Results
2014-22-03: We are superseding Airworthiness Directive (AD) 2012-14-11 for Arrow Falcon Exporters, Inc. (AFE), Rotorcraft Development [[Page 68109]] Corporation (RDC), and San Joaquin Helicopters (SJH) Model OH-58A, OH- 58A+, and OH-58C helicopters. AD 2012-14-11 required inspecting the main rotor mast (mast) for a crack. This new AD expands the mast inspection area, changes the inspection to a repetitive inspection, and removes the reporting requirement. The actions in this AD are intended to prevent failure of the mast and subsequent loss of control of the helicopter.
2014-22-01: We are superseding Airworthiness Directive (AD) 2012-26-16 for all PILATUS AIRCRAFT LTD. Models PC-12, PC-12/45, PC-12/47, and PC-12/ 47E airplanes. 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 a need to incorporate new revisions into the Limitations section, Chapter 4, of the FAA-approved maintenance program (e.g., maintenance manual). We are issuing this AD to require actions to address the unsafe condition on these products.
76-06-08: 76-06-08 HAWKER SIDDELEY AVIATION LTD: Amendment 39-2558. Applies to Model DH/BH 125 airplanes, all series, certificated in all categories. Compliance is required as indicated. To prevent possible failure of the original bolts attaching the outboard hinge bracket to the upper airbrake assembly (R/H and L/H) and loss of an airbrake in flight accomplish the following: (a) For upper airbrake assemblies, within the next 10 hours time in service after the effective date of this AD or upon the accumulation of 500 hours total time in service on the airbrake assembly, whichever occurs later, unless already accomplished, remove installed bolts P/N S21A102-4E or A102-4EDHS614D (3 per bracket) and install new bolts - (1) P/N S21A102-5E, or an FAA-approved equivalent, in accordance with paragraph A of the section entitled "Accomplishment Instructions" of Hawker Siddeley Aviation Limited Alert Service Bulletin 57 A49, dated July 1, 1975 or an FAA-approved equivalent; or(2) P/N S21DHS1432-5E or an FAA-approved equivalent in accordance with the section entitled "Accomplishment Instructions" (Part A) of Hawker Siddeley Aviation Ltd. Modification Service Bulletin 57 49 (2454) dated September 18, 1975, or an FAA-approved equivalent. (b) Replace P/N S21A102-5E or FAA-approved equivalent bolts, installed as replacements in accordance with paragraph (a)(1) of this AD, at intervals not to exceed 500 hours time in service from installation until the bolts have been replaced with P/N S21DHS1432-5E bolts (3 per bracket) in accordance with paragraph (a)(2) of this AD or an FAA-approved equivalent. (c) For upper airbrake assemblies held as spares, before installation on an airplane, replace the original bolts, P/N S21A102-4E or A102-4EDHS614D, (3 per bracket) with new bolts in accordance with paragraphs (a)(1) or (a)(2) of this AD. This amendment is effective upon publication in the FEDERAL REGISTER as to all persons except those persons to whom it was made immediately effective by the airmail letter, dated February 9, 1976, which contained this amendment.
2014-21-09: We are superseding Airworthiness Directive (AD) 2005-14-07 for certain The Boeing Company Model 727, 727C, 727-100, 727-100C, 727-200, and 727-200F series airplanes. AD 2005-14-07 required repetitive inspections of the carriage attach fittings on the inboard and outboard foreflaps of each wing for cracking and other discrepancies, and corrective actions if necessary. This new AD requires reducing certain repetitive inspection intervals for the inboard and outboard \n\n((Page 64307)) \n\ncarriage attach fittings for the outboard foreflaps, requires previously optional terminating actions which install improved outboard foreflap carriage attach fittings, and adds new initial and repetitive inspections of those fittings and corrective actions if necessary. This AD was prompted by a report of broken inboard and outboard carriage attach fittings of the outboard foreflaps found during an inspection. We are issuing this AD to detect and correct fatigue cracking of the attach fittingsof the foreflap carriage of the wings, which could result in partial or complete loss of the foreflap and consequent loss of controllability of the airplane.
