Results
98-13-20: This amendment adopts a new airworthiness directive (AD) that is applicable to Rolls-Royce Limited, Aero Division-Bristol, S.N.E.C.M.A, Olympus 593 series turbojet engines. This action requires a radiological inspection of the combustion chamber No. 2 outer cooling ring scoop circumferential and axial weld for weld quality, and reweld and reinspection, if necessary; and an inspection of the combustion chamber No. 2 inner and outer cooling ring web length, marking acceptable components with the letter "T" adjacent to the part number, and replacement of unacceptable components with serviceable parts. This amendment is prompted by reports of circumferential cracks at the No. 2 outer and inner rings of the combustor chamber, resulting in a section of the combustion chamber detaching and causing significant ignitor and low pressure turbine damage. The actions specified in this AD are intended to prevent combustion chamber detachment, which could result in an inflight engine shutdown or an engine fire. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of July 10, 1998. Comments for inclusion in the Rules Docket must be received on or before August 24, 1998.
84-13-04: 84-13-04 ALEXANDER SCHLEICHER: Amendment 39-4901. Applies to Model ASW 19 and ASW 19B gliders, serial numbers 19001 to 19402, except 19019 and 19210, certificated in all categories. Compliance required prior to next flight unless already accomplished. To prevent the occurrence of horizontal tailplane flutter, accomplish the following: 1. Apply a red radial line on the airspeed indicator at 108 Kts (200 km/h) to indicate the new Never Exceed Velocity (VNE). 2. Affix a placard stating "Maximum Speed 108 Kts (200 km/h)" placed to the airspeed indicator. 3. Enter a notation in the glider flight manual in the airspeed limitations section to read as follows: Max. Speed to 10,000' MSL 108 Kts (200 km/h) 10,001 to 16,400' MSL 80 Kts (155 km/h) 16,401 to 23,000' MSL 75 Kts (140 km/h) 23,001 to 29,500' MSL 65 Kts (120 km/h) 4. Compliance with this AD is not required when the elevator trailing edge contour change modification described in Alexander Schleicher Technical Note No. 17, dated March 27, 1984, is incorporated. 5. Alternate inspections, adjustment of the inspection interval, or other actions which provide an equivalent level of safety must be approved by the Manager, Brussels Aircraft Certification Office, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium, telephone 513.38.30 x2710. The Alexander Schleicher Technical Note No. 17, dated March 27, 1984, identified and described in this directive is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Alexander Schleicher Segelflugzeubau, D-6414 Poppenhausen, Federal Republic of Germany. These documents may also be examined at the Office of Regional Counsel, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment becomes effective September 13, 1984, as to all persons except those persons to whom it was made immediately effective by priority letter AD 84-13-04, issued June 26, 1984, which contained this amendment.
2012-22-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 747-400, -400D, and -400F series airplanes. This AD was prompted by reports of crown frame web cracking at left buttock line (LBL) 15.0, station (STA) 320. This AD requires measuring the web at STA 320 and, depending on findings, various inspections for cracks and missing fasteners, web and fastener replacement, and related investigative and corrective actions if necessary. We are issuing this AD to prevent complete fracture of the crown frame assembly, and consequent damage to the skin and in-flight decompression of the airplane.
98-13-31: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires repetitive visual inspections to detect discrepancies of the bushing installation of the aileron actuation fitting, and eventual installation of staked bushings in the fitting. Accomplishment of such installation terminates the repetitive inspections. This amendment also provides for an optional temporary preventive action, which, if accomplished, would allow the repetitive inspection intervals to be extended until the terminating action is accomplished. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent failure of the fitting lugs due to vibration caused by loose bushings in the fittings, and consequent reduced controllability of the airplane.
69-24-02 R3: 69-24-02 R3 BEECH: Amendment 39-879 as amended by Amendment 39-1222 is further amended by Amendment 39-4240 and Amendment 39-4607. Applies to all Beech 99 series (Serial Numbers U-1 thru U-49, U-51 thru U-131, U-133 thru U-145, and U-147) airplanes certificated in any category. EXCEPTIONS: Airplanes on which the elevator control system has been modified in accordance with optional Beech Kit 99-5011 or Kit 99-5014 are required to comply with paragraph C only. Compliance: Required as indicated, unless already accomplished. To prevent an unsafe condition, effective immediately, restrict the aircraft to a maximum speed of 174 knots Vmo and remark the airspeed aircraft to a maximum speed of 174 knots Vmo and remark the airspeed indicator at that speed until the following are accomplished: A. (1) Check for correct elevator control system rigging and if necessary, rerig in accordance with Beech Service Instruction 0309-364. (2) Subsequent to the requirements of Paragraph A(1), conduct a flight test in accordance with flight test procedures contained in Beech Service Instruction 0309-364, or equivalent procedures approved by the Manager, Aircraft Certification Branch, FAA, Central Region. B. On or before March 15, 1970, limit the upward travel of the leading edge of the stabilizer to a maximum of 3-1/2 degrees in accordance with Beech Service Instruction 0285-364 or Beech Service Instruction 0309-364. The downward travel of the stabilizer leading edge remains unchanged. C. (1) On or before March 15, 1970, (a) install an aural warning device which indicates that the stabilizer trim system is in motion, (b) install an out-of-trim warning system indicating that the aircraft is out-of-trim longitudinally prior to takeoff, (c) install a newly designed standby trim switch and guards to prevent inadvertent operation, and (d) revise the electrical circuitry to prevent unwanted circuit breaker activation in the event that pilot and co- pilot simultaneously call for opposite trim. These modifications must be accomplished in accordance with instructions and procedures set forth in Beech Service Instruction 0270-350 or equivalent modifications approved by the Manager, Aircraft Certification Branch, FAA, Central Region. (2) On or before March 15, 1970, relocate the trim release switch on the control wheel to make it more readily accessible to the crew, in accordance with instructions and procedures set forth in Beech Service Instruction 0249-156, or an equivalent method approved by the Manager, Aircraft Certification Branch, FAA, Central Region. (3) On or before March 15, 1970, revise Beech Model 99 Approved Airplane Flight Manual by incorporating revision C-1 dated November 14, 1969, and revise Beech Model 99A Approved Airplane Flight Manual by incorporating revision A-5 dated November 14, 1969. NOTE: When the revisions to the Approved Airplane Flight Manuals have been incorporated as required by paragraph C(3), the temporary amendments to the Airplane Flight Manuals in AD 69-18-06, as amended, may be deleted. D. When paragraphs A and B of this AD have been accomplished, the aircraft may be operated at a maximum speed not to exceed 200 knots Vmo. When the modifications required by paragraphs A, B, and C of this AD have been accomplished, the aircraft may be operated at a speed not to exceed 226 knots Vmo. E. Paragraphs A(1) and A(2) of this AD must be complied with irrespective of the speed limitation whenever the elevator control system is repaired or otherwise modified or the elevators are repaired or replaced. AD 69-24-02, Amendment 39-879, superseded AD 69-18-06 as amended insofar as it changes the maximum speed restrictions provided in AD 69-18-06 as amended. Amendment 39-879 became effective December 5, 1969, for all persons except those to whom it was made effective by first class letter dated November 21, 1969. Amendment 39-1222 became effective June 3, 1971. Amendment 39-4240 became effective October 16, 1981. This Amendment 39-4607 becomes effective April 11, 1983.
