P/N CEB993B. (c) Within the next 200 hours' time in service after the inspection required by (a), replace the actuator jack assembly, P/N's 21-59-0 or 21-59-1-A with a modified jack, P/N's 21- 59-0-E or 21-59-1-B. Upon accomplishing this replacement, the inspection requirements of (a) may be discontinued. This directive effective March 2, 1967.
2008-25-06: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: * * * Airbus has advised that an incorrect part number has been introduced in the IPC (illustrated parts catalog) * * * for the rear engine mount barrel nut. This problem affects Airbus A319, A320 and A321 models with IAE (International Aero Engine) V2500-A5 engines. The part number introduced in error is not certificated for the IAE V2500-A5 engine installation and, if installed, may fail in service. * * * * * Failure of the rear engine mount barrel nut could result in reduced structural integrity of the rear engine mount and possible separation of the engine from the airplane, and a consequent hazard to the airplane and persons and property onthe ground. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
2008-25-03: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: Bombardier Aerospace has completed a system safety review of the CL-600-2C10/CL-600-2D24 aircraft fuel system against new fuel tank safety standards. * * * The assessment showed that due to the close proximity of intrinsically safe fuel system wiring with other wiring, a single failure from wire chafing at various locations of the fuselage could result in an ignition source inside the fuel tank. In addition, chafing of the temperature sensor wiring against the high power wiring in the avionics compartment could lead to overheating of the temperature sensor and hot surface ignition. The presence of an ignition source inside the fuel tank could result in a fuel tank explosion. We are issuing this AD to require actions to correct the unsafe condition on these products.
66-27-04: 66-27-04 CANADAIR: Amdt. 39-299 Part 39 Federal Register November 2, 1966. Applies to Model CL-44D4 airplanes. Compliance required as indicated. To detect cracks in the main landing gear keel vertical posts and horizontal lower reinforcing angle members in the inboard nacelles, accomplish the following: (a) For airplanes incorporating the modification specified in Canadair Service Information Circular No. 247-CL44D4, dated April 9, 1963, within the next 50 hours' time in service after the effective date of this AD, unless already accomplished within the last 450 hours' time in service, and thereafter at intervals not to exceed 500 hours' time in service from the last inspection, visually inspect the vertical post members and nesting angles for cracks. (b) For airplanes other than those specified in (a), within the next 50 hours' time in service after the effective date of this AD, unless already accomplished within the last 450 hours' time in service, and thereafterat intervals not to exceed 500 hours' time in service from the last inspection, visually inspect the vertical post members and lower reinforcing angles for cracks, in accordance with paragraphs (a), (b), (c) and (d) of "Inspection Procedure" of Canadair Service Information Circular, Issue No. 2, No. 247-CL44D4, dated May 17, 1966, or later FAA-approved revision or an FAA-approved equivalent. (c) If a crack is found during an inspection required by (a) or (b), repair the crack in an FAA-approved manner, or replace the part with a part of the same part number that has been inspected in accordance with this AD and found free of cracks or with an equivalent part approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, before further flight, except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed. (d) The repetitive inspection of the vertical post members required by (a) and (b) may be discontinued whenthe vertical post members are modified with a modification approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (e) The repetitive inspection of the nesting angles required by (a) may be discontinued when the modification specified in Canadair Service Information Circular No. 247- CL44D4, dated April 9, 1963, is disassembled, the lower reinforcing angles and nesting angles are visually inspected for cracks and repaired or replaced as necessary in accordance with (c), and reinstalled in accordance with Canadair Service Bulletin CL44-452, dated July 8, 1966, or later FAA-approved revision, or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (f) The repetitive inspection of the lower reinforcing angle members required by (b) may be discontinued when the lower reinforcing angles are modified in accordance with Canadair Service Bulletin CL44-452, dated July 8, 1966, or later FAA-approved revision, or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (g) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. This directive effective November 12, 1966.
67-24-02: 67-24-02 PIPER: Amdt. 39-464, as amended by amendment 39-885. Applies to Type PA-22 Aircraft Serial Nos. 22-1 to 22-7642 Inclusive. Compliance required within the next 15 hours in service after December 6, 1969, the effective date of this AD as amended. To forestall the possibility of engine fuel starvation during takeoff operations, install a placard on the right fuel quantity gauge, as shown in Piper Service Bulletin No. 250 dated June 2, 1967. The placard shall read: "Right Tank Level Flight Only With Less Than 1/3 Tank." A 1/3 tank capacity equals 6 gallons. Aircraft equipped with a single fuel quantity gauge must also have the placard installed. 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. 522(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Piper Aircraft Corporation, Lock Haven, Pa. 17745. These documents may also be examined at the Office of Regional Counsel, Eastern Region, J.F.K. International Airport 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 Eastern Region in Jamaica, New York. Effective August 17, 1967. Revised December 6, 1969.