2014-17-51: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-2B16 airplanes. This emergency AD was sent previously to all known U.S. owners and operators of these airplanes. This AD requires inspecting the inboard flap fasteners of the hinge-box forward fitting at Wing Station (WS) 76.50 and WS 127.25 to determine the orientation and condition of the fasteners, as applicable, and replacement or repetitive inspections of the fasteners if necessary. This AD also provides for optional terminating action for the requirements of the AD. This AD was prompted by reports of fractured fastener heads on the inboard flap hinge-box forward fitting at WS 76.50 due to incorrect installation. We are issuing this AD to detect and correct [[Page 64089]] incorrectly oriented or fractured fasteners, which could result in premature failure of the fasteners attaching the inboard flap hinge-box forward fitting. Failure of the fasteners could lead to the detachment of the flap hinge box and the flap surface, and consequent loss of control of the airplane.
53-23-03: 53-23-03 REPUBLIC: Applies to All Model RC-3 Aircraft. Compliance required within the next 25 hours of operation but not later than December 1, 1953, and thereafter at each 25-hour period of operation or every 6 months, whichever occurs first. Cases of severe corrosion of the right and left upper and lower lift strut fittings, fuselage wing lift strut fittings and wing lift strut fittings have been reported. Since the strength of these fittings are of primary importance to the safe operation of the airplane, the following inspections should be made and corrective action taken. Fitting 17W22002 is located on the upper end of the lift strut and fitting 17W22003 is located on the lower end of the lift strut. Fitting 17F11013 is located in the fuselage and is attached to fitting 17W22003. Fitting 17W22004 is located in the wing and is attached to fitting 17W22002. Inspect thoroughly and test the fitting with a pointed instrument to determine whether corrosion is present. One 1 1/4-inch diameter inspection hole should be cut in the upper and one in the lower surface of the wing in accordance with Republic Aviation Service Bulletin No. 25, Supplement No. 2, in order to accomplish the inspection of the portion of the fitting 17W22004 which lies inside the wing skin. This inspection will require the aid of a light as well as a sharp-pointed instrument. The holes should be covered with United Carr Fastener Corp. Plug Button No. 51021, Seabee spare parts item No. 1379, or equivalent. A fitting may appear satisfactory but actually may be corroded under the surface. Such corrosion which may be intergranular in nature may actually result in a much greater loss of strength than would be indicated by the loss of metal from the surface. If the fitting has only slight surface corrosion, the corrosion should be carefully removed and the fitting should be suitably treated against further corrosion. Fittings which have deteriorated beyond slight surface corrosion should be replaced. (Republic Aviation Service Bulletin No. 25, including Supplements Nos. 1 and 2, covers this same subject in detail.) This supersedes AD 50-30-01. NOTE: This AD also applies to Model UC-1 aircraft (see note 3 on TCDS A6EA.)
2014-20-12: We are superseding Airworthiness Directive (AD) 75-20-06 for certain Alexandria Aircraft LLC (type certificate previously held by Bellanca Aircraft Corp., Viking Aviation, Inc., and Bellanca, Inc.) Models 14-19-3A, 17-30, 17-30A, 17-31, 17-31A, 17-31ATC, and 17-31TC airplanes. AD 75-20-06 required repetitively inspecting the aft fuselage structure near the top of the vertical side tubing, which connects the horizontal stabilizer carry-through to the upper fuselage longeron, for cracks and installing the manufacturer's service repair kit as a terminating action for the repetitive inspections to repair any cracks found. Since we issued AD 75-20-06, we have determined that installing the service kit has not prevented cracks from occurring. We have also determined that all affected airplane serial numbers should be included in the Applicability section. This AD requires continued repetitive inspections of the aft fuselage structure near the top of the vertical side tubing for cracks and making all necessary replacements of cracked parts. This AD also adds additional serial number airplanes to the Applicability section. We are issuing this AD to correct the unsafe condition on these products.
