Results
75-13-02: 75-13-02 BEECH: Amendment 39-2239. Applies to Models 95, B95, B95A, D95A, E95 (Serial Numbers TD-2 through TD-721) airplanes. Compliance: Required as indicated, unless already accomplished. To reduce the possibility of operations that may impose excessive propeller blade vibration stresses, within the next 100 hours' time in service after the effective date of this AD, accomplish the following: A) Install a Beech P/N 95-324079-1 operational limitation placard adjacent to the manifold pressure indicator which reads: "DO NOT EXCEED 23" HG M.P. BELOW 2300 RPM" and install a corresponding Beech P/N 95-590014-69 flight manual supplement in the airplane flight manual and operate the aircraft in accordance with this limitation. B) Remove, functionally test, and calibrate the tachometer(s) to obtain an accuracy of + or - 25 rpm at 2300 rpm and 2700 rpm. C) Remove, functionally test, and calibrate the manifold pressure indicator(s) to obtain an accuracyof + or - .4 inch Hg at 2300 inches Hg manifold pressure. D) Replace any instruments not meeting the tolerances specified in Paragraphs B and C above with instruments that meet those tolerances. E) Any alternate method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. NOTE: The FAA asks that an M or D Report be filed stating the amount and direction of any error on any tachometer or manifold pressure indicator not meeting the specified tolerances at the initial check. (Reporting approved by Office of Management and Budget under OMB No. 04-R0174.) Beechcraft Service Instruction No. 0723-241 and Hartzell Propeller Service Bulletin No. 107A dated January 27, 1975, refer to this subject. This amendment becomes effective June 23, 1975.
75-04-03: 75-04-03 MCDONNELL DOUGLAS: Amendment 39-2085. Applies to Douglas Model DC-10-10, -30, and -40 Series airplanes, certificated in all categories. \n\n\tCompliance required as follows unless already accomplished. \n\n\tTo improve cabin oxygen mask deployment reliability, accomplish the following: \n\n\tA.\tWithin the next 300 hours' additional time in service after the effective date of this AD, unless already accomplished within the last 300 hours' time in service: \n\n\t\t1.\tConduct an oxygen mask functional (drop) test, by cabin or by each section, in accordance with Paragraph IX of Douglas All Operators Letter No. 10-742, dated December 13, 1974, or 10-742A, dated January 24, 1975, or later FAA-approved revisions. \n\n\t\t2.\tOxygen compartments that fail to open must be inspected and modified in accordance with all of the applicable provisions of Douglas All Operators Letter No. 10-742, dated December 13, 1974, or 10-742A, dated January 24, 1975, or later FAA-approved revisions. \n\n\t\t3.Repeat the oxygen mask functional test until 100% mask drops are achieved by cabin or by each section. \n\n\tB.\tWithin the next 1500 hours' additional time in service after the effective date of this AD: \n\n\t\t1.\tModify partition oxygen compartment doors to insure 180 degrees opening in accordance with Douglas Service Bulletin No. 25-163, dated June 11, 1974, or later FAA-approved revisions. \n\n\t\t2.\tInstall pictorial warning placards on oxygen generator heat shields in accordance with Douglas Service Bulletin No. 35-12, dated April 26, 1974, or later FAA-approved revisions. \n\n\t\t3.\tModify oxygen compartments in passenger seat backs and partitions in accordance with Douglas Service Bulletin No. 35-16, dated August 19, 1974, or later FAA-approved revisions. \n\n\t\t4.\t(a)\tInspect and modify the cabin oxygen system, as applicable, in accordance with Douglas All Operators Letter No. 10-742, dated December 13, 1974, or 10-742A, dated January 24, 1975, or later FAA-approved revisions.(b)\tRepeat the oxygen mask functional test until 100% mask drops are achieved by cabin or by each section. \n\n\tC.\tResults of the functional tests required in paragraphs A. and B. must be forwarded within 30 days in a written report to the Chief, Aircraft Engineering Division, FAA Western Region. Recording approved by the Bureau of the Budget under Order BOB No. 04-R0174. \n\n\tD.\tThe Chief, Aircraft Engineering Division, FAA Western Region, may approve equivalent inspections and modifications upon submittal of substantiating data. \n\n\tE.\tAircraft may be flown to a base for accomplishment of the maintenance required by this AD per FAR's 21.197 and 21.199. \n\n\tThis amendment becomes effective February 14, 1975.
