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80-12-09: 80-12-09 GOVERNMENT AIRCRAFT FACTORIES (GAF): Amendment 39-3796. Applies to Models N22B (Serial Nos. N22B-5 and up) and N24A (Serial Nos. N24A-42 and up), certificated in all categories. Compliance required as indicated. To prevent possible failure of rudder trim tab control, accomplish the following: (a) Within the next 25 hours time in service after the effective date of the AD, unless already accomplished, visually inspect the rudder skin for loose rivets, and cracked intercostal flanges, and the rudder trim tab control rod for chafing, in accordance with GAF Nomad Alert Service Bulletin No. ANMD-55-11 (hereinafter referred to as the Service Bulletin) dated January 29, 1980, Part I and Part II, or an FAA-approved equivalent. (1) If no cracks are found as a result of the inspection required by paragraph (a) of this AD, within the next 100 hours time in service after the effective date of this AD, unless already accomplished, perform the following modificationin accordance with Part III of the Service Bulletin: (i) Replace rudder intercostal between W.L. 140.55 and W.L. 165.95 with an improved type or strengthen existing intercostal flanges. (ii) Rework lower rudder intercostal lightening hole and angle to prevent chafing of rudder trim tab control rod. (iii) Cut inspection hole in lower rib. (2) If loose rivets or cracked flanges are found, before further flight, accomplish the modifications required by sub-paragraph (1)(i), (ii) and (iii) of this paragraph. (b) Aircraft may be flown in accordance with FAR 21.197 and FAR 21.199 to a location where the modification can be performed. (c) For purposes of complying with this AD, an FAA-approved equivalent must be approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. NOTE: All persons affected by this directive who have not already received the Service Bulletin from the manufacturer, may obtain copies upon request to the Government Aircraft Factories, 226 Lorimer Street, Port Melbourne 3207 Vic., Australia. These documents may also be examined at the FAA, Engineering and Manufacturing District Office, 300 Ala Moana Blvd., Room 7321, Honolulu, Hawaii 96850, or Rules Docket, Room 916, FAA, 800 Independence Avenue, S.W., Washington, D.C. 20591. This amendment becomes effective June 23, 1980.
2000-25-02: This amendment supersedes Airworthiness Directive (AD) 98-05-04, which currently requires you to repetitively inspect the front and rear wood spars for damage (including installing any as-needed inspection holes) and repair or replace any damaged wood spar on certain American Champion Aircraft Corporation (ACAC) Model 8GCBC airplanes. Damage is defined as cracks, compression cracks, longitudinal cracks through the bolt holes or nail holes, or loose or missing nails. This AD retains the actions of AD 98-05-04 for the ACAC Model 8GCBC airplanes; extends the actions to all ACAC 7, 8, and 11 series airplanes (except the inspections are not repetitive for certain 7 and 11 series airplanes); incorporates alternative methods of accomplishing the actions; and requires reporting any damage found. This AD is the result of a review of the service history of the affected airplanes that incorporate wood wing spars where damage was found in this area and consideration of all public comments received. The actions specified by this AD are intended to detect and repair or replace damaged wood wing spars. Continued operation with such damage could progress to in-flight structural failure of the wing with consequent loss of control of the airplane
81-07-04: 81-07-04 LOCKHEED-CALIFORNIA: Amendment 39-4073. Applies to Model L-1011 series airplanes certificated in all categories. Compliance required as indicated. To prevent possible failures of the main landing gear wheels, accomplish the following: Inspect Goodrich main landing gear wheel assemblies for cracks, pits and corrosion in accordance with Lockheed Service Bulletin 093-32-184 dated February 5, 1981, and Goodrich Service Bulletin 393, Revision 1, dated February 20, 1981, in accordance with the following schedule: A. All in service Goodrich wheel assembly part numbers 3-1311-3 and 3-1365 (wheel outboard half P/N 10-1323 below change G, and wheel inboard half P/N 10-1324 and 10-1324-1 below change G) must be inspected at the next tire change on a main landing gear wheel/tire assembly after the effective date of this AD, and thereafter reinspected at each subsequent tire change. B. All in service Goodrich wheel assembly part numbers 3-1365 (wheel outboard half P/N10-1323 change G and up, and wheel inboard half P/N 10-1324-1 change G and up), 3-1365- 1, 3-1375, and 3-1375-1 must be inspected at the fourth tire change on a main landing gear wheel assembly after the effective date of this AD, and thereafter reinspected at each subsequent fourth tire change. C. Wheel halves found to have confirmed defects shall be replaced with an appropriate wheel-half inspected and found to be satisfactory in accordance with this AD or with a new or unused wheel half; except that if an inspected or new or unused wheel-half is not available, an uninspected wheel/tire assembly may be installed and the airplane operated to the next airport where the inspections can be accomplished, but in no event shall the uninspected wheel exceed ten landings. NOTE: Wheel half assemblies specified in paragraph B of this AD are considered new or unused until the fourth tire change. D. If a tire change is required at a station where the inspections required by this AD cannot be accomplished, an uninspected wheel/tire assembly may be installed and the airplane may be operated to the next airport where the inspections can be accomplished, but in no event shall the uninspected wheel exceed ten landings. