Results
75-07-08: 75-07-08 SIKORSKY AIRCRAFT: Amendment 39-2138. Applies to Models S-58 and S-58T series helicopters certificted in all categories, including Military Type HSS-1, HSS-1F, HSS-1N, HUS-1, HUS-1A, HUS-1AN, HUS-1G, HUS-1Z, H-34A, H-34C, H-34J, CH-34A, CH- 34C, HH-34F, SH-34G, SH-34H, SH-34J, UH-34D, UH-34E, UH-34G, UH-34J, VH-34C, VH- 34D, equipped with S1610-31100 series rotary rudder assemblies and S1610-34200 series counterweight assembly. Compliance required as indicated. To prevent excessive stresses and possible structural failures in the tail rotor blades, tail rotor gear box assembly, and tail pylon area, from tail rotor unbalance, accomplish the following: a. Within the next 10 hours time in service after the effective date of this AD unless already accomplished, lubricate and inspect the rotary assembly, P/N S1610-31100 series and counterweight assembly, P/N S1610-34200 series, for balance in accordance with Section 2, Paragraph A of Sikorsky Service Bulletin No. 58B15-14A, dated December 19, 1974, or later FAA approved revision. b. Thereafter, lubricate and inspect at intervals not to exceed 50 hours time in service from the last inspection, or earlier if unbalance is suspected, in accordance with Section 2, Paragraph B of Sikorsky Service Bulletin No. 58B15-14A or later FAA approved revision. c. If after two consecutive 50 hour repetitive inspections as specified in Paragraph b., the rotor balance remains unchanged, and within limits, these inspections can be extended to a 100 hour interval. These 100 hour intervals can be maintained if all subsequent 100 hour inspections show that rotor balance remains unchanged. d. Rotary rudder assemblies, P/N S1610-31100 series and counterweight assembly P/N S1610-34200 series which are not balanced as specified in Paragraphs a., b., and c must be either rebalanced or removed for corrective action in accordance with Section 2, Paragraphs A and B of Sikorsky Service Bulletin No. 58B15-14A or laterFAA approved revision, before further flight. e. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Sikorsky Aircraft, Division of United Aircraft Corporation, Stratford, Connecticut 06602. These documents may also be examined at the Office of the Regional Counsel, New England Region, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment becomes effective April 8, 1975.
61-16-05: 61-16-05 PIPER: Amdt. 311 Part 507 Federal Register July 26, 1961. Applies to PA-18/150, PA-18A/125, PA-18A/135, PA-18A/150, PA-18S/150, PA-18AS/150: Serial Numbers 18-1667, 18-1887, 18-4260, 18-4860, 18-6127, 18-6289, 18-6301, 18-6332, 18-6375, 18-6398, 18-6418, 18-6424, 18-6434, 18-6437, 18-6438, 18-6444, 18-6463, 18-6466, 18-6501, 18-6540, 18-6604, 18-6606, 18-6609, 18-6616, 18-6657, 18-6668, 18-6679, 18-6680, 18-6681, 18-6682, 18-6771, 18-6855, 18-6914, 18-6924, 18-6981, 18-6991, 18-6992, 18-6993, 18-7018, 18-7025, 18-7037, 18-7043, 18-7047, 18-7058, 18-7071, 18-7072, 18-7075, 18-7078, 18-7093. PA- 22/150, PA-22S/150, PA-22/160, PA-22S/160: Serial Numbers 22-6116, 22-6167, 22-6359, 22-6421, 22-6466, 22-6550, 22-6704, 22-6758, 22-6883. Compliance required within 10 hours' time in service after the effective date of this AD unless already accomplished. To preclude loss of control of the airplane as a result of failed nicopress sleeves in the control system, the following must be accomplished: (a) In PA-18 Series aircraft, aileron, lower elevator, and flap flexible stainless steel cable assemblies, P/N's 12794-03, 13271-02, 13745-02, 40123-44, 40123-86, 40123-87, and 10870-12 shall be replaced with respective assemblies P/N's 12794-00, 13271-00, 13745-00, 40123-03, 40123-77, 40123-76, and 10870-08, or respective assemblies P/N's 12794-04, 13271- 03, 13745-03, 14300-19, 14300-27, 14300-28, and 10870-17. (b) In PA-22 Series aircraft, aileron, lower elevator, flap and rudder cable assemblies P/N's 11527-03, 12515-04, 13108-03, 13109-13, 13109-15, 40123-83, 40123-84, and 40123-94 shall be replaced with respective cable assemblies P/N's 11527-02, 12515-03, 13108-02, 13109- 10, 13109-12, 40123-68, 40123-69, and 40123-93; or respective cable assemblies P/N's 11527- 04, 12515-06, 13108-11, 13109-18, 13109-19, 14300-24, 14300-25, and 14300-29. Standard landplane galvanized cables are satisfactory for continued or replacement use. (PiperService Bulletin Number 181 dated November 5, 1959, and Piper Service Letter No. 355, dated February 23, 1961, cover this same subject.) This supersedes AD 59-26-06. This directive effective July 26, 1961.
