Results
99-23-23: This amendment adopts a new airworthiness directive (AD) applicable to Bell Helicopter Textron, Inc. (BHTI) Model 412, 412EP, and 412CF helicopters. This action requires inspecting and measuring the thickness of certain main rotor yoke assemblies. This AD also requires adding 500 hours time-in-service (TIS) to the total time for main rotor yoke assemblies that measure below 0.478-inch thickness and noting the measurement and added TIS on the component history card or equivalent record. This amendment is prompted by a report of an emergency landing due to severe main rotor vibration on a BHTI Model 412 helicopter. Subsequent fatigue analysis indicates that the main rotor yoke assembly (yoke) does not have the anticipated service life when manufactured below 0.478-inch thickness. The actions specified in this AD are intended to prevent a fatigue failure of the yoke, loss of a main rotor blade, and subsequent loss of control of the helicopter.
61-23-03: 61-23-03 DE HAVILLAND: Amdt. 358 Part 507 Federal Register November 2, 1961. Applies to All Heron Model 114 Aircraft Without Modification No. 1454. Compliance required as indicated. (a) Cases have occurred of cracking of the wing rear false spar web adjacent to the wing-to-fuselage attachment P/N 14W253/4. To preclude failure of the spar, an X-ray or visual inspection for cracks must be conducted in accordance with de Havilland Technical News Sheet CT(114) No. W.10 Issue 2, within the next 250 hours' time in service after the effective date of this directive and at each 600 hours' time in service thereafter. If cracks are found, repair in accordance with de Havilland Drawing RD14W 224, Issue 4, or subsequent, within the time in service given in paragraphs (1), (2), and (3). (1) Cracks less than 1/2 inch in length must be repaired at the next wing removal and the inspection in (a) must be made every 600 hours' time in service in the interim between the inspection and the wing removal. (2) Cracks of 1/2 inch to 1 1/2 inches in length must be repaired within 300 hours' time in service after the inspection. (3) Cracks exceeding 1 1/2 inches in length must be repaired within the next 150 hours' time in service after the inspection. (b) The special inspection in (a) is no longer required when the repair per Drawing RD 14W 224, Issue 4, or subsequent, has been incorporated. (de Havilland Technical News Sheet CT(114) No. W.10 Issue 2 dated July 24, 1961, covers this subject.) This directive effective December 4, 1961.
85-17-08: 85-17-08 SIKORSKY AIRCRAFT: Amendment 39-5196. Applies to Model S-58A, B, C, D, E, F, G, H, J, BT, DT, ET, FT, HT, and JT helicopters certificated in any category and CH- 34 series HH-34 series, SH-34 series, UH-34 series, and VH-34 series helicopters certificated in the restricted category. Compliance is required as indicated, unless already accomplished. To prevent the separation of the stationary star and rotating star, accomplish the following: (a) Within the next 10 hours time in service after the effective date of this AD, and thereafter at intervals not to exceed 25 hours time in service from the last inspection, visually inspect the 12 main rotor head star assembly retainer bolts, Part Number lP/N) MS20074-04-04 or P/N AN74A4, for damage and for security of fastening in accordance with Section 2, Paragraph A, of Sikorsky Alert Service Bulletin (ASB) No. 58B10-19, dated May 31, 1985, or later FAA-approved revision. (b) If damaged or loose bolts are found, fluorescent magnetic particle inspect the bolts for cracks and visually inspect the bolt threads for crossed, stripped, or flattened threads in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 58B10-19, dated May 31, 1985, or later FAA-approved revision. (c) If a cracked bolt is found, or if the bolt threads are crossed, stripped, or flattened, replace with a new (zero time) bolt prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 58B10-19, dated May 31, 1985, or later FAA-approved revision. (d) Within the next 100 hours time in service after the effective date of this AD, unless already accomplished within the last 1,000 hours time in service, replace the 12 main rotor head star assembly retainer bolts with new (zero time) bolts in accordance with Section 2, Paragraphs A(7)(c), (d), and (e) and Paragraph B, of Sikorsky ASB No. 58B10-19, dated May 31, 1985, or later FAA-approved revision; thereafter, install new (zero time) main rotor head star assembly retainer bolts at intervals not to exceed 1,100 hours time in service. (e) Render unairworthy any bolts removed in accordance with the requirements of paragraph (d) of this AD by crushing the threads. (f) Aircraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the AD can be accomplished. (g) An alternate method of compliance or adjustments of the compliance time which provide an equivalent level of safety may be used when approved by the Manager, Boston Aircraft Certification Office, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone number (617) 273-7118. 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 Sikorsky Aircraft Division, United Technology Corporation, North Main Street, Stratford, Connecticut 06601. These documents also may be examined at the Rules Docket, Office of the Regional Counsel, FAA, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment becomes effective January 13, 1986, as to all persons except those persons to whom it was made immediately effective by priority letter AD No. 85-17-08, issued August 30, 1985, which contained this amendment.
