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83-07-15: 83-07-15 BRITISH AEROSPACE, AIRCRAFT GROUP, SCOTTISH DIVISION: Amendment 39-4617. Applies to Model HP.137 Jetstream MK.1 and Jetstream Series 200 airplanes certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent cracking of the front pressure bulkhead boundary angle and possible catastrophic depressurization, accomplish the following: a) For all affected airplanes with 12,900 or more flights, unless British Aerospace Modification No. 5115 is installed, within the next 100 flights after the effective date of this AD, apply a red line to the cabin differential pressure gauge to indicate that the maximum allowable differential pressure is 2.5 PSI. Extend the red line to the case of the instrument to avoid confusion if the glass cover is rotated. b) For all other affected airplanes with less than 12,900 flights, unless British Aerospace Modification No. 5115 is installed, within the next 100 flights after the effective date of this AD or upon the accumulation of 2,000 flights, whichever occurs later, apply a red line to the cabin differential pressure gauge to indicate that the maximum allowable differential pressure is 4.0 PSI except that in no case shall airplanes in this group be allowed to accumulate more than 13,000 flights without having the cabin differential pressure gauge remarked with a red line at 2.5 PSI in accordance with paragraph a) of this AD. Extend the red line to the case of the instrument to avoid confusion if the glass cover is rotated. c) For all aircraft modified per paragraphs a) and b) of this AD, fabricate a placard to read as follows and install it adjacent to the pressure differential gauge: 1) For aircraft affected by paragraph a) of this AD: "2.5PSI max. press. difference." 2) For aircraft affected by paragraph b) of this AD: "4.0PSI max. press. difference." d) Within the next 500 flights after the effective date of this AD on all affected airplanes, unless British Aerospace Modification No. 5115 is installed, modify the pressurization systems in accordance with British Aerospace, Aircraft Group Modification No. 5151, Issue 1, dated August 1981 or Modification No. 5152, Issue 1, dated August 1981, as appropriate to limit maximum pressurization to 4.0 PSI or 2.5 PSI respectively. e) For purposes of complying with this AD, subject to the acceptance by the assigned FAA maintenance inspector, the number of flights may be determined by multiplying each airplane's hours time-in-service by two. f) The modifications specified by paragraphs a), b) and (c), or d) of this AD are no longer applicable when British Aerospace Modification No. 5115 is installed. g) Aircraft may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. h) An equivalent method of compliance with this AD if used must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on April 14, 1983.
92-06-11: 92-06-11 BEECH: Amendment 39-8191; Docket 91-CE-78-AD. Applicability: The following Beech model airplanes that are equipped with Collins DPU-85N Display Processor Units (DPU) and Collins MPU-85N Multifunction Processor Units (MPU), certificated in any category: Model Serial Numbers B200 BB-1349 through BB-1409 300 FA-205 through FA-217 B300 FL-1 through FL-60 B300C FM-1 through FM-3 1900C UC-73, UC-139 and UC-147 Compliance: Required as indicated, unless already accomplished. To prevent incorrect pilot decisions based on undetected erroneous attitude information displayed by the Collins EFIS-85/86B (14) system or undesired autopilot movement of the airplane, accomplish the following: (a) Within the next 50 hours time-in-service (TIS) after the effective date of this AD, modify the airplane autopilot/flight director system wiring in accordance with the instructions in Beech Service Bulletin (SB) No. 2423, dated December 1991. (b) Within the next 6 calendar months after the effective date of this AD, modify the hardware of the Collins EFIS-85/86B (14) system, Collins DPU-85N Display Processor Unit, and Collins MPU-85N Multifunction Processor Unit in accordance with the instructions in Collins SB DPU-85N-34-51 and SB MPU-85N-34-51, both dated June 6, 1991. (c) 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. (d) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (e) The modifications required by this AD shall be done in accordance with Beech Service Bulletin No. 2423, dated December 1991, and Collins Service Bulletin DPU-85N-34-51 and Collins Service Bulletin MPU-85N-34-51, both dated June 6, 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 the Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201-0085; or Rockwell International, Collins General Aviation Division, 400 Collins Road, NE; Cedar Rapids, Iowa 52498. 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, 1100 L Street, NW; Room 8401, Washington, DC. (f) This amendment (39-8191) becomes effective on March 30, 1992.
99-15-09: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-600 series airplanes. This action requires revising the Airplane Flight Manual (AFM) to prohibit operation of the airplane under certain conditions; repetitive inspections of the tab mast fittings of the elevator tab assemblies to detect cracking; an elevator tab freeplay check; and corrective actions, if necessary. This AD also requires installing an additional fastener on the elevator tab mast fitting, which terminates the AFM revision and extends certain repetitive inspections. This AD also requires replacement of the elevator tab mast fitting with a new, improved fitting, which constitutes terminating action for the requirements of this AD. This amendment is prompted by a report of a severe vibration incident on a Boeing Model 737-800 series airplane; inspection revealed fracturing of the elevator tab mast fitting and excessive freeplay in the elevator tab. The actionsspecified in this AD are intended to prevent loss of controllability of the airplane due to excessive freeplay in the elevator tab or a free tab.
