Results
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.
2005-10-08: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier Model DHC-8-102, -103, -106, -201, -202, -301, -311, and -315 airplanes. This AD requires operators to install torque tube catchers on the control columns of the flight controls. This AD is prompted by the discovery that a single malfunction of the torque tube could result in both flight control columns being supported by only one self-aligning bearing. We are issuing this AD to prevent the torque tube from fouling against the underfloor control cables, which could result in reduced controllability of the airplane.
75-11-01: 75-11-01 CESSNA: Amendment 39-2204 as amended by Amendment 39-2548. Applies to all Cessna Model 320 Series airplanes. Compliance: Required as indicated, unless already accomplished. To detect and correct fuel line chafing or fuel leaks behind the engine firewalls, within 100 hours' time in service after the effective date of this AD, accomplish the following: A) Visually inspect fuel lines routed in the area behind the engine firewalls and the front wing spars for leaks or chafing in accordance with Cessna Service Letter ME70-39, Revision I, dated April 4, 1975, or subsequent revisions. B) If as a result of the inspection required by Paragraph A, fuel line leaks or damage is found, prior to further flight, replace the affected line with an airworthy part in accordance with Cessna Service Kit SK402-8C dated February 20, 1975, or subsequent revisions. C) On Models 320D, 320E and 320F airplanes, in addition to the inspection required in Paragraph A, add additional supporting clamps to the crossover fuel lines in accordance with Cessna Service Kit SK402-8C, dated February 20, 1975, or subsequent revisions. D) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Amendment 39-2204 became effective May 20, 1975. This amendment 39-2548 becomes effective March 22, 1976.
2001-11-08: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 747-400, 757-200, 767-200, and 767-300 series airplanes, that currently requires repetitive checks to detect certain failures in the warning electronic unit (WEU) or modular avionic warning electronic assembly (MAWEA); repetitive tests to detect any failure of tactile, visual, or aural alerts generated by the WEU or MAWEA; and corrective action, if necessary. This amendment makes these requirements applicable to other airplanes on which the defective power supplies may be installed, eliminates the repetitive tests for certain airplanes, and increases the interval for the repetitive tests for certain other airplanes. This amendment also requires replacing any subject power supply in the WEU or MAWEA with a new, modified, or serviceable power supply. The actions specified by this AD are intended to prevent failure of the WEU or MAWEA power supplies, which could result in loss of visual, aural, and tactile alerts to the flightcrew. Absence of such alerts could result in the flightcrew being unaware that an immediate or appropriate action should be taken in the event of an unsafe condition.
59-05-04: 59-05-04 ERCOUPE: Applies to All Models 415-C, 415-CD, 415-D, E and G. Compliance to be accomplished within the next 100 hours of operation. Frequent failures of the rear spar center section have been found on Ercoupe Model aircraft. These failures follow the same pattern in that the rear spar P/N 415-13048 L/R failed due to cracking of the upper flange in the area of the intersection of the rear spar with the fuselage side on either the right or left spar assembly. Repairs made in the field with gusset plates have been found to be only partially satisfactory and in most instances did not keep the crack from progressing into the spar web. This damage to the spar has been attributed to the following: (1) Rough landings coupled with a lack of fluid in the oleo struts. (2) Taxiing at high speeds over rough terrain. (3) A combination of (1) and (2) with the structure weakened by corrosion due to no protective coating of the spar. In view of the above, it is mandatory that therear spar on Forney (Ercoupe) aircraft be inspected and action taken as follows: (1) If no damage or cracks are found, the spar must be reinforced by stiffener angle P/N F-13109 or equivalent. The spar may be considered satisfactory if previously reinforced with P/N 415-13108 or equivalent. (2) If damage or crack exists but does not extend into the spar web, a repair may be made by the addition of stiffener P/N F-13109 or equivalent provided an inspection every 100 hours of service life thereafter discloses no further progressive damage. If damage is found to progress, then a new spar and stiffener must be installed. If damage or crack extends into the spar web, the spar must be replaced. (3) If a new spar and stiffener is installed, the 100-hour inspection requirement in (2) above may be eliminated. (Forney Manufacturing Company, Aviation Division, Fort Collins, Colorado, Service Memorandum 53A supersedes Service Memorandum 53 and covers this same subject.) This supersedes AD 57-13-03.
75-23-07: 75-23-07 CESSNA: Amendment 39-2418. Applies to Models 310Q/T310Q (Serial Numbers 310Q0401 thru 310Q1160) and 310R/T310R (Serial Numbers 310R0001 thru 310R0330) airplanes. Compliance: Required as indicated, unless already accomplished. To assure proper operation of the emergency exit window, within the next 50 hours' time in service after the effective date of this AD, accomplish the window modification, screw inspection, trim modification and placard installation, as applicable, in accordance with Cessna Service Letter ME75-26 or later approved revision or by any equivalent method of compliance approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective November 13, 1975.
2005-10-03: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 777-200 and -300 series airplanes. This AD requires modification of the operational program software (OPS) of the air data inertial reference unit (ADIRU). This AD is prompted by a report of the display of erroneous heading information to the pilot due to a defect in the OPS of the ADIRU. We are issuing this AD to prevent the display of erroneous heading information to the pilot, which could result in loss of the main sources of attitude data, consequent high pilot workload, and subsequent deviation from the intended flight path.
58-16-01: 58-16-01 PIPER: Applies to Model PA-20 Serial Numbers 20-1 to 20-1121 Inclusive, and Model PA-22 Serial Numbers 22-1 to 22-6087 Inclusive, That Have Cigar Lighters Installed. Compliance required within next 100 hours of operation. To preclude the possibility of blowing the master fuses, due to a short circuit caused by a faulty cigar lighter element, install a 15-ampere standard fuse, Bussman No. AGC-15 or equivalent, in the wire between the cigar lighter and the ammeter. (Piper Service Bulletin Number 163A covers the same subject.)
