Results
75-22-03: 75-22-03 HUGHES HELICOPTERS: Amendment 39-2387. Applies to all Hughes Model 369, 369A, 369H, 369HS, 369HE, 369HM, certificated in all categories and military OH- 6A and YOH-6A helicopters incorporating metal tail rotor blades Part Number 369A1613, or -3, in combination with tail rotor drive shaft, Part Number 369A5518. Compliance required within the next 50 hours time in service, after the effective date of this AD unless already accomplished. (a) To prevent possible emergency flight situations resulting from failure of the tail rotor drive shaft, replace drive shaft P/N 369A5518 with drive shaft P/N 369A5518-601 in accordance with Part I, B.3. of Hughes Service Information Notice No. HN-86, dated April 28, 1975, or later FAA-approved revisions. (b) Aircraft may be flown in accordance with FAR 21.197 to a base for the performance of the maintenance described herein. (c) Equivalent modification procedures may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. NOTE: Assure that the record to reflect compliance with this AD is made in aircraft records in accordance with FAR's 91.165 and 91.173. This amendment becomes effective October 24, 1975.
75-26-04: 75-26-04 BEECH: Amendment 39-2458. Applies to Model 200 (BB-2 through BB-71) airplanes. Compliance: Required as indicated, unless already accomplished. To prevent malfunction of the fuel transfer system, within the next 50 hours' time in service after the effective date of this AD, accomplish the following: A) Inspect the P/N 100-920067-1 fuel vent float valve for opening and operation in accordance with the criteria set forth in Beechcraft Service Instruction No. 0772-295 or later FAA-approved revision. (P/N 101-320035-1 baffle, if installed, must be removed to perform this inspection.) B) Replace any fuel vent float valve assembly that does not meet the criteria set forth in Beechcraft Service Instruction No. 0772-295 or later FAA-approved revision with new P/N 100-920067-1 fuel vent float valve assembly. C) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Thisamendment becomes effective December 19, 1975.
75-08-08: 75-08-08 MCDONNELL DOUGLAS: Amendment 39-2157. Applies to Douglas Model DC-10-10, -30, -30F, and -40 Series airplanes, certificated in all categories, with factory serial numbers as indicated in Douglas Service Bulletin No. 52-74, Revision 1, dated December 6, 1974, or later FAA-approved revisions. \n\n\tTo prevent inadvertent opening of Type A passenger doors (other than the most forward left and right doors) during taxi or in flight, accomplish the following: \n\n\t(a)\tPrior to taxi from the terminal area, after the effective date of this AD, and after all passengers have been seated, all Type A passenger door (other than the most forward left and right doors) arming handles must be cycled to the armed position; then returned to the fully disarmed position; and then returned to the fully armed position. If any difficulty in operating the arming handles is experienced, the airplane must be corrected by maintenance personnel prior to further flight. The flight crew will be advised of the check accomplishment prior to takeoff. Operators shall advise flight crews and cabin attendants of the foregoing procedure by the most immediate and practicable means. The procedure is to be accomplished until the modification described in (b), below, is performed on all airplanes in the operator's fleet. \n\n\t(b)\tOn or before July 1, 1975, unless already accomplished, modify the door operating mechanism per Douglas Service Bulletin No. 52-74, dated September 28, 1973, or later FAA-approved revisions, or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(c)\tAirplanes may be flown to a base for performance of maintenance required by this AD per FAR's 21.197 and 21.199, provided that, with respect to the door(s) identified as in need of maintenance, the emergency evacuation slides will either be removed or deactivated. \n\n\tThis supersedes the telegraphic AD, adopted March 1, 1975, and distributed by telegrams dated March 1, 1975. \n\n\tThis amendment becomes effective April 10, 1975.
75-09-11: 75-09-11 MCDONNELL DOUGLAS: Amendment 39-2184. Applies to Model DC-10-10 and DC-10-10F airplanes, certificated in all categories, with P/N ACG 7008-1, collar assembly - lower steering support installed. \n\n\tTo insure continued airworthiness of these assemblies, prior to the accumulation of 44,500 landings, remove from service collar assembly - lower steering support, P/N ACG 7008-1, and mark it permanently and conspicuously to prevent its inadvertent return to service. Replace it with a serviceable assembly having less than 44,500 landings. \n\n\tThe previously established service life of 231,666 landings is no longer applicable. \n\n\tThis amendment becomes effective April 28, 1975.
