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76-10-10: 76-10-10 BEECH: Amendment 39-2617. Applies to Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29), 65-90, 65-A90, B90 and C90 (Serial Numbers LJ-1 thru LJ-676), E90 (Serial Numbers LW-1 thru LW-163), 100 and A100 (Serial Numbers B-1 thru B- 225) and 200 (Serial Numbers BB-2 thru BB-111) airplanes. Compliance: Required as indicated, unless already accomplished. To preclude opening of the cabin door in flight, accomplish the following: I. Within 50 hours' time in service after the effective date of this AD, and within each 50 hours' time in service thereafter, check the cabin door for proper operation and rigging in accordance with the applicable Beech Shop/Maintenance Manual. When all the items specified in Paragraph II have been complied with the requirements of this paragraph (I) are no longer applicable. II. Within 50 hours' time in service after September 15, 1976, perform the following: A. On Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29) and 65-90, 65-A90, B90 (Serial Numbers LJ-1 thru LJ-351) airplanes, modify the cabin door installation in accordance with Beechcraft Service Instructions 0043-104 or 0016-105, Rev. I, or later approved revisions, as applicable. B. On Models 65-88 (Serial Numbers LP-1 thru LP-47 except LP-27 and LP-29), 65-90, 65-A90, B90 and C90 (Serial Numbers LJ-1 thru LJ-676), E90 (Serial Numbers LW-1 thru LW-163), 100 and A100 (Serial Numbers B-1 thru B-225), and 200 (Serial Numbers BB-2 thru BB-111) airplanes, perform the following in accordance with Beechcraft Service Instruction 0818-016 or later approved revisions: 1. Install Beech P/N 101-430124-1 and if fixed step door Beech P/N 101-430124-3 or -5 decals on the existing cabin door instruction plate and operate the cabin door accordingly. 2. Check cabin door latching mechanism and warning system for proper operation and rigging and rerig, if required, as instructed in the appropriate Beech Shop/Maintenance Manual. C. On the airplane models and serial numbers listed below add the indicated part number FAA-approved Airplane Flight Manual Supplement/Revision to the existing airplane pilot's operating manual or FAA-approved airplane flight manual: MODELS Beech Part Number (P/N) of FAA-Approved Airplane Flight Manual Supplement Revision dated November 14, 1975 or Subsequent 1) 65-88, 65-90, 65-A90, B90 and C90 (S/N LJ-502 thru LJ-624) and 100 (S/N B-2 thru B-89 and B-93) 1) P/N 131344 2) C90 (S/N LJ-625 thru LJ-676) 2) P/N 90-590010-53A6 3) E90 (S/N LW-1 thru LW-163) 3) P/N 90-590012-3A6 4) A100 (S/N B-1, B90 thru B-92, B-94 thru B-225) 4) P/N 100-590032-1A6 5) 200 (S/N BB-2 thru BB-111) 5) P/N 101-590010-3A4 III. 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 May 28, 1976.
81-02-10: 81-02-10 BELL: Amendment 39-4025. Applies to Models 214B and 214B-1 helicopters, serial numbers up to and including S/N 28049. Compliance is required as indicated unless already accomplished. To prevent a clutch failure which will result in the loss of engine power to the main rotor, accomplish the following: (a) The freewheeling clutch assembly, P/N 214-040-021-001, must be removed from service and P/N 214-040-021-103 clutch assembly installed according to the following schedule; (1) P/N 214-040-021-001 clutch assemblies with 290 or more hours' time in service on the effective date of this AD must be removed from service within the next 10 hours' time in service. (2) P/N 214-040-021-001 clutch assemblies with less than 290 hours' time in service on the effective date of this AD must be removed from service prior to attaining 300 hours' time in service. (3) P/N 214-040-021-001 clutch assemblies with unknown time in service must be removed within the next ten hours' time in service. NOTE: BHT Alert Service Bulletin No. 214-80-13, dated August 22, 1980, pertains to this subject. (b) Special flight permits may be issued in accordance with FAR 21.197 and FAR 21.199 to fly aircraft to a base where this AD can be accomplished. (c) Any alternate equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration. This AD supersedes AD 80-07-11 (Amdt. 39-3726, 45 FR 20778). 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 Bell Helicopter Textron, Product Support Department, Post Office Box 482, Fort Worth, Texas 73101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at the FAA 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 FAA at their headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. This amendment becomes effective January 26, 1981.
