Results
55-24-01: 55-24-01 LUSCOMBE: Published in 21 FR 9540 on December 4, 1956, and as amended in 22 FR 2416 on April 11, 1957, is further amended by Amendment 39-1565 and 39-1640. Applies to All 8 Series Aircraft Except Model 8-F with Serial Numbers S-1 and Up. To be accomplished by March 1, 1956, and thereafter at intervals not to exceed 12 calendar months from last inspection. Extreme surface corrosion has been found to exist inside the fuselage spar carry through structures P/N 28018 and 28019 of Luscombe Series 8 aircraft, particularly in those airplanes which are located near coastal areas. If allowed to progress, such corrosion could deteriorate the spar carry through members until a structural failure occurred. This corrosion is internal and cannot be detected by an external inspection. Therefore, the inside surfaces of the spar carry through members must be inspected. This may be accomplished by either of the two following acceptable methods: (1) Remove wings from the airplane and also the wing attachment fittings. The ends of both the front and rear spar superstructures will then be open so that an internal inspection of these hat-section members can be made. (2) Use of this method of inspection will not require the removal of the wings from the airplane. One-half inch holes may be drilled through the top wing skin directly over each spar carry through member so that a visual inspection can be made directly into the bottom of the hat sections. The airframe structure had adequate margins of safety in this area so that the existence of the 1/2-inch inspection holes will not impair the structural integrity of the airplane. Five of these 1/2-inch holes should be drilled over each of the spar carry through hat sections, one hole at the middle of each spar carry through, one hole 5 inches from each outboard end of the wing attachment fittings and one hole approximately centrally located between this latter hole and the middle hole. Thiswill provide a distance of approximately 7 1/2 inches between holes and should render it possible to inspect all of the internal surface of the hat-section spar carry through members. After the inspection has been made, the 1/2-inch holes must be covered with a small patch of aircraft fabric doped to the surface of the wing skin or by the insertion of a rubber or neoprene seal plug, or equivalent. This method will also provide a ready means of rechecking the spar carry through members for corrosion during the time of subsequent inspections. If any evidence of corrosion is found to exist, the affected spar carry through member should be removed and replaced with an identical new part. The above inspections may be discontinued if both spar carry through structures are replaced with new parts that are identical to the original and properly anodized and painted to prevent corrosion, or if an equivalent modification is approved by the Chief, Engineering and Manufacturing Branch,FAA Southern Region. Amendment 39-1565 became effective December 2, 1972. This Amendment 39-1640 becomes effective May 22, 1973.
2008-22-07: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Subsequent to accidents involving Fuel Tank System explosions in flight * * * and on ground, * * * Special Federal Aviation Regulation 88 (SFAR88) * * * required * * * a design review against explosion risks. * * * * * The unsafe condition is the potential of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
96-08-03: This amendment adopts a new airworthiness directive (AD), applicable to Flight Trails Helicopters, Inc. hardpoint assemblies, installed in accordance with Supplemental Type Certificate (STC) No. SH6080NM, or in accordance with Federal Aviation Administration (FAA) Form 337, "Major Repair and Alteration," approved on McDonnell Douglas Helicopter Systems (MDHS) Model 369D, 369E, 369F, 369FF, and 500N helicopters, that requires removing any Flight Trails Helicopters, Inc. hardpoint assembly not identified by part number (P/N) and serial number (S/N). This amendment is prompted by two incidents in which the hardpoint assembly used to support a search light or night vision system reportedly failed. The actions specified by this AD are intended to prevent failure of the hardpoint assembly, separation of the hardpoint assembly from the helicopter, and subsequent contact between the hardpoint assembly and the fuselage or rotor system of the helicopter.
2011-01-53: We are adopting a new airworthiness directive (AD) for the products listed above. This emergency AD was sent previously to all known U.S. owners and operators of these airplanes. This AD supersedes Emergency AD 2011-01-51, requires an immediate functional test of the fuselage drain holes, and requires sending a report of the results to the FAA. This AD also allows, with noted exceptions, for the return/ position of the airplane to a home base, hangar, maintenance facility, etc. This AD was prompted by reports of water accumulation in the belly of the fuselage that froze and caused the flight controls to jam. We are issuing this AD to prevent water or fluid from accumulating in the belly of the fuselage and freezing when the aircraft reaches and holds altitudes where the temperature is below the freezing point. This condition could cause the flight controls to jam with consequent loss of control.