82-16-05 R1: 82-16-05 R1 PIPER AIRCRAFT CORPORATION: Amendment 39-4459 as amended by Amendment 39-5278. Applies to Models PA-31 and PA-31-325 (Serial Numbers 31-2 through 31-8312019), PA-31-350 (Serial Numbers 31-5001 through 31-8452021), and PA-31-350-T1020 (Serial Numbers 31-8253001 through 31-8553002) equipped with Piper Part Numbers 455-301, 555-376, 555-511, or 555-366 turbocharger exhaust pipe couplings, certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent the possibility of an inflight powerplant fire due to a turbocharger exhaust pipe coupling failure, accomplish the following: (a) Within the next 100 hours time-in-service after the effective date of this AD or 100 hours time-in-service since the last inspection per this AD prior to its revision, whichever is first, and thereafter at intervals not exceeding 100 hours time-in-service, inspect the multi-segment Piper P/N 455-301, 555-376, 555-511, 555-366 turbocharger exhaust pipe couplings by accomplishing the following: (1) Gain access to the turbocharger exhaust systems. (2) Remove the turbocharger exhaust couplings and tailpipe. NOTE: Exercise caution to prevent spreading or forcing the coupling beyond its normal open position when removing or installing the coupling, (3) Using either a dye penetrant inspection method or a light and a 10-power magnifying glass, accomplish the following: (i) Inspect coupling for cracks, spreading of "V" band segments, failed spot welds, and indication of exhaust flanges bottoming in couplings. (ii) Inspect the condition of the coupling clamp for bending, overstress, thread damage, cracks or other obvious damage. (iii) Inspect turbocharger to turbocharger exhaust tailpipe connection area for proper mating of surfaces. (iv) Inspect tailpipe and turbocharger flanges for cracks and distortion. Remove carbon deposits from mating flanges before reassembly. (v) Reinstall serviceable couplings using the applicable torque and procedures described in paragraph (b). NOTE: Initial and repetitive inspection are not required for coupling Part Numbers 557- 584 and 557-369. (b) Prior to further flight, replace any cracked or otherwise damaged couplings found during any inspection required by paragraph (a) of this AD with applicable couplings specified below: MODEL PIPER COUPLING P/N (AEROQUIP P/N) TORQUE PA-31 455-301 (4404-376M) 555-376 (MVT68049-375H) (MVT68049-375D) 555-511 (MVT69861-377M) 557-584 (NH1005834-10) 40-50 in.-lbs. 40-50 in.-lbs. 40-50 in.-lbs. 30-35 in.-lbs. PA-31-325 555-511 (MVT69861-377M) 557-584 (NH1005834-10) 40-50 in.-lbs. 30-35 in.-lbs. PA-31-350 555-366 (MVT68049-450M) 557-369 (NH1005798-10) 45-55 in.-lbs. 30-35 in.-lbs. Install couplings in accordance with the instructions contained in Piper Service Bulletin No. 644C, dated December 3, 1985, ensuring that the tailpipe andturbocharger flanges are properly aligned and that the wrench socket is properly aligned to prevent bolt sideload. (c) Piper Aircraft Corporation Service Bulletin No. 644C dated December 3, 1985, pertains to the subject matter of this AD. (d) The time-in-service between the repetitive inspections required herein may be adjusted up to plus 25 percent of any specified inspection interval required by this AD to facilitate accomplishing these inspections concurrent with other scheduled maintenance on the airplane. (e) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (f) An equivalent method of compliance with this AD if used must be approved by the Manager, Atlanta Aircraft Certification Office, FAA, 1075 Inner Loop Road, College Park, Georgia 30337. All persons affected by this directive may obtain copies of the documents referred to herein upon request to Piper Aircraft Corporation, 2926 Piper Drive, Vero Beach, Florida 32960 or the FAA, Rules Docket, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. Amendment 39-4459 became effective September 15, 1982. This amendment, 39-5278, becomes effective April 11, 1986.
