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90-07-10 R1: 90-07-10 R1 CESSNA CITATION: Amendment 39-6599. Final copy of, and revision to, priority letter. Docket No. 90-NM-54-AD. Applicability: Model 550, 551, and S550 series airplanes, equipped with STC SA2698SW Freon Air Conditioners installed in the aft section of the baggage compartment, certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent failure of the elevator or rudder control cables, accomplish the following: A. Prior to further flight, deactivate the air conditioning system by pulling the air conditioner supply breakers on the main junction box and install tie wraps on the breakers. B. Within the next 10 hours time-in-service, perform a visual inspection of the elevator control cables under and in the vicinity of the aft baggage compartment floor for evidence of damage caused by electrical arcing and interference. 1. Replace any damaged cable prior to further flight. 2. If fasteners used to secure the compressor condenser unit mounting rails to the baggage floor interfere with the control cables, replace mounting hardware, prior to further flight, in accordance with Keith Service Bulletin No. 108, dated March 26, 1990. C. Prior to reactivating the air conditioning system, replace mounting hardware on the inboard mounting rail and replace existing ground wire, in accordance with Keith Service Bulletin No. 108, dated March 26, 1990. D. An 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, Special Programs Office, ASW-190, FAA, Southwest Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment, and then send it to the Manager, Special Programs Office, ASW-190. 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. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer, may obtain copies upon request to Keith Products, Inc., 4554 Claire Chennault, Dallas, Texas 75248. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or at the FAA, Southwest Region, Special Programs Office, 4400 Blue Mound Road, Fort Worth, Texas. Portions of this amendment were effective earlier to all recipients of Priority Letter AD 90-07-10, dated March 29, 1990. Airworthiness Directive 90-07-10 R1 supersedes Priority Letter AD 90-07-10, issued on March 29, 1990. Amendment 39-6599, AD 90-07-10 R1 became effective on May 29, 1990.
76-12-05: 76-12-05 AVCO LYCOMING: Amendment 39-2635. Applies to all Avco Lycoming T5508D model engines equipped with P/N 2-160-147-02 retaining plate spacers and P/N 2-161-182-01 compressor air bleed bands. Compliance required as indicated. To prevent partial power loss due to failure of the compressor air bleed band while the engine is operating at a power level which requires the bleed band to be closed, perform the following: (a) Prior to the first flight of each day, conduct an inspection of the compressor air bleed band for edge wear and cracks, and repair or replace as necessary in accordance with the instructions in Avco Lycoming Service Bulletin, Product Support No. 5508-0003, dated November 30, 1975, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. (b) Prior to July 1, 1976, replace P/N 2-160-147-02 retaining plate spacers with P/N 2-160-561-01 retainers and replace P/N 2-161-182-01 compressor air bleed band with P/N 2-161-182-03 compressor air bleed band, in accordance with the instructions in Avco Lycoming Service Bulletin, Product Support No. 5508-0007, dated April 30, 1976, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. (c) Equivalent methods of compliance may be approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. The manufacturer's Service Bulletins 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 Avco Lycoming Engine Group, Stratford Division, 550 South Main Street, Stratford, Connecticut 06497. These documents may also be examined at the Office of the Regional Counsel, New England Region, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the New England Regional Office in Burlington, Massachusetts. This amendment becomes effective June 23, 1976.
89-23-16: 89-23-16 CASA: Amendment 39-6391. Docket No. 89-NM-122-AD. Applicability: All Model C-212 series airplanes, certificated in any category. Compliance: Required within 60 days after the effective date of this AD, unless previously accomplished. To prevent, following an engine failure on takeoff, damage to or failure of the remaining engine caused by overtorquing, accomplish the following: A. Modify the Automatic Power Reserve (APR) system in accordance with CASA Service Bulletin 212-72-05, Revision 1, dated June 1, 1989. B. Upon accomplishment of the modification required by paragraph A., above, revise the FAA-approved Airplane Flight Manual to include the appropriate revision, which provides new procedures for setting power, as specified in Paragraph 1.B. of CASA Service Bulletin 212-72-05, Revision 1, dated June 1, 1989. C. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be usedwhen approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. D. 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. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Construcciones Aeronauticas, S.A. (CASA), Getafe, Madrid, Spain. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6391, AD 89-23-16) becomes effective on December12, 1989.
