Results
80-25-08: 80-25-08 BOEING VERTOL (VERTOL): Amendment 39-3987. Applies to Vertol Model 107-II helicopters certificated in all categories. Compliance required as indicated. To prevent fatigue failure of the main rotor tension-torsion strap assemblies, remove from service tension-torsion strap assemblies Part No. 107R2003-1 upon the accumulation of 27,800 hours in service and replace with an airworthy part that meets the requirement of this AD. This amendment is effective December 12, 1980.
2021-25-07: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model BD-100-1A10 airplanes. This AD was prompted by a discovery that a lockwire may not have been installed on the side stay actuator pin nut of the main landing gear (MLG). This AD requires inspecting the left-hand and right-hand MLG side stay actuator assembly pin nut for the presence of a lockwire, and installing a lockwire if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
73-16-06: 73-16-06 PITTS: Amdt. 39-1699. Applies to S-2A airplanes, Serial Numbers 2001 through 2059 certificated in all categories. Compliance required as indicated. To prevent the possibility of a power loss due to the failure of the induction air box flapper door, accomplish the following: (a) Within the next 50 hours time in service after the effective date of this airworthiness directive, unless already accomplished, install stiffener, Pitts Part Number 2-7112- 16, to the induction air box flapper door assembly, Pitts Part Number 2-7112-7, in accordance with Pitts Service Letter No. 2. Access to the induction air box flapper door assembly may be gained by removing the lower engine cowl. (b) Until the modification specified in paragraph (a) above has been accomplished, visually check the induction air box flapper assembly, Pitts Part Number 2-7112-7, for cracks in the welded areas upon the accumulation of 50 hours time in service or within 5 hours time in service, whichever is later, and thereafter at intervals not to exceed 5 hours time in service from the last check. The checks required by this AD may be performed by the pilot. If the flapper door in the area of the weld is found to be cracked, remove the flapper door assembly from the alternate air box, and add stiffener, Pitts Part Number 2-7112-16 in accordance with Pitts Service Letter No. 2, before further flight. (c) The installation of the stiffener, Pitts Part Number 2-7112-16, in accordance with Pitts Service Letter No. 2, will eliminate the necessity for the repetitive inspections required in paragraph (b). Pitts Service Letter No. 2, dated June 26, 1973, pertains to this subject. This amendment becomes effective August 13, 1973.
95-24-01: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas DC-10-10 series airplanes, that requires inspections of the wings to detect cracks in the aft spar lower cap, in certain stringer butterfly clips on the bulkheads, and in certain fastener holes; and repair, if necessary. This amendment also requires modification of those areas of the wings, which terminates the repetitive inspection requirements. This amendment is prompted by reports indicating that, during fatigue testing of the wing structure, cracks developed in the aft spar lower cap, in certain stringer butterfly clips, and in certain fastener holes due to fatigue-related stress. The actions specified by this AD are intended to prevent such fatigue- related cracking, which could lead to the failure of the aft spar cap and consequently could reduce the structural integrity of the wing.