2022-14-51: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Helicopters Model EC225LP helicopters. This AD was prompted by a report of a cracked main rotor hub (MRH) sleeve. This AD requires one-time visual inspections and, depending on the results, accomplishing additional inspections, repairing the MRH sleeve in accordance with a certain approval, and removing the MRH sleeve from service and installing an airworthy part. This AD also prohibits installing an MRH sleeve unless certain inspections have been accomplished. The FAA previously sent an emergency AD to all known U.S. owners and operators of these helicopters. The FAA is issuing this AD to address the unsafe condition on these products.
2013-24-11: We are adopting a new airworthiness directive (AD) for certain 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 a report of a disbonded doubler and a skin crack in section 41 of the fuselage, and multiple reports of cracked or missing fastener heads. This AD requires repetitive inspections for cracking of the fuselage skin, discrepant fasteners, and for disbonds at the doublers; and related investigative and corrective actions if necessary. For certain airplanes, this AD also requires a terminating repair for repair doublers. We are issuing this AD to prevent rapid decompression and loss of structural integrity of the airplane due to such disbonding and subsequent cracking of the skin panels.
99-15-07: This amendment adopts a new airworthiness directive (AD) that applies to all deHavilland Inc. (deHavilland) Models DHC-2 Mk. I, DHC-2 Mk. II, and DHC-2 Mk. III airplanes. This AD requires repetitively inspecting the rear fuselage bulkhead at Station 228 for cracks. This AD also requires repairing any crack found or replacing any cracked rear fuselage bulkhead in accordance with a repair or replacement scheme obtained from the manufacturer through the Federal Aviation Administration (FAA). This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Canada. The actions specified by this AD are intended to detect and correct cracking of the rear fuselage bulkhead at Station 228, which could result in structural damage of the fuselage to the point of failure with consequent loss of airplane control.
86-24-09: 86-24-09 BEECH AIRCRAFT CORPORATION: Amendment 39-5482. Applies to Model B-100 (all serial numbers) airplanes certificated in any category. Compliance: Required within the next 50 hours time-in-service after the effective date of this AD, unless already accomplished. To prevent engine flameout when in or departing an icing environment, accomplish the following: (a) Revise the airplane Pilot's Operating Handbook and Airplane Flight Manual (POH/AFM) by inserting Appendix 1 of this AD in the "LIMITATIONS" section of the POH/AFM. Appendix 1 procedures supersede any other POH/AFM procedures which may be contradictory. (b) The requirements of paragraph (a) of this AD may be accomplished by the holder of a pilot certificate issued under Part 61 of the Federal Aviation Regulations on any airplane owned or operated by him. The person accomplishing these actions must make the appropriate aircraft maintenance record entry as prescribed by FAR 91.173. (c) Airplanesmay be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (d) An equivalent method of compliance with this AD, if used, must be approved by contacting the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209; Telephone (316) 946-4400. All persons affected by this directive may obtain copies of the document referred to herein upon request to GTEC, P.O. Box 5217, Phoenix, Arizona 85010; Telephone (602) 231- 1000; or Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201; Telephone (316) 681- 9111; or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on December 15, 1986. APPENDIX 1 Supplement to the POH/AFM Beech Model B-100 Airplanes "The MAN FUEL/IGN switches shall be selected to ON during all operations in actual or potential icing conditions described herein: (1) During takeoff and climb out in actual or potential icing conditions. *(2) When ice is visible on, or shedding from propeller(s), spinner(s), or leading edge(s). *(3) Before selecting ANTI-ICE, when ice has accumulated. (4) Immediately, any time engine flameout occurs as a possible result of ice ingestion. (5) During approach and landing while in or shortly following flight in actual or potential icing conditions. * NOTE: If icing conditions are entered in flight without the engine anti-icing system having been selected, switch one ENGINE system to the ANTI-ICE ON position. If the engine runs satisfactorily, switch the second ENGINE system to the ANTI-ICE ON position and check that the second engine continues to run satisfactorily. CAUTION Flight in actual or potential icing conditions will be limited by duty cycle of the ignition system. Ignition system time limits must be observed to prevent exceeding duty cycle times. Operator should verify these limits for his particular installation. For the purpose of this supplement, the following definition applies: "Potential icing conditions in precipitation or visible moisture meteorological conditions: (1) Begin when the OAT is plus 5 degrees C (plus 41 degrees F) or colder, and (2) End when the OAT is plus 10 degrees C (plus 50 degrees F) or warmer." The procedures and conditions described in this appendix supersede any other POH/AFM procedures and conditions which may be contradictory.
2013-24-12: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 747-8 and 747-8F series airplanes. This AD requires repetitive ultrasonic or dye penetrant inspections for cracking of the barrel nuts and bolts, as applicable, on each forward engine mount, and related investigative and corrective actions if necessary. This AD was prompted by a report of cracked barrel nuts found on a forward engine mount. We are issuing this AD to detect and correct cracked barrel nuts on a forward engine mount, which could result in reduced load capacity of the forward engine mount, and could result in separation of an engine under power from the airplane, and consequent loss of control of the airplane.
