Results
59-06-07: 59-06-07 VICKERS: Applies to All Viscount Model 745D Aircraft. Compliance required by April 1, 1960. In order to preclude the possibility of foreign object jamming primary flight controls in the cockpit floor, cover plates must be installed over all openings, in accordance with Vickers- Armstrongs Modification Bulletin NR.D.2146 Parts (B), (C), (D), (E), (F), and (P), or equivalent. Modifications D.1185 and D.2272 also cover sealing openings in the cockpit floor area.
54-19-02: 54-19-02 SIKORSKY: Applies to All Model S-55 Helicopters. Compliance required before October 1, 1954. To insure against the loosening of the servo pilot valve locknut and subsequent improper servo operation, safety clips, Sikorsky P/N S-14-40-5194, should be installed and safetied to P/N S14-40-3227-24 lockwasher. (Sikorsky Service Information Circular 1440-458 covers this same subject.)
69-05-02: 69-05-02 FAIRCHILD: Amendment 39-729 as amended by Amendment 39-1588 is further amended by Amendment 39-2641. Applies to FH-227 Type Airplanes Certificated in all categories. Compliance required as indicated. To detect the development of cracks in the wing area, accomplish the following: (a) Within 25 hours time in service after the accumulation of the specified hours in service, unless already accomplished, inspect or continue to inspect in accordance with Fairchild Service Bulletin 51-1, as amended by Revision 6, of December 12, 1975 or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region or with an approved equivalent inspection. (b) Where a visual inspection may be accomplished in lieu of x-ray, at least a 10-power glass must be used. (c) For those aircraft incorporating Fairchild Service Bulletin 51-1, Appendix No. 1, dated January 5, 1973, or an approved equivalent special structural inspection and alteration, the inspection interval for the outer wing panel without cracks remains at 1200 hours. If any cracks were discovered prior to the alteration the inspection interval will be that specified in paragraph 1.D(1) of the Appendix. (d) If new cracks are found or if repaired cracks are found to be propagating, replace the cracked part with a part of the same part number or with an approved equivalent part, or incorporate an approved repair before further flight. However, upon request, with descriptive information of the crack and proposed operating limitations submitted through an FAA maintenance inspector, the flight of the airplane in accordance with FAR 21.197 to a base where the repair can be made, may be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. (e) Equivalent inspection, repairs or parts must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. (f) Upon request, with substantiating data submittedthrough an FAA maintenance inspector, the compliance times specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. Amendment 39-729 was effective March 12, 1969. Amendment 39-1588 was effective January 31, 1973. This amendment 39-2641 is effective June 18, 1976. NOTE: Amendment 39-2641 was distributed on June 7, 1976, as AD 76-12-04. The issuing office intended Amendment 39-2641 to be published as a revision to AD 69-5-2. Accordingly, the AD is being republished to reflect the correct AD number and AD 76-12-04 is cancelled.
68-05-02: 68-05-02 BRANTLY: Amdt. 39-572. Applies to Models B-2, B-2A and B-2B series helicopters with serial numbers 4 through 482 inclusive and Brantly Model 305 helicopters with serial numbers 1002 through 1045 inclusive. Compliance required as indicated. To prevent failure of the tail rotor drive shaft, accomplish the following: (a) Within 10 hours' time in service after receipt of this Airworthiness Directive, unless already accomplished, inspect the tail rotor drive shaft for corrosion using a borescope and treat it with a corrosion preventive compound in accordance with Brantly Service Bulletin, SB No. B2B-68-2, dated February 24, 1968, or FAA-approved equivalent. After treatment, and before the tail rotor drive shaft plug is installed, reinspect the tail rotor drift shaft, using a borescope, to assure that the entire internal surface of the shaft is completely covered with the corrosion preventive compound. (b) If corrosion is found as a result of the inspection required by paragraph (a) of this Airworthiness Directive which cannot be removed or leaves pits visible without magnification, disregard the treatment required by paragraph (a) of this Airworthiness Directive and prior to further flight replace the tail rotor drive shaft with a replacement part which complies with Brantly Service Bulletin, SB No. B2B-68-2, dated February 24, 1968, or FAA-approved equivalent. This amendment becomes effective March 30, 1968
57-03-03: 57-03-03 DOUGLAS AND LOCKHEED: Applies to Lockheed Constellation Series and Douglas DC-7 Series Airplanes. \n\n\tCompliance required by April 1, 1957, for Constellation aircraft and by May 1, 1957, for DC-7 aircraft. \n\n\tUnder certain cold weather operating conditions the perforated paper covering around the Purolator micronic filter elements are subject to accumulation of ice as a result of entrained water crystals in the fuel freezing on this covering. This interferes with proper fuel filtering by causing the fuel to pass through the bypass valve in the filter unit. To make the micronic filter less susceptible to clogging by ice, the perforated paper covering around the filter element is to be removed. Removal of the perforated cover does not affect the filtering characteristics of the filter element. Filter elements without the paper covering are identified as Purolator P/N 30868-3.