74-23-02: 74-23-02 BELL: Amendment 39-1999. Applies to Bell Model 204B and 205A-1 helicopters certificated in all categories. Compliance required as indicated. To detect cracks in the tail rotor pitch change link segments and to prevent possible failure of the tail rotor pitch change chains accomplish the following repetitive inspections on chains, P/N 205-001-721-1 and 205-001-748-1. (a) Within five hours time in service after June 7, 1974, unless already accomplished, and thereafter at intervals not to exceed 25 hours time in service from the last inspection, accomplish the following inspections: (1) Remove the chain or chains from the helicopter in accordance with the applicable maintenance manual. (2) Clean each chain with solvent and stiff bristle brush and air dry. (3) Roll each chain into a tight flat disc and inspect the links' outer segments, both sides, for cracks, using a light and a ten-power or higher magnifying glass. Unroll the chain and withteeth up inspect inner segments of the links at the radius between the teeth. Turn each chain over and inspect the opposite side of the link segments. (4) Replace chains with cracked segments before further flight. (5) Install chains with uncracked link segments in accordance with the pertinent maintenance manual and rig tail rotor controls in accordance with the pertinent maintenance manual. If 212-010-701 Tail Rotor Hub and Blade Assembly is installed on the Model 205A-1, rig the controls in accordance with the maintenance manual and as specified in Bell Helicopter Company Service Bulletin No. 205-05-74-1, Rev. A, dated June 24, 1974, or later approved revision. (b) Before the first flight of each day after June 7, 1974, accomplish the following repetitive inspections. (1) Remove the cover from the chain assembly. (2) Inspect each chain assembly for cracks in the link segments using a three-power or higher magnifying glass. Particular attention should be placed on the portion of each chain that travels over each sprocket and that extends three inches each side of this area or portion. (3) Remove chains with cracked segments before further flight in accordance with the applicable maintenance manual. (4) Install chains with uncracked segments in accordance with the applicable maintenance manual and rig the controls as specified in paragraph (a)(5) of this airworthiness directive. (c) Replace chains, P/N 205-001-721-1 having manufacturing dates of June 20, 1974, or later etched on the clevis, as follows: (1) Replace chains with more that 140 hours total time in service on the effective date of this Airworthiness Directive within ten hours time in service. (2) Replace chains with less than 140 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 150 hours total time in service. (d) Replace chains, P/N 205-001-721-1 having manufacturing dates prior to June 20, 1974, as follows: (1) Replace chains with more than 40 hours total time in service on the effective date of this Airworthiness Directive within ten hours time in service. (2) Replace chains with less than 40 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 50 hours total time in service. (e) Replace chains, P/N 205-001-748-1, with more than 140 hours total time in service on the effective date of this Airworthiness Directive within ten hours total time in service. Replace chains, P/N 205-001-748-1, with less than 140 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 150 hours total time in service. 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 themanufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Company, P. O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this Airworthiness Directive which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D. C. and at the FAA Southwest Regional Office, Fort Worth, Texas. (Bell telefax messages dated May 10, 1974; June 15, 1974; August 2, 1974; and September 25, 1974, pertain to this subject, and Information Letter to all Bell 204B/205A-1 operators, 35:WJD:jge-3152 dated August 12, 1974, also pertains to this subject.) This supersedes Amendment 39-1963 (39 F.R. 34054), A.D. 74-20-06. This amendment becomes effective November 3, 1974.
82-02-03: 82-02-03 BEECH: Amendment 39-4313. Applies to Model 76 (Serial numbers ME-1 through ME-435) airplanes certificated in any category. COMPLIANCE: Required as indicated unless already accomplished. To ensure the integrity of the elevator control cable and determine proper cable routing, accomplish the following: A) Prior to further flight, accomplish all of the following: 1. Remove the large inspection panel located in the center of the bottom skin just aft of the Station 68 frame. 2. Drill a 3/8-inch diameter inspection hole in the forward flange to which the inspection panel was attached as follows: Center the hole laterally on the elevator cable pulley and fore and aft on the flange. Use a thin bucking bar or wood block between the flange and the pulley and cable to prevent damage to the pulley and/or cable when drilling the inspection hole. NOTE: Airplane Serial Numbers ME-418 and after have the 3/8-inch diameter inspection hole already drilledin the forward flange. 3. Deburr the inspection hole and visually inspect the elevator down cable for broken or frayed strands and ensure the cable is on the pulley under all three guard pins. 4. If the cable is riding over any of the guard pins, replace the P/N NAS 427K12 guard pins over which the cable was routed, replace the P/N NAS 314-25-1411 cable, and rig in accordance with the Beech Model 76 Maintenance Manual. 5. Reinstall the inspection panel, and record compliance with this AD by an appropriate entry in the airplane maintenance records. B) Within 48 hours, report misrouted cables or other defects found as a result of any inspection required herein to the FAA via a Malfunction or Defect (M or D) Report (FAA Form 8010-4) or a letter to the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209. Describe the defect found, total time-in-service on the airplane or part at time of discovery, and the aircraft serial number. (Reporting approved by the Office of Management and Budget under OMB No. 04-R0174). C) Airplanes may be flown in accordance with FAR 21.197 to a location where the provisions of Paragraph A) of this AD can be performed provided the following is accomplished: 1. Remove the large inspection panel located in the center of the bottom skin just aft of the Station 68 frame. 2. Visually inspect the elevator down cable for broken or frayed strands while the elevator control column is slowly moved fore and aft. Pay particular attention to the cable strands near the elevator cable pulley. 3. If no frayed or broken cable strands are found, reinstall the inspection panel. D) Any equivalent method of compliance with this AD must be approved by the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 269-7000. Beechcraft Safety Communique 76-62 pertains to the subject matter of this AD. This amendment becomes effective February 14, 1982, to all persons except those to whom it has already been made effective by a priority mail letter from the FAA dated January 8, 1982.