61-22-08: 61-22-08 SUD AVIATION: Amdt. 354 Part 507 Federal Register October 28, 1961. Applies to All Model SE 3130 Alouette II Helicopters. Compliance required as indicated. To remove defective bolts and preclude the possibility of further failure of the tail rotor gear box housing attachment the following inspections are required: (a) The tail rotor gear box attachment bolts P/N 66.20.043 and tail rotor gear box guides P/N 66.20.213 shall be inspected within the next 10 hours' time in service in accordance with paragraphs (d) and (e) unless this inspection has already been complied with and/or the bolts replaced with parts that comply with paragraphs (d) and (e) subsequent to August 25, 1961. (b) Every 50 hours' time in service subsequent to the completion of inspection prescribed in paragraph (a), reinspect and check the torque of the bolts and nuts P/N 66.20.043. The torque should be between 10.1 and 12.3 ft. lbs. If the bolts and nuts do not meet the torque requirements, remove the tail rotor gear box and inspect in accordance with paragraphs (d) and (e). (c) Every 100 hours' time in service subsequent to completing inspection of paragraph (a), remove the tail rotor gear box and inspect the bolts and guides in accordance with paragraphs (d) and (e). (d) The bolts P/N 66.20.043 are to be checked for cracks, corrosion, peeling and surface finish. Surface finish inspection is applicable to paragraph (a) only. Remove the zinc chromate protective finish, if applicable, by using paint remover. The diameter of the bolt bearing area at the head and at the thread end shall be measured in two directions 90 degrees apart in order to detect any out of round condition. These diameters shall not be less than 8.29 mm. (0.326 inch). The surface finish of the central necked down area and the radius at each end shall be as follows: (1) The radius shall not be less than 1.61 mm. (0.0630 inch). (2) The maximum permissible surface roughnessshall not exceed 1 micron (39 micro-inches). (3) Localized defects no greater than 50 microns deep (1950 micro-inches) are permissible. If these standards are not met the bolts shall be replaced with bolts that do meet the standards. (e) Check the bore of the three tail rotor gear box guides P/N 66.20.213 by measuring the bore along two directions 90 degrees apart. The bore dimension shall not exceed 8.33 mm. (0.328 inch). If these standards are not met the tail rotor gear box shall be replaced with a new or overhauled unit or the gear box returned to the factory or approved overhaul agency for installation of new guides before reinstallation on the helicopter. (Sud Maintenance Manual Vol. I, Chapter 5, Pages 3 and 7, and Sud Service Bulletin No. 66-11-206 cover the same subject.) This directive effective October 28, 1961.
2018-02-15: We are superseding Airworthiness Directive (AD) 2007-08-06 for British Aerospace Regional Aircraft Models HP.137 Jetstream Mk.1, Jetstream Series 200, Jetstream Series 3101, and Jetstream Model 3201 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and address an unsafe condition on an aviation product. The MCAI describes the unsafe condition as the need for airworthiness limitations for critical components in the main and nose landing gear assemblies. We are issuing this AD to require actions to address the unsafe condition on these products.
2018-02-02: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Model AS350B, AS350BA, AS350B1, AS350B2, AS350B3, AS350D, AS355E, AS355F, AS355F1, AS355F2, AS355N, AS355NP, EC130B4, and EC130T2 helicopters. This AD requires inspecting the main rotor (M/R) mast jet oil lubrication hose (oil hose). This AD is prompted by a report of a blocked oil hose. The actions of this AD are intended to prevent an unsafe condition on these helicopters.
61-21-04: 61-21-04 DOUGLAS: Amdt. 347 Part 507 Federal Register October 13, 1961. Applies to All DC-8 Aircraft Equipped With JT3C, JT4 and Conway Engine Installations (Models DC-8-11, -12, -31, -32, -33, -41, -42 and -43). \n\n\tCompliance required as soon as the installation of available parts can be scheduled but not later than the next 600 hours' time in service after October 13, 1961, unless an operator has obtained approval from the Chief, Engineering and Manufacturing Branch, Federal Aviation Agency Western Region for an alternative compliance program. \n\n\tAs a result of numerous recent failures of the flexible hoses in the discharge lines of the engine-driven hydraulic pumps, unless already accomplished, certain hoses and clamps approved as part of the basic type design must be removed and replaced as follows: \n\n\t(a) Replace hoses an indicated or with FAA approved equivalents: \n\n\nAirplane Serial Nos. \nRemove Hose P/N\nInstall Hose P/N\n\n\n\nAll DC-8-11 and -12 aircraft(JT3-C engines)\n5654402-10-6129\n5765665-10-6129\nAll DC-8-21, -31, -32 and -33 aircraft (JT4 engines) except S/N 45422-45431, 45433 and 45602-45606\n5654402-10-6724\n5765665-10-6824\n45422-45431, 45433 and 45602-45606\n618-10-0676\n624663-10-0676\nAll DC-8-41, -42 and -43 aircraft (Conway engines)\n5654402-10-5700\n5765665-10-5700\n\n\t(b) Replace clamps, P/N 4365431D21C, as used with hoses removed per (a), with clamps listed in Douglas DC-8 Service Bulletin No. 29-37, Revision No. 4, dated November 6, 1961 (or later issue), or with equivalent FAA-engineering approved clamps. The replacement clamps shall be installed in the manner and positions described in Service Bulletin No. 29-37. \n\n\t(Douglas DC-8 Service Bulletin No. 29-37, pertains to this same subject.) \n\n\tThis directive effective October 13, 1961. \n\n\tRevised December 6, 1961.
