Results
2011-24-04: We are adopting a new airworthiness directive (AD) for certain Model DC-10-10, DC-10-10F, and MD-10-10F airplanes. This AD was prompted by reports of three instances of fuel leaks in the lower cap splice of the wing rear spar at station Xors=409. Investigation revealed the fuel leak was due to a crack in the lower cap. If not corrected, this condition could result in fuel leaks or cracking of the lower wing skin and structure, causing possible inability of the structure to sustain the limit load and adversely affecting the structural integrity of the airplane. This AD requires repetitive inspections for cracking on the lower cap of the rear spar of the left and right wings between stations Xors=417 and the outboard edge of the lower cap splice of the wing rear spar at station Xors=400; temporary and permanent repairs if necessary; and repetitive inspections of repaired areas, and corrective actions if necessary. We are issuing this AD to correct the unsafe condition on these products.
2012-02-07: We are superseding two existing airworthiness directives (ADs) for General Electric Company (GE) CF6-45 and CF6-50 series turbofan engines with certain low-pressure turbine (LPT) rotor stage 3 disks installed. The existing ADs currently require inspections of high- pressure turbine (HPT) and LPT rotors, engine checks, and vibration surveys. This new AD retains the requirements of the two ADs being superseded, adds an optional LPT rotor stage 3 disk removal after a failed HPT blade borescope inspection (BSI) or a failed engine core vibration survey, establishes a new lower life limit for the affected LPT rotor stage 3 disks, and requires removing these disks from service at times determined by a drawdown plan. This AD was prompted by the determination that a new lower life limit for the LPT rotor stage 3 disks is necessary. We are issuing this AD to prevent critical life- limited rotating engine part failure, which could result in an uncontained engine failure and damage to theairplane.
97-15-14: This amendment adopts a new airworthiness directive (AD) that applies to Industrie Aeronautiche e Meccaniche Rinaldo Piaggio S.p.A. (Piaggio) Model P-180 airplanes. This action requires inspecting for cracks around the vertical pin and the torque tube bottom flange of the rudder, and the fasteners that connect the torque tube to the bottom flange (torque tube bottom flange assembly). If cracks are not found, repetitively inspect until cracks are visible. If cracks are evident, this action requires modifying the rudder torque tube bottom flange assembly by replacing the cracked part with a part of improved design, which terminates the repetitive inspection. This AD is the result of several reports of fatigue cracks around the pin that vertically supports the rudder axle. The actions specified by this AD are intended to prevent fatigue cracks in the rudder torque tube bottom flange, which could result in loss of rudder control and possible loss of the airplane.
2012-02-11: We are superseding an existing airworthiness directive (AD) for [[Page 4649]] all RR RB211-535E4-37, -535E4-B-37, -535E4-B-75, and -535E4-C-37 turbofan engines. That AD currently requires performing initial and repetitive visual and fluorescent penetrant inspections (FPI) of the low-pressure (LP) turbine stage 1, 2, and 3 discs to detect cracks in the discs. This new AD continues to require those inspections and changes the definition of a shop visit to be less restrictive. This AD was prompted by our finding that the definition of shop visit in the existing AD was too restrictive. We are issuing this AD to revise the definition of shop visit and to detect cracks in the LP turbine stage 1, 2, and 3 discs, which could result in an uncontained release of LP turbine blades and damage to the airplane.
97-15-11: This amendment supersedes an existing airworthiness directive (AD), applicable to Avco Lycoming and Textron Lycoming reciprocating engines, that currently requires removal from service of defective piston pins, and replacement with serviceable parts. This amendment adds additional affected engine models that may have defective piston pins installed, and references a revised service bulletin. This amendment is prompted by the determination that additional engine models may have defective piston pins installed. The actions specified by this AD are intended to prevent piston pin failure, which could result in engine failure.
97-11-13: This document makes a correction to Airworthiness Directive (AD) 97-11-13, which was published in the Federal Register on May 29, 1997 (62 FR 28999), and concerns Fairchild Aircraft SA226 and SA227 series airplanes. The date of Fairchild Service Bulletin (SB) 227-24-008 is incorrectly referenced in paragraph (a) of this AD. All other reference is correct. The AD currently requires modifying the electrical power generation system. This action corrects the AD to reflect the right date for Fairchild SB 227-24-008 throughout the entire document.