92-03-12: 92-03-12 BOEING: Amendment 39-8169. Docket 91-NM-138-AD. Supersedes AD 91-11-06, Amendment 39-7002. \n\n\tApplicability: Model 707/720 series airplanes; as listed in Boeing Service Bulletin 3240, Revision 3, dated October 18, 1985; certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo ensure continued structural integrity of the wing rear spar upper chord, accomplish the following: \n\n\t(a)\tPerform a close visual inspection for cracks and corrosion of the wing rear spar upper chord from wing station (WS) 109.45 to WS 360 for Model 707-300 series airplanes; or from WS 180.71 to WS 360 for Model 720, 707-100, and 707-200 series airplanes; at rib and stiffener locations. Inspect in accordance with Boeing Service Bulletin 3240, Revision 3, dated October 18, 1985, prior to the later of the times specified in subparagraphs (a)(1) and (a)(2) of this AD, unless previously accomplished within the last 900 flight cycles or 335 days.Repeat the inspection at intervals not to exceed 1,000 flight cycles or one year, whichever occurs first. \n\n\t\t(1)\tWithin the next 30 days or 100 flight cycles after June 19, 1991 (the effective date of Amendment 39-7002, AD 91-11-06); or \n\n\t\t(2)\tPrior to the accumulation of 10,000 flight cycles. \n\n\t(b)\tIf cracks or corrosion areas are found, prior to further flight, accomplish either subparagraph (b)(1) or (b)(2) of this AD: \n\n\t\t(1)\tRepair, other than by stop drill procedure, in accordance with Part III, Figure 2, of Boeing Service Bulletin 3240, Revision 3, dated October 18, 1985 (this is considered the "final repair"), or \n\n\t\t(2)\tRepair in accordance with the stop drill procedures specified in Part III, Figure 2, of Service Bulletin 3240, Revision 3, dated October 18, 1985. This repair method may only be used provided that the limitations specified in Part III, Figure 2, Items 5a and 5b, of the service bulletin are met. \n\n\t\t\t(i)\tImmediately after stop drilling, conductan eddy current inspection of the stop drill hole in accordance with the instructions in Section 5-5-1 of D6-7170, Nondestructive Test Document, to ensure that the crack does not extend beyond the stop drill. Thereafter, inspect visually for crack growth beyond the stop drill at intervals not exceeding 300 flight cycles. \n\n\t\t\t(ii)\tIf crack growth beyond the stop drill occurs, prior to further flight, accomplish the final repair in accordance with paragraph (b)(1) of this AD. \n\n\t\t\t(iii)\tWithin 1,000 flight cycles or one year, whichever occurs first, after the stop drill has been accomplished, accomplish the final repair in accordance with paragraph (b)(1) of this AD. \n\n\t(c)\tIf previously stop drilled cracks are found as a result of the inspection required by paragraph (a) of this AD, conduct an eddy current inspection of the stop drill hole for crack growth beyond the stop drill, in accordance with the instructions in Section 5-5-1 of Boeing Document D6-7170, Nondestructive TestDocument. \n\n\t\t(1)\tIf growth beyond the stop drill has occurred, prior to further flight, repair in accordance with paragraph (b)(1) of this AD. \n\n\t\t(2)\tIf growth beyond the stop drill has not occurred, and the limitations specified in Part III, Figure 2, Items 5a and 5b, of Boeing Service Bulletin 3240, Revision 3, dated October 18, 1985, are met, prior to further flight accomplish either subparagraph (c)(1)(i) or (c)(1)(ii) of this AD: \n\n\t\t\t(i)\tRepair in accordance with paragraph (b)(1) of this AD; or \n\n\t\t\t(ii)\tReinspect visually for crack growth beyond the stop drill at intervals not exceeding 300 flight cycles.\n \n\t\t\t\t(A)\tIf crack growth beyond the stop drill occurs, prior to further flight, accomplish the final repair in accordance with paragraph (b)(1) of this AD. \n\n\t\t\t\t(B)\tWithin 1,000 flight cycles or one year, whichever occurs first after the initial inspection revealed the stop drill crack, accomplish the final repair in accordance with paragraph (b)(1) of this AD.(d)\tAfter each of the inspections and repairs required by this AD have been performed, apply BMS 3-23 corrosion inhibitor, or equivalent, to the affected areas. \n\n\t(e)\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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Seattle ACO. \n\n\t(f)\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\t(g)\tThe inspections and repairs shall be done in accordance with Boeing Service Bulletin 3240, Revision 3, dated October 18, 1985. 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, at 56 FR 25356. Copies may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. Copies may be inspected at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington, or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, D.C. \n\n\t(h)\tThis amendment (39-8169, AD 92-03-12) becomes effective on March 10, 1992.
89-24-07: 89-24-07 AEROSPATIALE: Amendment 39-6394. Docket No. 89-NM-126-AD. Applicability: All Model ATR42 series airplanes, certificated in any category. Compliance: Required within 60 days after the effective date of this AD, unless previously accomplished. To improve protection against loss of control when operating in icing conditions, including freezing rain, accomplish the following: A. Install an anti-icing advisory system in accordance with Aerospatiale Service Bulletins ATR42-27-0021, Revision 8, dated May 2, 1989; ATR42-30-0017, Revision 3, dated January 20, 1989; ATR42-30-0018, Revision 4, dated June 14, 1989; ATR 42-30-0021, Revision 3, dated July 20, 1989; ATR42-30-0024, Revision 1, dated February 13, 1989; and ATR42-30- 0027, Revision 1, dated February 21, 1989. B. Incorporate Revision 6 of the Airplane Flight Manual (AFM), dated August 1988, into the FAA-approved AFM. C. 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. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6394, AD 89-24-07) becomes effective on December 15, 1989.
82-10-02 R1: 82-10-02 R1 ROCKWELL INTERNATIONAL: Amendment 39-4376 as amended by Amendment 39-4436. Applies to Models NA-265-60 modified per STC SA687NW (S/N 306-5, - 6, -12, -24, -50, -71, -109, -116, -119, and -122); NA-265-65, S/N 465-1 through 465-76, and S/N 306-114 airplanes certificated in all categories. Compliance required as indicated unless already accomplished. To reduce the possibility of a flap screwjack actuator malfunction due to bearing failure, accomplish the following: A. Within the next 50 hours time-in-service after the effective date of this AD and within each additional 300 hours time-in-service, inspect each flap screwjack in accordance with Rockwell International Service Bulletin 81-14 dated December 28, 1981. Replace or repair any flap screwjack not meeting the acceptance criteria in this service bulletin. B. Model NA-265-65 Sabreliner airplanes are approved for take-offs and landings with zero degree flap settings and may, therefore, be operated with a defective flap screwjack, provided the flap system is deactivated and the airplane operated in accordance with the zero flap performance data in the Airplane Flight Manual SR-77-006. C. Model NA-265-60 airplanes may be operated with a defective flap screwjack, provided the flap system is deactivated and the following operating limitations are observed: 1. Takeoff must be conducted in compliance with Airplane Flight Manual zero flap requirements. 2. Landing weather minimums are one mile or RVR 5000. 3. Zero flap, dry runway landing distance must be determined by multiplying the factored landing distance shown in the Rockwell International Airplane Flight Manual, Supplement No. SR-81-018, by a factor of one point five. 4. Zero flap, wet runway landing distances must be determined by multiplying the distance obtained in "3." above by a factor of one point one five. 5. Thrust reversers must be operative prior to takeoff. D. Alternative means of compliance with this AD which provide an equivalent level of safety must be approved by the Chief, Wichita Aircraft Certification Office, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 269- 7000. E. Rework of all flap actuators in accordance with Rockwell International Service Bulletin 82-1 dated June 30, 1982, constitutes terminating action for the requirements of paragraph A, above. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request from Rockwell International, Sabreliner Division, 6161 Aviation Drive, St. Louis, Missouri 63134. These documents may also be examined at FAA Central Region, Wichita Aircraft Certification Office, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209. Amendment 39-4376 became effective May 17, 1982. This Amendment 39-4436 becomes effective September 13, 1982.
