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85-11-51 R1: 85-11-51 R1 SAAB-FAIRCHILD: Amendment 39-5145 as amended by amendment 39-5857. Applies to all Saab-Fairchild Model SF-340A airplanes certificated in any category. Compliance is required before further flight after the effective date of this airworthiness directive (AD). To prevent incorrect attitude indications, accomplish the following, unless previously accomplished: A. Incorporate the following information into the limitations section of the FAA- approved Airplane Flight Manual and provide to flight crews: "During the alignment or initialization period, an inertial system is susceptible to bus voltage transients and aircraft movement. The method traditionally used to initialize an inertial system is to apply power to the system and to keep the aircraft stationary until all errors in the system are biased to zero. Aircraft movement due to taxiing will cause inertial errors that are excessive. The following procedures must be used to correctly initialize inertially-base attitude heading reference systems (AHRS) to establish the correct attitude and heading references with respect to earth references: 1. AHRS initialization to be performed with both engines running, i.e., external power switched to off and both generators on line prior to applying power to the L and R avionics busses. 2. Approximately 70 seconds after avionics power is applied, the AHRS initialization is completed by the presentation of attitude on EFIS and the removal of the attitude flags from the displays. During initialization, ensure that the aircraft is not moved and there is no operation of brakes, flaps, nosewheel steering, i.e., the hydraulic pump is not to be operated. Also, there should be no changes made in engine power/prop settings. 3. After the system is initialized, as indicated by the attitude flag being out of view, the aircraft may be taxied and engine run-ups performed. Takeoff may not be made until the system has been operating atleast two minutes after initialization is completed, the attitude difference between the attitude displayed on both EFIS electronic attitude direction indicators (EADI's) and the standby attitude indicator is 3 degrees or less (either bank or pitch), and the heading on the compass card is not slewing away from the aircraft heading. 4. If the attitude error exceeds 3 degrees or if the compass heading slews away from the aircraft heading when checked in accordance with step 3., above, stop the aircraft, set brakes, stabilize engine power settings, and remove avionics power from the affected system by pulling the AHRS circuit breakers (AHC Avion and BAT) which will remove the attitude display from the EFIS. Reset circuit breakers and repeat initialization procedure and checks as in steps 2 and 3, above. NOTE: An incorrect initialization on the ground cannot be corrected by a reinitialization while airborne." B. "Accomplishment of Modification 1438 in accordance with SAAB Service Bulletin SF 340-34-038, dated October 24, 1986, or an equivalent production change constitutes terminating action for requirements of paragraph A. of this AD. Thereafter, the AHRS initialization shall be accomplished in accordance with SAAB Aircraft Operations Manual (AOM) Bulletin Number 24. A copy of AOM Bulletin Number 24 must be readily available to the crew during operations." C. Alternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: Compliance with paragraph A. of this directive may be effected by including a copy of this AD in the limitations section of the FAA-approved airplane flight manual and operating manual. All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Saab-Fairchild Product Support, S.58188, Linkoping, Sweden. These documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. Amendment 39-5145 became effective October 21, 1985. It was effective earlier to all recipients of telegraphic AD 85-11-51 issued May 31, 1985. This Amendment, 39-5857 becomes effective April 6, 1988.
2010-23-24: This amendment adopts a new airworthiness directive (AD) for the Sikorsky Model S-70A and S-70C helicopters. This AD requires an ultrasonic test (UT) inspection of the tail gearbox output bevel gear (gear) for a crack. If you find a crack, replacing the gear with an airworthy gear is required before further flight. This AD is prompted by three gear cracking incidents, one of which resulted in the tail rotor separating from the helicopter. The actions specified by this AD are intended to detect a crack in the gear to prevent a tail rotor separating, loss of tail rotor control, and subsequent loss of control of the helicopter.
2001-16-08: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 747 series airplanes equipped with General Electric Model CF6-45 or -50 series engines or Pratt & Whitney Model JT9D-3, -7, or -70 series engines; and all 747-E4B (military) airplanes. That AD currently requires repetitive inspections to detect cracking or fracture of the steel attachment fittings of the diagonal brace to the nacelle struts; and replacement of the attachment fittings with new steel fittings, if necessary. This amendment adds new repetitive inspections of the fasteners of the steel attachment fittings of the diagonal brace to the inboard and outboard nacelle struts to find discrepancies; and mandates certain one-time inspections of the existing attachment fittings, installation of new fasteners, and replacement or rework of the fittings, which terminates the repetitive inspections. This amendment is prompted by a report of fatigue cracking in a steel attachment fitting of the diagonal brace to the number 2 nacelle strut. The actions specified by this AD are intended to prevent such cracking or a fracture, which could result in failure of a nacelle strut diagonal brace load path and possible separation of the nacelle from the wing.