2018-02-14: We are adopting a new airworthiness directive (AD) for certain Honeywell International Inc. (Honeywell) TPE331 turboprop and TSE331 turboshaft engines. This AD was prompted by reports that combustion chamber case assemblies have cracked and ruptured. This AD requires inspection of the affected combustion chamber case assembly, replacement of those assemblies found cracked, and removal of affected assemblies on certain TPE331 and TSE331 engines. We are issuing this AD to address the unsafe condition on these products.
87-21-02 R1: 87-21-02 R1 CESSNA: Amendment 39-5740 as revised by Amendment 39-6215. Applicability: The following airplanes equipped with reciprocating engines certificated in any category: MODEL SERIAL NUMBER T303 T30300001 thru T30300301 310D 39032 thru 39299 310E 310M0001 thru 310M0036 310F 310-0001 thru 310-0156 310G thru 310R (Including T310P, T310Q and T310R) 310G0001 thru 310R2140 320 thru 320F 320-0001 thru 320F0045 335 335-0001 thru 335-0065 340 thru 340A 340-0001 thru 340A1543 401 thru 401B 401-0001 thru 401B0221 402 thru 402C 402-0001 thru 402C0653 404 404-0001 thru 404-0859 411 thru 411A 411-0001 thru 411A0300 414 thru 414A 414-0001 thru 414A0858 421 thru 421C 421-0001 thru 421C1257 Compliance: Required as indicated in the body of the AD, unless already accomplished per the unrevised version of this AD. To preclude misfueling of the airplane resulting in engine failure, accomplish the following: (a) Within the next 12 calendar months after the effective date of this AD, unless already accomplished, modify all fuel filler opening(s) in accordance with the instructions contained in Cessna Service Information Letter ME84-31 dated July 20, 1984. (b) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (c) In accordance with FAR Part 43, Appendix A, Item (c) 29, the modifications required by this AD (except installation of the SK303-29 kit) is preventative maintenance and may be performed by the holder of a pilot certificate issued under FAR Part 61 on airplanes owned or operated by him subject to the limitations of FAR 43.3(a). The maintenance record entries required by FAR 43.9 and FAR 91.173 must be accomplished. (d) An equivalent means of compliance with this AD may be used if approved by the Manager, Wichita Aircraft Certification Office, Federal Aviation Administration, 1801 Airport Road, Room 100, Wichita, Kansas 67209.All persons affected by this directive may obtain copies of the document(s) referred to herein upon request to Cessna Aircraft Company, Customer Services, Post Office Box 1521, Wichita, Kansas 67201; or may examine these documents at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD revises AD 87-21-02, Amendment 39-5740 which became effective on November 2, 1987. This amendment (39-6215, AD 87-21-02 R1) becomes effective on June 16, 1989.
2018-01-12: We are superseding Airworthiness Directive (AD) 2015-22-53 for Airbus Helicopters Model AS350B3 helicopters. AD 2015-22-53 required revising the rotorcraft flight manual (RFM) to perform the yaw load compensator check after rotor shut-down and to state that the yaw servo hydraulic switch must be in the ``ON'' position before taking off. Since we issued AD 2015-22-53, Airbus Helicopters developed a modification of the ACCU TST switch. This new AD retains the requirements of AD 2015-22-53 and requires modifying the yaw servo hydraulic switch (collective switch) and replacing the ACCU TST button. The actions of this AD are intended to address an unsafe condition on these products.
2018-01-09: We are superseding Airworthiness Directive (AD) 95-25-02, which applied to certain Fokker Services B.V. Model F28 Mark 0100 airplanes. AD 95-25-02 required inspection(s) to detect cracks of the fuselage-mounted half of hinge assemblies of the small cargo door, and replacement of any cracked hinge assembly with a new hinge assembly. This new AD was prompted by a report that the hinges of the small cargo door are made of a material that is sensitive to stress corrosion and fatigue cracking, and by the determination that the existing inspection program does not provide sufficient protection against fatigue-induced cracks. This AD requires contacting the FAA to obtain instructions for addressing the unsafe condition on these products, and doing the actions specified in those instructions. We are issuing this AD to address the unsafe condition on these products.