55-26-01: 55-26-01 DOUGLAS: Applies to A-26 Aircraft Having Rear Fuselage Fuel Tank Installation. To be accomplished prior to actuation of fuselage fuel tank. Because of an explosion in the air and loss of aircraft, instructions were issued October 12, 1955, to deactivate the rear fuselage fuel tank on the above aircraft until further notice. If the following modification, or its equivalent, is accomplished this fuel tank may be reactivated: 1. Provide fume tight closure and sealing of bulkheads at Stations 332 and 369. This should include tight fitting grommets or fairleads around control cables, or other members passing through bulkhead. 2. Remove all electrical equipment and oxygen tanks, if installed, from the tank compartment. Relocate elsewhere in the airplane as required. 3. Provide insulation around any electrical terminals in tank bay. 4. Provide adequate ventilation airscoop on top or side of tank compartment and exit vent on bottom. Provide drain holes in bottom of compartment to assure complete fuel drainage. 5. Ascertain that fuel tank, filler, cap, scupper, drain, and attaching lines and fittings are airworthy.
97-05-02: This amendment adopts a new airworthiness directive (AD), applicable to certain EMBRAER Model EMB-120 series airplanes, that requires removal of the upper channel fairings and their shims; and rework of the riveting holes, the aileron sealing canvas (aerodynamic seals), and the protective covers of the trim tab hinge fittings of the aileron and elevator. This amendment is prompted by reports of binding of the aileron due to water freezing between the upper channel fairings and the surface of the leading edge of the aileron. The actions specified by this AD are intended to prevent water from freezing these areas, which could result in binding of the aileron and subsequent reduced controllability of the airplane.
59-22-02: 59-22-02 VICKERS: Applies to All Viscount 810 Series Aircraft Which Do Not Embody Modification FG.1447. Compliance required as indicated. Service experience has shown that a gap of less than 0.25 inch between the end of No. 3 flap and the aileron may, under certain flight conditions, produce a condition where the flap could foul or contact the inboard end of the aileron (port and starboard wings). Within the next 500 flight-hours but not later than December 15, 1959, inspect for adequate clearance between the outboard end of the No. 3 flap at the No. 4 flap beam unit and the inboard end of the aileron on both the right and left sides. This inspection must also be carried out whenever a flap or an aileron is installed. Where the gap is found to be less than 0.25 inch the outboard end of the No. 3 flap must be modified to provide proper clearance. (Vickers-Armstrongs PTL No. 80 (800/810 Series) and Modification FG.1447 cover the same subject.)
56-08-01: 56-08-01 CURTISS: Applies to All Models C632S, C634D and C634S Propellers. Compliance required as indicated. I. Replace low pitch limit switches P/N 110425 at not more than the following intervals: Douglas DC-6 Series Aircraft 3,500 Hours Lockheed 749 Series Aircraft 3,500 Hours Lockheed 1049 Series Aircraft 3,500 Hours Convair 240 Series Aircraft 1,000 Hours (Curtiss Service Bulletin No. 52 dated December 29, 1949, and appropriate service manuals also cover these recommendations.) II. Disassemble and visually inspect low pitch limit switches P/N 154592 at not more than the same intervals. Inspect for proper switch operation and for mechanical and electrical condition. Failure or malfunction other than normal service wear shall be cause for replacement of switch parts or of the assembly. Switches not revealing adverse conditions may be returned to service. This supersedes Section III of AD 53-05-01.
57-09-01: 57-09-01 AERO COMMANDER: Applies to All Model 520 Aircraft, Serial Numbers 31 and Above, and to All Models 560, 560A and 680 Aircraft. Compliance required not later than the next 3 hours of flight or May 15, 1957, whichever occurs first and at 100-hour intervals thereafter. As a result of finding cracks in the aileron bellcrank casting in the vicinity of the aileron push-pull rod attach bolt, the following action is considered necessary unless already accomplished. Inspect, using dye penetrant or fluorescent methods, all aileron bellcrank castings P/N 3510005 on bellcrank assembly P/N 4510004-401 and 402 for cracks in upper or lower lugs to which the aileron push-pull rod attaches. Remove rod to make the inspection, replace all castings found defective and reattach push-pull rod, making certain no clearance exists between casting lugs and rod-end bearing inner race before tightening bolt. Use shim washers to eliminate clearance. The 100-hour reinspection of casing P/N 3510005 may be discontinued upon installation of revised casting under development by Aero Design. (Aero Design Service Bulletin No. 41, dated April 19, 1957, provides a sketch of the part and defines the area to be inspected.) This AD covers the same inspection required by CAA telegraphic instructions dated April 25, 1957.