90-02-18 R1: 90-02-18 R1 SOCATA GROUPE AEROSPATIALE: Amendment 39-6454 as revised by Amendment 39-6619. Docket No. 89-CE-37-AD. Applicability: Models TB 9 and TB 10, TB 20, and TB 21 (all serial numbers (S/N)) airplanes certificated in any category. Compliance: As indicated in the body of the AD, unless already accomplished per AD 90-02-18. To preclude loss of power due to contamination of the fuel system, accomplish the following: (a) Within the next 75 hours time-in-service (TIS) from February 6, 1990, except as indicated in paragraph (c) of this AD, modify the fuel system by the installation of the following applicable SOCATA modification kit, as described in SOCATA Service Bulletin (SB) Number 48/2, dated March 1990: Airplanes Kit Number All TB airplanes (S/N 1 through 822, 850 through 887, 889 and subsequent) 9154 Airplanes Kit Number TB 20 (S/N 823 through 849, and 888) 9155 (b) For Models TB 20 and TB 21 (S/N 1 through 730)airplanes (unless modified with SOCATA Modification Number 66), within the next 75 hours TIS from February 6, 1990, except as indicated in paragraph (d) of this AD, modify the fuel system by replacement of the installed Dukes fuel pump with a Weldon fuel pump and the addition of a check valve, in accordance with the instructions contained in SOCATA SB Number 47/1, dated October 1989. (c) If the required parts are not available to accomplish the modification specified in paragraph (a) of this AD, the airplane may continue operation for an additional 150 hours TIS after the compliance time specified in paragraph (a) of this AD provided the fuel tank sump is drained at intervals not to exceed each 50 hours TIS in accordance with the procedures specified in the DESCRIPTION section of SOCATA SB 48/2, dated March 1990. (d) If the required parts are not available to accomplish the modifications specified in paragraph (b) of this AD, the airplane may continue operation for an additional 150 hours TIS after the compliance time specified in paragraph (b) of this AD provided the following preflight actions are accomplished by the pilot prior to each engine start: (i) Select battery (main switch) ON. (ii) Advance the mixture control to FULL RICH. (iii) Select electric fuel boost pump ON. (iv) Advance the throttle until a positive fuel flow is observed on the fuel flow gauge, then retard the throttle and move the mixture control to IDLE/CUTOFF. (v) Select electric fuel boost pump OFF. (vi) Select battery (main switch) OFF. (vii) Visually inspect the electric fuel boost pump area for leaks. (viii) If no positive fuel flow is observed on the fuel flow gauge, or fuel leaks are detected from the electric fuel boost pump, repair or replace the defective component prior to further flight. NOTE: Avoid moving the propeller and standing in the propeller area while inspecting the engine. (e) Airplanes may be flown in accordance withFAR 21.197 to a location where this AD may be accomplished. (f) 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, Brussels Aircraft Certification Office, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, B-1000 Brussels, Belgium. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to SOCATA Groupe Aerospatiale, B.P. 38, 65001 Tarbes, Cedex, France; Telephone 62.51.73.00, or 62.51.73.55 (for Telefax); or the Product Support Manager, U.S., AEROSPATIALE, 2701 Forum Drive, Grand Prairie, Texas 73053; Telephone (214) 641-3614; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This AD revises AD 90-02-18, Amendment 39-6454. This amendment (39-6619, AD 90-12-18 R1) becomes effective on June 13, 1990.
