Results
93-25-17: 93-25-17 GENERAL ELECTRIC: Amendment 39-8781. Docket No. 92-ANE-08. Applicability: General Electric (GE) Models CT7-5A2, -7A, and -9C Turboprop Engines; and Models CT7-2A, and -6 Turboshaft Engines, incorporating gas generator turbine (GGT) disks and cooling plates, as listed by Part Number (P/N) and Serial Number (S/N) in GE CT7 Turboprop Service Bulletin (SB) A72-252, dated August 31, 1990; GE CT7 Turboshaft SB A72-17, and GE CT7 Turboshaft SB A72-18, both dated September 10, 1990, installed on, but not limited to, Saab 340A, Casa CN235-10, Bell 214ST, and European Helicopter Industries EH101. Compliance: Required as indicated, unless accomplished previously. To prevent fatigue cracks that can result in uncontained engine failure, accomplish the following: (a) Remove from service the affected GE Model(s) CT7-5A2, -7A, and -9C engines GGT rotor stage 1 and 2 disks, and stage 2 forward and aft cooling plates, prior to exceeding the "Max total allowable cycles"; as listed in GE CT7 Turboprop SB A72-252, dated August 31, 1990, and replace with serviceable parts. (b) Remove from service the affected GE Model(s) CT7-6 engine GGT rotor stage 1 disk and stage 2 forward cooling plate prior to exceeding the "Max total allowable cycles"; as listed in GE CT7 Turboprop SB A72-17, dated September 10, 1990, and replace with serviceable parts. (c) Remove from service the affected GE Model(s) CT7-2A engine GGT rotor stage 2 disk and stage 2 forward cooling plate prior to exceeding the "Max total allowable cycles"; as listed in GE CT7 Turboshaft SB A72-18, dated September 10, 1990, and replace with serviceable parts. (d) 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. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, EngineCertification 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. (e) 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. (f) The removals and replacements of the affected rotor stage disks and cooling plates shall be done in accordance with the following service bulletins: DOCUMENT NO. PAGE NO. ISSUE DATE GE CT7 Turboprop SB A72-252 1-13 Original Aug 31, 1990 TOTAL PGS: 13 GE CT7 Turboshaft SB A72-17 1-4 Original Sep 10, 1990 TOTAL PGS: 4 GE CT7 Turboshaft SB A72-18 1-4 Original Sep 10, 1990 TOTAL PGS: 4 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, 1000 Western Avenue, Lynn, Massachusetts 01910. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, Attn: Rules Docket No. 92-ANE-08, 12 New England Executive Park, Burlington, Massachusetts; or at the Office of the Federal Register, 800 North Capitol Street NW., suite 700, Washington, DC. (g) This amendment becomes effective on February 2, 1994.
90-25-13: 90-25-13 AEROSPATIALE (FORMERLY SUD AVIATION/SUD-SERVICE): Amendment 39-6820. Docket No. 90-NM-159-AD. Applicability: All Caravelle SE 210 Model I, III, and VIR series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To detect cracks in the rear spar upper cap at Rib 49 and to prevent reduced structural integrity of the wings, accomplish the following: A. Prior to the accumulation of 25,000 landings, or within 120 days after the effective date of this AD, whichever occurs later, perform an eddy current rototest inspection of hole No. 1 (left and right wings) at the level of the rear spar upper cap (angle extrusion) and the splice under Rib 49, in accordance with Aerospatiale Service Bulletin 57-69, dated March 12, 1990. If no crack is detected in hole No. 1, the airplane may be returned to service. B. If a crack is detected in the angle extrusion for hole No. 1, accomplish the following in accordance with Aerospatiale Service Bulletin 57-69, dated March 12, 1990: 1. Prior to further flight, repair the crack and perform an eddy current rototest inspection of holes No. 2 and 3 (left and right wings), in accordance with the service bulletin. 2. If no crack is found in holes No. 2 and 3, the airplane may be returned to service. Repeat the eddy current rototest inspection of hole No. 2 at intervals not to exceed 1,500 landings in accordance with the service bulletin. 3. If a crack is detected in holes No. 2 or 3, prior to further flight, repair in a manner approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. C. 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, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM-113, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence 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 Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington 98055-4056. This amendment (39-6820, AD 90-25-13) becomes effective on January 7, 1991.