2013-26-05: We are adopting a new airworthiness directive (AD) for all Dassault Aviation Model FAN JET FALCON, FAN JET FALCON SERIES C, D, E, F, and G airplanes; Model MYSTERE-FALCON 200 airplanes; and Model MYSTERE-FALCON 20-C5, 20-D5, 20-E5, and 20-F5 airplanes. This AD was prompted by reports of a manufacturing defect in the charge indicator on fire extinguisher bottles. This AD requires repetitive weighing of fire extinguisher bottles having a certain part number, and eventual replacement of those bottles to terminate the repetitive weighing. We are issuing this AD to detect and correct a dormant failure in the fire suppression system, which could result in the inability to put out a fire in an engine, auxiliary power unit, or rear compartment.
70-07-04: 70-07-04 BRITTEN-NORMAN, LTD: Amdt. 39-964. Applies to Models BN-2 and BN-2A airplanes which have not had Modification NB/M/404 incorporated. Compliance is required as indicated. (a) Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished, visually inspect the attachment channel brackets on the front face of the tail plane front spar for signs of movement or looseness of the blind bolts in accordance with Britten-Norman Service Bulletin No. BN-2/SB.27, dated February 12, 1970, or later ARB- approved issue,or an FAA-approved equivalent. (b) If signs of movement or looseness of the blind bolts are found during the inspection required by paragraph (a), before further flight repair the tail plane attachments in accordance with Britten-Norman Service Bulletin No. BN-2/SB.27, dated February 12, 1970, or later ARB-approved issue, or an FAA-approved equivalent. This amendment becomes effective March 31, 1970.
90-15-11: 90-15-11 PRATT & WHITNEY: Amendment 39-6474. Docket No. 89-ANE-33-AD. Applicability: Pratt & Whitney (PW) JT9D-7R4D1, -7R4E1, and -7R4H1 series turbofan engines installed on Airbus Industries A300/A310 aircraft. Compliance: Required as indicated, unless already accomplished. To prevent engine overtemperature and loss of power or engine inflight shutdown, accomplish the following: (a) Within the next 30 calendar days or 150 flight cycles, whichever occurs first after the effective date of this AD, accomplish the following: (1) Adjust the deceleration schedule of the fuel control unit (FCU), and reidentify the FCU, in accordance with the applicable instructions of Appendix 1 of this AD. If cycling bleeds occur as a result of the decel schedule uptrim, downtrim the decel schedule in accordance with the applicable instructions of Appendix 2 to this AD. (2) Modify the 3.0 bleed valve cylinder, Part Number (P/N) 806885 or P/N 774300, in accordance with the instructions of Appendix 3 or Appendix 4 of this AD, as applicable. (3) Inspect 3.0 bleed valve linkages for wear in accordance with the instructions of Appendix 5 to this AD, and accomplish the following: (i) Remove engines with worn 3.0 valve linkage prior to accumulating 5 cycles in service (CIS) since last inspection and replace with a serviceable engine. (ii) Reinspect linkages found serviceable in accordance with the inspection requirement of paragraph (a)(3) above, at intervals not to exceed 3,000 hours since last inspection. (b) Within the next 60 calendar days or 300 flight cycles, whichever occurs first after the effective date of this AD, accomplish the following: Adjust the engine vane and bleed control (EVBC), Hamilton Standard, P/N 776555-3, to the 1.27 engine pressure ratio (EPR) bleed trim and P/N 776555-5 to the 1.32 EPR bleed trim, in accordance with the instructions of Appendix 6, Appendix 7, or Appendix 8 to this AD, as applicable. Engines which have had no 3.0 bleed or valve schedule adjustments since last trimmed in the test cell (P/N 776555-3 EVBC trimmed to 1.27 EPR or P/N 776555-5 EVBC trimmed to 1.32 EPR) are exempt