91-18-16: 91-18-16 FOKKER: Amendment 39-8019. Docket No. 91-NM-15-AD. Applicability: Model F-27 series airplanes; Serial Numbers 10102 through 10684, 10686, 10687, and 10689 through 10692; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent elevator and trim tab flutter during flight and reduced controllability of the airplane, accomplish the following; (a) Within 100 hours time-in-service, unless accomplished within the previous 400 hours time- in-service, and thereafter at intervals not to exceed 500 hours time-in-service, perform a visual inspection to detect worn, loose, cracked, or broken parts in the elevator trim system, in accordance with the appropriate maintenance instructions referenced in the Fokker F27 Maintenance Circular 55-3, dated September 10, 1985. Repair any discrepant part(s) prior to further flight. (b) Within 18 months after the effective date of this AD, modify the elevator trim systemin accordance with the Accomplishment Instructions of Fokker Service Bulletin F27/27-130, dated September 11, 1990. Accomplishment of this modification constitutes terminating action for the repetitive visual inspections required by paragraph (a) of this AD. NOTE: This terminating action does not preclude the visual inspections of the elevator and trim tab that should be considered at the Check 4 or the 2C-check interval, which are recommended in Fokker Service Bulletin F27/27-130, dated September 11, 1990. (c) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. (e) The inspection requirement shall be done in accordance with Fokker F27 Maintenance Circular 55-3, dated September 10, 1985, and the modification requirement shall be done in accordance with Fokker Service bulletin F27/27-130, dated September 11, 1990. 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 Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-8019, AD 91-18-16) becomes effective on October 8, 1991.
78-19-03: 78-19-03 MESSERSCHMITT-BOLKOW-BLOHM GmbH: Amendment 39-3302. Applies to all Model BO-105 helicopters, all serial numbers, equipped with Autoflug safety belts types BAGU FAG-5 and FAG-7 with type GL-2 buckles and a date of manufacture prior to October 1, 1974. Compliance required as indicated unless already accomplished. To reduce the possibility of the safety belt release mechanism either jamming or requiring release forces beyond a specified range, while the belt is stressed under a load condition, accomplish the following: (a) Within the next 30 hours time in service after the effective date of this AD, inspect the safety belts and if the date of manufacture stamped on the belt's name tag is prior to October 1, 1974, before further use, except that the aircraft may be flown in accordance with FAR 21.197 and 21.199 to a place where repairs or replacement can be made - (1) Replace the safety belt assembly with a serviceable safety belt assembly of the same partnumber; or (2) Repair the GL-2 buckle assembly in accordance with Augoflug Technical Instruction No. 3/77, or an FAA-approved equivalent, and test the buckle and release mechanism for proper operation in accordance with paragraphs (a)(2)(i) through (iv) of this AD, or an FAA-approved equivalent (i) Checking the Locking and Release Action of the Buckle. Set the opening lever of the buckle to position "RELEASE" and press the lever against the stop. Check that the lever flange of the body is at least level with the upper edge of the cover and that it is slightly higher on the side opposite the fixed lug. Release the operating handle and check that it will automatically return to position "DON". Lift the operating lever position "RELEASE" and press it at right angles to its longitudinal axis to place the body in an inclined position. Check that the body tumbles freely in the cover. (ii) Checking the Arresting Mechanism of the Operation Lever. In position "LOCKED" the release force of the arresting mechanism of the operating lever - measured at the strap handle and at right angles to the lever - must be 4.4 + 1.4 lbs. (iii) Checking the Locking Mechanism of the Multiple-Point Buckle. Insert the belt lugs into the multiple-point buckle and subject each lug to a load of approximately 20 lbs. with the load applied at right angles to the buckle plane. Perform this test once with the cover of the buckle pointing upwards and once downwards, ensuring that the operating lever and body are unobstructed. The buckle must not open and must keep the belt lugs securely locked. (iv) Checking the Force to Release. With the safety belt fastened and positioned to simulate an inverted aircraft/seat position and while applying 250 lbs. load uniformly distributed across the buckle and belt webbing, measure the force necessary on the operating lever to actuate belt release. The measured force to actuate belt release must not be greater than 45 lbs. (b) Mark the safety belts repaired and tested to comply with paragraph (a)(2) of this AD with permanent legible marking either on the manufacturer's name tag or on the belt webbing near the name tag with the number of this AD and the date repaired. (c) If the safety belt is tested prior to repair in accordance with paragraph (a)(2) and complies with the limits specified in paragraphs (a)(2)(i) through (iv) of this AD, the compliance time to effect repairs is extended to 150 hours time in service after the effective date of this AD, at which time the GL-2 buckle assembly must be replaced or repaired and retested in accordance with paragraph (a) of this airworthiness directive. (d) Equivalent means of compliance with the AD must be approved by the Chief, Aircraft Certification Staff, Europe, Africa, and Middle East Region. This amendment becomes effective October 18, 1978.
