Results
46-33-02: 46-33-02 GLOBE: (Was Mandatory Note 2 of AD-766-5.) Applies to Models GC-1A, GC-1B Aircraft Serial Numbers 3 to 174 Inclusive. Compliance required prior to October 1, 1946. Install stiffeners, Globe P/N 11-213-1471-1 R & L, on the flange of the upper bulkhead at fuselage Station 167, with six rivets per stiffener, to prevent the formation of cracks originating at the joggles in the flange outboard of the stabilizer attachment points. Any crack should be stop drilled. If it extends into the web of the bulkhead an 0.040-inch 24ST alclad reinforcing plate extending to the flange should be installed on the web with rivets spaced not more than 3/4 inches apart. (Globe Customer Service Maintenance Bulletin No. 2 covers this same subject.)
86-18-06: 86-18-06 BOEING: Amendment 39-5386. Applies to all Model 747 series airplanes listed in Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, certificated in any category. To detect loose or broken bolts used for the forward attachment of the Numbers 1, 2, 3, 6, 7, and 8 trailing edge flap tracks to the wing lower surface, accomplish the following, unless already accomplished: \n\n\tA.\tPrior to the accumulation of 5,000 flight cycles or within the next 300 flight cycles after the effective date of this AD, whichever occurs later, conduct a one-time inspection for loose or broken bolts in accordance with the Accomplishment Instructions of Boeing Alert Service Bulletin 747- 57A2234, dated February 21, 1986, or later FAA-approved revisions. NOTE: "Proof Torque" inspections previously accomplished in accordance with Boeing Service Letters Numbers 747-SL-57-44, dated December 18, 1985, or 747-SL-57-44A, dated December 20, 1985, constitute an equivalent means of compliance with the above requirement. \n\n\tB.\tIf one bolt is found loose or broken, replace all eight bolts used for the forward attachment of the trailing edge flap track to the wing lower surface within the next 600 flight cycles in accordance with Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA- approved revision. \n\n\tC.\tIf two bolts are found loose or broken, replace all eight bolts used for the forward attachment of the trailing edge flap track to the wing lower surface within the next 300 flight cycles in accordance with Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA- approved revision. \n\n\tD.\tIf three or more bolts are found loose or broken, replace all eight bolts used for the forward attachment of the trailing edge flap track to the wing lower surface prior to further flight in accordance with Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA- approved revision. \n\n\tE.\tAlternate meansof compliance for paragraphs B., or C., above, is:\n \n\t\t1.\tIf broken bolts are found, their replacement may be deferred for 1,200 flight cycles, provided the remaining non-broken bolts are retorqued to installation requirements in accordance with Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA-approved revisions.\n \n\t\t2.\tIf loose bolts are found, retorque all bolts to installation requirements in accordance with the Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA- approved revisions. \n\n\t\t3.\tPerform a visual inspection of the H-11 bolts for broken bolts (evidenced by missing heads) within 15 months after the inspections of paragraph A., of this AD, and repeat the inspections thereafter at intervals not to exceed 15 months. \n\n\t\t4.\tReplacement of the bolts with new bolts in accordance with Boeing Alert Service Bulletin 747-57A2234, dated February 21, 1986, or later FAA-approved revisions is terminating action for therepetitive inspections required by paragraph E.3. \n\n\tF.\tAn 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tG.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. \n\n\tAll persons affected by this directive who have not already received this service information from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124-2207. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective September 11, 1986.
2021-04-03: The FAA is adopting a new airworthiness directive (AD) for certain Pilatus Aircraft Ltd. (Pilatus) Model PC-24 airplanes. This AD was prompted by 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 identifies the unsafe condition as improperly manufactured cockpit and cabin evaporator filters installed during production on some PC-24 airplanes. The FAA is issuing this AD to address the unsafe condition on these products.
2009-02-03: The FAA is superseding an existing airworthiness directive (AD) for Lycoming Engines IO, (L)IO, TIO, (L)TIO, AEIO, AIO, IGO, IVO, and HIO series reciprocating engines, TCM TSIO-360-RB reciprocating engines, and Superior Air Parts, Inc. IO-360 series reciprocating engines with certain Precision Airmotive LLC RSA-5 and RSA-10 series fuel injection servos. That AD currently requires inspecting servo plugs for looseness and damage on fuel injection servos that have a servo plug gasket, part number (P/N) 365533, installed, and if loose, inspecting the servo regulator cover threads for damage, inspecting the gasket for damage, reinstalling acceptable parts, and torquing the servo plug to a new, higher torque to maintain the proper clamp-up force between the plug and cover. This AD requires the same inspections, except if the plug is found loose, servo plug gasket, P/N 365533, must be replaced with a new, improved gasket, P/N 2577258. This AD also requires replacement by December 31,2009, of servo plug gaskets, P/N 365533, manufactured and made available on or after August 22, 2006, as mandatory terminating action to the repetitive inspections required by this AD. This AD also prohibits the installation of any servo plug gasket, P/N 365533. This AD also clarifies the TCM engine model applicability, and adds Bendix RSA-5 and RSA-10 series fuel injection servos to the applicability. This AD results from Precision Airmotive LLC introducing the installation of a new improved servo plug gasket, P/N 2577258, to the affected Precision Airmotive LLC RSA-5 and RSA-10 series, and Bendix RSA-5 and RSA-10 series, fuel injection servos. We are issuing this AD to prevent a lean running engine, which could result in a substantial loss of engine power and subsequent loss of control of the airplane.