91-14-12: 91-14-12 BRITISH AEROSPACE: Amendment 39-7053. Docket No. 91-NM-44-AD. Applicability: Model HS.125-600A and BH.125-600A series airplanes (Post Modification 252475) and Model HS.125-700A series airplanes (Post Modification 252509); as listed in British Aerospace Service Bulletin 24-279-3255A, dated November 16, 1990; certificated in any category. Compliance: Required within 160 days after the effective date of this AD, unless previously accomplished. To prevent loss of the standby constant frequency power system which provides the necessary back-up capability when the primary power system fails, accomplish the following: A. Install a partial cover above the standby inverter "TF" located between frames 22 and 23 LH if the converter is installed as depicted on pages 5-6 of the service bulletin, in accordance with British Aerospace Service Bulletin 24-279-3255A, dated November 16, 1990. B. An alternative method 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, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch. C. 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. D. The installation requirements shall be done in accordance with British Aerospace Service Bulletin 24-279-3255A, dated November 16, 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-7053, AD 91-14-12) becomes effective on August 6, 1991.
93-12-02: 93-12-02 SAAB-SCANIA: Amendment 39-8607. Docket 93-NM-02-AD. Applicability: Model SAAB SF340A series airplanes, serial numbers 004 through 159, inclusive; Model SAAB 340B series airplanes, serial numbers 160 through 252, inclusive, and 254 through 258, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the fuselage, accomplish the following: (a) Within 3,000 hours time-in-service after the effective date of this AD, install an additional protective shield between the existing heat protection and the air cycle machine, in accordance with SAAB 340 Service Bulletin SAAB 340-53-028, dated August 20, 1992. (b) 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 theirrequests 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. (c) 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. (d) The installation shall be done in accordance with SAAB 340 Service Bulletin SAAB 340-53-028, dated August 20, 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 SAAB-SCANIA AB, SAAB Aircraft Product Support, S-581.88, Linkoping, Sweden. 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. (e) This amendment becomes effective on July 22, 1993.
2014-12-07: We are adopting a new airworthiness directive (AD) for Agusta S.p.A. (Agusta) Model AB412 and AB412EP helicopters with a certain rotor brake pinion installed. This AD requires inspecting the rotor brake pinion for a crack, and replacing it if there is a crack. This AD is prompted by a report of a rotor brake pinion failure. These actions are intended to detect a crack on the rotor brake pinion and prevent failure of the rotor brake pinion, which could lead to detachment of parts inside the transmission and subsequent loss of control of the helicopter.
99-25-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Lockheed Model 382 series airplanes, that requires a one time visual inspection of the under floor to ring fittings at fuselage station 817E to verify installation of the correct sized fasteners; and follow on corrective actions, if necessary. This amendment is prompted by notification from the manufacturer indicating that during production incorrect sized fasteners were installed on the under floor to ring fittings at fuselage station 817E. The actions specified by this AD are intended to prevent fatigue cracking of the fastener holes and adjacent fuselage structure due to installation of the incorrect sized fasteners, which could result in reduced structural integrity of the airplane.
