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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.
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.