Results
71-18-02: 71-18-02 MCCAULEY: Amdt. 39-1274. Applies to Model 2AF34C55, 2AF34C55-A, B, C, D, E, F, G, H, HM, J, JM, K, KM, L, LM, M and N propellers. Compliance required as indicated, unless already accomplished. To prevent propeller counterweight assembly failures accomplish the following: (a) Propellers with 750 hours or more time in service as of the effective date of this airworthiness directive must be modified in accordance with paragraph (c) within the next 100 hours time in service. (b) Propellers with less than 750 hours in service as of the effective date of this airworthiness directive must be modified in accordance with paragraph (c) prior to the accumulation of 850 hours in service. (c) Modify propeller counterweight assembly in accordance with McCauley Service Bulletin Nos. 93 dated 21 April 1971 and 93-1 dated 30 June 1971 or later FAA approved revision or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, Eastern Region.NOTE: McCauley Service Bulletin 93-2 lists serial numbers of 2AF34C55N propeller hub assemblies which were in compliance with this airworthiness directive when shipped from the manufacturer. This amendment is effective August 31, 1971.
70-25-08: 70-25-08 PRATT & WHITNEY: Amdt. 39-1084. Applies to all Pratt & Whitney aircraft JT8D series turbofan engines which incorporate Part Number 500507 seventh stage compressor rotor disc with serial numbers listed in Pratt & Whitney Service Bulletin No. 2817, Rev. No. 1, dated 18 September 1970. Compliance required as indicated after the effective date of this airworthiness directive unless already accomplished. To preclude seventh stage compressor rotor disc failures as the result of suspected material deficiency, accomplish the following: 1. Replace discs with 2700 cycles or more in service within the next 30 cycles in service. 2. Replace discs with 2300 cycles, but less than 2700 cycles in service, within the next 100 cycles in service, but prior to accumulation of 2730 cycles. 3. Replace discs with 2000 cycles, but less than 2300 cycles in service, within the next 300 cycles in service, but prior to the accumulation of 2400 cycles. 4. Replace discs with less than 2000 cycles in service prior to the accumulation of 2300 cycles. The manufacturer's Service Bulletin identified and described in this directive is 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 this document from the manufacturer may obtain copies upon request from Pratt & Whitney Aircraft Division of United Aircraft Corporation, East Hartford, Conn. This document may also be examined at the FAA, Eastern Region, Federal Building, J. F. Kennedy International Airport, Jamaica, N.Y., 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 in the Eastern Region. This amendment is effective December 15, 1970 and was effective October 2, 1970, for all recipients of the telegram dated October 2, 1970, which contained this amendment.
71-17-02: 71-17-02 BOEING: Amendment 39-1268. Applies to Boeing Model 747 Series airplanes certificated in all categories. \n\n\tCompliance required as indicated. \n\tTo prevent body gear steer input on takeoff or landing, accomplish the following: \n\n\ta.\tWithin 100 hours' time in service after the effective date of this AD, unless already accomplished amend the Boeing B-747 Airplane Flight Manual, Certificate Limitations Section (used by each operator) by incorporating the following: \n\n\t"MISCELLANEOUS \n\tBODY GEAR STEERING. When aligned with the runway for takeoff and prior to advancing thrust levers, deactivate the body gear steering and leave deactivated until reaching taxi speed after landing, or refused takeoff." \n\n\tAs an interim acceptable procedure, deactivation of the body gear steering may be accomplished by disarming the circuit breaker identified on the circuit breaker panel as "Body Gear Steering Arm and Ind". \n\n\tb.\tOn or before January 1, 1972, unless already accomplished, install manually operated body gear steering arm/disarm switch per Boeing Service Bulletin 32-2113, dated July 23, 1971, or later FAA approved revision, or an equivalent installation approved by the Chief, Aircraft Engineering Division, FAA Western Region. Upon accomplishment of this installation, discontinue use of the circuit breaker to disarm the body gear per (a) above. The Airplane Flight Manual amendment per (a) above continues in full force and effect. \n\n\tThis amendment becomes effective August 17, 1971.
