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71-09-02: 71-09-02\tBOEING: Amendment 39-1197 as amended by Amendment 39-1225 and 39-1254 is further amended by Amendment 39-1276. Applies to Model 707/720 series airplanes equipped with 7079/T6 rudder hydraulic power actuator support fittings. \n\tCompliance required as indicated. \n\tTo detect cracks which might result in failure of the rudder hydraulic power actuator support fitting and to prevent additional cracking of the fitting in the vicinity of the actuator attachment holes, accomplish the following: \n\t(a)\tFor airplanes previously reworked in accordance with paragraph (b) of AD 69-13-02, as amended by Amendment 39-1174 effective March 18, 1971, within the next 100 hours' time in service after the effective date of this AD unless already accomplished in accordance with paragraph (d)(1) of that AD, perform either an ultrasonic inspection, or, after removal of bushings, an eddy current inspection to detect evidence of cracks in the support fitting. \n\t(b)\tUnless already accomplished withinthe last 300 hours' time in service prior to the effective date of this AD, within the next 100 hours' time in service after the effective date of this AD, perform either another ultrasonic inspection or, with bushings removed, an eddy current inspection of all fittings previously inspected by ultrasonic means. \n\t(c)\tAt intervals not to exceed 400 hours' time in service after the last ultrasonic inspection, reinspect by ultrasonic means all fittings previously inspected in that manner in compliance with (a) and (b), above, until an eddy current inspection, with bushings removed, is performed per (d), below. \n\t(d)\tWithin 1200 hours' time in service after the effective date of this AD but no later than 1200 hours' time in service after the last eddy current inspection with bushings removed, remove all bushings and perform an eddy current inspection of the fitting. \n\t(e)\tAfter accomplishment of the eddy current inspection per (b) or (d), above, and until affected fittings are replacedor modified per (f) or (i), below, inspect such fittings by ultrasonic and/or eddy current as follows: \n\t\t(1)\tInspect by ultrasonic means at intervals not to exceed 650 hours' time in service until the next such inspection after the effective date of this amendment. Thereafter, inspect either by ultrasonic means at intervals not to exceed 325 hours' time in service or by eddy current, with bushings removed, at intervals not to exceed 650 hours' time in service. \n\t\t(2)\tInspect by eddy current, with bushings removed, at intervals not to exceed 1200 hours' time in service until the next such inspection after the effective date of this amendment. Thereafter, inspect either by ultrasonic means at intervals not to exceed 325 hours' time in service or by eddy current, with bushings removed, at intervals not to exceed 650 hours' time in service. \n\t\t(3)\tAfter the next 100 hours' time in service following the effective date of this amendment, perform all ultrasonic and eddy current inspections with the equipment and procedures outlined in Boeing Service Bulletin No. 2903, Revision 6, dated June 4, 1971, or later FAA approved revision, or in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(f)\tWhen any fitting inspected in accordance with the foregoing paragraphs or paragraph (g), below, exhibits evidence of a crack which cannot be reworked within the hole oversize limits outlined in Boeing Service Bulletin 2903, Revision 6, dated June 4, 1971, or later FAA approved revision, either: replace the fitting prior to further flight with a new fitting made of 7075-T73 material; modify the fitting and install a steel replacement lug assembly in accordance with FAA- approved Boeing Service Bulletin 3042; or accomplish another replacement or modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(g)\tWhen any fitting inspected in accordance with paragraphs (a) through (e), above, or in accordance with this paragraph, exhibits evidence of a crack which can be reworked within the hole oversize limits outlined in Boeing Service Bulletin 2903, Revision 6, dated June 4, 1971, or later FAA-approved revision, the fitting may be returned to service, provided: \n\t\t(1)\tThe fitting is reworked and new bushings are fabricated in accordance with Part II of Boeing Service Bulletin 2903, dated June 2, 1969, or later FAA approved revision; \n\t\t(2)\tThe new bushings are installed in the fitting in accordance with (h), below; and \n\t\t(3)\tThe fitting is inspected thereafter by ultrasonic means or, with bushings removed, by eddy current at intervals not to exceed 325 hours' time in service. After the next 100 hours' time in service following the effective date of this amendment, perform all such inspections in accordance with (e)(3), above. The