97-08-06 R1: This document clarifies information in Airworthiness Directive (AD) 97-08-06 that applies to Louis L'Hotellier S.A. (L'Hotellier) ball and swivel joint quick connectors installed on gliders and sailplanes that are not equipped with a "Uerling" sleeve or an LS-safety sleeve. These connectors allow the operator of the gliders and sailplanes to quickly connect and disconnect the control systems during assembly and disassembly for storage purposes. AD 97-08-06 currently requires enlarging the safety pin guide hole diameter, and fabricating and installing a placard that specifies the requirement of securing the control system connectors with safety wire, pins, or safety sleeves prior to each flight. The actions specified in that AD are intended to prevent the connectors from becoming inadvertently disconnected, which could result in loss of control of the sailplane or glider. This document clarifies the applicability and modification instructions of AD 97-08-06 by including additional instructions to accomplish the same actions. This correction of the AD results from several operators expressing uncertainty about the applicability and modification instructions.
2011-18-12: We are superseding an existing airworthiness directive (AD) for the specified ECF model helicopters. This AD results from a mandatory continuing airworthiness information (MCAI) AD issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Union. The MCAI AD states that some cracks have been discovered in the spar of the upper fin on Model AS355N helicopters. Due to the fin design similarity between AS350 and AS355 helicopters, this AD action applies to both helicopter models. Modifying the upper and lower fin attachment is intended to prevent failure of a spar, loss of a fin, a separated fin hitting a rotor, and subsequent loss of control of a helicopter.
2006-11-19: We are adopting a new airworthiness directive (AD) for all DORNIER LUFTFAHRT GmbH Models 228-100, 228-101, 228-200, 228-201, 228- 202, and 228-212 airplanes. This AD requires you to repetitively inspect the wiring in the flight deck overhead panels (locations 5VE and 6VE) for chafing and damage and repair any chafed or damaged wires. Regardless of the results of each inspection, this AD requires you to assure correct installation of the wiring in the flight deck overhead panels by reattaching or replacing the wire tie attachment holders and securing any loose wires to the wire tie attachment holders with plastic wire ties. This AD results from mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. We are issuing this AD to detect, correct, and prevent chafed or damaged wires in the flight deck overhead panels, which could result in short-circuiting of related wiring. This condition could lead to electrical failure of affected systems and potential fire in the flight deck.
78-05-09: 78-05-09 BRITISH AIRCRAFT CORPORATION: Amendment 39-3154. Applies to BAC 1- 11, Model/Types 203/AE, 204/AF, 212/AR, 215/AU, 201/Z/AC, 401/AK, 410/AQ, 419/EP, and 412A/EB airplanes. Compliance is required as indicated. To prevent destruction of the fuel boost pump electrical connector accomplish the following: (a) Within the next 300 hours airplane time in service after the effective date of this AD, unless already accomplished, comply with the following: (1) Permanently install a placard at all cockpit fuel boost pump circuit breakers, using white letters on a red background, to read as follows: DO NOT RESET IN FLIGHT (2) Insert in the limitations section of the Airplane Flight Manual the statement: "Do not reset fuel boost pump circuit breakers until the affected boost pump electrical connector has been inspected in accordance with paragraph 2.3.7 of BAC Alert Service Bulletin 28-A-PM 5220, Issue 4, dated October 22, 1975, or an FAA-approved equivalent." (3) On airplanes equipped with a center fuel tank, permanently, install a placard in the cockpit next to the center fuel tank transfer pumps, using white letters on a red background, to read as follows: TURN TRANSFER PUMP OFF WHEN L.P. LIGHT ILLUMINATES (4) For airplanes equipped with a center fuel tank, insert in the limitations section of the Airplane Flight Manual the statement: "Turn center tank fuel transfer pumps "off" when the L.P. light illuminates". (b) For airplanes equipped with fuel boost pump and housing TRW P/N 249000-6 or Plessey P/N 570-8-21225 (which includes boost pump housing connector assembly TRW P/N 209237-1 or Plessey P/N 570-1-21227), comply with the following: (1) Prior to accumulation of 6,000 hours total time in service on the fuel boost pump housing connector assembly, or within the next 100 