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92-17-05:
92-17-05 DASSAULT AVIATION: Amendment 39-8334. Docket No. 92-NM-25-AD.
Applicability: Model Fan Jet Falcon Basic D, E, and F series airplanes; and Model Mystere- Falcon 20-C5, D5, E5, and F5 airplanes; certificated in any category.
Compliance: Required as indicated, unless accomplished previously.
To prevent reduced structural integrity of these airplanes, accomplish the following:
(a) Incorporate a revision into the FAA-approved maintenance inspection program that provides for inspection of the Significant Structural Items defined in Dassault Aviation Service Bulletin FJF-00-26 (FJF-730), Revision 1, dated December 12, 1990, at the later of the times specified in subparagraph (a)(1) or (a)(2):
(1) Prior to the accumulation of 20,000 landings or 30,000 hours time-in-service, whichever occurs first; or
(2) Within 6 months after the effective date of this AD.
(b) Report the results, positive or negative, of each inspection required by paragraph (a) of this AD to Dassault Aviation, in accordance with the instructions in Dassault Aviation Service Bulletin FJF-00-26 (FJF-730), Revision 1, dated December 12, 1990. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (44 U.S.C. 3501 et seq.) and have been assigned OMB Control Number 2120-0056.
(c) Cracked structures detected during the inspections required by paragraph (a) of this AD must be repaired or replaced, prior to further flight, in accordance with the instructions in Dassault Aviation Service Bulletin FJF-00-26 (FJF-730), Revision 1, dated December 12, 1990, or in accordance with other data meeting the certification basis of the airplane which is approved by the FAA or by the French Direction G n rale de l'Aviation Civile (DGAC).
(d) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch.
NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch.
(e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.
(f) The inspections, repairs, and replacement shall be done in accordance with Dassault Aviation Service Bulletin FJF-00-26 (FJF-730), Revision 1, dated December 12, 1990. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained fromFalcon Jet Corporation, Customer Support Department, Teterboro Airport, Teterboro, New Jersey 07608. 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., 7th Floor, Suite 700, Washington, DC.
(g) This amendment becomes effective on September 28, 1992.
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2018-25-09:
We are adopting a new airworthiness directive (AD) for all CFM International S.A. (CFM) LEAP-1B21, -1B23, -1B25, -1B27, -1B28, - 1B28B1, -1B28B2, -1B28B2C, -1B28B3, -1B28BBJ1, and -1B28BBJ2 turbofan engines. This AD requires removing certain electronic engine control (EEC) system operation (OPS) and engine health monitoring (EHM) software and installing versions eligible for installation. This AD was prompted by six aborted takeoffs on the similarly designed CFM LEAP-1A model turbofan engine after those engines did not advance to the desired takeoff fan speed due to icing in the pressure sensor line. We are issuing this AD to address the unsafe condition on these products.
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99-26-20 L:
99-26-20 MD HELICOPTERS INC.: Docket No. 99-SW-89-AD. \n\n\tApplicability: Model MD-900 helicopters, with Main Rotor Upper Hub (hub) Assembly, part number (P/N) 900R2101006-105 or 900R2101006-107, installed, certificated in any category. \n\n\tNOTE 1: This AD applies to each helicopter identified in the preceding applicability provision, regardless of whether it has been otherwise modified, altered, or repaired in the area subject to the requirements of this AD. For helicopters that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (c) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent failure of the hub assembly, loss of drive to the main rotor, and subsequent loss of control of the helicopter, accomplish the following: \n\n\t(a)\tFor the hub assembly, P/N 900R2101006-107, \n\n\t\t(1)\tWithin 6 hours time-in-service (TIS) or before flight after January 31, 2000, whichever occurs first, visually inspect the hub assembly drive plate attach flange (flange) for a crack and determine the torque of each flange attach nut (nut) in accordance with the Accomplishment Instructions, Part I, paragraph 2.A., steps (1) through (7) of MD Helicopter Inc. Service Bulletin SB900-072, dated December 10, 1999 (SB). If a crack is found, before further flight, remove and replace the hub assembly with an airworthy hub assembly. \n\n\t\t(2)\tWithin 25 hours TIS or before flight after January 31, 2000, whichever occurs first, conduct the Accomplishment Instructions, Part II, paragraph 2.B., steps (1) through (6), (8),and (9) of the SB. If a crack is found, before further flight, remove and replace the hub assembly with an airworthy hub assembly. \n\n\t(b)\tFor the hub assembly, P/N 900R2101006-105, \n\n\t\t(1)\tWithin 6 hours TIS or before flight after January 31, 2000, whichever occurs first, visually inspect the flange for a crack and determine the torque of each nut in accordance with the Accomplishment Instructions, Part I, paragraph 2.A., steps (1) through (7) of the SB. \n\n\tNOTE 2: The SB effectivity does not include hub assembly, P/N 900R2101006-105; however, for the requirements of this AD, certain provisions of the SB do apply to this P/N . \n\n\t\t(2)\tIf any nut has less than 180 inch pounds (20.34 Nm) of torque, before further flight, remove the hub assembly, disassemble, inspect, and reassemble in accordance with the procedures in paragraph (b)(4) of this AD. If a crack is detected, before further flight, remove and replace the hub assembly with an airworthy hub assembly. \n\n\t\t(3)\tWithin25 hours TIS or before flight after January 31, 2000, whichever occurs first, remove and visually inspect the hub assembly as follows: \n\n\t\t\t(i)\tIf present, remove sealant from the drive plate attachment to the hub assembly. \n\n\t\t\t(ii)\tMark the main rotor hub holes, bolts, and nuts to correspond with the drive plate hole numbers. (See Figure 1.) \n\n\t\t\t(iii)\tRemove the main rotor drive plate (drive plate) assembly and anti-fretting ring (fretting buffer). \n\n\t\t\t(iv)\tInspect drive plate to rotor hub assembly mating surfaces and the fretting buffer for fretting. \n\n\t\t\t(v)\tUsing paint stripper (C313 or equivalent) and cleaning solvent (C420 or equivalent), remove the paint from the upper mating surface of the hub assembly to enable an accurate visual inspection of the drive plate attachment bolt hole (bolt hole) area for cracking (Figure 1). Ensure the paint stripper and solvent DO NOT contaminate the upper bearing and upper grease seal areas. \n\n\t\t\t(vi)\tUsing a 10-power or highermagnifying glass and bright light, inspect the mating surface area and the area around and inside the 10 bolt holes of the hub assembly for a crack. If a crack is found, prior to further flight, replace the hub assembly with an airworthy hub assembly. \n\n\t\t\t(vii)\tIf no crack is found, remove fretting from the mating surfaces of the hub assembly and the drive plate assembly, reassemble, fillet seal (C211 or equivalent) the surface of the drive plate to fretting buffer to hub assembly mating lines, and seal all exposed unpainted upper surfaces of the hub assembly. \n\n\t\t\t(viii)\tReinstall the main rotor drive plate using 10 new sets of replacement attachment hardware. Torque the nuts to 160 inch pounds above locknut locking/run-on torque in the sequence shown (Figure 1). Record in the rotorcraft logbook, or equivalent record, the locknut locking/run-on torque for each nut. \n\n\t\t\t(ix)\tAfter the next flight, verify that the torque on each of the 10 nuts is at least 160 inch pounds above the locknut locking/run-on torque (minimum torque). Retorque as required without loosening nuts \n\n\t\t\t(x)\tThereafter, at intervals of at least 4 hours TIS, not to exceed 6 hours TIS, verify that the torque of each of the 10 nuts is at least the minimum torque. Retorque as required without loosening nuts. This torque verification is no longer required after the torque on each of the 10 nuts has stabilized at the minimum torque 160 inch pounds for each nut during two successive torque verifications. \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, FAA. Operators shall submit their requests through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles Aircraft Certification Office. \n\n\tNOTE 3: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Los Angeles Aircraft Certification Office. \n\n\t(d)\tIf any nut torque is below minimum torque and no hub assembly crack is found before disassembly inspection, after retorque in accordance with the applicable Maintenance Manual, a special flight permit for one flight below 100 knots indicated airspeed may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the helicopter to a location where the requirements of this AD can be accomplished. \n\n\t(e)\tCopies of the applicable service information may be obtained from MD Helicopters Inc., Attn: Customer Support Division, 5000 E. McDowell Rd., Mail Stop M615-GO48, Mesa, Arizona 85215-9797, telephone 1-800-388-3378 or 480-891-6342, datafax 480-891-6782. This information may be examined at the FAA, Office of the Regional Counsel, Southwest Region, 2601 Meacham Blvd., Room 663, Fort Worth, Texas. \n\n\t(f)\tEmergency Priority Letter AD 99-26-20, issued December 17, 1999, becomes effective upon receipt. \n\n\n\t\t\t\t\tFIGURE 1\n \t\t\t\t\t 99-26-20
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2006-06-13:
The FAA is adopting a new airworthiness directive (AD) for certain Airbus Model A330-200 and A330-300 series airplanes; and Model A340-200 and -300 series airplanes. This AD requires repetitive detailed inspections for discrepancies of the inboard and outboard actuator fittings of the aileron servo controls, corrective actions if necessary, and eventual replacement of all the attachment bolts of the aileron servo controls. This AD results from several cases of bushing migration on the inboard and outboard actuator fittings of the aileron servo controls; in one case the bushing had migrated completely out of the actuator fitting and the fitting was cracked. We are issuing this AD to prevent rupture of the inboard and outboard actuator fittings of the aileron servo controls, which could result in airframe vibration and consequent reduced structural integrity of the airplane.
