Results
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.)
2008-20-04: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued 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: High pressure (HP) turbine discs recently inspected in accordance with the Engine Manual have exhibited cracks in the disc rim. The discs have failed to meet the inspection acceptance criteria and have been returned to Rolls-Royce for engineering investigation. This investigation has concluded that the cracks have resulted from scores within the cooling air holes in the disc rim that could have been introduced during new part manufacture or during overhaul of the disc. The engineering investigation has concluded that if this cracking was undetected then it could result in uncontained disc failure and a potential unsafe condition for the aircraft. We areissuing this AD to prevent uncontained disc failure, possibly resulting in damage to the airplane.
47-43-01: 47-43-01 CESSNA: (Was Mandatory Note 12 of AD-768-5.) Applies to 120 and 140 Aircraft Serial Numbers Up to and Including 11842. Compliance required prior to January 1, 1948. Reroute the lower end of the primer line located on the left side of the firewall and rotate the strainer fitting so that it points downward and to the left at an angle of 60 degrees to the horizontal. Slip approximately 6 inches of vinylite tubing over the upper and lower ends of this primer line and install a shield around this line between the two pieces of vinylite tubing. This will preclude the possibility of fuel coming in contact with the left exhaust manifold in the event of a failure in this primer line. (Cessna Service Letter No. 34, dated March 24, 1947, covers this same subject.)
47-07-01: 47-07-01 BELLANCA: (Was Mandatory Note 4 of AD-773-5.) Applies to Models 14-13, 14-13-2 All Serial Numbers Up to and Including 1200. To be accomplished not later than next periodic inspection. Check fuel selector valve handle for proper indexing on valve by setting handle in L-ON and in R-ON position, by disconnecting the fuel line, and by blowing through line when there should be free passage of air. After tank positions have been set, the valve handle and shank should be permanently marked to identify the index position. Attach handle positively to shank by drilling through one side of the handle and halfway through the shank with a drill of number 53 size and inserting a pin of 1/16-inch diameter drill rod. (Bellanca Service Bulletin No. 4 covers inspection of the valve handle installation.)
62-10-01: 62-10-01 BELL: Amdt. 428 Part 507 Federal Register April 21, 1962. Applies to All Model 47J Helicopters Serial Numbers 1420 Through 1802 With Rollpins P/N 49-040-187-1750 or Clevis Pins P/N MS 20392-2-49 Installed In Elevator Spar, and With End Rib P/N 47-267- 453-1 (0.025-Inch Thick) Installed; Except Those Helicopters That Have Bell Kit No. 47-3746-1 or 47-3746-2 Installed. Compliance required as indicated. Numerous reports have been received of fatigue cracking of the tubular spar of both the right and left elevator at the Rollpin hole at B.L. 7.0, and fatigue cracking of the inboard rib of the elevators. To preclude failure of the elevator, the following shall be accomplished: (a) Within 25 hours' time in service after the effective date of this AD: (1) Remove the elevators from the tail boom in accordance with the Bell Maintenance Manual. (2) Clean the area around the Rollpin hole and remove any zinc chromate putty from any plugged hole in the tubular spar at B.L. 7.0 for both right and left elevators. (3) Inspect for cracks in the tubular spar of both elevators at the Rollpin hole at B.L. 7.0 using a 5-power or higher magnifying