hours airplane time in service after the effective date of this AD, whichever occurs later, unless already accomplished within the preceding 100 hours airplane time in service, and thereafter at intervals not to exceed 200 hours airplane time in service from the last inspection, until compliance with paragraph (g) of this AD, comply with paragraphs (d), (e) and (f) of this AD. (2) Prior to the accumulation of 6,000 hours total time in service on the fuel boost pump housing connector assembly, or within the next 1,500 hours airplane time in service after the effective date of this AD, whichever occurs later, unless already accomplished within the preceding 1,500 hours airplane time in service, and thereafter at intervals not to exceed 6,000 hours airplane time in service from the previous compliance, comply with paragraph (g) of this AD. (c) For airplanes equipped with fuel boost pump and housing, TRW P/N 249000-7 or Plessey P/N 570-8-23456 (which includes boost pump housing connector assembly TRW P/N 209237-1 or Plessey P/N 570-1-21227), comply with the following: (1) Prior to accumulation of10,000 hours total time in service on the fuel boost pump housing connector assembly, or within the next 100 hours airplane time in service after the effective date of this AD, whichever occurs later, unless already accomplished within the preceding 100 hours airplane time in service, and thereafter at intervals not to exceed 200 hours airplane time in service from the last inspection, until compliance with paragraph (g) of this AD, comply with paragraphs (d), (e), and (f) of this AD. (2) Prior to the accumulation of 10,000 hours total time in service on the fuel boost pump housing connector assembly, or within the next 1,500 hours airplane time in service after the effective date of this AD, whichever occurs later, unless already accomplished within the preceding 1,500 hours airplane time in service, and thereafter at intervals not to exceed 10,000 hours airplane time in service from the previous compliance, comply with paragraph (g) of this AD. (d) Visually inspect all fuel boost pumps on the airplane by operating them for 10 minutes and inspecting for signs of fuel leakage at the fuel pump conduit drains in accordance with paragraph 2.3.3 of BAC Alert Service Bulletin 28-A-PM 5220, Issue 4, dated October 22, 1975, (hereinafter referred to as BAC ASB), or an FAA-approved equivalent. (e) If, during any inspection required by this AD, fuel leakage is detected, identify the pump causing the leak and visually inspect its housing connector assembly, TRW P/N 209237-1 or Plessey P/N 570-1-21227, within the boost pump housing, and its associated fuel boost pump electrical connector for evidence of arcing, overheating, or deterioration of the 'O' ring in accordance with paragraph 2.3.7 of BAC ASB, or an FAA-approved equivalent. (f) Where evidence of arcing, overheating, or deterioration of the 'O' ring seal is found in either the boost pump housing or the electrical connector as a result of an inspection required by this AD, replace the affected partswith serviceable parts before further flight, except that the airplane may be flown in accordance with FAR Sections 21.197 and 21.199 to a base where the work can be performed, provided that the affected pumps are isolated by pulling and clipping the affected pump circuit breaker and placarding the affected pump ON-OFF switch in the cockpit third crewman's panel with the word, INOPERATIVE. (g) Disassemble the electrical connector assembly, perform the inspections, and replace the appropriate parts in accordance with paragraphs 2.3.10, 2.3.11, except for the replacement of "thin" nuts with "full" nuts, and 2.3.12 of BAC ASB, or an FAA-approved equivalent. (h) Upon the request of an operator, the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region, c/o American Embassy, APO New York 09667, may approve alternate repetitive inspection intervals and inspection procedures to facilitate continued operations by the operator, if data substantiating suchaction are submitted by the operator. This amendment becomes effective March 23, 1978.
69-26-01: 69-26-01 SHORT BROTHERS AND HARLAND LIMITED: Amdt. 39-893. Applies to Model SC-7, Series 2 and 3 Airplanes. Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished, replace the flap shroud to wing attachment lugs located along the bottom side of the wing with redesigned lugs in accordance with Short Brothers and Harland Service Bulletin No. 57-2, Series 2, dated 1 July 1969, or Service Bulletin No. 57-51, Series 3, dated 1 July 1969, or an FAA-approved equivalent. This amendment becomes effective December 16, 1969.