89-14-03: 89-14-03 LOCKHEED AERONAUTICAL SYSTEMS COMPANY: Amendment 39-6243. Applicability: Lockheed Model L-188A and L-188C series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent airplane control difficulties due to flap asymmetry, accomplish the following: A. Within 90 days after the effective date of this AD, inspect the flap universal joints to determine if they are splined without a stop. NOTE: Flap universal joints that are splined without a stop are easily detected by the lack of a screw and washer that holds the pins in place. If they do not have a stop, accomplish the following: 1. Remove from service any flap universal joints that are splined without a stop and install universal joints that are splined with a stop; or 2. Cut a circumferential groove in the splines of the torque tube shafts on each side of the universal joints (1 inch from the end of the shaft) and installa snap ring, in a manner approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region. The snap ring is to eliminate the possibility of spline disengagement. The snap ring should completely encircle the shaft; or 3. Modify each universal joint by drilling a hole and installing a 1/8-inch diameter steel pin through each side (two per joint). Locate the pin from .7 to .8-inch from each end, and locate so as to intersect the spline centerline. Peen both pin ends. This modification must be accomplished in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region. NOTE: Local spotfacing may be necessary on pins located at maximum dimension. B. Within 90 days after the effective date of this AD and thereafter at intervals not to exceed one year, verify proper torque tube-to-universal joint engagement of .375-inch, as specified in Detail E of the L-188 Maintenance Manual, Section 27-6-8, Figure 201; and verify that the distance between the edges of the torque tube shoulders at BL55 is 5.5 to 5.87 inches (universal joints with internal stop), or 6.0 to 6.6 inches for universal joints reworked with added stop pins. C. Within 90 days after the effective date of this AD and thereafter at intervals not to exceed one year, inspect the wing flap asymmetry shutoff valves, P/N 668225-1, to ascertain whether the valves hang up or respond slowly (greater than 1 second). 1. If the valves hang up or respond slowly, prior to further flight, install a functioning serviceable valve of the same part number (P/N 668225-1) and repeat the inspections at intervals not to exceed one year. 2. Installation of modified valve, P/N 668225-101, or a new valve P/N 668225-101, in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region, constitutes terminating action for the repetitive inspections of the wingflap asymmetry shutoff valves required by paragraph C. and C.1., above. D. Within 90 days after the effective date of this AD, rewire the flap asymmetry annunciator light in the cockpit to trigger the annunciator, in accordance with a method approved by the Manager, Los Angeles Aircraft Certification Office, Northwest Mountain Region, so that the light in the cockpit will illuminate when the asymmetry detector is tripped and not be dependent on the shutoff valve. E. 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, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Los Angeles Aircraft Certification Office. F. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a base to accomplish the actions required by this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Lockheed Aeronautical Systems Company, P.O. Box 551, Burbank, California 91520, Attn: L-188 Commercial Support Contracts, Dept. 63-11, Unit 33. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or 3229 East Spring Street, Long Beach, California. This amendment (39-6243, AD 89-14-03) becomes effective on July 24, 1989.
2017-19-22: We are superseding Airworthiness Directive (AD) 2014-07-09 for British Aerospace Regional Aircraft Jetstream Series 3101 and Jetstream Model 3201 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as both the need for newly added inspections for corrosion, which includes the door hinges/supporting structure and attachment bolts for the main spar joint and engine support, and inadequate existing instructions for inspection for corrosion for several areas including the rudder hinge location on the vertical stabilizer. We are issuing this AD to require actions to address the unsafe condition on these products.
2002-01-19: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F.28 Mark 0070 and 0100 series airplanes, that requires repetitive operational tests for discrepancies of the heating system of pitot tube #1, and replacement of the pitot tube, if necessary. This AD also requires eventual modification of the alternating current sensing circuit for pitot tube #1, which terminates the repetitive operational test requirement. This action is necessary to prevent failure of the heating system of pitot tube #1 due to a short circuit, which may go undetected and lead to the pilot receiving erroneous airspeed indications, resulting in reduced control of the airplane. This action is intended to address the identified unsafe condition.
2017-19-06: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-1A11 (CL-600), CL-600-2A12 (CL-601 Variant), and CL-600-2B16 (CL-601-3A, CL-601-3R, and CL-604 Variants) airplanes. This AD was prompted by a new life limitation that has been introduced for the side brace fitting shaft and side brace-to-airplane fitting pin of the main landing gear (MLG). This AD requires revising the maintenance or inspection program. This AD also requires an inspection to identify the serial number, to serialize, and to record the accumulated life of the side brace fitting shaft of the MLG. We are issuing this AD to address the unsafe condition on these products.
2011-01-04: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: It has been detected a short circuit in harness W101 due to its interference with the main door mechanism. Further analysis of the affected region has also revealed the possibility of chafing between the same harness and the oxygen tubing. The chafing of the wiring harness against the oxygen tubing could lead to a short circuit of the wiring harness and a subsequent fire in the airplane. Since this condition may occur in other airplanes of the same type and affects flight safety, a corrective action is required. Thus, sufficient reason exists to request compliance with this AD in the indicated time limit. We are issuing this AD to require actions tocorrect the unsafe condition on these products.
