Results
2000-25-07: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Boeing Model 737-100, -200, and -200C series airplanes. This action requires repetitive inspections of the aft end of each inboard flap track of the wing outboard flap, and corrective actions, if necessary. This action is necessary to detect and correct damage of the aft end of each flap track, which could result in loss of the outboard trailing edge flap and consequent loss of controllability of the airplane.
88-04-10: 88-04-10 DEHAVILLAND AIRCRAFT COMPANY OF CANADA, A DIVISION OF BOEING OF CANADA, LTD.: Amendment 39-5855. Applies to Model DHC-7 series airplanes, equipped with Modification No. 7/2414, certificated in any category. Compliance required as indicated, unless previously accomplished. To preclude the possibility of heat shield separation resulting from the failure of aluminum alloy washers, accomplish the following: A. Within 60 days or 500 flight hours, whichever occurs first after the effective date of this AD, replace aluminum alloy washers with stainless steel washers, in accordance with the Accomplishment Instructions of de Havilland DHC-7 Service Bulletin No. 7-57-29, dated August 1, 1986. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the 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 The de Havilland Aircraft Company of Canada, A Division of Boeing of Canada, Ltd., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This amendment becomes effective April 5, 1988.
2000-25-06: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Pratt & Whitney (PW) PW4000 turbofan engines with the current design low pressure turbine (LPT) 4th stage air seal installed. This action requires, based on engine model, replacement of the current design seal with a new design seal, or with a modified seal. This amendment is prompted by reports of cracks in LPT 4th stage air seals. The actions specified by this AD are intended to reduce stresses that could lead to LPT 4th stage air seal cracking, resulting in seal fracture, uncontained engine failure, and damage to the airplane.
2016-06-01: We are superseding an airworthiness directive (AD) 2007-06-06 for B-N Group Ltd. Models BN-2, BN-2A, BN-2A-2, BN-2A-3, BN-2A-6, BN- 2A-8, BN-2A-9, BN-2A-20, BN-2A-21, BN-2A-26, BN-2A-27, BN-2B-20, BN-2B- 21, BN-2B-26, BN-2B-27, BN2A MK. III, BN2A MK. III-2, BN2A MK. III-3 BN2A, BN2B, and BN2A MKIII (all models on TCDS A17EU and A29EU) airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as cracks in the inner shell of certain pitot/static pressure heads. We are issuing this AD to require actions to address the unsafe condition on these products.
52-17-02: 52-17-02 CURTISS-WRIGHT: Applies to all Model C-46 Series aircraft. Compliance required as noted. As a result of a number of failures the following precautionary measures must be taken: 1. The original compliance date set was not later than December 1, 1952, and at each No. 3 inspection thereafter, however, in order to convert the inspection interval to time in service, the following inspection shall be accomplished within 400 hours' time in service after the effective date of this amendment. Subsequent compliance required at each 400 hours' time in service after initial compliance. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southwest 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. Inspect the tail wheel shimmy damper support, P/N 20-360-3108-4, for cracks or damage which might lead to subsequent failure. Particular attention should be paid to the radius formed by the intersection of the vertical and lower horizontal surface on the forward side of the part. A general inspection of the shimmy damper assembly should be made and it should be determined that it is properly adjusted. All cracked or damaged parts must be replaced. The following method of adjustment of the shimmy damper as outlined in Air Force Technical Order AN 01-25LA-2, section IV, paragraph 8 (f) is quoted below for your convenience: (a) Position the shimmy damper on the mounting bracket and attach the two 3/8-inch bolts, nuts, and washers, using a 9/16-inch open-end wrench in conjunction with a 9/16- inch socket and a ratchet. (b) Before installing the link assembly to the shimmy damper arm, the arm must be positioned to allow for maximum travel of the unit in actual operation. This is accomplished by slowly actuating the arm toward the rear of the airplane until the movement stops. From this point reverse the action until the arm has made an arc of 67 degrees. It is now in a neutral position and will assure correct operation when completely installed. (c) With the arm in neutral and making sure not to move it from this position, attach the link assembly to it with the 5/16-inch bolt, nut, washer and cotter pin using a 1/2-inch open-end wrench, a 1/2-inch socket, and a ratchet. (d) Make sure that the tail wheel is locked in position. The pin on the centering disc will then be properly set for attaching the eyebolt of the link assembly. If the eyebolt hole does not match with the pin on the centering disc when the tail wheel is locked and the actuating arm is in neutral, adjustment must be made by loosening the locknut on the eyebolt and turning the eyebolt until the proper length is obtained. Be sure to tighten the locknut again.