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85-17-04: 85-17-04 EMPRESA BRASILEIRA DE AERONAUTICA S.A. (EMBRAER): Amendment 39-5126. Applies to Models EMB-110P1 and EMB-110P2 (all serial numbers) airplanes certificated in any category which have aluminum elevator control rod tubes installed. Compliance: Required as indicated, unless already accomplished. To prevent failure of the elevator control rod tube, accomplish the following: (a) Within the next 50 hours time-in-service after the effective date of this AD, visually inspect the elevator control rod tubes, P/N 4A-500-10-09-01, for evidence of corrosion or cracks. If corrosion or cracks are found, prior to further flight replace the control rod tube in accordance with Embraer Service Bulletin (S/B) No. 110-27-076, Revision 01, dated July 2, 1985. (b) Within 150 hours time-in-service or 30 (thirty) days, whichever occurs first, after the effective date of this AD, replace both left and right elevator aluminum control rod tubes P/N 4A-500-10-09-01 with steel control rod tubes P/N 110-500-10-00-04-01. Reidentify the elevator control rod assembly with the new P/N 110-500-10-00-09. (c) Airplanes may be flown in accordance with Federal Aviation Regulation 21.197 to a location where the AD may be accomplished. (d) An equivalent method of compliance with this AD may be used if approved by the Manager, Atlanta Aircraft Certification Office, FAA, 1075 Inner Loop Road, College Park, Georgia 30337; Telephone (404) 763-7428. All persons affected by this directive may obtain copies of the documents referred to herein upon request to Embraer, Post Office Box 343 - CEP 12.200 Sao Jose Dos Campos, Sao Paulo, Brazil, or FAA, Office of Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on August 30, 1985.
81-11-01: 81-11-01 GOVERNMENT AIRCRAFT FACTORIES (GAF): Amendment 39-4302. Applies to Nomad Models N22 and N24 series airplanes certificated in any category. COMPLIANCE: Required as indicated unless already accomplished. To assure the emergency exit can be opened when the emergency exit door handle is placed in the open position, within the next 10 hours time-in-service after the effective date of this AD, accomplish the following: A) Visually inspect the emergency exit door for proper operation by rotating the emergency exit door handle (interior) to the open position and open the door by physically pushing outward at the bottom of the exit door. It is only necessary to move the door 0.5 inches (12 millimeters) outward to assure proper emergency door operation and that the door latch pins are fully withdrawn. Further movement of the door beyond 0.5 inches could fracture the door trim panel at the lower edge. NOTE: Operation of the door-locking mechanism should not require undue force in complying with this paragraph. B) If the emergency exit door opens properly in accordance with paragraph A) of this AD, secure the door in the closed position and record compliance of this AD with an appropriate entry in the airplane maintenance records. C) If the emergency exit door handle will not rotate or the door will not open in accordance with paragraph A) of this AD, accomplish the following: 1. Remove the inner trim panel from the door in accordance with the manufacturer' s Maintenance Manual No. 25-20-00, page 201, paragraph 1B(3). 2. Visually inspect the door-locking mechanism, including linkage to door locking pins for proper assembly and also assure the door lock pins are free of corrosion and are not seized in the door frame or fuselage by performing the following as applicable: a) Proper assembly of the door lock assembly is assured if the forward door lock pin linkage rod is almost horizontal while the rear rod is angledupwards, thus as viewed from the inside, a clockwise rotation (towards open) of the exit door handle will withdraw the door locking pins. If the linkage is incorrectly assembled, refer to the manufacturer's Illustrated Parts Catalog No. 52-20-01, Figure 1, and: i. Disconnect both operating (linkage) rods (item 8) from the lever assembly (item12), by removing cotter pin, washer and connecting pins (items 9, 10 and 11). ii. Rotate the door lever (handle) assembly (item 12) so that the forward end is sloping upwards at approximately 45 degrees and reconnect the linkage rods using new cotter pins. iii. Rotate the door lever (handle) clockwise (open position) and confirm that the door lock pins can be withdrawn from the door frame and lubricate the locking pins with MIL-G-21164 grease or equivalent. b) If the door locking pins are corroded or seized in the door frame or fuselage, free the pins, remove any corrosion and lubricate the pins with MIL-G-21164 grease or equivalent. c) Reassemble the door and reinstall the trim panels. D) If any part of paragraph C) of this AD was completed, reaccomplish paragraph A) of this AD to assure the emergency exit opens properly and record compliance of this AD with an appropriate entry in the airplane maintenance records. E) Aircraft may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. F) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, P.O. Box 50109, Honolulu, Hawaii 96850. GAF-Nomad Telex Alert Service Bulletin ANMD-52-2 covers the subject matter of this AD. This amendment becomes effective on January 28, 1982, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated May 13, 1981, and identified as AD 81-11-01.
