Results
2002-01-18: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A319, A320, and A321 series airplanes, that requires replacement of the trigger spring of the slide bar on each of the passenger doors with a new, stronger trigger spring. This action is necessary to prevent corrosion of the trigger spring on the slide bar of the passenger doors, which could result in incorrect locking of the slide bar and, during deployment of the escape slide, lead to a delay in evacuating passengers in an emergency. This action is intended to address the identified unsafe condition.
76-10-10: 76-10-10 BEECH: Amendment 39-2617. Applies to Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29), 65-90, 65-A90, B90 and C90 (Serial Numbers LJ-1 thru LJ-676), E90 (Serial Numbers LW-1 thru LW-163), 100 and A100 (Serial Numbers B-1 thru B- 225) and 200 (Serial Numbers BB-2 thru BB-111) airplanes. Compliance: Required as indicated, unless already accomplished. To preclude opening of the cabin door in flight, accomplish the following: I. Within 50 hours' time in service after the effective date of this AD, and within each 50 hours' time in service thereafter, check the cabin door for proper operation and rigging in accordance with the applicable Beech Shop/Maintenance Manual. When all the items specified in Paragraph II have been complied with the requirements of this paragraph (I) are no longer applicable. II. Within 50 hours' time in service after September 15, 1976, perform the following: A. On Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29) and 65-90, 65-A90, B90 (Serial Numbers LJ-1 thru LJ-351) airplanes, modify the cabin door installation in accordance with Beechcraft Service Instructions 0043-104 or 0016-105, Rev. I, or later approved revisions, as applicable. B. On Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29), 65-90, 65-A90, B90 and C90 (Serial Numbers LJ-1 thru LJ-676), E90 (Serial Numbers LW-1 thru LW-163), 100 and A100 (Serial Numbers B-1 thru B-225), and 200 (Serial Numbers BB-2 thru BB-111) airplanes, perform the following in accordance with Beechcraft Service Instruction 0818-016 or later approved revisions: 1. Install Beech P/N 101-430124-1 and if fixed step door Beech P/N 101-430124-3 or -5 decals on the existing cabin door instruction plate and operate the cabin door accordingly. 2. Check cabin door latching mechanism and warning system for proper operation and rigging and rerig, if required, as instructed in the appropriate Beech Shop/Maintenance Manual. C. On the airplane models and serial numbers listed below add the indicated part number FAA-approved Airplane Flight Manual Supplement/Revision to the existing airplane pilot's operating manual or FAA-approved airplane flight manual: MODELS Beech Part Number (P/N) of FAA-Approved Airplane Flight Manual Supplement Revision dated November 14, 1975 or Subsequent 1) 65-88, 65-90, 65-A90, B90 and C90 (S/N LJ-502 thru LJ-624) and 100 (S/N B-2 thru B-89 and B-93) 1) P/N 131344 2) C90 (S/N LJ-625 thru LJ-676) 2) P/N 90-590010-53A6 3) E90 (S/N LW-1 thru LW-163) 3) P/N 90-590012-3A6 4) A100 (S/N B-1, B90 thru B-92, B-94 thru B-225) 4) P/N 100-590032-1A6 5) 200 (S/N BB-2 thru BB-111) 5) P/N 101-590010-3A4 III. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective May 28, 1976.
81-02-10: 81-02-10 BELL: Amendment 39-4025. Applies to Models 214B and 214B-1 helicopters, serial numbers up to and including S/N 28049. Compliance is required as indicated unless already accomplished. To prevent a clutch failure which will result in the loss of engine power to the main rotor, accomplish the following: (a) The freewheeling clutch assembly, P/N 214-040-021-001, must be removed from service and P/N 214-040-021-103 clutch assembly installed according to the following schedule; (1) P/N 214-040-021-001 clutch assemblies with 290 or more hours' time in service on the effective date of this AD must be removed from service within the next 10 hours' time in service. (2) P/N 214-040-021-001 clutch assemblies with less than 290 hours' time in service on the effective date of this AD must be removed from service prior to attaining 300 hours' time in service. (3) P/N 214-040-021-001 clutch assemblies with unknown time in service must be removed within the next ten hours' time in service. NOTE: BHT Alert Service Bulletin No. 214-80-13, dated August 22, 1980, pertains to this subject. (b) Special flight permits may be issued in accordance with FAR 21.197 and FAR 21.199 to fly aircraft to a base where this AD can be accomplished. (c) Any alternate equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration. This AD supersedes AD 80-07-11 (Amdt. 39-3726, 45 FR 20778). The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Bell Helicopter Textron, Product Support Department, Post Office Box 482, Fort Worth, Texas 73101. 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 the FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD, which includes the incorporated material in full, is maintained by the FAA at their headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. This amendment becomes effective January 26, 1981.