2022-08-02: The FAA is adopting a new airworthiness directive (AD) for all Airbus Helicopters Model EC 155B and EC155B1 helicopters. This AD was prompted by a report of a discrepancy in the rotorcraft flight manual (RFM) where the rotorcraft stay-up flying capabilities for Category B operation were provided through performance data only, not as airworthiness limitations that are dependent upon on the number of passengers on board. This AD requires revising the existing RFM for your helicopter, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2012-23-01: We are adopting a new airworthiness directive (AD) for all Cessna Aircraft Company (Cessna) Model 402C airplanes modified by Supplemental Type Certificate (STC) SA927NW and Model 414A airplanes modified by STC SA892NW. This AD was prompted by report of a Cessna Model 414A airplane modified by STC SA892NW that experienced an asymmetrical flap condition causing an uncommanded roll when the pilot set the flaps to the approach position. We are issuing this AD to prevent failure of the flap system, which could result in an asymmetrical flap condition. This condition could result in loss of control.
98-13-30: This amendment adopts a new airworthiness directive (AD), applicable to all Gulfstream Aerospace Corporation Model G-159 (G-I) airplanes, that requires revising the Airplane Flight Manual (AFM) to prohibit positioning the power levers below the flight idle stop. This amendment is prompted by incidents and accidents involving airplanes equipped with turboprop engines in which the ground propeller beta range was used improperly during flight. The actions specified by this AD are intended to prevent loss of airplane controllability or engine overspeed with consequent loss of engine power caused by the power levers being positioned below the flight idle stop while the airplane is in flight.
81-06-03: 81-06-03 SHORT BROTHERS LIMITED: Amendment 39-4057. Applies to Model SC-7 Series 3 airplanes, certificated in all categories, which have an autopilot installed. Compliance is required within the next 50 hours time in service after the effective date of this AD, unless already accomplished. To prevent aileron autopilot servo control cable breakage and possible loss of control of the airplane, accomplish the following: (a) Inspect the aileron autopilot servo control cable for proper positioning on the guide pulley, for integrity, and for evidence of damage in the vicinity of the guide pulleys at station 226 in accordance with paragraphs 2.A.1, 2, and 3 of Shorts Service Bulletin 22-A59, dated May 21, 1979 (hereinafter referred to as the service bulletin), or an FAA-approved equivalent. (b) If as a result of the inspection required in paragraph (a) of this AD: (1) A damaged cable is found, replace the cable and accomplish paragraphs 2.A.4, 5, and 6 of the service bulletin, or an FAA-approved equivalent. (2) An improperly positioned cable is found, reposition the cable after verification that no damage to the cable has occurred, and accomplish paragraph (d) of this AD. (c) Determine the clearance between the cable guard and the outer rim of the pulley at station 226. If the clearance exceeds 0.029 inches, reposition the guard in accordance with paragraph 2.A.4 of the service bulletin, or an FAA-approved equivalent, and accomplish paragraph (d) of this AD. (d) Check the aileron servo control cable for correct tension in accordance with paragraph 2.A.5 of the service bulletin, or an FAA-approved equivalent. (e) If an equivalent means of compliance is used in complying with this AD, that equivalent must be approved by the Chief, Aircraft Certification Staff, AEU-100, Federal Aviation Administration, Europe, Africa, and Middle East Office, Brussels, Belgium. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Short Brothers Limited, P.O. Box 241, Airport Road, Belfast BT3 9DZ, Ireland, Attention: Product Support Manager. These documents may be examined at FAA Headquarters, Room 916, 800 Independence Avenue, S.W., Washington, DC. This amendment becomes effective March 19, 1981.
58-19-01: 58-19-01 GRUMMAN: Applies to All TBF-1, TBF-1C, TBM-1, TBM-1C, TMB-3, TMB-3E Aircraft Certificated in the Limited or Restricted Category Compliance required as soon as possible but not later than December 1, 1958. As a result of a recent accident involving a powerplant fire caused by a broken aluminum flammable fluid carrying line, additional fire protection is required. In order to correct this condition, the following must be accomplished: Replace the two (2) carburetor vapor return lines, the one fuel pressure line, and the one oil pressure line within the powerplant zone with fire resistant flexible hose assemblies.
2012-10-53: We are adopting a new airworthiness directive (AD) for Eurocopter Deutschland GmbH (ECD) Model EC135 P1, EC135 P2, EC135 P2+, EC135 T1, EC135 T2, and EC135 T2+ helicopters. This is the Federal Register publication of an Emergency AD (EAD) that was previously sent to all known owners and operators of these helicopters. That EAD superseded an earlier related EAD. This AD requires, before further flight and at specified intervals, checking and inspecting the upper and lower main rotor hub (MRH) shaft flanges for a crack, and inspecting the lower hub-shaft flange bolt attachment areas for a crack. This AD is prompted by three reported incidents of cracking on the lower hub-shaft flanges of EC135 model helicopters. These actions are intended to detect a crack on the hub-shaft flange, which if not corrected could result in failure of the MRH and subsequent loss of control of the helicopter.
97-21-07 R1: This amendment revises an existing airworthiness directive (AD), applicable to AlliedSignal Inc. (formerly Textron Lycoming) Model T5313B, T5317A, and T53 series military turboshaft engines approved for installation on aircraft certified in accordance with Section 21.25 of the Federal Aviation Regulations (FAR), that currently requires a one-time visual inspection of accessory drive carrier assemblies for affected serial numbers (S/Ns) designating a defective assembly, and if the S/N is applicable, replacement with a serviceable assembly. This amendment adds military helicopter models and removes one civilian helicopter model to the sentence in the Applicability paragraph of the AD that provides guidance as to the helicopter models with the affected engines. This amendment is prompted by the need to revise the Applicability paragraph. The actions specified by this AD are intended to prevent accessory drive carrier assembly failure, which could result in an N2 overspeed andan uncontained engine failure.
