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98-03-05: This amendment adopts a new airworthiness directive (AD), applicable to all Airbus Model A330 and A340 series airplanes. This action requires removal of three electric motor-driven hydraulic pumps (EHP) and associated wiring, and installation of placards in the flight deck. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified in this AD are intended to prevent operation of the EHP, which could result in fire in the wheel well area, and consequent damage to airplane structure or injury to airplane occupants.
91-01-03: 91-01-03 PRATT & WHITNEY CANADA: Amendment 39-6843. Docket No. 90-ANE-38-AD. Applicability: Pratt & Whitney Canada (PWC) PW118, PW118A, PW120, PW120A, PW121, and JT15D- 5 model engines, installed on, but not limited to, DeHavilland of Canada DHC-8 Series 100, Embraer EMB120, Aerospatiale ATR-42, Beech Beechjet, Cessna T47A, and Siai-Marchetti S211 aircraft. Compliance: Required as indicated, unless already accomplished. To prevent a fire hazard in the engine nacelle, accomplish the following: (a) For engines equipped with Hamilton-Standard Model JFC118 hydromechanical fuel control units (HMU) identified in Table I of this AD, excluding HMU's marked "MS090-001", perform the following: (1) Perform an HMU leak check inspection in accordance with the applicable Accomplishment Instructions of Appendix I of this AD, within the next 15 hours time in service after the effective date of this AD. (2) Thereafter, reinspect the HMU for leakage in accordance with the applicable Accomplishment Instructions of Appendix I at intervals not to exceed 15 hours time in service since last inspection. (3) Remove from service, prior to further flight, HMU's exhibiting fuel leakage when inspected in accordance with (a)(1) or (a)(2) above. (4) X-ray or disassemble inspect the HMU for correct assembly in accordance with the Accomplishment Instructions of the applicable Hamilton-Standard (HS) service bulletin (SB) listed in Table I of this AD, at the next engine shop visit or HMU removal, or by June 30, 1991, whichever occurs first. (5) Remove from service, prior to further flight, HMU's confirmed incorrectly assembled when inspected in accordance with (a)(4) above. (6) For HMU's determined to be correctly assembled when inspected in accordance with (a)(4) above, the repetitive inspections of (a)(1) or (a)(2) above are no longer required. Table I HMU Model/P/N(s) JFC118-10/786390-3 HS SB JFC118-10-73-14, Revision 1 (Oct. 26, 1990) JFC118-11/786391-3 and 786391-5 JFC118-11-73-15, Revision 1 (Oct. 26, 1990) JFC118-12/786392-4 and 786392-6 JFC118-12-73-16, Revision 1(Oct. 26, 1990) JFC118-30/787230-1 JFC118-30-73-15, Revision 1 (Oct 26, 1990) JFC118-31/776660-3 and 790155-1 JFC118-31-73-14, Revision 1 (Oct. 26, 1990) (b) For the purpose of this AD, shop visit is defined as the induction of an engine into a shop for the conduct of maintenance. (c) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (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, Engine Certification Office, ANE-140, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts01803. The x-ray and disassembly inspection shall be done in accordance with the following HS documents: Document Page Revision Date HS SB JFC118-10-73-14 All 1 Oct. 26, 1990 HS SB JFC118-11-73-15 All 1 Oct. 26, 1990 HS SB JFC118-12-73-16 All 1 Oct. 26, 1990 HS SB JFC118-30-73-15 All 1 Oct. 26, 1990 HS SB JFC118-31-73-14 All 1 Oct. 26, 1990 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 the United Technologies Corporation, Hamilton-Standard Division, Technical Publications Department, One Hamilton Road, Windsor Locks, Connecticut 06096-1010. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, D.C. 20591. This amendment (39-6843, AD 91-01-03) becomes effective on January 07, 1991. APPENDIX I PART A: PW100 SERIES A. References: - Models PW118 (BS698)/PW118A (BS718): Maintenance Manual P/N 3034622. - Models PW120 (BS633/BS716) and PW121 (BS722/BS725): Maintenance Manual P/N 3034642. - Models PW120A (BS632) and PW121 (BS717): Maintenance Manual P/N 3034632. B. Accomplishment Instructions: Perform the following HMU leak check inspection. 1.) Perform a fuel system leak test in accordance with the applicable maintenance manual, or visually inspect the HMU for external fuel leaks within 30 minutes of shutdown. 2.) Ensure there is no fuel leakage at the HMU electrical connector area. 3.) If any fuel leak is observed, remove the HMU from service. 4.) Annotate engine log to include this AD inspection. NOTE: Information concerning this inspection can be found in Pratt & Whitney Canada (PWC) Service Bulletin (SB) No. 20951. PART B: JT15D SERIES A. References: -Model JT15D-5: Maintenance Manual P/N 3033442 B. Accomplishment Instructions: Perform the following HMU leak check inspection. 1.) Perform a fuel system leak test in accordance with the applicable maintenance manual, or visually inspect the HMU for external fuel leaks within 30 minutes of shutdown. 2.) Ensure there is no fuel leakage at the HMU electrical connector area. 3.) If any fuel leak is observed, remove the HMU from service. 4.) Annotate engine log to include this AD inspection. NOTE: Information concerning this inspection can be found in PWC SB No. A-7295.
