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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.
2013-24-17: We are adopting a new airworthiness directive (AD) for General Electric Company (GE) GE90-110B1 and GE90-115B turbofan engines with certain high pressure compressor (HPC) rotor stage 2-5 spools installed. This AD requires removing these spools from service at times determined by a drawdown plan. This AD was prompted by reports of cracks in HPC rotor stage 2-5 spool aft spacer arms. We are issuing this AD to prevent failure of a critical life-limited rotating engine part, which could result in an uncontained engine failure and damage to the airplane.
98-15-18 R1: We are revising Airworthiness Directive (AD) 98-15-18 that applies to certain Maule Aerospace Technology, Inc. Models M-4, M-5, M- 6, M-7, MT-7, MX-7, MXT-7, and M-8 airplanes that are equipped with rear wing lift struts, part number (P/N) 2079E, and/or front wing lift struts, P/N 2080E. AD 98-15-18 required repetitively inspecting certain wing lift struts for internal corrosion and replacing of any wing lift strut where corrosion was found. Since we issued AD 98-15-18, we were informed by the manufacturer that Model MXT-7-420 airplanes are no longer in existence, are no longer type certificated, and should be removed from the Applicability section. We were also informed that paragraph (b) in AD 98-15-18 had been misinterpreted and caused confusion. This AD removes Model MXT-7-420 airplanes from the Applicability section and clarifies the intent of the language in paragraph (b) of AD 98-15-18. This AD also retains all other requirements of AD 98-15-18. We are issuing this AD to correct the unsafe condition on these products.
2006-03-06: This amendment adopts a new airworthiness directive (AD), applicable to all Empresa Brasileira de Aeronautica S.A. (EMBRAER) Model EMB-135 airplanes, and EMB-145, -145ER, -145MR, -145LR, -145XR, - 145MP, and -145EP airplanes. This AD requires inspecting the pilot's and co-pilot's seat tracks for proper locking of the seats, and adjusting or replacing the seat tracks if necessary. This AD also requires replacement of the seat locking pin on certain SICMA-brand seats. The actions specified by this AD are intended to prevent uncommanded movement of the pilot's or co-pilot's seat, which could interfere with the operation of the airplane and consequent temporary loss of airplane control. This action is intended to address the identified unsafe condition.
95-24-07: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320-111, -211, and -231 series airplanes, that requires modification of the aileron support frame of the wings. This amendment is prompted by reports indicating that tensile cracks have been found at a certain mounting hinge of the aileron support frame during full scale fatigue testing of the test article due to fatigue-related stress. The actions specified by this AD are intended to prevent such fatigue-related cracking, which could result in loss of the aileron control surface and the inability of the pilot to control rolling moments of the airplane.
99-16-13: This amendment supersedes an existing airworthiness directive (AD), applicable to MDHI Model MD-900 helicopters, that currently requires applying specified serial numbers and establishing life limits for certain parts. This amendment is prompted by additional analysis that supports an increase in the life limit of certain parts. The actions specified by this AD are intended to increase the life limits for various parts.
2013-24-15: We are superseding Airworthiness Directive (AD) 2007-11-08 for all The Boeing Company Model 727 airplanes. AD 2007-11-08 required repetitive inspections of the in-tank fuel boost pump wiring, installation of sleeving over the in-tank fuel boost pump wires, repetitive inspections of a certain electrical wire, sleeve, and conduit, and applicable investigative and corrective actions; and repetitive engine fuel suction feed operational tests. This new AD also requires replacement of the wire bundles for the wing and center fuel boost pumps, installation of convoluted liners, and related investigative and corrective actions if necessary. This new AD also requires replacement of the fuel quantity indicating system (FQIS) wires, a low-frequency eddy current inspection for cracking, and repair if necessary. This new AD also requires revising the maintenance program to incorporate changes to the airworthiness limitations section. This AD was prompted by a report of damage found to thesleeve, jacket, and insulation on an electrical wire during a repetitive inspection. We are issuing this AD to prevent chafing of the fuel boost pump electrical wiring and leakage of fuel into the conduit, and to prevent electrical arcing between the wiring and the surrounding conduit, which could result in arc-through of the conduit, and consequent fire or explosion of the fuel tank.
99-16-06: This amendment adopts a new airworthiness directive (AD) that applies to certain The New Piper Aircraft, Inc. (Piper) Model PA-46-350P airplanes. This AD requires installing reinforcement plates to the wing forward and aft attach fittings. This AD is the result of a report that sheet steel material that is below design strength standards may have been utilized on the wing attach fittings on the Model PA-46-350P airplanes manufactured since January 1995. The actions specified by this AD are intended to prevent structural failure of the wing attach fittings caused by the utilization of substandard material, which could result in the wing separating from the airplane with consequent loss of control.
