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79-08-07: 79-08-07 CESSNA: Amendment 39-3451. Applies to Model 441 (S/N's 441-0001 through 441-0083 and 441-0085) airplanes equipped with a Cessna installed propeller anti-icing system. COMPLIANCE: Required as indicated unless already accomplished. To preclude the possible occurrence of smoke/fire behind the pilot's instrument panel, accomplish the following: A) Within the next 10 hours time-in-service, except for those airplanes previously modified, modify the propeller anti-icing electrical wiring in accordance with Cessna Propjet Service Information Letter PJ 79-2 dated March 5, 1979. B) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective on April 30, 1979 to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated March 16, 1979.
58-06-02: 58-06-02 HARTZELL: Applies to All HC-82XF/8833-0 (86 to 88 inch-diameter) Propellers, Hub Serial Numbers Between T-913 and T-2891 Except T-2564, T-2569, T-2594, T- 2595, T-2609, T-2648, T-2703 and T-2716. Compliance required prior to next flight. Two recent failures of A-159 split rings have occurred in HC-82XF/8833 propellers. In order to minimize the possibility of the occurrence of this type of serious failure, replace the present split rings with new split rings. The present rings are unmarked, but the new A-159 rings will be marked with the letter "N". (Hartzell Service Bulletin No. 57 covers this same subject.)
63-20-02: 63-20-02 CESSNA: Amdt. 628 Part 507 Federal Register October 3, 1963. Applies to All Models 190 and 195 Series Aircraft. Compliance required as indicated. As a result of a fatal accident involving fatigue failure of the front wing spar fuselage carry through lower cap the following corrective action is required: (a) Within 10 hours' time in service after the effective date of this AD, conduct the following visual and X-ray inspections in accordance with Cessna Service Letter 190/195-1 dated May 13, 1960, or FAA approved equivalent unless already accomplished within the last 100 hours' time in service: (1) Inspect using X-ray method for cracks in the lower cap of the front spar carry through member at both left and right sides of the fuselage in the area of the two most inboard steel rivets; (2) Remove rear spar to fuselage attachment bolts and visually inspect these bolts using at least a 4-power magnifying glass for evidence of wear or partial shear failures; (3) Visually inspect the rear spar to fuselage fittings of both the wing structure and carry through structure including the bolt bushings for wear or hole elongation; (4) Visually inspect the steel plates of the rear spar to fuselage attachment fittings on the wing for cracks using at least a 4-power magnifying glass; and (5) If any of the defects specified are found replace the parts with new parts or make an FAA approved repair and accomplish the modification required by (b) before further flight. (b) Within 10 hours' time in service after the effective date of this AD, accomplish the fuselage front spar modification specified in Cessna Service Letter 63-40, dated September 27, 1963. This specifies replacement of the five existing inboard lower cap high shear steel pins with NAS 464P7-A64 bolts and attachment of two 1/2 x l 3/4 4130 steel reinforcement bars on each side of the lower spar cap in accordance with Cessna Service Kits Nos. SK-195-3 and SK-195-4, dated September 23, 1963. (c) At each periodic inspection after the accomplishment of (b), visually inspect for cracks the visible portion of the lower spar cap and reinforcement bars at the inboard bolt location. If cracks are found, replace the cracked part with a new part or make an FAA approved repair before further flight. This supersedes AD 61-17-01. This directive effective October 8, 1963.