intervals within which the eddy current inspections must be performed may then be increased to 650 hours' time in service. \n\t(h)\tFittings inspected or reinspected by eddycurrent technique to comply with paragraphs (a) through (e) and (g), above, and fittings eligible for rework in accordance with (g), above, may be returned to service when the bushings are installed or reinstalled in the manner outlined in Boeing Service Bulletin 2903, Revision 5, dated February 3, 1971, or later FAA- approved revision. \n\t(i)\tBefore further flight after January 1, 1972, either: \n\t\t(1)\tReplace all 7079-T6 fittings with fittings made of 7075-T73 material; or \n\t\t(2)\tModify the 7079-T6 fitting and install a steel replacement lug assembly in accordance with FAA-approved Boeing Service Bulletin 3042; or \n\t\t(3)\tAccomplish another replacement or modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\tThe special inspections prescribed by this AD on any airplane are terminated when the fitting is replaced or modified in accordance with this paragraph. \n\t(j)\tWhen a fitting is found to exhibit evidence of a crack, the airplane may not be ferried. \n\t(k)\tAfter the effective date of this AD, actuator support fittings not previously reworked by the installation of aluminum-nickel-bronze bushings in accordance with paragraph (b) of AD- 69-13-02, effective June 6, 1969, must be inspected, reworked, or replaced as follows: \n\t\t(1)\tBefore further flight remove all bushings and perform an eddy current inspection of the support fitting. \n\t\t(2)\tBefore further flight, replace any fitting found to be cracked beyond rework limits, in accordance with (f), above. \n\t\t(3)\tAny fitting found cracked within rework limits may be returned to service if reworked in accordance with (g), above, and the new bushings are installed in accordance with (h), above. \n\t\t(4)\tBefore further flight, fittings inspected in accordance with (k)(1), above, and found to be uncracked must be modified to incorporate flanged aluminum-nickel-bronze bushings described by Paragraph C of Boeing Service Bulletin 2903, dated June 2, 1969, and by Part II - Bushing Replacement of that Service Bulletin, or in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region, unless the fitting is replaced or modified in accordance with (f) above. Installation of flanged bushings must be performed in accordance with (h), above. \n\t\t(5)\tAll fittings modified per (k)(4) to incorporate flanged aluminum-nickel-bronze bushing must be reinspected in the manner and within the corresponding intervals specified in (e), above, until replaced in accordance with (k)(6). \n\t\t(6)\tAll 7079-T6 fittings must be replaced before further flight after January 1, 1972, in accordance with (i), above. \n\t(l)\tFollowing each actual or simulated #3 or #4 engine power failure, or flight with #3 or #4 engine shutdown, or prior to ferry flight with #3 or #4 engine inoperative, perform either an ultrasonic inspection or, with bushings removed, an eddy current inspection before further flight to detect any evidence of a crack in the rudder actuator support fitting. Any fitting exhibiting evidence of a crack must be replaced per (f) above, or reworked per (g) above, before further flight. \n\tAD 71-09-02 Amendment 39-1197 supersedes amendment 39-786 (34 F.R. 9748), AD 69- 13-02, as amended by Amendment 39-800, (34 F.R. 12214), and Amendment 39-1174, (36 F.R. 5209). \n\tAmendment 39-1197 became effective April 27, 1971. \n\tAmendment 39-1225 became effective June 8, 1971 for all persons except those to whom it was made effective immediately by telegram dated May 14, 1971. \n\tAmendment 39-1254 became effective August 3, 1971. \n\tThis Amendment 39-1276 becomes effective September 2, 1971.
2016-05-08: We are superseding airworthiness directive (AD) 2006-23-17 for certain Turbomeca S.A. Turmo IV A and IV C turboshaft engines. AD 2006- 23-17 required repetitive inspections of the centrifugal compressor intake wheel (inducer) blades for cracks and corrosion, replacement of parts that fail inspection, and replacement of the TU 197 standard centrifugal compressor. This AD requires the same inspections, but at revised intervals, adds the replacement of the TU 215 standard centrifugal compressor, and requires replacement of parts that fail inspection. This AD was prompted by a centrifugal compressor inducer blade loss. This AD was also prompted by a Turbomeca S.A. review of the engine service experience and their determination that more frequent borescope inspections (BSIs) are required on engines not modified to the TU 191, TU 197, or TU 224 standard. We are issuing this AD to prevent failure of the centrifugal compressor inducer, which could lead to an uncontained blade release,damage to the engine, and damage to the airplane.