67-23-04: 67-23-04 FAIRCHILD-HILLER: Amdt. 39-458, Part 39, Federal Register August 5, 1967. Applies to Model F-27J Airplanes, Serial Numbers 111, and 113 through 121, inclusive and FH-227 Series Airplanes, Serial Numbers 501 through 520, inclusive. Incorporating Solar Auxiliary Power Unit, Model T62T25. Compliance required within the next 400 hours' time in service after the effective date of this AD, unless already accomplished. To prevent electrical damage to the A.C. engine-driven generator resulting from an unregulated and unloaded condition when the APU-driven generator (A.C.) is switched on the line, accomplish the following: (a) Rewire the engine-driven A.C. generator control sensing circuit in accordance with Fairchild-Hiller F-27 Service Bulletin 30-12, Revision No. 1, dated September 15, 1966, for F-27J aircraft and Fairchild-Hiller FH-227 Service Bulletin 30-1 dated August 23, 1966, for FH-227 aircraft, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, or perform an equivalent rewiring modification, approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (b) Upon request, with substantiating data submitted through an FAA maintenance inspector, an increase in the compliance time may be approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This amendment effective August 31, 1967.
55-15-04: 55-15-04 DOUGLAS: Applies to all DC-6 Series Airplanes Below fuselage No. 487 Equipped With Hamilton Standard Propellers. \n\n\tCompliance required by first scheduled engine change after November 1, 1955, but not later than November 1, 1956. \n\n\tTo increase the integrity of the propeller feathering system in the event of a powerplant fire, all existing propeller feathering lines located forward of the firewall must be replaced with lines and flexible hose assemblies which will meet current fireproof and fire resistant requirements. \n\n\t(a)\tReplace the existing flexible hose portion of the feathering line between the union on the forward pipe assembly and the elbow on the inner ring with a new hose assembly Aeroquip P/N 304004-10-17 1/2. Douglas General Service Letter DC-6 No. 206 dated August 26, 1954, covers this subject. Resistoflex SSFR-3800-10 hose assembly and Aeroquip 309009-10S hose assembly are also considered acceptable for this application. \n\n\t(b)\tThe existing flexible hose assembly connecting to the governor is not affected by this directive. \n\n\t(c)\tRemove the existing short 304 sleeves or flexible metal sleeve from the feathering pump supply line, Aeroquip P/N 304002-16D-12 3/8, and install a fireproof cover, Douglas P/N 3500614-1. Douglas General Service Letter DC-6 No. 206 dated August 26, 1954, covers this subject. Aeroquip 601000 hose assembly equipped with Aeroquip 304 full length protective sleeve or Aeroquip 680-16S hose assembly equipped with Aeroquip 304 short sleeves covering the end fittings are also considered acceptable for this application.
55-23-03: 55-23-03 LOCKHEED: Applies to Models 49, 149, 649, 749 and 1049 Aircraft, Serial Numbers 1963 Through 1980, 2021 Through 2088, 2503 Through 2590, 2610, 2611, and 2614 Through 2677, 4001 Through 4024, 4163 Through 4166, 4501 Through 4581, 4583 Through 4594, 4602 Through 4604, and 4613 Through 4615. Compliance required as indicated. 1. As soon as possible, but not later than next 250 hours conduct magnetic particle or magnaglow inspections on main landing gear downlock spring cylinder assembly rod end P/N 295168 for cracks in the threaded portion. If cracks are found replace the part immediately. Repeat inspection, magnetic particle or 20-power magnifying glass at 300-hour intervals until replacement in accordance with paragraph 2 is accomplished. 2. Replace downlock spring cylinder assembly P/N 270104 with new assembly P/N 475211 as soon as practical but not later than the next progressive or base overhaul period approximately 2,500 hours. Concurrently with this replacement, line ream the lugs on the downlock strut assembly to (0.3770 inch-0.3780 inch) diameter and replace the spacer P/N 268225-2 with bushing P/N LS3859-4-1094 and replace bolt AN 23-21 with AN 23-22 attaching the lower end of the spring cylinder assembly to the downlock strut. (Lockheed Service Bulletins 49/860 and 1049/2709 also cover this subject.)
71-05-03: 71-05-03 BEECH: Amendment 39-1159 as amended by Amendment 39-1222 is further amended by Amendment 39-1244. Applies to Models A60 (Serial Numbers P123, P127 through P167), and all 99 series airplanes (Serial Numbers U-1 through U-136) approved for flight into known icing conditions. Compliance: Required as indicated, unless already accomplished. To prevent operation under unsafe environmental conditions, accomplish the following: A) Effective immediately, operation of the airplane in known icing conditions is prohibited. B) Prior to further flight, modify the existing operating limitations placard located on the righthand side wall adjacent to the copilot oxygen outlet on Beech Model A60 airplanes and located on the overhead panel on all Beech 99 series airplanes to read as follows: THIS AIRPLANE IS NOT APPROVED FOR FLIGHT IN KNOWN ICING CONDITIONS. C) On Beech Models A60 airplanes, install either transducer unit, P/N 3E1793, P/N 190-1, or P/N 190-2 in accordance with Beech Service Instruction 0430-355, Rev. 1, or by the accomplishment of any equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. D) On all Beech 99 series airplanes, install transducer unit P/N 795-1 in accordance with Beechcraft Service Instructions No. 0378-355, or by the accomplishment of any equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. E) Upon accomplishment of the modifications provided in Paragraphs C or D, compliance with the provisions of Paragraphs A and B is no longer required. NOTE: The Airplane Flight Manuals and Type Certificate Data Sheets for these model airplanes will be amended to reflect these changes. Amendment 39-1159 became effective March 2, 1971. Amendment 39-1222 became effective June 3, 1971. This Amendment 39-1244 becomes effective July 16, 1971.