89-14-03: 89-14-03 LOCKHEED AERONAUTICAL SYSTEMS COMPANY: Amendment 39-6243. Applicability: Lockheed Model L-188A and L-188C series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent airplane control difficulties due to flap asymmetry, accomplish the following: A. Within 90 days after the effective date of this AD, inspect the flap universal joints to determine if they are splined without a stop. NOTE: Flap universal joints that are splined without a stop are easily detected by the lack of a screw and washer that holds the pins in place. If they do not have a stop, accomplish the following: 1. Remove from service any flap universal joints that are splined without a stop and install universal joints that are splined with a stop; or 2. Cut a circumferential groove in the splines of the torque tube shafts on each side of the universal joints (1 inch from the end of the shaft) and installa snap ring, in a manner approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region. The snap ring is to eliminate the possibility of spline disengagement. The snap ring should completely encircle the shaft; or 3. Modify each universal joint by drilling a hole and installing a 1/8-inch diameter steel pin through each side (two per joint). Locate the pin from .7 to .8-inch from each end, and locate so as to intersect the spline centerline. Peen both pin ends. This modification must be accomplished in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region. NOTE: Local spotfacing may be necessary on pins located at maximum dimension. B. Within 90 days after the effective date of this AD and thereafter at intervals not to exceed one year, verify proper torque tube-to-universal joint engagement of .375-inch, as specified in Detail E of the L-188 Maintenance Manual, Section 27-6-8, Figure 201; and verify that the distance between the edges of the torque tube shoulders at BL55 is 5.5 to 5.87 inches (universal joints with internal stop), or 6.0 to 6.6 inches for universal joints reworked with added stop pins. C. Within 90 days after the effective date of this AD and thereafter at intervals not to exceed one year, inspect the wing flap asymmetry shutoff valves, P/N 668225-1, to ascertain whether the valves hang up or respond slowly (greater than 1 second). 1. If the valves hang up or respond slowly, prior to further flight, install a functioning serviceable valve of the same part number (P/N 668225-1) and repeat the inspections at intervals not to exceed one year. 2. Installation of modified valve, P/N 668225-101, or a new valve P/N 668225-101, in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region, constitutes terminating action for the repetitive inspections of the wingflap asymmetry shutoff valves required by paragraph C. and C.1., above. D. Within 90 days after the effective date of this AD, rewire the flap asymmetry annunciator light in the cockpit to trigger the annunciator, in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region, so that the light in the cockpit will illuminate when the asymmetry detector is tripped and not be dependent on the shutoff valve. E. 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, Los Angeles Aircraft Certification Office, 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, Los Angeles Aircraft Certification Office. F. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a base to accomplish the actions required by 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 Lockheed Aeronautical Systems Company, P.O. Box 551, Burbank, California 91520, Attn: L-188 Commercial Support Contracts, Dept. 63-11, Unit 33. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or 3229 East Spring Street, Long Beach, California. This amendment (39-6243, AD 89-14-03) becomes effective on July 24, 1989.
2017-19-22: We are superseding Airworthiness Directive (AD) 2014-07-09 for British Aerospace Regional Aircraft Jetstream Series 3101 and Jetstream Model 3201 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) originated 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 both the need for newly added inspections for corrosion, which includes the door hinges/supporting structure and attachment bolts for the main spar joint and engine support, and inadequate existing instructions for inspection for corrosion for several areas including the rudder hinge location on the vertical stabilizer. We are issuing this AD to require actions to address the unsafe condition on these products.
2002-01-19: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F.28 Mark 0070 and 0100 series airplanes, that requires repetitive operational tests for discrepancies of the heating system of pitot tube #1, and replacement of the pitot tube, if necessary. This AD also requires eventual modification of the alternating current sensing circuit for pitot tube #1, which terminates the repetitive operational test requirement. This action is necessary to prevent failure of the heating system of pitot tube #1 due to a short circuit, which may go undetected and lead to the pilot receiving erroneous airspeed indications, resulting in reduced control of the airplane. This action is intended to address the identified unsafe condition.
2017-19-06: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-1A11 (CL-600), CL-600-2A12 (CL-601 Variant), and CL-600-2B16 (CL-601-3A, CL-601-3R, and CL-604 Variants) airplanes. This AD was prompted by a new life limitation that has been introduced for the side brace fitting shaft and side brace-to-airplane fitting pin of the main landing gear (MLG). This AD requires revising the maintenance or inspection program. This AD also requires an inspection to identify the serial number, to serialize, and to record the accumulated life of the side brace fitting shaft of the MLG. We are issuing this AD to address the unsafe condition on these products.
2011-01-04: 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: It has been detected a short circuit in harness W101 due to its interference with the main door mechanism. Further analysis of the affected region has also revealed the possibility of chafing between the same harness and the oxygen tubing. The chafing of the wiring harness against the oxygen tubing could lead to a short circuit of the wiring harness and a subsequent fire in the airplane. Since this condition may occur in other airplanes of the same type and affects flight safety, a corrective action is required. Thus, sufficient reason exists to request compliance with this AD in the indicated time limit. We are issuing this AD to require actions tocorrect the unsafe condition on these products.
