Results
86-17-02: 86-17-02 ROGERSON HILLER CORP.: Amendment 39-5367. Applies to Model UH-12D, UH-12E, and UH-12E4 series helicopters, including military Models OH-23F and OH-23G, and all those converted in accordance with STC's SH177WE and SH178WE, certificated in any category, equipped with main rotor blade fork P/N 52110-3. Compliance is required as indicated, unless already accomplished. To detect cracks and prevent failures which have occurred in the main rotor blade fork P/N 52110-3 at the outboard tension-torsion bar retention bolt hole, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, verify that each installed P/N 52110-3 fork has a serial number permanently displayed on its outer surface. Remove from service forks found not to be serialized and replace with serialized parts. (b) Perform a daily visual check of P/N 52110-3 forks for cracks in the area of the outboard tension-torsion bar retention bolt hole. Washers and nuts need not be removed for this inspection. This check may be performed by the pilot. NOTE: For the requirements regarding the listing of compliance and method of compliance with this AD in the aircraft's permanent maintenance record, see section 91.173. (c) On forks having 240 or more hours' time in service on the effective date of this AD, within the next 10 hours' time in service, unless already accomplished within the last 90 hours' time in service, and within each 100 hours' time in service thereafter from the last inspection, accomplish the inspection specified in (e). (d) On forks having less than 240 hours' time in service on the effective date of this AD, accomplish the inspection specified in (e) prior to the accumulation of 250 hours' time in service and within each 100 hours' time in service thereafter from the last inspection. (e) Perform a dye penetrant inspection, or FAA-approved equivalent, of the bolt hole and adjacent milled surfaces. For this inspection, remove the nut, washer, and pin. (f) Replace cracked rotor forks with like serviceable parts prior to further flight. (g) Special flight permits may be issued in accordance with sections 21.197 and 21.199 to operate aircraft to a base for the accomplishment of inspections required by this AD. (h) An alternate method of compliance which provides an equivalent level of safety may be approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. This amendment supersedes Amendment 39-3540 (44 FR 50035), AD 79-18-04. This amendment becomes effective August 22, 1986.
2009-12-11: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: Following a refined Finite Element Model (FEM) analysis of the Nose Landing Gear (NLG) actuator fitting installed on the roof panel of the NLG box of all A340-500/-600 aircraft, it has been demonstrated that potential fatigue cracks can be initiated on the NLG actuator fitting flanges. This situation, if not corrected, could lead to inadvertent extension of the NLG which could adversely affect the aircraft's continued safe flight or [could result in] failure to retract the NLG which, in combination with an engine failure, could adversely affect the aircraft's safe take off. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
80-26-06: 80-26-06 ROBINSON HELICOPTER: Amendment 39-3997. Applies to Model R-22 series helicopters certified in all categories, serial numbers 0002 through 0058. Compliance is required as indicated unless already accomplished. To prevent failure of the A093-2 forward tail rotor coupling flex plate and loss of power to the tail rotor, accomplish the following: (a) Prior to further flight from the effective date of this AD, conduct a one time inspection of the A193-2 forward tail rotor coupling flex plate. Inspect in accordance with the Robinson Service Letter No. 6, dated September 24, 1980. Remove from service any flex plate found cracked, bent, nicked, or with loose clamp-up. Replace the removed part with a like serviceable part. (b) Prior to further flight from the effective date of this AD, conduct a one time inspection of the A197-1 tail rotor drive shaft assembly in accordance with the Robinson Service Letter No. 7, dated September 27, 1980. Drive shafts withTIR (Total Indicated Runout) in excess of 0.015 inch or with excessive vibration must be removed from service and replaced with a like serviceable part. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate helicopters to a base for the accomplishment of inspections required by this AD. (d) 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. This amendment becomes effective December 29, 1980.
64-01-02: 64-01-02\tBOEING: Amdt. 668 Part 507 Federal Register January 7, 1964. Applies to All Models 707 and 720 Series Airplanes Listed in Boeing Service Bulletin No. 1788. \n\n\tCompliance required within 6,000 hours' time in service after the effective date of this AD unless already accomplished. \n\n\tInstances of engine burner case burn through have occurred in service resulting in damage to the nacelle and lower strut in which detection of the fire was delayed because there are no fire detector elements in the area on top of the engine. To correct this unsafe condition, the following must be accomplished:\n \n\tInstall additional fire detector elements (continuous or unit type) between the horizontal firewall and the top of the engine in the area between nacelle Stations 160 and 180. This additional detector installation shall be made in accordance with Boeing Service Bulletin No. 1788 dated July 11, 1963, or in accordance with an installation approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region. \n\n\t(Boeing Service Bulletin No. 1788 covers the installation applicable to the various fire detector systems as installed on different groups of airplanes.) \n\n\tThis directive effective February 7, 1964.