62-01-02: 62-01-02 LOCKHEED: Amdt. 384 Part 507 Federal Register January 9, 1962. Applies to All Models 49, 149, 649, 749, and 1049 Series Aircraft Incorporating Cleveland Pneumatic Tool Company Main Landing Gear Cylinders, P/N's 9040-2, 9040A-2, 9106-2, 9291-2, 9291B-2, 9291B-2B, or 9291D-2. Compliance required as indicated. As a result of failures of the cap weld on main landing gear cylinders which have been returned to service following compliance with the provisions of AD 59-07-03, as amended, all aircraft shall be inspected and/or modified as follows: (a) Within the next 400 hours' time in service after the effective date of this AD, unless already accomplished, all landing gear cylinders with cap welds which previously have been inspected by the double-wall gamma radiographic technique in compliance with AD 59-07-03 shall be reinspected in accordance with (c), or shall be replaced with cylinders which have been inspected in accordance with (c) and found to exhibit no evidence of cracks. (b) Within the next 1,200 hours' time in service after the effective date of this AD, unless already accomplished, all landing gear cylinders with cap welds which previously have been radiographically inspected by a method other than the double-wall gamma technique in compliance with AD 59-07-03, shall be reinspected in accordance with (c), or shall be replaced with cylinders which have been inspected in accordance with (c) and found to exhibit no evidence of cracks. (c) The inspection techniques and procedures to be applied to all cylinders noted in (a) and (b) shall be those prescribed by Section B of Lockheed Field Service Letter No. FS/254049L, dated August 29, 1961, or FAA engineering approved equivalent. The entire periphery of the cap weld shall be inspected. (d) Any cylinder with a cap weld which is found to exhibit any evidence of cracks when inspected as required by (a) or (b), shall be replaced prior to further flight of the aircraft. Cracked cylinders are not eligible for any further use in certificated aircraft. (e) The cap welds on all cylinders which have been installed or which have been inspected, found to exhibit no evidence of cracks, and returned to service in accordance with (a) or (b), shall be reinspected in accordance with (c), at periods thereafter not to exceed 5,000 hours' time in service. The first periodic reinspection of any cylinder in service which, prior to the effective date of this AD, was inspected in the manner prescribed by (c) shall be accomplished within 5,000 hours' time in service following the date of that inspection. (f) Cylinders with the cap welds which exhibit any evidence of cracks during any of these periodic reinspections shall be replaced prior to further flight of the aircraft. Cracked cylinders are not eligible for any further use in certificated aircraft. (g) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. (h) When the cylinders described are replaced with new cylinders, TPC Co. P/N's 9040A-2A, 9106-2A, and 9291D-2A, incorporating flash welded caps, the inspections specified in (a) through (g), may be discontinued. (Lockheed Field Service Letter FS/254049L covers this same subject.) This supersedes AD 59-07-03. This directive effective January 9, 1962.
82-05-02: 82-05-02 CANADAIR: Amendment 39-4327. Applies to Canadair Model CL-600-1A11 airplanes, serial numbers 1005 thru 1007, 1010 thru 1014, and 1016 thru 1022. To prevent loss of electrical power, unless already accomplished, accomplish the following: 1. Within the next 50 flight hours, or within the next 15 calendar days after the effective date of this Airworthiness Directive, whichever comes first, inspect, and repair if necessary, in accordance with accomplishment instructions contained in Canadair Alert Service Bulletin No. A600-0087, Revision 1, dated September 3, 1981, or a later revision approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Mountain Region. 2. Airplanes may be flown in accordance with FAR 21.197 to a maintenance base for accomplishment of the inspection required by this AD. 3. Alternate methods of compliance with this AD may be used when they provide an equivalent level of safety and are approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Mountain Region. 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 the addresses listed above. These documents may also be examined at FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective March 11, 1982.
2016-23-09: We are adopting a new airworthiness directive (AD) for various restricted category helicopters. This AD requires cleaning and visually inspecting certain main rotor (M/R) blades and, depending on the outcome of the inspections, repairing or replacing the M/R blades. This AD was prompted by a report of an M/R blade with multiple fatigue cracks around the blade retention bolt hole. The actions are intended to detect a crack in the M/R blade, and prevent failure of the M/R blade and subsequent loss of helicopter control.
62-12-06: 62-12-06 VICKERS: Amdt. 445 Part 507 Federal Register May 29, 1962. Applies to All Viscount Models 745D and 810 Series Aircraft. Compliance required as indicated. Failure of a main landing gear uplock mechanism resulted in inability to extend the gear. Inspection revealed that a cotter pin, which secures the connecting links to the center spindle, was missing. This permitted the clevis pin to drop into an inspection hole, and prevented the mechanism from moving to the unlock position. To prevent further cases of this malfunction, the following are required: (a) Within the next 150 hours' time in service after the effective date of this AD, unless already accomplished within the past 100 hours' time in service and at periods thereafter not exceeding 250 hours' time in service from the last inspection, visually inspect the main landing gear uplock mechanisms to ensure that the clevis pin is in the correct position and secured with a cotter pin that is not worn or damaged. Prior to further flight, properly reinstall clevis pins that have worked loose or shifted and replace worn or damaged cotter pins. (b) The inspections required by this airworthiness directive may be discontinued after replacement of the perspex covers with metal plates installed in accordance with Vickers Preliminary Technical Leaflet No. 236 Issue 3, Modification D.3049 (for 700 Series aircraft), or PTL No. 102 Issue 3 Mod. FG.1873 (for the 800 Series aircraft), or an FAA approved equivalent modification. (c) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, International Division, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. This directive effective June 29, 1962.
2016-23-03: We are adopting a new airworthiness directive (AD) for all Diamond Aircraft Industries GmbH Model DA 40 NG airplanes. 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 possible loss of engine power and emergency landing with consequent damage to the airplane and occupant injury caused by a manufacturing quality deficiency in a batch of V-clamps that could cause the V-clamp to crack and fail. We are issuing this AD to require actions to address the unsafe condition on these products.