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59-13-01:
59-13-01 ALLISON: Applies to Models 501-D13 and -D13A Engines.
Compliance required as indicated.
Seven recent cases of spalling of the propeller shaft roller bearing on Allison 501-D13 engines necessitates a continuity check using a probe be made of the reduction gear magnetic plug at least every 10 hours on engine installations not modified in accordance with Allison Commercial Engine Bulletin 72-74 and Lockheed Bulletin SB 262. This check is necessary due to high bearing loads caused by propeller during certain aircraft flight conditions. If metal chips are found as described in Allison Maintenance Manual, section 72-0, paragraph 8B, page 219, remove gearbox.
This supersedes AD 59-11-01.
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89-16-06:
89-16-06 BRITISH AEROSPACE (BAe) PLC: Amendment 39-6266.
Applicability: Jetstream Model 3101 (all serial numbers) airplanes which have Kit 3279A embodied as part of Omnibus Modification 7380, certificated in any category.
Compliance: Required within the next 100 hours time-in-service after the effective date of this AD, unless already accomplished.
To assure operation of the airplane within the design airspeed limitations, accomplish the following:
(a) Modify the pilot's and copilot's operating limitations placards and revise the Airplane Flight Manual in accordance with British Aerospace Alert Service Bulletin 11-A-JA880140, dated February 23, 1988.
(b) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished.
(c) An equivalent means of compliance with this AD may be used if approved by the Manager, Aircraft Certification Office, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, B-1000 Brussels, Belgium.
All persons affected by this directive may obtain copies of the document(s) referred to herein upon request to British Aerospace, Inc., Technical Librarian, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041; or may examine these documents at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106.
This amendment (39-6266, AD 89-16-06) becomes effective on August 29, 1989.
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2018-23-51:
We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 737-8 and -9 airplanes. This emergency AD was sent previously to all known U.S. owners and operators of these airplanes. This AD requires revising certificate limitations and operating procedures of the airplane flight manual (AFM) to provide the flight crew with runaway horizontal stabilizer trim procedures to follow under certain conditions. This AD was prompted by analysis performed by the manufacturer showing that if an erroneously high single angle of attack (AOA) sensor input is received by the flight control system, there is a potential for repeated nose-down trim commands of the horizontal stabilizer. We are issuing this AD to address the unsafe condition on these products.
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98-15-18:
This amendment supersedes Airworthiness Directive (AD) 95-26-18, which currently requires inspecting (one-time) certain wing lift struts for internal corrosion on certain Maule Aerospace Technology Corp. (Maule) M-4, M-5, M-6, M-7, MX-7, and MXT-7 series airplanes and Models MT-7-235 and M-8-235 airplanes, and replacing any wing lift strut where corrosion is found. That AD was the result of an accident where the wing separated from one of the affected airplanes. This AD retains the initial inspection and possible replacement requirements of AD 95-26-18, requires the inspections to be repetitive, and provides the option of using ultrasonic procedures to accomplish the inspection requirements. The actions specified by this AD are intended to prevent failure of the wing lift struts caused by corrosion damage, which could eventually result in the wing separating from the airplane. \n\n\tThe incorporation by reference of Maule Service Bulletin No. 11, dated October 30, 1995, as listed in the regulations was previously approved by the Director of the Federal Register as of January 26, 1996 (61 FR 623, January 9, 1996).
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84-26-03:
84-26-03 CFM INTERNATIONAL: Amendment 39-4976. Applies to all Model CFM 56-2 turbofan engines incorporating fuel injector P/N 9984M90G20.
Compliance is required as indicated unless already accomplished.
To prevent flameouts in engines equipped with fuel injectors without heat shields, accomplish the following:
Replace all fuel injectors, P/N 9984M90G20 with P/N 9984M90G28 or P/N 9984M90G14 fuel injectors in accordance with CFM56 SB 73-021, or approved fuel nozzles in accordance with SBs 73-024 and 73-034, or FAA approved equivalent.
(a) For engines being operated with JP-4 fuel (or equivalent) or with a mixture containing JP-4, compliance is required within 30 days of the effective date of this AD.
(b) For engines which are operated only with Jet A fuel (or equivalent), compliance is required by January 31, 1986.
Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations (FARs) 21.197 and 21.199 to a base where the AD canbe accomplished.
Upon request, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Engine Certification Office, Aircraft Certification Division, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803.
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 this document from the manufacturer may obtain copies upon request to CFM International, Neumann Way, Cincinnati, Ohio 45215. This document also may be examined at the Office of the Regional Counsel, FAA, New England Region, Attn: Rules Docket No. 83-ANE-30, 12 New England Executive Park, Burlington, Massachusetts 01803, weekdays, except Federal Holidays, between 8:00 a.m. and 4:30 p.m.
This amendment becomes effective on February 22, 1985.
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75-23-08 R5:
75-23-08 R5 CESSNA: Amendment 39-2419 as amended by Amendment 39-2687, 39-2830, 39-3298, and 39-3609, are further amended by Amendment 39-5451. Applies to T310, 320, 340, 401, and 411 series airplanes and Models 402, 402A, 402B, 414, 421, 421A, 421B and 421C (S/N 421C0001 through 421C0683) airplanes, certificated in any category.\n\n\tThis AD is issued to combine in one document inspections and parts replacements required by previous ADs and require modifications now available which will increase the reliability of the exhaust systems on the affected airplanes.\n\n\tCompliance: Required as indicated, unless previously accomplished.\n\n\tTo detect incipient failure and improve reliability of the engine exhaust systems installed on the above-noted airplanes, accomplish the following:\n\n\tI.\tRepetitive General Exhaust System Inspection\n\n\t\tA.\tOn all new and in-service airplanes listed above, visually inspect the exhaust system components per procedures hereafter specified in accordance with the following schedule:\n\n\t\t\t50 HOUR INSPECTION\n\n\t\t\tInspect components not listed in Table II or III within 50 hours time-in-service after the effective date of this AD or last inspection per AD 75-04-01 and within each 50 hours time-in-service thereafter.\n\n\t\t\t100 HOUR INSPECTION\n\n\t\t\tInspect components listed in Table II within 100 hours time-in-service after the effective date of this AD or last inspection in accordance with this AD and within each 100 hours time-in-service thereafter.\n\n\t\t\tEXEMPT COMPONENTS\n\n\t\t\tComponents listed in Table III are exempt from these inspection requirements.\n\n\t\t1.\tCLEANING - In order to properly inspect the exhaust system, components must be clean and free of oil, grease, etc. If required, clean as follows:\n\n\t\t\ta.\tSpray engine exhaust components with a suitable solvent (such as Stoddard Solvent), allow to drain, and then wipe dry with a clean cloth. \n\nWARNING \nNEVER USE HIGHLY FLAMMABLE SOLVENTS ON ENGINE EXHAUST SYSTEMS. \nNEVERUSE A WIRE BRUSH OR ABRASIVES TO CLEAN EXHAUST SYSTEMS OR MARK ON THE SYSTEM WITH LEAD PENCILS.\n\n\t\t\tb.\tRemove the heat shields from the turbocharger in accordance with heat shield removal procedures in the Service Manual.