Results
2003-04-15: This amendment adopts a new airworthiness directive (AD) for the specified Sikorsky Aircraft Corporation (Sikorsky) model helicopters. This action requires determining the manufacturer of a certain part-numbered rotor brake disc (RBD) and if the manufacturer is Parker Hannifin Corporation (PHC), re-identifying the RBD as appropriate. This action also requires before the first flight of the next day following any day in which a certain RBD was used, visually inspecting the RBD for a crack. If a crack is found, this AD also requires replacing the RBD with an airworthy RBD or deactivating it as applicable depending on the nature of the crack. This amendment is prompted by the discovery that certain RBDs manufactured by PHC were improperly heat treated resulting in "soft" RBDs that have an increased wear rate compared to those heat treated in accordance with the type design requirement. Further investigation reveals that "soft" RBDs develop cracks more frequently than previously manufactured RBDs. The actions specified in this AD are intended to prevent failure of the RBD, damage to the rotor blades and nearby hydraulic and fuel lines, and subsequent loss of control of the helicopter.
48-31-01: 48-31-01 GRUMMAN: Applies to G-44 and G-44A Aircraft. To be accomplished by September 15, 1948. Inspect upper terminal (P/N 17257-1) of stabilizer strut (P/N 17256) for cracks extending radially from the outside edge of the ears to the inside of the hole in which the shoulder bushings are pressed. Cracked terminals should be replaced with steel terminals. All terminals without cracks may be left in service if inspected every 100 hours. (Grumman Aircraft Engineering Corp. Service Bulletin No. 22 dated July 1, 1948, covers this same subject.)
2008-26-12: 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: This Airworthiness Directive (AD) is prompted by the discovery on L 23 SUPER-BLANIK sailplanes of cracks in zones where the front and aft control levers attach the connecting rod designated as "control bridge'' on the relevant Illustrated Parts Catalogues (IPC). If left uncorrected cracks could propagate and lead to the breakage of the connecting rod with subsequent loss of control of the sailplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
2016-05-12: We are superseding Airworthiness Directive (AD) 2012-15-13, for certain The Boeing Company Model 747-100B SUD, 747-300, 747-400, and 747-400D series airplanes; and Model 747-200B series airplanes having a stretched upper deck. AD 2012-15-13 required inspections for cracking and discrepancies of certain fasteners; modification of the frame-to-tension-tie joints; repetitive post-modification inspections; related investigative and corrective actions if necessary; and repetitive inspections for cracking in the tension tie channels, and repair if necessary. For certain airplanes, AD 2012-15-13 also required an inspection to determine if the angle is installed correctly, and re- installation if necessary; and an inspection at the fastener locations where the tension tie previously attached to the frame prior to certain modifications, and repair if necessary. This new AD adds a new inspection for cracking in the tension tie channels and post- modification inspections of the modifiedtension ties for cracking, and repair if necessary. This AD was prompted by an evaluation indicated that the upper deck is subject to widespread fatigue damage (WFD). We are issuing this AD to prevent fatigue cracking of the tension ties, shear webs, and frames of the upper deck, which could result in rapid decompression and reduced structural integrity of the airplane.
53-21-01: 53-21-01 de HAVILLAND: Applies to All Model DHC-2 (Beaver) Aircraft. Compliance required as indicated. Several cases have been reported where mechanics in the field upon assembling DHC-2 wings to fuselage have installed extra washers, packing, etc., to the rear spar wing bolt in order to take out any end play. It should be pointed out that a clearance is purposely provided in this fitting as the rear wing attachment is not designed to take drag loads. All Beaver aircraft should be inspected as soon as possible but not later than December 1, 1953, to see that no washers, bushings, etc., have been installed in this fitting and if found they should be removed immediately. In assembling the wing to fuselage, the front spar must be attached and then the rear wing bolt should be installed. It is quite normal that the rear spar wing fitting should not touch either inboard side of the fuselage fitting, but in most cases the wing fitting is almost against the forward side. The gaps in the fittings front and rear should not be packed with washers or spacers. The FAA concurs in this mandatory action by the Canadian Department of Transport. (de Havilland Technical News Sheet, Series B, No. 67, dated August 31, 1953, available from de Havilland Aircraft of Canada, Ltd., Toronto, Ontario, Canada, covers this same subject.)
