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87-18-03: 87-18-03 ROBINSON HELICOPTER COMPANY: Amendment 39-5875. Final copy of priority letter AD 87-18-03. Applies to Model R22 series helicopters, certificated in any category, Serial Numbers 0580 through 0644, and helicopters which have had the A258-1 main rotor pitch link assembly, A258-2 fitting, A258-3 link or A258-4 link assembly replaced between August 15, 1986, and May 7, 1987. Compliance required as indicated, unless already accomplished. (a) Prior to further flight after the effective date of this AD, visually inspect main rotor pitch link assembly A258-1 (two per aircraft) as follows: (1) Remove paint and visually inspect the A258-1 pitch link assembly employing a 3X or greater magnifying glass. (2) If a crack is found in the A258-2 upper fitting, replace it with a like serviceable part, prior to further flight. (3) If a crack is found in the A258-3 lower link, replace the A258-1 pitch link assembly with a like serviceable part, prior to further flight. NOTE: The A258-3 lower link and the B163-1 rod end make up the lower end A258-4 link assembly, as shown in Figure 2 of Robinson Helicopter Company Service Bulletin No. 57, dated May 11, 1987. (4) Replace the A258-2 upper fitting in the pitch link assembly using the following steps. NOTE: Reference the procedure given in Part B of Robinson Helicopter Company, Service Bulletin No. 57, dated May 11, 1987. (i) Color code swashplate arms, pitch horns, and both pitch control link rod ends. If matching color marks are lost or length of pitch link assembly changes, a track and balance of the main rotor blades per Section 10.200, Track and Balance, dated May 22, 1987, or Robinson Helicopter Company maintenance manual or an approved equivalent. (ii) Remove pitch links and measure overall length using a caliper or large micrometer. Record length to nearest 0.001 inch. (iii) Disassemble pitch link. Do not remove rod end attached to A258-3 link assembly. Replace A258-2 fitting with like serviceable part. (iv) Reassemble pitch link(s), keeping overall length to within plus minus 0.005 of length measured before disassembly. (v) Install on helicopter, and torque attaching bolts to 100 inch-lbs., plus nut drag per Section 8.412, Swashplate Installation, dated May 22, 1987, of Robinson Helicopter Company maintenance manual or an approved equivalent. (vi) Touch up paint using an epoxy primer and grey exterior paint. Mark all joints with a torque stripe. (b) If no cracks are found in a link assembly, upper fitting or lower link, after visual inspection in paragraph (a)(1), the helicopter may be returned to service for an interval of time not to exceed 10 hours time in service. (c) Prior to each flight during the time interval specified in paragraph (b), visually check for cracks in the link assembly. The checks required by this AD may be performed by the pilot and must be recorded in accordance with FAR Section 43.9. (d) Within the next 10 hours time in service after receipt of this AD, remove and replace A258-2 fittings with serviceable parts. Conduct a dye penetrant or equivalent inspection for crack indications in the A258-3 link. Install airworthy parts in accordance with paragraph (a) (4). (e) Replacement of all A258-2 fittings specified in this AD with serviceable fittings and inspection of the A258-3 links terminates action for these repetitive checks required by this AD. (f) An alternate method of compliance with this AD which provides an equivalent level of safety, may be approved by the Manager, Western Aircraft Certification Office, FAA, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009-2007. This amendment becomes effective on April 25, 1988, as to all persons except those persons to whom it was made immediately effective by priority letter AD 87-18-03, issued August 28, 1987, which contained this amendment.
90-25-07: 90-25-07 BOEING: Amendment 39-6788. Docket No. 89-NM-269-AD. \n\n\tApplicability: All Model 707/720 series airplanes, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tNOTE: This AD references Boeing Document Number D6-54928, "Aging Airplane Corrosion Prevention and Control Program, Model 707/720," Revision A, dated July 28, 1989, for inspection procedures, compliance times, and reporting requirements. In addition, this AD specifies inspection and reporting requirements beyond those included in the Document. Where there are differences between the AD and the Document, the AD prevails. \n\n\tTo control corrosion, accomplish the following: \n\n\tA.\tWithin one year after the effective date of this AD, revise the FAA-approved maintenance program to include the corrosion control program specified in Boeing Document Number D6-54928, "Aging Airplane Corrosion Prevention and Control Program, Model 707/720," Revision A, dated July28, 1989, (hereinafter referred to as "the Document"). \n\n\tNOTE: All structure found corroded or cracked as a result of an inspection conducted in accordance with this paragraph must be addressed in accordance with FAR Part 43. \n\n\tNOTE: Where non-destructive inspection (NDI) methods are employed, in accordance with Section 4.1 of the Document, the standards and procedures used must be acceptable to the Administrator in accordance with FAR 43.13. \n\n\tNOTE: Procedures identified in the Document as "optional" are not required to be accomplished by this AD. \n\n\tB.\t1.\tIf, as a result of any inspection conducted in accordance with the program required by paragraph A., above, Level 3 corrosion is determined to exist in any area, accomplish one of the following within 7 days after such determination: \n\n\t\ta.\tSubmit a report of any findings of Level 3 corrosion to the Manager of the Seattle Aircraft Certification Office (ACO) and inspect the affected area on all Model 707/720 aircraftin the operator's fleet; or \n\n\t\tb.\tSubmit for approval to the Manager of the Seattle ACO one of the following: \n\n\t\t\t(1)\tProposed adjustments to the schedule for performing the tasks in that area on remaining airplanes in the operator's fleet, which are adequate to ensure that any other Level 3 corrosion is detected in a timely manner, along with substantiating data for those adjustments; or\n\n\t\t\t(2)\tData substantiating that the Level 3 corrosion found is an isolated occurrence and that no such adjustments are necessary. \n\n\tNOTE: Notwithstanding the provision of Section 1.1. of the Document that would permit corrosion that otherwise meets the definition of Level 3 corrosion (i.e., which is determined to be a potentially urgent airworthiness concern requiring expeditious action) to be treated as Level 1 if the operator finds that it "can be attributed to an event not typical of the operator's usage of other airplanes in the same fleet," this paragraph requires that data substantiating any such finding be submitted to the FAA for approval. \n\n\tNOTE: As used throughout this AD, where documents are to be submitted to the Manager of the Seattle ACO, the document should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. The Seattle ACO will not respond to the operator without the PI's comments or concurrence. \n\n\t\t2.\tThe FAA may impose adjustments other than those proposed, upon a finding that such adjustments are necessary to ensure that any other Level 3 corrosion is detected in a timely manner. \n\n\t\t3.\tPrior to the compliance time specified for the first task required in the adjusted schedule approved under paragraph B.1. or B.2. of this AD, revise the FAA-approved maintenance program to include those adjustments. \n\n\tNOTE: The reporting requirements of this paragraph and of paragraph D., below, do not relieve operators fromreporting corrosion as required by FAR Section 121.703. \n\n\tC.\tTo accommodate unanticipated scheduling requirements, it is acceptable for a repeat inspection interval to be increased by up to 10% but not to exceed 6 months. The cognizant FAA Principal Inspector (PI) must be informed, in writing, of any extension. \n\n\tNOTE: Except as provided in this paragraph, notwithstanding Section 3.1., paragraph 4, of the Document, all extensions to any compliance time must be approved by the Manager of the Seattle ACO. \n\n\tD.\tReport forms for Level 2 corrosion and a follow-up report for Level 3 corrosion must be submitted at least quarterly in accordance with Section 5.0 of the Document. \n\n\tE.\tIf the repeat inspection or task intervals of an operator's existing corrosion inspection program are shorter than the corresponding intervals in Section 4.3 of the Document, they may not be increased without specific approval of the Manager of the Seattle ACO. \n\n\tF.\tBefore any airplane that is subject to this AD can be added to an air carrier's operations specifications, a program for the accomplishment of tasks required by this AD must be established in accordance with the following: \n\n\t\t1.\tFor airplanes that have previously been operated under an FAA- approved maintenance program, the initial task on each area to be accomplished by the new operator must be accomplished in accordance with the previous operator's schedule or with the new operator's schedule, whichever would result in the earlier accomplishment date for that task. After each task has been performed once, each subsequent task must be performed in accordance with the new operator's schedule. \n\n\t\t2.\tFor airplanes that have not previously been operated under an FAA- approved maintenance program, each initial task required by this AD must be accomplished either prior to the airplane's being added to the air carrier's operations specifications, or in accordance with a schedule approved by the Manager, Seattle ACO.G.\tIf corrosion is found to exceed Level 1 on any inspection after the initial inspection, the corrosion control program for the affected area must be reviewed and means implemented to reduce corrosion to Level 1 or better. \n\n\t\t1.\tWithin 60 days after such a finding, if corrective action is necessary to reduce future findings of corrosion to Level 1 or better, such proposed corrective action must be submitted for approval to the Manager, Seattle ACO. \n\n\t\t2.\tWithin 30 days after the corrective action is approved, revise the FAA- approved maintenance program to include the approved corrective action. \n\n\tH.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tI.\tSpecial 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. \n\n\tThe requirements shall be done in accordance with Boeing Document Number D6- 54928, "Aging Airplane Corrosion Prevention and Control Program, Model 707/720," Revision A, dated July 28, 1989. 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 Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. Copies may be inspected at the FAA, Transport Airplane Directorate, Northwest Mountain Region, 1601 Lind Avenue S.W., 5th Floor, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8301, Washington, D.C. \n\n\tThis amendment (39-6788, AD 90-25-07) becomes effective on December 31, 1990.