91-18-06: 91-18-06 BRITISH AEROSPACE: Amendment 39-8009. Docket No. 91-NM-81-AD. Applicability: All Viscount Model 744, 745D, and 810 series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent reduced structural integrity of the landing flaps, accomplish the following: (a) Within 180 days after the effective date of this AD, or prior to the accumulation of 500 landings after the effective date of this AD, whichever occurs first, and thereafter at intervals not to exceed 360 days, perform non-destructive testing (NDT) inspections to detect cracks in the wing flap guide rails, and to detect corrosion at the abutment face of flap guide rails and flap end ribs on all landing flaps on the left and right wings, in accordance with British Aerospace Preliminary Technical Leaflet (PTL) No. 301, Issue 2, dated November 2, 1989, or PTL No. 170, Issue 2, dated November 2, 1989, as applicable. (b) If cracks are found, prior to further flight, repair in a manner approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (c) If corrosion is found, prior to further flight, remove the guide rail from the landing flap end rib for visual confirmation, rectification rework, and reprotection of both the guide rail and flap end rib abutment surfaces in accordance with paragraph 2.3.1 of British Aerospace Preliminary Technical Leaflet (PTL) No. 301, Issue 2, dated November 2, 1989, or PTL No. 170, Issue 2, dated November 2, 1989, as applicable. (1) Local corroded areas must be blended out to a maximum depth of 0.06 inch. The blended areas should extend beyond the corroded area by a minimum of 0.2 inch where the depth of corrosion is less than 0.03 inch, and a minimum of 0.3 inch where the depth of corrosion is greater than 0.03 inch, but must not exceed 70 percent of the local abutment face width. (2) Following blending, perform a dye penetrant inspection of the abutment surfaces to ensure that all traces of corrosion have been removed. (3) Following repair, and prior to reinstallation of the guide rail, apply protective treatment in accordance with paragraph 2.5 of the appropriate PTL. (4) If corrosion found is in excess of the limitations specified in the appropriate PTL, prior to further flight, repair in a manner 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. NOTE: The request should 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 airplanes to a base in order to comply with the requirements of this AD. (f) The inspection and repair requirements shall be done in accordance with British Aerospace Preliminary Technical Leaflet (PTL) No. 301, Issue 2, dated November 2, 1989, or PTL No. 170, Issue 2, dated November 2, 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-8009, AD 91-18-06) becomes effective on September 23, 1991.
77-03-02 R2: 77-03-02 R2 BOEING: Amendment 39-2826 as amended by Amendment 39-2852 is further amended by Amendment 39-4572. Applies to Boeing Model 727 Series airplanes certificated in all categories that have engine forward fuel feed hose assemblies which have accumulated 5 years or more or 12,000 hours or more time in service whichever occurs first. Compliance required as indicated. \n\tTo prevent rupture of the engine fuel feed hose assemblies, accomplish the following: \n\tA.\tWithin the next 60 days, unless already accomplished within the past six months, inspect and replace as required with a like part or an equivalent hose assembly approved by the Manager, Seattle Aircraft Certification Branch, FAA, Northwest Mountain Region, the No. 1 and No. 2 and No. 3 engine forward fuel feed hose assemblies in accordance with Boeing S.B. 727-28-51, Figure 3, pages 26 and 27, Steps 1, 2, 3, and 4 issued 11/12/76, or later FAA approved revision. \n\tB.\tWithin 3,000 hours time in service, install hose clamps on Engine No. 1 and Engine No. 3 forward fuel feed hose assemblies in accordance with Boeing Service Bulletin 727-28-51, Figure 4, pages 28 and 29, steps 1, 2, and 4 issued November 12, 1976, or later FAA approved revision. \n\tC.\tAnnually reinspect all hose assemblies not replaced per paragraph A above or at an interval that is compatible with the airlines' inspection schedules and approved by the assigned FAA Principal Maintenance Inspector with prior approval of the Manager, Seattle Aircraft Certification Branch, FAA Northwest Mountain Region. \n\tD.\tIf all hose assemblies are replaced per paragraph A., above, with corrosion resistant steel (CRES) braided hoses, the repetitive inspections required by paragraph C., above, may be discontinued as the installation of CRES braided hoses is considered terminating action for this AD. \n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5U.S.C. 552(a)(1). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA, Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tAmendment 39-2826 became effective February 4, 1977. \n\tAmendment 39-2852 became effective March 30, 1977. \n\tThis Amendment 39-4572 becomes effective March 9, 1983.