75-02-04: 75-02-04 GENERAL DYNAMICS: Amendment 39-2069. Applies to Models 22, 22M, 30, and 30A series airplanes, certificated in all categories. Compliance required as indicated. To reduce potential fire hazard existing in lavatory waste containers of General Dynamics Models 22, 22M, 30, and 30A series airplanes, accomplish the following: (a) Within 300 hours time in service from the effective date of this airworthiness directive, unless already accomplished within the last 1,000 hours, perform a thorough inspection of all electrical appurtenances, including wiring, terminal boxes, switches and hot water heaters physically located within lavatory waste container areas for wear, abrasion and corrosion. Repair or replace as necessary. (b) By December 31, 1975, unless already accomplished, modify the existing lavatory waste containers in accordance with the following General Dynamics Service Bulletins, as applicable, or later FAA-approved revisions, or in a manner approved bythe Chief, Aircraft Engineering Division, FAA Western Region: MODEL SERVICE BULLETIN NUMBER 22 25-102A 22M 25-25A 30, 30A 25-34A (c) Aircraft may be operated to a base for accomplishment of maintenance required under this airworthiness directive, per FAR's 21.197 and 21.199. This amendment becomes effective January 24, 1975.
67-06-05: 67-06-05 SCHEIBE-FLUGZEUGBAU: Amdt. 39-351 Part 39 Federal Register February 24, 1967. Applies to Model Bergfalke II/55 Gliders, all Serial Numbers, and Model Bergfalke III Gliders, Serial Numbers 5500 through 5586. Compliance required as indicated. To detect cracks on the aileron bell crank, Drawing No. 104B.41-S3, located in the outer wing section, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, unless already accomplished within the last 40 hours' time in service and thereafter at intervals not to exceed 50 hours' time in service from the date of the last inspection, visually inspect the aileron bell cranks, Drawing No. 104B.41-S3, in the left and right outer wing section for cracks using a mirror, and at least a two-powered glass. If cracks are detected during any of these inspections, comply with (b) of this AD. (b) Replace aileron bell cranks, Drawing No. 104B.41-S3, found cracked during the inspectionprovided for in (a) before further flight with a new reinforced aileron bell crank, Drawing No. 104B.41-S3-E1. (c) The requirements of this AD are not applicable to those gliders that have been equipped with aileron bell cranks, Drawing No. 104B.41-S3-E1, in the outer wing section. (Scheibe-Flugzeugbau Technical Information Bulletin I/66 pertains to this subject.) This directive effective March 1, 1967.
75-15-02: 75-15-02 HUGHES HELICOPTERS: Amendment 39-2259. Applies to all Hughes Model 269A, 269A-1, and 269B helicopters, certificated in all categories, and military TH-55A helicopters, as indicated herein. Compliance required as indicated. To prevent possible loss of control of the throttle and loss of control of engine power due to cracking and separation of aluminum throttle gear sector, P/N 269A7223 (Basic), with nominal 129 degrees of teeth, accomplish the following: (A) Within 25 hours additional time in service, after the effective date this AD, unless already accomplished: (1) Gain access to the throttle sector gears in the collective sticks at the lower throttle control housings and inspect through the housing cover plates, or by equivalent means, to determine if the gear sectors are steel or aluminum. Observe the nominal included angle of the aluminum gear sector teeth. NOTE: The manufacturer has identified three types of throttle gear sectors in the Hughes 269 helicopter which are unidentified by part number on the helicopter. The three gear sectors can be identified as follows: (a) The P/N 269A7223-3 gear sector is a steel bevel gear having a nominal included angle between the gear teeth of 304 degrees. The gear has a pre-drilled hole approximately .05 inches minimum distance from the boss end to accept a spring pin (roll pin) and a cotter pin for assembly with the mating left hand or right hand factory pre-drilled aluminum shaft. (b) A P/N 269A7223 (Basic) gear sector is identical to the -3 gear, except the gear has been fabricated from aluminum and has a nominal included sector angle between the gear teeth of 304 degrees. (c) A P/N 269A7223 (Basic) gear sector is an aluminum gear having a nominal included angle between the gear teeth of 129 degrees and has a pre-drilled hole approximately .05 inches minimum distance from the boss end to accept a roll pin and cotter pin during assembly with the mating left handor right hand factory pre-drilled aluminum shaft. (2) Remove the left hand (applicable to dual controls) and