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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.
91-03-18: 91-03-18 BRITISH AEROSPACE: Amendment 39-6886. Docket No. 90-NM-198-AD. Applicability: All Model BAe/DH/HS/BH-125 series airplanes, equipped with Aircraft Products Company (APC) warming ovens having part numbers 255B-LH-28/115, 255B-RH-28/115, or 255-362, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent a warming oven fire and to provide additional overheat protection, accomplish the following: A. Within 60 days after the effective date of this AD, accomplish the following: 1. Inspect the wire insulation inside the warming oven in accordance with British Aerospace Service Bulletin 25-54-9758A&B, Revision 1, dated August 14, 1989. If the oven wiring or thermostat wiring is damaged, prior to further flight, replace the wiring and/or thermostat, as appropriate, in accordance with the service bulletin. 2. Modify the warming ovens by installing an additional thermostat and a new switch, in accordance with British Aerospace Service Bulletin 25-54-9758A&B, Revision 1, dated August 14, 1989. NOTE: British Aerospace Service Bulletin 25-54-9758A&B references Aircraft Products Company (APC) Service Bulletins 255B-25-002, Revision A, dated June 22, 1989, and 255-25-003, dated May 26, 1989, for additional instructions. 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, 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, 1601 Lind Avenue S.W., Renton, Washington. This amendment (39-6886, AD 91-03-18) becomes effective on March 11, 1991.
87-19-01: 87-19-01 MESSERSCHMITT-BOLKOW-BLOHM (MBB), BELL HELICOPTER TEXTRON, INC., and AEROSPATIALE: Amendment 39-5786. Applies to all MBB BO- 105, all series, with STC No. SH479GL installed; MBB Model BK-117, all series, with STC No. SH1853SO installed; Bell Model 222, all series, with STC No. SH1879SO installed; and Aerospatiale Model AS-355, all series, with STC No. SH673GL installed. Compliance is required before further flight, unless already accomplished. To prevent the ingestion of the labels (Facet P/N's 1741120 and 1741120-01) by the engines, accomplish the following: (a) Gain access to the engine compartment where the labels are located. (b) Remove the labels (Facet P/N's 1741120 and 1741120-01) from the airframe using methyl ethyl ketone (MEK) and a single-edge razor. (c) Clean area where label was installed. (d) Make appropriate logbook entry showing compliance with this AD. (e) An alternative method of compliance which provides an equivalent level of safety may be used when approved by the Manager, FAA Atlanta Aircraft Certification Office, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia 30349. (f) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and FAR 21.199 to a base where the AD can be accomplished. NOTE: 1. The information on the labels repeats information which is provided in the Rotorcraft Flight Manual Supplement. 2. Facet Service Bulletin No. 081987 dated August 19, 1987, refers to this subject. This Amendment 39-5786 becomes effective December 18, 1987, as to all persons except those persons to whom it was made immediately effective by priority letter AD 87-19-01, issued September 8, 1987, which contained this amendment.
98-19-19: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires a one-time inspection to detect discrepancies of the electrical harness of the propeller de-icing system and of the hydraulic pressure pipe from the engine driven pump (EDP); and follow-on corrective actions, if necessary. This action is prompted by the issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent chafing of the hydraulic pressure pipe of the EDP, which could result in charring of the hydraulic tube and consequent engine compartment fire.
98-19-11: This amendment adopts a new airworthiness directive (AD) that is applicable to Rolls-Royce Limited, Aero Division-Bristol/S.N.E.C.M.A. Olympus 593 series turbojet engines. This action requires initial and repetitive X-ray and ultrasonic inspections of exhaust diffuser vanes for corrosion and cracks, and, if necessary, removal from service of cracked exhaust diffusers and replacement with serviceable parts. This amendment is prompted by reports of 17 turbine exhaust diffuser modules with one or more exhaust diffuser vanes cracked. The actions specified in this AD are intended to prevent exhaust diffuser vane failure, which could result in an adverse effect on the engine oil and reheat systems, possibly causing an inflight engine shutdown or 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 30, 1998. Comments for inclusion in the Rules Docket must bereceived on or before November 16, 1998.