\n\n\t\t\tc.\tRemove shields around the exhaust bellows or slip joints, multi-segment "V" band clamps at joints, and other items which might hinder inspection of the system.\n\nNOTE \nDo not remove clamps.\n\n\t\t2.\tVISUAL INSPECTION OF COMPLETE SYSTEM \n\nNOTE \nConduct this inspection when the engine is cool\n\n\t\t\ta.\tVisually inspect exhaust stacks for burned areas, cracks and looseness. Insure that attach bolts are properly torqued in accordance with the aircraft Service Manual.\n\nNOTE\n\tDuring this inspection give special attention to condition of bellows and welded areas along seams and around bellows, and weld seams around exhaust system components.\n\n\t\t\tb.\tVisually inspect the flexible connection between the waste-gate and overboard duct (when applicable) for cracks and security.\n\n\t\t\tc.\tCorrect any major leaks in all areas of the exhaust system except waste-gate to tailpipe flexible coupling (when applicable) by replacement of defective parts or by repair in accordance with Part 43 of the FARs.\n\n\t\t\td.\tVisually inspect exhaust joint springs for correct compression. If the joint is disturbed or if springs are obviously loose, proceed with the following inspection. (See Figure 1).\n\n\n\n\t\t\t\ti.\tBefore removal of exhaust joint springs, measure installed length of each spring. Those compressed to less than 0.45 inch must be replaced.\n\n\t\t\t\tii.\tRemove all springs and measure free length. Springs having a free length of less than 0.57 inch must be replaced.\n\nNOTE\n\tAdd AN960-10 washers under head of joint bolts as required to obtain correct dimension. During installation, joint bolts should be tightened gradually and spring length checked frequently to prevent over compression of springs.\n\n\t\t\t\tiii.\tReinstall springs and measure installed length. Length must be 0.51 plus .00, -.03 inch.\n\n\t\t\te.\tIf installed, visually inspect the slipjoint(s) for bulges beyond normal manufacturing irregularities of .03 inches and/or cracks. If any bulges and/or cracks are present, replace bulged or cracked slipjoint(s) (refer to Service Manual). See Figure 2.\n\n\n\nNOTE\n\tThis inspection was previously required on Model 421B airplanes, S/N 421B0001 through 421B0935 by AD 75-17-38 effective August 22, 1975.\n\n\t\t\tf.\tInspection of multi-segment "V" band clamp(s) (Between engine and turbocharger).\n\n\t\t\t\ti.\tUsing crocus cloth, clean the outer band of the multi-segment "V" band clamp(s). Pay particular attention to the spot weld area on the clamp(s).\n\n\t\t\t\tii.\tWith clamp(s) properly torqued:\n\n\t\t\t\t\t(1)\tVisually inspect the outer band in the area of the spot weld for cracks (see Figure 3). If cracks are found, replace the clamp(s) with new multi-segment "V" band clamp(s).\n\n\t\t\t\t\t(2)\tVisually inspect the corner radii of clamp innersegments for cracks (See figure 3). This inspection requires careful use of artificial light and inspection mirrors.\n\n\n\n\nNOTE\n\tWhen replacement is required, install the new multi-segment "V" band clamp over the exhaust flanges and torque to the correct value (see Torque Value Chart, Table 1 to this AD). As the clamp is tightened, lightly tap it circumferentially in a radial direction with a rawhide or soft plastic mallet.\n\n\t\t\t\t\t(3)\tVisually inspect flatness of the outer band, especially within two inches of the spot welded tabs which retain the T-bolt fastener. This can be done by placing a straight edge across the flat part of the outer band as shown in Figure 4, then checking the gap between the straight edge and the outer band. This gap should be less than 0.062 inch. If deformation exceeds 0.062 inch limit, replace clamp(s) with new multi-segment clamp(s). (See Note preceding this step).\n\n\n\n\t\t\tg.\tOne-piece "V" band clamp inspection (Overboard exhaust to turbocharger).\n\n\t\t\tVisually inspect with a light and mirror, the clamp surfaces adjacent to the intersection of the "V" apex and bolt clips, and the entire length of the "V" apex of the clamp for signs of cracks or fractures. If cracks or fractures are visible, replace the clamp. (See Figure 5).\n\nNOTE\n\tThe above inspections, except for Paragraph e. and f., were previously required by AD 75-04-01 effective February 11, 1975. Inspections per Paragraph e. were previously required by AD 75-17-38 effective August 22, 1975. Inspections per Paragraph f. were previously required by AD 75-04-01 as amended by Amendment 39-2338 effective August 15, 1975.\n\n\tTABLE I\nEXHAUST COUPLING APPLICABILITY CHART\n\n\n\n\n\n\n\n\n\n\n\n\n\n\n\nAIRCRAFT APPLICABILITY\n\n\n\n\n\n\n\n\n\n\n\n\n\nSingle Piece "V" Coupling\nCoupling\nPart Number\nQty \nPer\nAcft\nTorque\nIn-Lbs\nT310P\nQ, R\n320\n320A,\nB,C\n320D\nE, F\n340\n340A\n401/402\nA, B\n414\n421\nA, B\n421\nC\nLocation\nLife Limited\nNH1000897-20\n4\n40*\n\n\n\n\nX\n\nX\n\n\nCollector (Wye)\nInlet\nNo\nNH1000897-30\n2\n40\n**\n\n\n\n\n\n\n\nX\n\nWastegate\nInlet\nNo\nNH1000897-40\n or\nV57A4234\n or\n41195AA423\n2\n40\n*\nX\n\n\nX\n\nX\n\n\n\nTurbine\nOutlet\nNo\nNH1000897-50\n or\nV57A5019\n or\n41195AA502\n2\n40\n*\n\n\n\n\nX\n\nX\nX\nX\nTurbine\nOutlet\nNo\n\n\n\n\n\n\n\n\n\n\n\n\n\nMulti-Segment "V" Band Clamps\n51394H250\n(51134-2505\nGasket)\n3\n35\n\nX\n\n\n\n\n\n\n\nCollector\nInlet\nYes\n838037\nAlternates:\n4309AL\n4309AF\n3\n35\n\n\nX\n\n\n\n\n\n\nCollector\nInlet\nYes\nMVT64832\nAlternates:\n4309AL\n4309F\n3\n35\nX\n\n\nX\n\nX\n\n\n\nCollector\nInlet\nYes\nMVT68892-250\nAlternate\n4301BT250\n4\n45\n\n\n\n\nX\n\nX\n\n\nCollector (Wye)\nInlet\nYes\n4256AB200\nAlternate\nMVT68892-200\n2\n45-50\n\n\n\n\nX\n\nX\nX\n\nWastegate Inlet 340A (414) & Wastegate Exit (421)\nYes\n\n4356AA300\n\n2\n\n70-90\n\n\n\n\n\n\n\n\nX\n\nWastegate Inlet (421)\nYes\n\n\t* Initial installation of single piece couplings must be madein strict adherence to the specified 40 inch-pounds torque value. However, periodic retorquing is required only if coupling shows torque values below 30 inch-pounds.\n\n\t**The 24096-300-N gasket is not required at the exhaust wastegate inlet of 421, 421A, and 421B aircraft when the NH 1000897-30 Single-piece coupling is installed.\n\nNOTES FOR TABLE II \nGENERAL NOTES:\n\t0850XXX and 5654XXX dash numbers shown in parentheses are earlier part numbers for the corresponding 9910XXX dash numbers..\n\n\tBall joint attaching parts (springs, bolts, nuts, washers and cotter pins) are subject to the 100-hour inspection instead of the 50-hour, only if installed in combination with the components listed in Table II.\n\n\t1.\tComponents incorporate formed sheet metal coupling flanges.\n\n\t2.\tComponents incorporate external safety ring.\n\n\t3.\tComponents incorporate internal safety sleeve.\n\n\t4.\tComponents require seals at coupling flange.\n\n\t5.\tAlternate for MVT68892-250 and 4301BT250 couplings.6.\tNH1000897-30 supersedes 4356AA300. Gasket P/N 24096-300-N is not required with the NH1000897-30 clamp.\n\n\t7.\tExhaust stack assemblies are sealed type, and the part numbers shown include risers for both cylinder banks and the crossover pipes.\n\nTABLE II\n100 HOUR INSPECTION EXHAUST SYSTEM COMPONENTS\n\n\nDescription &\nPart Number\nQty\nPer Eng\nT310 P,Q,R,\n320\nD,E,F\n340\n340A\n401/402\nA,B\n414\n421\nA,B\n421\nC\nNotes\nExhaust Stack Assy\n9910379-1 L.H. Eng\n(Inconel 601)\n\n1 (LH)\n\n\n\n\n\n\n\nX\n\n\n7\n9910379-2 R.H. Eng\n(Inconel 601)\n1 (RH)\n\n\n\n\n\n\nX\n\n7\n9910295-11 L.H.\n(5155184-1)\n1\n\n\n\n\n\n\n\nX\n\n9910295-12 R.H.\n(5155184-2)\n1\n\n\n\n\n\n\n\nX\n\nAft Slip Joints\n9910314-1\n(Inconel 601)\n\n1\n\n\n\n\n\n\n\nX\n\n\nAft Elbows\n9910299-3 L.H.\n(5654551-5)\n\n1\n\n\n\nX\n\nX\n\n\nX\n\n\n\n1 & 4\n9910299-4 R.H.\n(5654551-6)\n1\n\n\nX\nX\n\nX\n\n\n1 & 4\n9910301-1\n(0850712-39)\n1 L.H.\nX\nX\n\n\nX\n\n\n\n1\n9910301-3\n(0850712-41)\n1 R.H.\nX\nX\n\n\nX1\n9910301-5\n(0850712-40)\n1 L.H.\nX\nX\n\n\nX\n\n\n\n1\n9910379-19\n(Inconel 601)\n1 L.H.\n\n\n\n\n\n\nX\n\n\n9910379-20\n(Inconel 601)\n1 R.H.\n\n\n\n\n\n\nX\n\n\nWye Collector Assy\n9910299-8\n(Inconel 601)\n\n1\n\n\n\n\nX\n\n\n\n\n\n1,3, & 4\n9910301-4\n(0850732-3)\n1 L.H.