Results
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.
78-07-01: 78-07-01 CESSNA: Amendment 39-3163. Applies to TP206 (Serial Numbers P206-0191 thru P20600647), TU206 (Serial Numbers U206-0487 thru U20603693), T207 (Serial Numbers 20700001 thru 20700378) and T210 (Serial Numbers T210-0001 thru T210-0454 and Serial Numbers 21059200 thru 21061758) series airplanes certificated in all categories. To preclude engine oil pump failure due to contamination by the turbocharger thrust bearing anti-rotation pins and failure of the turbocharger shaft, within the next 25 hours' time-in-service after the effective date of this AD, accomplish the following: A. Check the turbocharger nameplate or aircraft permanent maintenance records to determine if the turbocharger serial number is prefixed by any of the following letter combinations: EF EFR FA FAR FH FHR EG EGR FB FBR FI FIR EH EHR FC FCR FJ FJR EI EIR FD FDR FK FKR EJ EJR FE FER FL FLR EK EKR FF FFR GA GAR EL ELR FG FGR GB GBR B. If the serial number on the turbocharger nameplate is not prefixed by any of the letter combinations set forth in Paragraph A, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. C. If the serial number on the turbocharger nameplate is prefixed by any of the letter combinations set forth in Paragraph A, check the aircraft permanent maintenance records to determine whether, when complying with AD 77-06-02, the turbocharger center housing was replaced by a mechanic or repair agency in accordance with Cessna Service Kit SK 210-75, dated February 24, 1977 (Reference Cessna Service Letter SE 77-3, Supplement #2 dated February 24, 1977) or by the turbocharger manufacturer (AiResearch). D. If the turbocharger center housing was replaced by AiResearch, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. E. If the turbocharger center housing was replaced by a mechanic or repair agency using instructions in Cessna Service Kit SK 210-75, visually inspect the turbocharger for signs of damage and proper compressor wheel attachment in accordance with Cessna Service Kit SK 210-78 dated November 15, 1977, or later revision (Ref. Cessna Service Letter SE 77-42 dated December 2, 1977, or later revisions) for damage which may have resulted from incomplete compressor wheel locknut torquing procedures prescribed in Cessna Service Kit SK 210-75. (1) If visual signs of damage are evident, return the turbocharger to AiResearch in accordance with Cessna Service Kit SK 210-78. (2) If no visual signs of damage are present but the compressor wheel attachment does not meet the criteria set forth in Cessna Service Kit SK 210-78, conduct additional inspections prescribed therein. Units found acceptable as a result of this inspection may be returned to service after reassembly per this kit. Return unacceptable units to AiResearch in accordance with instructions in Cessna Service Kit SK 210-78. (3) If no visual signs of damage are found and the compressor wheel attachment meets the criteria set forth in Cessna Service Kit SK 210-78, reassemble and identify the turbocharger in accordance with Cessna Service Kit SK 210-78. F. If the turbocharger center housing has not been replaced in accordance with AD 77-06-02, replace the turbocharger center housing in accordance with Cessna Service Kit SK 210-75B dated October 27, 1977, or later revision incorporating a compressor wheel seating procedure. (Ref. Cessna Service Letter SE 77-42.) G. Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. H. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment supersedes Amendment 39-2853 (42 FR 15894), AD 77-06-02. This amendment becomes effective April 6, 1978.
2001-26-53: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 2001-26-53, which was sent previously to all known U.S. owners and operators of Eurocopter France (ECF) Model AS350B, B1, B2, B3, BA, D, and AS355E helicopters by individual letters. This AD requires, before further flight, removing certain serial-numbered servocontrols. This AD is prompted by a report of manufacturing defects in a batch of main servocontrol rods. The actions specified by this AD are intended to prevent failure of a main servocontrol in the flight control system and subsequent loss of control of the helicopter.
2017-18-18: We are adopting a new airworthiness directive (AD) for all Airbus Model A350-941 airplanes. This AD requires repetitive on-ground power cycles to reset the internal timer. This AD was prompted by the in-service loss of communication between some avionics systems and the avionics network. We are issuing this AD to address the unsafe condition on these products.
89-25-08: 89-25-08 BEECH: Amendment 39-6410. Applicability: Models 65 (Serial Numbers (S/N) L-1, L-2, L-6, LF-7 through LF-76, and LC-1 through LC-180); 65-80 and 65-A80 (S/N LD-1 through LD-244); 65-A80 (S/N LD-245 through LD-269) when Beech Modification Kit No. 80-4004-1 or -3 is installed; and 65-B80 (all S/N) airplanes certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To detect possible fatigue cracking of the wing main spar lower cap and associated structure, accomplish the following: (a) Within the next 200 hours time-in-service (TIS) after the effective date of this AD, or upon accumulating 3000 hours TIS on Models 65-80 and 65-A80 airplanes, or upon accumulating 5000 hours TIS on Models 65 and 65-B80 airplanes, whichever occurs later, unless previously accomplished per AD 70-25-01, Amendment 39-1609, and thereafter at intervals not to exceed 1000 hours TIS (except as provided in paragraph (b) below) after the initial inspection, inspect the wing lower forward spar attach fittings, center section and outboard wing spar caps adjacent to the attach fittings by visual, fluorescent penetrant and eddy current methods as specified in the applicable section of Beech Structural Inspection and Repair Manual (SIRM), P/N 98-39006, Revision A4, dated May 1, 1987. NOTE 1: Beech offers a two-day training course free of charge to qualified personnel who have prior knowledge of eddy current inspection techniques. A listing of Beech Corporate maintenance facilities may be obtained from the sources contained in paragraph (e) of this AD. A listing of other facilities employing qualified inspectors is not available. (b) At each inspection required by paragraph (a) above, inspect any reinforcing strap installed per Supplemental Type Certificate (STC) SA1583CE for proper tension and condition in accordance with Aviadesign Engineering Order E.O. B-8001, Issue 3, dated May 30,1985. Correct any discrepancy prior to further flight. For airplanes equipped with STC SA1583CE and inspected in accordance with paragraph (a) above, the repetitive inspection interval of 1000 hours TIS in paragraph (a) above may be extended to 3000 hours TIS. (c) If any crack is found in a main spar lower cap or fitting, prior to further flight repair or replace the defective part using the instructions and limitations specified in the SIRM or other FAA approved instructions provided by Beech Aircraft Corporation. (d) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (e) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209; Telephone (316) 946-4400. NOTE 2: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or Western Aircraft Maintenance, 4444 Aeronca Street, Boise, Idaho 83705; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD supersedes AD 70-25-01, Amendment 39-1609. This amendment (39-6410, AD 89-25-08) becomes effective on January 4, 1990.