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.
2021-25-13: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model CL-600-1A11 (600), CL-600-2A12 (601), and CL-600-2B16 (601-3A, 601-3R, and 604 Variants) airplanes. This AD was prompted by a determination that new or more restrictive airworthiness limitations are necessary. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations. The FAA is issuing this AD to address the unsafe condition on these products.
98-03-11: This amendment adopts a new airworthiness directive (AD), applicable to all Airbus Model A300, A310, and A300-600 series airplanes. For certain airplanes, this amendment requires replacing the bearings of the throttle control levers with new sealed bearings. For certain other airplanes, this amendment requires replacing the throttle control assemblies with new assemblies. This amendment is prompted by the issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent asymmetric engine thrust on the airplane when the autothrottle is engaged, which could result in roll and yaw disturbances, and consequent reduced controllability of the airplane.
2021-24-14: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model 787-8, 787-9, and 787-10 airplanes. This AD was prompted by reports of damage to the thrust reverser (TR) translating sleeve secondary sliders due to contact between the slider and the slider track liner. This damage could reduce the fatigue life of the slider below its full design life for the TRs installed on certain engines. This AD requires determining the serial number of the TR and performing applicable on-condition actions; or replacing the TR with a serviceable TR. The FAA is issuing this AD to address the unsafe condition on these products.
89-10-11 R1: 89-10-11 R1 BELL HELICOPTER TEXTRON, INC.: Amendment 39-6211 as corrected by Amendment 39-6283. (Airworthiness Docket No. 88-ASW-55) Applicability: Bell Helicopter Textron, Inc., Model 206A and 206B helicopters certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent possible in-flight engine flameout, which could result in the loss of the helicopter, accomplish the following: (a) Within the next 20 days after the effective date of this AD, revise the following Rotorcraft Flight Manual and Rotorcraft Flight Manual Supplements, as applicable. (1) For Model 206A helicopters, revise the FAA approved Rotorcraft Flight Manual (RFM) No. BHT-206A-FM-1 Operating Limitations, Section 1, Type of Operation, by adding the following: The following equipment shall be installed when conducting flight operations in falling and/or blowing snow to prevent possibility of engine flameout: The Particle Separator Engine Air Induction System Kit and the Deflector Kit. (See BHT-206A-FMS-18 and BHT-206A-FMS-24.) or The Snow Winterization Air Induction System. (See BHT-206A-FMS- 11.) (2) For Model 206A helicopters, revise the FAA approved Rotorcraft Flight Manual Supplement (RFMS) Nos. BHT-206A-FMS-18 and BHT-206A-FMS-24 Operating Limitations, Section 1, Type of Operation, by adding the following: The FAA approved Engine Air Induction Deflector Kit No. 206-706-136 shall be installed in conjunction with the Particle Separator Air Inducting Kit No. 206-706-201-17 or 206-706-200-9 when conducting flight operations in falling and/or blowing snow, and the following limits apply: Hover flight in falling and/or blowing snow is limited to a 20-minute duration after which the helicopter must be landed and checked for snow and/or ice accumulation. Flight operations are prohibited when visibility in falling and/or blowing snow is less than one- half (1/2) statute mile. (3) For Model 206B (206B II) helicopters, revise the FAA approved RFM No. BHT-206B-FM-1 Operating Limitations, Section 1, Type of Operation, by adding the following: The following equipment shall be installed when conducting flight operations in falling and/or blowing snow to prevent possibility of engine flameout: The Particle Separator Engine Air Induction System Kit and The Deflector Kit. (See BHT-206B-FMS-15 and BHT-206B-FMS-18.) (4) For Model 206B (206B II) helicopters, revise FAA approved RFMS Nos. BHT-206B-FMS-15 and BHT-206B-FMS-18 Operating Limitations, Section 1, Type of Operation, by adding the following: The FAA approved Engine Air Induction Deflector Kit No. 206-706-136 shall be installed in conjunction with the Particle Separator Air Induction Kit No. 206-706-200-5 or 206-706-201-11 when conducting flight operations in falling and/or blowing snow, and the following limits apply: Hover flight in falling and/or blowing snow is limited to a 20-minute