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the inspection requirements of this AD. F. Alternative means of compliance, or other actions which provide an equivalent level of safety, may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Region. The manufacturer's specification and procedures identified and described in this directive and incorporated herein are made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents may obtain copies upon request to Lockheed-California Company, P.O. Box 551, Burbank, California 91520, Attention: Commercial Support Contracts. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108, or 15000 Aviation Boulevard, Hawthorne, California 90261. This amendment becomes effective April 15, 1981.
74-20-08: 74-20-08 CESSNA: Amendment 39-1975. Applies to all Models 310, 310A and 310B airplanes. Compliance: Required as indicated, unless already accomplished. To provide adequate inflight fire emergency procedure information, within the next 100 hours' time in service after the effective date of this AD, place in the cockpit either check list P/N 0811875-24 entitled "Aircraft Fire Procedures Checklist" or any equivalent checklist approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective October 1, 1974.
89-14-01: 89-14-01 TELEDYNE CONTINENTAL MOTORS (TCM): Amendment 39-6308. Final Rule of priority letter AD. Applicability: TCM Model TSIO-520BE engines (Serial Numbers 528001 through 528337) certificated in any category. Engines which have had the crankcase split and inspected and new bearings installed, since the accomplishment of AD 87-26-08, are exempt from the requirements of this AD. Compliance: Required as indicated, unless already accomplished. To prevent the possible loss of engine power, accomplish the following: (a) Prior to further flight and at intervals not to exceed 25 hours time-in-service, accomplish the crankshaft end play check in accordance with Section A of TCM Service Bulletin (SB) M89-14, dated June 29, 1989. If the crankshaft has no end play, the engine must be removed from service. (b) Prior to further flight, accomplish the thru-bolt torque check in accordance with Section B of TCM SB M89-14, dated June 29, 1989. If the force required to rotate the propeller is not within the range specified or if the force required to rotate the propeller after retorquing has changed from the previous value by more than 3 pounds or if the crankshaft has no end play, the engine must be removed from service. (c) Prior to further flight and at intervals not to exceed 200 hours time-in-service, accomplish the visual inspection of the number two crankcase main bearing in accordance with Section C of TCM SB M89-14, dated June 29, 1989. If there is any indication of bearing shift within the crankcase or crankshaft fillet/bearing contact or mismatch of bearing halves at the case split line, the engine must be removed from service. (d) The repetitive checks and inspections required by paragraphs (a) and (c) of this AD may be discontinued when the crankcase has been split and inspected and new bearings are installed. (e) Make appropriate log book entry showing compliance with this AD and record results of crankshaft end play and pounds required to rotate propeller. (f) Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations 21.197 and 21.199 to a base where this AD can be accomplished. (g) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, the Manager, Atlanta Aircraft Certification Office, Small Airplane Directorate, Aircraft Certification Service, Federal Aviation Administration, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia 30349, may approve an equivalent means of compliance or an adjustment of the compliance time schedule specified in this AD, which provides an equivalent level of safety. The checks and inspections shall be done in accordance with TCM SB M89-14, dated June 29, 1989. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Teledyne Continental Motors, P.O. Box 90, Mobile, Alabama 36601. Copies may beinspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the Office of the Federal Register, 1100 L Street, Room 8301, Washington, D.C. 20591. This amendment (39-6308, AD 89-14-01) becomes effective on September 22, 1989, as to all persons except those persons to whom it was made immediately effective by priority letter AD No. 89-14-01 issued on June 30, 1989, which contained this amendment.