2022-02-03: The FAA is superseding Airworthiness Directive (AD) 2013-26-01 for all CFM International, S.A. (CFM) CFM56-3 and CFM56-7B model turbofan engines with a certain accessory gearbox assembly (AGB) not equipped with a dynamic oil seal assembly in the handcranking pad. AD 2013-26-01 required an independent inspection to verify re-installation of the handcranking pad cover after removal of the pad cover for maintenance. This AD was prompted by a dual engine loss of oil event and 42 prior events of total loss of engine oil during flight. This AD requires independent inspection to verify re-installation of the AGB handcranking pad cover after maintenance. This AD also requires the replacement of the affected AGB as a terminating action to the inspection requirement. The FAA is issuing this AD to address the unsafe condition on these products.
97-13-11: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 97-13-11, which was sent previously to known U.S. owners and operators of certain Ayres Corporation (Ayres) S2R series airplanes. This AD requires inspecting the 1/4-inch and 5/16-inch bolt hole areas on the lower spar caps for fatigue cracking, and replacing any lower spar cap if fatigue cracking is found. This AD results from an accident on an Ayres S2R series airplane where the wing separated from the airplane in flight. The actions specified by this AD are intended to prevent fatigue cracking of the lower spar caps, which, if not detected and corrected, could result in the wing separating from the airplane with consequent loss of control of the airplane.
80-14-09: 80-14-09 EAGLE BALLOONS LTD. (SEMCO): Amendment 39-3825. Applies to Semco Hot Air Balloon, Model TC-4A, S/N SEM 81 and subsequent; Model T, S/N SEM 78 and subsequent Challenger, S/N SEM 25 and subsequent, equipped with tubular aluminum gondolas covered with chair duck canvas. \n\n\tCompliance required as indicated below after the effective date of this AD. \n\n\tTo preclude failure of the gondola structural fittings and to alter the gondola chair duck canvas, accomplish the following: \n\n\t1.\tBefore next flight, and each flight thereafter: \n\n\t\ta.\tVisually check all Diamond aluminum fittings for cracks, in the tongue radius area, on the following models: \n\n\n\nModel No.\n\nFitting \n\nDash No.\nPart\nQty. \n\n11\nNo. 150\n4 \n\n14\nNo. 156\n8 \n\n28\nNo. 115\n4 \nModel TC-4A Dug. No. 1\n4\nNo. 103\n8 \n\n4\nNo. 150\n4 \n\n5\nNo. 156\n8 \n\n3\nNo. 115\n4 \nModel T Dug. No. 1\n6\nNo. 103\n8 \n\n3\nNo. 150\n4 \n\n6\nNo. 156\n8 \n\n9\nNo. 115\n4 \nChallenger Dug. No. 1\n11\nNo. 1038\n\n\t\tb.\tReplace cracked parts with new parts before next flight. \n\n\t2.\tSecure the gondola chair duck canvas siding to the gondola floor using grommets in the lower portion of the canvas. Extend the existing canvas using a 3/8" french fell seam per Advisory Circular 43.13-1A, Chapter 3, Page 85. Hem the bottom of the canvas and install grommets as noted in sketch below. \n\nAD 80-14-09 \n\n\n\n\nRework existing plywood floor as shown below. \n\n\nAD 80-14-09 \n\n\nLace canvas to plywood floor using 1/4" diameter braided nylon line as shown. \n\n\nAD 80-14-09 \n\n\n\t3.\tWithin the next 100 hours or next annual inspection, whichever occurs first, accomplish the following: \n\n\t\ta.\tRemove Diamond aluminum slip-on fittings noted in paragraph 1.a. \n\n\t\tb.\tClean surfaces as necessary and visually inspect for cracks by dye penetrant with a glass of at least 10 power, or equivalent, particularly in the tongue radius area. \n\n\t\tc.\tIf no cracks are found, the Diamond aluminum slip-on fittings may be returned to service. \n\n\t\td.\tReplace cracked parts with unused parts prior to next flight. \n\n\t4.\tThe repetitive inspection in paragraph (3) is to be accomplished at intervals not to exceed 100 hours in service or annually thereafter, whichever occurs first. \n\n\t5.\tEquivalent inspections, alterations and replacement parts must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. \n\n\t6.\tUpon submission of substantiating data by an owner, or operator through an FAA Maintenance inspector, the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, may adjust the compliance times specified in this AD. \n\n\tThis amendment is effective July 7, 1980. \n\nAD 80-14-09