89-24-01 R1: 89-24-01 R1 TELEDYNE CONTINENTAL MOTORS (TCM): Amendment 39-6358 as revised by Amendment 39-6645. Docket No. 89-ANE-11. Applicability: TCM engines, Models TSIO-520B, BB, D, DB, E, EB, J, JB, K, KB, N, NB, UB, and VB equipped with scavenge oil pump gears, part numbers (P/N's) 635334, 639388, 649157, and 649159 except those engines equipped with starter adapter P/N's 642085A4 or 642085A5, those engines with serial numbers as shown in the Appendix of this AD, or those engines equipped with any starter adapter with a pulley for an air conditioner compressor drive, certificated in any category. NOTE: "Appendix" refers to TCM Service Bulletin M90-6 which is not reprinted in this AD. Compliance: Required within 500 flight hours after the effective date of this AD, or at the next maintenance event, after the effective date of this AD, during which the scavenge oil pump gears are removed from the engine, whichever occurs first, unless already accomplished. To prevent possible failure of scavenge oil pump gears which could result in total loss of engine power, accomplish the following: (a) Remove the scavenge oil pump gears from the scavenge oil pump housing and inspect the gear teeth for a drill point as shown in Figure 1 of the Appendix to this AD. (1) If the drill point is present, inspect the gears in accordance with the procedures outlined in the Appendix to this AD. (2) If the drill point is not present, a Rockwell hardness test may be conducted on the gear. If a Rockwell "A" value of 79 or greater or a Rockwell "C" value of 56 or greater is obtained, inspect the gear in accordance with the procedures outlined in the Appendix to this AD. (3) If the gears fail the inspection specified in (1) or (2) above, replace the gears with serviceable P/N 649157 or P/N 649159 gears having the drill point marking. NOTE: TCM SB M90-6 contains information relating to the requirements of this AD. (b) Make an appropriate log book entry showing compliance with this AD. (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) times specified in this AD may be approved by the Manager, Atlanta Aircraft Certification Office, Small Airplane Directorate, Aircraft Certification Service, Federal Aviation Administration, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia 30349. All persons affected by this AD who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Teledyne Continental Motors, P.O. Box 90, Mobile, Alabama 36601. This information may be examined at the FAA, Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New EnglandRegion, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803. This AD revises Amendment 39-6358, AD 89-24-01. This amendment (39-6645, AD 89-24-01 R1) becomes effective on August 10, 1990.
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.
76-16-01: 76-16-01 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-2683. Applies to the following groups of Model TPE331 series engines: Group A: TPE331-1-101B, S/N 93058 through 93061; TPE331-1-151A, S/N 92249 and 92336 through 92354; TPE331-1-151K, S/N 26001 through 26014; TPE331-1-151G, S/N 91193 through 91198; TPE331- 2-201A, S/N 90218 through 90278, TPE331-3U-303G or TPE331-3UW-303G, S/N 03108, 03109, and 03112 through 03180 and 05031 through 05042; TPE331-3U-307G, S/N 03001 and 03009; TPE331-5-251C, S/N 22006 through 22057; TPE331-5-251K, S/N 06113, 06190 through 06442 and 06444 through 06454; TPE331-6-251M or TPE3310-6-252M, S/N 20144 and 20182 through 20467 and 20469 through 20533; TPE331-6-252B, S/N 27001 and 27002, and any Model TPE331-1, -2, -3, -5, or -6 engines which have been modified in accordance with AiResearch Service Bulletin TPE331-72-0064 dated February 1, 1974 or subsequent revisions. Group B: TPE331-1-101B, S/N 93062, 93063; TPE331-1-151A, S/N 92355 through 92357; TPE331- 1-151K, S/N 26015 through 26023; TPE331-2-201A, S/N 90279 through 90296; TPE331-3U-303G or TPE331-3UW-303G, S/N 03181 through 03197, 05043 through 05052; TPE-5-251C, S/N 22058 through 22119; TPE331-5-251K, S/N 06443, 06455 through 06556; TPE331-6-251M or TPE331-6-252M, S/N 20534 through 20577. Group C: TPE331-1-101F, S/N 98017 through 98019; TPE331-1-151K, S/N 91194 and 91195; TPE331-2- 201A, S/N 90297 through 90303; TPE331-3U-303G or TPE331-3UW-303G, S/N 05053 through 05058, S/N 05016, S/N 03001, S/N 03010, S/N 03011, and S/N 03198 through 03205; TPE331-5-251K, S/N 06557 through 06576; TPE331-5-251C, S/N 22120 through 