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2017-05-08: We are adopting a new airworthiness directive (AD) for all Safran Helicopter Engines, S.A. Arriel 2B turboshaft engines. This AD requires removing any pre-modification (mod) TU 158 hydro-mechanical metering unit (HMU) and replacing with a part eligible for installation. This AD was prompted by a report of an uncommanded in- flight shutdown (IFSD) on a single-engine helicopter, caused by a low returning spring rate of the needle of the HMU. We are issuing this AD to correct the unsafe condition on these products.
78-02-02: 78-02-02 HUGHES HELICOPTERS: Amendment 39-3127. Applies to Model 269 Series helicopters, certificated in all categories including Model TH-55A. Compliance is required as indicated, unless already accomplished. To prevent failure of the tail rotor control pedals, accomplish the following: (a) Within the next 100 hours time in service from the effective date of this AD, unless already accomplished, and thereafter at intervals not to exceed 100 hours time in service perform the following: (1) On all models visually inspect the pilots' pedal arms P/N 269A7336 for cracks and corrosion in accordance with service information notice (SIN) N-121.1, Part I, Paragraph (b). (i) If visual inspection reveals evidence of incipient cracks, further inspect with dye penetrant and if confirmed replace with a serviceable pedal arm P/N 269A7336 before further flight. (ii) If corrosion is found, remove the corrosion in accordance with SIN N-121.1 before further flight.(iii) After corrosion removal of (1)(ii) above, inspect the wall thickness of the pedal arm in the areas of corrosion removal. If the wall thickness in the cylindrical section is less than .10 inches replace the pedal arm P/N 269A7336 with a serviceable pedal arm before further flight. If the corrosion removal on the pedal arm above the cylindrical section exceeds .005 inches in depth replace the pedal arm P/N 269A7336 with a serviceable pedal arm before further flight. (2) On Models 269A, A-1, and TH-55A, visually inspect the copilots' pedal arms P/N 269A7336 in accordance with paragraph (a)(1) above. (3) On Models 269B and C, remove the copilots' pedal arms P/N 269A7330 from the pedal sockets P/N 269A9973 or P/N 269A7334 and visually inspect for cracks and corrosion in accordance with SIN N-121.1, Part I, Paragraph (d). (i) If visual inspection reveals evidence of incipient cracks, further inspect with dye penetrant in accordance with SIN N-121.1 and ifconfirmed replace with a serviceable part or parts before further flight. (ii) If corrosion is found, remove the corrosion in accordance with SIN N121.1 before further flight. (iii) After corrosion removal of (3)(ii) above, inspect the wall thickness in the areas of corrosion removal. If the wall thickness of the pedal arm is less than .13 inches replace the pedal arm P/N 269A7330 before further flight. If the wall thickness of the socket is less than .10 inches replace the socket P/N 269A9973 or P/N 269A7334 before further flight. (b) Within the next 100 hours time in service, from the effective date of this AD unless already accomplished, and thereafter at intervals not to exceed 100 hours time in service, torque the pedal arm and/or socket bushing nuts to the limits specified by SIN N-121.1. (c) On the 269B and 269C helicopters only, within the next 100 hours time in service, from the effective date of this AD unless already accomplished, measure the copilots' pedal arm wall thickness above the quick release pin hole. If the wall thickness is less than 0.130 inches replace the pedal arm P/N 269A7330 before further flight. (d) On all Models except 269C, within the next 100 hours time in service, from the effective date of this AD unless already accomplished, rework the pilots' and copilots' left hand pedal arm and or sockets in accordance with SIN N-121.1 Part II. (e) Hughes service information notice (SIN) N-121.1, dated October 3, 1977, or later FAA approved revisions shall be used for compliance where indicated in this AD except for alternate inspection and rework methods approved under Paragraph (f). (f) Equivalent inspections, and reworks may be approved by Chief, Aircraft Engineering Division, FAA Western Region, Los Angeles, California. This amendment becomes effective January 24, 1978.