49-45-01: 49-45-01 LUSCOMBE: Applies to All Model 11A Aircraft. Compliance required as soon as possible but not later than next 25 hours operation time and at each 25-hour period thereafter until reinforcement of main landing gear aft canted fuselage bulkhead is accomplished. Inspect for buckling, cracks or other evidence of failure of permanent set of the main landing gear aft fuselage canted bulkhead in the web and/or flange in the area adjacent to the steel landing gear trunnion and fuel line. Inspect fuselage wing lift strut attach fitting for cracks in the radii of the flanges attaching it to each aft fuselage canted bulkhead. Usually evidence of failure of the aft canted bulkhead can be determined by a crack in the fuselage canted bulkhead web extending from the fuel line hole to the flange attaching the bulkhead to the belly skin and/or buckle in the cabin floor located approximately 1 inch directly aft of the bulkhead under the carpet flooring and/or loose rivets attaching theflange of the canted bulkhead to the belly skin. If the difficulties are not revealed as indicated, a 2-inch hole cut in the cabin floor located approximately 3 inches aft and inboard of that part of the canted bulkhead supporting the door will allow access for detailed examination of the aft side of the rear fuselage canted bulkhead. Removal of seat and floor carpet is necessary to accomplish this inspection. If loose rivets in the bulkhead flange at the attachment to the belly skin, cracks or permanent set in excess of 1/8 inch are found in the web of the bulkhead adjacent to the steel trunnion, the bulkhead must either be satisfactorily repaired or replaced. If noticeable permanent set in the web is apparent (under 1/8 inch), the web of the bulkhead may be reworked by straightening. If cracks are found in the fuselage wing lift strut attach fitting it should be replaced or the cracks should be stop drilled and the full length of each cracked flange reinforced with a 3/4 inchby 3/4 inch by 0.064 inch 24ST angle. In addition, the following modifications must be made: A collar must be incorporated on the front end of the hinge pin that passes through the front and rear main landing gear steel trunnions which are riveted to the two fuselage canted bulkheads. This tubular collar should be fabricated of 4230 steel and be at least 5/8-inch long and of sufficient thickness to effect a snug bearing fit against the forward end of the steel tube composing the socket of the forward steel trunnion. A 1/4-inch bolt should be used to attach the collar to the hinge pin using the existing 1/4-inch hole in the extreme forward end of the hinge pin. A curved doubler of 0.064 inch 24ST should be placed over the existing 0.040-inch floor skin connecting the flanges of the two main landing gear canted bulkheads. This doubler should pick up the existing floor skin and bulkhead top flange rivet pattern in the vicinity of the landing gear steel trunnion, extending in length at least 3 inches to either side of a vertical plane through the centerline of the landing gear hinge pin and picking up at least six of the existing rivets in each of the canted bulkheads. Blind type rivets may be used to attach this doubler. The rivet pattern attaching the flange of the aft canted fuselage bulkhead to the belly skin between the openings in the fuselage skin which allow entrance of the main landing gear legs should be inspected for rivet size and pattern. The first 20 rivets inboard from these openings must be 5/32-inch A17ST spaced approximately 1/2-inch apart. If the 2-inch inspection holes have been cut in the floor, they must be reinforced by at least a 4-inch diameter 0.040-inch 24ST doubler on the underneath side of the floor skin and a quick removable inspection cover placed on top side to be used for subsequent 25-hour inspections, if applicable. Any equivalent structural modification to preclude a failure, or permanent set in the aft canted bulkhead at the attachment of the main landing gear trunnion will be considered satisfactory.
2017-19-10: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 757-200, -200PF, and -200CB series airplanes. This AD was prompted by an analysis of the cam support assemblies of the main cargo door (MCD) that indicated that the existing maintenance program for the cam support assemblies is not adequate to reliably detect cracks before two adjacent cam support assemblies could fail. This AD requires an inspection to determine part numbers, repetitive inspections to detect cracking of affected cam support assemblies of the MCD, and replacement if necessary. We are issuing this AD to address the unsafe condition on these products.
2002-01-25: This amendment adopts a new airworthiness directive (AD), applicable to certain Bombardier Model DHC-8-100, -200, and -300 series airplanes, that requires repetitive inspections of the rudder pedal adjustment fittings for cracks and replacement of cracked fittings with new fittings. This amendment also provides an optional terminating action. This action is necessary to detect and correct cracking of the rudder pedal adjustment fittings, which could lead to deformation of the fittings, resulting in jammed rudder pedals and loss of rudder control, with consequent reduced controllability of the airplane. This action is intended to address the identified unsafe condition.
92-20-01: 92-20-01 DE HAVILLAND, INC.: Amendment 39-8375. Docket No. 92-NM-55-AD. Applicability: Model DHC-7 airplanes; on which stainless steel cables, Post- Modification Number 7/2609, have not been installed; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent a gear-up landing, accomplish the following: (a) Within 30 days after the effective date of this AD, perform a detailed inspection of the left- and right-hand main landing gear (MLG) emergency down release cables to detect corrosion, in accordance with de Havilland Alert Service Bulletin A7-32-94, dated September 3, 1991; or Revision A, dated November 15, 1991. (b) If any corrosion is detected as a result of the inspection required by paragraph (a) of this AD, prior to further flight, replace both cable assemblies with either stainless steel cables (Post-Modification Number 7/2609) or carbon steel cables (Pre-Modification 7/2609), in accordance with de Havilland Alert Service Bulletin A7-32-94, dated September 3, 1991; or Revision A, dated November 15, 1991. (c) Within 6 months after the effective date of this AD, or within 12 months after installing carbon steel cables, if installed in accordance with paragraph (b) of this AD: Install Post-Modification 7/2609 cables, in accordance with de Havilland Alert Service Bulletin A7-32-94, dated September 3, 1991; or Revision A, dated November 15, 1991. (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, New York Aircraft Certification Office (ACO), FAA, Engine and Propeller Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, New York ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York ACO. (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 inspection, replacement, and installation shall be done in accordance with de Havilland Alert Service Bulletin A7-32-94, dated September 3, 1991; or Revision A, dated November 15, 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 de Havilland, Inc., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (g) This amendment becomes effective on November 17, 1992.
