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93-17-12: 93-17-12 BELL HELICOPTER TEXTRON, INC. (BHTI): Amendment 39-8683. Docket No. 92-ASW-49. Supersedes AD 92-11-07, Amendment 39-8257. Applicability: Model 204B, 205A, 205A-1, 205B, 212, and 412 helicopters, with main rotor transmission lower planetary spider (spider), part number (P/N) 204-040-785-003, installed, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent fatigue failure of the spider, which could result in failure of the main transmission and subsequent loss of control of the helicopter, accomplish the following: (a) Within the next 600 hours' time-in-service (TIS) after the effective date of this AD, unless previously accomplished within the last 2,500 hours' TIS, and thereafter at intervals not to exceed 3,100 hours' TIS from the last magnetic particle inspection (MPI), remove the spider and perform an MPI for cracks in accordance with the pertinent BHTI maintenance, repair, and overhaul manuals. (b) Replace any cracked spider with an airworthy part prior to further flight. (c) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used when approved by the Manager, Rotorcraft Certification Office, FAA, Rotorcraft Directorate. Operators shall submit their requests through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Rotorcraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Manager, Rotorcraft Certification Office. (d) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the helicopter to a location where the requirements of this AD can be accomplished. (e) This amendment becomes effective on July 29, 1994.
48-09-01: 48-09-01 DOUGLAS: Applies to C-54 and DC-4 Aircraft. \n\n\tTo be accomplished not later than April 15, 1948. \n\n\tBecause of the hazards involved, the transfer of fuel between tanks must be prohibited. The following placard shall be installed in the cockpit in full view of the pilots: \n\n\t"Fuel cross-feed system not intended for transferring fuel from one tank to another and should not be used for this purpose. When using crossfeed system, turn off tank(s) not in use." \n\n\tIn addition to the placard, the FAA Approved Flight Manual must be revised to incorporate proper fuel system operation procedures in accordance with the above placard. Approved Flight Manual pages may be obtained from the airplane manufacturer.
63-15-02: 63-15-02 DE HAVILLAND: Amdt. 588 Part 507 Federal Register July 23, 1963. Applies to All Model DH 114 Aircraft. Compliance required within 200 hours' time in service after the effective date of this AD. Because of instances of fatigue failure of the bolt, P/N 4W.1121, which attaches the flap jack attachment shackle to the wing, replace the bolt with a new special bolt P/N 14W.5835 in accordance with a de Havilland Heron Modification No. 1498. (de Havilland TNS Series Heron (114) No. C.F.7 Issue 1 dated December 31, 1962, and de Havilland Modification News Sheet No. Heron 1498 dated January 1, 1963, cover this subject.) This directive effective August 22, 1963.
2008-21-04: The FAA is adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation Model 390 airplanes. This AD requires you to modify the cabin barometric pressure switch and cabin altitude high switch installations and perform a functional test of the switches and related systems. This AD results from the possibility of barometric pressure switch electrical connections being incorrectly connected or inadvertently disconnected. We are issuing this AD to modify the cabin barometric pressure switch and cabin altitude high switch to prevent them from becoming incorrectly connected or inadvertently disconnected, which may result in no CABIN ALT HI annunciation in the cockpit and no automatic deployment of the cabin oxygen masks. This failure could lead to incapacitation of the crew due to hypoxia with possible inability to control the airplane.
47-22-01: 47-22-01 LUSCOMBE: (Was Mandatory Note 12 of AD-694-4.) Applies Only to Model 8 Series Aircraft Equipped With Edo 60-1320 Floats. Compliance required immediately if possible but in any event not later than August 1, 1947. All seaplanes should be inspected to determine whether the bulkhead reinforcements of Luscombe Drawing 48701 are presently installed at fuselage Station 4 (rear float strut attachment. If not, those reinforcements shown on Luscombe Drawing 58730 must be installed to insure the structural integrity of the float installation. Each seaplane should also be inspected to determine conformity of Model 8A with Luscombe Drawing 58700 and Models 8C and 8D with Luscombe Drawing 58725
48-02-04: 48-02-04 DOUGLAS: Applies to DC-6 Serial Numbers 42854 Through 42880; 42882 Through 42896; 43000 Through 43017; 43035 Through 43038; 43055; 43056; 43062, and 43063. \n\nCompliance required by the next No. 3 inspection. \n\nTo prevent the brake lining from becoming wedged between brake disc and housing, replace the present adjustment pin Goodyear P/N 511940-1 and spring plate Goodyear P/N 512139 by the single piece adjusting pin Goodyear P/N 9510744. \n\n(Douglas Service Bulletin DC-6 No. 90 covers this same subject.)
