Results
87-06-10: 87-06-10 BOEING: Amendment 39-5578. Applies to all Model 727 series airplanes, certificated in any category. Compliance required as indicated, unless previously accomplished. \n\n\tPrior to the accumulation of 10,000 landings or 6 years from date of manufacture or prior replacement, or within 1 year after the effective date of this AD, whichever occurs latest, accomplish the following:\n \n\tA.\tReplace the sealed needle bearings in the downlock outer link of the side strut upper segment of the left and right main landing gear assemblies, part number BACB10B107J or BACB10CC10E, in accordance with the Boeing 727 Overhaul Manual, Subject 32-13-01, with new bearings with the same part number. Inspect retainer bolt for damage or corrosion. If damage or corrosion is detected, replace the bolt with a new bolt part number NAS1110-100DW or BACB30LT10D-100. Repeat the replacements and inspections at intervals not to exceed 6 years or 10,000 landings, whichever occurs first. \n\n\tB.\tTerminating action for the requirements of paragraph A., above, is replacement of the sealed needle bearings and the bolts in the downlock outer link of the side strut upper segment of the left and right main landing gear assemblies, part number BACB10B107J or BACB10CC10E, with the self-lubricating Karon lined Kamatics bearings and high strength chrome plated CRES bolts specified in Boeing Service Bulletin 727-32-0341, dated December 18, 1986, or later FAA-approved revision. \n\n\tC.\tAn alternate means of compliance or adjustment of compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tD.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. \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 for the airplane type. \n\n\tAll persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124-2207. This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective April 13, 1987.
2010-17-12R1: The FAA is revising an existing airworthiness directive (AD) for the products listed above. This AD revision results from the need to correct the applicability paragraph of that AD, and 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: Strip results from some of the engines listed in the applicability section of this AD revealed excessively corroded low- pressure turbine disks stage 2 and stage 3. The corrosion is considered to be caused by the environment in which these engines are operated. Following a life assessment based on the strip findings it is concluded that inspections for corrosion attack are required. The action specified by this European Aviation Safety Agency (EASA) AD 2008-0122 was intended to avoid a failure of a low- pressure turbine disk stage 2 or stage 3 due to potential corrosion problems which could result in uncontained engine failure and damage to the airplane. It has been later realized that the same unsafe condition could potentially occur on more serial numbers for the Tay 650-15 engines and on the Tay 651-54 engines. This AD, superseding EASA AD 2008-0122, retaining its requirements, is therefore issued to expand the Applicability in adding further engine serial numbers for the Tay 650-15 engines and in adding the Tay 651-54 engines. We are issuing this AD to detect corrosion that could cause the stage 2 or stage 3 disk of the LP turbine to fail, uncontained engine failure, and damage to the airplane.
2008-19-05: The FAA is adopting a new airworthiness directive (AD) for Lycoming Engines (formerly Textron Lycoming) models 320, 360, and 540 series, ``Parallel Valve'' reciprocating engines, with certain Engine Components, Inc. (ECi) cylinder assemblies, part number (P/N) AEL65102 series ``Titan'', installed. This AD requires initial and repetitive visual inspections and compression tests to detect cracks at the head- to-barrel interface, replacement of cylinder assemblies found cracked, and replacement of certain cylinder assemblies, at new reduced times- in-service. This AD results from reports of 45 failures with head separations of ECi cylinder assemblies. We are issuing this AD to prevent loss of engine power due to cracks at the head-to-barrel interface in the cylinder assemblies and possible engine failure caused by separation of a cylinder head, which could result in loss of control of the aircraft.
92-24-07: 92-24-07 BOEING: Amendment 39-8412. Docket No. 92-NM-38-AD. \n\n\tApplicability: Boeing Model 747-300, 747-400, and 747-100B series airplanes delivered with stretched upper decks, and Boeing Model 747 series airplanes modified to have stretched upper decks; as listed in Boeing Service Bulletin 747-53-2327, dated December 5, 1991; certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent reduced controllability of the airplane as a result of failure of the upper floor beams and probable interference with the control cables, accomplish the following: \n\n\t(a)\tConduct a high frequency eddy current inspection of the Body Station (BS) 860 and BS 980 upper deck floor beams to detect cracks, in accordance with Boeing Service Bulletin 747-53-2327, dated December 5, 1991, at the applicable time specified in subparagraph (a)(1) or (a)(2) of this AD. If cracks are found as a result of this inspection, prior to further flight, repair in a manner approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\t\t(1)\tFor Boeing Model 747-300, -400, and -100B series airplanes delivered with stretched upper decks: Prior to the accumulation of 20,000 flight cycles, or within 1,000 flight cycles after the effective date of this AD, whichever occurs later. \n\n\t\t(2)\tFor Boeing Model 747 series airplanes modified to have stretched upper decks: Prior to the accumulation of 20,000 flight cycles after incorporation of the stretched upper deck modification, or within 1,000 flight cycles after the effective date of this AD, whichever occurs later. \n\n\t(b)\tIf no cracks are found as a result of the inspection required by paragraph (a) of this AD, reinforce the BS 860 and BS 980 upper deck floor beams in accordance with Boeing Service Bulletin 747-53-2327, dated December 5, 1991, at the applicable time specified in subparagraph (b)(1) or (b)(2) of this AD. \n\n\t\t(1)\tFor Boeing Model 747-300, -400, and -100B series airplanes delivered with stretched upper decks: Prior to the accumulation of 20,000 flight cycles, or within 1,000 flight cycles after the effective date of this AD, whichever occurs later. \n\n\t\t(2)\tFor Boeing Model 747 series airplanes modified to have stretched upper decks: Prior to the accumulation of 20,000 flight cycles after incorporation of the stretched upper deck modification, or within 1,000 flight cycles after the effective date of this AD, whichever occurs later. \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, Seattle 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, Seattle ACO. \n\n\tNOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Seattle 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 and reinforcement shall be done in accordance with Boeing Service Bulletin 747-53-2327, dated December 5, 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 Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124-2207. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. \n\n\t(f)\tThis amendment becomes effective on December 15, 1992.
95-01-06 R1: This amendment revises an existing airworthiness directive (AD), applicable to certain Boeing Model 737-200 and -300 series airplanes, that currently requires inspections to detect cracking in the radii on the support angles on the lower jamb (latch lug fittings) of the main deck cargo door, and replacement of cracked parts. That amendment was prompted by reports of premature fatigue cracking on the support angles on the lower jamb of the main deck cargo door. The actions specified in that AD are intended to prevent in-flight separation of the main deck cargo door from the airplane due to fatigue cracking on the support angles on the lower door jamb. This amendment requires a change in the cognizant aircraft certification office for requesting approvals of alternative methods of compliance with the provisions of this AD.