duration after which the helicopter must be landed and checked for snow and/or ice accumulation. Flight operations are prohibited when visibility in falling and/or blowing snow is less than one- half (1/2) statute mile. (5) For Model 206B (206B III) helicopters, revise the FAA approved RFM No. BHT-206B3-FM-1 Operating Limitations, Section 1, Type of Operation, by adding the following: The following equipment shall be installed when conducting flight operations in falling and/or blowing snow to prevent possibility of engine flameout: The Particle Separator Engine Air Induction System Kit and the Deflector Kit. (See BHT-206B3-FMS-10 and BHT-206B3-FMS-12.) (6) For Model 206B (206B III) helicopters, revise the FAA approved RFM Nos. BHT-206B3-FMS-10 and BHT-206B3-FMS-12 Operating Limitations, Section 1, Type of Operation, by adding the following: The FAA approved Engine Air Induction Deflector Kit No. 206-706-136 shall be installed in conjunctionwith the Particle Separator Air Induction Kit No. 206-706-201 or 206-706-200 when conducting flight operations in falling and/or blowing snow, and the following limits apply: Hover flight in falling and/or blowing snow is limited to a 20-minute duration after which the helicopter must be landed and checked for snow and/or ice accumulation. Flight operations are prohibited when visibility in falling and/or blowing snow is less than one- half (1/2) statute mile. (b) Adding the following RFM and RFMS revisions to the basic RFM manual, as applicable, is an approved alternate method of compliance with paragraph (a) of this AD: BHT-206A-FM-1, Rev. D-39 BHT-206A-FMS-18, Rev. 6 BHT-206A-FMS-24, Rev. 3 BHT-206B-FM-1, Rev. B-39 BHT-206B-FMS-15, Rev. 4 BHT-206B-FMS-18, Rev. 3 BHT-206B3-FM-1, Rev. 18 BHT-206B3-FMS-12, Rev. 1 BHT-206B3-FMS-10, Rev. 1 (c) An alternate method of compliance, which provides an equivalent level of safety, may be used when approvedby the Manager, Helicopter Certification Branch, Federal Aviation Administration, Fort Worth, Texas 76193-0170. Airworthiness Directive 89-10-11 (Amendment 39-6211) became effective on June 8, 1989. This amendment (39-6283, AD 89-10-11 R1) becomes effective on July 31, 1989.
2010-20-20: This amendment supersedes an existing airworthiness directive (AD) for the specified Eurocopter France (Eurocopter) helicopters. That AD requires repetitively inspecting the main gearbox (MGB) planet gear carrier for a crack and replacing any MGB that has a cracked planet gear carrier before further flight. This action requires the same inspections required by the existing AD, but shortens the initial inspection interval. This AD is prompted by the discovery of another crack in a MGB planet gear carrier and additional analysis that indicates that the initial inspection interval must be shortened. The actions specified by this AD are intended to detect a crack in the web of the planet gear carrier, which could lead to a MGB seizure and subsequent loss of control of the helicopter.
87-24-11: 87-24-11 GARRETT TURBINE ENGINE COMPANY: Amendment 39-5781. Applies to Garrett Model GTCP 85 series Auxiliary Power Units equipped with one-piece cast turbine wheels, Part Nos. 968095-X, 3604604-X, 3606982-1, and 3842072-1. Compliance required as indicated, unless previously accomplished. To prevent turbine wheel separation and resulting containment shroud fragmentation, accomplish the following: A. Install the augmentation containment ring, part number 3612249-1, in Garrett Model GTCP 85 series auxiliary power units in accordance with the accomplishment instructions of Garrett Service Bulletin GTCP 85-49-5689, dated July 24, 1987, or later revisions approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region, as follows: 1. On in-flight operable units, within 18 months after the effective date of this AD. 2. On ground-operable-only units, within 36 months after the effective date of this AD. B. Installation ofthe Hastelloy S turbine shroud, in accordance with Garrett Service Bulletin 85-49-700, dated July 20, 1987, or later FAA approved revisions, is considered an acceptable alternate means of compliance with this AD. C. An alternate means of compliance with this AD which provides an acceptable level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a base 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 Garrett Aviation Services Company, Data Distribution, Department H64-5, P.O. Box 29003, Phoenix, Arizona 85038. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at 4344 Donald Douglas Drive, Long Beach, California. This amendment becomes effective January 13, 1988.