82-01-03: 82-01-03 AIRBUS INDUSTRIE: Amendment 39-4287. Applies to Model A300 series airplanes, certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent failure of the aft cargo door frames 67 and 69, accomplish the following: (a) Within the next 550 hours time in service after the effective date of this AD, unless already accomplished within the last 1450 hours time in service, inspect the rear cargo compartment door frame external flanges in accordance with paragraph 2.B.(1) of Airbus Industrie A300 Service Bulletin A300-53-108, Revision 1, dated October 15, 1979 (hereinafter referred to as the Service Bulletin), or an FAA-approved equivalent. (b) If no cracks are found during the inspection required by paragraph (a) of this AD: (1) The airplane may be returned to service; and (2) Repeat the inspection in paragraph (a) of this AD at intervals not to exceed 2000 hours time in service from the last inspection. (c) If, during the inspection required by paragraph (a) of this AD, the total number of cracks found does not exceed 5, or the total length of all cracks found does not exceed 2 inches: (1) The airplane may be returned to service; and (2) Repeat the inspection in paragraph (a) of this AD at intervals not to exceed 550 hours time in service from the last inspection. (d) If, during the inspection required by paragraph (a) of this AD, the total number of cracks found exceeds 5, or the total length of all cracks found exceeds 2 inches, inspect the adjacent internal flange on the forward side of frame 67 and the aft side of frame 69 in accordance with paragraph 2.B.(2)(b) of the Service Bulletin, or an FAA-approved equivalent. (e) If no cracks are found on the internal flanges during the inspection required by paragraph (d) of this AD: (1) The airplane may be returned to service; and (2) Repeat the inspections in paragraphs (a) and (d) of this AD atintervals not to exceed 550 hours time in service from the last inspection. (f) If, during inspection of the internal flanges required by paragraph (d) of this AD, the total number of cracks found does not exceed 5, or the total length of all cracks does not exceed 2 inches: (1) The airplane may be returned to service; and (2) Repeat the inspections in paragraphs (a) and (d) of this AD at intervals not to exceed 200 hours time in service from the last inspection. (g) If, during the inspection of the internal flanges required by paragraph (d) of this AD, the total number of cracks found exceeds 5, or the total length of all cracks exceeds 2 inches, before further flight, except as provided in paragraph (i) of this AD, replace the frame in accordance with paragraph 2, "Accomplishment Instructions," of Airbus Industrie A300 Service Bulletin A300-53-109, Revision 4, dated April 25, 1980, or an FAA-approved equivalent. (h) The inspections and repetitive inspections required by this AD may be discontinued when rear cargo door frames 67 and 69 have been replaced in accordance with Airbus Industrie A300 Service Bulletin A300-53-109, Revision No. 4, dated April 25, 1980, or an FAA- approved equivalent. (i) In accordance with FAR Sections 21.197 and 21.199 the airplane may be flown to a base where the maintenance required by this AD may be accomplished. (j) 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, AEU-100, Europe, Africa and Middle East Office, Federal Aviation Administration, c/o American Embassy, Brussels, Belgium. (k) Upon submission of substantiating data, through an FAA Aviation Safety Inspector, the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, Brussels, Belgium, may adjust the inspection intervals. 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 Airbus Industrie, Airbus Support Division, BP 33, 31700 Blagnac, France. These documents may be examined at FAA Headquarters, Room 916, 800 Independence Avenue, SW., Washington, D.C. 20591. This amendment becomes effective January 25, 1982.
