Results
75-09-10: 75-09-10 PIPER: Amendment 39-2183. Applies to Models PA-31P, Serial Nos. 31P-1 through 31P-7400213, PA-31 and PA-31-300, Serial Nos. 31-1 through 31-7401248, and PA-31- 350, Serial Nos. 31-5001 through 31-7405242 and 31-7405400 through 31-7405462 aircraft certificated in all categories. To prevent possible hazards in flight associated with loose elevator push-pull tube end fittings, accomplish the following: 1. Within the next 100 hours in service from the effective date of this AD unless already accomplished, inspect the elevator push-pull tube P/N 40847-04 for looseness in its attachment to the end fittings in accordance with the instructions contained in Piper Service Bulletin No. 409 dated June 7, 1974, for Model PA-31P or Piper Service Bulletin No. 433 dated December 17, 1974, for Models PA-31, PA-31-300, and PA-31-350 or an equivalent inspection. 2. If looseness is detected, replace with an acceptable elevator torque tube assembly P/N 40847-04 on Model PA-31Por P/N 40847-00 on Models PA-31, PA-31-300 and PA31-350. 3. Equivalent inspections must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective April 28, 1975.
83-04-05: 83-04-05 DeHAVILLAND: Amendment 39-4568. Applies to Model DHC-3 (U-1A) airplanes certificated in any category. COMPLIANCE: As prescribed in the body of the AD. To prevent loss of primary flight controls due to fracture of the control column lower weld assembly, accomplish the following: a. Within the next 25 hours time-in-service after the effective date of this AD, unless already accomplished within the last 75 hours, and at each 100 hours time-in-service thereafter, visually inspect the control column lower weld assembly, P/N C3CF 39-15 or C3CF 39-13 for cracks in accordance with the Accomplishment Section of DeHavilland Service Bulletin S/B No. 3/39 dated October 1, 1982. b. If no cracks are found, return the airplane to service and repeat the inspections required by paragraph a of this AD. c. If cracks are found in control column lower assembly P/N C3CF 39-13, replace the assembly prior to further flight with a serviceable assembly P/N C3CF 39-15 or P/N C3CF 39-21 (Modification 3/930) as specified in the Replacement Section of said Service Bulletin. d. If cracks longer than 1 inch are found in control column lower assembly P/N C3CF 39-15, repair or replace the assembly prior to further flight as instructed in the Accomplishment Instructions Section of said Service Bulletin. e. If cracks 1 inch or less in length are found in control column lower assembly P/N C3CF 39-15, repair or replacement may be delayed for a period of 150 hours time-in-service provided the inspection prescribed in paragraph a above is accomplished at 25-hour intervals. f. The repetitive inspections of paragraph a of this AD may be discontinued when control column lower assembly P/N C3CF 39-21 (Modification 3/930) is installed. g. The airplane may be flown in accordance with Federal Aviation Regulation 21.197 to a location where this AD can be accomplished. h. An equivalent method of compliance with this AD may be used if approved by theManager, New York Aircraft Certification Office, ANE-170, Federal Aviation Administration, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581; telephone (516) 791-6680. This amendment becomes effective on March 1, 1983.
81-18-04 R2: 81-18-04 R2 AVCO LYCOMING: Amendment 39-4199 as amended by Amendment 39-4258 is further amended by Amendment 39-4395. Applicability: All O-235, O-290-D, -D2, O-320, IO-320, AIO-320, AEIO-320, LIO-320, O-340, O-360, IO-360, AIO-360, AEIO-360, HO-360, HIO-360, LO-360, LIO-360, TIO-360, TO-360, LTO-360, VO-360, IVO-360, O-540 and IO-540 series engines, except for the following: O-320-H2AD, O-360-E1A6D, LO-360-E1A6D, TO-360-E1A6D, LTO-360-E1A6D, IO-540-P1A5, IO-540-R1A5, IO-540-S1A5 and; O-540/IO-540 series engines built with large capacity oil pumps and dual magnetos designated with "5D" in the model suffix (Example: IO-540-K1A5D). Compliance required as indicated unless already accomplished. To prevent failure of engine oil pumps which incorporate sintered iron impellers, accomplish the following: (a) Compliance is required within the next 25 hours in service after the effective date of this AD for all Lycoming HIO-360-D1A, -E1AD, -E1BD and -F1AD up to and including serial number L-22579-51A except the following: L-22311-51A thru L-22313-51A, L-22396-51A, L-22397-51A, L-22416-51A, L-22546-51A thru L-22549-51A, L-22563-51A, L-22568-51A thru L-22571-51A, and in addition all of the above engines that were overhauled in the field prior to April 1, 1981; all remanufactured engines of the above model shipped prior to April 1, 1981, regardless of serial numbers. (1) Replace the oil pump driven impeller and shaft with hardened steel impeller and shaft P/N LW-18110 and replace the driving impeller with impeller P/N LW-18109 in accordance with the