\nX\nX\n\n\nX\n\n\n\n1 & 2\n9910301-6\n(0850732-18)\n1 R.H.\nX\nX\n\n\nX\n\n\n\n1 & 2\n9910341-1\n(Inconel 601)\n1\n\n\nX\n\n\n\n\n\n1,3, & 4\n9910300-6\n(5155100-1)\n1\n\n\n\n\n\n\n\nX\n\nWastegate \nInlet Elbow\n9910299-5\n(5654551-4)\n\n1\n\n\n\n\nX\n\n\nX\n\n\n\nCouplings \n (One Piece)\nWastegate Inlet\nNH1000897-30\n\n1\n\n\n\n\n\n\n\nX\n\n\n6\nWye Collector Inlet\nNH1000897-20\n2\n\n\nX\nX\n\nX\n\n\n5\n\nNOTES FOR TABLE III \nGENERAL NOTES:\n\t5155XXX, 5355XXX, and 5654XXX dash numbers shown in parentheses are earlier part numbers for the corresponding 9910XXX dash numbers.\n\n\tBall joint attaching parts (springs, bolts, nuts, washers and cotter pins) are exempt from mandatory inspections only if installed in combination with the components listed in Table III.\n\n\t1.\tExhaust stack assemblies are seal-less type, and part numbers shown are complete bank assemblies.\n\n\t2.\tOn aircraft prior to 1979 Models, the initial replacement must include the aft elbow, wye collector assemblies and clamps since they are not interchangeable with earlier components.\n\n\t3.\tAttach 5155156-5 Elbow to overboard stack using one each 5155157-1 Flex Joint and two each U84C200 SH Clamps.\n\n\tII.\tParts Replacement\n\n\t\tA.\tReplacement of multi-segment "V" band clamps between aft engine cylinders and turbocharger inlet.\n\n\t\t\tOn all airplanes listed above within 50 hours time in service after February 11, 1975, for those clamps having more than 350 hours time in service as of that date or prior to 400 hours time in service for those clamps having less than 350 hours time in service as of that date, and at or prior to each additional 400 hours time in service thereafter, replace existing multi-segment "V"band exhaust system clamps located between aft engine cylinders and the turbocharger inlet (except for waste-gate to exhaust overboard pipe clamp on 421 airplanes) with new parts having Cessna part numbers in accordance with Attachment 4 to Cessna Service Letter ME-75-17 dated July 14, 1975, or later revisions. Use aircraft total time for clamp time in service unless aircraft maintenance records establish location and time in service on previously replaced clamps. See Figure 3 for multi-segment type clamp configuration.\n\nNOTE\n\t\tThe above clamp replacement was previously required by AD 75-04-01 effective February 11, 1975.\n\n\t\tB.\tReplacement of multi-segment turbocharger to overboard tail pipe clamps.\n\n\t\t\tWithin 200 hours time in service after August 15, 1975, replace presently installed multi-segment type turbocharger to overboard tailpipe clamps on Model T310, 320, 401, 402, 340, 414 and 421 airplanes with the clamps specified below:\n\nTABLE III\nEXEMPT EXHAUST SYSTEM COMPONENTS\n\n\nDescription &\nPart Number\nQty\nPer Eng\nT310 P,Q,R,\n320\nD,E,F\n340\n340A\n401/402\nA,B\n414\n421\nA,B\n421\nC\nNotes\nExhaust Stack Assy\n9910295-13 L.H. \n(Inconel 601)\nor\n9910295-9 L.H.\n(5155166-1)\n\n\n\n1\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\n\n\n1\n9910295-14 R.H.\n(Inconel 601)\nor\n9910295-10 R.H.\n(5155166-2)\n\n\n1\n\n\nX\n\n\nX\n\n\nX\n\n\nX\n\n\nX\n\n\nX\n\n\n\n\n1\n9910295-15\n(Inconel 601)\n1\n\n\n\n\n\n\n\nX\n\n9910295-16\n(Inconel 601)\n1\n\n\n\n\n\n\n\nX\n\nAft Slip Joints\n9910296-2\n(Inconel 601)\nor\n9910296-1\n(5355108-4)\n(Stainless Steel)\n\n\n\n2\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\n\nX\n\nAft Elbows\n(Inconel 601 mat'l, seal-less style machined coupling, flanges & internal safety sleeves)\n9910299-15 L.H.\n\n\n\n\n\n\n1\n\n\n\n\n\n\n\n\nX\n\n\n\n\n\n\nX\n\n\n\n\n\n\n\nX\n\n\n\n\n\n\n\n\n2\n9910299-16 R.H.\n1\n\n\nX\nX\n\nX\n\n\n2\n9910301-14\n1 L.H.\nX\nX\n\n\nX\n\n\n\n2\n9910301-16\n1 L.H.\nX\nXX\n\n\n\n2\n9910301-17\n1 R.H.\nX\nX\n\n\nX\n\n\n\n2\nWye Collector Assy\n(Inconel 601 mat'l & seal-less style maintained coupling flanges)\n99102099-9\n\n\n\n\n\n1\n\n\n\n\n\n\n\n\nX\n\n\n\n\n\n\nX\n\n\n\n\n\n\n\n2\n9910301-15\n1 L.H.\nX\nX\n\n\nX\n\n\n\n2\n9910301-18\n1 R.H.\nX\nX\n\n\nX\n\n\n\n2\n9910341-2\n1\n\n\nX\n\n\n\n\n\n2\n9910300-7\n1\n\n\n\n\n\n\n\nX\n\n\n\nDescription &\nPart Number\nQty\nPer Eng\nT310 P,Q,R,\n320\nD,E,F\n340\n340A\n401/402\nA,B\n414\n421\nA,B\n421\nC\nNotes\nOverboard Exhaust Stack Assembly\n0850711-33 & -31\n\n\n1 L.H.\n\n\nX\n\n\nX\n\n\n\n\nX\n\n\n\n\n0850711-34, -40, -42\n1 R.H.\nX\nX\n\n\nX\n\n\n\n\n9910299-1\n(5654551-7)\n1 L.H.\n\n\nX\nX\n\nX\n\n\n\n9910299-2\n(5654551-8)\n1 R.H.\n\n\nX\nX\n\nX\n\n\n\n5155156-3\n1 L.H.\n\n\n\n\n\n\nX\n\n\n5155156-4\n1 R.H.\n\n\n\n\n\n\nX\n\n\n9910300-1\n(5155100-7)\n1 L.H.\n\n\n\n\n\n\n\nX\n\n9910300-2\n(5155100-8)\n1 R.H.\n\n\n\n\n\n\n\nX\n\nWastegate Inlet Elbow\n9910299-10\n\n1\n\n\n\n\nX\n\n\nX\n\n\n\nWastegateOverboard Pipe\n0850713-1\n\n\n1\n\n\nX\n\n\nX\n\n\n\n\nX\n\n\n\n\n5355100-64\n1\n\n\nX\n\n\n\n\n\n\n9910299-6\n(5654551-3)\n1\n\n\n\nX\n\nX\n\n\n\nWastegate Outlet Elbow\n5155156-5\n\n\n1\n\n\n\n\n\n\n\n\nX\n\n\n\n3\n9910300-3\n(5155100-20)\n1\n\n\n\n\n\n\n\nX\n\nTurbo Shield\n0850902-1\n\n1\n\nX\n\nX\n\n\n\nX\n\n\n\n\n5155154-3\n1\n\n\n\n\n\n\nX\n\n\n5354005-1\n1\n\n\nX\nX\n\nX\n\n\n\nCouplings \n (One Piece)\n\nWye Collector Inlet \nNH1000897-60\n2 L.H.\n(All)\n2 R.H.\n(340,414)\n1 R.H.\n(T310, 320 & \n401/402)\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\nX\n\n\n\n\n\n2\nWastegate Inlet\nNH1000897-70\n1\n\n\n\nX\n\nX\n\n\n2\nTurbine Outlet\nNH1000897-40\n1\nX\nX\n\n\nX\n\n\n\n\nTurbine Outlet\nNH1000897-50\n1\n\n\nX\nX\n\nX\nX\nX\n\n\n\t\t\tOn T310 (all aircraft through S/N 310Q0600), 320D (all aircraft), 320E (all aircraft), 320F (all aircraft), 401 (all aircraft), and 402 (all aircraft through S/N 402B0300 series airplanes, replace the existing multisegment "V" band turbocharger to overboard tailpipe clamp with a new Part Number V57A4234 or 41195AA423 clamp. These new clamps are not "life limited".\n\n\t\t\tOn 340 (all aircraft through S/N 3400150, 414 (all aircraft through S/N 4140350), and 421 (all aircraft through 421B0396) series airplanes, replace the existing multi-segment "V" band turbocharger to overboard tailpipe clamp with a Part Number V57A5019 or 41195AA502 clamp. These new clamps are not "life limited".\n\nNOTE\n\t\t\tThe above clamp replacement was previously required by AD 75-04-01 as amended by Amendment 39-2338 effective August 15, 1975.\n\n\tIII.\tSpecial Inspections\n\n\t\tA.\tUnless previously accomplished, within 25 hours time in service after August 15, 1975, the inspection described below is to be conducted on the following aircraft on which the one-piece "V" band tail pipe clamps have been installed. (Clamps bearing a Cessna inspection stamp on the back side (a large C with a number) and installed on airplanes exempted from this inspectionon Attachment No. 2 of Cessna Service Letter ME-75-17 have been inspected by the manufacturer and comply with this requirement).\n\n\t\t\tT310 (All aircraft S/N 310P0001 through 310R0275)\n\t\t\t320D, 320E, 320F (All aircraft)\n\t\t\t340 (All aircraft through S/N 3400550)\n\t\t\t401 (All aircraft)\n\t\t\t402 (All aircraft S/N 4020001 through 402B0912)\n\t\t\t414 (All aircraft through S/N 4140648)\n\t\t\t421 (All aircraft S/N 4210001 through 421B0927)\n\n\t\t\t1.\tRemove upper cowling from LH and RH engine.\n\n\t\t\t\ta.\t(Aircraft Turbo 310, 320, 401, 402, and 421 series equipped with metal heat shield). Detach heat shield from turbocharger turbine housing by cutting safety wire and removing hinge pin located at top of shield assembly securing shield halves together.\n\n\t\t\t\tb.\tThe heat shield does not need to be removed on 320, 340 or 414 and any aircraft equipped with insulation blanket shield.\n\n\t\t\t2.\tClean and visually inspect with a light and mirror, the clamp surfaces adjacent to the intersection of the "V"apex and bolt clips (see Figure 5) for signs of cracks or fractures. If cracks or fractures are visible, replace clamp. (This step allows inspection of the clamp under tension).\n\n\n\n\t\t\t3.\tRemove nut, washer, and bolt from clamp, unseat coupling, and slide clamp down exhaust stack to gain adequate visibility. \n\nCAUTION: DO NOT REMOVE EXPANSION LIMITER.