(e) Secure the link assembly to the centering disc by installing the 1 1/8-inch castellated nut and washer drawing up the nut with a crescent wrench. Insert a cotter pin on this nut. 2. Compliance required as noted. Tail wheel locks for the subject model airplane have been manufactured which do not comply with the material specifications and in some cases physical dimensions of the approved drawings. The approved tail wheel lock P/N 20-360-1033, is a forged steel part of 8740 or N-S- 16 material heat treated to 150,000 pounds per square inch, with a Rockwell harness of C-33 to C-38; however, a cast steel lock complying with the physical dimensions, heat treat and hardness of the approved drawing will also be considered acceptable. The original compliance date for the following inspection was March 1, 1953; however, in view of a recent accident involving the installation of an unapproved tail wheel lock, it must be ascertained prior to February 15, 1956, that all tailwheel locks meet the stated specifications of an acceptable part. Any tail wheel locks which do not comply with the approved drawing, No. 20-360-1033, in regard to material, heat treat or physical dimensions as mentioned above, must be replaced with an approved part. It should be noted that some unsatisfactory parts in circulation bear the correct part number, therefore this fact cannot be considered a satisfactory means of determining that an approved part is installed in the airplane. Revised October 23, 1961. Revised November 23, 1961.
93-10-08: 93-10-08 DE HAVILLAND: Amendment 39-8588. Docket No. 91-CE-49-AD. Supersedes AD 73-19-06, Amendments 39-1800 and 39-1710. Applicability: Models DHC-6-1, DHC-6-100, DHC-6-200, and DHC-6-300 airplanes (serial numbers 1 through 420), certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent failure of the overhead control console, which could result in loss of control of the airplane, accomplish the following: (a) Within the next 25 hours time-in-service (TIS) after the effective date of this AD, unless already accomplished within the last 175 hours TIS (compliance with AD 73-19-06), visually inspect the overhead console control quadrants for cracks in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland Service Bulletin (SB) 6/298, Revision D, dated December 20, 1991. (1) If cracks are found that meet or exceed the requirements specified in paragraph 3.1 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, prior to further flight, replace the overhead control console quadrants, part number (P/N) C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298. (2) If cracks are found that meet or exceed the requirements specified in paragraph 3.2 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, within the next 200 hours TIS, replace the overhead control console quadrants, P/N C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298. (3) If cracks are found that are less than that specified in paragraph 3.2 of the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991, reinspect at intervals not to exceed 100 hours TIS until the modification specified in paragraph (b) of this AD is accomplished. (4) If no cracks are found, reinspect at intervals not to exceed 200 hours TIS until the modification specified in paragraph (b) of this AD is accomplished. (b) Within the next 2,400 hours TIS after the effective date of this AD, unless already accomplished as specified in paragraph (a)(1) or (a)(2) of this AD, replace the overhead control console quadrants, P/N C6-CE-1010, with P/N C6CE1421-27 in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of de Havilland SB No. 6/298, Revision D, dated December 20, 1991. (c) The installation of new overhead control console quadrants (Modification No. 6/1467) as specified in paragraphs (a)(1), (a)(2), and (b) of this AD is considered terminating action for the repetitive inspection requirement of this AD. (d) 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. (e) An alternative method of compliance or adjustment of the compliance times that provides an equivalent level of safety may be approved by the Manager, New York Aircraft Certification Office, 181 Franklin Avenue, Room 202, Valley Stream, New York 11581. The request shall be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, New York Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York Aircraft Certification Office. (f) The inspections and modification required by this AD shall be done in accordance with de Havilland Service Bulletin (SB) 6/298, Revision D, dated December 20, 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 de Havilland, Inc., 123 Garratt Boulevard, Downsview, Ontario, Canada, M3K 1Y5. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 800 North Capitol Street, NW, suite 700, Washington, DC. (g) This amendment (39-8588) supersedes AD 73-19-06, Amendments 39-1800 and 39-1710. (h) This amendment becomes effective on July 16, 1993.