2009-12-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: There has been one case reported of failure of a shaft (tailstock) on an elevator Power Control Unit (PCU), Part Number (P/ N) 390600-1007. Continued actuation of the affected PCU caused damage to the surrounding structure. * * * Each elevator surface has three PCUs, powered by separate independent hydraulic systems, and a single elevator PCU shaft failure may remain dormant. Such a dormant loss of redundancy, coupled with the potential for a failed shaft to produce collateral damage, including damage to hydraulic lines, could possibly affect the controllability of the aircraft. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
81-08-01 R1: 81-08-01 R1 LOCKHEED - CALIFORNIA: Amendment 39-4079 as amended by Amendment 39-4230. Applies to Lockheed Model L-1011-385 series airplanes certificated in all categories. Compliance required as indicated, unless already accomplished. To prevent possible loss of ground directional control resulting from unannunciated anti- skid systems failure or excessive pedal force due to parking brake circuit/supply valve failure, accomplish the following: (a) Commencing ten days from the effective date of this AD, prior to each flight, when "C" hydraulic system (ALT SYS "C") is used for anti-skid wheel braking, perform either a brake system check in accordance with item (1) below or an inspection in accordance with item (2) below: 1. Operations check: (i) Chock wheels and release parking brake. (ii) Pressurize both system B and C brake accumulators using AC motor pumps, then turn off pumps. NOTE: Hydraulic system B and C should not be pressurized while performing this check. (iii) Select alternate system C with brake system select switch on pilot's center instrument panel. (iv) Pump brake pedals. B system brake pressure, as indicated on pilot's center instrument panel gauge, should remain steady, and C system brake pressure should decay as pedals are pumped. If B system pressure decays, the NORM system B brake supply valve has failed and must be repaired prior to further flight. If unable to perform operations check, or if results are inconclusive or suspect, perform visual inspection as specified in item (2): 2. Visual inspection: (i) Pressurize both B and C hydraulic systems. (ii) Select ALT SYS C with brake system select switch on pilot's center instrument panel. (iii) Chock wheels and release parking brake. (iv) Gain access to both brake supply valve modules (Lockheed P/N 672173-119 (contacts normally open) and Lockheed P/N 672173-121 (contacts normally closed)) in hydraulic service center, located above B and C system accumulators in left and right forward corners of service center. (v) Verify that two each indicator pins (located adjacent to electrical connector on each valve) are protruding from both B and C brake supply valves. If both pins are not protruding, the valves have failed and must be repaired prior to further flight. (b) Within 1 year from the effective date of this amendment, modify the parking brake and anti-skid electrical circuits either in accordance with the Accomplishment Instructions of Part 2 of Lockheed Service Bulletin 093-32-145 dated May 1, 1980, or Part 2 of Lockheed Service Bulletin 093-32-179, Revision 1, dated April 7, 1981, or later revisions to either service bulletin approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Region. (c) The pre-flight check/inspections required by paragraph (a) of this AD are no longer required after accomplishment of paragraph (b) of this AD.(d) Alternative means of compliance or other actions which provide an equivalent level of safety may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Region. (e) 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 inspections and/or modifications required by this AD. 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 Lockheed-California Company, P.O. Box 551, Burbank, California 91520, Attention: Commercial Support Contracts, Dept. 63-11, U-33, B-1. These documents also may be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108, or Los Angeles Area Aircraft Certification Office, 4344 Donald Douglas Drive, Long Beach, California 90808. Amendment 39-4079 became effective April 16, 1981. This amendment 39-4230 becomes effective October 18, 1981.
63-27-04: 63-27-04 SUD AVIATION: Amdt. 664 Part 507 Federal Register December 25, 1963. Applies to All Model SE-210 Caravelle I, III and VIR Aircraft. Compliance required within 30 days after the effective date of this AD. To prevent engine flameout or engine damage from the bow wave caused by the nose wheel on flooded or slush covered runways, install Kleber-Colombes NO82E22 or Dunlop 082E22 nose wheel tires, with water deflectors, or FAA approved equivalents, unless a placard is installed in the airplane prohibiting operations on flooded or slush covered runways. This directive effective January 27, 1964.
62-14-06: 62-14-06 VICKERS: Amdt. 457 Part 507 Federal Register June 23, 1962. Applies to Viscount Models 744, 745D, and 810 Series Airplanes. Compliance required as indicated. Because of failures of the main landing gear retraction jack fork ends, the following is required. (a) Fork ends Part Numbers 70150-273, 74450-95, and 74450-411. (1) Fork ends which have not been reworked in accordance with (e) and have accumulated 3,500 or more landings as of the effective date of this AD shall be inspected in accordance with (d): (i) Within the next 50 landings if no inspection has been conducted subsequent to the accumulation of 3,500 landings and thereafter within each 800 landings. (ii) Within the next 50 landings if 750 or more landings have been made since the last inspection conducted subsequent to the accumulation of 3,500 landings and thereafter within each 800 landings. (iii) Prior to the accumulation of 800 landings if less than 750 landingshave been made since the last inspection conducted subsequent to the accumulation of 3,500 landings and thereafter within each 800 landings. (2) Fork ends which have not been reworked in accordance with (e) and have accumulated between 3,450 and 3,500 landings as of the effective date of this AD shall be inspected in accordance with (d) within the next 50 landings and thereafter within each 800 landings. (3) Fork ends which have