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77-05-04: 77-05-04 TELEDYNE CONTINENTAL MOTORS: Amendment 39-2848. Applies to the following Teledyne Continental Motors engines installed in but not limited to: IO-520-D installed on Cessna 185 and 188 aircraft serial numbers 18502839 through 18503234, 18503236 through 18503284, 18503287, 18503291, 18503293, 18802349 through 18802887, and 18802893. IO-520-F, TSIO-520-C, TSIO-520-G, TSIO-520-M and TSIO-520-R installed on Cessna U-206, T-206, and T-207 aircraft serial numbers U20603021 through U20603693, U20603695, U20603696, U20603699, U20603712, 20700315 through 20700378. IO-520-L, TSIO-520-H, TSIO-520-R, installed on Cessna 210 and T-210 aircraft serial numbers 21061040 through 21061736, 21061738 through 21061763, 21061766, 21061771, 21061773, 21061775 through 21061777, 21061789. IO-470-L installed on Beech Model 95-B55, aircraft serial numbers TC-2003 through TC-2053. Compliance required as indicated after the effective date of this AD unless already accomplished. To prevent crankshaft failure: A. Engines with less than 100 hours total time in service accomplish the following: (1) Within the next 10 hours time in service, check propeller operation in accordance with paragraph C or E as appropriate, and thereafter at intervals not to exceed 10 hours time in service from the previous check until 100 hours total time in service has been reached, whereupon this special check may be discontinued if no problem is evident. (2) Within the next 10 hours time in service examine the oil filter or screen as appropriate for each engine in accordance with paragraph F, and thereafter at intervals not to exceed 25 hours time in service until 100 hours total time in service has been reached, whereupon this special examination may be discontinued if no problem is evident. B. Engines with 100 hours or more total time in service, perform a one-time check or inspection of the following: (1) Within the next 10 hours time in service, check propeller operationin accordance with paragraph C or E as appropriate. (2) Within the next 10 hours time in service, examine the oil filter or screen, for each engine, in accordance with paragraph F. C. Check single engine airplane propeller operation as follows: (1) Ascertain that oil temperature is at or above the middle of the green arc on the oil temperature gage but in no case above the red line. CAUTION: Do Not Exceed Maximum Cylinder Head Temperatures. (2) With the propeller control in the low-pitch high RPM position, set engine speed with the throttle to 1700 RPM. (3) At 1700 RPM, pull the propeller control to the full high-pitch low RPM position until minimum governing RPM is observed, then push the control back to the low-pitch high RPM position. Repeat this procedure three times noting minimum governing RPM each time. Using this procedure the RPM should drop at least 400 RPM and should be reasonably smooth and consistent. (4) If a minimum drop of 400 RPM is obtained consistently in Paragraph C(3), proceed with paragraph F at time intervals specified. NOTE: Propeller operation checks C(1) through C(4) may be performed by the pilot; however, requirements of C(5)(a) through C(5)(c) and Paragraph F require appropriately authorized mechanic or repair station. (5) If a minimum drop of 400 RPM cannot be obtained or the propeller operation is not smooth and consistent, the following additional checks are to be accomplished. (Reference the applicable Aircraft Service Manual.) (a) Operation of the propeller control -- check routing, control clamping, rod end attachments, control rigging and adjustment. (b) Governor operation -- check security, signs of leakage, arm attachment and stop adjustment. (c) Propeller -- check for any external signs of leakage and/or damage. D. If any discrepancies are noted when accomplishing 5(a), (b) or (c), correct condition and recheck in accordance with Paragraph C(1) through C(4). If propeller operation is not in accordance with Paragraph C(3), proceed immediately to Paragraph F and prior to further flight, contact Mr. R.J. Moore, FAA-DER SO-260, Teledyne Continental Motors, P.O. Box 90, Mobile, Alabama 36601, (205) 438-3411, or his representative for disposition or other equivalent disposition method authorized by the Chief, Engineering and Manufacturing Branch, ASO-210, P.O. Box 20636, Atlanta, Georgia 30320. E. Check multiengine airplane propeller operation as follows: NOTE: Propeller checks in paragraphs E(1) and E(2) may be conducted by the pilot. (1) Ground run the engine at 1000 RPM until some oil temperature is indicated. Increase the engine RPM to 1700 until oil temperature has stabilized at or above the middle of the green arc but in no case above the red line. CAUTION: Do not exceed maximum cylinder head temperatures. (2) Place the propeller control in the low-pitch high RPM position and set engine speed to 900 RPM using the throttle. Note any tendency of the propeller to feather after engine speed has stabilized. (3) If feathering does occur in Step (2), the following additional checks are to be accomplished. (Reference the applicable aircraft Service Manual.) (a) Operation of the propeller control -- check routing, control clamping, rod end attachments, control rigging and adjustment. (b) Governor operation -- check security, signs of leakage, arm attachment and stop adjustment. (c) Propeller -- check for any external signs of leakage and/or damage. If any discrepancies are noted in accomplishing Paragraph 3(a), (b) or (c), correct condition and recheck in accordance with Paragraph E(1) through E(2). If feathering occurs, proceed immediately to Paragraph F, and prior to further flight contact Mr. R.J. Moore, FAA-DER SO-260, Teledyne Continental Motors, P.O. Box 90, Mobile, Alabama 36601, (205) 438-3411, or his representative for disposition or other equivalent disposition method authorized by the Chief,Engineering and Manufacturing Branch, ASO-210, P.O. Box 20636, Atlanta, Georgia 30320. F. Perform the following inspections of the oil filter or screen for evidence of metal contamination. (1) Remove and cut open the oil filter or remove and inspect the screen (whichever is applicable). (2) Visually inspect for abnormal amount of metal. (3) Some small quantity of minute metal particles is considered normal; however, should an abnormal amount be present it could be indicative of bearing distress; therefore, check the magnetic properties of the metal, and prior to further flight, request disposition in accordance with Paragraph D or E as appropriate. (4) If no abnormal amount of metal is present, and propeller control checks as outline in Paragraph D or E are satisfactory, the aircraft may remain in service. NOTE: For the requirements regarding the listing of compliance and method of compliance with this AD in the airplane's permanent maintenance record, see FAR 91.173. NOTE: Teledyne Continental Motors Service Bulletin M77-6, Supplement 1, pertains to this same subject. This amendment becomes effective March 11, 1977.
