Results
2009-21-06: We are adopting a new airworthiness directive (AD) for the products listed above. 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: A recent incident has been reported with a Dornier 328-100 aeroplane, where the right-hand (RH) power lever jammed in flight- idle position during the landing roll-out. The aeroplane was stopped by excessive braking. The investigation by the operator revealed that the cockpit door locking device * * * had fallen off the RH cockpit wall and blocked the RH power/condition lever pulley/cable cluster below the door. * * * This condition, if not corrected, could cause interference with the engine- and/or flight control cables, possibly resulting in reduced control of the aeroplane. * * * * * We are issuing this AD to require actions to correctthe unsafe condition on these products.
52-19-01: 52-19-01 CURTISS-WRIGHT: Applies to all Models of C-46 Series airplanes used in passenger operation under the provisions of Parts 41, 42, and 61 of the Civil Air Regulations as specified in 41.20(f), 61.31(b), and Amendment 42-2 dated November 1, 1946. Item (5) of this Directive must be accomplished by November 1, 1952. The other items should be accomplished as soon as possible but not later than April 1, 1953. (This note pertains only to the powerplant fire protection aspects of the above Regulations. AD's 49-18-01 and 49-25-01 cover the fire prevention for the cabin heater installation, and for the baggage and cargo compartments of the airplane respectively.) Recent experience with the C-46 aircraft in passenger operation has brought to light several instances in which the adequacy of the powerplant fire protection installation provided is questionable. This Directive, which cancels and supersedes AD 49-19-1, therefore is intended to correct those installations wherein such inadequacies may exist. (1) Shutoff Valves. Install fluid shutoff valves, which may be opened and closed in flight, aft of the firewall in all fuel, oil, and hydraulic lines. USAF Technical Order 01-25LA-190 covers this same subject. If propeller or carburetor anti-icing systems are employed and use alcohol or other flammable fluids as the anti-icing medium, the systems described in USAF T.O. AN 01-25LA-2, pages 458-464, are satisfactory except that shutoff valves or a selector valve which can be opened and closed in flight must be provided aft of the firewall, to shut off the flow to either engine. The system should be such as to shut off the pump automatically, or otherwise guard against hazardous pressures, when the flow to both nacelles is stopped. (No shutoff valve will be required for the feathering pump oil lines, see section (3) below.) (2) Engine Firewalls. Engine firewalls must be rendered fireproof by adequately sealing all openings such as the filtered air duct opening, the oil cooler control rod and filtered air control rod openings, other powerplant control openings, holes through the firewall for electric conduits, and any other firewall openings. (3) Propeller Feathering Pump Installation. The portion of the propeller feathering oil line forward of the firewall between the firewall and the pump shall be of steel or other fireproof material. The line between the pump and governor shall be of fire resistant material with coupled hose assemblies used in any flexible connections. Electrical conduit for the pump motor and other electrical components forward of the firewall which are essential for propeller feathering shall be fire resistant or protected in a manner to render them fire resistant. The feathering pump can be considered an adequate means of shutting off the flow of oil in the feathering line. (4) Fire Extinguisher System. (a) Carbon Dioxide Quantity and Rate. The fire extinguisher system must be improvedto provide at least 35# carbon dioxide per shot which must be discharged into the nacelle at a rate of not less than 17 1/2#/sec (the rate of discharge will be affected by the number of bottles, the discharge valve sizes, the line sizes, and the nozzle area). The portion of the two- shot fire extinguisher system which is described in USAF Technical Order 01-25LA-205 is satisfactory. (Note: Civil Aeronautics Board Draft Release No. 52-15, proposes to require a two- shot fire extinguisher system in these airplanes in the near future. Operators, therefore, may consider it more practical to accomplish both changes at the same time. (b) Distribution System. The piping and nozzle arrangements shall be such as to spray the bulk of the discharge in the power section with a smaller amount (approximately 7 percent) being sprayed over the oil cooler. In this regard, it will be acceptable, if desired, to split the power section portion of the discharge so as to spray approximately one-thirdof it from 5 nozzles located around the upper half of the engine mounting ring. (c) Nozzle Location and Spray Pattern. The location of the nozzles and pattern of their spray are important for effective fire extinguishing and shall be substantially as follows: 1. For the power section, a nozzle shall be located at the rear and the base of each cylinder, discharging the carbon dioxide in a fan spray radially away from the crankcase. The provisions of USAF Technical Order 01-25LA-162 pertaining to replacement of engine cylinder fire extinguisher nozzle brackets must be accomplished. 2. For the oil cooler, a nozzle or nozzles shall be located above the forward end of the cooler directing fan sprays down and forward on each side of the cooler and duct. 3. The five nozzles located around the upper half of the mounting ring, if used, shall be such as to direct a fan spray radially outward as well as a fan spray radially inward over the accessories. (5) Fire Detectors. The Fenwal continuous type fire detectors, which were originally provided must be removed and replaced with unit or continuous type fire detectors conforming with FAA Technical Standard Order, TSO-C11 or TSO-C11a. If unit type detectors are used, they shall be spaced as specified below. Continuous type detectors, if used, shall be so installed as to provide equivalent coverage. (a) Engine Nacelles. Fire detectors, spaced not over 7 inches apart, shall be installed on the lower half of the forward side of the firewall at its outer periphery, and along the horizontal diameter. (b) Engine Mount Ring and Oil Cooler Supports. Additional fire detectors, spaced not over 18 inches apart, shall be provided for the upper two-thirds of the engine mount ring. Also, a fire detector shall be installed on each oil cooler support approximately 2 to 3 inches above the oil cooler. (c) Warning Light Covers. Fire-warning-light covers or shutters which are capable of dimming or shutting off the light entirely, must be removed. (6) Engine Compartment Lines. The following lines carrying inflammable fluids or vapors in the engine compartment shall be fire resistant and items (a) through (g) inclusive shall also have fireproof firewall fittings. Flexible connections in lines attached to the engine or subject to relative motion or pressure shall employ fire resistant coupled hose assemblies: (a) carburetor bleed back lines, (b) cabin heater fuel lines, (c) oil dilution lines, (d) fuel pressure transmitter lines, (e) oil pressure transmitter lines, (f) manifold pressure lines, (g) all other hydraulic oil lines, (h) all engine fuel lines, (i) engine primer lines, (j) engine breather lines, (k) engine supercharger drain lines, (l) oil separator return lines, (m) vacuum system pressure lines, (n) all main oil lines, (o) engine oil cooler lines, (p) hydraulic pump drain lines, (q) exhaust collector drain lines, (r) oil tank vent lines, (s) fuel pump drainlines. The fire extinguisher distribution tubing and fittings ahead of the firewall must be of steel or other fireproof material. Flexible connections in the distribution tubing ahead of the firewall must be at least of fire resistant construction, and shall use coupled hose assemblies rather than hose clamp connections. (7) Airplane Flight Manual. Appropriate changes to the airplane flight manual shall be prepared to cover emergency procedures associated with the above changes. (8) More detailed information on methods of complying with this Directive is being furnished to CAA Agents.