91-16-02: 91-16-02 AIRBUS INDUSTRIE: Amendment 39-7092. Docket No. 91-NM-59-AD. Applicability: Model A320 series airplanes, on which Modification 22039 has not been accomplished, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To ensure complete closure of the low pressure fuel fire shut-off valve, accomplish the following: A. Within 350 hours time-in-service after the effective date of this AD, and thereafter at intervals not to exceed 350 hours time-in-service, perform a functional check of the low pressure fuel fire shut-off valve actuator, in accordance with Airbus Industrie All Operators Telex (AOT) 28-01, dated October 8, 1990. B. If any valve fails or indicates failure to open or close correctly, prior to further flight, replace the low pressure fuel fire shut-off valve actuator with either P/N HTE 190001 (PRE Airbus Industrie Service Bulletin A320-28-1028), or P/N HTE 190001-1 (POST Airbus Industrie ServiceBulletin A320-28-1028), in accordance with Airbus Industrie Service Bulletin A320-28-1028, Revision 1, dated November 23, 1990. Following actuator replacement, perform a functional test in accordance with AOT 28-01, dated October 8, 1990. Accomplishment of Modification 22039 (Service Bulletin A320-28-1028) constitutes terminating action for the repetitive functional checks required by paragraph A. of this AD. C. An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. E. The repetitive functional check requirements shall be done in accordance with Airbus Industrie All Operators Telex (AOT) 28-01, dated October 8, 1990. The replacement requirements shall be done in accordance with Airbus Industrie Service Bulletin A320-28-1028, Revision 1, dated November 23, 1990. 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 Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-7092, AD 91-16-02) becomes effective on September 11, 1991.
2022-26-05: The FAA is adopting a new airworthiness directive (AD) for all Rolls-Royce Deutschland Ltd & Co KG (RRD) TAY 620-15 and TAY 650-15 model turbofan engines. This AD was prompted by reports of cracks on the high-pressure turbine (HPT) stage 2 intermediate air seal attachment bolts (attachment bolts). This AD requires repetitive inspections of the HPT stage 2 intermediate air seal and attachment bolts and, depending on the results of the inspections, replacement of attachment bolts and the HPT stage 1 and stage 2 rotor disks, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2001-16-07: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-400 and 767 series airplanes, that requires modification of the core cowl assemblies of the engines. This action is necessary to prevent failure of the core cowl latches during an engine fire, and consequent in-flight separation of an engine core cowl and its strut fire barrier from the airplane. This action is intended to address the identified unsafe condition.
87-02-06: 87-02-06 GATES LEARJET: Amendment 39-5520. Applies to the following Gates Learjet series airplanes, models/serial numbers listed below, certificated in any category; except those airplanes equipped with Part Number (P/N) 2651034 forward engine mount assembly due to spare replacements: MODEL SERIAL NUMBER 35 001 through 522 36 001 through 053 55 001 through 107 Compliance required as indicated, unless previously accomplished. To ensure the structural integrity of the forward engine mounts, accomplish the following: A. Prior to the accumulation of 2,400 hours time-in-service or 2,400 landings (whichever occurs first), or within the next 75 hours time-in-service after the effective date of this AD, whichever occurs later, conduct a visual inspection of the installed left and right forward engine mounts in accordance with Paragraph 2A of Gates Learjet Service Bulletin 35/36-71-3 or 55-71-2, both dated January 5, 1987, or later FAA-approved revision, as appropriate. 1. If no cracks are found, repeat the visual inspection at intervals not to exceed 420 hours time-in-service. 2. If cracks are found, inspect or replace as indicated below: a. For total visible crack length (forward plus aft) of 1.0 inch or more, prior to further flight accomplish one of the following: (1) Replace cracked mount(s) with P/N 2651034 mount assembly; or (2) Conduct the magnetic particle inspection and disposition in accordance with paragraph B. of this AD. b. For total visible crack length (forward plus aft) of less than 1.0 inch, accomplish one of the following: (1) Replace the cracked mount(s) with P/N 2651034 mount assembly within the next 420 hours time-in-service; or (2) Conduct the magnetic particle inspection and the next 420 hours time-in-service. B. Prior to the accumulation of 2,400 hours time-in-service or 2,400 landings (whichever occurs first), or within the next 1,500 hours time-in-service after the effective date of this AD, whichever occurs later, conduct a magnetic particle inspection of the removed left and right engine mounts, in accordance with Paragraph 2B of Gates Learjet Service Bulletin 35/36- 71-3 or 55-71-2, both dated January 5, 1987, or later FAA-approved revisions, as appropriate. 1. If no cracks are found, repeat the inspection at intervals not to exceed 1,500 hours time-in-service. 2. If cracks are found, replace as indicated below: a. For total crack lengths (forward plus aft) of 3.0 inches or more, replace cracked mount(s) with P/N 2651034 mount assembly prior to further flight. b. For total crack lengths (forward plus aft) of less than 3.0 inches, replace cracked mount(s) with P/N 2651034 mount assembly within 420 hours time-in-service. C. The installation of a P/N 2651034 mount assembly constitutes terminating action for the repetitive inspections required by paragraphs A. and B. of this AD. D. Duplicate copiesof the Compliance Response form, included in Gates Learjet Service Bulletins 35/36-71-3 and 55-71-2, both dated January 5, 1987, used for reporting the results of the initial visual and magnetic particle inspections, must be submitted within one week after the inspection to the FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. E. 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. F. Alternate means of compliance, which provide an acceptable level of safety, may be used when approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. All persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Gates Learjet Corporation, P.O. Box 7707, Wichita, Kansas 67277. This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid- Continent Airport, Wichita, Kansas. This amendment becomes effective February 6, 1987.