86-09-07 R1: 86-09-07 R1 BOEING: Amendment 39-5306 as amended by Amendment 39-5580. Applies to Model 747 series airplanes listed in Boeing Alert Service Bulletin 747-53A2267, dated March 28, 1986, certificated in any category. \n\n\tTo prevent a condition that could lead to depressurization of the airplane accomplish the following, unless already accomplished: \n\n\tA.\tPerform an external visual inspection of lower longitudinal lap joint areas for corrosion in accordance with Boeing Alert Service Bulletin 747-53A2267, dated March 28, 1986, or later FAA- approved revision. \n\n\t\t1.\tInspect within the next 100 landings after the effective date of this amendment, or within 4,000 flight hours or 15 months from the last inspection of each lap joint area, provided no corrosion was found at that inspection, whichever is later. \n\n\t\t2.\tIf corrosion is found and is associated with cracks, skin penetration, or missing fasteners; or if it exceeds a total length of 20 inches along any 40-inch distance; repair before further flight in accordance with paragraph C., below. Reinspect airplanes with less severe corrosion at intervals not to exceed 1,000 flight hours or 6 months, whichever is sooner. \n\n\t\t3.\tIf no corrosion is found, reinspect at intervals not to exceed 4,000 flight hours or 15 months, whichever is sooner. \n\n\tB.\tIf corrosion is found in the longitudinal lap joint area, perform an internal inspection of the body frames for cracking in the vicinity of each corroded lap joint area before further flight in accordance with Boeing Alert Service Bulletin 747-53A2267, dated March 28, 1986, or later FAA- approved revision, unless already accomplished within the last 1,000 landings. Repeat the frame inspections thereafter at intervals not to exceed 1,000 landings until terminating action is performed on the lap joints in accordance with paragraph C., below. Repair cracks before further flight in accordance with an FAA-approved method. \n\n\tC.\tTerminating action for the repetitiveinspections of paragraphs A. and B., above, is repair of the lap joints in accordance with Boeing Service Bulletin 747-53A2267, dated March 28, 1986, or later FAA-approved revision. \n\n\tD.\tApply organic corrosion inhibiting compound in accordance with Boeing Service Bulletin 747-53A2267, Revision 1, dated September 25, 1986, or later FAA-approved revisions. \n\n\tE.\tFor the purposes of complying with this AD, the number of landings may be determined to equal the number of pressurization cycles where the cabin pressure differential was greater than 2.0 psi. \n\n\tF.\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, FAA, Northwest Mountain Region. \n\n\tG.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes unpressurized to a base for the accomplishment of inspections and/or modifications requiredby this AD. \n\n\tH.\tFor Boeing Model 747SR airplanes only, based on continued mixed operation of cabin pressure differentials, the reinspection intervals specified in paragraph B. of this AD may be multiplied by a 1.2 adjustment factor. \n\n\tAll persons affected by this proposal who have not already received the appropriate service information from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124-2207. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tAmendment 39-5306 superseded AD 76-05-05, Amendment 39-2538, as amended by Amendments 39-2698, 39-2961, and 39-3430. \n\n\tAmendment 39-5306 became effective May 27, 1986. \n\tThis amendment, 39-5580, becomes effective March 30, 1987.
95-23-09: This amendment supersedes an existing airworthiness directive (AD), applicable to McDonnell Douglas Model DC-10 series airplanes and KC-10A (military) airplanes, that currently requires the implementation of a program of structural inspections to detect and correct fatigue cracking in order to ensure the continued airworthiness of these airplanes as they approach the manufacturer's original fatigue design life goal. This amendment requires clarification of some Principle Structural Elements (PSE) and some non-destructive inspection (NDI) procedures. This amendment is prompted by new data submitted by the manufacturer indicating that certain revisions to the program are necessary in order to clarify some PSE's and some NDI procedures. The actions specified by this AD are intended to prevent fatigue cracking that could compromise the structural integrity of these airplanes.
87-26-02 R1: 87-26-02 R1 MESSERSCHMITT-BOLKOW-BLOHM (MBB): Amendment 39-5795 as revised by Amendment 39-6313. (Docket No. 87-SW-39-AD) \n\n\tApplicability: Model BO-105 series helicopters, certificated in any category, equipped with main rotor blade rotating control rod ends, P/N's 105-13141.01 and 105-13142.01. \n\n\tCompliance: Required as indicated unless already accomplished. \n\n\tTo detect and prevent fatigue cracks in the threads of the main rotor pitch links rod ends, accomplish the following: \n\n\t(a)\tBefore the first flight of each day after the effective date of this AD, check the main rotor control rods for binding of the spherical bearings as follows: \n\n\t\t(1)\tCheck the bearings on each control rod by rotating the rod about its longitudinal axis by hand. This check may be conducted by the pilot and must be recorded in accordance with Section 43.9. \n\n\tNOTE: The pilot, when complying, must make appropriate entries and the record must be maintained per Section 91.173 or Section 135.439. \n\n\t\t(2)\tInspect and rework control rods containing a binding bearing in accordance with paragraphs (b)(1) through (b)(7) before further use. \n\n\t(b)\tWithin the next 50 hours' time in service after the effective date of this AD, conduct the following main rotor control rod inspection for cracks and spherical bearing torque determination: \n\n\t\t(1)\tRemove the control rods from the helicopter. \n\n\t\t(2)\tRemove the rod ends from the control rods. \n\n\t\t(3)\tMeasure the force required to rotate the inner race of the rod end bearings in the circumferential direction around a bolt centerline (ref. figure No.1). \n\n\t\t(4)\tIf the bearing torque (force X moment arm) required to rotate the bearing inner race is more than 1.5 NM (13.3 in-lbs), accomplish either of the following-- \n\n\t\t\t(i)\tReplace the affected bearing and repeat the step in paragraph (b)(3); or \n\n\t\t\t(ii)\tReduce the bearing friction by installing a bolt through the bearing inner race and tighten with a nut. Then, using a drill, spin the bearing at 60 to 100 RPM for 1- or 2-minute periods until friction torque is reduced to 1.5 NM (13.3 in-lbs) or below. Monitor temperature constantly to avoid overheating. (Bearing axial and radial play is not allowed.) \n\n\t\t(5)\tInspect the rod end threads as follows: \n\n\t\t\t(i)\tWrap the spherical bearing and rod in suitable adhesive tape to prevent cleaning solvent and wet developer from entering the bearing (ref. figure No.2). \n\n\t\t\t(ii)\tRemove residual sealing compound, grease, and dirt from the threaded area of the rod ends. \n\n\t\t\t(iii)\tVisually inspect the threaded area of the rod ends for corrosion. \n\n\t\t\t(iv)\tInspect the threaded area of the rod ends for cracks by fluorescent magnetic particle or fluorescent dye penetrant inspection methods. \n\n\t\t(6)\tReplace before further flight any rod ends found to be cracked or corroded. Install serviceable