83-01-01: 83-01-01 GENERAL DYNAMICS (Convair): Amendment 39-4535. Applies to Model 240, and military models eligible or to be made eligible for civil use under Type Certificate A793, and all such model airplanes converted to turbopropeller power, certificated in all categories. Compliance required as indicated unless previously accomplished. To prevent fracture of the nose landing gear (NLG) axle caused by fatigue cracks, accomplish the following. NOTE: All references to hours of time in service apply to nose landing gear axle and not to the aircraft. (A) Within 150 hours time in service after the effective date of this AD, or prior to the accumulation of 10,150 hours time in service, whichever occurs later, unless accomplished within the last 1050 hours time in service, conduct an ultrasonic inspection of each NLG axle (P/N 240- 5210005 or P/N 240-7000475) in accordance with the applicable provisions of paragraph 2 entitled, "Accomplishment Instructions," of General Dynamics Convair Division Service Bulletin 600 (24OD) 32-6 dated February 21, 1980. If cracks are found, replace with a like serviceable part prior to return of aircraft to service. (B) Repeat the ultrasonic inspection specified in paragraph (A) or perform the magnetic particle inspection specified in paragraph (C) of this AD at intervals not to exceed 1200 hours of time in service. (C) Within 12,000 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 12,000 hours time in service, conduct a magnetic particle inspection of the NLG axle (P/N 240-5210005 or P/N 240-7000475) in accordance with the applicable provisions of paragraph 2, entitled "Accomplishment Instructions" of Service Bulletin 600 (24OD) 32-6. If cracks are found, replace NLG axle with a like serviceable part prior to return of aircraft to service. (D) Prior to issuance of a Certificate of Airworthiness for military aircraft being converted for civil certification,the airplane must be inspected in accordance with paragraph (A) or (C) of this AD. (E) 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 required by this AD. (F) Alternative inspections, modifications or other actions which provide an equivalent level of safety may be used when approved by the Manager, Western Aircraft Certification Field Office, FAA, Northwest Mountain Region, Hawthorne, California. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made 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 General Dynamics, Convair Division, P.O. Box 80877, San Diego, California 92138. These documents also may be examined at FAA, Northwest Mountain Region, 17900 PacificHighway South, Seattle, Washington, or Western Aircraft Certification Field Office, 15000 Aviation Boulevard, Hawthorne, California. This amendment becomes effective January 19, 1983.
90-10-10: 90-10-10 AVIONS MARCEL DASSAULT-BREGUET AVIATION (AMD-BA): Amendment 39-6598. Docket No. 89-NM-275-AD. Applicability: All Model Mystere Falcon 900 series airplanes equipped with thrust reverser synchronizing bell crank, Part Number F900HD3200240A2, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent failure of the thrust reverser door actuating system, accomplish the following: A. Prior to the accumulation of 3,000 landings on the thrust reverser synchronizing bell crank, Part Number F900HD3200240A2, or within the next 250 landings after the effective date of this AD, whichever occurs later, and thereafter at intervals not to exceed 3,000 landings, replace the bell crank, which is life limited to 3,000 landings, in accordance with the manufacturer's maintenance manual. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Northwest Mountain Region. 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. C. 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 Falcon Jet Corporation, Customer Support Department, Teterboro Airport, Teterboro, New Jersey 07608. 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 (39-6598, AD 90-10-10) becomes effective on June 15, 1990.