2008-25-07: We are adopting a new airworthiness directive (AD) for all Boeing Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747-200C, 747- 200F, 747-300, 747-400, 747-400D, 747-400F, and 747SR series airplanes. This AD requires repetitive inspections for cracks or fractures of the forward end attachment and the forward lower flange of the flap tracks of the trailing edge flaps, and corrective actions if necessary. For certain airplanes, this AD would also require modifying the fail-safe links of the main carriage. This AD results from a detailed structural analysis of the flap attach structural and fail-safe components, accomplished as a result of a dynamic stability and control analysis, which could not demonstrate continued safe flight and landing of the airplane after the loss of a trailing edge flap. We are issuing this AD to detect and correct cracks or fractures of the primary structural and fail-safe load paths of the inboard and outboard trailing edge flaps, which could resultin the loss of a flap during takeoff or landing, reducing flightcrew ability to maintain the safe flight and landing of the airplane.
2008-24-14: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: Cracks on the main landing gear trunnion fitting web have been discovered during fatigue testing. Failure of the main landing gear trunnion fitting web could compromise the structural integrity of the trunnion fitting and result in a main landing gear collapse. * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
49-23-01: 49-23-01 BOEING: Applies to All Model 75 Series Airplanes With Crop Dusting or Seeding Hopper Installations. Compliance required prior to next periodic inspection. Inspect to determine whether fuselage bottom truss was altered for installation of hopper throat. All alterations involving the removal or revision of the truss members require that equivalent structural strength be provided. One open bay in the bottom truss either immediately forward or immediately aft of the cross member (streamline tube) at Station 2L is permissible provided that it is limited to a rectangle bounded by the longerons, the above-mentioned cross member and a 7/8 x 0.035, or larger, x4130 tube parallel to the streamline tube and not more than 8 inches forward or aft thereof. The inside corners of this open rectangle should have 0.065-inch x 4130 gussets, or equivalent, extending along the longerons at least 2 inches. (Boeing Report No. WD-10645 covers this same subject and includes an alternate alteration recommended as preferable to the above. Copies of the report are obtainable from the Boeing Airplane Co., Wichita Division, Wichita 1, Kansas.) This supersedes AD 46-31-03.
2008-18-06: We are adopting a new airworthiness directive (AD) for McDonnell Douglas Model DC-9-10, DC-9-20, DC-9-30, DC-9-40, and DC-9-50 series airplanes, equipped with tail cone evacuation slide containers as specified above. This AD requires modifying the tail cone slide. This AD also requires additional tail cone drops and slide deployments, and repair if necessary. This AD results from several reports of inadvertent tail cone deployments in which the tail cone slide failed to deploy. We are issuing this AD to ensure that the tail cone evacuation slide deploys correctly. Failure of the slide to deploy during an emergency evacuation could result in injury to flightcrew and passengers.
2006-20-51 R1: We are revising an existing airworthiness directive (AD) that applies to certain Boeing Model 777-200LR and -300ER series airplanes. That AD currently requires revising the airplane flight manual to prohibit takeoffs at less than full-rated thrust. This new AD reduces the applicability of the existing AD. This AD results from a report of two occurrences of engine thrust rollback during takeoff. We are issuing this AD to prevent dual-engine thrust rollback during the takeoff phase of flight, which could result in the airplane failing to lift off before reaching the end of the runway or failing to clear obstacles below the takeoff flight path.
