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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 C) All performance and operating data contained in the Owners Manual for these model airplanes are no longer applicable.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.