73-17-02: 73-17-02 SLINGSBY: Amdt. 39-1702. Applies to all Model T.51 Dart Gliders which have metal reinforced wing spars. Compliance is required as indicated. To prevent possible loss of wing structural integrity due to corrosion of the wing spars, accomplish the following: Before July 14, 1973, unless already accomplished within the last year, and thereafter at intervals not to exceed one year since the last inspection - (a) Cut seven inspection holes in the lower surface skin of each wing in accordance with Slingsby Technical Instruction No. 58, Issue 1, dated May, 1973, or an FAA-approved equivalent; (b) Visually inspect the metal portion of the spars for corrosion; NOTE: During the inspection required by paragraph (b), particular attention should be directed to the bolted joints and rib attachment areas. (c) If corroded areas are found during an inspection required by paragraph (b), measure the depth of the corrosion in the affected areas; (d) If corrosion is found which exceeds a depth of 0.007 of an inch, before further flight, repair the corroded areas, or replace the corroded parts with serviceable identical parts or FAA-approved equivalent; and (e) Close the inspection holes cut in accordance with paragraph (a). (f) Notification in writing must be sent to the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region, American Embassy, APO New York, N.Y. 09667, stating the results, positive or negative, of each inspection required by this AD, within 10 days after such inspections. (Reporting approved by the Bureau of the Budget under BOB No. 04-R0174). This amendment is effective upon publication in the Federal Register as to all persons except those persons to whom it was made immediately effective upon receipt of the airmail letter dated June 27, 1973, which contained this amendment.
2024-24-08: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Canada Limited Partnership Model BD-500-1A10 and BD-500- 1A11 airplanes. This AD was prompted by multiple occurrences of pilot and co-pilot seats locking in a fore-aft position due to the seat fore- aft adjustment mechanism disconnecting, caused by a broken cotter pin in the seat base egress linkage. This AD requires modifying the pilot and co-pilot seats by replacing the hardware of the seat base egress linkage, as specified in a Transport Canada AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2024-24-05: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model 767-300F series airplanes. This AD was prompted by a determination that certain cargo compartment insulation blankets do not adequately fit some locations and allow smoke to migrate past the cargo compartment sidewall liners and upward into the main cabin. This AD requires replacing cargo compartment insulation blankets. The FAA is issuing this AD to address the unsafe condition on these products.
2016-05-07: We are adopting a new airworthiness directive (AD) for certain Engine Alliance (EA) GP7270 turbofan engines. This AD was prompted by reports of the installation of non-conforming honeycomb cartridges in the high-pressure compressor (HPC) adjacent to the HPC rotor stage 2 to 5 spool and stage 7 to 9 spool. This AD requires removal and replacement of the affected HPC rotor stage 2 to 5 and stage 7 to 9 spools and adjacent honeycomb cartridges. We are issuing this AD to prevent failure of the HPC rotor stage 2 to 5 and stage 7 to 9 spools, which could lead to uncontained engine failure and damage to the airplane.
58-10-03: 58-10-03 UNIVERSAL: Applies to All Models GC-1A and GC-1B Aircraft With Adel Precision Products Corporation Landing Gears With Forged Aluminum Torque Knees. (This AD is issued to clarify, add a 100-hour periodic inspection, and supersede AD 57- 13-07.) Compliance required as indicated. Failures have been reported of the stop ring brazed to the inner piston strut. Failures resulted in the piston sliding out of the strut and the torque knees assuming a straight position. This overextension of the strut precludes gear retraction into the wheel well. 1. Within the next 100 hours of operation, unless already accomplished, a suitable external stop or some other equivalent means should be provided which will function as a safety measure in case of failure of the internal stop ring. At the time of installation of the external stop, or within the next 100 hours of operation if not already accomplished, the Adel strut assembly stop ring is to be examined for condition and rejected if the stop ring is loose or shows signs of separation at the braze. After reassembly, the external stop must be installed so that there exists a clearance of 1/32 to 1/8 inch between the face of the stop and side of the strut cylinder with the gear fully extended. In the event there is insufficient clearance, the external stop must be reworked until the proper clearance is obtained. 2. At each subsequent 100 hours of operation, the clearance is to be rechecked with the strut fully extended. If there is no clearance between the external safety stop and the strut cylinder, it will be necessary to disassemble the strut and examine the internal stop ring for indications of failure. If failure of the stop ring is apparent, the inner cylinder assembly, Adel P/N 16084, must be replaced or suitably reworked. (Universal Aircraft Industries Customer Service Maintenance Bulletin No. 34 with Revision No. 1 covers this same subject.) This supersedes AD 57-13-07.