84-01-51: 84-01-51 SIKORSKY AIRCRAFT: Amendment 39-4826. Applies to Model S-61 series aircraft certificated in all categories. Compliance required within the next 25 hours time in service after the effective date of this AD, unless already accomplished. As a result of finding a crack in a rotary wing head shaft nut and finding the nut had improper Rockwell hardness and because of possible incomplete processing of these nuts, P/N S6110-21082-0, accomplish the following: (a) Fluorescent magnetic particle inspect and Rockwell hardness test the rotary wing head shaft nut, in accordance with Section B of Sikorsky Service Bulletin No. 61B10-14, dated November 3, 1969, as revised by Sikorsky Message Number CST-P-83-171, Revision 2, dated December 5, 1983. Visible magnetic particle inspection and Rockwell hardness test in accordance with Section B of the service bulletin may be used as an alternate to fluorescent magnetic particle inspection and Rockwell hardness test. (b) Anynut with cracks or not within the proper hardness range specified in Sikorsky service documents specified in paragraph (a) of this AD must be replaced before further flight with a nut that has been inspected in accordance with paragraph (a) of this AD and baked and marked in accordance with paragraph (c) of this AD or with a Sikorsky provided nut inspected and marked in accordance with this AD. (c) Any nut not cracked and with proper hardness as specified in the Sikorsky service documents must be baked at 375 degrees fahrenheit plus/minus 25 degrees fahrenheit for a minimum of three hours and marked in accordance with the Sikorsky service documents specified in paragraph (a) of this AD. (d) Alternate inspections, processes, or other means of compliance which provide an equivalent level of safety to this AD must be approved by the Manager, Boston Aircraft Certification Branch, New England Region, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. In accordance with FAR 21.197, flight is permitted to a base where the requirements of this AD -may be accomplished. Sikorsky Service Bulletin No. 61B10-14, dated November 3, 1969, and Sikorsky message Number CST-P-83-171, Revision 2, dated December 5, 1983, identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 551(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to United Technologies Corporation, Sikorsky Aircraft, North Main Street, Stratford, Connecticut 06602, ATTN: S-61 Commercial Product Support Department. These documents also may be examined in the Office of the Regional Counsel, Federal Aviation Administration, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment supersedes Amendment 39-889 (34 FR 19545), AD 69-25-05, which became effective December 10, 1969. This amendment becomes effective March 27, 1984, as to all persons except those persons to whom it was made immediately effective by the telegram dated December 30, 1983, which contained this amendment.
2002-21-12: This amendment adopts a new airworthiness directive (AD) for the specified Agusta S.p.A. (Agusta) helicopters that requires establishing or reducing the life limits of various parts listed in the airworthiness limitations section (ALS) of the maintenance manual. This amendment is prompted by the results of fatigue tests and analysis to determine life limits for various parts. The actions specified by this AD are intended to establish or reduce the life limits to prevent failure of specified parts and subsequent loss of control of the helicopter.
68-19-02: 68-19-02 SIAI-MARCHETTI: Amendment 39-654. Applies to Model S.205/22R airplanes, Serial Nos. 213, 370, 371, 372, 373, 374, 378, 379, 380, 381, 382, 384, 385, 386, 387, 388, 389, 390, 392, 393, 394, 395, 396, 398, 4-124, 4-125, 4-132, 4-142, 4-143, 4-145, 4-146, 4- 147, 4-148. Compliance required as indicated, unless already accomplished. To detect cracks in the lower part of fuselage frame No. 2, at the rudder control bellcrank brackets, and to detect cracks and deformation in the stringers of the fuselage belly skin which are located at the outboard rudder pedal support tube brackets, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, visually inspect, using a lamp and mirror, the lower portion of fuselage frame No. 2 for cracks and the lower fuselage stringers for cracks and deformation in accordance with Siai Service Bulletin No. 205B20, dated March 22, 1968, or later RAI-approved revision or an FAA-approved equivalent. If no cracks are found, comply with paragraph (d) within the next 100 hours' time in service after the effective date of this AD. (b) If cracks are found during the inspection required by paragraph (a) none of which exceed 3/16 inches in length, repeat that inspection at intervals not to exceed 20 hours' time in service from the last inspection and comply with paragraph (d) within the next 100 hours' time in service after the effective date of this AD. The repetitive inspection required in this paragraph may be discontinued upon compliance with paragraph (d). (c) If cracks are found during the inspection required by paragraph (a) which exceed 3/16 inches in length, comply with paragraph (d) before the next flight. (d) Modify fuselage frame No. 2 and stringers in accordance with Siai Service Bulletin No. 205B20, dated March 22, 1968, or later RAI-approved revision, or an FAA- approved equivalent. This amendment becomes effective October 11, 1968.
2001-02-09: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 757-200 series airplanes, that currently requires inspections to detect cracking on the free edge of the tang, if necessary, and of the fastener holes in the lower spar chord; and various follow-on actions. That AD also provides for an optional terminating action for the repetitive inspections. This amendment adds inspections to detect additional cracking of the fastener holes in the lower spar chord. This amendment also adds an optional terminating modification. This amendment is prompted by the issuance of new service information. The actions specified by this AD are intended to detect and correct fatigue cracking in the lower spar chord, which could result in reduced structural integrity of the engine strut.