2010-06-18: The FAA is adopting a new airworthiness directive (AD) for IAE V2500-A1, V2522-A5, V2524-A5, V2525-D5, V2527-A5, V2527E-A5, V2527M-A5, V2528-D5, V2530-A5, and V2533-A5 turbofan engines. This AD requires a onetime fluorescent penetrant inspection of certain vortex reducers for cracks. This AD results from reports of fractured vortex reducers found at shop visits. We are issuing this AD to inspect for cracks in the vortex reducer. Cracks in the vortex reducer could result in an uncontained failure of the high-pressure (HP) compressor stage 3-8 drum and subsequent damage to the airplane.
96-07-11: This amendment adopts a new airworthiness directive (AD), applicable to all Beech Model BAe 125-1000A and Hawker 1000 series airplanes, that requires a detailed visual inspection to detect chafing damage to the hydraulic pipes adjacent to the hydraulic module, and various follow-on actions. This amendment is prompted by reports of chafing damage between hydraulic pipes at three locations in the rear equipment bay adjacent to the hydraulic module. The actions specified by this AD are intended to prevent such chafing damage to the hydraulic pipe and subsequent hydraulic fluid leakage, which could lead to failure of essential airplane systems.
81-07-01 R1: 81-07-01 R1 GOVERNMENT AIRCRAFT FACTORIES (GAF): Amendment 39-4068 as amended by Amendment 39-4388. Applies to Nomad Model N22B (Serial Nos. N22B-5 and up) and N24A (Serial Nos. N24A-42 and up) airplanes, certificated in all categories. Compliance required within the next 25 hours time in service after the effective date of this AD, unless already accomplished. To prevent loss of structural rigidity which could result in hazardous tail flutter and subsequent loss of airplane control, accomplish the following (a) Modify the horizontal stabilizer to incorporate mass balance weights, improved riveting, addition of access panels, and structural reinforcements to the horizontal stabilizer in accordance with Part 2, "ACCOMPLISHMENT INSTRUCTIONS," of Government Aircraft Factories Nomad Service Bulletin NMD-55-10, Revision 1, dated May 2, 1980, or an FAA- approved equivalent. All AGS pop rivets removed as a result of this modification must be replaced with MS20470AD rivets.NOTE: AGS 2048-420-BS 1/8 inch pop rivets are acceptable in the access panel door as specified in Service Bulletin NMD-55-10, Rev. 1. (b) Install high strength steel trim tab control brackets, self-aligning control rod ends, and longer control rods in accordance with Part 2, "ACCOMPLISHMENT INSTRUCTIONS," of Government Aircraft Factories Nomad Service Bulletin NMD-55-8, Revision 3, dated September 1, 1980, or an FAA-approved equivalent. (c) Modify the horizontal stabilizer pivot bracket attachments in accordance with Part 2, "ACCOMPLISHMENT INSTRUCTIONS," of Government Aircraft Factories Nomad Alert Service Bulletin ANMD-55-13, Revision 1, dated August 22, 1980, or an FAA-approved equivalent. (d) In accordance with FAR 21.197 and 21.199 the airplane may be flown to a location where these modifications can be accomplished. (e) If an equivalent means of compliance is used in complying with paragraphs (a) through (c) of this AD, that equivalent means must be approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. 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 Government Aircraft Factories, 226 Lorimer Street, Port Melbourne 3207 Vic., Australia. These documents may be examined at FAA, Pacific-Asia Region, Engineering and Manufacturing District Office, 300 Ala Moana Blvd., Room 7321, Honolulu, Hawaii, and at FAA Headquarters, Room 916, 800 Independence Avenue, S.W., Washington, DC. Amendment 39-4068 became effective April 2, 1981. This amendment 39-4388 becomes effective on June 10, 1982.