\n\n\t\t\t4.\tVisually inspect for cracks and fractures the entire length of the "V" apex of the clamp (see Figure 5), and pay particular attention to the surfaces adjacent to the intersection of the "V" apex and bolt clips.\n\n\t\t\t\tNOTE: The use of artificial light and at least a 10-power magnifying glass is necessary to detect minute cracks or signs of incipient failure. \n\n\t\t\t\tDo not use dye penetrant inspection procedures since noncritical metal forming folds yield misleading failure indications.\n\n\t\t\t5.\tIf clamp is found to be defective, replace clamp.\n\n\t\t\t6.\tIf clamp has no defects, steel stamp or using a vibrating pencil, mark an "X" on the back side of metal torque tag.\n\n\t\t\t7.\tReinstall clamp on turbocharger turbine housing and overboard exhaust stack and install bolt, washer and nut. Torque nut to 40 inch-pounds. \n\nCAUTION: DO NOT EXCEED NUT TORQUE BY MORE THAN 5 INCH-POUNDS.\n\n\t\t\t8.\tResecure heat shield on turbocharger turbine housing when detached, by aligning shield hinge halves and reinstalling pin previously removed. Secure lower halves of shield using two pieces of .032 monel safety wire twisted together, through opposing eyelets of shield halves, draw shield snug on turbine housing, and safety by twisting end of wire together. Insure heat shield is snug on housing.\n\n\t\t\t9.\tReinstall upper cowling on LH and RH engine.\n\nNOTE\n\t\t\tThis inspection was previously required by AD 75-04-01 as amended by Amendment 39-2338 effective August 15, 1975. Previous inspection per this AD or accomplished by the manufacturer satisfies the above requirement (IIIA).\n\n\tIV.\tModifications\n\n\t\tA.\tOn the following aircraft within 50 hours time in service after the effective date of this AD inspect/modify the overboard exhaust stack flange (turbocharger and exhaust stack interface) in accordance with the instructions below.\n\n\t\t\tT310P0001 through T310R0230, T310R0232 through T310R0254, T310R0256, T310R0257, T310R0259 through T310R0263, T310R0265, T310R0266, T310R0268 through T310R0272, T310R0274, T310R0275, 320D0001 through 320F0045, 3400001 through 3400550, 4010001 through 401B0221, 4020001 through 402B0630, 402B0632 through 402B0909, 402B0911, 402B0912, 4140001 through 4140643, 4140645 through 4140648, 4210001 through 421B0889, 421B0891 through 421B0918, 421B0921, 421B0922, 421B0924, 421B0926 and 421B0927.\n\n\t\t\t1.\tRemove any existing overboard support clamps, and remove overboard exhaust stack. \n\nNOTE\n\t\t\tTake precautions to prevent collection of any dirt or foreign particles in opening of turbocharger turbine housing during modification.\n\n\t\t\t2.\tVisually inspect the mating surfacesof flanges of overboard exhaust stack and turbochargers for uniform contact over their entire perimeter.\n\n\t\t\t3.\tIf such contact is not evident, remove excess material (shaded area) of overboard exhaust stack as required, and as shown typical in Figure 6, by grinding, to allow mating surfaces of flanges of overboard exhaust stack and turbine housing to have uniform contact over their entire perimeter. Exercise care to prevent damage to flange surface of overboard exhaust stack when grinding off excess material.\n\n\n\nFIGURE 6\n\n\t\t\t4.\tReinstall overboard exhaust stack and secure clamps in accordance with turbine outlet exhaust stack one-piece "V" band clamp inspection.\n\nNOTE\n\t\t\tThis one-time inspection/modification may be performed in conjunction with the one-piece "V" bank clamp 25-hour inspection and redundant tail pipe support installation required by IV. B. below.\n\n\t\tB.\tOn the following aircraft, prior to but not later than December 31, 1975, install redundant tail pipe support clamps and brackets in accordance with Cessna Service kits specified below.\n\n\t\t\tT310 (All aircraft through S/N 310Q0600)\n\t\t\t320D0001 through 320F0045\n\t\t\t3400001 through 3400150\n\t\t\t4010001 through 401B0300\n\t\t\t4020001 through 402B0300\n\t\t\t4140001 through 4140350\n\t\t\t4210001 through 421B0300\n\n\t\t\t1.\tKit Number and Applicability\n\n\t\t\t\ta.\tSK402-31 - Turbo 310P0001 through Turbo 310Q0600\n\t\t\t\t\t320D0001 through 320F0045\n\t\t\t\t\t4010001 through 401B0300\n\t\t\t\t\t4020001 through 402B0300\n\n\t\t\t\tb.\tSK414-9-3400001 through 3400150\n\t\t\t\t\t4140001 through 4140350\n\n\t\t\t\tc.\tSK421-67-4210001 through 421B0300\n\n\tV.\tAny equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region.\n\n\tVI. The time in service between the repetitive inspections required herein may be adjusted up to plus 10 per cent of any specified inspection interval required by this AD to facilitate accomplishing these inspections concurrent with other scheduled maintenance on the airplane.\n\n\tCessna Service Letter ME75-17 dated July 14, 1975, Cessna Service Letter ME77-1 dated January 24, 1977, and Supplement 2 thereto dated June 26, 1978, and Cessna Multiengine Service Information Letter ME79-32 dated August 17, 1979, pertain to this subject.\n\n\tAmendment 39-2419 supersedes AD 75-04-01, Amendments 39-2083, 39-2130, and 39-2338, and AD 75-17-38, Amendment 39-2341.\n\n\tAmendment 39-2419 became effective November 17, 1975.\n\n\tAmendment 39-2687 became effective August 12, 1976.\n\n\tAmendment 39-2830 became effective February 17, 1977.\n\n\tAmendment 39-3298 became effective September 25, 1978.\n\n\tAmendment 39-3609 became effective November 13, 1979.\n\n\tThis amendment, 39-5451, becomes effective on November 4, 1986.
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88-03-01:
88-03-01 PRATT & WHITNEY CANADA: Amendment 39-6032. Final copy of priority letter AD issued February 2, 1988. Applies to Pratt & Whitney Canada (PWC) PT6B-36A turboshaft engines incorporating the third stage compressor stator assembly, Part Number (P/N) 3109163-01, installed in Sikorsky Aircraft S-76B helicopters. The third stage compressor stator assembly, P/N 3109163-01 is incorporated in engines, Serial Number PC-E36043 and subsequent, and those engines incorporating PWC Service Bulletin (SB) 11022, dated September 28, 1987.
Compliance is required as indicated, unless already accomplished.
To prevent failure of the third stage compressor stator vane assembly which could result in loss of engine power or an inflight engine shutdown, accomplish the following:
(a) Prior to further flight, accomplish the following:
(1) Attach a placard which reads "avoid N1 ranges, 78-80.5 percent, 95.6-97.6 percent", in full view of the pilot and copilot, as close as practicable to the N1 speed indicators.
(2) Place a copy of this AD in the appropriate rotorcraft flight manual.
(3) Operate the rotorcraft in accordance with these speed range restrictions.
(b) For engines which have accumulated 25 hours or less total time in service (TIS) from the effective date of this AD, accomplish the following:
(1) Install vibration damping device, Part Number (P/N) 3112128-01, on vane assembly, P/N 3109163-01, in accordance with PWC Alert SB A-11033, dated December 31, 1987, prior to accumulating 25 hours total TIS.
(2) Remove vane assembly, P/N 3109163-01, prior to accumulating 600 hours total TIS, and replace with a serviceable part.
(c) For engines which have accumulated more than 25 hours total TIS from the effective date of this AD, accomplish the following:
(1) Install vibration damping device, P/N 3112128-01, on vane assembly, P/N 3109163-01, in accordance with PWC Alert SB A-11033, dated December 31, 1987, within 25 hours TIS from the effectivedate of this AD.
(2) Remove vane assembly, P/N 3109163-01, within 100 hours TIS from the installation of damping device, P/N 3112128-01, and replace with a serviceable part.
(d) Thereafter, remove from service vane assembly, P/N 3109163-01, prior to accumulating 600 hours TIS since new, and replace with a serviceable part.
NOTE: When vane assembly, P/N 3109163-01, is installed as a serviceable part, it must include damping device, P/N 3112128-01.
(e) Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations 21.197 and 21.199 to a base where the AD can be accomplished.