2003-23-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-400, -400D, and -400F series airplanes, that requires reviewing airplane maintenance records; inspecting the yaw damper actuator portion of the upper and lower rudder power control modules (PCM) for cracking, and replacing the PCMs if necessary; and reporting airplane maintenance records review and inspection results to the manufacturer. This action is necessary to detect and correct cracking in the yaw damper actuator portion of the upper and lower rudder PCMs, which could result in an uncommanded left rudder hardover, consequent increased pilot workload, and possible runway departure upon landing. This action is intended to address the identified unsafe condition.
98-19-17: This amendment adopts a new airworthiness directive (AD) that applies to certain Glaser-Dirks Flugzeugbau GmbH (Glaser-Dirks) Model DG-400 gliders. This AD requires inspecting the powerplant mount and the propeller mount for any loose parts. This AD also requires modifying the starter motor, retrofitting the holder for the starter motor, and checking the engine ignition timing; either immediately or at a certain time depending on the results of the inspection. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. The actions specified by this AD are intended to prevent damage to the engine caused by vibration, which could result in loss of engine power during critical phases of flight.
87-08-02: 87-08-02 MCDONNELL DOUGLAS: Amendment 39-5606. Applies to McDonnell Douglas Model DC-9-10, -20, -30, -40, and C-9 (military) series airplanes, Fuselage Numbers 1 through 619, certificated in any category. Compliance required as indicated, unless previously accomplished. \n\n\tTo prevent outer skin cracks of the rudder and subsequent damage to adjacent structure, within 1,800 landings, or 9 months, after the effective date of this AD, whichever occurs earlier, accomplish the following, unless already accomplished within the last 1,200 landings: \n\n\tA.\tRadiographically inspect rudder ribs for cracks, in accordance with McDonnell Douglas DC-9 Service Bulletin 55-23, Revision 4, dated September 8, 1986, hereinafter referred to as S/B 55-23, or later FAA-approved revisions, and accomplish the following: \n\n\t\t1.\tIf no cracks are found, accomplish repetitive inspections at intervals not to exceed 3,000 landings, until such time as the requirements of paragraph A.3., below, are accomplished. \n\n\t\t2.\tIf cracks are found, accomplish one of the following, as applicable: \n\n\t\t\ta.\tFor cracks in rudder ribs only: \n\n\t\t\t\t(1)\tIf one rib is found cracked and the total length of crack does not exceed one-half the length of the cracked rib, perform repetitive inspections for rudder skin crack(s) in accordance with S/B 55-23, at intervals not to exceed 150 landings, until such time as the requirements of paragraph A.3., below, are accomplished. \n\n\t\t\t\t\t(a)\tIf the rib crack exceeds one-half the length of the cracked rib, accomplish the requirements of paragraph A.2.b.(1), below. \n\n\t\t\t\t\t(b)\tIf skin crack(s) are found, accomplish the requirements of paragraph A.2.b., below. \n\n\t\t\t\t(2)\tIf two adjacent ribs are found cracked and the total length of cracks for each rib does not exceed 6.0 inches, perform repetitive inspections for rudder skin cracks in accordance with S/B 55-23, at intervals not to exceed 150 landings, until such time as the requirements of paragraph A.3., below, are accomplished. \n\n\t\t\t\t\t(a)\tIf the rib crack exceeds 6.0 inches, accomplish the requirements of paragraph A.2.b.(1), below. \n\n\t\t\t\t\t(b)\tIf a skin crack(s) is found, accomplish the requirements of paragraph A.2.b., below. \n\n\t\t\t\t(3)\tIf two alternate ribs are found cracked, and the total length of the cracks does not exceed 16.0 inches, perform repetitive inspections for rudder skin cracks in accordance with S/B 55-23, at intervals not to exceed 150 landings until such time as the requirements of paragraph A.3., below, are accomplished. \n\n\t\t\t\t\t(a)\tIf the rib cracks exceed 16.0 inches, accomplish the requirements of paragraph A.2.b.(2), below. \n\n\t\t\t\t\t(b)\tIf a skin crack is found, accomplish the requirements of paragraph A.2.b., below. \n\n\t\t\t\t(4)\tIf more than two ribs are found cracked, notwithstanding the crack lengths, accomplish the requirements of paragraph 2.b.(1), below. \n\n\t\t\tb.\tFor cracks found in the rudder skin, or rudder rib and skin, accomplish the following:(1)\tBefore further flight, accomplish repairs to cracked rib(s) in accordance with S/B 55-23, or later FAA-approved revisions. \n\n\t\t\t\t(2)\tUpon completing repairs to cracked rib(s), accomplish skin repair in accordance with McDonnell Douglas DC-9 Structural Repair Manual, Section 55-03. \n\n\t\t3.\tInstallation of rib stiffeners in accordance with S/B 55-23, or replacement of all affected ribs with new production .040-inch thick 2024-T42 aluminum ribs, constitutes terminating action for the repetitive inspections required by this AD. \n\n\tB.\tAlternate means of compliance which provides an acceptable level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tC.\tUpon the request of an operator, an FAA Maintenance Inspector, subject to prior approval of the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region, may adjust the repetitive inspection intervals specified in this AD topermit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. \n\n\tD.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes with rudder rib cracks only (within the limits of this AD) to a base in order to comply with the requirements of this AD. For airplanes with rudder skin cracks, the rudder must be repaired or replaced prior to next flight. \n\n\tAll persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Los Angeles Aircraft Certification Office, 4344 Donald Douglas Drive, Long Beach, California. \n\n\tThis amendment becomes effective May 21, 1987.