2022-23-06: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Helicopters (Airbus) Model SA330J helicopters. This AD was prompted by a report of restricted movement of the collective lever caused by incidental contact of the secondary stop cover due to a loosened rivet. This AD requires removing the plate of the collective lever secondary stop and replacing it with self-adhesive tape to cover the stop support and decrease the risk of resistance on the rotor flight controls, 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.
47-10-14: 47-10-14 LOCKHEED: (Was Mandatory Note 16 of AD-763-3.) Applies to All Model 49 Serials Up to and Including 2088. Compliance required prior to April 15, 1947. Install a single flexible hose assembly between each fuel pump and flowmeter in place of the combination of short hose assembly and tube with hose couplings. (LAC Service Bulletin 49/SB-143 covers this same subject.)
2022-23-12: The FAA is adopting a new airworthiness directive (AD) for all The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747-200C, 747-200F, 747-300, 747-400, 747-400D, 747-400F, 747SR, and 747SP series airplanes. This AD was prompted by reports that high temperature composite trim air diffuser ducts (TADD) showed composite degradation and signs of hot air leakage. This AD requires a one-time low frequency eddy current (LFEC) inspection of certain center tank upper skin panels on the right and left side for any structural damage due to heat exposure, and repair if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
2000-25-07: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Boeing Model 737-100, -200, and -200C series airplanes. This action requires repetitive inspections of the aft end of each inboard flap track of the wing outboard flap, and corrective actions, if necessary. This action is necessary to detect and correct damage of the aft end of each flap track, which could result in loss of the outboard trailing edge flap and consequent loss of controllability of the airplane.
88-04-10: 88-04-10 DEHAVILLAND AIRCRAFT COMPANY OF CANADA, A DIVISION OF BOEING OF CANADA, LTD.: Amendment 39-5855. Applies to Model DHC-7 series airplanes, equipped with Modification No. 7/2414, certificated in any category. Compliance required as indicated, unless previously accomplished. To preclude the possibility of heat shield separation resulting from the failure of aluminum alloy washers, accomplish the following: A. Within 60 days or 500 flight hours, whichever occurs first after the effective date of this AD, replace aluminum alloy washers with stainless steel washers, in accordance with the Accomplishment Instructions of de Havilland DHC-7 Service Bulletin No. 7-57-29, dated August 1, 1986. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to The de Havilland Aircraft Company of Canada, A Division of Boeing of Canada, Ltd., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This amendment becomes effective April 5, 1988.
2000-25-06: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Pratt & Whitney (PW) PW4000 turbofan engines with the current design low pressure turbine (LPT) 4th stage air seal installed. This action requires, based on engine model, replacement of the current design seal with a new design seal, or with a modified seal. This amendment is prompted by reports of cracks in LPT 4th stage air seals. The actions specified by this AD are intended to reduce stresses that could lead to LPT 4th stage air seal cracking, resulting in seal fracture, uncontained engine failure, and damage to the airplane.
2016-06-01: We are superseding an airworthiness directive (AD) 2007-06-06 for B-N Group Ltd. Models BN-2, BN-2A, BN-2A-2, BN-2A-3, BN-2A-6, BN- 2A-8, BN-2A-9, BN-2A-20, BN-2A-21, BN-2A-26, BN-2A-27, BN-2B-20, BN-2B- 21, BN-2B-26, BN-2B-27, BN2A MK. III, BN2A MK. III-2, BN2A MK. III-3 BN2A, BN2B, and BN2A MKIII (all models on TCDS A17EU and A29EU) airplanes. 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 cracks in the inner shell of certain pitot/static pressure heads. We are issuing this AD to require actions to address the unsafe condition on these products.