2010-15-07: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Cracks on the lug of the rear attachment fitting of the horizontal stabilizer have been detected during the inspection of two PW-6U gliders operated by the same user. This condition, if not corrected, could result in no longer retaining the horizontal stabilizer in place and consequent loss of control of the aeroplane. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
90-12-05: 90-12-05 PRATT & WHITNEY CANADA: Amendment 39-6607. Docket No. 90-ANE-10. Applicability: Pratt & Whitney Canada (PWC) PT6A series turbopropeller engines repaired by Gregory Flying Service/Airforce Turbine Service (hereafter identified as Gregory Flying Service), located in Tynan, Texas. The affected engines are identified by model and serial number as follows: MODEL SERIAL NO. MODEL SERIAL NO. PT6A-15AG PCE 14012 PT6A-34AG PCE 56433 PT6A-15AG PCE 14018 PT6A-34AG PCE 56466 PT6A-15AG PCE 14028 PT6A-34AG PCE 56471 PT6A-34 PCE 38093 PT6A-34AG PCE 56516 PT6A-28 PCE 40023 PT6A-34AG PCE 56522 PT6A-27 PCE 41704 PT6A-34AG PCE 56527 PT6A-28 PCE 51327 PT6A-34AG PCE 56612 PT6A-29 PCE 51426 PT6A-34AG PCE 56628 PT6A-34 PCE 56089 PT6A-34AG PCE 56650 PT6A-34AG PCE 56381 PT6A-34AG PCE 56717 PT6A-34AG PCE 56405 PT6A-34AG PCE 56773 Compliance: Required as indicated, unless already accomplished. To prevent the development of an unsafe condition which could result in an uncontained engine failure, accomplish the following: (a) Remove from service those engines identified above within the next 30 calendar days after the effective date of this AD. (b) Return to service engines removed from service in accordance with paragraph (a) by accomplishing either one of the two following requirements: (1) Accomplish a complete engine overhaul in accordance with the requirements of the appropriate PWC overhaul manual. (2) Submit to the FAA, Manager, Engine Certification Office, through the cognizant FAA Airworthiness Inspector, all the pertinent records identified below for review, and obtain a written FAA approval prior to returning an engine to service: (i) For the parts, sub-assemblies, accessories, or components of engines that were subjected to repair or maintenance activities during the last shop visit at Gregory Flying Service, provide the following: (1) A list of all life-limited components by serial number, service history, and their current status as required by FAR Part 135, Section 439, or FAR Part 91, Section 173, as appropriate, and records establishing the origin and/or the prior service history of the subject parts. (2) Records of accomplishment of the required inspections, checks, tests, as applicable, in accordance with PWC engine manual requirements that establish the components' airworthiness. (ii) All repair, maintenance, and inspection records, concerning the last shop visit at Gregory Flying Service, as required by FAR Part 135, Section 439, or FAR Part 91, Section 173, as appropriate. (iii) Substantiating evidence that the work performed during the last shop visit at Gregory Flying Service was done in accordance with FAA approved data as required by the FAR's. (iv) Engine acceptance test data or engine installed test data, accomplished after the repair, whichever is applicable. (v) A list by part number or serial number of any engine, part, sub-assembly, accessory, or component acquired from Gregory Flying Service that has been involved in an accident. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (Pub. L. 96-511) and have been assigned OMB Control No. 2120-0056. (c) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (d) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance time specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment (39-6607, AD 90-12-05) becomes effective on May 30, 1990.
99-13-09: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 99-13-09 which was sent previously to all known U.S. owners and operators of MDHI Model 369D and E helicopters by individual letters. This AD requires, prior to further flight, inspecting and replacing, if necessary, a certain four-bladed tail rotor fork (fork) assembly. This AD also requires a repetitive inspection of certain fork assemblies at intervals not to exceed 50 hours time-in-service (TIS) and removing and replacing, if necessary, each unairworthy fork assembly with an airworthy fork assembly before further flight. This amendment is prompted by reports from the manufacturer of the discovery of a discrepant part. The actions specified by this AD are intended to prevent failure of certain fork assemblies, which could cause loss of a tail rotor blade and subsequent loss of control of the helicopter.
92-01-05: 92-01-05 BELL HELICOPTER TEXTRON, INC. (BHTI): Amendment 39-8128. Docket Number 89-ASW-54. Applicability: All Bell Helicopter Textron, Inc. Model 206A, 206B, 206L, 206L-1 and 206L-3 helicopters, certificated in any category, with main rotor blade part numbers (P/N) 206-015-001-001, 206-015-001-103, 206-015-001-105, or 206-010-200-033 installed. Compliance: Required before further flight, unless already accomplished. To prevent failure and separation of the main rotor blades, and subsequent loss of the helicopter, accomplish the following: (a) Visually inspect the Model 206A and 206B main rotor blades and determine if one of the following serial number (S/N) blades is installed: TAC-0089, TAC-0542, TAC-0607, TAC-0614, TAC-0624, TAC-1643, TAC-1749, TAC-1776, TAC-1831, TAC-1911, TAC-1922, TAC-2399, TAC-2768, TAC-5742, TKK-9794, TKK-9883, or TKK-9933. If any one of these main rotor blades is installed, remove and replace with a serviceable part prior to furtherflight. (b) Visually inspect the Model 206L, 206L-1, and 206L-3 main rotor blades and determine if one of the following S/N blades is installed: T-92, T-245, T-417, TLY-0075, TLY-0095, TLY-0764, TLY-0770, TLY-0973, TLY-1438, TLY-1619, TLY-1653, TLY-1697, TLY-1766, TLY-1801, TLY-1858, TLY-1953, TLY-1984, TLY-2031, TLY-2039, TLY-2064, TLY-2081, TLY-2148, TLY-2335, TLY-2337, TLY-2549, TLY-2603, TLY-2604, TLY-2625, TLY-2633, TLY-2648, TLY-2745, TLY-2786, TLY-2951, or TLY-2954. If any one of these main rotor blades is installed, remove and replace with a serviceable part prior to further flight. NOTE: The serial number may be found on the Bell Helicopter data plate located on top of the blade at the root end and is also marked on the root end of the lower grip plate in the l.5 inch radius. (c) If the serial number of the main rotor blade matches one listed in paragraph (a) or (b) of this AD, report the registration number and serial number of the affected helicopter and provide a copy of the parts tag with which the part was delivered, if available. Send the report to the Manager, Rotorcraft Certification Office, Southwest Region, Federal Aviation Administration, Fort Worth, Texas 76193-0170, telephone (817) 624-5170, within 10 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056.) (d) An alternative method of compliance which provides an equivalent level of safety, may be used if approved by the Manager, Rotorcraft Certification Office, Southwest Region, Federal Aviation Administration, Fort Worth, Texas 76193-0170, telephone (817) 624-5170. (e) This amendment (39-8128, AD 92-01-05) becomes effective on February 5, 1992, as to all persons except those persons to whom it was made immediately effective by Priority Letter AD's 89-22-01, and 89-22-01 R1 issued October 18, 1989, and November 21, 1989, respectively, which contained the requirements of this amendment.