48-18-03: 48-18-03 LOCKHEED: Applies to Model 49-46 Aircraft. Compliance required every 300 hours of operation. 1. Cabin supercharger drive shafts should be inspected at periods not to exceed 300 hours of operation in accordance with instructions and procedures specified in LAC Service Bulletin 49/SB-107, revised November 22, 1946. Concurrent with the foregoing inspection, the rear drive shaft universal joint, clutch end bearing, carbon oil seal and overriding clutch, should be completely overhauled. All defective parts are to be replaced and clutch end bearing 111GE is to be replaced regardless of condition. The sheet-metal retainer (LAC P/N 257643 is to be replaced as soon as practicable with bronze retainer (LAC P/N 299449). 2. The replacement of clutch end bearing 111GE will not be necessary if the supercharger is reworked to provide a double bearing support for the rear universal joint, and overrunning clutch assembly. This rework will also require replacing the present carbon faced oil seal with a slinger type and modify the supercharger housing to suit. The pre-flight inspections for oil seal damage can be dispensed with when slinger type seals have been installed. (LAC Service Bulletin 49/SB-393 covers this same subject.)
2018-01-05: We are adopting a new airworthiness directive (AD) for certain Fokker Services B.V. Model F28 Mark 0070 and 0100 airplanes. This AD requires contacting the FAA to obtain instructions for addressing the unsafe condition on these products, and doing the actions specified in those instructions. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the fuselage frames are subject to widespread fatigue damage (WFD). We are issuing this AD to address the unsafe condition on these products.
2003-21-08: This amendment adopts a new airworthiness directive (AD) for the specified Eurocopter France (Eurocopter) model helicopters that requires inspecting certain main rotor blades for disbonds, which may be indicated by cracking, and repairing or replacing each main rotor blade (MRB) as necessary. This amendment is prompted by the discovery of disbonded leading edge protective strips. The actions specified by this AD are intended to detect disbonding between the stainless steel protective strip and the MRB skin, which could cause loss of the protective strip, an out-of-balance condition, and subsequent loss of control of the helicopter.
61-21-03: 61-21-03 DOUGLAS: Amdt. 346 Part 507 Federal Register October 13, 1961. Applies to All DC-8 Series Aircraft. \n\n\tCompliance required within the next 150 hours' time in service after the effective date of this AD. \n\n\tAs a result of recent incidents which have shown the need for effective quantity indication of "reserve" hydraulic fluid in the system reservoir, the following must be accomplished: \n\n\tUnless already accomplished in accordance with FAA approved technical data, replace or modify in accordance with Douglas Service Bulletin No. DC-8 A29-40 (Reissue No. 1 or later) or FAA approved equivalent, the existing hydraulic quantity gage in the cockpit and the quantity transmitter in the fluid reservoir, to provide a system which will indicate the quantity of reserve hydraulic fluid in the reservoir down to approximately one gallon. The system indication error shall not exceed 10 percent. The dial on the quantity gage shall be marked in a manner which divides the total indicator range into three segments and identifies the three segments, as follows: \n\n\nSegment\nReserve Fluid Quantity in Reservoir\nNORMAL\nFull (11.5 gal.) to 4.9 gal.\nAUX.\n4.9 gal. to 1.5 gal.\nEMER.\n1.5 gal. to approx. 1.0 gal.\n\t\n\tThe exact quantities in gallons need not be marked on the quality indicator. "EMER." need not be shown provided that this segment is otherwise marked in a manner distinctly different from the other segments. \n\n\t(Douglas Service Bulletin No. DC-8 A29-40 (Reissue No. 1 or later) pertains to this same subject.) \n\n\tThis directive effective October 13, 1961.
