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86-05-09: 86-05-09 FAIRCHILD: Amendment 39-5252. Applies to all Model F27 and FH227 series airplanes certificated in any category. Compliance required as indicated. To detect cracks in the FH227 and F27 main landing gear drag strut assemblies due to the bores of some tubes having surface break-up generated during the drawing process, accomplish the following, unless previously accomplished: A. For airplanes which have accumulated 20,000 landings or more prior to the effective date of this AD, within 400 hours time-in-service after the effective date of this AD, using eddy current or ultrasonic inspection equipment and procedures, inspect the left and right upper main undercarriage drag stay for flaws, in accordance with the Accomplishment Instructions and sketches outlined in Dowty Rotol Service Bulletin 32-45N (for Model FH227 airplanes) or Service Bulletin 32-81C (for Model F27 airplanes), dated October 16, 1981. B. For airplanes which have accumulated more than 10,000landings but less than 20,000 landings prior to the effective date of this AD, within 800 hours time-in-service after the effective date of this AD, using eddy current or ultrasonic inspection equipment and procedures, inspect the left and right upper main undercarriage drag stay for flaws in accordance with the Accomplishment Instructions and sketches outlined in Dowty Rotol Service Bulletin 32-45N (for Model FH227 airplanes) or Service Bulletin 32-81C (for Model F27 airplanes), dated October 16, 1981. C. Within 2,000 hours time-in-service after the effective date of this AD, visually inspect the left and right upper main undercarriage stay in accordance with Dowty Rotol Service Bulletin 32-45N (for Model FH227 airplanes) or Service Bulletin 32-81C (for Model F27 airplanes), Item 2B(1)(2)(3)(4), dated October 16, 1981. D. If, as a result of the inspections referred to in paragraphs A. and B., above, flaws are detected, replace damaged part with a serviceable part prior tonext flight. E. Upon the request of an operator, an FAA Maintenance Inspector, subject to prior approval by the Manager, New York Aircraft Certification Office, FAA, New England Region, may adjust the inspection times specified in this AD to permit compliance at an established inspection period of that operator if the request contains substantiating data to justify the change for that operator. F. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the inspection requirements of this AD. G. Alternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. All persons affected by this proposal who have not already received these documents from the manufacturer may obtain copies upon request to Fairchild Industries, Inc., Fairchild Republic Division, Hagerstown, Maryland21740. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This amendment becomes effective April 13, 1986.
60-08-04: 60-08-04 LOCKHEED: Amdt. 132 Federal Register April 14, 1960. Applies to All Models 1049 Series and 1649A Aircraft. Compliance required as indicated. As a result of several cases of rudder torque tube failure (P/N 306525-2 and P/N 306525- 3), attributed to stress corrosion, the following inspection must be accomplished: Within the next 300 hours' time in service on all aircraft which have accumulated 5,000 hours' time in service, visually inspect for cracks, using a 10-power magnifying glass, the lower attachment portion of the upper rudder torque tube and the upper attachment portion of the lower rudder torque tube. If cracks or crack indications are found, reinspect the above area using dye penetrant or equivalent. (The cracks progress along the longitudinal direction from the edge of the tube to the first row of bolt holes and beyond. The cracks may also emanate from the second row of bolt holes.) Replace the torque tube prior to further flight if the tube is cracked beyond the first 3/8- inch bolt attachment hole or if the tube is cracked at the second row of 5/16-inch bolt attachment holes. If the tube is cracked from the edge to the first 3/8-inch bolt hole a repair may be used provided the repaired tube is reinspected in the above manner every 300 hours until replaced. The repair must be made prior to further flight and consists of the addition of two 1/4-inch blind lockbolts or Jobolts, located one on each side of the crack, spaced halfway between the existing 3/8-inch bolt holes and in line with these bolts. The maximum number of cracks permitted is two. If the crack progresses beyond the 3/8-inch bolt attachment hole, with the repair installed, replace the torque tube prior to further flight. If no cracks are found, reinspect visually every 1,300 hours' time in service. If torque tubes are replaced with new parts (P/N 306525-5 or P/N 306525-7), no further special inspections are required. (Lockheed Service Letter FS/237201 coversthis same subject.) Revised August 19, 1960.
