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84-10-51 R1: 84-10-51 R1 BOEING VERTOL COMPANY: Amendment 39-5024 as amended by Amendment 39-5352. Applies to Model 234 series helicopters, certificated in any category, equipped with forward main rotor drive shaft P/N 114D1245-7. Compliance is required as indicated, unless already accomplished. To prevent possible hazards in flight associated with cracking of the forward main rotor drive shaft, accomplish the following: (a) Within 70 hours time in service after the effective date of this amendment or upon accumulation of 1,600 total hours time in service on the rotor shaft, whichever occurs later, and thereafter at intervals not to exceed 35 hours time in service from the last inspection, inspect the rotor drive shaft in accordance with paragraph 3, "Accomplishment Instructions," Boeing Vertol Service Bulletin No. 234-63-1009, dated June 29, 1984, or Revision 1 dated May 1, 1986, or an FAA-approved equivalent. (b) Remove from service forward main rotor drive shafts havinga crack and replace with a serviceable part prior to further flight. (c) An equivalent method of compliance with this AD may be used when approved by the Manager, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581. (d) Upon submission of substantiating data by an owner through an FAA Maintenance Inspector, the Manager, New York Aircraft Certification Office, FAA New England Region, may adjust the compliance times specified in this AD. (e) 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 Boeing Vertol Company, Boeing Center, P.O. Box 16858, Philadelphia, Pennsylvania 19142. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, Federal Aviation Administration, 4400 Blue Mound Road, Fort Worth, Texas. Amendment 39-5024 (50 FR 15539) became effective April 22, 1985, as to all persons except those persons to whom it was made immediately effective by telegraphic AD No. T84-10- 51, issued May 3, 1984, which contained this amendment. This Amendment 39-5352 becomes effective July 24, 1986.
2013-19-01: We are adopting a new airworthiness directive (AD) for AgustaWestland S.p.A. (AgustaWestland) Model A119 and AW119 MKII helicopters to require inspecting the pilot and co-pilot doors to ensure that the windows are properly bonded within the doors. If the windows are not properly bonded, the AD requires applying bonding to the windows, the seals, and the window frames of the pilot and co-pilot doors. This AD was prompted by the loss of a pilot-door window during a test flight. The actions of this AD are intended to ensure the windows do not detach from the doors, potentially injuring persons on the ground and damaging the helicopter's tailboom and the tail rotor blades.
2013-18-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 767-200, -300, -300F, and -400ER series airplanes. This AD was prompted by reports of cracked and corroded nuts on an outboard flap support rib. This AD requires, for certain airplanes, repetitive inspections of the cap seal for damaged sealant on nuts common to certain outboard flap support ribs, and related investigative and corrective actions if necessary. For certain other airplanes, this AD also requires repetitive inspections of the cap seal for damaged sealant on nuts common to certain outboard flap support ribs, related investigative and corrective actions if necessary, and if necessary, a detailed inspection to determine the nut type installed in the outboard flap support rib and corrective actions. This AD also provides terminating action for the repetitive inspections under certain conditions. We are issuing this AD to detect and correct cracked and corroded nuts and bolts and the installation of incorrect nuts on certain outboard flap support ribs, which could lead to additional nut and bolt damage in the joint, result in loss of an outboard flap, and adversely affect continued safe flight and landing of the airplane.
99-07-12: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 747-100, -200, and -300 series airplanes, that requires repetitive inspections to detect cracking of certain lower lobe fuselage frames, and repair, if necessary. This amendment is prompted by reports indicating that fatigue cracks were found in lower lobe frames on the left side of the fuselage. The actions specified by this AD are intended to detect and correct fatigue cracking of certain lower lobe fuselage frames, which could lead to fatigue cracks in the fuselage skin, and consequent rapid decompression of the airplane.