76-20-06: 76-20-06 HILLER AVIATION: Amendment 39-2740. Applies to Model UH-12D and UH-12E Helicopters which have been converted to turbine power in accordance with Soloy Conversions, Limited, STC Nos. SH177WE and SH178WE respectively certificated in all categories. Compliance required as indicated. To prevent loss of helicopter control due to freezing of the governor cable, throttle cable, and/or anti-ice cable accomplish the following: (A) Within 15 days time in service after the receipt of this telegraphic AD, install a placard in view of the pilot which states: "Flight in outside air temperature of 32 degrees F. or lower is prohibited." (B) Within 60 days time in service after the receipt of this telegraphic AD, perform the modifications contained in Soloy Conversions, Ltd., Service Bulletin 01-560 dated September 3, 1976, or later FAA approved revisions. (C) The operation restriction prescribed in (A) above may be discontinued and the placard may be removed when the control cable modifications required by (B) above have been completed. (D) Equivalent procedures may be approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region, upon the submission of adequate substantiating data. 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 Soloy Conversions, Limited, P. O. Box 60, Chehalis, Washington 98532. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated September 8, 1976.
78-13-02: 78-13-02 BRITISH AEROSPACE: Amendment 39-3245. Applies to Hawker Siddeley Model BH/HS-125 Series 600A and 700A airplanes, certificated in all categories, that have either the RCA AVQ21 or Primus 40 weather radar systems installed. Compliance required as indicated. To prevent failure of the restraint provisions for the weather radar receiver/transmitter, accomplish the following: (a) Within 10 hours time in service after the effective date of this AD, unless already accomplished within the last 40 hours time in service, and thereafter at intervals not to exceed 50 hours time in service from the last inspection, inspect the support brackets for cracks and the "T" bolts for failure of the brazing in accordance with section 2, "Accomplishment Instructions" of British Aerospace Alert Service Bulletin 34-A134, dated April 1, 1978, or an FAA approved equivalent. (b) If a crack in a support bracket or a failure of the brazing of the "T" bolt is found during an inspection required by paragraph (a) of this AD, before further flight, except that the airplane may be flown in accordance with FAR 21.197 and 21.199 to a base where the replacement can be accomplished, replace the mounting tray with a serviceable part of the same part number, or repair the existing mounting tray in accordance with an FAA approved repair scheme and continue to inspect in accordance with paragraph (a) of this AD, or replace the mounting tray in accordance with paragraph (c) of this AD. (c) The inspections required by this AD may be discontinued upon replacement of the mounting tray with an improved standard tray, P/N 1719353-501 (Rev. E), in accordance with British Aerospace Modification 258171 or an FAA approved equivalent. This amendment becomes effective July 5, 1978.
89-25-10: 89-25-10 BEECH: Amendment 39-6409. Applicability: Models 65-90 and 65-A90 (Serial Number (S/N) LJ-1 thru LJ-317); 65-A90-1, 65-A90-2, 65-A90-3, 65-A90-4, B90, C90 (all S/N); C90A (S/N LJ-1063 thru LJ-1087, except LJ-1085); E90, 100, A100 and B100 (all S/N) airplanes certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To detect possible fatigue cracking of the wing main spar lower cap and associated structure, accomplish the following: (a) Within the next 200 hours time-in-service (TIS), after the effective date of this AD, or upon accumulating 3000 hours TIS, whichever occurs later, unless previously accomplished per AD 87-23-09, Amendment No. 39-5765, or AD 70-25-04, Amendment No. 39-1332, and thereafter at intervals not to exceed 1000 hours TIS (except as provided in paragraph (b) below) after the initial inspection, inspect the wing lower forward spar attach fittings, center sectionand outboard wing spar caps adjacent to the attach fittings by visual, fluorescent penetrant and eddy current methods as specified in the applicable section of Beech Structural Inspection and Repair Manual (SIRM), Part Number 98-39006, Revision A4, dated May 1, 1987. The inspection must be performed by personnel specifically trained by Beech Aircraft Corporation. NOTE 1: Beech offers a two-day training course free of charge to qualified personnel who have prior knowledge of eddy current inspection techniques. A listing of Beech Corporate maintenance facilities may be obtained from the sources contained in paragraph (g) of this AD. A listing of other facilities employing qualified inspectors is not available. (b) At each inspection required by paragraph (a) above, inspect any reinforcing strap installed per Supplemental Type Certificate (STC) SA1178CE or SA1583CE for proper tension and condition in accordance with Aviadesign Engineering Order E.O. B-8001, Issue 3, dated May30, 1985. Correct any discrepancy prior to further flight. For airplanes so equipped and inspected, the repetitive inspection interval of 1000 hours TIS in paragraph (a) above may be extended to 3000 