parts. \n\n\tNOTE: Operators are asked to submit rod ends removed as a result of this AD to MBB, DepartmentLV52, together with the helicopter serial number, flight hours, and service time, if known. \n\n\t\t(7)\tInstall serviceable main rotor control rods in accordance with the applicable maintenance manual instructions. \n\n\tNOTE: MBB ASB No. BO-105-10-103 dated October 28, 1987, pertains to this inspection and rework. \n\n\t(c)\tRepeat the inspections of paragraphs (b)(3) through (b)(6) after each rod end bearing replacement. \n\n\t(d)\tReport pitch link cracks by the inspections of paragraph (b) of this AD to the Manager, Aircraft Certification Division, Federal Aviation Administration, Fort Worth, Texas 76193-0100 within 10 days of the inspection. Provide aircraft serial number, total time, and time since the last pitch link rework, if any. (Reporting is approved by the Office of Management and Budget under OMB No. 2120-1156.) \n\n\t(e)\tUpon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Manager, Aircraft Certification Division, Federal Aviation Administration, Southwest Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period if the request contains substantiating data to justify the increase for that operator. \n\n\t(f)\tAn alternate method of compliance which provides an equivalent level of safety with the requirements of this AD may be used when approved by the Manager, Aircraft Certification Division, Federal Aviation Administration, Fort Worth, Texas 76193-0100 or by the Manager, Aircraft Certification Staff, AEU-100, Federal Aviation Administration, Europe, Africa, and Middle East Office, c/o American Embassy, Brussels, Belgium, APO NY 09667. \n\n\t(g)\tIn accordance with FAR Section 21.197 and 21.199, flight is permitted to a base where the inspections required by this AD may be accomplished. \n\n\tThis amendment (39-6313) revises Amendment 39-5795 (52 FR 46991; December 11, 1987), AD 87-26-02, which became effective on December 31, 1987. \n\n\tThis amendment (39-6313, AD 87-26-02 R1) becomes effective on September 26, 1989. \n\n\n\n\n\t\t\t\tBEARING FRICTION MEASUREMENT \n\t\t\t\t\tFIGURE NO. 1 \n\t\t\t\t\tAD 87-26-02 R1 \n\n\n\n\n\n\t\t\t\tROD END WITH BEARING TAPED \n\n\t\t\t\t\tFIGURE NO. 2 \n\t\t\t\t AD 87-26-02 R1
2010-22-07: This amendment supersedes an existing airworthiness directive (AD) for the Eurocopter Deutschland GmbH (ECD) Model MBB BK 117 C-2 helicopters. This amendment 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 there was an in- flight incident in which a dynamic weight broke off the control lever leading to considerable vibrations. A visual inspection revealed that the threaded bolt of the control lever had broken off. The actions specified by this AD are intended to prevent separation of dynamic weights, severe vibration, and subsequent loss of control of the helicopter.
2010-22-09: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: This Airworthiness Directive (AD) is prompted due to an occurrence when an aircraft had a partial in-flight separation of the aileron outboard bearing support. The aileron outboard bearing supports are attached with two forward attachment bolts and two aft attachment bolts. The forward attachment bolts are approximately 3.2 mm (0.125 inch) longer than the aft attachment bolts. If the aileron outboard bearing supports have been removed, it is possible that during the reinstallation of the aileron outboard bearing supports, the attachment bolts can be installed in wrong positions. Bolts that are installed in wrong positions can damage the threads in the rear attachment anchor nuts. Such a condition, if left uncorrected, could lead to in-flight separation of the aileron outboard bearing support, and as a consequence, the loss or limited controllability of the aircraft. We are issuing this AD to require actions to correct the unsafe condition on these products.
90-26-02: 90-26-02 BOEING: Amendment 39-6829. Docket No. 90-NM-244-AD. \n\n\tApplicability: Model 747-400 series airplanes listed in Boeing Alert Service Bulletin 747-26A2170, dated July 5, 1990; and Model 767 series airplanes listed in Boeing Alert Service Bulletin 767-26A0068, Revision 1, dated September 18, 1990; certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent insufficient cargo fire protection, which could result in significant damage to the airplane in the event of a cargo compartment fire, accomplish the following: \n\n\tA.\tFor Model 747-400 series airplanes, within 30 days after October 1, 1990 (the effective date of Amendment 39-6736, AD 90-19-11), inspect the cargo compartment fire extinguisher orifices in accordance with Boeing Alert Service Bulletin 747-26A2170, dated July 5, 1990. If the part number does not match any number specified as the correct number in the service bulletin, prior to further flight withcargo in that compartment, replace with a correct part. \n\n\tB.\tFor Model 767-300 series airplanes listed in Boeing Alert Service Bulletin 767- 26A0068, dated July 5, 1990: \n\n\t\t1.\tWithin 30 days after October 1, 1990 (the effective date of Amendment 39-6736, AD 90-19-11), inspect the forward cargo compartment fire extinguisher orifices to determine the part number installed, in accordance with Boeing Alert Service Bulletin 767- 26A0068, dated July 5, 1990. If the part number does not match any number specified as the correct number in that service bulletin, prior to further flight with cargo in that compartment, replace with a correct part. \n\n\t\t2.\tWithin 30 days after the effective date of this amendment, inspect the aft cargo compartment fire extinguisher orifices to determine the part number installed, in accordance with Boeing Alert Service Bulletin 767-26A0068, Revision 1, dated September 18, 1990. If the part number does not match any number specified as the correct numberin that service bulletin, prior to further flight with cargo in that compartment, replace with a correct part. \n\n\tC.\tFor Model 767 series airplanes listed in Boeing Alert Service Bulletin 767- 26A0068, Revision 1, dated September 18, 1990, other than those that are subject to paragraph B. of this AD, within 30 days after the effective date of this amendment, inspect the cargo compartment fire extinguisher orifices to determine the part number installed, in accordance with that service bulletin. If the part number in the service bulletin does not match any number specified as the correct number in the service bulletin, prior to further flight with cargo in that compartment, replace with a correct part.\n\n\tD.\tWithin 10 days after the inspection required by paragraph A., B., or C. of this AD, if configuration discrepancies are discovered, submit a report of findings to the Manager, Seattle Manufacturing Inspection District Office, ANM-108S, FAA, Transport Airplane Directorate, 1601Lind Avenue S.W., Renton, Washington 98055-4056. The report must include the airplane serial number. \n\n\tE.\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\tF.\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 information from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. This information maybe examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tAirworthiness Directive 90-26-02 supersedes AD 90-19-11, Amendment 39-6736. \n\tThis amendment (39-6829, AD 90-26-02) becomes effective on December 27, 1990.