77-02-03: 77-02-03 GENERAL ELECTRIC COMPANY: Amendment 39-2809. Applies to all General Electric Company CT64-820-4 model turboprop engines. Compliance required within the next 100 hours time in service after the effective date of this AD, unless already accomplished. To prevent the possibility of overstressing the Stage 4 turbine buckets, the engine output shaft overspeed limit is reduced from 1512 rpm to 1450 rpm. The original limit is given in Note 2 of Type Certificate Data Sheet E13EA-10, dated February 15, 1976. This amendment becomes effective February 9, 1977.
2006-18-06: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Model A318, A319, A320, and A321 airplanes. This AD requires revising the Limitations section of the airplane flight manual (AFM); performing a one-time hardness test of certain ribs of the left- and right-hand engine pylons, as applicable, which would terminate the AFM limitations; and performing related corrective actions if necessary. This AD results from a report that certain stainless steel ribs installed in the engine pylon may not have been heat-treated during manufacture, which could result in significantly reduced structural integrity of the pylon. We are issuing this AD to detect and correct reduced structural integrity of the engine pylon, which could lead to separation of the engine from the airplane.
88-01-12: 88-01-12 BRITISH AEROSPACE: Amendment 39-5824. Applies to all Model BAe 125-800 airplanes certificated in any category. Compliance required within 90 days after the effective date of this AD, unless previously accomplished. To reduce the risk of leakage from fuel, oxygen, and hydraulic systems and the resulting fire and functional risks that such leaks present, accomplish the following: A. Inspect the plumbing in the area of the ventral fuel tanks, and repair as necessary, in accordance with British Aerospace Service Bulletin 53-60, dated November 8, 1985. 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. 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 inspections and/or modifications required by this AD. All persons affected by this directive who have not already received the appropriate service document from the manufacturer may obtain copies upon request to British Aerospace, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective February 19, 1988.
88-06-05: 88-06-05 SUD AVIATION: Amendment 39-5873. Applies to Sud Aviation Caravelle series airplanes, listed in Sud Aviation Caravelle Service Bulletin No. 57-56, Revision 2, dated February 28, 1985, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent failure of the wing, accomplish the following: A. Within 100 landings after the effective date of this AD or prior to the accumulation of the number of landings specified in the schedule, below, whichever occurs later, accomplish the requirements of subparagraphs A.1. or A.2., below, as applicable. SCHEDULE I (1) 30,000 landings on airplanes whose maximum take-off weight * is 48 tons or less; (2) 20,000 landings on airplanes whose maximum take-off weight * exceeds 48 tons but is no more than 52 tons; (3) 10,000 landings on airplanes whose maximum take-off weight * exceeds 52 tons. * Maximum take-off weight as defined in the FAA-approved airplane flight manual. 1. For airplanes without external doubler, perform visual and nondestructive testing inspections (NDI) for cracking adjacent to wing rib No. 31 in accordance with Sud Aviation Caravelle Service Bulletin No. 57-56, Revision 2, dated February 28, 1985. 2. For airplanes with external doublers, perform visual and radiographic inspections for cracking adjacent to wing rib No. 31 in accordance with Sud Aviation Caravelle Service Bulletin No. 57-56, Revision 2, dated February 28, 1985. B. If no cracks are detected, repeat the inspections required by paragraph A., above, at the intervals specified in Sud Aviation Caravelle Service Bulletin No. 57-56, Revision 2, dated February 28, 1985, Figure 1, Wing Lower Surface Inspection Frequencies. C. Any cracks detected during the inspections required by paragraphs A. or B., above, must be repaired, prior to further flight, in accordance with Sud Aviation Caravelle Service Bulletin No. 57-56, Revision 2, dated February 28, 1985. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. E. 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 inspections and/or 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 copies upon request to Aerospatiale, 316 Rue de Bayonne, 31060 Toulouse Cedex 03, France. 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 becomes effective April 22, 1988.
90-20-07: 90-20-07 BRITISH AEROSPACE: Amendment 39-6729. Docket No. 90-NM-102-AD. Applicability: Model BAC 1-11 200 and 400 series airplanes, certificated in any category. Compliance: Required within 30 days after the effective date of this AD, unless previously accomplished. To provide auxiliary power unit (APU) air intake plenum overheat detection, accomplish the following: A. Install temperature sensors in the APU air intake plenum, in accordance with British Aerospace Alert Service Bulletin 49-A-PM5955, Issue 1, dated April 13, 1988. 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, 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. C. 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 British Aerospace PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. This amendment (39-6729, AD 90-20-07) becomes effective on October 23, 1990.
