Results
2013-25-12: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model DC-9-10, DC-9-30, and DC-9-40 series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the aft pressure bulkhead web area is subject to widespread fatigue damage (WFD). This AD requires modifying the aft pressure bulkhead. The modification includes inspecting for cracks around the rivet holes, and repair of any cracking. We are issuing this AD to prevent fatigue cracking of the aft pressure bulkhead, which could result in reduced structural integrity of the airplane.
99-18-01: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 737-700 and -800 series airplanes, that currently requires revising the Airplane Flight Manual (AFM) to prohibit operation of the airplane under certain conditions; repetitive inspections of the tab mast fitting of the elevator tab assemblies to detect cracking; an elevator tab freeplay check; and corrective actions, if necessary. That AD also provides for optional terminating action for certain repetitive inspections, and requires installing an additional fastener on the elevator tab mast fitting, which terminates the AFM revision and extends certain repetitive inspection intervals. This amendment continues to require certain actions, and revises and adds certain other requirements. This amendment is prompted by a report of a severe vibration incident on a Boeing Model 737-800 series airplane; inspection revealed fracturing of the elevator tab mast fitting and excessive freeplay in the elevator tab. The actions specified in this AD are intended to prevent loss of controllability of the airplane due to excessive freeplay in the elevator tab or a free tab.
80-25-06 R1: 80-25-06 R1 TELEDYNE CONTINENTAL MOTORS: Amendment 39-3984 as amended by Amendment 39-4061. Applies to Models GTSIO-520-L, serial numbers 608324 through 608627; GTSIO-520-M, serial numbers 606619 through 606890; and GTSIO-520-N, serial numbers 610001 through 610107, engines with 100 hours or less total time in service on the effective date of this AD, installed on but not limited to certain Cessna Models 404 and 421C airplanes certificated in all categories. Compliance required as indicated, unless already accomplished. To prevent engine failure due to loss of engine oil pressure, damage due to contaminated oil, and propeller shaft damage resulting from a malfunctioning thrust washer accomplish the following: (a) Before each flight and immediately after each flight until the accumulation of 100 hours total time in service, perform a special oil pressure check to determine the oil pressure with engine power at the same level as the magneto check. If oil pressure fluctuatesor is less than 30 psi, accomplish paragraphs (b)(1) and (b)(2) before further flight. This oil pressure check may be accomplished by the pilot as provided in FAR 43.3(h). (b) Prior to the next flight and at each oil change until the accumulation of 100 hours total time in service: (1) Remove the oil filter, disassemble the canister, and inspect the paper element between the pleats to determine the quantity of metallic material visually and by using a clean magnet. If total metallic contaminants are in excess of the quantity necessary to cover a 1/4 inch diameter surface, before further flight take the necessary maintenance action to replace those parts that are malfunctioning. NOTE: Exercise caution to prevent contamination of the filter element during disassembly. (2) Inspect to determine the end clearance (shaft end play) of the propeller drive shaft with engine at ambient temperature. If axial movement is in excess of .020 inch, before further flight take necessarymaintenance action to replace those parts that are malfunctioning. (c) Prior to the next flight, inspect the engine and airplane records and change oil if necessary to ensure that SAE No. 50 oil is installed for ambient temperatures above 40 degrees F or SAE No. 30 oil is installed for ambient temperatures below 40 degrees F. (d) Upon or before the accumulation of 25 hours, 50 hours and 100 hours total time in service, change oil and oil filter. Do not use multiviscosity oils within the first 100 hours time in service. (e) Make appropriate maintenance record entry when accomplishing each requirement of this AD. The airplanes equipped with affected engines may be flown in accordance with FAR 21.197 to a location where the AD compliance procedures can be accomplished. An equivalent method of compliance may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southern Region. NOTE: Continental Motors Service Bulletin NO. M80-30, dated December 10, 1980, pertains to this subject. Amendment 39-3984 became effective December 5, 1980. This amendment 39-4061 becomes effective March 23, 1981.
2013-25-09: We are adopting a new airworthiness directive (AD) for certain AgustaWestland S.p.A. (Agusta) Model AB139 and AW139 helicopters. This AD requires inspecting the nose landing gear (NLG) pin installations for incorrect assembly. This AD is prompted by reports of incorrectly installed pins discovered on in-service aircraft. These actions are intended to detect incorrectly installed pins, which could result in collapse of the NLG during taxi or landing.
2010-15-10: We are adopting a new airworthiness directive (AD) for certain Piper Aircraft, Inc. (Piper) PA-28, PA-32, PA-34, and PA-44 series airplanes. This AD requires you to inspect the control wheel shaft on both the pilot and copilot sides and, if necessary, replace the control wheel shaft. This AD results from two field reports of incorrectly assembled control wheel shafts. We are issuing this AD to detect and correct any incorrectly assembled control wheel shafts. This condition, if left uncorrected, could lead to separation of the control wheel shaft, resulting in loss of pitch and roll control.