91-06-20: 91-06-20 BRITISH AEROSPACE: Amendment 39-6937. Docket No. 90-NM-241-AD. Applicability: Model ATP series airplanes; Serial Numbers 2001 through 2020, inclusive; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent the malfunction of electrical equipment, overheat damage or fire, and the loss of electrical power, accomplish the following: A. Within 90 days after the effective date of this AD, inspect the electrical cable ring tongue terminal tags listed in Tables 1 to 21 of British Aerospace Service Bulletin ATP-24-21, Revision 2, dated April 24, 1990, for security and discoloration, in accordance with the service bulletin. NOTE: The service bulletin and this AD refer to "ring tongue terminal tags." That terminology is used to describe crimp-on wire terminals that have a loop on one end to fit over a terminal stud. B. If any terminal tag shows signs of being insecurely crimped or is discolored, replace it prior to further flight, in accordance with British Aerospace Service Bulletin ATP-24-21, Revision 2, dated April 24, 1990. C. Within 2,500 hours time-in-service, but not sooner than 2,000 hours time-in-service, following the inspection required by paragraph A. of this AD, perform a one-time repeat visual inspection on all listed terminal tags that have not been replaced or recrimped, in accordance with British Aerospace Service Bulletin ATP-24-21, Revision 2, dated April 24, 1990. 1. If any terminal tag shows signs of being insecurely crimped or is found discolored, replace it prior to further flight, in accordance with the service bulletin. 2. For those terminal tags which do not show signs of being insecurely crimped or discolored, no further action is required. D. 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, StandardizationBranch, 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. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the 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, 1601 Lind Avenue S.W., Renton, Washington. This amendment (39-6937, AD 91-06-20) becomes effective onApril 18, 1991.
2013-03-16: We are superseding an existing airworthiness directive (AD) for the Bell Helicopter Textron (Bell) Model 212 helicopters and adopting requirements for Bell Model 204B, 205A, 205A-1, 205B and 210 helicopters with certain part-numbered main rotor hub inboard strap fittings (fittings). This AD requires magnetic particle inspecting (MPI) the fittings for [[Page 9794]] a crack, and if a crack exists, replacing the fittings with airworthy fittings. This AD is prompted by reports of additional cracked fittings and the determination that additional part-numbered fittings may not have been manufactured in accordance with approved manufacturing processes and controls. These actions are intended to identify a crack in the fitting, which may lead to the fitting's failure, loss of a main rotor blade, and subsequent loss of helicopter control.
98-18-25: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 1000, 2000, 3000, and 4000 series airplanes, that requires replacement of certain hinges on the forward, center, and aft cargo doors with improved hinges. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent failure of the cargo door hinges caused by stress corrosion or fatigue cracks, which could result in decompression of the airplane, and possible in-flight separation of the cargo door.
98-18-12: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Textron Lycoming fuel injected reciprocating engines with certain Crane/Lear Romec "AN" rotary fuel pumps installed. This action requires initial and repetitive torque check inspections of pump relief valve attaching screws. In addition, if the torque remains within acceptable values after two inspections, the repetitive torque check inspections may be terminated. This amendment is prompted by reports of inflight engine fires caused by leaking rotary fuel pumps. The actions specified in this AD are intended to prevent rotary fuel pump leaks, which could result in an engine failure, engine fire, and damage to or loss of the aircraft.