82-24-03: 82-24-03 BOEING: Amendment 39-4496. Applies to all Model 707/720 series airplanes listed in Boeing Service Bulletin A3364, Revision 3, NSC-2 and later FAA approved revisions with nacelle strut diagonal braces and associated fittings which have accumulated 7500 or more landings. \n\n\tTo detect cracks in the nacelle strut diagonal brace and associated fittings accomplish the following: \n\n\tA.\tWithin 500 landings after the effective date of this AD unless already accomplished, inspect the nacelle strut diagonal braces and associated fittings in accordance with Boeing Service Bulletin A3364, Rev. 3, NSC2 or later FAA-approved revisions, and repeat thereafter at the intervals specified in Tables 1, 2, and 3 below. \n\n\tB.\tIf cracks are found, replace the cracked part prior to further flight or repair in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. Parts oversized beyond the limits specified in Boeing Service Bulletin A3364 must be replaced prior to further flight. \n\n\tC.\tFor Group III, IV, and VI airplanes, as specified in the service bulletin, replace the diagonal brace assembly if the outboard diagonal brace end fitting (forward or aft) attach holes have been oversized beyond the limits specified in Boeing Service Bulletin A3364, Rev. 3 or later FAA-approved revisions within 1000 landings since oversizing. \n\n\tD.\tAlternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t\t\t\tTABLE I\n\n\nForward Mating Fitting \n\nType Inspection\nRepeat Intervals Not To Exceed Landings\nVisual\t\n250\nHi frequency\n7500\nHi frequency followed by shot peening\t\n12000\n\n\t\t\t\t\tTABLE II \n\n\nAft Mating Fitting \n\nType Inspection\nRepeat Intervals Not To Exceed Landings\nVisual\n250\n1.\tHi frequency or dye penetrant of web. \n7500\n2.\tHi frequency of lug hole with bushing removed and hole oversized. \n\n1.\tHi frequency of lug hole\twith bushing removed, oversizing and peening of lug hole surface. \n12000\n2.\tHi frequency or dye penetrant of web and preening of web. \n\n\n\t\t\t\t\tTABLE III \n\n\nDiagonal Brace Assembly \n\nType Inspection\nRepeat Intervals Not To Exceed Landings \nHi frequency Interior only \n2500 \nHi frequency Exterior (fitting removal) with hole oversizing. \n7500\nHi frequency Exterior with hole oversizing \nand peening \n12000 \nHi frequency Exterior with holes previously \noversized to limit and peening.\n 7500 \n\n\tNOTE: Hole oversizing cannot exceed limits specified in Service Bulletin. \n\n\tThe manufacturer's specification 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 may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at the FAA, Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis AD supersedes AD 80-14-14. \n\n\tThis amendment becomes effective November 23, 1982.
92-27-18: 92-27-18 MCDONNELL DOUGLAS: Amendment 39-8453. Docket 92-NM-105-AD. \n\n\tApplicability: Model DC-10-10, -15, -30, -40, and KC-10 (Military) series airplanes; certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent failure of the main landing gear (MLG) piston and subsequent reduced controllability of the airplane during landing, accomplish the following: \n\n\t(a)\tFor Model DC-10-10 and DC-10-15 series airplanes that have not been modified in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, dated January 21, 1991: Within 30 months after the effective date of this AD, accomplish the requirements of paragraphs (a)(1) and (a)(2) of this AD: \n\n\t\t(1)\tConduct a visual inspection of the right and left MLG pistons to detect inward buckling. If inward buckling of the MLG piston is detected, prior to further flight, replace the MLG piston in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t\t(2)\tDetermine if the rebound check valve previously has been replaced in accordance with McDonnell Douglas Service Bulletin 32-60, Revision 3, dated August 7, 1974; or Revision 4, dated November 6, 1974; or Revision 5, dated May 1, 1975; or Revision 6, dated June 20, 1975; (hereafter referred to as "SB 32-60"). \n\n\t\t\t(i)\tIf the rebound check valve has not been replaced in accordance with SB 32-60: Within 30 months after the effective date of this AD, replace the rebound check valve in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t\t\t(ii)\tIf the rebound check valve has been replaced in accordance with SB 32-60: Within 30 months after the effective date of this AD, rework or replace the rebound check valve in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t(b)\tFor Model DC-10-10 and DC-10-15 series airplanes that have been modified in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, dated January 21, 1991; and for all Model DC-10-30 and DC-10-40 and KC-10 (military) series airplanes: Within five years after the effective date of this AD, accomplish the requirements of paragraphs (b)(1) and (b)(2) of this AD: \n\n\t\t(1)\tConduct a visual inspection of the right and left MLG pistons to detect inward buckling, in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. If inward buckling of the MLG piston is detected, prior to further flight, replace the MLG piston in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t\t(2)\tDetermine if the rebound check valve previously has been replaced in accordance with SB 32-60. \n\n\t\t\t(i)\tIf the rebound check valve has not been replaced in accordance with SB 32-60: Within five years after the effective date of this AD, replace the rebound check valve in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t\t\t(ii)\tIf the rebound check valve has been replaced in accordance with SB 32-60: Within five years after the effective date of this AD, rework or replace the rebound check valve in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. \n\n\t(c)\tAn 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, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Los Angeles ACO. \n\n\tNOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Los Angeles ACO. \n\n\t(d)\tSpecial 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. \n\n\t(e)\tThe inspection, rework, and replacement shall be done in accordance with McDonnell Douglas DC-10 Service Bulletin 32-227, Revision 1, dated April 30, 1992. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from McDonnell Douglas Corporation, P.O. Box 1771, Long Beach, California 90846-1771, Attention: Business Unit Manager, Technical Publications - Technical Administrative Support, C1-L5B. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Transport Airplane Directorate, Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California 90806-2425; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. \n\n\t(f)\tThis amendment becomes effective on February 25, 1993.
