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86-14-08: 86-14-08 BRITISH AEROSPACE: Amendment 39-5347. Applies to Model BAe 125- 800A series airplanes listed in BAe 125 Service Bulletin 27-136-(3059A), Revision 1, dated June 26, 1985, certificated in any category. To prevent loss of stall warning, accomplish the following within the next 60 days after the effective date of this AD, unless previously accomplished: A. Incorporate a new layshaft assembly in the stall identification system in accordance with the accomplishment instructions of British Aerospace 125 Service Bulletin 27- 136-(3059A), Revision 1, dated June 24, 1985. B. 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, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. All persons affected by this directive, who have not already received the appropriate service document from the manufacturer, may obtain copies upon request to British Aerospace, Inc., Librarian, Box 17414, Dulles International Airport, Washington, D.C. 20041. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective August 4, 1986.
86-20-07: 86-20-07 MCDONNELL DOUGLAS HELICOPTER COMPANY (Hughes Helicopters, Inc.): Amendment 39-5422. Applies to Model 369, 369A, 369D, 369E, 369H, 369HE, 369HM, and 369HS helicopters, including military Models YOH-6A and OH-6A, certificated in any category, equipped with tail rotor drive shaft flexible couplings, Part Number (P/N) 369A5501 or 369H92564. Compliance required as indicated unless already accomplished. To prevent failure of the tail rotor (T/R) drive shaft system and subsequent loss of T/R control, accomplish the following: (a) Within 100 hours' time in service after the effective date of this AD, install aft coupling failsafe device (P/N's 369D25530 bolt and 369D25531 socket) in accordance with Part I of the applicable Service Information Notices (SIN) DN-143, HN-2O6, or EN-31, each dated August 26, 1986. Installation of the failsafe device on military Models YOH-6A or OH-6A helicopters in civil use shall be accomplished in accordance with Part I of SIN HN-206. NOTE: The failsafe device required by paragraph (a) will be installed before delivery on all applicable Model 369E helicopters, Serial Number 0135E, and subsequent. (b) Within 100 hours' time in service after the effective date of this AD, install forward coupling failsafe device (P/N's 369D25530 bolt and 369D25531 socket) in accordance with Part I of SIN DN-95, dated August 7, 1981, or Part III, HN-173, dated November 2, 1981, as applicable. Installation of the coupling failsafe device on military Models YOH-6A or OH-6A helicopters shall be accomplished in accordance with Part III of SIR HN-173. (c) For all helicopters with tail rotor driveshaft flexible coupling failsafe devices installed, the T/R drive shaft forward and aft flexible couplings shall be checked as follows: (1) (At Each Preflight Check: Check for T/R backlash or looseness by rocking the T/R back and forth in its plane of rotation. The blade should not move in excess of 0.75 inch (1.93cm) atthe blade tip without rotation of the main rotor blades. (2) At Each Aircraft/Engine Shutdown: If thumping or rapping is heard from the T/R drive train during final revolutions of the T/R, check the T/R to assure that the T/R blade does not move in excess of 0.75 inch (1.93cm) at the blade tip without rotation of main rotor blades. (d) The checks required by this AD may be performed by the pilot and must be recorded in accordance with FAR Section 91.173. (e) If during the checks required by paragraph (c), the tail rotor blade tip movement exceeds the specified limits, prior to further flight, inspect and replace, as necessary, either or both fore and aft tail rotor drive shaft couplings. (f) Rotorcraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the modifications and inspections of paragraphs (a) and (b) of this AD can be accomplished. (g) An alternate method of compliance which provides an equivalent level of safety may be approved by the Manager, Western Aircraft Certification Office, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009-2007. The procedure shall be done in accordance with applicable parts of MDHC SIN's DN- 143, HN-206, EN-31, all dated August 26, 1986; MDHC SIN DN-95, dated August 7, 1981; MDHC SIN HN-173, dated November 2, 1981. The 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 McDonnell Douglas Helicopter Company, Centinela Avenue and Teal Street, Culver City, California 90230. These documents may be examined at the Office of the Regional Counsel, Federal Aviation Administration, Southwest Region, Room 158, Building 3B, 4400 Blue Mound Road, Fort Worth, Texas 76101, the Western Aircraft Certification Office, 15000 Aviation Boulevard, Hawthorne, California, or the Office of the FEDERAL REGISTER, 1100 L Street, NW., Room 8401,Washington, D.C. This amendment supersedes Amendment 39-4186 (46 FR 40868), AD 81-17-02, as amended by Amendment 39-4221 (46 FR 46566), AD 81-17-02R1. This amendment becomes effective October 24, 1986.