75-10-05: 75-10-05 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-2198 as amended by Amendment 39-2255. Applies to certain Models and serial numbers of TPE331 series engines as follows: Model Serial Number Effectivity TPE331-1-101B P-93058 through P-93064 TPE331-1-151A P-92249, P-92336 through P-92356 TPE331-1-151K P-91194, P-91195, P-26001 through P-26022 TPE331-1-151G P-91193, P-91196 through P-91198 TPE331-2-201A P-90281 through P-90294, P-90296 TPE331-3U/3UW-303G P-03108, P-03109, P-03112 through P-03193, P-03195, P-05031 through P-05048 TPE331-3U-307G/-303G P-03001, P-03009 TPE331-5-251C P-22006 through P-22103 TPE331-5-251K P-06113, P-06190 through P-06442, P-06444 through P-06537 TPE331-6-251M/-252M P-20144, P-20182 through P-20577 TPE331-6-252B P-27001, P-27002 In addition to the above serial number engines, also affected by this AD are any Model TPE331-1, -2, -3, -5, or -6 series engines which have been modifiedin accordance with AiResearch Service Bulletin TPE331-72- 0064, dated February 1, 1974 or subsequent FAA-approved revision. Compliance required as indicated. To detect, correct and prevent loosening of the torque sensor assembly mounting arm, accomplish the following: (a) Engines with less than 200 hours total time in service since new or last overhaul: Before exceeding 200 hours total time in service, or an additional 25 hours time in service after the effective date of this AD, whichever occurs later, inspect the fuel control assembly drive in accordance with the instructions contained in paragraph 2.C. of AiResearch Service Bulletin TPE331-72-0095, dated April 18, 1975, or later FAA-approved revision. If the fuel pump and control drive backlash exceed the limits specified in the above referenced service bulletin, modify the torque sensor assembly mounting arm per the instructions contained in paragraph 2.D. of that bulletin before further flight. (b) Engines with more than 200 hours total time in service since new or last overhaul: Before exceeding an additional 100 hours time in service after the effective date of this AD, inspect fuel control assembly drive, and modify the torque sensor assembly mounting arm per paragraph (a), above, as necessary. (c) If the fuel control assembly drive meets the backlash limits specified in the above referenced service bulletin, the inspections required in paragraphs (a), and (b), above, must be repeated at intervals not to exceed 400 hours thereafter until the torque sensor assembly mounting arm is modified per paragraph 2.D of the above referenced service bulletin. This modification must be incorporated before the affected engine exceeds the time in service since new or overhaul at the manufacturer's recommended mid-term inspection as defined in paragraphs 2.A., 2.C., or 2.D. of the Revision No. 7, dated March 20, 1975, or later revisions, of AiResearch Service Bulletin No. 606; or, if this mid-term inspection has already been accomplished prior to the effective date of this AD, before exceeding the recommended overhaul period as defined in the Service Bulletin No. 606, Revision 7, dated March 20, 1975, as revised. NOTE: Engine Models TPE331-6-252B and -252M are not specifically included in the above referenced Service Bulletin No. 606. Refer to paragraph 2.D of this service bulletin for mid-term inspection and overhaul times applicable to these engine models. (d) Equivalent procedures may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiation data. (e) Aircraft may be flown to a base for performance of inspections as required by paragraphs (a) and (b) above per FAR's 21.197 and 21.199. Amendment 39-2198 became effective May 15, 1975. This amendment 39-2255 becomes effective July 11, 1975.
97-14-06: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing 747 series airplanes, that currently requires replacement of certain fuse pins on the upper link of the inboard and outboard struts. That AD also requires inspections to detect corrosion or cracks of certain fuse pins, and replacement, if necessary. This amendment reduces the compliance times of actions associated with certain fuse pins and provides for optional terminating action for the requirements of this AD. This amendment is prompted by a report of fracturing of a bulkhead style fuse pin located in the inboard strut at the forward end of the upper link. The actions specified in this AD are intended to prevent failure of the strut and separation of an engine from the airplane due to fracturing of the fuse pins.
2012-01-10: We are adopting a new airworthiness directive (AD) for General Electric Company (GE) CF34-10E series turbofan engines. This AD was prompted by a report of heavy wear found on the seating surface of the center vent duct (CVD) (commonly referred to as center vent tube) support ring and on the inside diameter of the fan drive shaft at the mating location. This AD requires removing from service all CVD support assemblies and any fan drive shaft on the affected engines if wear is found on either the CVD support ring or the fan drive shaft. We are issuing this AD to prevent fan drive shaft failure, leading to uncontained engine failure and damage to the airplane.