2013-17-02: We are adopting a new airworthiness directive (AD) for certain Airbus Model A318-112, A319-111, A319-112, A319-115, A319-132, and A319-133 airplanes. This AD was prompted by a report that a fastener, which connects the cargo door keel beam foot to the circumferential butt-strap and the section 13-14 lower shell panel, was not installed on airplanes during production. This AD requires inspecting forward fuselage frame 24, stringer 39, right hand, to determine if the fastener is missing; measuring the hole dimensions of the five holes surrounding the missing fastener if necessary; and doing related investigative and corrective actions if necessary. We are issuing this AD to detect and correct the missing fastener, which could result in reduced structural integrity of the airplane.
2010-15-05: We are superseding an existing 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: Cracks on the stabilizer elevator inner hinges of seven L 23 SUPERBLAN K sailplanes have been detected during an inspection. This condition, if not corrected, could result in no longer retaining the elevator in place and in jamming of the Pilot's elevator control system, and subsequent loss of elevator control. We are issuing this AD to require actions to correct the unsafe condition on these products. DATES: This AD becomes effective August 30, 2010. As of April 26, 2010 (75 FR 17295, April 6, 2010), the Director of the Federal Register approved the incorporation by reference of Aircraft Industries, a.s. Mandatory Bulletin MB No.: L23/052a, dated March 2, 2010, listed in this AD.
99-15-06: This amendment adopts a new airworthiness directive (AD), applicable to AlliedSignal Inc. Model ALF502R-5 and ALF502R-3A turbofan engines, that requires incorporation of an improved fan core inlet anti-ice system. This amendment is prompted by reports of uncommanded reduction of engine thrust (rollback) and loss of thrust control in icing conditions. The actions specified by this AD are intended to prevent ice accretion on the fan core inlet stator vane surfaces, which can result in engine rollback and loss of thrust control in icing conditions.
65-23-01: 65-23-01 VICKERS: Amdt. 39-145 Part 39 Federal Register September 30, 1965. Applies to Viscount Models 744, 745D, and 810 Series Airplanes. Compliance required as indicated. As a result of reported cracks in the inner and outer wing spar attachment joint lugs, inspections were made and a number of airplanes were found to have cracks in the spar boom joint lugs. These cracks may occur on any one of the fingers of the joint lugs and appear to originate in the taper holes and then generally progress in a direction parallel to the joint lug. Accordingly, the following must be accomplished: (a) Unless already accomplished as required by AD 61-23-05, within the next 20 hours' time in service after the effective date of this AD, inspect for cracks using ultrasonic methods, or FAA-approved equivalent, all the outer wing to inner wing spar boom attachment joint lugs, top and bottom, right and left, and the inner wing to center section spar boom attachment joint lugs, top and bottom, right and left, in the region of the taper bolt holes. (1) If cracks are found in any of the lower spar boom joints, replace with new spar booms before further flight. (2) If no cracks are found in any of the lower spar boom joints, airplanes having cracks in the top spar boom joints, within the limits specified in (a)(2)(i) may be continued in service provided the inspection required by (a) is repeated on the affected top spar boom joints at intervals specified in (b)(4). (i) One crack is permitted in any of the top four joints, i.e., a total of four cracks per airplane. Permissible cracks are those extending completely between two adjacent holes in one lug only; or extending between the bolt hole nearest the end of one lug and the end of that lug; or between the bolt hole nearest the boom body and a line one inch from this hole toward the body of the boom, in one lug only. A number of small cracks confined within any one of these areas may be considered as one single crack for the purpose of this limitation. (3) If cracks beyond the limits specified in (a)(2)(i) are found in any of the top spar boom joints, replacement with a new top spar boom is required before further flight; except that, if there are no cracks in any of the lower spar boom joints and the extent of the cracking in the top spar boom joints has been reported to B.A.C. (Operating) Ltd. (Weybridge Division) for evaluation and the operator has obtained and presented to the FAA approval for flight from B.A.C. (Operating) Ltd. based upon such evaluation, the airplane may be flown in accordance with FAR 21.197 to a base where replacement with a new top spar boom can be accomplished. (b) Subsequent to the initial inspection required by (a) the following repetitive inspections must be accomplished. (1) AIRPLANES FITTED WITH DTD 363 BOTTOM SPAR BOOMS (744 AIRPLANES). (i) Accomplish ultrasonic or FAA-approved equivalent inspection on all the bottom spar boom joints not later than six months after the date of initial refitment of the taper bolts or subsequent reprotection of the spar boom joints. (ii) Bottom Outer to Inner Wing spar boom joints that have not been line reamed - Accomplish repetitive ultrasonic or FAA-approved equivalent inspection at intervals not exceeding 18 months, commencing from the date of the initial refitment of the taper bolts or subsequent reprotection of the spar boom joints. (iii) Bottom Outer to Inner Wing spar boom joints which have been line reamed, or Blueing Checks of the taper bolts has positively established the contact area between the bolt and the bolt hole to be not less than 95 percent including the web plates - Accomplish repetitive ultrasonic or FAA-approved equivalent inspection at intervals not exceeding two years, commencing after the date of initial refitment of the taper bolts or subsequent reprotection of the spar boom joints. (iv) Bottom Inner Wing to Center Section spar boom