2015-26-01: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Model AS332C1, AS332L1, AS332L2, EC225LP, AS-365N2, AS 365 N3, EC 155B, and EC155B1 helicopters with an energy absorbing seat (seat). This AD requires inspecting for the presence of labels that prohibit stowing anything under the seat. If a label is missing or not clearly visible to each occupant, we require installing a label. This AD was prompted by the discovery that required labels had not been systematically installed. The actions of this AD are intended to prevent objects from being stowed under the seat as these objects could reduce the energy-absorbing function of the seat, resulting in injury to the seat occupants during an accident.
64-28-01: 64-28-01 HARTZELL: Amdt. 39-14 Part 39 Federal Register December 16, 1964. Applies to Models HC-12X20-1, -2, -3, -5, and -7B Propellers Equipped with C-49-2B and C-49-2C Hub Spiders Having Serial Numbers 4220 through 5400 Installed on Downer (Republic) RC-3; Navion, Navion A; and Grumman G-44 Series Aircraft. Compliance required as indicated. As a result of loss of propeller blade due to failure of a hub spider, accomplish the following: (a) Visually inspect propeller hub spiders for cracks in accordance with Hartzell Service Bulletin No. 32 amended August 11, 1964, within 10 hours' time in service after the effective date of this AD, and at intervals thereafter not to exceed 25 hours' time in service from the last inspection until modification in accordance with Hartzell Service Bulletin No. 32 amended August 11, 1964, is accomplished. Replace cracked parts before further flight (b) Replace propeller hub spiders having accumulated less than 400 hours' time in service since new or last overhaul in accordance with Hartzell Service Bulletin No. 32 amended August 11, 1964, prior to the accumulation of 500 hours' total time in service since new or last overhaul. (c) Replace propeller hub spiders having accumulated 400 or more hours' time in service since new or last overhaul in accordance with Hartzell Service Bulletin No. 32 amended August 11, 1964, prior to the accumulation of 100 hours' time in service after the effective date of this AD. (Hartzell Service Bulletin No. 32 dated March 9, 1955, amended August 11, 1964, covers this subject.) This directive effective January 15, 1965 Revised April 8, 1965. Revised June 18, 1965.
83-23-05: 83-23-05 GENERAL DYNAMICS (Convair): Amendment 39-4773. Applies to Model 240 series and all military models eligible or to be made eligible for civil use under Type Certificate A- 793 and all such model airplanes converted to turbo propeller power, certificated in all categories, equipped with a forward right hand cabin "Main Entrance Door and Stairs - Mechanism Installation," P/N 240-3110695, regardless of whether or not the stairs have been replaced with a crew ladder. Compliance is required as indicated unless already accomplished. To prevent separation of the main entrance door in flight due to a malfunction of the door latching systems, accomplish the following: A. Within the next 30 days or 100 hours time in service, whichever occurs first after the effective date of this AD, perform a visual inspection and functional test of the forward right hand cabin main entrance door primary and secondary latch mechanisms, linkages, switch plungers, and switches to insurethat these elements function properly. Accomplish this using paragraphs (1) and (2), below, and the instructions as specified in paragraph 2, "Accomplished Instructions," of General Dynamics, Convair Division Service Bulletin 600 (240D) No. 53-6 dated May 25, 1983, or equivalent means approved by the Manager, Western Aircraft Certification Office, FAA, Northwest Mountain Region. (1) In the primary latching mechanism, the main entrance door latch link attaching arms must be thrown up, inboard, then down and over-center on both forward and aft door hooks. Each hook will move a plunger and lever to actuate separate warning switches to indicate that the primary latching mechanism is in the closed position. (2) In the secondary latching mechanism a forward moving linkage rod will guide separate pins through each door hook in the closed position, then up against stops in the aft and forward latch housings. This action simultaneously activates a door open warning switch to indicate that the secondary latching mechanism is in the closed position, and that the main entrance door is now securely closed and locked. NOTE 1. Failure of the primary latching mechanism to operate as designed may be caused by numerous conditions. One condition defeating the warning light exists when the main entrance door lock indicating pin or plunger is restricted in the depressed position. This may be due to rust, corrosion, a weak or broken pin spring, grease that has aged and hardened, or an accumulation of dirt and sand in the latch housing. Refer to appropriate Convair-Liner maintenance manual for other conditions. NOTE 