90-25-06: 90-25-06 CESSNA AIRCRAFT COMPANY: Amendment 39-6818. Docket No. 90-NM-253-AD. Applicability: Model 650 Series airplanes; Serial Nos. 650-0094 through -0096, -0098, -0136 through -0139, -0149, -0161, -0167, -0170, -0172, -0173, -0176 through 0182, -0184 through -0189, and -0192; equipped with Honeywell SPZ-8000 Digital Automatic Flight Control System; certificated in any category. Compliance: Required within 25 hours time-in-service after the effective date of this AD, unless previously accomplished. To prevent complete loss of the pilots' primary flight instrument displays, accomplish the following: A. Replace the distance measuring equipment (DME) antenna cable connectors in accordance with Steps l through 3 of the Accomplishment Instructions of Cessna Citation Alert Service Bulletin A650-34-68, dated November 2, 1990. B. Accomplish either subparagraph B.1. or B.2., below: 1. Check the electrical bonding of the DME antennas to airplane structurein accordance with Step 4 of the Accomplishment Instructions of Cessna Citation Alert Service Bulletin A650-34-68, dated November 2, 1990. a. If the resistance is greater than 0.010 Ohms, prior to further flight, rework the antennas installation in accordance with Step 4 of the service bulletin. b. If the resistance is equal to or less than 0.010 Ohms, restore the system for use in accordance with Steps 5 through 7 of the Accomplishment Instructions of the service bulletin. 2. Rework the antenna installation in accordance with Step 4 of Cessna Citation Alert Service Bulletin A650-34-68, dated November 2, 1990; and restore the system for use in accordance with Steps 5 through 7 of the Accomplishment Instructions of the service bulletin. C. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Wichita Aircraft Certification Office (ACO), ACE-130W, FAA, CentralRegion. NOTE: The request should be submitted directly to the Manager, Wichita ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Wichita ACO. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Cessna Aircraft Company, Customer Services, P.O. Box 1521, Wichita, Kansas 67201. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Wichita, Kansas. This amendment (39-6818, AD 90-25-06) becomes effective on December 11, 1990.
2016-25-22: We are adopting a new airworthiness directive (AD) for all Viking Air Limited Models DHC-2 Mk. I, DHC-2 Mk. II, and DHC-2 Mk. III airplanes that supersedes AD 2016-19-08. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as corrosion of the elevator control rod and of the elevator actuating lever on the control column, which could cause these components to fail. We are issuing this AD to require actions to address the unsafe condition on these products.
74-15-06: 74-15-06 SIKORSKY AIRCRAFT: Amendment 39-1890 as amended by Amendment 39-2134 and 39-2447 is further amended by Amendment 39-3089. Applies to all S-58 model series helicopters equipped with magnesium input and/or intermediate housings. Compliance required as indicated unless already accomplished. To preclude possible failure of the magnesium input and/or intermediate housings of the tail rotor gearbox assembly, P/N S1635-64000 series, accomplish the following: 1. Within the next three hours time in service after the effective date of this AD, or whenever unusual vibration in the tail rotor system is experienced, accomplish the following in accordance with the instructions set forth in Section 2, Paragraphs A and C of Sikorsky Service Bulletin No. 58B35-17C, dated November 24, 1975 (hereinafter Bulletin No. 58B35-17C). a. Inspect the nuts attaching the tail rotor gearbox assembly to the pylon fitting. b. Inspect the input and intermediate housings of the gearbox. c. Retorque the nuts that attach the input to intermediate housing. 2. After completion of the initial inspections required in Paragraph 1, conduct the repetitive inspections specified in a, b, and c below as applicable. Sections and paragraphs listed are part of Bulletin No. 58B35-17C. a. Every ten hours time in service inspect in accordance with Section 2, Parargraphs A(2) and C(2), (a) thru (d), if both Sikorsky Service Bulletins listed below are not complied with or lightweight input and/or intermediate tail rotor gearbox housings are installed: (1) 58B15-14B, Balance of Rotary Rudder Assembly. (2) 58B20-14B, Strengthening of Pylon Carry-Thru Structure of Rotary Rudder Assembly. b. Every 20 hours time in service, inspect in accordance with Section 2, Paragraphs C(2), (a) thru (d), if tail rotor gearbox S1635-64000-9 is installed and if both Sikorsky Service Bulletins 58B15-14B and 58B20-14B or their latest FAA approved revisions are compliedwith. c. Every 50 hours time in service, inspect in accordance with Section 2, Paragraphs C(2), (a) thru (d), if tail rotor gearbox S1635-64000-10 is installed and if both Sikorsky Service Bulletins 58B15-14B and 58B20-14B or their latest FAA approved revisions are complied with. 3. Within the next 25 hours time in service after the effective date of this AD, remove the tail rotor gearbox assembly (P/N S1635-64000 series) from the aircraft and accomplish the one-time inspection set forth in Section 2, Paragraph B, of Bulletin No. 58B35-17C. 4. Every 50 hours time in service after the inspection required in Paragraph 1 above, inspect the nuts attaching the tail gearbox to the pylon fitting in accordance with Section 2, Paragraph A(1) of Bulletin No. 58B35-17C. If heavy weight input housing, P/N S1635-64009-3N, and heavyweight intermediate housing, P/N S1635-64074, are both installed, inspect in accordance with Section 2, Paragraph D(2), of Bulletin No. 58B35-17C. 5. Tail rotor gearbox assembly (P/N S1635-64000 series) input and intermediate housings and pylon fittings (S1620-64129) found to have evidence of cracks shall be replaced with an acceptable assembly prior to further flight. All other defects found as a result of the inspections required by Paragraphs 1, 2, 3, and 4 shall be repaired in accordance with instructions contained in Bulletin No. 58B35-17C prior to further flight. 