52-17-02: 52-17-02 CURTISS-WRIGHT: Applies to all Model C-46 Series aircraft. Compliance required as noted. As a result of a number of failures the following precautionary measures must be taken: 1. The original compliance date set was not later than December 1, 1952, and at each No. 3 inspection thereafter, however, in order to convert the inspection interval to time in service, the following inspection shall be accomplished within 400 hours' time in service after the effective date of this amendment. Subsequent compliance required at each 400 hours' time in service after initial compliance. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southwest 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 such operator. Inspect the tail wheel shimmy damper support, P/N 20-360-3108-4, for cracks or damage which might lead to subsequent failure. Particular attention should be paid to the radius formed by the intersection of the vertical and lower horizontal surface on the forward side of the part. A general inspection of the shimmy damper assembly should be made and it should be determined that it is properly adjusted. All cracked or damaged parts must be replaced. The following method of adjustment of the shimmy damper as outlined in Air Force Technical Order AN 01-25LA-2, section IV, paragraph 8 (f) is quoted below for your convenience: (a) Position the shimmy damper on the mounting bracket and attach the two 3/8-inch bolts, nuts, and washers, using a 9/16-inch open-end wrench in conjunction with a 9/16- inch socket and a ratchet. (b) Before installing the link assembly to the shimmy damper arm, the arm must be positioned to allow for maximum travel of the unit in actual operation. This is accomplished by slowly actuating the arm toward the rear of the airplane until the movement stops. From this point reverse the action until the arm has made an arc of 67 degrees. It is now in a neutral position and will assure correct operation when completely installed. (c) With the arm in neutral and making sure not to move it from this position, attach the link assembly to it with the 5/16-inch bolt, nut, washer and cotter pin using a 1/2-inch open-end wrench, a 1/2-inch socket, and a ratchet. (d) Make sure that the tail wheel is locked in position. The pin on the centering disc will then be properly set for attaching the eyebolt of the link assembly. If the eyebolt hole does not match with the pin on the centering disc when the tail wheel is locked and the actuating arm is in neutral, adjustment must be made by loosening the locknut on the eyebolt and turning the eyebolt until the proper length is obtained. Be sure to tighten the locknut again.(e) Secure the link assembly to the centering disc by installing the 1 1/8-inch castellated nut and washer drawing up the nut with a crescent wrench. Insert a cotter pin on this nut. 2. Compliance required as noted. Tail wheel locks for the subject model airplane have been manufactured which do not comply with the material specifications and in some cases physical dimensions of the approved drawings. The approved tail wheel lock P/N 20-360-1033, is a forged steel part of 8740 or N-S- 16 material heat treated to 150,000 pounds per square inch, with a Rockwell harness of C-33 to C-38; however, a cast steel lock complying with the physical dimensions, heat treat and hardness of the approved drawing will also be considered acceptable. The original compliance date for the following inspection was March 1, 1953; however, in view of a recent accident involving the installation of an unapproved tail wheel lock, it must be ascertained prior to February 15, 1956, that all tailwheel locks meet the stated specifications of an acceptable part. Any tail wheel locks which do not comply with the approved drawing, No. 20-360-1033, in regard to material, heat treat or physical dimensions as mentioned above, must be replaced with an approved part. It should be noted that some unsatisfactory parts in circulation bear the correct part number, therefore this fact cannot be considered a satisfactory means of determining that an approved part is installed in the airplane. Revised October 23, 1961. Revised November 23, 1961.