instructions set forth in AVCO Lycoming Service Bulletin No. 454 dated April 10, 1981, or approved alternate method, unless it can be established that a sintered iron impeller is not installed (see NOTE below). (b) Compliance is required within the next 25 hours in service after the effective date of this AD for: Lycoming models O-360-A1LD S/Ns L-17555-36A through L-22462-36A, O-360-A1F6D S/Ns L-16685-36Athrough L-22582-36A, O-360-A5AD S/Ns L-17057-36A through L-20038-36A, IO-360-A1B6D/-A3B6D S/Ns L-9598-51A through L-16595-51A, L-17273-51A, L-17312-51A through L-17319-51A, L-17321-51A, L-17336-51A through L-17340-51A, L-17347-51A through L-17351-51A, L-17355-51A, L-17358-51A, L-17377-51A through L-17380-51A, IO-360-C1E6D, S/N L-14527-51A, TO-360-C1A6D S/Ns L-101-69A through L-243-69A, and in addition, all of the above model engines which were overhauled in the field between April 7, 1970, and October 15, 1976, regardless of serial numbers; and all of the above model engines which were remanufactured and shipped before April 1, 1981, regardless of serial numbers. (1) Replace the existing drive and driven impellers with a steel driving impeller P/N 60746 and an aluminum impeller and shaft assembly P/N LW-13775 in accordance with AVCO Lycoming Service Bulletin No. 455A dated April 24, 1981, or approved alternate, unless it can be established that a sintered iron impeller is not installed. (See NOTE below.) (c) For all other engines in the subject Applicability paragraph not specifically listed in Paragraphs (a) and (b) above, comply with Avco Lycoming Service Bulletin No. 456 dated August 21, 1981, or FAA approved revision or alternate, at 2000 hours since new or since last overhaul, whichever is later, or whenever the accessory section is removed. Those engines which have accrued 2000 hours or more on the effective date of this AD must comply within the next 100 hours in service. Compliance is required as described herein unless it can be established that a sintered iron impeller is not installed. (See NOTE below.) Alternate methods of compliance must be approved by the Chief, New York Aircraft Certification Office. Upon submission of substantiating data by an owner or operator through an FAA Maintenance inspector, the Chief, New York Aircraft Certification Office may adjust the compliance time specified in this AD. In accordance with FAR 21.197and 21.199, the aircraft may be flown to a location where the alterations required by this AD can be performed. NOTE: Engines originally manufactured prior to 1970 did not incorporate sintered iron impellers. For these engines, reference should be made to engine maintenance/overhaul logbook records, Lycoming build records, and pertinent Service Bulletins. Service Bulletin Nos. 381C and 385C describe a method to determine if the early design oil pump with aluminum/steel impellers is installed. Aluminum/steel impellers do not require replacement. Amendment 39-4199 became effective September 14, 1981. Amendment 39-4258 became effective November 19, 1981. This amendment 39-4395 becomes effective June 7, 1982.
75-08-03: 75-08-03 PIPER AIRCRAFT CORPORATION: Amendment 39-2148. Applies to PA- 28-140, PA-28-150, PA-28-151, PA-28-160, PA-28-S-160, PA-28-180, PA-28-S-180, airplanes serial numbers 28-03, 28-1 thru 28-7525201; PA-28R-180, PA-28R-200 airplanes serial numbers 28R-30002 thru 28R-7535143; certificated in all categories. To detect and correct unsecured fuel drain valves installed at the gascolator, accomplish the following or an equivalent method approved by Chief, Engineering and Manufacturing Branch, Southern Region: (a) Prior to next flight, unless already accomplished, check the fuel gascolator drain valve, (Piper Part No. 492-022) to determine if it is tight in the fitting by attempting to turn the valve counter-clockwise by hand. This check may be performed by the pilot. 1. If the valve is loose, tighten by turning clockwise, check the drain for proper operation and tighten the lock seal nut (Piper Pat No. 477-677), if installed, all in accordance with FAR 43. (b)Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished, secure the valve in the gascolator fitting as follows: 1. Remove the fuel gascolator drain valve from the gascolator AN 915-1D fitting, (Piper Part No. 458-946), if installed. NOTE: Do not remove the valve if Loc-tite has been previously applied. 2. Unless already accomplished, install a lock seal nut (Piper Part NO. 477- 677). Install with seal toward AN 915-1D fitting. 3. Apply one (1) drop of Loc-tite (88-31) sealant, coating only the 2nd, 3rd, and 4th threads evenly. NOTE: Do not allow Loc-tite to enter the Fuel System. 4. Install drain valve into AN915-1D fitting with a minimum of 2 1/2 turns and a maximum of 3 1/2 turns. 5. Secure valve in place with lock seal nut. Piper Service Bulletin 450 pertains to this subject. The amendment becomes effective April 4, 1975.