99-17-11: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A319, A320, and A321 series airplanes, that requires repetitive inspections to detect wear of the inboard flap trunnions, and to detect wear or debonding of the protective half-shells; and corrective actions, if necessary. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to detect and correct chafing and resultant wear damage on the inboard flap drive trunnions or on the protective half-shells, which could result in failure of the trunnion primary load path; this would adversely affect the fatigue life of the secondary load path and could lead to loss of the flap.
2022-15-05: The FAA is adopting a new airworthiness directive (AD) for certain Airbus SAS Model A318 series airplanes; Model A319-111, -112, - 113, -114, -115, -131, -132, and -133 airplanes; Model A320-211, -212, -214, -216, -231, -232, and -233 airplanes; and Model A321-111, -112, - 131, -211, -212, -213, -231, and -232 airplanes. This AD was prompted by a report that cracks were found on the web horizontal flange and inner cap on a certain frame (FR), left-hand (LH) and right-hand (RH) sides, at a certain stringer (STGR). This AD requires repetitive high frequency eddy current (HFEC) inspections for cracks on the web horizontal flange and inner cap, and applicable corrective actions, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
53-09-04: 53-09-04 CONTINENTAL: Applies to All Aircraft Equipped With Continental W-670-9A (Ordnance-Tank) Engines and Ground Adjustable Propellers Having Blades 11C1 (Hamilton Standard Model Designation) or 4350, 4350F, or 4350F1 (Navy Model Designation.) Compliance required not later than May 15, 1953. In the absence of suitable propeller vibration stress data, the following precautionary measures should be taken to minimize the possibility of propeller blade fatigue failures: (1) Disassemble propeller and inspect for cracks by etching the shank areas of the blades under the hub clamp rings. (2) Cut propellers to between 102 inches maximum and 100 inches minimum diameter. (3) Set blade angle so that static r.p.m. is between 1,500 and 1,975. (4) Install propeller on engine in the zero degree position (blades in line with crankthrow). (5) Placard airplane, "Do not exceed 1,900 r.p.m. for all operations except takeoff." (6) Remove all nicks and gouges from tip region andmaintain propeller blades as outlined in Civil Aeronautics Manual 18.
99-17-20: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 757-200 and -300 series airplanes, that requires modification of the off-wing emergency evacuation slide system. This amendment is prompted by reports that a certain type of off-wing escape slide aboard several airplanes separated from the airplane during flight. The actions specified by this AD are intended to prevent separation of the emergency evacuation slide from the airplane, which could result in damage to the fuselage and unavailability of an escape slide during an emergency evacuation.
92-27-08: 92-27-08 CESSNA: Amendment 39-8442. Docket No. 92-NM-228-AD. Applicability: Citation Model 650 series airplanes, having serial numbers -0001 thru -0219, inclusive, and -7001 thru -7013, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent the potential loss of elevator control, accomplish the following: (a) Within 25 hours time-in-service after the effective date of this AD: Perform an inspection of the inboard attach fittings of the left-hand and right-hand elevator torque tubes to determine minimum material wall thickness in accordance with Cessna Citation Alert Service Letter A650-27-30, dated November 12, 1992; and perform a visual inspection of the fittings to detect cracking. (1) If the material thickness of the inboard attach fitting, at all locations, is 0.045 inch or thicker, and if no cracked fitting is found, no further action is required by this AD. (2) If the materialthickness of the inboard attach fitting, at any location, is equal to or greater than 0.040 inch but less than 0.045 inch, and if no cracked fitting is found, within 150 flight hours, replace the inboard attach fitting in accordance with the service letter. (3) If the material thickness of the inboard attach fitting, at any location, is thinner than 0.040 inch, or if a cracked fitting is found, prior to further flight, replace the fitting with a fitting having part number 6234132-8, in accordance with the service letter. (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, Wichita 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, Wichita ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Wichita ACO. (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 inspection and replacement shall be done in accordance with Cessna Citation Alert Service Letter A650-27-30, dated November 12, 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 Cessna Aircraft Company, Citation Marketing Division, P.O. Box 7706, Wichita, Kansas 67277. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at FAA, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on January 4, 1993.
