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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.
60-10-04: 60-10-04 LOCKHEED: Amdt. 138 Part 507 Federal Register May 3, 1960. Applies to All Models 049, 149, 649, 749, 1049, and 1649 Series Aircraft. Compliance required as indicated. A crack was found in the segment ring of the fuselage aft pressure bulkhead. The crack, approximately 37 inches long, was in the top left section of the ring, extending from a point right of center. As a result of investigation of the failure, the following must be accomplished on all aircraft with more than 30,000 hours' time in service. (a) Within the next 180 hours' time in service, unless already accomplished within the last 4,500 hours' time in service, and every 4,500 hours' time in service thereafter, inspect the entire peripheral ring of the aft pressure bulkhead at the bend radius adjacent to the skin using one of the following methods or equivalent: (1) Radiographic inspection. (2) Pressurize the cabin to a minimum of 2 p.s.i. Apply soap solution to the rear face of thering and examine for leakage. This will require removal of the sealing compound. (b) If cracks are found, they must be repaired in accordance with FAA approved manufacturer's instructions. Pressurized flight is prohibited until cracks are repaired. (c) When any part of the aft pressure bulkhead peripheral ring is replaced by a new part, inspection of the new part per (a) is not required until the new part has accumulated 30,000 hours' time in service. (d) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, 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. (Lockheed Service Letter FS/240954 covers this subject.) Revised July 11, 1961. Revised March 9, 1962.
2009-11-10: We are adopting a new airworthiness directive (AD) for Eurocopter Deutschland GmbH (Eurocopter) Model EC135 helicopters. This AD results from a report of abnormal main rotor blade vibrations on a Eurocopter Model EC135 helicopter. This AD also results from mandatory continuing airworthiness information (MCAI) issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. The MCAI states that an operator reported unusual vibrations during the start phase of the main rotor blade on one helicopter. The vibrations stopped after the application of torque. Subsequent maintenance personnel found that six of the eight attachment screws of the lower hub-shaft bearing support were loose. This condition was discovered in two additional helicopters. Loose screws in the bearing support, if not detected and corrected, could result in abnormal main rotor blade vibrations and subsequent damage to the main transmission.
2009-11-07: 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: Resulting from the assessment of fuel tank wiring installations required by SFAR 88 (Special Federal Aviation Regulation) and equivalent JAA/EASA (Joint Aviation Authorities/European Aviation Safety Agency) policy, BAE Systems (Operations) Limited has revised the HS.748 Aircraft Maintenance Manual (AMM), now at Revision 19, to introduce Chapter 05-10-00 ``Critical Design Configuration Control Limitations (CDCCL)--Fuel System''. The CDCCLs provide instructions to retain critical ignition source prevention features during configuration changes that may be caused by modification, repair or maintenance actions. The CDCCLs have been identified as requirements for continued airworthiness to address the risk of fuel vapour ignition sources remaining undetected. This condition, if not corrected, could result in a fuel tank explosion and consequent loss of the aircraft. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
61-18-01: 61-18-01 BOEING: Amdt. 326 Part 507 Federal Register August 24, 1961 as amended by amendment 39-1002. Applies to Boeing 707 Series Aircraft Serial Nos. 17586-17652, 17658- 17690, 17692-17712, 17718-17724, 17903-17906, 17918-17919, 17925-17930, 18012 and 18054 and Boeing 720 Series Aircraft Serial Nos. 17907-17917, 18013-18020, 18023, and 18041 as indicated. \n\n\tDue to failure in a main landing gear trunnion support, the following inspections, contained in paragraphs (a), (b), and (c), are required on all specified 707 Series aircraft until paragraph (d) has been accomplished. Paragraph (e) is required on all specified 707 Series aircraft after paragraph (d) is accomplished and on all specified 720 Series aircraft. \n\n\t(a) Within the next 200 hours' time in service, unless already accomplished within the last 150 hours' time in service, and thereafter at every 350 hours' time in service: \n\n\t\t(1) Clean the web, upper and lower chord areas and aft flanges on the inboard and outboard side of 220,000 p.s.i. heat treat steel main landing gear trunnion support rib and conduct a visual inspection of the cleaned areas for evidence of cracks.