Results
74-23-06: 74-23-06 MCDONNELL DOUGLAS: Amendment 39-2005 as amended by Amendment 39-2999 is further amended by Amendment 39-3031. Applies to McDonnell Douglas DC-8 series airplanes, certificated in all categories, incorporating Pratt and Whitney JT3D engines. \n\n\tWithin 24 hours after receipt of this telegram, incorporate the operating limitations and procedures, set forth in paragraphs (1), (2), and (3), below, into the Douglas DC-8 FAA approved Airplane Flight Manual. Make appropriate notations on the log of pages. Operators shall promptly implement these limitations and procedures. \n\n\t(1)\tRevise Section I, Limitations, to include a new item relative to fuel management. \n\n\t\tFUEL BOOST AND/OR FEED PUMP OPERATION \n\n\tPrior to descent, except during landing, the main tank pumps must be in the boost and feed position. \n\n\t(2)\tRevise Section I, Limitations, to include a new item relative to engine operation at idle power. \n\n\t\tFLIGHT IDLE OPERATION \n\n\tInflight at altitudes above 6,000 feet MSL and at indicated airspeeds below 200 knots, a minimum N2 engine rotor speed of 62 percent must be maintained except during landing. \n\n\t(3)\tRevise Section II, Emergency Procedures, to include a new item relative to recovery from a condition where an engine(s) fail to accelerate in flight after the throttle levers are advanced from the idle settings: \n\n\t\tENGINE RESPONSE TO THROTTLE LEVER(S) \n\n\tIf the engine(s) fail to accelerate in flight after the throttle lever(s) are advanced, the following procedure should be used after it has been determined that a flameout has not occurred. This procedure need not be continued and affected systems may be reactivated if engine operation has been restored to normal. \n\t\n\nPHASE I - II\n\nThrottle Lever(s) affected engine(s)\nAdvance\nEngine Anti-Ice \nOff \nAirspeed\nIncrease (as practicable)\n\t\t \t \nPHASE III \n\nIgnition \nOverride/Both \nMain Tank Pumps \nBoost and Feed\nPneumatic Bleed (Affected Engine(s))\nReduce \nElectrical Load\nRemove Non-required \n\t\n\tNote: One reservoir feed pump may be inoperative provided: \n\n\t(a)\tThe affected reservoir feed pump is placarded inoperative at the feed pump switch position. \n\n\t(b)\tEstablished maintenance procedures for this item are followed. \n\n\t(c)\tSufficient fuel is carried in the associated tank to provide a minimum of 2000 pounds of additional fuel in excess of the fuel (including reserves) needed for the flight. \n\n\t(d)\t'FUEL LOADING AND MANAGEMENT' is in accordance with the FAA Approved Airplane Flight Manual. \n\n\tAmendment 39-2005 became effective November 14, 1974, for all persons except those to whom it was made effective immediately by telegram dated October 11, 1974. \n\n\tAmendment 39-2999 became effective August 10, 1977. \n\n\tThis amendment 39-3031 becomes effective September 15, 1977.
2011-20-08: This amendment adopts a new airworthiness directive (AD) for the specified Agusta model helicopters. This action requires inspecting certain modules and related connectors for corrosion. If there is corrosion on the connectors, this AD requires cleaning the connectors before further flight. If there is corrosion on a module, before further flight, this AD requires replacing the module with an airworthy module. This AD also requires modifying the Number 2 Modular Avionic Unit (MAU) ventilation duct. This amendment is prompted by some in- flight emergencies due to internal corrosion of the MAU circuit card assemblies. The actions specified in this AD are intended to detect corrosion of certain modules to prevent the display of misleading data to the flight crew, disengagement of the flight director modes of the autopilot or other alert system, increased workload of the flight crew, and subsequent loss of control of the helicopter.
97-06-12: This amendment supersedes two existing airworthiness directives (AD), applicable to British Aerospace Model BAe 146 and Avro 146-RJ series airplanes, that currently require inspections to detect cracking of the upper main fitting of the nose landing gear (NLG), and replacement or repair of cracked parts, if necessary. Those actions were prompted by reports of cracking in the main fittings of the NLG. This amendment requires that, for certain airplanes, the inspections be accomplished at reduced intervals. This amendment is prompted by the results of new analyses of the cracking that were conducted by the manufacturer of the NLG. The actions specified by this AD are intended to prevent failure of the main fitting, which could lead to collapse of the NLG during landing.