(f) Upon request from an operator, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Engine Certification Office, Engine & Propeller Directorate, Aircraft Certification Service.
(g) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, the Manager, Engine Certification Office, Engine & Propeller Directorate, Aircraft Certification Service, may adjust the compliance time specified in this AD.
PWC Alert SB 11033, dated December 31, 1987, and PWC SB 11022, dated September 28, 1987, identified and described in this document, are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Pratt & Whitney Canada, Box 10, Longueuil, Quebec, Canada J4K 4X9. These documents may also be examined at the Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, Rules Docket No. 88-ANE-06, 12 New England Executive Park, Burlington, Massachusetts 01803, Room 311.
This amendment, 39-6032, becomes effective November 29, 1988, as to all persons except those persons to whom it was made immediately effective by individual priority letter AD 88-03-01, issued February 2, 1988, which contained this amendment.
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2018-25-07:
We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model BD-700-1A10 and BD-700-1A11 airplanes. This AD was prompted by reports of drainage holes on the belly fairing forward and middle access panels being obstructed with sealant. This AD requires inspecting for and removing all sealant blocking the drainage holes on the belly fairing forward and middle access panels. We are issuing this AD to address the unsafe condition on these products.
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86-14-05:
86-14-05 SHORT BROTHERS, LTD.: Amendment 39-5344. Applies to Model SD3-60 airplanes, serial numbers SH3601 through SH3676 inclusive, certificated in any category. Compliance is required within 90 days after the effective date of this AD. To maintain the structural integrity of the horizontal stabilizer, accomplish the following, unless previously accomplished:
1. Modify the horizontal stabilizer lower skin to spar attachment in accordance with Short Brothers, Ltd., Service Bulletin SD36-55-06, Revision 1, dated May 1985.
2. 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.
3. 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 Shorts Aircraft, 1725 Jefferson Davis Highway, Suite 510, Arlington, Virginia 22202. 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.
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2000-16-12:
This amendment adopts a new airworthiness directive (AD) that is applicable to General Electric Company (GE) CF6-45, -50, -80A, -80C2, and -80E1 turbofan engines with certain high pressure compressor rotor (HPCR) stage 3-9 spools installed. This action requires initial ultrasonic and eddy current inspections of certain HPCR stage 3-9 spools for cracks. This amendment is prompted by an uncontained failure of an HPCR 3-9 spool. The actions specified in this AD are intended to detect cracks which can cause separation of the HPCR stage 3-9 spool and result in an uncontained engine failure.
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92-03-02:
92-03-02 MCDONNELL DOUGLAS: Amendment 39-8156. Docket 92-NM-02-AD. \n\n\tApplicability: Model DC-9-81, -82, -83, and -87 series airplanes, and Model MD-88 airplanes, certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo prevent loss of engine thrust, accomplish the following: \n\n\t(a)\tWithin 10 days after the effective date of this AD, revise the Limitations Section of the FAA-approved Airplane Flight Manual (AFM) to include the following. This may be accomplished by inserting a copy of this AD in the AFM. \n\n\t"Ice on Wing Upper Surfaces\t\n\n CAUTION\n\n\t\t Ice shedding from the wing upper surface during takeoff can cause severe damage to one or both engines, leading to surge, vibration, and complete thrust loss. The formation of ice can occur on wing surfaces during exposure of the airplane to normal icing conditions. Clear ice can also occur on the wing upper surfaces when cold-soaked fuel is in the main wing fuel tanks,and the airplane is exposed to conditions of high humidity, rain, drizzle, or fog at ambient temperatures well above freezing. Often, the ice accumulation is clear and difficult to detect visually. The ice forms most frequently on the inboard, aft corner of the main wing tanks. (END OF CAUTIONARY NOTE) \n\nThe wing upper surfaces must be physically checked for ice when the airplane has been exposed to conditions conducive to ice formation. Takeoff may not be initiated unless the flight crew verifies that a visual check and a physical (hands-on) check of the wing upper surfaces have been accomplished, and that the wing is clear of ice accumulation when any of the following conditions occur: \n\n(1)\twhen the ambient temperature is less than 50 degrees F and high humidity or visible moisture (rain, drizzle, sleet, snow, fog, etc.) is present; \n\n(2)\twhen frost or ice is present on the lower surface of either wing; \n\n(3)\tafter completion of de-icing. \n\nWhen tufts and triangular decals are installed in accordance with McDonnell Douglas MD-80 Service Bulletin 30-59, the physical check may be made by assuring that all installed tufts move\nfreely.\n\n\t\t\t\t\tNOTE\n\n \tThis limitation does not relieve the requirement that aircraft surfaces are free of frost, snow, and ice accumulation, as required by Federal Aviation Regulations Sections 91.527 and 121.629. (END OF NOTE)" \n\n\t(b)\tWithin 10 days after the effective date of this AD, revise the Configuration Deviation List (CDL) Appendix of the AFM to include the following. This may be accomplished by inserting a copy of this AD in the AFM.\n \n\t"30-80-01 Triangular Decal and Tuft Assemblies \n\n\tUp to two (2) decals or tufts per side may be missing, provided: \n\n\t\t\ta)\tAt least one decal and tuft on each side is located along the aft spar line; and \n\n\t\t\tb)\tThe tufts are used for performing the physical check to determine that the upper wing is free of ice by observing that the tufts move freely. \n\n\t\tUp to eight (8)decals and/or tufts may be missing, provided: \n\n\t\t\ta)\tTakeoff may not be initiated unless the flight crew verifies that a physical (hands-on) check is made of the upper wing in the location of the missing decals and/or tufts to assure that there is no ice on the wing when icing conditions exist;\n\n\t\t\t\t\t OR \n\n\t\t\tb)\tWhen the ambient temperature is more than 50 degrees F." \n\n\t(c)\tWithin 30 days after the effective date of this AD, install tufts and triangular decals on the inboard side of the wings' upper surfaces, in accordance with McDonnell Douglas Service Bulletin 30-59, dated September 18, 1989; Revision 1, dated January 5, 1990; or Revision 2, dated August 15, 1990. \n\n\t(d)\tAn alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA PrincipalMaintenance or Operations Inspector, as appropriate, who may concur or comment and then send it to the Manager, Los Angeles ACO. \n\n\t(e)\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(f)\tThe installation required by this AD shall be done in accordance with McDonnell Douglas Service Bulletin 30-59, dated September 18, 1989; Revision 1, dated January 5, 1990; or Revision 2, dated August 15, 1990. \n\n\tThis 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 90846-0001, Attention: Business Unit Manager, Technical Publications, C1-HDR (54-60). Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. \n\n\t(g)\tThis amendment (39-8156, AD 92-03-02) becomes effective on January 17, 1992.
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88-25-51 R1:
88-25-51 R1 McDONNELL DOUGLAS: Amendment 39-6128. Final copy of, and revision to, telegraphic AD T88-25-51 which was issued December 16, 1988. \n\n\tApplicability: McDonnell Douglas Model DC-9-81, -82, -83, -87, and MD-88 series airplanes, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo ensure that the airplane horizontal stabilizer is configured correctly for takeoff, accomplish the following: \n\n\tA.\tWithin 5 calendar days after the effective date of this AD and daily thereafter, perform a functional test of the horizontal stabilizer takeoff position indicator assembly by selecting a center of gravity setting of 13 and a flap setting of 8, and verifying that the longitudinal trim is correctly displayed as 6 plus or minus 1/2 degrees. If the longitudinal trim displayed is not correct, check the adhesive bond strength of the center of gravity and flap indicator dials as indicated in paragraph B., below. \n\n\t\t1.\tIf both indicator dials are firmly bonded to their respective wheels, re-rig the horizontal stabilizer takeoff position indicator assembly in accordance with McDonnell Douglas MD-80 Maintenance Manual, Chapter 27-40-12, step 3, paragraph A, "Test Module." \n\n\t\t2.\tIf either the flap or center of gravity indicator dials show signs of loose bonding, or if the assembly cannot be re-rigged to bring it within tolerance, replace the horizontal stabilizer takeoff position indicator assembly prior to further flight. \n\n\tB.\tWithin 25 calendar days after the effective date of this AD and thereafter at intervals not to exceed 25 days, accomplish both of the following: \n\n\t\t1.\tCheck the center of gravity and flap indicator dial adhesive bond strength in accordance with the "Accomplishment Instructions," with the exception of paragraphs F and J, of McDonnell Douglas Service Bulletin A27-304, Revision 1, dated December 23, 1988. If the indicator dials are dislodged or loosened, replace the horizontalstabilizer takeoff position indicator assembly prior to further flight. \n\n\t\t2.\tPerform an adjustment/test of the horizontal stabilizer takeoff position indicator assembly in accordance with McDonnell Douglas MD-80 Maintenance Manual, Chapter 27-40-12, step 3, paragraph A, "Test Module." If the adjustment/test is not successful, check the adhesive bond strength of the center of gravity and flap indicator dials as indicated in paragraph B.1., above. \n\n\t\t\ta.\tIf both indicator dials are firmly bonded to their respective wheels, re-rig the horizontal stabilizer position indicator assembly in accordance with McDonnell Douglas MD-80 Maintenance Manual, Chapter 27-40-12, step 3, paragraph A, "Test Module." \n\n\t\t\tb.\tIf either the flap or center of gravity indicator dials show signs of loose bonding, or if the assembly cannot be re-rigged to bring it within tolerance, replace the horizontal stabilizer takeoff position indicator assembly prior to further flight. \n\n\tC.\tModification of theflap and center of gravity dial indicators in the horizontal stabilizer takeoff position indicator assembly, in accordance with "Repair Action II" of McDonnell Douglas Service Rework Drawing SRO9270010, New, dated December 17, 1988, constitutes terminating action for the requirements of this AD. \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 inspection requirements of this AD. \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, Los Angeles Aircraft Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Los Angeles Aircraft Certification Office. \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: Director of Publications, CL-100 (54-60). These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California. \n\n\tThis AD is the final copy of, and revision to, Telegraphic AD T88-25-51. Portions of this amendment were effective earlier to all recipients of Telegraphic AD T88-25-51, issued December 16, 1988, which was effective immediately upon receipt. \n\n\tThis amendment (39-6128, AD 88-25-51 R1) becomes effective February 15, 1989.