89-16-09: 89-16-09 LOCKHEED: Amendment 39-6288. Applicability: Model 382 series airplanes, serial numbers 3946 through 4637, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent failure of the nose landing gear strut assembly, accomplish the following: A. Within the next 250 hours time-in-service after the effective date of this AD, and thereafter at intervals not to exceed 3,150 hours time-in-service, perform an inspection for corrosion of the nose landing gear strut assembly, part number 388071-3, in accordance with Lockheed Service Bulletin 382-32-31, dated September 5, 1979. 1. If corrosion is detected which is within the limits specified in the service bulletin, rework in accordance with the service bulletin prior to further flight. 2. If corrosion is detected which is outside the limits specified in the service bulletin, prior to further flight, accomplish one of the following: a. Replace the defective cylinder assembly, part number 371675-1 with a serviceable cylinder assembly; or b. Modify the nose landing gear strut and components in accordance with Lockheed Service Bulletin 382-32-23, dated March 17, 1978. B. Modification of the nose landing gear strut and components in accordance with Lockheed Service Bulletin 382-32-23, dated March 17, 1978, constitutes terminating action for the inspection requirements of this AD. 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, Atlanta Aircraft Certification Office, FAA, Central Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Atlanta Aircraft Certification Office. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a basein order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Lockheed Aeronautical Systems Company, 86 South Cobb Drive, Marietta, Georgia 30063. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Atlanta Aircraft Certification Office, FAA, Central Region, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia. This amendment (39-6288, AD 89-16-09) becomes effective on September 7, 1989.
80-13-14: 80-13-14 CESSNA: Amendment 39-3844. Applies to the following Model 310R, 335, 340A, 402C, 404, 414A, and 42lC airplanes certificated in all categories: Model Serial Number 310R 1865 335 0044, 0049, 0052, and 0059 340A 0959, 0961, 0962, 0967, 0972, 0978, 0982, 0983, 0984, 0985, 0998, 1002, 1003, 1012, 1014, and 1019 402C 0259, 0263, 0275, 0283, and 0288 404 0612, 0620, 0623, and 0628, 414A 0457, 0458, 0463, 0464, 0466, 0469, 0470, 0472, 0475, 0476, 0482, 0483, 0486, 0487, 0490, 0495, and 0497 421C 0823 through 0866, 0869, and 0870 COMPLIANCE: Required as indicated unless already accomplished. To preclude failure of the fuel flow transducer and resultant leakage of fuel within the engine compartment, prior to further flight accomplish the following on both engines: A) Remove engine cowling and inspect Cessna P/N 9910395-9 (Aerosonics Corp 33184-2) fuel flow transducer on each engine for serial number. B) If the P/N 9910395-9 transducer has a serial number included in the block 2364 through 2930, replace the transducer prior to further flight with a transducer whose serial number is not included in the block 2364 through 2930, reinstall the engine cowling, and accomplish Item D of this AD. C) If the P/N 9910395-9 transducer serial number is not included in block 2364 through 2930, reinstall the engine cowling, and accomplish Item D of this AD. D) Make the prescribed entry in the aircraft maintenance record indicating compliance with this AD when B, or C has been accomplished. E) Any equivalent method of compliance with this AD must be approved by the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. This amendment becomes effective on July 22, 1980, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated June 24, 1980.
77-15-05: 77-15-05 LET N.P.: Amendment 39-2974. Applies to Blanik 13 gliders, certificated in all categories, with serial numbers 173205 through 174230, inclusive. Compliance is required as indicated. To prevent structural failure of the fin top rib, accomplish the following: (a) Within the next 10 hours time in service after the effective date of this AD, unless already accomplished within the last 90 hours time in service, and, thereafter, at intervals not to exceed 100 hours time in service from the last inspection, visually inspect the fin top rib and central stiffener (fuselage stiffener) with a 5 power magnifying glass in accordance with the accomplishment instructions set forth in paragraph A of LET N.P. UH HRADTSTE-KUNOVICE (LET N.P.) Mandatory Bulletin No. L13/040, undated, or an FAA-approved equivalent. (b) If cracks are found in the fin top rib of less than 5 mm in length, or of any length in the central stiffener, before further flight, either - (1) Repair the fin top rib or central stiffener as necessary, in accordance with the accomplishment instructions set forth in paragraph B of LET N.P. Mandatory Bulletin No. L13/040, undated, or an FAA-approved equivalent; or (2) Comply with paragraph (e) of this AD. (c) If any cracks are found in the fin top rib which exceed 5 mm in length, before further flight, comply with