Results
47-07-03: 47-07-03 FAIRCHILD: (Was Service Note 4 of AD-724-2.) Applies to M-62 Series Aircraft. Prior to original certification and at each periodic inspection thereafter, and as otherwise noted, make the following inspections: (1) Inspect the wing center section bottom surface for cracks. This inspection should be made after each severe landing. Cracks extending into the spar flange area indicate cracked spar flanges and should be investigated very thoroughly. (2) Inspect the butt ends of the spars to assure that the butt plates are in place and properly attached. (3) Inspect the strap hinge fittings for looseness. Clearance between the spar webs and hinge plates is not critical as long as the plates are bolted tight to the bushings if the bushings protrude. If bushings are loose, replace. (4) Inspect the plywood spar webs for checks or cracks. This inspection should always be made after any damage to the landing gear. Cracks other than those parallel to the face grain generally indicate serious damage to the spar web. (5) Inspect the trailing edge of the wing center section and outer panel over flap area for deterioration due to accumulated moisture. (6) Inspect the forward face of front spar and belly skin at engine cutout in wing center section for oil soaking and skin separation. (These inspections and methods of repair are covered in greater detail in Fairchild Service Bulletin 47-62-1 dated January 24, 1947. Owners may obtain copies from Fairchild Personal Planes Division of Fairchild Engine and Airplane Corp., Hagerstown, Md.)
47-21-02: 47-21-02 FAIRCHILD: (Was Mandatory Note 8 of AD-707-2 and Mandatory Note 5 of AD-706-1.) Applies to 24R-46 and -46S, and 24W-46 and -46S Aircraft. Compliance required prior to July 1, 1947. Replace the landing light fuse with one of 20-ampere capacity. (Fairchild Service Bulletin 47-24-1 dated January 8, 1947, covers this same subject.)
47-43-05: 47-43-05 CESSNA: (Was Service Note 3 of AD-768-5.) Applies to 120 and 140 Aircraft Serial Numbers 8001 to 13780, Inclusive. Inspection required upon each 100 hours of operation until reinforcing channels are installed at all hinge fittings. Inspect for fatigue cracks in the elevator spar web at the hinges. These cracks start either at the rivets or at an edge of the fitting and progress around the fitting until the elevator breaks loose from the hinge fitting. If cracks less than 1/2 inch in length are found a reinforcing channel, Cessna P/N 0434151 at the outboard hinge or 0434152 at the inboard hinge, should be installed on the aft side of the spar with the flanges riveted between the spar flanges and the skin with two AN 455AD3 rivets per flange. Four AN 442AD4 rivets should be used to attach each fitting to the spar web and reinforcing channel. If any cracks are longer than 1/2 inch the spar should be replaced and the reinforcing channels added. (Cessna Service Letter No. 46 dated July 31, 1947, covers this same subject.)
47-51-05: 47-51-05 CURTISS-WRIGHT Applies to Model C-46 Series Aircraft Equipped with Aileron Horn Assembly, Curtiss-Wright P/N 20-050-5715. Compliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished. The aileron horns part 20-050-5715 have failed due to cracking of the horn between the attaching bolt holes and the outer edge. Inspection should be made to determine if this part has been replaced by P/N SK-10213. If not, part 20-050-5715 which is a casting should be replaced by a machined horn manufactured from 24ST material in accordance with Curtiss-Wright Drawing No. SK-10213. (Army Technical Order 01-25L-102 also covers this same subject.) Revised December 28, 1964.
50-25-01: 50-25-01\tSTINSON: Applies to All Model 108 Series Aircraft.\n\tCompliance required not later than September 1, 1950.\n\tReports have been received of fuel seepage into the space between the inner cabin trim and the outer fabric covering of the fuselage. This results in soaking of insulating material in the cabin wall. The source of the fuel can be spillage during filling of tanks, thermal expansion of fuel in full tanks, or tank leakage. This fuel runs to the under surface of the wing, adhering to the lower curved surface of the trailing edge of the wing at the flap well, thence inboard to the fuselage and across the rear window. Since the window seal is often not perfectly tight the fuel may then enter the cabin wall.\n\tTo preclude the fire hazard of fuel soaked insulation within the cabin wall due to these causes, a drip strip similar to that shown in Figure 1 should be installed on the underside of each wing. This drip strip will prevent fuel from flowing from the wing to the fuselage.\n\t(Piper Service Bulletin No. 115, dated March 31, 1950, covers this same subject.)