2. Failure of the secondary latching mechanism to operate as designed may also be caused by numerous conditions. One condition exists when a linkage rod fails to move the pins completely through the door hooks due to a bent rod, loose adjustment nuts or actuating linkage out of adjustment. Refer to appropriate Convair-Liner maintenance manualfor other conditions. B. Defective units of the primary and secondary main entrance door latch locking mechanism discovered during accomplishment of paragraph A, above, must be repaired or replaced prior to further flight. C. Repeat the inspections and tests specified in paragraph A. of this AD at intervals not to exceed twelve calendar months since the last such inspection. D. Prior to issuance of a Certificate of Airworthiness for military aircraft being converted for civil certification, the airplane must be inspected and tested in accordance with paragraph A. of this AD. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes unpressurized to a base in order to comply with the inspections and tests required by paragraph A. of this AD. F. Alternative means of compliance providing an equivalent level of safety may be used when approved by the Manager, Western Aircraft Certification Office, FAA, Northwest MountainRegion, Hawthorne, California. All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to General Dynamics, P.O. Box 80877, San Diego, California 92138, ATTN: Mr. Larry Hayes, Manager, Product Support, Convair Division. These documents also may be examined at Regional Rules Docket, Office of Regional Counsel, FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or Western Aircraft Certification Office, FAA, Northwest Mountain Region, 15000 Aviation Blvd., Hawthorne, California. This amendment becomes effective November 28, 1983.
80-17-05: 80-17-05 BELL: Amendment 39-3875. Applies to Models 206A, 206B, 206A-1, 206B-1, and 206L helicopters equipped with tail rotor blades, P/N 206-010-750-005 and -007, certificated in all categories (Airworthiness Docket No. 80-ASW-18). Compliance required as indicated. To prevent possible failure of tail rotor blades, P/N 206-010-750-005 and -007, due to fatigue cracks, accomplish the following: a. Before the first flight of each day after the effective date of this AD, visually check for chordwise cracks in the tail rotor blade skin surfaces in the area between Blade Station 7.1 and 11.1 using a three-power or higher magnifying glass. (Blade Station 0 is the center of the tail rotor yoke.) b. Replace tail rotor blades having cracks before further flight. c. Blades with 450 or more hours' time in service (as calculated in paragraph (e) below) on the effective date of this AD must be removed from service within the next 50 hours' time in service. d. Blades with less than 450 hours' time in service (as calculated in paragraph (e) below) on the effective date of this AD must be removed from service prior to or on attaining 500 hours' time in service. e. For purposes of this AD, hours' time in service is calculated by the following formula: Time on 206A/B Series + Time on 206L = Calculated Time in Service 2.4 f. The check required by paragraph (a) of this AD may be performed by the pilot, providing the pilot's logbook has been endorsed by a properly rated mechanic stating that the pilot has been trained to conduct the daily check in accordance with this AD. NOTE: For the requirements regarding the listing of compliance with this AD in the aircraft maintenance record, see FAR 91.173. (Bell Helicopter Textron Operations Safety Notice No. 206-79-5/206L-79-2, dated December 4, 1979; Alert Service Bulletin Nos. 206-80-6 dated February 22, 1980, and 206L-80- 8, Rev. A, dated June 3, 1980; and Technical Bulletin Nos. 206-78-3 dated July 18, 1978, and 206L-79-38 dated September 28, 1979, pertain to this subject.) This amendment becomes effective September 10, 1980.
2015-24-02: We are adopting a new airworthiness directive (AD) for Viking Air Limited Model DHC-3 Airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as corrugation cracking found at various wing stations and on the main spar lower cap. We are issuing this AD to require actions to address the unsafe condition on these products.
84-06-03: 84-06-03 SHORT BROTHERS LTD: Amendment 39-4833. Applies to Model SD3-30 series airplanes as listed in Short Brothers Ltd. Service Bulletin SD3-24-18 dated November 1983, certificated in all categories. Compliance is required as indicated unless previously accomplished. To prevent short circuiting of the ECE141CC01A line contactor, accomplish the following: A. Within the next 60 days or 300 hours time in service, whichever occurs first after the effective date of this AD, modify the ECE141CC01A line contactor on the 1C and 2C panels in accordance with Short Brothers Ltd., Service Bulletin SD3-24-18 dated November 1983. B. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. This amendment becomes effective April 9, 1984.