77-26-03: 77-26-03 MCCAULEY PROPELLERS: Amendment 39-1301 as amended by Amendment 39-3232. Applies to the following three bladed constant speed Model D3A34C401 and D3A34C402 propellers installed on but not limited to the Cessna Model TU206G, T207A, and T210M aircraft: Affected Propeller Serial Numbers 765815 765820 765821 765822 765823 765824 765828 766622 766624 766625 766626 766627 766638 766639 766640 766641 766642 766643 766644 767451 767452 767453 767454 767455 767456 767457 767458 767471 767472 767473 767474 767475 767476 767477 767479 767480 767481 767661 767662 767663 767664 767665 767666 767667 767668 767678 767679 767680 767682 767683 767684 767776 767777 767778 767799 767800 767801 767802 767803 767804 767805 767806 767822 767823 767824 767825 767826 767827 767828 767829 767830 767831 767832 767833 767834 767835 767837 767838 767839 767879 767880 768190 768191 768192 768193768194 768195 768196 768197 768199 768202 768254 768335 768336 768337 768338 768339 768340 768341 768342 768343 768344 768345 768346 768347 768348 768349 768350 768351 768352 768353 768354 768560 768888 768889 768890 768891 768892 768893 768970 768971 769002 769007 769008 769009 769092 769093 769094 769209 769210 769211 769212 769214 769215 769216 769217 769229 769230 769264 769265 769266 769267 769269 769288 769289 769291 769292 769449 769450 769451 769452 769453 769454 769455 769456 769457 769459 769460 769548 769549 769550 769551 769552 769553 769555 769556 769557 770212 770213 770214 770216 770217 770218 770219 770220 770221 770222 770223 770224 770225 770226 770390 770391 770392 770394 770489 770490 770491 770492 770583 770584 770585 770608 770609 770610 770611 770612 770613 770614 770615 770623 770624 770625 770626 770627 770628 770629 770630 770631 770632 770633 770634 770635 770636 770637 770638 770639 770640 770641 770642 770721 770722 770723 770724 770725 770864 770865 770866 770867 770868 770869 770870 770871 770872 770873 770874 770875 770876 770877 770878 770879 770880 770881 770882 770883 770884 770885 771659 771660 771661 771662 771663 771664 771665 771676 771677 771679 771680 771681 771682 771683 771684 771685 771813 771844 771845 771846 771987 771988 772004 772005 772006 772007 772008 772010 772011 772012 772013 772014 772015 772017 772041 772042 772204 772205 772206 772207 772208 772209 772210 772250 772251 772252 772253 772254 772255 772256 772257 772506 772507 772509 772510 772518 772519 772520 772521 772683 772694 772696 772698 772699 772700 772701 772702 772703 772704 772705 772731 772732 772733 772735 772865 772866 772940 772941 772942 772943 772945 772946 772974772975 772977 772981 772983 772984 773011 773093 773094 773095 773096 773097 773098 773101 773192 773193 773195 773196 773197 773198 773199 773200 773201 773202 773205 773207 773342 773343 773533 773709 773775 773780 773783 773785 773787 773918 773924 NOTE: Serial numbers are stamped on the side of the propeller hub. These propellers are equipped with Model 90DFA-() blades (usually - 10 cutoff). Compliance required before further flight, except that the airplane may be flown in accordance with FAR 21.197 to a Federal Aviation Administration Certificated Propeller Repair Station. To prevent possible blade pitch control failures, accomplish the following: (a) Replace blade actuating pin screws, P/N A-1635-104 (cadmium plated), with new screws, P/N A-1635-108 (black oxide) in accordance with McCauley Service Bulletin No. 129, dated October 7, 1977, and Service Manual No. 761001, or later Federal Aviation Administration approved revisions. (b) Replacement of the above parts must be accomplished by a Federal Aviation Administration Certificated Propeller Repair Station, since it is considered a major repair. (c) When the affected propellers are approved for return to service, compliance with this airworthiness directive shall be noted in the Aircraft's Records. (Cessna Service Letter SE-77-37 dated October 10, 1977, also pertains to this subject.) The manufacturer's specifications and procedures identified in this directive are incorporated herein and made part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by the directive who have not already received these documents from the manufacturer, may obtain copies upon request to McCauley Accessory Division, Cessna Aircraft Company, Box 7, Roosevelt Station, Dayton, Ohio 45417. These documents may also be examined at the Great Lakes Regional Office, 2300 East Devon Avenue, Des Plaines, Illinois 60018, and at FAA Headquarters, 800Independence Avenue, S.W., Washington, D.C. 20591. A historical file on this airworthiness directive which includes incorporated material in full is maintained by the FAA at its Headquarters in Washington, D.C., and the Great Lakes Region. Amendment 39-3159 became effective December 28, 1977. This amendment 39-3233 becomes effective on June 7, 1978.
2017-05-02: We are adopting a new airworthiness directive (AD) for certain Airbus Model A318-112 airplanes; Model A319-111, -112, -115, -132, and -133 airplanes; Model A320-214, -232, and -233 airplanes; and Model A321-211, -212, -213, -231, and -232 airplanes. This AD was prompted by a quality control review on the final assembly line, which determined that the wrong aluminum alloy was used to manufacture several structural parts. This AD requires a one-time eddy current conductivity measurement of certain cabin and cargo compartment structural parts to determine if an incorrect aluminum alloy was used, and replacement if necessary. We are issuing this AD to address the unsafe condition on these products.
2001-20-11: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 757 series airplanes, that currently requires repetitive freeplay checks of the elevator, and replacement of worn elevator power control actuator (PCA) reaction link rod-end bearings and the PCA rod-end bearing, if necessary. That AD also provides for an optional terminating action for the repetitive checks. This amendment removes the optional terminating action provided by the existing AD, expands the applicability of the existing AD to include additional airplanes, and requires repetitive freeplay checks of the elevator at a revised repeat interval and repetitive lubrication of bearings of the elevator actuator load loop and hinge line. The actions specified by this AD are intended to prevent unacceptable airframe vibration during flight, which could lead to excessive wear of bearings of the elevator PCA load loop and hinge line and result in reduced controllability of theairplane. This action is intended to address the identified unsafe condition.
