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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.