47-13-02: 47-13-02 TAYLORCRAFT: (Was Service Note 1 of AD-696-3.) Applies to Models BC-65, BCS-65, BC12-65, BCS12-65, BC12-D, BCS12-D, BCS12-D1 Aircraft. Inspection required each 25 hours of engine operation. This inspection applies only to fuel hose bearing white dash lines and having end fittings marked "CAA, SNA (date)". Examine the two flexible fuel lines to determine whether the hose inner liner has collapsed or failed thus causing a restriction to the flow of fuel. Particular attention should be given to the hose close to the fittings on the fuel strainer. Defective hose appears soft or spongy when squeezed with the fingers. Any defective hose is to be replaced immediately. (This information is contained, in part, in Taylorcraft Service Bulletin No. 60 dated June 14, 1946.)
48-02-05: 48-02-05 DOUGLAS: Applies to DC-6 Serial Numbers 42854 Through 42880; 42882 Through 42884; 43000 and 43001. \n\nCompliance required by the first engine change after March 1, 1948. \n\nTo prevent the hot exhaust burning through the exhaust stack recess sheet on the upper and lower outboard accessory cowling, remove the present shield on the inboard side of the cowling and install a screw fastened exhaust chute of 0.042-thickness corrosion resistant steel sheet on the outboard side of the recess sheet. An air gap must exist between the exhaust chute and the recess sheet to allow a flow of ram air for heat dissipation. \n\n(Douglas Service Bulletin DC-6 No. 30 covers this same subject.)
2008-21-05: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 767-200, -300, and -400ER series airplanes. That AD currently requires an inspection to determine if the door-mounted escape slide/rafts have certain part numbers. This new AD does not retain that requirement. This new AD continues to require an inspection for excessive tension of the firing cable, and procedures for providing slack in the firing cable or rerouting the firing cable if necessary. For certain airplanes, this new AD also requires a review of the airplane maintenance records to determine if a certain service bulletin has been incorporated, or an inspection to determine if certain door-mounted escape slide/rafts are installed. This new AD also requires modification of certain escape slide/rafts. This AD results from reports of uncommanded inflation inside the airplane of a door-mounted escape slide/raft located in the passenger compartment. We are issuing this ADto prevent injury to maintenance personnel, passengers, and crew during otherwise normal operating conditions and to prevent interference with evacuation of the airplane during an emergency, due to uncommanded inflation of a door-mounted escape slide/raft. \n\n\nDATES: This AD becomes effective November 13, 2008. \n\tThe Director of the Federal Register approved the incorporation by reference of a certain publication listed in the AD as of November 13, 2008. \n\tOn June 30, 2005 (70 FR 34638, June 15, 2005), the Director of the Federal Register approved the incorporation by reference of a certain other publication.
70-05-05: 70-05-05 AMERICAN AVIATION: Amdt. 39-950. Applies to Model AA-1 aircraft, Serial Nos. AA-1-0001 through AA-1-0159. Compliance required within the next 10 hours in service after the effective date of this Airworthiness Directive, unless already accomplished, and thereafter at intervals not to exceed 50 hours in service from the last inspection, except as provided in paragraph 3. To preclude the possibility of exhaust fumes from entering the cabin heat system due to undetected cracks in the muffler, accomplish the following: 1. Inspect muffler and shroud assembly (P/N 14-504001) for cracks particularly in the area adjacent to all welds inside the shroud at the transition between the muffler and the tailpipe. If visual inspection is not possible, pressure test for leaks in accordance with AC 43.13- 1, Chap 14, Section 3, paragraph 287D. If cracks are found in the muffler tailpipe or the muffler shroud they should be repaired by an inert gas-shielded arc welding process such as Heliarc. (Muffler material is AISI 321 stainless steel and shroud material is AISI 304 stainless steel). Accomplish above inspection and necessary repairs in accordance with Advisory Circular 43.13- 1, Chap. 14, Section 3, paragraph 387 and 388. 2. Check alignment between rigid brace P/N 503008-501 and tailpipe to insure that tailpipe is not stressed when brace is installed. 3. The repetitive 50 hour inspection requirement of the exhaust system may be omitted if the aircraft has been altered by installation of a new muffler, Turbo system P/N 099001-113, with the rigid brace. (American Aviation Service Bulletin No. 116 dated 9 January 1970 covers this same subject) This amendment is effective March 13, 1970.
46-41-01: 46-41-01 BELLANCA: (Was Mandatory Note 2 of AD-773-5.) Applies to Models 14-13, 14-13-2 Serial Numbers 1060 to 1111, Inclusive. Compliance required prior to November 15, 1946. Replace rudder bellcrank (Bellanca P/N 9817) located at the left and right ends of the rudder torque tube with parts furnished by the manufacturer which are stamped "heat-treat" in ink. (Bellanca Service Bulletin No. 2 dated August 26, 1946, covers this same subject.)