66-03-01: 66-03-01 BRISTOL SIDDELEY: Amdt. 39-186 Part 39 Federal Register January 28, 1966. Applies to de Havilland Model Gipsy Queen Series 30 Engines With Crankshafts With Modification 2036 or 2162 Installed. Compliance required as indicated. To prevent further failures of the crankshaft front end that could result in loss of the propeller, accomplish the following: (a) Unless already accomplished within the last 375 hours' time in service, inspect crankshaft for cracks within the next 25 hours' time in service after the effective date of this AD in accordance with Bristol Siddeley Technical News Sheet (TNS) GQ30 No. 70, dated November 16, 1965, or later ARB-approved issue, or an FAA- approved equivalent, and thereafter at intervals not to exceed 400 hours' time in service from the last inspection until modified in accordance with (c). (b) Replace cracked crankshafts before further flight with an uncracked crankshaft. (c) Incorporate Modification 2925 (which introduces shot peening underneath the chrome to improve the fatigue strength) on crankshaft within the next 1,600 hours' time in service after the effective date of this AD unless already accomplished. This directive effective January 28, 1966.
61-14-02: 61-14-02 BELL: Amdt. 300 Part 507 Federal Register July 1, 1961. Applies to All Models 47J, 47J-2, 47G-3, and 47G-2A Helicopters With Float Gear Kit No. 47-706-661 installed. Compliance required as indicated. Several cases of failures have been reported of the forward cross tube failing near the left attaching point (cross tube to centerframe) through drilled holes which have been welded closed and filed smooth. To remove defective tubes and preclude the possibility of float cross tube No. 47-512-111-9 failing during landing, the following one-time inspection is required: (a) Within the next 10 hours' time in service after the effective date, remove paint from around the P/N 47-512-111-9 cross tube for an area approximately 2 inches on each side of the stop plates, P/N 47-500-024-1 and visually inspect for any indication of welded holes in this area. (b) If no evidence of welds is found, prime and repaint as necessary. (c) If any indication of a weldexists, replace the cross tube with an FAA approved part within the next 5 hours' time in service following the above inspection. (Bell Service Bulletin No. 132 dated May 23, 1961 covers this subject.) This directive effective July 12, 1961.
2005-10-18: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 747 series airplanes. That AD currently requires a one-time inspection to determine the material type of the stop support fittings of the main entry doors (MEDs). That AD also currently requires repetitive detailed inspections to detect cracks of certain stop support fittings of the MEDs, and replacement of any cracked stop support fitting with a certain new stop support fitting. This new AD adds new inspections, and replacement if necessary, of the stop support fittings of MED 3, and adds airplanes to the applicability. This AD is prompted by reports of MED 3 having certain stop support fittings that are susceptible to stress corrosion cracking. We are issuing this AD to detect and correct stress corrosion cracking of the stop support fittings of the MEDs, which could result in damage to the adjacent forward edge frame of the door and consequent loss of a MED and rapid decompression of the airplane. \n\nDATES: This AD becomes effective June 23, 2005. \n\n\tThe incorporation by reference of Boeing Service Bulletin 747-53- 2358, Revision 1, dated April 19, 2001; and Boeing Special Attention Service Bulletin 747-53-2485, dated January 8, 2004; as listed in the AD, is approved by the Director of the Federal Register as of June 23, 2005. \n\n\tOn January 25, 1999 (63 FR 70316, December 21, 1998), the Director of the Federal Register approved the incorporation by reference of Boeing Service Bulletin 747-53-2358, dated August 26, 1993.
