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64-21-06 R2: 64-21-06 R2 BOEING VERTOL: Amendment 813 Part 507 Federal Register September 10, 1964 as amended by Amendment 39-2156 is further amended by Amendment 39-3827. Applies to Model 107-II helicopters. Compliance required as indicated. As a result of a fatigue failure in service of the rotor pitch housing, accomplish the following: (a) Unless already accomplished within the last 30 hours' time in service, before further flight, inspect the lug areas of all rotor pitch housings P/N's 107R2553-1, -2, -3, -4, -5, and -6 and blade sockets P/N's 42R1043-7 and -8 with 260 or more hours' time in service using magnetic particle inspection method or FAA approved equivalent. To accomplish the inspection, remove rotor blades and rotor hub pitch bearing assemblies. Repeat this inspection at intervals not to exceed 30 hours' time in service. (b) Inspect, using magnetic particle inspection, the lug areas of all rotor pitch housings P/N's 107R2553-1, -2, -3, -4, -5, and -6 and blade sockets P/N's 42R1043-7 and -8 with less than 260 hours' time in service in accordance with (a) prior to the accumulation of 260 hours' time in service. (c) Conduct a daily visual inspection for cracks in the lug areas of all rotor pitch housings P/N's 107R2553-1, -2, -3, -4, -5, and -6 and blade sockets P/N's 42R1043-7 and -8. This may be accomplished without disassembly from the helicopter. (d) Unless already accomplished, within the next 50 hours in service on pitch housing 107R2553-8, -10, -14, -16, with 1000 hours or more in service and within the next 100 hours in service on pitch housing 107R2553-7, -9, -13, 15, with 2000 hours or more in service install crack detector wire in accordance with Part I "Installation Procedure" of Boeing Service Bulletin No. 107-343 dated March 10, 1980, or equivalent. (1) Inspect for cracks in accordance with Part II "Inspection Procedures" of the above Bulletin, or equivalent, the lug area of pitch housings 107R2553-8, -10, -14, -16 with 1000 hours or more in service within the next 50 hours in service and thereafter at intervals not to exceed 25 hours in service, and pitch housing 107R2553-7, -9, -13, -15 with 2000 hours or more in service within the next 100 hours in service and thereafter at intervals not to exceed 50 hours in service. (2) Unless already accomplished, install crack detector wire in accordance with Part I "Installation Procedure" of the above Bulletin, or equivalent on pitch housings 107R2553-8, -10, -14, -16 with less than 1000 hours in service prior to the accumulation of 1050 hours in service, and on pitch housings 107R2553-7, -9, -13, -15 with less than 2000 hours in service prior to the accumulation of 2100 hours in service. (3) Inspect pitch housings 107R2553-8, -10, -14, -16 with less than 1000 hours in service in accordance with (1) prior to accumulation of 1050 hours in service. Inspect pitch housings 107R2553-7, -9, -13, -15 with less than 2000 hours in service in accordance with (1) prior to accumulation of 2100 hours in service. (4) Conduct a visual inspection for cracks in the lug area of blade sockets 42R1043-11, -12, -13 and -14 at intervals not to exceed 50 hours in service. This may be accomplished without disassembly from the helicopter. (e) If any cracks are found replace the part before further flight with a part found serviceable in accordance with this AD. (f) Upon request with substantiating data submitted through an FAA Maintenance Inspector, the compliance times specified in this AD may be adjusted by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (g) Retire from service all rotor pitch housings P/N's 107R2553-7, -8, -9, -10, -13, - 14, -15, d -16 upon the accumulation of 5,000 hours' total time in service. This supersedes AD 64-06-08. Amendment 39-813 was effective September 10, 1964. Amendment 39-2156 was effective April 9, 1975. This Amendment 39-3827 is effective July 7,1980.