88-09-01: 88-09-01 BELL HELICOPTER TEXTRON, INC.: Amendment 39-5874. Applies to Bell Helicopter Textron, Inc., Model 214ST helicopters certificated in any category. (Airworthiness Docket No. 87-ASW-64.) Compliance is required as indicated, unless already accomplished. (a) Within the next 25 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 25 hours' time in service, perform an inspection of the tail rotor intermediate gearbox, P/N 214-040-009-001, as follows: (1) Remove the fairings and covers as required to gain access to the intermediate gearbox. (2) Visually inspect the three gearbox mounting lugs for cracks. If a crack indication is found, inspect using a fluorescent or dye penetrant method. If a crack is found, remove and replace the intermediate gearbox case with a serviceable part before further flight and accomplish the requirements of paragraphs (b)(3), (b)(5), and (b)(6) of this AD. (3) If no cracks are found, verify that the washer stack-up is in accordance with figure 4, page 16, Section 65-20-00 of the Bell 214ST Maintenance Manual, Rev. 22, dated May 14, 1987, and determine that the torque on the three attachment bolts, P/N AN5H15A, is 100-140 in-lbs. Resecure the attachment bolts with MS20995C32 safety wire or equivalent. (b) Within the next 250 hours' time in service after the effective date of this AD, remove and inspect the tail rotor intermediate gearbox, P/N 214-040-009-001, as follows: (1) Remove the fairings and covers as required to gain access to the intermediate gearbox. (2) Remove the intermediate gearbox and visually inspect the three gearbox mounting lugs for cracks. If a crack indication is found, inspect using a fluorescent or dye penetrant method. If a crack is found, remove and replace the intermediate gearbox with a serviceable part before further flight. (3) Inspect the intermediate gearbox and intermediate gearbox support areaof the tailboom for evidence of barrier tape. If barrier tape is present, it must be removed before further flight. (4) Inspect the intermediate gearbox mounting lugs for evidence of excessive wear and bolt hole elongation. If wear or bolt hole elongation is beyond the limits specified on pages 35 and 36 in Section 65-20-07 or figure 26, page 48, Section 65-20-08, of the Bell 214ST Component Repair and Overhaul Manual, Rev. 12, dated April 15, 1986, remove and replace the intermediate gearbox case with a serviceable part before further flight. (5) Inspect the intermediate gearbox tailboom attachment points for excessive wear or damage. If the wear or damage is beyond the limits specified in figure 16, page 40B, Section 53-11-00 of the Bell 214ST Maintenance Manual, Rev. 22, dated May 14, 1987, remove and replace the worn or damaged parts with serviceable parts before further flight. (6) Clean the mounting surfaces of the intermediate gearbox and tailboom using methyl-ethyl-ketone (MEK), or equivalent safety solvent. Apply epoxy polyamide primer or zinc chromate primer to the mounting surfaces. Do not use barrier tape during reassembly. Install the gearbox while the primer is wet. The washer stack-up must comply with Figure 4, page 16, Section 65-20-00 of the Bell 214ST Maintenance Manual, Rev. 22, dated May 14, 1987. Torque the three attachment bolts, P/N AN5H15A, to 100-140 in-lbs. Secure the bolts with MS20995C32 safety wire or equivalent. (c) Upon complying with paragraph (b) of this AD, the requirements in paragraph (a) of this AD no longer apply. (d) In accordance with FAR 21.197 and 21.199, the helicopter may be flown to a base where the requirements of this AD can be accomplished. (e) An alternate method of compliance which provides an equivalent level of safety, may be used when approved by the Manager, Helicopter Certification Branch, Federal Aviation Administration, Fort Worth, Texas 76193-0170. NOTE: Compliance with Parts I and II of Bell Helicopter Textron Alert Service Bulletin No. 214ST-87-40, dated 9/17/87, constitutes compliance with this AD. The procedure shall be done in accordance with page 16, Section 65-20-00 and page 40b, Section 53-11-00 of the Bell 214ST Maintenance Manual, Rev. 22, dated May 14, 1987; and pages 35 and 36, Section 65-20-07 and page 48, Section 65-20-08 of the Bell 214ST Component Repair and Overhaul Manual, Rev. 12, dated April 15, 1986. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a)(1) and 1 CFR Part 51. Copies may be obtained from Bell Helicopter Textron, Inc., P.O. Box 482, Fort Worth, Texas 76101, Attention: Customer Support. Copies may be inspected at the Office of Regional Counsel, FAA, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, D.C. This amendment 39-5874 becomes effective April 22, 1988.
79-23-05: 79-23-05 BEECH: Amendment 39-3611. Applies to Model 77 (S/N's WA-1 through WA-39, WA-41, WA-42 and WA-44 through WA-46) airplanes certificated in all categories. COMPLIANCE: Required as indicated unless already accomplished. To assure structural integrity of the attachment of the left and right wing to the fuselage carry-thru structure, accomplish the following: A) Prior to the next flight, in accordance with instructions in (1) Beech Mailgram sent to owners on October 3, 1979 or (2) Beechcraft Service Instructions No. 1088 (whichever is available): 1. Remove the four NAS1112-11 or -15 wing main spar to fuselage carry- thru structure attachment bolts and visually inspect (1) the threads on the bolts and (2) the threads in the Part Number 108-120010-5 internal nut plate for each bolt, for stripped threads or other damage. 2. Replace any damaged NAS1112-11 or -15 bolts or any Part Number 108-120010-5 internal nut plates with new components. NOTE: Ifnew NAS1112 bolts are obtained locally, the heads must have a .093 to .103 inch diameter hole drilled in them to accommodate MS20995-NC32 safety wire. Locate the hole in the bolt head at approximately the same location as the safety wire hole in the bolt removed. 3. Modify, apply zinc chromate primer to, and install the correct number of AN960-1216 washers under the NAS1112-11 or -15 bolts. NOTE: The Beech Part Number 108-120013-1 radius washer must be installed between the AN960-1216 washers and the main spar tube on each of the four NAS1112-11 or - 15 bolts. The total shank length of a NAS1112-11 bolt is 1.26 inches and for the NAS1112-15 1.51 inches. B) Aircraft may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished provided that inspection of each of the four NAS1112-11 or -15 wing main spar to fuselage carry-thru attachment bolts shows that all four bolts are (1) in place, (2) safety wired in place and(3) secure when an attempt is made to turn or pull the bolts by hand. This inspection must be accomplished by an FAA certificated aircraft mechanic or persons authorized under Federal Aviation Regulations 43.3. C) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective on November 19, 1979, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated October 5, 1979.