55-25-04: 55-25-04 SIKORSKY: Applies to All Model S-55 Helicopters. Compliance required as indicated. A recent failure of the tail cone skin in the area where the tail cone is spliced to the fuselage has been reported in a Model S-55 helicopter. Since this is the means of attaching the tail cone to the fuselage and to preclude further failures, the following inspections and reinforcements for the splice area are considered necessary. 1. Tail cones with less than 1,500 hours, add reinforcements at the next major inspection or within the next 150 hours. 2. Tail cones with 1,500 hours to 2,000 hours, inspect the skin for cracks at the next major inspection or within the next 150 hours. If no cracks are found, add reinforcements. If cracks are found, replace the skin and add reinforcements. 3. Tail cones with 2,000 hours or more, inspect the skin for cracks at the next intermediate inspection or within the next 50 hours. If no cracks are found, add reinforcements. Ifcracks are found, replace the skin and add reinforcements. 4. The inspection procedures to be followed, the reinforcements to be added and the method of attaching the reinforcements are explained in detail in Sikorsky Information Circular No. 1420-632.
61-12-01: 61-12-01 DOUGLAS: Amdt. 293 Part 507 Federal Register June 3, 1961. Applies to All Model DC-8 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tAn instance has been reported wherein the left wing flight spoilers extended during takeoff resulting in a left wing heavy condition. The loss of spoiler control was attributed to a sheared rivet in rod assembly P/N 4719278 of the aileron shift mechanism in the left wheel well. To preclude further difficulties with this part which can result in a hazardous condition, the following inspection, operational checks or rework must be accomplished: \n\n\t(a)\tPrior to each flight, conduct an operational check of the outboard spoiler control system, observing the pressure drop in the spoiler hydraulic system during the check. During rapid rotation of the aileron control wheel a pressure drop of approximately 200 p.s.i. is normal. If a drop in hydraulic pressure does not occur, inspect the spoiler system for a sheared rivet in rod assembly P/N 4719278 of the spoiler shifting mechanism. If the rivet is not sheared a more thorough examination of the spoiler system should be made to determine and correct the cause of the malfunction since operation of the system should result in some pressure drop. In addition, after any maintenance work on the outboard spoilers or outboard spoiler control system and prior to return of the aircraft to service, the spoiler system should be thoroughly checked to ascertain that the rivet in rod assembly P/N 4719278 is not sheared and that the system operates normally. If sheared, the rivet must be replaced or "Rod Assembly-Lateral Control Spoiler" P/N 4719278-1 and "Spring" P/N 4771197-1B installed per Douglas Drawing No. 5718924 "J" change, or FAA approved equivalent, prior to next flight. Pilots and flight engineers, in addition to authorized maintenance personnel, are authorized to perform the operational check. \n\n\t(b)\tThe operational check and inspection described in (a) may bediscontinued when rework per Douglas Drawing No. 5718924 "J" change, or FAA approved equivalent, is accomplished. \n\n\t(Douglas Service Bulletin No. A27-104 dated March 27, 1961, covers this subject.) \n\n\tThis directive effective June 9, 1961.
2005-10-19: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 747-100, 747-100B, 747-200B, 747-300, 747-400, 747-400D, 747SR, and 747SP series airplanes. This AD requires replacing or modifying the control panels for the galley cart lift and modifying related electrical cable assemblies, as applicable. This AD is prompted by reports of injuries to catering personnel and flight attendants who were loading or unloading galley carts on one deck when, due to a disabled or malfunctioning safety interlock door switch, the galley cart lift unexpectedly moved when it was activated from the control panel on the other deck. We are issuing this AD to ensure that the galley cart lift can be sent only from the deck on which it is in use, which will prevent unexpected movement of the cart lift that could result in possible injury to catering personnel or flight attendants.