94-23-05: This amendment adopts a new airworthiness directive (AD) that is applicable to AlliedSignal Inc., (formerly Garrett Engine Division) TFE731-3A-200G and -3AR-200G model turbofan engines. This action requires removing from service certain low pressure turbine (LPT) disks, imposing an hourly life limit on the first stage and second stage LPT disks, performing a dimensional inspection of second stage LPT disks at repetitive intervals, and incorporating honeycomb material in the second stage LPT nozzle air seal. This amendment is prompted by reports of LPT disk web separations. The actions specified in this AD are intended to prevent LPT disk web separations, which can result in an uncontained engine failure and damage to the aircraft.
47-42-17: 47-42-17\tDOUGLAS: (Was Service Note 3 of AD-781-1.) Applies to the Following DC-6 Aircraft Serial Numbers: Douglas 43061; AAL 42854 to 42865, Inclusive; 42879 to 42880, Inclusive; 42882 to 42896, Inclusive; and 43035 to 43044, Inclusive; UAL 42866 to 42875, Inclusive; and 43000 to 43024, Inclusive; Panagra 42876 to 42878, Inclusive; National 43055 to 43058, Inclusive; Sabena 43062 to 43064, Inclusive; Braniff 43105, 43106; KLM 43111 to 43112, Inclusive; and AAF 42881. \n\n\tInspection required at each 300 hours (or at each 150 hours for non-air-carrier operations).\n \n\tInspect the center spar web between Stations 167 and 184 for cracks in the web along the lower row of rivets which attach the spar web to the leg of the upper spar cap. For aircraft with the 10-tank fuel system this inspection can be properly made only by removing the fuel tank inspection opening near the affected area, since the spar web attaches to the forward side of the spar cap leg and small cracks in the web cannot be detected without close examination. If cracks are found during this inspection or, if between the inspections, leaks occur which are caused by cracks in the center spar web between Stations 167 and 184, the spar web must be reinforced by installing a doubler in accordance with Douglas Drawing 5356664.\n\n\tWhen the spar web reinforcement has been incorporated the special inspection required by this Note may be eliminated. All DC-6 aircraft not mentioned above will be reinforced at the factory.\n \n\t(Douglas Service Bulletin DC-6 No. 29, "Rework Center Spar Web, Stations 167-184, Integral Wing Fuel Tank DC-6 Airplane", covers the same reinforcement as described on Drawing 5356664.)
95-12-21: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A340-211 and 311 series airplanes. This action requires the installation of doublers on certain stringers located in the center fuselage. This amendment is prompted by the results of the manufacturer's full-scale fatigue test which indicate that fatigue cracking can occur at these stringer locations. The actions specified in this AD are intended to prevent reduced structural integrity of the fuselage due to the problems associated with fatigue cracks in the subject stringers.
71-09-06: 71-09-06 MARTIN AIRCRAFT: Amdt. 39-1200. Applies to Models 202, 202A and 404 airplanes certificated in all categories. Compliance required as indicated. To detect cracks or corrosion in the main landing gear aluminum gland nut, P/N 202SD81548, (Menasco Drawing No. 511034), accomplish the following: (a) Within the next 25 hours in service, after the effective date of this AD, unless already accomplished within the last 475 hours in service, and thereafter at intervals not to exceed 500 hours in service from the last inspection, comply with (b). (b) Unscrew the left and right main landing gear aluminum gland nut, P/N 202SD81548, so that all threads are visible. Inspect the nut for evidence of cracks or corrosion, using dye penetrant in conjunction with at least a 10-power glass or an equivalent inspection. Cracked or corroded parts must be replaced with an unused part or an equivalent before further flight, except that the airplane may be flown in accordance withFAR 21.197 to a base where the repair can be performed. (c) The repetitive inspection specified in (a) may be discontinued when the aluminum gland nut is replaced by a part made of steel, as per ECD 32590 to Menasco Drawing No. 511034, change L, dated 8 March 1961, or replaced by a steel gland nut in accordance with Eastern Air Lines Drawing No. 204-8125, or with an equivalent part. (d) Upon submission of substantiating data by an owner or operator through an FAA Maintenance Inspector, the Chief, Engineering & Manufacturing Branch, FAA, Eastern Region, may adjust the initial inspection and the repetitive inspection interval specified in this AD. Equivalent parts and inspections must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective May 4, 1971.