80-19-12 R1: 80-19-12 R1 BEECH: Amendment 39-3916 as amended by Amendment 39-4404. Applies to Model 76 (Serial Numbers ME-1 through ME-415) airplanes certificated in any category, that have 100 hours or more time-in-service. COMPLIANCE: Required as indicated unless already accomplished. To ensure the integrity of the engine mount structure, accomplish the following: A) Within the next 50 hours time-in-service after the effective date of this AD and thereafter at intervals not to exceed 100 hours time-in-service, accomplish the following: 1. Remove the engine cowlings in accordance with instructions in the Aircraft Manufacturer's Maintenance Manual. 2. Perform a dye penetrant inspection of the two lower tubes of each engine mount assembly which extend from the firewall to the lower engine-mounting pads in accordance with FAA Advisory Circular AC 43.13-1A, paying particular attention to the area just aft of the gusset on the forward end of the tubes. NOTE: Thoroughly clean and degrease the area to be inspected; however, it is not necessary to remove the paint. B) Each 50 hours time-in-service after the dye penetrant inspections required by Paragraph A), visually inspect the same areas of the lower engine mount tubes. C) Prior to further flight, repair any cracks detected in the engine mount lower tubes as a result of any inspection required by this AD. Contact the Sales and Service Department, Beech Aircraft Corporation, Liberal, Kansas 67901, telephone (316) 624-1613, for repair instructions. D) The time-in-service for the inspections required herein may be adjusted up to 10 hours to facilitate accomplishing the inspections concurrent with other scheduled maintenance on the airplane. E) The repetitive inspections required by paragraphs A) and B) of this AD may be discontinued when engine mount assemblies are replaced with P/N 105-910019-1 engine mount assemblies in accordance with Beechcraft Service Instructions No. 1147, Revision 1. F) Any equivalent means 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) 942-4285. Amendment 39-3916 became effective September 25, 1980. This Amendment 39-4404 becomes effective on June 14, 1982.
78-26-08: 78-26-08 GENERAL ELECTRIC COMPANY: Amendment 39-3376. Applies to all model CT58-100-2, CT58-110-1, CT58-110-2, CT58-140-1 series A&L, CT58-140-1 and CT58-140-2 turboshaft engines. Compliance required as indicated, unless already accomplished. To prevent failure of the lube pump drive shaft and resultant loss of oil pressure, perform the following in accordance with General Electric Alert Service Bulletin CEB-253 ((CT58) A72-157) dated October 6, 1978, or later FAA approved revision. 1. Engines with new lube and scavenge pumps P/N 4000T98P02, Serial Number NMA-05635, 5724, 5729, 5736, and 5938 through 06151, or with any lube and scavenge pump P/Ns 37D400035P101, 4000T98P01, or 4000T98P02 overhauled after March 1976. (A) Remove from service pumps with 50 hours or less time in service since new or since overhaul within 4 hours operating time after the effective date of this AD. Replace with serviceable pump. (B) Remove from service pumps with over 50 but less than 500hours time in service since new or since overhaul within 50 hours operating time after the effective date of this AD. Replace with serviceable pump. 2. Inspect removed pumps, spare pumps overhauled after March 1976 and new spare pumps in the above serial number range, prior to installation on an engine, in accordance with Section 2B of the bulletin. Remove lip seals marked Viton and replace with new unmarked Sirvene seals and inspect adjacent parts for damage in accordance with the instructions in Section 2B(7) of the bulletin. NOTE: New pumps in the affected serial number range were shipped on new production CT58-140-1 engines, S/Ns 295218 through 295233, 295235 through 295241, 295243 through 295249, and 295255. The manufacturer's service bulletin identified and described in this directive is 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 this document from the manufacturer mayobtain copies upon request to Customer Service and Support Manager, General Electric Company, 1000 Western Avenue, Lynn, Massachusetts 01910. This document may also be examined at Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, and FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. This amendment becomes effective December 19, 1978.
