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2020-03-16: The FAA is adopting a new airworthiness directive (AD) for all Textron Aviation Inc. (Textron) (type certificate previously held by Cessna Aircraft Company) Models 210G, T210G, 210H, T210H, 210J, T210J, 210K, T210K, 210L, T210L, 210M, and T210M airplanes. This AD requires visual and eddy current inspections of the carry-thru spar lower cap, corrective action if necessary, application of a protective coating and corrosion inhibiting compound (CIC), and reporting the inspection results to the FAA. This AD was prompted by the in-flight break-up of a Model T210M airplane in Australia, due to fatigue cracking that initiated at a corrosion pit, and subsequent reports of other Model 210-series airplanes with widespread and severe corrosion. The FAA is issuing this AD to address the unsafe condition on these products.
84-09-04: 84-09-04 MCDONNELL DOUGLAS: Amendment 39-4855. Applies to McDonnell Douglas Model DC-8-70 series airplanes equipped with P/N C24466000 Thrust Reverser Hydraulic Control Units (HCU's) on the inboard thrust reversers, certificated in all categories. \n\n\tCompliance required as indicated in the body of this AD, unless previously accomplished: \n\n\tA.\tReplace P/N C24466000 HCU's on both inboard (#2 & #3) engine thrust reversers with P/N C24466000-/-2, P/N C24466001-2, or other FAA approved units, within 30 days after the effective date of this AD. The AFM limitations required by AD 82-19-51 R1, Amendment 39-4714 may be removed after HCU's are replaced on all aircraft in each operator's fleet. \n\n\tNote: P/N C24466000 HCU's can be modified to P/N C24466000-/-2 in accordance with CFMI Service Bulletin (CFM-56-2) 78-057, dated March 30, 1982, or later FAA approved revisions. \n\n\tB.\tAlternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tC.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplanes to a base to comply with the requirements of this AD. \n\n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). These documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Los Angeles Aircraft Certification Office, 4344 Donald Douglas Drive, Long Beach, California. \n\n\tThis supersedes Amendment 39-4714 (48 FR 40212), AD 82-19-51. \n\n\tThis amendment becomes effective June 1, 1984.
74-17-04: 74-17-04 PIPER: Amendment 39-1919 as amended by Amendment 39-2154 is further amended by Amendment 39-2865. Applies to the following Models which are covered with cotton or linen at the critical area on top of the windshield where the fabric attaches to the channel: PA-12, Serial Numbers 12-2904 and higher, except 12-2907, 12-2911, 12-2914, 12-2915, 12-2917, 12-2925, 12-2950, 12-3028-S and 12-3901 through 12-3903; PA-14; PA-15; PA-16; PA-17; PA-20; and PA-22 airplanes, Serial Numbers 22-1 and up, certificated in all categories. To prevent sudden failure of the fabric at the top of the windshield where the fabric attaches to the channel, accomplish the following: 1. For all airplanes, unless already accomplished, the indicated Piper Kits or equivalent parts approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, must be installed within the next 25 hours in service after the effective date of this AD on airplanes which have fabric exceeding three years since installation: (a) Kit 760 799 effective on: PA-12, Serial numbers 12-2904 and higher, except 12-2907, 12-2911, 12-2914, 12-2915, 12-2917, 12-2925, 12-2950, 12-3028-S and 12-3901 through 12-3903; all PA-12S and PA-14 airplanes. (b) Kit 754 404 effective on: All PA-15, PA-16, PA-17 and PA-20 and PA-22 Airplane Serial Numbers 22-1 and up. 2. For all airplanes at every 100 hours in service after accomplishment of 1(a) or 1(b), remove the metallic strip and inspect the fabric over the top surface. (a) If no cracks or fraying are found, reinstall the metallic strip. (b) If any cracks or fraying are found, prior to further flight, add a fabric reinforcement strip (pinked-tape) starting from a line formed by the intersection of the windshield and the leading edge of the channel and extending aft at least three inches from the trailing edge of the channel. Reinstall the metallic strip. 3. For all airplanes which have fabric installed within thelast three years, every 100 hours in service after the effective date of this AD, until three years are accumulated, inspect the fabric over the top surface. If any cracks or fraying are found, prior to further flight, install Piper Kits or equivalent parts approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, as in 1(a) or 1(b) as appropriate. Upon request and appropriate substantiating data submitted through an FAA maintenance inspector, the compliance time specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. Piper Service Letter No. 362A pertains to this subject. This AD supersedes AD 61-06-06. Amendment 39-1919 was effective August 16, 1974. Amendment 39-2154 was effective April 9, 1975. This amendment 39-2865 is effective April 11, 1977.