2022-19-09: The FAA is adopting a new airworthiness directive (AD) for all Airbus Canada Limited Partnership Model BD-500-1A10 and BD-500-1A11 airplanes. This AD was prompted by reports of in-service findings of corrosion on the flange of the main landing gear (MLG) lower spindle pin. This AD requires repetitive inspections of the left and right MLG lower spindle pins to detect corrosion, and applicable repair or replacement if necessary, as specified in a Transport Canada Civil Aviation (TCCA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2017-05-07: We are adopting a new airworthiness directive (AD) for The Boeing Company Model 777-200 and -300 series airplanes equipped with Rolls-Royce Model Trent 800 engines. This AD was prompted by reports of damage to the upper bifurcation forward fire seal and seal deflector, and localized damage to the insulation blanket installed just aft of the fire seal. This AD requires installing serviceable thrust reverser (T/R) halves on the left and right engines. We are issuing this AD to address the unsafe condition on these products.
2010-24-11: The FAA is adopting a new airworthiness directive (AD) for certain Model 737-600, -700, -700C, -800, and -900 series airplanes. This AD requires sealing the fasteners on the front and rear spars inside the main fuel tank and on the lower panel of the center fuel tank, inspecting the wire bundle support installation in the equipment cooling system bays to identify the type of clamp installed and determine whether the Teflon sleeve is installed, and doing related corrective actions if necessary. This AD results from a design review of the fuel tank systems. We are issuing this AD to prevent arcing at certain fuel tank fasteners in the event of a lightning strike or fault current event, which, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
2001-20-03: This amendment adopts a new airworthiness directive (AD) for Bell Helicopter Textron Canada (BHTC) Model 206L-4 helicopters that requires installing a high altitude tail rotor static stop yield indicator (indicator) to allow operators to detect excessive bending loads sustained by the tail rotor yoke. A preflight check of the indicator is also required. This amendment is prompted by a determination that a tail rotor yoke with a high altitude rotor system is susceptible to a static and dynamic overload. Static overload could occur after the tail rotor yoke sustains an excessive bending load due to a strike from a ground vehicle. Dynamic overload could occur as a result of a hard landing. The actions specified by this AD are intended to prevent failure of the tail rotor yoke in flight and subsequent loss of control of the helicopter.
86-13-03: 86-13-03 AVIONS MARCEL DASSAULT - BREGUET AVIATION: Amendment 39-5337. Applies to Model Falcon 10 series airplanes, serial numbers 1 through 206 inclusive, certificated in any category. Compliance is required within 60 days after the effective date of this AD. To detect an incorrectly wired automatic leading edge slat selector valve, accomplish the following in accordance with AMD-BA Service Bulletin F10-27-31(-247), dated July 18, 1984, unless previously accomplished. A. Test the automatic leading edge slat extension system, and revise the leading edge slat selector valve wiring, if necessary. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspectionsand/or modifications required by this AD. All persons affected by this directive, who have not already received the appropriate service document from the manufacturer, may obtain copies upon request to the AMD-BA Representative, 40 J.J.C., Teterboro Airport, Teterboro, New Jersey 07608. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective July 24, 1986.
90-12-06: 90-12-06 GENERAL ELECTRIC COMPANY: Amendment 39-6591. Docket No. 89-ANE-07. Applicability: General Electric Company (GE) CF6-50/-45 turbofan engines installed on, but not limited to, McDonnell Douglas DC-10, Airbus A300, and Boeing B747 aircraft. Compliance: Required at the next shop visit after the effective date of this AD, but not later than January 1, 1992, unless already accomplished. To prevent uncontained low pressure turbine (LPT) failure due to a fire in the LPT rotor cavity, accomplish the following: (a) Install the center vent tube extension assembly and associated hardware in accordance with GE CF6- 50 Series Service Bulletin (SB) 72-395, Revision 3, dated June 30, 1989. NOTE: Shop visit for the purpose of this AD is defined as the induction of the engine into the shop for performance of maintenance. (b) 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. (c)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 (schedule) times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. The installation of the center vent tube extension assembly and associated hardware shall be done in accordance with GE CF6-50 Series SB 72-395, Revision 3, dated June 30, 1989, which incorporates the following list of effective pages: Page Number Revision Number Date Pages 1 - 6 Revision 3 June 30, 1989 Page 7 Revision 1 April 29, 1977 Pages 8 - 26 Revision 3 June 30, 1989 This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552 (a) and 1 CFR Part 51. Copies may be obtained from General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the Office of the Federal Register, 1100 L Street, NW, Room 8301, Washington, DC 20591. This amendment (39-6591, AD 90-12-06) becomes effective on June 25, 1990.
