Results
59-10-06: 59-10-06 KAMAN: Applies to All K-240 Helicopters (Military HTK-1). Due to associated adverse service experience, the wood type cheek plate rotor blades, P/N K411042-1, -2, -7, and -8 are ineligible for use. These blades modified to conform to, or replaced by, fiberglass type cheek plate rotor blades, P/N K411042-77 and -78, are acceptable with a 2,500-hour retirement life.
75-22-09: 75-22-09 ROCKWELL: Amendment 39-2396. Applies to Model 112, S/N 3 through 380, certificated in all categories. Compliance required as indicated: (a) Before further flight and prior to each flight thereafter until complete inspection and modifications are accomplished, visually inspect the airplane structure at all aileron hinge positions. If visual inspection reveals distortion of the aileron skin or readily visible cracks in hinge doublers, comply with subparagraph (b)(3) below: (b) Within 10 hours' time in service, after the effective date of this AD, accomplish the following: (1) Replace all outboard aileron doublers in accordance with Rockwell International Service Bulletin No. SB-112-35 dated October 1, 1975, or later approved revision or by an equivalent method approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration, Fort Worth, Texas. (2) Inspect all inboard aileron doublers for cracks, proper doubler thickness (.040 inches), and preloading of hinge doublers in accordance with Rockwell International Service Bulletin No. SB-112-35 dated October 1, 1975, or later approved revision, or by an equivalent method approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration, Fort Worth, Texas. (3) If cracks, improper thickness, or preloading are found, replace doublers in accordance with Rockwell International Service Bulletin No. SB-112-35 dated October 1, 1975, or by an equivalent method approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration, Fort Worth, Texas, before further flight. This amendment becomes effective November 4, 1975.
2012-10-06: We are adopting a new airworthiness directive (AD) for all Saab AB, Saab Aerosystems Model SAAB 2000 airplanes. This AD was prompted by reports that environmentally friendly de-icing agents used on certain electrical connectors and braids could cause corrosion damage. This AD requires performing, in certain locations, a detailed inspection for corrosion of the electrical and electronics installation, and if corrosion is found repairing each affected harness braid or replacing each affected component and/or wiring harness. We are issuing this AD to detect and correct corrosion of critical system wiring, which could result in arcing and, in combination with other factors, a fire and consequent damage to the airplane.
75-05-09: 75-05-09 BOEING: Amendment 39-2105. Applies to all model 737 series airplanes, certificated in all categories. Compliance required as indicated. \n\t(1)\tWithin the next 100 hours' time in service after the effective date of this AD, unless already accomplished, inspect the engine fuel shutoff and crossfeed valve wire bundles for chafing in accordance with Boeing Service Bulletin No. 737-28-1022, dated December 20, 1974, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\t(2)\tIf wire or wire insulation chafing or clamp damage is found, replace and modify the wiring and replace the wiring harness clamps in accordance with Boeing Service Bulletin No. 737-28-1022, dated December 20, 1974, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region.\n \t(3)\tWithin the next 1000 hours' time in service after the effective date of this AD, for airplanes with no wire or wire insulation chafing or clamp damage, replace and modify the wiring and replace the wiring harness clamps in accordance with Boeing Service Bulletin No. 737-28- 1022, dated December 20, 1974, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tThe 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). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P. O. Box 3707, Seattle, Washington 98124. The documents may be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective March 24, 1975.
63-25-01: 63-25-01 BEECH: Amdt. 652 Part 507 Federal Register December 5, 1963. Applies to Model 35 Aircraft Serial Numbers D-1 through D-1500, Model 35R Aircraft Serial Numbers D- XXR1 and up (35R Aircraft are Remanufactured Model 35 Aircraft and Retain the Original Serial Number in Addition to the Appropriate 35R Serial Number), and Model Super V Conversions of the Standard Beech Models 35, A35, or B35 Serial Numbers SV-XXX-D-1 through SV-XXX-D-1500. Compliance required within 25 hours' time in service after the effective date of this amendment unless already accomplished within the last 75 hours' time in service and thereafter within 100 hours' time in service from the last inspection. Inspections required by AD 62-02-01 have not been adequate to detect all fatigue cracks in the steel center section front trusses prior to failure. To preclude these failures, modify the fuselage and inspect the front and rear steel trusses in accordance with (a) and (b). In order to gain access to the front and rear trusses, remove the front seat bottom, rear seat, front and rear spar forward partitions, and all floorboards adjacent to the front and rear spars. Also, disconnect the air duct on the right side of the forward spar and remove any other adjacent installations as found necessary for access. (a) Front truss and fuselage. Modify the fuselage in accordance with (a)(1) prior to inspecting. Inspect in accordance with (a)(2) within the next 25 hours' time in service unless the aircraft has been so inspected within the last 75 hours' time in service and thereafter within 100 hours' time in service and thereafter within 100 hours' time in service from the last inspection. (1) Cut two 3 1/2 inch diameter inspection openings in the fuselage skin just under the forward centersection steel truss at right and left butt stations 16.50 inches, in accordance with Beech Service Bulletin 35-24, as revised November 5, 1963, or FAA approved equivalent. (NOTE: In addition to these two openings, any or all of the three inside inspection openings defined in Service Bulletin 35-24 may be incorporated at the owner's option.) (2) Inspect the front truss for cracks, using the magnetic particle inspection procedures outlined in Beech Service Bulletin 35-24 as revised November 5, 1963, or FAA approved equivalent. (3) Cracked trusses shall be replaced or repaired in accordance with Beech Service Bulletin 35-24, as revised November 5, 