83-13-06: 83-13-06 McCAULEY ACCESSORY DIVISION: Amendment 39-4667. Applies to McCauley Model 3AF32C504, 3AF32C505, 3AF32C506, 3AF32C507, 3AF32C508, and 3AF32C509 full feathering propellers with specific serial numbers listed in McCauley Service Bulletin No. 147 installed on, but not limited to, Piper PA-34-220T, Cessna T303, T310P, T310Q, T310R, 320D, 320E, 320F, 335, 340Z 340A, 401, 401A, 401B, 402, 402A, 402B, 402C, 414, and 414A type aircraft certificated in all categories. Compliance required within the next 30 days after the effective date of this AD, unless already accomplished A. To prevent possible failures of the counterweight bolts, accomplish the following: 1. Remove propeller spinner (shell). 2. Remove propeller counterweight bolt, P/N A-1635-125, from each blade, and install new P/N A-1635-125 bolt(s), identified with the letter "M" stamped on the head, torqued to 65-60 lb.-ft. in accordance with paragraphs 3 and 4 of McCauley Service Bulletin 147 dated March 4, 1983, or FAA approved equivalent. 3. Reinstall propeller spinner (shell). B. A special flight permit may be used in accordance with Federal Aviation Regulations 21.197 and 21.199 to operate the aircraft to a base where the AD can be accomplished. Upon request of the operator, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Chicago Aircraft Certification Office, FAA, 2300 East Devon Avenue, Des Plaines, Illinois 60018. Portions of the McCauley Service Bulletin No. 147 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 this document from the manufacturer may obtain a copy upon request to McCauley Accessory Division, Cessna Aircraft Company, 3535 McCauley Drive, P.O. Box 430, Vandalia, Ohio 45377. This document also may be examined at Rules Docket, Office of Regional Counsel,FAA, Attn: Rules Docket No. 83-ANE-17, 12 New England Executive Park, Burlington, Massachusetts 01803, and may be examined weekdays, except Federal holidays, between 8:00 am and 4:30 pm. This amendment becomes effective July 5, 1983.
2014-12-04: We are superseding Airworthiness Directive (AD) 2003-01-04 for BHTI Model 204B, 205A, 205A-1, 205B, and 212 helicopters. AD 2003-01-04 required inspecting the main rotor grip (grip) and reporting certain inspection results to the FAA. AD 2003-01-04 also required performing additional inspections, repair, or replacement depending on whether a crack or delamination was found, and determining and recording the hours time-in-service (TIS) and the engine start/stop cycles for each grip on a component history card or equivalent record. This new AD requires the same actions as AD 2003-01-04 but adds a retirement life to certain grips and expands the applicability to include the Model 210 helicopter and additional part-numbered grips. This AD was prompted by the discovery of additional cracked grips. We are issuing this AD to prevent failure of a grip, separation of a main rotor blade, and subsequent loss of control of the helicopter.
2014-12-08: We are superseding Airworthiness Directive (AD) 2004-11-10 for Przedsiebiorstwo Doswiadczalno-Produkcyjne Szybownictwa ``PZL-Bielsko'' Model SZD-50-3 ``Puchacz'' sailplanes. 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 fatigue damage of the welded joint between the airbrake torque tube and the airbrake control system lever located inside the fuselage. We are issuing this AD to require actions to address the unsafe condition on these products.
99-25-05: This amendment adopts a new airworthiness directive (AD) that is applicable to Hartzell Propeller, Inc., Model HD-E6C-3( ) series propellers, installed on Fairchild Dornier 328-110 series and 328-120 series airplanes. This action supersedes telegraphic AD T99-06-51 that currently requires initial and repetitive inspections of the propeller hub for cracks or grease leaks, and replacement of the hub if any cracks are found. This amendment requires an initial and repetitive inspections of Hartzell propeller hub, part number (P/N) D-5108-1, for cracks or grease leaks, replacement of the hub if any cracks are found, and allows the installation of propeller hub, P/N D-5108-5, as a terminating action for the inspection requirements. This amendment is prompted by the addition of propeller hub P/N D-5108-5 as a terminating action for the inspection requirements and by the removal of the inspection requirements for Hartzell propeller hub, P/N D-5108-5. The actions specified by this AD are intended to prevent severe vibration due to cracks in the propeller hub that could result in propeller blade loss, loss of control, and possible damage to the airplane.
