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90-15-04: 90-15-04 BRITISH AEROSPACE: Amendment 39-6652. Docket No. 90-NM-41-AD. Applicability: Model BAC 1-11 200 and 400 series airplanes, pre-modification PM5384, certificated in any category. Compliance: Required within 2,400 hours time-in-service or two years after the effective date of this AD, whichever occurs first, unless previously accomplished within the past 2,400 hours time-in-service or within the past two years; and thereafter at intervals not to exceed 4,800 hours time-in-service or four years, whichever occurs first. To prevent tailplane trim gearbox oil from being contaminated with water, accomplish the following: A. Remove the tailplane trim gearbox from the airplane, drain the oil, flush and refill with clean oil, and replace the filler plug and wire lock, in accordance with paragraph 2.2 of British Aerospace Alert Service Bulletin 27-A-PM5384, Issue 1, dated July 24, 1989. Reinstall the gearbox in the airplane and test in accordance with Maintenance Manual Chapter 27-40. 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, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM-113, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence 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 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. 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 (39-6652, AD 90-15-04) becomes effective on August 14, 1990.
2001-17-26 R1: This document corrects and clarifies information in an existing airworthiness directive (AD) that applies to certain Raytheon Model DH.125, HS.125, BH.125, and BAe.125 (U-125 and C-29A) series airplanes; Model Hawker 800, Hawker 800 (U-125A), Hawker 800XP, and Hawker 1000 airplanes. That AD currently requires an inspection for cracking or corrosion of the cylinder head lugs of the main landing gear actuator and follow-on/corrective actions. This document corrects and clarifies the affected airplane serial numbers. This correction is necessary to ensure that operators do not misinterpret which airplanes are subject to the requirements of this AD. The incorporation by reference of certain publications listed in the regulations was approved previously by the Director of the Federal Register as of October 3, 2001 (66 FR 45575, August 29, 2001).
85-25-03: 85-25-03 SIKORSKY AIRCRAFT: Amendment 39-5172. Applies to Model S-64E helicopters, certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent operation with a cracked main rotor head torque tube inner bracket, accomplish the following: (a) Prior to the first flight of each day, after the effective date of this AD, visually inspect with a 10-power or higher magnifying glass the main rotor head torque tube inner bracket assembly, Part Number S1510-21332-0, for cracks and/or corrosion in accordance with Section 2, Paragraph A, of Sikorsky Alert Service Bulletin (ASB) No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (b) If the torque tube inner bracket assembly is cracked, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (c) If the torque tube inner bracket assembly is corroded, determine the extent and limits of the corrosion prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB-No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. If the extent or limits of the corrosion are exceeded, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. Otherwise, rework the torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph B, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (d) Aircraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the AD can be accomplished. (e) Upon request, an alternative means of compliance with the requirements of this AD which provide an equivalent level of safety maybe used when approved by the Manager, Boston Aircraft Certification Office, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone (617) 273-7112. Sikorsky ASB No. 64B10-4A, dated July 17, 1985, identified and described in this directive is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received this document from the manufacturer may obtain copies upon request to Sikorsky Aircraft Division, United Technologies Corporation, North Main Street, Stratford, Connecticut 06601. These documents may also be examined at the Office of the Regional Counsel, FAA, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment becomes effective December 30, 1985.