from this requirement. (c) Incorporate the following modifications to upgrade the EVBC to Hamilton Standard P/N 776555-5, within one year from the effective date of this AD, by accomplishing the following: (1) Incorporate the fluid drain between sensor servo piston chevron seals, in accordance with the instructions of Appendix 9 to this AD. (2) Incorporate the pilot valve spring, P/N 801040-1, in accordance with the instructions of Appendix 10 to this AD. (3) Incorporate the decel bleed reset piston spring, P/N 801073-1, in accordance with the instructions of Appendix 11 to this AD. (4) Incorporate the 3.0 bleed cam, P/N 765357-11, and recalibrate the control, in accordance with the instructions of Appendix 12 to this AD. (5) Incorporate the actuator valve, P/N 800997-1, in accordance with the instructions of Appendix 13 to this AD; or remove actuator valve, P/N 728149-3, and replace with a new or serviceable actuator valve, P/N 728149-3. Replacement actuator valve, P/N 728149-3, must be removed from service at or prior to accumulating 10,000 hours since new. (6) Adjust EVBC to 1.32 EPR bleed trim, in accordance with paragraph (b)(1) above. (d) Modify FCU, Hamilton Standard P/N 782960-1, and P/N 782960-2, within one year from the effective date of this AD by accomplishing the following: (1) Incorporate the decel cam, P/N 774538-15, in accordance with the instructions of Appendix 14 of this AD. (2) Revise the test procedure for the decel bleed override, in accordance with the instructions of Appendix 15 of this AD. (e) Install 3.0 bleed dampers, in accordance with PW Service Bulletin (SB) JT9D-7R4-72-336, Revision 5, dated June 23, 1988, at the next shop visit. Installation of 3.0 bleed dampers terminates the reinspection requirements of paragraph (a)(3)(ii) above. NOTE: For the purpose of this AD, the definition of "shop visit" is any time the engine or module is in a maintenance shop capable of complying with the PW SB instructions, regardless of the planned maintenance action or the reason for engine removal. (f) Restore the leading edge of the first stage compressor blades in accordance with PW Engine Manual, P/N 785058, Chapter/Section 72-31-02, Repair-19, Pages 901 through 912, dated June 15, 1990, at the next fan module overhaul, after the effective date of this AD and thereafter at every fan module overhaul. NOTE: For the purpose of this AD, the definition of "fan module overhaul" is anytime the fan module is disassembled, inspected, and repaired, in accordance with PW Engine Manual. (g) 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. (h) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. First stage compressor blade restoration and 3.0 bleed damper installation shall be done in accordance with the following PW documents: Document No. Page Number Issue/Revision Date Engine Manual 901 thru 914 ---- June 15, 1990 -PW785058 JT9D-7R4-72-336 1, 3, 4, 8, 9, 5 June 23, 1988 10, 11 thru 20 4 August 28, 1987 2 3 June 8, 1987 5, 6, 7 Original March 2, 1987 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 Pratt& Whitney Publications Department, P.O. Box 611, Middletown, Connecticut 06457. Copies may be inspected at the Regional Rules Docket, Office of the Assistant Chief Counsel, Federal Aviation Administration, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts 01803, or at the Office of the Federal Register, 1100 L Street, NW, Room 8301, Washington, DC 20591. This amendment (39-6474, AD 90-15-11) becomes effective on July 16, 1990. Appendix 1 NOTE: This appendix consists of material from: Pratt & Whitney (PW) Special Instructions (SI) 59F-89 (applicable to JT9D-7R4D1, E1 (AI-500 series)), dated July 10, 1989, and SI 61F-89 (applicable to JT9D-7R4E1, H1 (AI-600 series)), dated July 10, 1989. Appendix 2 NOTE: This appendix consists of