82-23-03 R1: 82-23-03 R1 GARRETT TURBINE ENGINE COMPANY: Amendment 39-4488 as amended by Amendment 39-4761. Applies to all Garrett TFE731-2, -3, -3R, -3A, -3AR, -3B and -3BR series turbofan engines. Compliance required as indicated unless already accomplished. (a) To prevent failure of high pressure compressor impellers due to cracks originating from a double radius in the seal relief region or a manufacturing notch in the forward balance ring, inspect and rework compressor impeller, Part Number 3070274-1 or 3072639-1, in accordance with the procedures contained in the accomplishment instructions section of Garrett Service Bulletin Number TFE731-72-3239RWK, dated September 13, 1982, or equivalent means approved by the Manager, Western Aircraft Certification Field Office, ANM- 170W, Northwest Mountain Aircraft Certification Division. Inspect in accordance with the following schedule: (1) Impellers which have been operated in TFE731-3, -3R, -3A, -3AR, -3B, or -3BR engines during any portion of their service lives must be inspected prior to exceeding 5,100 cycles in service. NOTE: The engine life limited parts log card for each impeller is located with the impeller or in the engine log book. This card lists the engine serial number of each engine in which the impeller has operated. The 5 digit serial numbers of the TFE731-3 series engines begin with the 2 digits 75, 76, 77, 78, 80, 82, 83, 84, 85, 87, 90, or 94. (2) Impellers which have been operated in TFE731-2 engines, only, must be inspected prior to exceeding 6,200 cycles in service. NOTE: The 5 digit serial numbers of the TFE731-2 series engines begin with the 2 digits 73, 74, 81, 86, 88, or 89. Impellers found to have crack indications are to be removed from, and not returned to, service. Impellers not found to have crack indications may be returned to service after having the seal relief region and notch in the forward balance ring recontoured in accordance with the procedures contained in the accomplishment instructions section of Garrett Service Bulletin Number TFE731-72-3239 RWK, dated September 13, 1982, or equivalent means approved by the Manager, Western Aircraft Certification Field Office, ANM-170W, Northwest Mountain Aircraft Certification Division. (b) To prevent cracking and possible failure of the following listed fan and compressor rotor discs, used in TFE 731-2 series engines, the life limits on these parts have been reduced. Unless already accomplished, remove rotor discs from service prior to reaching the life limits shown below or, before accumulation of an additional 30 cycles in service after March 7, 1975, whichever occurs later: Component Part Number Life Limit in Cycles Fan Disc 3072162 10,000 First Stage Compressor 3072190 3,000 Second Stage Compressor 3072191 3,700 Third Stage Compressor 3072192 1,200 Fourth Stage Compressor 3072193 1,200 NOTE: For the purpose of this AD, a cycle is considered as any engine operating sequence involving an engine start, at least one acceleration to 80 percent low pressure rotor speed or above, and shutdown. NOTE: Garrett FAA approved Service Bulletin TFE731-72-3001, Revision 15, dated June 30, 1983, provides the cyclic life limits for all life limited components not covered by this AD. This bulletin may be revised in the future, with the approval of the FAA, in order to provide life limit increases if appropriate. (c) Service life limits have been assigned to the following specific parts used in the TFE731-2 series engines: (1) Replace the High Pressure Turbine Blades, Part Number 3072111-1, (used in high pressure turbine rotor assembly, part number 3070098), with serviceable turbine blades before exceeding 1,000 hours total time in service, or before exceeding 200 additional hours time in service after March 7, 1975, whichever occurs later. (2) Model TFE731-2 engines Serial Numbers P-74101 through P-74113 and Serial Numbers P- 73106 through P-73184 not modified by incorporation of Power Section Change Number 22: Replace the Pinion Gear Assembly, Part Number 3071626-1, and Sun Gear, Part Number 3071598-1, with a serviceable Pinion Gear Assembly and Sun Gear before exceeding 500 hours total time in service, or before exceeding an additional 50 hours time in service after March 7, 1975, whichever occurs later. (3) Model TFE731-2 engines Serial Numbers P-74101 through P-74138 and Serial Numbers P- 73106 through P-73209 not modified by incorporation of Power Section Change Number 41 or 44: Replace the Third Stage Compressor Stator Assembly, Part Number 3070279-7, -8, or -10, with a serviceable Third Stage Compressor Stator Assembly before exceeding 500 hours total time in service, or before exceeding an additional 50 hours time in service after March 7, 1975, whichever occurs later. NOTE: Power Section Change Numbers are annotated on the Engine Data Plate affixed to the engine. 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). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Garrett Turbine Engine Company, P.O. Box 5217, Phoenix, Arizona 85010. These documents also may be examined at FAA New England Region, Office of the Regional Counsel, 12 New England Executive Park, Burlington, Massachusetts. This Amendment 39-4761 amends Amendment 39-4488 (47 FR 50462), AD 82-23-03, and supersedes Amendment 39-2116 (40 FR 8939), AD 75-05-12. Amendment 39-4488 became effective November 15, 1982. This Amendment 39-4761 becomes effective December 9, 1983.