2019-25-10: We are adopting a new airworthiness directive (AD) for all Fokker Services B.V. Model F28 Mark 0070 and 0100 airplanes. This AD was prompted by reports of fuselage bottom [[Page 437]] skin exfoliation corrosion, fuselage skin bulging and cracking, and missing fastener heads. This AD requires a detailed inspection of the fuselage bottom skin for corrosion; skin cracks or bulges; and missing, loose, or broken fasteners; and, depending on the findings, accomplishment of applicable repairs; as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. We are issuing this AD to address the unsafe condition on these products.
81-18-01: 81-18-01 BELL: Amendment 39-4192. Applies to Models 206A, 206B, 206A-1, 206B-1, 206L, and 206L-1 helicopters, equipped with main rotor trunnions, P/N 206-010-104-3, 206-011- 113-001, 206-011-120-001, 206-011-113-103, and 206-011-120-103, certificated in all categories (Airworthiness Docket No. 81-ASW-27). Compliance required as indicated. To prevent possible failure of the main rotor trunnion, P/N's 206-010-104-3, 206-011- 113-001, 206-011-120-001, 206-011-113-103, and 206-011-120-103, due to fatigue cracks, accomplish the following, unless already accomplished: a. Main rotor trunnions, P/N 206-011-120-001, with 1,100 or more hours' time in service on the effective date of this AD must be retired from service within the next 100 hours' time in service. b. Main rotor trunnions, P/N 206-011-120-001, with less than 1,100 hours' time in service on the effective date of this AD must be retired from service prior to or on attaining 1,200 hours' time in service.c. Main rotor trunnions, P/N 206-010-104-3, 206-011-113-001, and 206-011-120- 103, with 2,300 or more hours' time in service on the effective date of this AD must be retired from service within the next 100 hours' time in service. d. Main rotor trunnions, P/N 206-010-104-3, 206-011-113-001, and 206-011-120- 103, with less than 2,300 hours' time in service on the effective date of this AD must be retired from service prior to or on attaining 2,400 hours' time in service. e. Main rotor trunnions, P/N 206-011-113-103, with 4,700 or more hours' time in service on the effective date of this AD must be retired from service within the next 100 hours' time in service. f. Main rotor trunnions, P/N 206-011-113-103, with less than 4,700 hours' time in service on the effective date of this AD must be retired from service prior to or on attaining 4,800 hours' time in service. g. The helicopter may be flown in accordance with FAR 21.197 to a base where compliance with this AD can be performed. h. The retirement times, for the main rotor trunnions, established by this AD, are as follows: Part Number Service Life-Hours 206-011-120-001 1,200 206-010-104-3 2,400 206-011-113-001 2,400 206-011-120-103 2,400 206-011-113-103 4,800 (Bell Helicopter Textron Alert Service Bulletins 206-80-7, Rev. B, dated October 15, 1980, and 206L-80-9, Rev. B, dated October 15, 1980, pertain to this subject.) This amendment becomes effective September 22, 1981.
80-13-01: 80-13-01 STRONG ENTERPRISES: Amendment 39-3793. Applies to all angled plastic parachute ripcord handles (P/N 1034) to which the ripcord cables are attached through only one leg of the handle and not attached through the drilled reinforcing crossbar in a lengthwise direction (see Figure 1). These handles were manufactured by Strong Enterprises in accordance with FAA Technical Standard Order (TSO) C-23b, Parachutes, for use on Strong Enterprises "Pop-Top" Chest-Mounted Reserve Parachutes (P/N 1023), but may be found on parachutes of other makes, models or types. \n\n\tCompliance required as indicated unless already accomplished. \n\n\tTo prevent the possible nondeployment of a parachute canopy due to separation of the plastic handle from the ripcord cable when subjected to the deployment pull force, accomplish the following: replace the plastic handle shown in Figure 1 with a metal handle (P/N 1025) supplied by Strong Enterprises prior to the parachute being made available for any parachute jump. \n\n\tCompliance with the provisions of this AD may be accomplished in an equivalent manner approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. \n\n\tThis amendment becomes effective June 16, 1980.