6. Remove from service magnesium input and intermediate housings P/N's S1635-64009-3N, S1635-64074, S1635-64009, S1635-64010 as follows: a. Prior to July 31, 1978, for all turbine engine powered S-58T helicopters. b. Prior to December 31, 1978, for all piston engine powered S-58 helicopters. Replace with aluminum input and intermediate housings S1635-64009-101 and S1635-64074-103 respectively. NOTE: Sikorsky Service Bulletin No. 58B35-25 covers this subject. 7. Upon submittal of substantiating data through a Federal Aviation Administration inspector, the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region, may adjust the compliance time. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Sikorsky Aircraft, Division of United Aircraft Corporation, Stratford, Connecticut 06602. These documents may also be examined at the Office of the Regional Counsel, New England Region, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803 and at Federal Aviation Administration Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the Federal Aviation Administration at its headquarters in Washington, D.C., and at the New England Regional Office in Burlington, Massachusetts. Amendment 39-1890 became effective July 27, 1974. Amendment 39-2134 became effective April 3, 1975. Amendment 39-2447 became effective December 10, 1975. This amendment 39-3089 becomes effective December 30, 1977.
73-26-01: 73-26-01 PIPER: Amdt. 39-1754 as amended by Amendment 39-1772. Applies to Model PA-25 airplanes Serial Numbers 25-1 through 25-1999, PA-25-235 and PA-25-260, Serial Numbers 25-02 and 25-2000 through 25-74005573, certificated in all categories. Compliance required as indicated. For Model PA-25-235 and Model PA-25-260 airplanes with forward wing spar(s) with 2000 hours or more time in service on the effective date of this AD, compliance is required within 5 hours time in service after the effective date of this AD, unless already accomplished within the last 295 hours time in service, and thereafter at intervals not to exceed 300 hours time in service from last inspection. For Model PA-25-235 and Model PA-25-260 airplanes with forward wing spar(s) with less than 2000 hours time in service on the effective date of this AD, compliance is required upon accumulation of 2000 hours time in service or within the next 5 hours time in service, whichever is later, unless already accomplished within the last 295 hours time in service, and thereafter at intervals not to exceed 300 hours time in service from last inspection. For Model PA-25 (150 h.p.) airplanes, compliance is required upon accumulation of 2000 hours wing spar time in service or within the next five hours time in service after the effective date of this AD, whichever is later, unless already accomplished. To detect cracks in the forward wing spar, accomplish the following: (a) Remove left and right wing from fuselage. (b) Remove wing attach fitting located on each forward wing spar root. (c) Using standard dye or fluorescent penetrant inspection procedures, or equivalent approved by the Chief, Engineering and manufacturing Branch, FAA Southern Region, inspect the following areas for cracks: 1. The inboard end of left and right forward spar web and doubler in the area of wing spar attach fitting bolt holes. 2. The left and right forward spar lower cap rear flange aroundthe inboard four skin attachment holes. (d) If cracks are found as a result of the inspections required in paragraph (c), the affected parts must be replaced with serviceable parts of the same part number or repaired in accordance with the instructions contained in Piper Service Bulletin No. 414, before further flight. (e) When Piper Kit No. 760840, Spar Web Reinforcement Plate or equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Southern Region, is installed immediately after inspection and repair as required by this AD, repetitive inspections at 300 hour intervals are no longer necessary. (f) Report in writing within 10 days any cracks found during the inspections required by this AD to Chief, Engineering and Manufacturing Branch, ASO-210, FAA, Southern Region, P.O. Box 20636, Atlanta, Georgia 30320. Each report must include aircraft model, serial and registration numbers, location of cracks, length of cracks and number of hours wing spar time in service. (Reporting approved by the Bureau of the Budget under BOB No. 04-R0174.) Piper Service Bulletin No. 410 pertains to the inspections required by this AD and Piper Service Bulletin No. 414 pertains to Piper Kit No. 760840, Spar Web Reinforcement Plate. Amendment 39-1754 became effective December 17, 1973 for all persons except those to whom it was made effective upon receipt of the airmail letter dated December 7, 1973 which contained this amendment. The Amendment 39-1772 becomes effective January 28, 1974.
2001-16-19: This amendment adopts a new airworthiness directive (AD), applicable to all Boeing Model 747-100 and -200 series airplanes modified by supplemental type certificate ST00196SE, that requires modification of the in-flight entertainment (IFE) system and revision of the Airplane Flight Manual. This action is necessary to ensure that the flight crew is able to remove electrical power from the IFE system when necessary and is advised of appropriate procedures for such action. Inability to remove power from the IFE system during a non-normal or emergency situation could result in inability to control smoke or fumes in the airplane flight deck or cabin. This action is intended to address the identified unsafe condition.