2000-24-26: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Rolls-Royce plc RB211 Trent 800 series turbofan engines, that currently requires initial and repetitive ultrasonic inspections of fan blade roots for cracks, and replacement, if necessary, with serviceable parts. This amendment requires the reduction of the initial cyclic compliance threshold and repetitive inspection intervals. This amendment also allows inspections to be accomplished within 100 cycles-in-service if the initial or repetitive thresholds are exceeded on the effective date of this AD. This amendment is prompted by an improved understanding of the crack propagation mechanism and the latest service operational data. The actions specified by this AD are intended to detect and prevent fan blade failure, which could result in multiple fan blade releases, uncontained engine failure, and possible damage to the airplane.
2016-05-09: We are adopting a new airworthiness directive (AD) for certain MDHI Model 369A (Army OH-6A), 369H, 369HE, 369HM, 369HS, 369D, 369E, 369F, 369FF, and 500N helicopters. This AD requires inspecting the auxiliary fuel pump (fuel pump) wire routing in the left-hand fuel cell and corrective action, if necessary. This AD also requires installing a warning decal on the left-hand fuel cell access cover. This AD was prompted by accidents resulting from incorrectly positioned fuel pump wiring within the fuel tank interfering with the operation of the fuel quantity sensor float, which caused an erroneous fuel quantity indication in the cockpit. The actions are intended to detect and correct routing of the fuel pump wiring to prevent interference with the fuel quantity sensor float, an erroneous fuel quantity indication in the cockpit, and subsequent fuel exhaustion and emergency landing.
2009-03-03: We are adopting a new airworthiness directive (AD) for certain McDonnell Douglas airplanes listed above. This AD requires installing a dam assembly for the container of the fuel boost pump of the center tank located in the right main tank, and doing the related investigative actions, and corrective actions if necessary. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent the center tank fuel boost pump from operating in a fuel vapor zone and becoming a potential ignition source in the right main tank, potentially resulting in a fuel tank explosion and consequent loss of the airplane.
73-20-07 R2: 73-20-07 R2 BEECH: Amendment 39-1728 as amended by Amendment 39-3732 is further amended by Amendment 39-4461. Applies to Models A23-19, 19A, M19A, B-19 (Serial Numbers MB-1 through MB-708); Models 23, A23, A23A, B23, C23 (Serial Numbers M-1 through M-1576); Models A23-24, A24 (Serial Numbers MA-1 through MA-368); and Models A24R, B24R (Serial Numbers MC-2 through MC-282) certified in all categories. Compliance: Required as indicated. To detect cracks or other structural damage to the forward wing attach point brackets (P/N 169-400013-3 and -5(LH) and 169-400013-4 and -6 (RH)), within the next 25 hours' time in service after the effective date of this AD unless previously accomplished on airplanes having more than 100 hours' time in service and thereafter at each normal annual, progressive or 100 hour inspection interval as required by FAR 91.169, accomplish the following: (a) Remove the seats and sidepanels and visually inspect the forward wing attach brackets, P/N169-400013-3 and -5 (LH) and 169-400013-4 and -6 (RH), to determine if cracks or other structural damage exists in the area around the attach bolt hole. This inspection is to be performed in accordance with Beech Service Instructions 0042-031, Rev. II. (b) If as a result of any inspection required herein cracks or other structural damage is discovered, prior to further flight repair the forward wing attach brackets in accordance with Beechcraft Service Instructions 0042-031, or later FAA-approved revisions, or any other repair approved by the Chief, Wichita Aircraft Certification Office, FAA, Wichita, Kansas. To accomplish the repair required herein the aircraft may be flown in accordance with FAR 21.197 to a base where the repair may be performed. (c) The inspection intervals required in this AD may be adjusted 15 plus or minus hours where required to fit users maintenance cycles if authorized by an FAA Flight Standards Inspector. (d) When airplanes have been modified by installing all of the new wing attach structural parts listed under materials in Beechcraft Service Instructions No. 0042-031, Rev. II, further compliance with this AD is not required. (e) Any equivalent method of compliance with this AD must be approved by the Chief, Wichita Aircraft Certification Office, Federal Aviation Administration, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. Amendment 39-1728 became effective October 5, 1973. Amendment 39-3732 became effective March 21, 1980. This Amendment 39-4461 became effective on September 3, 1982.