93-16-13: 93-16-13 AEROSTAR INTERNATIONAL, INC.: Amendment 39-8697. Docket No. 93-CE-44-AD. Applicability: Models RX-6, RX-7, RX-8, RXS-8, S-40A, S-49A, S-50A, S-52A, S-55A, S-57A, S-57S, S-60S, S-60A, S-66A, S-71A, S-77A, 78C, 90C, QUBE-80, CTS, W100LB, 110P, and SPII balloons (all serial numbers), certificated in any category, that are equipped with hoses identified with one of the following: o AEROQUIP FC321-06 UL 5/16 LP-GAS HOSE 350 MAX. OPER. PSI 1Q92; o AEROQUIP FC321-06 UL 5/16 LP-GAS HOSE 350 MAX. OPER. PSI 2Q92; o AEROQUIP FC321-06 UL 5/16 LP-GAS HOSE 350 MAX. OPER. PSI 3Q92; o AEROQUIP FC321-06 UL 5/16 LP-GAS HOSE 350 MAX. OPER. PSI 4Q92; or o If hose identification is not legible. Compliance: Required as indicated after receipt of this AD, unless already accomplished. To prevent an uncontained fire in the balloon basket caused by a leaking fuel hose, accomplish the following: (a) Prior to further flight, remove each fuel hose fromthe leather sleeve, and, at system pressure, sniff test the entire length of the hose for signs of leakage. These signs could include frosting or chilling. (1) If any sign of fuel leakage is found, prior to further flight, replace the entire fuel hose/manifold assembly with an approved assembly that includes hoses with markings different than that specified in the Applicability section of this AD. (2) If no sign of fuel leakage is found, perform this fuel leakage test prior to each flight thereafter until the replacement required by paragraph (b) of this AD is accomplished. (b) Within the next 10 hours time-in-service after the effective date of this AD, unless already accomplished in accordance with paragraph (a)(1) of this AD, replace the entire fuel hose/manifold assembly with an approved assembly that includes hoses with markings different than that specified in the Applicability section of this AD. (c) Replacing the entire fuel hose/manifold assembly as required by either paragraph (a)(1) or (b) of this AD eliminates the repetitive test requirement of this AD. NOTE 1: Aerostar Service Bulletin No. 132, dated August 12, 1993, references the actions required by this AD. For the sake of inclusiveness, the procedures presented in this service bulletin have been incorporated into this AD. (d) The test and fuel hose/manifold assembly replacement required by this AD may be performed by the owner/operator holding at least a private pilot certificate as authorized by Federal Aviation Regulations (FAR) 43.7, and must be entered into the aircraft records showing compliance with this AD in accordance with FAR 43.11. (e) An alternative method of compliance or adjustment of the compliance times that provides an equivalent level of safety may be approved by the Manager, Chicago Aircraft Certification Office, FAA, 2300 East Devon Avenue, Room 232, Des Plaines, Illinois 60018. The request shall be forwarded through an appropriate FAA Maintenance Inspector, who may concur or comment and then send it to the Manager, Chicago Aircraft Certification Office. NOTE 2: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Chicago Aircraft Certification Office. (f) Replacement fuel hose/manifold assemblies and copies of the service bulletin referenced in NOTE 1 of this AD may be obtained from Aerostar International, Inc., 1812 E. Avenue, P.O. Box 5057, Sioux Falls, South Dakota 57117-5057. The service information may also be examined at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. (g) This amendment becomes effective on October 15, 1993, to all persons except those persons to whom it was made immediately effective by priority letter AD 93-16-13, issued August 18, 1993, which contained the requirements of this amendment.
95-10-06: This amendment adopts a new airworthiness directive (AD), applicable to certain Bombardier Model CL-600-1A11, -2A12, -2B16, and -2B19 series airplanes, that requires an inspection to verify the proper operation of the uplock latch of the air driven generator (ADG), and replacement of the uplock latch with a serviceable part, if necessary. This amendment also requires replacing the uplock assembly with a modified uplock assembly, and performing a rigging inspection. This amendment is prompted by a report indicating that, upon operation of the manual release system, the ADG did not deploy due to failure of the shaft pin. The actions specified by this AD are intended to prevent failure of the shaft pin, which could lead to the inability of the pilot to manually deploy the ADG when necessary (i.e., when an airplane's primary electrical power sources are lost and the ADG fails to deploy automatically).