22139, TPE331-5-251M, S/N 28001 through 28005; TPE331-6-251M, S/N 20468, S/N 20017, and S/N 20018; TPE331-6-252B, S/N 27003 and S/N 27004, and any Model TPE331-1, -2, -3, -5, or -6 engines which have been modified in accordance with AiResearch Service Bulletin TPE331-72-0064, dated February 1, 1974, or subsequent revisions. NOTE: Operators of airplanes incorporating the engines affected by this airworthiness directive are advised to examine the applicability section, Configuration Identification, Table 1, or AiResearch Service Bulletin TPE331-72-0092, Revision 5, dated June 8, 1976, or later FAA-approved revision, to identify the extent of work required by this airworthiness directive which may have been previously performed. Compliance required as indicated. To detect and prevent fatigue failures of the high speed pinion (HSP) bearing oil transfer tube, supply tube, and lubricating adapter, and to detect, correct, and prevent loosening of the HSP bearing carrier bolts, accomplish the following: (a) For engines in Group A, above, within the next 100 hours time in service after December 30, 1974 (the effective date of Amendment 39-2054, AD 74-26-11), unless already accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, dated December 9, 1974, or later FAA-approved revision: (1) Replace the two high speed pinion bearing carrier bolts, P/N MS21279-07, with two bolts, P/N MS9489-07, and lockplate, P/N 3101483-1, (2) Inspect to insure proper torque on bolt, P/N MS21279-10, securing lube nozzle, P/N 3101209, and (3) If aluminum lube adapter, P/N 3101210, is not replaced as described in paragraph (d) of this directive, inspect to insure proper torque on bolt, P/N MS21297-07, securing this lube adapter. (b) For engines in Group A and Group B, above, which incorporate the oil transfer tube, P/N 3101187-1, within the next 100 hours time in service after December 30, 1974, unless already accomplished within the last 100 hours time in service prior to December 30, 1974, and thereafter at intervals not to exceed 200 hours time in service from last inspection, until the modifications of paragraph (c) are accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, dated December 9, 1974, or later FAA-approved revision: (1) Inspect the integral support bracket on oil transfer tube, P/N 3101187-1, for cracks or separation, and (2) If the oil transfer tube bracket is cracked or separated, before further flight, replace both the oil transfer tube and oil supply tube with new tubes, oil transfer tube, P/N 3101475-1, and oil supply tube, P/N's 3101473-1 or 3101473-2, and associated hardware. (c) For engines in Group A and Group B, above, which incorporate the oil transfer tube, P/N 3101187-1, before accumulating a total of 1800 hours time in service after December 30, 1974, unless already accomplished, replace the oil transfer tube, P/N 3101187-1, with one of the following serviceable oil transfer tubes in accordance with AiResearch Service Bulletin TPE331-72-0092, Revision 1, dated, January 1, 1975, or later FAA-approved revision: (1) P/N 3101187-2 which has been heat-treated and reidentified as P/N 3101187-3, (2) P/N 3101187-3, or (3) P/N3101475-1 (d) For engines in Group A, Group B, and Group C, above, within the next 200 hours time in service after the effective date of this AD, unless already accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, Revision 5, dated June 8, 1976, or later FAA-approved revision: (1) Replace the aluminum main journal bearing lubricating adapter, P/N's 3101210-1 or 3101210- 2, with a new steel lubricating adapter, P/N's 3101474-1 or 3101474-2, and (2) If two carrier bolts, P/N MS9489-07, and lockplate, P/N 3101483-1, are not replaced, inspect these carrier bolts manually for any indication of looseness. If loose bolts are found, before further flight install a new lockplate, P/N 3101483-1, and retorque the two carrier bolts, P/N MS9489-07 (e) Equivalent procedures, inspections, modifications, or parts may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. (f) Aircraft may be flown to a base for accomplishment of inspections and modifications required by this AD per FAR's 21.197 and 21.199. This supersedes Amendment 39-2054 (39 F.R. 4439), AD 74-26-11, as amended by Amendments 39-2092 (40 F.R. 6771), 39-2214 (40 F.R. 22126), 39-2254 (40 F.R. 28605), 39-2367 (40 F.R. 42740), and 39-2382 (40 F.R. 48499). This amendment becomes effective August 6, 1976.