90-10-01: 90-10-01 AEROSPATIALE (FORMERLY SUD AVIATION/SUD-SERVICE): Amendment 39-6585. Docket No. 90-NM-03-AD. Applicability: All Aerospatiale Caravelle Model SE 210 Model series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To detect fatigue cracks in the wing-to-fuselage junction and fuselage lower section components, accomplish the following: A. Prior to the accumulation of 29,000 landings, or within 2,000 landings after the effective date of this AD, whichever occurs later, perform one of the following inspections in accordance with Aerospatiale Service Bulletin 53-51, Revision 6, dated July 31, 1986: 1. Visual inspection of the wing-to-fuselage junction fittings at Frames 31 and 35, in accordance with the service bulletin. Repeat this inspection thereafter at intervals not to exceed 2,500 landings; or 2. Visual and eddy current inspection of the wing-to-fuselage junction fittings at Frames 31and 35, and a magnetic particle inspection of all removed horizontal bolts, in accordance with the service bulletin. Repeat this inspection thereafter at intervals not to exceed 7,500 landings. B. If cracks or sheared bolts or rivets are detected during the inspection required by paragraph A., above, replace with serviceable parts prior to further flight, in accordance with Aerospatiale Service Bulletin 53-51, Revision 6, dated July 31, 1986. Repeat inspections thereafter at intervals specified in paragraph A., above. C. Prior to the accumulation of 29,000 landings, or within 2,000 landings after the effective date of this AD, whichever occurs later, and thereafter at intervals not to exceed 2,500 landings, perform a visual and dye penetrant inspection of the fuselage lower section beam under Rib 52, the attachment fittings at the rear section of Rib 52, and the front pick-up fittings between the beam under Rib 52 and the canted frames at Frame 30, in accordance with Aerospatiale Service Bulletin 53-52, Revision 7, dated March 22, 1988. D. If sheared-off attachment hardware is detected during the inspection required by paragraph C., above, prior to further flight replace with serviceable parts in accordance with the Structural Repair Manual, Chapter 57-1-33. Repeat the inspection identified in paragraph C., above, at intervals not to exceed 2,500 landings. E. If cracks are detected during the inspection required by paragraph C., above, in Fittings 210.12.52.382/383 and 210.22.53.012 at the attachment level, prior to further flight, repair in a manner approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. Repeat the inspection identified in paragraph C., above, at an interval to be approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. F. If cracks are detected during the inspection required by paragraph C., above, in Fittings 210.12.52.382/383 and 210.22.53.012 outside the attachment areas, prior to further flight, replace with new fittings, in accordance with Aerospatiale Service Bulletin 53-52, Revision 7, dated March 22, 1988. Repeat the inspection identified in paragraph C., above, at intervals not to exceed 2,500 landings. G. 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. 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. H. 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 mayobtain copies upon request to Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. 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 (39-6585, AD 90-10-01) becomes effective on June 1, 1990.
47-11-02: 47-11-02 NAVION: (Was Mandatory note 2 of AD-782-3.) Applies to All Models Equipped With Hartzell HC12x20-1 Propeller Hubs and 8628 Blades. To be accomplished not later than May 1, 1947. Vibration tests of the Hartzell HC12x20-1 propeller with these airplanes indicate that the propeller diameter should be reduced from 86 inches to 84 inches. This is accomplished by cutting 1 inch from the tip of each 8628 blade, and making the shortened blade 18428R. This blade rework must be performed either by the Hartzell factory or by a certified propeller repair agency. (Par. B of North American Field Service Bulletin No. 20 dated January 28, 1947, covers this rework.) Upon compliance with this AD, the presently required placard against engine operation between 1,950 and 2,150 r.p.m. and over 2,250 r.p.m. may be removed.