88-12-03: 88-12-03 BELL HELICOPTER TEXTRON, INC. (BHTI): Amendment 39-6539. Final copy of priority letter AD. Docket No. 88-ASW-32. Applicability: Bell Model 214ST helicopters, serial numbers 28101 through 28195 and 18402, certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent failure of the main rotor (M/R) drag brace assembly which could result in subsequent loss of the helicopter, accomplish the following: (a) For each M/R drag brace assembly, part number (P/N) 214-010-113-105, with 250 or more hours' time in service, inspect within the next 10 hours' time in service and thereafter at intervals not to exceed 250 hours' time in service as follows: (1) Remove drag brace assembly to the blade attachment bolt, P/N 20-057- 16-68D, and the drag brace to the grip attachment bolt, P/N 20-057-16-92D. (2) Measure length of each drag brace assembly, then disassemble and thoroughly clean all parts. (3) Inspectthe barrel, P/N 214-010-120-107; the clevis, P/N 214-010-121- 105; and the nuts, P/N 214-010-198-101, for corrosion pitting and mechanical damage. Remove the part from service if pitting is found in the threads. NOTE: If corrosion pitting or mechanical damage is superficial, it may be polished out. (4) Remove paint and primer for inspection. (5) Inspect each drag brace barrel for cracks using a magnetic particle or fluorescent penetrant method. Pay special attention to the threaded area. (b) If cracks are found, or if corrosion pitting or mechanical damage is found which exceeds a maximum depth of 0.010 inch, replace with serviceable parts. (c) Assemble and install the drag brace as follows: (1) Threads of barrel which have cadmium plating removed during disassembly must be brush cadmium plated or protected. Coat threads of the drag brace barrel, nuts, and the inner surface of clevis with either corrosion preventive compound MIL-C-16173, Grade 2 orGrade 3, or an equivalent. (2) Reassemble and reset the length of the drag brace as measured in the requirements of paragraph (a)(2). Torque inboard nut 375 to 425 foot pounds, while holding outboard nut. (3) Reinstall the drag brace assembly and torque inboard and outboard attachment bolt nuts, 225 to 400 foot pounds. (4) After final drag brace adjustment, if required, clean the exterior of the drag brace assembly. Coat the threads, nuts, and clevis area with MIL-C-16173, Grade 1 (drying), corrosion preventive compound or equivalent. (d) Whenever a new design drag brace assembly, P/N 214-010-191-101, is installed in accordance with BHTI Alert Service Bulletin (ASB) 214ST-88-47, Revision A, dated February 20, 1989, this AD no longer applies. (e) An alternate method of compliance or adjustment of the compliance time, which provides an equivalent level of safety, may be used if approved by the Manager, Rotorcraft Certification Office, FAA, Southwest Region,4400 Blue Mound Road, Fort Worth, Texas. (f) In accordance with FAR Section 21.197 and 21.199, the helicopter may be flown to a base where the inspection may be accomplished. This amendment becomes effective April 9, 1990, as to all persons except those persons to whom it was made immediately effective by Priority Letter AD 88-12-03, issued on June 3, 1988, which contained this amendment.
90-15-18: 90-15-18 BEECH: Amendment 39-6667. Docket No. 90-CE-13-AD. Applicability: Models 99, 99A, and A99A (Serial Numbers (S/N) U-1 through U-147, except U-146); and B99 (S/N U-146 through U-151, except U-147) airplanes certified in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To prevent loss of structural integrity in the vertical stabilizer (fin) main spar, accomplish the following: (a) For airplanes that have accumulated 2,000 or more hours TIS on the effective date of this AD, within the next 50 hours TIS, unless already accomplished within the last 450 hours TIS per AD 73-03-04, and thereafter at intervals not to exceed 500 hours TIS, visually inspect utilizing a three to five power magnifying glass the vertical fin main spar at each side of the bend location for cracks or nicks as shown in Figure 3 of Beechcraft Service Instructions No. 0530-134, Revision 1, dated June 1975. (b) If, during any inspection required herein, a crack that does not exceed 0.25 inches in length is found in either a spar flange or in an angle doubler, and no such cracks are found in both members on the same side, prior to further flight either: (1) Repair the spar by installing a plate doubler in accordance with Beechcraft Service Instructions No. 0530-134, Revision 1, dated June 1975, and reinspect at 500 hour intervals thereafter per Paragraph (a); or (2) Replace the spar with an equivalent airworthy part and reinspect per the requirements of this AD. (c) If, during any inspection required herein, a crack is found in both the spar flange and angle doubler flange on the same side, or if a crack exceeds 0.25 inch in length, prior to further flight replace the vertical fin assembly with a Part Number (P/N) 115-640000-605 or -607 or -651 vertical fin. (d) Within the next 500 hours TIS after the effective date of this AD, or upon the accumulation of 20,000 hours TIS on theoriginal vertical stabilizer, whichever occurs later, either: (1) Replace the vertical fin with a serviceable P/N 115-640000-605 or -607 or -651 vertical fin; or (2) Verify that no cracks have ever been detected in the affected structure and install a plate doubler per Beechcraft Service Instructions No. 0530-134, Rev. 1, dated June 1975. (e) The inspections specified in Paragraph (a) of this AD are no longer required in a P/N 115-640000-605 or -607 or -651 vertical fin or if a plate doubler per Paragraph (d)(2) above has been installed. A doubler installed over previously cracked structure does not comply with this paragraph. (f) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (g) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the document referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or may examine this document at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD supersedes AD 73-03-04, Amendment 39-3695. This amendment (39-6667, AD 90-15-18) becomes effective on September 11, 1990.