60-12-01: 60-12-01 CONTINENTAL: Amdt. 153, Part 507, Federal Register May 17, 1960. Applies to E185-8, E185-9, E185-11, E225-4 and E225-8 Engine Models. Compliance required at next periodic inspection, but not later than September 1, 1960. An inflight failure has indicated that additional information regarding engines affected by replacement requirements of AD 56-06-01 should be provided. This AD is therefore issued to supply specific serial numbers of engines that were rebuilt (remanufactured) between April 1, 1954 and May 1, 1955. The affected engines are: Remanufactured E185-8, -9 and -11 Engines: 25015, 25037, 25044, 25048, 25057, 25065, 25071, 25075, 25086, 15087, 25137, 25141, 25162, 25189, 25202, 25210, 25220, 25234, 25243, 24254, 25269, 25288, 25307, 25320, 25325, 25333, 25376, 25379, 25381, 25387, 25422, 25426, 25464, 25518, 25526, 25545, 25562, 25575, 25578, 25611, 25649, 25718, 25754, 15761, 25766, 25767, 25783, 25790, 25795, 25819, 25834, 25897, 25930, 25950, 25957, 25958, 26003, 26088, 26095, 26104, 26121, 26138, 26304, 26321, 26327, 26343, 26352, to 26412 inclusive. Remanufactured E225-4 and -8 Engines: 30122, 30391, 30454, 32154, 35001, 35082, 35086, 35095, 35113, 35128, 35132, 35133, 35135, 35137, 35138, 35139, 35144, 35145, 35151, to 35254 inclusive. The above engines may have piston pin assembly P/N 530845, which is satisfactory, or piston pin assembly P/N 535145, which is unsatisfactory. Unless previously accomplished per Continental Service Bulletin No. M56-2 dated February 14, 1956, including Supplement No. 1 dated March 12, 1956, or AD 56-06-01, replace piston pin assembly P/N 535145 with P/N 539467. Use the applicable method of inspection outlined below to determine which piston pin assembly is installed in the above remanufactured engines: (a) If none of the cylinders on the engine in question have been removed in the field since the engine was shipped from the factory, remove and inspect the piston pin assembly in any one of the cylinders. Continental Motors Corporation procedures provide that all cylinders will have the same piston pin assembly. (b) If the engine in question has had any of the cylinders removed in the field since the engine was shipped from the factory, inspect those cylinders and also at least one of the factory installed cylinders which has not been disturbed. This supplements AD 56-06-01 and supersedes AD 59-10-04.
85-22-08 R2: 85-22-08 R2 GATES LEARJET CORPORATION: Amendment 39-5166 as amended by Amendment 39-5223 is further amended by Amendment 39-5296. Applies to Model 55 airplanes, serial numbers 55-003 through 55-086, except those incorporating Service Bulletin 55- 27-7A, Airplane Modification Kit 55-84-7B, or Airplane Accessory Kit 55-83-4. Compliance required as indicated after the effective date of this AD, unless already accomplished. To prevent use of incorrect landing distances, accomplish the following: A. Within the next 10 hours time-in-service after the effective date of this AD, insert the following information in the "FAA Approved Airplane Flight Manual for Gates Learjet 55," dated 3/17/81 (on 10/4/85, changes through Change 20, dated 5/10/84, were current): 1. Replace the paragraph "SPOILERS/AUTOSPOILERS/ SPOILER0NS" (pages 1-14) with: "SPOILERS/SPOILERONS (a) Spoilers (1) During landings, a time delay in the spoiler circuit will cause increased spoiler extension times. For a normal landing (spoilerons operative with flaps below 25 degrees), full spoiler extension will require approximately 5 seconds. If the spoilerons are inoperative (AUG AlL light illuminated) or the SPOILER0N circuit breaker (co-pilot's AC bus) is pulled, spoiler extension will require approximately 11 seconds. To account for the increase spoiler deployment times, the following corrections must be applied to the Actual Landing Distance obtained from the ACTUAL LANDING DISTANCE chart in Section V: Normal Landing: Multiply distance by 1.04. Spoilerons Inoperative: Multiply distanced 1.15. (2) If the spoilers are inoperative during flight, the maximum operating altitude is limited to 41,000 feet. (3) Do not extend spoilers with flaps extended while airborne. (4) Do not extend spoilers, or operate with spoilers deployed, at speeds above VM0/MM0. (b) Spoilerons (1) Spoilerons may be inoperative. If spoileron preflight check fails, the SP0ILERON circuit breaker (co-pilot's AC bus) must be pulled. When the SP0ILERON circuit breaker is pulled, the spoilers and spoilerons will be inoperative in flight and landing spoiler deploy times will increase. With spoilerons inoperative, the maximum operating altitude will be limited to 41,000 feet, and the Actual Landing Distance obtained from Section v must be increased by 15 percent (multiply by 1.15)." 