87-01-05: 87-01-05 EMPRESA BRASILEIRA DE AERONAUTICA S.A. (EMBRAER): Amendment 39-5490. Applies to Models EMB-110P1 and EMB-110P2 (Serial Numbers 110001 through 110467 inclusive) airplanes certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To preclude excessive vibration in the flight control surfaces and possible loss of control of the airplane, accomplish the following: (a) Within the next 100 hours time-in-service (TIS) after the effective date of this AD, visually inspect for jamming or seizure of all bearings installed in the aileron trim tab bellcrank, actuator eyelets and the terminals of control rods for the elevator, rudder and aileron trim tab control systems in accordance with Section 2, "ACCOMPLISHMENT INSTRUCTIONS" of EMBRAER Service Bulletin (S/B) 110-27-036, Change 02, dated December 3, 1981. If a jammed or seized bearing is found, prior to further flight; (1) Remove and replace the defective part with a serviceable part of the same Part Number (P/N), or (2) Modify, inspect, replace and/or repair as required, (i) the aileron and rudder trim tab control systems in accordance with paragraph (c)(1) of this AD, and (ii) the elevator trim tab control system in accordance with paragraph (c)(2) of this AD. (b) When the modifications and actions specified in paragraph (a)(2) of this AD have been accomplished, the actions required by paragraph (c) of this AD are no longer required. (c) Within the next 250 hours TIS after the effective date of this AD, (1) Modify the aileron and rudder trim tab control systems in accordance with Section 2. "ACCOMPLISHMENT INSTRUCTIONS" of EMBRAER S/B 110-027-0060, Change 02, dated July 3, 1986. (2) Modify, inspect, replace and/or repair the elevator trim tab control system and elevator, (i) on airplanes without dual control rods as described in Section 2. "ACCOMPLISHMENT INSTRUCTIONS",paragraphs 2.1 through 2.1.18 of EMBRAER S/B 110-027-0068, dated April 9, 1986, or (ii) on airplanes with dual control rods installed as described in Section 2, "ACCOMPLISHMENT INSTRUCTIONS", paragraphs 2.2 through 2.2.10 of EMBRAER S/B 110-027-0068, dated April 9, 1986. (d) Aircraft may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (e) An equivalent method of compliance with this AD may be used if approved by the Manager, Atlanta Aircraft Certification Office, ACE-115A, FAA, Central Region, 1075 Inner Loop Road, College Park, Georgia 30337; Telephone (404) 763-7428. All persons affected by this directive may obtain copies of the documents referred to herein upon request to Empresa Brasileira de Aeronautica S.A. (EMBRAER), Post Office Box 343-CEP 12.200 Sao Jose dos Campos, Sao Paulo, Brazil; or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on January 22, 1987.
89-03-11: 89-03-11 BRITISH AEROSPACE: Amendment 39-6134. Applicability: Model BAC 1-11 200 and 400 series airplanes, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent a hazardous landing condition, accomplish the following: A. For Model BAC 1-11 200 series airplanes [pre-modification PM3070; Modification PM3070 with main support beams, Part No. EDO3-1001/2; or Modification PM5928]: At or prior to the accumulation of 50,000 landings or within the next 1,500 landings after April 25, 1988, (which is the effective date of AD 88-07-01, Amendment 39-5878), whichever occurs later, perform an eddy current or dye penetrant inspection for cracks in the rear pintle support beam, in accordance with paragraph 2.1.1 of the accomplishment instructions of British Aerospace BAC 1-11 Alert Service Bulletin 57-A-PM5896, Issue Number 4, dated February 17, 1988. Thereafter, repeat the inspection at intervals not to exceed 3,200 landings. NOTE: Airplanes that complied with paragraph A. of AD 88-07-01, Amendment 39-5878, are considered to have met the initial inspection requirements of this paragraph, and the inspection is to be repeated at intervals not to exceed 3,200 landings. B. For Model BAC 1-11 400 series airplanes [pre-modification PM3070; Modification PM3070 with main support beams, Part No. EDO3-1001/2; or Modification PM5928]: At or prior to the accumulation of 15,000 landings, or within 1,500 landings after March 17, 1986 (which is the effective date of AD 86-04-07, Amendment 39-5235), whichever occurs later, perform an eddy current or dye penetrant inspection for cracks in the rear pintle support beam, in accordance with paragraph 2.1.1 of the accomplishment instructions of British Aerospace BAC 1-11 Alert Service Bulletin 57-A-PM5896, Issue Number 4, dated February 17, 1988. Thereafter, repeat the inspection at intervals not to exceed 3,200 landings. NOTE: Airplanes that complied with paragraph A. of AD 86-04-07, Amendment 39-5235, are considered to have met the initial inspection requirements of this paragraph, and the inspection is to be repeated at intervals not to exceed 3,200 landings. C. For Model BAC 1-11 200 and 400 series airplanes [Modification PM3070 with main support beams, Part No. ENO3-1259/60]: At or prior to the accumulation of 7,500 landings or within 90 days after the effective date of this AD, whichever occurs later, perform an eddy current or dye penetrant inspection for cracks in the rear pintle support beam, in accordance with paragraph 2.1.1 of the accomplishment instruction of British Aerospace BAC 1-11 Alert Service Bulletin 57-A-PM5896, Issue