98-13-08: This amendment adopts a new airworthiness directive (AD) that applies to certain Pilatus Aircraft Ltd. (Pilatus) Model PC-12 airplanes. This AD requires replacing and re-routing the power return cables on the starter generator and generator 2, inserting a temporary revision to the pilot operating handbook (POH), and installing a placard near the standby magnetic compass. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Switzerland. The actions specified by this AD are intended to prevent directional deviation on the standby magnetic compass caused by an overload of electrical current in the airplane structure, which could result in flight-path deviation during critical phases of flight in icing conditions and instrument meteorologic conditions (IMC).
98-13-04: This amendment adopts a new airworthiness directive (AD) that applies to all Glaser-Dirks Flugzeugbau GmbH (Glaser-Dirks) Models DG-100 and DG-400 gliders. This AD requires repetitively inspecting the airbrakes to assure they retract at their outboard end first, and repairing the airbrakes if they do not retract at their outboard end first; and repetitively inspecting the airbrake torque tube in the fuselage for cracks or deformations, and reinforcing or replacing, as necessary, if cracks or deformations are found in the airbrake torque tube. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. The actions specified by this AD are intended to prevent overloading of the airbrake control system caused by free play between the bellcrank and airbrake plate, which could result in failure of the operating lever of the airbrake torque tube in the fuselage.
81-23-01 R1: 81-23-01 R1 BEECH: Amendment 39-4289. Applies to the following model airplanes regardless of the category or categories of airworthiness certification: MODELS SERIAL NUMBER (S/N)* 65, A65 & A65-8200 LC-181 through LC-335 70 LB-1 through LB-35 65-A80, 65-A80-8800 & 65-B80 LD-151 through LD-511 and LD-34, LD-46, LD-119 65-A88, 65-88 LP-1 through LP-54 65-90, 65-A90, B90 & C-90 LJ-1 through LJ-929 E90 LW-1 through LW-342 99, 99A, B99 U-1 through U-164 100 & A100 B-1 through B-247 B100 BE-1 through BE-102, and BE-104 Military: L23F** ** LF-7 through LF-76** 65-A90-1 LM-1 through LM-144 65-A90-2 LS-1, -2, -3 65-A90-3 LT-1, -2 65-A90-4 LU-1 through LU-16 NU-8F LG-1 *Except that airplanes which have installed BEECHCRAFT Kit No. 90-4077-1 S, BEECHCRAFT Kit No. 99-4023-1 S, or Aviadesign Supplemental Type Certificate SA1178CE or SA1583CE are not affected by this AD. **Except that Model L23F airplanes which do not have a preload indicating washer assembly (i.e., one with radial holes in a center ring) are not affected by this AD. COMPLIANCE: Required as indicated, unless already accomplished. In order to assure integrity of bolts and nuts at the lower forward attachments of outer wing panels to the wing center section, accomplish the following: A) Prior to next flight, accomplish all of the following: 1. Remove all bolts, washers, and nuts from each lower forward wing attachment and thoroughly clean each removed part. Throughout all action required by this AD: a. Use procedures in the applicable Beech Maintenance Manual except where other procedures are specified by this AD, b. Unless different instruction from Beech Aircraft Corporation is obtained and followed, reposition wing, as necessary, to remove or reinstall bolt by hand without using any tool, c. Keep parts of each preload indicating washer assembly together so that parts of one assembly cannot be intermingled with parts of another assembly, d. Clean each removed part with naphtha or methyl ethyl ketone (MEK) using a bristle brush, and repeat this cleaning as necessary prior to each subsequently specified action until lubricant is applied, and e. Accomplish all of the specified actions on both (i.e., left and right) sides of the airplane. 2. Visually inspect each bolt and nut for reddish rust. Do not classify copper residue over cadmium plating as rust. For a bolt, rust is acceptable only on the end (including not more than one thread) farthest from the head and within counterboard recess between wrench serrations of the bolt head. For compliance with Paragraph A)6 and C), below, classify a bolt as rusted if rust is found elsewhere. Classify a nut as rusted if rust is found anywhere. 3. Visually inspect each bolt and nut for a pit or crack in steel (not cadmium or copper plating) material. Use 10X or stronger magnifying glass. For each bolt, pay particular attention to the fillet and shank, including threads. For each nut, pay particular attention to the chamfer (that faces the bolt head when installed) and perceptible threads adjacent to this chamfer. (Refer to Paragraphs A)6 and C) below.) 4. Bake each bolt and nut continuously for 23 hours at 350 degrees to 400 degrees Fahrenheit and cool in still air. 5. After accomplishment of Paragraph A)4, above, use a magnetic particle method of Advisory Circular AC43.13-1A to inspect each bolt and nut for a crack, paying particular attention to locations specified in Paragraph A)3, above. For each bolt, use a fluorescent particle method with 5250 to 6750 ampereturns in a coil to produce longitudinal magnetization in each bolt. (6,000 ampereturns means 2,000 amperes in a 3-turn coil or 1,000 amperes in a 6-turn coil, etc.) For each nut, use any magnetic particule method with 500 to 700 amperes through a central conductor of at least 0.6-inch diameter through two nuts to produce circular magnetization. Demagnetize each bolt and nut after the above inspection. 