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2024-23-10:
The FAA is adopting a new airworthiness directive (AD) for certain ATR--GIE Avions de Transport R[eacute]gional Model ATR42 and ATR72 airplanes. This AD was prompted by a report that for airplanes converted from passenger to cargo configuration using certain supplemental type certificates, no height limitation for the cargo, when loaded in the cargo compartment, is defined, and that as a consequence, cargo might be loaded up to the ceiling of the cargo compartment. This AD requires modification of the cargo compartment and implementation of updated cargo loading procedures. The FAA is issuing this AD to address the unsafe condition on these products.
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2016-06-03:
We are adopting a new airworthiness directive (AD) for all Airbus Model A319-131, -132, and -133 airplanes; Model A320-232 and - 233 airplanes; and Model A321-131, -231, and -232 airplanes. This AD was prompted by reports of forward engine mount attachment pins that were manufactured from discrepant raw material. This AD requires identification and replacement of affected forward engine mount attachment pins. We are issuing this AD to prevent failure of a forward engine mount attachment pin, possible loss of an engine in-flight, and consequent reduced controllability of the airplane.
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53-26-01:
53-26-01 BRIGGS AND STRATTON: Applies to All Airplanes Equipped With AAF Type B-5 Ignition Switch Manufactured by Briggs and Stratton. Affected Airplanes Include Beech Models D18S (Serial Numbers A-1 Through A-537), D18C, D18C-T, C18S and AT-11 Having the AAF Type B-5 Briggs and Stratton Switch Installed.
Compliance required not later than April 1, 1954.
Design of the Briggs and Stratton AAF Type B-5 switch will permit foreign objects to enter the master ignition switch portion of the switch assembly. As a result, the magnetos of both engines connected to the switch may accidentally become grounded. Gaps in the enclosure of the Briggs and Stratton master switch can allow entry of foreign objects; whereas, AAF Type B-5 switches produced by other manufacturers are tightly sealed. These latter switches are not considered hazardous.
Briggs and Stratton AAF Type B-5 switches are identified by:
(1) The letters "AAF TYPE B-5" on the face of the switch and
(2)The words "Briggs & Stratton Corp., Milwaukee, Wis., U.S.A." stamped on the master switch enclosure. Determination of whether the switch carries the designation of (2) will probably necessitate examination with a flashlight and mirror or removal of the switch from its mount in the airplane.
If AAF Type B-5 switches manufactured by Briggs and Stratton are installed, accomplish either of the following:
(a) Replace the Briggs and Stratton AAF Type B-5 switch with a Type B-5 switch having the master ignition switch portion adequately sealed against entry of foreign objects.
(b) Remove the master ignition and battery switch portion of the Briggs and Stratton AAF Type B-5 in the following manner:
1. Remove the switch assembly from its mount on the airplane.
2. Drill out the three rivets attaching the master switch portion to the face plate of the ignition switch assembly.
3. The six electrical wires connecting the master switch portion to the threaded terminals are to be disconnected at the threaded terminals and discarded with the master switch portion.
4. Clean the ignition switch brass ground strip to make a good electrical contact and rivet it to the case with AN rivet.
5. Reinstall the modified Type B-5 ignition switch assembly in the airplane. Minor rework of the airplane electrical system may be required if the electrical master switch was connected through the Type B-5 master ignition and battery switch.
NOTE: Proper precautions should be observed when the ignition switch is removed or disconnected since the engine magnetos are not grounded.
(Beech Service Bulletin; Model D18S, D-18C, D-18C-T, C18S, AT-11; No. 64, issued November 10, 1953, covers this same subject.)
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2003-09-11:
This amendment adopts a new airworthiness directive (AD) that applies to certain Pilatus Aircraft Ltd. (Pilatus) Models PC-12 and PC-12/45 airplanes. This AD requires you to inspect the pedestal leg assembly on aft facing passenger seats for correct configuration. If incorrectly configured, this AD requires you to modify to the correct configuration. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Switzerland. The actions specified by this AD are intended to detect and correct pedestal leg assemblies on aft facing passenger seats that are in nonconformance with manufacturing standards. Nonconforming passenger seats could result in passenger injury in an emergency situation.
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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.
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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.)
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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.
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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.
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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.)
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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.
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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.
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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.
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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.
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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.
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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.
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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.
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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.
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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.
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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.)
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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.
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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.
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51-13-01:
51-13-01 MARTIN: Applies to All Models 202 and 202A Series Aircraft as Noted.
Compliance required as indicated.
1. Applies to all Serial Numbers except 9125 through 9127 and 9129 through 9131.
A. Compliance required every other No. 3 inspection period, approximately 620 hours flight time, or 6 months, whichever occurs first.
(1) Inspect the wing top cover splice angles with 2- to 6-power glass (may be conducted with paint on). On Model 202 Series aircraft, except Serial Numbers 9149 and 9150, these angles are P/N U23435 through U23442 (wedge assemblies Nos. 2021A12633 and 2021A12624). On Model 202 aircraft, Serial Numbers 9149 and 9150, and on all Model 202A Series aircraft, these angles are P/N A12025, A12026, U29403, U29404, U29407, and U29408 (wedge assemblies Nos. 2021A12023 and 2021A12024).