2022-02-12: The FAA is adopting a new airworthiness directive (AD) for all Leonardo S.p.a. Model AB139 and AW139 helicopters. This AD was prompted by the identification of certain parts needing maintenance actions, including life limits and maintenance tasks. This AD requires incorporating into maintenance records requirements (airworthiness limitations), as specified in a European Aviation Safety Agency (now European Union Aviation Safety Agency) (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2005-13-33: The FAA is adopting a new airworthiness directive (AD) for certain Airbus Model A300 B2 and B4 series airplanes. This AD requires modifying the wiring of the autopilot pitch torque limiter switch. This AD is prompted by several reports of pitch trim disconnect caused by insufficient length in the wiring to the pitch torque limiter lever. We are issuing this AD to prevent possible trim loss when the flightcrew tries to override the autopilot pitch control, which could result in uncontrolled flight of the airplane.
95-06-03: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 95-06-03 which was sent previously to all known U.S. owners and operators of Robinson Helicopter Company (Robinson) Model R22 helicopters by individual letters. This AD requires an inspection and modification of the main rotor (M/R) gearbox. This amendment is prompted by a report of an incident involving a Model R22 helicopter in which the two M/R mast spanner nuts (nuts) became loose, resulting in failure of the M/R mast support structure. The actions specified by this AD are intended to prevent M/R separation and subsequent loss of control of the helicopter.
95-12-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Aerospatiale Model ATR72 series airplanes. This action requires repetitive inspections to detect displacement of the rear hinge bush, and to detect cracking or rupture of the rear hinge pin on the main landing gear (MLG) leg; and the correction of any discrepancies. This amendment is prompted by a report of the failure of this hinge pin on an in-service airplane. The actions specified in this AD are intended to prevent failure of the hinge pin, which can lead to failure of the MLG leg or MLG attachment assembly.
2011-21-17: We are adopting a new airworthiness directive (AD) for all General Electric Company (GE) CT7-8A, CT7-8A1, CT7-8E, and CT7-8F5 turboshaft engines with a fuel filter differential pressure switch, part number (P/N) TD028VF0H7Y5 (part of the fuel filter assembly, P/N 4110T53P06) installed. This AD requires daily visual inspections of the fuel filter differential pressure switch for fuel leaks and for excessive cracking of the switch mounting flanges due to stress- corrosion. This AD also requires the installation of a collar kit over the fuel filter differential pressure switch as terminating action to the daily inspections. This AD was prompted by reports of 47 fuel filter differential pressure switches found with stress-corrosion cracking of the mounting flanges. We are issuing this AD to prevent unrecoverable in-flight engine shutdown, engine bay fire due to fuel leakage, and forced landing or accident.
97-06-09: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-300, -400, and -500 series airplanes. This AD requires replacing certain aileron/rudder trim control modules with an improved module that contains an improved rudder trim switch that precludes the problems of sticking associated with the existing switch. This amendment is prompted by reports of sticking conditions in the rudder trim switch. The actions specified by this AD are intended to prevent such sticking, which could result in uncommanded movement of the rudder and consequent deviation of the airplane from its set course.
69-26-06: 69-26-06 MCDONNELL DOUGLAS: Amendment 39-902. Supersedes Amendment 39-738, AD 69-06-03. Applies to all McDonnell Douglas Model DC-9 Series aircraft. \n\n\tCompliance required as indicated, unless already accomplished. \n\n\tTo prevent heat damage to the H.F. (if installed) and V.H.F. coaxial cables and other wiring located in the tail compartment of DC-9 Series aircraft, accomplish the following: \n\n\t(A)\tWithin the next 200 hours' time in service after the effective date of this AD, unless already accomplished, perform the following: \n\n\t\t(1)\tDetermine that the H.F. (if installed) and V.H.F. coaxial cables located in the tail compartment of the aircraft adjacent to the 8th stage bleed duct have not deteriorated due to excessive heat. The determination may be accomplished by the use of electrical tests such as fault finder pulse indications, reflectometer measurements, or X-ray inspections or a satisfactory equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. If the electrical