53-01-01: 53-01-01 BELL: Applies to Model 47D1 Helicopters, Serial Numbers 477 to 625, Inclusive. Compliance required at next 300-hour teardown inspection but not later than February 28, 1953. In order to improve the method of mounting the tail rotor gearbox assembly and to avoid the possibility of distorting the S10R bearing when tightening the existing clamp, install sleeve P/N 47-640-058-1 on assembly 47-640-044-3. Clamping ring P/N 47-644-197-1 and clamp P/N 47-640-046-1 are replaced by the riveted sleeve. (Mandatory Service Bulletin No. 90, Revision B, dated December 8, 1952, covers this same subject. Service Bulletin No. 90, Revision A, also covers the same subject but Revision B simplifies the installation and completes the parts called out.)
70-14-03: 70-14-03\tBOEING: Amdt. 39-1024 as amended by Amendment 39-1093 and 39-1123 is further amended by Amendment 39-1144. Applies to all Model 747 airplanes certificated in all categories. \n\tCompliance required within the next 10 landings after the effective date of this AD, unless already accomplished within the last 90 landings and thereafter at intervals not to exceed 100 landings from the last inspections. \n\tTo detect cracks of the flap track, accomplish the following: \n\t(a)\tWithin 10 landings after the effective date of this AD, unless already accomplished within the last 90 landings, visually inspect all flap tracks for cracks in accordance with Boeing Alert Service Bulletin 57-2011, Revision 1, dated June 6, 1970, or later FAA approved revisions, or in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t\t(1)\tIf a crack is found, replace flap track or repair in accordance with a method approved by the Chief, Aircraft Engineering Division, FAA Western Region, prior to further flight. Repeat inspections, per (a) (2) below. \n\t\t(2)\tIf no cracks are found, repeat the inspection for cracks at intervals not to exceed 100 landings. \n\t(b)\tWithin 10 hours after the effective date of this AD, unless already accomplished, install a placard as noted below in full view of the captain and first officer, or provide an equivalent procedure acceptable to the cognizant Air Carrier District Office. \n\tPlacard wording is as follows: \n\n\t\t\t\t Recommended \n\t\t Flap\t\t Flap Operating Speed \n\t\tPosition\t\t(Knots IAS)\t \n\t\t 25\t\t 170 \n\t\t 30\t\t 140 \n\n\t(c)\tReport in the airplane log all instances when the flap speeds in (b) above are exceeded. \n\t(d)\tUpon accomplishment of (c), visually inspect all inboard flap tracks for cracks per (a), prior to further flight. \n\t\t(1)\tIf a crack is found, replace or repair the flap track per (a)(1). Repeat inspections per (a)(2). \n\t\t(2)\tIf nocracks are found, repeat inspections per (a)(2). \n\t(e)\tInstall placard advising pilot of reporting requirements specified in paragraph (c). \n\t(f)\tThe existing flap tracks may be replaced with redesigned flap tracks in accordance with Boeing Service Bulletin 57-2100, Revision 4, dated November 25, 1970, or later FAA approved revision or in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. Upon completion of this modification, the inspection and placard installation requirements of paragraphs (a) through (e) are no longer applicable. \n\tAmendment 39-1024 effective July 21, 1970. \n\tAmendment 39-1093 effective October 22, 1970. \n\tAmendment 39-1123 effective December 11, 1970. \n\tThis Amendment 39-1144 becomes effective January 19, 1971.
2016-06-02: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 737-300, -400, and -500 series airplanes. This AD requires repetitive inspections for cracking in the horizontal and vertical flanges of the rear spar upper chord of the horizontal stabilizer, and related investigative and corrective actions if necessary. This AD was prompted by a report of cracking in the center section of the horizontal stabilizer. We are issuing this AD to detect and correct cracking of the rear spar center section of the horizontal stabilizer that could lead to departure of the horizontal stabilizer from the airplane.