70-03-04 R1: 70-03-04 R1 CESSNA: Amendment 39-933 as amended by Amendment 39-5564. Applies to the following serial numbered Models T310P, T310Q, 320D, 320E and 320F airplanes equipped with Teledyne Continental Model TSIO-520B engines, and to the following serial numbered Models 401, 401A, 401B, 402, 402A and 402B airplanes equipped with Teledyne Continental Model TSIO-520E engines: MODEL SERIAL NUMBER T310P T310P0001 thru T310P0240 T310Q T310Q0001 thru T310Q0291 320D 320D0001 thru 320D0130 320E 320E0001 thru 320E0110 320F 320F0001 thru 320F0045 401 4010001 thru 4010322 401A 401A0001 thru 401A0132 401B 401B0001 thru 401B0121 402 4020001 thru 4020322 402A 402A0001 thru 402A0129 402B 402B0001 thru 402B0122 Compliance: Within 25 hours time-in-service (TIS) after the effective date of this AD on airplanes with turbosupercharger turbine housings having 400 hours or more TIS, or at or before 425 hours TIS on turbosupercharger turbine housings having less than 400 hours TIS and thereafter at intervals not to exceed 100 hours TIS, unless already accomplished. To detect incipient failure of turbosupercharger turbine housings installed in the above airplanes, accomplish the following: (a) Remove the engine top cowling and the turbosupercharger turbine insulation blanket and visually inspect the complete surface of the turbine housing of the TCM turbosupercharger assembly P/N 632729 (AID P/N 406610) for cracks, bulges and burnt areas. Remove and reinstall the turbosupercharger insulation blanket in accordance with applicable Cessna Service Manuals. (b) If cracks, bulges or burnt areas are found during the inspection required by paragraph (a) of this AD, prior to further flight, replace the defective part with an airworthy part. (c) Replacement of the turbosupercharger turbine insulation blanket with stainless steel heat shields in accordance with Cessna Multi-engine Service Letter ME72-4, dated March 24, 1972, will terminate further time interval repetitive inspections required by this AD. However, the inspection cited in paragraph (a) above and any necessary corrective action in paragraph (b) above must be completed at the time of the heat shield installation. (d) Any equivalent method of compliance with this AD, if used, must be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209. All persons affected by this directive may obtain copies of the document(s) referred to herein upon request to Cessna Aircraft Company, Customer Services, P.O. Box 1521, Wichita, Kansas 67201; or may examine the document(s) referred to herein at the FAA, Rules Docket, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. NOTE 1: Cessna Multi-engine Service Letter No. ME70-3, dated January 9, 1970, ME70-3 Supplement I, dated February 9, 1970, and ME72-4, dated March 24, 1972, relate to this subject.NOTE 2: Time-in-service on turbosupercharger turbine housings may be determined from the engine maintenance records. This amendment revises AD 70-03-04, Amendment 39-933. This amendment becomes effective March 28, 1987.
80-15-51: 80-15-51 PRATT & WHITNEY AIRCRAFT: Amendment 39-3898. Applies to JT8D-1, -1A, -1B, -5, -7, -7A, -7B, -9, -9A, -11, -15, -17, -17R, -209, and -209A model turbofan engines. Compliance is required as indicated. To prevent uncontained failure of engine rear compressor front hubs, accomplish the following: Inspect, in accordance with the following schedule, unless already accomplished, all eighth stage rear compressor front hubs, P/Ns 690908, 701308, 717608, 717708, and 738308, which contain balance cuts, for cracks in accordance with the ultrasonic or fluorescent magnetic particle procedures contained in Pratt & Whitney Aircraft Alert Service Bulletin ASB 5154, Revision No. 1, dated July 16, 1980, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region: ENGINE MODEL HUB TOTAL CYCLES AS OF JULY 21, 1980 INSPECTION REQUIRED JT8D-1, -1A, -1B, -5, -7, -7A, and -7B A) Less than 13,000. A) Before accumulation of 15,000 cycles. B) Within the next 2,000 than 15,000. B) 13,000 or more and less cycles. C) 15,000 or more. C) Before accumulation of 17,000 cycles or within the next 500 cycles, whichever is later. JT8D-9, -9A, -11, -15, -209, and -209A A) Less than 10,000. A) Before accumulation of 12,000 cycles. B) 10,000 or more and less than 12,000. B) Within the next 2,000 cycles. C) 12,000 or more and less than 14,000. C) Before accumulation of 14,000 cycles or within the next 500 cycles, whichever is later. D) 14,000 or more and less than 15,000. D) Within the next 500 cycles. E) 15,000 or more and less than 16,000. E) Within the next 300 cycles. F) 16,000 or more. F) Within the next 200 cycles. JT8D-17 and -17R A) Less than 11,000. A) Before accumulation of 12,000 cycles. B) 11,000 or more and less than 12,000. B) Within the next 1,000 