2019-23-16: The FAA is adopting a new airworthiness directive (AD) for all The Boeing Company Model 737-100, -200, -200C, -300, -400, and -500 series airplanes. This AD was prompted by a report of a fuel leak resulting from a crack on the left in-spar upper wing skin. This AD requires repetitive surface high frequency eddy current (HFEC) inspections of the left and right upper wing skin for any crack, repetitive general visual inspections of the upper wing skin in the adjacent rib bay areas for any crack, and applicable on-condition actions. The FAA is issuing this AD to address the unsafe condition on these products.
81-16-05: 81-16-05 SLICK ELECTRO, INC.: Amendment 39-4173. Applies to the following Slick Magneto models and serial numbers: Magneto Model Numbers* 4250, 4250R 4251, 4251R 4216, 4216R 4230, 4230R 4252, 4252R 6210, 6210R 4201, 4201R 4281, 4281R 6214, 6214R *All 4200 series magnetos use Slick Coil part number M-3114 All 6200 series magnetos use Slick Coil part number M-3009 Serial Numbers** 8100000-8109999 9050000-9059999 9110000-9119999 8110000-8119999 9060000-9069999 9120000-9129999 8120000-8129999 9070000-9079999 0010000-0019999 9010000-9019999 9080000-9089999 0020000-0029999 9020000-9029999 9090000-9099999 0030000-0039999 9030000-9039999 9100000-9109999 0040000-0049999 9040000-9049999 **Year and month of manufacture is given by first three numbers of serial number; for example 9032576 was manufactured March 1979. The above magneto models are installed on, but not limited to, the following engines: Teledyne-Continental A-65 C-90 O-470 A-75 O-200 IO-470 C-75 IO-360 IO-520 C-85 TSIO-360 TSIO-520 Lycoming O-235-C2C O-320-D3G O-360-A4M O-235-H2C O-320-E1J O-360-C1E O-235-K2C O-320-E2D O-360-C1F O-235-L2C O-320-E2G O-360-C2E O-320-A2D O-320-E3D O-360-F1A6 O-320-D1D AEIO-320-E1B AEIO-360-B1G6 O-320-D2G AEIO-320-E2B AEIO-360-H1A O-320-D2J O-360-A4K To prevent magneto failure due to cracked coil, accomplish the following within the next 25 hours time in service after the effective date of this AD unless already accomplished: A. Remove magneto and visually inspect coil for cracks in accordance with Slick Electro, Inc. Service Bulletin No. 1-81 revised June 29, 1981. B. Replace cracked coils and coils with less than 250 hours time in service with serviceable coils manufactured prior to October 1, 1978, or subsequent to April 30, 1980. The date of manufacture is stamped on each coil. C. Accomplishment of this AD should be indicated by stampingthe letter "C" into the metal name plate following the last digit of the magneto serial number, as well as the appropriate logbook entry. Alternate methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA Great Lakes Region. This amendment becomes effective August 6, 1981.
79-18-08: 79-18-08 MCDONNELL DOUGLAS: Amendment 39-3534. Applies to McDonnell Douglas Model DC-10-10, DC-10-10F, and DC-10-40 series airplanes certificated in all categories. \n\n\tCompliance required as indicated. To prevent failure of the aft engine mount barrel nuts on No. 1, No. 2 and No. 3 engines which could result in the loss of engine support, accomplish the following unless already accomplished subsequent to July 6, 1979: \n\n\ta)\tBefore further flight after August 29, 1979, visually inspect the aft engine mount barrel nuts on No. 1, No. 2 and No. 3 engines for evidence of cracking with particular regard to the area adjacent to the protruding bolt shank. \n\n\tb)\tBefore further flight after August 29, 1979, verify torque values of the aft engine mount bolts which mate with the barrel nuts on No. 1, No. 2 and No. 3 engines to the minimum value specified in the DC-10 Maintenance Manual. \n\n\tc)\tRevenue flight to the first base for the necessary equipment to accomplish (b) of this AD is authorized if such equipment is not available at the location of the aircraft on August 29, 1979. \n\n\td)\tReport results of all inspections to the Chief, Aircraft Engineering Division, FAA Western Region, within twenty-four hours of accomplishment in the following format: \n\n\t\t1)\t"N" number. \n\n\t\t2)\tHours' time-in-service and inspection. \n\n\t\t3)\tThe results of inspection by reference to specific paragraph of this AD. \n\n\te)\tIf cracks are discovered per paragraph (a) or if minimum torque values cannot be achieved per paragraph (b), replace the barrel nut with like serviceable part prior to further flight. \n\n\tThis amendment becomes effective August 29, 1979, and was effective earlier for all recipients of the telegram dated July 15, 1979.