80-02-01 R2: 80-02-01 R2 BOEING: Amendment 39-3659 as amended by Amendment 39-3864 is further amended by Amendment 39-3969. Applies to Model 727 series airplanes certificated in all categories which have an operative Boeing designed auxiliary body fuel system installed. Compliance required within 600 flight hours from the effective date of this AD unless this initial inspection and test has been accomplished within the past 300 flight hours. To prevent failure of the fuel system and unwanted fuel transfer to the auxiliary body fuel tanks, accomplish the following: \n\n\tA.\tInspect and test the auxiliary body fuel tank installation as required and as outlined in Boeing Service Bulletin 727-28-A62 Revision 2 or later FAA approved revisions. If the shrouds are found defective, replace with a serviceable Boeing part or deactivate the system prior to further flight. \n\n\tB.\tRepeat the required inspections and tests every 1500 flight hours after the initial inspection unless the system has been deactivated. If shrouds are found defective, replace with a serviceable Boeing part, or deactivate the system prior to further flight.\n\n \tThe recurring inspection interval may be adjusted upon request of the operator. The FAA Maintenance Inspector, subject to prior approval by the Chief, Engineering and Manufacturing Branch, Northwest Region, may adjust the repetitive inspection interval specified in this AD if the request contains substantiating data to justify the increase. \n\n\tC.\tPrior to reactivating a previously deactivated auxiliary body fuel system, whether deactivated voluntarily or due to defects found in the auxiliary tank fuel system through inspections and tests accomplished in paragraphs A. and B. above, the auxiliary body fuel system must be repaired as required and shown to meet the requirements of paragraph A. and must be reinspected on a recurring basis per paragraph B. \n\n\tD.\tInstallation of the following Boeing final design part numbered refuel and vent line shrouds for the corresponding original and or interim design Boeing part number refuel and vent line shrouds will relieve the operator from compliance with the inspection and test requirements outlined in Paragraphs A, B and C above. This constitutes terminating action for the AD. \n\n\n\nOriginal Design\tPart Numbers\nInterim Design Part Numbers\nFinal Design Part Numbers \nBE-650-12-78-6*\n10-61707-43 \n10-61707-431 \n -528-6 \n -44\n -441 \n -223-6\n -45\n -451 \n -450-6\n -46\n -461 \n -334-6\n -47\n -471 \n -240-6\n -48\n -481 \n10-61707-8\t\t\t\n10-61707-49 \n10-61707-491 \n -9\n-50\n-501 \n-10\n-51\n-511\n-11\n-52\n-521\n-12\n-53\n-531\n-13\n-54\n-541\n-14\n-55\n-551\n-15\n-56\n-561\n-16\n-57\n-571\n-17\n-58\n-581\n-18\n-59\n-591\n-19\n-49\n-491\n-20\n-50\n-501\n-21\n-51\n-511\n-22\n-52\n-521\n-23\n-53\n-531\n-24\n-54\n-541\n-25*\n-55\n-551\n-26\n-56\n-561\n-27\n-57\n-571\n-28\n-58\n-581\n-29\n-59\n-591\n-30\n-60\n-601\n-31\n-61\n-611\n-32\n-62\n-621\n-33\n-63\n-631\n-34\n-64\n-641\n-41-71\n-711\nBE-650-12-78-6*\n-74\n-741\n10-61707-25*\t\n-75\n-751 \n\n\t*Originally, 10-61707-25 and BE-650-12-78-6 served as both a vent shroud and a fuel shroud (used separately). In the final design, they are identified separately, and the drawings so indicate. \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 and special tools from the manufacturer, may obtain same upon request to Boeing Commercial Airplane Company, Post Office Box 3707, Seattle, Washington 98124. The service bulletin may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tAmendment 39-3659 became effective January 22, 1980.\n\n \tAmendment 39-3864 became effective August 12, 1980. \n\tThis Amendment 39-3969 becomes effective November 20, 1980.
2010-22-01: We are superseding an existing airworthiness directive (AD) for the products listed above. That AD currently requires repetitive inspections for fatigue cracking and corrosion of the upper link fuse pin of the nacelle struts, and related investigative and corrective actions if necessary. The existing AD also provides terminating action for the repetitive inspections. This AD revises certain criteria for the terminating action. This AD was prompted by two reports of cracked upper link fuse pins. We are issuing this AD to prevent fatigue cracking or corrosion of the upper link fuse pin, which could result in failure of the fuse pin and consequent reduced structural integrity of the nacelle strut and possible separation of the strut and engine from the airplane during flight.
79-04-01 R3: 79-04-01 R3 BOEING: Amendment 39-3410 as amended by Amendment 39-3577 is further amended by Amendment 39-3889 and Amendment 39-4000. Applies to all Boeing Model 727 series airplanes certificated in all categories. Compliance required as indicated below. To prevent main landing gear-up landings as a result of structural failure in the lock system, accomplish the following: \n\n\tA.\tUnless already accomplished, within the next 1,500 landings after March 12, 1979, or prior to the accumulation of the threshold listed in the table below, whichever occurs later, replace the components listed in the table in accordance with Boeing Service Bulletins Nos. 727-32-211, Revision 4, or 727-32-237, Revision 2, or 727-32-286, Revision 1, or later FAA approved revisions, or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. As an alternate to replacement, the applicable components may be inspected for cracks in accordance with inspection methodsspecified in Boeing Service Bulletin No. 727-32-211, Revision 4, or 727-32-286, Revision 1, or later FAA approved revisions, or other methods approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region at the intervals specified in the table. Cracked parts must be replaced prior to further flight.\n \tB.\tUnless already accomplished, within the next 3,000 landings after March 12, 1979, replace (1) the left and right hand main gear manual extension gearbox horizontal supports, P/N 65-24575-1, with P/N 65-69156-1 and (2) the left and right hand main gear manual extension support yokes, P/N 65-26300- 1/-2 or 65-26300-7/-8, 65-81412-1/-2, with P/N 65-26300-21/-22 or 65-26300-23/-24 in accordance with the applicable procedures of Boeing Service Bulletin Numbers 727-32-164, Revision 2, and 727-32-204, Revision 1, or later FAA approved revisions, or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. Replacement of the gearbox housing is not required if the bushing/fastener configuration described in Boeing Service Bulletin No. 727-32-164, Revision 2, or later FAA approved revisions, or equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, is installed in each horizontal support attach hole for the gearbox housing. As an alternate to replacement of items (1) and (2) above, within the next 3,000 landings after March 12, 1979, and thereafter at intervals of 3,000 landings, eddy current inspect the horizontal supports and penetrant or eddy current inspect the support yokes in accordance with Boeing Service Bulletin No. 727-32-164, Revision 2, and Boeing Service Bulletin No. 727-32-204, Revision 1, or later FAA approved revisions, or in an equivalent manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. Cracked parts must be replaced prior to further flight. \n\tC.