77-07-10: 77-07-10 BRITTEN NORMAN, LTD.: Amendment 39-2869. Applies to Models BN-2A Islander and BN-2A Mk III Trislander airplanes, all series, certificated in all categories, except those airplanes incorporating Britten Norman Modification NB/M878. Compliance is required as indicated, unless already accomplished. To detect internal corrosion and prevent possible failure of the aileron mass balance support arm, accomplish the following: (a) For ailerons modified in accordance with Britten-Norman, Ltd. Modification NB/M/336, within the next 25 hours time in service after the effective date of this AD or prior to 3 years since new, whichever occurs later, accomplish the following: (1) Gain access to and inspect the interior surface of the aileron mass balance support arm in accordance with steps 1 thru 4 of Part 1 and Figure 1 of Britten-Norman, Ltd. Service Bulletin BN-2/SB.98, Issue 1, dated October 27, 1976 (hereinafter S.B. BN- 2/SB.98), or an FAA-approved equivalent.(2) If no corrosion or acceptable corrosion (as defined in Table 1 of Figure 1 of S.B. BN-2/SB.98) is found as a result of the inspection required by paragraph (a)(1) of this AD, clean and protect the interior surface of the support arm and the shank of the bobweight in accordance with subpart (a) of step 5, and reassemble in accordance with steps 6 through 9, of Part 1 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (3) If marginal corrosion (as defined in Table 2 of Figure 1 of S.B. BN- 2/SB.98) is found as a result of the inspection required by paragraph (a)(1) of this AD, before further flight, clean, protect, and reassemble the support arm, in accordance with paragraph (a)(2) of this AD. Thereafter, within the next 300 hours time in service or 3 months, whichever occurs sooner, replace the support arm with a new part of the same part number or an FAA-approved equivalent, and reassemble in accordance with paragraph (a)(2) of this AD. (4) If, during the inspection required by paragraph (a)(1) of this AD, corrosion is found beyond acceptable limits (as defined in Table 3 of Figure 1 of S.B. BN- 2/SB.98), before further flight, replace the support arm with a new part of the same part number, or an FAA-approved equivalent, and reassemble in accordance with paragraph (a)(2) of this AD. (b) For ailerons not modified in accordance with Britten-Norman, Ltd. Modification NB/M/336, within the next 25 hours time in service after the effective date of this AD or prior to 3 years since new, whichever occurs later, accomplish the following: (1) Gain access to and inspect the interior surface of the mass balance support arm in accordance with steps 1 thru 4 of Part 2 and Figure 2 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (2) If no corrosion or acceptable corrosion (as defined in Table 1 of Figure 1 of S.B. BN-2/SB.98) is found as the result of the inspection required by paragraph (b)(1) of this AD, clean and protect the interiorsurface of the support arm and the shank of the bobweight in accordance with subpart (a) of Step 5, Part 1, and repair the support arm in accordance with Figure 2 and reassemble in accordance with steps 6 thru 9, of Part 2 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (3) If marginal corrosion (as defined in Table 2 of Figure 1 of S.B. BN- 2/SB.98) is found as a result of the inspection required by paragraph (b)(1) of this AD, before further flight, clean, protect, repair, and reassemble the support arm and shank of the bobweight in accordance with paragraph (b)(2) of this AD. Thereafter within the next 300 hours time in service or 3 months, whichever occurs sooner, replace the support arm in accordance with the repair and reassembly instructions specified in Part 3 and Figure 3, of S.B. BN-2/SB.98, or an FAA-approved equivalent. (4) If, during the inspection required by paragraph (b)(1) of this AD, corrosion is found beyond acceptable limits (as defined in Table 3 of Figure 1 of S.B. BN- 2/SB.98), before further flight, replace the support arm in accordance with Part 3 and Figure 3 of S.B. BN-2/SB.98, or an FAA-approved equivalent. This amendment becomes effective April 18, 1977.