not been reworked in accordance with (e) and have accumulated less than 3,450 landings as of the effective date of this AD shall be inspected in accordance (d) prior to the accumulation of 3,500 landings and thereafter within each 800 landings. (b) Fork ends Part Numbers 72450-315 and 74450-499. (1) Fork ends which have not been reworked in accordance with (e) and have accumulated 5,750 or more landings as of the effective date of this AD shall be inspected in accordance with (d) within the next 50 landings and thereafter within each800 landings. (2) Fork ends which have not been reworked in accordance with (e) and have accumulated less than 5,750 landings as of the effective date of this AD shall be inspected in accordance with (d) prior to the accumulation of 5,800 landings and thereafter within each 800 landings. (3) If two successive 800 landing inspections are accomplished without evidence of cracks, subsequent inspections may be made at intervals not exceeding 1,600 landings. (c) All fork ends which have been reworked in accordance with (e) shall be inspected in accordance with (d) within each 800 landings after rework, except that reworked fork ends which have accumulated more than 750 landings, as of the effective date of this AD shall be inspected in accordance with (d) within the next 50 landings and each 800 landings thereafter. If two successive 800 landing inspections are accomplished without evidence of cracks, subsequent inspections may be made at intervals not exceeding 1,600landings. (d) Remove and inspect using magnetic particle inspection or FAA-approved equivalent in accordance with British Aircraft Corporation (B.A.C.) Ltd. Preliminary Technical Leaflet (PTL) No. 171 Issue 6 (for 744 and 745D) or later ARB-approved issue; or PTL 31 Issue 6 (for 810) or later ARB-approved issue. Parts showing evidence of cracks shall be replaced or reworked in accordance with paragraph (e) before further flight. (e) Parts showing evidence of cracks may be reworked once in accordance with British Aircraft Corporation (B.A.C.) Ltd. Preliminary Technical Leaflet (PTL) No. 171 Issue 6 (for 744 and 745D) or later ARB-approved issue; or PTL 31 Issue 6 (for 810) or later ARB- approved issue. Any parts showing evidence of cracks after reworking must be rejected. (f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Certification Staff, FAA Europe, Africa, Middle East Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. (g) For the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing each airplane's hours' time in service by the operator's fleet average time from takeoff to landing for the airplane type. This supersedes AD 60-11-11. This directive effective July 24, 1962. Revised March 17, 1966.
62-23-01: 62-23-01 AUTO CRAT MANUFACTURING COMPANY SAFETY BELT: Amdt. 409 Part 507 Federal Register October 26, 1962. Applies to All Aircraft Equipped With Auto Crat Models BN 1-2001, BN 3-1501, BN 3-6001, or BN 3-7001 Series Safety Belts. Compliance required within the next 25 hours' time in service after the effective date of this AD. It has been determined that Auto Crat Models BN 1-2001, BN 3-1501, BN 3-6001, or BN 3-7001 Series safety belts, not intended for aircraft use and which do not fully comply with the requirements of Technical Standard Order C22, have been improperly marked with TSO-C22 identification. Accordingly, Auto Crat belt assemblies identified with these model numbers shall be replaced with belt assemblies that conform to TSO-C22 standards. This directive effective November 19, 1962.
2009-11-04: The FAA is adopting a new airworthiness directive (AD) for Rolls-Royce Corporation (RRC) AE 2100D2, AE 2100D2A, AE 2100D3, and AE 2100J turboprop engines with certain propeller gearbox (PGB) shaft-and- carrier assemblies installed. These engines are U.S. type-certificated but as of the effective date of this AD are only installed on military airplanes. This AD requires monitoring a certain population of PGB shaft-and-carrier assemblies for vibration during flight, and borescope-inspecting the PGB shaft for cracks if vibration is experienced. This AD would also require removing the affected population of PGB shaft-and-carrier assemblies from service and installing serviceable PGB shaft-and-carrier assemblies. This AD results from a report of a crack found in the forward cone of a PGB shaft in an RRC AE 2100D3 turboprop engine that was removed from service due to high vibration. We are issuing this AD to prevent separation of the propeller from the airplane, which could result in injury, and damage to the airplane.
64-11-01: 64-11-01\tBOEING: Amendment 39-763. Applies to models 707 and 720 Series Aircraft. \n\tCompliance required as indicated. \n\n\tCracks have been discovered in the front and rear upper and lower wing spar chords on the 707/720 Series aircraft. Accordingly, in the interest of safety accomplish the following or an equivalent approved by Aircraft Engineering Division, FAA Western Region: \n\n\t(a)\tParagraph (a) Superseded by AD 77-02-01. \n\n\t(b)\tOn all 707 and 720 Series aircraft delivered prior to October 1962, accomplish the following: \n\n\t\t(1)\tOn 707 and 720 Series aircraft which have been inspected in accordance with AD 64-03-02, paragraph (b), visually reinspect thereafter at intervals not to exceed 6,000 hours' time in service for spanwise cracks in the wing upper and lower front spar chords between front spar Station 727 and the production break fittings. \n\n\t\t(2)\tIf a crack is detected, repair the cracked spar chord in accordance with Boeing front spar repair drawing 65-40144 oran Aircraft Engineering Division, FAA Western Region, approved equivalent before further flight. The repetitive inspections required by subparagraph (1) may be discontinued on any chord repaired in accordance with this subparagraph. \n\n\t(c)\tUpon request of an operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\tNOTE. - This AD supplements but does not cancel AD 64-03-02. \n\n\t(Boeing Service Bulletins No. 1964 (R-3) and No. 1964(R-3)A cover this same subject.) \n\n\tThis directive effective May 21, 1964. \n\n\tRevised August 22, 1964. \n\n\tRevised March 5, 1965. \n\n\tRevised May 1, 1965. \n\n\tRevised August 20, 1968. \n\n\tRevised October 29, 1968. \n\n\tRevised March 22, 1969.Revised May 9, 1969.