96-07-10: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-100, -200, and -300 series airplanes, that requires an inspection to determine if hinge bolts and nuts are installed in the overhead stowage bins, and the installation of hinge bolts and nuts, if necessary. This amendment is prompted by reports that overhead stowage bins in the passenger compartment have fallen out of position due to missing hinge bolts. The actions specified by this AD are intended to ensure that hinge bolts are installed in the overhead stowage bins. Missing hinge bolts could result in the overhead stowage bins falling out of position and injuring airplane occupants.
2011-03-05: We are 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: The TC Holder received from operators, whose fleets are operated in demanding operating-conditions and with very frequent Short Take- Off and Landing (STOL) operations, reports of cracks located in the web of fuselage frame 19. On 05 February 2007, EASA issued Airworthiness Directive (AD) 2007-0028 which mandated Alert Service Bulletin (ASB) 228-266 and required an inspection of the frame 19 on all Dornier 228 aeroplanes. In addition, the TC Holder also initiated a flight-test campaign including strain measurements as well as finite element modelling and fatigue analyses to better understand the stress distribution onto the frame 19 and the associated structural components. The results of these investigations confirmed that STOL operations diminish extensively the fatigue life of the frame 19. Fuselage frame 19 supports the rear attachment of the Main Landing Gear (MLG). This condition, if not corrected, could cause rupture of frame 19, leading to subsequent collapse of a MLG. We are issuing this AD to require actions to correct the unsafe condition on these products.
2010-06-07: We are adopting a new airworthiness directive (AD) for the specified model helicopters. This AD results from a mandatory continuing airworthiness information (MCAI) AD issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. The MCAI AD states that the AD is issued following a manufacturing nonconformity found on one batch of the servo-control caps. With a defective servo-control, rotation of the distributor might not be stopped mechanically since only friction of inner seals holds the distributor sleeve in its position. The AD actions are intended to address the unsafe condition created by a manufacturing nonconformity found on one batch of servo-control caps. If not corrected this condition could cause untimely movements of servo-controls, which are used on main and anti-torque rotors, and lead to the loss of control of the helicopter.
2011-03-04: We are superseding an existing airworthiness directive (AD) for the products listed above. AD 2009-09-09 currently requires repetitive inspections of the rudder hinges and the rudder hinge brackets for damage, i.e., cracking, deformation, and discoloration. If damage is found during any inspection, AD 2009-09-09 also requires replacing the damaged rudder hinge and/or rudder hinge bracket. This new AD retains the inspection requirements of AD 2009-09-09, adds airplanes to the Applicability section, and adds a terminating action for the repetitive inspection requirements. This AD resulted from the manufacturer developing a modification that terminates the repetitive inspections and from the manufacture adding airplane serial numbers into the Applicability section. We are issuing this AD to detect and correct damage in the rudder hinges and the rudder hinge brackets, which could result in failure of the rudder. This failure could lead to loss of control.