2007-23-04: 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: It has been discovered in several cases that clamp bolts of the elevator spring tab mechanism were not installed in the correct orientation. Bolts have been found installed with bolt heads on the lower position and in two cases, some bolts, nuts and washers [hardware] were found to be loose or missing. Detachment of an elevator spring tab mechanism clamp bolt could lead to jamming of the elevator control system and reduced controllability of the aircraft. We are issuing this AD to require actions to correct the unsafe condition on these products.
2007-23-02: 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 ground inspection of an A340-311 aircraft, it has been discovered that 5 fasteners were missing between Frame (FR) 18 and FR19 on longitudinal joint at stringer 28RH (right hand). Further investigations have revealed that the missing fasteners have not been installed in production due to incorrect production instructions. If not corrected, this situation could affect the structural integrity of the aircraft in the area of stringer 28 between FR18 and FR19 at longitudinal joint. We are issuing this AD to require actions to correct the unsafe condition on these products.
72-22-05: 72-22-05 PIPER: Amendment 39-1545 as amended by Amendment 39-2052. Applies to PA-24, PA-24-250 and PA-24-260 airplanes certificated in all categories. To prevent possible adverse airplane vibration effects, accomplish the following: 1. Within the next 10 hours in service after the effective date of this Airworthiness Directive, unless already accomplished, attach the following operating limitation placard near the airspeed indicator in full view of the pilot: a. For PA-24 type airplanes, "Do not exceed 188 mph cas (Vne)". b. For PA-24-250 and PA-24-260 type airplanes, "Max. structural cruising: 167 mph cas (Vno). Do not exceed 188 mph cas (Vne)." 2. Within three (3) months after the effective date of this Airworthiness Directive, accomplish either: a. An alteration of the red radial Vne line and the cautionary yellow arc of the airspeed indicator to reflect the airspeeds noted in 1. above in accordance with an FAA-approved alteration; or b. An alteration of the rudder in accordance with Piper Service Kit No. 760705 or an FAA-approved equivalent alteration and an alteration of the airspeed instrument in accordance with an FAA-approved alteration to reflect the following speed restrictions: Vne of 202 mph (cas) for PA-24; of 203 mph (cas) for PA-24-250 and PA-24-260 Vno of 180 mph (cas) for PA-24-250 and PA-24-260. 3. For PA-24-250 and PA-24-260 type airplanes, a Vne of 227 mph (CAS) may be used upon altering the stabilator in accordance with Piper Service Kit No. 760747, or an approved equivalent, and by altering the rudder as in paragraph 2(b) above. (Piper Service Bulletin No. 687, dated June 19, 1974, refers to this subject.) 4. FAA approved alterations must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. Amendment 39-1545 was effective October 31, 1972. This amendment 39-2052 is effective December 26, 1974.
78-08-09: 78-08-09 GRUMMAN AMERICAN: Amendment 39-3191. Applies to Models G-164 (S/N 1 thru 400), G-164A (S/N 401 thru 1674), and G-164B (S/N 1B thru 79B) airplanes certified in all categories. Compliance required as indicated. To prevent collection of water in the bottom of the rudder main tubular spar (P/N A1203- 11) of the rudder assembly and the resulting corrosion, accomplish the following: 1. Within the next 50 hours in service after the effective date of this AD, unless already accomplished, perform the inspection and corrosion protection specified in Steps (1) thru (6) of Grumman American Service Bulletin No. 61 dated June 6, 1977, or equivalent. 2. Within the next 300 hours in service after compliance with paragraph (1) of this AD, and within every 300 hours in service thereafter, visually inspect the exterior of the main tubular spar for corrosion. If corrosion is noted, comply with Steps (5) and (6) of the bulletin or equivalent. 3. If repairs are required in (1)and (2) of this AD, they shall be in accordance with Advisory Circular 43.13-1A, Paragraph 74, Figure 2.7 or equivalent. 4. Upon request, with substantiating data submitted through an FAA Maintenance Inspector, the compliance times specified in this AD may be adjusted by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. 5. Equivalent methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective April 21, 1978.