64-23-01: 64-23-01\tBOEING: Amdt. 819 Part 507 Federal Register October 14, 1964. Applies to Models 707 and 720 Series Aircraft. \n\n\tCompliance required within the next 65 hours' time in service after the effective date of this AD, unless already accomplished. \n\n\t(a)\tWith the rudder full right, inspect the installation of the rudder control link arm assembly P/N 65-12703 to ensure: \n\n\t\t(1)\tthat clearance exists between the rudder control link arm assembly and the lower pivot clevis for that arm assembly; \n\n\t\t(2)\tthat clearance exists between the arm assembly and the rudder spar web; and \n\n\t\t(3)\tthat there is no evidence of previous interference. \n\n\t(b)\tIf clearance does not exist, rework the existing parts in accordance with Boeing Alert Service Bulletins 2052 and 2052A, or in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region, or install new parts which provide clearance. \n\n\t(c)\tIf there is evidence of previous interference, replace the lower arm assembly pivot bolt. \n\n\t(Boeing Alert Service Bulletins 2052 and 2052A cover this same subject.) \n\n\tThis directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated September 25, 1964.
2009-12-15: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: The manufacturer has advised of receiving a report from a G 120A operator of an electrical fire caused by a chafed/scorched cable loom. It has been found that the RH main power distribution cable chafed on the instrument panel combing. It is likely that vibrations made the wiring to chafe. The chafing caused eventually electrical arcing and subsequently an in-flight fire that damaged partially the instrument panel cover. This AD requires actions that are intended to address the unsafe condition described in the MCAI.
2009-12-05: We are adopting a new airworthiness directive (AD) for certain Boeing Model 737-300, -400, and -500 series airplanes. This AD requires modifying the control power wiring of the normal supply fan and the low flow sensor for the equipment cooling system of the electronic flight instrument system (EFIS). This AD results from a report of loss of both the normal EFIS cooling supply and the indication of EFIS cooling loss due to a single failure of the battery bus, causing eventual power-down of the EFIS displays; the standby attitude indication is also powered by this battery bus. We are issuing this AD to prevent loss of all attitude indications from both the standby indicator and EFIS displays, which could decrease the ability of the flightcrew to maintain the safe flight and landing of the airplane.
89-06-04: 89-06-04 McDONNELL DOUGLAS: Amendment 39-6152. \n\n\tApplicability: Model DC-9-10 through -30 series and C-9 (Military) series airplanes, equipped with a non-ventral aft pressure bulkhead, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent cracks which could result in structural failure of the non-ventral aft pressure bulkhead, accomplish the following: \n\n\tA.\tPrior to the accumulation of 25,000 landings or within 500 landings after the effective date of this AD, whichever occurs later, inspect the aft pressure bulkhead attach tee section, in accordance with the following procedures. \n\n\t\t1.\tRemove any sealant from inspection area of the tee section that might hinder optically aided and high frequency eddy current inspections. Clean dirt, grease, and all foreign materials from inspection area using lint-free wipers and 1,1,1 trichloroethane solvent or equivalent; \n\n\t\t2.\tUsing optically aided visual inspection technique, inspect the 5910163-89, -93, -94, and -95 attach tees from the aft side of the bulkhead, in accordance with McDonnell Douglas Alert Service Bulletin A53-231, dated February 24, 1989 (hereinafter referred to as ASB53-231). Repeat this inspection thereafter at intervals not to exceed 1,500 landings; and \n\n\t\t3.\tUsing a high frequency eddy current inspection technique, in accordance with ASB53-231, inspect the 5910163-91 and -92 attach tees from the aft side of the bulkhead. Repeat this inspection thereafter at intervals not to exceed 500 landings. \n\n\tB.\tIf cracks are found prior to further flight, replace the cracked tee cap or repair by splicing in a section of tee cap with a new like or improved part, in accordance with McDonnell Douglas Service Rework Drawings SRO9530001, Revision C, dated August 18, 1987, and SRO9530001, Revision "Advance D", dated October 29, 1987. Prior to the accumulation of 25,000 landings after the repair or replacement, resume the repetitive inspections in accordance with paragraph A., above. \n\n\tC.\tCompliance with the requirements of this AD constitutes terminating action for the requirements of AD 88-13-09, Amendment 39-5954, relating to airplanes equipped with non-ventral aft pressure bulkheads. \n\n\tNOTE: The requirements of AD 88-13-09 relating to airplanes with ventral aft pressure bulkheads are not affected by this AD. \n\n\tD.\tAn 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, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments then send it to the Manager, Los Angeles Aircraft Certification Office. \n\n\tE.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes unpressurized to a base in order to comply with the requirements of this AD. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director of Publications, C1-LOO (54-60). These documents may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or 3229 East Spring Street, Long Beach, California. \n\n\tThis amendment (39-6152, AD 89-06-04) becomes effective March 24, 1989.