71-24-09: 71-24-09 BEECH: Amendment 39-1347. Applies to Model 56TC (Serial Numbers TG-1 thru TG-76) Airplanes. Compliance: Required as indicated, unless already accomplished. To provide information reflecting applicable operating limitations and margin between maximum structural cruising speed and never exceed speed, within the next 50 hours' time in service after the effective date of this AD, revise placards and change airspeed indicator marking as follows: 1) Install placard at lefthand cabin side, adjacent to ignition switch panel reading: "This airplane must be operated as a Normal Category airplane in compliance with the operating limitations stated in the form of placards, markings and manuals (Pilot's Check List). Occupied seats must be in upright position during takeoff and landing. Maximum weight 5990 lb. No acrobatic maneuvers including spins approved. Max. speed w/landing gear extended (normal) (TG-1 thru TG-71) - 165 m.p.h. (143 knots) (TG-72 and up) - 175 m.p.h. (152 knots) Max. speed with flaps extended (15 degrees down) - 175 m.p.h. (152 knots) Max. speed with flaps extended (normal) - 144 m.p.h. (125 knots) Max. design maneuver speed - 183 m.p.h. (159 knots) Minimum control speed single engine - 97 m.p.h. (84 knots) Max. structural cruising speed (S.L. to 20,000 ft. alt.) - 233 m.p.h. (202 knots) Max. structural cruising speed (25,000 ft. alt.) - 222 m.p.h. (193 knots) Max. structural cruising speed (30,000 ft. alt.) - 214 m.p.h. (186 knots) Never exceed speed (S.L. to 20,000 ft. alt.) - 262 m.p.h. (227 knots) Never exceed speed (25,000 ft. alt.) - 249 m.p.h. (216 knots) Never exceed speed (30,000 ft. alt.) - 240 m.p.h. (208 knots) 2) Install placard on floating instrument panel near airspeed indicator reading: "See limitations placard for 'max. structural cruise' and 'never exceed limits'." 3) Re-mark airspeed indicator to extend yellow arc from 240 m.p.h. to 233m.p.h. so that green arc does not enter this range. Beechcraft Service Instruction No. 0173-016 considers this subject. This amendment becomes effective November 30, 1971.
63-03-01: 63-03-01 BOEING AND DOUGLAS: Amdt. 532 Part 507 Federal Register February 5, 1963. Applies to Boeing Models 707-100B, 707-300B, and 720-000B Series Aircraft, and to Douglas DC-8-50 Series Aircraft With Pratt & Whitney JT3D Series Engines. \n\n\tCompliance required within the next 4,000 hours' time in service after the effective date of this AD, unless already accomplished. \n\n\tClogging of engine main oil filters by foreign matter has caused lubrication system malfunctions which have resulted in engine mechanical failures affecting safety of flight. To prevent such failures, accomplish the following: \n\n\t(a)\tFor Pratt & Whitney JT3D Series engines with serial numbers listed in Pratt & Whitney Engine Service Bulletin No. 327 dated January 8, 1962: \n\n\t\t(1)\tModify the engine oil filter assembly to provide for the installation of a differential pressure switch between the bypass port and the filter drain port, and provide an additional spring in the bypass valve to increase the pressure at which bypass occurs, in accordance with Service Bulletin No. 327, or FAA approved equivalent. \n\n\t\t(2)\tInstall a pressure switch across the engine main oil system filter, set to be actuated when the differential pressure between the inlet and outlet ports reaches a value of approximately 50 p.s.i. This change shall be accomplished in accordance with Boeing Service Bulletin No. 1586 dated April 11, 1962, for Boeing aircraft, and in accordance with Douglas Service Bulletin No. 79-11 (to be issued later) for DC-8 aircraft, or FAA approved equivalent. Prior or concurrent incorporation of (a)(1) is required with this change. \n\n\t(b)\tFor Boeing Models 707-100B, 707-300B, and 720-000B Series aircraft with serial numbers listed in Boeing Service Bulletin No. 1586 dated April 11, 1962, and for Douglas DC-8-50 Series aircraft listed in Douglas Service Bulletin DC-8 No. 79-11 (to be issued later): \n\n\t\t(1)\tProvide means in the cockpit to give corresponding indication of the actuationof the differential pressure switch on each engine in accordance with Boeing Service Bulletin No. 1586 for Boeing aircraft, and in accordance with Douglas Service Bulletin 79-11 for DC-8 aircraft, or FAA approved equivalent. \n\n\tNOTE: Any person may submit an equivalent means of compliance with the objective of this directive. Such equivalent means shall be submitted to FAA, Western Region, Attention, Chief, Engineering and Manufacturing Branch, for evaluation and approval. Adequate substantiation of equivalency will be required. If approved, the equivalent means, when accomplished, shall be deemed as compliance with (a) and (b). The objective of this directive is to provide means of preventing serious mechanical damage to engines which would affect safety of flight as a result of lubrication failure of engine main bearings. \n\n\t(c)\tWhen the modifications prescribed in (a) and (b) are accomplished or when an equivalent means of compliance is approved and accomplished, the engine oil filter inspections prescribed by AD 61-24-01 are no longer required. \n\n\t(d)\tAppropriate revisions to the FAA Airplane Flight Manual covering procedures required in connection with devices installed shall be prepared and submitted to FAA, Western Region, Attention, Chief, Engineering and Manufacturing Branch, for approval. \n\n\tThis directive effective March 7, 1963.
