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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.