86-21-09: 86-21-09 BRITISH AEROSPACE (BAe): Amendment 39-5447. Applies to BAe Jetstream Model 3101 (all serial numbers) airplanes certificated in any category which have incorporated Arkansas Modification Center (AMC) Supplemental Type Certificate (STC) No. SA5900SW baggage pod serial numbers 002 through 044 inclusive. Compliance: Required within the next fifty (50) landings after the effective date of this AD, unless already accomplished. NOTE: If landings are not recorded, one hour time-in-service (TIS) equals two landings. To prevent possible chafing of the flap jack lock valve pressure pipe and loss of hydraulic fluid, accomplish the following: (a) Visually inspect flap jack lock valve hydraulic pressure pipe BAe Part Number (P/N) 616302 for evidence of chafing in accordance with Section 2. "ACCOMPLISHMENT INSTRUCTIONS," Paragraph B "ACCOMPLISHMENTS," of BAe Alert Service Bulletin (ASB) No. 27-A-JA860226, dated August 11, 1986. (1) If chafing has occurred, before further flight, (i) Replace hydraulic pressure pipe P/N 616302 with a serviceable airworthy part, in accordance with BAe S/B No. 27-A-JA860226, and (ii) Modify pod by incorporating an aperture in accordance with Section 2., "ACCOMPLISHMENT INSTRUCTIONS," paragraphs 1 through 8 of revised AMC Service Bulletin (S/B) No. 25-0002(-2), dated September 11, 1986. (iii) Accomplish paragraph (b) of this AD. (2) If no chafing has occurred, before further flight, (i) modify pod by incorporating an aperture in accordance with Section 2., "ACCOMPLISHMENT INSTRUCTIONS," paragraphs 1 through 8 of AMC S/B No. 25-0002(-2), and (ii) accomplish paragraph (b) of this AD. (b) Measure the clearance distance between the flap jack lock valve hydraulic pressure pipe, P/N 616302 and the pod membrane smoke detector mounting screws, over the whole flap position operating range, in accordance with Section 2., "ACCOMPLISHMENT INSTRUCTIONS". If the clearance at all flap positions is: (1) 0.50 inches or more, install the aperture cover plate provided in AMC Kit P/N 31-5179-37 and accomplish paragraphs 14 through 17 of Section 2., "ACCOMPLISHMENT INSTRUCTIONS," of AMC S/B No. 25-0002(-2), and return the airplane to service. (2) Less than 0.50 inches, (i) Modify the pod in accordance with paragraphs 18 through 29 of Section 2., "ACCOMPLISHMENT INSTRUCTIONS," of AMC S/B No. 25-0002(-2), and (ii) Install baggage pod and repeat the actions specified in paragraph (b) of this AD. (c) Aircraft may be flown in accordance with FAR 21.197 to a location where this Airworthiness Directive (AD) can be accomplished. (d) An equivalent means of compliance with this AD may be used if approved by the Manager, FAA, Southwest Region, Special Programs Branch, ASW-190, 4400 Blue Mound Road, Post Office Box 1689, Fort Worth, Texas 76101. All persons affected by this AD may obtain copies of Arkansas Modification Center, Inc., (AMC) Service Bulletin (S/B) No. 25-0002(-2) dated July 22, 1986, Revised September 11, 1986, referred to herein upon request to the Arkansas Modification Center, Inc., Post Office Box 3356, Adams Field, Little Rock, Arkansas 72203, and Alert Service Bulletin No. 27-A- JA860226, dated August 11, 1986, referred to herein upon request to the British Aerospace, Engineering Department, Post Office Box 17414, Dulles International Airport, Washington, D.C. 20041; or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective October 29, 1986.
2022-15-02: The FAA is adopting a new airworthiness directive (AD) for certain Cameron Balloons Ltd. (Cameron) Stratus double burner assemblies installed on hot air balloons. This AD was prompted by reports from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI identifies the unsafe condition as fatigue cracking of the weld on Stratus double burner hangers. This AD requires repetitively inspecting certain Stratus double burner hangers and replacing certain Stratus double burners, and prohibits installing certain parts. The FAA is issuing this AD to address the unsafe condition on these products.
2013-24-16: We are adopting a new airworthiness directive (AD) for Schempp-Hirth Flugzeugbau GmbH Model Duo Discus T gliders. 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 the instructions provided to inspect the propeller hub and blades are insufficient for detecting cracks and/or other damage, and other operating instructions provided by the flight and maintenance manual are incorrect and insufficient. We are issuing this AD to require actions to address the unsafe condition on these products.
99-17-14: This amendment adopts a new airworthiness directive (AD), applicable to certain Bombardier Model DHC-8 series airplanes, that requires a one-time inspection of the spring assemblies located in the rudder control feel unit to verify that dual rate configuration springs are installed; and revising the Airplane Flight Manual to prohibit airplane operation from runways less than 75 feet wide, if necessary. This amendment also requires eventual replacement of any single rate configuration springs with dual rate configuration springs, which terminates the requirement for the AFM revision. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent an asymmetric rudder force condition, which could result in reduced controllability of the airplane and consequent potential for center line deviation.