89-13-02: 89-13-02 TEXTRON LYCOMING: Amendment 39-6378. Final Rule of Priority Letter AD. Applicability: Textron Lycoming Model HIO-360-D1A engines, certificated in any category, equipped with Bendix Model RSA-7AA Servo Fuel Injectors, P/N 2524347, with the following serial numbers: 70001101 through 70001110 70001401 through 70001404 70001602 through 70001604 70001701 through 70001710 70007001 through 70007110 (except 70007107, 70007108) 70010104 through 70010110 (except 70010107) 70013701 through 70013710 (except 70013706) 70013802 through 70013810 (except 70013803, 70013809) 70013903 through 70013910 (except 70013909) 70014001 through 70014009 (except 70014002, 70014005, 70014007, 70014008) 70014101 through 70014104 (except 70014102, 70014103) Serial numbers followed by the letter "P" are in compliance with this AD. These fuel injectors are installed on, but not necessarily limited to, Schweizer Aircraft Corp. Helicopter Model 269C. Compliance: Required within 5 hours time in service after the effective date of this AD, unless already accomplished. To prevent loss of engine power or engine stoppage, accomplish the following: (a) Remove the fuel injector unit from the engine and replace with an unaffected unit or one that has been repaired in accordance with the accomplishment instructions in Appendix I to this AD. NOTE: Injector units may be returned to Precision Airmotive Corp. or an authorized facility listed in Appendix II to this AD. (b) Install unaffected fuel servo unit or a unit repaired in accordance with Appendix I to this AD, as follows: (1) Disconnect fuel injection lines from fuel injection nozzles. (2) Carefully remove inserts from nozzles and identify to allow the inserts to be reinstalled in the proper location. (3) Use CLEAN shop air to blow out all fuel injection lines, inserts, and nozzle bodies (still installed in cylinder). (4) Visually inspect inserts for cleanliness.(5) Reinstall fuel injector and reconnect fuel lines to fuel injector but do not reconnect to nozzles. (6) Purge fuel injector system by flowing fuel through the lines for 10 seconds with boost pump on, injector in full rich and full throttle. (7) Reinstall inserts in original nozzle locations and torque fuel injector line nuts to 40-50 in.-lbs. (8) Make appropriate mixture and idle adjustments per aircraft manufacturer's applicable Maintenance Manual and/or appropriate service publications. NOTE: Precision Airmotive Service Bulletin (SB) PRS-92, dated May 26, 1989; Precision Airmotive Service Information Letter #31, Revision 1, dated May 26, 1989; Schweizer Aircraft Corp. Special Advisory, dated May 25, 1989; Schweizer Aircraft Corp. Service Notice N-202, dated February 23, 1989; and Textron Lycoming SB No. 485, dated June 1, 1989, contain related information. (c) Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations 21.197 and 21.199 to a base where the AD can be accomplished. (d) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance schedule specified in this AD may be approved by the Manager, New York Aircraft Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, Room 202, 181 South Franklin Avenue, Valley Stream, New York 11581. This amendment (39-6378, AD 89-13-02) becomes effective on November 30, 1989, as to all persons except those persons to whom it was made immediately effective by priority letter AD No. 89-13-02, issued on June 22, 1989, which contained this amendment. APPENDIX I NOTE: Precision Airmotive Corporation Service Information Letter #31, Rev. 1, dated May 25, 1989, pertains to this inspection. APPENDIX II NOTE: List of Precision Airmotive Corporation Product Support Centers, pertains to this AD.