hours TIS. (c) If any crack is found in a main spar lower cap or fitting, prior to further flight repair or replace the defective part using the instructions and limitations specified in the Beech SIRM or other FAA approved instructions provided by Beech Aircraft Corporation. (d) Within one week after completion of any inspection required by paragraph (a) or (b) of this AD, complete the reporting form included with this AD as Figure 1 and mail it to the address shown thereon (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056). (e) The initial and repetitive inspections specified in this AD are no longer required when the airplane is modified by Beech Wing Modification Kit No. 90-4077-1S or 100-4007-1S. (f) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (g) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209; Telephone (316) 946-4400. NOTE 2: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, Wichita, Kansas 67201-0085; or Western Aircraft Maintenance, 4444 Aeronca Street, Boise, Idaho 83705, or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. ThisAD supersedes AD 87-23-09, Amendment 39-5765, and AD 70-25-04, Amendment 39-1332. This amendment (39-6409, AD 89-25-10) becomes effective on January 4, 1990. REPORTING FORM - 89-25-10 Airplane Model No. ___________________________________________________ Airplane Serial No.____________________________________________________ Date of inspection per this AD____________________________________________ Airframe total hours time-in-service________________________________________ Were any fatigue cracks found? No Yes __ If "Yes" was checked above, complete the following: Location of crack_____________________________________________________ Was crack removable by reaming or grinding? No Yes __ Additional Comments __________________________________________________ Mailing Address: FAA, Wichita ACO Airframe Branch, Room 100 1801 Airport Road Wichita, KS 67209 FIGURE 1 - 89-25-10
76-07-04: 76-07-04 HAWKER SIDDELEY AVIATION LTD: Amendment 39-2563. Applies to de Havilland Model DH-114 "Heron" airplanes certificated in all categories which have not been altered in accordance with Heron Modification 1612. Compliance is required as indicated. To detect cracks in the nose landing gear inner casing, and prevent the possible collapse of the nose landing gear upon landing, accomplish the following: (a) Within the next 50 hours time in service after the effective date of this AD, unless already accomplished within the preceding 600 hours time in service, and thereafter, at intervals not to exceed 600 hours time in service from the last inspection, inspect the nose landing gear inner casing for cracks in accordance with paragraphs 3.1 and 3.2 of section 3 entitled "Inspection" of Hawker Siddeley Aviation Ltd., Technical News Sheet No. U.17, Issue 1, dated September 17, 1973, or an FAA-approved equivalent. (b) If any cracks are found during an inspection required by paragraph (a) of this AD, before further flight, except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed, replace the inner casing with a new part of the same part number or a serviceable used part of the same part number that has been inspected and found to be free of cracks in accordance with the inspection prescribed in paragraph (a) of this AD. Continue to inspect the replacement casing for cracks in accordance with Hawker Siddeley Aviation Ltd., Technical News Sheet, No. U.17, Issue 1, dated September 17, 1973, or an FAA-approved equivalent at intervals not to exceed 600 hours time in service from replacement. This amendment becomes effective April 12, 1976.
77-14-04: 77-14-04 HAWKER SIDDELEY AVIATION, LTD.: Amendment 39-2952. Applies to DH-104 "Dove" and DH-114 "Heron" airplanes. Compliance is required within the next 500 hours time in service after the effective date of this AD unless already accomplished. To prevent the possibility that a loss of generated electrical power would be undetected by the flight crew, accomplish the following: (a) Alter the electrical system to incorporate a bus bar low voltage sensing unit, a bus bar low voltage warning light, and an essential service switch, designed and installed in accordance with paragraphs 5 and 6 of Hawker Siddeley Aviation, Ltd., Technical News Sheet, Series: Heron (114), No. N. 6., Issue 3 (for DH-114 "Heron") and CT104, No. 227, Issue 3 (for DH-104 "Dove"), both dated July 3, 1972, as amended to November 20, 1972, or FAA-approved equivalent of either. (b) Amend the "Normal and Emergency Procedures", Part B, of the "Operating Procedures" section, Section II, of the applicable Airplane Flight Manual by adding the electrical system operation information contained in paragraphs 7 and 8 and Figure 1 of the applicable Technical News Sheet, referred to in paragraph (a) of this AD, or an FAA-approved equivalent. (c) Check the condition of the electrical distribution and generator system in accordance with paragraph 6 of the applicable Technical News Sheet, referred to in paragraph (a) of this AD, or an FAA-approved equivalent, and repair, as necessary. The checks required by this paragraph may be performed by persons authorized to perform preventive maintenance under FAR 43. This amendment becomes effective August 5, 1977.