74-16-01: 74-16-01 GENERAL DYNAMICS: Amendment 39-1904 as amended by Amendment 39-3206. Applies to General Dynamics Model 340 and 440 series airplanes, certificated in all categories, including those modified for turbo-propeller power. Compliance required as indicated. To prevent loss of structural integrity of the frame, due to longitudinal cracks in the stringer or beltframe at the beltframe attach points, accomplish the following: (a) For airplanes with more than 52,000 hours time in service as of the effective date of this AD, perform an initial inspection of the stringer to frame attachments per (g)(1), within 300 hours additional time in service, unless already accomplished within the last 4,700 hours time in service, in addition: (1) Aircraft certificated at a maximum gross weight of 55,000 pounds or less, and without a main deck cargo door, within 5,000 hours time in service from the initial inspection, inspect per paragraph (g)(2), and thereafter: (i) Within 5,000 hours time in service from the inspection accomplished per paragraph (g)(2), and thereafter at intervals not to exceed 5,000 hours time in service from the last inspection, inspect per paragraph (g)(3); and, (ii) Within 7,500 hours time in service from the inspection accomplished per paragraph (g)(2), and thereafter at intervals not to exceed 7,500 hours time in service from the last inspection, inspect per paragraph (g)(4). (2) Aircraft with a main deck cargo door and aircraft certificated at a maximum gross weight of more than 55,000 pounds, within 5,000 hours time in service from the initial inspection and thereafter at intervals not to exceed 5,000 hours time in service from the last inspection, inspect per paragraph (g)(2). (b) For airplanes with more than 42,000 hours time in service, up to and including 52,000 hours time in service as of the effective date of this AD, perform an initial inspection of the stringer to frame attachments per (g)(1), within 1,500 hours additional time in service, unless already accomplished within the last 3,500 hours time in service, and reinspect in accordance with paragraph (a)(1) or (a)(2). (c) For airplanes with more than 35,000 hours' time in service up to and including 42,000 hours' time in service as of the effective date of this A.D., perform an initial inspection of the stringer to frame attachments per (g)(1), below, within the next 3,000 hours' time in service, unless already accomplished within the last 4,500 hours' time in service, and thereafter perform the inspection described in (g)(2), below, at intervals not to exceed 7,500 hours' additional time in service. See (f), below, for reduction in repetitive inspection interval. (d) For airplanes with more than 25,000 hours' time in service up to and including 35,000 hours' time in service as of the effective date of this A.D., perform an initial inspection of the stringer to frame attachment per (g)(1), below, within the next 3,000 hours' time in service, unless already accomplished within the last 7,000 hours' time in service, and thereafter perform the inspection described in (g)(2), below, at intervals not to exceed 10,000 hours' time in service. See (f), below, for reduction in repetitive inspection interval. (e) For those airplanes with less than 25,000 hours' time in service on the effective date of this A.D., and have not previously been inspected for cracks, perform the initial inspection described in (g)(1), below, within 3,000 hours' additional time in service after accumulating 25,000 hours' time in service. Repetitive inspections must be performed per (g)(2), below, at the intervals required by (a) through (d), above. For those aircraft which have been previously inspected for cracks per (g)(1), below, accomplish the repetitive inspections per (g)(2), below, at the intervals required by (a) through (d), above. The first inspection per (g)(2) must be accomplished before accumulating 28,000 hours' total time in service or 10,000 additional flight hours from the inspection per (g)(1), whichever occurs later. (f) The second inspection, for airplanes affected by paragraphs (c) and (d), above, will be within the 7,500 or 10,000 hours' time in service as specified; however, for subsequent inspections, the interval will be based upon the airplane time in service at the time of the last inspection rather than the total time in service which results in the airplane reaching the next category. If the last inspection was accomplished when the airplane had more than 42,000 hours' time in service, all subsequent inspections must be at intervals not to exceed 5,000 hours' additional time in service. (g) (1) Perform a close, visual, above the floor level, inspection of the stringers and beltframes for evidence of cracks at all stringer to beltframe attach points between stations 140 and 889. (2) Perform a close visual inspection of all stringer to frame attachment (above and below floor level) for evidence of cracks between stations 140 and 889. (3) Perform a close visual inspection of all stringer to frame attachment for evidence of cracks below the floor between stations 681 and 889. (4) Perform a close visual inspection of all stringer to frame attachment for evidence of cracks between stations 140 and 889 above the floor and between stations 140 and 681 below the floor. (5) If cracks are found in any of these inspections, repair in accordance with applicable structural repair manual prior to further flight. (h) Equivalent inspections and repairs may be approved by the Chief, Aircraft Engineering Division, upon the submission of adequate substantiating data. (i) This A.D. may be amended to modify the repetitive inspection intervals if substantiating data is presented to the Chief, Aircraft Engineering Division, FAA Western Region. (j) If cracks are found as a result of the inspections performed in compliance with this A.D., report the findings to the Chief, Aircraft Engineering Division, FAA Western Region, P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009. (Reporting approved by the Bureau of Budget under BOB No. 04-R-0174.) Identify part number, total hours' time in service on the part and time in service since last inspection; number, size and location of crack(s); total time on the aircraft, serial number of the aircraft. (k) Airplanes may be flown to a base for the accomplishment of maintenance required by this A.D. per FAR's 21.197 and 21.199. Amendment 39-1904 became effective on August 2, 1974. This amendment 39-3206 becomes effective May 4, 1978.