53-06-01: 53-06-01 CESSNA: Applies to All Model T-50 Aircraft. \n\n\tCompliance required as soon as possible, but not later than July 1, 1953. \n\n\tTo guard against the possibility of fire or smoke due to inadequate electrical protection of the landing light motor circuits, install a 10-ampere fuse or circuit breaker, in the manner shown in Figure 2 for each motor circuit. \n\n\n\n\n\nFIGURE 2 \nAD 53-06-01
60-26-03: 60-26-03 BRANTLY: Amdt. 235 Part 507 Federal Register December 20, 1960. Applies to All Brantly Model B-2 Helicopters Prior To Serial No. 44 Not Previously Modified In Accordance With Brantly Service Bulletin No. 4. Compliance required within 25 hours' time in service after the effective date of this directive. (a) Remove the B2-108-34 drive shaft extension and the B2-108-35 coupling from between the transmission and the oil cooler fan drive pulley. Using Brantly aligning jig No. AT- 108-31 and following instructions in Brantly Service Bulletin No. 4, check alinement of the tail rotor drive shaft installation. Adjust bearing positions as necessary per instructions to obtain satisfactory alinement. (b) Using Brantly drill jig No. MT-339-16 and in accordance with instructions in Service Bulletin No. 4, install an additional flexible coupling on the aft end of the B2-108-34 drive shaft extension. Brantly parts B2-108-45 (one), B2-14-14 (two), plus associated standard attachments prescribed in Service Bulletin No. 4 are required for this modification. Orientation of the two flexible couplings on the shaft with respect to each other in accordance with Modification Step No. 15 of Service Bulletin No. 4 is imperative. This directive effective December 20, 1960.
2015-23-13: We are adopting a new airworthiness directive (AD) for all Airbus Model A318, A319, A320, and A321 series airplanes. This AD was prompted by a determination that, in specific flight conditions, the allowable load limits on the vertical tail plane could be reached and possibly exceeded. Exceeding allowable load could result in detachment of the vertical tail plane. This AD requires modification of the pin programming flight warning computer (FWC) to activate the stop rudder input warning (SRIW) logic; and an inspection to determine the part numbers of the FWC and the flight augmentation computer (FAC), and replacement of the FWC and FAC if necessary. We are issuing this AD to prevent detachment of the vertical tail plane and consequent loss of control of the airplane.
2015-23-11: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747-300, 747SR, and 747SP series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that certain fuselage skin lap joints are subject to widespread fatigue damage (WFD). This AD requires repetitive post-modification inspections for cracking of the skin or internal doubler along the edge fastener rows of the modification, and repair if necessary. We are issuing this AD to detect and correct fatigue cracking in certain fuselage skin lap joints, which could result in rapid depressurization of the airplane.
64-09-02: 64-09-02 BOEING: Amdt 717 Part 507 Federal Register April 21, 1964. Applies to All Models 707 and 720 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tThere have been instances of above normal control column forces and sluggish airplane response. In all cases, the airplanes involved had been exposed to heavy rain prior to takeoff or had taken off in heavy rain, with comparatively low freezing altitudes prevailing. This can result in ice formation in the elevator balance panel hinge or around the balance panel track grip. \n\n\tIn order to preclude this unsafe condition, accomplish the following: \n\n\tWithin 600 hours' time in service after the effective date of this AD, and thereafter at intervals not exceeding 1,200 hours' time in service, accomplish the following or equivalent approved by Engineering and Manufacturing Branch, FAA Western Region: \n\n\t(a) Mix MIL-G-25760 or MIL-G-7118 grease or equivalent approved by Engineering and Manufacturing Branch, FAA WesternRegion, with Methyl Ethyl Ketone or aliphatic naptha (Varsol), or equivalent approved by Engineering and Manufacturing Branch, FAA Western Region, to a consistency suitable for application with a squirt type oil can, and apply a good coverage of lubricant to the following areas: \n\n\t\t(1) Upper surface of the piano hinge for each balance panel, and \n\n\t\t(2) On either side of both balance panel track grips. \n\n\t(b) Upon request of the operator, and FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Boeing Service Bulletins Nos. 1865 and 1865A cover this same subject.) \n\n\tThis directive effective May 19, 1964. \n\n\tRevised September 5, 1964.