93-05-15: 93-05-15 CESSNA AIRCRAFT COMPANY: Amendment 39-8519. Docket 92-NM-155-AD. Applicability: Citation Model 500/501 series airplanes, unit numbers -0001 through -0689, inclusive, equipped with thrust reversers; Citation Model 550/551 series airplanes, unit numbers -0002 through -0678, inclusive, equipped with thrust reversers; Citation Model S550 series airplanes, unit numbers -0001 through -0160, inclusive; Citation Model 560 series airplanes, unit numbers -071A, -092A, -109A, and -0001 through -0118, inclusive; Citation Model 650 series airplanes, serial numbers -0001 through -0217, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent severely reduced controllability of engine power authority, accomplish the following: (a) Within 150 hours time-in-service after the effective date of this AD, modify the thrust reverser throttle load limiter in accordance with Cessna Citation Service Bulletin SB500-78-11, dated September 13, 1991 (for Model 500/501 series airplanes); Cessna Citation Service Bulletin SB550-78-03, dated September 13, 1991 (for Model 550/551 series airplanes); Cessna Citation Service Bulletin SBS550-78-04, dated September 13, 1991 (for Model S550 series airplanes); Cessna Citation Service Bulletin SB560-78-02, dated September 13, 1991 (for Model 560 series airplanes); or Cessna Citation Service Bulletin SB650-78-05, Revision 1, dated June 12, 1992 (for Model 650 series airplanes); as applicable. (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 modification shall be done in accordance with the following Cessna Citation Service Bulletins, as applicable, which contain the specified effective pages: Service Bulletin Referenced and Date Page Number Revision Level Shown on Page Shown on Page 500-78-11, September 13, 1991 1-7 Original September 13, 1991 SBS550-78-04, September 13, 1991 1-7 Original September 13, 1991 SB550-78-03, September 13, 1991 1-7 Original September 13, 1991 SB560-78-02, September 13, 1991 1-7 Original September 13, 1991 SB650-78-05, Revision 1, June 12, 1992 1-3 4-6 1 Original June 12, 1992 February 14, 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 the FAA, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on April 23, 1993.
2013-24-14: We are adopting a new airworthiness directive (AD) for Diamond Aircraft Industries GmbH Models DA40 and DA40F 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 fatigue strength found in the aft main spar does not ensure unlimited lifetime structural integrity. We are issuing this AD to require actions to address the unsafe condition on these products.
2010-15-04: We are adopting a new airworthiness directive (AD) for the specified ECF Model EC225LP helicopters. This AD results from a mandatory continuing airworthiness information (MCAI) AD issued by the aviation authority of the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. The MCAI AD states there have been a "few'' reports of cracks and failure of the main rotor hub (MRH) cone restrainer support lugs at their attachment points on the reinforcement ring where the dome fairing is secured. Also, cracks on the dome fairing support have been reported. Failure of the cone restrainer support or the dome fairing support attachment lugs may lead to loss of the dome fairing, damage to the rotor blades, and subsequent loss of control of the helicopter.
99-16-02: This amendment supersedes an existing airworthiness directive (AD), applicable to certain British Aerospace Model BAC 1-11 200 and 400 series airplanes, that currently limits the number of operations at increased cabin pressure differential, and requires repetitive structural inspections for cracking of the fuselage, and repair or replacement of parts, if necessary. This amendment requires additional repetitive inspections for cracking of the fuselage. This amendment is prompted by the determination that airplanes operating at increased cabin pressure differential are more likely to develop fatigue cracking earlier in their service lives than those airplanes operating at normal cabin differential pressures. The actions specified by this AD are intended to detect and correct fatigue cracking of the airplane fuselage, which could result in reduced structural integrity of the airplane.
2013-22-19: We are adopting a new airworthiness directive (AD) for all Gulfstream Aerospace Corporation Model GV and GV-SP airplanes. This AD was prompted by reports of two independent types of failure of the fuel boost pump with overheat damage found on the internal components and external housing on one of the failure types, and fuel leakage on the other. This AD requires inspecting to determine if fuel boost pumps having a certain part number are installed, replacing the fuel boost pumps having a certain part number, and revising the airplane maintenance program to include revised instructions for continued airworthiness. We are issuing this AD to prevent fuel leakage in combination with a capacitor clearance issue, which could result in an uncontrolled fire in the wheel well.