\t1.\tWithin the next 600 landings after October 9, 1979, unless accomplished within the last 900 landings preceding October 9, 1979, in accordance with the original issue of this AD, and thereafter at intervals not to exceed 1,500 landings, accomplish the rotational force tests on the main landing gear uplock hook assemblies (2 per airplane) in accordance with the procedures specified in Paragraph III of Boeing Service Bulletin No. 727-32-212, Revision 2, or later FAA approved revisions, for assembly P/Ns 65-24485-3, -4, and -6, and Boeing Service Bulletin No. 727-32-245, Revision 4, for assembly P/Ns 65-24485-7, or later FAA approved revisions, or alternate procedures approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. Rework and/or replace uplock hook assembly components as required to obtain the acceptable hook rotational forces specified in the service bulletins. \n\tC.\t2.\tOn or before July 1, 1982, accomplish the main landing gear uplock assembly modification specified in Boeing Service BulletinNo. 727-32-245, Revision 4, dated August 31, 1979, or later FAA approved revisions, or an alternate approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. This modification constitutes terminating action for the requirements of Paragraph C.1 above. \n\tD.\tUnless already accomplished, within the next 3,000 landings after March 12, 1979, accomplish the inspection/cleanup and replacement, as necessary, of the main landing gear manual extension support yoke attach bolts (2 per side) in accordance with Figures 1 and 2, as applicable, of Boeing Service Bulletin No. 727-32-251, or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. \n\tE.\tWithin the next 1,500 landings after March 12, 1979, or prior to the accumulation of 6,000 landings whichever occurs later, and thereafter at intervals not to exceed 1,500 landings, magnetic particle or eddy current inspect the main landing gear uplock lower shaft assemblies, P/Ns 65-24489-3 and 65- 43772-1/-2 in accordance with Boeing Service Bulletin No. 727-32-257, Revision 1, or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. \n\t\tCracked parts must be replaced with serviceable parts prior to further flight. Terminating action consists of replacement of the uplock lower shaft assembly with an improved shaft in accordance with Figure 2 of Boeing Service Bulletin No. 727-32-257, Revision 1, or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. \n\tF.\tAccomplish one of the following: \n\t\t1.\tOn or before July 1, 1982, install the main landing gear safety bar mechanism, LH and RH sides, in accordance with Boeing Service Bulletin No. 727-32-275, dated March 28, 1980, or later FAA approved revisions, or an equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. If Boeing Service Bulletin Numbers 727-32-237, Revision2, dated March 9, 1979; 727-32-251 dated March 11, 1977; and 727-32-257, Revision 1, dated July 21, 1978; or later FAA approved revisions, or equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, have been accomplished, the improved safety bar mechanism or FAA approved equivalent need not be installed until December 31, 1983. \n\t\t2.\tOn or before July 1, 1982, replace the main landing gear lock system components with improved components in accordance with Boeing Service Bulletin No. 727-32-237, Revision 2, dated March 9, 1979, or later FAA approved revisions, or equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, and thereafter overhaul and maintain the main landing gear lock system in accordance with a supplemental overhaul/maintenance program acceptable to the assigned FAA maintenance inspector and approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. (Recurring action taken under the overhaul/maintenance program shall be recorded to show the status of compliance with this AD.) \n\tAccomplishment of Paragraph F.1. constitutes terminating action to Paragraphs A, D, and E above. If Paragraph F.2. is accomplished, the requirements of F.2. supersede Paragraphs A, D, and E. \n\tG.\tOn or before July 1, 1982, (1) replace the main landing gear manual extension system gearbox horizontal supports, LH and RH sides, Boeing P/N 65-24575-1, with Boeing P/N 65-69156-1 in accordance with Boeing Service Bulletin No. 727-32-164, Revision 2, dated June 22, 1979, or later FAA approved revisions, or an equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, (2) replace the main landing gear manual extension system support yokes, LH and RH sides, Boeing P/Ns 65-26300-1/-2, 65-26300-7/-8, 65-81412-1/-2, 65-26300-11/-12, and 65-26300- 17/-18 with Boeing P/Ns 65-26300-21/-22 or 65-26300-23/-24 in accordance with Boeing Service Bulletin No. 727-32-204, Revision 3, dated December 7, 1979, or later FAA approved revisions, or an equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, and (3) replace the main landing gear manual extension system gearbox housing, Boeing P/N 65-27485-1/-2, LH and RH sides, with Boeing P/N 65-27485-11/-12 in accordance with Boeing Service Bulletin No. 727-32-279, dated June 22, 1979, or later FAA approved revisions, or an equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. The replacement accomplished per this paragraph constitutes terminating action to Paragraph B of this AD. \n\tH.\tAirplanes may be ferried to a maintenance base for replacement of parts in accordance with FAR 21.197. \n\tI.\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Seattle Area Aircraft Certification Office, FAA Northwest 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 that operator.\n \tJ.\tFor the purpose of this AD, when conclusive records are not available to show the total number of landings accumulated by a particular part (or assembly), the number of landings may be computed by dividing the airplane time-in-service since the part (or assembly) was installed in the airplane by the operator's fleet average time per flight for his Model 727 series airplanes. \n\tFor those operators who are unable to determine the total time-in-service or landings accumulated on a part (or assembly) since new, the total number of landings of the airplane on which it is installed may be used. This applies only to the establishment of the total landings for the initial compliance threshold. \n\tNOTE: The overhaul/maintenance program specified in ATA Report 32-30-1, Revision 4, dated September 9, 1980, is an FAA approved program for the purpose of complying with the requirements of paragraph F.2. of this AD. \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\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. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tAmendment 39-3410 became effective March 12, 1979. \n\tAmendment 39-3577 became effective October 9, 1979. \n\tAmendment 39-3889 became effective September 25, 1980. \n\tThis amendment 39-4000 becomes effective January 25, 1981. \n\n\n\nComponent\nPart No.\nReplacement or Initial Inspection Threshold\t (Landings)\t\t\nRepeat Inspection Interval Not To Exceed (Landings)\n\n\n\n\nDownlock Torque Shaft\n65-78698-1,-2\n35,000\n3,000\n\n-5,-6\n\n\n\n-7,-8\n\n\n\n\n\n\nDownlock Rod Assy.\n69-20527-2 \n\n\n\n69-33654-1\n12,000\n1\n\n69-33654-2\n\n\n\n69-33654-3\n\n\n\n69-33654-4\n\n\n\n69-33654-5\n35,000\n3,000\n\n\n\n\nDownlock Torque Tube \nAssy. \n65-26921-17\n37,000\n3,000\n\n65-26921-18\n\n\n\n\n\n\nUplock\tUniversal Block\n65-24488-1\n10,000 flt hrs or 4 years, whichever occurs first.\n1,500 \n\n-4\n48,000 landings\n3,000\n\n\n\n\nUplock\tUniversal Bolt\nNAS 1106-44D\n11,000\nTo be replaced \n\n69-47743-1\n\n\n\n\n\n\nUplock\tUpper Shaft Assy.\n65-25851-1\n46,000\n3,000 \n\n65-25851-2\n\n\n\n65-25851-5\n\n\n\n65-25851-6\n\n\n\n\n\n\nUplock Lower Crank\n65-49325-1,-2\n\n1,500 \n\n65-49325-5,-6\n3,000\n\n\n65-49325-7,-8\n\n\n\n\n\n\nDownlock Crank 2\n69-25028-1\n\n\n\n-2\n\n\n\n-3\n25,000\n1,500\n\n-4\n\n\n\n\n\t\t\n\n\t\t\n\n1\tWithin 1500 landings from the last inspection or within 1500 landings from October 9, 1979, whichever occurs first,thereafter not to exceed 500 landings. \n\n2\tInitially inspect or replace prior to accumulation of 25,000 landings or within 750 landings from the effective date of this amendment, whichever occurs later. Replacement with steel downlock crank, Boeing Part Number 69-20528-5 or -6 or an equivalent approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, terminates the downlock crank inspection requirement of this table. \n\n\n\nComponent\nPart No.\nReplacement or Initial Inspection Threshold (Landings)\nRepeat Inspection Interval Not To Exceed (Landings) \nUplock\tLower Crank Bolt \nNAS 1105-28\t\t\n7,000\nTo be replaced\n\nBACP18T5-( )\t\t\n7,000 \n\n\nMS20392-5C( )\t\t\n(oversize option) \n12,000\n\n\nMS20392-6C( )\t\t \n(oversize option) \n50,000\n\n\n\n\n\nDownlock Rod Bolt (Inboard)\nNAS 1105-13DW\n20,000\nTo be replaced \n\nBACB30LJ5U13\t\tOptional Bolts:\nBACB30LJ5U15 \nBACB30LJ5-16 \nBACB30GE5-16 \n20,000\nTo be replaced \n\n\n\n\nDownlock Rod Bolt (Outboard)\nNAS 1105-13DW\t\t\t\t\t\t\n20,000\nTo be replaced \n\nBACB30LM5DU12\n20,000\nTo be replaced\n\nOptional Bolts: \nBACB30NE5D12 \nBACB30GE5D12 \nNAS 1305-12D
95-24-11: This amendment adopts a new airworthiness directive (AD) that applies to certain Fairchild Aircraft SA226 and SA227 series airplanes. This action requires installing foreign object damage (FOD) barriers in the floorboards of the cockpit between the pedestal and floor from Fuselage Station (FS) 79.38 to FS 88.06 and on the outboard forward edge of the left-hand and right-hand cockpit forward floorboards at FS 79.38. Two incidents of objects falling through openings in the cockpit floor and jamming the elevator controls and the yoke prompted this action. The actions specified by this AD are intended to prevent airplane flight control jammings caused by objects falling through the cockpit floor openings.
80-25-08: 80-25-08 BOEING VERTOL (VERTOL): Amendment 39-3987. Applies to Vertol Model 107-II helicopters certificated in all categories. Compliance required as indicated. To prevent fatigue failure of the main rotor tension-torsion strap assemblies, remove from service tension-torsion strap assemblies Part No. 107R2003-1 upon the accumulation of 27,800 hours in service and replace with an airworthy part that meets the requirement of this AD. This amendment is effective December 12, 1980.
2021-25-07: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model BD-100-1A10 airplanes. This AD was prompted by a discovery that a lockwire may not have been installed on the side stay actuator pin nut of the main landing gear (MLG). This AD requires inspecting the left-hand and right-hand MLG side stay actuator assembly pin nut for the presence of a lockwire, and installing a lockwire if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
73-16-06: 73-16-06 PITTS: Amdt. 39-1699. Applies to S-2A airplanes, Serial Numbers 2001 through 2059 certificated in all categories. Compliance required as indicated. To prevent the possibility of a power loss due to the failure of the induction air box flapper door, accomplish the following: (a) Within the next 50 hours time in service after the effective date of this airworthiness directive, unless already accomplished, install stiffener, Pitts Part Number 2-7112- 16, to the induction air box flapper door assembly, Pitts Part Number 2-7112-7, in accordance with Pitts Service Letter No. 2. Access to the induction air box flapper door assembly may be gained by removing the lower engine cowl. (b) Until the modification specified in paragraph (a) above has been accomplished, visually check the induction air box flapper assembly, Pitts Part Number 2-7112-7, for cracks in the welded areas upon the accumulation of 50 hours time in service or within 5 hours time in service, whichever is later, and thereafter at intervals not to exceed 5 hours time in service from the last check. The checks required by this AD may be performed by the pilot. If the flapper door in the area of the weld is found to be cracked, remove the flapper door assembly from the alternate air box, and add stiffener, Pitts Part Number 2-7112-16 in accordance with Pitts Service Letter No. 2, before further flight. (c) The installation of the stiffener, Pitts Part Number 2-7112-16, in accordance with Pitts Service Letter No. 2, will eliminate the necessity for the repetitive inspections required in paragraph (b). Pitts Service Letter No. 2, dated June 26, 1973, pertains to this subject. This amendment becomes effective August 13, 1973.
95-24-01: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas DC-10-10 series airplanes, that requires inspections of the wings to detect cracks in the aft spar lower cap, in certain stringer butterfly clips on the bulkheads, and in certain fastener holes; and repair, if necessary. This amendment also requires modification of those areas of the wings, which terminates the repetitive inspection requirements. This amendment is prompted by reports indicating that, during fatigue testing of the wing structure, cracks developed in the aft spar lower cap, in certain stringer butterfly clips, and in certain fastener holes due to fatigue-related stress. The actions specified by this AD are intended to prevent such fatigue- related cracking, which could lead to the failure of the aft spar cap and consequently could reduce the structural integrity of the wing.