paragraph (e) of this AD. (d) If no cracks are found, within the next 200 hours time in service after the effective date of this AD, comply with paragraph (e) of this AD. (e) Replace both the fin top rib and central stiffener in accordance with the accomplishment instructions set forth in paragraphs C and D of Let N.P. Mandatory Bulletin No. L13/040, undated, or an FAA-approved equivalent. NOTE: To assist in the dismantling of the rudder the following information is provided: (1) It is recommended that, in loosening the split pin securing the slotted nut in the bottom rudder hinge, the bent ends of the split pin be straightened first using a suitable pointed tool and then the pin be removed. (2) The rudder dismantling procedures are as follows: (i) The nut of the bottom rudder hinge accessible through the mounting opening on the fuselage part should be released and unscrewed. (ii) The cloth blinds on the rudder edge curve and leading edge should be removed. The top rudder hinge stud bolt should be released and unscrewed. (iii) The rudder should be lifted slightly and its top edge should be pulled slightly backwards to be released from the top hinge. (iv) The rudder should be held in the inclined position and carefully lifted from the step ball bearing. (3) The bottom rivets of the intermediate stiffener supporting the rudder bearing should be routed by an extended drill rod through the 60 mm. diameter opening in the rear fuselage partition as recommended by the Note in paragraph D of LET N.P. Mandatory Bulletin L13/040. Routingof the rear fuselage part in position 14 of the partition is not recommended. (4) A modified shank should be used for riveting a new fin rib. For this purpose, a 35 mm. diameter opening should be made in the fin spar (See LET N.P. Mandatory Bulletin L13/040, paragraph C). (f) The inspections required by paragraph (a) of this AD may be discontinued when the fin top rib and central stiffener have been replaced in accordance with paragraph (e) of this AD. This amendment supersedes Amendment 39-2333 (40 FR 33007), AD 75-17-28, as amended by Amendments 39-2379 (40 FR 45802) and 39-2498 (41 FR 2375). This amendment becomes effective August 4, 1977.
2022-11-17: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model BD-700-1A10 and BD-700-1A11 airplanes. This AD was prompted by reports of internal corrosion on the inboard flaps found prior to regularly scheduled maintenance checks. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate a certain aircraft maintenance manual (AMM) task. The FAA is issuing this AD to address the unsafe condition on these products.
99-07-11: This amendment adopts a new airworthiness directive (AD) that applies to all SOCATA - Groupe Aerospatiale (SOCATA) Model TBM 700 airplanes. This AD requires inspecting the left-hand and right-hand outboard hinge fittings of the horizontal stabilizer for cracks, and replacing any cracked fitting. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for France. The actions specified by this AD are intended to prevent structural damage to the stabilizer caused by outboard hinge fitting cracks, which could result in uncontrolled flight if the hinges break.
2013-17-08: We are superseding airworthiness directive (AD) 2010-20-08, which applied to certain The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747-200B, 747-200C, 747-200F, 747-300, 747-400, 747-400D, 747-400F, and 747SR series airplanes. AD 2010-20-08 required repetitive inspections to find cracking of the web, strap, inner chords, inner chord angle of the forward edge frame of the number 5 main entry door cutouts; the frame segment between stringers 16 and 31; repair if necessary; and repetitive inspections for cracking of repairs. This new AD expands the previous fuselage areas that are inspected for cracking. This AD was prompted by multiple reports of cracking outside of the previous inspection areas and a report of a crack that initiated at the aft edge of the inner chord rather than initiating at a fastener location. We are issuing this AD to detect and correct such cracks, which could cause damage to the adjacent body structure and could result in depressurization of the airplane in flight.
2006-14-05: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Bombardier Model CL-600-2C10 (Regional Jet Series 700 and 701) and CL-600-2D24 (Regional Jet Series 900) series airplanes. That AD currently requires repetitive detailed inspections for cracking or deformation, or pulled or missing fasteners, on the lower panel of the left- and right-hand main landing gear (MLG) doors, as applicable, and corrective actions if necessary. This new AD reduces the repetitive inspection interval for certain airplanes. This new AD also adds airplanes to the applicability. This AD results from a report of a MLG door departing from an airplane. We are issuing this AD to prevent failure of the lower panel of the MLG door, departure of the lower panel from the airplane, and consequent damage to airplane structure, which could adversely affect the airplane's continued safe flight and landing.