93-07-10: 93-07-10 MCDONNELL DOUGLAS HELICOPTER COMPANY AND HUGHES HELICOPTERS, INC.: Amendment 39-8542. Docket Number 92-ASW-42. Applicability: Model 369D, 369E, 369F, 369HE and 369HS series helicopters, equipped with Gajon Associates, LTD. (Viking Helicopters Limited) Supplemental Type Certificate (STC) No. SH1134EA external cargo container kit (baggage pod), with or without the auxiliary fuel system, certificated in any category. Compliance: Required within 10 days or 25 hours' time-in-service, whichever occurs first, after the effective date of this AD, unless accomplished previously. To prevent hazardous yaw oscillations during descents, which could result in loss of control of the helicopter, accomplish the following: (a) Install a durable placard on the instrument panel as close as is practical to the airspeed indicator and that is legible to the pilot that reads: BAGGAGE POD INSTALLED Vne-90 KIAS IN POWERED DESCENT (>1000 fpm) OR IN AUTOROTATION(b) Insert the following statement into the Operating Limitations Section of the flight manual supplement for the baggage pod: AIRSPEED LIMITS Vne is 90 KIAS in moderate rates of powered descent (greater than 1,000 fpm) or in autorotation with baggage pod installed. (c) Insert the following statement into the Emergency and Malfunction Procedures Section of the flight manual supplement for the baggage pod: ENGINE FAILURE AT HIGH CRUISE SPEED NOTE: At speeds in excess of 90 KIAS in stabilized moderate descents (greater than 1,000 fpm) or in autorotation, the lateral directional handling of the helicopter is degraded. Yaw oscillation may occur and persist, and there is a tendency for the pilot to overcontrol. (1) Adjust collective pitch according to altitude and airspeed to maintain rotor speed between 410 and 508 RPM. (2) Apply pedal pressure as necessary to control aircraft yaw. (3) Adjust cyclic control as necessary to reduce airspeed to90 KIAS or less as collective is lowered and stabilized autorotation is achieved. NOTE: See basic rotorcraft flight manual for recommended minimum rate of descent and maximum glide distance power-off speeds. (d) Compliance with Paragraphs (b) and (c) above may be accomplished by attaching a copy of the appropriate AD paragraphs to the Operating Limitations Section and Emergency Procedure Section of the flight manual supplement. (e) 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, New York Aircraft Certification Office, FAA, 181 South Franklin Avenue, room 202, Valley Stream, New York. Operators shall submit their requests through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, New York Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliancewith this AD, if any, may be obtained from the Manager, New York Aircraft Certification Office. (f) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the helicopter to a location where the requirements of this AD can be accomplished. (g) This amendment becomes effective on July 6, 1993.
94-24-01: This amendment supersedes Airworthiness Directive (AD) 87-04-04 and AD 89-16-02, which currently require the following on Jetstream Aircraft Limited (JAL) HP 137 Mk1, Jetstream series 200, and Jetstream Models 3101 and 3201 airplanes: repetitively inspecting the universal joints and universal rivets, replacing any damaged part, and limiting the in-service life of the torque tube shaft assembly. JAL has introduced an improved flap torque tube shaft assembly that includes universal joints that are not life limited. The Federal Aviation Administration's policy on aging commuter-class aircraft is to eliminate or, in certain instances, reduce the number of certain repetitive short-interval inspections when improved parts or modifications are available. This action requires installing this improved flap torque shaft assembly in place of the repetitive inspection and life limit requirements of the two existing AD's. The actions specified by this AD are intended to prevent failure of the flap torque assembly, which could result in asymmetric flap deployment and loss of control of the airplane.
85-22-05: 85-22-05 BEECH: Amendment 39-5146. Applies to Beech airplanes listed in Table I below, certificated in any category, upon accrual of five years in service. This AD does not apply to those airplanes in which bolts and nuts made of Inconel have been installed in the wing attachment joints that are specified in Table I below. 85-22-05 TABLE I BEECH MODEL (MILITARY MODEL) SERIAL NUMBER JOINTS (1) INSTRUCTION (2) 34C GP-1 and up LF,UF,UR,LR T-34C-1-0083 T-34C GL-1 and up LF,UF,UR,LR T-34C-0158R2 T-34C-1 GM-1 and up LF,UF,UR,LR T-34C-1-0083 50 H-1 thru H-11 LF* P/N 98-39006 50 (L-23A,U-8A) LH-1 thru LH-55 LF* See Note 3a B50 CH-12 thru CH-110 LF* P/N 98-39006 B50 (L-23B) LH-56 thru LH-95 LF* See Note 3b C50 CH-111 thru CH-360 LF* P/N 98-39006 D50 (L-23E, L-23G) DH-1 thru DH-154 LF* P/N 98-39006 D50A, D50B, D50C DH-155 thru DH-300 LF* P/N 98-39006 D50E, D50E-5990 DH-301 thru DH-347 LF* P/N 98-39006 E50 EH-1 thru EH-70 LF* P/N 98-39006 E50 (L-23D, U-8D) LH-96 and up LF* See Note 3b E50 (RL-23D, RU-8D) RLH-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) LHC-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) LHD-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) RLHE-1 and RLHE-2 LF* See Note 3b E50 (RL-23D, RU-8D) LHE-3 and up LF* See Note 3b F50 FH-71 thru FH-96 LF* P/N 98-39006 G50 GH-94 thru GH-119 LF* P/N 98-39006 H50 HH-120 thru HH-149 LF* P/N 98-39006 J50 JH-150 thru JH-176 LF* P/N 98-39006 60, A60, B60 P-4 and up LF, UF, LR P/N 60-590001- 25A13 65 LC-1 thru LC-239 LF* P/N 98-39006 65 (L-23F, U-8F) L-1 thru L-6 LF* See Note 3b 65 (L-23F, U-8F) LF-7 and up LF* See Note 3b A65, A65-8200 