99-17-08: This amendment adopts a new airworthiness directive (AD) that applies to certain Pilatus Aircraft Ltd. (Pilatus) Models PC-12 and PC-12/45 airplanes. This AD requires modifying the generator 2 excitation by removing certain diodes and installing a new 5-amp circuit breaker and suppression filter. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Switzerland. The actions specified by this AD are intended to prevent damage to electrical components if the generator 2 is not switched off prior to engine shutdown and it overheats, which could result in loss of electrical power to certain critical airplane components.
99-17-05: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Israel Aircraft Industries Model Astra SPX series airplanes. This action requires repetitive inspections to detect cracking of the main fuel tube assemblies of the left and right engines, and corrective action, if necessary. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified in this AD are intended to detect and correct fuel line fractures, which could result in in-flight engine shutdowns or an increased risk of engine nacelle fires.
78-08-13: 78-08-13 PRATT & WHITNEY AIRCRAFT: Amendment 39-3195 as amended by Amendment 39-3710. Applies to all Pratt & Whitney Aircraft Models JT9D-3, -3A, -7, -7A, -7H, -7AH, -7F, and -7J turbofan engines. Compliance required as indicated, unless already accomplished. To prevent main gearbox fires due to internal bearing failures, perform the following in accordance with the provisions of Pratt & Whitney Alert Service Bulletin 4854 dated March 27, 1978, or later FAA approved revision: 1. All gearboxes having more than 5,000 hours time in service since new or more than 5,000 hours time in service since the last bearing inspection, in accordance with the JT9D Engine Manual, P/N 646028, Section 72-09-50, must have the magnetic plug inspected within the next 300 hours time in service after the effective date of this AD, and every 300 hours time in service thereafter. 2. All gearboxes having 5,000 hours or less time in service since new or 5,000 hours or less time in service since the last bearing inspection, in accordance with the JT9D Engine Manual, P/N 646028, Section 72- 02-50, must be inspected with either (a) or (b) whichever occurs earlier: a. Within the next 650 hours and every 650 hours thereafter until a total gearbox time of 5,000 hours has been accumulated. b. At a total gearbox time of 5,300 hours. 3. If metal particles (spalled pieces), as defined in Paragraph 2A of Alert Service Bulletin 4854, are found on the magnetic plug, perform the following: Reinspect the magnetic plug within the next 50 hours time in service: a. If no additional spalled particles are found resume the repetitive inspection of paragraph 1 or 2 as applicable. b. If spalled particles are found, reinspect within the next 25 hours time in service. (1) If no additional spalled particles are found resume the repetitive inspection of paragraph 1 or 2 as applicable. (2) If spalled particles are found, remove the gearbox from service prior to further flight. Upon request of the operator, an equivalent method of compliance with the requirements of this AD may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, New England Region, may adjust the repetitive inspection interval 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 that operator. The manufacturer's specifications and procedures identified and 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 Pratt & Whitney Aircraft, Division of United Technologies Corporation, 400 Main Street, East Hartford, Connecticut 06108. These documents may also be examined at the Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD, which includes the incorporated material in full, is maintained by the FAA at its headquarters in Washington, D.C. and at the New England Region. Amendment 39-3195 became effective upon publication in the Federal Register. This amendment 39-3710 becomes effective April 18, 1980.