95-18-09: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 0100 series airplanes, that requires modification of the rear spar-to-fuselage attachment. This amendment is prompted by a report indicating that, during full-scale fatigue tests on a Model F28 Mark 0100 test article, cracking was found in the coupling plate and web plate of the rear spar end fitting at the attachment to the main frame at fuselage station 17011 due to fatigue-related stress. The actions specified by this AD are intended to prevent fatigue-related cracking in the rear spar-to-fuselage attachment which, if not detected and corrected in a timely manner, could result in reduced structural integrity of the wing.
91-26-08: 91-26-08 AIRBUS INDUSTRIE: Amendment 39-8121. Docket No. 91-NM-87-AD. Applicability: Model A310 and A300-600 series airplanes, on which Modification 7769 has not been accomplished, certificated in any category. Compliance: Required within 180 days after the effective date of this AD, unless accomplished previously. To prevent loss of the Generator Control Units (GCU) and AC electrical power, accomplish the following: (a) Remove the three currently installed GCU's and replace them with three modified GCU's having protective covers, in accordance with Airbus Industrie Service Bulletins A310-24-2040, Revision 1, dated January 28, 1991 (for Model A310); A300-24-6029, Revision 1, dated February 22, 1991 (for Model A300-600); and Sundstrand Service Bulletin 735226/740206/740120-24-9 (for Models A310 and A300-600), dated June 15, 1989 (Modification 7769); as applicable. (b) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA Principal Avionics Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. (c) 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. (d) The installation required by this AD shall be done in accordance with Airbus Industrie Service Bulletin A310-24-2040, Revision 1, dated January 28, 1991 (for Model A310); Airbus Industrie Service Bulletin A300-24-6029, Revision 1, dated February 22, 1991 (for Model A300-600); and Sundstrand Service Bulletin 735226/740206/740120-24-9 (for Models A310 and A300-600), dated June 15, 1989 (Modification 7769). This incorporation by reference was approved by the Director of the Federal Registerin accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. (e) This amendment (39-8121, AD 91-26-08) becomes effective on February 14, 1992.
2013-17-07: We are adopting a new airworthiness directive (AD) for certain General Electric Company (GE) GE90-76B, -85B, -90B, -94B, -110B1, and - 115B turbofan engines. This AD was prompted by multiple reports of distress of certain stage 1 high-pressure turbine (HPT) stator shrouds due to accelerated corrosion and oxidation, including one engine in- flight shutdown (IFSD) caused by failure of the HPT stator shrouds. This AD requires initial and repetitive on-wing 360-degree borescope inspections (BSIs) for corrosion and oxidation of stage 1 HPT stator shrouds. If a shroud is found to be distressed, this AD requires reinspection at a reduced interval or removal from service before further flight. We are issuing this AD to prevent failure of stage 1 HPT stator shrouds, resulting in an IFSD of one or more engines, loss of thrust control, and damage to the airplane.
51-25-02: 51-25-02 MARTIN: Applies to All Models 202 and 202A Airplanes. Compliance required by January 1, 1952. Install cover over D7231-125 inverter remote reset circuit breaker terminals to prevent inadvertent shorting between "MAIN" and "CONTROL" terminals. Compliance may be made in same manner or equivalent to Martin Service Instruction Letter No. 8 dated September 10, 1951. This supersedes the first item in AD 51-11-01 requiring the redesign of the d.c. input circuit breaker mounting.
2013-18-07: We are superseding revised Airworthiness Directive (AD) 76-12- 07 for all Bell Model 204B and certain serial-numbered Model 205A-1 helicopters with a certain tail rotor pitch control chain (chain) installed. AD 76-12-07 required visually inspecting the chain to detect a crack in the link segments and, for affected Model 205A-1 helicopters, replacing the chain and cable control system with a push- pull control system. This new AD requires, for Bell Model 204B, inspecting certain chains at specified intervals, revising the inspection procedures, installing a tail rotor cable and chain damper kit (damper kit), and revising the maintenance manual or Instructions for Continued Airworthiness (ICAs) to include the inspection intervals. This new AD also requires, for certain Bell Model 205A-1 helicopters, replacing the chain and cable control system with an airworthy tail rotor push-pull control system kit. This AD was prompted by the rapid growth of a crack leading to premature chainfailure. The actions are intended to prevent failure of the chain, loss of tail rotor blade pitch control, and subsequent loss of control of the helicopter.