Number 4, dated February 17, 1988. Thereafter, repeat the inspection at intervals not to exceed 3,200 landings. NOTE: Model BAC 1-11 200 series airplanes that complied with paragraph A. of AD 88-07-01, Amendment 39-5878; and Model BAC 1-11 400 series airplanes that complied with paragraph A. of AD 86-04-07, Amendment 39-5235; are considered to have met the initial inspection requirements of this paragraph, and the inspection is to be repeated at intervals not to exceed 3,200 landings. D. If cracks are discovered during the inspections required by paragraph A., B., or C., above, accomplish the following: 1. If cracks are less than or equal to 0.6 inch on pre-modification PM3070 airplanes, or less than or equal to 0.2 inch on post-modification PM3070 airplanes, repair or replace the cracked part prior to further flight, in accordance with paragraph 2.5.1 of the Accomplishment Instructions of British Aerospace BAC 1-11 Alert Service Bulletin 57-A- PM5896, Issue Number 4, dated February 17, 1988. If cracks are repaired in accordance with the service bulletin, continue to perform eddy current or dye penetrant inspections at intervals not to exceed 800 landings, providing that, in the case of pre-modification PM3070 airplanes, if cracks are presentin the aft half beam, cracked parts must be replaced prior to further flight. 2. If cracks exceed 0.6 inch in length on pre-modification PM3070 airplanes, or exceed 0.2 inch in length on post-modification PM3070 airplanes, repair or replace in a manner approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. If the part is repaired, further inspection requirements will be determined by the Manager, Standardization Branch, at the time of repair approval. If the main landing gear pintle support beam is replaced, the inspection must be repeated at intervals not to exceed 3,200 landings. E. Prior to June 17, 1986, (which is 90 days after the effective date of AD 86-04-07, Amendment 39-5235), inspect for damage of the toggle links' special bolt assembly, part number AB44A1275, in accordance with the accomplishment instructions of British Aerospace BAC 1-11 Alert Service Bulletin 32-A-PM5872, dated July 25, 1983. If the special bolt assembly is found damaged, replace with a serviceable part before further flight. NOTE: Airplanes that have complied with paragraph B. of AD 86-04-07, Amendment 39-5235, are considered to have met the requirements of this paragraph. F. 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, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Standardization Branch, ANM-113. G. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of the modifications required by this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copiesupon request to British Aerospace, Inc., Librarian for Service Bulletins, P. O. Box 17414, Dulles International Airport, Washington, D. C. 20041. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment supersedes AD 86-04-07 (Amendment 39-5235) which was effective March 17, 1986, and AD 88-07-01 (Amendment 39-5878) which was effective April 25, 1988. This amendment (39-6134, AD 89-03-11) becomes effective March 15, 1989.
2017-26-10: We are superseding Airworthiness Directive (AD) 2015-08-01, which applied to certain The Boeing Company Model 757 airplanes. AD 2015-08-01 required, depending on \n\n((Page 61676)) \n\nairplane configuration, installing new relays and bracket assemblies, inspecting to ensure that the new relays do not contact adjacent wire bundles, torqueing the bracket assembly installation nuts and ground stud nuts, retesting the bond resistance between the bracket assemblies and the terminal lugs on the ground studs, and doing related investigative and corrective actions if necessary. This AD does not retain any requirements, and instead requires deactivating the spoiler control module relays and capping and stowing the associated wiring on airplanes on which the actions required by AD 2015-08-01 have been done. This AD was prompted by a report of an uncommanded spoiler movement during flap configuration just before landing, on an airplane on which the actions required by AD 2015-08-01 hadbeen done. We are issuing this AD to address the unsafe condition on these products.
88-19-08: 88-19-08 BRITISH AEROSPACE: Amendment 39-6010. Applies to Model BAC 1-11 series airplanes equipped with R.F.D. AES-12B inflatable escape slides, identified in British Aerospace BAC 1-11 Service Bulletin 25-PM5943, Revision 1, dated May 8, 1987, certificated in any category. Compliance is required within 5 months after the effective date of this AD, unless previously accomplished. To prevent failure of the emergency escape slide deployment system, accomplish the following, unless previously accomplished: A. Modify the R.F.D. Type AES-12B emergency escape slide system in accordance with BAC 1-11 Service Bulletin 25-PM5943, Revision 1, dated May 8, 1987. B. 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, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments and then send it to the Manager, Standardization Branch, ANM-113. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of the modification 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 British Aerospace, Inc., Librarian, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment, 39-6010, becomes effective October 10, 1988.