74-13-09: 74-13-09 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-1882 is further amended by Amendment 39-1903. Applies to Model TFE731-2-1C, and -2-2B engines installed in, but not limited to AMD Falcon 10 aircraft, certificated in all categories. Compliance required within the next 50 hours time in service after the effective date of this AD, unless already accomplished, and prior to installing replacement fuel pump assemblies. To detect the improper configuration of the fuel pump, accomplish the following: (a) Inspect the fuel pump assembly in accordance with the instructions contained in AiResearch Service Bulletin TFE731-73-3004, dated June 14, 1974, or later FAA-approved revisions. (b) The inspection prescribed in paragraph (a), above, need not be accomplished prior to the installation of replacement fuel pumps identified as P/N 3070851-7 and -8, or later dash number designations. (c) Equivalent procedures may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiation data. (d) Aircraft may be flown to a base for performance of maintenance required by this AD per FAR's 21.197 and 21.199. Amendment 39-1882 became effective June 28, 1974. This Amendment, 39-1903 becomes effective July 31, 1974.
75-09-16: 75-09-16 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-2192. Applies to Lockheed-California Company Model L-1011-385-1 series airplanes certificated in all categories with hydraulic dampers identified by Part Number 672470-101, -103, -105. To prevent occurrence of dangerous flutter which could result in damage or loss of elevators, outboard ailerons and rudder control surfaces, and/or the associated lifting surface accomplish the following: (a) Within 300 flight hours of the effective date of this AD unless previously accomplished within 200 hours flight time prior to the effective date of this AD, and at intervals not to exceed 500 flight hours thereafter, accomplish the inspections, servicing and replacement of dampers if required, on the elevators of aircraft with Serial Number 193X-1102 and subsequent, per Lockheed Alert Service Bulletin 093-27-A126, dated March 28, 1975 or later FAA-approved revisions. (b) Within 500 flight hours of the effective date of this AD, and at intervals not to exceed 500 flight hours thereafter, accomplish the inspections, servicing and replacement of dampers as required on the outboard ailerons per Lockheed Alert Service Bulletin 093-27-A126, dated March 28, 1975, or later FAA-approved revisions. (c) Within 1500 flight hours after the effective date of this AD, and at intervals not to exceed 1500 flight hours thereafter, accomplish the inspections, servicing and replacement of dampers as required on the rudder per Lockheed Alert Service Bulletin 093-27-A126, dated March 28, 1975, or later FAA-approved revisions. (d) If damper replacement is required per instructions of above Lockheed Alert Service Bulletin 093-27-A126, dated March 28, 1975, or later FAA-approved revisions, replace the damper with (-105) damper configuration only, as defined in the Lockheed Service Bulletin 093-27-088, Revision No. 1 dated February 28, 1974, or later FAA-approved revisions. (e) Equivalent inspections and replacements may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (f) Airplanes may be flown to a base for the accomplishment of the inspections required by the AD, per FAR's 21.197 and 21.199. This amendment becomes effective May 2, 1975.
71-23-02: 71-23-02 GRUMMAN: Amdt. 39-1328. Applies to all Model G-159 airplanes. Compliance required as indicated. To detect cracking in the wing to fuselage attachment fittings at butt line 9 of Grumman Model G-159 airplanes, accomplish the following: a. Within six months time in service after the effective date of this AD, unless already accomplished, inspect the wing to fuselage attachment fittings, P/Ns 159WM10064 and 159WM10065 (P/N 159WM10223 assembly), and P/N 159WM10045 at butt line 9 left and right, wing front beam for cracks, deformation, gaps or improper shimming in accordance with Grumman Gulfstream I Aircraft Service Change No. 190, dated June 28, 1971, or later FAA approved revision or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Southern Region. b. If cracks, deformation, gaps or improper shimming are found when conducting the inspection required by paragraph a., within 100 hours time in service after detection correct in accordance with Aircraft Service Change 190 or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Southern Region. c. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southern Region, may adjust the inspection time to coincide with inspections for wing corrosion required by AD 67-04- 01. This amendment becomes effective November 26, 1971.