93-22-05: 93-22-05 TELEDYNE CONTINENTAL MOTORS: Amendment 39-8744. Docket 93-ANE-54. Applicability: Teledyne Continental Motors (TCM) Model O-200A reciprocating engines with Engine Serial Numbers 256030 through 256037; and TCM C85, C90, O-200, and O-240 series reciprocating engines with carburetor air intake housing assemblies, Part Numbers (P/N) CE11141, CE11142, 639814, 639815, 641534, and Repair Kit Assemblies, P/N 641689, purchased after August 31, 1991, without a permanent ink stamp "CSB 93-13" located on the inside of the housing assembly. These engines are installed on but not limited to the following aircraft: American Champion Models 7BCM, 7CCM, S7CCM, 7DC, S7DC, 7EC, S7EC, 7FC, 7JC, 7ECA, 11BC, S11BC, 11CC, S11CC, and 402; Anderson Greenwood Model 14; Cessna Model 120, 140, 140A, 150, 150A-M, and A150K-M; Luscombe Model 8E, 8F, and T-8F; McClish (Funk) Model B85C; Piper Model PA-18 and PA-19; Reims Model F150G, H, J, K, L, M, FA150K, L, FRA150L, and M; Spinks Model Lark 95; Superior (Culver) Model V and V-2; Taylorcraft Model 19 and F-19; and Univair (Erco, Forney, Alon, Mooney) Model 415E, 415G, F-1, F-1A, A-2, and M-10. Compliance: Required as indicated, unless accomplished previously. To prevent engine failure due to a cracked air valve in the carburetor air intake housing assembly, accomplish the following: (a) Within the next 5 hours time in service (TIS) after the effective date of this AD, inspect the carburetor air intake housing assembly in accordance with paragraph 2 of the Inspection Procedure section of TCM Critical Service Bulletin (CSB) No. 93-13, dated August 12, 1993. (1) If the carburetor air intake housing assembly meets the requirements of Paragraph 2A of the Inspection Procedures of TCM CSB No. 93-13, dated August 12, 1993, no further action is required. (2) If the carburetor air intake housing assembly meets the requirements of paragraph 2B of TCM CSB No. 93-13, dated August 12, 1993, inspect the carburetor air intakehousing assembly for cracks. If cracks are found anywhere in the assembly, prior to further flight replace with a serviceable assembly. (b) Thereafter, for assemblies that meet the requirements of paragraph 2B of TCM CSB No. 93-13, dated August 12, 1993, inspect the carburetor air intake housing assembly for cracks in accordance with Paragraphs 3 and 4 of the Inspection Procedure of TCM CSB No. 93-13, dated August 12, 1993, at intervals not to exceed 25 hours TIS since the last inspection. If cracks are found anywhere in the assembly, prior to further flight replace with a serviceable assembly. (c) Inspect uninstalled carburetor air intake housing assemblies in accordance with paragraph (a) of this AD prior to installation. (d) For the purpose of this AD, a serviceable carburetor air intake housing assembly is defined as: (1) An assembly purchased on or before August 31, 1991; or (2) An assembly that meets the inspection criteria of paragraph (a)(1) of this AD; or(3) An assembly with the following P/N's: (i) 653661, which supersedes CE11142; (ii) 653670, which supersedes 639815; (iii) 653675, which supersedes 641534; (iv) 653657, which supersedes 641689; or (4) An assembly, P/N 641534, with a permanent ink stamp "CSB 93-13" located on the inside of the housing assembly. NOTE: The assemblies, P/N's CE11141 and 639814, have not been superseded, as these are assemblies with the air filter included, corresponding to airboxes, P/N's CE11142 and 639815. (e) Replacement with a serviceable carburetor air intake housing assembly constitutes terminating action to the inspection requirements of this AD. (f) 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, Atlanta Aircraft Certification Office. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send itto the Manager, Atlanta Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Atlanta Aircraft Certification Office. (g) 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. (h) The inspections, and replacement, if necessary, shall be done in accordance with the following service document: Document No. Pages Revision Date TCM CSB No 93-13 1-3 Original August 12, 1993 Total pages: 3 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 Teledyne Continental Motors, P.O. Box 90, Mobile, AL 36601; telephone (205) 438-3411. Copies may be inspected at the FAA, New England Region, Office of theAssistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (i) This amendment becomes effective December 14, 1993, to all persons except those persons to whom it was made immediately effective by priority letter AD 93-22-05, issued November 4, 1993, which contained the requirements of this amendment.