tests indicate any coaxial cable impedance change in the areas of the 8th stage bleed duct or the APU exhaust shroud, or the X-ray inspections show physical change, such as noticeable drift of the center conductor, or any unsatisfactory condition in these areas, replace the damaged coaxial cables or repair the damaged areas of the cables by use of proper connectors and new coaxial sections, in conjunction with (2) and (3), below. In lieu of electrical testing of X-ray inspection an operator may replace the cables within this 200 hour period. \n\n\t\t(2)\tProvide maximum possible clearance (at least one inch) between the H.F. (if installed) and V.H.F. conduits, and the right engine 8th stage bleed duct by rotating the conduit clamps and reworking the spacers as necessary. NOTE: McDonnell Douglas DC-9 Alert Service Bulletin A23-24, dated February 21, 1969 describes this work. \n\n\t\t(3)\t(a)\tVisually inspect the APU exhaust shroud for any indications of overtemperature condition, such as shroud discoloration or exterior airframe paint discoloration around the shroud outlet. If the APU exhaust duct has been deformed or leaks, and continued use of the APU is desired, replace the duct in accordance with McDonnell Douglas S.B. 49-8, dated May 2, 1966, and Service Letters AOL-9 No. 74, dated February 6, 1967, and AOL-9 No. 139, dated September 29, 1967, or later FAA approved revisions, or an equivalent duct replacement approved by the Chief, Aircraft Engineering Division, FAA, Western Region. \n\n\t\t\t(b)\tInspect all wiring adjacent to the APU exhaust shroud for heat damage. Replace or repair to an airworthy condition all wiring found damaged. \n\n\t\t(4)\tPressure test the pneumatic duct installation in the DC-9 tail cone area in accordance with the DC-9 Maintenance Manual Temporary Revision 36-19, dated December 3, 1969, or the subsequent equivalent revision, or an equivalent pressure test approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t\t(5)\tSteps (1) and (3) above, must be repeated prior to any further IFR operation after every report of a pneumatic duct malfunction or an APU exhaust duct failure until (A)(7) or (B), below, has been accomplished. \n\n\t\t(6)\tStep (4), above, must be repeated whenever pneumatic duct maintenance is performed in the tail cone area. \n\n\t\t(7)\tSteps (1) through (5), above, need not be accomplished if the operator accomplishes Step B below, within 200 hours' time in service from the effective date of this AD. \n\n\t(B)\tWithin the next 2000 hours' time in service from the effective date of this AD, unless already accomplished, perform the following in accordance with McDonnell Douglas Service Bulletin 23-24, Rev. 2, dated June 23, 1969; S.B. 23-28, dated December 3, 1969, and S.B. 27-104, Rev. 2, dated April 15, 1969, or later FAA approved revisions, or an equivalent installation and modification approved by the Chief, Aircraft Engineering Division, FAA, Western Region: \n\n\t\t(1)\tInstall an insulation blanket on the 8th stage bleed duct adjacent to the HF and VHF coaxial cable conduits. \n\n\t\t(2)\tReroute the HF (if installed) and VHF coaxial cables, and the other wiring bundle (flight recorder, interphone wiring, etc.) away from the APU exhaust shroud area where they are now located. \n\n\t\t(3)\tReplace the sections of the polyethylene dielectric type VHF and HF (if installed) coaxial cables with sections of electrically equivalent polytetrafluoroethylene (teflon) dielectric type coaxial cables from just forward of the pressure dome feed-through throughout the tail compartment, or from between just aft of the pressure dome feed-through to just aft of the exhaust duct from the air condition pack heat exchangers, and apply PF105-700 glass fiber batt and CRS-102 silicon wrap heat insulation material over all exposed low temperature cable which is not installed in conduit. NOTE: No additional rerouting or repositioning, other than that specified inparagraph (A) (2) and (B)(2), is required. \n\n\t\t(4)\tAdd a metal heat shield between the eighth stage pneumatic duct and electrical wire bundle in the tail cone R.H. side just aft of pressure panel and forward of the eighth stage bleed duct. \n\n\t\t(5)\tReposition the wire harnesses FBC and DDC, containing overheat sensor wiring and APU generator control wiring located in the tailcone L.H. side to a new position more outboard of the wing ice protection duct. \n\n\tNOTE: Compliance with the coaxial cable separation and rerouting modification also provided in AOL No. 9-333, dated August 27, 1969, and Service Bulletin No. 23-28, dated December 3, 1969, is optional. \n\n\tThis AD supersedes amendment 39-738, (34 F.R. 5427) AD 69-6-3. \n\n\tThis amendment becomes effective on December 30, 1969.