cycles. C) 12,000 or more and less than 13,000. C) Before accumulation of 13,000 cycles or within the next 500 cycles, whichever is later. D) 13,000 or more and less than 13,500. D) Within the next 500 cycles. E) 13,500 or more and less than 14,000. E) Before accumulation of 14,000 cycles or within the next 100 cycles, whichever is later. F) 14,000 or more. F) Within the next 100 cycles, unless already accomplished. All models Inspect thereafter at intervals not to exceed 6,000 cycles in service since last inspection. (NOTE: Established life limits are not to be exceeded.) Remove from service, before further flight, any rear compressor front hub having a crack indication. Report defects found to Chief, Engineering and Manufacturing Branch, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, within 5 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 04-R0174.) Airplanes may be ferried in accordance with the provisions of FAR 21.197 to a base where the AD can be accomplished. Upon request of the operator an equivalent means of compliance with the requirements of this AD may be approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to Pratt & Whitney Aircraft, East Hartford, Connecticut 06108. These documents may also be examined at FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD is maintained by the FAA at its headquarters in Washington, D.C., and at the FAA, New England Region Headquarters, Burlington, Massachusetts. This AD supersedes AD T80-14-51 issued July 3, 1980. This Amendment 39-3898 becomes effective August 21, 1980, as to all persons except those persons to whom it was made immediately effective by the telegram of July 16, 1980.
2022-08-14: The FAA is adopting a new airworthiness directive (AD) for all The Boeing Company Model 747-8F series airplanes. This AD was prompted by reports of fuselage crown stringer cracking between station (STA) 740 and STA 1000, stringer (S)-7 to S-12. This AD requires repetitive detailed inspections for cracking of fuselage crown stringers and applicable on-condition actions. The FAA is issuing this AD to address the unsafe condition on these products.
2010-12-04: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: This Airworthiness Directive (AD) is prompted due to the discovery of corrosion at the bonding strap connections on the left and right lower longerons between fuselage frames 1 and 1A. The possibility of corrosion is increased because of the high electrical current flow between the tinned copper terminal lug of the bonding strap and the aluminum longeron. Such a condition, if left uncorrected, could lead to failure of the longeron and will prejudice the structural integrity of the aircraft. We are issuing this AD to require actions to correct the unsafe condition on these products.
98-19-10: This amendment adopts a new airworthiness directive (AD) that is applicable to CFM International (CFMI) CFM56-3, -3B, and -3C series turbofan engines. This action requires, on aircraft with two affected engines installed, removal of one affected engine from an aircraft, and replacement with a serviceable engine, or replacement of a suspect accessory gearbox (AGB) starter gearshaft with a serviceable gearshaft within 350 hours time in service (TIS) after the effective date of this AD, or by September 1, 1998, whichever occurs first. This action also requires, on aircraft with only one affected engine installed, removal of the affected engine from the aircraft, and replacement with a serviceable engine, or replacement of the suspect starter gearshaft with a serviceable gearshaft within 2,100 hours TIS after the effective date of this AD, or by February 1, 1999, whichever occurs first. This amendment is prompted by reports of two inflight engine shutdowns caused by an AGB starter gearshaft failure. The actions specified in this AD are intended to prevent an AGB starter gearshaft failure, which can result in an inflight engine shutdown, and on aircraft with two affected engines installed, possible dual inflight engine shutdown and forced landing.
2022-08-12: The FAA is superseding Airworthiness Directive (AD) 2020-21- 17, which applied to all The Boeing Company Model 757 airplanes. AD 2020-21-17 required repetitive inspections for skin cracking and shim migration at the upper link drag fittings, diagonal brace cracking, and fastener looseness; and applicable on-condition actions. This AD was prompted by reports of bolt rotation in the engine drag fitting joint and fastener heads and cracks found in the skin of the fastener holes, and the need to reduce the compliance time for certain groups. This AD retains the requirements of AD 2020-21-17 with reduced compliance times for certain airplane groups. The FAA is issuing this AD to address the unsafe condition on these products.