93-11-04: 93-11-04 FOKKER: Amendment 39-8595. Docket 92-NM-195-AD. Applicability: Model F27 series airplanes (except Model F27 Mark 050 series airplanes); serial numbers 10102 and 10105 through 10240, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent premature failure of the outer flap center hinge due to fatigue, which could result in reduced structural integrity of the outer flap, accomplish the following: (a) For airplanes on which Fokker Service Bulletin SBF27/57-22 has not been accomplished and that have accumulated 72,000 or more total landings as of the effective date of this AD: Within 1,000 landings or 6 months after the effective date of this AD, whichever occurs first, perform a high frequency eddy current inspection of the outer flap center hinges to detect cracks, in accordance with Part 1 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991.(1) If no crack is found, repair the outer flap center hinge within 2,800 landings or 2 calendar years following the inspection required by paragraph (a) of this AD, whichever occurs first, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991. (2) If any crack is found, repair the outer flap center hinge, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991, at the times specified in paragraph (a)(2)(i) or (a)(2)(ii) of this AD. (i) If the crack length inside the bore is less than 5 mm, and if no crack is found on any of the faces around the bore: Repair within 300 landings following the inspection required by paragraph (a) of this AD. (ii) If the crack length inside the bore is equal to or more than 5 mm; or if the crack length inside the bore is less than 5 mm, and the crack is also present in one of the faces around the bore: Repair prior to further flight. (b) For airplanes on which Fokker Service Bulletin SBF27/57-22 has been accomplished and that have accumulated 55,000 or more total landings as of the effective date of this AD: Within 1,000 landings or 6 months after the effective date of this AD, whichever occurs first, perform a high frequency eddy current inspection of the outer flap center hinges to detect cracks, in accordance with Part 1 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991. (1) If no crack is found, repair the outer flap center hinge within 2,800 landings or 2 calendar years following the inspection required by paragraph (b) of this AD, whichever occurs first, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991. (2) If any crack is found, repair the outer flap center hinge in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-66, dated October 11, 1991, at the times specified in paragraph (b)(2)(i) or (b)(2)(ii) of this AD. (i) If the crack length inside the bore is less than 5 mm, and if no crack is found on any of the faces around the bore: Repair within 300 landings following the inspection required by paragraph (b) of this AD. (ii) If a crack inside the bore is located on the grease nipple bore half of the hinge; or if the crack length inside the bore is equal to or more than 5 mm; or if the crack length inside the bore is less than 5 mm and the crack is also present in one of the faces around the bore: Repair prior to further flight. (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, 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. (d) 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. (e) The inspections and repairs shall be done in accordance with Fokker Service Bulletin F27/57-66, dated October 11, 1991. 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 Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. 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. (f) This amendment becomes effective on July 22, 1993.
99-22-01: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 99-22-01, which was sent previously to all known U.S. owners and operators of Eurocopter Deutschland GmbH (ECD) Model EC135 P1 and T1 helicopters by individual letters. This AD requires, before further flight and at specified time intervals until a modified tail boom connecting frame flange (frame flange) is installed, inspecting and replacing, if necessary, the frame flange. This AD also requires, within 7 days, installing an additional bearing support on the frame flange. Thereafter, this AD requires visually inspecting the frame flange for cracks or misalignment of the slippage marks at specified time intervals. This amendment is prompted by the discovery of a crack in the frame flange at the attachment points of the tail rotor drive shaft bearing support. The actions specified by this AD are intended to prevent a fracture of the bearing frame flange, failure of the tailrotor drive shaft, and subsequent loss of control of the helicopter.
2014-12-03: We are adopting a new airworthiness directive (AD) for all Rolls-Royce Deutschland Ltd & Co KG (RRD) BR700-725A1-12 turbofan engines. This AD requires removal of affected fuel metering units (FMUs) on RRD BR700-725A1-12 engines. This AD was prompted by reports of wear on the receptors of the double-ended unions in the FMU housing on BR700-725A1-12 engines causing fuel leakage. We are issuing this AD to prevent failure of the FMU, which could lead to damage to one or more engines and damage to the airplane.