93-10-08: 93-10-08 DE HAVILLAND: Amendment 39-8588. Docket No. 91-CE-49-AD. Supersedes AD 73-19-06, Amendments 39-1800 and 39-1710. Applicability: Models DHC-6-1, DHC-6-100, DHC-6-200, and DHC-6-300 airplanes (serial numbers 1 through 420), certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent failure of the overhead control console, which could result in loss of control of the airplane, accomplish the following: (a) Within the next 25 hours time-in-service (TIS) after the effective date of this AD, unless already accomplished within the last 175 hours TIS (compliance with AD 73-19-06), visually inspect the overhead console control quadrants for cracks in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland Service Bulletin (SB) 6/298, Revision D, dated December 20, 1991. (1) If cracks are found that meet or exceed the requirements specified in paragraph 3.1 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, prior to further flight, replace the overhead control console quadrants, part number (P/N) C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298. (2) If cracks are found that meet or exceed the requirements specified in paragraph 3.2 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, within the next 200 hours TIS, replace the overhead control console quadrants, P/N C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298. (3) If cracks are found that are less than that specified in paragraph 3.2 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, reinspect at intervals not to exceed 100 hours TIS until the modification specified in paragraph (b) of this AD is accomplished. (4) If no cracks are found, reinspect at intervals not to exceed 200 hours TIS until the modification specified in paragraph (b) of this AD is accomplished. (b) Within the next 2,400 hours TIS after the effective date of this AD, unless already accomplished as specified in paragraph (a)(1) or (a)(2) of this AD, replace the overhead control console quadrants, P/N C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991. (c) The installation of new overhead control console quadrants (Modification No. 6/1467) as specified in paragraphs (a)(1), (a)(2), and (b) of this AD is considered terminating action for the repetitive inspection requirement of this AD. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (e) An alternative method of compliance or adjustment of the compliance times that provides an equivalent level of safety may be approved by the Manager, New York Aircraft Certification Office, 181 Franklin Avenue, Room 202, Valley Stream, New York 11581. The request shall be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, New York Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York Aircraft Certification Office. (f) The inspections and modification required by this AD shall be done in accordance with de Havilland Service Bulletin (SB) 6/298, Revision D, dated December 20, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from de Havilland, Inc., 123 Garratt Boulevard, Downsview, Ontario, Canada, M3K 1Y5. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 800 North Capitol Street, NW, suite 700, Washington, DC. (g) This amendment (39-8588) supersedes AD 73-19-06, Amendments 39-1800 and 39-1710. (h) This amendment becomes effective on July 16, 1993.
2000-24-26: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Rolls-Royce plc RB211 Trent 800 series turbofan engines, that currently requires initial and repetitive ultrasonic inspections of fan blade roots for cracks, and replacement, if necessary, with serviceable parts. This amendment requires the reduction of the initial cyclic compliance threshold and repetitive inspection intervals. This amendment also allows inspections to be accomplished within 100 cycles-in-service if the initial or repetitive thresholds are exceeded on the effective date of this AD. This amendment is prompted by an improved understanding of the crack propagation mechanism and the latest service operational data. The actions specified by this AD are intended to detect and prevent fan blade failure, which could result in multiple fan blade releases, uncontained engine failure, and possible damage to the airplane.
2016-05-09: We are adopting a new airworthiness directive (AD) for certain MDHI Model 369A (Army OH-6A), 369H, 369HE, 369HM, 369HS, 369D, 369E, 369F, 369FF, and 500N helicopters. This AD requires inspecting the auxiliary fuel pump (fuel pump) wire routing in the left-hand fuel cell and corrective action, if necessary. This AD also requires installing a warning decal on the left-hand fuel cell access cover. This AD was prompted by accidents resulting from incorrectly positioned fuel pump wiring within the fuel tank interfering with the operation of the fuel quantity sensor float, which caused an erroneous fuel quantity indication in the cockpit. The actions are intended to detect and correct routing of the fuel pump wiring to prevent interference with the fuel quantity sensor float, an erroneous fuel quantity indication in the cockpit, and subsequent fuel exhaustion and emergency landing.