glass. (4) Inspect the inboard rib for cracks using a 5-power or higher magnifying glass. (b) If cracks are found in the tubular spar modify the elevator with Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator," or FAA engineering approved equivalent prior to further flight. (c) If no cracks are found in the tubular spar, install clevis pin in accordance with subparagraphs (1) through (4) or Bell Service Letter No. 56 and reinspect in accordance with subparagraph (5). (1) Position coupling assembly P/N 47-267-483-1 on elevators and line drill through Rollpin holes with a "D" (0.2460-inch diameter) drill. Remove sharp edges from holes. Install MS 20392-3-49 clevis pins, AN 960-4162 washers, and AN 381-3-6 cotter pins. A finger tight slip fitof the clevis pins is desired, approximately 0.0005 inch loose. (2) Reinstall the elevator on the helicopter, shim as required to prevent preload or end play at bearings. (3) Check clearance between skin and end of clevis pins. Trim skin, if necessary, to obtain clearance. (4) Rerig elevator in accordance with the Bell Maintenance Manual. (5) Reinspect in accordance with (a)(1) through (a)(3) within each succeeding 50 hours' time in service until Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator", or FAA approved equivalent is installed. (d) If cracks are found in the inboard rib, repair the elevator as specified below, or modify with Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, or FAA engineering approved equivalent prior to further flight. (1) Remove the inboard rib by drilling out the rivets and remove the Bell P/N 47-267-404-7 shoulder from the rib by drilling out the rivets. (2) Add a doubler of 0.032 thickness, or a new rib of 0.032 thickness, material aluminum alloy 2024-0, or a Bell rib P/N 47-267-453-7 (one required per elevator). (3) Rivet Bell P/N 47-267-404-1 shoulder to the old rib and new doubler or the new rib. Use the rivet pattern in the shoulder with AN 470-AD3 or -4 rivets. (4) Install the rib assembly, using the rivet pattern in the elevator skin with MS 20600 AD4 or -5 rivets. (e) If no cracks are found in the inboard rib: (1) Reinstall the elevator on the helicopter in accordance with Bell Maintenance Manual. (2) Reinspect rib for cracks in accordance with (a)(4) within each succeeding 50 hours' time in service until Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator", or FAA engineering approved equivalent is installed. (f) Upon request of the operator, an FAA maintenance inspector subject to prior approval of the Chief, Engineering and Manufacturing Branch, Southwest 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. (Bell Service Bulletin No. 135 SB dated July 27, 1961, covers this same subject. Bell's Service Letter No. 56 covers an acceptable fix for paragraphs (c)(1) through (c)(4) of this AD.) This directive effective May 22, 1962.
2008-19-11: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) provided by the European Aviation Safety Agency (EASA) to identify and correct an unsafe condition on Turbomeca S.A. Arrius 2B1, 2B1A, 2B2, and 2K1 turboshaft engines. The MCAI describes the unsafe condition as: A short circuit of some tantalum capacitors inside certain electronic control (EEC) units may lead to an in-flight shutdown on one of the two engines resulting from: --Direct activation of the overspeed electronic protection; --Non-direct activation of the electronic overspeed protection by lowering the threshold, --Spurious activation of the starting sequence; or --Loss of power control with no freeze of the fuel-metering valve. We are issuing this AD to prevent in-flight engine shutdowns and possible forced autorotation landing or accident.