85-24-04: 85-24-04 CESSNA AIRCRAFT COMPANY: Amendment 39-5231. Applies to Cessna Model S550 airplanes, Unit Numbers -0001 through -0075, certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent hazardous accumulation of ice on the inboard wing leading edge during flight in icing conditions, accomplish the following: A. Prior to further flight: 1. Fabricate and install on the instrument panel in clear view of the pilot the following placard, using letters of a minimum of 0.10 inch in height: "FLIGHT INTO KNOWN OR FORECAST ICING PROHIBITED" and operate the airplane accordingly. 2. Revise the Airplane Flight Manual, Section II, Operating Limitations - Operations Authorized, to read: "This airplane is approved for day and night, VFR, IFR flight. Flight into known or forecast icing conditions prohibited." This may be accomplished by inserting a copy of this AD in the Airplane Flight Manual. B. Within the next 15flight hours or 15 days after the effective date of this AD, whichever occurs first, inspect and replace, as necessary, the fluid ice protection system proportioning units as described below: 1. Locate the fluid anti-ice proportioning units as follows: a. Remove the access panel from the lower surface of the left wing root leading edge fairing to gain access to the left outboard wing proportioning unit. b. Remove the right wing root leading edge fairing to gain access to the right outboard wing and left and right inboard wing-engine proportioning units. c. Remove access panel 340BL or 340BR (located below the horizontal stabilizer on the vertical stabilizer) to gain access to the tail proportioning unit. 2. Locate the identification plate installed on each proportioning unit and identify that the correct part number proportioning unit is installed in the proper location (refer to TABLE 1, below). TABLE 1 PROPORTIONING UNIT PART NUMBER/LOCATIONPART NUMBER LOCATION PU306DW38 Left and Right Wing Root (Left and Right Outboard Wing) PU306DC39 Right Wing Root (Left and Right Inboard Wing-Engine) PU306DT41 Vertical Stabilizer (Tail) Refer to Model S550 Maintenance Manual, Revision 30-1, dated May 1, 1985, for detailed installation information. 3. If the correct part number proportioning units are installed in all locations, reinstall all access plates using existing hardware and make an entry in the aircraft maintenance record indicating that the requirements of this AD have been complied with. 4. If the incorrect part number proportioning unit(s) is installed in any location(s), remove and replace the proportioning unit(s) with the correct part number proportioning unit. Refer to the Model S550 Maintenance Manual, Revision 30-1, dated May 1, 1985, Chapter 30, Removal/ Installation Proportioning Unit, for removal and replacement instructions. C. Within 5 days after the required inspection,report any defects found to the Manager, Wichita Aircraft Certification Office, FAA, Central Region, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. D. The requirements of paragraphs A.1. and A.2., above, may be accomplished by the holder of a pilot certificate issued under Part 61 of the Federal Aviation Regulations (FAR) on any airplane owned or operated by him. The person accomplishing these actions must make the appropriate aircraft maintenance record entry as prescribed by FAR 91.173. E. The requirements of paragraphs A.1. and A.2., above, are no longer required after the requirements of paragraph B. of this AD have been accomplished. F. Airplanes may be flown in accordance with FAR 21.197 and 21.199 to a location where the inspection/modification requirements of the AD can be accomplished. G. Alternate means of compliance with this AD which provide an acceptable level of safety may be used if approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Cessna Aircraft Company, P.O. Box 7704, Wichita, Kansas 67277. These documents also may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas. This Amendment becomes effective February 24, 1986, as to all persons, except those persons to whom it was made immediately effective by Priority Letter AD 85-24-04, issued December 3, 1985.
91-10-03 R1: 91-10-03 R1 GENERAL ELECTRIC COMPANY: Amendment 39-9186. Docket 90-ANE- 25. Revises AD 91-10-03, Amendment 39-6956. Applicability: General Electric Company (GE) CF6-45 and CF6-50 series turbofan engines installed on, but not limited to, McDonnell Douglas DC-10 series, Boeing 747 series, and Airbus A300 series aircraft. NOTE: This AD applies to each engine identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For engines that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must use the authority provided in paragraph (e) to request approval from the FAA. This approval may address either no action, if the current configuration eliminates the unsafe condition, or different actions necessary to address the unsafe condition described in this AD. Such a request should include an assessment of the effect of the changed configuration on the unsafe condition addressed by this AD. In no case does the presence of any modification, alteration, or repair remove any engine from the applicability of this AD. Compliance: Required as indicated, unless accomplished previously. To prevent a high pressure compressor (HPC) rear shaft fracture, which could result in an inflight engine shutdown and an uncontained engine failure, accomplish the following: (a) Fluorescent penetrant inspect HPC rear shafts, Part Numbers (P/N) 9127M58P03, 9079M63P12, 9079M63P15, 9079M63P16, 9079M63P17, 9079M63P18, and 9079M63P19, in accordance with the Accomplishment Instructions of GE Service Bulletin (SB) No. 72-958, Revision 1, dated October 18, 1990, as follows: (1) For HPC rear shafts currently installed with hook bolts, P/N 9012M99G10, 9114M95G07, and 9114M95G10, inspect in accordance with the following schedule: (i) For shafts which have not been previously inspected and have 10,000 cycles since new (CSN) or greater on the effective date of this airworthiness directive (AD), inspect within the next 1,500 cycles in service (CIS) after the effective date of this AD. (ii) For shafts which have not been previously inspected and have less than 10,000 CSN on the effective date of this AD, inspect within the next 2,500 CIS from the effective date of this AD, or before accumulating 7,500 CSN, whichever occurs later. However, no shaft may exceed 11,500 CSN prior to inspection. (iii) For shafts that have been previously inspected and have 3,000 cycles since last inspection (CSLI) or less on the effective date of this AD, reinspect within 4,500 CSLI, or before accumulating 7,500 CSN, whichever occurs later. (iv) For shafts that have been previously inspected and have greater than 3,000 CSLI on the effective date of this AD, reinspect within the next 1,500 CIS from the effective date of this AD, or before accumulating 7,500 CSN, whichever occurs later.