98-19-05: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 757-200 series airplanes, that requires the application of a sealant, secondary fuel barrier, and corrosion-inhibiting compound to certain portions of the wing center section. This amendment is prompted by reports indicating that, during manufacture, the secondary fuel barrier was not applied to certain portions of the wing center section. The actions specified by this AD are intended to prevent leakage of fuel through the fasteners, sealant, or structural cracks in the center section structure, which could result in fuel or fuel vapors entering the cargo or passenger compartment of the airplane.
98-19-01: This amendment adopts a new airworthiness directive (AD) that applies to certain Stemme GmbH & Co. KG (Stemme) Model S10 sailplanes. This AD requires replacing the O-ring that is installed in the mounting part of the pitot tube (in the propeller dome) with one of improved design. 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 failure of the pitot tube O-ring caused by an ineffective design, which could result in the pitot tube falling out and the sailplane pilot losing airspeed indications.
92-11-03: 92-11-03 BRITISH AEROSPACE: Amendment 39-8253. Docket No. 91-NM-280-AD. Applicability: British Aerospace Model DH/BH/HS 125 series airplanes, excluding Models 125-600A, 700A, 800A, and 1000A series airplanes; as listed in British Aerospace Service Bulletin S.B. 57-75, dated July 30, 1991; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the wings, accomplish the following: (a) Within 3 months after the effective date of this AD, and thereafter at intervals not to exceed 4 years or 2,200 flights, whichever occurs first, perform an eddy current inspection on specified areas of the left and right wing upper skins to detect cracks in countersunk bolt holes in the wing skins and in the internal stringers, in accordance with British Aerospace Service Bulletin S.B. 57-75, dated July 30, 1991. (b) If cracks are discovered as a result of the eddy current inspection requiredby paragraph (a) of this AD, prior to further flight, perform a dye penetrant inspection, in accordance with British Aerospace Service Bulletin S.B. 57-75, dated July 30, 1991. (c) If cracks are discovered as a result of either the eddy current inspections required by paragraph (a) of this AD, or the dye penetrant inspection required by paragraph (b) of this AD, prior to further flight, repair the crack(s) as follows: (1) Cracks that do not exceed the limits specified in British Aerospace Service Bulletin S.B. 57-75, dated July 30, 1991, must be repaired in accordance with the procedures in the Service Bulletin. (2) Cracks that exceed the limits specified in British Aerospace Service Bulletin S.B. 57-75, dated July 30, 1991, must be repaired in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (d) An alternative method 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. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. (e) 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. (f) The inspections and repairs shall be done in accordance with British Aerospace Service Bulletin S.B. 57-75, dated July 30, 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 British Aerospace, PLC. Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington D.C. 20041-0414. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. (g) This amendment becomes effective on July 9, 1992.
2013-03-08: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-1A11 (CL-600), CL-600-2A12 (CL-601), and CL-600-2B16 (CL-601-3A, CL-601-3R, & CL-604 Variants) airplanes. This AD was prompted by reports of cracking found on the upper and lower Web of the engine support beam. This AD requires revising the maintenance program. We are issuing this AD to detect and correct fatigue cracking of the engine support beam, which could result in failure of the engine support beam and affect the structural integrity of the airplane.
98-19-02: This amendment adopts a new airworthiness directive (AD) that is applicable to Superior Air Parts, Inc., piston pins installed on Teledyne Continental Motors reciprocating engines. This amendment requires removal from service of defective piston pins, and replacement with serviceable parts. This amendment is prompted by reports of numerous piston pin fractures. The actions specified by this AD are intended to prevent a piston pin failure from causing secondary engine damage resulting in loss of oil or total power failure, and from causing jamming of the engine crankshaft resulting in a catastrophic engine failure.