91-15-25: 91-15-25 GENERAL ELECTRIC COMPANY (GE): Amendment 39-7090. Docket No. 90- ANE-34. Applicability: GE CF6-80A series and CF6-80C2 series engines, installed on, but not limited to, Airbus A300 and A310 and Boeing 747 and 767 aircraft. Compliance: Required as indicated, unless already accomplished. To prevent an uncontained engine failure, accomplish the following: (a) Inspect high pressure compressor rotor (HPCR) stages 11-14 spool-shafts for vane to spool rubs within 500 cycles in service (CIS) after the effective date of this AD, or prior to accumulating 8000 cycles since new, whichever occurs later, according to the following: (1) Inspect CF6-80A series engines, Serial Numbers (S/N) 580-101 through 580-319, and S/N 585-101 through 585-222, installed with an HPCR stages 11-14 spool-shaft, Part Number (P/N) 9225M37G11, 9225M37G14, 9225M37G16, 9225M37G19, 9225M37G20, or 9225M37G21, in accordance with the Accomplishment Instructions in GE CF6-80A Service Bulletin (SB) 72-459, Revision 2, dated June 14, 1989. (2) Inspect CF6-80C2 series engines, S/N 690-101 through 690-181, S/N 695-101 through 695-150, and S/N 705-101 through 705-112, installed with an HPCR stages 11- 14 spool-shafts, P/N 9380M30G07, 9380M30G08, 9380M30G09, 9380M30G10, or 1531M21G01, in accordance with the Accomplishment Instructions in GE CF6-80C2 SB 72-130, Revision 2, dated October 18, 1989. (3) Inspect in accordance with the applicable requirements of paragraph (a)(1) or (a)(2) of this AD, HPCR stages 11-14 spool-shafts which were installed in a high pressure compressor (HPC) at an engine shop visit, with stages 10 through 13 vane radiiless than the values indicated in Table 1 of this AD for CF6-80A service engines, and Table 2 of this AD for CF6-80C2 series engines. (b) The inspection requirements of paragraph (a) of this AD are not applicable to HPCR stages 11-14 spool-shafts visually inspected at the piece-part level, determined not to have vane to spool rub damage,and were installed in an HPC with stages 10 through 13 vane radii greater than or equal to the values indicated in Table 1 of this AD for CF6-80A series engines, and Table 2 of this AD for CF6-80C2 series engines. (c) Remove from service within 500 CIS after the effective date of this AD or prior to accumulating 8,000 cycles since new, whichever occurs later, HPCR stages 11-14 spool-shafts with vane to spool rub damage. NOTE: CF6-80A SB 72-460 and CF6-80C2 SB 72-131 introduce an FAA approved rework procedure to increase the FAA approved life limit for HPCR stages 11-14 spool-shafts with vane to spool rub damage. (d) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (e) Upon submission of substantiating data by an owner or operator through an FAA Inspector (maintenance, avionics or operations, as appropriate), an alternate method of compliance with the requirements of this AD or adjustments to thecompliance times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. (f) The inspections shall be done in accordance with the following GE documents: DOCUMENT PAGE NO. ISSUE/REVISION DATE GE CF6-80A 2-9 Original 11/25/86 SB 72-459 1 Rev. 2 6/14/89 Total Pages: 9 GE CF6-80C2 2-9 Original 11/25/86 SB 72-130 1 Rev. 2 10/18/89 Total Pages: 9 This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, Room 311, 12 New England Executive Park, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, D.C. TABLE 1 CF6-80A HPC Stator Vane Minimum Radii Vane Stage Position Minimum Vane Radius (in) Stage 10 Vanes 1-6, 35-80 11.416 7-8, 33, 34 11.417 9, 10, 31, 32 11.418 11, 12, 29, 30 11.419 13-16, 25-28 11.420 17-24 (or 1-80 if round grind) 11.421 Stage 11 Vanes 1-6, 35-80 11.623 7, 8, 33, 34 11.624 9, 10, 31, 32 11.625 11, 12, 29, 30 11.626 13, 16, 25-28 11.627 17-24 (or 1-80 11.628 if round grind) Stage 12 Vanes with Liners 1-10, 31-80 11.769 11, 12, 29, 30 11.770 13-15, 26-28 11.771 16-25 (or 1-80 if round grind) 11.772 Stage 12 Vanes without