49-09-02: 49-09-02 NAVION: Applies to Airplanes Equipped With Adel Electric Booster Pumps. The Following Adel Pumps Do Not Require Modification in Accordance With This Directive: (1) Pumps With Serial Numbers above 2451, (2) Pumps Having a Red Painted Band on the Pump Housing, (3) Pumps Having the Letters "G" or "S" Suffixed to the Pump Serial Number. To be accomplished as soon as possible but not later than April 1, 1949. Several instances of air leakage into the fuel system have been reported on Navions equipped with Adel electric booster pumps. It has been determined that air can enter the fuel system through the 0.062-inch diameter hole in the plate at the rear of the Adel pump inlet chamber. This hole was originally provided to prevent overboard drainage of fuel through a faulty pump shaft seal while the pump was running. All of the pumps affected require blocking of the hole at the rear of the pump inlet chamber. This is accomplished in the field by means of an Adel manufactured wire plug which is inserted into the hole through the pump inlet port. Pumps with Serial Numbers below 1600 which do not have the letter "R" suffixed to the serial number also require replacement of the pumpshaft running seal spring. Adel Accessories Service Bulletin No. 147-49 describes these changes. The required plug and spring and copies of the Adel Bulletin and Ryan's covering Service Letter No. 57, may be obtained from the Ryan Aeronautical Co., San Diego, California.
47-49-04: 47-49-04 LOCKHEED: (Was Mandatory Note 37 of AD-763-3.) Applies to Model 49 Serials 2068 to 2088, Inclusive. Compliance required within next 50 hours of operation unless the 1 3/16-inch headless drive pin has been installed. Inspect attachments of rudder pedal lever arms to the clip assembly in the 284587 rudder pedal slot cover guide assemblies to determine whether it is possible for the flat head pin to cause jamming of the system. If any possibility of jamming exists, the flat head pin should be replaced with a headless drive pin 1 3/16 inches long. (LAC Service Bulletin 49/SB-260 covers this same subject.)
47-06-01: 47-06-01 GLOBE: (Was Service Note 1 of AD 766-5.) Applies to Models GC-1A and GC-1B Aircraft. To be accomplished prior to April 1, 1947, and upon each 100 hours operation thereafter. Inspect main landing gear retraction system to determine that adjustments are as follows: (1) When the side brace is against the down stop the middle joint should be 1/8 inch to 1/4 inch above dead center (3/16 inch to 5/16 inch if measured from edges of links in accordance with Globe Customer Service Maintenance Bulletin No. 7). (2) When the side brace is against the down stop and the down lock plunger is fully extended, covering at least 1/2 of the adjustment screw head, the clearance between the plunger and the screw head should be from 0.001 inch to 0.005 inch. (3) When the side brace is against the down stop the limit switch plunger should be depressed approximately 1/32 inch beyond the cutoff point. (4) The turnbuckle in the emergency extension cable should be adjusted so that on manual extension of the gear both down locks operate before the handcrank has been wound to the full down position. After it has been determined that the turnbuckle adjustment is satisfactory in this respect it should be determined also that with the handcrank wound to the full up position the cable length is sufficient to permit the up limit switches to cut off. (Globe Customer Service Maintenance Bulletin No. 7 covers this same subject.)