78-03-05: 78-03-05 MITSUBISHI HEAVY INDUSTRIES, LTD: Amendment 39-3137. Applies to models MU-2B, MU-2B-10, MU-2B-15, MU-2B-20, MU-2B-25 and MU-2B-26 airplanes with serial numbers up through and including 347 except 313 and 321 and models MU-2B-30, MU-2B-35, and MU-2B-36 airplanes with serial numbers up through and including 696 except 652 and 661. NOTE: This AD is not applicable to MU-2B series airplanes having serial numbers with the suffix "SA." Compliance is required as indicated. To prevent failure of the cowling latches between the engine nacelle upper door and side doors, subsequent separation of the upper cowling panel, and possible loss of control of the airplane, accomplish the following: Within the next 25 hours time in service after the effective date of this AD, unless already accomplished, replace the cowling latch links between the engine nacelle upper door and side doors in accordance with the instructions contained in Mitsubishi MU-2 Service Bulletin No. 171A datedJuly 14, 1975, as supplemented by Mitsubishi MU-2 Service Bulletin No. 180 dated August 26, 1977, or Mitsubishi Service Bulletin No. 180A dated November 17, 1977, or an FAA- approved equivalent, approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. This supersedes amendment 39-2695, (41 FR 34009), AD 76-16-05. This amendment becomes effective February 23, 1978.
76-18-12: 76-18-12 GRUMMAN AMERICAN AVIATION CORPORATION: Amendment 39- 2721. Applies to Model G-1159 airplanes certificated in all categories, serial numbers 1 through 154 and 775. Compliance required within the next 100 hours time in service after the effective date of this AD, unless already accomplished. To detect loose terminal connections at the generator terminal boards and to prevent the loosening of these connections, accomplish the following: 1. Modify each engine electrical junction box to provide an access hole for inspecting the two generator terminal boards and the eight associated wiring connections. Grumman American Aircraft Service Change No. 203 Amendments 1 and 2 provide the information for accomplishing this modification. 2. If the inspection reveals that all generator terminals are secure, i.e., that all lock washers are compressed and torque stripes not broken, the connections are satisfactory. 3. Reinspect in accordance with A.S.C. No. 203 Amendments 1 and 2 in intervals of 300 hours time in service until the basic A.S.C. No. 203 has been accomplished. 4. If the lock washer on any terminal is found not to be compressed, or evidence of arcing at any connection is noted, the affected engine shall be removed and the basic A.S.C. No. 203 accomplished to the corresponding engine junction box. 5. The reinspection procedure must continue for the remaining engine junction box until the basic A.S.C. No. 203 has been accomplished. This service change requires the removal of a plain nut, a plain washer, and lockwasher from each generator terminal, and replacing them with a self-locking nut and a plain washer. 6. Compliance with the basic A.S.C. No. 203 must be accomplished at the next engine removal, if not done prior to that time. This Airworthiness Directive may be accomplished by any other means approved by the Chief, Engineering and Manufacturing Branch of the Southern Region, Atlanta, Georgia. This amendmentbecomes effective September 22, 1976.
77-14-01: 77-14-01 AGUSTA: Amendment 39-2949. Applies to Model A-109A helicopters equipped with main rotor hub flap hinge bearings P/N SJ7355/IR7355. Compliance is required as indicated, unless already accomplished. To prevent possible improper operation of the main rotor flaps due to premature failure of any one of the main rotor hub flap hinge bearings, accomplish the following: (a) For helicopters with serial numbers up to and including S/N 7120, except S/N 7119 - (1) Within the next 10 hours time in service after the effective date of this AD, replace the main rotor hub flap hinge bearings P/N SJ7355/IR7355 in accordance with Part I of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent; and (2) Within 100 hours time in service after complying with paragraph (a)(1) of this AD, and thereafter at intervals not to exceed 100 hours time in service, or any time abnormal oil leaks occur from flap hinges, perform inspections and replace bearings, as necessary, in accordance with Part II of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent. (b) For helicopters with serial numbers S/N 7119, S/N 7121 and up, within the next 25 hours time in service after the effective date of this AD, except for those which have been inspected within the previous 75 hours, and thereafter at intervals not to exceed 100 hours time in service or any time abnormal oil leaks occur from flap hinges, perform inspections and replace bearings, as necessary, in accordance with Part II of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent. (c) Equivalent methods of complying with this AD must be approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region, c/o American Embassy, APO New York, N.Y. 09667. This amendment becomes effective July 15, 1977.