72-20-02: 72-20-02 CONTINENTAL: Amendment 39-1522 as amended by Amendment 39-1613 as amended by Amendment 39-1774 is further amended by Amendment 39-1974. Applies to Models IO-470-D, IO-470-E, IO- 470-F, IO-470-H, IO-470-L, IO-470-M, IO-470-N, IO-470-S, IO-470-U, IO-470-V, and TSIO-470-B, TSIO-470- C, TSIO-470-D engines which have installed "Non-H" cylinder assemblies, Part Number 626820, and "H" cylinder assemblies manufactured or remanufactured prior to April 1963. NOTE 1: "H" type cylinders are those having the letter H impression stamped on the top edge of the rocker box flange over the exhaust valve. "Non-H" type cylinders do not have this identification. Compliance: Required as indicated unless already accomplished. To preclude additional inflight failures of Part Number 626820 "Non-H" type cylinder assemblies and "H" type cylinder assemblies manufactured or remanufactured prior to April 1963, accomplish the following or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, Southern Region, Atlanta, Georgia: A. Within the next 20 hours' time in service after the effective date of this airworthiness directive and thereafter at intervals not exceeding 20 hours' time in service from the last inspection: (1) Visually inspect the circumference of these cylinder assemblies at the junction of the aluminum head and steel barrel for oil leaks and/or combustion product stains. Engine permanent maintenance record entry must be made to reflect airworthiness directive compliance. NOTE 2: In order to perform the inspection required by Paragraph A(1) of this airworthiness directive, it may be necessary to remove engine cowling or access doors to permit visual examination with mirrors or other visual aids of the prescribed area of the cylinder. If the engine is clean and free of oil in the area to be inspected, the inspection required by Paragraph A(1) may be performed without further cleaning of the engine. If oil leakage froman unknown source has caused a generally oily condition, the engine should be washed down and run up to normal operating conditions prior to the inspection required by Paragraph A(1). During the inspection it may also be helpful to rotate the propeller to detect significant differences in compression between cylinders or audible compression leakage through a crack in the cylinder barrel. (2) The inspection(s) required by Paragraph A(1) must be performed by those persons authorized to perform inspections under Federal Aviation Regulation 43.3 except that on aircraft not utilized in air carrier services the inspection(s) may be performed by the holder of a pilot's certificate issued under Part 61 of the Federal Aviation Regulations on any aircraft owned or operated by him. (3) If oil leaks and /or combustion product stains are found at the junction of the cylinder head and barrel during any inspection required by Paragraph A(1), before further flight, a certificated powerplantmechanic shall investigate and establish the source of these conditions. If a cylinder barrel crack is found, the cracked cylinder must be replaced with an airworthy part. B. At next engine overhaul replace Part Number 626820 "Non-H" cylinder assemblies and "H" cylinder assemblies manufactured or remanufactured prior to April 1963 (as identified by the date of manufacture impression stamped in the machined area beneath the valve rocker shaft supports) with Part Number 626820 cylinder assemblies manufactured or remanufactured after April 1963 and having the letter "H" impression stamped on the top edge of the rocker box flange over the exhaust valve. NOTE 3: On engines manufactured or remanufactured by Teledyne Continental Motors during 1964 or later, as indicated by the year suffix on the serial number, it may be assumed without further verification that "H" type cylinders installed were manufactured subsequent to April 1963 if maintenance records do not indicate a cylinder exchange. On other engines having "H" type cylinders installed, it will be necessary to establish the cylinder's date of manufacture by removing the rocker box cover and inspecting the area beneath the rocker arm for the impression stamped manufacture date. Information on the location of this stamp was given on Page 115 of the August 1971 General Aviation Inspection Aids Summary. However, in the event the manufacture date appears ambiguous or illegible Teledyne Continental Service Bulletin M73-2 should be used to identify cylinders. NOTE 4: The term "Overhaul" as used in this AD is taken to include both major and top overhaul. When considering top overhauls, the AD applies equally to individual cylinder top overhauls as well as a complete engine assembly. C. Inspections as outlined in Paragraph A(1) are no longer required when Paragraph B has been accomplished. This airworthiness directive Amendment 39-1522 supersedes Amendment 39-509, AD 67-31-5. Amendment 39-1522 became effective September 25, 1972. Amendment 39-1613 became effective April 2, 1973. Amendment 39-1774 became effective February 1, 1974. This Amendment 39-1974 becomes effective October 2, 1974.
99-16-03: This amendment adopts a new airworthiness directive (AD), applicable to certain Learjet Model 23, 24, 25, 28, 29, 31, 55, and 60 series airplanes, that requires a one-time detailed visual inspection of the electrical wire leads of the horizontal stabilizer anti-ice system to verify that the numbers on the wire leads correctly correspond to the numbers on the connected airframe wiring; installation of a wire ID strap on the left- and right-hand sides of each terminal block; and installation of a warning placard. This amendment is prompted by a report of severe flight control buffeting of a Learjet Model 55 series airplane due to a malfunction of the horizontal stabilizer anti-ice system. The actions specified by this AD are intended to prevent undetected accretion of ice on the leading edge of the horizontal stabilizer, which could result in the loss of pitch control and consequent reduced controllability of the airplane.