91-05-21: 91-05-21 GENERAL ELECTRIC COMPANY: Amendment 39-6900. Docket No. 90-ANE- 28. Applicability: General Electric Company (GE) CF6-80C2A5 and CF6-80C2B6 engines, Serial Numbers (S/N) 690-101 through 690-369, and S/N 695-101 through 695-423; and CF6- 80C2B6F and CF6-80C2D1F engines, S/N 702-101 through 702-470, and S/N 703-101 through 703-136, which do not incorporate the increased shroud cooling design features of paragraph (b) of this AD, installed on, but not limited to, Airbus A300, Boeing 767, and McDonnell Douglas MD-11 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent high pressure turbine (HPT) failure and possible aircraft damage, accomplish the following: (a) Borescope inspect engines in accordance with sections 2.B., 2.C., and 2.D of the Accomplishment Instructions in GE CF6-80C2 Service Bulletin (SB) 72-473, Revision 1, dated September 21, 1990, unless previously accomplished, according to the following schedule based upon cycles since new (CSN) on the effective date of this AD: (1) Inspect within 10 cycles in service (CIS) after the effective date of this AD or prior to accumulating 520 CSN, whichever occurs later, for CF6-80C2A5 and CF6- 80C2B6 engines, S/N 690-101 through 690-369, and S/N 695-101 through 695-350; and CF6- 80B6F engines, S/N 702-101 through 702-315, and S/N 702-317 through 702-321. (2) Inspect within 10 CIS after the effective date of this AD or prior to accumulating 1,250 CSN, whichever occurs later, for CF6-80C2A5 and CF6-80C2B6 engines, S/N 695-351 through 695-423; and CF6-80C2B6F and CF6-80C2D1F engines, S/N 702-316, 702-322 through 702-470, and S/N 703-101 through 703-136. (3) Remove from service or reinspect in accordance with the following: (i) Remove from service prior to further flight, engines with at least one Category 4 shroud. (ii) Remove from service within 25 hours time in service (TIS) since last inspection (SLI), engines with no Category 4 shrouds, but at least one Category 3 shroud. (iii) Borescope reinspect at intervals not to exceed 125 hours TIS SLI, engines with no Category 3 or 4 shrouds, but at least one Category 2 shroud. (iv) Borescope reinspect at intervals not to exceed 300 hours TIS SLI, engines with no Category 2,3, or 4 shrouds, but at least one Category 1 shroud. (v) Borescope reinspect at intervals not to exceed 520 CIS SLI, engines with no Category 1, 2, 3, or 4 shrouds. (b) Replace the HPT stator stage one shroud support assemblies, Part Numbers (P/N) 9381M61G06 and 9381M61G07; the HPT stator support hanger assemblies, P/N 9397M73G05 and 9397M73G06; and the HPT stage one shrouds, P/N 1333M75P05, 1333M75P06, 1333M75P07, 1333M75P08, 1333M75P09, and 1333M75P10 in accordance with the Accomplishment Instructions in GE CF6-80C2 SB 72-474, Revision 1, dated December 11, 1990, at the next HPT module exposure after the effective date of this AD, but prior to December 31, 1994. (c) For the purpose of this AD, HPT module exposure is defined as the separation of the HPT stator support case from the compressor rear frame. (d) For the purpose of this AD, the shroud Categories are defined in GE CF6-80C2 SB 72-473, Revision 1, dated September 21, 1990. (e) 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. (f) 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 times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. The borescope inspections and installation of improved HPT hardware shall be done in accordance with the following documents: Document Page Revision Date CF6-80C2SB 72-473 5, 6, 7, Original 7/3/90 10-23 1, 2, 3, Rev. 1 9/21/90 4, 8, and 9 CF6-80C2 SB 72-474 1-24 Rev. 1 12/11/90 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 General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the Office of the Assistant Chief Counsel, FAA, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-6900, AD 91-05-21) becomes effective on March 27, 1991.