74-23-02: 74-23-02 BELL: Amendment 39-1999. Applies to Bell Model 204B and 205A-1 helicopters certificated in all categories. Compliance required as indicated. To detect cracks in the tail rotor pitch change link segments and to prevent possible failure of the tail rotor pitch change chains accomplish the following repetitive inspections on chains, P/N 205-001-721-1 and 205-001-748-1. (a) Within five hours time in service after June 7, 1974, unless already accomplished, and thereafter at intervals not to exceed 25 hours time in service from the last inspection, accomplish the following inspections: (1) Remove the chain or chains from the helicopter in accordance with the applicable maintenance manual. (2) Clean each chain with solvent and stiff bristle brush and air dry. (3) Roll each chain into a tight flat disc and inspect the links' outer segments, both sides, for cracks, using a light and a ten-power or higher magnifying glass. Unroll the chain and withteeth up inspect inner segments of the links at the radius between the teeth. Turn each chain over and inspect the opposite side of the link segments. (4) Replace chains with cracked segments before further flight. (5) Install chains with uncracked link segments in accordance with the pertinent maintenance manual and rig tail rotor controls in accordance with the pertinent maintenance manual. If 212-010-701 Tail Rotor Hub and Blade Assembly is installed on the Model 205A-1, rig the controls in accordance with the maintenance manual and as specified in Bell Helicopter Company Service Bulletin No. 205-05-74-1, Rev. A, dated June 24, 1974, or later approved revision. (b) Before the first flight of each day after June 7, 1974, accomplish the following repetitive inspections. (1) Remove the cover from the chain assembly. (2) Inspect each chain assembly for cracks in the link segments using a three-power or higher magnifying glass. Particular attention should be placed on the portion of each chain that travels over each sprocket and that extends three inches each side of this area or portion. (3) Remove chains with cracked segments before further flight in accordance with the applicable maintenance manual. (4) Install chains with uncracked segments in accordance with the applicable maintenance manual and rig the controls as specified in paragraph (a)(5) of this airworthiness directive. (c) Replace chains, P/N 205-001-721-1 having manufacturing dates of June 20, 1974, or later etched on the clevis, as follows: (1) Replace chains with more that 140 hours total time in service on the effective date of this Airworthiness Directive within ten hours time in service. (2) Replace chains with less than 140 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 150 hours total time in service. (d) Replace chains, P/N 205-001-721-1 having manufacturing dates prior to June 20, 1974, as follows: (1) Replace chains with more than 40 hours total time in service on the effective date of this Airworthiness Directive within ten hours time in service. (2) Replace chains with less than 40 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 50 hours total time in service. (e) Replace chains, P/N 205-001-748-1, with more than 140 hours total time in service on the effective date of this Airworthiness Directive within ten hours total time in service. Replace chains, P/N 205-001-748-1, with less than 140 hours total time in service on the effective date of this Airworthiness Directive prior to attaining 150 hours total time in service. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from themanufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Company, P. O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this Airworthiness Directive which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D. C. and at the FAA Southwest Regional Office, Fort Worth, Texas. (Bell telefax messages dated May 10, 1974; June 15, 1974; August 2, 1974; and September 25, 1974, pertain to this subject, and Information Letter to all Bell 204B/205A-1 operators, 35:WJD:jge-3152 dated August 12, 1974, also pertains to this subject.) This supersedes Amendment 39-1963 (39 F.R. 34054), A.D. 74-20-06. This amendment becomes effective November 3, 1974.
82-02-03: 82-02-03 BEECH: Amendment 39-4313. Applies to Model 76 (Serial numbers ME-1 through ME-435) airplanes certificated in any category. COMPLIANCE: Required as indicated unless already accomplished. To ensure the integrity of the elevator control cable and determine proper cable routing, accomplish the following: A) Prior to further flight, accomplish all of the following: 1. Remove the large inspection panel located in the center of the bottom skin just aft of the Station 68 frame. 2. Drill a 3/8-inch diameter inspection hole in the forward flange to which the inspection panel was attached as follows: Center the hole laterally on the elevator cable pulley and fore and aft on the flange. Use a thin bucking bar or wood block between the flange and the pulley and cable to prevent damage to the pulley and/or cable when drilling the inspection hole. NOTE: Airplane Serial Numbers ME-418 and after have the 3/8-inch diameter inspection hole already drilledin the forward flange. 3. Deburr the inspection hole and visually inspect the elevator down cable for broken or frayed strands and ensure the cable is on the pulley under all three guard pins. 4. If the cable is riding over any of the guard pins, replace the P/N NAS 427K12 guard pins over which the cable was routed, replace the P/N NAS 314-25-1411 cable, and rig in accordance with the Beech Model 76 Maintenance Manual. 5. Reinstall the inspection panel, and record compliance with this AD by an appropriate entry in the airplane maintenance records. B) Within 48 hours, report misrouted cables or other defects found as a result of any inspection required herein to the FAA via a Malfunction or Defect (M or D) Report (FAA Form 8010-4) or a letter to the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209. Describe the defect found, total time-in-service on the airplane or part at time of discovery, and the aircraft serial number. (Reporting approved by the Office of Management and Budget under OMB No. 04-R0174). C) Airplanes may be flown in accordance with FAR 21.197 to a location where the provisions of Paragraph A) of this AD can be performed provided the following is accomplished: 1. Remove the large inspection panel located in the center of the bottom skin just aft of the Station 68 frame. 2. Visually inspect the elevator down cable for broken or frayed strands while the elevator control column is slowly moved fore and aft. Pay particular attention to the cable strands near the elevator cable pulley. 3. If no frayed or broken cable strands are found, reinstall the inspection panel. D) Any equivalent method of compliance with this AD must be approved by the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 269-7000. Beechcraft Safety Communique 76-62 pertains to the subject matter of this AD. This amendment becomes effective February 14, 1982, to all persons except those to whom it has already been made effective by a priority mail letter from the FAA dated January 8, 1982.
90-02-18 R1: 90-02-18 R1 SOCATA GROUPE AEROSPATIALE: Amendment 39-6454 as revised by Amendment 39-6619. Docket No. 89-CE-37-AD. Applicability: Models TB 9 and TB 10, TB 20, and TB 21 (all serial numbers (S/N)) airplanes certificated in any category. Compliance: As indicated in the body of the AD, unless already accomplished per AD 90-02-18. To preclude loss of power due to contamination of the fuel system, accomplish the following: (a) Within the next 75 hours time-in-service (TIS) from February 6, 1990, except as indicated in paragraph (c) of this AD, modify the fuel system by the installation of the following applicable SOCATA modification kit, as described in SOCATA Service Bulletin (SB) Number 48/2, dated March 1990: Airplanes Kit Number All TB airplanes (S/N 1 through 822, 850 through 887, 889 and subsequent) 9154 Airplanes Kit Number TB 20 (S/N 823 through 849, and 888) 9155 (b) For Models TB 20 and TB 21 (S/N 1 through 730)airplanes (unless modified with SOCATA Modification Number 66), within the next 75 hours TIS from February 6, 1990, except as indicated in paragraph (d) of this AD, modify the fuel system by replacement of the installed Dukes fuel pump with a Weldon fuel pump and the addition of a check valve, in accordance with the instructions contained in SOCATA SB Number 47/1, dated October 1989. (c) If the required parts are not available to accomplish the modification specified in paragraph (a) of this AD, the airplane may continue operation for an additional 150 hours TIS after the compliance time specified in paragraph (a) of this AD provided the fuel tank sump is drained at intervals not to exceed each 50 hours TIS in accordance with the procedures specified in the DESCRIPTION section of SOCATA SB 48/2, dated March 1990. (d) If the required parts are not available to accomplish the modifications specified in paragraph (b) of this AD, the airplane may continue operation for an additional 150 hours TIS after the compliance time specified in paragraph (b) of this AD provided the following preflight actions are accomplished by the pilot prior to each engine start: (i) Select battery (main switch) ON. (ii) Advance the mixture control to FULL RICH. (iii) Select electric fuel boost pump ON. (iv) Advance the throttle until a positive fuel flow is observed on the fuel flow gauge, then retard the throttle and move the mixture control to IDLE/CUTOFF. (v) Select electric fuel boost pump OFF. (vi) Select battery (main switch) OFF. (vii) Visually inspect the electric fuel boost pump area for leaks. (viii) If no positive fuel flow is observed on the fuel flow gauge, or fuel leaks are detected from the electric fuel boost pump, repair or replace the defective component prior to further flight. NOTE: Avoid moving the propeller and standing in the propeller area while inspecting the engine. (e) Airplanes may be flown in accordance withFAR 21.197 to a location where this AD may be accomplished. (f) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety may be approved by the Manager, Brussels Aircraft Certification Office, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, B-1000 Brussels, Belgium. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to SOCATA Groupe Aerospatiale, B.P. 38, 65001 Tarbes, Cedex, France; Telephone 62.51.73.00, or 62.51.73.55 (for Telefax); or the Product Support Manager, U.S., AEROSPATIALE, 2701 Forum Drive, Grand Prairie, Texas 73053; Telephone (214) 641-3614; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This AD revises AD 90-02-18, Amendment 39-6454. This amendment (39-6619, AD 90-12-18 R1) becomes effective on June 13, 1990.