2010-17-12: The FAA is superseding an existing 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: Strip results from some of the engines listed in the applicability section of this AD revealed excessively corroded low- pressure turbine disks stage 2 and stage 3. The corrosion is considered to be caused by the environment in which these engines are operated. Following a life assessment based on the strip findings it is concluded that inspections for corrosion attack are required. The action specified by this European Aviation Safety Agency (EASA) AD 2008-0122 was intended to avoid a failure of a low- pressure turbine disk stage 2 or stage 3 due to potential corrosion problems which could result in uncontained engine failure and damage to theairplane. It has been later realized that the same unsafe condition could potentially occur on more serial numbers for the Tay 650-15 engines and on the Tay 651-54 engines. This AD, superseding EASA AD 2008-0122, retaining its requirements, is therefore issued to expand the Applicability in adding further engine serial numbers for the Tay 650-15 engines and in adding the Tay 651-54 engines. We are issuing this AD to detect corrosion that could cause the stage 2 or stage 3 disk of the LP turbine to fail and result in an uncontained failure of the engine.
69-05-04: 69-05-04 SCHLEICHER: Amdt. 39-730. Applies to Schleicher Model ASK-13 gliders, Serial Numbers 13000 through 13104, 13108, 13109, except Serial Numbers 13071 and 13096. Compliance required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. To prevent the landing gear end buffer plate from slipping inside the rubber buffer, replace the 2.76-inch diameter buffer plate located on the left and right landing gear with a 3.55- inch diameter buffer plate, in accordance with Schleicher Modification No. 3, dated September 4, 1968, or later LBA-approved issue or an FAA-approved equivalent. This amendment becomes effective April 5, 1969.
2005-10-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A319, A320, and A321 series airplanes. This AD requires repetitive inspections of the left- and right-side main landing gear (MLG) side-stay cuff lugs and down-lock spring attachments for evidence of cracked or fractured side-stay cuff lugs or down-lock spring attachments, and repair if necessary. This AD also provides for optional terminating action for the repetitive inspections. This action is necessary to prevent failure of the MLG side-stay cuff lugs or down-lock spring attachments, which could result in improper down-lock of the MLG during a freefall extension, and possible collapse of the MLG. This action is intended to address the identified unsafe condition.
57-13-08: 57-13-08 PIPER: Applies to Model PA-23 Serial Numbers 23-1 to 23-729, Inclusive. Compliance required by September 1, 1957. To prevent inadvertent retraction of the landing gear due to malfunction of the landing gear hydraulic system, install an antiretraction device (Piper Kit No. 754140 or equivalent) that will prevent start of the retraction cycle and consequent landing gear collapse while the aircraft is on the ground. (Piper Service Bulletin No. 145 of December 31, 1956, covers this subject.)
97-11-06: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Pratt & Whitney PW4164 and PW4168 series turbofan engines. This action requires initial and repetitive visual inspections of the fan blades for surface damage and cracks, initial and repetitive lubrication of the fan blade part span shrouds, a one time ultrasonic inspection (UI) of the fan blade root attachment area for cracks, and a one time fan blade root attachment front corner radii inspection for proper dimension. Also, this AD requires visual inspection of the fan blades and removal of fan blades damaged by a bird strike as well as removal of blades immediately adjacent to damaged blades. In addition, this AD requires installation of an improved fan blade assembly as terminating action to the inspection requirements of this AD. This amendment is prompted by a report of a high N1 rotor imbalance and liberation of the fan containment system causing loss of structural support of theengine inlet cowl, following loss of a fan blade during a test. The actions specified in this AD are intended to prevent fan blade failure and separation at the root section, which could result in high N1 rotor imbalance, and liberation of the fan containment system, which can hazard the aircraft.
71-24-05: 71-24-05 PRATT & WHITNEY: Amendment 39-1137 as amended by Amendment 39-1847. Applies to all Pratt & Whitney Aircraft JT12A series turbojet and JFTD12A series turboshaft engines which incorporate any of the compressor rotor discs with serial numbers listed in Pratt & Whitney Aircraft Turbojet Engine Service Bulletin No. 3421 dated 4 August 1971 as amended by Revision No. 3, dated December 10, 1973. Compliance required as indicated. (a) To preclude compressor rotor disc failure as the result of reduced life from traces of lead - (1) For those discs listed in Pratt & Whitney Aircraft turbojet Engine Service Bulletin No. 3421, dated August 4, 1971, remove from service the listed discs prior to reaching the revised life limit or within the next 30 cycles in service after November 26, 1971, whichever comes later. (2) For those discs not listed in the above Service Bulletin but listed in Revision No. 3, dated December 10, 1973, remove from service the listed discs prior to reaching the revised life limit or within the next 30 cycles in service after the effective date of this amendment, whichever comes later. The manufacturer's Service Bulletin identified and described in this directive is incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received this document from the manufacturer may obtain copies upon request from Pratt & Whitney Aircraft Division of United Aircraft Corporation, East Hartford, Connecticut 06108. This document may also be examined at the FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts or at the FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on the AD which includes the incorporated material in full is maintained by the FAA in its headquarters in Washington, D.C., and in the New England Region. Upon submission of substantiating data through an FAA Maintenance Inspector by an owner or operator to the Chief, Engineering and Manufacturing Branch, FAA New England Region, compliance time may be adjusted. The incorporation by reference provision in this document was approved by the Director of the Federal Register on June 19, 1967. Amendment 39-1337 was effective November 26, 1971. This Amendment 39-1847 becomes effective May 31, 1974.