74-08-11: 74-08-11 DEHAVILLAND: Amendment 39-1819. Applies to DHC-6 airplanes Serial Numbers 136 to 350 inclusive, 353, 356, 360 to 366 inclusive, and 368 certificated in all categories. To prevent rivet failures in the side seat rail attachment structure of inward facing seats which are forward of the right-hand cabin door between Fuselage Stations 239.0 and 290.3, accomplish the following in accordance with de Havilland Service Bulletin No. 6/302, dated November 8, 1973: 1. For airplanes that presently have inward facing seats installed, within the next 200 hours in service after the effective date of this AD, unless already accomplished, inspect in accordance with the Accomplishment Instructions of the Service Bulletin, and alter, as necessary, in accordance with the Service Bulletin or in accordance with an approved equivalent alteration. 2. For future initial installation of inward facing seats, attachments must be in accordance with Details "A", "B", "C", and "D" of theService Bulletin, or approved equivalent attachments. 3. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 522(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to de Havilland Aircraft of Canada, Ltd., Attention: Product Support Department, Downsview, Ontario, Canada. These documents may also be examined at the Eastern Region, Engineering and Manufacturing Branch, Federal Building, J.F.K. International Airport, Jamaica, N.Y. 11430, and at the FAA headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Eastern Region. 4. Equivalent alterations and attachments must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. 5. The compliance time may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, upon receipt of substantiating data submitted through a local FAA maintenance inspector. This amendment is effective April 18, 1974.
72-13-02: 72-13-02 GENERAL DYNAMICS: Amdt. 39-1460 as amended by Amendment 39- 1476. Applies to Model 22 and 22M airplanes certificated in all categories which have accumulated 20,000 or more hours' time in service. Compliance required as indicated. To prevent failures of the main landing gear beams (P/N 22-16732-5 or -6) as a result of cracks, accomplish the following: (a) Within the next 25 hours' time in service after the effective date of this AD and thereafter at intervals not to exceed 25 hours' time in service until (c), below, is accomplished, visually inspect the exposed surface of the upper and lower beam caps as described in General Dynamics Service Bulletin 880 SB No. 57-23C, paragraph 2.IA(1) or 880M SB No. 57-12C, paragraph 2.IA(1), dated May 12, 1972, or later FAA-approved revisions. If cracks are found during these inspections, comply with (c), below, before further flight. (b) Within the next 200 hours' time in service after the effective date of this AD and thereafter at intervals not to exceed 200 hours' time in service until (c), below, is accomplished, inspect the exposed surface of the upper and lower beam caps as described in SB No. 57-23C, paragraph 2.IA(2) or SB No. 57-12C, paragraph 2.IA(2), or later FAA-approved revisions. If cracks are found during these inspections, comply with (c), below, before further flight. (c) Within the next 1000 hours' time in service after the effective date of this AD, unless already accomplished, remove the trailing edge panels and rework the trailing edge panel attachment holes per SB No. 57-23C, paragraph 2.IA(3) or SB No. 57-12C, paragraph 2.IA(3), or later FAA-approved revisions. Inspect the screw holes using eddy current or equivalent procedures immediately following this rework and thereafter until (d), below, is accomplished, at not more than 3000 hours' time in service following this rework. Repeat the inspection at or before 4000 and 5000 hours time in service following this rework.If cracks are found during any of these inspections, comply with (d), below, before further flight. If bushings have been installed in accordance with previous issues of SB No. 57-23 or SB No. 57-12, reinspect the bushing fit for tightness and inspect the surface around the bushings for cracks using eddy current or equivalent procedures at intervals not exceeding 3000 hours' time in service following this initial inspection. If loose bushings or cracks are found, repair in accordance with instructions approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight. For inspection intervals following repair, see (e) below. (d) On or before the accumulation of 6,000 hours' time in service following the rework described in (c), above, unless already accomplished, again remove the trailing edge panels and rework holes per SB No. 57-23C, paragraph 2.IA(5) or SB No. 57-12C, paragraph 2.IA(5), or later FAA-approved revisions. Inspect the screw holes usingeddy current or equivalent procedures immediately following this