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85-03-03 R1:
85-03-03 R1 ROLLS-ROYCE LIMITED: Amendment 39-4987 as amended by Amendment 39-5197. Applies to Rolls-Royce Dart engines series 506, 510, 511, 514, 525, 526, 527, 528, 529, 530, 531, 532, 535, 542, and all variants of these series.
Compliance is required as indicated unless already accomplished.
To prevent possible uncontained failure of the low pressure impeller, accomplish the following:
Inspect certain low pressure impellers in accordance with mandatory Rolls-Royce Alert Service Bulletin (SB) Da72-A488, Revision 1, dated October 18, 1984, and Rolls-Royce SB Da72-480, Revision 4, dated December 1984, or FAA approved equivalent. Remove from service all impellers not ,meeting the inspection criteria per Rolls-Royce SB Da72-480.
(a) Inspect low pressure impellers with serial numbers listed in Appendix 1 of Rolls-Royce SB DA72- A488, Revision 1, which have accumulated less than 2,000 flights since new, within the next 200 flights, but not later than March 31, 1985.(b) Inspect low pressure impellers with serial numbers listed in Appendix 2 of Rolls-Royce SB Da72- A488, Revision 1, which have accumulated less than 2,000 flights since new, within the next 600 flights or next shop visit, whichever comes first, but not later than September 30, 1985.
NOTE: Shop Visit (as defined in the World Airlines Technical Operations Glossary) is defined as the input of an engine to a repair shop where the subsequent engine maintenance entails:
(a) Separation of a major engine flange (lettered or numbered) other than flanges mating with major sections of the nacelle or reverser. Note: Separation of flanges purely for purposes of shipment, without subsequent internal maintenance, is not a "Shop Visit."
(b) Removal of a disk or hub or spool.
(c) Removal of the main or angle gearbox.
(d) Removal of the fuel nozzles.
(c) Inspect low pressure impellers with serial numbers listed in Appendix 1 of Rolls-Royce SB Da72- A488, Revision 2, which have not been inspected per (a) above, which have accumulated less than 3,500 flights since new, within the next 100 flights, but not later than February 28, 1986.
(d) Inspect low pressure impellers with serial numbers listed in Appendix 1 of Rolls-Royce SB Da72- A488, Revision 2, which have not been inspected per (a) above, which have accumulated 3,500 or more flights since new, within the next 500 flights, but not later than April 30, 1986.
For purpose of this AD the term "repair shop" refers to a maintenance station where low compressor overhaul facilities exist.
Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations (FARs) 21.197 and 21.199 to a base where the AD can be accomplished.
Upon request, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Engine Certification Office, Aircraft Certification Division, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803.
Rolls-Royce SB Da72-A488, Revision 2, dated July 25, 1985, is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received this document from the manufacturer may obtain Rolls-Royce Limited, East Kilbride, Glasgow G74 4PY, Scotland. This document also may be examined at the Office of the Regional Counsel, FAA, New England Region, Attention: Rules Docket No. 84-ANE-19, 12 New England Executive Park, Room Number 311, between the hours of 8:00 a.m. and 4:30 p.m., Monday through Friday except Federal holidays.
Amendment 39-4987 became effective on February 11, 1985.
This Amendment 39-5197 becomes effective on January 6, 1986.
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2018-25-02:
We are adopting a new airworthiness directive (AD) for certain Airbus SAS Model A318, A319, A320, and A321 series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the principal structural elements and certain life- limited parts are subject to widespread fatigue damage (WFD). This AD requires revising the existing maintenance or inspection program to incorporate new or more restrictive airworthiness limitations. We are issuing this AD to address the unsafe condition on these products.
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91-21-05:
91-21-05 MCDONNELL DOUGLAS: Amendment 39-8052. Docket No. 91-NM-174-AD.\n\n\tApplicability: Model DC-10-10, -10F, and -15 series airplanes, fuselage numbers through 379; and Model DC-10-30, -30F, and -40 series airplanes, fuselage numbers through 275; certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished.\n\n\tTo ensure the structural integrity of these airplanes, accomplish the following:\n\n\t(a)\tExcept as provided in paragraphs (b) through (e) of this AD, prior to the accumulation of 7,000 landings or within 30 days after the effective date of this AD, conduct the initial inspections specified in paragraphs (a)(1), (a)(2), or (a)(3) of this AD.\n\n\t\t(1)\tConduct an eddy current inspection of the wing rear spar lower cap aft tang and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424 in accordance with Option III of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. In addition, within 1,500 landings after performing these eddy current and dye penetrant inspections, conduct the inspections specified in paragraph (a)(2) or (a)(3) of this AD and repeat thereafter as indicated. Or\n\n\t\t(2)\tConduct an ultrasonic inspection of the area around the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424 in accordance with Option II of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 1,900 landings. Or\n\n\t\t(3)\tConduct an eddy current inspection of the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424 in accordance with Option I of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 3,300 landings.\n\n\t(b)\tThe requirements of paragraph (c) of this AD apply to airplanes on which the following actions have been accomplished:\n\n\t\t(1)\tThe dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 422 has been accomplished prior to the effective date of this AD, in accordance with McDonnell Douglas Service Bulletin 57-61, Revision 2, dated August 15, 1990; and\n\n\t\t(2)\tThe eddy current inspection of the wing rear spar lower cap aft tang has been accomplished prior to the effective date of this AD per DC-10 Supplemental Inspection Document, Principal Structural Element (PSE) 57.10.007 and 57.10.008, in accordance with McDonnell Douglas Service Bulletin 57-61, Revision 2, dated August 15, 1990.\n\n\t(c)\tFor airplanes specified in paragraph (b) of this AD, conduct the initial inspections specified in either paragraph (c)(1) or (c)(2) of this AD within 1,500 landings after the inspections (eddy current and dye penetrant) specified in paragraph (b)(1) of this AD, or within 30 days after the effective date of this AD, whichever occurs later.\n\n\t\t(1)\tConduct an ultrasonic inspection of the area around the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424, in accordance with Option II of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 1,900 landings. Or\n\n\t\t(2)\tConduct an eddy current inspection of the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424, in accordance with Option I of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 3,300 landings.\n\n\t(d)\tThe requirements of paragraph (e) of this AD apply to airplanes on which the following actions have been accomplished: \n\n\t\t(1)\tThe dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 422 has been accomplished prior to the effective date of this AD, in accordance with McDonnell Douglas Service Bulletin 57-61, Revision 2, dated August 15, 1990; and\n\n\t\t(2)\tThe eddy current inspection of the wing rear spar lower cap aft tang fastener holes located between stations Xors= 417 and Xors= 422 has been accomplished prior to the effective date of this AD per DPS 4.735-9 in accordance with McDonnell Douglas Service Bulletin 57-61, Revision 2, dated August 15, 1990.\n\n\t(e)\tFor airplanes specified in paragraph (d) of this AD, conduct the initial inspections specified in either paragraph (e)(1) or (e)(2) of this AD within 3,300 landings after the accomplishment of the inspection specified in paragraph (d)(1) of this AD, or within 30 days after the effective date of this AD, whichever occurs later.\n\n\t\t(1)\tConduct an ultrasonic inspection of the area around the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424 in accordance with Option II of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 1,900 landings. Or\n\n\t\t(2)\tConduct an eddy current inspection of the six wing rear spar lower cap aft tang fastener holes and a dye penetrant inspection of the wing trailing edge access door sill located between stations Xors= 417 and Xors= 424 in accordance with Option I of McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 1991. Repeat these inspections at intervals not to exceed 3,300 landings.\n\n\t(f)\tIf any cracks are found as a result of the inspections conducted in accordance with this AD, prior to further flight, repair ina manner approved by the Manager of the Los Angeles Aircraft Certification Office, FAA, Transport Airplane Directorate.\n\n\t(g)\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate to a base in order to comply with the requirements of this AD.\n\n\t(h)\tAn alternative method of compliance or adjustment time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate.\n\n\tNOTE: The request should be forwarded through the FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles ACO.\n\n\t(i)\tThe inspection requirements shall be done in accordance with McDonnell Douglas Alert Service Bulletin A57-123, dated July 25, 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 McDonnell Douglas Corporation, P.O. Box 1771, Long Beach California 90806-0001, Attention: Business Unit Manager of Technical Publications, Technical Administrative Support C1-L5B (54-60). Copies may be inspected 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; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C.\n\n\tThis amendment (39-8052, AD 91-21-05) becomes effective on October 23, 1991.
