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98-24-26: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-400 series airplanes, that requires replacing the cam assembly, cam bellcrank assembly, and thrust reverser control switch actuator on all four thrust levers with new components. This amendment is prompted by a report of an uncommanded automatic retraction of the leading edge flaps during takeoff. The actions specified by this AD are intended to prevent such uncommanded automatic retraction, which would seriously degrade liftoff and climb capabilities, and could result in near-stall conditions at a critical phase of the flight.
60-04-04: 60-04-04 LOCKHEED: Amdt. 101 Part 507 Federal Register February 13, 1960. Applies to All Models 049, 149, 649 and 749 Series Aircraft Which Have the Cleveland Pneumatic Model 8298 Series Main Landing Gear Struts Installed With the Removable Side Brace Attachment Collar. Compliance required as indicated. Due to fatigue failures found in the above main landing gear outer cylinder, the following inspections and rework must be accomplished on all main landing gears which have accumulated 25,000 or more hours' time in service. (a) Unless already accomplished in the last 1,000 hours' time in service, within 400 hours' time in service inspect for cracks in the main landing gear outer cylinder surface, at the 0.125-inch radius of the shoulder against which the drag strut-side brace collar retaining nut bears, by means of one of the three methods in (b). Reinspect every 1,000 hours' time in service thereafter, until the rework in (c)(2) is accomplished. Outer cylinders with cracks must be replaced prior to further flight. Cracked cylinders may be returned to service after repair and rework is accomplished in accordance with (c). Rework on all uncracked outer cylinders must be accomplished in accordance with (c)(2) not later than the total accumulated hours' time in service indicated in (b). (b) Inspection and rework: (1) ULTRASONIC SHEAR WAVE DETECTION METHOD. This procedure may be used on cylinders with piston and oil in the cylinder or the cylinder only. Rework in accordance with (c)(2) must be accomplished within 4,000 hours' time in service if the ultrasonic method is used. (2) MAGNETIC PARTICLE DETECTION METHOD. This method requires removing and dismantling of the strut assembly. Rework in accordance with (c)(2) must be accomplished within 4,000 hours' time in service if the magnetic particle method is used. (3) RADIOGRAPHIC METHOD. This method requires the removal of the piston from the cylinder and complete 360 degree coverage. Reworkin accordance with (c)(2) must be accomplished within 3,000 hours' time in service if the radiographic method is employed. (c) Repair and rework instructions: (1) Outer cylinders with cracks in the radius described in (a) and for a distance of 0.5 inch below the radius tangency point circumferentially around the cylinder may be repaired by grinding out to a maximum depth of 0.017 inch. Complete removal of cracks must be verified by magnetic particle inspection or equivalent. If cracks are completely removed as verified by such inspection, remove an additional 0.008 inch of material from the repaired area. Rework may be acceptable on outer cylinders with cracks that cannot be removed by grinding to a depth of 0.017 inch. Such cases may be submitted to the FAA for evaluation of the extent of cracking and to determine if rework is possible. Rework accomplished subsequent to such an evaluation must be in accordance with FAA approved repair instructions. (2) On all cylinders,whether cracks are found or not, rework the area described in (c)(1) as follows: (i) Clean and polish the above cylinder area to remove all tool marks and corrosion. (ii) Shotpeen the above area using steel shot 0.019-0.033-inch diameter to an intensity of 0.012-0.016 A(subscript 2) ALEMENT (Reference LAC Process Bulletin 217M, Revision 1). (d) Upon completion of the rework described in (c)(2), all Model 8298 cylinders shall be reinspected for cracks at periods not to exceed 9,000 hours' time in service using one of the inspection methods noted in (b). Cracked cylinders must be replaced prior to further flight. Cracked cylinders may be returned to service after repair and rework is accomplished in accordance with (c). (Lockheed Service Letter FS/239304 covers this same subject.) Revised April 9, 1960. Revised September 15, 1960.
62-20-02: 62-20-02 GENERAL DYNAMICS/CONVAIR: Amdt. 482 Part 507 Federal Register September 8, 1962. Applies to All Models 22, 22M and 30 Series Aircraft. Compliance required within the next 350 hours' time in service from the effective date of this AD, unless already accomplished. To prevent the aileron override spring from jamming in the compressed position if full opposition loads are applied to the lateral control system, the aileron override spring assembly shall be modified to incorporate a stop in accordance with General Dynamics/Convair Service Bulletins A27-56, A27-25, or A27-26 for the Models 22, 22M or 30, respectively, or in accordance with an FAA engineering approved equivalent. (General Dynamics/Convair Alert Service Bulletin A27-56 for the Model 22, A27-25 for the Model 22M, and A27-26 for the Model 30 cover this same subject.) This directive effective October 9, 1962.