76-07-11: 76-07-11 ALEXANDER SCHLEICHER: Amendment 39-2572. Applies to Model ASW-17 gliders, Serial Numbers 17001 through 17043, certificated in all categories. Compliance is required as indicated. To prevent water ballast from leaking into the fuselage area which could result in a loss of elevator control, accomplish the following: (a) Before further flight, install a placard in the vicinity of the water ballast system control to read as follows: "USE OF WATER BALLAST SYSTEM PROHIBITED." (b) Within the next 10 hours time in service after the effective date of this AD, unless already accomplished, comply with the following: (1) Install a standard 25 mm (1 inch) diameter hose clamp on each of the two polyvinylchloride (PVC) hoses to secure the two PVC hoses to the brass tubes that feed the water ballast exhausts which are located at the bottom of the fuselage behind the landing gear. (2) Inspect for clamp security by exerting a 10 pound pull on each PVC hose after securing clamps. (3) Install new clamps in place of those clamps which come loose following the inspection required by paragraph (b)(2) of this AD and inspect the replacement clamps in accordance with paragraph (b)(2) of this AD. (c) Upon compliance with paragraph (b) of this AD, the placard required by paragraph (a) of this AD may be removed. This amendment becomes effective April 22, 1976.
90-24-08: 90-24-08 FOKKER: Amendment 39-6805. Docket No. 90-NM-140-AD. Applicability: Model F-28 series airplanes, Serial Numbers 11003 to 11241, inclusive, and 11991 and 11992, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent reduced structural integrity of the fuselage, accomplish the following: A. For airplanes Serial Numbers 11008 through 11241: Inspect the fuselage lap joint at stringer 73 between frames 5305 and 9305, in accordance with Part 1 of Fokker Service Bulletin F-28/53- A94, Revision 1, dated July 5, 1989, and with the following schedule: 1. For airplanes that accumulated 32,000 landings or more as of August 18, 1989 (the effective date of Amendment 39-6284, AD 89-17-02), inspect within 2 days after August 18, 1989. 2. For airplanes that have accumulated fewer than 32,000 landings as of August 18, 1989, inspect within 60 days after August 18, 1989, or prior to the accumulation of 20,000 landings, whichever occurs later. B. Repeat the inspections required by paragraph A. of this AD at intervals not to exceed 1,000 landings. C. If cracks are found, repair prior to further flight, in accordance with Part 2 of Fokker Service Bulletin F28/53-A94, Revision 1, dated July 5, 1989. After repair, continue to inspect in accordance with Part 1 of the service bulletin, at intervals not to exceed 1,000 landings. D. Replace the lap joint at stringer 73 between frames 4900 through 9805, in accordance with the accomplishment instructions in Fokker Service Bulletin F28/53-95, Revision 1, dated February 16, 1990, according to the schedule below. Accomplishment of this modification terminates the requirement for the repetitive inspections required by paragraphs B. and C. of this AD. 1. For airplanes, Serial Numbers 11008 through 11241, inclusive: prior to the accumulation of 30,000 landings, or within one year after the effective date of this AD, whichever occurslater. 2. For airplanes, Serial Numbers 11003, 11004, 11006, 11991, and 11992: prior to the accumulation of 50,000 landings, or within one year after the effective date of this AD, whichever occurs later. 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, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM- 113, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Standardization Branch, ANM-113. F. 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 documents from the manufacturer may obtain copies upon request to Fokker USA, Inc., 1199 N. Fairfax Street, Alexandria, Virginia 22314. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton Washington. Airworthiness Directive 90-24-08 supersedes AD 89-17-02 (Amendment 39-6284) which superseded AD 89-07-01 (Amendment 39-6157). This amendment (39-6805, AD 90-24-08) becomes effective on December 17, 1990.
94-21-02: This amendment adopts a new airworthiness directive (AD) that is applicable to all Dornier Model 328-100 airplanes. This action requires repetitive tightening of the screws and quick-release fasteners on the wing/body fairing panels. This amendment is prompted by reports of loosened wing/body fairing panels. The actions specified in this AD are intended to prevent structural damage to the horizontal or vertical stabilizer and potential injury to persons on the ground due to loosened wing/body fairing panels that may separate from the airplane.
92-07-12: 92-07-12 BOEING: Amendment 39-8208. Docket No. 91-NM-164-AD.\n\n\tApplicability: Model 767 series airplanes, as listed in Boeing Service Bulletin 767-25-0137, Revision 1, dated May 9, 1991, certificated in any category.\n\n\tCompliance: Required within the next 18 months after the effective date of this AD, unless accomplished previously.\n\n\tTo ensure proper deployment of the off-wing escape system, accomplish the following:\n\n\t(a) Modify the off-wing escape system in accordance with Boeing Service Bulletin 767-25-0137, Revision 1, dated May 9, 1991.\n\n\t(b) 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Seattle ACO.\n\n\t(c) Special flight permits maybe issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.\n\n\t(d) The modification shall be done in accordance with Boeing Service Bulletin 767-25-0137, Revision 1, dated May 9, 1991. 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 Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC.\n\n\t(e) This amendment becomes effective on May 4, 1992.