rework and thereafter, until (e), below, is accomplished, at not more than 3000 hours' time in service following this rework. Repeat the inspection at or before 4000 and 5000 hours' time in service following this rework. If cracks are found during any of these inspections, comply with (e), below, before further flight. (e) On or before the accumulation of 6,000 hours' time in service following the rework described in (d), above, unless already accomplished, again remove the trailing edge panels and rework holes per SB No. 57-23C, paragraph 2.IA(7) or SB No. 57-12C, paragraph 2.IA(7), or later FAA-approved revisions. Inspect the screw holes using eddy current or equivalent procedures immediately following this rework and thereafter at not more than 2000 hours' time in service following this rework. Repeat the inspection at or before 3000, 4000, 5000 and 6000 hours' time in service following this rework. If cracks arefound during any of these inspections, repair in accordance with instructions approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight, or replace the beam per (f), below. After a beam is repaired, the interval inspections of the screw holes (or of the surface around the bushings), following repair and until the beam is replaced, depend upon the repair accomplished, as follows: (1) Reinforcing straps (steel or aluminum) added to upper or lower cap; 3000 hours' time in service. (2) Any plugged hole in caps greater than .375 inch diameter but less than .50 inch diameter and then reamed or bushed; 2000 hours' time in service. (3) Any plugged hole in caps greater than .50 inch diameter and then reamed or bushed; 1000 hours' time in service. (f) On or before the accumulation of 6,500 hours' time in service following the rework described in (e), above, and at intervals not exceeding 500 hours' time in service thereafter, reinspect the screw holes using eddy current or equivalent procedures until the beam is replaced by a new beam of the same type, or an improved type approved by the Chief, Aircraft Engineering Division, FAA Western Region. If cracks are found during these inspections, replace the beam before further flight. After the beam is replaced by a new beam of the same type, the inspections and rework specified in this AD are applicable only after the accumulation of 20,000 hours' time in service on the new beam. After the beam is replaced by a new beam of an improved type, this AD is not applicable. (g) The Chief, Aircraft Engineering Division, FAA Western Region, may approve equivalent inspections, rework procedures, and replacement beams for this AD. This amendment becomes effective June 15, 1972. Amendment 39-1460 became effective June 15, 1972. Amendment 39-1476 becomes effective June 30, 1972.
66-08-05: 66-08-05 MORANE-SAULNIER: Amdt. 39-211 Part 39 Federal Register March 17, 1966. Applies to Models M.S. 760, M.S. 760 A, and M.S. 760B Airplanes. Compliance required within the next 300 hours' time in service after the effective date of this AD unless already accomplished. Replace aluminum alloy rudder control system rod, P/N 0176-27.1.191, located between Stations 5 and 10, with steel rod, P/N 0176-27.1.218. (Morane-Saulnier Service Bulletin No. 45 pertains to this subject.) This directive effective April 16, 1966.
67-17-06: 67-17-06 SIKORSKY: Amdt. 39-415 Part 39 Federal Register May 16, 1967. Applies to Model S-55 Series Helicopters Equipped with S14-20-5300 Series Tail Cone and Pylon Assembly (Straight Tail Cone). Compliance required as indicated. To prevent operation with fatigue cracks in the pylon lower ring accomplish the following: (a) Within the next 10 hours' time in service and thereafter at intervals not to exceed 10 hours' time in service from the last inspection, remove the pylon lower access cover and visually inspect for cracks pylon lower ring, P/N S14-20-3013, with 3,500 or more hours' time in service on the effective date of this AD. NOTE: - Where excess paint that may hinder the visual inspection is present, remove with MIL-R-8633 water-soluble paint remover and apply a light coat of MIL-P-8585 zinc chromate primer. (b) If a crack is found, replace the pylon lower ring assembly, P/N S14-20-3052, before further flight. (c) Remove from service pylonlower ring assembly, P/N S-14-20-3052, with 3,400 or more hours' time in service on the effective date of this AD within the next 100 hours' time in service. (d) Remove from service all other pylon lower ring assemblies, P/N S14-20-3052, before the accumulation of 3,500 hours' time in service. (e) Operators who have not kept records of hours of time in service on pylon lower ring assembly, P/N S14-20-3052, shall substitute helicopter hours of time in service in lieu thereof. (Sikorsky Service Bulletin 55B20-1E covers this subject.) This directive effective May 21, 1967.