2017-19-09: We are superseding Airworthiness Directive (AD) 2014-25-01, which applied to certain Bombardier, Inc., Model DHC-8-400 series airplanes. AD 2014-25-01 required modifying the nose landing gear (NLG) trailing arm and installing a new pivot pin retention mechanism. This AD instead requires modifying the NLG shock strut assembly. This AD was prompted by reports of discrepancies of a certain bolt at the pivot pin link, resulting in corrosion of the bolt. We are issuing this AD to address the unsafe condition on these products.
2017-16-01: We are adopting a new airworthiness directive (AD) for certain Ameri-King Corporation emergency locator transmitters (ELTs) as installed on various aircraft. This AD was prompted by multiple reports of ELT failure and a report of noncompliance to quality standards and manufacturer processes related to Ameri-King Corporation ELTs. This AD requires repetitive inspections of the ELT for discrepancies; repetitive checks, tests, and verifications, as applicable, to ensure the ELT is functioning; and corrective actions if necessary. This AD also allows for optional replacement of affected ELTs and, for certain aircraft, optional removal of affected ELTs. We are issuing this AD to address the unsafe condition on these products.
2017-19-11: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model BD-700-1A10 and BD-700-1A11 airplanes. This AD was prompted by a determination that a certain task in the aircraft maintenance manual (AMM) will not accomplish the intent of a candidate certification maintenance requirement (CCMR) for detecting dormant failures of the pitch feel (PF) and rudder travel limiter actuator (RTLA) back-up modules. This AD requires doing an operational test of the flight control unit (FCU) back-up modules, and repair if necessary. We are issuing this AD to address the unsafe condition on these products.
2002-01-16: This amendment supersedes Airworthiness Directive (AD) 86-24-11 and AD 86-25-04, which require you to incorporate, into the Limitations Section of the pilot's operating handbook and airplane flight manual (POH/AFM) of Fairchild Aircraft, Inc. (Fairchild Aircraft) SA226 and SA227 series airplanes, procedures for preventing an engine flameout while in icing conditions. This AD retains the POH/AFM requirements from the above-referenced AD's and requires a modification to the torque sensing system to allow the igniters to automatically turn on when an engine senses low torque. This AD is the result of two instances of a dual engine flameout on the affected airplanes. When the torque sensing system modification is incorporated, the POH/AFM requirements are no longer necessary. The actions specified by this AD are intended to prevent a dual engine flameout on the affected airplanes by providing a system that automatically turns on the engine igniters when low torque is sensed. A dual engine flameout could result in failure of both engines with consequent loss of control of the airplane.
84-09-51 R1: 84-09-51 R1 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-4903. Applies to Lockheed Model L-1011-385 series airplanes, certificated in all categories. Compliance required as indicated, unless previously accomplished. To prevent depressurization of the aircraft due to the failure of the negative pressure relief valve, accomplish the following: A. Within 100 flight hours after the receipt of this AD, unless previously accomplished, inspect the negative pressure relief valve mounting adapters on each aircraft identified as follows: 1. Lockheed Serial Numbers 1175 through 1250. 2. Lockheed Serial Numbers 1001 through 1174. Aircraft which have had either of the two negative pressure relief valve adapters replaced by the operator. B. Inspection and repair procedures: 1. After instituting all the preliminary safety precautions, gain access to aft side of the pressure bulkhead through fire bottle inspection panel 317 BB (see Lockheed Maintenance Manual MM6-40-00), and determine type of adapter installed. If adapter does not have circumferential spin marks and the serial number is 229 or lower, no further inspection is required. If serial number is 230 or higher, or spin marks are found, then additional inspection must be performed. NOTE: Adapter serial numbers are identified by the last three digits in the serial number block on the decal located at the 12 o'clock position aft of the pressure bulkhead. Spin marks are best determined by shining indirect light, such as flash light, on the adapter side wall. 2. On aircraft requiring further inspection: a. Gain access to the forward side of the negative pressure relief valve (see Lockheed Maintenance Manual MM25-42-00). b. On the forward side of the pressure bulkhead, LBL/RBL20, WL300, visually inspect the adapter part of each negative pressure relief valve in the area of the flange radius. Clean the area with solvent prior to inspection. 3. If a crack is found, replace or repair the adapter before the next flight. Replacement adapter must be inspected prior to installation and is subject to the requirements of this AD. 4. Installation of repair or reinforcement clips for adapter: Install 300 series stainless steel clips, either 0.050 1/4 hard or 0.040 1/2 hard material. Dimensions are 0.95-inch wide with 0.89 and 1.50-inch flanges, and 85 degrees bend, with 0.12-inch bend radius. Install clips using existing fastener holes through the aft pressure bulkhead and two additional fasteners through the side of the adapter, install one at minimum of 0.34 inches and the other at minimum of 0.95 inches from end of long leg of clip. Fasteners are to be MS20470AD5 rivets, or NAS1398M5 rivets, or structural equivalent. a. For repair of a cracked adapter, stop-drill crack 1/4-inch diameter, and install a minimum quantity of 40 clips per adapter, using 2 of each 3 attachments through the pressure bulkhead. b. For reinforcement ofuncracked adapters, install a minimum quantity of 20 clips per adapter, using 1 of each 3 attachments through pressure bulkhead. 5. If no cracks are found, repeat the inspection per B.1. through B.4., above, at intervals not to exceed 25 landings. 6. Replace or reinforce the adapter within 350 flight hours after the last inspection or after the receipt of this AD, whichever is later. 7. With adapter reinforced with the 20 clips per paragraph B.4.b., above, the reinspection intervals may be extended from 25 landings to 1000 landings. 8. For adapters repaired with the 40 clips per paragraph B.4.a., above, reinspect from the aft side for cracks at the aft fastener, common to the clip and adapter, every 500 landings. 9. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. This amendment becomes effective September 11, 1984, and was effective earlier to those recipients of telegraphic AD T84-09-51, dated April 19, 1984.