91-14-05: 91-14-05 AVCO LYCOMING: Amendment 39-6955. Docket No. 91-ANE-09. Applicability: Avco Lycoming ALF 502L series turbofan engines installed on, but not limited to the Canadair Challenger CL601 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent complete loss of engine power, uncontained disk failure, and possible damage to the aircraft, accomplish the following: (a) Clean, visually inspect, coat with "Sermetel W" or "ALSEAL-518" and reidentify third stage high compressor (HC) rotor disks in accordance with Textron Lycoming Service Bulletin (SB) ALF 502L 72-203, Revision 3, dated December 7, 1990, as follows: (1) For those third stage HC disks with 7,000 cycles since new (CSN) or greater on the effective date of this AD, within the next 500 cycles in service, not to exceed 8,500 CSN. (2) For those third stage HC disks with less than 7,000 CSN on the effective date of this AD, at or prior to accumulating a total of 7,500 CSN. (b) Remove, prior to further flight, third stage HC disks found with evidence of corrosion pitting or cracks. (c) 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. (d) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. The inspection procedures shall be done in accordance with the following Textron Lycoming service bulletin: Document No. Page No. Issue/Revision Date ALF 502L 72-203 1, 2, 4 3 December 7, 1990 ALF 502L 72-203 3, 6 2 August 8, 1990 ALF 502L 72-203 5 1 June 16, 1989This 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 Textron Lycoming, 550 Main Street, Stratford, Connecticut 06497. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street NW, Room 8401, Washington, D.C. This amendment (39-6955, AD 91-14-05) becomes effective on July 15, 1991.
2001-16-17: This amendment adopts a new airworthiness directive (AD), applicable to all Boeing Model 767-300 series airplanes modified by supplemental type certificate (STC) SA5765NM or SA5978NM, that requires removal or modification of the in-flight entertainment (IFE) system installed by those STCs. This action is necessary to prevent the inability of the flight crew to remove power from the IFE system when necessary. Inability to remove power from the IFE system during a non-normal or emergency situation could result in inability to control smoke or fumes in the airplane flight deck or cabin. This action is intended to address the identified unsafe condition.
2016-25-05: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747- 200C, 747-200F, 747-300, 747-400, 747-400D, 747-400F, 747SR, and 747SP series airplanes. This AD was prompted by reports of cracking found in the splice plates, hinge fittings, terminal fittings, the upper skin of the outboard \n\n((Page 93586)) \n\nand center sections, upper chord, and rear spar webs before reaching the inspection interval specified in AD 2006-10-16. Cracked and fractured Maraging steel fasteners were also found. This AD requires repetitive inspections for cracking, an inspection to determine whether fasteners are magnetic, repetitive ultrasonic inspections for cracking and fractures of affected fasteners, and related investigative and corrective actions if necessary. We are issuing this AD to address the unsafe condition on these products.
2001-16-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F27 Mark 050 series airplanes equipped with certain Pratt & Whitney Canada Model PW127B engines. This action requires replacing both torque sensor No. 1 and the electrical connectors on the wiring harness between torque sensor No. 1 and the auto-feathering unit (AFU). This action is necessary to prevent inadvertent autofeathering of the propellers, due to interruption of the torque signal between torque sensor No. 1 and the AFU, which could result in loss of engine power and loss of control of the airplane. This action is intended to address the identified unsafe condition.