material from: PW SI 60F-89 (applicable to JT9D-7R4D1, E1 (AI-500 series)), dated July 10, 1989, and 62F-89 [applicable to JT9D-7R4E1, H1 (AI-600 series)), dated July 10, 1989. Appendix 3NOTE: This appendix consists of material from: PW Service Bulletin (SB) JT9D-7R4-75-98, dated January 19, 1988. Appendix 4 NOTE: This appendix consists of material from: PW SI 58F-89, dated June 16, 1989. Appendix 5 NOTE: This appendix consists of material from: PW Maintenance Manual, Chapter/Section 72-00-00 Inspection/Check -06, dated September 15, 1986. Appendix 6 NOTE: This appendix consists of material from: AI A300-600/A310 AMM Chapter/Section 71-00-00, Test Number 10, Temporary Revision No. 71-029, dated January 10, 1989. Appendix 7 NOTE: This appendix consists of material from: PW SI 82F-88, Revision A, dated November 16, 1988. Appendix 8 NOTE: This appendix consists of material from: PW SI 87-88, dated August 26, 1988. Appendix 9 NOTE: This appendix consists of material from: Hamilton Standard (HS) SB GTA9-75-9, Revision 1, dated October 14, 1984. Appendix 10 NOTE: This appendix consists of material from: HS SB GTA9-75-16, dated March 20, 1988. Appendix 11 NOTE: This appendix consists of material from: HS SB GTA9-75-17, dated March 31, 1988. Appendix 12 NOTE: This appendix consists of material from: HS SB GTA9-75-18, Revision 1, dated January 15, 1989. Appendix 13 NOTE: This appendix consists of material from: HS SB GTA9-75-19, dated August 12, 1988. Appendix 14 NOTE: This appendix consists of material from: HS SB JFC68-10-10, Revision 1, dated November 15, 1989. Appendix 15 NOTE: This appendix consists of material from: HS SB JFC68-10-9, dated March 31, 1988.
68-19-07: 68-19-07 SIKORSKY: Amendment 39-657 as amended by Amendment 39-2450. Applies to S-61 Series helicopters. Compliance required as indicated. To preclude the possibility of failure of the main rotor blade spindle, accomplish the following: (a) Before further flight, remove main rotor blade spindles P/Ns S6110-23325-1, S6110-23325-2, and S6112-23025-1 that either have been "salvaged" in accordance with procedures set forth in paragraph entitled "Salvage of Spindle" contained in Sikorsky Aircraft Overhaul Manual for the pertinent helicopter model, or have accumulated 3000 or more hours' time in service on the effective date of this AD, and replace with blade spindles of the same part number that have not been "salvaged" and that have less than 3000 hours' time in service. (b) Replace main rotor blade spindles P/Ns S6110-23325-1, S6110-23325-2, and S6112-23025-1, that have not been "salvaged" and have less than 3000 hours' time in service on the effective date of thisAD, before the accumulation of 3000 hours' time in service with main rotor blades spindles of the same part number that have not been "salvaged" and have less than 3000 hours' time in service. (c) Before further flight, remove from service main rotor blade spindles P/N's S6110-23325-1, S6110-2325-2, and S6112-23025-1, serial numbers AX51, AX54, F2148, F2444, F2485, F2207, F1406, F1416, F1415, F1399, B-35, and F2451. Amendment 39-657 was effective September 27, 1968. This amendment 39-2450 becomes effective December 23, 1975.
2014-16-12: We are adopting a new airworthiness directive (AD) for all Dassault Aviation Model FALCON 2000EX airplanes. This AD was prompted by a revision to the airplane airworthiness limitations to introduce a corrosion prevention control program, among other changes, to the maintenance requirements and airworthiness limitations. This AD requires revising the maintenance or inspection program to include the maintenance tasks and airworthiness limitations specified in the airworthiness limitations section of the airplane maintenance manual. We are issuing this AD to prevent reduced structural integrity of the airplane.