62-27-04: 62-27-04 DOUGLAS: Amdt. 520 Part 507 Federal Register December 20, 1962. Applies to DC-8 Standard Leading Edge Aircraft Powered With Pratt & Whitney JT3C, JT4A or Conway Engines. \n\n\tNOTE: Does not apply to aircraft with extended leading edge and to JT3D powered aircraft with standard leading edge. \n\n\tCompliance required as indicated. \n\n\tAs a result of failure of the upper inboard spar cap structure of the outboard pylons, accomplish the following: \n\n\t(a) Unless already accomplished within the last 425 hours' time in service, within the next 25 hours' time in service, inspect upper inboard spar cap structure of the outboard pylon for evidence of cracks. Gain access to the area to be inspected by removing the pylon leading edge nose cap between Station YOP 214 and 244 and access doors numbers 110, 113, 411, and 414. Using close visual or dye penetrant methods, inspect the upper inboard cap and adjacent structure for cracks in the area of Station YOP 230 and at the edgesof support fitting P/N 3647306-501. \n\n\t(b) If cracks are found, repair in accordance with Douglas Drawing 5776811 or FAA approved equivalent, prior to further flight. \n\n\t(c) If no cracks are found the inspections outlined in (a) must be repeated at periods not to exceed 500 hours' time in service from the last inspection. \n\n\t(d) The repetitive inspections may be discontinued on aircraft repaired in accordance with Douglas Drawing 5776811 and on aircraft modified to incorporate preventive rework accomplished in accordance with FAA engineering approved technical data. \n\n\t(e) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Douglas DC-8 Alert Service Bulletin A54-33, Revision No. 2, dated January 24, 1964, covers this same subject.) \n\n\tThis directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated November 21, 1962. \n\n\tRevised June 8, 1963. \n\n\tRevised June 23, 1964.
67-01-03: 67-01-03 PRATT & WHITNEY: Amdt. 39-336 Part 39 Federal Register January 4, 1967. Applies to Model JT4A Series Turbojet Engines. Compliance required as indicated unless already accomplished. To prevent failure of the fuel manifold assembly, accomplish the following: (a) Within the next 3,400 hours' time in service after the effective date of this AD, inspect all P/N's 378155, 378156, 391957, 391959, 447330 and 447339 fuel manifold assemblies for cracks using the fluorescent penetrant inspection procedures outlined in Pratt & Whitney Aircraft JT4A Overhaul Manual. (1) If cracks are found, replace before further flight the fuel manifold assembly with one of the fuel manifold assemblies listed above or with a P/N 572766 or 572767 fuel manifold assembly. (2) If no cracks are found, glass bead peen all cluster tee fillets (eight places per each fuel manifold assembly) prior to return to service and thereafter at each fuel mainfold assembly overhaul in accordance with Pratt& Whitney Aircraft JT4A Overhaul Manual Temporary Revision No. 73-9 dated October 20, 1966, or Pratt & Whitney Aircraft JT4A Overhaul Manual Revision No. 39 which includes Temporary Revision No. 73-9. (b) At each fuel manifold assembly overhaul, inspect and glass bead peen in accordance with (a) all fuel manifold assemblies P/N's 572766 and 572767. NOTE: Fuel Manifold Assemblies P/N's 572766 and 572767 were glass bead peened during initial fabrication. (c) Prior to use, inspect and glass bead peen in accordance with (a) all spare fuel manifold assemblies P/N's 378155, 378156, 391957, 391959, 447330, and 447339. (Pratt & Whitney Aircraft letter dated March 23, 1966, to all operators of JT4A turbojet engines covers this subject.) This directive effective January 5, 1967.