92-16-19: 92-16-19 LOCKHEED AERONAUTICAL SYSTEMS COMPANY: Amendment 39-8329. Docket No. 92-NM-125-AD. Supersedes AD 92-10-51, Amendment 39-8271. Applicability: All Model L-1011-385 series airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent degradation of airplane pitch control, accomplish the following: (a) Prior to the accumulation of 12,000 landings, or within 3 days after June 24, 1992 (the effective date of AD 92-10-51, Amendment 39-8271), whichever occurs later, accomplish the procedures specified in paragraphs (a)(1), (a)(2), and (a)(3) of this AD: (1) Gain access to the lower end of the stabilizer hydraulic actuators (four per airplane) where they attach to the front spar of the horizontal stabilizer center box structure at Fuselage Station FS 1875. (2) Inspect for missing, sheared, or deformed stabilizer lower actuator attach pins, part number 1563117-101 (one per actuator). (3)If any pin is missing, sheared, or deformed, replace the pin prior to further flight. (b) At the applicable time specified in paragraph (b)(1), (b)(2), or (b)(3) of this AD, either perform a one-time magnetic particle inspection to detect cracks on each of the four lower horizontal stabilizer actuator pins (part number 1563117-101), or replace each of the four actuator pins (part number 1563117-101) with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. (1) For airplanes that have accumulated less than 12,000 landings as of the effective date of this AD: Prior to the accumulation of 12,000 landings, or within 60 days after the effective date of this AD, whichever occurs later. (2) For airplanes that have accumulated at least 12,000 landings but not more than 20,000 landings as of the effective date of this AD: Within 45 days after the effective date of this AD. (3) For airplanes that have accumulated more than20,000 landings as of the effective date of this AD: Within 30 days after the effective date of this AD. (c) If a magnetic particle inspection is performed in accordance with paragraph (b) of this AD, accomplish the procedures specified in paragraph (c)(1) or (c)(2), as applicable: (1) If any actuator pin is found to be cracked: Prior to further flight, replace the cracked pin with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, part number 1563117-101, it must be replaced with a new pin. (2) If no actuator pins are found to be cracked: Prior to the accumulation of 1,000 landings after accomplishing the magnetic particle inspection, replace the pin with a new pin, in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992. Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, part number 1563117-101, it must be replaced with a new pin. (d) If each of the four actuator pins, part number 1563117-101, are replaced with a new pin in accordance with paragraph (b) of this AD: Thereafter, prior to the accumulation of 12,000 landings on each actuator pin, it must be replaced with a new pin. (e) 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, Atlanta Aircraft Certification Office (ACO), FAA, Small 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, Atlanta ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Atlanta ACO. (f) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a locationwhere the requirements of this AD can be accomplished. (g) The inspection and replacement shall be done in accordance with Lockheed Service Bulletin 093-27-304, Revision 1, dated May 28, 1992, which contains the following list of effective pages: Page Number Revision Level Date 1 - 8 1 May 28, 1992 9 - 11 Original May 15, 1992 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 Lockheed Western Export Company (LWEC), Dept. 693, Zone 0755, 86 South Cobb Drive, Marietta, Georgia 30063. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the FAA, Atlanta Aircraft Certification Office, Suite 210C, 1669 Phoenix Parkway, Atlanta, Georgia; or at the Office of the Federal Register, 800 North Capitol Street NW., 7th Floor, Suite 700, Washington, DC. (h) This amendment becomes effectiveon September 1, 1992.
2000-24-24: This amendment adopts a new airworthiness directive (AD) that is applicable to Rolls- Royce plc RB211 Trent 768-60, Trent 772-60, and Trent 772B-60 series turbofan engines having common nozzle assembly part number (P/N) FK16544 or FK16558. This action requires initial and repetitive visual inspections of the inner and outer skins of the common nozzle assembly and specifies allowable limits for cracks, loose rivets, and missing rivets. This action also requires repair if the common nozzle assembly damage exceeds allowable limits. This amendment is prompted by two reports of in-flight inner skin detachment. The actions specified in this AD are intended to detect cracks, loose rivets, and missing rivets, which could result in inner skin detachment, release of common nozzle assembly debris from the engine, and possible damage to the airplane control surfaces.