2015-23-01: We are adopting a new airworthiness directive (AD) for Sikorsky Aircraft Corporation (Sikorsky) Model 269A, 269A-1, 269B, 269C, 269C-1, 269D, and TH-55A helicopters. This AD requires repetitively inspecting and lubricating the tail rotor (T/R) driveshaft splined fittings. This AD was prompted by a report that the T/R driveshaft can disconnect due to deterioration of the splined coupling. The actions are intended to detect and prevent excessive wear of the splined coupling, which could lead to failure of the T/R driveshaft and subsequent loss of control of the helicopter.
82-03-02: 82-03-02 DETROIT DIESEL ALLISON: Amendment 39-4305. Applies to all Model 250-C28B engines installed in, but not limited to, Bell Model 206L-1 rotorcraft certificated in all categories. Compliance is required as indicated, unless previously accomplished. To preclude possible engine power loss resulting from intermittent or spurious activation of the N2 (Electronic) Overspeed Control System, accomplish the following before further flight: (1) Disarm the aircraft installed engine N2 Overspeed Circuit Breaker by pulling the N2 Overspeed Circuit Breaker and secure by wrapping with tape or placing a Ty-wrap (or equivalent plastic tie strap) around the breaker stem. (2) Install placard which states: "ENG OVSP CIRCUIT DEACTIVATED" in 1/4" or larger letters adjacent to N2 Overspeed Circuit Breaker. NOTE: The engine electronic overspeed test and the engine electronic overspeed trip system outlined in the rotorcraft Flight Manual will no longer function with the circuit breaker deactivated. (3) Install a P/N AN814-4DL or Alternate Plug and P/N AS3084-04 Packing in the outlet port of the N2 Overspeed Solenoid Valve. Tighten and safety with lockwire. NOTE: This action further assures deactivation of the N2 (Electronic) Overspeed Control System. Primary overspeed protection is still provided by the mechanical power turbine fuel governor. Upon request of the operator, an equivalent means of compliance with the requirements of this AD may be approved by the Chief, Chicago Aircraft Certification Office, Federal Aviation Administration, Central Region. NOTE: Detroit Diesel Allison Commercial Engine Alert Bulletin CEB-A-73-2020 refers to this subject. This amendment becomes effective February 4, 1982.
79-21-09 R1: 79-21-09 R1 CONSOLIDATED AERONAUTICS: Amendment 39-3589 is further amended by Amendment 39-4128. Applies to Colonial C-1, Colonial C-2, Lake LA-4, Lake LA-4A, Lake LA-4P, and Lake LA-4-200 airplanes with an anchor light or bilge pump installed with a power line which bypasses the battery relay. Compliance required within the next 50 hours time in service unless already accomplished. To preclude burning of power wires in case of an electrical short, install a Cole-Hersee 3031-D fuse holder and a Buss AG5A 5 ampere fuse, or FAA approved equivalents, within 4 inches of the battery relay, in the power lines for the anchor light and bilge pump. (Lake Aircraft Division, Consolidated Aeronautics Service Bulletin B-62 pertains to the installation required by this AD.) Upon request, an equivalent method of compliance with the requirements of this AD may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region. Amendment 39-3589 became effective upon publication in the FEDERAL REGISTER. This Amendment, 39-4128, becomes effective June 8, 1981.
2015-06-02 R2: We are adopting a new airworthiness directive (AD) for GA 8 Airvan (Pty) Ltd Model GA8-TC320 airplanes. This AD revises AD 2015-06- 02 R1, which required inspection to detect and correct the omission of steel washers at each isolator mount location. This AD retains the actions of AD 2014-06-02 R1 but corrects the AD number in the parenthetical of the compliance time in paragraph (f)(1) of the AD. This AD was prompted by reports of missing required engine mount fire seal washers, which could reduce the engine retention capability in the event of a fire. We are issuing this AD to require actions to address the unsafe condition on these products.