69-04-02: 69-04-02 CESSNA: Amdt. 39-722. Applies to Model A188 airplanes with Continental Motors Corporation IO-520-D (300 HP) engines, Serial Numbers 188-0001 through 188-0478 with the exception of Serial Numbers 188-0441, 188-0443, 188-0444, 188-0447, 188-0448, 188- 0453, 188-0460, 188-0462 through 188-0465, 188-0468 through 188-0476. Compliance: Required as indicated. To prevent failure of the throttle control bellcrank and the throttle and mixture support shaft, unless already accomplished, perform the following: (1) Within 10 hours' time in service after the effective date of this AD, visually inspect the throttle control bellcrank P/N 0750163-1 for cracks or broken welds around the welded center boss. If cracks or a broken weld is found, replace the bellcrank with a strengthened Cessna P/N 0750163-4 in accordance with Cessna Service Letter SE68-28, Supplement No. 1, or an equivalent modification approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. If this strengthened part P/N 0750163-4 is not available, P/N 0750163-1 may be installed. The aircraft may then be operated not to exceed 100 hours at which time P/N 0750163-4 must be installed prior to further flight. (2) If any cracks are found during the inspection required by Paragraph (1), a description of such cracks and the time in service as of the date of said inspection should be reported to the local General Aviation District Office on FAA Form FAA-1226. (Reporting approved by the Bureau of the Budget under BOB No. 04-R0174.) (3) Within 10 hours' time in service after the effective date of this AD, replace the throttle and mixture control support shaft with P/N 0750173-2 and roll pin P/N MS-16562-16 in accordance with Cessna Service Letter SE68-28, dated October 8, 1968, or equivalent modification approved by the Chief, Engineering & Manufacturing Branch, FAA, Central Region. This amendment becomes effective March 3, 1969.
2010-11-01: We are adopting a new airworthiness directive (AD) for the products listed above. 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: During aircraft full scale fatigue test, it has been found the occurrence of cracks in the cockpit windshield post lower eyelet fitting at the attachment of the center post on the forward fuselage (SSI 53-10-19). Further analysis of this cracking resulted in modifications on the aircraft Airworthiness Limitation Items (ALI), to include new inspection tasks and its respective intervals. Undetected fatigue cracking in this area could adversely affect the structural integrity of these airplanes. We are issuing this AD to require actions to correct the unsafe condition on these products.
2010-10-22: The FAA is superseding an existing airworthiness directive (AD), which applies to Model BAe 146 airplanes and Model Avro 146-RJ airplanes. That AD currently requires revising the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness to incorporate life limits for certain items and inspections to detect fatigue cracking in certain structures. This new AD requires incorporating new and more restrictive life limits for certain items and for certain inspections to detect fatigue cracking in certain structures. This AD also requires revising the airworthiness limitations to include critical design configuration control limitations for the fuel system. This AD results from issuance of a later revision to the airworthiness limitations. We are issuing this AD to ensure that fatigue cracking of certain structural elements is detected and corrected, and to prevent ignition sources in the fuel tanks; fatigue cracking of certain structural elements couldadversely affect the structural integrity of these airplanes.
68-07-08: 68-07-08 MAULE: Amdt. 39-578. Applies to Models M-4, Serial Numbers 3 thru 94, M- 4C, Serial Numbers 1C thru 10C, M-4S, Serial Numbers 1S thru 3S, M-4T, Serial Numbers 1T thru 3T, M-4-210, Serial Numbers 1001 thru 1045, M-4-210C, Serial Numbers 1001C thru 1064C, M-4-220S, Serial Numbers 2001S thru 2003S, and M-4-220C, Serial Numbers 2001C thru 2006C, series aircraft. Compliance: Required as indicated, unless already accomplished. To prevent loss of rudder trim tab control and resultant interference within aileron operation, accomplish the following: Within the next 50 hours' time in service after the effective date of the airworthiness directive, modify the rudder trim tab hinges in accordance with either Maule Aircraft Corporation Service Letter Number 14, dated February 19, 1968, or any other method approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Central Region. This amendment becomes effective April 11, 1968.
2010-10-15: We are adopting a new airworthiness directive (AD) for the specified ECF model helicopters. This AD results from a mandatory continuing airworthiness information (MCAI) AD issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. The MCAI AD states that with certain pilot and copilot seats in the rear high position and seat backrest fully tilted the seat shoulder harness could become jammed between the seat and bulkhead. This condition, if not corrected, could result in the shoulder harness binding and causing the inertial reel to malfunction and no longer retain the flight crew member in the seat in the event of an emergency or hard landing.
60-06-01: 60-06-01 BOEING: Amdt. 111 Part 507 Federal Register March 8, 1960. Applies to Models 707 and 720 Series Airplanes Equipped with Nose Gear Outer Cylinder, Boeing P/N 5-83025- 3005. \n\n\tCompliance required as indicated. \n\n\tDue to failures found in the nose landing gear outer cylinder as a result of initial defects, the following inspections are required unless already accomplished:\n \n\t(a) Conduct a daily visual inspection with a 10-power magnifying glass for cracks in the outer surface of the nose gear outer cylinder in the upper barrel area of the trunnion, and in the area below the towing collar. Outer cylinders with cracks appearing on the outer surface must be replaced prior to further flight. Further inspection is not required on replacement part. This inspection must be continued until item (b) is accomplished. \n\n\t(b) Within 20 calendar days, inspect the areas described in (a) with probe type ultrasonic shear wave equipment or equivalent. Outer cylinders with cracks on the outer surface or with crack indications that exceed 1 inch in length must be replaced prior to further flight. Cylinders with crack indications 1 inch or less in length may be retained in service, provided: \n\n\t\t(1) A daily visual inspection as in (a) is continued and the ultrasonic inspections or equivalent are repeated every 20 landings. \n\n\t\t(2) A dye penetrant inspection or equivalent of the interior surface of the outer cylinder is accomplished within 550 hours' time in service and every 550 hours' time in service thereafter. \n\n\t(c) If no cracks or crack indications are found by inspections (a) and (b), the nose landing gear may revert to normal inspection procedures and periods except for accomplishing (d). \n\n\t(d) Conduct a dye penetrant inspection or equivalent for cracks on the interior surface of all nose gear outer cylinders within the next 3,000 hours' time in service. Outer cylinders with cracks on the interior surface that exceed 1 inch in length or that extend to the outer surface must be replaced prior to further flight. Cylinders with cracks 1 inch or less in length on the interior surface may be retained in service provided the cracks do not extend to the outer surface and the cylinders are inspected in accordance with (b)(1) and (b)(2). Operators may accomplish this inspection within the next 1,500 landings occurring subsequent to March 8, 1960, in lieu of the next 3,000 hours' time in service. (It will be necessary for operators to maintain a record of landings to ascertain compliance. If past records are unavailable, the number of prior landings may be estimated.) \n\n\t(Boeing Wire Service Bulletin No. 739 pertains to the above subject.) \n\n\tRevised November 18, 1960.\n \n\tRevised May 14, 1966.
91-03-13: 91-03-13 BOEING: Amendment 39-6882. Docket No. 89-NM-261-AD. \n\n\tApplicability: Model 747 series airplanes, as listed in Boeing Service Bulletin 747-57-2253, Revision 1, dated July 5, 1990, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo verify proper application of the center wing fuel tank secondary fuel barrier and prevent fuel or fuel vapors from entering the cargo or passenger compartments, accomplish the following: \n\n\tA.\tWithin the next 30 months after the effective date of this AD, inspect the center wing fuel tank secondary fuel barrier application, in accordance with Boeing Service Bulletin 747-57-2253, Revision 1, dated July 5, 1990. If the barrier has been improperly applied, as specified in the service bulletin, repair prior to further flight, in accordance with the service bulletin. \n\n\tB.\tWithin 30 days after accomplishing the inspection required by paragraph A. of this AD, submit a report of the complete findings of inspections from which it is determined that the secondary fuel barrier is not properly applied, to: Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, 1601 Lind Avenue SW., Renton, Washington 98055; rapid fax: (206) 227-1181; telex 756366. \n\n\tC.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tD.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington. \n\n\tThis amendment (39-6882, AD 91-03-13) becomes effective on March 11, 1991.