89-03-06: 89-03-06 FOKKER: Amendment 39-6122. Applicability: Model F-27 series airplanes, Serial Numbers 10623 through 10692, in Mark 500 configuration, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent loss of electrical power and eliminate a potential fire hazard, due to chafing wire bundles, accomplish the following: A. Within 60 days after the effective date of this AD, perform a one-time inspection of the upper fuselage cableloom for chafing and insufficient clamping, and repair, if necessary, in accordance with Fokker Service Bulletin F27/24-77, dated June 7, 1988. B. 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, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who may add any comments 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 airplanes to a base for the accomplishment of inspections and/or modifications required by 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 Fokker Aircraft USA, Inc., 1199 N. Fairfax Street, Alexandria, Virginia 22314. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6122, AD 89-03-06) becomes effective March 6, 1989.
2022-23-10: The FAA is superseding Airworthiness Directive (AD) 2021-06- 03, which applied to all The Boeing Company Model 777F series airplanes. AD 2021-06-03 required deactivating the potable water system. This AD was prompted by a report of a water supply line that detached at a certain joint located above an electronic equipment (EE) cooling filter, leading to water intrusion into the forward EE bay. This AD retains the actions required by AD 2021-06-03 and requires installing a shroud to the water supply line in the forward cargo compartment, and performing a leak test of the potable water system. For certain airplanes, this AD also requires replacing tubes and hoses from the water supply line and installing a shroud to the water return line. The FAA is issuing this AD to address the unsafe condition on these products.
2010-21-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD requires installing a support bracket and coupler on the left and right wing-to-fuselage transition, and metallic overbraid on the left and right leading edge wire assembly. This AD was prompted by fuel system reviews conducted by the manufacturer, as well as reports that the fuel quantity system was affected by lightning-induced transients. We are issuing this AD to prevent lightning-induced transients to the fuel quantity indication system, which could cause voltage levels to go beyond original design levels between fuel tank probes and structure, and become a potential ignition source at the fuel tank, which, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
83-04-02: 83-04-02 FOKKER-VFW B.V.: Amendment 39-4569. Applies to Model F28 airplanes serial number 11003 and on. To prevent loss of aileron control, accomplish the following, unless already accomplished within the last three months prior to the effective date of this AD: 1. Within the next 50 hours time in service or one month, whichever occurs first after the effective date of this AD, perform a one time inspection of the aileron control cables in both the R.H. and L.H. wings for airplanes with more than one year's time from the date of manufacture in accordance with the following instructions: a. Open the main landing gear doors and flap shroud doors 81j, 81k, 82j, and 82k. b. Rotate the aileron control wheel in the cockpit to the full left-hand position. c. With the control wheel in this position, inspect the four aileron input cables at location of cable pulleys between stations 3100 and 3600 (illustrated parts catalog, chapter 27-10-01-05). Inspection must be performed with the aid of a mirror and a very strong light otherwise corrosion may remain undetected. d. If wires are broken or there are definite signs of corrosion in the cables, accomplish the actions of paragraph 2. e. Rotate the aileron control wheel in the cockpit to the full right-hand position and repeat the inspections described in paragraphs (c) and (d). 2. If corrosion on the cables is evident from the inspections of paragraph 1, above, accomplish the following: a. Disconnect the cable at the turnbuckle in the wheelwell per the F28 Maintenance Manual, chapter 27-11. b. Pull out the cable at the cable pulleys between stations 3100 and 3600. c. Carefully bend the cable one time at the location of the corrosion over a radius of one inch. If any wire breaks while being bent, replace the cable prior to further flight. d. If no wires failed during the bending test of paragraph 2(c), repeat the test by bending the cable in the opposite direction. If any wire breaks, replace the cable before further flight. e. If the cable is not rejected by the bending tests of paragraphs 2(c) or 2(d), inspect the cable for interior corrosion in accordance with the F28 Maintenance Manual, chapter 20-30-9. f. If interior corrosion is found, replace the cable before further flight. g. Remove any exterior corrosion and treat the cable in accordance with F28 Maintenance Manual, Chapter 20-30-9. h. Reconnect the turnbuckles in the L.H. and R.H. wheelwells and adjust the cable system in accordance with F28 Maintenance Manual, Chapter 27-11-00, and check the aileron for free and smooth operation. 3. 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. 4. 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. 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). This amendment becomes effective March 7, 1983.