2006-13-03: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 757 airplanes. This AD requires, for certain airplanes, reworking the spar bonding path and reapplying sealant; and, for certain other airplanes, testing the electrical bond between the engine fuel feed hose and the wing front spar and, if applicable, reworking the spar bonding path and reapplying sealant. This AD also requires, for all airplanes, an inspection to ensure the electrical bonding jumper is installed between the engine fuel feed tube and the adjacent wing station. This AD also requires operators that may have installed an incorrect O-ring to install the correct part and do a re- test. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent arcing or sparking at the interface between the bulkhead fittings of the engine fuel feed tube and the front spar during a lightning strike, which could provide a possible ignition source for the fuelvapor inside the fuel tank and result in a fuel tank explosion.
97-19-10: This amendment supersedes an existing airworthiness directive (AD), applicable to Sikorsky Aircraft-manufactured Model S-64E and S-64F helicopters, that currently requires initial and repetitive inspections of the main gearbox assembly second stage lower planetary plate (plate) for cracks, and removal and replacement of the plate if cracks are found; and daily inspections of certain main transmission oil filter packs for magnesium chips, and removal and replacement of the main transmission if chips are found. The AD also provides for reworking and re-identifying the plate, as well as establishes a retirement life for the plate, including those that have been reworked and re-identified. This amendment requires, for Model S-64E helicopters, inspections and rework of the plate and establishes a new retirement life for the plate. This amendment is prompted by the type certificate holder's reports that four plates were discovered to have cracks, three of which had been reworkedin accordance with the existing AD. The actions specified by this AD are intended to add another plate to the applicability of the AD, remove the requirements of AD 77-20-01 for the Model S-64F and prevent failure of the plate on the Model S-64E due to fatigue cracking, which could lead to failure of the main gearbox and subsequent loss of control of the helicopter.
84-25-01: 84-25-01 SIKORSKY AIRCRAFT DIVISION: Amendment 39-4961. Applies to all Sikorsky Model S-61 series helicopters, certified in all categories, equipped with Part Number (P/N) S6135-20608-1 main transmission power input spur gears. Compliance is required as indicated (unless already accomplished). To intercept pending failure of either spur gear, P/N S6135-20608-1, in the main gearbox, accomplish the following: (a) On all helicopters equipped with a functioning cockpit main gearbox chip detecting system, visually inspect the main gearbox oil scavenge screen for steel particles within the next 10 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 10 hours time in service from the last inspection. (b) On all helicopters not equipped with a cockpit main gearbox chip detecting system, visually inspect both the main gearbox chip detector and the oil scavenge screen for steel particles within the next 10 hours time in service after the effective date of this AD and thereafter at intervals not to exceed 10 hours time in service from the last inspection. (c) Any magnetic steel particles found by inspecting either the chip detector or the scavenge screen will require replacement of the main gearbox before further flight unless the particles are (1) fine hairline particles, or (2) confirmed as originating from components other than the spur gear. (d) If the inspections of paragraph (c) are inconclusive, the gearbox must be replaced or, at the option of the operator, the following maintenance tests may be conducted to evaluate the condition of the spur gear. (1) Drain and refill the gearbox with an approved lubricant. (2) Operate the helicopter at a safe height (below 5 feet), at a gross weight not less than 16,000 pounds, with a nominal neutral center of gravity, at 100 percent main rotor speed for one-half hour with one engine at 100 percent torque and the remaining engine at a torquelevel required to sustain safe hover. Repeat the procedure for one-half hour with the other engine at 100 percent torque and the remaining engine at a torque level required to sustain safe hover. (3) Inspect the gearbox scavenge screen and the magnetic chip detector and apply the inspection criteria of paragraph (c) and, if appropriate, reconduct the maintenance tests of paragraph (d) one time. Apply the inspection criteria of paragraph (c). NOTE: This procedure is not intended to authorize continued operation of the helicopter if any questionable safety condition is exposed by debris found when conducting these checks. (e) Replace spur gear P/N S6135-20608-1 with spur gear P/N S6135-20608-3 and comply with Sikorsky Overhaul and Repair Instruction 6135-342, Revision A, or later revision, or FAA-approved equivalent before further flight after December 30, 1986. The inspections of paragraphs (a), (b), and (c) may be discontinued for helicopters modified as required bythis paragraph. (f) Upon request, with substantiating data submitted through an FAA maintenance inspector, equivalent methods of compliance, adjustment in the inspection intervals, and adjustment in the replacement date may be approved by the Manager, Boston Aircraft Certification Office, FAA, New England Region. This amendment supersedes AD 83-17-04, Amendment 39-4706. This amendment becomes effective on January 19, 1985.