2008-01-04: We are superseding an existing 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: * * * * * The Bombardier CL-600-2B19 airplanes have had a history of flap failures at various positions for several years. Flap failure may result in a significant increase in required landing distances and higher fuel consumption than planned during a diversion. * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
98-15-07: This amendment adopts a new airworthiness directive (AD), applicable to certain Dassault Model Mystere-Falcon 50 series airplanes, that requires installation of a reinforcement fitting at the junction of the baggage floor and frame 35 on both the left- and right-hand sides of the airplane. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent fatigue cracking in the subject area, which could result in reduced structural integrity of the airframe.
77-07-03: 77-07-03 PIPER AIRCRAFT CORPORATION: Amendment 39-2860. Applies to models PA-31, PA-31-325, S/N's 31-7612090, 31-7612091, 31-7612092, 31-7612096, 31-7612098, 31-7612101 thru 31-7612105, 31-7612107 thru 31-7612110, and 31-7712002 thru 31-7712006; PA-31-350, S/N's 31-7652142 thru 31-7652147, 31-7652149, 31-7652151, 31-7652152, 31-7652154 thru 31-7652162, 31-7652164, 31-7652165, 31-7652167, 31-7652168, 31-7652171 thru 31-7652174, 31-7652176, 31-7752002 thru 31-7752007, 31-7752009 thru 31-7752011, 31-7752013, 31-7752014 and 31-7752017 thru 31-7752026. Compliance required within 100 hours in service. To prevent electrical terminals form shorting against the relay cover, accomplish the following: (a) Replace the spacer bushings in accordance with the instructions given in Service Bulletin No. 535, dated December 15, 1976, or with an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (b) Upon request, with substantiating data submitted through an FAA Maintenance Inspector, the compliance time specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This Amendment becomes effective April 4, 1977.
85-14-06 R1: 85-14-06 R1 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE (SNIAS): Amendment 39-5089 as amended by Amendment 39-5121. Applies to all Aerospatiale Model AS 350 and AS 355 series helicopters certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent possible failure of the fuselage to tailboom interface frame, accomplish the following: (a) For helicopters which have 1,100 hours or more time in service on the effective date of this AD, inspect in accordance with paragraph (d) within the next 100 hours time in service. (b) For helicopters which have 900 hours or more but less than 1,100 hours time in service on the effective date of this AD, inspect in accordance with paragraph (d) before reaching 1,200 hours time in service. (c) For helicopters which have less than 900 hours time in service on the effective date of this AD, inspect in accordance with paragraph (d) within the next 300 hours time in service. (d) Inspect the bolts for torque and, if necessary, the frame for cracks at the fuselage-to-tailboom interface in accordance with: (1) Service Bulletin No. 05.16 for Model AS 350 series helicopters. (2) Service Bulletin No. 05.14 for Model AS 355 series helicopters. (e) In addition for AS 355 series helicopters, conduct the following initial visual inspection within the next 100 hours time in service for helicopters which have 1,100 hours or more time in service on the effective date of this AD or before reaching 1,200 hours time in service for those helicopters having less than 1,100 hours total time in service on the effective date of this AD: (1) Remove the tailboom from the fuselage in accordance with the Model AS 355 maintenance manual, or FAA-approved equivalent, as appropriate. Prior to tailboom removal, inspect the bolts for torque readings in accordance with paragraph (d). (i) Visually inspect the aft fuselage frame at the fuselage tailboom interface for cracks. Conduct the visual inspection on all accessible frame areas with special emphasis in frame flange radii and at bolt holes. (ii) Conduct dye penetrant inspections of areas of suspected cracks that cannot be verified by a visual inspection. (2) If all the bolt torque readings from the inspection of SB No. 05.14 are 26.5-inch-pounds or greater, the following RH upper quadrant (looking forward) frame inspection may be conducted in lieu of the full frame inspection of paragraph (1): (i) Remove the bolts common to the tailboom, fuselage frame, and RH fuselage frame radius block. (ii) Remove the RH radius block after grinding off the three rivet heads which retain the radius block. The radius block is shown as Item 21 of detail C of page 10 of Aerospatiale Repair Manual 53.10.22, Volume 1. (iii) Visually inspect the forward side of the RH aft fuselage frame for cracks. Conduct the visual inspection on all accessible frame areas with special emphasis in frame flange radii and at bolt holes. (iv) Conduct dye penetrant inspections of areas of suspected cracks that cannot be verified by a visual inspection. (v) Apply zinc chromate primer to the aft surface of the radius block; replace it using the original bolts, but do not re-rivet to the frame. (3) Report cracks and bolt torque values measured before tailboom or radius block removal to the Manager, Aircraft Certification Division, Federal Aviation Administration, P.O. Box 1689, Fort Worth, Texas 76101 within 10 days of the inspection. Use a copy of View F of Service Bulletin No. 05.14 or No. 05.16 to show the locations of cracks or loose fasteners (those below 26.5 inch-pounds of measured torque). If all fasteners