63-10-01: 63-10-01 FAIRCHILD: Amdt. 563 Part 507 Federal Register May 7, 1963. Applies to all C-82A and C-82A Jet Packet Aircraft equipped with Pratt & Whitney R-2800 Military or Civil C, CA, or CB Series engines and Hamilton Standard 33E60/6491-0 or 33E60/6801-0 Propellers. Compliance required as indicated. Premature wear in the bushings in the original rear (heavyweight crankshaft vibration dampers in R-2800 "C" Series engines reduces crankshaft damping characteristics such as to cause propeller blade vibration stresses above the safe limit for continuous service. In addition, flight operation at grass weights in excess of 36,500 pounds with unmodified Hamilton Standard 6491-0 propeller blades installed, imposes blade vibration stresses above the safe limit for continuous service. To prevent propeller blade fatigue failures from these causes, the following modifications and engine operating limitations are required: (a) On airplanes equipped withany of the engines specified herein, prior to 100 hours' time in service after the effective date of this AD, accomplish the following: (1) Install the following placard on the instrument panel in full view of the pilot: "Avoid operation between 1,550 and 1,650 r.p.m." (2) Apply red arc markings on the tachometers over this restricted speed range. (b) On airplanes equipped with R-2800-C Series engines having the original C series type crankshaft with heavyweight front and rear vibration dampers accomplish one of the following within the next 300 hours' time in service after June 6, 1963, for engines which had 350 or more hours' time in service since new or overhaul on June 6, 1963, or prior to the accumulation of 650 hours' time in service since new or overhaul for engines which had less than 350 hours' time in service since new or overhaul on June 6, 1963: (1) Replace the crankshaft assembly with crankshaft assembly, P/N 326054, in accordance with Pratt & Whitney Aircraft Service Bulletin No. 1687 dated April 22, 1958. Identify engines so modified by adding suffix letter " D" to the engine model designation on the nameplate. (2) Replace present engines with engines incorporating the CA or CB type crankshaft. These engines are identified by the suffix letter H or D on the engine model designation on the nameplate. (c) On 33E60/6491-0 propellers, prior to 300 propeller hours' time in service after the effective date of this AD, accomplish one of the following: (1) Replace the 6491-0 propeller blades with 6801-0 blades. (2) Replace present 6491-0 blades with 6491-0 blades which have had the blade shanks cold rolled and the inner portion of the blades shotpeened by Hamilton Standard or by a method approved by the FAA. Note. - Hamilton Standard has advised that only blades which have had no time in service and which are in good condition otherwise, will be accepted for this rework. (d) On 33E60/6491-0 or 33E60/6801-0 propellers, prior to 300 propeller hours' time in service after the effective date of this AD, the barrel bolt lugs on the 33E60 hubs shall be reworked and shotpeened and the inside of the barrel arms shotpeened in accordance with Detailed Instruction No. 6 of Hamilton Standard Bulletin No. 628 dated January 30, 1962. (e) The propeller modifications specified in (c) and (d) are not required provided the operating limitations of this airplane are revised to limit the operating gross weight to 36,500 pounds and to limit the maximum takeoff power to 2,100 hp. This directive effective June 6, 1963. Revised January 27, 1965.
79-24-07: 79-24-07 DETROIT DIESEL ALLISON: Amendment 39-3624. Applies to Aeroproducts Models A6441FN-606 and -606A propeller blades. Compliance is required as indicated. To preclude propeller blade failure due to fatigue, complete an interim inspection of the blades in accordance with Detroit Diesel Allison Commercial Service Letter 501-D13/-606 CSL-240 dated March 1, 1976 thru Revision 3 dated July 26, 1977 as follows: NOTE: The fluorescent magnetic particle method must be used. a. If blades have more than 4000 hours since the last interim inspection or overhaul, complete an inspection within the next 150 hours time in service. b. If blades have 3150 to 4000 hours since the last interim inspection or overhaul, complete an inspection within the next 450 hours, or prior to exceeding 4150 hours, whichever occurs first. c. If blades have 2550 to 3149 hours since the last interim inspection or overhaul, complete an inspection within 600 hours, or prior to exceeding 3600 hours, whichever occurs first. d. If blades have less than 2550 hours since the last interim inspection or overhaul, complete an inspection prior to exceeding 3150 hours. e. If blade cracks are found during these inspections and/or if spalled areas are found exceeding limits as per CSL-240, entire propeller hub must also be inspected per CSL-240. f. Report the results of all inspections to Chief, Engineering and Manufacturing Branch, AGL-210, FAA, 2300 East Devon Avenue, Des Plaines, Illinois 60018. (Reporting approved by Office of Management and Budget under OMB No. 04-RO-174.) g. All subsequent interim inspections shall be carried out at intervals not to exceed 3150 hours time in service. h. Alternate inspections or other actions which provide equivalent levels of safety may be used when approved by the Chief, Engineering and Manufacturing Branch, FAA, Great Lakes Region. This amendment becomes effective December 5, 1979.