2. Replace the first NOTE under the BEFORE TAXI Spoileron System Check (page 2-20) with the following: "In the event the spoilerons fail the following check, SP0ILERON circuit breaker (co-pilot's AC bus) - Pull. AUG AlL will illuminate. The spoilers and spoilerons will be inoperative in flight. The spoiler deploy time will be increased for landing. In this event, the maximum operating altitude is limited to 41,000 feet and the Actual Landing Distance obtained from Section V must be increased 15 percent (multiply by 1.15)." 3. The following factors must be applied to the Actual Landing Distance obtained from figure 5-57 (page 5-74): "Normal Landing: Multiply distance by 1.04 Spoilers Inoperative: Multiply distance by 1.15." B. In order to comply with the requirements of paragraph A., above, a copy of this AD may be used as a temporary amendment to the Airplane Flight Manual (AFM) and carried in the airplane as part of the AFM until replaced by Gates Learjet-published Temporary Flight Manual (TFM) Changes TFM 85-17, TFM 85-18, and TFM 85-19. C. Temporary Flight Manual (TFM) Changes TFM 85-17, TFM 85-18, and TFM 85-19 may be removed from the AFM upon incorporation of either Gates Learjet Service Bulletin (SB) 55-27-7A (dated December 12, 1985), Airplane Modification Kit (AMK) 55-84-7B (dated December 12, 1985), or Airplane Accessory Kit (AAK) 55-83-4. D. For airplanes equipped with thrust reversers modified in accordance with Gates Learjet SB 55-27-7 or AMK 55-84-7A, that modification (SB 55-27-7 or AMK 55-84-7A) must be removed within the next 30 days after the effective date of this AD. E. For airplanes not equipped with thrust reversers, modification in accordance with Gates Learjet SB 55-27-7 or AMK 55-84-7A constitutes terminating action for the requirements of this AD. F. Alternate means of compliance with provide an acceptable level of safety may be used when approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Gates Learjet Corporation, P.0. Box 7707, Wichita, Kansas 67277. These documents also may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Mid-Continent Airport, Wichita, Kansas. Amendment 39-5166 became effective November 22,1985. Amendment 39-5223 became effective February 6, 1986. This Amendment 39-5296 becomes effective May 19, 1986.
2016-12-13: We are superseding Airworthiness Directive (AD) 2000-05-17 and AD 2001-04-12, which apply to Eurocopter France (now Airbus Helicopters) Model EC120B helicopters. AD 2000-05-17 and AD 2001-04-12 required repetitive visual checks of the engine-to-main gearbox (MGB) coupling tube assembly (coupling tube) for a crack and replacing any cracked tube with an airworthy tube. This new AD requires removing certain engine mount parts from service, measuring the height of the engine mounting base for certain helicopters, replacing the engine mount if a certain height is exceeded, inspecting the flared coupling on certain helicopters for a crack, and replacing the coupling if it is cracked. Since we issued AD 2000-05-17 and AD 2001-04-12, there have been reports of additional cracks in coupling tubes. These actions are intended to prevent coupling tube failure, loss of engine drive, and a subsequent forced landing of the helicopter.
2016-12-15: We are superseding Airworthiness Directive (AD) 2016-07-30 for all Airbus Model A330-200, -200 Freighter, and -300 series airplanes, and all Airbus Model A340-200, -300, -500, and -600 series airplanes. For certain airplanes, AD 2016-07-30 required replacing certain Angle of Attack (AOA) sensors (probes) with certain new AOA sensors. For certain other airplanes, AD 2016-07-30 also required inspections and functional heat testing of certain AOA sensors for discrepancies, and replacement if necessary. This new AD requires the same actions as AD 2016-07-30. This new AD was prompted by a report of a typographical error in the regulatory text of AD 2016-07-30. We are issuing this AD to prevent erroneous AOA information and Alpha Protection (Alpha Prot) activation due to blocked AOA probes, which could result in a continuous nose-down command and consequent loss of control of the airplane.
2016-12-12: We are superseding Airworthiness Directive (AD) 2008-05-18 R1 for certain Fokker Services B.V. Model F.27 Mark 050, 200, 300, 400, 500, 600, and 700 airplanes. AD 2008-05-18 R1 required revising the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness to incorporate new limitations for fuel tank systems. This new AD requires a new maintenance or inspection program revision to incorporate the revised Airworthiness Limitation Items (ALIs) and critical design configuration control limitations (CDCCLs). This new AD also adds certain airplanes to the applicability. This AD was prompted by the issuance of revised service information to update the Fuel ALIs and CDCCLs that address fuel tank system ignition sources. We are issuing this AD to prevent 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.