62-27-02: 62-27-02 BELL: Amdt. 522 Part 507 Federal Register December 28, 1962. Applies to All Model 47 Series Helicopters Equipped With P/N 47-642-020-1 Wood Tail Rotor Blades. Compliance required as indicated. There have been several failures of wood tail rotor blades resulting from wood deterioration. To preclude further wood blade failures the following must be accomplished: (a) Within 50 hours' time in service after the effective date of this AD: (1) Remove wood tail rotor blades in accordance with the applicable Bell Maintenance and Overhaul (M&O) Manual. (2) Remove the fiberglass wrapping from the root end area of blades and remove the fiberglass blade covering from areas underneath the wrapping in accordance with the applicable Bell M&O instructions for repair of wood main rotor blades. Cut blade covering by lightly sanding cover as a knife or other sharp instrument can cause damage. (3) Inspect root end of blades from root end of blade to 6 inchesoutboard for: (i) Elongated bolt holes. Maximum allowable diameter 0.260 inch. (ii) Decay of wood. Detection of decay can be made visually by noting discoloration of the basic material. (Generally decay will start as a grayish discoloration and deepens to a brown color during the later stages.) (iii) Cracks in the stainless steel leading edge strip and grip plates using at least a 5-power magnifying glass. (4) Blades found with bolt hole diameters exceeding 0.260 inch, with decay, or with any cracks, shall be removed from service prior to further flight. (5) Blades without defects may be returned to service after: (i) Recovering the blade root area in accordance with patching procedures given in the applicable Bell M&O Manual; and (ii) Rewrapping the root area with two pieces of MIL-P-8013 No. 181 fiberglass cloth 2 x 27 inches in accordance with Bell Service Bulletin No. 75 dated September 17, 1951. (b) Blades returned to serviceafter compliance with (a) shall be retired from service prior to the accumulation of 200 hours' time in service since reinstallation in accordance with (a). This directive effective January 29, 1963.
68-02-04: 68-02-04 FAIRCHILD-HILLER: Amendment 39-544. Applies to Type FH-1100 Helicopters, Serial Numbers 9 through 49. Compliance required as indicated. To prevent fatigue failures of the Cyclic Input Swashplate Ring, P/N 24-34205-3, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, unless already accomplished, and thereafter at intervals not to exceed 25 hours' time in service from the last inspection, visually inspect the cyclic input swashplate ring, P/N 24-34205-3, in accordance with Part A (excluding paragraph 5) of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967, or later revisions approved by the Chief, Engineering & Manufacturing Branch, Federal Aviation Administration, Eastern Region. Equivalent inspections may be approved by an FAA maintenance inspector. (b) If a crack is found, remove the ring from service prior to further flight. (c) Accomplish the following on rings that have not been reworked in accordance with Part B of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967: (1) Remove from service or rework in accordance with Part B of the aforementioned Letter rings with 75 or more hours time in service on the effective date of this AD within the next 25 hours' time in service. (2) Remove from service or rework in accordance with Part B of the aforementioned Letter all other rings before the accumulation of 100 hours' time in service. (d) Rings which have been modified in accordance with Part B of Fairchild Hiller Service Information Letter No. 2 dated August 11, 1967, or in accordance with any other method approved by the Chief, Engineering & Manufacturing Branch, Federal Aviation Administration, Eastern Region may be continued in service until the accumulation of 750 hours' time in service. The 25-hour repetitive inspection of (a) may be discontinued on modified rings when a satisfactory inspection for crackshas been accomplished on the ring after it has been modified, by the dye penetrant method or an equivalent approved by an FAA maintenance inspector. This AD is effective January 27, 1968.