2009-03-03: We are adopting a new airworthiness directive (AD) for certain McDonnell Douglas airplanes listed above. This AD requires installing a dam assembly for the container of the fuel boost pump of the center tank located in the right main tank, and doing the related investigative actions, and corrective actions if necessary. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent the center tank fuel boost pump from operating in a fuel vapor zone and becoming a potential ignition source in the right main tank, potentially resulting in a fuel tank explosion and consequent loss of the airplane.
73-20-07 R2: 73-20-07 R2 BEECH: Amendment 39-1728 as amended by Amendment 39-3732 is further amended by Amendment 39-4461. Applies to Models A23-19, 19A, M19A, B-19 (Serial Numbers MB-1 through MB-708); Models 23, A23, A23A, B23, C23 (Serial Numbers M-1 through M-1576); Models A23-24, A24 (Serial Numbers MA-1 through MA-368); and Models A24R, B24R (Serial Numbers MC-2 through MC-282) certified in all categories. Compliance: Required as indicated. To detect cracks or other structural damage to the forward wing attach point brackets (P/N 169-400013-3 and -5(LH) and 169-400013-4 and -6 (RH)), within the next 25 hours' time in service after the effective date of this AD unless previously accomplished on airplanes having more than 100 hours' time in service and thereafter at each normal annual, progressive or 100 hour inspection interval as required by FAR 91.169, accomplish the following: (a) Remove the seats and sidepanels and visually inspect the forward wing attach brackets, P/N169-400013-3 and -5 (LH) and 169-400013-4 and -6 (RH), to determine if cracks or other structural damage exists in the area around the attach bolt hole. This inspection is to be performed in accordance with Beech Service Instructions 0042-031, Rev. II. (b) If as a result of any inspection required herein cracks or other structural damage is discovered, prior to further flight repair the forward wing attach brackets in accordance with Beechcraft Service Instructions 0042-031, or later FAA-approved revisions, or any other repair approved by the Chief, Wichita Aircraft Certification Office, FAA, Wichita, Kansas. To accomplish the repair required herein the aircraft may be flown in accordance with FAR 21.197 to a base where the repair may be performed. (c) The inspection intervals required in this AD may be adjusted 15 plus or minus hours where required to fit users maintenance cycles if authorized by an FAA Flight Standards Inspector. (d) When airplanes have been modified by installing all of the new wing attach structural parts listed under materials in Beechcraft Service Instructions No. 0042-031, Rev. II, further compliance with this AD is not required. (e) Any equivalent method of compliance with this AD must be approved by the Chief, Wichita Aircraft Certification Office, Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. Amendment 39-1728 became effective October 5, 1973. Amendment 39-3732 became effective March 21, 1980. This Amendment 39-4461 became effective on September 3, 1982.
92-16-19: 92-16-19 LOCKHEED AERONAUTICAL SYSTEMS COMPANY: Amendment 39-8329. Docket No. 92-NM-125-AD. Supersedes AD 92-10-51, Amendment 39-8271. Applicability: All Model L-1011-385 series airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent degradation of airplane pitch control, accomplish the following: (a) Prior to the accumulation of 12,000 landings, or within 3 days after June 24, 1992 (the effective date of AD 92-10-51, Amendment 39-8271), whichever occurs later, accomplish the procedures specified in paragraphs (a)(1), (a)(2), and (a)(3) of this AD: (1) Gain access to the lower end of the stabilizer hydraulic actuators (four per airplane) where they attach to the front spar of the horizontal stabilizer center box structure at Fuselage Station FS 1875. (2) Inspect for missing, sheared, or deformed stabilizer lower actuator attach pins, part number 1563117-101 (one per actuator). (3)If any pin is missing, sheared, or deformed, replace the pin prior to further flight. (b) At the applicable time specified in paragraph (b)(1), (b)(2), or (b)(3) of this AD, either perform a one-time magnetic particle inspection to detect cracks on each of the four lower horizontal stabilizer actuator pins (part number 1563117-101), or replace each of the four actuator pins (part number 1563117-101) with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. (1) For airplanes that have accumulated less than 12,000 landings as of the effective date of this AD: Prior to the accumulation of 12,000 landings, or within 60 days after the effective date of this AD, whichever occurs later. (2) For airplanes that have accumulated at least 12,000 landings but not more than 20,000 landings as of the effective date of this AD: Within 45 days after the effective date of this