2000-15-52: This document publishes in the Federal Register an amendment adopting superseding Airworthiness Directive (AD) 2000-15-52, which was sent previously to all known U.S. owners and operators of Bell Helicopter Textron, Inc. Model (BHTI) Model 204B, 205A, 205A-1, 205B, and 212 helicopters by individual letters. This AD reduces the retirement index number (RIN) life limit for the main rotor mast (mast); increases the RIN factor for masts and main rotor trunnions (trunnions); applies standard RIN factors for all external load lifts; and requires a one-time inspection of the snap ring groove area of the mast. This AD also establishes RIN factors for masts and trunnions that have been previously installed on military or restricted category helicopters and removes from service those masts that have been previously installed with a hub spring. This amendment is prompted by an occurrence of a cracked mast at a lower value than the established RIN life limit. The actions specified by thisAD are intended to preclude the occurrence of fatigue cracks in the damper clamp splined area of a mast. A crack in the damper clamp splined area could result in failure of a mast or trunnion, separation of the main rotor system, and subsequent loss of control of the helicopter.
2016-05-01: We are superseding Airworthiness Directive (AD) 96-12-12, which applies to certain Piper Aircraft, Inc. Models PA-31, PA-31-300, PA-31-325, and PA-31-350 airplanes. AD 96-12-12 requires a one-time inspection of the bulkhead assembly at fuselage station (FS) 317.75 for cracks and the installation of one of two reinforcement kits determined by whether cracks were found during the inspection. This new AD requires repetitive inspections of the bulkhead assembly at FS 317.75 for cracks, repair of cracks as necessary, and the installation of a reinforcement modification. This AD was prompted by cracks found in the FS 317.75 upper bulkhead. We are issuing this AD to correct the unsafe condition on these products.
67-08-04: 67-08-04 FAIRCHILD-HILLER: Amdt. 39-364 Part 39 Federal Register March 7, 1967. Applies to Models F-27 and FH-227 Series Airplanes. Compliance required as indicated. To detect cracks in the elevator torque tube and supporting rings, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, unless already accomplished, visually inspect the inside area of the elevator torque tube P/N 27-223002-3 between Elevator Stations 11.811 (inboard end rib) and 15.748 (next rib outboard) and the elevator end (inboard) rib and internal elevator structure adjacent to the elevator torque tube for cracks using a mirror and light or an equivalent inspection approved by an FAA maintenance inspector. If a crack is found, comply with (c) before further flight. (b) Within the next 75 hours' time in service after the effective date of this AD, unless already accomplished, inspect the elevator torque tube P/N 27-223002-3 between Elevator Stations 11.811 and 15.748 and the torque tube supporting rings P/N 27-223006-9 (inboard end rib) and P/N 27-223006-3 (next rib outboard) for cracks using X-ray or dye-penetrant with a glass of at least 10-power or an equivalent inspection approved by an FAA maintenance inspector. If a crack is found comply with (c) before further flight. (c) Replace any part found cracked with a part of the same part number that has been inspected for cracks in accordance with (b) or with an equivalent part approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (d) Upon request of the operator an FAA maintenance inspector may, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Eastern Region, increase the initial compliance times specified in this AD by not more than 5 hours' time in service if the request contains substantiating data to justify the increase for that operator. (Fairchild-Hiller Alert Service Bulletins No. F-27-55-13A and FH-227-55-2A dated February 21, 1967, pertain to this subject.) This directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated March 1, 1967.
2000-25-04: This amendment adopts a new airworthiness directive (AD), applicable to certain Raytheon (Beech) Model MU-300, MU-300-10, 400, 400A, and 400T series airplanes, that requires a one-time inspection to detect hydraulic fluid leakage from the B-nut area, which attaches a hydraulic tube to the anti-skid valve assembly, and corrective actions, if necessary; and installation of an additional support for the hydraulic tube. This amendment is intended to prevent an asymmetric braking condition and a longer stopping distance due to sudden loss of normal braking to the left wheel. Such loss of normal braking could result in the airplane overrunning the runway surface. This action is intended to address the identified unsafe condition.