73-05-01: 73-05-01 AVCO LYCOMING: Amendment 39-1595 as amended by Amendment 39-1843 is further amended by Amendment 39-1889. Applies to VO-540 Series engines beginning with S/N 101-43 thru 2299-43, IVO-540 series engines beginning with S/N 101-60 thru 150-60, TVO-540 Series engines beginning with S/N 101-53 thru 103-53 and TIVO-540 Series engines beginning with S/N 101-57 thru 163-57 and all similar series engines overhauled (also known as remanufactured) by Lycoming before January 2, 1974 except the following Serial Numbers: RL-376-43, RL-483-43, RL-693-43, RL-695-43, RL-1091-43, RL-1205-43, RL-1429-43, RL- 1753-43, RL-1960-43, RL-2017-43, RL-2032-43, RL-2125-43, RL-2139-43, RL-2244-43, RL-2245-43 and RL-2275-43. Compliance required as indicated after the effective date of this AD, unless already accomplished or unless higher crush bearing P/N LW13212 has been installed in accordance with Lycoming Service Bulletins No. 303E or 303F. To prevent failures of P/N 71947, P/N 73174, P/N75548 and P/N LW10776 connecting rod assemblies and P/N 77450 connecting rod assembly with P/N 75547 connecting rod bearing installed in accordance with Lycoming Service Bulletins No. 303B, 303C or 303D, accomplish the following: a. Engines with connecting rod assemblies that have accumulated 400 hours or more in service since new or overhaul must have the connecting rod assemblies replaced with new connecting rod assembly, P/N LW13422 and new connecting rod bearing P/N LW13212, within the next 50 hours in service. b. Engines that have connecting rod assemblies with less than 400 hours in service must have the connecting rod assemblies replaced with new connecting rod assembly, P/N LW13422 and new connecting rod bearing, P/N LW13212 prior to 450 hours in service. c. Engines that have been modified in accordance with Lycoming Service Bulletins No. 303B, 303C or 303D must comply with items (a) or (b) above, except that the P/N 77450 connecting rod, free of any galling maybe reused with P/N LW12596 connecting rod bolts and identified as assembly number LW13422. NOTE: Lycoming Service Bulletin No. 371A applies to this subject. Amendment 39-1595 superseded AD 65-13-07 and became effective February 28, 1973. Amendment 39-1843 was effective May 17, 1974. This Amendment 39-1889 is effective July 9, 1974.
67-25-07: 67-25-07 VICKERS: Amdt. No. 39-473, Part 39, Federal Register September 7, 1967. Applies to Viscount Models 744, 745D and 810 Series Airplanes. Within the next 1,000 hours' time in service after the effective date of this AD, unless already accomplished: Replace Graviner D.1870 or D.1870N/1 loop type overheat detectors with Graviner D.5352N/1 coil overheat detectors. Replace Graviner point type detectors as shown in the following table: Existing P/N Replacement P/N Graviner 4D/3 Graviner 134D/3 Graviner 13D/3 Graviner 135D/3 Graviner 68D/3 Graviner 136D/3 Graviner 150D/07/180 Graviner 150D/07/210 Detectors which are equivalent to the specified Graviner detectors may be used as replacements, subject to the approval of the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region. Replace the detectors in accordance with British Aircraft Corporation Limited (BAC) Modification Bulletin No. D.3187, Issue 2, (700 Series), and Modification Bulletin No. FG.2055, Issue 2, (810 Series), or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region. This amendment effective September 7, 1967.
67-12-06: 67-12-06 PIPER: Amdt. 39-386 Part 39 Federal Register April 4, 1967. Applies to Models PA-28 and PA-32 Series Airplanes as follows: GROUP I - PA-28-140, Serial Nos. 28-20000 through 28-20622; PA-28-150-160-180, Serial Nos. 28-2 through 28-35; 28-37 through 28-497; 28-499 through 28-543; 28-545 through 28-1307; 28-1309 through 28-2136; PA-28-235, Serial Nos. 28-10000 through 28-10590. GROUP II - PA-28-140, Serial Nos. 28-20623 through 28-21383; PA-28-150-160-180, Serial Nos. 28-2137 through 28-3021; PA-28-235, Serial Nos. 28-10591 through 10719; PA-32-260, Serial Nos. 32-1 through 32-307. GROUP III - PA-28-140, Serial Nos. 28-21384 through 28-21764; 28-21767 through 28-21786; 28-21788 through 28-21807; 28-21809 through 28-21835; 28-21837 through 28-21856; 28-21858 through 28-21877; 28-21879 through 28-22003; 28-22005 through 28-22010; 28-22012 through 28-22026; 28-22028 through 28-22032; 28-22035 through 28-22040; 28-22043 through 28-22050; 28-22052 through 28-22056; 28-22058, 28-22059, 28-22061 through 28-22064; 28-22066, 28-22071, 28-22074, 28-22080, 28-22082, 28-22088. PA-28-150-160-180, Serial Nos. 28-3022 through 28-3499; 28-3501 through 28-3503; 28-3505 through 28-3508; 28-3510; 28-3512 through 28-3528; 28-3530 through 28-3537; 28-3541, 28-3543, 28-3545 through 28-3549; 28-3552; 28-3555 through 28-3557; 28-3562. PA-28-235, Serial Nos. 28-10720, 28-10721, 28-10732. PA-32-260, Serial Nos. 32-308 through 32-540; 32-542 through 32-558; 32-560 through 32-570, 32-572 through 32-677; 32-679 through 32-683; 32-685, 32-689, 32-690; 32-692 through 32-698; 32-700 through 