joints - Accomplish repetitive ultrasonic or FAA-approved equivalent inspection at intervals not exceeding two years, commencing from the date of initial refitment of the taper bolts or subsequent reprotection of the spar boom joints. (2) AIRPLANES FITTED WITH L.65 BOTTOM SPAR BOOMS (745D & 810 AIRPLANES). (i) Accomplish repetitive ultrasonic or FAA-approved equivalent inspection at intervals not exceeding three years, commencing from the date of the last ultrasonic inspection or the subsequent six yearly reprotection of the spar boom joints. (ii) If cracks are found in any of the lower spar boom joints, replacement with new spar booms is required prior to further flight. (3) ALL AIRPLANES: TOP SPAR BOOM JOINTS. (i) Inspect all top spar boom joints, in which the taper bolts have not been retensioned to the revised instruction detailed in the applicable PTL referenced herein, for cracks using ultrasonic methods, or FAA-approved equivalent, at intervals notexceeding six months commencing from the date of the last ultrasonic or FAA-approved equivalent inspection. Airplanes found to have cracks in the top spar boom joints which are within the limits specified in (a)(2)(i) and Sect. II, Part 2 "Limitations" of the applicable PTL referenced herein may be continued in service and must be reinspected in accordance with (b)(4). When cracks are found that exceed the limits of (a)(2)(i) or Section II Part 2 "Limitations" of the applicable referenced PTL, the spar boom must be replaced before further flight in accordance with (a)(3). (ii) Inspect top spar boom joints in which the taper bolts have been retensioned to the revised instruction detailed in the applicable referenced PTL, for cracks using ultrasonic methods or FAA-approved equivalent, at intervals not exceeding two years commencing from the date of refitment of the taper bolts. (iii) If, for any reason, an additional ultrasonic inspection has been carried out after retensioning, but before expiration of the two year period, the next ultrasonic inspection must be accomplished within not later than two years from the date of the subsequent inspection. Airplanes found to have cracks in the top spar boom joints which are within the limits specified in (a)(2)(i) and Section II Part 2 "Limitations" of the applicable PTL referenced herein may continue in service and must be reinspected in accordance with (b)(4). When cracks are found that exceed the limits in (a)(2)(i) and Section II Part 2 "Limitations" of the applicable PTL the spar boom must be replaced before further flight in accordance with (a)(3). (4) On airplanes having cracks in the top spar boom joints which are within the permissible limits of (a)(2)(i) and Section II Part 2 "Limitations" of the applicable PTL referenced herein inspect the affected joint using ultrasonic methods or FAA-approved equivalent as follows: (i) Where the acceptable defect has a reflectivity of at least 30 percent but not more than 100 percent of the appropriate 5/64 inch flat bottom hole standard, at intervals not exceeding 12 months, or at such periods as are approved by the airplane manufacturer, or the Chief, Aircraft certification Staff, FAA European Region, commencing from the date of the last ultrasonic inspection. (ii) Where the acceptable defect has a reflectivity above 100 percent but not exceeding 200 percent of the appropriate 5/64-inch flat bottom hole standard, at intervals not exceeding six months or at such periods as are approved by the airplane manufacturer, or the Chief, Aircraft Certification Staff, FAA European Region, commencing from the date of the last ultrasonic inspection. (iii) Where the acceptable defect has a reflectivity more than 200 percent of the appropriate 5/64-inch flat bottom hole standard, at intervals not exceeding three months or at such periods as are approved by the airplane manufacturer, or by the Chief, Aircraft Certification Staff, FAAEuropean Region, commencing from the date of the last ultrasonic inspection. (iv) When cracks exceed the limits of (a)(2)(i) and Section II Part 2 "Limitations" of the applicable reference PTL the spar booms must be replaced before further flight in accordance with (a)(3). (5) In addition to the ultrasonic inspection requirements of (a) and (b)(1) through (4) accomplish ultrasonic or FAA-approved equivalent inspection of the top and bottom spar boom joints when a replacement outer wing is fitted to the airplane. No ultrasonic inspection is necessary when an outer wing is removed and refitted to the same airplane between the specified ultrasonic inspection intervals, provided that bolt tensioning has previously been accomplished in accordance with the applicable PTL referenced herein and provided that the correct ultrasonic inspection cycles have been maintained. (c) REPROTECTION OF SPAR BOOM JOINTS. Reprotection of wing spar joints at controlled intervals is required and must be accomplished in accordance with the procedure in the applicable referenced PTL. This procedure necessitates the removal of all bolts in each joint to enable a thorough visual inspection for corrosion and cleanliness, both of the bolts and holes. On re-assembly a set of new or reclaimed bolts must be fitted in accordance with Section IV of the subject PTL. After inspection as required by this AD and repair in accordance with the applicable referenced PTL the bolts and joints are to be reprotected and assembled in accordance with the applicable referenced PTL. Following the fitting of the bolts a torque loading check must be accomplished in accordance with Section V of the applicable referenced PTL. Refitment of the bolts must be followed by an inspection of the joint for cracks, using ultrasonic methods or FAA-approved equivalent, prior to further flight. Reprotection of all top and bottom spar boom joints must be carried out on all airplanes when they achieve five years of age, dating from the time of manufacture. On Type 744 airplanes (DTD 363 bottom spar booms) compliance was required by December 31, 1961. On Types 745D and 810 airplanes (L65 bottom