68-08-01: 68-08-01 MCCAULEY AIRCRAFT PROPELLERS: Amdt. 39-581 as amended by Amendment 39-1314 is further amended by Amendment 39-1377. Applies to the following two- and three-bladed constant speed propeller models with hub serial numbers indicated below: PROPELLER MODELS 2D34C8 C2A36C32 D2A34C58-B 3A32C76-S D2A34C78-K 2D34C8-A C2A36C32-A D2A34C58-J 3A32C76-T D2A34C78-L 2D34C8-J C2A36C32-D D2A34C58-K 3A32C76-AD D2A34C78-M 2D34C8-K D2A36C33 D2A34C58-L 3A32C76-AS D3A32C79 2D34C8-M D2A36C33-D D2A34C58-M 3A32C76-AT D3A32C79-A 2A36C23-C D2A36C45 2A34C66 3A32C76-FD D3A32C79-B 2A36C23-CD D2A36C45-D 2A34C66-A 3A32C76-FS D3A32C79-F 2A36C23-CH D2A34C49 2A34C66-B 3A32C76-FT D3A32C79-J 2A36C23-CJ D2A34C49-A 2A34C66-C 3A32C76-JD D3A32C79-K 2A36C23-CP D2A34C49-B 2A34C66-J 3A32C76-JS 2A36C82-T 2A36C23-CS D2A34C49-J 2A34C66-K 3A32C76-JT 2A36C82-DT 2A36C23-DD D2A34C49-K 2A34C66-L 3A32C76-KD D3A32C88 2A36C23-DH D2A34C49-L 2A34C66-M 3A32C76-KS D3A32C88-A 2A36C23-DJ D2A34C49-M E2A34C70 3A32C76-KT D3A32C88-F 2A36C23-DP 2A34C50 E2A34C70-A D3A32C77 D3A32C88-J 2A36C29 2A34C50-A E2A34C70-J D3A32C77-A D3A32C88-K 2A36C29-A 2A34C50-B E2A34C70-K D3A32C77-F D3A32C90 2A36C29-D 2A34C50-J E2A34C70-M D2A32C77-J D3A32C90-A B2A36C31 2A34C50-K E2A34C73 D3A32C77-K D3A32C90-B B2A36C31-A 2A34C50-L E2A34C73-A D2A34C78 D3A32C90-C B2A36C31-D 2A34C50-M E2A34C73-J D2A34C78-A D3A32C90-F D2A36C31-A D2A34C58 E2A34C73-K D2A34C78-B D3A32C90-J D2A36C31-D D2A34C58-A E2A34C73-M D2A34C78-J D3A32C90-K 3A32C76-D HUB SERIAL NUMBERS 59000 up to and including 712778 except 700492, 700500 thru 700558; 700561 thru 700568; 700570 thru 700594; 700596 thru 701050 and 701053 Compliance required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. To prevent failure of the propeller cylinder attach screws, accomplish the following: Modify propeller cylinder attachment in accordance with McCauley Service Bulletin No. 92, dated April 21, 1971, or later FAA-approved revision. However, for propellers used on Bellanca Aircraft Models 17-30 and 17-30A modify propeller cylinder attachment in accordance with McCauley Service Bulletin No. 94, dated July 28, 1971, or later FAA-approved revision instead of Service Bulletin No. 92. Equivalent methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. Amendment 39-581 was effective April 11, 1968. Amendment 39-1314 was effective October 14, 1971. This Amendment 39-1377 is effective January 21, 1972.
2024-24-03: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model MD-11 and MD-11F airplanes. This AD was prompted by a report of a Model MD-11F airplane experiencing an uncommanded deployment of a thrust reverser in flight at low altitude. This AD requires initial and repetitive detailed inspections and repetitive wire integrity tests of the engine pylon thrust reverser control system wire harnesses, junction box assembly and junction box cover, left-side and right-side thrust reverser electrical harnesses, core (engine compartment) miscellaneous wire harness assembly, and 30- degree bulkhead wire harness assembly; and applicable on-condition actions. This AD also requires reporting inspection results. The FAA is issuing this AD to address the unsafe condition on these products.
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.