(v) Remove from service, HPC rear shaft hook bolts identified in (a)(1) of this AD, after any inspection performed in accordance with paragraph (a)(1) of this AD, and replace with new tapered turn-around bolts, P/N 1375M69P01 or VCD0016. (2) For HPC rear shafts installed with turn-around bolts, P/N 9249M54P01, or tapered turn-around bolts, P/N 1375M69P01 or VCD0016, inspect in accordance with the following schedule: (i) For shafts which have not been previously inspected and have 6,500 CSN or greater on the effective date of this AD, inspect within the next 2,500 CIS after the effective date of this AD. (ii) For shafts which have not been previously inspected and have less than 6,500 CSN on the effective date of this AD, inspect prior to accumulating 9,000 CSN. (iii) For shafts that have been previously inspected and have 3,500 CSLI or less on the effective date of this AD, reinspect within 6,000 CSLI, or before accumulating 9,000 CSN, whichever occurs later.(iv) For shafts that have been previously inspected and have greater than 3,500 CSLI on the effective date of this AD, reinspect within the next 2,500 CIS from the effective date of this AD, or before accumulating 9,000 CSN, whichever occurs later. (v) Remove from service, HPC rear shaft turn-around bolts identified in paragraph (a)(2) of this AD, after any inspection performed in accordance with paragraph (a)(2) of this AD, and replace with new tapered turn-around bolts, P/N 1375M69P01 or VCD0016. NOTE: Information concerning the tapered turn-around bolt noted in paragraph (a) of this AD can be found in GE SB No. 72-877. (b) Remove from service, prior to further flight, any shafts found cracked at inspection. (c) Thereafter, for shafts which have been inspected in accordance with paragraph (a) of this AD, reinspect in accordance with the Accomplishment Instructions of GE SB No. 72- 958, Revision 1, dated October 18, 1990, at intervals not to exceed 6,000 CSLI. (d) Compliance with paragraph (a) of AD 91-10-03 satisfies the corresponding requirements of paragraph (a) of this AD. (e) An alternative method of compliance or adjustment of the initial compliance time that provides an acceptable level of safety may be used if approved by the Manager, Engine Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Engine Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Engine Certification Office. (f) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the aircraft to a location where the requirements of this AD can be accomplished. (g) The actions required by this AD shall be done in accordance with the following service document: Document No. Pages Revision Date GE SB No. 72-958 1-2 1 October 18, 1990 3-6 Original August 15, 1990 Total pages: 6. 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 as of June 17, 1991. Copies may be obtained from General Electric Company, Technical Publications Department, 1 Neumann Way, Cincinnati, OH 45215. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (h) This amendment becomes effective on April 28, 1995.
90-21-16: 90-21-16 SAAB-SCANIA: Amendment 39-6763. Docket No. 90-NM-111-AD. Applicability: Model SF-340A series airplanes, Serial Numbers 004 through 159, certified in any category. Compliance: Required as indicated, unless previously accomplished. To prevent loss of climb performance during single engine operation (i.e. one engine failed), and latent failures in the engine control unit, accomplish the following: A. Within 60 days after November 20, 1989 (the effective date of Amendment 39-6361, AD 89- 22-08), install modified electrical control units (ECU) in accordance with General Electric Service Bulletin 74-16, dated March 22, 1989. B. Within 60 days after the effective date of this amendment, remove any relays and associated wiring previously installed in accordance with SAAB-Scania Service Bulletins SAAB SF340- 61-020 and SAAB SF340-77-003, in accordance with the procedures specified in SAAB-SCANIA Service Bulletin SF340-77-006, dated March 9, 1990. C. After completion of paragraphs A. and B. of this AD, delete steps referencing "Applicable With Mod. No. 1931" in the Emergency Procedures Section and Normal Procedures Expanded Checklist of the FAA-approved Airplane Flight Manual (AFM). 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, 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. 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 SAAB-SCANIA, Aircraft Division, S.58188, Linkoping, Sweden. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. Airworthiness Directive 90-21-16 supersedes AD 89-22-08, Amendment 39- 6361. This amendment (39-6763, AD 90-21-16) becomes effective on November 14, 1990.
2017-04-03: We are adopting a new airworthiness directive (AD) for Pilatus Aircraft Ltd. Models PC-6, PC-6-H1, PC-6-H2, PC-6/350, PC-6/350-H1, PC- 6/350-H2, PC-6/A, PC-6/A-H1, PC-6/A-H2, PC-6/B-H2, PC-6/B1-H2, PC-6/B2- H2, PC-6/B2-H4, PC-6/C-H2, and PC-6/C1-H2 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as certain combinations of the aileron counterweight and the attaching parts possibly resulting in reduced thread engagement and leading to disconnection of the aileron counterweight from the aileron. We are issuing this AD to require actions to address the unsafe condition on these products.
2001-20-02: This amendment adopts a new airworthiness directive (AD), that is applicable to Pratt & Whitney (PW) PW4000 series turbofan engines with 2nd stage high pressure turbine (HPT) air seal assembly part number (P/N) 50L976 or P/N 50L960 installed. This amendment requires operators to recalculate 2nd stage HPT air seal assembly cycles-in-service, based on flight hour-to-cycle ratio usage. This amendment also requires upon recalculation, initial and repetitive on-wing borescope inspections of 2nd stage HPT air seal assemblies for cracks based on the newly calculated service life. This amendment also requires the removal from service of any cracked seal assemblies, and the removal of seal assemblies at or before newly calculated service life limits. This amendment is prompted by reports that thirteen 2nd stage HPT air seal assemblies have been found cracked in the rim area. Although these thirteen air seals were operating in the hottest configuration design, which is no longer in service, the current design 2nd stage HPT air seal assemblies are still operating in a temperature environment that is hotter than anticipated. The actions specified by this AD are intended to prevent 2nd stage HPT air seal assembly fracture that could result in an uncontained engine failure.