spar booms) if this age was achieved before June 30, 1962, compliance was required by that date. Subsequently, repeat the reprotection at intervals of six years. If the spar bolts were retensioned before receipt of Vickers-Armstrongs Cable SS6952 or Issue 4 of the referenced PTL (for 744 and 745D) or Issue 3 of the referenced PTL (for 810) reprotection of all spar joints must be accomplished within six years after the date of the initial retentioning and at subsequent intervals of six years. On airplanes manufactured since January 1961, reprotection of all spar joints is required six years after date of manufacture and repeated at intervals of six years. (d) INTERFERENCE FITTING OF WING SPAR TAPER BOLTS. Refitment of the bolts must be accompanied by ultrasonic or FAA-approved equivalent inspection. Before being considered serviceable for further use, all taper bolts removed from the airplanes must be treated in accordance with the reclaiming procedure Section VII of the applicable reference PTL. The torque loading figures obtained during refitment of the spar taper bolts must be recorded for each airplane since these figures form the datum point for subsequent periodic torque loading inspections. However, where the original torque loading figures have not been recorded it will be necessary to accept the minimum figures quoted in the applicable referenced PTL for datum purposes at the next check inspection, at which time the torque loading figures must be recorded. Suitable charts on which to record the torque loading values are contained in the applicable referenced PTL's. Following the installation and torque loading check of new or reclaimed bolts as detailed in the applicable referenced PTL the inspections detailed below must be accomplished. (e) TORQUE LOADING INSPECTIONS OF SPAR BOOM TAPER BOLTS. (1) OUTER TO INNER WING SPAR BOOM JOINTS - PRIOR TO LINE REAMING THE JOINTS. (i) Visually inspect the bolts for security not less than 100 hours' nor more than 500 hours' time in service after the time of fitment to the joints. (ii) Conduct a torque loading inspection not later than six months after the date of refitment of the bolts to ensure that the torque loading values have not deteriorated to a value less than the figure previously recorded. If the torque values are not less than the recorded figures, accomplish next repetitive torque loading inspection within 18 months after the date of refitment of the taper bolts. (iii) If, on the second repetitive inspection, the torque values are still not less than the recorded figures accomplish next inspection not later than 42 months after the date of refitment of the taper bolts. (iv) Subsequently, provided the torque values are still not less then the recorded figures accomplish the next repetitive inspection within 66 months from the date of refitment of the taper bolts. (v) The cycle of inspections required by this paragraph applies only when the torque values remain at not less than the recorded figures at each check. If the readings of any one or more bolts are less than the recorded figure, accomplish rectification in accordance with Section VI of the applicable PTL referenced herein. (2) OUTER TO INNER SPAR BOOM JOINTS - SUBSEQUENT TO LINE REAMING THE JOINTS. (i) Visually inspect the bolts for security not less than 100 hours' nor more than 500 hours' time in service after the time of fitment to the joints. (ii) Conduct a torque loading inspection not later than six months after the date of refitment of the bolts to ensure that the torque loading values have not deteriorated to a value less than the figures previously recorded. If the torque values are not less than the recorded figures, accomplish next repetitive torqueloading inspection within two years after the date of refitment of the taper bolts. (iii) Subsequently, if torque values remain at not less than the recorded figures accomplish repetitive inspection at intervals of two years. (iv) The cycle of inspections required by this paragraph applies only when the torque values remain at not less than the recorded figures at each check. If the readings of any one or more bolts are less than the recorded figured, accomplish rectification in accordance with Section VI of the applicable referenced PTL. (v) The repetitive torque loading inspections required by (d)(2) (ii) and (iii) also apply to those airplanes that have been subject to Blueing Checks of the taper bolts in the holes and it has been positively established that the contact area between the bolt and bolt holes in the top and bottom outer to inner wing joints is NOT LESS THAN 95 PERCENT INCLUDING THE WEB PLATES. (3) INNER WING TO CENTER SECTION SPAR BOOM JOINTS.(i) Visually inspect the bolts for security not less than 100 hours' nor more than 500 hours' time in service after the time of fitment to the joints. (ii) Conduct a torque loading inspection not later than six months after the date of refitment of the bolts to ensure that the torque loading values have not deteriorated to a value less than the figure previously recorded. If the torque values are not less than the recorded figures, accomplish next repetitive torque loading inspection within two years after the date of refitment of the taper bolts. (iii) Subsequently, if the torque values remain at not less than the recorded figures accomplish repetitive inspections at intervals of two years. (iv) The cycle of inspections required by this paragraph applies only when the torque values remain at not less than the recorded figures at each check. If the reading of any one or more bolts is less than the recorded figure, accomplish rectification in accordance with Section VI of the applicable referenced PTL. (British Aircraft Corporation (Weybridge Division) Preliminary Technical Leaflet No. 230 Issue 8 (700 Series) and corrigendum (for 744 and 745D airplanes), No. 97 Issue 7 (800/810 Series) and corrigendum (for 810 Series airplanes) or later ARB-approved issue cover this subject.) This supersedes AD 61-23-05. This directive effective October 29, 1965.