2011-02-07: We are superseding an existing airworthiness directive (AD) for General Electric Company (GE) CF6-45 and CF6-50 series turbofan engines with certain low-pressure turbine (LPT) rotor stage 3 disks installed. That AD currently requires initial and repetitive borescope inspections of the high-pressure turbine (HPT) rotor stage 1 and stage 2 blades for wear and damage, including excessive airfoil material loss. That AD also requires fluorescent-penetrant inspection (FPI) of the LPT rotor stage 3 disk under certain conditions and removal of the disk from service before further flight if found cracked. That AD also requires repetitive exhaust gas temperature (EGT) system checks (inspections). This AD requires HPT rotor stage 1 and stage 2 blade inspections and EGT system inspections. This AD also requires FPI of the LPT rotor stage 3 disk under certain conditions, removal of the disk from service before further flight if found cracked, and an ultrasonic inspection (UI) of the LPT rotor stage 3 disk forward spacer arm. This AD also requires initial and repetitive engine core vibration surveys and reporting to the FAA any crack findings, disks that fail the UI, and engines that fail the engine core vibration survey. This AD was prompted by reports received of additional causes of HPT rotor imbalance not addressed in AD 2010-12-10, and two additional LPT rotor stage 3 disk events. We are issuing this AD to prevent critical life-limited rotating engine part failure, which could result in an uncontained engine failure and damage to the airplane.
2011-01-14: We are 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: The current Aircraft Maintenance Manual (AMM) of PC-6 B2-H2 and B2-H4 models does not include a Chapter 04 in the Airworthiness Limitations Section (ALS). For PC-6 models other than B2-H2 and B2- H4, no ALS at all is included in the AMM. With the latest Revision 12 of the AMM, a new Chapter 04 has been introduced in the AMM for PC-6 B2-H2 and B2-H4 models. For PC-6 models other than B2-H2 and B2-H4, a new ALS document has been implemented as well. These documents include the Mandatory Continuing Airworthiness Information (MCAI) which are maintenance requirements and/or airworthiness limitations developed by Pilatus Aircraft Ltd and approved by EASA. Failure to comply with these MCAI constitutes an unsafe condition. We are issuing this AD to require actions to correct the unsafe condition on these products.
96-07-03: This amendment adopts a new airworthiness directive (AD), applicable to Societe Nationale Industrielle Aerospatiale and Eurocopter France (Eurocopter France) Model AS 350B, BA, B1, B2, and D, and Model AS 355E, F, F1, F2, and N helicopters, without an autopilot installed, that requires a visual inspection to determine whether the cyclic pitch change control rod (rod) end fittings were safetied, and removal and replacement of the rod if the rod end fittings were not safetied. This amendment is prompted by a manufacturer's report that some of the rod end fittings had not been safetied at the factory. The actions specified by this AD are intended to prevent loss of tightening torque on the adjustment nuts of the rod, shifting of the neutral point of the cyclic stick, reduction in the amount of available movement of the cyclic stick in the roll axis, and subsequent reduction in the controllability of the helicopter.
2021-26-12: The FAA is adopting a new airworthiness directive (AD) for certain Stemme AG Model Stemme S 12 gliders. This AD was prompted by mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as the incorrect installation of an axle connecting the main landing gear (MLG) to the center steel frame. This AD requires inspecting the MLG installation and repairing if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
77-23-11: 77-23-11 CESSNA: Amendment 39-3084. Applies to Model 182 Series (Serial Nos. 18260797 thru 18265965) plus all other Cessna Model 182 Series incorporating ELT installations accomplished in accordance with Cessna Service Letter SE73-41 dated December 12, 1973, and references Kits AK150-97F or AK150-104A. Compliance: Required as indicated unless already accomplished. To preclude the possibility of an in-flight fire due to a loose ELT antenna coaxial cable connector making contact with the terminals of the battery relay and thereby shorting the electrical power to ground: A) Before the next flight, visually inspect the ELT antenna coaxial cable for security, and correct as necessary in accordance with Cessna Service Letter SE73-41, or later approved revisions and/or the applicable Cessna Service Manual, by accomplishing the following: 1. Assure that the ELT antenna RF connector is solidly attached at the ELT and is spring loaded in the detent position. If not correctly attached and/or spring loaded, either properly secure the connector, or, if damaged, replace the complete antenna coaxial cable assembly. 2. Assure that the plastic cable clamp, Cessna P/N S1155-3 or equivalent, is properly installed on the ELT coaxial cable and secured to structure above the ELT transmitter and mount (applicable to S/N's 33000 thru 18265175). If the coaxial cable is improperly secured, install a new plastic cable clamp, Cessna P/N S1155-3 or equivalent, if necessary, and properly secure. 3. Assure that the length of excess coaxial cable is coiled and secured to the mounted ELT transmitter by the required plastic sta-strap which has been properly installed. If not, properly secure with a new plastic sta-strap or equivalent. B) If operations are being conducted with the ELT removed, as provided by FAR 91.52(f), ascertain that the antenna cable is secured away from the battery relay. C) Airplanes may be flown, in accordance with FAR 91.52(e)(2), to a location where this AD can be accomplished, provided the operator disconnects the ELT antenna cable and removes it from the aircraft. D) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering & Manufacturing Branch, FAA, Central Region. This amendment becomes effective November 28, 1977, to all persons except those to whom it has already been made effective by air mail letter from the FAA dated October 28, 1977.