2001-08-17: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model DC-8 series airplanes, that requires an inspection of the antifogging or heating wiring to detect chafing or damage, as applicable; inspection of the insulation blankets to detect damage; and repair, if necessary. This amendment also requires revising the clearview window heating wiring installations. This amendment is prompted by a report of an electrical short that resulted in damage to the antifogging circuit wiring and insulation blanket above the Captain's clearview window. The actions specified by this AD are intended to prevent chafed and damaged wires as a result of a sharp bend and restricted space between the fuselage frame and the clearview window in the full open position, which could result in an electrical short, damage to the antifogging circuit wiring and insulation blanket, and consequent smoke and fire in the flight deck.
2016-11-14: We are adopting a new airworthiness directive (AD) for all Fokker Services B.V. Model F.28 Mark 1000, 2000, 3000, and 4000 airplanes. This AD was prompted by a design review that revealed no controlled bonding provisions are present on a number of critical locations inside the fuel tanks or connected to the walls of the fuel tanks. This AD requires installing additional and improved bonding provisions in the fuel tanks and revising the airplane maintenance or inspection program, as applicable, by incorporating fuel airworthiness limitation items and critical design configuration control limitations (CDCCLs). We are issuing this AD to prevent an ignition source in the fuel tank vapor space, which could result in a fuel tank explosion and consequent loss of the airplane.
90-15-15 R1: 90-15-15 R1 AMERICAN CHAMPION AIRCRAFT (BELLANCA, CHAMPION): Amendment 39-6671 as amended by Amendment 39-6785; Docket No. 90-CE-25-AD. Applicability: Model 8KCAB airplanes (all serial numbers) that are equipped with upper wing front spar fittings part number (P/N) 2-1976, certificated in any category. Compliance: Required as indicated after the effective date of this AD. To prevent failure of the upper wing front spar strut fittings, P/N 2-1976, that could result in an in-flight separation of the wing, accomplish the following: (a) Within the next 25 hours time-in-service (TIS) after the effective date of this AD or prior to the accumulation of 500 hours TIS on the front spar strut fittings (P/N 2-1976), whichever occurs later, unless previously accomplished within the last 250 hours TIS, and thereafter at intervals not to exceed 250 hours TIS from the last inspection, accomplish the following: NOTE: Operators who have not kept records of hours TIS on individual front spar strut fittings (P/N 2-1976) may substitute airplane hours TIS instead. (1) Remove the front spar strut fittings (P/N 2-1976) and strip all paint with a chemical stripper. Clean and prepare the fittings for a magnetic particle inspection. (2) Conduct a magnetic particle inspection of the fittings for cracks, paying close attention to the areas near the welds. (3) If cracks are not found, prior to further flight, clean the fittings and apply a spray coat or a dip coat of zinc chromate primer, reinstall the fittings, and return the airplane to service. (b) If cracks are found as a result of the inspection required by paragraph (a)(2) of this AD, prior to further flight, replace any cracked fittings with one of the following: (1) A new or serviceable fitting (P/N 2-1976) that has been inspected and treated per the requirements of paragraph (a) of this AD. (2) A new American Champion Aircraft fitting (P/N 3-1658) that is installed inaccordance with the instructions in American Champion Aircraft Service Kit 302, revised October 1, 1990. (3) A new Safe Aircraft Repair, Inc. fitting (P/N SAR2-1976) and stiffener (P/N SAR2-5001) that are installed in accordance with the instructions in STC SA1514GL, issued to Safe Aircraft Repair, Inc. on August 27, 1990. (c) Upper wing front spar strut fittings (P/N 2-1976) may be replaced with new parts in accordance with paragraphs (b)(2) or (b)(3) of this AD regardless of whether cracks are found during the inspection required by paragraph (a) of this AD. (d) Replacement of the upper wing front spar strut fittings (P/N 2-1976) with new parts in accordance with paragraphs (b)(2) or (b)(3) of this AD constitutes terminating action for the repetitive inspections required by paragraph (a) of this AD. (e) 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. (f) An alternate 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, Chicago Aircraft Certification Office, 2300 E. Devon Avenue, Des Plaines, Illinois 60018. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Chicago Aircraft Certification Office. (g) All persons affected by this directive may obtain copies of the documents referred to herein upon request to American Champion Aircraft, P.O. Box 37, Rochester, Wisconsin 53167; Telephone (414) 534-6315; or Safe Air Repair, Inc., 3325 Bridge Avenue, Albert Lea, Minnesota 56007; Telephone (507) 373-5408; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. Airworthiness Directive 90-15-15 R1 revises AD 90-15-15, Amendment 39-6671. This amendment (39-6785, AD 90-15-15 R1) becomes effective on May 15, 1991.