right hand aluminum gear sector assemblies (throttle sector gearshaft assemblies) from the helicopter which have an included angle between the gear sector teeth of 129 degrees and identify the gear sector portion per paragraph (A)(7). (3) Install new steel type throttle gearshaft assemblies having the following part numbers: P/N 269A7707-3(L.H.) or P/N 269A7707-7(L.H.) in the left throttle housing position and a P/N 269A7269-3(R.H.) in the right throttle housing position. (4) If the preceding part number throttle gearshaft assemblies cannot be procured from the manufacturer for compliance with this AD, accomplish the installation required in paragraph (A)(3) in accordance with the following procedure: (a) Procure a new P/N 269A7223-3 steel gear sector. (If a P/N 269A7223-3 steel gear sector cannot be procured from the manufacturer, only because of lackof availability of this part, the P/N 269A7223 (Basic) aluminum gear sector which has 304 degrees included angle between the gear sector teeth may be used as no service problems pertaining to cracking with this gear sector have been reported.) (b) Remove the P/N 269A7223 (Basic) aluminum gear sector having an included angle between the gear teeth of 129 degrees from the mating aluminum shaft by removing the cotter pin and the roll pin. The aluminum shaft is comprised of two different part number shafts, differing essentially in length. The P/N 269A7708(L.H.) shaft is located in the left throttle housing for dual control helicopters and a P/N 269A7271(R.H.) shaft is located on the right throttle housing. (c) Using a magnifying glass having at least 10X power, perform a close visual inspection of the shaft hole for cracks, corrosion, wear, scoring, hole elongation at the inside hole diameter which accepts the roll pin, or other defects. (d) Measure and record the left hand (if applicable) and right hand shaft hole inside diameter which accepts the roll pin. The acceptable dimensions for the inside hole diameter are .156 inches to .160 inches. (e) Measure and record the shaft outside diameter. The acceptable dimensions for the outside diameter of the left hand and right hand shaft are .6240 inches to .6250 inches. (f) Measure and record the existing inside diameter of the hole in the P/N 269A7223-3 steel sector gears and confirm that it measures .156 inches to .160 inches. (g) If the existing shaft and new sector gear are found acceptable, install the P/N 269A7223-3 steel gear sector on the mating aluminum shaft. Assure that the center line of the hole in the gear and the center line of the hole in the shaft are in accurate alignment. While maintaining alignment, press the P/N NAS561-5-14 roll pin through the gear and shaft hole with the chamfered end of the roll pin in the starting position and secure with a P/N AN381-3-20 cotter key. Seal with zinc chromate primer. If excessive force is required to press the roll pin in place, misalignment between the holes may have occurred and will require re-inspection for possible damage to the aluminum shaft hole. Tapering the roll pin beyond the existing chamfered end to allow for easy insertion of the roll pin is unacceptable. NOTE: The manufacturer has introduced a two piece optional roll pin configuration into production for the left hand throttle gearshaft assembly to improve the procedures for adjusting the AN932-2 pipe plug. Field fabrication of this dual roll pin design is not permissible in the field unless procedures are approved by the Chief, Aircraft Engineering Division, FAA Western Region. (h) Identify by ink stamping the letter "A" on the outer gear sector diameter of the gears. (i) Any shafts in stock which have not been pre-drilled by Hughes Helicopters may not be installed. Drilling of shafts in the field is not permissible unless special equipment and procedures are approved by the Chief, Aircraft Engineering Division, FAA Western Region. (5) Inspect on a one-time basis, those gear sectors which have been field assembled per paragraph (A)(4), after the accumulation of 250 hours additional time in service and before accumulating 300 hours additional time in service per the following procedures: Inspect the gearshaft assemblies for corrosion or other defects and determine if radial or axial play exists between the gear and the shaft. If any perceptible play exists, the shaft hole may be elongated or other defects may exist which will require disassembly of the gear from the shaft and inspection per paragraph (A)(4) or replacement per paragraph (A)(3). (6) Record inspections and modifications by paragraph numbers in compliance with this AD in the Aircraft Maintenance Records in accordance with FAR 91.173. (7) Identify in a conspicuous