LC-240 thru LC-335 LF* P/N 98-39006 65-80, 65-A80, 65-A80-8800 LD-1 thru LD-269 LF* P/N 98-39006 65-B80 LD-270 and up LF* P/N 98-39006 65-88 LP-1 thru LP-47 LF* P/N 98-39006 65-A90-1 (U-21A, RU-21A, RU-21D, JU-21A, U-21G, RU-21H) LM-1 and up LF* See Note 3c 65-A90-2 (RU-21B) LS-1 and up LF* See Note 3c 65-A90-3 (RU-21C) LT-1 and up LF* See Note 3c 65-A90-4 (RU-21E, RU-21H) LU-1 and up LF* See Note 3c 70 LB-1 thru LB-35 LF* P/N 98-39006 65-90, 65-A90, B90, C90 LJ-1 thru LJ-993 LJ-995 thru LJ-1007 LJ-1009 thru LJ-1034 LJ-1037 & LJ-1039 thru LJ-1044 LF* LF* LF* LF* LF* P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 E90 LW-1 thru LW-347 LF* P/N 98-39006 F90 LA-2 thru LA-90 LA-92 thru LA-156 LA-158 thru LA-169 LA-171 thru LA-173 LA-175 thru LA-182 LA-185, LA-187 LA-189 thru LA-191 LA-193 thru LA-196 and LA-199 LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 H-90 (T-44A) LL-1 thru LL-18 LL-20 thru LL-40 LL-42 thru LL-48 LL-50 thru LL-61 LF LF LF LF T-44A-0049R1 T-44A-0049R1 T-44A-0049R1 T-44A-0049R1 99, 99A, 99A (FACH) U-1 thru U-49 and LF* P/N 98-39006 A99, A99A, & B99 U-51 thru U-164 LF* P/N 98-39006 C99 U-50 and U-165 thru U-179 U-181 thru U-184 U-186 thru U-192 U-194 thru U-196 LF,UF LF,UF LF,UF LF,UF LF,UF P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 100, A100 & A100A B-1 thru B-247 LF* P/N 98-39006 B100 BE-2 thru BE-131, and BE-135 LF* LF* P/N 98-39006 P/N 98-39006 A100-1 BB-2 thru BB-342 LF,UF P/N 98-39006 (RU-21J), 200, BB344 thru BB-983 LF,UF P/N 98-39006 and B200 BB-985 thru BB-1038 LF,UF P/N 98-39006 B200 BB-1040 thru BB-1045 BB-1047 thru BB-1049 BB-1053 thru BB-1078 & BB-1080 LF,UF LF,UF LF,UF LF,UF P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 200C, B200C BL-1 thru BL-51 LF,UF P/N 98-39006 (C-12F) BL-53, BL-55 LF,UF P/N 98-39006 200 CT BN-1 LF,UF P/N 98-39006 200 T BT-1 thru BT-22 LF,UF P/N 98-39006 A200 (C-12A, C-12C)BC-1 thru BC-75 BD-1 thru BD-30 LF,UF LF,UF P/N 98-39006 P/N 98-39006 A200C (UC-12B) BJ-1 thru BJ-47 LF,UF P/N 98-39006 A200CT (C-12D,FWC-12D) BP-1 thru BP-27 LF,UF P/N 98-39006 (*-See Note 4) NOTE 1: Wing attachment joints, on left and right sides of each airplane, are abbreviated as: LF=lower forward, UF=upper forward, UR=upper rear, LR=lower rear. NOTE 2: T-34C-1-0083-1, T-34C-0158 Rev. 2, and T-44A-0049 Rev. 1 are Beech Service Instructions. Cited Beech Manuals, and their earliest applicable revision dates are: PART NUMBER NAME DATE 60-590001-25 Maintenance Manual June 13, 1984 98-39006 Structural Inspection and Repair Manual December 20, 1984 NOTE 3: Apply the following portions of P/N 98-39006 manual even though applicability to military models is not shown within the P/N 98-39006 manual: NOTE MANUAL SECTION REFERENCE FIGURE BOLT P/N NUT P/N 3a 57-10-00 209 NAS495-14-27 EB-144 3b 57-11-00210 MS20014-29 EB-144 3c 57-13-00 212 LWB-14-32 FN22-1414 Compliance: Required initially, upon accrual of five years after first airworthiness certification or within 60 days after the effective date of this AD (whichever is later), unless already accomplished, and thereafter at intervals which do not exceed five years. To assure structural integrity of attachments of outer wing panels to the wing center section, use procedures in instructions identified in Table I of this AD to accomplish the following at each wing attachment joint that is specified for a particular airplane by Table I of this AD: (a) Remove each steel nut and each steel tension bolt. Use visual and magnetic particle methods to inspect the bolt and nut for cracking and corrosion in parent steel, and replace each bolt and nut found cracked or corroded. NOTE 4: In lower forward joints that are asterisked in Table I of this AD, while bolts are removed for accomplishment of Paragraph (a), above, it is recommended that inboard and outboard fittings be inspected, by a fluorescent penetrant method, for fatigue cracks in washer face areas of the fittings. For some of the asterisked joints, inspections of fittings are required by other AD's, but inspections of fittings are not required by this AD. (b) During reassembly of each joint, coat the bolt, nut, and adjacent parts with MIL- C-16173 Grade 2 corrosion preventative compound. (c) Within the next 150 hours of flight time, check joint tightness, and tighten as necessary. (d) Inject MIL-C-16173 Grade 2 corrosion preventative compound into a lubrication fitting on each barrel nut, (wherever a barrel nut is used) when joint tightness is checked per Paragraph (c), above, and thereafter at intervals which do not exceed one year. (e) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (f) For destructive or nondestructive examination and any specified return, each nut and each bolt that is replaced in response to this AD must be identified with the related years in service, joint, and airplane serial number and sent to FAA/AVN-112, Room 203, Aviation Records Building, Mike Monroney Aeronautical Center, 6500 South MacArthur Boulevard, Post Office Box 26460, Oklahoma City, Oklahoma 73125. Parts so sent will be destroyed if return to a specified address is not requested. Reporting requirements approved by OMB pursuant to clearance No. 2120 0056. (g) An equivalent means of compliance with this AD may be used if approved by the Manager, FAA, Wichita Aircraft Certification Office, Room 100, 1801 Airport Road, Wichita, Kansas 67209; telephone (316) 946-4400. All persons affected by this directive may obtain copies of the documents referred to herein upon request to Beechcraft Aero and Aviation Centers; Beech Aircraft Corporation, 9709 East Central, Post Office Box 85, Wichita, Kansas 67201, or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on November 12, 1985.