76-13-12: 76-13-12 BRITTEN-NORMAN LTD: Amendment 39-2660. Applies to Islander Models BN-2, BN-2A, BN-2A-2, BN-2A-3, BN-2A-6, BN-2A-8, BN-2A-9, BN-2A-20 and BN-2A-21 airplanes, certificated in all categories, which incorporate Britten-Norman Modification No. NB/M/430. Compliance is required within the next 100 hours time in service after the effective date of this AD, unless already accomplished. To prevent inadvertent retraction of the flaps beyond the normal 6 degree droop position, accomplish the following: (a) Perform a visual inspection of the flap retract position in accordance with the Inspection paragraph of Britten-Norman Service Bulletin No. BN-2/SB.66, dated August 3, 1973, or an FAA-approved equivalent. (b) If the flap retract position is found to be out of tolerance, before further flight, adjust the flap actuator in accordance with Rectification paragraphs 1, 2, 3, and 4 of Britten-Norman Service Bulletin No. BN-2/SB.66, dated August 3, 1973, or an FAA-approved equivalent. NOTE: Material L65, specified in Fig. 1 of Britten-Norman Service Bulletin No. BN-2/SB.66, dated August 3, 1973, is the same as Alum. Alloy U.S. QQA-261 (AL 14S). This amendment becomes effective July 15, 1976.
2013-13-01: We are adopting a new airworthiness directive (AD) for certain Piper Aircraft, Inc. Models PA-46-310P, PA-46-350P, PA-46R-350T, and PA-46-500TP airplanes. This AD requires inspecting the fuel vent valves to identify if the nitrile parts are installed and modifying and eventually replacing the fuel vent valves if the nitrile parts are installed. This AD was prompted by nitrile fuel vent valves not providing the correct ventilation. If not corrected, this unsafe condition may lead to structural damage of the wings, which could result in loss of control. We are issuing this AD to correct the unsafe condition on these products.
78-09-05: 78-09-05 CESSNA: Amendment 39-3202. Applies to the following models and serial number airplanes certificated in all categories: \n\n\n\nMODELS\nSERIAL NUMBERS\n336\n336-0001 through 336-0195 \n337, 337A, 337B, T337B,\t\n 337C, T337C, 337D and T337D\n337-0001 through 337-1193\n \n337E, T337E, 337F, T337F,\n 337G, 337H and T337H \n33701194 through 33701852\nT337G and P337H\nP3370001 through P3370313\nM337B\nAll serial numbers\n\n\tCompliance: Required as indicated in accordance with Compliance Table I set forth in this AD or as otherwise set forth herein, unless already accomplished. \n\n\tTo detect cracking of the wing front and rear spar lower caps, front spar web and web doubler - In accordance with instructions set forth herein and in Cessna Multi-Engine Service Letter ME78-2, dated February 13, 1978, or later revisions: \n\n\tI.\tAt time intervals noted in Table I of this AD, inspect the right and left wing front and rear spar lower cap and front spar web and web doubler onairplanes having model and serial numbers shown below; \n\n\n\n\n\n\n\n\n\nTABLE 1\nCOMPLIANCE TIMES\n\nAirplane Type/\nOperation\nTotal Time in\nService for each of the following:\n\nInspection Times for Spar\nCap - Inspections required\nby Paragraph I.A.\n\nInspection Times for \nFront Spar Web - Inspections\nRequired by Paragraph I.B. \n\n\n1) Front Spar\n Lower Cap\nInitial Inspection\nin accordance\nwith this AD \nInterval for\nRepetitive\nInspections\nInitial\nInspection\nInterval for\nRepetitive\nInspections*\n\n2) Rear Spar\n Lower Cap\n3) Front Spar\n Web & Web\n Doubler\n\n\n\n\n\n\n\n\n\n\nNon-Pressurized\n(See Note 1)\n0 to 4999\nNone\nNone\nNone\nNone \n\n5,000 & up\nWithin 25 hours time-in-service after the effective date\nof this AD for those airplanes which \nhave not yet been inspected in accordance with \nAD 76-10-11\nor 73-04-02 or;\nwithin 500 hours\ntime-in-service \nafter the last\ninspection in\naccordance with\nAD 76-10-11 or\n73-04-02.