77-07-04: 77-07-04 GRUMMAN AMERICAN AVIATION CORPORATION: Amendment 39- 2861. Applies to GAAC Model AA-5, Serial Numbers 0641 through 0834, and to Model AA-5A, Serial Numbers 0001 through 0321, airplanes certificated in all categories. \n\n\tCompliance required as indicated below after the effective date of this AD unless already accomplished. \n\n\t(a)\tWithin 10 hours of flight after the effective date of this AD, in order to prevent possible failure of the carburetor heat valve assembly, remove the lower cowl and inspect the carburetor heat valve assembly for configuration as shown in Figure 1. \n\n\t\t(1)\tIf Configuration B is installed, reinstall lower cowl, and the aircraft may be approved for return to service. \n\n\t\t(2)\tIf Configuration A is installed, remove the carburetor heat valve assembly and inspect for cracks in the bend radius. If cracks are found, remove valve assembly from service and replace with a new valve assembly Part Number 5503006-505. Do not reuse the removed carburetor heat valve assembly which must be discarded. \n\n\t(b)\tIf no cracks are found in Configuration A carburetor heat valve assembly, valve assembly may be reinstalled, the lower cowl reinstalled, and after compliance with (c) below, the aircraft may be approved for return to service for twenty-five (25) hours of operation. \n\n\t\t(1)\tAt the end of the first period not to exceed twenty-five hours of operation after the initial inspection required in (a)(2), repeat the initial inspection procedure required in (a)(2). \n\n\t\t(2)\tIf no cracks are found at this second inspection of the carburetor heat valve assembly, the valve assembly may be reinstalled, the cowl replaced, and after compliance with (c) below, the aircraft may be approved for return to service for a second and final operational period not to exceed 25 hours. \n\n\t\t(3)\tNo Configuration A carburetor heat valve assembly may be continued in service in excess of 50 hours after the initial inspection required in (a)(2) above. \n\n\t(c)\tTo insure adequate carburetor heat rise, after the removal and reinstallation of the carburetor heat valve assembly, the following checks must be made prior to flight: \n\n\t\t(1)\tAfter carburetor heat valve assembly is installed into air box assembly, temporarily install air box assembly onto lower cowl. Remove air filter and check forward and aft gap between the valve assembly and carburetor heat box/lower cowl contact points. Maximum gap is 0.120 inches at both ends of valve with carburetor heat in the on and off position. If excessive gap exists, remove air box assembly and crimp edge of valve assembly up or down as required to obtain a gap less than 0.120 inch as specified in the GAAC Service Bulletin No. 159. \n\n\t\t(2)\tFollowing cowl installation, perform engine run up to check carburetor heat drop (50 RPM drop minimum). If drop does not meet minimum requirements, rework valve per subparagraph (c)(1) above. \n\n\t\t(3)\tCarburetor heat rigging to be accomplished in accordance with the AA-5 series Service Manual. \n\n\tGrumman American Aviation Corporation Service Bulletin No. 159 dated February 25, 1977, or later approved revisions, pertains to this subject. \n\n\tEquivalent methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southern Region, 3400 Whipple Street, East Point, Georgia 30344. \n\n\tThis amendment becomes effective April 11, 1977.
86-05-10: 86-05-10 MCDONNELL DOUGLAS: Amendment 39-5256. Applies to McDonnell Douglas Model DC-10-10, -10F, -15, -30, -30F, -40, and KC-10A (Military) series airplanes, certificated in any category, equipped with anti-skid valve assemblies, Goodyear Aerospace Corporation Part Numbers 6000189, 6000189-2, and 6000189-3, having serial number dates of JUL81-8519 through serial numbers dated April 1985, or any valves overhauled during January 1981 through April 1985 having end cap screws replaced with screws obtained from Goodyear Aerospace Corporation. Compliance required as indicated, unless previously accomplished. \n\n\tTo preclude the potential of hydraulic system fluid loss with subsequent loss of braking system performance on anti-skid valve assemblies, accomplish the following: \n\n\tA.\tWithin 12 months after the effective date of this airworthiness directive (AD), modify and reidentify the anti-skid valve assemblies as outlined in the Accomplishment Instructions of Goodyear Service Bulletin DC10-10-32-29 and DC10-30/40-32-41, dated May 3, 1985, or later revisions approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: Removal and installation of skid control manifold or skid control valve may be accomplished as outlined in McDonnell Douglas Service Bulletin 32-209, dated June 27, 1985. \n\n\tB.\tAlternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tC.\tSpecial 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. \n\n\tAll persons affected by this proposal who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director,Publications and Training, C1-750 (54-60); or Goodyear Aerospace Corporation, 1210 Massillon Road, Akron, Ohio 44315. These documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Los Angeles Aircraft Certification Office, 4344 Donald Douglas Drive, Long Beach, California. \n\n\tThis amendment becomes effective April 21, 1986.
51-21-01: 51-21-01 CESSNA: Applies to Models 120 and 140 Aircraft, Serial Numbers 8001 to 10650, Inclusive, and/or Any Other Serial Numbers Not Having Steel Reinforcing Channel, Cessna P/N 0433131 or Equivalent Installed. Compliance required prior to December 1, 1951. Inspect bottom rib of rudder for kinks or cracks in the rib flanges just aft of the rudder horn fitting. If there is any damage, the bottom rib assembly should be replaced with Cessna P/N 0433106, since the damage may progress until the rib breaks in two with the loss of rudder control. On installing P/N 0433106, which is an assembly of the rib, the control horn and a steel reinforcing channel (P/N 0433131), AD-4 protruding head type rivets should be used wherever rivets or spotwelds were used on the original installation; six additional rivets for attachment of aft part of steel channel to bottom rib and four AD-4 protruding head type or Cherry 163-4-4 rivets added to attach the skin to each flange of the steel reinforcing channel. If no damage to the bottom rib assembly is detected, installation of the steel reinforcing channel only is necessary. This may be accomplished by drilling out existing rivets for attachment of control horn and installing the steel reinforcing channel (P/N 0433131), above the bottom rib with flanges up, using existing rivet holes. The completed installation should be the same as for the installation of the complete lower rib outlined above. The reinforcing channel, P/N 0433131, is 4 17/32 long with 5/8 flanges, planform to fit inside lower rib installed as near horn flange as possible. Flanges removed on forward inch of channel. Material 0.036 1025 steel or 0.051 24ST, ALCLAD or equivalent. It is recommended that the length of the chains to the steerable tail wheel be so adjusted that under static conditions the coil springs are not extended more than 1/8 inch, since excessive tautness of the chains contributes to the rib failures. (Cessna Service Letter No.46, dated July 31, 1947, covers this same subject.) This supersedes AD 47-43-07.