97-06-03: This amendment supersedes an existing airworthiness directive (AD), applicable to Bell Helicopter Textron, Inc. (BHTI) Model 214ST helicopters, that currently establishes a mandatory retirement life of 50,000 high-power events for the main rotor mast (mast). This amendment requires changing the retirement life for the mast from high-power events to a maximum accumulated Retirement Index Number (RIN) of 140,000 and applying this RIN to an additional part-numbered mast. This amendment is prompted by fatigue analyses and tests that show certain masts fail sooner than originally anticipated because of an unanticipated high number of takeoffs and external load lifts in addition to the deterioration in strength that occurs under other operating conditions. The actions specified by this AD are intended to prevent fatigue failure of the mast, which could result in failure of the main rotor system and subsequent loss of control of the helicopter.
51-20-01: 51-20-01 CURTISS-WRIGHT Applies to Model C-46 Series aircraft equipped with C-4 Cockpit Light Assembly. \n\n\tCompliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished. \n\n\tTo eliminate an electrical fire hazard existing in the pilot's compartment, the following rework must be accomplished: \n\n\t1. Rework the C-4 cockpit light rheostat assembly in accordance with Figure 2. Fabricate insulator strip per Figure 2(a), and install as shown in Figure 2(b). \n\n\n\n\nSECTIONAL VIEW OF COCKPIT LIGHT CASE - TYPE C-4A & C-4, WITH RHEOSTAT INSULATOR INSTALLED \nFIGURE 2 \n\n\t(U.S.A.F.T.O. No. 03-5G-12, dated September 8, 1950, covers this same subject.) \n\n\t2. Revise the wiring to the C-4 lamp assembly as shown in Figure 3. \n\n\n\n\t(Page 410D, Figure 301B of AN 01-25L-2 covers this same subject.) \n\n\tRevised December 28, 1964.
2014-16-27: We are adopting a new airworthiness directive (AD) for certain Dassault Aviation Model FALCON 900EX airplanes. This AD was prompted by our determination of the need for a revision to the airplane airworthiness limitations to introduce a corrosion prevention control program, among other changes, to the maintenance requirements and airworthiness limitations. This AD requires revising the maintenance or inspection program, as applicable, to include the maintenance tasks and airworthiness limitations specified in the Airworthiness Limitations section of the airplane maintenance manual. We are issuing this AD to prevent reduced structural integrity of the airplane, and prevent reduced controllability of the airplane.
51-05-01: 51-05-01 PRATT & WHITNEY: Applies to Douglas DC-6 and Convair 240 Aircraft Equipped With R-2800-34M1, -83AM3, -83AM4, and Double Wasp CA Series Engines Using Antidetonant Injection (Wet Power) for Takeoff. Compliance required as soon as possible but not later than February 15, 1951. A. Each operator of an airplane covered shall select a power which he undertakes to maintain. If that power is less than the corresponding value available during the type certification tests of the airplane, the operating weights of the entire fleet shall be reduced to values such as will enable the airplanes to perform as indicated in the approved airplane flight manual with the power selected. B. The power actually develop by each engine shall be measured each time it reaches each of the following stages. 1. Upon installation of overhauled engines in aircraft. 2. At the No. 3 inspection nearest to the midpoint of the authorized service time between overhauls. 3. At thenearest No. 3 inspection or some convenient point near or at the end of the authorized service time between overhauls. C. The procedures and methods employed in making these power measurements shall be acceptable to the FAA. 1. The frequency of the power measurements should be contained as indicated above until the results obtained on each operator's fleet have been evaluated for the purpose of establishing whether more frequent or less frequent measurements are warranted. 2. Operators not employing line maintenance practices which will reasonably insure the continued availability of the selected power will start this program making more frequent power measurements than indicated above. D. An airplane incorporating an engine which at any of the required power measurements, fails to develop the selected power shall not be dispatched unless: 1. The power is restored to the selected value, or 2. The engine is replaced by one developing the selected power, or 3. The operating weights of the individual airplane are reduced as specified in A. E. If, on a fleet-wide basis, the initial powers measured during any individual power measurement are consistently below the selected power, the operator shall: 1. Initiate or improve line maintenance to the extent necessary to give reasonable assurance that the selected power is continuously available, or 2. Make more frequent measurements of power, or 3. Select a lower value of power representative of the initially measured values and reduce operating fleet weights as specified in A. F. An acceptable method for power measurement and data correction utilizing static ground runups in the aircraft is described by AAL in their instructions on "Convair Ground Power Check". This information has been distributed by ATA to all the airlines involved. Alternative methods providing equivalent or greater accuracy will be acceptable. G. Results of the above power checksare to be submitted regularly and promptly to the assigned FAA Field Agents.