62-25-02: 62-25-02 DOUGLAS: Amdt. 510 Part 507 Federal Register November 28, 1962. Applies to All Models DC-6 and DC-7 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tDue to failure of a main gear shock strut cylinder and numerous cases of cracks in the 0.125-inch radii next to the torque link lugs on the cylinders, and on the piston tube axle fittings, the following shall be accomplished. \n\n\t(a) DC-6 Series Aircraft. \n\n\t\t(1) With 30,000 or more hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter within each 300 hours' time in service from the last inspection. \n\n\t\t(2) With less than 30,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) priorto the accumulation of 30,200 hours' time in service, and thereafter within each 300 hours' time in service. \n\n\t(b) DC-7 Series aircraft. \n\n\t\t(1) With more than 15,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 100 hours' time in service, and thereafter within each 300 hours' time in service from the last inspection. \n\n\t\t(2) With less than 15,000 hours' time in service as of the effective date of this AD shall be inspected and reworked in accordance with paragraphs (c) and (d) prior to the accumulation of 15,200 hours' time in service, and thereafter within each 300 hours' time in service. \n\n\t(c) Inspect, using dye penetrant, or magnetic particle, or FAA approved equivalent, for cracks in the 0.125-inch radii at the edges of the torque link lugs in the main landing gear shock strut cylinder and the piston tube axle fitting. \n\n\t(d) If cracks are found, they may be removed by reworking the 0.125-inch radius in accordance with the instructions contained in Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto or Douglas Service Engineering Letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D and 534E attached thereto. If cracks cannot be removed without exceeding limits specified in the Douglas sketches, the gear must be replaced prior to further flight. Parts that can be reworked, and those in which no cracks are found, must be repainted with zinc chromate primer and aluminized lacquer before they are returned to service. \n\n\t(e) When the 0.125-inch radii at the edges of the torque link lugs on the strut cylinders and axle fittings have been enlarged to 0.250-inch radii, holding the tolerances described in Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto, or Douglas Service Engineering letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D and 534E attached thereto, and the parts are refinished as described in (d), the repetitive inspections required herein may be discontinued. \n\n\t(f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering & Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Douglas Service Engineering letter C1-78-M1281/DJW dated April 20, 1962, and sketches 498A and 498B attached thereto, or Douglas Service Engineering letter C1-78-140/DJW, dated January 30, 1963, and sketches 534A, 534B, 534C, 534D, and 534E attached thereto, covers thissame subject.) \n\n\tThis directive effective November 28, 1962. \n\n\tRevised April 4, 1963.
2008-17-13: We are adopting a new airworthiness directive (AD) for certain Boeing Model 737-100, -200, -200C, -300, -400, and -500 series airplanes. This AD requires replacing the existing straight-to-90- degree hose assembly for the Lavatory "A'' water supply. The replacement is a new straight hose assembly and a separate 90-degree elbow fitting. This AD results from a report of a separated hose assembly for the passenger water system. We are issuing this AD to prevent a water leak into the flight deck ceiling, which could result in an electrical short and possible loss of several functions essential to safe flight.
2008-17-19: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: One ATR 42-300 experienced a collapse of the Right (RH) Main Landing Gear (MLG) when taxiing, caused by failure of the side brace assembly. Investigations revealed a crack propagation that occurred from a corrosion pit, in a very high stressed area of the upper arm. * * * * * * * * The unsafe condition is cracking of the upper arms of the secondary side brace assemblies of the MLG, which could result in collapse of the MLG during takeoff or landing, damage to the airplane, and possible injury to the flightcrew and passengers. We are issuing this AD to require actions to correct the unsafe condition on these products.