72-13-03: 72-13-03 BELLANCA: Amdt. 39-1461. Applies to Model 17-30 Airplanes. Compliance: As indicated below, unless already accomplished. To prevent possible engine flooding when using the fuel boost pump, accomplish either Part A or Part B as applicable: PART A 1. On those airplanes equipped with Airborne Model 2B6-9 fuel boost pumps (Airplane serial numbers 30217 through 30262 were delivered from the factory with this model pump installed) which have not been modified in accordance with AD 71-13-04: a. Within 50 hours' time in service after June 26, 1971, modify the fuel boost pump electrical circuit by installing a three (3) position toggle switch, a three (3) ohm twenty (20) watt resistor, a switch guard and a five (5) amp circuit breaker between the bus and the switch in accordance with Bellanca Service Letter No. 61A, Revision A, dated April 26, 1971, or later FAA approved revisions, and Bellanca Drawing SK-2-1040, Sheet 1, Revision D. Do not connect any other equipment to the fuel boost pump circuit. b. Within 50 hours' time in service after June 26, 1971, insert Airplane Flight Manual, Revision No. 13, dated May 26, 1971, in the Airplane Flight Manual. (Revision No. 13 is included in Bellanca Service Kit SK-2-1040 referred to in Service Letter No. 61A, Revision A.) c. Any alternate equivalent method of compliance with Paragraphs a and b above must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region, except that an equivalent five (5) amp circuit breaker may be utilized. 2. On those airplanes equipped with Airborne Model 2B6-9 fuel boost pumps (Airplane serial numbers 30217 through 30262 were delivered from the factory with this model pump installed) which have complied with AD 71-13-04: a. Within 50 hours' time in service after the effective date of this AD and install a five (5) amp circuit breaker between the bus and the fuel boost pump switch in accordance with Bellanca Drawing SK-2-1040, Sheet 1, Revision D. Do not connect any other equipment to the fuel boost pump circuit. b. Any alternate equivalent method of compliance with Paragraph a above must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region, except that an equivalent five (5) amp circuit breaker may be utilized. PART B On those airplanes equipped with Weldon fuel boost pump Models 4020-A2A or 10050- A (Airplane serial numbers 30002 through 30216) were delivered from the factory with one of these model pumps installed: 1. Effective immediately, do not operate the fuel boost pump any longer than is necessary to achieve required fuel pressure. (Continued use of the fuel boost pump may cause engine flooding under certain operating conditions.) 2. Within 10 hours' time in service after September 14, 1971, install a placard beneath or adjacent to the fuel boost pump switch to read as follows: "TO PREVENT ENGINE FLOODING TURN OFF FUEL BOOST PUMP IMMEDIATELY AFTER FUEL PRESSURE IS RESTORED." NOTE: The operator may make and install this placard using letters approximately 1/8 inch in height. 3. Within 50 hours' time in service after the effective date of this AD: a. Install a two (2) position spring loaded switch and a five (5) amp circuit breaker between the bus and the fuel boost pump switch in accordance with Bellanca Service Letter No. 71, dated February 16, 1972, or later FAA approved revisions, and Bellanca Drawing SK-2-1040, Sheet 2, Revision A. Do not connect any other equipment to the fuel boost pump circuit. b. Insert Airplane Flight Manual, Revision No. 14, dated April 17, 1972, in the Airplane Flight Manual. (Revision No. 14 is included in kit referred to in Service Letter No. 71.) c. Remove placard installed under Part B, Paragraph 2, and install, in same location, placard Bellanca Part No. SK-2-1043 which reads as follows: "AUX FUEL PUMP USE TO RESTORE FUEL PRESSURE RELEASE TO PREVENT ENGINE FLOODING" 4. Any alternate equivalent method of compliance with Part B must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region, except that an equivalent five (5) amp circuit breaker may be utilized. This AD supersedes AD 71-13-04. This amendment becomes effective June 20, 1972.