89-15-09: 89-15-09 BRITISH AEROSPACE PLC: Amendment 39-6263. Applicability: Jetstream Model 3101 (Serial numbers 696 through 794, 796 through 799, 801 through 804, 806 through 809, 811 through 813, 815 through 817, and 820) airplanes certificated in any category. Compliance: Required within the next 75 hours time-in-service (TIS) after the effective date of this AD, unless already accomplished. To prevent fire or damage to the airplane electrical system, accomplish the following: (a) Install a 7.5 ampere circuit breaker in accordance with British Aerospace (BAe) Mandatory Alert Service Bulletin (ASB) Jetstream 24-A-JA7672A, dated November 2, 1988, and BAe ASB Erratum No. 1, dated December 2, 1988, in the electrical power output line of : (1) The main 26 volt a.c. inverter; and (2) The essential 26 volt a.c. inverter. (b) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (c) An equivalent means of compliance with this AD may be used if approved by the Manager, Aircraft Certification Office, AEU-100, Europe, Africa, Middle East Office, FAA, c/o American Embassy, B-1000 Brussels, Belgium. All persons affected by this directive may obtain copies of the documents referred to herein upon request to British Aerospace, Inc., Technical Librarian, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041; or may examine these documents at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment (39-6263, AD 89-15-09) becomes effective on August 14, l989.
99-15-03: This amendment adopts a new airworthiness directive (AD) that applies to certain Stemme GmbH & Co. KG (Stemme) Model S10-VT sailplanes. This AD requires modifying the wastegate control in order to eliminate heat damage. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. The actions specified by this AD are intended to prevent the wastegate control from malfunctioning because of heat damage, which could result in loss of automatic manifold pressure control and engine damage.
2022-14-07: The FAA is adopting a new airworthiness directive (AD) for all Bombardier, Inc., Model BD-700-1A10 and BD-700-1A11 airplanes. This AD was prompted by reports that some oxygen box assemblies had their piston ejected during the mask deployment test. This AD requires a one- time inspection of each passenger oxygen box dual manifold assembly to find and replace affected parts. This AD also prohibits installing affected parts. The FAA is issuing this AD to address the unsafe condition on these products.
99-15-05: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model DC-9-10, -20, -30, -40, and -50 series airplanes, and C-9 (military) airplanes, that requires a one-time visual inspection to determine if all corners of the aft lower cargo doorjamb have been previously modified. This amendment also requires low frequency eddy current inspections to detect cracks of the fuselage skin and doubler at all corners of the aft lower cargo doorjamb, various follow-on repetitive inspections, and modification, if necessary. This amendment is prompted by fatigue cracks found in the fuselage skin and doubler at the corners of the aft lower cargo doorjamb. The actions specified by this AD are intended to detect and correct such fatigue cracking, which could result in rapid decompression of the fuselage and consequent reduced structural integrity of the airplane.
69-19-04: 69-19-04 SLINGSBY: Amdt. 39-841. Applies to Model T.53B Gliders. To ensure that the glider can be trimmed when being aero-towed, within the next 10 hours' time in service after the effective date of this AD, unless already accomplished, check to determine that the tailplane incidence is 1 degree 50' + 20'. If it is incorrect change the incidence of the tailplane to 1 degree 50' + 20' measured from the aircraft datum in accordance with Slingsby Aircraft Company Ltd., Technical Instruction No. 40, dated July 1969 or later ARB issue or an FAA approved equivalent. This amendment becomes effective September 15, 1969.
2022-14-10: The FAA is superseding Airworthiness Directive (AD) 2018-13- 08, which applied to certain Airbus SAS Model A318 series airplanes; Model A319-111, -112, -113, -114, -115, -131, -132, and -133 airplanes; Model A320-211, -212, -214, -216, -231, -232, and -233 airplanes; and Model A321-111, -112, -131, -211, -212, -213, -231, and -232 airplanes. AD 2018-13-08 required repetitive inspections for cracking of the radius of the front spar vertical stringers and the horizontal floor beam on frame (FR) 36, repetitive inspections for cracking of the fastener holes of the front spar vertical stringers on FR 36, and repair if necessary, and, for certain airplanes, a potential terminating action modification of the center wing box area. This AD was prompted by a determination that additional airplanes are subject to the unsafe condition. This AD revises the applicability by adding airplanes and retains the requirements of AD 2018-13-08; 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.