70-17-06: 70-17-06 MORANE SAULNIER: Amdt. 39-1068. Applies to Models MS. 880B, MS. 885, and MS. 894A airplanes. To prevent the possibility of flames or harmful gases passing into the cabin from the engine compartment, within the next 100 hours' time in service after the effective date of this AD, unless already accomplished, replace the existing firewall sealant with "STABOND HT-4", Specification LAC-40-475 fire-resistant sealant manufactured by American Latex Product Corporation, or other FAA-approved fire-resistant sealant. This amendment becomes effective September 12, 1970.
75-03-07: 75-03-07 LOCKHEED: Amendment 39-2081. Applies to Model L-1011-385-1 airplanes certificated in all categories. Compliance required as indicated. To prevent possible malfunctions resulting in unuseable passenger evacuation slides and slide/rafts and improve the overall reliability of the passenger evacuation system, accomplish the following: (a) Within the next 30 days' calendar time after the effective date of this AD, unless already accomplished, rework the escape slides and slide/raft primary valise release cables and inspect slide/raft packs for steel reenforcing plate delamination in accordance with Lockheed Service Bulletin 093-25-206, dated January 10, 1975, or later FAA-approved revisions, or equivalent inspections and/or modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. (b) Within the next 180 days' calendar time after the effective date of this AD, unless already accomplished, replace escape slide release pin assemblies in accordance with Air Cruisers Service Bulletin 25-16, dated March 22, 1974, or later FAA-approved revisions, or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. (c) Within the next 180 days' calendar time after the effective date of this AD, unless already accomplished, inspect for proper escape slide girt extension and pack height in accordance with Lockheed Service Bulletin 093-25-205, dated January 10, 1975, or later FAA-approved revisions, or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. (d) Within the next 180 days' calendar time after the effective date of this AD, unless already accomplished, install heat shrinkable tubing on escape slide and slide/raft release pin assemblies in accordance with Lockheed Service Bulletin 093-25-212, dated January 10, 1975, or later FAA-approved revisions, or equivalent modifications approved by the Chief, Aircraft Engineering Division,FAA Western Region. (e) This paragraph is superseded by AD 76-01-07, paragraph (d). (f) Aircraft may be flown to a base for performance of the maintenance required per this AD in accordance with FAR's 21.197 and 21.199. This amendment becomes effective March 5, 1975.
76-05-04: 76-05-04 BEECH: Amendment 39-2536 as amended by Amendment 39-2662. Applies to Models 35, 35R, A35 and B35 (Serial Numbers D-1 thru D-2680) and Pine Air Model Super-V (Serial Numbers SV-XXX-D-1 thru SV-XXX-D-2680) airplanes, having 1,000 or more hours' time in service. Compliance required as indicated, unless already accomplished per AD 75-20-04. To prevent possible stabilizer loss or failure, on those airplanes having the P/N 35- 405130 stabilizer attach fitting, within 50 hours' time in service after the effective date of this AD and thereafter at intervals not to exceed 1,000 hours' time in service, accomplish the following in accordance with Beechcraft Service Instructions 0729-130, Rev. I or later FAA-approved revisions: A) Remove stabilizer attach bolts, plates and other components necessary to provide access to the stabilizer attach fitting and then remove said fitting. B) Inspect the stabilizer attach fitting by visual and dye penetrant methods in accordance with the procedures specified in FAA Advisory Circular (AC) 43.13-1A. C) If as a result of any inspection required herein, a stabilizer attach fitting is found cracked, prior to further flight, replace it with a new or airworthy P/N 35-405130, P/N 35- 650044-1 or P/N 35-405130-3 stabilizer attach fitting. D) When P/N 35-650044-1 or P/N 35-405130-3 stabilizer attach fitting has been installed the requirements of this AD no longer apply. E) If a crack is found as a result of any inspection required herein, provide the FAA with written notification thereof utilizing Malfunction and Defect Report (FAA Form 8330-2) stating the location and length of any crack discovered and the total operating time of the airplane or part of the time of discovery. (Reporting approved by the Office of Management and Budget under ONB No. 04-R0174.) F) Aircraft may be flown in accordance with FAR 21.197 to a base where the repair can be performed. G) Any alternate method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Amendment 39-2536 supersedes AD 75-20-4, Amendment 39-2370. Amendment 39-2536 became effective March 12, 1976. While the effective date of this amendment 39-2662 is July 12, 1976, the effective date for determining compliance with AD76-05-04 remains March 12, 1976.