47-42-09: 47-42-09 DOUGLAS: (Was Mandatory Note 6 of AD-781-1.) Applies to DC-6 Serial Numbers 42854, 42855, 42858 Through 42865, 42869 Through 42880, 42882 Through 42891, 43000 Through 43009, 43055 and 43056. \n\nTo be accomplished not later than next No. 3 inspection (or not later than next 150 hours for noncarrier operations). \n\nCertain cases have been found where the aileron hinge plates at wing Stations 421, 485, 585, and 675 were fabricated from overgage stock resulting in an interference fit between the plate and the clevis fitting. The following hinge plates and fittings should be inspected to determine whether or not they conform with the tolerances listed below. If plates are found which exceed the widths noted below, they should be reworked with emery cloth to specified limits and touched up with zinc chromate primer. Fittings which have been installed over an oversize plate should be anodized and carefully inspected before being reinstalled. \n\n\n\nHinge\n\n\n\n\nStation\nNo.\nPlate\nThickness\nFitting\nWidth\n421\n1\n3320118\n0.249 \n4334619\n0.249\n\n\n\n0.237\n\n0.254\n485\n2\n3323460\n0.238*\n4345756\n0.334\n\n\n\n0.243\n\n0.350\n585\n3\n3323461\n0.311\n4345755\n0.311\n\n\n\n0.297\n\n0.316\n575\n4\n3323462\n0.249\n4345754\n0.249\n\n\n\n0.237\n\n0.254\n\n*Thickness of 0.093 angle not included.\nTotal thickness should not exceed 0.334. \n\n(Douglas Service Letter A-214-529.004/RLT dated July 21, 1947, and attached sketches cover this same subject.)
47-33-02: 47-33-02\tDOUGLAS: (Was Mandatory Note 14 of AD-618-3, Supplement 2; and Mandatory Note 15 of AD-669-3, Supplement 2.) Applies to All DC3 Series Aircraft. \n\nTo be accomplished not later than the first engine change after September 1, 1947, but in any event not later than December 1, 1947. \n\nIn order to preclude cowl flap hydraulic line failures and possible subsequent fires, replace grommets and lines forward of the firewall with AN 833-4 elbows and AN 924-4 nuts, or equivalent, and new fire resistant flexible hose assemblies of proper length.
47-27-04: 47-27-04 DOUGLAS: (Was Mandatory Note 18 of AD-762-7.) Applies to All C-54 and DC-4 Series Airplanes Having Exhaust Collector Rings Made Up of Top Segments P/N 5174842-56 L.H. and 5174529-56 R.H. \n\n\tTo be accomplished not later than the first engine change subsequent to July 15, 1947, but in any event not later than October 15, 1947. \n\n\tSeveral reports have been received of cracking failure of the collector ring "Y" outlet assembly due to breathing of the exhaust stack. This induces failure which creates a fire hazard. This type of exhaust collector "Y" is not reinforced with a flange and is shown on page 4, Douglas Service Bulletin No. DC-4 No. 31. To correct this condition weld a scalloped stiffener flange on the exhaust collector aft of the "Y" outlet assembly. \n\n\t(Douglas Service Bulletin No. DC-4 No. 68 covers this same subject.) \n\n\tUntil this repair is accomplished, inspection for cracks should be made immediately and at periods not to exceed 50 hours of operation.
47-07-04: 47-07-04 FAIRCHILD: (Was Mandatory Note 6 of AD-724-2.) Applies to M-62 Series Aircraft. Compliance required at next periodic inspection. In order to eliminate the possibility of the control sticks becoming disengaged, both front and rear control sticks should be reworked by drilling through stick and socket, and installing and safetying an AN393-51 clevis pin. Washer AN 960-10L and cotter AN 380-2-2. (Fairchild Service Bulletin 44-62-2 dated October 31, 1944, covers this same subject.)