\n \n\t\t(2) Clean and visually, or radiographically, inspect the forward trunnion support fitting for evidence of cracks on the forward and aft side in the region of the bearing collar. \n\n\t\t(3) If evidence of cracks is found in (a)(1) or (a)(2) above, conduct more detail inspections using fluorescent dye penetrant at temperatures of 50 degrees F. or above, X-ray, or equivalent. \n\n\t(b) Within the next 35 hours' time in service, unless already accomplished within the last 30 hours' time in service and thereafter at every 125 hours' time in service, clean the main landing gear aft trunnion bearing support, paying particular attention to the areas listed below, and conduct a fluorescent dye penetrant inspection or equivalent for cracks; \n\n\t\t(1) Area around the barrel nut hole, both forward and aft sides. \n\n\t\t(2) A strip 1/2-inch wide around upper bearing support, from the upper barrel nut to lower 1.31 diameter inboard (tension) bolt hole, on aft side. \n\n\t\t(3) A strip 1/2-inch wide around upper bearing, from upper barrel nut to trunnion support rib, on forward side. \n\n\t(c) If cracks exceeding allowable lengths specified in the latest revision of Boeing Service Bulletin No. 859 (R-2 or later) are found during inspections (a) and (b), the affected components must be replaced or repaired in accordance with FAA approved Boeing procedures prior to further flight. When cracks less than the maximum allowable lengths specified in S.B. 859 (R-2 or later), are found, the following shall be accomplished: \n\n\t\t(1) Stop drill in accordance with S.B. 859 (R-2 or later) instructions and inspect for crack progression at each 350 hours' time in service after stop drilling. If cracks progress beyond the stop drilled hole, contact Boeing for FAA approved Boeing repair procedures to be incorporated prior to further flight. \n\n\t\t(2) If cracks are not accessible for stop drilling prior to further flight FAA approved Boeing instructions must be obtained for the required inspection intervals and procedures for the specific crack location and length. \n\n\t(d) Unless already accomplished at the factory or by the operator within the next 3,500 hours' time in service conduct the following detail inspections and rework as indicated: \n\n\t\t(1) Remove the main landing gear and trunnion in accordance with BAC Maintenance Manual Procedure. \n\n\t\t(2) Remove all nuts and washers along the periphery of the trunnion support rib. \n\n\t\t(3) Rework the main landing gear trunnion support fittings per the latest revision of FAA approved S.B. No. 874 (August 9, 1960, or later). \n\n\t\t(4) Following the rework, clean the aft and forward trunnion support fittings and perform a thorough magnetic particle and visual inspection for cracks. \n\n\t\t(5) Conduct a thorough visual inspection for evidence of cracks in the main landing gear trunnion support rib and flanges using a low power (2- or 3- power) wide-field (at least 2 1/2-in diameter field of view) magnifying glass or FAA approved equivalent, and covering every square inch of exposed area (both sides) with special emphasis around each and every bolt hole on all flanges and boundaries. Any suspected discrepancy should be confirmed with dye penetrant or equivalent after paint removal. \n\n\t\t(6) If cracks are found during inspections (d)(1) through (d)(5), the affected components must be replaced or repaired in accordance with FAA approved Boeing procedures prior to further flights. \n\n\t\t(7) After reinstalling nuts and washers in accordance with Part I, Subparagraph (e), S.B. 859 (R-2 or later), measure the gap between the upper and lower flanges and skin at several points along the forward 8 inches of each flange using a thickness gage. See latest revision of FAA approved S.B. 859 (R-2 or later) Part I, Subparagraph (f) for instructions if any gap exceeds 0.02 inch. \n\n\t(e) The following repetitive inspections are required on all specified 707 Series aircraft upon completion of inspections and rework outlined in (d) and on all specified 720 Series aircraft. These provisions of paragraph (e), (1) and (2) may be deleted from the 720 aircraft inspection intervals, provided that no cracks have been found in the steel main landing gear trunnion support rib. If cracks are found or have been found in the rib assembly and repaired per Service Bulletin No. 859 (R-2) the repetitive inspections of paragraph (e), (1) and (2) shall apply. \n\n\t\t(1) Every 420 hours' time in service, visually inspect the forward and aft trunnion support fittings for cracks. \n\n\t\t(2) Every 840 hours' time in service visually inspect the web and flanges on the inboard and outboard sides of the trunnion support rib for cracks. \n\n\t\t(3) Every 5,000 hours' time in service clean all areas of the main landing gear trunnion support assembly ofdirt and grease using Naphtha TT-N-95 or BNS 3-2. After cleaning, together with mirror and lighting as required, and using a low powered (2- or 3-power) wide-field (at least 2 1/2-inch diameter field of view) magnifying glass, or FAA approved equivalent, conduct the following inspections as indicated: \n\n\t\t\t(i) Visually inspect the web and flanges on the inboard and outboard sides of the trunnion support rib for evidence of cracks. \n\n\t\t\t(ii) Visually inspect the forward trunnion support fitting for evidence of cracks of the forward and aft sides in the region of the bearing collar. \n\n\t\t\t(iii) Visual inspect the aft trunnion support fitting for evidence of cracks. \n\n\t\t\t(iv) Crack indications found in (i), (ii), and (iii) should be confirmed by dye penetrant inspection. Allowable crack limits and FAA approved rework information for (i) is shown in FAA approved Service Bulletin No. 859 (R-2 or later) (Figure 2). \n\n\t\t(4) If cracks are found in the inspections of (e), the affected components must be replaced or repaired in accordance with FAA approved Boeing procedures prior to further flight. \n\n\t\t(5) Upon request of the operator, an FAA maintenance inspector, subject to approval of the Chief, Engineering and Manufacturing Branch, 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\t(Boeing Service Bulletin Number 859 (R-2) pertains to this subject.) \n\n\tThis supersedes AD 60-08-01. \n\n\tAmendment 326 effective September 23, 1961. \n\n\tRevised January 26, 1962. \n\n\tRevised October 26, 1963. \n\n\tThis amendment (39-1002) becomes effective June 6, 1970.