89-10-04: 89-10-04 McDONNELL DOUGLAS: Amendment 39-6204. \n\n\tApplicability: Model DC-8 series airplanes, equipped with left (LH) or right (RH) main landing gear (MLG) attach fitting, P/N(s) 5611425-1 through -508, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent severe structural damage to the airplane during takeoff or landing due to stress corrosion failure of the MLG attach fittings, accomplish the following: \n\n\tA.\tWithin the next 12 calendar months after the effective date of this AD, unless already accomplished within the last 12 calendar months, and thereafter at intervals not to exceed 24 calendar months, except as provided below, perform a visual inspection of the MLG attach fittings for cracks at locations in accordance with Figure 1. of McDonnell Douglas DC-8 Service Bulletin 57-94, Revision 1, dated June 23, 1987 (hereafter referred to as the Service Bulletin). \n\n\tB.\tIf no cracks are found, apply LPS-3 corrosion inhibiting oil to the fitting in accordance with the Service Bulletin, and repeat inspections for cracks in accordance with paragraph A. of this AD. \n\n\tC.\tIf cracks are found, accomplish the following: \n\n\t\t1.\tIf cracks are located in area 1 or 3, as defined in Figure 1. of the Service Bulletin, before further flight, replace the fitting, P/N(s) 5611425-1, -2, -501, -502, -503, -504, -505, -506, -507, or -508, with respective P/N(s) 5893930-1, -2, -1, -2, -509, -510, -507, -508, -505, or -506. \n\n\t\t2.\tIf cracks are located in area 2, as defined in Figure 1. of the Service Bulletin, accomplish the following: \n\n\t\t\ta.\tIf cracks are within limits as prescribed by Figure 1. of the Service Bulletin, apply LPS-3 corrosion inhibiting oil to the fitting in accordance with the Service Bulletin, and repeat visual inspections for crack development at intervals not to exceed 7 calendar days, in accordance with the Service Bulletin. \n\n\t\t\tb.\tIf cracks exceed limits as prescribed by Figure 1. of the Service Bulletin, replace the fitting in accordance with paragraph C.1. of this AD before further flight. \n\n\tD.\tReplacement of both the LH and RH MLG attach fittings in accordance with paragraph C.1. of this AD constitutes terminating action for the inspection requirements of this AD. \n\n\tE.\tAn alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspection (PMI), who may add any comments and then send it to the Manager, Los Angeles Aircraft Certification Office. \n\n\tF.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director of Publications, C1-LOO (54-60). These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or at 3229 East Spring, Long Beach, California. \n\n\tThis amendment (39-6204, AD 89-10-04) becomes effective on May 29, 1989.
72-04-01: 72-04-01 BOEING: Amendment 39-1392 as amended by Amendment 39-1407. Applies to 747-100 and 747-200B Series airplanes. \n\tCompliance required as indicated. \n\tTo prevent unscheduled stabilizer trim and to maintain stabilizer control capability, accomplish the following: \n\t(a)\tFor airplanes incorporating stabilizer trim modules Boeing P/N 60B80027-2 and 60B80027-3 and/or stabilizer trim drive motors P/N 60B00250-1, within 100 hours time in service after effective date of this A.D., and, thereafter at intervals not to exceed 100 hours time in service from the last inspection, test the stabilizer trim system components per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent tests approved by the Chief, Aircraft Engineering Division, FAA Western Region, until modified in accordance with paragraph (c), below. \n\t(b)\tReplace, or modify, prior to further flight, stabilizer trim system components which are found defective bythe inspections per paragraph (a), above, in accordance with Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent replacements or modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(c)\tThe inspections required per (a), above, may be discontinued after accomplishment of the following: \n\t\t(1)\tReplace stabilizer trim modules, Boeing P/N 60B80027-2 and 60B80027-3, with stabilizer trim modules modified with improved arming and control valve seals and steel retainer caps, per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\tNOTE: Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, incorporates LTV Electrosystems Service Bulletins 27-6, 27-8, and 27-9 in "Part II, Terminating Action." \n\t\t(2)\tReplace stabilizer trimmotor, Boeing P/N 60B00250-1, without suffix "D" identification following unit serial number identification, with stabilizer trim motors modified with solid locking pins, per Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, or later FAA approved revisions or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\tNOTE: Boeing Service Bulletin 27-2054, Revision 3, dated February 15, 1972, incorporates Vickers Service Bulletin 910274-4, dated February 10, 1972. \n\t(d)\tWithin 100 hours time in service after the effective date of this A.D., unless already accomplished, incorporate in the FAA approved Airplane Flight Manual, "Emergency Procedures" (Section 2), the following procedures: \n\t"UNSCHEDULED STABILIZER TRIM \n\tRecall \n\tStabilizer Trim Hydraulic Switches - \n\tCUTOUT \n\tReference \n\tAutopilot - DISENGAGE \n\tControl column movement in opposition of trim will stop unscheduled trim caused by electrical fault. \n\tMoving stabilizer trim hydraulic switches to CUTOUT will stop any unscheduled trim. Allow sufficient time for valves to operate. \n\tA portion of the system may be determined to be usable by moving one stabilizer trim hydraulic switch at a time to NORM. If trim is normal that system may be used to adjust trim as required." \n\tAmendment 39-1392 became effective February 11, 1972. \n\tThis amendment 39-1407 becomes effective March 14, 1972.