2010-21-09: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: A damaged fuel heater caused a fuel leakage in the engine nacelle; investigation revealed that the damage to the fuel heater was due to chafing with an oil cooling system hose. Piaggio Aero Industries (PAI) issued Service Bulletin (SB) 80- 0175, which was applicable to all aeroplanes and contained instructions for a repetitive inspection of the affected parts and, if necessary, their replacement and/or for the repositioning of oil/ fuel tubing if minimum clearances were not found. We are issuing this AD to require actions to correct the unsafe condition on these products.
90-23-05: 90-23-05 GENERAL ELECTRIC COMPANY: Amendment 39-6773. Docket No. 90-ANE-08. Applicability: General Electric Company (GE) CF6-80A3 turbofan engines installed on, but not limited to, Airbus A310-200 aircraft. Compliance: Required at the next engine removal or within 18 months after the effective date of this AD, whichever occurs first, unless already accomplished. To prevent failure of the engine aft mount, which could result in engine separation, accomplish the following: (a) Conduct an "in shop" dip etch and fluorescent penetrant inspection of the engine aft upper mount beam, Part Number (P/N) 224-1606-501 or 224-1606-503, and engine aft lower mount beam, P/N 224-1607-501, in accordance with the accomplishment instructions contained in Part 2 of GE CF6-80A Series Service Bulletin (SB) 71-053, Revision 2, dated June 26, 1990. (b) Remove from service prior to further flight, engine aft upper and lower mounts with crack indications and replace with serviceable parts.(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, Engine Certification Office, ANE-140, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. The fluorescent penetrant inspection of engine aft mount beam assemblies shall be done in accordance with the following GE document: DOCUMENT PAGE REVISION DATE GE SB 71-053 1, 2 2 June 26, 1990 GE SB 71-053 3-8 1 February 8, 1990 GE SB 71-053 9, 10, 11 2 June 26, 1990 GE SB 71-053 12 1 February 8, 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 General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the Office of the Federal Register, 1100 L Street, NW, Room 8301, Washington, DC 20591. This amendment (39-6773, AD 90-23-05) becomes effective on December 3, 1990.
75-22-04: 75-22-04 HUGHES HELICOPTERS: Amendment 39-2289. Applies to Hughes Model 369, 369A, 369H, 369HM, 369HS, and 369HE helicopters certificated in all categories, including military YOH-6A and OH-6A equipped with fiberglass tail rotor blades P/N 369A1710, 369A1710-9, 369A1710-11, 369-6120, 369A1607, and 369CSK22. Compliance required as indicated. To detect possible corrosion, cracks, or other defects, inspect by visual, X-ray, or other specified means, the affected tail rotor blades and replace or rework in accordance with the instructions specified in Hughes Service Information Notice (SIN) No. HN-88, dated August 28, 1975, or later FAA-approved revisions, as follows: (a) For blades with 500 or more hours time in service on the effective date of this AD, perform the visual and X-ray inspection, corrosion removal, casting procedure, metal treatment procedure, corrosion protection procedure, and fiberglass inspection - repair/spar exterior inspection procedure, set forth atParts I through VIII of the Hughes SIN, referenced above, within the next 100 hours additional time in service or within six calendar months from the effective date of the AD, whichever occurs first, unless already accomplished. (b) For blades with less than 500 hours time in service on the effective date of this AD, perform the visual and X-ray inspections, corrosion removal, casting procedures, metal treatment procedure, corrosion protection procedure, and fiberglass inspection - repair/spar exterior inspection procedure, set forth at Parts I through VIII of the Hughes SIN, referenced above, prior to accumulating 600 hours total time in service or within six calendar months from the effective date of this AD, whichever occurs first, unless already accomplished. (c) After the effective date of this AD, perform the inspections and procedures described at Parts I through VIII of the Hughes SIN, referenced above, prior to the installation of spare blades or rotors on the aircraft. (d) After the effective date of this AD, for all blades, perform the visual and X-ray inspections described at Part X of the Hughes SIN, referenced above, at intervals not to exceed 12 calendar months from the last inspection. (e) After the effective date of this AD, repair or rework eligible blades as specified in the Hughes SIN, referenced above, as necessary, prior to further flight. Reinstall blades in accordance with Part IX of the Hughes SIN. Blades that exceed limits specified in the Hughes SIN and are therefore not repairable, must be marked in a conspicuous manner or destroyed so as to prevent inadvertent return to service. (f) Paragraphs (a), (b), and (c), above, do not have to be accomplished on blades marked with a green dot or white dot per the preface, Hughes SIN. After the effective date of this AD, perform the visual and X-ray inspections for corrosion described at Part X of the Hughes SIN, referenced above, within twelve months after putting theblades into service, and at intervals not to exceed twelve months thereafter. (g) Equivalent inspections and rework may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (h) Special flight permits may be issued for operating aircraft to a base for performance of the inspections and repairs or rework required by paragraphs (a) and (b), above, of this AD, per FAR's 21.197 and 21.199. This amendment becomes effective October 23, 1975.