2012-10-05: We are adopting a new airworthiness directive (AD) for all Fokker Services B.V. Model F.28 Mark 0070 and 0100 airplanes. This AD was prompted by an in-flight failure of the hydraulic control panel, which resulted in the absence of pressure and quantity indication of the hydraulic system and accompanying alerts for ``hydraulic system 1 low quantity'' and ``hydraulic system 2 low quantity.'' This AD requires implementing new abnormal procedures for hydraulics in the airplane flight manual (AFM). We are issuing this AD to prevent loss of control of the airplane due to incorrect hydraulic system failure information being provided to the flightcrew, followed by application of inappropriate procedures.
74-24-07: 74-24-07 CHROMALLOY: Amendment 39-2021 as amended by Amendment 39-2080. Applies to Chromalloy Model RLB-6 Series rescue locator beacons incorporating battery packs, P/N A3-01-0175, installed in emergency equipment (including, but not limited to, slide-raft combinations and life rafts). Compliance required as indicated unless already accomplished: (a) On or before December 2, 1974, remove from service all Chromalloy Model RLB-6 Series rescue locator beacons or the battery packs incorporated in those units. (b) Contrary provisions of Sections 121.339, 121.353, and 135.163 of the Federal Aviation Regulations notwithstanding, Part 121 and 135 operators of aircraft, equipped with life rafts or slide-raft combinations incorporating Chromalloy Model RLB-6 Series rescue locator beacons may, after removal of such beacons or their battery packs pursuant to paragraph (A), continue to operate the aircraft (1) After replacing the battery packs with Chromalloy battery packs,P/N P4-01-0044, marked "FAA approved for use only as a replacement battery under AD 74-24-07."; (2) After replacing the affected beacons with FAA approved replacement beacons; or (3) Pending compliance with subparagraphs (1) or (2) of this paragraph, after installing in the aircraft an FAA approved portable type emergency locator transmitter that is accessible to the flight crew. (Note: Chromalloy Model RLB-6 Series rescue locator beacons that incorporate battery packs, P/N P4-01-0044, meet the performance requirements of TSO-C91 except their "low operating temperature" is +29 degrees F). Amendment 39-2021 was effective upon publication in the Federal Register as to all persons except those persons to whom it was made immediately effective by the telegram, dated November 9, 1974, which contained this amendment. This amendment 39-2080 is effective upon publication in the Federal Register as to all persons except those persons to whom it was made immediately effective by the telegram, dated November 15, 1974, which contained this amendment.
2012-09-03: We are adopting a new airworthiness directive (AD) for all Saab AB, Saab Aerosystems Model SAAB 2000 airplanes. This AD was prompted by reports of hydraulic accumulator failure. This AD requires replacing certain hydraulic accumulators with stainless steel hydraulic accumulators, and structural modifications in the nose landing gear bay. We are issuing this AD to prevent failure of hydraulic accumulators, which may result in damage to the airplane and injury to occupants.
74-02-01: 74-02-01 BELL: Amdt. 39-1769. Applies to Model 205A-1 and 212 helicopters, certificated in all categories, equipped with tail rotor hub assembly, P/N 212-010-701. Compliance required within the next 25 hours' time in service after the effective date of this A.D., unless already accomplished within the last 25 hours' time in service, and thereafter at intervals not to exceed 25 hours' time in service from the last inspection. To detect possible cracks or excessive looseness in the tail rotor trunnion bearings, accomplish the following: (a) Using a 3 power magnifying glass, inspect the trunnion bearing (P/N 212-010- 723-1 or P/N 212-010-768-1) exposed outer races for cracking. (b) To detect any axial looseness in the trunnion bearings, grasp the hub assembly and attempt to move it towards and away from the tail boom fin assembly. If trunnion bearing movement is considered excessive, remove the tail rotor hub and blade assembly. Measure the amount of axial playin each trunnion bearing using a dial indicator. (c) If there is evidence of either trunnion bearing outer race cracking or axial play in excess of the maximum allowable .015 inch, replace the affected bearing in accordance with the instructions contained in Bell Model 212 Overhaul Manual, dated May 1, 1972, Chapter-Section 65-20-01. (Bell Helicopter Company Service Bulletin Nos. 205-05-73-5, dated August 23, 1973, and 212-05-73-4, dated August 20, 1973, pertain to this subject.) The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Company, P. O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA,4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this A.D. which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. This amendment becomes effective February 22, 1974.