68-25-05: 68-25-05 SCHLEICHER: Amendment 39-692. Applies to Schleicher Model AS-K13 Gliders, Serial Nos. 13000 through 13091. Compliance required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. To prevent the wheel brake cable from fouling with the release lever of the CG coupling, install a fairlead for the wheel brake cable in accordance with Schleicher Modification No. 2 dated May 30, 1968, or later LBA approved issue or a Federal Aviation Administration approved equivalent. This amendment becomes effective January 10, 1969.
71-24-06: 71-24-06 FAIRCHILD HILLER ROTORCRAFT: Amdt. 39-1338. Applies to FH1100 Type Helicopters Certificated in all Categories. Compliance required as indicated. To preclude fatigue failure of the rear attachment lug on engine mount strut P/N 24-63110-1 accomplish the following: a. Within the next 25 hours' time in service after the effective date of this AD unless already accomplished, inspect and replace if necessary, engine mount strut P/N 24-63110-1 in accordance with Section 2, subsection A.1. and A.2., Accomplishment Instructions, Fairchild Hiller Service Bulletin SB FH1100-71-2 dated 30 September 1971 or later FAA-approved revision or an alternate method approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. b. Within the next 100 hours' time in service after the effective date of this AD unless already accomplished, alter engine mount strut P/N 24-63110-1 in accordance with Section 2, subsection A.3. through A.8., Accomplishment Instructions, F/H Service Bulletin SB FH1100-71-2 dated 30 September 1971 or later FAA-approved revision or an alternate method approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective November 26, 1971.
75-17-08: 75-17-08 PILATUS AIRCRAFT LTD: Amendment 39-2310. Applies to B4-PC11 gliders, certificated in all categories, equipped with canopy limiting lock link. Compliance is required as indicated, unless already accomplished. To ensure canopy emergency jettisoning capability, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, disconnect the limiting cable. (b) Within the next 50 hours' time in service after the effective date of this AD, alter the canopy limiting cable attachment in accordance with the accomplishment instructions of Pilatus Service Bulletin No. 1001, dated April 1974, or an FAA-approved equivalent. This amendment becomes effective August 18, 1975.
75-26-06: 75-26-06 AIR CRUISERS COMPANY: Amendment 39-2456. Applies to Life Raft Systems, P/N Series D23835, 17D23336, 21D23548, 21D23541, 12D11751, 18D23350, and Life Raft Assembly P/N 22D23585 with dates of manufacture from January, 1971, through August 13, 1975, inclusive. Compliance is required, unless already accomplished, to eliminate the possibility of separation at the hose connection fitting-body juncture braze of inlet port assembly, P/N 15C18082. No later than 90 days after the effective date of this AD, accomplish the inspection, replacement, where required, and marking of the above-mentioned part numbers in accordance with Air Cruisers Company Service Bulletin 111-74-1, Rev. No. 1, dated August 12, 1975, or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective December 17, 1975.
73-25-04: 73-25-04 BEECH: Amendment 39-1751 as amended by Amendment 39-1797. Applies to Model B19 (Serial Numbers MB-481 through MB-616) airplanes. Compliance: Required as indicated, unless already accomplished. To assure the takeoff and climb capability of these aircraft meet the certification requirements, accomplish the following: A) Effective immediately, operation of the airplane at a gross weight of 2000 pounds and in excess of three occupants is prohibited. B) Within the next 10 hours' time in service or ten calendar days, whichever comes first, after the effective date of this AD: 1) In place of the existing normal category placard entry which reads "MAXIMUM DESIGN WEIGHT 2250 POUNDS" substitute in wear resistant form a placard entry which reads "MAXIMUM DESIGN WEIGHT 2000 POUNDS" and 2) By appropriate entries and calculations amend the airplane weight and balance records to reflect a maximum design weight of 2000 pounds, c.g. locations between 109.9and 118.3 inches and a maximum of three occupants. C) All performance and operating data contained in the Owners Manual for these model airplanes are no longer applicable. D) As an alternate means of compliance with this AD, for operation with four occupants and a maximum certificated gross weight of 2150 pounds, install Beech Kit 23-9014-1 S in accordance with Beechcraft Service Instruction 0616-010 or any equivalent modification approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This information will be reflected in a forthcoming Type Certification Data Sheet revision. Amendment 39-1751 became effective December 14, 1973, to all persons except those to whom it was made effective by air mail letter dated November 7, 1973. This Amendment 39-1797 becomes effective March 18, 1974.