54-02-01: 54-02-01 LYCOMING: Applies to All Model O-290-D2 Engines. Compliance required by February 1, 1954, and at each subsequent 100-hour inspection. In order to preclude the possibility of serious engine damage from incipient detonation under certain atmospheric and altitude conditions, the magneto timing of all Lycoming O-290-D2 engines must be set at 18 degrees BTC. To insure that the magneto timing does not change substantially from this setting, it should be checked and reset, if necessary, at each subsequent 100-hour inspection. (Lycoming Service Bulletin No. 169 covers this same subject, but recommends that the timing be checked every 50 hours.)
2005-10-12: This amendment adopts a new airworthiness directive (AD) for the Schweizer Aircraft Corporation (Schweizer) Model 269C, C-1, and D helicopters. This action requires inspecting, modifying, and testing the lateral control trim actuator assembly on certain serial-numbered helicopters. This amendment is prompted by three reported incidents of the inner spring tube separating from the lateral trim control housing resulting in a lateral cyclic control restriction. The actions specified in this AD are intended to prevent separation of the inner spring tube from the lateral trim control housing and the associated loss of trim control, increased local resistance to right cyclic stick movement, and an emergency landing or subsequent loss of control of the helicopter.
57-04-01: 57-04-01 CESSNA: Applies to All Model 310 Aircraft Prior to and Including Serial Number 35392, Except Those Incorporating Aileron Hinges Having P/N 0824006-1 and -2. Compliance required not later than April 15, 1957, and at 100-hour intervals of operation thereafter. Inspect the outboard aileron hinge for cracks in the upper flange, adjacent to the bearing. Any hinges found cracked should be replaced with the new, redesigned outboard aileron hinge (P/N 0824006-1 left, and 0824006-2 right). When the redesigned outboard aileron hinge is installed, this special inspection is no longer required. To facilitate hinge replacement, it is permissible to cut a standard size inspection opening in the wing lower skin, below the outboard hinge, attachment point. (Cessna Service Bulletin No. 310-7 covers this same subject.)
61-08-01: 61-08-01 DOUGLAS: Amdt. 273 Part 507 Federal Register April 8, 1961. Applies to All DC-6 and DC-7 Series Airplanes Which Do Not Have Antifuel Transfer Check Valves Installed In The Fuel System. \n\n\tCompliance required within the next 330 hours' time in service. \n\n\t(a)\tInstall a placard in the flight compartment in full view of the pilot to read as follows: "Fuel Transfer in Flight Prohibited." \n\n\t(b)\tRevise the operations procedures section covering fuel system management in the FAA approved airplane flight manual to incorporate fuel system operation procedures in accordance with the above placard and delete the item which states that check valves are installed. Flight manual revisions must be approved by FAA. \n\n\tWhen check valves to prevent transfer of fuel are installed, this AD is no longer applicable. \n\n\t(Douglas Service Bulletin No. DC-6-294 dated November 19, 1948, revised February 17, 1950, covers an acceptable installation of check valves.) \n\n\tThis directive effective May 9, 1961.
57-03-02: 57-03-02 DOUGLAS AND LOCKHEED: Applies to Lockheed Constellation Series Airplanes and Douglas DC-7 Series Airplanes With Wright Engines.\n\n\tCompliance required not later than December 1, 1957.\n\n\tUnder certain cold weather operating conditions it is possible for the fuel inlet strainer and other parts in the engine master control to become clogged with ice as a result of entrained water in the fuel freezing on the screen. This has caused a loss of power on all engines simultaneously.\n\n\tTo relieve strainer icing, a screen having a bypass valve Bendix P/N 366204 should be installed in replacement of the screen not having a bypass valve. Bendix Service Bulletin No. 797 covers this subject. It should be noted that the incorporation of the screen with bypass will not positively prevent power loss from fuel ice; therefore, work is continuing to develop a means to protect other portions of the master control downstream of the inlet strainer. If necessary, a supplement to this notewill be published when additional information becomes available.