57-06-04: 57-06-04 WRIGHT: Applies to All Aircraft Incorporating C18CA, C18CB, TC18DA, and TC18EA Series Engines. Compliance required as indicated. Results of recent investigations indicate that the engine front section bearing durability can be improved by accomplishing the following: 1. At next engine overhaul the engine front section must be assembled with the propeller shaft thrust bearing (ball bearing) behind the radial bearing (roller bearing) as viewed from the propeller end of the engine. (Wright Aeronautical Division Service Bulletins Nos. C18C-252, TC18D-255, and TC-18E-66 cover this same subject.) Propellers must be balanced in accordance with instructions contained in applicable propeller manufacturer's recommendations. 2. Engines overhauled after April 15, 1957, must incorporate engine propeller shaft thrust bearings (ball bearings) and radial bearings (roller bearings) that have been inspected for proper internal bearing clearances in accordance with instructions issued by the Wright Aeronautical Division in their Service Letter dated March 22, 1957.
99-24-18: This amendment supersedes an existing airworthiness directive (AD), applicable to Eurocopter France Model AS-350B, B1, B2, B3, BA, and D, and AS-355E, F, F1, F2, and N helicopters, that requires inspecting certain versions of the tail rotor spider plate bearing (bearing) for the proper rotational torque, axial play, and any brinelling of the bearing. This amendment has the same inspection requirements as the current AD. Also, this AD expands the applicability to include additional part numbers (P/N's) and reduces the initial and recurring inspection compliance times. This amendment is prompted by additional reports of deterioration of the bearing. The actions specified by this AD are intended to prevent seizure of the bearing, loss of tail rotor control, and subsequent loss of control of the helicopter.
99-24-16: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747 series airplanes, that requires removal of cable guards in the lateral control system and replacement with new, improved cable guards. This amendment is prompted by reports of high control wheel forces and restricted control wheel movement. The actions specified by this AD are intended to prevent deterioration of cable guards in the lateral control system, which could result in a jam of the lateral control system and consequent reduced lateral controllability of the airplane.
83-13-05: 83-13-05 AIRBUS INDUSTRIE: Amendment 39-4675. Applies to the model A300 series airplanes, certificated in all categories. To prevent failure of certain components of the main and nose landing gears, within 250 landings after the effective date of this AD or prior to the accumulation of the number of landings specified in each paragraph below, whichever occurs later, accomplish the following, unless previously accomplished: A. Reinforce the lower torque link pins of the main landing gear in accordance with the instructions of Messier-Hispano-Bugatti (MHB) Service Bulletin 470-32-065, dated September 3, 1976, prior to the accumulation of 12,000 landings on B1 model aircraft, and 16,000 landings on B2-100, B2-200, B2-300, B4-100, B4-200 and C4-200 model aircraft, having the serial numbers specified in Airbus Industrie (AI) Service Bulletin A300-32-077, Revision 1, dated February 19, 1980. B. Modify the main landing gear drag strut of aircraft with serial numbers specifiedby AI Service Bulletin A300-32-114, dated January 13, 1978, in accordance with the instructions of MHB Service Bulletin 470-32-110, dated November 28, 1977, prior to the accumulation of 20,000 landings. C. Modify the actuating cylinder of the main landing gear for aircraft with serial numbers specified by AI Service Bulletin A310-32-116, Revision 6, dated September 17, 1982, in accordance with the instructions of MHB Service Bulletin 470-32-108, Revision 1, dated November 24, 1978, according to the following schedule: 1. For aircraft with modifications AI 1799 and 2025 originally installed, prior to the accumulation of 32,000 landings on B2 models or 22,000 landings on B4 models. 2. For aircraft with modifications AI 1799 and 2025 not originally installed, prior to the accumulation of 12,000 landings on both models B2 and B4. D. Incorporate modifications MHB 151 and 161 on the shock strut connecting rod of the main landing gear of aircraft with serial numbers specified in AI Service Bulletins A300-32- 096, Revision 4, and A300-32-076, Revision 3, both dated September 17, 1982, in accordance with the instructions of MHB Service Bulletins 470-32-039 and 470-32-040, Revision 1, both dated October 31, 1980, prior to the accumulation of 12,000 landings. E. Incorporate modification MHB 59 on the brace assembly actuating cylinder of the main landing gear of aircraft with serial numbers specified in AI Service Bulletin A300-32-036, Revision 2, dated September 17, 1982, in accordance with the instructions of MHB Service Bulletin 470-32-031, dated January 8, 1976, prior to the accumulation of 17,500 landings. F. Replace the actuating cylinder piston of the nose landing gear of aircraft with serial numbers specified in AI Service Bulletin A300-32-904, Revision 5, dated September 17, 1982, in accordance with the instructions of MHB Service Bulletin 470-32-033, dated January 9, 1976, prior to the accumulation of 14,000 landings. G. For the purpose of this AD, and when approved by an FAA maintenance inspector, the number of landings may be computed by dividing each airplane's time in service by the operator's fleet average time from takeoff to landing for the aircraft type. H. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. I. 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. This amendment becomes effective July 7, 1983.