93-19-02 R1: 93-19-02 R1 PRATT & WHITNEY: Amendment 39-9038. Docket 92-ANE-33. Revises AD 93-19-02, Amendment 39-8695. Applicability: Pratt & Whitney (PW) JT9D-3A, -7, -7H, -7A, -7AH, -7F, -7J, -20, and -20J turbofan engines installed on but not limited to Boeing 747 series, Airbus A300 series, and McDonnell Douglas DC-10 series aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent diffuser case rupture and an uncontained engine failure, accomplish the following: (a) For those diffuser cases that have not been inspected in accordance with PW Alert Service Bulletin (ASB) No. 6076, Revision 1, dated August 20, 1992, initially inspect the diffuser case for cracks in accordance with the intervals and requirements described in paragraphs (d), (f), (g), (i), (j), (k), or (l) of this AD, as applicable. (b) For those diffuser cases that have not been inspected in accordance with PW ASB No. 6076, Revision 1, dated August 20, 1992, inspect the diffuser case rear rail along the shell wall at Boss 6 for weld repair size in accordance with PW ASB No. 6076, Revision 1, dated August 20, 1992, at the next M flange separation of the high pressure turbine case after the effective date of this AD. Diffuser cases with weld repairs in the rear rail along the shell wall of axial length greater than or equal to 1.5 inches at Boss 6 must not be returned to service. If the weld length is less than 1.5 inches, inspect in accordance with the new criteria, improved technique, intervals, and requirements defined in the Accomplishment Instructions of PW Service Bulletin (SB) No. 5591, Revision 7, dated August 25, 1992. NOTE: Additional information regarding weld repair requirements for the diffuser case rear rail is contained in PW JT9D Engine Manual, Part Number 686028, dated September 1, 1993. (c) For those diffuser cases that have been inspected in accordance with PW ASB No. 6076, Revision 1, dated August 20, 1992, accomplish the following:(1) For diffuser cases that have weld repairs in the rear rail along the shell wall at Boss 6 of axial length greater than or equal to 1.5 inches, remove from service and replace with a serviceable part prior to further flight. (2) For diffuser cases that have weld repairs in the rear rail along the shell wall at Boss 6 of axial length less than 1.5 inches, initially inspect the diffuser case for cracks in accordance with the intervals and requirements described in paragraphs (d), (f), (g), (i), (j), (k), or (l) of this AD, as applicable. (3) For diffuser cases that have no weld repairs in the rear rail along the shell wall at Boss 6, initially inspect the diffuser case for cracks in accordance with the intervals and requirements described in paragraphs (e), (g), (h), (i), (j), (k), or (l) of this AD, as applicable. (d) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, that contained rear rails with no cracks at any boss location at the last ECI, and have a weld repair in the rear rail along the shell wall at Boss 6, perform an initial ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, as follows: (1) For diffuser cases with greater than 275 cycles in service (CIS) since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, on the effective date of this AD, perform an ECI in accordance with the new criteria and improved technique defined in the Accomplishment Instructions PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 500 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, or prior to accumulating 75 CIS after the effective date of this AD, whichever occurs first. (2) For diffuser cases with less than or equal to 275 CISsince the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, on the effective date of this AD, perform an ECI in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 350 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986. (e) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, that contained rear rails with no cracks at any boss location at the last ECI, and have no weld repairs in the rear rail along the shell wall at Boss 6, perform an ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 500 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986. (f) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, that contained rear rails with "A" cracks at Boss 6 at the last ECI, and have a weld repair in the rear rail along the shell wall at Boss 6, perform an ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 300 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, or prior to accumulating 60 CIS after the effective date of this AD, whichever occurs first. (g) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, that contained rear rails with "A" cracks at any boss location other than at Boss 6 at the last ECI, with or without weld repairs in the rear rail along the shell wall at Boss 6, perform an ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 300 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 16, 1986. (h) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, that contained rear rails with "A" cracks at Boss 6 at last ECI, and have no weld repairs at Boss 6, perform an ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 300 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986. (i) For those diffuser cases that havebeen inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, and contained rear rails with "B" cracks at Boss 6 at last ECI, with or without weld repairs in the rear rail along the shell wall at Boss 6, remove from service and replace with a serviceable part prior to accumulating 5 CIS after the effective date of this AD. (j) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, and contained rear rails with "B" cracks at any boss location other than Boss 6 at last ECI, with or without weld repairs in the rear rail along the shell wall at Boss 6, perform an ECI of the diffuser case rear rail for cracks in accordance with the new criteria and improved technique defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992, prior