78-07-01: 78-07-01 CESSNA: Amendment 39-3163. Applies to TP206 (Serial Numbers P206-0191 thru P20600647), TU206 (Serial Numbers U206-0487 thru U20603693), T207 (Serial Numbers 20700001 thru 20700378) and T210 (Serial Numbers T210-0001 thru T210-0454 and Serial Numbers 21059200 thru 21061758) series airplanes certificated in all categories. To preclude engine oil pump failure due to contamination by the turbocharger thrust bearing anti-rotation pins and failure of the turbocharger shaft, within the next 25 hours' time-in-service after the effective date of this AD, accomplish the following: A. Check the turbocharger nameplate or aircraft permanent maintenance records to determine if the turbocharger serial number is prefixed by any of the following letter combinations: EF EFR FA FAR FH FHR EG EGR FB FBR FI FIR EH EHR FC FCR FJ FJR EI EIR FD FDR FK FKR EJ EJR FE FER FL FLR EK EKR FF FFR GA GAR EL ELR FG FGR GB GBR B. If the serial number on the turbocharger nameplate is not prefixed by any of the letter combinations set forth in Paragraph A, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. C. If the serial number on the turbocharger nameplate is prefixed by any of the letter combinations set forth in Paragraph A, check the aircraft permanent maintenance records to determine whether, when complying with AD 77-06-02, the turbocharger center housing was replaced by a mechanic or repair agency in accordance with Cessna Service Kit SK 210-75, dated February 24, 1977 (Reference Cessna Service Letter SE 77-3, Supplement #2 dated February 24, 1977) or by the turbocharger manufacturer (AiResearch). D. If the turbocharger center housing was replaced by AiResearch, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. E. If the turbocharger center housing was replaced by a mechanic or repair agency using instructions in Cessna Service Kit SK 210-75, visually inspect the turbocharger for signs of damage and proper compressor wheel attachment in accordance with Cessna Service Kit SK 210-78 dated November 15, 1977, or later revision (Ref. Cessna Service Letter SE 77-42 dated December 2, 1977, or later revisions) for damage which may have resulted from incomplete compressor wheel locknut torquing procedures prescribed in Cessna Service Kit SK 210-75. (1) If visual signs of damage are evident, return the turbocharger to AiResearch in accordance with Cessna Service Kit SK 210-78. (2) If no visual signs of damage are present but the compressor wheel attachment does not meet the criteria set forth in Cessna Service Kit SK 210-78, conduct additional inspections prescribed therein. Units found acceptable as a result of this inspection may be returned to service after reassembly per this kit. Return unacceptable units to AiResearch in accordance with instructions in Cessna Service Kit SK 210-78. (3) If no visual signs of damage are found and the compressor wheel attachment meets the criteria set forth in Cessna Service Kit SK 210-78, reassemble and identify the turbocharger in accordance with Cessna Service Kit SK 210-78. F. If the turbocharger center housing has not been replaced in accordance with AD 77-06-02, replace the turbocharger center housing in accordance with Cessna Service Kit SK 210-75B dated October 27, 1977, or later revision incorporating a compressor wheel seating procedure. (Ref. Cessna Service Letter SE 77-42.) G. Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. H. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment supersedes Amendment 39-2853 (42 FR 15894), AD 77-06-02. This amendment becomes effective April 6, 1978.
2001-26-53: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 2001-26-53, which was sent previously to all known U.S. owners and operators of Eurocopter France (ECF) Model AS350B, B1, B2, B3, BA, D, and AS355E helicopters by individual letters. This AD requires, before further flight, removing certain serial-numbered servocontrols. This AD is prompted by a report of manufacturing defects in a batch of main servocontrol rods. The actions specified by this AD are intended to prevent failure of a main servocontrol in the flight control system and subsequent loss of control of the helicopter.
2017-18-18: We are adopting a new airworthiness directive (AD) for all Airbus Model A350-941 airplanes. This AD requires repetitive on-ground power cycles to reset the internal timer. This AD was prompted by the in-service loss of communication between some avionics systems and the avionics network. We are issuing this AD to address the unsafe condition on these products.