66-22-06: 66-22-06 VICKERS: Amdt. 39-282 Part 39 Federal Register September 7, 1966. Applies to Viscount Models 744 and 745D Series Airplanes. Within the next 1,000 hours' time in service after the effective date of this AD, unless already accomplished, install replacement access panels and cowls for the inverters in accordance with British Aircraft Corporation Ltd. Modification Bulletin No. D.3157 or later ARB-approved issue or an equivalent approved by the Chief, Aircraft Certification Staff, Europe, Africa, Middle East Region. This directive effective October 7, 1966.
74-18-12: 74-18-12 AIRESEARCH MANUFACTURING COMPANY of ARIZONA: Amendment 39-1941. Applies to AiResearch Model TFE 731-2-2B engines installed in, but not limited to Lear-Gates Learjet model 35/36 Aircraft, certificated in all categories. (A) Before further flight, unless previously accomplished, and prior to the installation of replacement fuel control assemblies, replace the hydromechanical fuel control orifice assembly in accordance with AiResearch Alert Service Bulletin TFE 731-A73-3006, dated August 12, 1974, or later FAA approved revisions. (B) Equivalent procedures may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiation data. (C) Aircraft may be flown to a base for performance of maintenance required by this AD per FAR's 21.197 and 21.199. This amendment is effective September 3, 1974, for all persons except those to whom it was made effective immediately by telegram dated August 5, 1974.
2012-08-11: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model DHC-8-400 series airplanes. This AD was prompted by test reports that showed that failure of a retract port flexible hose of a main landing gear (MLG) retraction actuator could cause excessive hydraulic fluid leakage. This AD requires a detailed inspection for defects and damage of the retract port flexible hose on the left and right MLG retraction actuator and replacement of the flexible hose if needed. We are issuing this AD to detect and correct defects and damage of the retract port flexible hose which could lead to an undamped extension of the MLG and could result in MLG structural failure, leading to an unsafe asymmetric landing configuration.
50-41-01: 50-41-01 TAYLORCRAFT: Applies to All Model B Series Aircraft, Serial Numbers 1001 and Up. Compliance required not later than November 15, 1950. Reports have been received of interference between the elevator horn bolt and the fin cover plate apparently caused by improper field installation of the cover plate through bolt. Cases are known where the bolt has worn through the cover plate and such interference may result in jamming of the elevator control system. An inspection of the parts should be made and if evidence of interference is noted, suitable means of preventing the cover plates from interfering with the elevator horn bolt should be incorporated; a spacer bushing at least 1/4 x 0.028 x 1 1/4 inches installed around the cover plate through bolt is considered satisfactory. (Taylorcraft, Inc. Service Bulletin 65 covers this same subject.)
68-11-01: 68-11-01\tBOEING: Amendment 39-605. Applies to Model 707 Series airplanes equipped with Collins Model 51RV-1 navigation receivers and Pratt & Whitney P/N 563586 or General Laboratories P/N 42721 solid state engine ignitor systems. \n\tCompliance required as indicated. \n\tReports indicate a localizer course indicator error may exist when the Pratt & Whitney P/N 563586 or General Laboratories P/N 42721 solid state engine ignitors are turned "ON". Since this unsafe condition is likely to exist in other Boeing Model 707 Series airplanes, the Director determined that: \n\tA.\tFlight operations shall not be predicated on the use of localizer receiver information when engine ignition systems are "ON". \n\tB.\tThe operating limitations specified in subparagraph A of this paragraph must be placed in the aircraft in the form of a placard in clear view of the pilot. For purposes of this AD any documents including a pilot check list including these limitations constitute a placard within the meaningof this paragraph. \n\tC.\tThis limitation does not apply when a 1000 microfarad 50 volt capacitor has been installed across the 28 volt DC input lead to the Collins 51RV-1 receiver in the radio junction box in accordance with Boeing telegraphic Alert Bulletin 2752 dated May 10, 1968, or an equivalent suppression system approved by the Western Region Aircraft Engineering Division. \n\tThis amendment becomes effective on May 31, 1968, for all persons except those to whom it was made effective immediately by telegram dated May 10, 1968.