2010-10-23: The FAA is superseding an existing airworthiness directive (AD) for the Dowty Propellers, propeller models listed above. That AD currently requires, for all Dowty Rotol propellers, visual inspections for seizure and for cadmium plating of the blade pitch change operating links and eyebolt fork assemblies. That AD also requires replacement or heat-treatment of the blade pitch change operating links and eyebolt fork assemblies, if necessary. This AD requires the same actions, but only for certain propeller models. This AD results from the FAA determining that AD 70-16-02 does not apply to all propellers, since current Dowty propellers are differently designed. We are issuing this AD supersedure to specify the affected propeller models, and to prevent seizure or embrittlement and cracking of the blade pitch change operating links and eyebolt fork assemblies, which could result in reduced controllability of the airplane.
62-25-01: 62-25-01 BOEING: Amdt. 509 Part 507 Federal Register November 21, 1962. Applies to Models 707 and 720 Series Aircraft. \n\n\tCompliance with certain service bulletins pertaining to the Model 707 and Model 720 flight control systems is considered necessary to provide for significant improvement in the safety and reliability of operation of those aircraft models. Accordingly, the aircraft models listed below shall be inspected and/or modified within the compliance times and in accordance either with the service bulletins as indicated or with equivalent methods approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region. Airplanes modified in accordance with later FAA approved revisions of the service bulletins listed below will be considered to have complied with the appropriate provisions of this AD. \n\n\t(a) Compliance required within the next 400 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.1. Stabilizer trim actuator auxiliary brake retaining nut. \n707 & 720\n984 \n\n\t(b) Compliance required within the next 650 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.\n1. Rudder pedal push-rod attachment. \n707\n337 \n2. Inboard aileron tab nose weight attachment screws. \n707\n860 \n3. Spoiler and emergency flap switch placard installation. \n707 & 720\n*1524 \n\n\t*BAC P/N 10-60424-621 (Type I) and P/N 10-60424-184 (Type I) are approved equivalents. \n\n\t(c) Compliance required within the next 2,700 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.\n1. Bearing retainer installation for center and inboard hinges for inboard aileron tab.\n707\n307 (R-1) and 307 (R-1)A.\n2. Guard installation for chain in stabilizer trim unit.\n707\n655 and 655B\n3. Flap drive torque tube guard installation in wheel well area.\n707\n680 and 680A\n4. Rudder controlinput stop modification and directional bushing replacement.\n707\n735 (R-1)\n5. Replacement of stabilizer trim actuator.\n707 & 720\n889\n6. Flap takeoff warning switch relocation.\n707 & 720\n1016 (R-1) and 1016 (R-1)C\n7. Stabilizer trim actuator motor replacement.\n707\n1247\n8. Emergency flap switch installation.\n707 & 720\n1251\n9. Rudder power con- trol unit replacement, "Extension Sleeve Revision".\n707 & 720\n1479 (R-1) Part I only\n10. Rudder pressure control valve bypass installation.\n707 & 720\n1482 (R-1).\n11. Rudder control centering spring cable modification.\n707 & 720\n1625\n12. Rudder control centering spring cable guard installation.\n707 & 720\n1680 & 1680A\n \n\n\t(d) Compliance required within the next 3,500 hours time in service following the effective date of the AD. \n\n\nModification\nModel\nService Bulletin No. \n1. Inboard aileron centering spring cartridge. \n707 & 720\n1344 \n2. Control wheel stabilizer trim switch installation. \n707 & 720\n1410** and 1410B\n3. Replacement of rudder hydraulic system solenoid valve. \n707 & 720\n1490 (R-1) \n\n\t**BAC P/N's 10-3265-6 and 10-60705-1 are approved equivalents. \n\n\t(e) Compliance required within the next 5,000 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No. \n1. Stabilizer trim actuator brake unlock gear ball bearing adapter addition. \n707 & 720\n1128 & 1128A \n2. Stabilizer trim actuator brake pawl spring. \n707 & 720\n1237\n3. Outboard spoiler shutoff valve consolidation.\n707 & 720\n1336 (R-1) and 1336 (R-1)B. \n4. Replacement of spoiler hydraulic system shutoff valve. \n707 & 720\n1484\n\n\tThis directive effective December 20, 1962. \n\n\tRevised December 24, 1963.
84-02-05: 84-02-05 BOEING: Amendment 39-4798. Applies to Boeing Model 747 series airplanes certificated in all categories. Compliance required as indicated, unless already accomplished. \n\n\tA.\tTo clarify the operation of the anti-icing system, emphasize the need to maintain the specified minimum N1 engine rpm during icing conditions, and expand the definition of icing conditions, accomplish the following: Within 120 days from the effective date of this AD, unless already accomplished, revise the FAA approved Airplane Flight Manual (AFM) CERTIFICATE LIMITATIONS SECTION by adding: \n\n\t\t\t\t"ENGINE ANTI-ICE SYSTEM \n\n\tWhen penetrating or operating in Icing Conditions, maintain a minimum of 50 percent N1 rpm at 10,000 feet and above, and 45 percent N1 rpm for Pratt & Whitney JT9D and General Electric CF6 engines, and 42 percent N1 for Rolls Royce RB211 engines, below 10,000 feet altitude, except as required for landing. \n\n\tNacelle anti-ice must be ON during all ground and flight operationswhen icing conditions exist or are anticipated, except during climb and cruise when the temperature is below -40 degrees C SAT. Nacelle anti-ice must be ON prior to and during descent in all icing conditions, including temperatures below -40 degrees C SAT. \n\n\t\tNOTE: Icing Conditions - Icing Conditions exist when the OAT on the ground and for takeoff, or TAT inflight is 10 degrees C or below and visible moisture in any form is present (such as clouds, fog with visibility of one mile or less, rain, snow, sleet, and ice crystals). \n\n\tIcing conditions also exist when the OAT on the ground and for takeoff is 10 degrees C or below when operating on ramps, taxiways or runways where surface snow, ice, standing water, or slush may be ingested by the engines or freeze on engines, nacelles or engine sensor probes." \n\n\tB.\tTo alert the flight crew of engine operation at a lower N1 than required for icing condition, install a LOW N1 rpm caution indication system as follows: \n\n\t\tWithin 24 months from the effective date of this AD, unless already accomplished, provide "LOW N1" indication that will alert the flight crew that the nacelle anti-ice is "ON and N1 is less than 45 percent N1 (42 percent N1 for RB211 engines) below 10,000 feet, and is less than 50 percent N1 above 10,000 feet altitude. \n\n\t\tNOTE: The LOW N1 indication may be provided by incorporating Boeing Service Bulletin S/B 747-77-2060 for the JT9D Pratt & Whitney powered airplanes and S/B 747-77-2063 for General Electric CF6 and Rolls Royce RB211 powered airplanes. \n\n\t\tBoth service bulletins have been approved by the FAA and were released on February 14, 1983. The service bulletins may be obtained from the Boeing Company at the following address: The Boeing Company, P.O. Box 3707, Seattle, Washington 98124. \n\n\tC.\tAlternate means of compliance with the AD which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, 9010 East MarginalWay South, Seattle, Washington. \n\n\tD.\tA special flight permit may be issued in accordance with FAR 21.197 and 21.199 for the purpose of flying the aircraft which has exceeded the compliance period to a maintenance facility where the modification can be performed. \n\n\tThis amendment becomes effective on March 2, 1984.