74-18-11: 74-18-11 LOCKHEED: Amendment 39-1944. Applies to L-1011 airplanes certificated all categories. Compliance required as indicated, unless already accomplished: To prevent inadvertent activation of the automatic ground speed brakes, accomplish the following: (1) Prior to further flight, deactivate the automatic ground speed brake (AGSB) system and install placards, per (a), or (b) below. (a) Remove both DLC/AUTO-SPLR engage switch light mechanisms in the flight station overhead Flight Control Electronics System (FCES) control panel or accomplish Lockheed Alert Service Bulletin 093-22-A066. Install a placard on Captain's and F/O instrument panels directly below vertical speed indicators stating: "DLC/AUTO SPOILERS INOPERATIVE. REFER TO AFM OR EQUIVALENT." (b) Equivalent de-activation of AGSB system procedures and placards and equivalent system modifications may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (c) Aircraft may be flown to a base where maintenance required by this AD may be performed, per FAR's 21.197 and 21.199. (2) An operator may reactivate the DLC/AGSB system and remove the placards required by this AD on his fleet of L-1011 airplanes after all of the following are accomplished: (a) All FCES computers both in service and spares in an individual operators fleet are modified per Lockheed Service Bulletin 093-22-067, dated April 22, 1974, or later FAA- approved revisions, and that parts pooling is controlled such that no spares, aside from the type listed, will be installed. (b) All proximity logic boxes both in service and spares in an individual operator's fleet are modified per Lockheed Service Bulletin 093-31-029, dated August 8, 1974, or later FAA-approved revisions, and that parts pooling is controlled such that no spares, aside from the type listed, will be installed. (c) The L-1011 aircraft maintenance manual 22-00-00 is revised to incorporate the intent of temporary revision 4N, for use at the interval defined in the MRB document. (d) All Airplane Flight Manuals in an individual operator's fleet are revised to incorporate pages 3-28, 3-28.1, 3-28.2 and 3.29 of AFM LR 25225 and pages 3-29, 3-30, 3-30.1 and 3-30.2 of AFM LR 25925 dated August 15, 1974, or later FAA-approved revisions and aircraft operating procedures are amended and adopted immediately to include flight crew monitoring of the speed brake lever for proper DLC operation when moving the wing flap control lever beyond the 30 degree flap lever position. (3) An operator may reactivate the DLC/AGSB system and remove the placards required by this AD after making equivalent modifications approved by the Chief, Aircraft Engineering Division, FAA Western Region. This supersedes Amendment 39-1814 (39 F.R. 13258), AD 74-08-08. This amendment becomes effective September 3, 1974.
2005-10-11: The FAA is superseding an existing airworthiness directive (AD), which applies to all Boeing Model 737-300, -400, and -500 series airplanes. That AD currently requires repetitive inspections of certain connectors located in the main wheel well to detect discrepancies, and corrective action if necessary. This new AD instead mandates a modification. This AD is prompted by the development of a modification intended to address the unsafe condition. We are issuing this AD to prevent discrepancies of certain connectors located in the main wheel well. Those discrepancies could result in electrical arcing of the connectors, uncommanded closure of the engine fuel shut-off valves, and consequent in-flight loss of thrust or engine shutdown from lack of fuel.
60-24-03: 60-24-03 PIPER: Amdt. 225 Part 507 Federal Register November 17, 1960. Applies to PA-24 and PA-24 "250" Aircraft Serial Numbers 24-1 to 24-2161 Inclusive. Compliance required within the next 10 hours of operation or at the next periodic inspection, whichever occurs first, after the effective date of this directive. To prevent any interruption in fuel flow should the vent tubes become obstructed, the two fuel cell vent tubes which are located under the wings shall be modified in the following manner: Drill an 0.098-inch diameter hole (#40 drill) in the aft side of each tube three-fourths of an inch from the end. (Piper Service Bulletin No. 193 covers this subject.) This directive effective December 19, 1960.