2016-07-25: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 787-8 airplanes. This AD was prompted by reports indicating that the ram air turbine (RAT) assembly may fail to operate if deployed at low airspeeds. This AD requires replacing either the RAT pump and control module assembly or the entire RAT assembly. We are issuing this AD to prevent failure of the RAT assembly to operate at low air speeds. The volume fuse on the RAT assembly may be activated in-flight before the RAT is deployed. This may lead to improper pump hydraulic pressure offloading when the RAT is needed. Failure of the RAT to operate in an all engine out event would result in loss of control of the airplane.
2001-02-04: This amendment adopts a new airworthiness directive (AD) that applies to all Pilatus Aircraft LTD (Pilatus) Model PC-6 airplanes that are equipped with a certain stabilizer trim actuator. This AD requires you to inspect the lower lug of the actuator for cracks, damage, or distortion; verify that the staked bearing is correctly installed in the bore of the lug; and repair any cracked, damaged, or distorted parts and reassemble any incorrectly installed staked bearing, as necessary. 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 detect and correct damage, distortion, or cracks in the lower lug assembly, which could result in failure of the lower lug. Such failure could lead to loss of the stabilizer trim actuator with consequent loss of control of the airplane.
79-12-04: 79-12-04 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-3486. Applies to all AiResearch Model TPE 331-10-501G engines incorporating the forged third stage turbine wheel, Part Number 3101516-1, installed in the Swearingen Merlin IIIB airplane. NOTE: The engine and airplane manufacturer's records indicate that the subject forged third stage turbine wheels, Part Number 3101516-1, are contained in the following serial numbered TPE 331-10-501G turboprop engines and are installed on the following Swearingen Merlin IIIB airplanes: Engine Airplane Serial Number Serial Number Airplane Registration P35003 T292 LVBHP P35004 T292 (Argentina) P35008 T294 N800TA P35010 T294 (Teterbury Aviation, NJ) P35009 T295 N4442F P35012 T295 (Fairchild) P35011 T296 VHSWK P35013 T296 (Stillwell; in Australia) P35017 T298 N5495M P35019 T298 (Stillwell; in Australia) P35014 T297 N5465M P35015 T297 (Planes, Inc.; Atlanta, GA)P35016 T299 N81CH P35020 T299 (Duncan Aviation, NB) P35022 T300 N29TA P35023 T300 (Teterbury Aviation, NJ) P35021 T301 N5652M P35024 T301 (Western Airmotive, Hillsboro, OR) P35025 T302 N5497M P35026 T302 (Houston, TX) P35007 T293 DIBB (Cologne, Germany) P35018 T276 N5111B (Philadelphia) Compliance required as indicated. To reduce the possible hazard from the total loss of power of one or more engines during a single flight resulting from the failure of the forged third stage turbine wheel blade, accomplish the following. (a) Within the next 10 operating hours time in service after May 10, 1979, on at least one of the affected TPE 331-10-501G engines on each affected airplane, unless already accomplished, remove the forged third stage turbine wheel, Part Number 3101516-1, and replace it with a cast third stage turbine wheel, Part Number 3102655-1, in accordance with existing maintenance instructions. NOTE: Suitable instructions may be found in TPE 331-10 Engine Maintenance Manual, AiResearch Part Number 72-00-23. (b) Within the next 50 operating hours time in service after May 10, 1979, on all TPE 331-10-501G engines installed on the Merlin IIIB airplane, unless already accomplished, remove the forged third stage turbine wheel, Part Number 3101516-1, and replace it with a cast third stage turbine wheel, Part Number 3102655-1, in accordance with instructions referenced in Paragraph (a) of this airworthiness directive. (c) Alternative actions which provide an equivalent level of safety may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. (d) 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 the actions required by this AD. (e) Upon request of operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region may adjust the initial and repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. This amendment becomes effective upon publication in the Federal Register as to all person except those persons to whom it was made immediately effective by telegram dated May 10, 1979, which contains this amendment.