98-11-32: This amendment adopts a new airworthiness directive (AD) that is applicable to Allison Engine Company AE 3007A and AE 3007C series turbofan engines. This action supersedes priority letter AD 98-02-09, that currently requires certain checks of the center sump magnetic chip collector plug for paste. Engines found with paste are required to be removed from service. This action references revisions of the applicable Alert Service Bulletins (ASB) providing clarifications of check procedures. This amendment is prompted by a change in the part number applicability, a change in the check interval, and the publication of these revised ASBs. The actions specified by this AD are intended to prevent No. 4 bearing failure due to excessive bearing wear, which can result in an inflight engine shutdown. DATES: Effective June 18, 1998 The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of June 18, 1998. Comments for inclusion in the Rules Docket must be received on or before August 3, 1998.
99-02-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A330-301, -321, -322, -341, -342, and A340-211, -212, -213, -311, -312, and -313 series airplanes. This action requires repetitive high-frequency eddy current (HFEC) inspections to detect cracking of the inner flange of the rear fuselage frame FR73A, between beams 5 and 6; and corrective actions, if necessary. This amendment also provides for optional terminating action for the repetitive inspections. 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 detect and correct fatigue cracking of the inner flange of the rear fuselage frame FR73A, which could result in reduced structural integrity of the fuselage.
99-02-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320 and A321 series airplanes. This amendment requires modification of the slat and flap control computer (SFCC) in the aft electronics rack. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent failure of the SFCC caused by computer software anomalies or contamination by conductive dust. This condition, if not corrected, could result in uncommanded slat retraction during takeoff and consequent insufficient wing lift available to complete a successful takeoff.
97-10-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A310 series airplanes, that requires repetitive inspections to detect discrepancies or damage of the steady bearing assemblies of the flap transmission system, and replacement of any discrepant or damaged assembly with a new, like assembly. This amendment also requires eventual replacement of all the steady bearing assemblies with new, improved assemblies, which terminates the repetitive inspection requirements. This amendment is prompted by reports of cracking of the hardened steel inner race, and broken or missing inner races of the steady bearing assemblies. The actions specified by this AD are intended to prevent such discrepancies and damage of the shafts of the steady bearing assemblies, which could cause the shafts to fail; failure of the steady bearing shafts during a subsequent asymmetric stop could result in an uncommanded asymmetric retraction of the flap, and subsequentreduced controllability of the airplane.
75-01-03: 75-01-03 ROCKWELL INTERNATIONAL: Amendment 39-2061. Applies to all NA- 265, NA-265-20, -30, -40, -60, -70 and -80 model airplanes, certificated in all categories. Compliance required as indicated. To provide temporary operating limitations on airplanes affected, pending modification of the landing gear warning system to ensure continuous functioning of the aural warning device under the conditions of CAR 4b, accomplish the following: (1) The following operating limitation is hereby adopted effective ten days after the effective date of this AD, applicable to NA-265-60, -70 and -80 model airplanes: "MAXIMUM TAKEOFF AND LANDING PRESSURE ALTITUDE - 8,000 FEET." (2) For NA-265-60, -70 and -80 model airplanes, within ten days after the effective date of this AD, unless already accomplished, install a placard: "MAXIMUM TAKEOFF AND LANDING PRESSURE ALTITUDE - 8,000 FEET. GEAR WARNING HORN MAY NOT SOUND ABOVE 125 KIAS WITH FLAPS LESS THAN 80%." (3) For NA-265, NA-265-20, -30 and -40 model airplanes, within ten days after the effective date of this AD, unless already accomplished, install a placard: "GEAR WARNING HORN MAY NOT SOUND ABOVE 125 KIAS WITH FLAPS LESS THAN 80%." (4) Within 9 months after the effective date of this AD, unless already accomplished, remove and replace the altitude and airspeed switch, in accordance with Rockwell International Sabreliner Service Bulletin 74-32, dated December 18, 1974, or later FAA-approved revisions. (5) Equivalent installations may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiating data. (6) After accomplishing the work required by paragraph 4, above, or FAA-approved equivalent per paragraph 5, the operating limitation imposed by paragraph 1, above, will no longer apply and the placards specified in paragraphs 2 and 3, above, must be removed. (7) Airplanes may be flown to a base for accomplishment of the installation required by paragraph 4, above, per FAR's 21.197 and 21.199. This amendment becomes effective January 6, 1975.