64-10-02: 64-10-02\tBOEING: Amdt. 726 Part 507 Federal Register May 6, 1964. Applies to Models 707 and 720 Series Aircraft Serial Numbers 17586 through 17612, 17614 through 17652, 17658 through 17690, 17692 through 17724, 17903 through 17906, 17918 through 17930, 18012 through 18037, 18054 through 18063, 18067 through 18071, 18083 through 18087, 18165 through 18167, 18245 through 18251, 18334 through 18339, 18351 through 18357, 18372 through 18375, 18378, 18381 through 18395, 18411, 18414 through 18422, 18424, 18425, 18451 through 18452.\n \n\tCompliance required as indicated unless already accomplished. \n\n\tTo provide improved airplane controllability in case of horizontal stabilizer electrical trim malfunction, and to provide a switch with sufficient operating torque to assure switch operation necessary for horizontal stabilizer electrical trim operation in both directions, accomplish the following: \n\n\t(a)\tWithin 3,000 hours' time in service after the effective date of this AD on all theabove listed airplanes, except Serial Numbers 18393 through 18395, 18451, and 18452, remove the existing stabilizer limit switches (Boeing P/N 66-11056 or Control Company of America P/N 1HS6) and associated mounting brackets and install new limit switches Control Company of America P/N H10-1001 and associated mounting brackets and cams. This installation shall be accomplished in accordance with Boeing Service Bulletin No. 1635(R-1), Paragraph 3, Part I, or an FAA approved equivalent. \n\n\t(b)\tWithin 1,000 hours' time in service after the effective date of this AD on airplane Serial Numbers 18393 through 18395, 18451 and 18452, replace the four horizontal stabilizer electrical trim limit switches Control Company of America P/N H10-59 with Control Company of America P/N H10-1001 switches in accordance with Boeing Service Bulletin No. 1635 (R-1), Paragraph 3, Part II, or an FAA approved equivalent. \n\n\t(Boeing Service Bulletin No. 1635(R-1) covers this same subject.) \n\n\tThis directiveeffective June 5, 1964.
61-04-04: 61-04-04 LOCKHEED: Amdt. 251 Part 507 Federal Register February 10, 1961. Applies to All Models PV-1 and B-34 Aircraft With Main Landing Gear Drag Strut Attach Plates, P/N 112900 or P/N 12665. Compliance required as indicated. (It will be necessary for operators to maintain a record of landings in order to ascertain compliance with this AD.) Investigation has shown that failure of the main landing gear drag strut attach plate, P/N 112900 has caused the main landing gear to collapse. The B-34 main landing gear drag strut attach plate, P/N 12665, is the same as the PV-1. As a result of this service experience the following shall be accomplished: (a) Inspection in accordance with either (1) or (2) is required: (1) Within the next 25 landings and every 50 landings thereafter, inspect the main landing gear drag strut attach plates, P/N 112900, or P/N 12665, in place on the aircraft for cracks and looseness.* Dye penetrant or FAA approved equivalent must be used for crackdetection. If the drag strut attach plate is loose, then it must be visually inspected for elongation of the bolt holes and the attaching bolts visually inspected for cracks, or deformation. If cracks or elongation of the bolt holes are found in the drag strut attach plate, it must be replaced prior to the next flight. Deformed or cracked bolts must be replaced prior to the next flight. (2) Within the next 25 landings, unless already accomplished within the last 125 landings and every 150 landings thereafter, remove the main landing gear drag strut attach plates, P/N 112900, or P/N 12665, and conduct an X-ray or dye penetrant, or FAA approved equivalent, inspection for cracks. The bolt holes in the drag strut attach plate must be visually inspected for elongation and the attaching bolts inspected for cracks or deformation. If cracks or elongation of the bolt holes are found in the drag strut attach plate, it must be replaced prior to the next flight. Deformed or cracked bolts mustbe replaced prior to the next flight. (b) The special inspections contained in this AD may be discontinued after the Howard Aero, Inc., replacement main landing gear drag strut attach plates, P/N 5-302006, or FAA approved equivalents, are installed using 180,000 p.s.i. minimum ultimate tensile strength attaching bolts. This directive effective March 14, 1961. *A method of inspection for looseness is to check for gaping of 0.010 inch or greater, between the drag strut plate and the backing plate at a point approximately 4.5 inches above the lower edge of the attach plate. Insert a feeler gage between the attach plate and the backing plate from the outboard side to make this check.