32-702; 32-704 through 32-711; 32-714, 32-716, 32-717, 32-719 through 32-722; 32-725, 32-727, 32-730, 32-733; 32-740 through 32-742; 32-744 through 32-746; 32-749. Compliance required as follows: Group I - Prior to but not later than July 31, 1967. Group II - Prior to but not later than July 31, 1968. Group III - Prior to but not later thanJuly 31, 1969. At the option of the local FAA General Aviation District office inspector, these compliance times may be extended, for a period not to exceed 30 days, to coincide with the annual inspection for the aircraft involved. Due to the possibility of internal corrosion resulting from inadequate corrosion protection of certain open end steel tube assemblies, accomplish the inspections described below on the following parts: PART NO. NOMENCLATURE MODELS AFFECTED 62369-0 Aileron Balance Weight PA-28-140-150-160-180-235 62369-1 Aileron Balance Weight PA-32-260 63546 Rudder Horn Assembly PA-28-140-150-160-180-235 PA-32-260 63578 Balance Weight Assembly-Stabilator PA-28-140-150-160-180 65310 Balance Weight Assembly-Stabilator PA-28-235 68432 Balance Weight Assembly-Stabilator PA-32-260 (a) Conduct a close visual inspection of the interior of the open end tubes specified above for a protective coating and evidence of corrosion,in accordance with Inspection Procedure, Piper Service Bulletin No. 240, dated December 13, 1966, or later FAA-approved revision, or by a method approved by the Chief, Engineering and Manufacturing Branch, FAA Southern Region. Inspect both right and left aileron balance weight assemblies in accordance with Inspection of Aileron Balance Weight assembly provision of this Service Bulletin. (b) If there is evidence of a protective coating on the interior of the tube, and there is no evidence of corrosion, further inspection is not required. (c) If there is no evidence of a protective coating on the interior of the tube and evidence of corrosion, or if there is a protective coating, and evidence of corrosion, accomplish the following: (1) Remove all corrosion from the interior of the tube in accordance with the instructions in Piper Service Bulletin No. 240, dated December 13, 1966, or later FAA-approved revision, or by a method approved by the Chief, Engineering and Manufacturing Branch, FAA Southern Region. (2) Replace all corroded parts with a new part of the same part number if the corrosion cannot be removed as provided for in (c) (1). If doubt exists as to whether or not the corrosion involved can be removed without replacing the part, the matter must be referred to the local FAA General Aviation District office for assistance in making a determination. (3) If the extent of corrosion does not require replacement of the part in accordance with (c)(2), and the corrosion has been removed in accordance with (c)(1), apply a zinc chromate primer, Spec. MIL-P-8585, or an FAA-approved equivalent, to the inside of the tube to prevent further corrosion. (d) Further inspection is not required after accomplishing (c)(1), (c)(2), or (c)(3).
2008-24-01: 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: Three instances have occurred in which the aircraft took off with pre-mod 6/1676 flight control gust locks still installed, sometimes with disastrous results. Based on investigation, the FAA and National Transportation Safety Board (NTSB) believe that an attempted takeoff with the gust locks installed could be the cause of a recent accident in Hyannis, Massachusetts. We are issuing this AD to require actions to correct the unsafe condition on these products.
71-14-03: 71-14-03 INTERNATIONAL INFLATABLES COMPANY: Amdt. 39-1239. Applies to aircraft incorporating International Inflatables Company Regulator, P/N 68240. Compliance required as indicated. To prevent unwanted bottle and regulator separation of P/N 68240, accomplish the following: (a) Within the next 30 days after the effective date of this A.D., unless already accomplished within the last 30 days, and thereafter at intervals not to exceed 60 days from the last inspection, inspect the International Inflatables Company Regulator, P/N 68240, if installed in an aircraft, in accordance with International Inflatables Company Service Bulletin No. 33-102, Volume 1, No. 101, dated June 11, 1971, or later FAA approved revisions, or an equivalent inspection procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region. If there is any evidence of corrosion, replace the regulator prior to further flight with a previously inspected (per this A.D.) and corrosion-free regulator. Do not return any regulator exhibiting evidence of corrosion to service. (b) After the effective date of this AD, and prior to the installation of an International Inflatables Company Regulator, P/N 68240 on an aircraft, inspect that regulator per (a) above. NOTE: An improved regulator is under development. If and when approved by FAA, this AD will be amended to require installation within a prescribed time period. This amendment becomes effective July 2, 1971.