72-24-03: 72-24-03 TELEDYNE CONTINENTAL: Amendment 39-1554. Applies to Models IO-346-A, IO-520-B and -C and TSIO 520-B, -E, and -J engines having Teledyne Continental Part Number 631645 oil filter adapter installed, used with AC OF-9-A oil filter assembly. Compliance required as indicated, unless already accomplished. To prevent loss of oil accomplish the following or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, Southern Region, Atlanta, Georgia: (A) Not later than next engine overhaul, replace Teledyne Continental Part Number 631645 oil filter adapter with strengthened Teledyne Continental Part Number 632881 (Kit Part Number 637584A1) in accordance with Teledyne Continental Service Bulletin M70-16 or AC Part Number 5579663 (Package No. 6437861) in accordance with AC instruction sheet 6439067. NOTE: The correct AC part may be identified by a cast-in insignia "A" (raised letter A with a raised bar over the letter A). All other AC adapters areineligible, including those identified by a plain raised letter "A" and a raised letter "A" and a raised dot (A) directly above. The Teledyne Continental adapter is identified by part number only. (B) Within 25 hours' time in service from the effective date of the amendment, unless already accomplished under AD 71-11-04, paragraph (B) inspect the base plate to determine whether the gasket retaining seat is wedge-shaped or rectangular. If the gasket seat is wedge-shaped, replace this part with improved Teledyne Continental Part Number 633750 or AC Part Number 6437508 (Package No. 6436627) base plate having a rectangular shaped gasket retaining seat. NOTE: The required base plate can be identified by the presence of a thin sheet metal square shouldered retaining ring spot welded around the gasket groove to hold the gasket in place. (C) Unless already accomplished within the last 50 hours' of service prior to date of this amendment, accomplish the following within the next 25hours' time in service and at every oil filter element change thereafter: (1) Visually inspect the upper surface of the oil filter adapter face using a light and mirror for indications of radial cracks inward from the outer edge. Replace any cracked adapters with serviceable parts prescribed in paragraph (A). (2) After placing filter element in housing in accordance with oil filter element manufacturers' instructions, install assembled housing and base plate to adapter as follows: a. Clean all gasket and seal surfaces. b. Lubricate new gasket well on both sides using engine oil. c. Install assembly on adapter and turn center stud to a light seal contact by hand. d. Visually inspect base plate to adapter seal for proper positioning and seating. e. Torque center studs to 15 - 18 lb. ft. If torque wrench is not available or center stud is inaccessible to torque wrench, tighten center stud 1 3/4 turns beyond point of initial seal contact. f. Re-attach upper bracket and resafety. g. Operate engine for approximately 5 minutes at 1000-2000 RPM. Check for oil leaks and proper assembly using a light and mirror if necessary. If a leak appears between top of housing and stud, remove stud and check for nicks or visual damage at sealing surface. Correct any damage and re-install using a new copper gasket. DO NOT INCREASE TORQUE TO STOP LEAKS. Continental Service Bulletin M66-6, dated April 28, 1966, refers to this subject. (D) The requirements of Paragraph (C) are no longer applicable when Paragraph (A) has been complied with. This airworthiness directive supersedes Amendment 39-1215, (36 F.R. 9241, 9242), A.D. 71-11-04, as amended by Amendments 39-1256, (39 F.R. 14127) and 39-1289, (39 F.R. 18373). This amendment becomes effective November 15, 1972.
90-20-04: 90-20-04 WYTWORNIA SPRZETU KOMUNEKACYJNEGO PZL-MIELEC: Amendment 39-6722. Docket No. 90-CE-17-AD. Applicability: Models M18 and M18A (Dromader) (Serial Numbers 1Z001-01 through 1Z021-20) airplanes, certificated in any category. Compliance: Required within the next 100 hours time-in-service after the effective date of this AD, unless already accomplished. To prevent failure of the engine push-pull cables and loss of engine control, accomplish the following: (a) Remove the engine throttle control and the propeller governor push-pull control and replace those cables in accordance with the instructions and part numbers referenced in PZL- Mielec Mandatory Engineering Bulletin No. K/02.127/89, dated February 1990. (b) The airplane may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (c) An alternate method of compliance or adjustment of the compliance time, which provides an equivalent level of safety, may be approvedby the Manager, Brussels Aircraft Certification Office, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, B- 1000 Brussels, Belgium; Telephone 322-513.38.30 extension 2710/2711. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. All persons affected by this directive may obtain copies of the document referred to herein upon request to Wytwornia Sprzetu Komunekacyjnego PZL-Mielec 39-301 Mielec, Poland; or may examine this document at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This amendment (39-6722, AD 90-20-04) becomes effective on October 17, 1990.
66-21-01: 66-21-01 PRATT & WHITNEY: Amdt. 39-279 Part 39 Federal Register August 31, 1966. Applies to Model JT3C-7 and JT3C-12 Turbojet Engines. Compliance required as indicated. Within the next 6,000 hours' time in service after the effective date of this AD or at the next compressor disassembly, whichever occurs first, unless already accomplished, replace the first, fifth and sixth stage compressor rotor disc spacer assemblies, P/N's 359410, 359414, and 359415, with spacer assemblies, P/N's 460599, 460604, and 460606, respectively. (Pratt & Whitney Aircraft Turbojet Engine Service Bulletin No. 689, Revision No. 3, dated November 16, 1964, pertains to this subject.) This directive effective September 30, 1966.
2015-02-14: We are superseding Airworthiness Directive (AD) 2009-20-05 for certain Airbus Model A318, A319, A320, and A321 series airplanes. AD 2009-20-05 required one-time inspections for cracking, damage, correct installation, and correct adjustment of the main landing gear (MLG) door hinge and actuator fittings on the keel beam, and corrective actions if necessary. This new AD expands the applicability, reduces the compliance time, and requires repetitive inspections instead of the one-time inspection. This AD also requires revising the maintenance or inspection program. This AD was prompted by reports of cracks on fittings that had successfully passed certain required inspections. We are issuing this AD to detect and correct cracking on the MLG door hinge fitting and actuator fitting on the keel beam, which could lead to in-flight detachment of an MLG door, possibly resulting in injury to persons on the ground and/or damage to the airplane.
2016-25-18: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model BD-700-1A10 and BD-700-1A11 airplanes. This AD requires an inspection for discrepancies of the attachment points of the links between the engine rear mount assemblies, and corrective actions if necessary. This AD was prompted by a report indicating that during maintenance, an engine mount pin was found backed out of the rear mount link, and the associated retaining bolt was also found fractured. We are issuing this AD to address the unsafe condition on these products.