47-51-16: 47-51-16 DOUGLAS: Applies to DC-6 Serial Numbers 42854 Through 42880, Inclusive; 42882 Through 42891, Inclusive; and 43000 Through 43009, Inclusive; 43055; and 43056. \n\nTo be accomplished not later than August 1, 1948. \n\nIn order to increase the strength of the flap support assembly at wing Station 378 and to replace the temporary rework outlined in Douglas Co. Service Letter of May 12, 1947, which was necessitated by failure of the flap hinge support assembly on an airplane in flight, the following must be accomplished: \n\n(a)\tRemove the two Shafer bearings, P/N AB-4A from flap link assembly, P/N 4325008 and press in new Shafer bearings, P/N AB-5A and stake in place. \n\n(b)\tRemove outboard flap link support assembly, P/N 5107188, and line ream (0.312-0.313) diameter through to permit use of 5/16-inch bolt for attachment of upper end of link assembly, P/N 4325008. Assembly, P/N 5107188 becomes P/N 5107188-500 after rework. \n\n(c)\tPress out old bushing, P/N 1338719, two places in flap hinge bracket assembly, P/N 3320998, and press in new bushings, P/N 1338719-500. \n\n(d)\tAfter replacing P/N 5107188-500 replace P/N 4325008, using bolts, P/N 2356375-22; washers, P/N 124682-5-12-6 and P/N AN 960-516; nut, P/N AN 310-5 and cotter pin, P/N AN 380-2-3. \n\n(Douglas Service Bulletin DC-6 No. 66 covers this same rework.)
93-19-01: 93-19-01 Dowty Rotol (now Dowty Aerospace Gloucester Ltd.): Amendment 39-8694. Docket 93-ANE-10. Supersedes AD 87-21-51, Amendment 39-5929. Applicability: Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Propeller Models (c)R354/4-123-F/13, (c)R354/4-123-F/20, and (c)R375/4-123-F/21 installed on SAAB-SF340A, and SAAB-SF340B series aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent possible loss of the propeller, accomplish the following: (a) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/13 propellers, perform Dowty Aerospace Gloucester Ltd. Modification (C) VP3336 by installing interface shim Part Number (P/N) 660712669 in accordance with Dowty Aerospace Gloucester Ltd. Service Bulletin (SB) No. SF340-61-57, dated February 15, 1991, within 500 hours time in service (TIS) since the last torque check and inspections accomplished in accordance with AD 87-21-51. (b) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model R354/4- 123-F/13 propellers, perform a torque check of the propeller retention bolts for low torque and a magnetic particle inspection of the propeller retention bolts for cracks; perform dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. If propeller retention bolts or hubs are found to have cracks, remove from service prior to further flight, and replace with serviceable propeller retention bolts and hubs, within 500 hours TIS since the last torque check and cracking inspections accomplished in accordance with AD 87-21-51. These actions must be accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991. (c) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/20 and (c)R375/4-123-F/21 propellers, perform Dowty Aerospace Gloucester Ltd. Modification (C) VP3336 by installing interface shim P/N 660712669 in accordancewith Dowty Aerospace Gloucester Ltd. SB No. SF340-61-57, dated February 15, 1991, within 100 hours TIS after the effective date of this AD; or 500 hours TIS since the last torque check and inspections accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991, or Dowty Aerospace Gloucester Ltd. SB No. SF340-61-21, Revision 4, dated October 1, 1987; or 500 hours TIS since new, whichever occurs latest. (d) For Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) Model (c)R354/4- 123-F/20 and (c)R375/4-123-F/21 propellers, perform a torque check of the propeller retention bolts for low torque; a magnetic particle inspection of the propeller retention bolts for cracks; and dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. Remove from service prior to further flight cracked propeller retention bolts and hubs, and replace with serviceable propeller retention bolts or hubs, within 100 hours TIS after the effective date of this AD; or 500 hours TIS since the last torque check and inspections accomplished in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991, or Dowty Rotol (now Dowty Aerospace Gloucester Ltd.) SB No. SF340-61-21, Revision 4, dated October 1, 1987; or 500 hours TIS since new, whichever occurs latest. These actions must be accomplished in accordance with Dowty Aerospace Gloucester Ltd, SB No. SF340-61-58, Revision 1, dated July 18, 1991. (e) Thereafter, for