71-02-02: 71-02-02 PRATT & WHITNEY: Amdt. 39-1143 as amended by Amendment 39-1201. Applies to all Pratt & Whitney Aircraft JT4A series turbojet engines which incorporate the following serial number 12th stage compressor rotor disc part number 360112 1F-6145 3F-8849 4F-3831 4F-8300 5F-8122 1F-6150 3F-8850 4F-4569 4F-8301 5F-8123 1F-6151 through 4F-4570 5F-8342 2F-2758 3F-8855 4F-4642 4F-8458 5F-8346 2F-2759 4F-4643 5F-6445 5F-8405 2F-2764 3F-8856 4F-4645 5F-6446 5F-8510 2F-3781 3F-8920 4F-4648 5F-6447 5F-8512 3F-8921 5F-6449 2F-3785 3F-8923 4F-4747 5F-6452 5F-8522 2F-3788 3F-8924 4F-4748 5F-7606 5F-8529 2F-3926 3F-8926 4F-4749 5F-7748 5F-8589 2F-3927 3F-8927 4F-4750 5F-7749 through 3F-8928 4F-4752 5F-7835 2F-3933 3F-9114 4F-4968 5F-7837 2F-9674 3F-9115 4F-5029 5F-7838 2F-9675 3F-9116 through 5F-7839 2F-9676 4F-0454 4F-5035 5F-7840 2F-9677 4F-0461 4F-50595F-7867 3F-0551 4F-0465 through 5F-7872 through 4F-0467 4F-5063 5F-7875 3F-0558 4F-0471 4F-5087 5F-7985 4F-0473 4F-5089 5F-7987 3F-0559 4F-0475 4F-5090 5F-8031 3F-0560 4F-0477 4F-5091 5F-8032 3F-0563 4F-3606 5F-8035 4F-3690 4F-5092 4F-3692 through 5F-8039 4F-5099 5F-8042 4F-3699 5F-8119 4F-3700 4F-3828 4F-3830 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 5000 cycles or more in service within the next 25 cycles in service after the effective date of this AD replace the suspect serial number disc. b. With 5000 cycles or less in service prior to the accumulation of 5025 cycles 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 39-1143 was effective January 14, 1971 and was effective as to all recipients of the telegram dated 10 December 1970 which contained this airworthiness directive. This Amendment 39-1201 is effective May 4, 1971.
68-09-02: 68-09-02 LOCKHEED: Amdt. 39-593. Applies to Model 1329 airplanes, serial numbers 5001 through 5029 and 5031 through 5055. Compliance required as indicated. Lockheed Model 1329 and military C-140 series airplanes have experienced a number of stress corrosion cracks in the wing skin risers at the riser-skin intersection. To detect cracking of risers on these airplanes, accomplish the following: (a) Within 15 days after the effective date of this AD, inspect all aircraft in accordance with the instructions contained in Lockheed-Georgia Company Service Bulletin No. 329-257, dated April 11, 1968, and Service Bulletin Supplement No. 329-257A, dated April 26, 1968, or later FAA approved revision, or in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. NOTE. - Aircraft which have been inspected by representatives of Lockheed-Georgia Company are considered to comply with paragraph (a) above. (b) If cracks are detectedin the riser-skin intersections, reinspect at intervals not to exceed 180 days in accordance with the instructions contained in Lockheed-Georgia Company Service Bulletin No. 329-257, dated April 11, 1968, and Service Bulletin Supplement No. 329- 257A dated April 26, 1968, or later FAA approved revision, or in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. (c) If, as a result of the initial inspection or any subsequent inspections, it is determined that more frequent inspections are required, individual airplane owners will be so notified by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. (d) If no cracks are found, or if cracks are repaired in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region, the repetitive inspections required by paragraphs (b) and (c), above, may be discontinued. (e) Repeat the inspection required by paragraph (a), above, at each annual inspection, or each twelve months. These inspections need only consist of Phase I and Phase II inspections as explained in Lockheed-Georgia Company Service Bulletin No. 329-257, dated April 11, 1968, and Service Bulletin Supplement No. 329-257A dated April 26, 1968, or later FAA approved revision, or in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. This amendment becomes effective May 7, 1968. Revised Nov. 19, 1968.