manner that the part is not serviceable to prevent inadvertent return to service, those throttle gear sectors, shafts, or throttle gearshaft assemblies that have been removed from service due to the provisions of this AD. (8) For additional procedures concerning removal, installation and inspection of throttle gearshaft assemblies refer to Hughes Model 269 Series Helicopter Basic Handbook of Maintenance Instructions, issued April 1, 1974, Revision No. 2, January 1974, or later revisions. Operators are cautioned to carefully observe the following manual requirements: (a) Proper shimming between the gear sector and shaft bearing to stay within allowable backlash limits. (b) Proper fit and seating of the bearings against the bore shoulder and proper application of loctite. (c) Proper determination that gears do not bind and checking for damage if binding occurs. (d) Proper tightening of the pipe plug for a push fit with zero play between rod and gearshaft. (e) Proper alighment of heel edge of the gear teeth on the sector gear and the pinion. (f) Proper rigging of the gear sector position in relation to the throttle pinion and the pilot's throttle grip. (B) Within 25 hours additional time in service, after the effective date of this AD, unless already accomplished, operators of: (1) Helicopters which, as of the effective date of this AD, have had the factory installed aluminum throttle gear sectors replaced with either steel or aluminum gear sectors as replacements; or (2) Helicopters which, as of the effective date of this AD, incorporate throttle gearshaft assemblies of any type which were assembled in the field with shafts drilled at other than the Hughes production facility; shall perform the inspections and replacements described in paragraphs (A)(3) through (A)(8), above. (C) Equivalent inspection and modification procedures for the throttle gear sector, shaft and gearshaft assembly may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (D) Aircraft may be operated to a base for accomplishment of that maintenance required by this AD, per FAR's 21.197 and 21.199. This amendment becomes effective July 15, 1975.
72-23-04: 72-23-04 NORTH AMERICAN ROCKWELL: Amdt. 39-1553. Applies to Sabreliner Model NA-265-40, Serial Numbers 282-1 thru 282-97; Model NA-265-50, Serial Number 287-1; and Model NA-265-60, Serial Numbers 306-1 thru 306-63. Compliance required within the next 100 hours' time in service, but not later than 90 days after the effective date of this A.D., whichever occurs first, unless already accomplished within the last 2300 hours' time in service, and thereafter at intervals not to exceed 2400 hours' time in service or 2400 landings, whichever occurs first, from the last inspection. To detect and remove cracks, tool marks, nicks, scratches, rust or other minor surface blemishes, inspect and rework the main landing gear outer strut cylinder and inner trunnion (cone shaped) surface in accordance with the inspection and rework provisions of North American Rockwell Sabreliner Service Bulletin No. 72-15 dated 30 October 1972 or later FAA-approved revision, or an equivalent procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region. This amendment becomes effective November 11, 1972.
66-24-03: 66-24-03 BRITISH AIRCRAFT: Amdt. 39-292 Part 39 Federal Register October 4, 1966. Applies to Model BAC 1-11 200 and 400 Series Airplanes. Compliance required as indicated. To prevent fatigue failures of the Belleville washer stack of the nose undercarriage up/down lock jack, P/N AB44A39 (200 Series) and P/N AK44A39 (400 Series), accomplish the following: (a) For airplanes with Belleville washer stacks, P/N AB44B67, with less than 2,300 landings on the effective date of this AD, remove stacks from service before the accumulation of 2,500 landings. (b) For airplanes with Belleville washer stacks, P/N AB44B67, with 2,300 or more landings on the effective date of this AD, remove stacks from service within the next 200 landings. (c) For airplanes with Belleville washer stacks, P/N AB44-1791, (with BAC Modification PM2437 washers) with less than 7,800 landings on the effective date of this AD, remove stacks from service before the accumulation of 8,000 landings.(d) For airplanes with Belleville washer stacks, P/N AB44-1791, (with BAC Modification PM2437 washers) with 7,800 or more landings on the effective date of this AD, remove stacks from service within the next 200 landings. (e) For 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 time from takeoff to landing for the airplane type. (British Aircraft Corporation (B.A.C.) Ltd. Alert Service Bulletin No. 32-A-PM2437 pertains to this subject.) This directive effective November 3, 1966.