2000-23-32: This amendment adopts a new airworthiness directive (AD) that applies to certain DG Flugzeugbau GmbH (DG Flugzeugbau) Models DG-500 Elan Series, DG-500M, and DG-500MB sailplanes. This AD requires you to visually inspect the elevator control system for proper movement, obtain and incorporate a repair scheme if improper movement is found, and modify and install resin thickened cottonflock reinforcements to the elevator control system as a way to increase the stiffness of the elevator control support stand. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for the Federal Republic of Germany. The actions specified by this AD are intended to detect and correct improper movement in the elevator control system and to increase the stiffness of the elevator control support stand. Without accomplishing these actions, the pilot's capability to use full elevator control deflection could be limited, which could require increased force in moving the elevator control with a consequent potentially uncontrolled flight condition.
2015-21-04: We are adopting a new airworthiness directive (AD) for certain Pratt & Whitney (PW) PW4164, PW4168, PW4168A, PW4164-1D, PW4168-1D, PW4168A-1D, and PW4170 turbofan engines. This AD was prompted by crack finds in the 6th stage low-pressure turbine (LPT) disk. This AD requires [[Page 64313]] removal of the affected 6th stage LPT disks. We are issuing this AD to prevent failure of the 6th stage LPT disk, which could lead to an uncontained disk release, damage to the engine, and damage to the airplane.
81-08-51: 81-08-51 DEHAVILLAND AIRCRAFT OF CANADA: Amendment 39-4263. Applies to DeHavilland DHC-7 series airplanes, serial numbers 3 through 42 inclusive. 1. To prevent failure of the main landing gear to extend, unless already accomplished, accomplish the following within the next ten hours time-in-service: A. Using a suitable sharp pointed tool, attempt to rotate the thickest washer under the head of each bolt attaching the main landing gear actuator brackets P/N's 75710543-003 and 75710542-003 to the wing skin on each inboard nacelle. B. If rotation is observed, replace existing barrel nuts DSC 100-8, retainers and washers with Esna barrel nuts P/N 2452-080 and washers P/N's MS20002-C8 and MS20002- 8 in accordance with DeHavilland Engineering Order 72136 prior to further flight. Torque bolts to 850-900 inch-lbs. and wirelock. 2. To prevent possible malfunction of the flaps, unless already accomplished, accomplish the following within the next ten hours time-in-service: A. Using a 9/16 inch, 12 point socket wrench with standard length handle, attempt to tighten all bolts attaching the number 4 flap track bracket to each wing. Ref. PSM 1-7- 4, Chapter 57-52-00, Fig. 15, Item 160, for bolt location. B. If movement of the bolt head is observed, replace existing barrel nuts DSC 100-8 with Esna barrel nuts P/N 2452-080 using additional MS20002-C8 washers, as required, to ensure correct bolt engagement in accordance with Engineering Order 72139 prior to further flight. Torque bolts to 900-1000 inch-lbs. 3. For aircraft Serial Numbers 28 through 42 inclusive, repeat inspections specified in paragraphs 1 and 2 above daily until replacement of existing barrel nuts DSC 100-8 is accomplished, in accordance with subparagraphs 1B and 2B, above. 4. For aircraft Serial Numbers 28 through 42 inclusive, replacement of existing barrel nuts DSC 100-8 is required within 50 hours time-in-service following receipt of this AD. 5. Alternatemeans of compliance or other actions which provide an equivalent level of safety may be used when approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Mountain Region. 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 the addresses listed above. These documents may also be examined at FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective November 25, 1981 and was effective earlier to those recipients of Telegraphic AD T-81-08-51 dated April 10, 1981.