53-09-01: 53-09-01 PRATT & WHITNEY: Applies to R-2000 Series Engines Overhauled by San Antonio Air Depot Between July 1, 1952, and April 9, 1953. Compliance required prior to further carriage of passengers or cargo in aircraft with such engines installed. Aircraft may be ferried to base where inspection is to be conducted. Several operators of C-54 aircraft utilizing military overhauled engines, have experienced failures of link (knuckle) pins in the subject engines due to improper overhaul, inspection, or assembly procedures. The failures have occurred in comparatively low-time engines, and cracked pins have been found in engines with zero TSO. To preclude the possibility of further failures of this nature, engines falling within the category as noted must be dissassembled and the link pins magnetically inspected, eliminating any found with cracks, then reassembled in accordance with manufacturer's instructions, prior to further flight. Date of overhaul and identification ofoverhaul base stamped on exterior surface of engine nose housings.
77-10-01: 77-10-01 MCDONNELL DOUGLAS: Amendment 39-2895. Applies to DC-4 and C-54 series airplanes, certificated in all categories, using Hamilton Standard propellers. \n\n\tCompliance required by the first engine change after the effective date of this AD or October 1, 1977, whichever occurs earlier, unless previously accomplished in accordance with AD 55-15-03 as amended by Amendment 55-24 to Part 507. \n\n\tTo increase fire resistance integrity of the propeller feathering system against damage by a powerplant fire, all flexible hose components of propeller feathering lines forward of the firewall must be replaced with lines and fittings which will meet the fire resistance requirements of the hose assemblies specified in (a) through (f) herein. However, if the feathering lines in Zone I include a section of steel tubing, flexible hose assemblies located forward of the cylinders and connecting to the governor are not affected by this directive. \n\n\tThe following flexible hose assemblies are acceptable for use in this application: \n\n\t(a)\tResistoflex SSFR-3800-10 hose assemblies. \n\n\t(b)\tAeroquip 680-10S hose assemblies with Aeroquip 304 protective sleeves over end fittings (Aeroquip Assembly P/N 304000). \n\n\t(c)\tAeroquip 309009 hose assemblies. \n\n\t(d)\tAeroquip 309009-8S hose assemblies (where feathering system requires this size). \n\n\t(e)\tAeroquip 634000-8 or -10, as applicable, hose assemblies. \n\n\t(f)\tHose assemblies that fully comply with FAR 37.140 (TSO-C42) and have a pressure rating equal to or greater than that of the propeller feathering system installed on the airplane. \n\n\t(g)\tEquivalent hose assemblies or other means of compliance may be used when approved by the Chief Aircraft Engineering Division, FAA Western Region. \n\n\tSpecial flight permits may be issued in accordance with FAR's 21.97 and 21.199 to operate the airplane to a base for the accomplishment of this AD. \n\n\tThis supersedes AD 55-15-03, as amended by Amendment 55-24 to Part 507. \n\n\tThis Amendment becomes effective June 16, 1977.
96-16-03: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320-200 series airplanes, that requires modification of the shock absorber sub-assembly of the main landing gear (MLG). This amendment is prompted by reports of internal damage to the shock absorber sub-assembly due to loose screws in the upper bearing dowels. The actions specified by this AD are intended to prevent such damage, which could result in the overextension of the shock absorber and failure of the torque link. This situation may lead to the inability of the MLG to retract and subsequent collapse of the MLG.
2009-25-10: This amendment adopts a new airworthiness directive (AD) for the Sikorsky Model S-92A helicopters. This action requires a one-time visual inspection of the main gearbox (MGB) lube system filter assembly for oil filter damage. This action also requires if either the primary or secondary oil filter is damaged, replacing both filters, all packings, and the studs before further flight. This AD also requires replacing the oil filter bowl within 30 days after replacing a damaged filter and a daily leak inspection for an oil leak (no oil leaks allowed) during that 30-day interim period. This amendment is prompted by three reports of damaged oil filters or packings resulting from installing the filter assembly with an oversized packing possibly because of incorrect part numbers in the maintenance manual. Based on a previous accident investigation, failure of the oil filter bowl or mounting studs can result in sudden and complete loss of oil from the MGB. The actions specified in thisAD are intended to prevent complete loss of oil from the MGB, failure of the MGB, and subsequent loss of control of the helicopter.
90-04-09: 90-04-09 BRITISH AEROSPACE: Amendment 39-6511. Docket No. 89-NM-216-AD. Applicability: Model BAe 146-200A and -300A series airplanes, as listed in British Aerospace Service Bulletin 53-84-00737D, Revision 1, dated August 22, 1989, certificated in any category. Compliance: Required prior to the accumulation of 3,000 landings since new, or within 30 days after the effective date of this AD, whichever occurs later, unless previously accomplished. To prevent reduced structural integrity of the fuselage, accomplish the following: A. Modify the fuselage rear section by adding eight rivets to the Stringer 21P end termination area, in accordance with British Aerospace Service Bulletin 53-84-00737D, Revision 1, dated August 22, 1989. 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 will either concur or comment 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 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 British Aerospace, 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, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6511, AD 90-04-09) becomes effective on March 19, 1990.