61-20-02: 61-20-02 PIPER: Amdt. 336 Part 507 Federal Register September 20, 1961. Applies to Model PA-24 "250" Aircraft Serial Numbers 24-103 to 24-1629 Inclusive, Which Do Not Have a Reinforcing Plate Welded to the Stack in the Area Where the Rear Engine Cylinder Stack is Welded to the Exhaust Stack Assembly. Compliance required as indicated. Due to incidents of cracks occurring in the exhaust stack assembly, right side, P/N 21664-03, the following inspections and reinforcement must be accomplished: (a) Within 25 hours' time in service after effective date of this AD, remove the carburetor heat shroud assembly, P/N 21664-03, for any indication of cracks or deterioration particularly in the area where the rear engine cylinder exhaust stack is welded to the exhaust stack assembly. If evidence of cracks or deterioration is noted, the assembly must be replaced with a new assembly prior to further flight. The provisions of this paragraph shall be reaccomplished at intervals of50 hours' time in service until such time as the installation in paragraph (b) is accomplished. (b) Within 100 hours' time in service after initial compliance with paragraph (a), a clamp-on reinforcement Piper Kit No. 754396 or equivalent, shall be installed on the exhaust stack assembly, P/N 21664-03. After installation of the clamp-on reinforcement, the provisions of paragraph (a) are no longer applicable. (Piper Service Bulletin 202 dated May 22, 1961, applies to this subject.) This directive effective October 20, 1961.
61-16-01: 61-16-01 AERONCA: Amdt. 320 Part 507 Federal Register August 8, 1961. Applies to All Model 15 Series Aircraft. Compliance required as indicated. As a result of cracks found on the wing lift strut fittings, P/N 5-463-2, the following inspection shall be accomplished. For those aircraft with 1,000 or more hours' time in service, the inspection shall be accomplished within 25 hours' time in service after the effective date of this AD, unless accomplished within the last 25 hours' time in service, and at each periodic inspection thereafter. For those aircraft with less than 1,000 hours' time in service, the inspection shall be accomplished at the next periodic inspection after the effective date of this AD, and at each periodic inspection thereafter. (a) Detach the wing lift strut assembly and remove the upper and lower 1 1/2-inch diameter, 2024 tubular fittings, P/N 5-463-2. Clean, remove the paint and polish lightly with crocus cloth so as to increase the contrastfor inspection. Using the dye penetrant method and a 10-power glass, or equivalent, inspect for cracks on the inside and outside of the fitting surfaces, particularly in the area of the bolt holes and 90 degrees to the centerline of the bolt holes. All cracked tubes and all tubes with an outside diameter exceeding 1.515 inches at any point must be replaced prior to further flight. (b) Fittings being reinstalled shall be finished with the following or equivalent: Two coats of zinc chromate primer on the inside and outside of the fitting and finish enamel on the outside of the fitting. The 1/4-inch bolts attaching the fitting to the strut and the 1/2-inch bolt attaching the strut assembly to the airplane shall be torqued only enough to bring the respective parts into contact. The fitting must not be forced out of round. (c) When it is necessary to install new fittings, Aeronca P/N 5-463-2, Champion Aircraft Corporation P/N 1-9280, Prentice Aircraft, Inc. P/N 61-16-1, orFAA approved equivalent may be used. Existing AN 4 bolts, if not damaged, may be reinstalled in holes that match. No elongation of old holes in struts is permissible. (d) The following rework may be used if the holes in new fittings and old struts do not match to allow proper installation. AN5 bolts may be installed in all holes or in any combination. (1) If holes do not match, line-drill the assembled strut and fitting (19/64 inch) and ream to accept an AN 5 bolt. (.3125 inch minimum, .3135 inch maximum diameter.) In the event that portable equipment must be used, drill the hole 19/64 inch and accomplish the final sizing with a tapered hand reamer to assure alignment. (2) Install AN 5-22A bolts each with two AN 960-516 washers, and AN 365-524 nuts in holes that require rework. (e) When Prentice Aircraft, Inc., P/N 61-16-1 or FAA approved equivalent is installed, the repetitive inspections required by this directive may be discontinued. (Aeronca Service Helps and Hints No. 59 and Champion Aircraft Corporation Service Helps and Hints No. 59 cover the same subject.) This directive effective August 18, 1961. Revised November 28, 1961. Revised April 10, 1962.
61-16-07: 61-16-07 SUD AVIATION: Amdt. 321 Part 507 Federal Register August 5, 1961. Applies to All Alouette II SE 3130 Helicopters. Compliance required as indicated. As a result of two cases of cracks in the main rotor hubs, all main rotor hubs P/N 3130. S12.20.001 must be retired from service upon accumulation of 660 hours' time in service, except that main rotor hubs which have already accumulated 650 or more hours' time in service on the effective date of this directive must be retired from service within the next 10 hours' time in service. (Sud Aviation Helicopters Service Alouette II SE 3130 No. AL 12.11.204 covers this subject.) This directive effective August 5, 1961.