75-11-06: 75-11-06 BELLANCA: Amendment 39-2209 as amended by Amendment 39-2242. Applies to the following airplanes: 17-31 : S/N 32-1 17-31TC : All serials 17-31A : S/N 32-21 through 75-32-159 except S/N 32-25 17-31ATC : S/N 31004 through S/N 75-31116 Compliance: Required within the next 25 hours time in service after the effective date of this AD. The airplane may be flown to a facility where the modification can be accomplished after the expiration of the 25 hours time in service after the effective date of the AD. To prevent possible engine power failure accomplish the following as appropriate: A. Models Up To But Not Including 1973 Models* The check valve (P/N 19121-206**) is located on the diagonal on the forward wing truss under R/H front seat; the following procedure is to be used. 1. Remove right front seat, lift up carpet and remove plywood cover. 2. Remove check valve. 3. Disassemble the check valve, remove the internal check ball and reassemble the valve without the internal check ball. Remove the Bellanca placard showing the valve as P/N 19121-206 and remark the valve housing with a yellow stripe through the AN part number. 4. Reinstall the valve in the vapor return line. 5. Make appropriate log book entries. B. 1973 Model Aircraft* The check valve (19121-206**) is located under right floorboard; the following procedure is to be used: 1. Remove: right front seat forward stop and seat, floor mat, right kick panel, right half of heel plate, metal cover over fuel selector, and right floorboard. 2. Remove check valve. 3. Disassemble the check valve, remove the internal check ball and reassemble the valve without the internal check ball. Remove the Bellanca placard showing the valve as P/N 19121-206 and remark the valve housing with a yellow stripe through the AN part number. 4. Reinstall the valve in the vapor return line. 5. Reverse Step 1 to complete operation.6. Make appropriate log book entries. C. 1974 and 1975 Model Aircraft* The check valve (19121-206**) is located under right floorboard; the following procedure is to be used: 1. Remove: right front seat forward stop and seat, floor mat, right cabin heat fresh air deflector, seat adjuster mechanism, metal cover over fuel selector, and right floorboard. 2. Remove check valve. 3. Disassemble the check valve, remove the internal check ball and reassemble the valve without the internal check ball. Remove the Bellanca placard showing the valve as P/N 19121-206 and remark the valve housing with a yellow stripe through the AN part number. 4. Reinstall the valve in the vapor return line. 5. Reverse Step 1 to complete procedure. 6. Make appropriate log book entries. *Identify aircraft model year by referring to first two digits in serial number for 1973, 1974 and 1975 aircraft; aircraft serial numbers prior to 1973 were not coded to year.**The check valve is a small 13/16 inch diameter by 1 3/4 inch long (plus fitting) AN valve located in the vapor return line between the fuel selector and the firewall. The valve may be positively identified by noting the Bellanca placard rework number 19121-206 located after AN in lieu of the AN number. Amendment 39-2209 became effective May 22, 1975. This amendment 39-2242 becomes effective June 25, 1975.
56-18-01: 56-18-01 CONVAIR: Applies to All 240 Series Aircraft. Compliance required as indicated. Instances of complete electric power system failure have occurred inadvertently upon failure of a single generator or engine. Due to high overloading that may be imposed on the remaining single generator, this generator is subject to failure unless prompt action is taken to reduce electrical loads within the generator's rated capacity. In order to improve electric power system reliability, the following shall be provided on aircraft in which a probable combination of electric utilization loads can exceed the continuous rating of one generator: 1. Generator Inoperative Warning Light (at least one, located for reliable warning), to be installed by April 1, 1957. 2. Monitoring System (add relays and circuitry to automatically disconnect buffet power and one inverter in case of loss or disconnection of one generator, with monitor override switch optional), to be installed by September 1, 1957. The automatic monitoring system is not required if it can be shown that the crew can manually reduce the total utilization load to the rating of one generator within 15 seconds after a generator or engine failure during any flight condition. A longer time interval than 15 seconds may be accepted if substantiated by demonstrations on representative generators which have reached approximately full overhaul time. (Convair Service Newsletter No. 352 dated June 16, 1956, contains preliminary technical information, including schematic wiring diagrams, relative to this same subject.) Final reworks must be in accordance with FAA-approved technical data.