71-09-04: 71-09-04 BOEING: Amdt. 39-1199. Applies to Model 727 airplanes listed in Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revisions. \n\tCompliance required as indicated. \n\tTo detect cracking of the main landing gear actuator beam, accomplish the following: \n\tA.\tUnless already accomplished within the last 300 hours' time in service preceding the effective date of this AD, within the next 300 hours' time in service after the effective date of this AD or prior to the accumulation of 5300 hours' time in service, whichever occurs later, accomplish one of the following: \n\t\t1.\tVisually inspect the main landing gear actuator beam for any evidence of cracking in accordance with Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revision, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat the visual inspection at intervals not to exceed 600 hours' time in service, or \n\t\t2.\tUltrasonically inspect the main landing gear actuator beam for cracks in accordance with Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revision, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat the ultrasonic inspection at intervals not to exceed 1500 hours' time in service. \n\tB.\tIf cracks are found replace the beam with a serviceable beam. \n\tC.\tWhere records maintained by the operator are such as will permit a clear determination of the number of hours' time in service accumulated by the main landing gear actuator beam, P/N 65-17658-11, installed on the airplane, the inspection times prescribed by this AD may be applied to the beam rather than to the airplane. \n\tD.\tInspections prescribed by this AD do not apply to new replacement beams, P/N 65- 57153-3, or to P/N 65-17658-11 beam until 5300 hours' time in service is reached. \n\tE.\tAirplanes having cracked main landing gear actuator beams which require replacement under this AD may, in accordance with FAR 21.197, be flown with the landing gear extended to a base where the replacement can be accomplished. \n\tThis AD becomes effective May 28, 1971.
90-23-17: 90-23-17 BRITISH AEROSPACE: Amendment 39-6800. Docket No. 90-NM-122-AD. Applicability: Model DH.125-1A series airplanes, equipped with Hawker Siddeley Dynamics Air Conditioning System and Rolls Royce Viper Engines, certified in any category. Compliance: Required as indicated, unless previously accomplished. To prevent chafing between the air conditioning duct and the rear pressure bulkhead, and subsequent rapid decompression of the airplane, accomplish the following: A. Within 30 days after the effective date of this AD, perform a detailed visual inspection for chafing on the aft face of the rear pressure bulkhead, in accordance with the Accomplishment Instructions of British Aerospace Service Bulletin 53-71, dated November 1, 1989. B. If defects are found, prior to further flight, perform a dye penetrant inspection to detect cracks in the vicinity of the affected area; and perform a dial test indicator measurement to determine the depth of damage in therear pressure bulkhead, in accordance with British Aerospace Service Bulletin 53-71, dated November 1,1989. 1. If the damage to the rear pressure bulkhead is less than 0.003 inch deep, prior to further flight, carefully blend out, polish, and then restore protective treatment in accordance with the service bulletin. 2. If the damage to the rear pressure bulkhead is greater than 0.003 but less than 0.010 inch deep, within 100 landings, repair in accordance with Appendix B of the Service Bulletin. 3. If the damage to the rear pressure bulkhead is greater than 0.010 inch deep, prior to further flight, repair in accordance with Appendix B of the Service Bulletin. C. Within 30 days after the effective date of this AD, adjust the clearance between the air conditioning duct clamp and the rear pressure bulkhead so there is at least a 3/4-inch clearance. This can be accomplished by rotating and adjusting the duct position. D. 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 Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Standardization Branch, ANM-113. E. 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 British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue, S.W., Renton, Washington. This amendment (39-6800, AD 90-23-17) becomes effective on December 11, 1990.
77-20-06: 77-20-06 PILATUS AIRCRAFT, LTD. AND FAIRCHILD HILLER: Amendment 39-3050. Applies to Model PC-6 airplanes (all variants) manufactured by Pilatus Aircraft, Ltd., up through S/N 724 and to Model PC- 6 airplanes (all variants) manufactured by Fairchild Hiller, S/N's 2001 through 2047, certificated in all categories. Compliance is required within the next 25 hours time in service after the effective date of this AD, unless already accomplished within the last 75 hours time in service, and thereafter at intervals not to exceed 100 hours time in service from the last inspection, until the conditions of paragraph (c) are met. (a) To prevent a hazardous degree of corrosion from developing inside the wing struts, accomplish the following: (1) Visually inspect the internal surface of each wing strut for corrosion in accordance with paragraph 2.1 of Pilatus Aircraft Ltd., Service Bulletin No. 105, dated May 1971, (hereinafter S.B. No. 105) for Pilatus manufactured airplanes or paragraph2A of Fairchild Hiller Service Bulletin PC6-57-3, dated July 15, 1971 (hereinafter S.B. PC6-57-3), for Fairchild Hiller airplanes, or an FAA-approved equivalent. (2) If only light corrosion (corrosion which has not caused surface blistering) is found during an inspection required by paragraph (a)(1) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, remove the corrosion from, and apply an anticorrosive treatment to, the inside of the wing strut in accordance with paragraph 2.3 of S.B. No. 105 or paragraph 2D of S.B. PC6-57-3, as applicable, or an FAA-approved equivalent. (3) If corrosion is found during an inspection required by parargraph (a)(1) of this AD which has resulted in exceeding the limits prescribed in paragraph (a)(2), within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, replace the wing strut with a serviceable strut of the same part number that has had anticorrosive treatment applied to the inside surface in accordance with paragraph 2.3 of S.B. No. 105 or paragraph 2D of S.B. PC6-57-3, as applicable, or an FAA- approved equivalent. (b) For Pilatus Aircraft, Ltd., Model PC-6 airplanes, S/N's 338 through 701, to prevent a hazardous degree of corrosion from developing on the wing strut attachment brackets, acocmplish the following: (1) Visually inspect each wing strut attachment bracket for corrosion in accordance with paragraph 2.1 of Pilatus Aircraft, Ltd., Service Bulletin No. 93, dated June 1969, (hereinafter S.B. NO. 93) or an FAA-approved equivalent. (2) If only light corrosion (corrosion which has caused 2% to 10% reduction in cross-section per paragraph 2.2 of S.B. No. 93) is found during an inspection required by paragraph (b)(1) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, remove the corrosion and apply an anti-corrosive treatment to the wing strut attachment bracket in accordance with paragraph 2.4 of S.B. No. 93, and reinstall the bracket in accordance with paragraph 2.5.1 of S.B. No. 93 or an FAA-approved equivalent. (3) If corrosion is found during an inspection required by paragraph (b)(1) of this AD which has resulted in exceeding the limits prescribed in paragraph (b)(2) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, replace the wing strut attachment bracket with a new wing strut attachment bracket (P/N 111.35.06.055 (left) or 111.35.06.056 (right)) in accordance with paragraph 2.5.2 of S.B. No. 93 or an FAA-approved equivalent. (c) The inspections required by paragraph (a)(1) or (b)(1) of this AD may be discontinued, in accordance with the following: (1) Inspection of the wing strut when the strut has had light corrosion removed and has had the anticorrosive treatment in accordance with paragraph (a)(2) or when the strut has been replaced in accordance with paragraph (a)(3) of this AD. (2) Inspection of the wing strut attachment bracket when the bracket has had light corrosion removed and has had the anticorrosive treatment in accordance with paragraph (b)(2), or when the bracket has been replaced in accordance with paragraph (b)(3) of this AD. This amendment becomes effective November 3, 1977.