86-25-03: 86-25-03 THE DEHAVILLAND AIRCRAFT COMPANY OF CANADA, A DIVISION OF BOEING OF CANADA, LTD.: Amendment 39-5478. Applies to all Model DHC-7 and DHC-8 series airplanes, certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent the elevator trim-tab from becoming inoperable due to accumulated water freezing in the trim-tab jack-screw housing, accomplish the following, unless already accomplished. A. For Model DHC-7 airplanes, within three months after the effective date of this AD, incorporate de Havilland Modification Number 7/2489 to provide a drain hole in the elevator trim-tab jack-screw housing in accordance with the "Accomplishment Instructions" contained in de Havilland Service Bulletin No. 7-27-76, dated January 17, 1986. B. For Model DHC-8 airplanes, within three months after the effective date of this AD, incorporate de Havilland Modification Number 8/0415 to provide a drain hole in the elevator trim-tab jack-screw housing in accordance with the "Accomplishment Instructions" contained in de Havilland Service Bulletin No. 8-27-15, dated January 17, 1986. C. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain upon request to de Havilland Aircraft Company of Canada, a Division of Boeing of Canada Ltd., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This amendment becomes effective January 12, 1987.
84-08-04: 84-08-04 BEECH: Amendment 39-4931. Applies to Model A36TC (S/Ns EA-1 thru EA-241 ((modified in accordance with Beech Service Instructions No. 1191)), and EA-243 thru EA-272), and Model B36TC (S/Ns EA-242, EA-273 thru EA-382, EA-384 thru EA-396, EA-398 thru EA- 404, EA-406 thru EA-409, EA-411 thru EA-415, and EA-418) airplanes certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent loss of engine oil due to blockage of the turbocharger oil return line check valve, accomplish the following: (a) Prior to further flight after the effective date of this AD: (1) Inspect the turbocharger oil return line check valve per Beechcraft Service Bulletin No. 2027 to determine if a P/N 101-389011-55 check valve is installed. This valve may be identified by the blue color and the P/N on the valve outlet end fitting. (i) If a Beech P/N 101-389011-55 check valve is not installed, the airplane may be returned to service without further action. (ii) If a Beech P/N 101-389011-55 check valve is installed, prior to return to service and within each additional five hours time-in-service until replaced, inspect and clean this valve in accordance with Beech Service Bulletin No. 2027. (b) Within the next 25 hours time-in-service after the effective date of this AD on airplanes on which a Beech P/N 101-389011-55 check valve is installed, modify this part to, or replace it with, a Beech P/N 101-389011-69 check valve in accordance with Beech Service Bulletin No. 2027. (c) The repetitive inspections in paragraph (a)(ii) are not required when a Beech P/N 101-389011-69 check valve is installed. (d) An equivalent method of compliance with this AD may be used if approved by the Manager, Wichita Aircraft Certification Office, Room 100, 1801 Airport Road, Mid- Continent Airport, Wichita, Kansas 67209, telephone (316) 946-4400. This amendment becomes effective on October 15, 1984, to all personsexcept those to whom it has already been made effective by priority letter from the FAA dated April 25, 1984.
2015-15-06: We are superseding Airworthiness Directive (AD) 2003-13-01 for certain The Boeing Company Model 767 airplanes. AD 2003-13-01 required an inspection to detect cracks and fractures of the outboard hinge fitting assemblies on the trailing edge of the inboard main flap, and follow-on and corrective actions if necessary. For certain airplanes, AD 2003-13-01 required an inspection to determine if a tool runout option has been performed in the area. This new AD reduces certain compliance times, adds airplanes to the applicability, and provides optional terminating action for certain inspections. This AD was prompted by reports of hinge assembly fractures found before certain required compliance times in AD 2003-13-01. We are issuing this AD to prevent the inboard aft flap from separating from the wing and potentially striking the airplane, which could result in damage to the surrounding structure and potential personal injury.