are found to have a torque of 26.5 inch-pounds or greater, a statement of such is sufficient without a marked-up View F. Provide aircraft serial numbers, total time, and time since tailboom removal, if any. (Reporting is approved by the Office of Managementand Budget under OMB No. 2120-1156.) NOTE: The initial visual inspection of paragraph (e) and reporting of results are required for all Model AS 355 helicopters even if the bolt torque values measured during the inspections of paragraph (d) are 26.5 inch-pounds or greater. (f) Replace any cracked frames or repair in accordance with Service Bulletin No. 05-14 or No. 05-16. (g) Reinstall the tailboom in accordance with the appropriate Model AS 350 or AS 355 maintenance manual, or FAA-approved equivalent, if removed during the inspections and rework of paragraphs (d), (e), and (f). (h) Repeat the inspections required in paragraph (d) at intervals not to exceed 1,200 hours time in service from the last inspection. (i) An alternate method of compliance with this AD and adjustment of repetitive compliance times may be approved by the Manager, Aircraft Certification Division, Federal Aviation Administration, 4400 Blue Mound Road, Fort Worth, Texas 76106 or by the Manager,Aircraft Certification Office, AEU-100, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, Brussels, Belgium. Adjustments in compliance time may be approved upon recommendations of the FAA aviation safety inspector. (j) In accordance with FAR Sections 21.197 and 21.199, flight is permitted to a base where the inspections required by this AD may be accomplished. The manufacturer's specifications and procedures identified and described in this directive are incorporated by reference 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 Aerospatiale Helicopter Corporation, 2701 Forum Drive, Grand Prairie, Texas 75051, Attention: Customer Support. These documents may also be examined in the Rules Docket at the Office of the Regional Counsel, Southwest Region, Federal Aviation Administration Room 156, Building 3B, 4400 Blue Mound Road, Fort Worth, Texas 76106. Amendment 39-5089 became effective July 26, 1985. This amendment 39-5121 becomes effective September 12, 1985.
86-11-05: 86-11-05 GARRETT TURBINE ENGINE COMPANY (formerly the AiResearch Manufacturing Company of Arizona): Amendment 39-5325. Applies to all Model TFE731-2, -3, and -3R turbofan engines installed in aircraft certified in all categories equipped with fan rotor discs, part number 3073436-1, -2, -3, and -4 and 3072162-1, -2, -3, and -4. Compliance is required as indicated unless already accomplished. To prevent failure of the fan rotor disc due to fatigue cracking, accomplish the following: (a) For fan rotor discs with 4000 or less operating cycles on the effective date of this AD, remove fan rotor disc from service before reaching 4100 total operating cycles. (b) For fan rotor discs with more than 4000 but with 6000 or less operating cycles on the effective date of this AD, remove fan rotor disc from service within the next 100 operating cycles or prior to reaching 6005 total operating cycles whichever occurs first. (c) For fan rotor discs with more than 6000 but with 8000 or less operating cycles on the effective date of this AD, remove fan rotor disc from service within the next 5 operating cycles or prior to reaching 8001 total operating cycles whichever occurs first. (d) For fan rotor discs with more than 8000 operating cycles on the effective date of this AD, remove fan rotor disc from service prior to further flight. NOTES: (1) An operating cycle is considered as any engine operating sequence involving an engine start, at least one acceleration to 80 percent or more low pressure rotor speed, and an engine shutdown. An alternate method of determining cycles is to consider each airplane landing as one cycle. Fan rotor discs which have had their cyclic lives calculated by any other method are to have their cyclic lives recalculated using one of the two methods specified above. (2) It has come to the attention of the FAA that some life limit log cards in the engine log book were incorrectly filled out in that the "dash number" of the disc part numbers (i.e., -1/-2/-3/-4) may have been left off. This should be corrected when the disc is next available to read the entire part number. All fan rotor discs are -1, -2, -3, or -4, and are affected by this AD. Fan rotor discs may be returned to GTEC for evaluation. Shipping information is contained in Garrett Alert Service Bulletin No. TFE731-A72-3328 Revision 2 dated April 4, 1986. Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. Upon request, an equivalent means of compliance with the requirements of this AD may be approved by the Manager, Western Aircraft Certification Office, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009. Upon submission of substantiating data by an owner or operator through an FAA maintenance inspector, the Manager, Western Aircraft Certification Office, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009-2007 may adjust the compliance time specified in this AD. This amendment supersedes priority letter AD 86-04-02 issued February 14, 1986. There are two effective dates for this superseding AD. For those operators who received priority letter AD 86-04-02, the effective date of this AD is the date of the priority letter AD (issued February 14, 1986). The effective date of this AD is June 16, 1986, for operators with fan rotor disc part number 3073436-4 or 3072162-4 installed and for operators who did not receive the priority letter AD.