93-09-02: 93-09-02 AIRBUS INDUSTRIE: Amendment 39-8569. Docket 92-NM-226-AD. Applicability: Model A300, A310, and A300-600 series airplanes, as listed in Airbus Industrie All Operator Telex (AOT) 29-07, dated August 28, 1992; airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent jamming of the hydraulic fire shut off valve (FSOV) actuator, which could result in the inability to isolate hydraulic fluid from an engine fire, accomplish the following: (a) For airplanes having any hydraulic FSOV that has not ever been functionally tested in accordance with Maintenance Planning Document (MPD) Task Number 29-11-33-0503-1 (for Model A300 series airplanes) or MPD Task Number 29-11-33-01-1 (for Model A310 and A300-600 series airplanes): Within 450 hours time-in-service after the effective date of this AD, perform a one-time operational test to detect potential bearing failure of the electric motor that drives thehydraulic FSOV actuator, and to detect FSOV malfunction, in accordance with Airbus Industrie All Operator Telex (AOT) 29-07, dated August 28, 1992. (1) If any FSOV fails the operational test, prior to further flight, replace that FSOV in accordance with the AOT. (2) If a FSOV passes the operational test, no further action is necessary for that FSOV. (b) For airplanes having any hydraulic FSOV that has been functionally tested only once in accordance with MPD Task Number 29-11-33-0503-1 (for Model A300 series airplanes) or MPD Task Number 29-11-33-01-01 (for Model A310 and A300-600 series airplanes): Within 900 hours time-in-service after the effective date of this AD, perform a one-time operational test to detect potential bearing failure of the electric motor that drives the hydraulic FSOV actuator, and to detect FSOV malfunction, in accordance with Airbus Industrie All Operator Telex (AOT) 29-07, dated August 28, 1992. (1) If any FSOV fails the operational test, prior tofurther flight, replace that FSOV in accordance with the AOT. (2) If a FSOV passes the operational test, no further action is necessary for that FSOV. (c) For airplanes having any hydraulic FSOV that has been functionally tested two or more times in accordance with MPD Task Number 29-11-33-0503-1 (for Model A300 series airplanes) or MPD Task Number 29-11-33-01-01 (for Model A310 and A300-600 series airplanes): No further action is required by this AD. (d) 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. (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 tests shall be done in accordance with Airbus Industrie All Operator Telex (AOT)-29-07, dated August 28, 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 Airbus Industrie, 1 Rond Point Maurice Bellonte, 31707 Blagnac Cedex, France. 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. (g) This amendment becomes effective on June 9, 1993.
75-17-05 R1: 75-17-05 R1 BRITISH AIRCRAFT CORPORATION: Amendment 39-2303 as amended by Amendment 39-4197. Applies to BAC Model 1-11, 200 Series airplanes. Compliance is required as indicated. To detect and repair cracks in the nose landing gear sustaining ram, accomplish the following: (a) Within the next 500 landings after the effective date of this AD or before the accumulation of 20,500 landings, whichever occurs later, unless already accomplished within the last 1,000 landings, and thereafter at intervals not to exceed 1,500 landings from the last inspection, inspect the nose landing gear sustaining ram P/N AB44A-1399, for cracks using a permanent magnet or suitable electro-magnetic process in accordance with paragraph 2.1.1 of British Aerospace, Aircraft Group, Alert Service Bulletin 32-A-PM5070, Issue No. 3, dated March 29, 1979 (hereinafter referred to as the Service Bulletin), or an FAA-approved equivalent. (b) If cracks are found during an inspection required by paragraph (a) of this AD which exceed the limits specified in paragraph 2.1.7 of the Service Bulletin, before further flight, replace the nose landing gear sustaining ram, P/N AB44A-1399, with a serviceable part of the same part number. (c) If cracks are found during an inspection required by paragraph (a) of this AD which do not exceed the limits specified in paragraph 2.1.7 of the Service Bulletin, before further flight, either - (1) Replace the nose landing gear sustaining ram, P/N AB44A-1399, with a serviceable part of the same part number; or (2) Remove the cracks in accordance with instructions in paragraph 2.1.4 of the Service Bulletin, or an FAA-approved equivalent, and reprotect the ram in accordance with the instructions in either paragraph 2.1.5 or 2.1.6 of the Service Bulletin, or an FAA-approved equivalent. (d) For the purpose of complying with this AD, subject to acceptance by the assigned FAA Maintenance Inspector, the number of landings may be determined bydividing each airplane's hours' time in service by the operator's fleet average time from takeoff to landing for the BAC 1-11 airplane. (e) If an equivalent means of compliance is used in complying with this AD, that equivalent means must be approved by the Chief, Aircraft Certification Staff, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, Brussels, Belgium. 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 British Aerospace, Inc., Box 17414, Dulles International Airport, Washington, D.C. 20041. These documents may be examined at FAA Headquarters, Room 916, 800 Independence, SW., Washington, D.C. 20591. Amendment 39-2303 became effective September 1, 1975. This amendment 39-4197 becomes effective September 8, 1981.