\n500 hoursWithin 25 hours time-in-service\nafter the effective date of this AD.\n\n500 hours\n\n\n\n\n\n\nPressurized\n(See Note 1)\n0 to 9,999\n\nNone\nNone\nNone\nNone \n\n10,000 & up\nWithin 25 hours\ntime-in-service\nafter the effective\ndate of this AD\nfor those airplanes\nwhich have not yet\nbeen inspected in\naccordance with \nAD 76-10-11 or\n73-04-02 or;\nwithin 500 hours\ntime-in-service \nafter the last inspections\nin accordance with\nAD 76-10-11 or 73-04-02.\n500 hours\nWithin 25 hours time-in-service\nafter the effective date of this AD.\n\n\n500 hours\n\n\n\n\n\nNOTE 1: \nFor those airplanes\nwhich have engaged\n in contour or\nterrain following\noperations at low\naltitudes, such \nas power/pipeline\npatrol, fish or\ngame spotting,\naerial applications,\npolice patrol, live-\nstock management,\netc., Cessna \nrecommends and FAA strongly urges\ninspections at \nintervals shown to\nthe right of this\nnote.\n0 to 2,999\n\n3,000 & up\nNone\n\nWithin 25 hours\ntime-in-service\nafterthe effective\ndate of this AD\nfor those airplanes which have\nnot yet been \ninspected in \naccordance with the\nsuggestion in \nAD 76-10-11 or\n73-04-02 or, within 300 hours time-\nin-service after the last inspection in accordance with\nAD 76-10-11 or\n73-04-02.\nNone\n\n300 hours\nNone\n\nWithin 25 hours time-in-service\nafter the effective date of this AD.\n\nNone \n\n300 hours\n\n\n*After initial inspection in\naccordance\nwith this AD the compliance\ntime for\nrepetitive\ninspections\nmay be adjusted to \nallow compliance at the same\ntime as the\ninspections\nrequired by\nparagraph\n1A of this AD.\n\n\t\tA.\tFront and rear spar lower cap inspection;\n\n\nModel\nSerial Numbers\n336\n336-0001 through 336-0195\n337, 337A, 337B, T337B, 337C,\n T337C, 337D, and T337D\n337-0001 through 337-1193\n337E, T337E, 337F, T337F, and 337G\n33701194 through 33701548\nT337G\nP3370001 through P3370138\nM337B\nAll serial numbers\n\n\t\t\t1.\tFront spar lower cap inspection. \n\t\t\t\n\t\t\t\ta.\tInspect three fastener holes on the left wing and three fastener holes on the right wing for wing spar cracks using eddy current inspection procedures outlined in the above noted Cessna Service Letter. Figure 1 shows the area involved and the three fasteners (two NAS 221 screws at W.S. 64.41 and the jack point bolt hole at W.S. 68.45) that are to be inspected. \n\t\t\t\n\t\t\t\tb.\tRemove the two NAS 221 screws at Wing Station 64.41 one at a time for this inspection and hold the boom fairing firmly against underlying thicknesses of material to insure proper eddy current probe depth settings during this inspection. A cross-section with the screws at W.S. 64.41 removed is shown in Figure 2. Figure 3 shows the spar assembly (less boom and wing skins) and the relationship of the lower cap to the other parts at the strut attachment. Figure 4 shows the spar cross-section thru the jack point bolt hole. \n\n\n\n\n\n\n\n\n\n\n\t\t\t\tc.\tIf cracks are found in either the right or left wing front spar lower cap during any inspection required by this AD, prior to approving the airplane for return to service replace the front spar lower cap in both the right and left wing with new spar caps. \n\n\t\t\t2.