76-18-01: 76-18-01 HUGHES HELICOPTERS: Amendment 39-2707. Applies to Model 269 series (including military Model TH-55A) helicopters certificated in all categories equipped with tail boom support strut assemblies P/N 269A2015 or P/N 269A2015-5, tail boom center attach fitting P/N 269A2324 and center frame aft cluster fittings P/N 269A2234(LH) and 269A2235(RH). Compliance required as indicated, unless already accomplished. To prevent fatigue failure of the tail boom support strut installation and resultant loss of the helicopter flight controllability accomplish the following: (a) For helicopters equipped with tail boom support strut assemblies P/N 269A2015 or 269A2015-5, within the next 50 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 50 hours time in service from the last inspection until the modifications of paragraph (c) are accomplished: (1) Support the tail boom and remove the tail boom support strut assemblies P/N 269A2015 on Hughes Model 269A, 269A-1, 269B and military TH-55A helicopter or P/N 269A2015-5 on Hughes Model 269C helicopter in accordance with the Hughes 269 Helicopter Basic Handbook of Maintenance Instructions Section 13; (2) Visually inspect the tail boom support strut aluminum end fittings for deformation or damage and visually inspect the tail boom support strut aluminum end fittings for cracks using the dye penetrant method in accordance with Hughes Service Information Notice N- 109.2 dated September 1, 1976 or later FAA-approved revisions. (b) If deformation, damage, or cracks are found during the inspections required by paragraph (a), before further flight, accomplish the modifications of either paragraph (c)(1) or (c)(2). (c) Within the next 500 hours time in service or one year whichever comes first, after the effective date of this AD: (1) Modify the tail boom support strut assemblies P/N 269A2015 and 269A2015-5 by replacing the aluminum end fittingswith stainless steel end fittings and attach bolts per Hughes Service Information Notice N-109.2 dated September 1, 1976 or later FAA- approved revisions, or (2) Replace the P/N 269A2015 strut assemblies with Hughes P/N 269A2015-9 and replace the P/N 269A2015-5 strut assemblies with Hughes P/N 269A2015-11. (d) For Hughes Model 269C helicopters, within the next 100 hours time in service, after the effective date of this AD, serialize the tail support strut assemblies P/N 269A2015-5 and P/N 269A2015-11 in accordance with Hughes Service Information Notice N-108 dated May 21, 1973 or later FAA-approved revisions. (e) Within 200 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 200 hours time in service from the last inspection until the modifications of paragraph (f)(2) are accomplished, inspect the tail boom center attach fittings P/N 269A2324 and center frame aft cluster fittings P/N 269A2234 (LH) and P/N 269A2235(RH) fordamage in accordance with Hughes Service Information Notice N-82.2 dated September 1, 1976 or later FAA-approved revisions. (f) If damaged parts are found during the inspections required by paragraph (e), before further flight: (1) Replace the damaged part with a serviceable part of the same part number, or (2) Replace damaged tail boom center attach fittings P/N 269A2324 with Hughes P/N 269A2324-7 and replace damaged center frame aft with cluster fittings P/N 269A2234(LH) and P/N 269A2235(RH) with Hughes P/N 269A2234-3(LH) and P/N 269A2235- 3(RH). (g) Equivalent procedures, inspections repairs, tail boom support strut assemblies, tail boom center attach fittings or center frame aft cluster fittings may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (h) Special flight permits may be issued to authorize operation of helicopter to a base for accomplishment of the inspections required by this AD per FAR's 21.197 and 21.199. This supersedes amendment 39-1587 (38 F.R. 2331), AD 73-03-01. This amendment becomes effective September 7, 1976.
79-10-12: 79-10-12 CANADAIR: Amendment 39-3471. Applies to all Canadair Models CL-44D4 and CL-44J certificated in all categories. Compliance required prior to application for U.S. registration and airworthiness certification. To preclude possible failure of the main landing gear actuating system, accomplish the following: a) Inspect the end caps of the main landing gear actuator, Jarry Hydraulics, P/N 3650-3 or -7 (Canadair Assembly numbers 44-75129-800 or -802 respectively) for cracks using the dye penetrant method or an FAA approved equivalent inspection. b) If cracks are found, replace end cap with a crack free cap prior to next flight. This amendment is effective May 18, 1979.