94-08-02: This amendment adopts a new airworthiness directive (AD), applicable to certain Jetstream Model ATP series airplanes, that requires modification of the wiring for the electric-powered disconnect unit for the elevator control system and a subsequent functional test of the elevator control system. This amendment is prompted by an in-service report of damaged wire insulation in the electrical power circuit for the elevator disconnect unit, which resulted in grounding of the circuit and consequent uncommanded operation of the disconnect unit. The actions specified by this AD are intended to prevent uncommanded operation of the elevator disconnect unit, which would result in single elevator operation and, consequently, reduced controllability of the airplane.
78-22-05: 78-22-05 BEECH: Amendment 39-3329. Applies to the following models and serial number airplanes, certificated in all categories: \n\n\nMODEL\nSERIAL NUMBERS\n50\nH-1 through H-11\nB50, C50\nCH-12 through CH-360\nD50, D50A, D50B, D50C and D50E\nDH-1 through DH-347 \nE50\nEH-1 through EH-70\nF50\nFH-71 through FH-93, FH-95, FH-96\nG50\nGH-94, GH-97 through GH-119\nH50\nHH-120 through HH-149\nJ50\nJH-150 through JH-176\n65, A65, A65-8200\nL-1, L-2, L-6, LF-7, LF-8, LC-1 through LC-335\n70\nLB-1 through LB-35\n65-80, 65-A80, 65-A80-8800, 65-B80\nLD-1 through LD-511 \n65-88\nLP-1 through LP-28, LP-30 through LP-47\n65-90, 65-A90, B90, C90\nLJ-1 through LJ-797 \nE90\nLW-1 through LW-297\n99, 99A, A99A, B99\nU-1 through U-164\n100, A100\nB-1 through B-247 \nB100\nBE-1 through BE-55\n\n\tCOMPLIANCE: Required as indicated unless already accomplished. \n\n\tTo prevent partial loss of elevator control due to possible failure of elevator control push rods, in accordance with instructions set forth herein and in Beechcraft Service Instructions No. 0334-152, Revision III, or later approved revisions, accomplish the following: \n\n\tA)\tOn the following airplanes: \n\n\nMODEL\nSERIAL NUMBERS\n50\nH-1 through H-11\nB50, C50\nCH-12 through CH-360 \nD50, D50A, D50B, D50C and D50E\nDH-1 through DH-347 \nE50\nEH-1 through EH-70\nF50\nFH-71 through FH-93, FH-95, FH-96\nG50\nGH-94, GH-97 through GH-119\nH50\nHH-120 through HH-149\nJ50\nJH-150 through JH-176\n65, A65, A65-8200\nL-1, L-2, L-6, LF-7, LF-8, LC-1 through LC-335\n70\nLB-1 through LB-35\n65-80, 65-A80, 65-A80-8800 and 65-B80\nLD-1 through LD-511 \n65-88\nLP-1 through LP-28 and LP-30 through LP-47\n65-90, 65-A90, B90 and C90\nLJ-1 through LJ-716 \nE90\nLW-1 through LW-222\n99, 99A, A99A and B99\nU-1 through U-164\n100, A100\nB-1 through B-233\nB100\nBE-1 through BE-23\n\n\tUpon the accumulation of 1,050 hours total time-in-service for airplanes with less than 1,000 hours total time-in-service on the effective date of this AD,or \n\n\tWithin 50 hours time-in-service after the effective date of this AD, for airplanes with more than 1,000 hours total time-in-service on the effective date of this AD, that have not complied with AD 70-18-02, or \n\n\tWithin 500 hours time-in-service after the last inspection in accordance with AD 70-18- 02, for airplanes with more than 1,000 hours total time-in-service that have complied with AD 70-18-02: \n\n\t\t1.\tReview the airplane maintenance records for entries showing that Beechcraft Service Instructions No. 0334-152, Revision II, has been complied with. If Revision II of the Service Instructions has not been complied with, accomplish Paragraph A)2 of this AD. If Revision II of the Service Instructions has been complied with, accomplish Paragraph A)3 of this AD within an additional 100 hours time-in-service after this finding. \n\n\t\t2.\tGain access to the elevator control push rods (reference Figures I and II of this AD) and visually inspect for the words "CORROSION PROOFED"ink stamped near the center of the rods. If the words "CORROSION PROOFED" are ink stamped on a rod, no further action on that rod is required by this AD. If the words "CORROSION PROOFED" are not ink stamped on the rods, remove the elevator control push rods, clean, dye penetrant and visually inspect, check for bows, apply corrosion preventative treatment, identify and reinstall, all in accordance with Part I of Beechcraft Service Instructions No. 0334-152, Revision III, or later approved revisions. \n\n\t\t3.