80-09-09: 80-09-09 CESSNA: Amendment 39-3768. Applies to the following airplanes, equipped with Cessna installed dual flight director indicators, unless modified to provide independent power sources, distribution and bus systems for the dual flight director installation or by relocating the copilot's second attitude indicator to the top center position of the copilot panel instrument "T". Serial Numbers 500-0066 500-0303 500-0178 500-0304 500-0184 500-0342 500-0275 500-0343 500-0276 500-0344 500-0278 500-0346 500-0279 500-0359 500-0292 500-0364 500-0293 500-0371 500-0294 500-0370 500-0299 500-0373 500-0300 500-0379 Model 500 Serial Numbers 501-0016 501-0063 501-0024 501-0066 501-0027 501-0067 501-0034 501-0112 501-0052 501-0121 Model 501 Serial Numbers 550-0005 550-0079 550-0010 550-0091 550-0012 550-0095 550-0025 550-0098 550-0032 550-0100 550-0034 550-0106 550-0038 550-0110 550-0044 550-0114 550-0046 550-0116 550-0047 550-0121 550-0056 550-0124 550-0062 550-0126 550-0063 550-0130 550-0064 550-0134 550-0066 550-0136 550-0067 550-0138 550-0069 550-0141 550-0073 550-0152 550-0076 Model 550 Serial Numbers 551-0016 551-0024 Model 551 551-0061 COMPLIANCE: Required as indicated, unless previously accomplished. To preclude the loss of attitude information to the crew, within the next 100 hours time-in-service after the effective date of this AD, accomplish the following: A) Insert the following information in the applicable FAA Approved Airplane Flight Manual, Section II, Operating Limitations, under the heading Dual Flight Director Installation and operate the airplane in accordance with this insertion: "The copilot's second attitude indicating system must be installed, operational, and remain operating throughout theflight for those airplanes equipped with the Dual Flight Director Installation." B) Relocate the copilot's second attitude indicator from the center instrument panel or right side instrument panel if applicable to one of the following positions on the copilot instrument panel. 1) Two instrument positions below the copilot airspeed indicator, or 2) One instrument position below the copilot's altimeter, or 3) Two instrument positions below the copilot's altimeter. C) Use Paragraph A of this AD or a duplicate thereof, as an amendment to the FAA Approved Airplane Flight Manual until replaced by the following applicable FAA Approved Cessna revision: Model Revision No. FAA Approved Date 500 43 April 15, 1980 501 13 April 18, 1980 550 13 April 15, 1980 551 10 April 18, 1980 D) Any equivalent method of compliance with this Airworthiness Directive must be approved by the Chief, Wichita Engineering and Manufacturing District Office, Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. This amendment becomes effective May 12, 1980.
2009-12-03: We are adopting a new airworthiness directive (AD) for certain Boeing Model 757-200, -200CB, and -300 series airplanes. This AD requires a detailed inspection for damage of the wire bundle of the right recirculation fan, and repair if necessary. This AD also requires re-routing the wire bundle of the right recirculation fan. This AD results from a report indicating that, during landing of a Model 757 airplane, an overheat warning and smoke occurred in the main cabin, and the right recirculation fan stopped operating. We are issuing this AD to prevent damage of the wiring bundle of the right recirculation fan. Such damage could result in a short circuit and possible fire in the mix bay or smoke in the main cabin.
2009-12-08: The FAA is adopting a new airworthiness directive (AD) for all Boeing Model 747 airplanes. This AD requires inspecting for cracks in the left- and right-side Stringer 11 longeron adjacent to the horizontal stabilizer pivot bulkhead, and related investigative and corrective actions if necessary. This AD results from a report of a crack found in the right-side Stringer 11 longeron horizontal flange, adjacent to the horizontal stabilizer pivot bulkhead, during a routine maintenance inspection. We are issuing this AD to detect and correct fatigue cracking of the longeron, which can propagate and cause damage to the adjacent horizontal stabilizer pivot bulkhead. This damage could result in loss of structural integrity and consequent inability of the bulkhead to carry flight loads, which could adversely affect controllability of the airplane.
61-21-05: 61-21-05 LOCKHEED: Amdt. 342 Part 507 Federal Register October 3, 1961. Applies to All Model 188 Aircraft. Compliance required as indicated. Several instances of inadvertent tripping of compass circuit breakers have occurred due to bumping the ganging bar by the occupant of the observers seat. Due to the effect of the loss of heading information upon operational safety, accomplishment of the following is required: Within the next 100 hours' time in service after the effective date of this AD, inspect the pilot's and copilot's compass circuit breakers, and if the circuit breakers are ganged together, either remove the ganging bars, or install a protective guard which precludes inadvertent operation of the ganging bar. (These circuit breakers are located in the lowest circuit breaker panel beside the observers seat on the left side of the flight deck.) This directive effective November 2, 1961.