99-25-02: This amendment adopts a new airworthiness directive (AD) that is applicable to all Boeing Model 737-100, -200, -300, -400, and -500 series airplanes. This action requires a one-time inspection to verify correct installation of the fastener that connects the input rod of the spoiler mixer mechanism to the torque tube crank, and corrective actions, if necessary. For certain airplanes, this action requires replacement of the nut, bolt, and cotter pin that connects the input rod of the spoiler mixer mechanism to the torque tube crank with a new or serviceable nut, bolt, and cotter pin. This amendment is prompted by reports indicating numerous discrepancies in the installation of the fastener that connects the input rod of the spoiler mixer mechanism to the torque tube crank. The actions specified in this AD are intended to prevent the linkage between the ratio changer input rod and the aft aileron control quadrant from becoming disconnected, which could result in reduced controllability of the airplane.
82-18-01: 82-18-01 EMBRAER: Amendment 39-4440. Applies to EMB-110P1 and EMB-110P2 models (S/Ns 110001 through 110329 and 110331 through 110339), airplanes certificated in any category. Compliance: Required on or before December 31, 1982, unless previously accomplished. To prevent leakage of water into the fuel tank, accomplish the following: (a) Defuel the airplane in accordance with the EMB-110 maintenance manual. (b) Remove the fuel filler neck components from each tank and install EMBRAER Kit S.B. 110-28-020 in accordance with the instructions contained in EMBRAER Service Bulletin 110-28-020, dated July 2, 1981. (c) An equivalent method of compliance may be used, if approved by the Chief, Atlanta Aircraft Certification Office, ACE-115A, Federal Aviation Administration, P.O. Box 20636, Atlanta, Georgia 30320. This amendment becomes effective September 27, 1982.
2022-18-05: The FAA is adopting a new airworthiness directive (AD) for all Airbus SAS Model A318, A319, A320, and A321 series airplanes. This AD was prompted by unclear and incomplete placard instructions for the doghouse door lock. This AD requires installing improved handling instruction placards on affected doghouses and re-identifying the doghouses, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. This AD also prohibits the installation of affected doghouses under certain conditions. The FAA is issuing this AD to address the unsafe condition on these products.
2014-12-01: We are superseding Airworthiness Directive (AD) 2013-11-05 for Bell Model 214B, 214B-1, and 214ST helicopters with certain tail rotor [[Page 32860]] hanger bearings (bearing) installed. AD 2013-11-05 required inspecting the bearing to determine whether an incorrectly manufactured seal material is installed on the bearing. This new AD retains the repetitive inspection of the bearings and also requires replacing the defective bearings. This AD was prompted by a report that certain bearings were manufactured with an incorrect seal material that does not meet Bell specifications. We are issuing this AD to prevent failure of the bearing and subsequent loss of control of the helicopter.
56-27-02: 56-27-02 HILLER: Applies to All UH-12, UH-12A and UH-12B Helicopters Including Spares. Compliance required as soon as possible but not later than February 28, 1957. Investigation has revealed that defective welds may exist at the clamp lugs on the four upper lord mount supports on P/N 63100-2 lower frame assembly (engine mount), or on mounts, P/N 63100-2M, modified in accordance with Hiller Service Bulletin No. 51. Failure of this weld has resulted in tilting of the rotor mast and loss of collective pitch control. The following one-time inspection is required on the above mounts to detect possible defective welds which must be reworked as indicated. 1. If the engine mount is cadmium plated, no inspection of the weld will be required, since these lower frame assemblies have been fabricated subsequent to the period of questionable weldments. 2. If the engine mount is not cadmium plated, remove the paint from all four lord mount supports in the area of the clamplugs and inspect for identification markings in or around the weld at the clamp lugs. If the weld is stamped with either a 7 or 8 or no stamp at all, it will be necessary to remove the mount from service until such time as the lugs can be removed and rewelded to CAM 18 standards. (Hiller Service Information Letter No. 111 covers this subject.)