98-13-15: This amendment adopts a new airworthiness directive (AD), applicable to all Dassault Model Mystere-Falcon 200, Fan Jet Falcon, and Mystere-Falcon 20 series airplanes, that requires repetitive inspections to detect cracks at the attaching holes of the wing-to-fuselage fairings and to ensure tightness of the attaching screws; and repair of any discrepancy. This amendment also requires installation of cupwashers under the vertical seams of the upper fairings. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent loss of the wing-to-fuselage upper fairings during flight, which could result in the fairings impacting the engines or tail sections, and consequent reduced controllability of the airplane.
95-18-06 R1: 95-18-06 R1 Hamilton Standard: Priority Letter issued on August 30, 1995. Docket No. 95-ANE-50. Revises Priority Letter 95-18-06, issued August 28, 1995. Applicability: Hamilton Standard Models 14RF-9, 14RF-19, 14RF-21, and 14SF-5, 14SF-7, 14SF-11, 14SFL11, 14SF-15, 14SF-17, 14SF-19, and 14SF-23; and Hamilton Standard/British Aerospace 6/5500/F propellers installed on but not limited to Embraer EMB-120 and EMB 120-RT; SAAB-SCANIA SF 340B; Aerospatiale ATR42-100, ATR42-300, ATR42-320, ATR72; DeHavilland DHC-8-100 series, DHC-8-300 Series; Construcciones Aeronauticas SA (CASA) CN-235 series and CN-235-100; Canadair CL-215T and CL-415; and British Aerospace ATP airplanes. NOTE: This AD applies to each propeller identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For propellers that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must use the authority provided in paragraph (h) to request approval from the Federal Aviation Administration (FAA). This approval may address either no action, if the current configuration eliminates the unsafe condition, or different actions necessary to address the unsafe condition described in this AD. Such a request should include an assessment of the effect of the changed configuration on the unsafe condition addressed by this AD. In no case does the presence of any modification, alteration, or repair remove any propeller from the applicability of this AD. Compliance: Required as indicated, unless accomplished previously. To prevent separation of a propeller blade due to cracks initiating in the blade taper bore, that can result in aircraft damage, and possible loss of aircraft control, accomplish the following: (a) For Hamilton Standard Model 14RF-9 propeller blades, installed on Embraer EMB-120 series aircraft, within the next 10 flight cycles after the effective date of this AD, remove from service propeller blades that have been ultrasonically shear wave inspected in accordance with AD 94-09-06 or AD 95-05-03, removed from service due to crack indications, and subsequently reworked and returned to service. These propeller blades include, but are not limited to, the following serial numbers: 847598 851646 852085 852561 853151 854530 854535 854838 855014 855042 855196 855859 857375 858696 859824 860589 867590 876707 880245 (b) For Hamilton Standard Models 14RF-19, 14RF-21, and 14SF-5, 14SF-7, 14SF-11, 14SFL11, 14SF-15, 14SF-17, 14SF-19, and 14SF-23; and Hamilton Standard/British Aerospace 6/5500/F propeller blades, installed on aircraft other than Embraer EMB-120 series aircraft, within the next 10 flight cycles after the effective date of this AD, unless inspected previously in accordance with Telegraphic AD T95-18-51, perform an ultrasonic shear wave inspection for cracks in the blade taper bore of propeller blades that have been ultrasonically inspected in accordance with AD 94-09-06 or AD 95-05-03, removed from service due to crack indications, and subsequently reworked and returned to service. Thereafter, at intervals not to exceed 1,250 flight cycles since last inspection, perform an ultrasonic shear wave inspection for cracks in the blade taper bore of propeller blades. Perform the ultrasonic shear wave inspection in accordance with the Accomplishment Instructions of the following Hamilton Standard Alert Service Bulletins (ASB's), as applicable: No. 14RF-21-61-A68, No. 14SF-61-A88, No. 14RF-19-61-A49, No. 6/5500/F-61-A36; all dated August 25, 1995. Remove propeller blades with crack indications from service and replace with serviceable parts. (c) For Hamilton Standard Model 14RF-9 propeller blades, installed on Embraer EMB-120 series aircraft, not