1963, before further flight. After accomplishment of these repairs the inspection specified in paragraph (2) shall be continued. (b) Rear truss. Inspect in accordance with either (1) or (2). (1) Visual inspection. Within 500 hours' time in service since the last visual inspection, performed in accordance with Beech Service Bulletin 35-24 as revised December 1961, and continually thereafter within 500 hours' time in service from the last inspection, conduct a thorough visual inspection for cracks with adequate lighting, a 3-power magnifying glass, and mirror. (2) Magnetic particle inspection. Within 1,000 hours' time in service since the last magnetic particle inspection performed in accordance with Beech Service Bulletin 35-24 as revised December 1961, and continually thereafter within 1,000 hours' time in service from the last inspection, inspect for cracks using the magnetic particle inspection procedures outlined in Beech Service Bulletin 35-24 as revised November 5, 1963, or FAA approved equivalent. (3) Cracked trusses shall be replaced or repaired in accordance with Beech Service Bulletin 35-24 as revised November 5, 1963, before further flight. If the truss is repaired, the next inspection shall be within 100 hours' time in service after the repairs were accomplished. (Any cracks that may develop because of the localized heating during repair should be detectable by this time.) Following this inspection, subsequent inspections shall be at500 or 1,000 hours' time in service, in accordance with (1) or (2), depending on whether a visual or a magnetic particle inspection was performed. (c) If the front truss is replaced with a new heavier steel truss (Beech P/N 35- 410030-17), the requirements specified in (a) shall become applicable 2,000 hours' time in service after installation of this truss. NOTE: Model 35 airplanes whose steel carry-through trusses have been modified in accordance with an STC that prescribes inspection intervals that differ from those prescribed in this AD, shall be inspected in accordance with the inspection intervals of the STC. (d) If the front truss is replaced with a new aluminum truss installed in accordance with Beech Kit No. 35-694, the requirements specified in (a) are not applicable. (e) An appropriate entry in the airplane log shall be made showing whether the front, rear or both truss(es) were inspected and the type of inspection on the rear truss. This supersedes AD 62-02-01. This directive effective December 10, 1963. Revised April 24, 1964. Revised February 16, 1965. Revised January 19, 1967.
2012-10-04: We are adopting a new airworthiness directive (AD) for certain Cessna Aircraft Company Models 210G, T210G, 210H, T210H, 210J, T210J, 210K, T210K, 210L, T210L, 210M, T210M, 210N, T210N, P210N, 210R, T210R, and P210R airplanes. This AD requires an inspection(s) of the left and right wing lower main spar caps for cracks and either replacing cracked wing lower main spar caps, wing spars, or wings (as applicable) with serviceable spar caps, spars, or wings that are found free of cracks or incorporating an FAA-approved modification. This AD also requires reporting the results of the inspections to the FAA. This AD was prompted by reports of cracks found in the wing lower main spar caps on the above-referenced airplanes with cantilever metal wings. We are issuing this AD to correct the unsafe condition on these products.
54-05-01: 54-05-01 CURTISS-WRIGHT: Applies to all C-46 Series aircraft equipped with Landing Gear Retraction Cylinder Heads, Curtiss-Wright P/N S20-313-3044. \n\n\tCompliance required as soon as practicable but not later than July 15, 1954.\n \n\tAs a result of several failures of the landing gear retraction cylinder head emanating from fatigue cracks occurring in the radius of the O ring groove, it has been found necessary to inspect and rework the landing gear retraction cylinder head to provide a more serviceable part. \n\n\tRemove the landing gear retraction cylinder head and inspect the upper and lower radii of the O ring groove for cracks by means of a dye penetrant. If cracks are found the part should be replaced. In the event no cracks are present the upper and lower radii of the O ring groove should be increased to 0.040 inch as shown in Figure 4 and again inspected by means of a dye penetrant. If no cracks are found after this inspection, the part may be returned to service; however,if cracks are evident, the part should be replaced. \n\n\n\n\tNote: Smooth machine finish (RMS-32 or better) with no tool marks or discontinuities. No change in size of O ring necessary.
2012-09-13: We are adopting a new airworthiness directive (AD) for certain Airbus Model A330-200 freighter series airplanes; Model A330-200 and - 300 series airplanes; and Model A340-200 and -300 series airplanes. This AD was prompted by a report of corrosion found on the main fitting of the nose landing gear (NLG) leg in the vicinity of the dowel pin bushes retaining the lower steering flange. This AD requires modifying the NLG main fitting by adding primer paint to the cadmium around the dowel bush holes. We are issuing this AD to prevent NLG main fitting rupture, which could result in an NLG collapse.
2012-10-02: We are adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation Models 58 and G58 airplanes. This AD was prompted by installation of oversized clamps on fuel vapor return and/ or fuel vent lines in the outboard sections of the left and right wings. This AD requires inspecting for oversized or deformed fuel hose clamps and replacing as necessary. We are issuing this AD to correct the unsafe condition on these products.
2012-09-08: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 767-200 and -300 series airplanes. This AD was \n\n((Page 28241)) \n\nprompted by reports of multiple site damage cracks in the radial web lap and tear strap splices of the aft pressure bulkhead at station (STA) 1582 due to fatigue. This AD requires repetitive inspections for cracking of the aft pressure bulkhead at STA 1582, repair or replacement of any cracked bulkhead, and eventual replacement of the aft pressure bulkhead at STA 1582 with a new bulkhead. Accomplishing the replacement terminates the repetitive inspections required by this AD. We are issuing this AD to prevent fatigue cracking of the aft pressure bulkhead, which could result in rapid decompression of the airplane and possible damage or interference with the airplane control systems that penetrate the bulkhead, and consequent loss of controllability of the airplane.