96-07-01: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model DC-10 series airplanes and KC-10A (military) airplanes, that requires visual inspections to detect failure of the attachments located in the banjo No. 4 fitting of the vertical stabilizer. This amendment also requires an eddy current inspection to detect cracking of the flanges and bolt holes of that fitting, and repair or replacement of attachments. This amendment is prompted by reports of failed attachments of the vertical stabilizer; the failures are attributed to fatigue. The actions specified by this AD are intended to prevent loss of the fail safe capability of the vertical stabilizer due to cracking of its attachments.
95-17-09 R1: This amendment revises Airworthiness Directive (AD) 95-17-09, which requires relocating the left-hand (LH) and right-hand (RH) essential bus current limiters (225 amp) to the battery bus (main bus tie) on certain Fairchild Aircraft SA226 and SA227 series airplanes. The Federal Aviation Administration (FAA) has determined that the applicability of the current AD should be changed to reflect a different serial number range and model designation of certain SA227 series airplanes. This action retains the essential bus current limiter relocations required by AD 95-17-09, and revises the Applicability section of that AD. The actions specified by this AD are intended to prevent failure of the LH and RH essential bus when engine failure results in a blown generator current limiter, which could result in loss of airplane electrical power.
50-26-01: 50-26-01 CURTISS-WRIGHT Applies to Model C-46 Series Aircraft equipped with landing gear side braces, Curtiss-Wright P/N's 20-310-1028 and 20-310-1029. Compliance required within 100 hours' time in service after the effective date of this amendment, unless already accomplished within the last 400 hours' time in service, and at intervals thereafter not to exceed 500 hours' time in service from the last inspection. Thoroughly inspect the landing gear side braces P/N 20-310-1028 and 20-310-1029 for cracks in the vicinity of the welds at either end of the struts, using magnetic or X-ray inspection. If cracks are found, the following will apply: 1. For one crack only in the weld proper less than 1/2 inch in length and 0.060-inch deep that does not penetrate into the tube member itself, stress relieve by grinding out the crack and polishing to remove all grinding marks. No rewelding required. 2. For more than one crack in the weld proper or cracks larger thanthose mentioned in item 1 that do not penetrate into the tube member itself, repair by grinding out the cracks and rewelding in a welding jig (using the oxyacetylene torch method) and reheat treat the tube assembly to 180,000 p.s.i. and Rockwell C-38. 3. If cracks are found in the tube member itself, the part should be replaced by a completely new assembly or repaired by replacing the tube and refabricating to the original specifications. 4. P/N S51E105 is considered a satisfactory replacement for P/N 20-310-1028. When P/N S51E105 is installed it should be inspected by magnetic particle or X-ray method of inspection prior to initial installation and at periods not to exceed 1,000 hours of operation thereafter. 5. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southern Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase. This supersedes AD 50-19-1. Revised December 28, 1964.