2011-23-05: We are superseding an existing airworthiness directive (AD) for certain Model 737-300, -400, and -500 series airplanes. That AD currently requires repetitive inspections for cracking of the 1.04-inch nominal diameter wire penetration hole, and applicable related investigative and corrective actions. This AD reduces the compliance times for those actions. This AD was prompted by reports of cracking in the frame, or in the frame and frame reinforcement, common to the 1.04- inch nominal diameter wire penetration hole intended for wire routing; and recent reports of multiple adjacent frame cracking found before the compliance time required by the existing AD. Such cracking could reduce the structural capability of the frames to sustain limit loads, and result in cracking in the fuselage skin and subsequent rapid depressurization of the airplane. We are issuing this AD to correct the unsafe condition on these products.
97-06-05: This amendment adopts a new airworthiness directive (AD) that applies to Avions Pierre Robin Model R2160 airplanes. This action requires repetitively inspecting the weld area between the strut and the lower plate of the nose landing gear leg for cracks, and replacing the strut when cracks are found. The AD is the result of several reports of cracks in the weld securing the nose wheel steering bottom bracket to the nose landing gear leg on the affected airplanes. The actions specified by this AD are intended to prevent nose landing gear failure caused by cracks in the weld area between the strut and the lower plate of the nose landing gear leg, which could result in loss of control of the airplane during landing operations.
77-16-05: 77-16-05 CESSNA: Amendment 39-2998. Applies to Models 210-5(205), 206, P206/TP206, U206/TU206, 207/T207, and 210/T210 series airplanes of the serial numbers specified below. \n\n\tCompliance: Required within the next 25 hours time-in-service after the effective date of this AD, unless already accomplished. \n\n\tTo prevent malfunction of the P/N C291503-0101 or P/N 1216100-1 fuel selector valve, accomplish the following: \n\n\tA.\tOn Model 210-5(205) series (serial numbers 205-0481 thru 205-0577), Model 206 series (serial numbers 206-0001 thru 206-0275), Model P206/TP206 series (serial numbers P206-0001 thru P20600647), Model U206/TU206 series (serial numbers U206-0276 thru U20603123), Model 207/T207 series (serial numbers 20700001 thru 20700322), Model 210/T210 series (serial numbers 21058221 thru 21061154, and T210-0001 thru T210-0454) airplanes, examine the aircraft maintenance records to determine whether the fuel selector valve has been changed subsequent to December 19, 1975.If the valve has not been changed, make an entry in the maintenance record indicating this AD is not applicable to the airplane and no further action is necessary. Examination of the records and the record entry may be accomplished by the owner/operator. \n\n\tIf the valve has been changed subsequent to December 19, 1975, accomplish a fuel valve inspection and, if indicated, replacement in accordance with paragraph C. below. \n\n\tB.\tOn Model U206/TU206 series (serial numbers U20603124 thru U20603712, U20603714 thru U20603791, U20603793 thru U20603797, U20603799 thru U20603803, U20603805, U20603808 thru U20603812, U20603814 thru U20603846, U20603848, U20603850, U20603851, U20603853 thru U20603857, U20603861, U20603862, U20603867 thru U20603871, U20603875, U20603876, U20603882 and U20603886), Model 207/T207 series (serial numbers 20700323 thru 20700373, 20700375 thru 20700394), and Model 210/T210 series (serial numbers 21061155 thru 21061731, 21061733 thru 21061766, 21061768 thru 21061860, 21061862 thru 21061881, 21061883 thru 21061987, 21061990 thru 21061993, 21061995 thru 21062005, 21062007 thru 21062009, 21062012 thru 21062017, 21062020 thru 21062022, 21062024 thru 21062027, 21062029, 21062033 thru 21062037, 21062043) airplanes, accomplish a fuel selector valve inspection and, if indicated, replacement in accordance with paragraph C. below. \n\n\tC.\t1.\tPlace fuel selector valve in OFF position. \n\n\t\t2.\tRemove selector valve handle and associated parts. Obtain access to the valve by removal of the control valve pedestal and selector valve access plate from the floor. \n\n\t\t3.\tCheck the selector valve serial number. If the serial number is 1421 thru 3269 inclusive, accomplish the pull test inspection described in paragraph C.4. below. If the serial number is not within this block, the valve is not affected, and the aircraft may be returned to service after reassembly. \n\n\t\t4.\tUsing tools fabricated in accordance with Figure 1, or an equivalent test arrangement, accomplish a pull test on the selector valve in accordance with the following procedures: \n\n\t\t\ta.\tRemove safety wire, roll pin, and valve handle shaft from valve control yoke. \n\n\t\t\tb.\tFeed cable through hole in yoke and crimp cable securely. Place bar in loop in other end of cable. \n\n\t\t\tc.\tRemove four screws securing selector valve cover. Break the seal between the cover and body and rotate the cover to positively assure it is free. \n\n\t\t\td.\tAttach a tensiometer to the cable in the section between loops and from a seated or squatting position, with one foot on each side of the valve, using the legs and arms to lift, apply an upward force of 130 pounds minimum to 150 pounds maximum directly in line with the shaft while observing the shaft for movement. Exert upward force gradually. If any upward movement of valve shaft is noted, release force immediately. Drain fuel from airplane and install a new valve in accordance with the aircraft's Service Manual Instructions. \n\nCAUTIONIf the shaft pulls out of the valve body, fuel spillage may result. Therefore, all precautions applicable to working with or around open fuel should be observed. \n\n\t\t\te.\tIf no upward movement of the valve shaft is noted or evident, reassemble the valve and aircraft and return to service. \n\n\n\n\nAD 77-16-05 \n\n\tD.\tAny equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. \n\n\tCessna Service Letter SE77-22, dated June 27, 1977, or later approved revisions pertain to the subject matter of this AD. \n\n\tThis amendment becomes effective August 11, 1977.