88-17-05: 88-17-05 SCHEMPP-HIRTH GMBH: Amendment 39-5982. Applies to Models Nimbus-2B, Janus B, and Mini-Nimbus B gliders certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent failure of the elevator drive bracket and loss of pitch control accomplish the following: (a) Within the next 10 hours time in service (TIS) after the effective date of this AD unless already accomplished, and thereafter at intervals not to exceed 50 hours TIS from the last inspection, visually inspect the elevator drive bracket, P/N 30.055, using a 5-power or greater magnifying glass for cracks in the area adjacent to the welds on both sides of the bracket as shown in Sketch 1 or Sketch 3 of Schempp-Hirth Flugzeugbau GmbH Technical Note (TN) No. 286-24, dated August 14, 1987, for Nimbus-2B gliders, TN No. 295-19, dated August 14, 1987, for Janus B gliders, and TN No. 328-8, dated August 14, 1987, for Mini-Nimbus B gliders. (b) If cracks are found in the elevator drive bracket, replace the elevator drive bracket with elevator drive bracket, Part Number (P/N) 30.055, modification "a", dated August 24, 1987, before further flight. (c) Prior to November 30, 1988, unless already accomplished, replace elevator drive bracket P/N 30.055 with elevator drive bracket P/N 30.055, modification "a", dated August 24, 1987. NOTE: Installation of P/N 30.055, modification "a" is shown in Schempp-Hirth drawing HS5-30.055/1 Elevator-to-Stabilizer Assembly. (d) Upon request an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Brussels Aircraft Certification Office, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, 15 Rue de la Loi B-1040 Brussels, Belgium; telephone 513.38.30, Ext. 2710; or the Manager, New York Aircraft Certification Office, Aircraft Certification Division, Federal Aviation Administration, New England Region, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581; telephone (516) 791- 6680. (e) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, the Manager, Brussels Aircraft Certification Office, or the Manager, New York Aircraft Certification Office may adjust the compliance times specified in this AD. Schempp-Hirth GmbH TN No. 286-24 with 4 attached sketches, TN No. 295-19 with 4 attached sketches, and TN No. 328-8 with 4 attached sketches and 1 attached drawing, all dated August 14, 1987, identified and described in this document, 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 Messrs. Schempp-Hirth, Flugzeugbau GmbH, Krebenstr 25, PostFach 14 43, D-7312 Kircheim/Teck, Federal Republic of Germany. These documents may also be examined at the Office of the Regional Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, Room 311, Rules Docket 88-ANE-27, between the hours of 8:00 a.m. and 4:30 p.m., Monday through Friday, except federal holidays. This amendment, 39-5982, becomes effective on August 19, 1988.
69-25-08: 69-25-08 LYCOMING ENGINES: Amdt. 39-892 as amended by Amendment 39-938. Applies to Lycoming GO-435-C2A, GO-480, GSO-480 and IGSO-480 type engines using spline type reduction gear assemblies 72875 and 72879, GO-480 and GSO-480 type engines using flange type reduction gear assemblies 69346, 70412 and 71803 and IGO-540 and IGSO-540 type engines using reduction gear assemblies 72782, 74900, 75679, 76494 and 77731. Compliance required unless already accomplished prior to the accumulation of six hundred (600) hours in service or 100 hours after the effective date of this AD whichever comes later for the GO-435, IGO-540 and IGSO-540 type engines and prior to the accumulation of seven hundred (700) hours in service or 100 hours after the effective date of this AD whichever comes later for GO-480, GSO-480 and IGSO-480 type engines. To prevent failure of the reduction gear assembly a one time inspection is to be accomplished as follows: a. Check torque on reduction gearing pinioncage attaching nut. 1. If found to be in excess of 50 ft.-lbs. remove, inspect for damage and replace the oil retaining housing if necessary. During assembly comply with part (b) using a new lockplate and a pinion cage attaching nut torque of 400 ft.-lbs. When the assembly is disassembled replace thrust bearing oil slinger P/N 68319 with P/N 77454. 2. If found to be less than 50 ft.-lbs. disassemble the assembly and insure that all components meet approved specifications and replace the thrust bearing oil slinger P/N 68319 and P/N 77454 and the oil retaining housing. During assembly comply with part (b) using a new lockplate and a pinion cage attaching nut torque of 400 ft.-lbs. b. If the pinion cage is secured with a spline nut it must be replaced with an approved hexagonal nut. c. If the assembly has been rebuilt using thrust bearing oil slinger P/N 77454, stamp the suffix (-1) to the part number on the reduction gear housing. Lycoming Service BulletinNo. 319A and Lycoming Service Instruction No. 1210A pertain to this subject. Amendment 39-892 effective December 17, 1969. This Amendment (39-938) is effective February 19, 1970.