2012-08-13: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 777-200 and -300 series airplanes. This AD was prompted by reports of two failures of the single-tabbed bracket on the rudder. This AD requires replacing certain single-tabbed bonding brackets in the airplane empennage with two-tabbed bonding brackets. This AD also requires, for certain airplanes, installing new bonding jumpers, and measuring the resistance of the modified installation to verify resistance is within specified limits. We are issuing this AD to prevent failure of the bonding jumper bracket, which could result in loss of lightning protection ground path, which could \n\n((Page 24358)) \n\nlead to increased lightning-induced currents and subsequent damage to composite structures, hydraulic tubes, and actuator control electronics. In the event of a lightning strike, loss of lightning ground protection could result in the loss of control of the airplane.
71-19-02: 71-19-02 SOCIETE NATIONAL INDUSTRIELLE AEROSPATIALE (S.N.I.A.S.): Amendment 39-1281. Applies to Sud Model SE.210, MK. V1-R "Caravelle" airplanes. To prevent a possible fire due to unnoticed overheating of a hydraulic system, within the next 500 hours' time in service after the effective date of this AD, unless already accomplished, incorporate S.A. Modification 1262 by installing a Green and Blue Hydraulic System Fluid Overheat Detection System in accordance with Sud Service-Caravelle Bulletin No. 29-70, Revision 10, dated October 12, 1970, or later SGAC-approved issue or an FAA-approved equivalent. This amendment becomes effective September 27, 1971.
68-20-06: 68-20-06 NAVION: Amendment 39-663. Applies to Navion through Navion H airplanes. Compliance required within the next 50 hours time in service after the effective date of this AD, and each annual inspection thereafter. To prevent failure of the Rudder Horn, P/N 145-24401, accomplish the following: Inspect for horizontal cracks or corrosion in the edge of the rudder horn. These cracks would appear as delaminations or swelling under the paint. Replace corroded or cracked rudder horns with new or unused part of the same part number or Federal Aviation Administration approved equivalent part before further flight. This amendment becomes effective October 3, 1968.
67-15-03: 67-15-03 ROLLS-ROYCE: Amdt. 39-399 Part 39 Federal Register April 14, 1967. Applies to All Dart Series 525, 526, 527, 528, 529, 531, and 532-7 Series Engines with Rolls-Royce Modification 529 (Part 2) Bearing in the Rear Position. Compliance required as indicated. (a) Inspect the oil filter on all Dart 525, 526, 527, 528, 529, 531, and 532-7 Series engines modified in accordance with Rolls-Royce Modification 529 Part 2 Standard at the following times: (1) Within 25 hours' time in service after the effective date of this AD, unless already accomplished within the last 25 hours, and thereafter at intervals not exceeding 50 hours' time in service from the last inspection; and (2) Before further flight, when an increase in oil consumption or a drop in oil pressure is reported. (b) If metal particles are found in the filters, remove the engine from service and further inspect to determine whether repairs are required. (c) If the inspection in (b) indicates that repairs are required, modify the rear bearing in accordance with Rolls Royce Dart Service Bulletin No. Da. 72-232, by implementing any of the following modifications as applicable (1) Mod. 1023, Mod. 1030 and DRS. 411; (2) Mod. 1106 or Mod. 1109; or (3) Mod. 1167. (d) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Certification Staff, Europe, Africa and Middle East 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. This supersedes AD 63-21-7. This directive effective May 14, 1967.
2012-07-06: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 777 airplanes. This AD was prompted by a new revision to the airworthiness limitations of the maintenance planning document. This AD requires revising the maintenance program to update inspection requirements to detect fatigue cracking of principal structural elements (PSEs). We are issuing this AD to ensure that fatigue cracking of various PSEs is detected and corrected; such fatigue cracking could adversely affect the structural integrity of these airplanes.