6. Replace each rusted, pitted, and/or cracked nut and bolt with a new Part Number (P/N) as follows: a. If new preload indicating (PLI) washer assembly is to be used in accordance with Paragraph A)9, below, nut P/N is 72789-1414, 72789M-1414, FN22-1414, or FN22M-1414. ("M" in P/N denotes black coating. All eligible nuts have a locking feature which necessitates use of a wrench for full engagement with bolt.) b. If a used PLI washer assembly is reinstalled in accordance with Paragraph A)9, below, nut P/N is 72789-1414 or FN22-1414. c. Bolt is P/N LWB 22-14-XX or VEP 220121-14-XX where XX is 31 for airplanes with S/N LD-34, LD-46, LD-119, and LJ-1 through LJ-67, and XX is 32 for all other airplanes affected by this AD. Replace preload indicating washer with new P/N 61475-14-43.5 assembly (not any other P/N) if this assembly is available. Obtain new parts only from BeechService Centers or Beech Aircraft Corporation. (Neither baking nor field inspection of new parts is necessary.) Do not replate any part. 7. Clean the bore and recessed washer seat area of the outboard and inboard wing fittings with naphtha or methyl ethyl ketone (MEK). Visually inspect these areas for corrosion, burrs, gouges and coining. If any defect is found, contact Beech Aircraft Service Department, 9709 East Central, Wichita, Kansas 67201; telephone (316) 681-7261, 7278, or 7352, for rework disposition. Also, if any defect is found, treat the bore and recessed washer seat areas of the inboard and outboard wing fittings with Alodine 1200, 1200S, or 1201. Allow the alodine coating to dry for 5 minutes. Wash the coating with water and blow dry with air without wiping. Paint treated washer seat areas with zinc chromate primer (obtain locally) and allow primer to dry. 8. Coat the inspected areas of the wing fittings, all of each bolt, all of each nut, and all of each preload indicating washer assembly with either clean MIL-C-16173, Grade 2 corrosion preventative compound or clean General Electric G322L Versilube Silicone Lubricant. 9. Install removed or new parts using standard procedures except as follows: a. Preload indicating (PLI) washer assembly may be reused with P/N 72789-1414 and/or P/N FN22-1414 nuts, only. b. Ascertain that a radius of the adjacent washer is next to the fillet under the bolt head and next to the outer edge of the recess in each wing fitting. Position wing as necessary to allow bolt to slide into fitting without use of any tool. c. Tighten the joint by rotating the nut (do not turn the bolt). Use standard procedure if new PLI washer assembly is installed. If used PLI washer assembly is reinstalled, make necessary correction for any torque wrench adapter and apply 3250 to 3400 inch-pounds torque, but install new PLI washer assembly if center ring of the used assembly turns after 3400 inch-pounds torque is applied. Do not allow wrench to bear against fitting. d. Coat entire portion of bolt that projects beyond nut, using a material that is specified in Paragraph A)8, above. e. Make aircraft maintenance record entry showing work accomplished, especially procedure used for tightening nut, and whether new or used PLI washer was installed. Indicating washer assembly with either clean MIL-C-16173, Grade 2 corrosion preventative compound or clean General Electric G322L Versilube Silicone Lubricant. B) Between 90 and 110 hours time-in-service after accomplishment of action specified by Paragraph A) of this AD, check nut tightness, using the same procedure that was used for accomplishment of Paragraph A)9c, above. C) Within 3 days after replacing a part in accordance with Paragraph A)6, above, or noting a defect when complying with this AD, submit a written report to the Federal Aviation Administration via an FAA M or D Report (FAA Form 8330-2) or a letter to the office specified in Paragraph E), below and send the replaced part(s) to Beech Aircraft Corporation. In the submitted report, please advise date of last previous bolt removal. D) A special flight permit in accordance with Federal Aviation Regulation 21.197 for flight to the nearest base is permitted in order to accomplish Paragraph A) of this AD. The nearest FAA Flight Standards District Office may be contacted to obtain a telegraphic special flight permit. E) 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, 7001, or 7002. This amendment becomes effective on January 4, 1982, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated October 31, 1981.
98-13-05: This amendment adopts a new airworthiness directive (AD) that applies to certain Alexander Schleicher Segelflugzeugbau (Alexander Schleicher) Model AS-K13 sailplanes. This AD requires inspecting the main spar fitting for excessive tolerance, traces, movement, etc., and repairing the main spar fitting if any of the above conditions exist. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. The actions specified by this AD are intended to prevent failure of the main spar caused by excessive movement of the main spar fitting, which could result in loss of control of the sailplane.
2012-22-15: We are superseding an existing airworthiness directive (AD) for all Fokker Services B.V. Model F.28 Mark 0070 and 0100 airplanes. That AD currently requires revising the airworthiness limitations section (ALS) of the instructions for continued airworthiness for certain airplanes, and the FAA-approved maintenance program for certain other airplanes, to incorporate new limitations. This new AD requires revising the maintenance program to incorporate the limitations, tasks, thresholds, and intervals specified in certain revised Fokker maintenance review board (MRB) documents. This AD was prompted by a revised Fokker 70/100 MRB document with revised limitations, tasks, thresholds, and intervals. We are issuing this AD to reduce the potential of structural failures or of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane.