\tGain access to the elevator control push rods (reference Figures I and II of this AD) and visually inspect for the words "CORROSION PROOFED" ink stamped near the center of each rod. If the words "CORROSION PROOFED" are ink stamped on a rod, then no further action on that rod is required by this AD. If the words "CORROSION PROOFED" are not ink stamped on any rod and new rods were not installed during compliance with Revision II of the Service Instructions, paint a black stripe around the center of each rod in accordance with Part II of Beechcraft Service Instructions No. 0334-152, Revision III, or later approved revisions. If any new rods were installed during compliance with Revision II of the Service Instructions and the new rods are not ink stamped "CORROSION PROOFED", on each new rod so installed, remove the elevator control push rod, disassemble, inspect for corrosion preventative treatment and alignment, and if not found accomplished, apply the corrosion preventative treatment, identify, reassemble and reinstall, all in accordance with Part III of Beechcraft Service Instructions No. 0334-152, Revision III, or later approved revisions. \n\n\tB)\tOn the following airplanes: \n\n\nMODEL\nSERIAL NUMBERS\nC90\nLJ-717 through LJ-797\nE90\nLW-223 through LW-297\nA100\nB-234 through B-247 \nB100\nBE-24 through BE-55\n\n\tWithin 100 hours time-in-service after the effective date of this AD: \n\n\t\t1.\tGain access to the elevator control push rods (reference Figures Iand II of this AD) and visually inspect for the words "CORROSION PROOFED" ink stamped near the center of the rods. If the words "CORROSION PROOFED" are ink stamped on a rod, no further action on that rod is required by this AD. If the words "CORROSION PROOFED" are not ink stamped on the rods, remove the elevator control push rod tubes, disassemble, inspect for corrosion preventative treatment and alignment, and if not found accomplished, apply the corrosion preventative treatment, identify, reassemble and reinstall, all in accordance with Part III of Beechcraft Service Instructions No. 0334-152, Revision III, or later approved revisions. \n\n\tC)\tOn all airplanes affected by this AD, replace any elevator control push rods found cracked or otherwise unserviceable during any inspection required by this AD with serviceable rods that comply with the requirements of this AD prior to returning the airplane to service. \n\n\tD)\tAny equivalent means of compliance with this AD must be approvedby the Chief, Engineering and Manufacturing Branch, FAA, Central Region. \n\n\tThis AD supersedes AD 70-18-02, Amendment 39-907 (35 FR 305) as amended by Amendments 39-1074 (35 FR 13722 and 13723) and 39-1879 (39 FR 21120). \n\n\tThis amendment becomes effective November 6, 1978.
76-13-07: 76-13-07 CESSNA: Amendment 39-2656. Applies to Models 401, 402, 411, 414 and 421 Series Airplanes. Compliance: Required as indicated, unless already accomplished. To prevent failure of the fork bolt located at the aft end of the main landing gear retraction system torque tube connecting the torque tube to the outboard push/pull tube, accomplish the following: A) Within 200 hours' time in service on those airplanes having 1,800 or more hours' time in service or prior to 2,000 hours' time in service on those airplanes having less than 1,800 hours' time in service, and at each subsequent 2,000 hours' time in service thereafter, replace right and left 1/2-inch diameter P/Ns 0843518-1, 0843518-2, 0843500-35, 0843500-54, 5243518-1 and 5243518-3 fork bolts with new P/N 5243518-3 or FAA-approved superseding part number fork bolts on the airplanes specified below except airplanes on which P/N 5141052-1 fork bolts have been installed as field replacements: 401 - 401-0001 thru 401B0053 402 - 402-0001 thru 4-2B0035 411 - All Serials 414 - 414-0001 thru 414-0098 421 - 421-0001 thru 421A0158 B) Within 200 hours' time in service on those airplanes having 4,800 or more hours' time in service or prior to 5,000 hours' time in service on those airplanes having less than 4,800 hours' time in service, and at each subsequent 5,000 hours' time in service thereafter, replace right and left 5/8-inch diameter P/N 5141052-1 fork bolts with new P/N 5141052-1 or FAA-approved superseding part number fork bolts on the airplanes specified below or any lower airplane serial numbers on which these fork bolts have been installed as field replacements. 401 - 401B0054 and on 401 - 402B0036 and on 414 - 414-0099 and on 421 - 421B0001 and on C) Fork bolt life limits set by this AD may be extended 25 hours, up to 2,025 hours for 1/2-inch diameter fork bolts and 5,025 hours for 5/8-inch diameter fork bolts, to allow replacement at regular scheduled maintenance or inspections. D) Aircraft may be flown in accordance with FAR 21.197 to a base where this AD may be accomplished. E) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Cessna Service Letter ME75-23 or later approved revisions refers to this subject. This amendment becomes effective July 7, 1976.