75-22-11: 75-22-11 GENERAL ELECTRIC: Amendment 39-2398. Applies to Models CF6-6D and CF6-6D1 Turbofan Engines. Compliance required as indicated. To prevent possible disintegration of the high pressure turbine thermal shield accomplish the following: (a) Within the next 400 operating cycles after the effective date of this Airworthiness Directive, unless already accomplished, and every 400 cycles thereafter, borescope inspect the high pressure turbine thermal shield on all CF6-6D and CF6-6D1 engines except those noted in paragraph (b) below in accordance with the instructions of General Electric Service Bulletin (CF6-6) 72-603 dated October 1, 1975 or subsequent FAA Approved revision. (b) Inspection is not required: (1) On engines containing thermal shields replaced or modified in accordance with General Electric Service Bulletin (CF6-6) 72-442, dated October 8, 1973, Revision 1 or subsequent FAA Approved revision, or General Electric Service Bulletin (CF6-6) 72-502 dated October 23, 1974 or subsequent FAA Approved revision. (2) On engines incorporating thermal shields, General Electric part number 9687M67P08, assembly number 9687M67G12; part number 9687M67P09, assembly number 9687M67G13; part number 9687M67P12, assembly number 9687M67G16; or part number 9687M67P17, assembly number 9687M67G21. (c) For the purposes of this Airworthiness Directive, the definition of a "cycle" is the definition appearing in the General Electric CF6-6 Shop Manual GEK 9266, Section 72-00-00, Page 301 dated August 1, 1975 or subsequent FAA Approved revision. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to General Electric Company, Cincinnati, Ohio 45215. These documents may also be examined at the FAA Great Lakes Region, 2300 E. Devon Avenue, Des Plaines, Illinois 60018 and at FAA headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Great Lakes Region. This amendment becomes effective October 28, 1975.
60-16-03: 60-16-03 DOUGLAS: Amdt. 188 Part 507 Federal Register August 6, 1960. Applies to All DC3 Series Aircraft With Geared Rudder Tab Installations Based On Data Approved Prior to the Effective Date of This Airworthiness Directive. \n\n\tCompliance is required as indicated. \n\n\t(a)\tIn order to correct rudder force reversal tendencies on existing installations, the following shall be accomplished: \n\n\t\t(1)\tWithin two weeks after the effective date of this directive and until the aircraft has been flight tested or modified in accordance with this directive, a placard shall be placed in the aircraft in full view of the pilot which reads as follows: \n\n\t\t\t"Possible rudder force reversal and/or rudder lock may be experienced in this aircraft if rudder application is not coordinated with lateral control. Avoid yawed flight." \n\n\t\t\tThis placard shall be retained in the airplane and complied with until either of the applicable procedures described in (2) have been accomplished. \n\n\t\t(2)\tToremove the placard, either of the following procedures must be accomplished: \n\n\t\t(i)\tINSPECTION AND TEST OF THE GEARED TAB INSTALLATION. \n\n\t\t\t(a)\tCheck the rigging of the geared rudder tab installation in accordance with the manufacturer's approved installation data to prove conformity of this installation prior to the required flight test below. The results of the rigging check must be recorded in the aircraft logbook and signed by the individual making the check. \n\n\t\t\t(b)\tContact the nearest FAA Regional Office and make arrangements through the Flight Test Branch for having the aircraft tested. The results of this flight test must be recorded in the aircraft logbook and signed by the individual conducting the flight test. \n\n\t\t\t(c)\tIf the rudder control characteristics in the flight test are found to meet the requirements of Civil Air Regulations, Part 4a, Sections 4a.758-T (or Civil Air Regulations, Part 4b, Section 4b.157), the placard in paragraph (1) may be removed.(d)\tIf the rudder control characteristics in the flight test are found not to meet the requirements of Civil Air Regulations, Part 4a, Section 4a.758-T (or Civil Air Regulations, Part 4B, Section 4b.157), the placard may not be removed until a corrective design modification has been made, officially inspected and flight tested, and found to comply with the above regulations. \n\n\t\t(ii)\tREPLACEMENT WITH AN APPROVED NEW OR MODIFIED GEARED TAB INSTALLATION. \n\tAt such time as a "fix" or a new design installation has been developed, officially inspected and flight tested, and found to comply with the regulations, such an FAA approved modification or design may be installed in accordance with the manufacturer's specifications, a rigging and installation check made and recorded in the aircraft logbook by the individual who made the check. No mandatory flight tests will be necessary for such installations and the above-mentioned placard may be removed at this time. \n\n\t\t(b)\tTo precludethe installation on other aircraft of geared tabs of the same design which may have rudder force reversal tendencies, the following shall be accomplished prior to each approval: \n\n\t\t(1)\tAn official flight test shall be arranged with the nearest FAA Regional Office to determine that the installation complies with the regulations. The results of this flight test, as well as the prior inspection for conformity with approved installation data, must be recorded in the aircraft logbook and signed by the individuals conducting the installation inspection and flight test. \n\n\tThis directive shall become effective 30 days after the date of its publication in the Federal Register.