84-17-02: 84-17-02 GATES LEARJET: Amendment 39-4902. Applies to Gates Learjet Model 24 series airplanes certificated in all categories. Compliance required within eighteen (18) months after effective date unless previously accomplished or prior compliance with AD 81-16-08, Amendment 39-4546. Accomplish the requirements of this AD at an FAA approved repair station. The modification and inspection of the horizontal stabilizer trim actuator must be performed by an FAA certificated repair agency authorized to overhaul and test the Gates Learjet Horizontal Stabilizer Trim Actuator. The necessary shop equipment or the equivalent, as referenced in Learjet Repair Manual Number 1711-9, is listed in Attachment I hereto. A. Modify Learjet Model 24 airplane flight control systems, stall warning, and control wheel by incorporating the airplane modification kit listed in Table I. Table I MODIFICATION KIT MODEL AMK 82-5 24, 24A AMK 82-1 24B, 24B-A AMK 81-18 24D, 24D-A AMK81-13 24E, 24F, 24F-A NOTE: Modification of JET Autopilot Controllers and Computers as required by the instructions in the above kits must be performed in an FAA approved facility for maintenance of the JET FC-110 autopilot. B. Required Airplane Maintenance Record entry must be accomplished by the facility performing its portion of the AD as prescribed in paragraph A. of this AD. C. After airplane modification, verify that the Airplane Flight Manual (AFM) contains the revision/change listed in Table II, below, or later FAA approved revision/change. Upon completion of the modifications required by paragraph A. of this AD, the more restrictive paragraphs A.2., A.5., and A.6. of AD 80-19-11 are no longer applicable. Table II Gates Learjet Airplane Flight Manual/Supplement Revision/Change Modification kit 1. 24 AFM, FM- 004 Revision 32 AMK82-5 2. 24 RAS AFM, W0159 Revision 4 AMK82-5 3. 24A AFM, FM-005 Revision 18 AMK82-5 4. 24E CR 736 AFM, FM-008 Revision 7 AMK82-5 5. FC-110 Autopilot, W1037 New AMK82-5 6. 24B AFM, FM-006 Revision 21 AMK82-1 7. 24B-A AFM, FM-007 Revision 21 to 24B AMK82-1 8. 24B RAS AFM, W0157 Revision 3 AMK82-1 9. FC-110 Autopilot, W1041 New AMK82-1 10. 24D AFM, FM-009 Change 18 AMK81-18 11. 24D-A AFM, FM-010 Change 18 to 24D AMK81-18 12. 24D RAS AFM, W-0119 Change 3 AMK81-18 13. FC-110 Autopilot, W1030 New AMK81-18 14. 24E AFM, FM-011 Change 9 AMK81-13 15. 24F AFM, FM-012 Change 7 AMK81-13 16. FC-110 Autopilot, W1018 New AMK81-13 D. Prior to accomplishing the modification required by paragraph A. of this AD, contact the Wichita Aircraft Certification Office, FAA, Central Region, telephone (316) 269- 7008, if any modification or alteration has been performed on the affected airplane systems, for further instruction relative to the compatibility of the modification and this AD. E. Alternate methods of compliance with this AD may be used if they are approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. This amendment becomes effective October 5, 1984. ATTACHMENT I The stabilizer actuator test stand (P/N ST-00463) is used to functionally test the stabilizer actuator after overhaul. The physical structure of the test stand must be capable of withstanding a minimum load of 2500 lbs. without any bending or deformation. The stabilizer actuator is vertically mounted on the test stand with one end stationary and the other end movable through a hydraulic actuator. The test stand consists of the following components: a. Hydraulic Actuator - The hydraulic actuator is capable of applying a regulated load of 0 to 2500 lbs. on the stabilizer actuator during the entire extend or retract cycles. b. Hydraulic Pressure Regulator - The pressure regulator is used to select hydraulic pressures applied to the stabilizer actuator during the functional test. c. Hydraulic Pressure Gauge - The hydraulic pressure gauge is used to monitor hydraulic pressure applied to the stabilizer actuator. The gauge must be certified at least monthly. d. Digital Position Readout - The digital position readout indicator is used to monitor the travel of the stabilizer actuator. Signals to the indicator are picked up from a rigid mounted linear potentiometer and movable wiper attached to the hydraulic actuator. The digital readout is accurate to 1/1000th of an inch. e. Linear Scale - A linear scale, graduated in 100th of an inch, is permanently mounted on the test stand to verify the digital readout. A tool of known length is used to verify the linear scale and digital readout before the stabilizer actuator functional test is performed. The tool length must be certified at least yearly. f. Lapse Timer - A lapse timer is coupled to the control switches and the stabilizer actuator to monitor travel time during the extend and retract cycles. The lapse timer must measure seconds to beaccurate to 1/100th of a second. g. Trim Controller - The trim controller is used to simulate two-speed input to the stabilizer actuator primary motor. The trim controller part number is EM 2079-6. h. Pre-Select Timer - The pre-select timer is used to check stabilizer actuator travel vs. time, voltage, and amperage inputs in accordance with the functional test. i. Power Supply - The power supply is variable through 0-30 volts DC and 0-30 amperes DC. j. DC Voltmeter - The DC voltmeter must be capable of measuring 0-30 volts DC and must be certified at least yearly. The voltmeter is used to monitor the voltage inputs to the stabilizer actuator in accordance with the functional test. k. DC Ammeter - The DC ammeter must be capable of measuring 0-30 amperes DC and must be certified at least yearly. The ammeter is used to monitor the amperes inputs to the stabilizer actuator in accordance with the functional test. l. Millivolt Meter - The millivolt meter is used to monitor the stabilizer actuator linear potentiometer for a smooth and steady signal output. The meter is 0-50 volts graduated in 100 mv increments. m. Switches - Necessary switches installed to operate the stabilizer actuator primary and secondary motors to extend or retract. n. A digital or Simpson 260 meter, not a part of the test stand, is used to verify the resistance of the stabilizer actuator linear potentiometer. The digital or Simpson 260 meter must be certified at least every 90 working days.
2022-08-07: The FAA is adopting a new airworthiness directive (AD) for all Embraer S.A. Model ERJ 170-100 STD, -100 LR, -100 SU, and -100 SE; ERJ 170-200 STD, -200 LR, -200 SU; ERJ 190-100 STD, -100 LR, -100 IGW, and -100 ECJ; and ERJ 190-200 STD, -200 LR, and -200 IGW airplanes. This AD was prompted by a report of the failure of the inner pane of certain passenger windows to meet maximum operating pressure and lack of fail- safe design. This AD requires determining if certain NORDAM passenger windows are installed, and performing corrective actions if any affected part is installed. This AD also prohibits the installation of affected parts. The FAA is issuing this AD to address the unsafe condition on these products.
2010-11-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: During ERJ 170 airplane full scale fatigue test, cracks were found in some structural components of the airplane. Analysis of these cracks resulted in modifications on the airplane Airworthiness Limitation Items (ALI), to include new inspections tasks or modification of existing ones and its respective thresholds and intervals. Failure to inspect these components according to the new tasks, thresholds and intervals, could prevent a timely detection of fatigue cracks. Undetected fatigue cracks in these areas could adversely affect the structural integrity of these airplanes. * * * * * We are issuing this AD to require actions to correct the unsafecondition on these products.