AD. (3) For airplanes that have accumulated more than20,000 landings as of the effective date of this AD: Within 30 days after the effective date of this AD. (c) If a magnetic particle inspection is performed in accordance with paragraph (b) of this AD, accomplish the procedures specified in paragraph (c)(1) or (c)(2), as applicable: (1) If any actuator pin is found to be cracked: Prior to further flight, replace the cracked pin with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, part number 1563117-101, it must be replaced with a new pin. (2) If no actuator pins are found to be cracked: Prior to the accumulation of 1,000 landings after accomplishing the magnetic particle inspection, replace the pin with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, part number 1563117-101, it must be replaced with a new pin. (d) If each of the four actuator pins, part number 1563117-101, are replaced with a new pin in accordance with paragraph (b) of this AD: Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, it must be replaced with a new pin. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Atlanta Aircraft Certification Office (ACO), FAA, Small Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Atlanta ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Atlanta ACO. (f) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a locationwhere the requirements of this AD can be accomplished. (g) The inspection and replacement shall be done in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992, which contains the following list of effective pages: Page Number Revision Level Date 1 - 8 1 May 28, 1992 9 - 11 Original May 15, 1992 This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Lockheed Western Export Company (LWEC), Dept. 693, Zone 0755, 86 South Cobb Drive, Marietta, Georgia 30063. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the FAA, Atlanta Aircraft Certification Office, Suite 210C, 1669 Phoenix Parkway, Atlanta, Georgia; or at the Office of the Federal Register, 800 North Capitol Street NW., 7th Floor, Suite 700, Washington, DC. (h) This amendment becomes effectiveon September 1, 1992.
2000-24-24: This amendment adopts a new airworthiness directive (AD) that is applicable to Rolls- Royce plc RB211 Trent 768-60, Trent 772-60, and Trent 772B-60 series turbofan engines having common nozzle assembly part number (P/N) FK16544 or FK16558. This action requires initial and repetitive visual inspections of the inner and outer skins of the common nozzle assembly and specifies allowable limits for cracks, loose rivets, and missing rivets. This action also requires repair if the common nozzle assembly damage exceeds allowable limits. This amendment is prompted by two reports of in-flight inner skin detachment. The actions specified in this AD are intended to detect cracks, loose rivets, and missing rivets, which could result in inner skin detachment, release of common nozzle assembly debris from the engine, and possible damage to the airplane control surfaces.
89-03-06: 89-03-06 FOKKER: Amendment 39-6122. Applicability: Model F-27 series airplanes, Serial Numbers 10623 through 10692, in Mark 500 configuration, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent loss of electrical power and eliminate a potential fire hazard, due to chafing wire bundles, accomplish the following: A. Within 60 days after the effective date of this AD, perform a one-time inspection of the upper fuselage cableloom for chafing and insufficient clamping, and repair, if necessary, in accordance with Fokker Service Bulletin F27/24-77, dated June 7, 1988. B. An 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, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Standardization Branch, ANM-113. C. Special 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. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Fokker Aircraft USA, Inc., 1199 N. Fairfax Street, Alexandria, Virginia 22314. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6122, AD 89-03-06) becomes effective March 6, 1989.
2022-23-10: The FAA is superseding Airworthiness Directive (AD) 2021-06- 03, which applied to all The Boeing Company Model 777F series airplanes. AD 2021-06-03 required deactivating the potable water system. This AD was prompted by a report of a water supply line that detached at a certain joint located above an electronic equipment (EE) cooling filter, leading to water intrusion into the forward EE bay. This AD retains the actions required by AD 2021-06-03 and requires installing a shroud to the water supply line in the forward cargo compartment, and performing a leak test of the potable water system. For certain airplanes, this AD also requires replacing tubes and hoses from the water supply line and installing a shroud to the water return line. The FAA is issuing this AD to address the unsafe condition on these products.