81-19-04: 81-19-04 RAJAY INDUSTRIES, INC.: Amendment 39-4214. Applies to all affected aircraft, certificated in all categories modified per Rajay Supplemental Type Certificates Aircraft Model Rajay STC No. Aerocommander 500-A STC SA683WE Aerocommander 500-B, 500-S & 500-U STC SA529WE Beech H35, J, K, M, N & P, and 35-33, 35-A33, 35-B33, 35-C33, E33 & F33 STC SA1252WE Beech S35, V35, V35A, V35B, 35-C33A, E33A, E33C, F33A, F33C STC SA2556WE Beech 95, B95, B95A, D95A & E95 STC SA153SO Britten Norman Islander, BN-2, BN-2A, BN-2A-6, BN-2A-8, BN-2A-9 STC SA2243WE Cessna 180-A, B, C, D, E, F, G, H & J STC SA1157WE Cessna 182-A, B, C, D, E, F, G, H, J, K, L, M, N & P STC SA1032WE Cessna 210-A, B, C, 210-5(205), 210-5A(205A) STC SA1098WE Cessna 310-C, D, E, F & G STC SA2383SO Cessna 310-I and J STC SA181SO Evangel 4500-300 Evangel 4500-300, Series II STC SA2657WE Helio Courier H-295 STC SA156SO Helio Courier H-395 STC SA125SO Lake 4 and 4A STC SA2270WE Lake 4-200 STC SA2990WE Mooney 20A, M20B, M20C, M20D & M20G STC SA1156WE Mooney 20E & M20F STC SA1411WE Mooney 20J STC SA3555WE Piper PA-23-160, Apache STC SA1260WE Piper PA-23-160, Apache STC SA4-1637-WE Piper PA-23-235, -250, Aztec "B" STC SA539WE Piper PA-23-250, Aztec "C" STC SA840WE Piper PA-24-250, Comanche STC SA811WE Piper PA-24-400, Comanche STC SA2359WE Piper PA-30 & 39, Twin Comanche STC SA787WE Piper PA-32-260, Cherokee STC SA1557WE Piper PA-32R-300, Lance STC SA3513WE Piper PA-34-200, Seneca STC SA2937WE Compliance required as indicated, unless already accomplished. To prevent failure of the powerplant hoses carrying air, fuel and/or oil and resultant fire hazard, accomplish the following: (a) Within 100 hours' time in service from the effective date of this AD, or prior to return to service after the next annual inspection, whichever occurs first, visually inspect the powerplant fuel, air and oil hose assemblies listed in RajayService Letter No. 28 dated August 3, 1981 to determine the general condition and age of the hose assemblies based upon the metal plate attached to the hose, and; 1) If the hose assembly is five years old or older, replace with like serviceable part prior to further flight. 2) If the hose assembly does not have a metal tag and the age cannot be determined, replace with like serviceable part prior to further flight. 3) Record hose ages in the aircraft engine log book and establish a replacement schedule for affected hoses such that a five year life will not be exceeded. 4) If the hose assembly is deteriorated (regardless of age), replace with like serviceable part prior to further flight. (b) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate aircraft to a base for the accomplishment of inspections or modifications required by this AD. (c) Alternative inspections, modifications or other actions which provide an equivalent level of safety may be used when approved by the Chief, Engineering and Manufacturing Branch, FAA Western 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. 553(a)(1). All persons affected by this directive, who have not already received these documents from the manufacturer, may obtain copies upon request to Rajay Industries, Inc., 2600 East Wardlow Road, P.O. Box 207, Long Beach, California 90801, telephone (213) 426-0346. A historical file on this AD, which includes the incorporated material in full, is maintained by the FAA at its Headquarters in Washington, D.C. and at FAA Western Region Office. This amendment becomes effective September 17, 1981.
2022-27-09: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Helicopters Model EC130T2 helicopters. This AD was prompted by a crack in the tailboom. This AD requires repetitively inspecting the vibration level on the tail rotor drive shaft and, depending on the results, taking corrective action. This AD also requires reporting information and prohibits installing certain rotor drive shafts unless the inspection is done, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2010-24-03: This amendment adopts a new airworthiness directive (AD) for Robinson Model R22, R22 Alpha, R22 Beta, and R22 Mariner helicopters, and Model R44 and R44 II helicopters. This AD requires visually inspecting each tail rotor (T/R) control pedal bearing block support (support) for a crack, measuring the thickness of each support, installing support safety tabs on certain supports, and replacing supports of a certain thickness during the next 2,200 hour overhaul. This amendment is prompted by two reports of Model R22 helicopters experiencing broken supports during flight, which resulted in the T/R control pedals becoming jammed. The actions specified by this AD are intended to prevent the supports from breaking, which can bind the T/R control pedals, resulting in a reduction of yaw control and subsequent loss of control of the helicopter.
2001-17-09 R1: This amendment rescinds Airworthiness Directive (AD) 2000-17-09, which is applicable to certain McDonnell Douglas Model MD-11 series airplanes. That AD requires an inspection of the upper avionics circuit breaker panel at the main observer's station to detect damage of the wires and to verify the correct routing of the wire bundles; corrective actions, if necessary; and installation of a new clamp, spacer, and sta-straps. The requirements of that AD were intended to prevent chafing in the upper avionics circuit breaker panel of the main observer's station, which could result in arcing and consequent smoke and/or fire in the cockpit. Since the issuance of that AD, the FAA has determined that the improper procedures specified by the service bulletin referenced in that AD could lead to wiring pre-load conditions and consequent wire damage, and arcing in the upper avionics circuit breaker panel. Such conditions could result in arcing and consequent smoke and/or fire in the cockpit.
2001-19-02: This amendment adopts a new airworthiness directive (AD) that is applicable to GE CF34-3A1, -3B, and -3B1 turbofan engines with scavenge screens part numbers (P/N's) 4047T95P01 and 5054T86G02 installed in the B-sump oil scavenge system. This action requires initial and repetitive visual inspections and cleaning of the B-sump scavenge screens. This amendment is prompted by five reports of B-sump oil scavenge system failure causing engine in-flight shutdowns. The actions specified in this AD are intended to prevent B-sump scavenge screen blockage due to coking, which could result in ignition of B-sump oil in the secondary air system, fan drive shaft separation, and uncontained engine failure.