62-26-05: 62-26-05 PIPER: Amdt. 511 Part 507 Federal Register December 5, 1962. Applies to All Models PA-24 and PA-24 "250" Aircraft. Compliance required within the next 50 hours' time in service after the effective date of this AD, unless already accomplished. To preclude hazardous carbon monoxide contamination in the cockpit and engine power loss, caused by cracked muffler heater shrouds, accomplish the following: (a) For aircraft Serial Numbers 24-1 to 24-2298 inclusive, equipped with channel reinforced muffler P/N's 22594-00, 22594-02 on PA-24, and P/N's 22593-00, 23159-00 on PA-24 "250" installed as service replacements: (1) Remove the tail pipe, the right-hand exhaust stack, and carburetor heat shroud and inspect for cracks and hot spots. Pay particular attention to the junction of rear cylinder exhaust tube with the stack assembly. (2) Remove the muffler and muffler heater shroud. Carefully inspect the muffler for visible cracks, particularly in the areanear the tail pipe opening and examine the internal baffle and perforated tube. Submerge the muffler in water and pressure test at 10 p.s.i. (3) Replace the muffler prior to further flight if cracks, heat deterioration, defects, or wrinkles formed in the perforated tube are observed or if leaks are detected during the pressure test. (4) Rework the muffler heater shroud by: (i) enlarging the opening in the shroud in accordance with the Piper template; (ii) installing the muffler reinforcement tube, P/N 23482-00 using 20 rivets PDR 134A-6, or FAA approved equivalent; and (iii) installing cover plate P/N 23498-00 using 11 rivets AN 426A3-4, or FAA approved equivalent, in accordance with Piper Immediate Action Service Bulletin No. 210 (Kit P/N 754 484). (5) Reinstall the muffler exhaust stacks, tailpipe, and air ducts on the airplane. (b) For aircraft Serial Numbers 24-2299 to 24-3284 inclusive, equipped with channel reinforced muffler P/N's 22594-00, 22594-02 on PA-24, and P/N's 22593-00, 23159-00 on PA-24 "250", except aircraft Serial Numbers 24-2876, 24-2929, 24-2949, 24-2967, 24-2990, 24-3033, 24-3095, 24-3114, 24-3130, 24-3150, 24-3155, 24-3173, 24-3191, 24-3193, 24-3194, 24-3196, 24-3198, 24-3203, 24-3204, 24-3222, 24-3233, 24-3234, 24-3241, 24-3244, 24-3248, 24-3254, 24-3257, 24-3258, 24-3265, 24-3268, 24-3270, 24-3273, 24-3274, 24-3276, 24-3277, 24-3278, 24-3279, 24-3280, 24-3282, 24-3283, which have been modified: (1) Perform inspections required by (a)(1) and (a)(2), and the replacement required by (a)(3), if necessary. (2) Install new cabin heater shroud, P/N 23507-00 on PA-24, and P/N 23489-00 on PA-24 "250". Center the tailpipe in the shroud tailpipe opening. (3) Reinstall the muffler exhaust stacks, tailpipe, and air ducts on airplane. NOTE: PA-24 and PA-24 "250" mufflers have been manufactured incorporating two different styles of tailpipe reinforcement brackets. This AD requires modification of one style only - those with channel style reinforcement. See Sketch A of Piper Service Bulletin No. 210 for further identification. Both types of mufflers have been sold as service replacements. It will therefore be necessary to examine aircraft Serial Numbers 24-1 to 24-2587 inclusive, if the original muffler has been replaced, to determine if the modification is required. Aircraft Serial Numbers 24-2588 through 24-3284, were manufactured with the channel shaped reinforcement and will require modification except those already modified as indicated. (Use Piper Service Letter No. 324B as a guide for inspections in addition to Service Bulletin No. 210.) This directive effective December 5, 1962.
62-24-04: 62-24-04 DOUGLAS: Amdt. 504 Part 507 Federal Register November 8, 1962. Applies to All Model DC-8 Aircraft With Midwing Flap Actuating Cylinder Douglas P/N 3643685. These Cylinders Can Be Identified as Having an Outside Diameter of 3.810-3.820 Inches at the Forward End of the Barrel Where the Cylinders Attach to the Wing Flap Crank. \n\n\tCompliance required as indicated. \n\n\t(a)\tOn aircraft incorporating flap travel limit stops per paragraphs (a) and (b) of AD 62-16-02, the wing flaps shall be lowered with hydraulic pressure from the auxiliary pump and a close visual inspection of the forward 1/4 inch of length and around the entire periphery of both midwing flap actuating cylinder barrels shall be made daily for evidence of cracks or fluid leakage. Barrels exhibiting leakage or evidence of cracks shall be replaced prior to further flight. \n\n\t(b)\tOn aircraft not incorporating flap travel limit stops per AD 62-16-02, the inspections prescribed by (a) shall be accomplished prior to each flight. \n\n\t(c)\tThe inspections prescribed by (a) and (b) may be discontinued when the midwing wing flap cylinders P/N 3643685 are inspected and reworked in the manner described in Figure 1 of Douglas DC-8 Service Bulletin No. 27-134 for the outboard wing flap cylinders. Midwing wing flap cylinders inspected and reworked by operators in this manner will be subject to the inspection requirements prescribed for the outboard wing flap cylinders by paragraph 1.D(2) of Service Bulletin 27-134. Cylinders reworked by the operator shall in addition to the identification prescribed by Service Bulletin 27-134, be further identified by a color code or FAA approved equivalent. \n\n\t(d)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, May adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Douglas DC-8 Service Bulletin No. 27-134 covers this same subject.) \n\n\tThis directive effective upon publication in the Federal Register for all person except those to whom it was made effective immediately by telegram dated October 19, 1962.