61-14-03: 61-14-03 CONVAIR: Amdt. 301 Part 507 Federal Register July 7, 1961. Applies to All Model 22 (880) Aircraft Having Cartridge Assembly-Rudder Spring P/N 22-46223 Installed. Compliance required as indicated. One instance has occurred on the Model 22 (880) aircraft in which the cartridge assembly-rudder spring P/N 22-46223 seized due to corrosion and galling, resulting in limited rudder travel. The effects of this malfunction upon operational safety are such as to require accomplishment of the following: (a) Within the next 100 hours' time in service unless already accomplished in the last 220 hours' time in service, and at each 320 hours' time in service thereafter, conduct the following operational check: Lock the rudder flight tab and rudder in neutral position and operate the rudder pedals right and left until the cockpit stops are contacted 5 or more times. If any roughness in operation is noted, remove cartridge assembly-rudder spring P/N 22-46223, and replace with aP/N 22-46223 which has been inspected and found free of corrosion or galling and operationally checked, or a part reworked in accordance with the provisions of Convair Service Bulletin No. 27-33 and reidentified as P/N 22-46223-1, prior to further flight. (b) When the cartridge assembly-rudder spring is replaced with an assembly reworked in accordance with provisions of Convair Service Bulletin 27-33 and reidentified as P/N 22-46223-1, the inspection specified in (a) may be discontinued. (Convair Service Bulletin No. 27-33 covers this same subject.) This directive effective July 7, 1961.
98-24-04: This amendment adopts a new airworthiness directive (AD) that applies to certain SOCATA - Groupe AEROSPATIALE (SOCATA) Model TBM 700 airplanes. This AD requires repetitively inspecting (using visual methods) the web of the left and right flap carriage for cracks, and replacing any cracked flap carriage with one of improved design. The proposed AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for France. The actions specified by this AD are intended to detect and correct cracks in a flap carriage, which could result in loss of the flap function with consequent reduced and/or loss of airplane control.
83-07-06: 83-07-06 BOEING VERTOL COMPANY AND KAWASAKI HEAVY INDUSTRIES, LTD.: Amendment 39-4600. Applies to Boeing Vertol Model 107-II and Kawasaki Model KV107-II and KV107-IIA helicopters certificated in all categories equipped with No. 5 Synchronizing Shaft, P/N 107D3340-1 or 107D3140-1. Compliance is required as indicated on all No. 5 Synchronizing Shaft Assemblies P/N 107D3340-1 or 107D3140-1 with 1200 hours' or more total time in service. To prevent failure of the steel synchronizing shaft due to improperly drilled rivet holes, accomplish the following: A. Within 50 hours' time in service after the effective date of this AD (unless already accomplished within the past 250 hours' time in service) and every 300 hours' time in service thereafter, visually inspect the interior and exterior surfaces of the No. 5 Synchronizing Shaft Assembly P/N 107D3340-1 or 107D3140-1 for cracks, damage or defects in the area adjacent to the adapter-to-tube rivets at both ends, particularly at the inboard row of rivets. The visual inspection shall be either a lighted borescope inspection using at least 2X magnification or dye penetrant inspection methods combined with at least 2X visual magnification. All paint on surfaces to be inspected must be removed prior to inspection. B. Before the accumulation of 2000 hours' time in service, or before the accumulation of 2000 hours' time in service since last magnetic particle inspection, whichever is less, and at each 2000-hour interval thereafter, magnetic particle inspect the entire shaft assembly, P/N 107D3340-1 or 107D3140-1 in accordance with Boeing Vertol Overhaul Manual 107-5. C. Replace any cracked or otherwise unserviceable part found during the inspections of Paragraphs A or B with serviceable parts prior to further flight. D. An equivalent method of compliance with this AD may be used when approved by the Manager, New York Aircraft Certification Office, New England Region. This amendment becomeseffective April 4, 1983.
2005-25-17: The FAA is adopting a new airworthiness directive (AD) for certain EMBRAER airplanes listed above. This AD requires modifying the drain system of the auxiliary power unit (APU) by installing a scavenge pump and, for certain airplanes, replacing the APU exhaust assembly. This AD results from a report of fuel leaking from the APU feeding line and accumulating inside the APU compartment because the drain system is inadequate when the APU is running. We are issuing this AD to prevent fuel accumulation and subsequent flammable fuel vapors in the APU cowling, which, combined with an ignition source, could result in a fire or explosion.
