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98-05-14 R1: This amendment clarifies information contained in Airworthiness Directive (AD) 98-05-14, which currently requires revising the FAA-approved Airplane Flight Manual (AFM) to specify procedures that would prohibit flight in severe icing conditions (as determined by certain visual cues), limit or prohibit the use of various flight control devices while in severe icing conditions, and provide the flight crew with recognition cues for, and procedures for exiting from, severe icing conditions on certain Cessna Aircraft Company (Cessna) Models T210N, P210N, and P210R airplanes. That publication incorrectly references the possibility of certain ice accumulation on the "lower" surface of the wing, instead of the "upper" surface of the wing while operating with the flaps extended. This incorrect statement may result in pilot misinterpretation of the icing effects with the flaps extended, and lead to an incorrect action. This document replaces the word "lower" with "upper" in this sentence. The actions specified in this AD are intended to continue to minimize the potential hazards associated with operating these airplanes in severe icing conditions by providing more clearly defined procedures and limitations associated with such conditions.
2013-04-01: We are superseding an existing airworthiness directive (AD) for all Rolls-Royce plc (RR) RB211-524 series turbofan engines. That AD currently requires removal and repair of certain thrust reverser units (TRUs) prior to reinstallation. This AD requires the same actions for an expanded population of TRUs and extends the compliance time for repairing certain TRUs. This AD was prompted by additional engineering evaluation of TRUs, as a result of a translating cowl gearbox stubshaft failure and subsequent repair. We are issuing this AD to prevent failure of the attachment rivets, which may result in release of the TRU from the engine.
98-02-04: This document makes a correction to Airworthiness Directive (AD) 98-02-04 applicable to CFM International (CFMI) CFM56-5B/2P series turbofan engines that was published in the Federal Register on January 21, 1998 (63 FR 3031). The low pressure turbine (LPT) case part number (P/N) in the compliance section is incorrect. This document corrects that P/N. In all other respects, the original document remains the same.
91-20-02: 91-20-02 GENERAL ELECTRIC COMPANY: Amendment 39-8036. Docket Number 91-ANE-04. Applicability: General Electric Company (GE) CF6-80C2 series turbofan engines, installed on, but not limited to, Airbus A300 and A310, Boeing 747 and 767, and McDonnell Douglas MD-11 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent engine fire, accomplish the following: (a) Visually inspect left-hand fuel manifolds, Part Numbers (P/N) 130 3M31G06, 1303M31G07, and 1303M31G08, and right-hand fuel manifolds, P/N 1303M32G06, 1303M32G07, and 1303M32G08, in accordance with the Accomplishment Instructions contained in GE CF6-80C2 Service Bulletin (SB) 73-115, Revision 1, dated February 5, 1991, as follows: (1) Inspect the engine drain mast for fuel leakage every day of operation, after the effective date of this AD. (2) Inspect the circumferential fuel supply manifold at the next scheduled core cowl opening, after the effective date of thisAD; and thereafter, reinspect at intervals not to exceed 500 hours time-in-service since last inspection. (3) Remove from service prior to further flight, fuel manifolds that exhibit leakage, and replace with serviceable parts. (b) Replace left-hand fuel manifolds, P/N 1303M31G06, 1303M31G07, and 1303M31G08, with left- hand fuel manifold, P/N 1303M31G10, and replace right-hand fuel manifolds, P/N 1303M32G06, 1303M32G07, and 1303M32G08, with right-hand fuel manifold, P/N 1303M32G10, in accordance with the Accomplishment Instructions of GE SB 73-114, Revision 2, dated February 4, 1991, at next engine removal, after the effective date of this AD, but no later than June 30, 1993. (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 Inspector (maintenance, avionics, or operations, as appropriate), 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. (e) The inspection shall be done in accordance with the following General Electric documents: DOCUMENT NO. PAGE NO. ISSUE/REVISION DATE GE CF6-80C2 3 and 4 Original 12/4/90 SB 73-114 2 Revision 1 12/6/90 1, 5-31 Revision 2 2/4/91 Total Pages: 31 GE CF6-80C2 2-7 Original 12/5/90 SB 73-115 1 Revision 1 2/5/91 Total Pages: 7 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 beinspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 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 (39-8036, AD 91-20-02) becomes effective on November 25, 1991.