2016-04-12: We are adopting a new airworthiness directive (AD) for certain Turbomeca S.A. Arriel 2B, 2B1, 2C, 2C1, 2C2, 2D, 2E, 2S1, and 2S2 turboshaft engines. This AD requires inspection, and, depending on the results, removal of the engine accessory gearbox (AGB). This AD was prompted by a report of an uncommanded in-flight shutdown (IFSD) of an Arriel 2 engine caused by rupture of the 41-tooth gear, which forms part of the bevel gear in the engine AGB. We are issuing this AD to prevent failure of the engine AGB, which could lead to in-flight shutdown, damage to the engine, and damage to the aircraft.
91-05-14: 91-05-14 AIRBUS INDUSTRIE: Amendment 39-6912. Docket No. 90-NM-92-AD. Applicability: Model A300 B2 and B4 series airplanes, certificated in any category, equipped with the flap ball screwjack no-back mechanisms, as listed in Airbus Industrie Service Bulletin A300-27-172, dated April 10, 1984, which have not been modified in accordance with Modification AI 5240, as described in Airbus Industrie Service Bulletin A300-27-173, dated May 2, 1984. Compliance: Required as indicated, unless previously accomplished. To prevent excessive wear of the carbon friction disc of the no-back assemblies, which could lead to an asymmetric flap condition in the event of flap transmission shaft failure, accomplish the following: A. For airplanes on which flap jamming with one or more affected flap screwjacks has occurred, perform a jackhead axial backlash measurement on the affected flap ball screwjacks in accordance with Airbus Industrie Service Bulletin A300-27-172, dated April 10, 1984, at the earlier of the following: 1. Prior to 13,000 landings; or 2. Within 1,000 landings after November 3, 1987 (the effective date of Amendment 39-5736, AD 87-21-03). B. For airplanes on which flap jamming with any of the affected flap screwjacks has not occurred, perform a jackhead axial backlash measurement on the affected flap ball screwjacks in accordance with Airbus Industrie Service Bulletin A300-27-172, dated April 10, 1984, at the earlier of the following: 1. Prior to 13,000 landings; or 2. Whichever is the later of the following: a. Within 1,000 landings after the effective date of this AD; or b. Within the next "N" number of landings after the effective date of this AD, as determined by the following formula: N = 3,600 - 0.2 x (number of accumulated landings). C. If backlash is found, repeat the measurement required by paragraph A. or B., of this AD, at the following intervals: 1. If the backlash is less than or equalto 0.33 mm, prior to the accumulation of 3,600 landings after the last measurement. 2. If the backlash is more than 0.33 mm but less than or equal to 0.40 mm, prior to the accumulation of 2,000 landings after the last measurement. 3. If the backlash is more than 0.40 mm, but less than or equal to 0.56 mm, prior to the accumulation of 1,000 landings after the last measurement. D. Replace the flap ball screwjack within the next 250 landings when a measurement required by paragraph A. or B. of this AD indicates that the backlash is greater than 0.56 mm. E. Within 18 months after the effective date of this AD, modify the flap ball screwjack assemblies by installing a new carbon friction disc, a modified drive shaft, and a collar to the no-back mechanism, in accordance with Airbus Industrie Service Bulletin A300-27-173, dated May 2, 1984. Installation of this modification constitutes terminating action for the repetitive measurements required by paragraph C. of this AD.NOTE: This Airbus Service Bulletin references Lucas Aerospace Service Bulletin No. 1058-27-1100 for additional modification instructions. F. 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, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM-113, and a copy sent to the cognizant Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Standardization Branch, ANM-113. G. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the 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 Airbus Industrie, AirbusSupport Division, Avenue Didier Daurat, 31700 Blagnac, France. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. Airworthiness Directive 91-05-14 supersedes AD 87-21-03, Amendment 39-5736. This amendment (39-6912, AD 91-05-14) becomes effective on April 1, 1991.