69-23-03: 69-23-03 PIPER: Amendment 39-871. Applies to Piper PA-18-150 type airplanes, S/N 18-8809 through 18-8851, except S/N 18-8844 and 18-8849, certificated in all categories. To determine if the lower longerons are made from 1025 steel or the approved 4130 steel, accomplish before further flight the following inspection: (a) Remove the rear seat, look downward through the no sag springs and locate the lower fuselage longerons on the left and right sides of the aircraft. (b) From a point adjacent to the aileron cable pulley measure aft 10 inches on the top of the longeron and perform the following: (1) Using longitudinal strokes sand the top surface of the right and left lower longerons with No. 80 emery cloth, for an area approximately 3 inches in length and 1/2 inch wide. NOTE: No chemical reaction will occur unless this area is thoroughly cleaned down to the bare metal. Use only longitudinal strokes to remove the paint, primer, metalizing or corrosion from this area. (2) Wipe specified area with a clean cloth. (3) Apply one drop of concentrated nitric acid on the sanded area. (4) If the longeron is 1025 steel, the acid drop will boil and turn a yellowish orange or brown. (5) Wipe off the acid spot and neutralize the area with baking soda paste. (6) Wipe surface and wash with water. (7) Prime and paint the test area. (8) If the test indicates (as described above) the longeron is 1025 steel, the aircraft is considered unairworthy and must be grounded until the aircraft is altered in accordance with an alteration approved by the Chief, Engineering and Manufacturing Branch, Eastern Region. (9) The aircraft is considered airworthy, relative to the longeron being 4130 if there is no reaction to the aforementioned acid test, and can be returned to service. Piper Service Bulletin No. 302, dated October 23, 1969 covers this subject. This amendment is effective November 15, 1969 and was effective upon receipt for all recipients of the telegram dated October 24, 1969 which contained this amendment.
69-12-01: 69-12-01 PIPER: Amdt. 39-773 applies to PA-28R-180 series airplanes, Serial Numbers 28-30004 through 28-31092 and to PA-28R-200 series airplanes, Serial Numbers 28-35000 through 28-35237, 28-35239 through 28-35251 and 38-35253 through 28-35256. Compliance required as indicated. To prevent the possibility of restricting air flow through the engine air induction system, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this airworthiness directive, unless already accomplished, remove air induction inlet hose Piper Part Number 63633-43 and inspect for deterioration, broken or frayed cord wrapping, evidence of hose partially collapsed, indication of reinforcement wire slippage or evidence of wire not being bonded properly to the hose interior wall, especially in the vicinity of the Koroseal ties or clamps. (b) If the hose shows any deterioration, as noted in (a), remove it from service. Install an airworthy hose, Piper P/N 63633-43, or Aero-Duct SCAT 3-1/4 O.D. hose, or MIL-H- 8796B Type 1 Class 3 hose, 22-1/2 long, making sure that the wire reinforcing coils extend over the pilot tubes at each end and under the clamps. Piper Kit 760 328 may be installed in lieu of duct replacement. (See Item (d) below). (c) Repeat the inspections of item (a) above at intervals not to exceed 10 hours' time in service. These inspections need not be continued after installation of Piper Kit 760 328. (d) Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished, install a new Piper air induction hose installation Kit No. 760 328 in accordance with the instructions included with the kit. (e) The inspection or kit installation time intervals (other than for the initial inspection) may be adjusted by an FAA inspector up to a maximum of 5 hours to coincide with aircraft annual or 100-hour scheduled inspections. Other means of compliance with the requirements of this directive may be utilized, if approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. (Piper Service Bulletin No. 297 covers this same subject.) This amendment becomes effective 10 June 1969.
2009-01-02: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 737-600, -700, -700C, -800 and -900 series airplanes. This AD requires a one-time general visual inspection of frames between body station (BS) 360 and BS 907 to determine if certain support brackets of the air conditioning (A/C) outlet extrusions are installed; medium- and high-frequency eddy current inspections for cracking of the frames around the attachment holes of the subject brackets; and repair if necessary. This AD also requires installing new, improved fittings for all support brackets of the A/C outlet extrusions between BS 360 and BS 907. This AD results from numerous reports of multiple cracks in the frames around the attachment holes of certain support brackets of the A/C outlet extrusions. We are issuing this AD to prevent frame cracking, which, if not corrected, could lead to a severed frame that, combined with cracking of the skin lap splice above stringer 10, could result in rapiddecompression of the airplane.
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.