86-14-08: 86-14-08 BRITISH AEROSPACE: Amendment 39-5347. Applies to Model BAe 125- 800A series airplanes listed in BAe 125 Service Bulletin 27-136-(3059A), Revision 1, dated June 26, 1985, certificated in any category. To prevent loss of stall warning, accomplish the following within the next 60 days after the effective date of this AD, unless previously accomplished: A. Incorporate a new layshaft assembly in the stall identification system in accordance with the accomplishment instructions of British Aerospace 125 Service Bulletin 27- 136-(3059A), Revision 1, dated June 24, 1985. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. All persons affected by this directive, who have not already received the appropriate service document from the manufacturer, may obtain copies upon request to British Aerospace, Inc., Librarian, Box 17414, Dulles International Airport, Washington, D.C. 20041. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective August 4, 1986.
86-20-07: 86-20-07 MCDONNELL DOUGLAS HELICOPTER COMPANY (Hughes Helicopters, Inc.): Amendment 39-5422. Applies to Model 369, 369A, 369D, 369E, 369H, 369HE, 369HM, and 369HS helicopters, including military Models YOH-6A and OH-6A, certificated in any category, equipped with tail rotor drive shaft flexible couplings, Part Number (P/N) 369A5501 or 369H92564. Compliance required as indicated unless already accomplished. To prevent failure of the tail rotor (T/R) drive shaft system and subsequent loss of T/R control, accomplish the following: (a) Within 100 hours' time in service after the effective date of this AD, install aft coupling failsafe device (P/N's 369D25530 bolt and 369D25531 socket) in accordance with Part I of the applicable Service Information Notices (SIN) DN-143, HN-2O6, or EN-31, each dated August 26, 1986. Installation of the failsafe device on military Models YOH-6A or OH-6A helicopters in civil use shall be accomplished in accordance with Part I of SIN HN-206. NOTE: The failsafe device required by paragraph (a) will be installed before delivery on all applicable Model 369E helicopters, Serial Number 0135E, and subsequent. (b) Within 100 hours' time in service after the effective date of this AD, install forward coupling failsafe device (P/N's 369D25530 bolt and 369D25531 socket) in accordance with Part I of SIN DN-95, dated August 7, 1981, or Part III, HN-173, dated November 2, 1981, as applicable. Installation of the coupling failsafe device on military Models YOH-6A or OH-6A helicopters shall be accomplished in accordance with Part III of SIR HN-173. (c) For all helicopters with tail rotor driveshaft flexible coupling failsafe devices installed, the T/R drive shaft forward and aft flexible couplings shall be checked as follows: (1) (At Each Preflight Check: Check for T/R backlash or looseness by rocking the T/R back and forth in its plane of rotation. The blade should not move in excess of 0.75 inch (1.93cm) atthe blade tip without rotation of the main rotor blades. (2) At Each Aircraft/Engine Shutdown: If thumping or rapping is heard from the T/R drive train during final revolutions of the T/R, check the T/R to assure that the T/R blade does not move in excess of 0.75 inch (1.93cm) at the blade tip without rotation of main rotor blades. (d) The checks required by this AD may be performed by the pilot and must be recorded in accordance with FAR Section 91.173. (e) If during the checks required by paragraph (c), the tail rotor blade tip movement exceeds the specified limits, prior to further flight, inspect and replace, as necessary, either or both fore and aft tail rotor drive shaft couplings. (f) Rotorcraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the modifications and inspections of paragraphs (a) and (b) of this AD can be accomplished. (g) An alternate method of compliance which provides an equivalent level of safety may be approved by the Manager, Western Aircraft Certification Office, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009-2007. The procedure shall be done in accordance with applicable parts of MDHC SIN's DN- 143, HN-206, EN-31, all dated August 26, 1986; MDHC SIN DN-95, dated August 7, 1981; MDHC SIN HN-173, dated November 2, 1981. The incorporation by reference was approved by the Director of the FEDERAL REGISTER in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from McDonnell Douglas Helicopter Company, Centinela Avenue and Teal Street, Culver City, California 90230. These documents may be examined at the Office of the Regional Counsel, Federal Aviation Administration, Southwest Region, Room 158, Building 3B, 4400 Blue Mound Road, Fort Worth, Texas 76101, the Western Aircraft Certification Office, 15000 Aviation Boulevard, Hawthorne, California, or the Office of the FEDERAL REGISTER, 1100 L Street, NW., Room 8401,Washington, D.C. This amendment supersedes Amendment 39-4186 (46 FR 40868), AD 81-17-02, as amended by Amendment 39-4221 (46 FR 46566), AD 81-17-02R1. This amendment becomes effective October 24, 1986.
2010-21-09: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: A damaged fuel heater caused a fuel leakage in the engine nacelle; investigation revealed that the damage to the fuel heater was due to chafing with an oil cooling system hose. Piaggio Aero Industries (PAI) issued Service Bulletin (SB) 80- 0175, which was applicable to all aeroplanes and contained instructions for a repetitive inspection of the affected parts and, if necessary, their replacement and/or for the repositioning of oil/ fuel tubing if minimum clearances were not found. We are issuing this AD to require actions to correct the unsafe condition on these products.