82-06-51: 82-06-51 SIKORSKY: Amendment 39-4387. Applies to Sikorsky Model S-76A series helicopters certificated in all categories equipped with P/N 76351-09000-061 or prior dash-numbered main gearbox assemblies. To prevent possible failure of a main gearbox shaft and spur pinion P/N 76351-09012-103 due to fatigue cracks propagating in the bottom threaded area, accomplish the following: 1. Within the next 50 hours' time in service, unless already accomplished within the last 50 hours' time in service, and thereafter every 100 hours' time in service from the last inspection, inspect all shaft and spur pinions, P/N 76351-09012-103, with 1,500 or more hours' time in service since new on the effective date of this AD, in accordance with Sikorsky Alert Service Bulletin No. 76-66-16A or later FAA approved revision, or FAA approved equivalent. 2. Within the next 100 hours' time in service and thereafter every 100 hours' time in service from the last inspection, inspect all shaft and spur pinions, P/N 76351-09012-103, with 1,100 or more but less than 1,500 hours' time in service since new on the effective date of this AD, in accordance with Sikorsky Alert Service Bulletin No. 76-66-16A or later FAA-approved revision, or FAA approved equivalent. 3. Inspect shaft and spur pinions with less than 1,100 hours' time in service since new on the effective date of this AD before the accumulation of 1,200 hours' time in service, and thereafter every 100 hours' time in service in accordance with Sikorsky Alert Service Bulletin No. 76-66-16A or later FAA-approved revision, or FAA approved equivalent. 4. If a crack is found, replace the shaft and spur pinion with a serviceable part prior to further flight. 5. Replace the shaft and spur pinions, P/N 76351-09012-103, with a serviceable part prior to the accumulation of 1,500 hours' time in service since new or within 200 hours' time in service after the effective date of this AD, whichever occurs later. Thereafter, replace shaft and spur pinions with serviceable parts prior to accumulation of 1,500 hours' time in service. 6. All shaft and spur pinions whose hours' time in service cannot be established are to be inspected in accordance with paragraph 1 and are to be replaced with a serviceable part within 200 hours' time in service from the effective date of this AD. Equivalent means of compliance may be approved by the Chief, Boston Aircraft Certification Branch, Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment becomes effective June 3, 1982, to all persons except those persons to whom it was made immediately effective by telegraphic AD T82-06-51, issued March 5, 1982, which contained this amendment.
79-10-03 R2: 79-10-03 R2 AVCO LYCOMING: Amendment 39-3462 as amended by Amendment 39-3628 is further amended by Amendment 39-3813. Applies to O-360-E1A6D Series engines, Serial Number L-101-77 thru L-347-77, L-352-77 and LO-360-E1A6D Series engines, Serial Number L-101-72 thru L-319-72, L-321-72 thru L-324-72, L-326-72 thru L-339-72, L-341-72 thru L-348-72, L-350-72 installed in the Piper Model PA-44 aircraft, and O-320-H2AD Series engines, Serial Number L-101-76 thru L-5707-76 or any engine remanufactured prior to January 4, 1979, installed in the Cessna Model 172N and Partenavia Model P-66C. Unless already accomplished, compliance required before further flight for O-360-E1A6D/LO-360-E1A6D engines, and within 50 hours after the effective date of this AD for O-320-H2AD model engines. Aircraft may be flown in accordance with FAR 21.197 to a base where the required inspection can be performed. To prevent loss of integrity of the engine to aircraft mounting due to loosening of the engine mounting bracket attaching bolts, inspect the eight part number LW38-2.75 mounting bolts for correct torque. Mounting bolts found to be 200 inch-pounds or less of torque when measured in the tightening direction must be replaced and torqued to 360 inch-pounds. Those mounting bolts found to be less than 360 inch-pounds but greater than 200 inch-pounds must be retorqued to 360 inch-pounds. Lycoming Service Instruction No. 1380, dated 6-22-79, applies to this subject. Upon submission of substantiating data by an owner or operator through an FAA maintenance inspector, the Chief, Engineering & Manufacturing Branch, FAA Eastern Region, May adjust the compliance time specified in this AD. Amendment 39-3462 was effective May 9, 1979. Amendment 39-3628 was effective December 7, 1979. This Amendment 39-3813 is effective July 1, 1980, and was effective immediately for all recipients of the airmail letter dated December 10, 1979, pertaining to this same matter.
98-24-18: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Bombardier Model DHC-8-102 and -103 series airplanes, that currently requires a one-time inspection to detect disbonding of the upper and lower skin panels of the horizontal stabilizer, and repair, if necessary. This amendment establishes repetitive intervals for the inspection to detect disbonding of the upper and lower skin panels of the horizontal stabilizer. This amendment also revises the applicability of the existing AD to include certain additional airplanes, and to exclude certain other airplanes. 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 reduced strength capability and consequent failure of the horizontal stabilizer, which could result in loss of controllability of the airplane.