59-22-03: 59-22-03 MOONEY: Applies to All M-18 Series Aircraft. Compliance required within the next 10 flight-hours but not later than December 15, 1959. The following inspections, repairs, and replacements shall be accomplished: (a) Empennage (1) Remove and disassemble empennage. Remove control surfaces and hinge brackets from fin and horizontal stabilizer. Remove bolts through stabilizer main spar attach blocks. Disassemble stabilizer and fin from empennage truss and each other. (2) Inspect all bolted joints for the following items: (i) Wear on bolts (ii) Wear on bolt holes in fittings and lugs (iii) Wear on bushing (iv) Wear of bushing on fittings, lugs and wood. (Replace parts as necessary). (3) Remove all fabric from stabilizer and fin. Inspect all wood and glue joints including attachment of leading edge skin to main spar for deterioration. (4) At center section of stabilizer spar inspect glue joint between attach blocks and spar for deterioration and inspect spar and blocks for cracks. Inspect fin spar for cracks at attach bolts. (5) Any defective wood parts shall be replaced or repaired in accordance with CAM 18 and/or manufacturer's recommendations. When the fin and stabilizer are satisfactory, reinforcement of the stabilizer main spar center section and the fin and stabilizer center hinge rib-rear spar attachment shall be accomplished in accordance with Mooney M-18 Service Letter No. 16. (Kits of reinforcement parts are available from Mooney Aircraft, Inc.) (6) Clean all empennage drain holes, and see that they are located as specified in Mooney M-18 Service Letter 16. (7) Inspect welds at rudder and elevator hinges and control horns and at all joints on the tail truss for inadequate welds (i.e. weld which does not fill fillet cross section area) and for cracks using either method (i) or (ii) below. (i) Magnetic particle or X-ray inspection. (ii) Remove paint and primer and visually inspect welds with a 10- power glass. Parts with defective welds are to be replaced or repaired. A joint may be rewelded providing the old weld is removed and the surfaces thoroughly cleaned. (8) Remove upper tail truss attach fittings from aft fuselage bulkhead and inspect as described in item (2). Inspect bulkhead front and back for cracks in area of these fittings. Inspect glue joint between bulkhead and aft fuselage skin and longerons for deterioration or separation. Repair in accordance with Mooney M-18 Service Letter No. 17. Examine trim linkage attached to lower part of aft bulkhead for worn bolts. Replace bolts as necessary. (9) Reassemble and install empennage making sure all bolts are tight. Block airplane solidly at tail skid and inspect for empennage play as follows: (i) Stabilizer - Move up and down at one tip and measure at opposite tip. Total allowable play 1/2-inch up and down. (ii) Stabilizer - Move fore and aft at one tip and measure at opposite tip. Total allowable play 3/4-inch fore and aft. (iii) Fin - Move fore and aft at top leading edge and measure at bottom rudder trailing edge. Total allowable play 1/2-inch up and down. (b) Aft Fuselage (1) Inspect wood around forward fuselage tubular structure attach fittings for deterioration. Clean all drain holes. Inspect all glue joints for deterioration. See that drain holes are located as specified in Mooney M-18 Service Letter 16. (c) Wing (1) Remove-seat, auxiliary fuel tank and belly access panel. Inspect ribs, skin and both spars at lower center section and around fuselage fittings for wood and glue joint deterioration. (2) Inspect all wood and glue joints in wheel well area for deterioration. Inspect both spars for cracks in area of the gear attachments. (3) Inspect interior of wing in areas having access openings. (4) Remove aileron and inspect hinges and control horn in accordance with part (a), item (7). (5) Remove wing fabric locally in area of aileron hinges and at inboard corner of aileron cutout and check condition of wood and glue joints. If evidence of deterioration is found remove fabric further as necessary for complete examination of forward area of wing trailing edge. Check security of attachment of wing trailing edge in aileron area. (6) Clean all drain holes in wing, and see that they are located as specified in Mooney M-18 Service Letter 16. (d) Control Systems (1) Inspect all control systems (aileron, trim, rudder, and elevator). (i) Visually inspect all welds for cracks and inadequate welds (i.e. weld which does not completely fill fillet cross section area). (ii) Check security of all bolted hinge and fitting attach points. (Mooney M-18 Service Letters Nos. 16 and 17 pertains to this same subject.)
2016-07-22: We are adopting a new airworthiness directive (AD) for all Airbus Model A300 B4-600, B4-600R, and F4-600R series airplanes, Model A300 C4-605R Variant F airplanes (collectively called Model A300-600 series airplanes), and Model A310 series airplanes. This AD was prompted by reports of insufficient clearance for the electrical wiring bundles in the leading and trailing edges of the right-hand (RH) and left-hand (LH) wings. This AD requires modifying the electrical routing installation at the RH and LH wings. We are issuing this AD to prevent insufficient clearance of electrical wiring bundles located in the leading and trailing edges of the RH and LH wings, which could lead to chafing damage and arcing, possibly resulting in an on-board fire.