75-09-04 R1: 75-09-04 R1 BOEING: Amendment 39-2174 as amended by Amendment 39-4926. Applies to Boeing Model 727 series airplanes, certificated in all categories, listed in Boeing Service Bulletin 727-55-62, or later FAA approved revisions (line numbers 1 through 641, inclusive). Compliance required as indicated. \n\n\tTo detect cracks in the horizontal stabilizer rear spar center section fitting, accomplish the following: \n\n\tA.\tWithin the next 750 flight hours after the effective date of this AD, unless accomplished within the last 2250 flight hours, inspect the horizontal stabilizer rear spar center section fitting in accordance with Paragraph III of Boeing Service Bulletin 727-55-62, or later FAA approved revisions or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t1.\tIf no cracks are found in the fitting, repeat the inspections at intervals not to exceed 3000 flight hours, until replaced per Paragraph B. \n\n\t\t2.\tIf a crackis found at the upper or lower flanges of the fitting and is within the allowable limits specified in Figure 2 of Boeing Service Bulletin 727-55-62, or later FAA approved revisions, prior to further flight, stop drill the crack per the service bulletin. Inspect the stop drilled areas for crack growth at 1500 flight hour intervals, until repaired or replaced in accordance with Boeing Service Bulletin 727-55-62, or later FAA approved revisions, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t3.\tIf a crack is found in any of the hinge lugs and is within the allowable limits specified in Boeing Service Bulletin 727-55-62, or later FAA approved revisions, before further flight, repair the fitting in accordance with Figure 3 of the service bulletin, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. Inspect the repaired fitting thereafter at intervals not to exceed 1500 flight hours. \n\n\t\t4.\tIf a crack at any location is beyond the allowable limits specified in Boeing Service Bulletin 727-55-62, or later FAA approved revisions, before further flight, repair in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, or replace the fitting with a new improved fitting per Paragraph B. \n\n\tB.\tAs terminating action for this AD, replace the horizontal stabilizer rear spar center section fitting with a new improved 7075-T73 aluminum allow fitting. \n\n\tC.\tAirplanes having cracked horizontal stabilizer rear spar center section fittings which require replacement under this AD may be flown in accordance with FAR 21.197 to a base where the replacement can be accomplished. \n\n\tD.\tUpon request of an operator, an FAA Maintenance Inspector, subject to the prior approval of the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, may adjust the inspection intervals in this AD, if the request contains substantiating data to justify the increase for that operator. \n\n\tThe 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). \n\n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. The documents may also be examined at FAA Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington. \n\n\tAmendment 39-2174 became effective April 29, 1975. \n\n\tThis Amendment 39-4926 becomes effective January 22, 1985.
90-23-12: 90-23-12 BOEING: Amendment 39-6799. Docket No. 90-NM-134-AD. \n\n\tApplicability: Model 737-300 and 737-400 series airplanes, listed in Boeing Service Bulletin 737-49-1071, dated May 10, 1990, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent auxiliary power unit (APU) rotor failure resulting from an undetected EGT overtemperature condition, accomplish the following: \n\n\tA.\tFor Model 737-400 series airplanes: Within 1,000 hours time-in-service after May 29, 1990 (the effective date of Amendment 39-6583, AD 90-09-05), modify the APU instrumentation wiring in a manner that will assure continuous flight-compartment APU exhaust gas temperature (EGT) indication following a shutdown. The modification must be accomplished in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Transport Airplane Directorate; or in accordance with Boeing Service Bulletin 737-49-1071, dated May 10, 1990. \n\n\tB.\tFor Model 737-300 series airplanes: Within 1,000 hours time-in-service after the effective date of this amendment, modify the APU instrumentation wiring in accordance with Boeing Service Bulletin 737-49-1071, dated May 10, 1990. \n\n\tC.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tD.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer mayobtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tAirworthiness Directive 90-23-12 supersedes AD 90-09-05, Amendment 39-6583. \n\tThis amendment (39-6799, AD 90-23-12) becomes effective on December 11, 1990.
59-12-05: 59-12-05 COLONIAL: Applies to Models C-1 and C-2 Aircraft Serial Numbers 1 Through 132. Compliance required as indicated. Due to a recent incident where the plastic lock for the control surface hinge pin cracked, thus making it possible for the hinge pin to work out, the following inspection and replacement of all plastic locks is required. Prior to next flight inspect the control surface hinge pin locks. (1) If made of metal, no further action necessary. (2) If made of plastic material inspect for cracks. Parts found cracked must be replaced with locks fabricated of 0.025 2024-T3 aluminum alloy material or equivalent before further operation. (3) All plastic locks must be replaced within the next 10 hours of operation with metal locks fabricated of 0.025 2024-T3 aluminum alloy material or equivalent. (Colonial Service Bulletin No. 15 covers this same subject.)
2010-10-10: We are adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation Model 390 airplanes. This AD requires you to inspect the essential bus lightning strike protection for proper installation of metal oxide varistor (MOV) and spark gap wiring. This AD also requires you to rework the wiring as necessary to achieve the required lightning strike/surge protection. This AD results from a report that the wires to the MOV and spark gap were swapped. We are issuing this AD to detect and correct improper installation of the MOV and spark gap wiring, which could result in overload of the MOV in a lightning strike and allow electrical energy to continue to the essential bus and disable equipment that receives power from the essential bus. The disabled equipment could include the autopilot, anti-skid system, hydraulic indicator, spoiler system, pilot primary flight display, audio panel, or the 1 air data computer. This failure could lead to a significant increase in pilot workload during adverse operating conditions.
91-13-06: 91-13-06 McDONNELL DOUGLAS: Amendment 39-7037. Docket No. 91-NM-03-AD. Supersedes AD 90-03-17. \n\n\tApplicability: Model DC-9 series, Model DC-9-80 series, C-9 (Military), and Model MD-88 airplanes, equipped with Westinghouse bus control unit (BCU) Part Number 947F946-2, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent the loss of generator electrical power, accomplish the following: \n\n\tA.\tWithin 30 days after February 15, 1990 (the effective date of AD 90-03-17, Amendment 39-6496), add the following to the LIMITATIONS section of the approved Airplane Flight Manual (AFM). This may be accomplished by inserting a copy of this AD into the AFM. \n\n\t\t1.\tOn takeoff with both engine generators and APU generator operating and APU bus switches selected to the ON position, the AC BUS X-TIE switch must be placed in the OPEN position. At or above 10,000 feet MSL with both engine generators operating, the APU may be shut down and the AC BUS X-TIE switch placed in the AUTO position. \n\n\t\tNOTE: In the event of an in-flight failure of an engine generator that results in the APU generator powering an AC bus, de-activate all galley power and place the other APU BUS switch in the ON position. \n\n\t\t2.\tOn takeoff with the APU generator inoperative, or an engine generator inoperative, dispatch is permitted in accordance with the present MMEL conditions except that the AC BUS X-TIE switch must be in the OPEN position. Verify that all transformer rectifiers (TR's) are operating and place the DC BUS X-TIE switch in the CLOSE position. Takeoff minimums are restricted to ceiling 1,000 foot and visibility 3 miles. The Captain must make the takeoff with his instrument incandescent flood lights adjusted to a level which would adequately compensate for the subsequent potential loss of his instrument integral lights. At or above 10,000 feet MSL, place the DC BUS X-TIE switch in the OPEN position and the AC BUSX-TIE switch in the AUTO position. \n\n\t\tNOTE: In the event of an in-flight failure of an engine generator following dispatching with the APU generator powering an AC bus, de-activate all galley power and place the other APU BUS switch in the ON position. \n\n\t\t3.\tWhen operating with the AC BUS X-TIE switch in the AUTO position, if rapid cycling of the AC cross-tie relay occurs, manifested by a buzzing/chattering sound from the electrical power center and any combination of random circuit breaker trips, inappropriate aural warning messages, loss of some flight instruments, and/or flashing cockpit annunciators, place the AC BUS X-TIE switch to the OPEN position. If a generator trips off-line, it may be reset only once. If the engine generator fault cannot be cleared, the APU should be utilized, if available. \n\n\t\t4.\tPrior to the approach with both engine generators operating, start the APU and place both APU BUS switches to the ON position. Place the AC BUS X-TIE switch to the OPENposition after APU electrical power becomes available. \n\n\t\t5.\tPrior to the approach with only two generators operating, place the AC BUS X-TIE switch in the OPEN position. Landing minimums are restricted to Category I and the Captain must make the approach with his instrument incandescent flood lights adjusted to a level which would adequately compensate for the subsequent potential loss of his instrument integral lights. \n\n\t\t6.\tIn the event of an in-flight failure that results in an AC bus not powered, place the DC BUS X-TIE switch in the CLOSE position. \n\n\t\t7.\tIn the event of an in-flight failure that results in both AC BUSES being powered by only one generator, the landing minimums are restricted to ceiling 1,000 feet and 3 miles visibility. The Captain must make the approach and landing. \n\n\t\t8.\tAutoland is permitted with two engine generators operating and APU generator operating with both APU BUS switches in the ON position and the AC BUS X-TIE switch in the OPEN position. Reconfirm APU generator availability after "AUT LND/AUT LND" is indicated on the Flight Mode Annunciator (FMA). An autoland approach must be discontinued following a failure of an engine generator. \n\n\tB.\tWithin 2 years after the effective date of this AD, replace the Westinghouse Bus Control Unit, Part Number 947F946-2 with the Westinghouse Bus Control Unit, Part Number 947F946-3, in accordance with McDonnell Douglas DC-9 Service Bulletin 24-119 dated January 24, 1990. The limitations on the electrical operating procedures and restricted operation of the automatic landing system required by paragraph A. of this AD may be removed from the AFM when the modified Westinghouse BCU is installed. \n\n\tC.\tAn alternate means of compliance and adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Los Angeles ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Los Angeles ACO. \n\n\tD.\tSpecial 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 requirement of this AD. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Business Unit Manager, Technical Publications, C1-HCW (54-60). These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington, or the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California. \n\n\tThis amendment supersedes Amendment 39-6496, AD 90-03-17. \n\n\tThis amendment(39-7037, AD 91-13-06) becomes effective on July 15, 1991.