90-08-17: 90-08-17 BEECH: Amendment 39-6564. Docket No. 89-CE-39-AD. Applicability: Models 65-90, 65-A90, B90, C90 and C90A (Serial Numbers (S/N) LJ-1 through LJ-1222); E90, F90, H90, 100, A100 and B100 (all S/N); 200 and B200 (S/N BB-2 through BB-1344); 200C, B200C, 200CT, B200CT, 200T, B200T, A200, A200C and A200CT (all S/N); and 300 (S/N FA-1 through FA-204 and FF-1 through FF-19) airplanes certificated in any category. Compliance: Required within the next 100 hours time-in-service after the effective date of this AD, unless already accomplished. To prevent water accumulation in the aft fuselage which can freeze and restrict control movement, accomplish the following in accordance with the instructions in Beech Service Bulletin Number 2312, dated December 1989: (a) Inspect the aft fuselage moisture drain system. (1) If the drain system is unobstructed and the drain openings are not undersized, no further action is required. (2) If any obstruction of thedrain system or any undersized opening is found, prior to further flight remove the obstruction or enlarge the opening as required; and (3) Report, in writing, any defects found to the Manager, Wichita Aircraft Certification Office (address below) within 7 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056.) (b) Airplanes may be flown in accordance with FAR 21.197 to a location where the AD may be accomplished. (c) An alternate method of compliance or adjustment of the compliance time 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. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the document referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; Telephone (316) 681-7111; or may examine this document at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment (39-6564, AD 90-08-17) becomes effective on May 7, 1990.
92-10-07: 92-10-07 ROLLS-ROYCE PLC: Amendment 39-8241. Docket No. 87-ANE-05. Applicability: Rolls-Royce plc (R-R) Viper Mk. 521, Mk. 522, and Mk. 601 turbojet engines installed on but not limited to Hawker-Siddeley HS125 and HS125-600 aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent failure of high pressure compressor (HPC) rotor disks, causing an engine failure, accomplish the following: (a) Remove from service HPC stage 5 and 6 rotor disks with the following part numbers (includes all Modification Standards): Viper Mk. 521: V.20492, V.20494, V.27430, V.27433, V.39638, V.39639, V.39642, V.39643, V.39677, V.39678, V.39681, V.39682, V.42057, V.42059, V.46180, V.46181, V.46182, V.46204, V.46205, and V.46206; Viper Mk. 522: V.20492, V.20494, V.27430, V.27433, V.39638, V.39639, V.39642, V.39643, V.39677, V.39678, V.39681, V.39682, V.42057, V.42059, V.46180, V.46181, V.46182, V.46204, V.46205, and V.46206; Viper Mk. 601: V.42057, V42059, V.44768,V.44769, V.44780, V.44781, V.46182, and V.46206, and replace with a serviceable part, as follows: (1) HPC stage 5 and 6 rotor disks that have accumulated 12,500 or more cycles since new on the effective date of this AD, within the next 100 cycles in service. (2) HPC stage 5 and 6 rotor disks that have accumulated less than 12,500 cycles since new on the effective date of this AD, at or prior to accumulating 12,600 cycles since new. NOTES: (1) Further information pertaining to disk life limits can be obtained in Chapter 5 of the R-R Viper Mk. 521, Mk. 522, and Mk. 601 Maintenance and Shop Manuals. (2) Reference R-R plc Service Bulletins Number 72-A154 for Viper Mk. 601, and Number 72-A372 for Viper Mk. 521 and Viper Mk. 522 engines. (b) 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. (c) 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, Engine Certification Office, FAA, Engine and Propeller Directorate. The request should be forwarded through an FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Engine Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Engine Certification Office. (d) This amendment becomes effective on September 21, 1992.