83-07-15: 83-07-15 BRITISH AEROSPACE, AIRCRAFT GROUP, SCOTTISH DIVISION: Amendment 39-4617. Applies to Model HP.137 Jetstream MK.1 and Jetstream Series 200 airplanes certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent cracking of the front pressure bulkhead boundary angle and possible catastrophic depressurization, accomplish the following: a) For all affected airplanes with 12,900 or more flights, unless British Aerospace Modification No. 5115 is installed, within the next 100 flights after the effective date of this AD, apply a red line to the cabin differential pressure gauge to indicate that the maximum allowable differential pressure is 2.5 PSI. Extend the red line to the case of the instrument to avoid confusion if the glass cover is rotated. b) For all other affected airplanes with less than 12,900 flights, unless British Aerospace Modification No. 5115 is installed, within the next 100 flights after the effective date of this AD or upon the accumulation of 2,000 flights, whichever occurs later, apply a red line to the cabin differential pressure gauge to indicate that the maximum allowable differential pressure is 4.0 PSI except that in no case shall airplanes in this group be allowed to accumulate more than 13,000 flights without having the cabin differential pressure gauge remarked with a red line at 2.5 PSI in accordance with paragraph a) of this AD. Extend the red line to the case of the instrument to avoid confusion if the glass cover is rotated. c) For all aircraft modified per paragraphs a) and b) of this AD, fabricate a placard to read as follows and install it adjacent to the pressure differential gauge: 1) For aircraft affected by paragraph a) of this AD: "2.5PSI max. press. difference." 2) For aircraft affected by paragraph b) of this AD: "4.0PSI max. press. difference." d) Within the next 500 flights after the effective date of this AD on all affected airplanes, unless British Aerospace Modification No. 5115 is installed, modify the pressurization systems in accordance with British Aerospace, Aircraft Group Modification No. 5151, Issue 1, dated August 1981 or Modification No. 5152, Issue 1, dated August 1981, as appropriate to limit maximum pressurization to 4.0 PSI or 2.5 PSI respectively. e) For purposes of complying with this AD, subject to the acceptance by the assigned FAA maintenance inspector, the number of flights may be determined by multiplying each airplane's hours time-in-service by two. f) The modifications specified by paragraphs a), b) and (c), or d) of this AD are no longer applicable when British Aerospace Modification No. 5115 is installed. g) Aircraft may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. h) An equivalent method of compliance with this AD if used must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on April 14, 1983.
92-06-11: 92-06-11 BEECH: Amendment 39-8191; Docket 91-CE-78-AD. Applicability: The following Beech model airplanes that are equipped with Collins DPU-85N Display Processor Units (DPU) and Collins MPU-85N Multifunction Processor Units (MPU), certificated in any category: Model Serial Numbers B200 BB-1349 through BB-1409 300 FA-205 through FA-217 B300 FL-1 through FL-60 B300C FM-1 through FM-3 1900C UC-73, UC-139 and UC-147 Compliance: Required as indicated, unless already accomplished. To prevent incorrect pilot decisions based on undetected erroneous attitude information displayed by the Collins EFIS-85/86B (14) system or undesired autopilot movement of the airplane, accomplish the following: (a) Within the next 50 hours time-in-service (TIS) after the effective date of this AD, modify the airplane autopilot/flight director system wiring in accordance with the instructions in Beech Service Bulletin (SB) No. 2423, dated December 1991. (b) Within the next 6 calendar months after the effective date of this AD, modify the hardware of the Collins EFIS-85/86B (14) system, Collins DPU-85N Display Processor Unit, and Collins MPU-85N Multifunction Processor Unit in accordance with the instructions in Collins SB DPU-85N-34-51 and SB MPU-85N-34-51, both dated June 6, 1991. (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) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. (e) The modifications required by this AD shall be done in accordance with Beech Service Bulletin No. 2423, dated December 1991, and Collins Service Bulletin DPU-85N-34-51 and Collins Service Bulletin MPU-85N-34-51, both dated June 6, 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 the Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201-0085; or Rockwell International, Collins General Aviation Division, 400 Collins Road, NE; Cedar Rapids, Iowa 52498. 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, 1100 L Street, NW; Room 8401, Washington, DC. (f) This amendment (39-8191) becomes effective on March 30, 1992.
99-15-09: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-600 series airplanes. This action requires revising the Airplane Flight Manual (AFM) to prohibit operation of the airplane under certain conditions; repetitive inspections of the tab mast fittings of the elevator tab assemblies to detect cracking; an elevator tab freeplay check; and corrective actions, if necessary. This AD also requires installing an additional fastener on the elevator tab mast fitting, which terminates the AFM revision and extends certain repetitive inspections. This AD also requires replacement of the elevator tab mast fitting with a new, improved fitting, which constitutes terminating action for the requirements of this AD. 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 actionsspecified 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.