propeller models identified in paragraphs (a) and (c) of this AD, perform a torque check of the propeller retention bolts for low torque; a magnetic particle inspection of the propeller retention bolts for cracks; and dye penetrant, ultrasonic, and eddy current inspections of the propeller hub backface for cracks. Remove from service prior to further flight cracked propeller retention bolts and hubs, and replace with serviceable propeller retention bolts or hubs at intervals not to exceed 1500 hours TIS since the last torque check and inspections performed in accordance with paragraphs (b) or (d), as applicable, of this AD. These actions must be accomplished in accordance with Dowty Gloucester Aerospace Ltd. SB No. SF340-61-58, Revision 1, dated July 18, 1991. (f) Installation of newly designed Dowty Aerospace Gloucester Ltd. propeller hub assembly P/N 660714241 and remarking of the propeller in accordance with Dowty Aerospace Gloucester Ltd. SB No. SF340-61-61, Revision 1, dated October 19, 1992, constitutes terminating action to the inspection requirements of this AD. (g) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Boston Aircraft Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, BostonAircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Boston Aircraft Certification Office. (h) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (i) The inspections shall be done in accordance with the following Dowty Aerospace Gloucester service bulletins: Document No. Pages Revision Date SF340-61-57 1-5 Original February 15, 1991 Total pages: 5. SF340-61-58 1-2 1 July 18, 1991 3-5 Original February 15, 1991 Appendix A 1-3 Original February 15, 1991 Appendix B 1-3 Original February 15, 1991 Appendix C 1-2 1 July 18, 1991 3-4 Original February 15, 1991 Total pages: 15. SF340-61-61 1-3 1 October 19, 1992 4-8 Original September 9, 1992Total pages: 8. 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 Dowty Aerospace Gloucester Ltd., Anson Business Park, Cheltenham Road East, Gloucester GL2 9QN England. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (j) This amendment becomes effective on October 29, 1993.
94-17-05: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Airbus Model A340-211 and -311 series airplanes. This action requires replacement of certain circuit breakers for the toilet system vacuum generator. This amendment is prompted by reports of excessive vacuum generator overloads or blocked rotor conditions, in which the thermal protection device and/or related aircraft circuit breakers do not interrupt the power supply. The actions specified in this AD are intended to prevent a fire in the event of excessive current in the electrical circuit due to overheating of the vacuum generator.
95-12-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Model A320-231 series airplanes, that requires repetitive functional checks to detect leakage of the distribution piping of the engine fire extinguishing system, and repair, if necessary; and modification of the piping, which would terminate the inspection requirements. This amendment is prompted by reports of cracking of the engine fire extinguisher pipe, which resulted in leakage of the fire extinguisher agent. The actions specified by this AD are intended to prevent leakage of the fire extinguishing agent, which could prevent the proper distribution of the agent within the nacelle in the event of a fire.
95-24-06: This amendment supersedes an existing airworthiness directive (AD), applicable to Bell Helicopter Textron, A Division of Textron Canada (BHT), Model 206B and 206L helicopters, that currently requires a revision to the Limitations section of the FAA-approved Supplemental Type Certificate (STC) Rotorcraft Flight Manual Supplement (RFMS) until replacement of the engine power-out warning sensor on BHT Model 206B and 206L helicopters equipped with Allison 250-C20R engines in accordance with certain supplemental type certificates. This amendment requires a revision to the Limitations section of the STC RFMS, but removes the requirement for replacement of the engine power-out warning sensor. This amendment is prompted by a reevaluation of the need for an engine power-out warning sensor based on the lack of reported operational occurrences of the false engine-out warnings. The actions specified by this AD are intended to maintain a heightened pilot awareness that false engine-out warnings may occur when practicing autorotations and could result in an unnecessary emergency autorotative landing.