Liners 1-10, 31-80 11.786 11, 12, 29, 30 11.787 13-15, 26-28 11.788 16-25 (or 1-80 if round grind) 11.789 Stage 13 Vanes 1-10, 31-80 11.903 11, 12, 29, 30 11.904 13-15, 26-28 11.905 16-25 (or 1-80if round grind) 11.906 NOTES: (1) Vane radius is measured from vane tip at vane centerline to stator case centerline. (2) Vane positions are numbered clockwise, aft looking forward, starting with No. 1 at the left hand horizontal split line upper stator case. (3) These revised minimum vane radii were incorporated into the CF6-80A Engine Manual, GEK 72501, in Revision 23. TABLE 2 CF6-80C2 HPC Stator Vane Minimum Radii Vane Stage Position Minimum Vane Radius (in) Stage 10 Vanes 1-6, 35-80 11.409 7, 34 11.410 8, 9, 32, 33 11.411 10, 31 11.412 11, 12, 29, 30 11.413 13, 14, 27, 28 11.414 15, 16, 25, 26 11.415 17-24 (or 1-80 if round grind) 11.416 Stage 11 Vanes 1-6, 35-80 11.617 7, 34 11.618 8, 9, 32, 33 11.619 10, 31 11.620 11, 12, 29, 30 11.621 13, 14, 27, 28 11.622 15, 16, 25, 26 11.623 17-24 (or 1-80 if round grind) 11.624 Stage 12 Vanes 1-10,32-84 11.780 11, 31 11.781 12, 13, 29, 30 11.782 14-16, 26-28 11.783 17-25 (or 1-84 if round grind) 11.784 Stage 13 Vanes 1-10, 31-80 11.906 11, 12, 29, 30 11.907 13, 14, 27, 28 11.908 15, 16, 25, 26 11.909 17-24 (or 1-80 if round grind) 11.910 NOTES: (1) Vane radius is measured from vane tip at vane centerline to stator case centerline. (2) Vane positions are numbered clockwise, aft looking forward, starting with No. 1 at the left hand horizontal split line upper stator case. (3) These revised minimum vane radii were incorporated into the CF6-80C2 Engine Manual, GEK 92451, in Revision 7. This amendment (39-7090, AD 91-15-25) becomes effective on September 6, 1991.
62-08-03: 62-08-03 BEECH: Amdt. 421 Part 507 Federal Register April 17, 1962 as amended by Amendment 39-1019. Applies to Models 35-33, 35-A33 and 35-B33, Serial Numbers prior to CD-803, except CD-745 and CD-789; Models 35, A35, B35, C35, D35, E35, F35, G35, H35, J35, K35, M35, N35, and P35, Serial Numbers prior to D-6952; Model 50, Serial Numbers H-1 through H-11; Models B50 and C50, Serial Numbers CH-12 through CH-360; Models D50, D50A, D50B, D50C and D50E, Serial Numbers prior to DH-327, except DH-323 and DH-324; Model E50, Serial Numbers EH-1 through EH-70; Model F50, Serial Numbers FH-71 through FH-96 except FH-94; Model G50, Serial Numbers GH-94 and GH-97 through GH-119; Model H50, Serial Numbers HH-120 through HH-149; Model J50, Serial Numbers JH-150 through JH- 163; Models 95-55 and 95-A55, Serial Numbers prior to TC-276 except TC-235, TC-245, TC- 266, TC-273 and TC-274; Model 65, Serial Numbers prior to LC-141 except LC-125; Models 95, B95 and B95A, Serial Numbers prior to TD-499 airplanes. Compliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished. As a result of cracks in and one failure of the white plastic rams horn control wheel, accomplish the following: (a) Within the next 5 hours' time in service after the effective date of this AD and thereafter at each 100 hours' time in service or twelve calendar months, whichever occurs first, visually inspect the white plastic rams horn control wheels for cracks. Give particular attention to the area on the forward side of the hub and in the area of the attachment pin. (b) If cracks are found, replace the control wheel prior to further flight. (c) The inspections required by this AD may be discontinued when the proper replacement control wheel as specified in Beech Service Bulletin dated March 1970, titled "Control Wheel, Clock and Map Light, Inspection of Plastic Rams Horn Type Control Wheels" has been installed or an FAA approved equivalent is installed. (Beech Service Bulletin dated March 23, 1962, titled "Inspection of Plastic Rams Horn Type Control Wheels" covers this same subject.) Amendment 421 effective April 30, 1962. Revised February 16, 1965. This Amendment (39-1019) becomes effective July 9, 1970.