58-01-05: 58-01-05 LOCKHEED: Applies to All Models 49-46, 149, 649, 649A, 749, 749A, and 1049-54 Aircraft. Compliance required as indicated. As a result of cracks discovered in Lockheed 749A wing skin and stringers, the following inspections shall be accomplished on the various model aircraft as indicated, and if any cracks are discovered, they must be repaired prior to further operation. Any FAA/LAC approved repair may be used. Inspect and reinspect for cracks in the lower wing skin and stringers, left and right, from wing Station 125 through Station 215 between the front and rear beams. Inspections to be conducted at the following specified times and intervals using X-ray and visual, or visual means. The X-ray inspection method is recommended if equipment is available, since cracks under the stringers would be detected. I. For Models 649, 649A, 749, 749A and 1049-54: A. The first inspection should be performed before 20,500 hours have been accumulated on the aircraft. For aircraft on which inspections of STA 191 through 206 have already been made in accordance with AD 56-03-01, initial inspections of additional indicated areas need not be earlier than and may be correlated with reinspections required by B(1), B(2), B(3), and B(4). B. Reinspections must be accomplished in accordance with one of the following programs: (1) X-ray at 2,500 hours (maximum) intervals without opening the fuel tanks following the recommendations and technique outlined on Lockheed Sketch No. 101057 or a FAA/LAC approved equivalent. In addition to the X-ray inspection at this time, the bottom side of the wing skin must be visually inspected from front to rear beam beneath the nacelle to wing fillets on the inboard and outboard sides of No. 2 and No. 3 nacelles. This necessitates opening the kidney plate inspection holes in these fillets and/or removal of the tail cone assembly. See Lockheed Sketch No. 101057 for location of cracks which have previously been discovered; or (2) X-ray at 3,200 hours (maximum) intervals by opening the fuel tanks and following the technique outlined on Lockheed Sketches No. 101057 and No. 101058, or FAA/LAC approved equivalent. In addition to the X-ray inspections at 3,200 hours, aircraft with over 20,000 hours must be visually inspected at 200-hour (maximum) intervals as follows: Inspect the bottom side of lower wing skin, for leaks resulting from cracks, from front beam to rear beam between W.S. 125 and W.S. 145 and between W.S. 191 and W.S. 215. This necessitates opening the kidney plate inspection holes in the nacelle to wing fillets on the inboard and outboard sides and/or removal of the tail cone assembly of nacelles No. 2 and No. 3. This area should be given special attention. If leaks are discovered and cracks suspected, tanks must be opened and stripped of sealant to visually inspect upper side of skin. Inspect the upper side of lower wing skin for cracks in the dry area from front beam to rear beam between W.S. 145 and W.S. 191. See Lockheed Sketches No. 101057 and No. 101058 for location of cracks which have been previously discovered; or (3) When X-ray equipment is not available, a visual inspection must be made at 800-hour (maximum) intervals after opening the fuel tanks and removing the sealant from the designated areas. (It should be noted that cracks under stringers cannot be detected by the visual inspection method); or (4) X-ray at 2,800 hours (maximum) intervals by opening the fuel tanks and following the technique outlined on Lockheed Sketches No. 101057 and No. 101058, or FAA approved equivalent. In addition to the X-ray inspections at 2,800 hours, visually inspect at 350-hour (maximum) intervals as follows: Inspect the bottom side of lower wing skin, for leaks resulting from cracks, from front beam to rear beam between W.S. 125 and W.S. 145 and between W.S. 191 and W.S. 215. This necessitates opening the kidney plate inspection holes in the nacelle to wing fillets on the inboard and outboard sides and/or removal of the tail cone assembly of nacelles No. 2 and No. 3. This area should be given special attention. If leaks are discovered and cracks suspected, tanks must be opened and stripped of sealant to visually inspect upper side of skin. Inspect the upper side of lower wing skin for cracks in the dry area from front beam to rear beam between W.S. 145 and W.S. 191. See Lockheed Sketches No. 101057 and No. 101058 for location of cracks which have been discovered previously. C. The reinspections required as per paragraphs B(1), B(2), B(3), or B(4) may be discontinued when permanent reinforcement per Lockheed Drawing No. 550236 has been accomplished, except that: In the area from W.S. 125 to W.S. 191 where the size and kind of material remains unchanged (i.e., the old material is merely replaced with new) the reinspection program noted above must be reinstated not later than 20,000 hours after rework. D. Lockheed Drawing Nos. 11755, 490668, 492806, and 493312, describe approved permanent repairs for individually affected areas in which cracks have been previously discovered. Reinspections in the area between W.S. 191 and W.S. 215 may be discontinued if permanent repair is made per Lockheed Drawing No. 11755. The reinspection program must be reinstated not later than 20,000 hours after rework is accomplished in the individually affected areas per drawing numbers 490668, 492806, or 493312. E. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. II. For Models 49-46 and 149: A. The first inspection should be performed before 25,500 hours have been accumulated on the aircraft. For aircraft on which inspections of STA 191 through 206 have already been made in accordance with AD 56-03-01, initial inspections of additional indicated areas need not be earlier than and may be correlated with reinspection required by B(1), B(2), B(3), and B(4). B. Same as paragraph I.B. 1. Same as paragraph I.B.(1). 2. Same as paragraph I.B.(2) (except substituted 25,000 hours for 20,000 hours). 3. Same as paragraph I.B.(3). 4. Same as paragraph I.B.(4). C. The reinspections required as per paragraphs B(1), B(2), B(3), or B(4) may be discontinued when permanent reinforcement per Lockheed Drawing No. 550236 has been accomplished, except that: In the area from W.S. 125 to W.S. 191 where the size and kind of material remains unchanged (i.e., the old material is merely replaced with new) the reinspection program noted above must be reinstated not later than 25,000 hours after rework. D. Lockheed Dwg. Nos. 11755, 490191, 490668, 492806, and 493312 describe approved permanent repairs for individually affected areas in which cracks have been previously discovered. Reinspections in the area between W.S. 191 and W.S. 215 may be discontinued if permanent repair is made per Lockheed Drawing No. 11755 or 490191. The reinspection program must be reinstated not later than 25,000 hours after rework is accomplished in the individually affected areas per drawing Nos. 490668, 492806, or 493312. E. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. This supersedes AD 56-03-01. Revised September 26, 1963.
2017-18-17: We are superseding Airworthiness Directive (AD) 2004-23-20, which applied to certain Airbus Model A300, A300 B4-600, and A300 B4- 600R series airplanes; and Model A300 F4-605R and A300 C4-605R Variant F airplanes. AD 2004-23-20 required, for certain airplanes, repetitive inspections for cracking around certain attachment holes, installation of new fasteners for certain airplanes, and follow-on corrective actions if necessary. AD 2004-23-20 also required modifying certain fuselage frames, which terminated certain repetitive inspections. This new AD reduces certain compliance times, expands the applicability, and requires an additional repair on certain modified airplanes. This AD was prompted by a report indicating that the material used to manufacture the upper frame feet was changed and negatively affected the fatigue life of the frame feet. We are issuing this AD to address the unsafe condition on these products.
2017-19-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 727 airplanes. This AD was prompted by analysis of the cam support assemblies of the main cargo door (MCD) that indicated the repetitive high frequency eddy current (HFEC) inspections required by the existing maintenance program are not adequate to detect cracks before two adjacent cam support assemblies of the MCD could fail. This AD requires repetitive ultrasonic inspections for cracking of the cam support assemblies of the MCD and replacement if necessary. We are issuing this AD to address the unsafe condition on these products.
90-15-04: 90-15-04 BRITISH AEROSPACE: Amendment 39-6652. Docket No. 90-NM-41-AD. Applicability: Model BAC 1-11 200 and 400 series airplanes, pre-modification PM5384, certificated in any category. Compliance: Required within 2,400 hours time-in-service or two years after the effective date of this AD, whichever occurs first, unless previously accomplished within the past 2,400 hours time-in-service or within the past two years; and thereafter at intervals not to exceed 4,800 hours time-in-service or four years, whichever occurs first. To prevent tailplane trim gearbox oil from being contaminated with water, accomplish the following: A. Remove the tailplane trim gearbox from the airplane, drain the oil, flush and refill with clean oil, and replace the filler plug and wire lock, in accordance with paragraph 2.2 of British Aerospace Alert Service Bulletin 27-A-PM5384, Issue 1, dated July 24, 1989. Reinstall the gearbox in the airplane and test in accordance with Maintenance Manual Chapter 27-40. 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, 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. 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. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. 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 (39-6652, AD 90-15-04) becomes effective on August 14, 1990.
2001-17-26 R1: This document corrects and clarifies information in an existing airworthiness directive (AD) that applies to certain Raytheon Model DH.125, HS.125, BH.125, and BAe.125 (U-125 and C-29A) series airplanes; Model Hawker 800, Hawker 800 (U-125A), Hawker 800XP, and Hawker 1000 airplanes. That AD currently requires an inspection for cracking or corrosion of the cylinder head lugs of the main landing gear actuator and follow-on/corrective actions. This document corrects and clarifies the affected airplane serial numbers. This correction is necessary to ensure that operators do not misinterpret which airplanes are subject to the requirements of this AD. The incorporation by reference of certain publications listed in the regulations was approved previously by the Director of the Federal Register as of October 3, 2001 (66 FR 45575, August 29, 2001).