77-13-02: 77-13-02 HAWKER SIDDELEY AVIATION, LTD: Amendment 39-2926. Applies to Model DH/BH-125 airplanes, all series, certificated in all categories. Compliance is required as indicated. To prevent the failure of the knife edges of the brake control valve, P/N AC. 61520, and the possible complete loss of braking on one side of the airplane with no advance warning to the flight crew, accomplish the following: (a) Comply with paragraph (b) or (c) of this AD as follows, and, thereafter, continue to comply with paragraph (b) or (c) of this AD at intervals not to exceed 4,500 landings since last compliance: (1) For airplanes having brake control valve knife edges that have accumulated less than 4,300 landings on the effective date of this AD, compliance is required prior to the accumulation of 4,500 landings. (2) For airplanes having brake control valve knife edges that have accumulated 4,300 or more landings, but less than 5,900 landings, on the effective date ofthis AD, compliance is required prior to the accumulation of an additional 200 landings. (3) For airplanes having brake control valve knife edges that have accumulated 5,900 or more landings on the effective date of this AD, compliance is required prior to the accumulation of 6,100 landings or an additional 100 landings whichever occurs later. (4) For airplanes for which no records exist that indicate the number of landings the brake control valve knife edges have accumulated, compliance is required prior to the accumulation of 100 landings after the effective date of this AD. (b) Replace the knife edges with new parts, P/Ns ACO. 34629, ACO. 34630, and ACO. 36133, in accordance with Paragraph A, of Section 2, titled "Accomplishment Instructions," of Hawker Siddeley Aviation, Ltd. Service Bulletin 32-166, dated January 27, 1976, or an FAA- approved equivalent. (c) Replace the brake control valve, P/N AC. 61520, with a valve of the same part number that incorporates knife edges having part numbers specified in paragraph (b) of this AD in accordance with Paragraph B, of Section 2, titled "Accomplishment Instructions," of Hawker Siddeley Aviation, Ltd. Service Bulletin 32- 166, dated January 27, 1976, or an FAA-approved equivalent. This amendment becomes effective July 20, 1977.
77-10-12: 77-10-12 MCDONNELL DOUGLAS: Amendment 39-2906. Applies to Model DC-8 Series airplanes, certificated in all categories. \n\n\tCompliance required within the next 300 hours time in service or 30 days after the effective date of this AD, whichever comes first, unless already accomplished within the past 30 days. \n\n\tTo detect cracks and prevent failure or jamming of the elevator geared tab crank arm and gust lock assemblies, comply with the following: \n\n\t(a)\tVisually inspect both the left and right side inboard and outboard elevator geared tab crank arm assemblies, P/N's 4710541 and 4710542 for failure and/or cracks. \n\n\t(b)\tVisually inspect the P/N 5644178 gust lock assembly, and specifically, the torque tube section of the P/N 4644181 crank assembly, for failure and/or cracks. \n\n\t(c)\tFailed or cracked parts must be replaced with like serviceable parts before further flight. \n\n\t(d)\tVerify that clearance exists between the crank assemblies, P/N 4710541 and 4710542, and the boxsection for all positions along elevator travel. \n\n\tNote: McDonnell Douglas DC-8 Alert Service Bulletin A27-262, dated April 28, 1977, covers the same subject. \n\n\t(e)\tSpecial flight permits may be issued in accordance with FAR's 21.197 and 21.199 to authorize operation of an airplane to a base to perform the inspections required by this AD. \n\n\tThis amendment becomes effective May 26, 1977.
78-20-03: 78-20-03 HUGHES: Amendment 39-3307 as amended by Amendment 39-3599. Applies to Hughes Model 369D and 369H helicopters which have either the C-500 aerial spray system, STC No. SH184NW or the Chadwick C-500 fire suppression kit, STC No. SH567NW installed. To prevent hydraulic fluid from entering the engine after a seal failure of the Chadwick hydraulic motor, accomplish the following: Within the next 100 hours time in service or 30 days, whichever comes first, after the effective date of this AD, install a drain line on the engine mounting flange, Chadwick P/N 500- 60405-1, in accordance with Chadwick Service Bulletin 500-78-02 dated September 1, 1978, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. To prevent possible fatigue cracking of the helicopter undercarriage, accomplish the following: Within the next 50 hours flying time or six months, whichever comes first, after the effectivedate of this amended AD, install a doubler, P/N 500-60487-1, and grommet, P/N AN931-7-11, in accordance with Chadwick Service Bulletin 500-79-01 dated September 27, 1979, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest 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 Chadwick Inc., 11969 SW Herman Road, Sherwood, Oregon 97140. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. Amendment 39-3307 became effective November 2, 1978. This Amendment 39-3599 becomes effective November 7, 1979.
93-08-08: 93-08-08 DE HAVILLAND, INC.: Amendment 39-8555. Docket 92-NM-144-AD. Applicability: Model DHC-8 series airplanes equipped with microwave landing system (MLS) provisions, as listed in de Havilland Service Bulletin S.B. 8-34-60, Revision 'C,' dated November 1, 1991; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent the airplane from landing short of the runway, accomplish the following: (a) Prior to installation of an MLS, or within 45 days after the effective date of this AD, whichever occurs later, perform a general visual inspection to determine whether the vertical deviation wiring has been properly installed and/or connected to link the MLS with the ground proximity warning system (GPWS), in accordance with de Havilland Service Bulletin S.B. 8-34-60, Revision 'C,' dated November 1, 1991. (1) If any vertical deviation wiring has been found that has not been properly installed and/or connected to link the MLS with the GPWS, prior to further flight, correct the wiring and perform a functional test of the system in accordance with the service bulletin. (2) If vertical deviation wiring has been found that has been properly installed and/or connected to link the MLS with the GPWS, no further action is necessary. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, New York Aircraft Certification Office (ACO), FAA, Engine and Propeller Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, New York ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York ACO. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The inspection, wiring correction, and functional test shall be done in accordance with de Havilland Service Bulletin S.B. 8-34-60, Revision 'C,' dated November 1, 1991. 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 de Havilland, Inc., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Engine and Propeller Directorate, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on May 26, 1993.