88-23-01: 88-23-01 MITSUBISHI HEAVY INDUSTRIES, LTD. (MHI Amendment 39-6056. Applies to Models MU-2B, MU-2B-10, -15, -20, -25, -26, -26A, -30, -35, -36, -36A, -40, and -60 airplanes (all serial numbers with or without the SA suffix) certificated in any category. Compliance: As indicated in the AD, unless already accomplished. To preclude flap control failure, accomplish the following: (a) On airplanes with more than 4,000 hours Time-in-Service (TIS), within the next 100 hours TIS, and on airplanes with less than 4,000 hours TIS, before reaching 4,100 hours TIS: (1) Inspect torque tube joint Part Number (P/N) 010A-61254, of torque tube assembly P/N 010A-61250 for cracks, and replace a cracked joint with a serviceable unit. If no crack is found, repeat the inspection at intervals not to exceed 100 hours TIS, as follows: (i) For airplanes type certificated in accordance with U. S. Type Certificate (TC) A2PC (those without the serial number suffix SA) as described in MHI MU-2B Service Bulletin (S/B) No. 189B, dated May 27, 1988. (ii) For airplanes type certificated in accordance with U.S. TC A10SW (those airplanes with the serial number suffix SA) as described in MHI MU-2B S/B No. 067/27-008, dated November 16, 1987. (2) Inspect all other torque tube joints listed as "existing joints" in Figure 1 of the applicable S/B for cracks, and replace a cracked joint with a serviceable unit. If no crack is found, repeat the inspection thereafter at intervals not to exceed 500 hours TIS, as follows: (i) For TC A2PC airplanes as described in MHI S/B 189B. (ii) For TC A10SW airplanes as described in MHI S/B No. 067/27-008. (b) Accomplish the inspections and replacement requirements of paragraph (a) of this AD, as applicable, prior to 4,100 hours TIS on each replacement joint listed as an "existing joint," as follows: (1) For TC A2PC airplanes (without S/N suffix SA) in Figure 1 of MHI S/B No. 189B. (2) For TC A10SW airplanes (with S/N suffix SA) in Figure 1of MHI S/B No. 067/27-008. (c) Repetitive inspections of a torque tube joint are terminated upon installation of a joint listed as an "improved joint," as follows: (1) For TC A2PC airplanes (without S/N suffix S/A) in Figure 1 of MHI S/B No. 189B. (2) For TC A10SW airplanes (with S/N suffix SA) in Figure 1 of MHI S/B No. 067/27-008. (d) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (e) An equivalent method of compliance with this AD may be used on the A2PC airplanes, if approved by the Manager, Los Angeles Aircraft Certification Office, ANM-120L, FAA, 3229 East Spring Street, Long Beach, California 90806-2425, Telephone (213) 988-5200, and on the A10SW airplanes, if approved by the Manager, Wichita Aircraft Certification Office, ACE-115W, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209; Telephone (316) 946-4400. All persons affected by this directive may obtain copies of the documentsreferred to herein upon request to the Beech Aircraft Corporation (Licensee to Mitsubishi), P. O. Box 85, Wichita, Kansas 67201; Telephone (316) 681-9111; or may examine these documents at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment, 39-6056, becomes effective on December 4, 1988.
2009-11-11: We are adopting a new airworthiness directive (AD) for certain McDonnell Douglas Model MD-90-30 airplanes. This AD requires installing fuses and wire protection in certain wing and fuel tank spars. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent possible damage to the fuel level float or pressure switch wires. Such damage could become a potential ignition source inside the fuel tank, and, when combined with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
79-17-06: 79-17-06 BEECH: Amendment 39-3532. Applies to Model 76 (Serial Numbers ME-1 through ME-228) Airplanes certificated in all categories. COMPLIANCE: Required as indicated unless already accomplished. To prevent possible collapse of the main landing gear, accomplish the following: A) Within the next twenty-five (25) hours time-in-service after the effective date of this AD, install a new part number MS20392-2C67 clevis pin, common to the main landing gear fork and the fork pin, in the left and right main landing gear in accordance with procedures in Beechcraft Service Instructions Number 1073. NOTE: Installation of the new larger clevis pins requires a 3/16 (.1875) inch diameter carbide tip drill bit, drill press and a suitable drill press vise or "V" blocks and clamps to hold the fork pin while enlarging the clevis pin hole. B) Aircraft may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished. 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 August 27, 1979.
79-07-03: 79-07-03 BOEING: Amendment 39-3443. Applies to all Model 737 series airplanes, except airplanes with the engine pressure ratio (EPR) activated takeoff warning system. Compliance required as indicated. \n\n\tWithin 2,000 hours time in service or one year after the effective date of this AD, whichever comes first, unless already accomplished, set the thrust lever operated switches, S283 and S133, to provide actuation down to and including -65 degrees F in accordance with the applicable part of Boeing Service Bulletin 737-31-1026, or later FAA approved revisions, or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Northwest Region. \n\tThe 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 Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA, Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108.\n \n\tThis amendment becomes effective May 4, 1979.