97-13-02: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 97-13-02, which was sent previously to all known U.S. owners and operators of Diamond Aircraft Industries, Inc. (Diamond) Model DA 20-A1 airplanes. This AD requires fabricating and installing a placard and inserting limitations into the airplane s flight manual limitations section prohibiting spin maneuvers until a modification is installed. This AD results from an occurrence where a pilot s shoe jammed between the rudder control pedal and the firewall during a spin recovery in a Canadian registered HOAC-Austria Model DV 20 KATANA airplane. The actions specified by this AD are intended to prevent the pilot s shoe from becoming jammed between the rudder pedal and firewall which could result in loss of control of the airplane.
77-12-03: 77-12-03 HAWKER SIDDELEY AVIATION, LTD: Amendment 39-2920. Applies to Model DH-114 airplanes, certificated in all categories, that are not equipped with engine mount stay struts marked as Issue 2 or subsequent or marked "DOI No. 68640." Compliance is required within the next 50 hours time in service after the effective date of this AD, unless already accomplished. To detect internal corrosion of the engine mount stay strut, P/N 14 EM.337A, accomplish the following: (a) Remove the stay strut from the aircraft and inspect in accordance with the X-ray procedure of Appendix I of Hawker Siddeley Technical News Sheet No. E.8, dated March 19, 1973, or an FAA-approved equivalent. (b) Replace any strut for which internal corrosion is detected by the inspection required by paragraph (a) of this AD with a new part or otherwise serviceable spare part of the same part number determined to be free of internal corrosion by the inspection required by paragraph (a) of this AD oran FAA-approved equivalent. (c) It is requested that copies of the X-rays, or results of an FAA-approved equivalent inspection, be forwarded to the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region, c/o American Embassy, APO New York, N.Y. 09667. (Reporting approved by the Office of Management and Budget under OMB No. 04-RO-174). This amendment becomes effective July 18, 1977.
2010-19-06R1: We are revising an existing airworthiness directive (AD) for the products listed above. This AD was prompted by Turbomeca restoring all or part of the life limits of the affected discs, and European Aviation Safety Agency's (EASA) issuance of AD 2010-0101R2, dated March 24, 2011, to do the same. Turbomeca has introduced a reinforced eddy- current inspection (ECI) which, combined with a revised analysis, allows the life limit of the affected discs to be extended. We are issuing this revision to prevent failure of the gas generator (GG) second stage turbine disc which could result in the release of high energy debris and damage to the helicopter.
97-14-04: This amendment supersedes two existing airworthiness directives (AD), applicable to certain Boeing Model 737 series airplanes, that currently require tests of the main rudder power control unit (PCU) to detect excessive internal leakage of hydraulic fluid, stalling, or reversal, and to verify proper operation of the PCU; and replacement of the PCU with a unit having a different part number, if necessary. This amendment adds requirements for replacement of the PCU and the vernier control rod bolts with newly designed units. This amendment also adds a requirement for leak tests of the PCU, and replacement of the PCU with a serviceable or newly designed unit, if necessary. This amendment is prompted by reports of fracturing of the vernier control rod bolts as a result of the shank of the bolt running into the threads on the nutplate during installation of the rod. The actions specified by this AD are intended to prevent such fracturing, which could result in uncommanded movements of the rudder, and consequent reduced controllability of the airplane.