85-25-03: 85-25-03 SIKORSKY AIRCRAFT: Amendment 39-5172. Applies to Model S-64E helicopters, certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent operation with a cracked main rotor head torque tube inner bracket, accomplish the following: (a) Prior to the first flight of each day, after the effective date of this AD, visually inspect with a 10-power or higher magnifying glass the main rotor head torque tube inner bracket assembly, Part Number S1510-21332-0, for cracks and/or corrosion in accordance with Section 2, Paragraph A, of Sikorsky Alert Service Bulletin (ASB) No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (b) If the torque tube inner bracket assembly is cracked, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (c) If the torque tube inner bracket assembly is corroded, determine the extent and limits of the corrosion prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB-No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. If the extent or limits of the corrosion are exceeded, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. Otherwise, rework the torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph B, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (d) Aircraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the AD can be accomplished. (e) Upon request, an alternative means of compliance with the requirements of this AD which provide an equivalent level of safety maybe used when approved by the Manager, Boston Aircraft Certification Office, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone (617) 273-7112. Sikorsky ASB No. 64B10-4A, dated July 17, 1985, 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 may obtain copies upon request to Sikorsky Aircraft Division, United Technologies Corporation, North Main Street, Stratford, Connecticut 06601. These documents may also be examined at the Office of the Regional Counsel, FAA, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment becomes effective December 30, 1985.
91-05-21: 91-05-21 GENERAL ELECTRIC COMPANY: Amendment 39-6900. Docket No. 90-ANE- 28. Applicability: General Electric Company (GE) CF6-80C2A5 and CF6-80C2B6 engines, Serial Numbers (S/N) 690-101 through 690-369, and S/N 695-101 through 695-423; and CF6- 80C2B6F and CF6-80C2D1F engines, S/N 702-101 through 702-470, and S/N 703-101 through 703-136, which do not incorporate the increased shroud cooling design features of paragraph (b) of this AD, installed on, but not limited to, Airbus A300, Boeing 767, and McDonnell Douglas MD-11 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent high pressure turbine (HPT) failure and possible aircraft damage, accomplish the following: (a) Borescope inspect engines in accordance with sections 2.B., 2.C., and 2.D of the Accomplishment Instructions in GE CF6-80C2 Service Bulletin (SB) 72-473, Revision 1, dated September 21, 1990, unless previously accomplished, according to the following schedule based upon cycles since new (CSN) on the effective date of this AD: (1) Inspect within 10 cycles in service (CIS) after the effective date of this AD or prior to accumulating 520 CSN, whichever occurs later, for CF6-80C2A5 and CF6- 80C2B6 engines, S/N 690-101 through 690-369, and S/N 695-101 through 695-350; and CF6- 80B6F engines, S/N 702-101 through 702-315, and S/N 702-317 through 702-321. (2) Inspect within 10 CIS after the effective date of this AD or prior to accumulating 1,250 CSN, whichever occurs later, for CF6-80C2A5 and CF6-80C2B6 engines, S/N 695-351 through 695-423; and CF6-80C2B6F and CF6-80C2D1F engines, S/N 702-316, 702-322 through 702-470, and S/N 703-101 through 703-136. (3) Remove from service or reinspect in accordance with the following: (i) Remove from service prior to further flight, engines with at least one Category 4 shroud. (ii) Remove from service within 25 hours time in service (TIS) since last inspection (SLI), engines with no Category 4 shrouds, but at least one Category 3 shroud. (iii) Borescope reinspect at intervals not to exceed 125 hours TIS SLI, engines with no Category 3 or 4 shrouds, but at least one Category 2 shroud. (iv) Borescope reinspect at intervals not to exceed 300 hours TIS SLI, engines with no Category 2,3, or 4 shrouds, but at least one Category 1 shroud. (v) Borescope reinspect at intervals not to exceed 520 CIS SLI, engines with no Category 1, 2, 3, or 4 shrouds. (b) Replace the HPT stator stage one shroud support assemblies, Part Numbers (P/N) 9381M61G06 and 9381M61G07; the HPT stator support hanger assemblies, P/N 9397M73G05 and 9397M73G06; and the HPT stage one shrouds, P/N 1333M75P05, 1333M75P06, 1333M75P07, 1333M75P08, 1333M75P09, and 1333M75P10 in accordance with the Accomplishment Instructions in GE CF6-80C2 SB 72-474, Revision 1, dated December 11, 1990, at the next HPT module exposure after the effective date of this AD, but prior to December 31, 1994. (c) For the purpose of this AD, HPT module exposure is defined as the separation of the HPT stator support case from the compressor rear frame. (d) For the purpose of this AD, the shroud Categories are defined in GE CF6-80C2 SB 72-473, Revision 1, dated September 21, 1990. (e) 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. (f) 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 times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. The borescope inspections and installation of improved HPT hardware shall be done in accordance with the following documents: Document Page Revision Date CF6-80C2SB 72-473 5, 6, 7, Original 7/3/90 10-23 1, 2, 3, Rev. 1 9/21/90 4, 8, and 9 CF6-80C2 SB 72-474 1-24 Rev. 1 12/11/90 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 Office of the Assistant Chief Counsel, FAA, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-6900, AD 91-05-21) becomes effective on March 27, 1991.