93-02-10: 93-02-10 BRITISH AEROSPACE: Amendment 39-8493. Docket 92-NM-188-AD. Supersedes AD 87-12-08, Amendment 39-5652. Applicability: Model BH/DH/HS/BAe 125 series airplanes; as listed in British Aerospace Alert Service Bulletin S.B. 24-A261, Revision 1, dated August 17, 1987; certificated in any category. Compliance: Required as indicated, unless accomplished previously. NOTE: Paragraphs (a) and (b) of this AD restate the requirements of AD 87-12-08, Amendment 39-5652, paragraphs (a) and (b). As allowed by the phrase, "unless accomplished previously," if the requirements of AD 87-12-08 have been accomplished previously, paragraph (a) of this AD does not require that the initial inspection be repeated. To prevent a circuit from overheating and a resultant fire, accomplish the following: (a) Within 10 days after July 7, 1987 (the effective date of AD 87-12-08, Amendment 39- 5652): Inspect the battery cables to detect chafing and local damage, in accordance with British Aerospace Telex Alert Service Bulletin S.B. 24-A261, dated March 6, 1987; or British Aerospace Alert Service Bulletin S.B. 24-A261, dated March 9, 1987, or Revision 1, dated August 17, 1987. If chafing or damage is found, prior to further flight, replace the affected cable, in accordance with the applicable service bulletin. (b) Repeat the inspection required by paragraph (a) of this AD at intervals not to exceed one year, and, if chafing or damage is found, replace the affected cable prior to further flight. (c) Modification of the wiring installation at Panel ZL (Modification No. 253204A or B), in accordance with British Aerospace Service Bulletin SB.24-261-3204A&B, Revision 1, dated March 24, 1988, constitutes terminating action for the repetitive inspections required by paragraph (b) of this AD. (d) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (f) The modification shall be done in accordance with British Aerospace Service Bulletin SB.24-261-3204A&B, Revision 1, dated March 24, 1988. (NOTE: The issue date of this British Aerospace Service Bulletin SB.24-261-3204A&B is only indicated on "page 1 of 12"; no other page is dated.) This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, DC 20041-0414. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (g) This amendment becomes effective on March 26, 1993.
2016-07-14: We are adopting a new airworthiness directive (AD) for certain Airbus Model A319, A320, and A321 series airplanes. This AD is intended to complete certain mandated programs intended to support the airplane reaching its limit of validity (LOV) of the engineering data that support the established structural maintenance program. This AD was prompted by fatigue testing that determined fatigue damage could appear on clips, shear webs, and angles at certain rear fuselage sections and certain frames. This AD requires replacing the clips, shear webs, and angles, including doing all applicable related investigative actions, and repair if necessary. We are issuing this AD to prevent fatigue damage on the clips, shear webs, and angles; such damage could affect the structural integrity of the airplane.
2001-02-02: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Bombardier Model DHC-8-200, and -300 series airplanes. This action requires repetitive inspections to detect chafing or arcing damage to the cable/wire and fuel tube assemblies on the right hand side of each engine, and replacement with new components, if necessary. This action also provides an optional terminating action for the repetitive inspections required by this AD. This action is necessary to prevent chafing of the cable/wire bundles against the fuel line, which could result in arcing and a consequent fire or explosion. This action is intended to address the identified unsafe condition.
2000-25-52: This document publishes in the Federal Register an amendment adopting Emergency Airworthiness Directive (AD) 2000-25-52 which was sent previously to all known U.S. owners and operators of MD Helicopters, Inc. (MDHI) Model 369A, H, HE, HM, HS, D, E, FF, and 500N helicopters. This amendment supersedes an existing emergency AD that requires, before further flight, performing a tap inspection on both the upper and lower surfaces of each main rotor blade (blade). If any voids are detected that exceed specified inspection requirements, the emergency AD also requires replacing the unairworthy blade with an airworthy blade before further flight. This amendment requires the same actions as the emergency AD and corrects the applicability to include the appropriate serial numbers. This amendment is prompted by the discovery of an error in the emergency AD. The actions specified by this AD are intended to prevent failure of a blade and subsequent loss of control of the helicopter.
2016-07-21: We are superseding Airworthiness Directive (AD) 2015-20-13 for certain Piper Aircraft, Inc. Models PA-28-161, PA-28-181, and PA-28R- 201 airplanes. AD 2015-20-13 required inspecting the right wing rib at wing station 140.09 for cracks and taking necessary corrective action. This AD retains the actions for AD 2015-20-13 and adds airplanes to the applicability. This AD was prompted by reports that additional airplanes have been found with the same cracks. We are issuing this AD to correct the unsafe condition on these products.
50-07-01: 50-07-01 ERCO: Applies to Ercoupe Models 415C (Which Incorporate Adjustable Elevator Trim Tabs), 415CD, and 415D Airplanes. To be accomplished by September 1, 1950. To preclude the possibility of elevator flutter in the event the elevator trim tab control wire fails, elevator trim tab stop and spring, Erco P/N 415-SK-287 and 415-22035 should be installed. (Engineering and Research Corp. "Ercoupe Service Memorandum No. 55 and 55A" cover this same subject.)