67-05-03: 67-05-03 MCDONNELL DOUGLAS: Amend. 39-347 as amended by amendment 39-1160. Applies to Model DC-8 Series (except DC-8-61) and DC-8F Series Airplanes as Indicated Herein. \n\n\tCompliance required as indicated. \n\n\tNumerous reports have been received concerning cracks in the inboard and outboard pylons and pylon stub wing structure. These cracks have been found in the skins, bulkheads, and spar caps. The cracks have varied in length, and in some cases, the parts have cracked completely through. To prevent further failure of this nature, accomplish the following, or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region: \n\n\t(a)\tFor all inboard and outboard pylons that have accumulated a total of 6,000 or more hours' time in service at the effective date of this AD, unless already accomplished within the last 2,000 hours' time in service, and for all inboard and outboard pylons accumulating a total of 6,000 hours' time in service after the effective date of this AD, within the next 1,000 hours' time in service, conduct a visual inspection as follows: \n\n\t\t(1)\tInspect outboard pylons in the area of the pylon stub wing lower skin and doubler for cracks in the cutout for the leading edge lower slot door on airplanes with Serial Numbers 45256-45272, 45274-45277, 45282, 45284-45289, 45292-45300, 45304-45306, 45376=45382, 45384-45390, 45392, 45393, 45408-45413, 45418, 45419, 45421, 45425-45427, 45429-45431, 45433-45437, 45445, 45565-45567, 45569, 45570, 45596, 45597, 45602, 45603, 45605, 45606, 45609-45612, 45626, 45627, 45638, 45645, 45646, 45649, 45650, 45672, 45673, 45687-45690, 45806-45808, 45815, 45877. Refer to Service Sketch No. 754 of Douglas Engineering Service Letter C1-78-2016/DBA, October 18, 1966, or later revisions for location of the cracks. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. If cracks arefound during this inspection or any reinspection, the defective parts must be replaced before further flight with uncracked parts, or parts reworked in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended as a redesign to prevent further cracking of these areas and approved as such, the repetitive inspection requirements may be discontinued. \n\n\t\t(2)\tInspect at Station YOP 229 the outboard pylon inboard skin in the vicinity of the slot leading edge on airplanes with Serial Numbers 45252-45272, 45274-45283, 45289, 45291-45300, 45304-45306, 45376-45382, 45384-45393, 45418, 45419, 45421-45427, 45429-45431, 45433-45437, 45442-45445, 45565-45567, 45569, 45570, 45588-45597, 45602, 45603, 45605, 45609-45612, 45626, 45627, 45638, 45645, 45646, 45649, 45650. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during the inspection or any reinspection must be reworked before further flight in accordance with Douglas Service Bulletins Nos. 54-30, Kits J and K, dated November 27, 1962, 54-31, Kits L and M dated November 30, 1962, or 54-32, Kits J and K, dated December 28, 1962, or later FAA-approved revisions, whichever is applicable to the serial number of the airplanes listed therein. After rework in accordance with this AD, the repetitive inspection requirement may be discontinued. \n\n\t\t(3)\tInspect at YOP Station 252 to YOP Station 262 the aft inboard flex panel on the outboard pylons on airplanes with Serial Numbers 45252-45263, 45274-45276, 45278-45283, 45289, 45291-45297, 45376-45379, 45384-45387, 45391, 45392, 45416, 45418, 45419, 45422-45427, 45429, 45442-45445, 45567, 45569, 45588-45595, 45598-45600, 45602, 45603. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during the inspection or any reinspection must be repaired in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight or replaced before further flight in accordance with Douglas Service Bulletins Nos. 54-10 Kit F, dated August 17, 1960, 54-14, Kit B, dated November 3, 1960, or later FAA approved revisions, whichever is applicable to the serial number of the airplane listed therein. After repair or replacement in accordance with this AD, the repetitive inspection requirement may be discontinued. \n\n\t\t(4)\tInspect the inboard side of the inboard pylon in the area of the D-duct spar angle and the pylon side skin on airplanes with Serial Numbers 45253-45272, 45274-45283, 45289, 45291-45300, 45304-45306, 45376-45382, 45384-45393, 45418, 45419, 45421-45427, 45429-45431, 45433-45437, 45442-45445, 45565-45567, 45569, 45570, 45588-45597, 45602, 45603, 45605,45606, 45609-45612, 45626, 45627, 45638. Refer to Service Sketch No. 756, Douglas Service Engineering Letter C1-78-2106/DBA, October 18, 1966, or later revisions for location of cracks. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during this inspection or any reinspection, must be replaced before further flight with uncracked parts or parts reworked in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended as a redesign to prevent further cracking of these areas, and approved as such, the repetitive inspections required may be discontinued. \n\n\t(b)\tOn all airplanes having Serial Numbers 45252-45272, 45274-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45427, 45429-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45614, 45616-45630, 45632-45638, 45640-45651, 45653, 45655-45663, 45665-45673, 45676, 45684, conduct an inspection for cracks on all upper inboard and outboard spar caps of the outboard pylons in the area between Sta. YOP 214 and Sta. YOP 255, and repair or replace if necessary as follows: \n\n\t\t(1)\tUnless already accomplished within the last 475 hours' time in service within the next 25 hours' time in service after the effective date of this AD, visually inspect the outboard pylons that have accumulated a total of 8,000 hours' time in service as of the effective date of this AD; or \n\n\t\t(2)\tFor outboard pylons that have not accumulated a total of 8,000 hours' time in service as of the effective date of this AD, visually inspect them within 25 hours' time in service after 8,000 hours' time in service is reached. \n\n\t\t(3)\tIf no cracks are found during the inspections required by paragraphs (b)(1) or (b)(2), the inspections required therein must be repeated at intervals not to exceed 500 hours' time in service from the date of the last inspection. If cracks are found during this inspection or any reinspection before further flight, replace or repair in accordance with Douglas Service Bulletins Nos. 54-33, dated March 15, 1963, 54-35, dated April 9, 1965, or later FAA-approved revisions, whichever is applicable to the affected airplane. Upon replacement or repair as provided for in this AD, the 500 hours' repetitive inspection may be discontinued. \n\n\t\t(4)\tThe inspections required in Paragraph (b)(1) and (b)(2) do not apply to airplanes modified in accordance with McDonnell Douglas Service Bulletin Nos. 54-33, dated March 15, 1963; 54-35, dated April 9, 1965; or later FAA-approved revisions. Each service bulletin contains the airplane serial numbers to which it is applicable. \n\n\tHowever, all airplanes modified in accordance with S.B. Nos. 54-33 or 54-35 must be reinspected per Paragraph (b) above within the next 1500 hours' time in service after the effective date of this AD amendment, unless already accomplished within the last 1500 hours' time in service, and thereafter at intervals not to exceed 3000 hours' time in service. Upon installation of an improved spar kit per McDonnell Douglas S.B. 54-57, Revision 1, dated December 