2010-21-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD requires installing a support bracket and coupler on the left and right wing-to-fuselage transition, and metallic overbraid on the left and right leading edge wire assembly. This AD was prompted by fuel system reviews conducted by the manufacturer, as well as reports that the fuel quantity system was affected by lightning-induced transients. We are issuing this AD to prevent lightning-induced transients to the fuel quantity indication system, which could cause voltage levels to go beyond original design levels between fuel tank probes and structure, and become a potential ignition source at the fuel tank, which, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
83-04-02: 83-04-02 FOKKER-VFW B.V.: Amendment 39-4569. Applies to Model F28 airplanes serial number 11003 and on. To prevent loss of aileron control, accomplish the following, unless already accomplished within the last three months prior to the effective date of this AD: 1. Within the next 50 hours time in service or one month, whichever occurs first after the effective date of this AD, perform a one time inspection of the aileron control cables in both the R.H. and L.H. wings for airplanes with more than one year's time from the date of manufacture in accordance with the following instructions: a. Open the main landing gear doors and flap shroud doors 81j, 81k, 82j, and 82k. b. Rotate the aileron control wheel in the cockpit to the full left-hand position. c. With the control wheel in this position, inspect the four aileron input cables at location of cable pulleys between stations 3100 and 3600 (illustrated parts catalog, chapter 27-10-01-05). Inspection must be performed with the aid of a mirror and a very strong light otherwise corrosion may remain undetected. d. If wires are broken or there are definite signs of corrosion in the cables, accomplish the actions of paragraph 2. e. Rotate the aileron control wheel in the cockpit to the full right-hand position and repeat the inspections described in paragraphs (c) and (d). 2. If corrosion on the cables is evident from the inspections of paragraph 1, above, accomplish the following: a. Disconnect the cable at the turnbuckle in the wheelwell per the F28 Maintenance Manual, chapter 27-11. b. Pull out the cable at the cable pulleys between stations 3100 and 3600. c. Carefully bend the cable one time at the location of the corrosion over a radius of one inch. If any wire breaks while being bent, replace the cable prior to further flight. d. If no wires failed during the bending test of paragraph 2(c), repeat the test by bending the cable in the opposite direction. If any wire breaks, replace the cable before further flight. e. If the cable is not rejected by the bending tests of paragraphs 2(c) or 2(d), inspect the cable for interior corrosion in accordance with the F28 Maintenance Manual, chapter 20-30-9. f. If interior corrosion is found, replace the cable before further flight. g. Remove any exterior corrosion and treat the cable in accordance with F28 Maintenance Manual, Chapter 20-30-9. h. Reconnect the turnbuckles in the L.H. and R.H. wheelwells and adjust the cable system in accordance with F28 Maintenance Manual, Chapter 27-11-00, and check the aileron for free and smooth operation. 3. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. 4. Special 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. 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. 552(a)(1). This amendment becomes effective March 7, 1983.
2000-15-52: This document publishes in the Federal Register an amendment adopting superseding Airworthiness Directive (AD) 2000-15-52, which was sent previously to all known U.S. owners and operators of Bell Helicopter Textron, Inc. Model (BHTI) Model 204B, 205A, 205A-1, 205B, and 212 helicopters by individual letters. This AD reduces the retirement index number (RIN) life limit for the main rotor mast (mast); increases the RIN factor for masts and main rotor trunnions (trunnions); applies standard RIN factors for all external load lifts; and requires a one-time inspection of the snap ring groove area of the mast. This AD also establishes RIN factors for masts and trunnions that have been previously installed on military or restricted category helicopters and removes from service those masts that have been previously installed with a hub spring. This amendment is prompted by an occurrence of a cracked mast at a lower value than the established RIN life limit. The actions specified by thisAD are intended to preclude the occurrence of fatigue cracks in the damper clamp splined area of a mast. A crack in the damper clamp splined area could result in failure of a mast or trunnion, separation of the main rotor system, and subsequent loss of control of the helicopter.