89-25-11: 89-25-11 PRATT & WHITNEY: Amendment 39-6352. Applicability: Pratt & Whitney (PW) JT8D-9, -9A, -11, -15, -15A, -17, -17A, -17R, and -17AR turbofan engines. Compliance: Required as indicated, unless already accomplished. To prevent fire, inflight shutdown, engine cowl release, or airframe damage associated with a first stage fan blade liberation, remove certain first stage fan blade retaining plates in accordance with the Accomplishment Instructions of PW Alert Service Bulletin (ASB) 5841, dated February 15, 1989, and replace with an improved design retaining plate as follows: (a) Replace retaining plate Part Numbers (P/N) 520451, 616645, or 639616 with retaining plate P/N 803996 at the next shop visit but no later than: (1) Two years or 4,000 hours in service after the effective date of this AD, whichever occurs later, for wing-mounted engines. (2) Four years or 8,000 hours in service after the effective date of this AD, whichever occurs later, forfuselage-mounted engines. (b) Replace retaining plate P/N 760297, 793935, or 802710 with retaining plate P/N 803996 at the next shop visit but no later than five years or 10,000 hours in service after the effective date of this AD, whichever occurs later. NOTES: (1) A shop visit occurs following engine removal where the subsequent engine maintenance entails the following: (a) Separation of a major engine flange (lettered or numbered), other than flanges mating with major sections of the nacelle or reverser. Separation of flanges purely for purposes of shipment, without subsequent internal maintenance, is not a "shop visit." (b) Removal of a disk, hub, or spool. (2) FAA approved first stage fan blade retaining plate designs may be used in lieu of P/N 803996 retaining plate as an alternate method of compliance. (c) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD may be accomplished. (d)Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD, may be approved by the Manager, Engine Certification Office, ANE-140, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. The inspection/replacement procedures shall be done in accordance with PW ASB 5841, dated February 15, 1989. 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 inspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts, 01803, or at the Office of the Federal Register, 1100 L Street NW,Room 8301, Washington, DC 20591. This amendment (39-6352, AD 89-25-11) becomes effective on January 15, 1990.
2009-25-06: We are superseding an existing airworthiness directive (AD) for the products listed above. 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: * * * [T]he FAA has published SFAR 88 (Special Federal Aviation Regulation 88). * * * Under this regulation, all holders of type certificates for passenger transport aeroplane * * * are required to conduct a design review against explosion risks. One of the consequences of the Airbus design review is the modification of the fuel pump wiring to provide protection against chafing of the fuel pump cables. This condition, if not corrected, could generate short circuits leading to fuel pump failure and arcing. These could become a potential ignition source inside the fuel tank which, in combination with flammable fuel vapours (if present), could result in a fuel tank explosion and consequent loss of the aeroplane. To address this unsafe condition, EASA [European Aviation Safety Agency] issued AD 2007-0066 that required this modification [of the fuel pump against short circuit] in accordance with Airbus Service Bulletin (SB) A300-24-0103 Revision 01. Airbus subsequently introduced an additional modification of the electrical wiring of the outer fuel pump and the landing lights of the left (LH) and the right (RH) side in Revision 02 of the SB A300-24-0103, leading to the issuance of EASA AD 2008-0188 which superseded EASA AD 2007-0066 and required the additional work. More recently, Airbus introduced some additional protection to routes 1P and 2P harnesses in zone 571 and 671 of the aeroplane. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
2009-20-12: We are adopting a new airworthiness directive (AD) for certain Boeing Model 747 airplanes identified above. This AD requires replacing the inboard trailing edge (TE) flap transmission carbon disk no-back brakes with skewed roller no-back brakes at the TE flap transmission, positions 4 and 5. This AD results from reports of the inboard TE flaps blowing back due to the failure of a transmission carbon disk no-back brake. The no-back brake did not hold the TE flaps in the commanded position. We are issuing this AD to prevent a decrease of the aerodynamic controllability of the airplane, which could adversely affect the airplane's continued safe flight and landing.