91-10-05: 91-10-05 O2 CORPORATION (FRANK MCGOWAN CO.): Amendment 39-6989. Docket No. 90-CE-43-AD. Applicability: The following Mask Presentation Unit Part Numbers that are installed on, but not limited to, British Aerospace 125-800A airplanes; Challenger CL600-1A11, CL600-2B16, and CL600-2A12 airplanes; Gulfstream G-1159, G-1159A, G-1159B, and G-IV airplanes; and Falcon 20 airplanes, certificated in any category: 121-040-04 150-004-03 151-020 150-002 150-004-04 151-020-02 150-002-01 150-004-05 151-020-04 150-002-02 150-004-06 152-001 150-002-03 150-004-07 152-001-01 150-002-04 150-004-08 152-001-04 150-002-05 150-004-12 152-001-05 150-002-08 150-005 152-001-08 150-003T 150-006 152-001-13 150-003-04T 150-022 152-003 150-004 151-010 152-004 150-004-01 151-010-02 152-004-05 150-004-02 151-010-04 Compliance: Required within the next 3 calendar months after the effective date of this AD, unless already accomplished. To prevent malfunctioning of the lanyard release pin that could prevent the flow of oxygen to a passenger in an emergency situation, accomplish the following: (a) With the oxygen system activated, perform a test of the lanyard release pins by accomplishing the following: (1) Open the passenger mask presentation units of the airplane and allow the mask assemblies to drop out. (2) Make up a 7.5 pound weight with an attached string and hook, (e.g., spring, clip, etc.). (3) Attach the hook to the lanyard attaching point of each actuator pin without dropping the weight and allow the weight to hang from the lanyard attaching point. (b) If the pin pulls free from the oxygen actuator valve at 7.5 pounds or less of hanging weight, then the pin is satisfactory and the unit may be returned to service. (c) If the pin does not pull free from the oxygen actuator valve using the test required by paragraph (a) of this AD, prior to further flight accomplish the following:(1) Replace the pin with either part number 100-111-2 or 100-111-3, which has a 20-degree angle and a rounded nose. NOTE 1: The pin is available from the manufacturer by contacting Mr. Burt Parry, O2 Corporation, 3522 N. Comotara, Wichita, Kansas 67226; Telephone (316) 634-1240; Facsimile (316) 634-1061. (2) Test the replacement pin installation in accordance with the test requirements of paragraph (a) of this AD to assure that the lanyard pin can be removed with a pull of 7.5 pounds or less. If the pin pulls free from the oxygen actuator valve at 7.5 pounds or less of hanging weight, then the pin is satisfactory and the unit may be returned to service. (d) An alternate method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas; Telephone (316) 946-4419. The request should be forwardedthrough an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (e) All persons affected by this directive may obtain copies of any information that is applicable to this AD from the FAA, Central Region, Office of the Assistant Chief Counsel, Attention: Rules Docket No. 90-CE-43-AD, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. Replacement parts that might be needed to complete the actions of this AD may be obtained from Mr. Burt Parry, O2 Corporation, 3522 N. Comotara, Wichita, Kansas 67226; Telephone (316) 634-1240; Facsimile (316) 634-1061. This amendment (39-6989, AD 91-10-05) becomes effective on June 3, 1991.
81-22-01: 81-22-01 McDONNELL DOUGLAS: Amendment 39-4244. Applies to all McDonnell Douglas Model DC-9-80 Series airplanes, certificated in all categories. \n\n\tCompliance required as indicated in the body of this AD, unless previously accomplished. To prevent failure of the slat drive mechanism, accomplish the following: \n\n\tA.\tUpon the accumulation of 600 landings, or within the next 200 landings after the effective date of this AD, whichever occurs later perform a one-time inspection of the slat control valve assembly to verify "correct" flow direction installation of check valves and restrictors as shown on Douglas Service Sketch 3303 of Douglas DC-9 Super 80 Alert Service Bulletin A27-225, Revision 1, dated September 22, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA, Northwest Mountain Region. Correct the installation of check valve/restrictor if necessary. \n\n\tNOTE: Directional arrow on restrictors and check valves may be in a position requiring additional visual aids to view. \n\n\tB.\tIf slat control valve assembly check valves and restrictors are installed properly, no further action is required. \n\n\tC.\tIf slat control valve assembly check valves and/or restrictors were improperly installed, accomplish the following in accordance with Douglas DC-9 Super 80 Alert Service Bulletin A27-225, Revision 1, dated September 22, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region, unless already accomplished: \n\n\t\t1.