2019-24-18: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model 727 airplanes, Model 757 airplanes, and Model 767-200, -300, -300F, and -400ER series airplanes. This AD was prompted by reports of nuisance stick shaker activation while the airplane accelerated to cruise speed at the top of climb. This AD was also prompted by an investigation of those reports that revealed that the angle of attack (AOA) (also known as angle of airflow) sensor vanes could not prevent the build-up of ice, causing the AOA sensor vanes to become immobilized, which resulted in nuisance stick shaker activation. This AD requires a general visual inspection of the AOA sensors for certain AOA sensors, and replacement of affected AOA sensors. The FAA is issuing this AD to address the unsafe condition on these products.
74-22-05: 74-22-05 PIPER: Amendment 39-1991. Applies to Piper Models PA-23 and PA-23-160. Serial Nos. 23-1 to 23-2046 incl. and PA-23-235, PA-23-250 and PA-E23-250 Serial Nos. 27-1 to 27-3943 incl. except 27-3837. Compliance required within 100 hours or ninety days after the effective date of this AD, whichever occurs first. In order to prevent the heater fuel valve stem and cap nut from backing off and spilling fuel, accomplish the following: a. Inspect the stem and cap nut on Valves P/N's 17781-00 and 19460-00 for proper safetying and secureness in accordance with Apache Service Manual No. 752422, Section 13-138 or Aztec Service Manual No. 753564, Section 13-135 as applicable, and Piper Service Bulletin No. 401 dated May 21, 1974, or subsequent approved revisions, or equivalent methods approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. b. If the stem and nut are not secure or properly safetied, disassemble and inspect the stem, seat body and threads for damage in accordance with Section 13-135 or 13-138 of the applicable airplane Service Manual. Any part found damaged must be replaced with a serviceable part. Also, replace the cap nut (or the complete assembly if necessary) with a like part which has been drilled for MS20995C41 safety wire, or has equivalent safetying provisions approved by the Chief, Engineering & Manufacturing Branch, FAA, Eastern Region. c. Re-install valve assembly, safety wire cap nut and pressure check fuel system in accordance with Section 13-136 or 13-139 of applicable Airplane Service Manual and Piper Service Bulletin No. 401, or equivalent methods approved by the Chief, Engineering & Manufacturing Branch, FAA, Eastern Region. d. Aircraft may be flown to a base where the maintenance required by this Airworthiness Directive may be performed per FAR's 21.197 and 21.199. This amendment is effective October 23, 1974.
2007-22-01: We are adopting a new airworthiness directive (AD) for Bell Helicopter Textron Canada (Bell) Model 206A and 206B series helicopters. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority to identify and correct an unsafe condition on an aviation product. The aviation authority of Canada, with which we have a bilateral agreement, states in the MCAI: Transportation Safety Board of Canada (TSB) investigation into an accident involving Model 206B has revealed that the Spindle repaired by Cadorath Aerospace Inc., failed during flight resulting in loss of control of the helicopter. A similar repair was performed by H-S Tools & Parts Inc. This AD requires actions that are intended to address this unsafe condition related to certain repaired transmission pylon support spindles.
83-24-09: 83-24-09 DeHAVILLAND AIRCRAFT OF CANADA, LTD.: Amendment 39-4781. Applies to DeHavilland Model DHC-7 airplanes, Serial Numbers 3 through 92 inclusive. Compliance is required within the next 750 flight hours or three months after effective date of this AD, whichever occurs first, unless previously accomplished. To protect against loss of trimability and partial loss of pitch control, accomplish the following: 1. Incorporate Modification Number 7/2209 on each elevator and Modification Numbers 7/2211 and 7/2216 on the tailplane in accordance with instruction contained in DeHavilland Service Bulletin No. 7-55-6, Rev. B, dated September 17, 1982. 2. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. 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 as required by this AD. This amendment becomes effective December 21, 1983.