89-25-08: 89-25-08 BEECH: Amendment 39-6410. Applicability: Models 65 (Serial Numbers (S/N) L-1, L-2, L-6, LF-7 through LF-76, and LC-1 through LC-180); 65-80 and 65-A80 (S/N LD-1 through LD-244); 65-A80 (S/N LD-245 through LD-269) when Beech Modification Kit No. 80-4004-1 or -3 is installed; and 65-B80 (all S/N) airplanes certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To detect possible fatigue cracking of the wing main spar lower cap and associated structure, accomplish the following: (a) Within the next 200 hours time-in-service (TIS) after the effective date of this AD, or upon accumulating 3000 hours TIS on Models 65-80 and 65-A80 airplanes, or upon accumulating 5000 hours TIS on Models 65 and 65-B80 airplanes, whichever occurs later, unless previously accomplished per AD 70-25-01, Amendment 39-1609, and thereafter at intervals not to exceed 1000 hours TIS (except as provided in paragraph (b) below) after the initial inspection, inspect the wing lower forward spar attach fittings, center section and outboard wing spar caps adjacent to the attach fittings by visual, fluorescent penetrant and eddy current methods as specified in the applicable section of Beech Structural Inspection and Repair Manual (SIRM), P/N 98-39006, Revision A4, dated May 1, 1987. NOTE 1: Beech offers a two-day training course free of charge to qualified personnel who have prior knowledge of eddy current inspection techniques. A listing of Beech Corporate maintenance facilities may be obtained from the sources contained in paragraph (e) of this AD. A listing of other facilities employing qualified inspectors is not available. (b) At each inspection required by paragraph (a) above, inspect any reinforcing strap installed per Supplemental Type Certificate (STC) SA1583CE for proper tension and condition in accordance with Aviadesign Engineering Order E.O. B-8001, Issue 3, dated May 30,1985. Correct any discrepancy prior to further flight. For airplanes equipped with STC SA1583CE and inspected in accordance with paragraph (a) above, the repetitive inspection interval of 1000 hours TIS in paragraph (a) above may be extended to 3000 hours TIS. (c) If any crack is found in a main spar lower cap or fitting, prior to further flight repair or replace the defective part using the instructions and limitations specified in the SIRM or other FAA approved instructions provided by Beech Aircraft Corporation. (d) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (e) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209; Telephone (316) 946-4400. NOTE 2: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or Western Aircraft Maintenance, 4444 Aeronca Street, Boise, Idaho 83705; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD supersedes AD 70-25-01, Amendment 39-1609. This amendment (39-6410, AD 89-25-08) becomes effective on January 4, 1990.
2002-01-12: This amendment adopts a new airworthiness directive (AD), that is applicable to General Electric Company (GE) GE90 series turbofan engines. This amendment requires removing from service high pressure turbine (HPT) interstage seals, identified by GE as the pre-life-improved rotor (pre-LIR) configuration, and installing a new design, identified by GE as the life improved rotor (LIR) configuration seal. This amendment also requires a new lower life limit for the LIR configuration seal. This amendment is prompted by an uncontained engine failure which occurred during a factory development engine ground test. The actions specified by this AD are intended to prevent failure of the HPT interstage seal that could result in an uncontained engine failure and damage to the airplane.
87-11-05 R4: 87-11-05 BOEING VERTOL COMPANY (VERTOL) AND KAWASAKI HEAVY INDUSTRIES, LTD.: Amendment 39-5591. Applies to Model 107-II helicopters and Kawasaki Model KV107-II and IIA helicopters, certificated in any category. Compliance is required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. (a) To prevent hazards in flight associated with the fatigue failure of life limited components, the following retirement lives are imposed: LIFE LIMIT COMPONENT PART PART NUMBER HOURS APPENDIX Alum Synch 107D3141 9,940 Shaft Assembly 107D3341 9,940 Steel Synch 107D3140 7,850 Shaft Assembly 107D3340 7,850 A15D3840-1 24,750 Synchronizing 107D3154-2 500 No. 1 Shaft Splined 107D3154-3, -4 3,000 Adapter 