2001-01-11: This amendment adopts a new airworthiness directive (AD) that applies to certain Rolladen Schneider Flugzeugbau GmbH (Rolladen Schneider) Models LS 4 and LS 4a sailplanes. This AD requires you to inspect the airbrake system for damage and proper rigging, with correction, repair, or replacement, as necessary. This AD also requires you to report any damage found to the Federal Aviation Administration (FAA). 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 detect and correct damage to the airbrake locking bracket caused by asymmetric loads. This condition could result in the pilot's inability to operate the airbrake controls, with consequent loss of sailplane control.
81-23-03 R2: 81-23-03 R2 CESSNA: Amendment 39-4318 as amended by Amendment 39-4491. Applies to Model P210N (Serial Numbers P21000001 through P21000811) airplanes certificated in any category with 25 or more total hours time-in-service. COMPLIANCE: Required as indicated unless already accomplished. To ensure the integrity of the engine exhaust system, accomplish the following: A) Prior to further flight on all airplanes with 25 hours or more time-in-service unless previously accomplished in accordance with AD 81-23-03: 1. Remove the engine cowl side and front panels and visually inspect the entire engine exhaust system for cracks. The cabin heat shroud does not have to be removed for this inspection. 2. Replace any cracked exhaust system components with serviceable parts. NOTE: Particular attention should be given to the exhaust system components attached to and between the two forward most (Nr. 5 and Nr. 6) cylinders. B) Within the next 10 hours time-in-service after the effective date of this AD and at each 5O hours time-in-service thereafter except as required by paragraph B)3: 1. Remove the engine cowl side and front panels as well as the cabin heater muff. 2. Visually inspect the entire engine exhaust system for cracks and bulges in the exhaust stacks, manifold, turbocharger and wastegate assemblies in accordance with Cessna SIL SE82-3. 3. Replace any cracked or bulged exhaust system components with serviceable parts in accordance with Cessna SIL SE82-3. If the exhaust system components have been removed and replaced with serviceable parts, or repaired by welding, reinspect within 25 hours time-in-service and replace or repair cracked or bulged components in accordance with Cessna SIL SE82-3. If no defects are found, the 50-hour inspection interval may be reinstated. C) Compliance with this AD is not required when complete Inconel Exhaust Systems consisting of Cessna Part Numbers 2154000-53 Left Hand Assembly, 2154000-54 Right Hand Assembly, and 2154000-68 Forward Crossover Assembly are installed. D) The inspection required by paragraph A)1 of this AD may be accomplished by the holder of at least a private pilot certificate issued under Part 61 of the Federal Aviation Regulations on any airplane owned or operated by that person (provided the airplane is not used in air carrier service). Make the prescribed entry in the aircraft maintenance record appropriate to this inspection. E) Any equivalent method of compliance with this AD must be approved by the Chief, Wichita Aircraft Certification Office (formerly Aircraft Certification Program), Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 269-7000. Cessna Single-engine Customer Care Service Information Letters SE79-52, Revision 1, dated November 5, 1979, SE81-21, dated May 18, 1981, and SE82-3, dated January 25, 1982, pertain in part to the subject matter of this AD. Amendment 39-4318 became effective on February 18, 1982, to all persons except those to whom it has already been made effective by airmail letter from the FAA dated October 31, 1981, as corrected by letter dated November 5, 1981. This amendment 39-4491 becomes effective on November 15, 1982.
2016-07-09: We are superseding Airworthiness Directive (AD) 2011-21-06 for all BAE SYSTEMS (Operations) Limited Model 4101 airplanes. AD 2011-21- 06 required revising the maintenance program. This new AD requires a new revision of the maintenance or inspection program. This AD was prompted by a determination that the life limit of certain main landing gear components must be reduced, and certain post-repair inspections of critical structure are necessary. We are issuing this AD to prevent failure of certain structurally significant items, including the main landing gear and nose landing gear, which could result in reduced structural integrity of the airplane; and to prevent fuel vapor ignition sources, which could result in a fuel tank explosion and consequent loss of the airplane.
2010-21-14: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Due to a manufacturing error, some rivets, required by drawings, were not installed in the joints between two ceiling beams and the rear pressurized bulkhead. If left uncorrected, long term fatigue stress could locally weaken the structure, compromising the fuselage structural integrity. We are issuing this AD to require actions to correct the unsafe condition on these products.
2001-01-03: This amendment adopts a new airworthiness directive (AD) that applies to all British Aerospace HP137 Mk1, Jetstream series 200, and Jetstream Models 3101 and 3201 airplanes. This AD requires you to remove the nose landing gear steering actuator and install one that incorporates a modified piston rod. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for the United Kingdom. The actions specified by this AD are intended to prevent failure of the nose landing gear steering actuator because of problems with the current design piston rod. Continued operation with the current design piston rod could result in loss of nose wheel steering and possible loss of control of the airplane during takeoff, landing, and taxi operations.