affected by paragraph (a) of this AD, perform ultrasonic shear wave inspections in accordance with the Accomplishment Instructions of Hamilton Standard ASB No. 14RF-9-61- A85, dated August 28, 1995. Remove propeller blades with crack indications from service and replace with serviceable parts: (1) For propeller blades with 1,250 or more flight cycles since last ultrasonic shear wave inspection on the effective date of this AD, or that have not been ultrasonically shear wave inspected, perform an ultrasonic shear wave inspection for cracks within the next 50 flight cycles after the effective date of this AD. (2) For propeller blades with less than 1,250 flight cycles since last ultrasonic shear wave inspection on the effective date of this AD, perform an ultrasonic shear wave inspection for cracks within the next 50 flight cycles after the effective date of this AD, or prior to accumulating 1,250 flight cycles, whichever occurs later. (3) Thereafter, perform repetitive ultrasonic shear wave inspections at intervals not toexceed 1,250 flight cycles since last inspection. (d) For Hamilton Standard Models 14RF-19, 14RF-21, and 14SF-5, 14SF-7, 14SF-11, 14SFL11, 14SF-15, 14SF-17, 14SF-19, and 14SF-23; and Hamilton Standard/British Aerospace 6/5500/F propeller blades; identified by serial number in the ASB's listed in this paragraph, installed on aircraft other than Embraer EMB-120 aircraft, and not affected by paragraph (b) of this AD, perform ultrasonic shear wave inspections in accordance with the Accomplishment Instructions of Hamilton Standard ASB's, as applicable: No. 14RF-21-61-A69, No. 14SF-61-A89, No. 14RF-19- 61-A50, No. 6/5500/F-61-A37; all dated August 28, 1995. Remove propeller blades with crack indications from service and replace with serviceable parts: (1) For propeller blades with 1,250 or more flight cycles since last ultrasonic shear wave inspection on the effective date of this AD, or that have not been ultrasonically shear wave inspected, perform an ultrasonic shear wave inspection for cracks within the next 150 flight cycles after the effective date of this AD. (2) For propeller blades with less than 1,250 flight cycles since last ultrasonic shear wave inspection on the effective date of this AD, perform an ultrasonic shear wave inspection for cracks within the next 150 flight cycles after the effective date of this AD, or prior to accumulating 1,250 flight cycles, whichever occurs later. (3) Thereafter, perform repetitive ultrasonic shear wave inspections at intervals not to exceed 1,250 flight cycles since last inspection. (e) No ultrasonic shear wave inspections are required for Hamilton Standard Models 14RF- 19, 14RF-21, and 14SF-5, 14SF-7, 14SF-11, 14SFL11, 14SF-15, 14SF-17, 14SF-19, and 14SF-23; and Hamilton Standard/British Aerospace 6/5500/F propeller blades, that have been shotpeened in the taper bore during manufacture, and not identified by serial numbers in the ASB's listed in paragraph (b) of this AD. (f) Propeller blades removed from service in accordance with this AD may not be returned to service. (g) For the purpose of this AD, a flight cycle is defined as one takeoff and the next landing of an aircraft. In addition, each touch and go is defined as a flight cycle, and each water load pick up for amphibian aircraft operation is defined as a flight cycle. (h) 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, Boston Aircraft Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Boston Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Boston Aircraft Certification Office. (i) Special flight permits may be issued in accordance withsections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the aircraft to a location where the requirements of this AD can be accomplished. (j) Copies of the applicable service information may be obtained from Hamilton Standard, One Hamilton Road, Windsor Locks, CT 06096-1010; telephone (203) 654-6876. This information may be examined at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA. (k) Priority Letter AD 95-18-06 R1, issued August 30, 1995, becomes effective upon receipt. (l) Priority Letter AD 95-18-06 R1 revises priority letter 95-18-06 issued August 25, 1995. This Revised Priority Letter AD 95-18-06 R1 revised Priority Letter AD 95-18-06 issued on August 28, 1995, and supersedes the Telegraphic AD T95-18-51 issued on August 25, 1995.