97-19-05: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes. This amendment requires replacement of the ignition exciter in the auxiliary power unit (APU) with a part that is designed to operate better in cold weather. This amendment is prompted by two occurrences of the APU failing to start after flight in cold soak conditions. The actions specified by this AD are intended to prevent such APU failure, which could result in the inability of the APU to restart the engines in the event both engines quit operating during flight.
92-16-05: 92-16-05 SAAB-SCANIA: Amendment 39-8314. Docket No. 92-NM-54-AD. Applicability: Model SAAB 340B series airplanes; serial numbers 240 through 299, inclusive; airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent the accumulation of ice and subsequent reduced controllability of the airplane and stall margins, accomplish the following: (a) Within 60 days after the effective date of this AD, modify the stabilizer de-icer boot system in accordance with SAAB-SCANIA Service Bulletin 340-30-039, dated December 16, 1991. (b) 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, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch. (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 modification shall be done in accordance with SAAB-SCANIA Service Bulletin 340-30-039, dated December 16, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from SAAB-SCANIA AB, Product Support, S-581.88, Linkoping, Sweden. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. (e) This amendment becomes effective on September 8, 1992.
62-27-03: 62-27-03 CURTISS-WRIGHT: Amdt. 525 Part 507 Federal Register December 28, 1962. Applies to all C-46 Series aircraft which have rudder assemblies P/N 20-140-5110 installed. Compliance required as indicated. A recent inflight failure of the rudder spring tab rod terminal resulted in the free flutter of the rudder spring tab which forced the rudder to oscillate. To preclude the failure of this rod terminal and the resulting loss of control of the aircraft, the following shall be accomplished: (a) (1) Aircraft with less than 10,000 hours' time in service shall be inspected in accordance with (b) prior to the accumulation of 10,200 hours' time in service and thereafter every 400 hours' time in service. (2) Aircraft with 10,000 or more hours' time in service shall be inspected in accordance with (b) within 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 200 hours' time in service, and thereafter every 400hours' time in service from the last inspection. (b) (1) Remove the clevis, P/N 20-530-5133-2, and AN 316-5R nut from the rudder spring tab rod assembly, P/N 20-530-5134-1. (2) Inspect for cracks in the shoulder and threaded shank of the rod terminal, P/N 20-530-5132-2, using a dye penetrant, magnetic particle, or an FAA approved equivalent inspection method in conjunction with at least a 5-power magnifying glass. (3) Replace cracked rod terminals, P/N 20-530-5132-2, prior to further flight. (c) Rod terminals, P/N 20-530-5132-2, or rod assembly, P/N 20-530-5134-1, having zero or a known time in service not exceeding 5,000 hours, may be installed in compliance with (b)(3). Such parts shall be inspected and replaced in accordance with the following: (1) Parts having less than 4,800 hours' time in service shall be inspected in accordance with paragraphs (b)(1) and (2) prior to the accumulation of 5,000 hours and thereafter every 400 hours' time in service.(2) Parts having 4,800 or more hours' time in service shall be inspected in accordance with paragraphs (b)(1) and (2) within the next 200 hours' time in service and thereafter every 400 hours' time in service. (3) Cracked parts shall be replaced prior to further flight. (d) The inspection required by paragraph (b) is no longer required when rod assembly, P/N 20530-5134-1, is replaced with a new Component Air, Inc., rod assembly, P/N CAI-46003-1, or with an FAA approved rod terminal and clevis, or a rod assembly incorporating a clevis with 3/8-24 UNF or 7/16-20 UNF threads. (e) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southern Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. This directive effective January 29, 1963.
2012-03-09: We are adopting a new airworthiness directive (AD) for all Model 747SP series airplanes. This AD was prompted by a report of a rudder hard-over event on a Model 747-400 series airplane, caused by a rudder power control module (PCM) manifold cracking and separating in the area of the yaw damper cavity end-cap. This condition could result in a hard-over of the rudder surface leading to an increase in pilot workload and a possible high-speed runway excursion upon landing, in the event of failure of the lower or upper rudder PCM manifold. This AD requires replacing or modifying the upper and lower rudder PCMs. We are issuing this AD to correct the unsafe condition on these products.
97-19-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires a one-time inspection of the hydraulic tubes and electrical harness wires of the wing rear access door for chafing, leakage, or wear damage; repair of any discrepancy found; and modification of the wing rear access door. This amendment is prompted by reports of interference between the wing rear access door and the hydraulic tubes and electrical harnesses, and chafing damage to the hydraulic tubes. The actions specified by this AD are intended to prevent such interference or chafing damage, which could lead to failure of the number 2 hydraulic system or loss of certain electrical and landing systems, and resultant reduced controllability of the airplane.