(2) If cracks are found, the top cover wedge assembly must be replaced as outlined in item B.
(3) Continue the inspections outlined in item A. until the replacementaction outlined in item B. is accomplished.
B. Compliance required not later than January 1, 1953.
(1) Replace all wing top cover wedge angles with new angles as follows: On Model 202 Series aircraft, except Serial Numbers 9149 and 9150, the new angles are P/N 202A3000068-1 and -2 through 202A3000071-1 and -2 (wedge assemblies Nos. 2021C12336-9 and -10). On Model 202A aircraft, Serial Numbers 9149 and 9150, and on all Model 202A Series aircraft, the new angles are P/N 202A3000072-1 and -2 through 202A3000074-1 and -2 (wedge assemblies Nos. 2021C12090-9 and -10).
(2) At the time of wedge assembly replacement, the outer wings must be reinstalled with the engines removed, and the mating faces shimmed to a maximum permissible gap of 0.020 inch. The rear spar fitting must also be shimmed to a zero gap for approximately the upper half of its area.
(3) Each airplane must have a torque check on the attach angle bolts immediately after the first flight after any outer wing installation. If the torque is within 15 percent of the installation torque, it is satisfactory.
(4) After the new wedge angles are incorporated, the inspection outlined in item A. is no longer applicable to that airplane.
2. Applies to Serial Numbers 9125 through 9127 and 9129 through 9131.
A. Compliance required every 170 hours flight time. Conduct the inspection outlined in above item 1.A. on angles P/N U23435 through U23442 (wedge assemblies Nos. 2021A12623 and 2021A12624) until cracks are found and/or the replacement action outlined in item B. below is accomplished.
B. Compliance required not later than May 15, 1952. Replace all wing top cover angles with new angles P/N 202A3000068-1 and -2 through 202A3000071-1 and -2, wedge assemblies Nos. 2121C12336-9 and -10, and when accomplished, the inspections outlined in above item A. are no longer applicable to that airplane.
3. Applies to all Model 202 and 202A Series.
A. Compliance required every 12,000 flighthours. Replace all wing top cover attach angle bolts and spar web splice bolt.
(Martin 202/202A Service Bulletins Nos. 180, 184, and 187 cover this same subject.)
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2001-13-03 R1:
This amendment revises an existing airworthiness directive (AD) for Kaman Aerospace Corporation (Kaman) Model K-1200 helicopters that currently requires reducing the life limit of the rotor shaft and teeter pin assembly and establishing a life limit for the flap clevis. This amendment retains those requirements but removes a flap clevis part number from the applicability and, as a result of a comment, changes the application of the life limit from the flap clevis to the flap clevis assembly. This amendment is prompted by the determination after an analysis of testing results that a certain flap clevis assembly should have an unlimited life. The actions specified by this revision are intended to remove the life limit for a specified flap clevis assembly. The actions specified by this AD are intended to prevent fatigue failure of the rotor shaft, teeter pin assembly, and flap clevis assembly, and subsequent loss of control of the helicopter.
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72-25-06:
72-25-06 BEECH: Amdt. 39-1572. Applies to Beech Models C-45G, TC-45G, C-45H, TC-45H, TC-45J (SNB-5), RC-45J (SNB-5P) D18C, D18S, E18S, E18S-9700, G18S, H18, JRB- 6, 3N, 3NM and 3TM Aircraft certificated in all categories with STC SA4-1531, STC SA111WE, STC SA1832WE or any other STC modification incorporating the provisions of the Volpar TriGear Installation.
Compliance required as indicated.
To prevent possible failure of the nose landing gear to extend from the retracted position on those aircraft with Volpar TriGear installed, accomplish the following:
For airplanes incorporating Volpar Tricycle Landing Gear, within the next 100 hours' time in service after the effective date of this AD unless already accomplished, modify the Volpar nose walking beam assembly, P/N 261, in accordance with Volpar Service Bulletin No. 21, as revised 7 August 1972, or later FAA-approved revision, or an equivalent modification approved by the Chief, Aircraft Engineering Division, FAA Western Region.
NOTE: For bulletins, parts or service information contact:
Volpar Incorporated
16300 Stagg Street
Van Nuys, California 91406
Phone 213-787-4393
NOTE: Pilots are cautioned to conduct a through pre-flight inspection in this area. A decal should be located on the nose gear shock strut housing listing the recommended strut inflation pressure.
This amendment becomes effective December 9, 1972.
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2024-24-10:
The FAA is superseding Airworthiness Directive (AD) 2023-05- 02, which applied to certain Airbus SAS Model A318, A319, A320, and A321 series airplanes. AD 2023-05-02 required revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations. This AD was prompted by a determination that new or more restrictive airworthiness limitations are necessary. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate additional new or more restrictive airworthiness limitations, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
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