74-20-09: 74-20-09 SIKORSKY AIRCRAFT: Amendment 39-1972 as amended by Amendment 39-2034 is further amended by Amendment 39-2829. Applies to all Sikorsky S-61A, S-61L, S- 61N, S-61R and S-61NM helicopters certificated in all categories that are equipped with S6115- 20501 series, S6115-20601 series, S6117-20101 series and S6188-15001 series main rotor blades. Compliance required as indicated. To prevent operation with loose or missing screws, or with cracks in the outboard end of the main rotor blades, including the tip cap and tip cap bracket, accomplish the following: (a) Inspect main rotor blades that do not have tip cap bracket doubler P/N 61070- 15120 installed, and have less than 2,000 hours time in service on the effective date of this AD, for loose or missing screws or for cracks at the outboard end of the main rotor blade spar, tip cap, or tip cap bracket, in accordance with Section 2., paragraphs A and B., of Sikorsky Service Bulletin No. 61B15-9F, dated November 2, 1976 orlater FAA approved revisions within the next 30 hours time in service after the effective date of this AD, unless already accomplished, and at intervals thereafter within 30 hours time in service from the last inspection. If screws are missing or loose, and cannot be secured, or if there is motion in the joint between the tip cap and the blade, or if cracks are found, replace the main rotor blade or repair it in accordance with Section 2., paragraph C., of the above service bulletin prior to further flight. If a crack is found in the tip cap, tip cap attachment land, or tip cap bracket, replace or repair the main rotor blade in accordance with Section 2., paragraph B., of the above service bulletin prior to further flight. (b) Inspect main rotor blades that do not have tip cap bracket doubler P/N 61070- 15120 installed, and have 2000 hours or more time in service on the effective date of this AD, for loose or missing screws or for cracks at the outboard end of the main rotor blade spar, tip cap, or tip cap bracket, in accordance with Section 2., paragraphs A and B, of the above Service Bulletin prior to the first flight of each day. If screws are missing or loose, and cannot be secured, or if there is motion in the joint between the tip cap and the blade, or if cracks are found, replace the main rotor blade or repair it in accordance with paragraph 2C of the above bulletin prior to further flight. If a crack is found in the tip cap, tip cap attachment land, or tip cap bracket replace or repair the main rotor blade, in accordance with Section 2., paragraph B., of the above service bulletin, prior to further flight. (c) Inspect main rotor blades that have tip cap bracket P/N 61070-15120 installed for cracks at the outboard end of the main rotor blade spar, tip cap, and tip cap bracket, in accordance with Section 2, paragraph B., of the above bulletin, within the next 50 hours time in service after the effective date of this AD, unless already accomplished,and at intervals thereafter within 50 hours time in service from the last inspection. If a crack is found in the tip cap, tip cap attachment land, or tip cap bracket, replace or repair the main rotor blade prior to further flight, in accordance with Section 2., paragraph B., of the above bulletin. (d) For helicopters operating at 19,500 pounds gross weight and below, inspect the outboard end of the main blades, series S6115-20501, series S6115-20601, series S6117-20101, and series S6188-15001, using the dye penetrant method, in accordance with Section 2., paragraph D., of the above Service Bulletin, within the next 200 hours time in service after the effective date of this AD, unless already accomplished, and at intervals thereafter within 200 hours time in service from the last inspection. If a crack is found, replace the main rotor blade prior to further flight. (e) For helicopters operating above a gross weight of 19,500 pounds, inspect the outboard end of the main rotorblades using the dye penetrant method, in accordance with Section 2., paragraph D., of the above Service Bulletin, within the next 50 hours time in service after the effective date of this AD, unless already accomplished, and at intervals thereafter within 50 hours time in service from the last inspection. If a crack is found, replace the main rotor blade prior to further flight. (f) Upon request of the operator, equivalent methods of compliance with the inspection and repair requirements of this AD may be approved by the Chief, Engineering and Manufacturing Branch, New England Region. Repetitive inspection intervals specified in this AD may be adjusted, by the Chief, Engineering and Manufacturing Branch, New England Region, to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. Amendment 39-1972 supersedes Amendment 39-1583 (38 F.R. 1501), AD 73-02-02. Amendment 39-1972 became effective October 4, 1974. Amendment 39-2034 became effective December 12, 1974. This Amendment 39-2829 becomes effective February 25, 1977.