68-01-04: 68-01-04 FOUND BROTHERS AVIATION LTD: Amendment 39-561. Applies to FBA-2C aircraft. Compliance required as indicated. To preclude the failure of the forward wing to fuselage attachment either by the failure of the attachment bolt or by the cracking of the wing root rib web around the attachment fitting, accomplish the following: (a) Replace the wing to fuselage forward attachment NAS 146-42 bolt with an unused bolt of the same part number or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, within the next 150 hours' time in service after the effective date of this AD, unless already accomplished within the last 350 hours' time in service, and thereafter at intervals not to exceed 500 hours' time in service from the last replacement. (b) Within the next 150 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter at intervalsnot to exceed 250 hours' time in service from the last inspection, visually inspect for cracks the wing root ribs repaired in accordance with either Found Brothers repair 2C39-18, Issue 2, or 2C39-19, Issue 2, or equivalent repair approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (c) Upon request with substantiating data submitted through an FAA maintenance inspector, the compliance times specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This amendment is effective January 9, 1968.
75-26-02: 75-26-02 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE, (S.N.I.A.S., formerly Sud Aviation). Amendment 39-2454. Applies to Aerospatiale Alouette III Helicopter Models SA-315B, SE-3160, SA-316B, SA-316C, and SA-319B, certificated in all categories, incorporating main rotor heads P/N's 3160S. 12.10.000.11 through .14, P/N's 3160S. 12.20.000.4 through .7, P/N's 3160S.12.10.000.1 through .10 modified in accordance with Modification No. S296-AM 1108 or Alouette Service Bulletin No. 65-52, or P/N's 3160S.12.20.000.1 through .3 modified in accordance with Modification No. S296-AM 1108 or Alouette Service Bulletin No. 65-52. Compliance is required as indicated, unless already accomplished. To prevent failure of the fixed levers of the main rotor head hydraulic drag dampers, accomplish the following: (a) Upon the effective date of this AD, and thereafter once on each day of operation, until accomplishment of paragraph (c) of this AD, visually inspect each of the three hydraulicdamper fixed levers for cracks in the area of the eccentric attachment hole. (b) If cracks are found in any hydraulic damper fixed levers, before further flight, replace the cracked hydraulic damper fixed lever with a serviceable unit of the same part number. (c) Within the next 100 hours' time in service after the effective date of this AD, remove the three hydraulic drag dampers from the main rotor head, inspect, rectify as necessary, and reinstall in accordance with subparagraph 1C of Lama Service Bulletin No. 65.15, dated September 23, 1974, for Model SA-315B, or subparagraph 1C of Alouette Service Bulletin No. 65.101, dated September 23, 1974, for the other designated models, or an FAA-approved equivalent of the applicable Service Bulletin. This amendment becomes effective December 23, 1975.
58-07-03: 58-07-03 VICKERS: Applies to All Viscount 700 Series Aircraft. Compliance required before accumulation of 6,000 flight hours. Investigations have proved it is necessary to replace the 1/2-inch diameter bolts securing the top inboard attachment fittings of the inner nacelles to the leading edge member at Station 96 at 6,000 hours. The P/N's of the bolts, which are to be replaced at 6,000 hours are as follows: 70103-4405 (Mod. D.1031 embodies); 80203-2405 (Mod. D.1327 or D.2025 embodied). Vickers Mod. D.2581 introduces redesigned nuts and bolts as direct replacement for the above bolts. This design ensures that any bending moments present will be taken by the full shank diameter of the bolts. The modified bolt assemblies are split pinned. Vickers-Armstrong has issued PTL 179, Issue 2, and Modification D.2581 covering this same subject. The British Air Registration Board considers this mandatory. The FAA concurs with this action and considers compliance therewith mandatory.
97-17-06: This amendment adopts a new airworthiness directive (AD), applicable to Bell Helicopter Textron, Inc. (BHTI) Model 214ST helicopters, that requires replacement of each emergency float inflation solenoid valve (valve). This amendment is prompted by two inadvertent inflations of emergency float systems that resulted from self-activations of the valves. The actions specified by this AD are intended to prevent self-activation of the valves, and subsequent inadvertent inflation of the emergency float system, which could lead to loss of control of the helicopter.