\tPerform dye penetrant or eddy current inspections of the slat drive cylinder support brackets per Paragraph 2.E of the service bulletin. \n\n\t\t2.\tConduct a visual inspection of the slat drive mechanism per Paragraph 2.F of the service bulletin. \n\n\t\t3.\tIf no cracks or other indications of distress are detected, conduct repetitive inspections per Paragraph 2.G of the service bulletin, at intervals not to exceed 350 landings.4.\tIf cracked or failed support brackets or other indications of distress in the slat drive mechanism are detected, remove the slat drive mechanism per Paragraph 2.H of the service bulletin. \n\n\t\t5.\tReplace cracked support brackets per Paragraph 2.I of the service bulletin. \n\n\t\t6.\tIf a slat drive cylinder support bracket is failed, or other indications of distress are detected, accomplish the following: \n\n\t\t\ta.\tInspect and repair per the instructions of paragraph 2.J(1) of the service bulletin. \n\n\t\t\tb.\tInspect wing spar per the instructions of paragraph 2.J(2) of the service bulletin. \n\n\t\t7.\tIf inspection of the wing spar reveals any indication of damage, repair in a manner approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region. \n\n\t\t8.\tInstall repaired slat drive mechanism per the installation instructions of Paragraph 2.K of the service bulletin. \n\n\t\t9.\tFor slat drive mechanisms which have not had both slat drive support cylinder support brackets replaced, conduct repetitive inspections per paragraph C.3 of this AD. \n\n\t\t10.\tInstallation of new slat drive cylinder support brackets terminates the requirements for repetitive inspections per this AD. \n\n\tD.\tSpecial flight permits may be issued in accordance with FARs 21.197 and 21.199 to operate airplanes to a base in order to comply with the inspection requirements of this AD. \n\n\tE.\tIf the number of landings on an aircraft is not known, for the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing the airplane's hours time in service by the operator's fleet average time from takeoff to landing for the airplane type. \n\n\tF.\tAlternative means of compliance with this AD which provide an equivalent level of safety may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tThe 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 McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). These documents also may be examined at FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108, or 4344 Donald Douglas Drive, Long Beach, California 90808. \n\n\tThis amendment becomes effective November 5, 1981.
77-22-04: 77-22-04 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE (S.N.I.A.S.): Amendment 39-3068. Applies to Models SA341G "Gazelle" helicopters, certificated in all categories, equipped with any of the following landing gear: Low-skid, low-frequency landing gear P/Ns 341A.41.5200.00 to .04 not incorporating Mod. AMS 07.1349/S.371. P/Ns 341A.41.5200.05 to .11 incorporating Mod. AMS 07.1349 but not incorporating Mod. AMS 07.1578/S.440. High-skid, low-frequency landing gear P/N 341A.41.5300.00 not incorporating Mod. AMS 07.1350. P/Ns 341A.41.5300.01 to .11 incorporating Mod. AMS 07.1350 but not incorporating Mod. AMS 07.1579. Compliance is required within the next 500 hours time in service after the effective date of this AD, but not later than January 31, 1978, unless already accomplished. To prevent possible excessive corrosion remove the landing gear, inspect for corrosion, and if corrosion is found, replace, or turn 180 degrees and thereafter replace, the landing gear members, as necessary and reinstall all in accordance with Paragraph C of "Gazelle" Service Bulletin No. 01.10 dated September 12, 1975, revised December 13, 1976 and the Models SA341G "Gazelle" Repair Manual or an FAA approved equivalent. This Amendment becomes effective November 28, 1977.
84-09-02: 84-09-02 AIRBUS INDUSTRIE: Amendment 39-4853. Applies to Model A300 B2 and B4 series airplanes, certificated in all categories. Compliance is required within the next 750 hours time in service after the effective date of this AD, unless already accomplished. To prevent breakage of the fuel and hydraulic line attaching clamps in the engine pylons, which could result in line breakage and a consequent fire hazard, accomplish the following: A. Remove the existing fuel and hydraulic line attaching clamps and brackets, and install new reinforced teflon-lined clamps and redesigned brackets in accordance with paragraph 2, "ACCOMPLISHMENT INSTRUCTIONS," of Airbus Industrie Service Bulletin A300-54-007, Revision 5, dated December 22, 1978. 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 June 1, 1984.