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72-06-04:
72-06-04 HAWKER SIDDELEY: Amdt. 39-1405. Applies to Hawker Siddeley Model DH-114, Series 2, "Heron" airplanes which do not incorporate Heron Modification 1592, Part A.
Compliance is required as indicated.
To prevent failure of the main undercarriage down lock operating lever assembly, accomplish the following:
(a) For airplanes with main undercarriage down lock operating lever assembly, P/N 14-2U.181A (pre Modification 608), within the next 150 hours' time in service after the effective date of this AD unless accomplished within the last 150 hours' time in service, remove the lever assembly from the airplane, remove the protective coating and paint from the assembly and inspect for cracks, using a dye penetrant method, in accordance with Hawker Siddeley Technical News Sheet Series: Heron (114) No. U.12, Issue 1, dated April 13, 1970, or later ARB-approved issue or FAA-approved equivalent. If no cracks are found visually inspect the lever assembly for corrosion.
(b) If cracks, or corrosion that cannot be removed by cleaning are found during the inspection required by paragraph (a), before further flight replace the affected part with a serviceable part of the same part number or replace the lever assembly with a new lever assembly, P/N 14-2U.181A/2, in accordance with Hawker Siddeley "Heron" Modification 1592.
(c) For airplanes with main undercarriage down lock operating lever assembly, P/N 14-2U.181A/1 (post Modification 608), within the next 150 hours' time in service after the effective date of this AD, or within 300 hours' time in service from the last inspection, whichever occurs later, and thereafter at intervals not to exceed 300 hours' time in service from the last inspection, remove the lever assembly from the airplane, remove the protective coating and paint from the assembly and inspect for cracks, using a dye penetrant method, in accordance with Hawker Siddeley Technical News Sheet Series: Heron (114) No. U.12, Issue 1, dated April 13, 1970, or later ARB-approved issue or FAA-approved equivalent.
(d) If cracks are found during an inspection required by paragraph (c), before further flight replace the lever assembly with a new lever assembly, P/N 14-2U.181A/2 in accordance with Hawker Siddeley "Heron" Modification 1592.
(e) The repetitive inspections required by paragraph (c) may be discontinued after lever assembly, P/N 14-2U.181A/2, has been installed in accordance with Hawker Siddeley "Heron" Modification 1592.
(f) Replacement parts and serviceable parts that are reinstalled must be protected with a coat of lanolin, or FAA-approved equivalent prior to their installation in the airplane.
This amendment becomes effective April 6, 1972.
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2006-06-02:
This amendment supersedes an existing airworthiness directive (AD) that currently applies to Eurocopter France (ECF) Model SA 365N, N1, and AS 365N2 helicopters. That AD currently requires inspecting the main gearbox (MGB) suspension diagonal cross-member (diagonal cross-member) for cracks and replacing it with an airworthy part if any crack is found. This amendment requires more frequent inspections of the diagonal cross-member and adding the Model SA-366G1 helicopters to the applicability. This amendment is prompted by several reports of cracks in the diagonal cross-member. The actions specified by this AD are intended to prevent failure of the diagonal cross-member, pivoting of the MGB, severe vibrations, and a subsequent forced landing.
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89-08-01:
89-08-01 McDONNELL DOUGLAS: Amendment 39-6178. \n\n\tApplicability: Model DC-8 series airplanes, equipped with control columns, P/N 5614272-1 and/or -2, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent the loss of airplane control in critical flight regimes due to fatigue failure of the control column, accomplish the following: \n\n\tA.\tWithin the next 1,250 hours time-in-service after September 25, 1974 (the effective date of Amendment 39-1967), unless already accomplished within the last 1,250 hours time-in-service, and thereafter at intervals not to exceed 2,500 hours time-in-service, except as provided below, conduct a dye penetrant or eddy current inspection of the control columns in accordance with the instructions in McDonnell Douglas All Operators Letter 8-632, issued October 11, 1972, or equivalent inspection technique approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t1.\tIf surface indications of cracking exist, consisting of small specks not yet joined to form a linear crack of at least 1/8-inch in length, no rework is required, but the inspection interval is thereafter reduced to 2,000 hours time-in-service. \n\n\t\t2.\tIf linear cracks of 1/8-inch or more exist, blendout may be accomplished, in lieu of replacement, subject to the following qualifications: Blendout shall not exceed .030-inch in depth from the original surface and shall be blended over an area 10 times the depth. The defect shall not exceed an initial length of 1/4-inch. No more than two defects can occur in the same horizontal plane, and the defects shall be separated by at least 2-inch center-to-center spacing. Additional defects may be blended if the vertical distance between horizontal planes is at least 1/4-inch and the 2-inch center-to-center spacing requirement in the same horizontal plane is observed. After this rework, inspect at intervals not to exceed 2,500 hours time-in-service. \n\n\t\t3.\tIf cracks which exceed the limits described in paragraph A.2 of this AD are discovered as a result of any inspection, remove and replace the control columns, P/N's 5614272-1 and/or -2, in accordance with paragraph C. of this AD, or rework in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tB.\tWithin the next 1,250 hours time-in-service after the effective date of this AD, unless already accomplished within the last 1,250 hours time-in-service, conduct a dye penetrant or eddy current inspection of the control columns, in accordance with the instructions in McDonnell Douglas DC-8 Alert Service Bulletin A27-267, dated February 18, 1987, Revision 1, dated May 22, 1987, or Revision 2, dated February 14, 1989, or equivalent inspection technique approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t1.\tIf no cracks are found, accomplish repetitive inspections at intervals not to exceed 2,500 hours time-in-service. \n\n\t\t2.\tIf cracks are found, prior to further flight, remove and replace the control column in accordance with paragraph C. of this AD. \n\n\tAccomplishment of the provisions of paragraph B., constitutes terminating action for the inspection requirements of paragraph A. of this AD. However, for areas specified in the McDonnell Douglas All Operators Letter 8-632, dated October 11, 1972, allowable limits and blendout criteria established in accordance with paragraph A. of this AD still apply. \n\n\tC.\tReplacement of both pilot's and copilot's control columns, P/N's 5614272-1 or 5614272-2, respectively, with new control columns, SB 09270288-3, 5614272-501, or -503 (pilot's), and SB 09270288-4, 5614272-502, or -504 (co-pilot's), in accordance with McDonnell Douglas DC-8 Service Bulletin 27-267, issued January 20, 1988, or Revision 1, dated February 17, 1989, constitutes terminating action for the requirements of this AD. \n\n\tD.\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, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Los Angeles Aircraft Certification Office. \n\n\tE.\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 McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director of Publications, C1-LOO (54-60). These documents may be examined at the FAA, NorthwestMountain Region, 17900 Pacific Highway South, Seattle, Washington, or at 3229 E. Spring Street, Long Beach, California. \n\n\tThis amendment supersedes AD 73-07-09, (Amendment 39-1967) which was effective on September 25, 1974. \n\n\tThis amendment (39-6178, AD 89-08-01) becomes effective May 2, 1989.
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2018-24-03:
We are adopting a new airworthiness directive (AD) for all Dassault Aviation Model Falcon 10 airplanes. This AD was prompted by a determination that new and more restrictive maintenance requirements and airworthiness limitations are necessary. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive maintenance requirements and airworthiness limitations. We are issuing this AD to address the unsafe condition on these products.
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2000-05-30:
This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747 series airplanes, that requires repetitive inspections to detect discrepancies of the cables, fittings, and pulleys of the engine thrust control cable installation, and replacement, if necessary. This AD also requires certain preventative actions on the engine thrust control cable installation for certain airplanes. This amendment is prompted by reports of failure of engine thrust control cables. The actions specified by this AD are intended to prevent such failures, which could result in a severe asymmetric thrust condition during landing, and consequent reduced controllability of the airplane.
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86-19-01:
86-19-01 BOEING: Amendment 39-5394. Applies to Model 747-100SR series airplanes listed in Section 3.0 of Boeing Document No. D6-35655 "Supplemental Structural Inspection Document" (SSID), approved March 22, 1986, certificated in any category. Compliance is required as indicated in the body of the AD.\n\n\tTo ensure the continuing structural integrity of these airplanes, accomplish the following, unless already accomplished:\n\n\tA. Within one year after the effective date of the AD, incorporate a revision into the FAA-approved maintenance inspection program which provides no less than the required damage tolerance rating (DTR) for each Structural Significant Item (SSI) listed in Boeing Document D6-35655, approved March 22, 1986, or later FAA-approved revisions. The required DTR value for each SSI is listed in the document. The revision to the maintenance program must include and be implemented in accordance with the procedures in Sections 5.0 and 6.0 of the SSID.\n\n\tB. Cracked structure must be repaired before further flight in accordance with an FAA-approved method.\n\n\tC. 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.\n\n\tD. 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region.\n\n\tE. Operators who have acceptably incorporated Boeing Document No. D6-35655, approved March 22, 1986, or later FAA-approved revisions, into their approved maintenance program are exempt from 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 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\tThis amendment becomes effective September 22, 1986.
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