93-05-04: 93-05-04 AIRBUS INDUSTRIE: Amendment 39-8509. Docket 92-NM-112-AD. Applicability: Model A300 B2-1C, B2-203, B2K-3C, B4-2C, B4-103, and B4-203 series airplanes; equipped with General Electric (GE) engines on which engine Pylon Modification No. 2434, as described in Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, has not been installed; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the engine pylon, accomplish the following: (a) Prior to the accumulation of 22,400 total hours time-in-service, or within 1,000 hours time-in-service after the effective date of this AD, whichever occurs later, perform an initial visual and eddy current inspection to detect wear or cracks of the inner doubler on the pylon side panel around the fire extinguisher access doors, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992. If cracks are detected, prior to further flight, accomplish the requirements of either paragraph (a)(1) or (a)(2) of this AD, as applicable: (1) If cracks less than 5 mm (0.197 inch) are detected, prior to further flight, repair in an accordance with a method approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. (2) If cracks greater than or equal to 5 mm (0.019 inch) are detected, accomplish either paragraph (a)(2)(i) or (a)(2)(ii), as applicable: (i) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. (ii) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD; and repeat those inspections thereafter in accordance with paragraphs (b), (c), (d), or (e) of this AD, as applicable. (b) For airplanes having Configuration No. 1.A.: Repeat the visual and eddy current inspections to detect doubler wear and cracks, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,400 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,400 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 1,800 hours time-in-service. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Accomplish either paragraph (b)(5)(i), (b)(5)(ii), or (b)(5)(iii) of this AD. (i) Repeat the visual and eddy current inspection at intervals not to exceed 350 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (b) of this AD. (6) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 6 mm (0.236 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (c) For airplanes having Configuration No. 1.B: Repeat the visual and eddy current inspections for doubler wear and cracks, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detectedby any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,400 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,400 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Accomplish either paragraph (c)(4)(i), (c)(4)(ii), or (c)(4)(iii) of this AD. (i) Repeat the visual and eddy current inspection at intervals not to exceed 300 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after the replacement date of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (c) of this AD. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 4 mm (0.157 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (d) For airplanes having Configuration No. 1.A. on which the modification described in Airbus Industrie Service Bulletin No. A300-54-008 (door with 4 latches) has been accomplished; and for airplanes having Configuration No. 2.A.: Repeat the visual and eddy current inspections for doubler wear and cracks at zones D1, D2, D3, and D4, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than or equal to 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 5,800 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 5,800 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,300 hours time-in-service. (5) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Repeat the visual and eddy current inspection at intervals not to exceed 2,300 hours time-in-service. (6) For airplanes on which the measured doublerwear, as detected by any visual inspection required by this AD, is greater than 6 mm (0.236 inch) and less than 8 mm (0.315 inch): Accomplish either paragraph (d)(6)(i), (d)(6)(ii), or (d)(6)(iii) of this AD. (i) Repeat the visual and eddy current inspections at intervals not to exceed 450 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (d) of this AD. (7) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 8 mm (0.315 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (e) For airplanes having Configuration No. 1.B. on which the modification described in Airbus Industrie Service Bulletin No. A300-54-008 (door with 4 latches) has been accomplished; and for airplanes having Configuration No. 2.B.: Repeat the visual and eddy current inspections for doubler wear and cracks at zones D1, D2, D3, and D4, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,500 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,500 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,300 hours time-in-service. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Repeat the visual and eddy current inspections at intervals not to exceed 1,800 hours time-in-service. (6) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 6 mm (0.236 inch) and less than 8 mm (0.315 inch): Accomplish either paragraph (e)(6)(i), (e)(6)(ii), or (e)(6)(iii) of this AD. (i) Repeat the visual and eddy current inspections at intervals not to exceed 400 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (e) of this AD. (7) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 8 mm (0.315 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (f) Accomplishment of the replacement of the inner doubler on the pylon side panel around the fire extinguisher access doors; and installation of the modification (screwed doors) in accordance with Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982, and Change Notice, dated July 8, 1985; constitutes terminating action for the visual and eddy current inspections required by this AD. (g) 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. (h) 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. (i) The modification shall be done in accordance with Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982; and Change Notice, dated July 8, 1985, for Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982. The inspections shall be done in accordance Airbus Industrie Service Bulletin No. A300-54-070, Revision 1, dated March 17, 1992, which contains the following list of effective pages: Page Number Revision Level Shown on Page Date Shown on Page 1-10, 31-32 1 March 17, 1992 11-30, 33-34 Original February 6, 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. (j) This amendment becomes effective on April 29, 1993.
90-05-05: 90-05-05 BOEING: Amendment 39-6516. Docket No. 89-NM-166-AD. \n\n\tApplicability: Model 767 series airplanes equipped with Kidde engine fire and overheat detection systems, as listed in Boeing Service Bulletin 767-26-0037, Revision 1, dated June 1, 1989, certificated in any category. \n\n\tCompliance: Required within 6 months after the effective date of the AD, unless previously accomplished. \n\n\tTo prevent false engine fire and overheat warnings, which could result in unnecessary engine in-flight shutdowns and airplane diversions that unduly jeopardize continued safe operation of the airplane, accomplish the following: \n\n\tA.\tModify the engine fire and overheat detection system on each engine in accordance with Boeing Service Bulletin 767-26-0037, Revision 1, dated June 1, 1989. \n\n\tB.\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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Seattle Aircraft Certification Office.\n \n\tC.\tSpecial 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. \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 Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment (39-6516, AD 90-05-05) becomes effective on March 27, 1990.
90-26-08: 90-26-08 BOEING: Amendment 39-6831. Docket No. 90-NM-252-AD. \n\n\tApplicability: Model 747-400 series airplanes, listed in Boeing Alert Service Bulletin 747-26A2171, dated October 4, 1990, certificated in any category. \n\n\tCompliance: Required within 45 days after the effective date of this airworthiness directive, unless previously accomplished. \n\n\tTo prevent the failure of the fire detection warning system to annunciate a fire, accomplish the following: \n\n\tA.\tModify the wiring of the aural and master warning indication systems of the fire detection system, in accordance with Boeing Alert Service Bulletin 747-26A2171, dated October 4, 1990. \n\n\tB.\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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tC.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tThis amendment (39-6831, AD 90-26-08) becomes effective on December 26, 1990.
90-15-05: 90-15-05 BOEING: Amendment 39-6654. Docket No. 90-NM-24-AD. \n\n\tApplicability: Model 767 series airplanes, identified in paragraphs A. and B., below, certificated in any category. \n\n\tCompliance: Required within the next 20 months after the effective date of this AD, unless previously accomplished. \n\n\tTo provide satisfactory reliability of the evacuation system, accomplish the following: \n\n\tA.\tFor airplanes identified in Boeing Service Bulletin 767-25-0120, dated December 14, 1989: Modify the off-wing evacuation system (compartment door closed proximity sensor installation) in accordance with that service bulletin. \n\n\tB.\tFor airplanes identified in Boeing Alert Service Bulletin 767-25A0131, Revision 1, dated November 9, 1989: Modify the off-wing evacuation system by replacing the escape system actuator in accordance with Boeing Alert Service Bulletin 767-25A0131, Revision 1, dated November 9, 1989, or Boeing Service Bulletin 767-25-0011, Revision 2, dated October 12, 1989 or Revision 3, dated December 14, 1989.\n \n\tC.\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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant Principal Inspector (PMI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tE.\tSpecial 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. \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 Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport AirplaneDirectorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThe amendment (39-6654, AD 90-15-05) becomes effective on August 20, 1990.