79-09-01: 79-09-01 BOEING: Amendment 39-3452. Applies to Model 727-100 series airplanes certificated in all categories listed in Boeing 727 Service Bulletin No. 24-30, revised May 18, 1966, that complied with FAA AD 66-30-02 by having been modified to Option No. 2 in said service bulletin. Compliance required as indicated. To prevent generator electrical lead damage and possible hydraulic fire from cable clamp failure allowing the electrical leads to chafe against hydraulic lines or components, accomplish the following: \n\t1.\tWithin the next 100 hours time-in-service after the effective date of this AD, and every 2,000 hours time-in-service thereafter, inspect the No. 1 generator electrical leads from the pressure feed through fittings (below the floor level) to the engine strut feed through, all generator electrical lead clamps, hydraulic systems, and airframe for routing separation and insulation chafing. Repair or replace any electrical lead, hydraulic line, airframe part or clamp damage as required in accordance with approved maintenance procedures. \n\tThe repetitive 2,000 hour inspection interval may be adjusted by FAA air carrier maintenance inspectors to the nearest scheduled maintenance inspection period. \n\t2.\tIf the generator electrical leads are not routed with at least three (3) inches separation from hydraulic lines or components, provide additional physical protection with aircraft quality Skydrol resistant insulation wrap and clamp as required. \n\t3.\tIf generator electrical leads are routed with at least three (3) inches separation from hydraulic lines or components, no further action is required under this AD. \n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis amendment becomes effective May 8, 1979.
78-03-04: 78-03-04 BOEING: Amendment 39-3136. Applies to all Boeing 707-100/-100B/-200 series airplanes, certificated in all categories upon the accumulation of 20,000 landings. \n\tAccomplish the following: \n\tA.\tWithin the next 750 landings, unless accomplished within the last 750 landings, and thereafter at intervals not to exceed 1500 landings, X-ray inspect the upper rear spar chord horizontal leg and adjacent wing skin from the side of the body at BBL 70.5 to WS274 in accordance with Boeing Service Bulletin 3304, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. Wing skins or rear spar chords found cracked are to be repaired prior to further flight in accordance with paragraph B. \n\tB.\tRepair in accordance with 1 or 2 below as applicable, prior to further flight except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed: \n\t\t1.\tAirplanes with skin cracks near the rear spar which do not extend beyond the fastener pattern of stringer No. 1 may continue in service for a maximum of 750 additional landings, subject to the following conditions: \n\t\t\ta.\tCrack ends must be stop drilled per Boeing Service Bulletin 3304. \n\t\t\tb.\tIf crack ends in a fastener hole, the hole must be inspected per Boeing Service Bulletin 3304 to assure there is no crack progression beyond fastener hole, then an additional 1/16" oversize must be made and an oversize fastener installed. \n\t\t\tc.\tEddy current inspection per Boeing Service Bulletin 3304 of crack ends must be conducted at intervals not to exceed 50 landings. Any crack progression requires repair in accordance with 2 below prior to further flight. \n\t\t\td.\tCracks must be permanently repaired within 750 landings in accordance with 2 below. \n\t\t2.\tRepair in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tC.\tFor the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing each airplane's hours time-in-service by the operator's fleet average from takeoff to landing for the airplane type. \n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis AD supersedes AD 77-06-03. \n\tThis amendment becomes effective March 1, 1978.