79-10-06 R1: 79-10-06 R1 ENSTROM HELICOPTER CORPORATION: Amendment 39-3465 as amended by Amendment 39-4342. Applies to all Model F-28C and 280C helicopters. Compliance required as indicated. To prevent tail rotor failures as a result of tail rotor blade grip cracks, accomplish the following: A) Prior to next flight after receipt of this AD and prior to each flight thereafter, visually check the tail rotor blade grips in the vicinity of the blade retention bolt holes for any evidence of cracks with at least a 10X glass. Pilot may make this check. If any cracks are found, the blade and grip unit must be replaced with a serviceable unit before further flight. B) Prior to the next 50 hours' time in service after the effective date of this AD, unless already accomplished, remove the tail rotor blades from the blade grips and examine the grips in the vicinity of the blade retention bolt holes using standard dye penetrant inspection methods. Caution - care must be taken not tointermix blades and grips as they are match drilled sets. If any cracks are found, before further flight, remove the blade-and-grip unit and replace with a serviceable unit having either P/N 28-150013-1 or 28-150044-1 grips. Install replacement grips in accordance with paragraph C) of this AD. C) Install serviceable replacement P/N 28-150013-1 or P/N 28-150044-1 grips in accordance with applicable Enstrom Service Directive 0048, dated April 5, 1979, or 0048, Revision A, dated September 8, 1980, as outlined below: (1) Install P/N 28-150013-1 grips, in accordance with Enstrom Service Directive 0048, dated April 5, 1979, as follows: (a) By hand with the use of a 100 degrees - 1/2 inch back countersink (#AT4021-4) and a 3/16 inch pilot (#AT404-4), or equivalent tools, chamfer the edges (8 per grip) of the retention bolt holes in the blade grip .015 x 40 degrees. Repeat the same operation on each tail rotor blade retention bolt hole (4 places). After chamfering, thoroughly inspect the grips and blades for any nicks, burrs, or sharp edges. If any are found, they should be blended out by crocus cloth. (b) Replace the close tolerance bolts using a lubriplate compound and retorque to 50-75 in. lbs. (2) Install, P/N 28-150044-1 grips in accordance with Enstrom Service Directive 0048, Revision A, dated September 8, 1980, as follows: (a) Tail rotor assemblies incorporating Spindle P/N 28-150014-13 only are eligible for this alternate means of compliance. The part number is etched on the side of each spindle. Spindle P/N 28-150014-13 may be further identified by their shoulder-to-shoulder dimension and the rotor assembly's overall Tip-to-Tip length which are 3.46 + .01 and 56 7/16 inches, respectively. (b) Installation of Tail Rotor Blades on Tail Rotor Blade Grips P/N 28-150044-1 to comprise Blade and Grip Assemblies, P/N 28-150001-5 must be accomplished by Enstrom Customer Service. (c) Operators must send the old Tail Rotor Blade and Grip Assemblies P/N 28-150001-3 to Enstrom Customer Service Center for rework. D) Replace the close tolerance bolts using a lubriplate compound and retorque to 50 - 75 in. lbs. E) Preflight inspections required by paragraph A) of this AD may be discontinued after the installation of P/N 28-150044-1 grips. Enstrom Service Directive Bulletin No. 0048 also applies to the subject matter of this AD. Amendment 39-3465 became effective upon publication in the Federal Register, as to all persons except those to whom it was made immediately effective by the airmail letter dated April 9, 1979, which contained this amendment. This Amendment 39-4342 becomes effective March 19, 1982.