96-24-01: This amendment adopts a new airworthiness directive (AD), applicable to all Fokker Model F27 Mark 100, 200, 300, 400, 500, 600, and 700 series airplanes, that requires replacement of certain rudder horn assemblies with a new assembly. For certain airplanes, the amendment also requires replacement of certain rudder control rods with a new rod. This amendment is prompted by reports of cracked rudder horns and a cracked rudder control rod, caused by impact overload. The actions specified by this AD are intended to prevent such an overload and consequent cracking of the subject parts, which could result in reduced structural integrity of the rudder horn assembly or loss of rudder control; this condition could lead to reduced controllability of the airplane.
2014-16-24: We are superseding Airworthiness Directive (AD) 2012-10-53 for Eurocopter Deutschland GmbH (ECD) (now Airbus Helicopters) Model EC135P1, EC135P2, EC135P2+, EC135T1, EC135T2, and EC135T2+ helicopters. AD 2012-10-53 required, before further flight and at specified intervals, checking and inspecting the upper and lower main rotor hub (MRH) shaft flanges for a crack, and inspecting the lower hub-shaft flange bolt attachment areas for a crack. Since we issued AD 2012-10- 53, it has been determined that it is safe to increase the visual inspection intervals of the MRH shaft flanges from 10 hours time-in- service (TIS) to 50 hours TIS and remove the inspection of the lower MRH shaft flange bolt attachment areas. This new AD continues to require checking and inspecting the upper and lower MRH shaft flanges for a crack. These actions are intended to detect a crack on the MRH shaft flange, which if not corrected, could result in failure of the MRH and subsequent loss of control of the helicopter.
56-23-03: 56-23-03 SIKORSKY: Applies to All Model S-55 Helicopters. Compliance required as indicated. Due to additional failures, it has been found necessary to lower the retirement time specified in AD 56-10-02. Accordingly, the horizontal hinge pin assembly, P/N S10-10-3331 and S10-10-3331-1 with 750 hours service must be replaced by December 1, 1956, or prior to the accumulation of 850 hours, whichever occurs first. Parts with 850 hours or more must be replaced prior to further service. Thereafter, P/N S10-10-3331 and S10-10-3331-1 are to be retired at not more than 750 hours. This supersedes AD 56-10-02.
2014-16-07: We are superseding Airworthiness Directive (AD) 2011-15-09 for certain Bombardier, Inc. Model DHC-8-400, -401, and -402 airplanes. AD 2011-15-09 required repetitive inspections for proper operation of the main landing gear (MLG) alternate extension system (AES), and corrective actions if necessary. This new AD requires, for certain airplanes, new repetitive inspections for proper operation of the MLG AES, and corrective actions if necessary. This new AD also requires eventually replacing the MLG AES cam mechanism assembly with a new assembly, which terminates the repetitive inspections for those airplanes. This AD was prompted by a determination that, for certain airplanes not affected by AD 2011-15-09, a different MLG AES cam mechanism assembly was installed, resulting in input lever fractures and inability to open the MLG door; those assemblies could be subject to the same unsafe condition in AD 2011-15-09. We are issuing this AD to prevent improper operation of the cam mechanism or rupture of the door release cable, which [[Page 48969]] could result in loss of control of the airplane during landing.