68-17-08: 68-17-08\tBOEING: Amendment 39-640 as amended by Amendments 39-670, 39-852, 39- 891, and 39-923 is further amended by Amendment 39-1420. Applies to all Boeing 707/720 series aircraft listed in Boeing Service Bulletin 1995 (Revision 5) dated September 28, 1967 or later FAA approved revision. \n\tCompliance required as indicated. \n\tTo detect cracking and prevent failure of the lower wing skin at front spar station 392, accomplish the following: \n\t(a)\tInspect the lower wing skin of aircraft which have not been repaired by installation of the small repair doubler (identified on Page 25, Boeing Service Bulletin 1995, Revision 5), for cracks emanating from the two outboard fasteners of the splice plate tab as noted in Figure 1 of Boeing Service Bulletin 1995, (Revision 5 or later FAA approved revisions) by the use of either a dye penetrant or an eddy current inspection technique, or an equivalent inspection method approved by the Chief, Aircraft Engineering Division, FAA Western Region, atthe times specified in (h), (i), (j) or (k) as appropriate and, if cracks are found, repair prior to further flight per (f) or (g). \n\t(b)\tInspect the lower wing skin of aircraft which have been repaired by installation of the small repair doubler in accordance with Boeing Service Bulletin 1995, within 1600 hours (for 720 Series) or 2000 hours (for 707 Series) after installation or within the next 400 hours (for 720 Series) or 600 hours (for 707 Series) time in service after the effective date of this AD, unless inspected within the previous 1200 (for 720 Series) or 1400 (for 707 Series) hours time in service and at intervals thereafter not to exceed 1600 (for 720 Series) or 2000 (for 707 Series) hours time in service, per (e). \n\t(c)\tInspect the lower wing skin for cracks emanating from the attachments of the front spar support fitting as noted in Figure 1 of Boeing Service Bulletin 1995 (Revision 7) dated 20 August 1969 or later FAA-approved revisions, at the times specified in (h),(i), (j), or (k) as appropriate, and, if cracks are found, repair prior to further flights per (g). The initial inspection must be accomplished either by means of a dye penetrant technique or in accordance with the eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region. The eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revision, or an equivalent FAA-approved inspection technique, must be used for all inspections thereafter. \n\t(d)\tOn those aircraft which have not had the drag fitting trimmed and the fairing attach angle modified in accordance with Boeing Service Bulletin 1995 (Revision 7) or later FAA- approved revisions, within the next 400 hours (for 720 Series) or 600 hours (for 707 Series) time in service after the effective date of this AD and thereafter at intervals not to exceed 800 hours (for 720 Series) or 1,200 hours (for 707 Series) time in service, inspect for cracks in the lower wing skin, emanating from the forward fastener for the drag fitting and from the fasteners for the fairing attach angle as noted in Figure 1 of Boeing Service Bulletin 1995 (Revision 7) or later FAA-approved revisions, at the threshold times as specified in (h), (i), (j), or (k) as appropriate. The initial inspection must be accomplished either by means of a dye penetrant technique or by use of eddy current inspection techniques described in S.B. 1995 (Revision 7), or later FAA- approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region. The eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revision, must be used for all inspections thereafter. If cracks are found around the fairing attach angle or emanating aft from the drag fitting fastener, rework the drag fitting, double, andskin prior to further flight in accordance with (g). If cracks are found emanating forward from the drag fitting fastener, rework the drag fittings, doubler and skin, prior to further flight in accordance with Boeing Service Bulletin 1995 (Revision 7) or later FAA-approved revision, or in accordance with an equivalent rework or modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(e)\tInspect the lower wing skin covered by the small repair doubler for cracks by use of the x-ray inspection techniques noted in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) or an equivalent inspection technique approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat inspections at intervals not to exceed 1600 hours (for 720 Series) or 2000 hours (for 707 Series) time in service. If crack growth is found, repair prior to further flight in accordance with (g). \n\t(f)\tIf the cracks fall within the crack length limitsoutlined in the paragraph titled "Installation of the small repair doubler," (Part II, Boeing Service Bulletin 1995, Revision 5 or later FAA-approved revisions) repair in accordance with that section of the bulletin or later FAA- approved revisions. Within 1600 hours (for 720 Series) or 2000 hours (for 707 Series) after installation of the doubler, inspect in accordance with (e). \n\t(g)\tUpon completion of any of the following modifications, the inspections required by this AD may be discontinued: \n\t\t(1)\tInstallation of 65-56257-1-2 or 65-57788-1-2 doublers as appropriate, per Boeing Service Bulletin 1995 (Revision 5 or later FAA-approved revisions). \n\t\t(2)\tBoeing Service Bulletin 2484. \n\t\t(3)\tBoeing Service Bulletin 2487. \n\t\t(4)\tInstallation of the 720 Wing Structural Improvement Program (Per Boeing Document 65-12700) accomplish at Boeing's Wichita facility. \n\t\t(5)\tAn equivalent installation approved by the Chief, Aircraft Engineering Division, Western Region, FAA. \n\t(h)\tFor those airplanes listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part I, and having less than 6000 (for 720 Series) or less than 10,000 (for 707 Series) hours time in service on the effective date of this AD, prior to the accumulation of 6800 (for 720 Series) or 11,200 (for 707 Series) hours time in service, respectively, and thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service from the last inspection. \n\t(i)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part I, and having 6000 or more (in the case of 720 Series aircraft) or 10,000 or more (in the case of 707 Series aircraft) hours time in service on the effective date of this AD, within the next 400 (for 720 Series) or 600 (for 707 Series) hours time in service, unless accomplished within the last 400 (for 720 Series) or 600 (for 707 Series) hours time in service, and at intervals thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service. \n\t(j)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part II, and having less than 10,000 (in the case of 720 Series) or less than 15,000 (in the case of 707 Series) hours time in service on the effective date of this AD, prior to the accumulation of 10,800 or 16,200 hours time in service, respectively, and thereafter at intervals not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service from the last inspection. \n\t(k)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part II, and having 10,000 or more (in the case of the 720 Series) or 15,000 or more (in the case of the 707 Series) hours time in service on the effective date of this AD, within the next 400 (for 720 Series) or 600 (for 707 Series) hours time in service unless accomplished within the last 400 (for 720 Series) or 600 (for 707 Series) hours time in service, and at intervals thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service. \n\t(l)\tAirplanes having cracks which require rework under this AD may be flown in accordance with FAR 21.197 with the concurrence of Chief, Aircraft Engineering Division, FAA Western Region, to a base where the rework can be accomplished. \n\t(m)\tUpon request of an operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\tAmendment 39-640 became effective September 23, 1968. \n\tAmendment 39-670 became effective October 18, 1968. \n\tAmendment 39-852 became effective November 1, 1969. \n\tAmendment 39-891 became effective December 16, 1969. \n\tAmendment 39-923became effective January 17, 1970. \n\tThis Amendment 39-1420 becomes effective April 1, 1972.