75-05-10: 75-05-10 BOEING: Amendment 39-2107. Applies to Boeing Model 727-200 series airplanes certificated in all categories, and Boeing Model 727-100, Serial Numbers 18877, 18878, 18879, 19281 and 19279 only. Compliance required as indicated. \n\tTo prevent escape slide latch cable failure and ensuing inability to open emergency exit and deploy escape slide, accomplish the following: \n\tA.\tWithin 25 hours time in service after effective date of this AD, unless already accomplished, inspect door mounted escape slide latch cable at both entry and service doors (4 doors). If corrosion or breakage is found, replace prior to further flight with a carbon steel assembly of the same type design and reinspect at intervals not to exceed 500 hours time in service, or replace with corrosion resistant assembly, per Boeing Alert Service Bulletin 727-25- 223, or later FAA approved revisions, or replace in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tB.\tBySeptember 1, 1975, unless already accomplished, replace carbon steel cable assembly with corrosion resistant assembly per Boeing Alert Service Bulletin 727-25-223, or later FAA approved revisions, or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \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. The documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective March 25, 1975.
2012-06-06: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 757 airplanes. This AD requires replacing the power control relays for the fuel boost pumps and override pumps with new relays having a ground fault interrupter (GFI) feature. This AD also requires an electrical bonding resistance measurement for certain GFI relays to verify that certain bonding requirements are met. This AD also requires, for certain airplanes, an inspection to ensure that certain screws are properly installed, and installing longer screws if necessary. This AD was prompted by fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent damage to the fuel pumps caused by electrical arcing that could introduce an ignition source in the fuel tank, which, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
72-16-02 R3: 72-16-02 R3 BEECH: Amendment 39-1494 as amended by Amendment 39-1549 and 39-2211 is further amended by Amendment 39-4085. Applies to Beech Models C45G, TC-45G, C-45H, TC-45H, TC-45J (SNB-5), RC-45J (SNB-5P), D18C, D18S, E18S, E18S-9700, G18S, H18, JRB-6, 3N, 3NM and 3TM Aircraft certificated in all categories with STC SA4-1531, STC SA111WE, STC SA1832WE or any other STC modification incorporating the provisions of the Volpar Tri-Gear Installation. \n\n\tCompliance required as indicated. \n\n\t1.\tNose Landing Gear Fork \n\n\t\ta.\tFor airplanes incorporating Volpar nose landing gear fork P/N 347 perform the following: \n\n\t\t\t(i)\tWithin the next 50 hours time in service or 25 landings, whichever occurs earlier, after the effective date of this amendment to AD 72-16-02, unless already accomplished within the last 50 hours time in service or 25 landings, and thereafter at intervals not to exceed 100 hours time in service or 70 landings, whichever occurs earlier, from the last inspection, inspect the fork for cracks using dye penetrant or fluorescent penetrant inspection methods in accordance with Volpar Service Bulletin No. 17, as revised July 29, 1969, or later FAA-approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region, until modified in accordance with paragraph 1b. below. \n\n\t\t\t(ii)\tBefore each flight conduct a visual check of fork P/N 347 for cracks until modified in accordance with 1b. below. This visual check may be performed by the pilot in command and shall be recorded in the appropriate aircraft records per FAR 91.173. \n\n\t\tb.\tIf cracks are found by the inspections or checks per paragraph 1.a.(i) or 1.a.(ii) above, replace fork prior to further flight with Volpar P/N 884. \n\n\t\tc.\tThe inspections and checks required per paragraphs 1.a.(i) and 1.a.(ii) may be discontinued when Volpar fork P/N 884 is installed. \n\n\t2.\tNOSE LANDING GEAR TRUNNION \n\n\t\ta.