91-15-20: 91-15-20 BEECH: Amendment 39-7084. Docket No. 91-CE-08-AD. Applicability: Models 95-C55, 95-C55A, D55, D55A, E55, and E55A airplanes (serial number (S/N) TC350, and S/N TE-1 through TE-1201), and Models 58 and 58A airplanes (S/N TH-1 through TH-1610), certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent severe engine vibration and possible separation of the engine from the airplane caused by cracked engine mounts, accomplish the following: (a) Upon the accumulation of 600 hours time-in-service (TIS) or within 100 hours TIS after the effective date of this AD, whichever occurs later, inspect each engine mount in accordance with the instructions in Beech Service Bulletin (SB) No. 2362, Revision 1, dated February 1991. (1) If no cracks are found, repeat the inspection on each engine mount thereafter at intervals not to exceed 100 hours TIS. (2) If a crack is found on an engine mount, prior to further flight, remove the engine from the affected engine mount, remove the mount from the airplane, and magnetic particle inspect the engine mount to determine the length of the crack in accordance with the instructions in Beech SB No. 2362, Revision 1, dated February 1991. (i) If the length of the crack is .52 inches (true) or less, repair and reinforce the engine mount using Beech Kit 58-9007-1S in accordance with the instructions in Beech SB 2362, Revision 1, dated February 1991. (ii) If the length of the crack is greater than .52 inches (true), replace the cracked engine mount with a part number (P/N) 96-910010-67 engine mount in accordance with the instructions in Beech SB No. 2362, Revision 1, dated February 1991. (3) The repetitive inspections specified in paragraph (a)(1) of this AD may be terminated on an engine mount that has been repaired and reinforced with Beech Kit 58-9007- 1S or if a P/N 96-910010-67 engine mount has been installed. (b) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a location where the requirements of this AD can be accomplished. (c) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Mid-Continent Airport, Wichita, Kansas 67209. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (d) The inspections, modifications, and replacements required by this AD shall be done in accordance with Beech Service Bulletin No. 2362, Revision 1, dated February 1991. 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 the Beech Aircraft Corporation,P.O. Box 85, Wichita, Kansas 67201-0085. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-7084, AD 91-15-20) becomes effective on September 3, 1991.
2011-02-05: We are adopting a new airworthiness directive (AD) for the products listed above. This AD requires inspections for scribe lines in the fuselage skin at skin lap joints and butt joints, the skin at certain external approved repairs, the skin around external features such as antennas, and the skin at decals and fairings; and related investigative and corrective actions if necessary. This AD was prompted by reports of scribe lines found at skin lap joints and butt joints, around external repairs and antennas, and at locations where external decals had been cut. We are issuing this AD to detect and correct scribe lines, which can develop into fatigue cracks in the skin and cause rapid decompression of the airplane.
2022-02-02: The FAA is superseding Airworthiness Directive (AD) 2021-15- 51, which applied to Bell Textron Inc. (type certificate previously held by Bell Helicopter Textron Inc.) Model 204B, 205A, 205A-1, 205B, and 212 helicopters. AD 2021-15-51 required removing certain main rotor hub strap pins (pins) from service and prohibited installing them on any helicopter. Since the FAA issued AD 2021-15-51, it was determined that a defective pin could also be installed on Bell Textron Inc. Model 210 helicopters. This AD continues the required actions in AD 2021-15- 51 and expands the applicability to add Model 210 helicopters. The FAA is issuing this AD to address the unsafe condition on these products.
96-06-06: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-100, -200, -300, and SP series airplanes, that requires revising the Airplane Flight Manual (AFM) to prohibit the use of the autoland function. This amendment also requires installation of a diode and a marker on certain shelves and making wiring changes to the flight mode annunciator of the autopilot/flight director system, which terminates the requirements for the AFM limitation. This amendment is prompted by a report that the flightcrew was unaware of the configuration of the autoland system during landing. The actions specified by this AD are intended to ensure flightcrew awareness of the configuration of the autoland system in the event of a change from fail-operational to fail-passive mode.
2011-02-06: We are adopting a new airworthiness directive (AD) for the products listed above. This AD requires repetitive inspections for cracks in the fuselage skin and backup structure at the lower VHF antenna cutout at station 1197 + 99 between stringers 39 left and 39 right, and corrective actions if necessary. Certain repairs terminate certain inspection requirements. This AD was prompted by reports of cracking found in the section 46 fuselage lower skin around the periphery of the VHF antenna baseplate at station 1197 + 99. We are issuing this AD to detect and correct fatigue cracks in the fuselage skin and internal backup structure, which could result in rapid decompression of the airplane.