2010-26-52: We are publishing in the Federal Register an amendment which was sent previously to all known U.S. owners and operators that supersedes an existing airworthiness directive (AD) for the specified Bell Helicopter Textron, Inc. (BHT) Model helicopters with certain tail rotor blades (blades). The superseded AD requires, before further flight, replacing certain blades with airworthy blades. This AD retains the requirements of the superseded AD but adds new blade part numbers (P/Ns) and serial numbers (S/Ns) to the applicability. This AD was prompted by another incident in which the blade tip weight separated from a blade during flight, causing vibration. This incident led to the determination that additional blades could be affected, and should be added to the applicability. We are issuing this AD to prevent loss of the blade tip weight, loss of a blade, and subsequent loss of control of the helicopter.
65-06-03: 65-06-03\tBOEING: Amdt 39-49 Part 39 Federal Register March 17, 1965. Applies to Models 707 and 720 Aircraft with the Serial Numbers Listed in Boeing Service Bulletin No. 2029. \n\tCompliance required as indicated. \n\tCracks have recently been discovered on two airplanes in the aft elevator control quadrant, P/N 50-3119, around the upper two bolts, P/N NAS 1105, which attach the quadrant to the elevator control torque tube. In both instances, Boeing Service Bulletin No. 1259 had not been accomplished, and an interference fit existed between the quadrant and torque tube. As a result of these cracks, the following or an equivalent approved by the Aircraft Engineering Division, FAA Western Region, shall be accomplished: \n\t(a)\tWithin the next 600 hours' time in service after the effective date of this AD, unless already accomplished within the last 600 hours' time in service, and thereafter at periods not to exceed 1,200 hours' time in service from the last inspection, inspect by eddy current or fluorescent dye penetrant the aft elevator quadrant, P/N 50-3119, in accordance with paragraph 3.c. of Service Bulletin No. 2029 or later FAA-approved revisions for evidence of cracks in the upper hub and in the vicinity of the two upper bolts which attach the quadrant to the elevator control torque tube. \n\t(b)\tIf a crack is found, replace the elevator control quadrant with one of the same part number before further flight. The inspections in (a) also apply to the replacement part. \n\t(c)\tWithin the next 6,000 hours' time in service after the effective date of this AD, accomplish rework and dimensional checks in accordance with paragraph 3 of Boeing Service Bulletin 2029 or later FAA-approved revisions. \n\t(d)\tWhen the rework and dimensional checks required by (c) are accomplished, the repetitive inspections specified in (a) may be discontinued. \n\t(e)\tUpon request of an operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\t(Boeing Service Bulletin No. 2029 covers this same subject.) \n\tThis directive effective April 16, 1965.
97-05-10: This amendment adopts a new airworthiness directive (AD) that is applicable to all Boeing Model 737 series airplanes. This action requires removal of the main rudder power control unit (PCU) and replacement with a serviceable unit. This amendment is prompted by a report of the installation of an incorrect bolt on the main rudder PCU. The actions specified in this AD are intended to prevent cracking of the bearing of the main rudder PCU due to installation of an incorrect bolt; such cracking could result in seizure of the bearing and resultant uncommanded rudder movement.
97-05-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-200, -300, and -400 series airplanes, that requires repetitive inspections to detect cracking of the front spar web of the center section of the wing, and repair, if necessary. This amendment is prompted by reports of fatigue cracking found in the front spar web. The actions specified by this AD are intended to prevent the leakage of fuel into the forward cargo bay, as a result of fatigue cracking in the front spar web, which could result in a potential fire hazard.
94-05-06: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model MD-11 series airplanes, that requires modification or replacement of designated passenger cabin floor panels. This amendment is prompted by a report that, during manufacture, the inserts that attach the floor panels to the seat tracks and floor beams were installed using sealant rather than required adhesive. The actions specified by this AD are intended to prevent loss of the passenger cabin floor capability to support the airplane interior inertia loads under emergency landing conditions.