99-14-08: This amendment supersedes an existing airworthiness directive (AD), applicable to Pratt & Whitney (PW) JT9D series turbofan engines, that currently requires initial and repetitive inspections of the sixth stage low pressure turbine (LPT) inner airseal, and modification of the sixth stage LPT inner airseal to reduce the potential for two failure modes. This amendment requires additional repetitive borescope inspections for sixth stage LPT inner airseals found with cracks less than one inch in length. This amendment is prompted by the publication of a revision to a PW service bulletin that introduces the new borescope inspections. The actions specified by this AD are intended to prevent an uncontained failure of the sixth stage LPT inner airseal, which can result in damage to the aircraft. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of September 13, 1999.
90-07-09: 90-07-09 SUD-SERVICE (FORMERLY SUD AVIATION): Amendment 39-6555. Docket No. 89-NM-246-AD. Applicability: Sud-Service Caravelle SE 210 Model III and VIR series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent failure of the left-hand forward passenger door frame and subsequent decompression of the airplane, accomplish the following: A. Prior to the accumulation of 20,000 landings, or within 50 landings after the effective date of this AD, whichever occurs later, perform one of the following inspections in accordance with Sud- Service Service Bulletin 53-56, dated October 21, 1988: 1. Perform a visual inspection of the skin plating, stamping, and visible edge of the fitting, plus a dye penetrant inspection of the fitting edge; or 2. Perform an X-ray inspection in accordance with the following maintenance manuals: a. For Mark III airplanes - Chapter 53-1-1, Figure 606; b. For Mark VIR airplanes - Chapter 53-12-1, Figure 604; or 3. Perform an inspection of holes by defectometer, after removal of fasteners identified on Figure 606 (Mark III) or Figure 604 (Mark VIR), as applicable; or 4. Perform an inspection of holes by rototest, after removal of fasteners identified on Figure 606 (Mark III) or Figure 604 (Mark VIR), as applicable. B. If no cracks are found, repeat the inspections at the following intervals: a. If the immediately preceding inspection was performed by visual method and dye check, the next inspection must be performed within 100 landings. b. If the immediately preceding inspection was performed by X-ray, the next inspection must be performed within 300 landings. c. If the immediately preceding inspection was performed by defectometer, the next inspection must be performed within 2,000 landings. d. If the immediately preceding inspection was performed by rototest, the next inspection must be performed within 4,000 landings. C. If cracks are found, repair prior to further flight, in accordance with Sud-Service Service Bulletin 53-56, dated October 21, 1988. Thereafter, repeat visual inspections of the fasteners around the edge at intervals not to exceed 500 landings, and replace with a new fitting prior to the accumulation of 5,000 landings, in accordance with the service bulletin. D. Upon the installation of a new fitting, perform the initial inspection prior to the accumulation of 20,000 landings, in accordance with paragraph A., above, and thereafter at intervals specified in paragraph B., above. E. 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 will either concur or comment and thensend it to the Manager, Standardization Branch, ANM-113. F. 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 Aerospatiale, 316 Rue de Bayonne, 31060 Toulouse, Cedex 03, France. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6555, AD 90-07-09) becomes effective on May 1, 1990.
92-22-12: 92-22-12 BRITISH AEROSPACE: Amendment 39-8398. Docket No. 92-NM-50-AD. Applicability: Model BAe 146-100A series airplanes, Constructor's Nos. E1002 and subsequent; Model BAe 146-200A series airplanes, Constructor's No. E2008 and subsequent; and Model BAe 146-300A series airplanes, Constructor's Nos. E3118 and subsequent; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent loss of normal braking during ground operations, accomplish the following: (a) For airplanes having pre-Modification HCM00716B configuration: Within 6 months after the effective date of this AD, remove the anti-skid control box, and install a new anti-skid control box, Modification HCM70491A; and perform an operational test on the anti-skid control box; in accordance with British Aerospace BAe 146 Service Bulletin SB.32-124-70491A&B, Revision 1, dated November 25, 1991; or Revision 2, dated June 30, 1992. (b) For airplanes having post-modification HCM00716B configuration: Within 6 months after the effective date of this AD, remove the anti-skid control box, and install a new anti-skid control box, Modification HCM70491B; and perform an anti-skid braking system integrity test; in accordance with British Aerospace BAe 146 Service Bulletin SB.32-124-70491A&B, Revision 1, dated November 25, 1991; or Revision 2, dated June 30, 1992. (c) 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, Standardization Branch, ANM-113, FAA, Transport 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, Standardization Branch, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113.(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) The replacement, operational test, and integrity test shall be done in accordance with the following British Aerospace BAe 146 service bulletins, as applicable, which contain the specified effective pages: Service Bulletin Referenced and Date Page Number Revision Level Shown on Page Date Shown on Page SB.32-124-70491A&B 1, 3, 6 1 November 25, 1991 Revision 1, November 25, 1991 2, 4-5, 7-10 Original (Not dated) SB.32-124-70491A&B 1, 3, 6 2 June 30, 1992 Revision 2, June 30, 1992 2, 4-5, 7-10 Original (Not dated) 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, DC. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (f) This amendment becomes effective on December 3, 1992.