2019-24-12: The FAA is adopting a new airworthiness directive (AD) for certain De Havilland Aircraft of Canada Limited Model DHC-8-401 and - 402 airplanes. This AD was prompted by a report that certain fuselages were delivered with nonconforming keel tension fittings and stringer end fittings. This AD requires a detailed visual inspection of stringer end fittings and keel fittings for loose or working fasteners, signs of wear, and corrosion, and repair if necessary; and a general visual inspection of the keel tension fitting and stringer end fittings, as applicable, and repairs and replacement of the keel and stringer end fittings if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
72-14-05: 72-14-05 PIPER: Amendment 39-1480 as amended by Amendment 39-1943. Applies to PA-23-250 and PA-E23-250 series aircraft Serial Nos. 27-2505 and up equipped with non-supercharged engine and certificated in all categories. Compliane required within the next 50 hours time in service after the effective date of this AD, unless already accomplished within the last 50 hours and thereafter at intervals not to exceed 100 hours. a. For airplanes Serial Nos. 27-7304959 and up, and earlier serial numbered aircraft which have field replacements, equipped with exhaust stack assemblies P/N's 33419-2, -3, 33420-2, -3 and stack support kit No. 760702, visually inspect both engine exhaust system stacks, manifolds, support tubes and brackets, and the slip joint inside the alternate air heat shroud. Inspect for cracks, flaking, burning, distortion and exhaust gas leaks at flanges and joints. Parts found cracked, flaked, burned, distorted or which allow leakage must be replaced prior to further flight. b. For airplanes Serial Nos. 27-2505 to 27-7304958 inclusive equipped with original type exhaust system configuration or field replacement components as listed under Piper Service Letter No. 533, Service Spares Letter No. SP-301 or Service Bulletin No. 319: Visually inspect in accordance with a. above and adjust and align ball joint stack assemblies and supports in accordance with Piper Service Bulletin No. 319 dated April 19, 1971 or later approved revision. c. Equivalent inspections or modifications may be approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. Aircraft may be flown to a base where the maintenance required by this airworthiness directive may be performed as permitted by FAR's 21.197 and 21.199. (Piper Service Spares Letter No. SP-301 dated November 11, 1969, and Service Bulletin No. 319 dated April 19, 1971, pertain to this subject) Amendment 39-1480 supersedes AD 69-23-01. Amendment 39-1480 was effective July 12, 1972. This Amendment 39-1943 is effective September 3, 1974.
80-17-12: 80-17-12 MCDONNELL DOUGLAS: Amendment 39-3890. Applies to DC-10 Series Airplanes fuselage numbers 1 through 243. \n\n\tCompliance required as indicated, unless already accomplished. To prevent restriction of travel of the wing engine emergency fire shut off handle and loss of associated fire fighting safeguards, accomplish the following: \n\n\tWithin 300 hours additional time in service after the effective date of this AD, accomplish paragraph (1) or (2) below. \n\n\t(1)\tInspect and seal cable guard pin P/N AA 2603-9 in place in accordance with McDonnell Douglas Alert Service Bulletin A76-27 dated 6/24/80 and within 1500 hours additional time in service after the effective date of this AD accomplish the modification described in McDonnell Douglas Service Bulletin 76-27 dated 7/16/80; or \n\n\t(2)\tAccomplish the modification described in McDonnell Douglas Service Bulletin 76-27 dated 7/16/80. \n\n\t(3)\tAccomplishment of the modification described in McDonnell Douglas Service Bulletin 76-27 dated 7/16/80 is terminating action for this AD. Service Bulletin 76-27 calls for the replacement of the existing guard pin with a new guard pin incorporating a lip retaining head. \n\n\t(4)\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of modification required by this AD. \n\n\t(5)\tAlternative inspections, modifications or other actions which provide an equivalent level of safety may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\tThis amendment becomes effective August 28, 1980.
2007-21-08: The FAA is adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Model Hawker 800XP airplanes. This AD requires doing an inspection of panel DA wiring for clearance and for signs of chafing or exposed conductors, and repairing or replacing the wires and cable ties if necessary. This AD results from reports of wire bundle interference in the DA panel, chafed wire bundles, and exposed conductors. We are issuing this AD to prevent chafing of wire bundles, which could cause an electrical short and consequent loss of several functions essential for safe flight and smoke or fire in the flight compartment and main cabin.
83-24-03: 83-24-03 BOEING: Amendment 39-4774. Applies to Model 707 and 720 series aircraft certificated in all categories. Accomplish the following within 120 days after the effective date of this AD, unless already accomplished within the last 21 months, and at intervals thereafter not to exceed 24 months: \n\n\tA.\tVisually inspect the lower rudder tab control rod assemblies, including all dash numbers, which have been in service more than 5 years, in accordance with Boeing Service Bulletin 3424, dated July 1, 1983, or later FAA approved revisions. \n\n\tB.\tReplace any rod assembly exhibiting cracks or corrosion with a new or reconditioned rod assembly.\n \n\tC.\tAlternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tD.\tAircraft may be ferried to a base for maintenance in accordance with Section 21.197 and 21.199 of the Federal Aviation Regulations. \n\n\tE.Upon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of an operator, if the request contains substantiating data to justify the adjustment period. \n\n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at the FAA, Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective January 3, 1984.