86-23-07: 86-23-07 BRITISH AEROSPACE: Amendment 39-5464. Applies to Model B.121 Series I, II, and III (all serial numbers) airplanes certificated in any category. Compliance: Required initially within 50 hours time-in-service (TIS) for airplanes having or upon accumulating 2,450 hours or more TIS, and thereafter at intervals of 50 hours TIS, unless already accomplished. To assure the integrity of the vertical fin/fuselage attachment structure, accomplish the following: (a) Visually inspect for cracks in the following areas: (1) Center Angle, Part Number (P/N) BE-10-10085 in accordance with paragraph 3. "ACTION" subparagraph (c) of British Aerospace (BAe) Service Bulletin (S/B) No. B121/86, dated March 29, 1984. (i) If cracks are found that equal or exceed the conditions shown in paragraph 3. "ACTION" subparagraph (c) of BAe S/B No. B121/86, prior to further flight, repair in accordance with repair instructions obtained from the manufacturer, British Aerospace, and approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium (hereinafter referred to as "Manager, AEU-100"). (ii) If no cracks are found or if cracks do not exceed the limits shown in paragraph 3. "ACTION" subparagraph (d) of BAe S/B No. B121/86, repeat the inspection at intervals not exceeding 50 hours TIS. (2) The underside of Diaphragm Decking upper, P/N BE-10-10155/1, in accordance with paragraph 3. "ACTION" subparagraph (d) of BAe S/B No. B121/86, dated March 29, 1984. (i) If cracks are found, prior to further flight, repair in accordance with repair instructions obtained from the manufacturer, British Aerospace, and approved by the Manager, AEU-100. (ii) If no cracks are found, repeat the inspection at intervals not exceeding 50 hours TIS. (3) The heel of the side skin attachment flange (left and right) adjacent to the tailplane front spar attachment bolts in accordance with paragraph 3. "ACTION" subparagraph (e) of BAe S/B No. B121/86, dated March 29, 1984. (i) If cracks are found, prior to further flight, repair in accordance with repair instructions obtained from the manufacturer, British Aerospace, and approved by the Manager, AEU-100. (ii) If no cracks are found, repeat the inspection at intervals not exceeding 50 hours TIS. (b) Extension or elimination of the repetitive inspections specified in this AD may be included as part of the FAA-approved repair obtained in accordance with paragraphs (a)(1)(i), (a)(2)(i), and (a)(3)(i) of this AD. (c) Aircraft may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished. (d) An equivalent means of compliance with this AD may be used if approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. All persons affected by this directive may obtain a copy of the document referred to herein upon request to British Aerospace, Engineering Department, Post Office Box 17414, Dulles International Airport, Washington, D.C. 20041; Telephone (703) 435-9100, or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on December 18, 1986.
93-16-04: 93-16-04 SHORT BROTHERS, PLC: Amendment 39-8661. Docket 93-NM-18-AD. Applicability: Model SD3-SHERPA series airplanes; serial numbers SH3201 through SH3205 inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent overheating and burning of the earth wires and a resultant fire, accomplish the following: (a) Within 6 months after the effective date of this AD, replace the existing 24-gauge earth wires on electrical panels 29C and 30C on the bottom of frame 74 with 16-gauge earth wires, in accordance with Shorts Service Bulletin SD3 SHERPA-24-1, dated May 1992. (b) 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. (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) The replacement shall be done in accordance with Shorts Service Bulletin SD3 SHERPA- 24-1, dated May 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 Short Brothers, PLC, 2011 Crystal Drive, Suite 713, Arlington, Virginia 22202-3719. 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. (e) This amendment becomes effective on October 4, 1993.
98-19-12: This amendment adopts a new airworthiness directive (AD) that is applicable to Rolls-Royce, plc RB211 Trent 700 series turbofan engines. This action requires repositioning of the oil metering jet up into the oil distributor within the bevel gearshaft, followed by repetitive inspections of the Magnetic Chip Detector (MCD). Evidence of driving bevel gearshaft ball bearing failure requires replacement of the Step Aside Gearbox (SAGB). This amendment is prompted by reports of uncommanded engine rundowns caused by failure of the SAGB driving bevel gearshaft ball bearing due to oil starvation. This causes a loss of drive to the external gearbox and accessories, resulting in an inflight engine shutdown. The actions specified in this AD are intended to prevent inflight engine shutdowns caused by SAGB driving bevel gearshaft ball bearing failure. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Registeras of October 1, 1998. Comments for inclusion in the Rules Docket must be received on or before November 16, 1998.