\tFor airplanes incorporating Volpar nose landing gear trunnion P/N 271 with outside boss diameter of 1.01 + .01 in. (color coded clear) within the next 50 hours time in service after the effective date of this AD, unless already accomplished within the last 950 hours time in service, and thereafter at intervals not to exceed 1000 hours time in service from the last inspection, inspect the trunnion for cracks using dye penetrant or fluorescent penetrant inspection methods in accordance with Volpar Service Bulletin No. 19, dated 16 January 1970, or later FAA-approved revision, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region, until modified in accordance with paragraph 2b. below. \n\n\t\tb.\tIf cracks are found by the inspection per paragraph 2a. above, replace trunnion prior to further flight with Volpar P/N 271 \n\n\t\tc.\tThe inspections required per paragraph 2a. may be discontinued upon accomplishment of paragraph 2b. above. \n\n\t3.\tMain Landing Gear Cylinder and Top Brace Assembly \n\n\t\ta.\tFor airplanes with Volpar tri-gear which do not incorporate the Volpar P/N 859 strap reinforcement on Beech main landing gear cylinder and top brace assembly P/N 404-188406, inspect the cylinder and top brace assembly for cracks within 50 hours' time in service after the effective date of amendment 39-1594 to AD 72-16-02, unless already accomplished. For inspection purposes accomplish the following: \n\n\t\t\t(1)\tSupport aircraft on jacks. \n\n\t\t\t(2)\tRemove main wheel and brake assembly. \n\n\t\t\t(3)\tRemove main landing gear shock strut assembly from aircraft. \n\n\t\t\t(4)\tClean surfaces and inspect the cylinder and top brace assembly for cracks using magnetic particle inspection method per MIL-I-6868 or dye penetrant inspection method "C", type II, per MIL-I-6866. \n\n\t\tb.\tIf cracks are found by the inspections per paragraph 3a. above, repair in accordance with FAR Part 43 prior to accomplishing modification per paragraph 3c. below. \n\n\tNOTE: The repair is restricted to theareas shown on the attached Figure No. 2. If cracks are found in areas other than shown, disassemble shock strut and replace cylinder and top brace assembly as follows: \n\n\t\t\t(1)\tRelease air charge and remove AN 6286 valve from main landing gear shock strut. \n\n\t\t\t(2)\tRemove the following components from shock strut: \n\n\t\t\t\t(a)\tCylinder cap assembly P/N 414-188438 \n\n\t\t\t\t(b)\tBracket P/N 709 \n\n\t\t\t\t(c)\tTorque links P/N 738 and P/N 706 \n\n\t\t\t(3)\tDrain oil from the cylinder. \n\n\t\t\t(4)\tRemove the AN 365-820 nut from the lower end of the piston at the P/N 426 fork. \n\n\tNOTE: Care must be taken to avoid shearing the roll pin installed on the E-G-H18 aircraft metering rod assembly. Use a 3/4" socket to hold the upper end of the metering rod. On C-45 and D18 aircraft, a slotted screw driver is used to hold the metering rod. \n\n\t\t\t(5)\tRemove the P/N 426 fork from the piston by pressing off. Heat may be used on the fork to facilitate removal. Heat to a maximum of 300 degrees F - 350 degrees F. \n\n\t\t\t(6)\tRemove the P/N 275 stud from the bottom of the piston and slide piston, metering rod, inner cylinder and seals from the outer cylinder assembly. \n\n\t\t\t(7)\tReverse the above procedure for the assembly of shock strut using a cylinder and top brace assembly that has been inspected and modified in accordance with paragraph 3c. below. \n\n\t\t\t(8)\tComplete a landing gear operational check before returning the aircraft to service. \n\n\tCAUTION: (a)\tThe AN 936-816 lock washer should be installed on to the threaded portion of the metering rod between the P/N 275 stud and the base of the piston. \n\n\t\t(b)\tThe AN 6227-7 "O" ring should be installed in groove on metering rod before installation in the piston. \n\n\t\t(c)\tThe 426 fork should not be driven or pressed on to piston with the AN 365 nut. Heat should be used on the P/N 426 fork. Cool piston with ice to allow slide fit, then torque AN 365 nut in place on stud. \n\n\tc.\tIf no cracks are found by the inspections perparagraph 3a. above, modify cylinder and top brace assembly with Volpar P/N 859 strap reinforcement prior to further flight in accordance with the attached Figure No. 1. \n\n\tNOTE: Following the installation of the reinforcement, reinspect the top brace assembly for cracks using magnetic particle inspection method per MIL-I-6868. If cracks are found, repair in accordance with FAR Part 43 prior to further flight. \n\n\tAmendment 39-1494 became effective August 3, 1972. \n\n\tAmendment 39-1549 became effective November 3, 1972. \n\n\tAmendment 39-2211 became effective May 27, 1975. \n\n\tThis amendment 39-4085 becomes effective April 16, 1981.