76-16-08: 76-16-08 PITTS AVIATION ENTERPRISES, INC.: Amendment 39-2692. Applies to Model S-1S, Serial Numbers 1-0001 through 1-0043, and Model S-2A, Serial Numbers 2001 through 2122, airplanes certificated in all categories. Compliance required as indicated, unless already accomplished: To prevent partial loss of rudder control due to control cable slippage resulting from improper swaging of the Nicopress sleeves, accomplish the following: A. Within 10 hours' time in service after the effective date of this airworthiness directive or by September 13, 1976, whichever occurs first, inspect the 3 swaged grooves of the Nicopress part number 18-3-M sleeves in the rudder control system to assure that the major axis diameter is between .348 and .353 inches. The Nicopress "go" gage for this sleeve is part number 64-CGMP and has an opening of .353 inches. The correct opening for checking these sleeves is marked "M." The Model S-2A airplane has a total of 8 sleeves to be checked and the Model S-1S has 4. If improperly swaged sleeves are found, before further flight, reswage to the proper limits or replace as required. B. Until compliance with paragraph A is accomplished, acrobatic flight (including but not limited to, maneuvers delineated in the respective FAA Approved Airplane Flight Manuals) is prohibited and prior to further flight a placard must be installed on the instrument panel in full view of the pilot which reads: "ACROBATIC FLIGHT PROHIBITED." C. Upon compliance with paragraph A, the placard required by paragraph B may be removed. This amendment becomes effective August 20, 1976.
90-10-07: 90-10-07 BOEING: Amendment 39-6592. Docket No. 89-NM-102-AD. \n\n\tApplicability: Model 767 series airplanes, line numbers 1 through 243, 257, 262, 263, 269, 273 through 275, 282, 285, 289, and 291, certificated in any category. \n\n\tCompliance: Required within the next 24 months after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent uncommanded extension of three flight spoilers on one wing, due to a failure of a spoiler wheel command unit, accomplish the following: \n\n\tA.\tFor Group 1 airplanes (as listed in the Boeing service bulletin): Replace both spoiler wheel command units in accordance with Boeing Service Bulletin 767-27-0085, Revision 1, dated November 30, 1989. \n\n\tB.\tFor Group 2 airplanes (as listed in the Boeing service bulletin): Replace the left side spoiler wheel command unit in accordance with Boeing Service Bulletin 767-27-0085, Revision 1, dated November 30, 1989. \n\n\tC.\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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Seattle Aircraft Certification Office. \n\n\tC.\tSpecial 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. \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 Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal WaySouth, Seattle, Washington. \n\n\tThis amendment (39-6592, AD 90-10-07) becomes effective on June 12, 1990.
77-14-13: 77-14-13 HAWKER SIDDELEY AVIATION, LTD: Amendment 39-2958. Applies to Model BH-125 Series 600A airplanes, S/N's 25/6001-6004, 6007, 6009-6011, 6013, 6014, 6016, 6018, 6020, 6022-6026, 6032, 6034, 6038, 6040, 6044, 6046, certificated in all categories. Compliance is required within the next 300 hours time in service after the effective date of this AD, unless already accomplished. To prevent possible unwanted rolling of the airplane when operating at buffet onset, add vortex generators to the leading edge of each wing by incorporating Hawker Siddeley Aviation, Ltd., Modification No. 252442 in accordance with Section 2 entitled "Accomplishment Instructions" of Hawker Siddeley Aviation, Ltd., Service Bulletin 57-48-(2442), dated June 25, 1975, including Revision 1, dated July 23, 1975, or an FAA-approved equivalent. This amendment becomes effective August 8, 1977.