98-14-51: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) T98-14-51 that was sent previously to all known U.S. owners and operators of CFM International CFM56-7B series turbofan engines by individual telegrams. This AD requires checks of the Accessory Gearbox (AGB)/Transfer Gearbox (TGB) Magnetic Chip Detector (MCD) for abnormal magnetic particles that indicate a pending starter gearshaft failure, and, removal from service of suspect starter gearshafts and replacement with serviceable parts. This amendment is prompted by reports of 2 inflight engine shutdowns due to uncontained failures of the AGB starter gearshafts. The actions specified by this AD are intended to prevent a dual inflight engine shutdown event, which could result in a forced landing and loss of the aircraft.
79-12-06: 79-12-06 CESSNA: Amendment 39-3492 as amended by Amendment 39-3757. Applies to Model 500 (Serial Numbers 500-0001 through 500-0349), Models 500 and 501 (Unit Numbers -0350 through -0520) and Models 550 and 551 (Unit Numbers -0002 through -0114) airplanes with 600 or more hours time-in-service, except those which have been modified in accordance with one of the following Cessna Citation Service Bulletins: SB57-10, Revision 1, dated March 28, 1980; SB57-11, Revision 1, dated March 28, 1980; or SB550-57-3, Revision 1, dated March 28, 1980, as applicable. COMPLIANCE: Required as indicated unless already accomplished. A) Within the next 100 hours time-in-service after the effective date of this AD and thereafter within each 600 hours time-in-service until the number of landings or time-in-service threshold for dye penetrant inspection in Table I is reached, visually inspect the upper and lower spar cap stems at Wing Station 37.0 in the areas specified in Cessna Citation Service Letter SL 57-2, Revision 2, dated May 1, 1979, or SL 550-57-1, Revision 1, dated May 1, 1979, as applicable for cracks and/or: B) Upon reaching or if at or beyond the dye penetrant inspection threshold specified in Table I, within 100 hours time-in-service or 100 landings, whichever occurs first, and thereafter within 600 hours time-in-service or 600 landings, whichever occurs first, dye penetrant inspect the upper and lower spar cap stem at Wing Station 37.0 in the areas specified in Cessna Citation Service Letter SL 57-2, Revision 2, dated May 1, 1979, or SL 550-57-1, Revision 1, dated May 1, 1979, as applicable, for cracks. TABLE I MODEL SERIAL/UNITS NOS DYE PENETRANT INSPECTION THRESHOLD (LANDINGS OR TIME-IN-SERVICE-HOURS) 500 S/N 500-0001 thru 500-0349 4300 500 & 501 Unit -0350 thru -0520 1300 550 & 551 Unit -0002 thru -0114 4300 C) If, as a result of any inspection required by Paragraph A) or B) a crack of less than .3 inches in length is found in the upper spar cap stem, reduce the dye penetrant inspection interval on this spar cap to 200 landings or 200 hours time-in-service, whichever occurs first. D) If, as a result of any inspection required by Paragraph A) or B) a crack of less than .3 inches in length is found in the lower spar cap stem, reduce the dye penetrant inspection interval on this spar cap to 100 landings or 100 hours time-in-service, whichever occurs first. NOTE: The FAA encourages compliance with the manufacturer's request that all cracks be reported to the Cessna Customer Service Department. E) If, as a result of any inspection required by Paragraph A), B), C) or D), a crack of .3 inches in length or longer is found in either spar cap, prior to further flight, repair or modify the wing in accordance with instructions obtained from Cessna Aircraft Company, Jet Marketing Division, Customer Service, P.O. Box 7706, Wichita, Kansas 67277. F) The time-in-service for initial andbetween repetitive inspections required herein may be adjusted up to 10 hours to facilitate accomplishing them concurrent with other scheduled maintenance on the airplane. G) Aircraft may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished. H) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing District Office, Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. Cessna Citation Service Letters SL57-2, Revision 2, dated May 1, 1979; and SL550-57-1, Revision 1, dated May 1, 1979; and Cessna Citation Service Bulletins SB57-10, Revision 1, dated March 28, 1980; SB57-11, Revision 1, dated March 28, 1980; and SB550-57-3, Revision 1, dated March 28, 1980, cover the subject matter of this AD. Amendment 39-3492 became effective June 21, 1979. This Amendment 39-3757 becomes effective April10, 1980.