to accumulating 75 CIS since the last ECI performed in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986. (k) For those diffuser cases that have been inspected in accordance PW SB No. 5591, Revision 4, dated March 6, 1986, and contained rear rails with "C" cracks at Boss 6 at last ECI, with or without weld repairs in the rear rail along the shell wall at Boss 6, remove from service and replace with a serviceable part prior to further flight. (l) For those diffuser cases that have been inspected in accordance with PW SB No. 5591, Revision 4, dated March 6, 1986, and contain rear rails with "C" cracks at any boss location other than Boss 6 at last ECI, with or without weld repairs in the rear rail along the shell wall at Boss 6, remove from service and replace with a serviceable part as follows: (1) For shell wall cracks of greater than or equal to 2 inches, remove from service and replace with a serviceable part prior to further flight. (2) For shell wall cracks of less than 2 inches, remove from service and replace with a serviceable part within 5 CIS after the effective date of this AD. (m)Thereafter, perform repetitive ECI of the diffuser case rear rail for cracks in accordance with the new criteria, improved technique, intervals, requirements, and removal from service criteria defined in the Accomplishment Instructions of PW SB No. 5591, Revision 7, dated August 25, 1992. (n) For those diffuser cases that have been weld repaired at any boss location, at the next K flange separation of the diffuser case after the effective date of this AD, perform a one-time x-ray inspection of the diffuser case rear rail and sides of all bosses for weld quality in accordance with PW SB No. 6088, dated August 5, 1992, prior to installation of the diffuser case. Remove any weld defects within the inspection zone in accordance with PW SB No. 6088, dated August 5, 1992, prior to installation of the diffuser case. (o) For those diffuser cases with rear rails that have been weld repaired at any boss location, incorporate the modifications described in PW SB No. 5805, Revision 6, dated September 15, 1993, at the next removal of the diffuser case for repair after the effective date of this AD. (p) Installation of an improved diffuser case in accordance with PW SB No. 6105, Revision 2, dated May 14, 1993, constitutes terminating action to the inspections and modifications required by this AD. (q) 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, Engine Certification Office. The request should be forwarded through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Engine Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Engine Certification Office. (r) Except for diffuser cases that have cracks that require removal prior to further flight, special flight permits may beissued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. For diffuser cases that have cracks that require removal prior to further flight, on aircraft that are eligible for an engine-inoperative ferry, special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished with one engine inoperative. (s) The inspections and modifications shall be done in accordance with the following PW service bulletins: Document No. Pages Revision Date SB No. 5591 1-3 7 August 25, 1992 4-9 6 August 14, 1992 10 7 August 25, 1992 11-12 6 August 14, 1992 13 7 August 25, 1992 14-15 6 August 14, 1992 16 7 August 25, 1992 17-19 6 August 14, 1992 Total pages: 19 SB No. 5805 1-4 6 September 15, 1993 5 Original April 20, 1988 6-72 6 September 15, 1993 Total pages: 72 ASB No. 6076 1-5 1 August 20, 1992 6-19 Original July 31, 1992 Total pages: 19 SB No. 6088 1-11 Original August 5, 1992 Total pages: 11 SB No. 6105 1 2 May 14, 1993 2-7 Original January 15, 1993 8 1 April 14, 1993 9 2 May 14, 1993 10-15 Original January 15, 1993 16 2 May 14, 1993 17-18 Original January 15, 1993 19 2 May 14, 1993 20-46 Original January 15, 1993 47 1 April 14, 1993 48 2 May 14, 1993 49-56 Original January 15, 1993 Total pages: 56. This incorporation by reference was approved previously by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51 as of October 18, 1993 (58 FR 51212, October 1, 1993). Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Burlington, MA; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC.
2022-15-01: The FAA is adopting a new airworthiness directive (AD) for certain The Boeing Company Model 787-8, 787-9, and 787-10 airplanes. This AD was prompted by a report that during a C-check, corrosion was found in the vertical fin tension bolt hole located in the aluminum crown frames at a certain section. This AD requires inspecting certain vertical fin tension bolt holes; reviewing the bolt sealant application installation procedure in the existing maintenance or inspection program, as applicable; checking maintenance records to determine the replacement status of vertical fin tension bolts; and doing applicable on-condition actions. The FAA is issuing this AD to address the unsafe condition on these products.
99-17-02: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 777 series airplanes, that requires repetitive inspections of the safety spring wear plate doublers attached to the auxiliary power unit (APU) firewall, measurement of wear of the doublers, and follow-on actions, if necessary. For certain airplanes, this amendment also requires a one-time inspection to detect improper clearance between the safety spring wear plate doubler and the APU firewall, and corrective action, if necessary. This amendment also provides for optional terminating action for the repetitive inspections. This amendment is prompted by reports indicating that excessive wear was found on the safety spring wear plate doublers on the APU firewall of Boeing Model 777 series airplanes. The actions specified by this AD are intended to detect and correct wear of the safety spring wear plate doublers on the APU firewall, which could result in a hole in the APU firewall, and consequent decreased fire protection capability.