\tRear spar lower cap inspection. \n\n\t\t\t\ta.\tDye penetrant inspect the lower spar cap area between two and three inches (2nd rivet outboard of W.S. 66.00 rib) outboard of wing station 66.00 rib for spar cap cracks originating in the rivet hole in accordance with inspection provisions in the above noted Cessna Service Letter. Figures 5, 6 and 7 show the location of the area on the rear spar lower cap to be inspected. \n\n\n\n\n\n\n\n\n\n\t\t\t\tb.\tIf cracks are found in either the right or left wing rear spar lower cap during any inspection required by this AD, prior to approving the airplane for return to service replace the rear spar lower cap in both the right and left wing. \n\t\n\t\tB.\tFront spar web and web doubler inspection; \n\n\nModel\nSerial Numbers\n336\n336-0001 through 336-0195\n337, 337A, 337B,T337B, 337C,\n T337C, 337D and T337D\n337-0001 through 337-1193\n337E, T337E, 337F, T337F, 337G,\n 337H and T337H\n33701194 through 33701852\nT337G and P337H\nP3370001 through P3370313\nM337B\nAll serial numbers\n\n\t\t\t1.\tRemove wing root access panels and wing root fairings. \n\n\t\t\t2.\tVisually inspect the radii of both the spar web and web doubler for cracks in the shaded critical area shown in Figure 8 of this AD. \n\n\t\t\t3.\tIf cracks are found in either the right or left wing front spar web or web doubler during any inspection required by this AD, prior to approving the airplane for return to service, replace the discrepant components. \n\n\tII.\tAirplanes found to have cracked spar caps, webs or web doublers during inspections required by this AD may be flown in accordance with FAR 21.197 to a base where the component replacement can be accomplished. \n\n\tIII.\tAccomplish the repetitive inspections made mandatory by this AD on those spar caps, webs and web doublers replaced in accordance with this AD upon accumulation of the total times-in-service shown in Table I of this AD. Repetitive inspections "strongly urged" via Note 1 in Table I of this AD should also be accomplished on spar caps, webs and web doublers replaced in accordance with this AD upon accumulation of the total times-in-service shown in Note 1, Table I, of this AD. \n\n\tIV.\tNotify in writing the Chief, Engineering and Manufacturing Branch, FAA, Central Region, of the location and length of any crack found during inspections required by this AD and also the total time in service of the component at the time the crack was discovered. (Reporting approved by the Office of Management and Budget under OMB NO. 04-R0174.) \n\n\tV.\tThe time interval for repetitive inspections required by this AD, after compliance with initial inspection requirements, can be adjusted up to 25 hours to allow accomplishment of these inspections at regular scheduled maintenance periods. \n\n\tVI.\tEquivalent methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. \n\n\tThis AD supersedes AD 76-10-11, Amendment 39-2621 (41 FR 22045, 22046, 22047 and 22048). \n\n\tThis Amendment becomes effective May 11, 1978.
2010-14-20: The FAA is adopting a new airworthiness directive (AD) for McCauley Propeller Systems model 4HFR34C653/L106FA propellers. This AD requires a onetime fluorescent penetrant inspection (FPI) and eddy current inspection (ECI) of the propeller hub for cracks. This AD results from reports of 10 hubs found cracked during propeller overhaul. We are issuing this AD to prevent failure of the propeller hub, which could cause blade separation, damage to the airplane, and loss of control of the airplane.