89-21-02: 89-21-02 BOEING: Amendment 39-6342. (Docket No. 89-NM-171-AD) \n\tApplicability: All Model 727 and Model 737-100, -200, and -200C series airplanes certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent engine separation, accomplish the following: \n\n\tA.\tWithin the next 5 days after the effective date of this AD, revise the FAA-approved maintenance program to indicate that only FAA-approved engine mount cone bolt nuts specified in the table below shall be installed on the engine mount cone bolt on Boeing Models 727 and 737-100, -200, - 200C airplanes. \n\n\n\nBoeing Model 727: \n\nLine No.\nFAA-approved engine mount cone bolt nut \n1 - 193\nBACN10B-10L or BACN10GW-10 or NAS1804-10\nor LHEB220-108 or 69-59074-1 \n194 - 692\nLHEB220-108 or 69-59074-1 \n693 - 1832\n69-59074-1 \n\n\n\nBoeing Model 737-100, -200, -200C: \n\nLine No.\nFAA-approved engine mount cone bolt nut \n1 - 124\nLHEB220-108 or 69-59074-1\n125 -1585\n69-59074-1\n\t\t\t \n\tNo substitute shall be used for the LHEB220-108 or Boeing 69-59074-1 engine mount cone bolt nut. \n\n\tB.\tWithin 60 days after the incorporation of paragraph A., above, conduct an inspection to verify that each installed engine mount cone bolt nut conforms to the approved type design as described in paragraph A., above, and, if the 69-59074-1 is installed, verify its authenticity. If the authenticity of the engine mount cone bolt nut 69-59074-1 cannot be verified, replace it with an authentic 69-59074-1 engine mount cone bolt nut prior to further flight. The authenticity of the FAA-approved 69-59074-1 engine mount cone bolt nut is determined by the following method: \n\n\t\tThe only FAA-approved sources for the 69-59074-1 engine mount cone bolt nuts are The Boeing Company and Standard Press Steel. The authentic part markings of the engine mount cone bolt nuts are labeled on the vertical rim of the nut base at the largest diameter; nuts produced prior to 1969 may have the part marking on the sloping surface in lieu of the vertical rim. The authentic cone bolt nut is identified as "-SPS 69-59074-1-" with a single space between the "S" and "6" and absolutely no other markings. \n\n\tC.\tWithin 10 days after completion of the inspection required by paragraph B., above, for each airplane, submit a report of findings of counterfeit engine mount cone bolt nuts installed on the airplane to the Manager, Manufacturing Inspection Office, ANM-108, FAA, Transport Airplane Directorate, 17900 Pacific Highway South, C-68966, Seattle, Washington 98168. This report must include the model of the airplane inspected, the date of inspection, and the number of cycles flown since the last engine maintenance where engine mount cone bolts were installed. \n\n\tNOTE: The report should be forwarded through the assigned air carrier Principal Maintenance Inspector (PMI), who will then send it to the Manager, Manufacturing Inspection Office, ANM-108. \n\n\tD.\tAn alternatemeans of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: 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, Seattle Aircraft Certification Office. \n\n\tE.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment (39-6342, AD 89-21-02) becomes effective on October 16, 1989.
50-04-01: 50-04-01 BRIGGS AND STRATTON: Applies to All Aircraft Equipped With Army Air Force Type A-8 Ignition Switches Manufactured by Briggs and Stratton. \n\n\tInitial compliance required not later than March 1, 1950, and every 100 hours operation thereafter. \n\n\tA serious hazard may exist on this type switch after considerable use has worn the internal switch lever stops, allowing overtravel past the "OFF" position. Such overtravel may allow the magneto ground to be broken and permit the engine to fire when the switch is in the "OFF" position. \n\n\tType A-8 ignition switches manufactured by Briggs and Stratton can be identified by the name Briggs and Stratton stamped on the rear of the switch case. Another distinguishing feature of this switch is a formed sheet metal lever which is not found on other makes of type A-8 switch. \n\n\t1.\tInspection should consist of the following: Check switch lever for overtravel past the "OFF" position. Figure 1 shows the location of the switch lever in the "OFF" position. The pointer projecting from the lever points to the middle "F" in the word "OFF". When the lever can be turned to a point beyond the centerline of the "O" in the word "OFF", the rotation stops have becomes worn and the switch should be replaced. \n\n\n\n\n\t2.\tThis inspection must be repeated at 100-hour intervals. \n\n\t3.\tInspection may be discontinued if switch is replaced by Type A-8 of another make or by some other satisfactory type ignition switch.
87-24-08: 87-24-08 BOEING: Amendment 39-5775. Applies to Model 757 series airplanes, line position 0002 through 0138, certificated in any category. Compliance required within the next one year after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent shutdown of the power transfer unit and inability to extend the landing gear, accomplish the following: \n\n\tA.\tModify the hydraulic system in accordance with Boeing Alert Service Bulletin 757-29A0035, Revision 1, dated September 10, 1987, or later FAA-approved revision. \n\n\tB.\tAn 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tC.\tSpecial 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.\n \n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124-2207. These documents 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. \n\n\tThis amendment becomes effective December 29, 1987.