75-16-22: 75-16-22 DeHAVILLAND DH-114: Amendment 39-2298. Applies to all DeHavilland Model DH-114 airplanes modified in accordance with Supplemental Type Certificate (STC) SA1685WE. Compliance required within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 1500 hours' time in service, and thereafter at intervals not to exceed 1500 hours' time in service from the last inspection. To prevent excessive wear of the counterweight bushings and subsequent ineffectiveness of the counterweight function, accomplish the following: Inspect and replace, if required, crankshaft counterweight pins and bushings in accordance with Teledyne Continental Overhaul Manual X-30039 or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, ASO-210, P.O. Box 20636, Atlanta, Georgia 30320. This amendment becomes effective August 8, 1975.
69-13-01: 69-13-01 BRITTEN NORMAN LTD: Amdt. 39-783. Applies to Britten Norman Models BN-2 and BN-2A Aircraft Serial Numbers 3 through 43 and Serial Numbers 45 and 46. Compliance required as indicated unless already accomplished. To prevent a possible failure of the aileron, rudder, or nose wheel steering control system, accomplish the following: (a) Within the next 25 hours' time in service, inspect the threaded female portion of the fork-ends, P/N NB 45-B-879, of each turnbuckle assembly in the aileron, rudder, and nose wheel steering systems for evidence of thread defects in accordance with Britten-Normal Service Bulletin BN-2/SB.15, dated April 16, 1969, or later ARB-approved issue or later FAA-approved equivalent. (b) If the threaded female portion of any turnbuckle fork-end is found to be defective during the inspection required by paragraph (a), replace each defective fork-end with a serviceable fork-end of the same part number before further flight. This amendment becomes effective June 23, 1969.
54-26-01: 54-26-01 GRUMMAN: Applies to All Models G-44 and G-44A Aircraft. Compliance required by June 15, 1955. There have been reported numerous instances of the landing gear locking mechanism failing because of either hydraulic system leaks or failure of the mechanical locks. These malfunctions have been reported in both the up and down position of the landing gear. To prevent future similar malfunctions, provide a more positive means of holding the gear in its locked position, both in the fully extended and fully retracted positions. Grumman Service Bulletin No. 24, October 18, 1954, accomplishes this by providing a closed center hydraulic system. This arrangement provides hydraulic pressure to hold the gear in the selected position and unwanted extension or retraction is prevented even though the mechanical locks may fail or leaks develop in the hydraulic system. This supersedes AD 48-05-05.
52-11-02: 52-11-02 CONVAIR: Applies to All Model 240 Aircraft. Compliance required not later than the first major engine overhaul after February 1, 1953. To improve further the engine nacelle fire resistance of 240 aircraft, steel facings must be installed over certain aluminum alloy components of the engine cowl panels, the oil cooler duct, and the nacelle structure forward of the firewall. (Convair Service Bulletin No. 240-425, Revision 2, describes these changes in detail. Preliminary information on this modification is contained in Convairogram No. 30, dated April 8, 1952.)
68-06-03: 68-06-03 HAWKER SIDDELEY: Amdt. 39-566. Applies to Model DH. 125 airplanes, Series 1A, 1A/522 and 3A. Compliance required as indicated. To prevent a fully asymmetric flap condition in the lift dump position, within the next 150 hours' time in service after the effective date of this AD, replace the flap center hinge bolt, P/Ns 25CF71, 25CF1837, 25CF2387 and 25CF2357, with a self-retaining bolt, P/N 3110-7681, in accordance with Hawker Siddeley Service Bulletin 27-49-(1894) Revision 2, dated November 27, 1967, or later ARB-approved revision or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region. This amendment becomes effective April 20, 1968.
75-04-07: 75-04-07 PRATT & WHITNEY AIRCRAFT: Amendment 39-2084. Applies to all Pratt & Whitney Aircraft JT3D-3, JT3D-3B, and JT3D-7 turbofan engines containing tenth stage compressor disk, P/N 701810. Compliance required as indicated. To ensure adequate life limit margin for tenth stage compressor disk, P/N 701810, the cyclic life limits on these disks have been reduced below the figures currently approved. Unless already accomplished, remove from service the tenth stage compressor disk prior to exceeding the revised life limit listed below or within the next 25 cycles in service after the effective date of this AD, whichever comes later. Engine Model Previous Life Limit (Cycles) Revised Life Limit (Cycles) JT3D-3 30,000 25,000 JT3D-3B 30,000 25,000 JT3D-7 25,000 23,000 If a disk has been used in more than one engine model, the disk is limited to the lowest cyclic life permitted for the engine models in which it has been exposed. This amendment becomes effective February 19, 1975.