75-17-06: 75-17-06 SIAI-MARCHETTI: Amendment 39-2307. Applies to Models F.260 and F.260B, all serial numbers, certificated in all categories. Compliance is required within the next 10 hours time in service after the effective date of this AD, unless already accomplished. To provide instructions for the correct operation of the emergency landing gear control, replace the "Emergency Landing Gear Extension" procedure set forth on page 9 of the airplane Flight Manual with a copy of the procedure set forth in SIAI Service Bulletin S.B.No.260B5, dated December 23, 1970, or an FAA-approved equivalent. Retain a copy of the procedure of Service Bulletin S.B.No. 260B5, or an FAA-approved equivalent, in the airplane Flight Manual until the procedure has been incorporated by the approved revision dated August 9, 1973 into the airplane Flight Manual. This amendment becomes effective August 15, 1975.
75-07-09: 75-07-09 MCDONNELL DOUGLAS: Amendment 39-2143. Applies to all McDonnell Douglas Models DC-10-10, DC-10-10F, DC-10-30, DC-10-30F and DC-10-40 Airplanes, certificated in all categories. \n\n\tCompliance required within the next 300 hours' time in service, after the effective date of this AD, unless already accomplished. \n\n\tTo prevent the possibility of electrical shock resulting from electrical short circuits to insufficiently grounded overhead and engineer's panels in the flight deck, install bonding jumpers in accordance with McDonnell Douglas Service Bulletin 24-45, Revision 2, dated December 2, 1974, and 24-62, dated July 15, 1974, or later FAA-approved revisions, or equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\tAn airplane may be flown to a base for the performance of the work required by this AD per FAR's 21.197 and 21.199. \n\n\tThis amendment becomes effective May 1, 1975.
58-18-01: 58-18-01 HAMILTON STANDARD: Applies to All Hamilton Standard Propellers Controlled by 5U18 Governors and Installed on TC18DA and TC18EA Series Engines. Compliance required at first governor overhaul after December 1, 1958, but not later than June 1, 1959. Adverse environmental conditions in the propeller governor resulting from certain types of engine failure have caused improper operation of the governor in such a manner as to result in propeller overspeeding, failure to feather and reversing when feathering was initiated. In order to minimize the possibility of such occurrences, provide a means for feathering that will be independent of the low pressure relief and pilot valves incorporated in the 5U18 governors. Installation of the Deterjet Model DJ-1025 governor bypass valve or replacement of the 5U18 governor by the Hamilton Standard 5AA22 governor are considered acceptable means to accomplish the desired objective. (installation of Deterjet Model DJ-1025 on Douglas DC-7C approved under Supplemental Type Certificate SA4-507.) (Deterjet Service Bulletin No.1 and Hamilton Standard Bulletin No. 561 cover this same subject.)
73-26-10: 73-26-10 GENERAL DYNAMICS: Amdt. 39-1765. Applies to all Models 22, 22M, and 30A airplanes certificated in all categories incorporating Hamilton Standard Electric Freon Packs Part Numbers 574056 or 573970 (Model 22), P/N 573971 (Model 22M), P/N 574057 (Model 30A). Compliance required as indicated. To prevent partial or total loss of generator power due to electric freon pack compressor motor failure, accomplish the following: (a) Model 22 airplanes. Within 50 hours' time in service after the effective date of this AD, unless already accomplished, replace the freon pack fuses with LPN 110 fuses, per paragraph 2B., Accomplishment Instructions, General Dynamics Service Bulletin 880 S.B. No. 24-56, dated November 21, 1973, or later FAA-approved revisions. Airplanes with existing fuses rated at 110 amperes or less need not accomplish the foregoing installation. (b) Models 22, 22M and 30A airplanes. Within 300 hours' time in service after the effective dateof this AD, unless already accomplished, modify the freon pack compressor motor protectors, P/N 573667, per paragraph 2A., Accomplishment Instructions General Dynamics Service Bulletins 880 S.B. No. 24-56, dated November 21, 1973 (for Model 22), 880M S.B. No. 24-15, dated November 21, 1973 (for Model 22M), and 990 S.B. No. 24-16, dated November 21, 1973, (for Model 30A), or later FAA- approved revisions to the bulletins. (c) Equivalent installations may be approved by the Chief, Aircraft Engineering Division, FAA Western Region, upon submission of adequate substantiating data. (d) Airplanes may be operated to a maintenance base under special flight permits per Sections 21.197 and 21.199 of the Federal Aviation Regulations for the purpose of accomplishing this AD. This amendment becomes effective February 1, 1974.
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.