62-02-04: 62-02-04 DOUGLAS: Amdt. 392 Part 507 Federal Register January 24, 1962. Applies to DC-8 Aircraft Serial Numbers 45252-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45419, 45421-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45614, 45616-45628 and 45636-45638. \n\n\tCompliance required as indicated. \n\n\tTo prevent aileron tab lockout mechanism bracket assembly failure, resulting in partial or complete loss of control force to one aileron, the following shall be accomplished: \n\n\t(a)\tAt periods prescribed in (b), conduct close visual inspection, using low-power magnifying glass or equivalent means, for evidence of cracking of the left and right side aileron tab lockout bracket assemblies, P/N 4643350. The critical areas to be inspected are shown in Douglas Service Bulletin 27-115, Figure 1, Step 3. Any evidence of cracking shall be verified by dye check or equivalent means, with the tab lockout cylinder disconnected from the bracket assembly, within the next 10hours' time in service following the detection of such evidence of cracking. Any part found to be cracked shall be replaced prior to further flight with an assembly of the same part number which has been inspected in accordance with the provisions of this paragraph and found to be free of cracks or with assembly P/N 3773970-1. \n\n\t(b)\tThe initial and repetitive inspections of assemblies, P/N 4643350, shall be conducted at the following times: \n\n\t\t(1)\tOn assemblies which have accumulated a total time in service of less than 3,000 hours as of the effective date of this AD: Initial inspection within next 350 hours' time in service, but in no event to exceed 3,100 hours' assembly total time in service; repetitive inspections thereafter at intervals not to exceed 350 hours' time in service except that after the assembly total time in service reaches 3,000 hours the repetitive intervals shall not exceed 100 hours' time in service. \n\n\t\t(2)\tOn assemblies which have accumulated a total time in service of 3,000 hours or more as of the effective date of this AD: Initial inspection within next 100 hours' time in service; repetitive inspections thereafter at intervals not to exceed 100 hours' time in service. \n\n\t(c)\tWhen assembly, P/N 3773970-1 is installed in place of P/N 4643350, the repetitive inspections may be discontinued. \n\n\t(d)\tWhen assembly P/N 4643350 is replaced with an assembly of the same part number which has been inspected in accordance with (a) and found to be free of cracks, the replacement part shall be reinspected in accordance with the provisions of (b). \n\n\t(e)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator.(Douglas Service Bulletin 27-115, Revision No. 1, dated October 25, 1961, pertains to this same subject.) \n\n\tThis directive effective January 24, 1962.
72-07-04: 72-07-04 HAWKER SIDDELEY AVIATION: Amdt. 39-1409. Applies to Models DH.125-1A, -1A/522, -1A/R-522, -1A/S-522, -3A, -3A/R, -3A/RA, and -400A airplanes which have been modified in accordance with Hawker Siddeley Modification 252052. To prevent possible seizure of the windscreen de-icing handpump No. M.2604, within the next 100 hours' time in service after the effective date of this AD, unless already accomplished, accomplish either of the following: (a) Modify the pump by replacing the washers and seal with pre-Modification 252052 parts in accordance with Hawker Siddeley Service Bulletin No. 30-24-(2194), dated December 23, 1970, or later ARB-approved issue or FAA-approved equivalent; or (b) Replace the pump with a new pump No. M.2601/1 in accordance with Hawker Siddeley Service Bulletin No. 30-24-(2194), dated December 23, 1970, or later ARB-approved issue or FAA-approved equivalent. This amendment becomes effective April 15, 1972.
50-23-02: 50-23-02 DOUGLAS: Applies to All Model DC-6 Aircraft. \n\n\tTo be accomplished as indicated below: \n\n\t1.\tAll P/N 5245424 and P/N 5248748 nose gear yoke end fittings which have not been shotpeened in the journal radius prior to original installation or by subsequent rework should be removed for inspection after being in service for a period not to exceed 6,000 hours. Nose gear yoke end fittings which have already accumulated service time in excess of 6,000 hours should be removed for inspection as soon as practical but not later than September 1, 1950. Shotpeening can be distinguished by the dull gray color and coarse surface of the shotpeened area. \n\n\t2.\tFittings removed at the 6,000-hour period may be used for an additional 4,000 hours or a total service life of 10,000 hours if inspected and reworked as follows: \n\n\t\t(a)\tStrip anodic surface from part, and subject to Zyglo inspection paying particular attention to the journal radius. If no cracks are found, the radius should be polished to remove all blemishes and then shotpeened. This inspection and shotpeening must be done by the Douglas Aircraft Co., an agency approved by that company, or by a method that has been substantiated as being equivalent to the procedure recommended by the Douglas Co. \n\n\t\t(b)\tInspect the base radius of the spot faces of the six mounting holes. Parts having zero radius (sharp corner) to 0.031 radius at this point must be reworked to obtain an 0.062 spot face radius. It will be permissible to increase the original spot face diameter of 1 1/8 inches to 1 1/4 inches to obtain the 0.062 radius. Parts having 0.031 or better radius need not be reworked. Parts should be reanodized after completion of all work. \n\t\t(c)\tInspect the inside diameter of the 2103390 ring. All sharp edges should be given a 0.031 radius. \n\t\t(d)\tInspect the inside diameter of the flanged end of the 2333253 bushing to see that it has a 1/8-inch radius and rework if necessary. \n\n\t3.\tFittings shotpeened at time of original installation may be operated for a maximum service period of 10,000 hours provided they do not have the zero spot face radius at the mounting holes. Parts falling in this category should be removed at the normal gear overhaul period of 8,000 hours for rework of the spot face radius. \n\n\t4.\tAll fittings should be scrapped after reaching a total service life of 10,000 hours. \n\n\t(Douglas General Service Letter DC-6 No. 26 dated April 7, 1950, covers the same subject.)