86-18-09: 86-18-09 GROB WERKE GMBH (BURKHART GROB): Amendment 39-5410. Applies to Model G103 TWIN ASTIR gliders serial numbers 3000 through 3291, and 3000-T-1 through 3284-T-44 certificated in any category. Compliance is required as indicated unless already accomplished. To prevent failure of the rudder control rear parallel lever P/N II 103-4320.05 which could result in loss of rudder control, accomplish the following: (a) 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 after the last inspection, until compliance with Paragraph (c) is accomplished, visually inspect the rear parallel lever in the area of the left and right boreholes, using a 10 power or greater magnifying glass, for cracks in accordance with Part 1 of the "Instructions" section of Grob Technical Information No. TM 315- 30, dated October 1, 1985. (b) If a cracked lever is found during the inspection required by Paragraph (a) of this AD, before further flight, replace the rear parallel lever with a stronger new parallel lever in accordance with part 2 of the "Instructions" section of Grob Technical Information No. TM 315- 30, dated October 1, 1985, and Grob Repair Instructions No. 315-30, dated October 1, 1985. (c) Within the next 30 hours time-in-service but no later than 60 days after the effective date of this AD, replace any rear parallel lever not replaced in accordance with Paragraph (b) of this AD, with a stronger rear parallel lever in accordance with Part 2 of the "Instructions" section of Grob Technical Information No. TM 315-30, dated October 1, 1985, and Grob Repair "Instructions" No. 315-30, dated October 1, 1985. NOTE: Stronger rear parallel lever does not have a new part number. It can be identified as it is made of stock aluminum, not a casting as the original part. Upon request, an equivalent means of compliance with the requirements of this AD may be approvedby the Manager, Brussels Aircraft Certification Office, AEU-100, Europe, Africa, and Middle East Office, FAA, c/o American Embassy, 15 Rue de la Loi B-1040 Brussels, Belgium, Telephone No. 513.38.30 ext. 2710 or the Manager, New York Aircraft Certification Office, Aircraft Certification Division, FAA, New England Region, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581, Telephone No. 516-791-6680. Upon submission of substantiating data by an owner or operator through an FAA maintenance inspector, the Manager, Brussels Aircraft Certification Office, or the Manager, New York Aircraft Certification Office, may adjust the compliance time specified in this AD. Grob Technical Information No. 315-30 dated October 1, 1985, and Grob Repair Instructions No. 315-30 dated October 1, 1985, identified and described in this document, 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 alreadyreceived these documents from the manufacturer may obtain copies upon request to Grob Systems, Inc., Aircraft Division, I-75 and Airport Drive, Bluffton, Ohio 45817. These documents also may be examined at the Office of Regional Counsel, ANE-7, FAA New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, Room 311, Rules Docket 86-ANE-36, between the hours of 8:00 am and 4:30 pm; Monday thru Friday, except Federal holidays. This amendment becomes effective on September 15, 1986.
86-21-04: 86-21-04 BRITISH AEROSPACE: Amendment 39-5419. Applies to Model BAe-146 series airplanes, with serial numbers as listed in BAe Service Bulletin 32-18, Revision 1, dated November 28, 1984, certificated in any category. Compliance is required within 60 days after the effective date of this AD. To prevent structural failure of the main landing gear, accomplish the following, unless previously accomplished: 1. Inspect and repair, if necessary, the main landing gear main fittings in accordance with BAe Service Bulletin 32-18, Revision 1, dated November 28, 1984. 2. 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. 3. 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 thisAD. All persons affected by this directive who have not already received the appropriate service bulletin from the manufacturer may obtain copies upon request to British Aerospace, Inc., Librarian, Box 17414, Dulles International Airport, Washington, D.C. 20041. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This Amendment becomes effective October 20, 1986.
99-16-12: This amendment adopts a new airworthiness directive (AD) that applies to certain Raytheon Aircraft Company (Raytheon) Model Beech 1900D airplanes that are equipped with the electric elevator trim option. This AD requires installing electric elevator trim servo covers. This AD is the result of reports of the affected airplanes leaving the factory without electric elevator trim servo covers installed. If the covers are not installed, moisture could freeze on parts of the electric actuator. The actions specified by this AD are intended to prevent failure of the electric elevator trim and difficulty operating the manual elevator trim caused by moisture freezing on parts of the electric actuator installation, which would result in the pilot having to apply constant pressure to the control wheel during flight.
79-23-04: 79-23-04 GENERAL ELECTRIC COMPANY: Amendment 39-3610. Applies to all General Electric CT58 turboshaft engines which are presently in use or have been used in repetitive heavy-lift operation. Compliance required as indicated, unless already accomplished. To prevent low cycle fatigue initiated failure, revise the total recorded operating cycles of all life-limited rotating components, on the effective date of this AD, and remove these components from service in accordance with the multiplying factors and retirement lives contained in General Electric Alert Service Bulletin CT58 (A72- 162) CEB-258, dated July 9, 1979. Later FAA approved revisions or equivalent means may be approved by the Chief, Engineering and Manufacturing Branch, New England Region. Hourly limits are not affected by this AD. Components with revised total recorded operating cycles in excess of the limits or within 600 cycles or 100 hours of the limits in Tables I, II, or III of General Electric Alert Service Bulletin CT58 (A72-162) CEB-258, on the effective date of this AD, must be removed from service prior to the accumulation of 600 additional cycles or 100 hours, whichever comes first. NOTE: Repetitive heavy-lift operations are considered to be those operations during which a lift-carry-drop cycle is repeated more than 10 times per hour without landing. This activity is typical of logging operations and may also include some construction or utility operations. 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 General Electric Company, 1000 Western Avenue, Lynn, Massachusetts, 01910. These documents may also be examined at the Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD, which includes the incorporated material in full, is maintained by the FAA at its Headquarters in Washington, D.C., and at FAA, New England Region Headquarters, Burlington, Massachusetts. This amendment becomes effective upon publication in the Federal Register.