91-15-16: 91-15-16 GENERAL ELECTRIC COMPANY: Amendment 39-7080, Docket No. 91-ANE- 01. Applicability: General Electric Company (GE) CF6-45/-50 series turbofan engines, installed on, but not limited to, Airbus A300, Boeing 747, and McDonnell Douglas DC-10-15 and DC-10-30 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent uncontained engine failure, accomplish the following: (a) Eddy current inspect affected high pressure turbine (HPT) thermal shields, Part Numbers (P/N's) 9045M31P04, 9045M31P05, 9045M31P07, 9045M31P08, 9045M31P09, 9045M31P10, 9045M31P12, 9045M31P13, 9143M71P01, 9143M71P02, 9155M16P01, 9155M16P02, 9155M16P03, 9155M16P04, 9181M64P01, 9181M64P02, 9181M64P07, 9181M64P08, 9181M64P10, 9186M96P02, and 9186M96P03, in accordance with the Accomplishment Instructions contained in GE CF6-50/-45 Service Bulletin (SB) 72-879, Revision 6, dated October 30, 1990, as follows: (1) Inspect prior to accumulating 800 cycles since lastHPT overhaul or 400 cycles in service after the effective date of this AD, whichever occurs later. (2) Thereafter, reinspect at intervals not to exceed 400 cycles since last inspection. (3) Remove cracked HPT thermal shields from service prior to further flight and replace with a serviceable part. (b) Affected HPT thermal shields stated in paragraph (a) of this AD, would also be assembled into one of the following HPT thermal shield assembly P/N's: 9045M53G04, 9045M53G05, 9045M53G07, 9045M53G08, 9045M53G09, 9045M53G10, 9045M53G12, 9045M53G13, 9045M53G14, 9186M78G01, 9186M78G02, 9186M78G03, 9186M78G04, 9186M78G05, 9186M78G06, 9186M78G07, 9186M78G08, 9208M76G02, 9208M76G03, and 9208M76G04. (c) For the purpose of this AD, an HPT overhaul is defined as the induction of the engine into a shop where the subsequent maintenance entails HPT disassembly. (d) The eddy current inspection requirements of paragraph (a) of this AD are not applicable to engines incorporating an affected P/N thermal shield that has operated exclusively with an interstage seal, P/N 9315M16G14, 9315M16G15, or 9315M16G17, provided the owner or operator submits to their Airworthiness Inspector the configuration documentation substantiating that the affected thermal shield has never been operated with a P/N 9045M23G07, 9045M23G08, 9045M23G09, 9045M23G10, 9045M23G11, or 9045M23G12 interstage seal. (e) Prior to further flight, remove from service HPT stage one disks which have operated in engines containing an HPT thermal shield cracked through its forward flange. These removed HPT stage one disks may not be returned to service. (f) Prior to further flight, remove from service HPT stage two disks which have operated in engines containing an HPT thermal shield cracked through its rear flange. HPT stage two disks may be returned to service if no cracks are detected when inspected in accordance with Appendix I. (g) 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. (h) Upon submission of substantiating data by an owner or operator through an FAA Inspector (maintenance, avionics, or operations, as appropriate), 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. (i) The eddy current inspections shall be done in accordance with the following GE CF6-50/-45 SB 72-879: PAGE NO. ISSUE/REVISION DATE 3-14, 17-26, 29-31 Rev. 5 1/6/88 1, 2, 16, 28 Rev. 6 10/30/90 15, 27 Orig. 1/24/86 Total Pages: 31 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 FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, D.C. APPENDIX I 1. Reference: CF6-50 Shop Manual Document No. GEK-50481. 2. Accomplishment Instructions: A. Clean, etch, and fluorescent penetrant inspect (FPI) the high pressure turbine rotor (HPTR) stage two disk according to Chapter 72-53-04, High Pressure Turbine Rotor Stage 2 Disks - Inspection, paragraph 2, Fluorescent - Penetrant Inspect Disk, of the reference shop manual. NOTE: The immersion ultrasonic inspection (Subtask 72-53-04-270-051) may be used in lieu of an FPI for the stage 2 disk dovetail serrations only. B. Clean, etch, and eddy current inspect (ECI) the HPTR stage two disk dovetail slot bottoms according to Chapter 72-53-04, High Pressure Turbine Rotor Stage 2 Disk- Inspection, paragraph 5, Special Inspection of Dovetail Slot Bottoms, of the reference shop manual. NOTE: If ECI capability is not available, the disk must be recleaned in accordance with paragraph 2.A (starting with Subtask 72-53-04-140-051). The slot bottoms must then be fluorescent penetrant inspected in accordance with paragraph 2.D (Subtask 72-53-04-230-001- 057) paying special attention to slot bottom corners. This amendment (39-7080, AD 91-15-16), becomes effective on September 4, 1991.