49-45-01: 49-45-01 LUSCOMBE: Applies to All Model 11A Aircraft. Compliance required as soon as possible but not later than next 25 hours operation time and at each 25-hour period thereafter until reinforcement of main landing gear aft canted fuselage bulkhead is accomplished. Inspect for buckling, cracks or other evidence of failure of permanent set of the main landing gear aft fuselage canted bulkhead in the web and/or flange in the area adjacent to the steel landing gear trunnion and fuel line. Inspect fuselage wing lift strut attach fitting for cracks in the radii of the flanges attaching it to each aft fuselage canted bulkhead. Usually evidence of failure of the aft canted bulkhead can be determined by a crack in the fuselage canted bulkhead web extending from the fuel line hole to the flange attaching the bulkhead to the belly skin and/or buckle in the cabin floor located approximately 1 inch directly aft of the bulkhead under the carpet flooring and/or loose rivets attaching theflange of the canted bulkhead to the belly skin. If the difficulties are not revealed as indicated, a 2-inch hole cut in the cabin floor located approximately 3 inches aft and inboard of that part of the canted bulkhead supporting the door will allow access for detailed examination of the aft side of the rear fuselage canted bulkhead. Removal of seat and floor carpet is necessary to accomplish this inspection. If loose rivets in the bulkhead flange at the attachment to the belly skin, cracks or permanent set in excess of 1/8 inch are found in the web of the bulkhead adjacent to the steel trunnion, the bulkhead must either be satisfactorily repaired or replaced. If noticeable permanent set in the web is apparent (under 1/8 inch), the web of the bulkhead may be reworked by straightening. If cracks are found in the fuselage wing lift strut attach fitting it should be replaced or the cracks should be stop drilled and the full length of each cracked flange reinforced with a 3/4 inchby 3/4 inch by 0.064 inch 24ST angle. In addition, the following modifications must be made: A collar must be incorporated on the front end of the hinge pin that passes through the front and rear main landing gear steel trunnions which are riveted to the two fuselage canted bulkheads. This tubular collar should be fabricated of 4230 steel and be at least 5/8-inch long and of sufficient thickness to effect a snug bearing fit against the forward end of the steel tube composing the socket of the forward steel trunnion. A 1/4-inch bolt should be used to attach the collar to the hinge pin using the existing 1/4-inch hole in the extreme forward end of the hinge pin. A curved doubler of 0.064 inch 24ST should be placed over the existing 0.040-inch floor skin connecting the flanges of the two main landing gear canted bulkheads. This doubler should pick up the existing floor skin and bulkhead top flange rivet pattern in the vicinity of the landing gear steel trunnion, extending in length at least 3 inches to either side of a vertical plane through the centerline of the landing gear hinge pin and picking up at least six of the existing rivets in each of the canted bulkheads. Blind type rivets may be used to attach this doubler. The rivet pattern attaching the flange of the aft canted fuselage bulkhead to the belly skin between the openings in the fuselage skin which allow entrance of the main landing gear legs should be inspected for rivet size and pattern. The first 20 rivets inboard from these openings must be 5/32-inch A17ST spaced approximately 1/2-inch apart. If the 2-inch inspection holes have been cut in the floor, they must be reinforced by at least a 4-inch diameter 0.040-inch 24ST doubler on the underneath side of the floor skin and a quick removable inspection cover placed on top side to be used for subsequent 25-hour inspections, if applicable. Any equivalent structural modification to preclude a failure, or permanent set in the aft canted bulkhead at the attachment of the main landing gear trunnion will be considered satisfactory.
2017-19-10: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 757-200, -200PF, and -200CB series airplanes. This AD was prompted by an analysis of the cam support assemblies of the main cargo door (MCD) that indicated that the existing maintenance program for the cam support assemblies is not adequate to reliably detect cracks before two adjacent cam support assemblies could fail. This AD requires an inspection to determine part numbers, repetitive inspections to detect cracking of affected cam support assemblies of the MCD, and replacement if necessary. We are issuing this AD to address the unsafe condition on these products.
2002-01-25: This amendment adopts a new airworthiness directive (AD), applicable to certain Bombardier Model DHC-8-100, -200, and -300 series airplanes, that requires repetitive inspections of the rudder pedal adjustment fittings for cracks and replacement of cracked fittings with new fittings. This amendment also provides an optional terminating action. This action is necessary to detect and correct cracking of the rudder pedal adjustment fittings, which could lead to deformation of the fittings, resulting in jammed rudder pedals and loss of rudder control, with consequent reduced controllability of the airplane. This action is intended to address the identified unsafe condition.