2008-24-51: This document publishes in the Federal Register an amendment adopting airworthiness directive (AD) 2008-24-51 that was sent previously to all known U.S. owners and operators of Boeing Model 737- 600, -700, -700C, -800, and -900 series airplanes by individual notices. This AD requires accomplishing a wiring test of the autoshutoff system to verify continuity and a visual verification that the wiring is correctly installed; doing corrective actions, if necessary; and doing a functional test of the autoshutoff system, and applicable maintenance actions. This AD is prompted by a report of a failure of the left-hand fuel pump of the center wing tank to shut off after being selected "OFF'' by the flightcrew during flight on a Boeing Model 737-700 series airplane. Subsequent to that report, the failure was found on two additional airplanes. We are issuing this AD to prevent extended dry-running of the fuel pump, which could lead to localized overheating of parts inside the fuel pump,and which could produce an ignition source inside the fuel tank.
2000-22-51: This document publishes in the Federal Register an amendment adopting superseding Airworthiness Directive (AD) 2000-22-51, which was sent previously by individual letters to all known U.S. owners and operators of Model HH-1K, TH-1F, TH-1L, UH-1A, UH-1B, UH-1E, UH-1F, UH-1H, UH-1L, and UH-1P; and Southwest Florida Aviation SW204, SW204HP, SW205, and SW205A-1 helicopters manufactured by Bell Helicopter Textron Inc. (BHTI) for the Armed Forces of the United States. This AD requires establishing a retirement life for certain main rotor masts, creating a component history card or equivalent record, and identifying certain masts as unairworthy. This AD also requires removing the hub spring, if installed, and determining whether a main rotor mast (mast) has ever been installed on a helicopter while operated with a hub spring. Conducting certain inspections based on the retirement index number (RIN) and on whether the helicopter was ever operated with a hub spring is also required.Replacing any mast that has inadequate radius or a burr in the damper clamp splined area is also required. Finally, this AD requires sending information concerning the mast to the FAA. This amendment is prompted by the discovery of a crack in a mast with a lower RIN value than the established life limit. This action is necessary to preclude the occurrence of a fatigue crack in the damper clamp splined area of a mast. This condition, if not corrected, could result in failure of a mast or main rotor trunnion (trunnion), separation of the main rotor system, and subsequent loss of control of the helicopter.
93-23-08: 93-23-08 CORPORATE JETS LIMITED (FORMERLY BRITISH AEROSPACE): Amendment 39-8742. Docket 92-NM-245-AD. Applicability: Model DH/BH/HS/BAe 125 series airplanes, excluding Model BAe 125-1000A series airplanes; equipped with Garrett Model TFE 731-3 series engines; as listed in Corporate Jets Limited Service Bulletin SB.24-293-3501A,B,C,D,E,F, & G, Revision 2, dated March 31, 1993; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent overheating of the battery contactors and emergency contactors and a potential fire in the rear equipment bay, accomplish the following: (a) For Model BAe 125-800A series airplanes: Within 6 months after the effective date of this AD, modify the mounting arrangements of the battery contactors and emergency contactors in the rear equipment bay, Modification Number 253501, Parts A, B, F, & G, as appropriate; and prior to further flight, perform a functional test; in accordance withCorporate Jets Limited Service Bulletin SB.24-293-3501A,B,C,D,E,F, & G, Revision 2, dated March 31, 1993. (b) For all other airplanes: Within 6 months after the effective date of this AD, modify the mounting arrangements of the battery contactors and emergency contactors in the rear equipment bay, Modification Number 253501, Parts C, D, & E, as appropriate; and prior to further flight, perform a functional test; in accordance with Corporate Jets Limited Service Bulletin SB.24-293-3501A,B,C,D, & E, Revision 1, dated February 4, 1993; or SB.24-293-3501A,B,C,D,E,F, & G, Revision 2, dated March 31, 1993. (c) 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, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (d) 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. (e) The modifications and tests shall be done in accordance with Corporate Jets Limited Service Bulletin SB.24-293-3501A,B,C,D, & E, Revision 1, dated February 4, 1993; or Corporate Jets Limited Service Bulletin SB.24-293-3501A,B,C,D,E,F, & G, Revision 2, dated March 31, 1993; as applicable. (NOTE: The issue date of Service Bulletins SB.24-293-3501A,B,C,D, & E and SB.24-293-3501A,B,C,D,E,F, & G is indicated only on page 1 of each document; no other page of these documents is dated.) 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 Corporate Jets, Inc., 22070 Broderick Drive, Sterling, Virginia 20166. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (f) This amendment becomes effective on December 29, 1993.