81-14-03 R1: 81-14-03 R1 BRITISH AEROSPACE: Amendment 39-4151 as amended by Amendment 39- 4241. Applies to British Aerospace (formerly Armstrong Whitworth) Model AW-650 series 100 and series 200 aircraft, serial numbers 6651, 6652, 6653, 6654, 6656, 6660, 6801, 6802, 6083, and 6805, certified in all categories. Compliance required as indicated. To prevent possible loss of the vertical fin, accomplish the following: Within the next ten days after the effective date of this AD, unless previously accomplished, visually inspect the tail boom frame 28 on the left and right booms for cracks in accordance with paragraph 2C(1) of British Aerospace Service Bulletin Number 55/39 dated June 1981 or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA, Northwest Mountain Region. A. If no cracks are found, reinspect at intervals not to exceed 75 flight hours after the initial inspection by visual and eddy current methods in accordance with paragraphs 2C(1) and 2C(2)of British Aerospace Service Bulletin Number 55/39 dated June 1981 or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA, Northwest Mountain Region. B. Within 150 flight hours from the time of the initial visual inspection and thereafter at intervals not to exceed 150 flight hours, X-ray inspect in accordance with paragraph 2C(3) of British Aerospace Service Bulletin Number 55/39 dated June 1981 or in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA, Northwest Mountain Region. C. If cracks are found, evaluate the damage in accordance with paragraph 2E of British Aerospace Service Bulletin Number 55/39 dated June 1981. If the airplane can continue to fly, reinspect in accordance with the inspection intervals contained in paragraph A and B of this ZD. If the cracks are such that the airplane cannot continue to fly after evaluation of the damage in accordance with paragraph 2E of the service bulletin, the modification as described in paragraph 2F of the service bulletin must be made prior to further flight. D. Prior to the accumulation of 600 flight hours after the initial visual inspection or within 50 flight hours after the effective date of this Amendment, whichever comes later, the modification described in paragraph 2F of British Aerospace Service Bulletin 55/39 dated June 1981 must be accomplished. Upon the accomplishment of the modification, no further AD inspections are required. E. Airplanes may be flown in accordance with FAR 21.197 and 21.199 to a maintenance base for accomplishment of the inspection required by this AD. F. Alternate methods of compliance may be used which provide an equivalent level of safety 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. Amendment 39-4151 became effective July 8, 1981. This amendment 39-4241 becomes effective November 12, 1981.
2010-14-08: We are adopting a new airworthiness directive (AD) for certain Model 747-400, 747-400D, and 747-400F series airplanes. For all airplanes, this AD requires installing new pump control and time delay relays, doing related investigative and corrective actions if necessary, and changing the wiring for the center and main fuel tank override/jettison fuel pumps; and, for certain airplanes, installing new relays and wiring for the horizontal stabilizer override/jettison fuel pumps. This AD also requires a revision to the maintenance program to incorporate Airworthiness Limitation No. 28-AWL-24 and No. 28-AWL- 26. For certain airplanes, this AD also requires installing an automatic shutoff system for the horizontal stabilizer tank fuel pumps and installing new integrated display system software. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent uncommanded operation of certain override/jettison pumps which could cause overheat, electrical arcs, or frictional sparks, and could lead to an ignition source inside a fuel tank. This condition, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
99-07-10: This amendment adopts a new airworthiness directive (AD) that applies to all Industrie Aeronautiche e Meccaniche (I.A.M.) Model Piaggio P-180 airplanes. This AD requires inspecting the upper and lower engine nacelle inner panels for any loose or partially detached inner film, and removing any loose or partially detached inner film. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Italy. The actions specified by this AD are intended to prevent the accumulation of loose particles on the engine inlet screen caused by film delamination, which could result in reduced engine power and possible loss of airplane control.
60-18-02: 60-18-02 VICKERS: Amdt. 197 Part 507 Federal Register August 24, 1960. Applies To All 745D and 810 Series Aircraft. Compliance required as indicated. Conduct inspections of the brake accumulator systems as specified in Vickers Preliminary Technical Leaflet (PTL) 222 (700 Series) and PTL 87 (800/810 Series) within the next 300 hours' time in service and at subsequent periodic intervals of 800 hours' time in service. These inspections are not mandatory when filters, Dunlop ACM 18308 or equivalent, are installed in accordance with Vickers Modification Bulletins D.2994 (700 Series) and FG.1796 (800/810 Series). This amendment shall become effective 30 days after date of its publication in the Federal Register.
99-07-05: This amendment adopts a new airworthiness directive (AD), applicable to all Lockheed Model L-1011-385 series airplanes, that requires repetitive external visual inspections and internal borescope inspections to detect discrepancies of the elevator assembly; and either repair or repair/modification of certain identified discrepancies. This amendment is prompted by a report of fretting at the diagonal truss to web joint of the elevator and cracking in the cap fillet radius adjacent to the joint, apparently due to loose fasteners as a result of local vibration. The actions specified by this AD are intended to detect and correct such fretting and cracking, which could result in reduced structural integrity of the elevator and consequent flutter instability if coupled with other structural failures.