2005-10-13: The FAA is adopting a new airworthiness directive (AD) for Rolls-Royce Corporation (RRC) (formerly Allison Engine Company) 250-B17B, - B17C, -B17D, -B17E, -C20, -C20B, -C20F, -C20J, -C20S, and -C20W turboprop and turboshaft engines that do not have turbine energy absorbing ring, RRC part number (P/N) 23035175, or an equivalent FAA- approved serviceable turbine energy absorbing ring, installed. This AD requires installation of a turbine energy absorbing ring in the plane of the 1st stage turbine wheel. This AD results from an unacceptable rate of uncontained 1st stage turbine wheel failures. We are issuing this AD to minimize the risk of uncontained 1st stage turbine wheel fragments from causing damage to the aircraft or damage to the second engine on twin-engine installations, which could lead to loss of control and loss of the aircraft.
2015-25-06: We are superseding Airworthiness Directive (AD) 2010-06-04, for certain Airbus Model A300 B2-1C, B2-203, B2K-3C, B4-103, B4-203, B4-2C airplanes; Model A310 series airplanes; Model A300 B4-600 series airplanes; and Model A300 B4-600R series airplanes. AD 2010-06-04 required repetitive inspections to detect cracks of the pylon side panels (upper section) at rib 8; and corrective actions if necessary. This new AD continues to require repetitive inspections for cracking of the pylons 1 and 2 side panels (upper section) at rib 8 with reduced compliance times, and corrective actions if necessary. This AD also requires repetitive post-repair and post-modification inspections and repair if necessary. This AD also removes certain airplanes having a certain modification from the applicability. This AD was prompted by reports of cracks found on pylon side panels at rib 8 and a fleet survey and updated fatigue and damage tolerance analyses. We are issuing this AD to detect and correct cracking of pylon side panels (upper section) at rib 8, which could lead to reduced structural integrity of the pylon primary structure, which could cause detachment of the engine from the fuselage.
73-24-01: 73-24-01 ROCKWELL INTERNATIONAL: Amdt. 39-1743. Applies to Rockwell Commander Model 112 airplanes, Serial Numbers 3 through 120, certificated in all categories. Compliance required before further flight unless already accomplished. To prevent failure of the aileron hinges and/or the elevator trim tab hinges, accomplish the following: (a) Inspect all hinge halves of both ailerons and both elevator trim tabs from underside of aircraft to determine if the hinges are of the formed type made by rolling the edge of a 0.040 inches thick flat sheet or of the extruded type 0.060 inches thick. (b) If extruded hinge halves are found in all locations, no further action is required. (c) If a formed hinge piece is found, that complete hinge must be replaced with Rockwell Commander Part No. 42251-1 for the aileron hinges or Rockwell Commander Part No. 44020-5 for the elevator trim tab hinges before further flight. If no cracks are visually evident in any formed hinges,the airplane may be flown in accordance with FAR 21.197 to a base where the replacement can be performed. Rockwell International Service Bulletin No. SB-112-6 pertains to this same subject. This amendment becomes effective November 19, 1973.