94-07-07: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model MD-11 and MD-11F airplanes, that requires modification of the fuel crossfeed low level dump system shutoff. This amendment is prompted by an FAA determination that, in the event of a failure of the number 2 bus tie relay and the subsequent loss of the number 2 electrical power source, an all-engine flameout event could occur due to fuel starvation during or shortly after a fuel dumping operation. The actions specified by this AD are intended to prevent loss of the fuel dump system shutoff due to a failure of the number 2 DC bus electrical relay and the subsequent loss of the number 2 electrical power source.
91-17-01: 91-17-01 BEECH: Amendment 39-7099. Docket No. 90-CE-71-AD. Applicability: The following Model airplanes, certificated in any category: Models Serial Numbers 35-33, 35-A33, 35-B33, CD-1 through CD-981 and 35-C33, E33, F33, G33 CD-983 through CD-1304 35-C33A, E33A, F33A CE-1 through CE-235, CE-249, CE-250, CE-256, CE-260, CE-264 through CE-268, and CE-270 through CE-1565 E33C, F33C CJ-1 through CJ-179 36, A36 E-1 through E-2103, E-2105 through E-2110 A36TC, B36TC EA-1 through EA-319, and EA-321 through EA-388 T34C-1 GM-1 through GM-142 34C GP-1 through GP-50 T-34C GL-1 through GL-353 45 G-3 through G-6 A45 G-7 through G-156, G-257 through G-306, G-696 through G-845, CG-1 through CG-47, CG-58 through CG-60, CG-68, CG-73, CG-75, CG-78, CG-79, CG-105, CG-106, CG-108, CG-111 through CG-179, CG-200 through CG-223, CG-279 through CG-319 D45 BG-1 through BG-423 95 TD-2 through TD-302 B95 TD-303 through TD-452 B95A TD-453 through TD-533 D95A TD-534 through TD-707 E95 TD-708 through TD-721 95-55 TC-1 through TC-190 95-A55 TC-191 through TC-349, TC-351 through TC-370, and TC-372 through TC-501 95-B55, 95-B55A TC-371, TC-502 through TC-2456 95-C55 TC-350 95-C55A TE-1 through TE-49, and TE-51 through TE-451 D55, D55A TE-452 through TE-767 E55, E55A TE-768 through TE-1201 56TC TG-2 through TG-83 A56TC TG-84 through TG-94 58, 58A TH-1 through TH-1388, and TH-1390 through TH-1395 58P, 58PA TJ-3 through TJ-435, and TJ-437 through TJ-443 58TC, 58TCA TK-1 through TK-150 Compliance: Required the next time the elevator trim tab actuators are removed for any reason, but no later than 12 calendar months after the effective date of this AD, unless already accomplished. To prevent loss of control of the airplane because of interchanging the right-hand and left-hand elevator trim tab actuators, accomplish the following: (a)Paint a stripe on each stabilizer rear spar (right-hand black; and left-hand blue) in accordance with the Accomplishment Instructions of Beech Service Bulletin No. 2399, dated March 1991. (b) Remove the cover over the actuator inspection hole on each stabilizer and paint the inspection hole ledges (right-hand black; and left-hand blue) in accordance with the Accomplishment Instructions of Beech Service Bulletin No. 2399, dated March 1991. (c) Paint a stripe .50 by 1 inch on each actuator housing through the inspection holes (right-hand black; and left-hand blue) in accordance with the Accomplishment Instructions in Beech Service Bulletin No. 2399, dated March 1991. Actuators must not be removed to paint the .50 by 1 inch stripe on the housing. NOTE: A left-hand trim tab actuator will have threads on its actuator screw that will rotate clockwise when screwed into the actuator assembly, and a right-hand trim tab actuator will have threads on its actuator screw that willrotate counterclockwise when screwed into the actuator assembly. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a location where the requirements of this AD can be accomplished. (e) An alternative method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (f) The modification required by this AD shall be done in accordance with Beech Service Bulletin No. 2399, dated March 1991. 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 the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-7099, AD 91-17-01) becomes effective on November 25, 1991.
95-12-03: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 0100 series airplanes, that requires installation of reinforcement plates at certain fuselage stations. This amendment is prompted by a report indicating that cracks were found in the frame strips at certain fuselage stations on a Model F28 Mark 0100 series airplane test article due to fatigue-related stress. The actions specified by this AD are intended to prevent such fatigue-related cracking, which could result in reduced structural integrity of the fuselage pressure vessel.