9, 1969, or later FAA approved revision, these repetitive inspections may be discontinued. \n\n\t(c)\tOn all airplanes with Serial Number 45252-45272, 45274-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45427, 45429-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45614, 45616-45630, 45632-45638, 45640-45651, 45653, 45655-45694, 45750, 45752-45760, 45762-45769, 45800, 45809, 45814-45821, 45824, 45850-45862, 45877-45882, 45886, 45916, conduct an inspection on the cant bulkhead fittings in all outboard pylons, and on the cant bulkhead fittings in all inboard pylons for the referenced serial numbers, with the addition of Serial Number 45883, in the area shown on Service Sketch No. 669, Douglas Engineering Service Letter C1-78-2016/DBA, October 18, 1966, or later FAA-approved revision, and rework or replace, if necessary, as follows - \n\n\t\t(1)\tUnless already accomplished within the last 2,000 hours' time in service, within the next 1,000 hours' time in service after the effective date of this AD, visually inspect the inboard and outboard pylons that have accumulated 10,000 or more hours' time in service as of the effective date of this AD; or \n\n\t\t(2)\tFor those inboard and outboard pylons referenced in paragraph (c) that have not accumulated a total of 10,000 hours' time in service as of the effective date of this AD, visually inspect them within 1,000 hours' time in service after 10,000 hours' time in service is reached. \n\n\t\t(3)\tIf no cracks are found during the inspection required by paragraphs (c)(1) and (2), repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. If cracks are found during inspection or reinspectionbefore further flight, replace the defective parts with uncracked parts, or rework in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended to prevent further cracking of the structure and is approved as such, the repetitive inspection requirement may be discontinued. \n\n\t(d)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, 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 the operator. \n\n\t(Douglas Service Engineering Letter C1-78-2016/DBA dated October 18, 1966, or later revisions, covers this subject.) \n\n\tAmendment 39-347 effective February 4, 1967. \n\n\tThis Amendment 39-1160 becomes effectiveApril 4, 1971.
82-04-02: 82-04-02 MCDONNELL DOUGLAS: Amendment 39-4317. Applies to all McDonnell Douglas Model DC-9 and C-9 series airplanes, certificated in all categories with rudder pedal arm P/N 3616012 installed with more than 13,500 hours time in service. (Note: Time in service on the rudder pedal arm may be used, if the operator has records to substantiate it.) Compliance required as prescribed herein. To detect fatigue cracking and possible structural failure of the rudder pedal arms, P/N 3616012, accomplish the following, unless already accomplished: \n\n\tA.\tWithin the next 2,000 landings or six months after the effective date of this AD, whichever occurs first, perform ultrasonic and dye penetrant inspections on rudder pedal arm assemblies, P/N 3616012, as outlined in Service Sketch 3251 and Accomplishment Instructions of McDonnell Douglas DC-9 Service Bulletin 27-209 dated May 29, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region. \n\n\tB.\tIf no cracks are found, replace the rudder pedal arms with new P/N 3953505 aluminum rudder pedal arm assemblies or retain the 3616012 parts and repeat ultrasonic and dye penetrant inspections at intervals not to exceed 4,000 landings or one year, whichever occurs first. Replacement with aluminum rudder pedal arm assemblies constitutes terminating action for this AD. \n\n\tC.\tIf cracks are found, prior to further flight replace the rudder pedal arms with: \n\n\t\t1.\tNew P/N 3953505 aluminum rudder pedal arm assemblies and thereby terminate the repetitive inspection requirements of this AD, or \n\n\t\t2.\tReplace with new P/N 3616012 magnesium rudder pedal arm assemblies, and repeat inspections specified in paragraph B above. \n\n\tD.\tSpecial 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. \n\n\tE.\tFor the purposes of complying with this AD, subject to acceptance by the assigned FAA Maintenance Inspector, the number of landings may be determined by dividing each airplane's number of hours time in service by the operator's fleet average time from takeoff to landing. \n\n\tF.\tUpon the request of an operator, an FAA Maintenance Inspector, subject to prior approval by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region, may adjust the inspection times specified in this AD to permit compliance at an established inspection period of that operator if the request contains substantiating data to justify the change for that operator. \n\n\tG.\tAlternative means of compliance with this AD which provide an equivalent level of safety may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region. \n\n\tThe 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 proposal who have not already received these documents from the manufacturer may obtain copies upon request to the McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). These documents also may be examined at the FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108, or 4344 Donald Douglas Drive, Long Beach, California 90808. \n\n\tThis airworthiness directive becomes effective March 21, 1982.
80-09-05: 80-09-05 BOEING: Amendment 39-3760. Applies to all 737-100 and -200 series airplanes certificated at takeoff weights in excess of 97,800 pounds and containing the horizontal stabilizer trim actuator P/N 10-61326-4 or P/N 10-61326-5. \n\n\tNOTE: The 737 airplanes from line 482 and on were delivered with horizontal stabilizer trim actuator P/N 10-61326-6. This AD will apply to those 737 airplanes line number 482 and on if P/N 10-61326-4 or -5 have been exchanged for P/N 10-61326-6.\n \n\tCompliance is required as follows: To assure sufficient horizontal stabilizer trim capability, accomplish either A, B or C below within the next six (6) months after the effective date of this AD, unless already accomplished. \n\n\tA.\tFor horizontal stabilizer trim actuators having the P/N 10-61326-4 or P/N 10-61326- 5, test the actuator stall torque in accordance with Boeing Service Bulletin No. 737-27-1101, dated February 1, 1980. The actuators found to have less than 350 inch pounds of torque must be replaced with a serviceable actuator P/N 10-61326-4, -5, or -6. Thereafter, for the actuators P/N 10-61326-4 and -5, conduct repetitive torque test per the "Actual Stall Torque/Maximum Test Interval" chart, Figure 2, of that service bulletin at intervals not to exceed the maximum test intervals (hours) indicated by the curve on the chart. \n\n\tB.\tReplace the horizontal stabilizer trim actuator, Boeing P/N 10-61326-4 or -5 with stabilizer trim actuator, Boeing P/N 10-63126-6 in accordance with Boeing Service Bulletin No. 737-27-1101, dated February 1, 1980. Replacing the stabilizer trim actuator, Boeing P/N 10- 61326-4 or -5, with Boeing P/N 10-61326-6 is the terminating requirement for this AD. \n\n\tC.\tPerform an equivalent inspection and/or installation approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereofpursuant 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 Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA, Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis amendment supersedes AD 70-25-10.\n \n\tThis amendment becomes effective May 6, 1980.