60-10-01: 60-10-01 BELL: Amdt. 146 Part 507 Federal Register May 10, 1960. Applies to All Helicopter Models: 47B, 47B3, 47D, 47D1, 47G, and 47H1, all Serial Numbers; 47G2 Serial Numbers 1327 Through 2467, 2469, 2470, 2472 Through 2477, 2556 Through 2558; 47J Serial Numbers 1420 Through 1776 (Except For Helicopters On Which Kit No. 47-3410-1 (333SI) Has Been Installed); 47E, and 47K. Compliance required as indicated except Model 47G2, Serial Numbers 2451, 2452, 2457, 2459 through 2467, 2469, 2470, 2472 through 2477, 2556 through 2558, for which compliance date is September 2, 1960. As the result of a number of recent failures of the scissor lever pivot bolts due to excessive wear, the following is required unless already accomplished. (a) Prior to June 30, 1960, except 47E and 47K as to which compliance is required prior to August 15, 1960, inspect the scissor lever pivot bolts, AN 174-31, and bolt holes in the brackets of the collective pitch sleeve weld assembly, P/N 47-150-117-5 for wear. Wear limits and reinspection intervals are specified in the following items (1), (2), (3), and (4). (1) If the diameter of the two AN 174-31 bolts is less than 0.2465 inch in any area, bolts must be replaced prior to next flight. (2) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is 0.2550 inch or more, install four bushings, P/N 47-150-260-3 or equivalent, and new AN 174-31 bolts within the next 25 hours' time in service. (3) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is between 0.2500 and 0.2550 inch, the bolts and bolt holes must be reinspected dimensionally every 25 hours' time in service until bushings P/N 47-150-260-3 are installed. (4) If the diameter of the bolt holes in the brackets of the collective pitch sleeve assembly is 0.2500 inch or less, the bolts and bolt holes must be reinspected dimensionally every 100 hours' time in service until bushings P/N 47-150-260-3 are installed. (b) Upon installation of the bushings P/N 47-150-260-3, the bolts and bushing holes must be inspected every 300 hours' time in service thereafter. (c) Upon installation of the four bushings, P/N 47-150-260-3, designate the reworked collective pitch sleeve weld assembly as P/N 47-150-117-21. (Bell Service Bulletin No. 129SB, dated March 18, 1960, covers this same subject.) Revised July 15, 1960. Revised August 19, 1960.
2005-10-10: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier Model CL-600-2B19 (Regional Jet Series 100 & 440) airplanes. This AD requires revising the Airworthiness Limitations section of the Instructions for Continued Airworthiness of the Canadair Regional Jet Maintenance Requirements Manual by incorporating new repetitive detailed inspections of the secondary load path indicator for the horizontal stabilizer trim actuator (HSTA). This AD is prompted by a report of a potential failure of the horizontal stabilizer trim actuator (HSTA) secondary nut in conjunction with a latent failure of the HSTA primary load path discovered during sampling program activities. We are issuing this AD to detect and correct latent failure of the primary load path of the HSTA, which, in conjunction with a failure of the HSTA secondary nut, could result in loss of horizontal trim control and consequent reduced controllability of the airplane.
70-25-09: 70-25-09 AEROSTAR: Amdt. 39-1127. Applies to Model 601, S/N's 61-001 through 61- 0070, and to all Model 600 airplanes equipped with the Aerostar Model 601 oxygen system. Compliance required within the next 10 hours time in service after the effective date of this AD unless already accomplished. To prevent possible short circuiting of the microphone jack connection by contact with the oxygen outlet receptacle due to their proximity to each other, accomplish the relocation specified in Aerostar Aircraft Corporation Service Bulletin No. S.B. 600-26 dated November 6, 1970, or later FAA-approved revisions, or other equivalent modification approved by the Chief, Engineering and Manufacturing Branch, Southwest Region, FAA. This amendment becomes effective to all known owners of Aerostar Model 601 airplanes and Model 600 airplanes equipped with the Aerostar Model 601 oxygen system upon receipt of individual copies mailed December 8, 1970 and to all other persons on December18, 1970.
96-12-07: This amendment supersedes an existing airworthiness directive (AD), applicable to Teledyne Continental Motors (TCM) (formerly Bendix) S-20, S-1200, D-2000, and D-3000 series magnetos equipped with impulse couplings, that currently requires inspections for wear, and replacement, if necessary, of the impulse coupling assemblies. This amendment requires replacement, if necessary, of worn riveted impulse coupling assemblies with serviceable riveted impulse couplings or snap ring impulse couplings. This amendment is prompted by the availability of an improved design for the impulse coupling assembly. The actions specified by this AD are intended to prevent magneto failure and subsequent engine failure.