2001-15-23: This amendment adopts a new airworthiness directive (AD), applicable to certain BAe Systems (Operations) Limited Model BAe 146 and Avro 146-RJ series airplanes, that requires identifying the discharge valves and cabin pressure controllers, and replacing them with new parts if necessary. The actions specified by this AD are intended to prevent the installation of incorrect pressurization discharge valves and cabin pressure controllers, which could subject the airframe to excess stress and adversely affect the airframe fatigue life. This action is intended to address the identified unsafe condition.
88-17-09 R1: 88-17-09 R1 MCDONNELL DOUGLAS HELICOPTER COMPANY (MDHC) (HUGHES HELICOPTERS, INC.): Amendment 39-5964 as revised by Amendment 39-6400. Docket No. 88-ASW-24. Applicability: Model 369D, E, F, and FF helicopters, certificated in any category, which have tail rotor transmission/tail boom extension mounting studs, Part Number (P/N) MS51992A803-13 or -14, installed. Compliance: Required as indicated, unless already accomplished. (a) To prevent failure of the tail rotor transmission mounting studs (P/N MS51992A803-13 or -14) on MDHC Model 369 D/E helicopters, accomplish the following: (1) Prior to further flight after the effective date of this AD, conduct an initial inspection of the tail rotor transmission attachment to the tail boom casting for any indications of relative motion between the parts in accordance with MDHC Service Information Notice (SIN) DN-151/EN-39/FN-28, Part I, paragraphs a, b, c, d, and e, dated October 10, 1987. (2) Within the next 25hours' time in service after the effective date of this amended AD, and at intervals not to exceed 25 hours' time in service from the last check, conduct a check for security in accordance with MDHC SIN DN-151/EN-39/FN-28, Part II, paragraph a, dated October 10, 1987. The checks required by this paragraph may be performed by the pilot and must be recorded in accordance with FAR Section 43.9. (3) If there are indications of relative motion between the tail rotor transmission and the tail boom casting found by the inspection and check of paragraphs (a)(1) and (a)(2), prior to further flight, remove the tail rotor transmission and replace all four mounting studs in accordance with MDHC SIN DN-151/EN-39/FN-28, Part III, paragraphs a thru j, dated October 10, 1987. (4) At intervals not to exceed 100 hours' time in service from the last inspection after the effective date of this AD, conduct repetitive inspections of the torque of each mounting nut and reapply torque stripe paint in accordance with MDHC SIN DN-151/EN-39/FN- 28, Part IV, paragraphs a and b, dated October 10, 1987. (b) To prevent failure of the tail boom extension mounting studs (P/N MS51992A803-13 or -14) on MDHC Model 369 F/FF helicopters, accomplish the following: (1) Prior to further flight after the effective date of this AD, conduct an initial inspection of the tail boom extension attachment to the tail boom casting for any indications of relative motion between the parts in accordance with MDHC SIN DN-151/EN- 39/FN-28, Part I, paragraphs a, b, c, d, and e, dated October 10, 1987. (2) Within the next 25 hours' time in service after the effective date of this amended AD, and at intervals not to exceed 25 hours' time in service from the last check, conduct a check of the tail boom extension installation for security in accordance with MDHC SIN DN- 151/EN-39/FN-28, Part II, paragraph a, dated October 10, 1987. The checks required by this paragraph may be performed by the pilot and must be recorded in accordance with FAR Section 43.9. (3) If there are indications of relative motion of fretting products between the tail boom casting and the tail boom extension found by the above inspections or checks, prior to further flight, remove tail boom extension and replace all four studs in accordance with MDHC SIN DN-151/EN-39/FN-28, Part III, paragraphs b thru j, dated October 10, 1987. (4) At intervals not to exceed 100 hours' time in service from the last inspection after the effective date of this AD, conduct repetitive inspections of the torque of each mounting nut and reapply torque stripe paint in accordance with MDHC SIN DN-151/EN-39/FN- 28, Part IV, paragraphs a and b, dated October 10, 1987. (c) An alternate means of compliance which provides an equivalent level of safety may be used when approved by the Manager, Western Aircraft Certification Office, P.O. Box 92007, Worldway Postal Center, Los Angeles, CA 90009-2007. (d) In accordance with Sections 21.197 and 21.199, the helicopters may be flown to a base where compliance may be accomplished. These inspections and check procedures shall be done in accordance with MDHC Mandatory SIN DN-151/EN-39/FN-28, dated October 10, 1987. 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 MDHC Technical Publications, Building 543/D214, McDonnell Douglas Helicopter Company, 5000 E. McDowell Road, Mesa, Arizona 85205-9797, telephone (602) 891-6484. A copy may also be inspected at the Office of the Assistant Chief Counsel, Federal Aviation Administration, 4400 Blue Mound Road, Room 158, Building 3B, Fort Worth, Texas, or at the Office of the Federal Register, 1100 L Street NW, Room 8401, Washington, D.C. This AD revises Amendment 39-5964 (53 FR 30023; August 10, 1988), AD 88-17-09, which became effective on August 24, 1988. This amendment (39-6400, AD 88-17-09 R1) becomes effective on December 26, 1989.