91-08-05: 91-08-05 CONSTRUZIONI AERONAUTICHE GIOVANNI AGUSTA: Amendment 39- 6959. Docket No. 91-ASW-10. Applicability: All Model A109A and A109AII helicopters, certificated in any category, with main rotor blades, part number (P/N) 109-0103-01 (all dash numbers), with serial numbers 378 through and including 1519, installed. Compliance: Required prior to further flight for blades with 300 or more hours total time in service on the effective date of this AD, and thereafter, for the remaining blades upon attaining 300 hours time in service, unless already accomplished. To prevent possible fatigue failure of the main rotor blade and subsequent loss of the helicopter, accomplish the following: (a) Before further flight, inspect each main rotor (M/R) blade for cracks using a dye penetrant or equivalent inspection method and again within the next 10 hours time in service as follows: (1) Identify the designated area of the M/R blade in accordance with the requirements of paragraph 6.1.2 of Agusta Service Document BTT No. 109-6, Rev. B, dated November 2, 1989. (2) Prepare the designated areas and conduct dye penetrant inspections in accordance with paragraphs 6.1.3 through 6.1.6 of Agusta BTT No. 109-6, Rev. B, dated November 2, 1989. (3) Remove the blade from the helicopter for the initial inspection. (4) After completing the inspection, protect the designated area in accordance with Agusta BTT No. 109-6, Rev. B, dated November 2, 1989, paragraph 6.1.8, or other FAA-approved methods. NOTE: If an eddy current inspection, in accordance with Agusta BTT No. 109-6, Rev. B, dated November 2, 1989, is accomplished before further flight, the requirements of this paragraph are satisfied. (b) Within the next 20 hours time in service after the effective date of this AD, and thereafter at intervals not to exceed 25 hours time in service, inspect each M/R blade in the designated area identified in paragraph (a) for cracks using an eddy current inspection method in accordance with the requirements of paragraph 6.2 and subparagraphs 6.2.1 through 6.2.7.2 of Agusta BTT No. 109-6, Rev. B, dated November 2, 1989, as follows: (1) Remove the blade from the helicopter for the inspections. (2) After completing the inspection, protect the designated area in accordance with Agusta BTT No. 109-6, Rev. B, dated November 2, 1989, paragraph 6.1.8, or other FAA-approved methods. (c) If a crack is detected during the inspection of paragraph (a) or (b) above, remove the blade from service and replace with an airworthy M/R blade. (d) In accordance with FAR Sections 21.197 and 21.199, the helicopter may be flown to a base where the inspections required by the AD may be accomplished. (e) An alternate method of compliance or adjustment of the compliance times, which provides an equivalent level of safety, may be used if approved by the Manager, Rotorcraft Standards Staff, FAA, Fort Worth, Texas 76193-0110, telephone (817) 624-5110. (f) Report cracks found to the manager identified in paragraph (e) within 10 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056.) (g) The inspection requirements of paragraphs (a) and (b) do not apply to serviceable blades which have been inspected and reidentified in accordance with Part II of Agusta Technical Bulletin No. 109-79, dated July 27, 1990. Serviceable blades will be reidentified by adding "T" after the serial number on the data plate. The inspection procedures shall be done in accordance with Agusta Technical Bulletin No. 109-6, Rev. B, dated November 2, 1989, and Agusta Technical Bulletin No. 109-79, dated July 27, 1990, which incorporates Report No. 109-02-79, dated July 15, 1990. The 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 Agusta Aerospace Corporation, 3050 Red Lion Road, Philadelphia, PA 19114. Copies may be inspected at the Office of the Assistant Chief Counsel, FAA, 4400 Blue Mound Road, Fort Worth, TX, or at the Office of the Federal Register, 1100 L Street, N.W., Room 8401, Washington, D.C. Airworthiness Directive 91-08-05, supersedes priority letter AD 89-23-08, issued November 2, 1989. This Amendment (39-6959, AD 91-08-05) becomes effective on May 7, 1991.
76-07-10: 76-07-10 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE: Amendment 39- 2573. Applies to Model SA341G helicopters, certificated in all categories. Compliance is required as indicated. To prevent possible inflight failure of main gear box (MGB) attachment, accomplish the following: a) Within the next 200 hours time in service after the effective date of this AD, unless already accomplished, remove, inspect, reinstall or rectify as prescribed, and reidentify the following components in accordance with the applicable provisions of the "Description" paragraph 1C of Gazelle Service Bulletin No. 01.02, as amended October 31, 1973, or an FAA- approved equivalent: (1) MGB forward and rear "A" frames, P/N 341A.38.1019.00, .02 and .04, and P/N 341A.38.1016.00, .02 and .04, respectively. (2) MGB forward and rear shackles, P/N 341A.38.1017.00 and .01, and P/N 341A.38.1079.00 and .01, respectively, and associated pins, P/N 341A.31.4169.20 and .21. b) Within the next 3000 hours time in service after the accomplishment of paragraph (a) of this ad or within the next 3000 hours time in service after the accomplishment of the "Description" paragraph 1C of Gazelle Service Bulletin No. 01.02, as amended October 31, 1973, whichever occurs sooner, and thereafter at intervals not to exceed 3000 hours time in service from time of replacement, replace the MGB forward and rear shackles, P/N's 341A.38.1017.00 and .01 and P/N's 341A.38.1079.00 and .01, respectively, with new shackles of the same part number. This supersedes Amendment 39-1629 (39 FR 9660), AD 73-09-05. This amendment becomes effective April 22, 1976.
2000-15-04: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 747-200 and -300 series airplanes, that currently requires various inspections and functional tests to detect discrepancies of the thrust reverser control and indication system, and correction of any discrepancy found. This amendment requires installation of a terminating modification, and repetitive functional tests of that installation, and repair, if necessary. This amendment is prompted by the results of a safety review of the thrust reverser systems on Model 747 series airplanes. The actions specified by this AD are intended to ensure the integrity of the fail safe features of the thrust reverser system by preventing possible failure modes in the thrust reverser control system that can result in inadvertent deployment of a thrust reverser during flight.
2009-24-21: The FAA is superseding an existing airworthiness directive (AD) that applies to all McDonnell Douglas Model DC-9-14, DC-9-15, and DC-9-15F airplanes; and McDonnell Douglas Model DC-9-20, DC-9-30, DC-9- 40, and DC-9-50 series airplanes. That AD currently requires repetitive inspections for cracks of the main landing gear (MLG) shock strut cylinder, and related investigative and corrective actions if necessary. This AD adds more work on airplanes that have main landing gear shock struts with certain identified part numbers. This AD results from two reports of a collapsed MLG and a report of cracks in two MLG cylinders. We are issuing this AD to detect and correct fatigue cracks in the shock strut cylinder of the MLG, which could result in a collapsed MLG during takeoff or landing, and possible reduced structural integrity of the airplane.