107D3144-1 14,615 Mix Box 107D2066-10 11,075 Collector Gear thru -28 Fwd Rotor Shaft A02D1259-3 2,100 and Carrier Assy A02D1269-1 2,100A07D1269-1 2,100 A02D1259-3SP 13,580 A02D1269-1SP 13,580 A07D1269-1SP 13,580 A02D1269-2 13,580 Quill Shaft 107D2067-1, 5,620 No. 2 -3, -5 Aft Transmission 107D2419 or 8,700 Planet Carrier A02D2419 Aft Rotor Drive Shaft upper extension A02D3147 7,200 Aft Rotor Shaft 107D3171-2, -3, -4, -5 15,860 Lower Section A2D3171-3, -4, -6 15,860 Aft Rotor Shaft 107D3151 19,140 Center Section A02D3151 19,140 Fwd Hub 107R2550 2,500/1,200 No. 3 Aft Hub 107R2550 10,460/3,000 Fwd and Aft 107R2551-1 or 6,610 Connecting Link A02R2551-8 Fwd and Aft Hub A02R2589-3 25,140 Horizontal Hinge Pin 107R2589-3 25,140 Fwd and Aft Lag Damper Piston 107R2601-1 4,670 Rod End APPENDICES: No. 1. Synchronizing shaft splined adapter, Part Number (P/N) 107D3154-2, shall be repetitively inspected as specified in Boeing Vertol (BV) Service Bulletin 107-116 (R-1), Revision B, dated February 21, 1983, or FAA-approved equivalent. P/N 107D3154-3 and -4 require repetitive visual and dye check inspections as specified in the BV107-6 Maintenance Schedule, Section 2, dated February 7, 1986, and Service Bulletin 107-116 (R-1)B. No. 2. Quill shaft, P/N 107D2067-1, has a life limit of 120 hours unless the modifications specified in paragraph (e) of AD 63-24-04 revised October 15, 1964, Amendment 648 of Part 507 (28 FR 12614) as amended by Amendment 821 of Part 507 (29 FR 14169) are accomplished. (Also, reference Boeing Service Bulletins 107-113, Revision A, dated November 22, 1963, and 107-182, Revision B, dated July 26, 1965.) No. 3. The life limit for the forward rotor hub spline is 2,500 hours. After 2,500 hours of service, the hub is to be magnetic particle inspected. If found to be free of cracks and not having a wear step on the profile face of the spline in excess of 0.002 inch, the hub may be inverted and installed on the aft rotor head for an additional 3,000 hours of use. On completion of the additional 3,000 hours, the hub is to be retired from service. The life limit for the aft rotor hub spline is 10,460 hours. After 10,460 hours of use, the hub is to be magnetic particle inspected. If found to be free of cracks and not having a wear step on the profile face of the spline in excess of 0.002 inch, the hub may be inverted and installed on the forward rotor head for an additional 1,200 hours of use. On completion of the additional 1,200 hours, the hub is to be retired from service. (An acceptable procedure for measuring the 0.002-inch step wear is contained in Boeing Vertol Overhaul Manual 107-5, Chapter 60-20-1.) Inverting and switching of main rotor hubs are permitted to the life limit stated above any number of times provided accurate records are maintained of the total hours the hubs are installed on an aft head and on a forward head. (b) An alternate method of compliance or adjustment of the compliance time, which provides an equivalent level of safety, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581. (c) Aircraft may be ferried to a base in accordance with FAR Sections 21.197 and 21.199 where compliance can be accomplished. The repair and inspection procedures in Boeing Vertol (BV) Service Bulletin (SB) 107- 113, Revision A, dated November 22, 1963; BV SB 107-116 (R-1), Revision B, dated February 21, 1983; BV SB 107-182, Revision B, dated July 26, 1965, and BV 107-6 Maintenance Schedule, Section 2, temporary Revision 31, dated February 7, 1986, incorporated by reference in this directive were approved by the Director of the FEDERAL REGISTER pursuant to 5 U.S.C. 552(a)(1). Copies may be obtained from Boeing Vertol Company, Boeing Center, P.O. Box 16858, Philadelphia, Pennsylvania 19142. These documents may be examined at the Office of the Regional Counsel, FAA, Southwest Region 4400 Blue Mound Road, Fort Worth, Texas 76106 or the Office of the Federal Register, 1100 L Street, NW., Room 8401, Washington, D.C. This amendment supersedes the following: Amendment 656 of Part 507 (28 FR 13931) as amended by Amendment 39-107 (30 FR 8963), AD 63-26-04; Amendment 39-2993 (42 FR 38803), AD 77-16-03; Amendment 39-3746 (45 FR 25050), AD 80-08-11; Amendment 39-4522 (47 FR 57484), AD 82-27-06; and Amendment 39-4757 (48 FR 50069), AD 83-22-02. This amendment, 39-5591, becomes effective May 14, 1987.