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2002-01-16: This amendment supersedes Airworthiness Directive (AD) 86-24-11 and AD 86-25-04, which require you to incorporate, into the Limitations Section of the pilot's operating handbook and airplane flight manual (POH/AFM) of Fairchild Aircraft, Inc. (Fairchild Aircraft) SA226 and SA227 series airplanes, procedures for preventing an engine flameout while in icing conditions. This AD retains the POH/AFM requirements from the above-referenced AD's and requires a modification to the torque sensing system to allow the igniters to automatically turn on when an engine senses low torque. This AD is the result of two instances of a dual engine flameout on the affected airplanes. When the torque sensing system modification is incorporated, the POH/AFM requirements are no longer necessary. The actions specified by this AD are intended to prevent a dual engine flameout on the affected airplanes by providing a system that automatically turns on the engine igniters when low torque is sensed. A dual engine flameout could result in failure of both engines with consequent loss of control of the airplane.
84-09-51 R1: 84-09-51 R1 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-4903. Applies to Lockheed Model L-1011-385 series airplanes, certificated in all categories. Compliance required as indicated, unless previously accomplished. To prevent depressurization of the aircraft due to the failure of the negative pressure relief valve, accomplish the following: A. Within 100 flight hours after the receipt of this AD, unless previously accomplished, inspect the negative pressure relief valve mounting adapters on each aircraft identified as follows: 1. Lockheed Serial Numbers 1175 through 1250. 2. Lockheed Serial Numbers 1001 through 1174. Aircraft which have had either of the two negative pressure relief valve adapters replaced by the operator. B. Inspection and repair procedures: 1. After instituting all the preliminary safety precautions, gain access to aft side of the pressure bulkhead through fire bottle inspection panel 317 BB (see Lockheed Maintenance Manual MM6-40-00), and determine type of adapter installed. If adapter does not have circumferential spin marks and the serial number is 229 or lower, no further inspection is required. If serial number is 230 or higher, or spin marks are found, then additional inspection must be performed. NOTE: Adapter serial numbers are identified by the last three digits in the serial number block on the decal located at the 12 o'clock position aft of the pressure bulkhead. Spin marks are best determined by shining indirect light, such as flash light, on the adapter side wall. 2. On aircraft requiring further inspection: a. Gain access to the forward side of the negative pressure relief valve (see Lockheed Maintenance Manual MM25-42-00). b. On the forward side of the pressure bulkhead, LBL/RBL20, WL300, visually inspect the adapter part of each negative pressure relief valve in the area of the flange radius. Clean the area with solvent prior to inspection. 3. If a crack is found, replace or repair the adapter before the next flight. Replacement adapter must be inspected prior to installation and is subject to the requirements of this AD. 4. Installation of repair or reinforcement clips for adapter: Install 300 series stainless steel clips, either 0.050 1/4 hard or 0.040 1/2 hard material. Dimensions are 0.95-inch wide with 0.89 and 1.50-inch flanges, and 85 degrees bend, with 0.12-inch bend radius. Install clips using existing fastener holes through the aft pressure bulkhead and two additional fasteners through the side of the adapter, install one at minimum of 0.34 inches and the other at minimum of 0.95 inches from end of long leg of clip. Fasteners are to be MS20470AD5 rivets, or NAS1398M5 rivets, or structural equivalent. a. For repair of a cracked adapter, stop-drill crack 1/4-inch diameter, and install a minimum quantity of 40 clips per adapter, using 2 of each 3 attachments through the pressure bulkhead. b. For reinforcement ofuncracked adapters, install a minimum quantity of 20 clips per adapter, using 1 of each 3 attachments through pressure bulkhead. 5. If no cracks are found, repeat the inspection per B.1. through B.4., above, at intervals not to exceed 25 landings. 6. Replace or reinforce the adapter within 350 flight hours after the last inspection or after the receipt of this AD, whichever is later. 7. With adapter reinforced with the 20 clips per paragraph B.4.b., above, the reinspection intervals may be extended from 25 landings to 1000 landings. 8. For adapters repaired with the 40 clips per paragraph B.4.a., above, reinspect from the aft side for cracks at the aft fastener, common to the clip and adapter, every 500 landings. 9. Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. This amendment becomes effective September 11, 1984, and was effective earlier to those recipients of telegraphic AD T84-09-51, dated April 19, 1984.
58-01-05: 58-01-05 LOCKHEED: Applies to All Models 49-46, 149, 649, 649A, 749, 749A, and 1049-54 Aircraft. Compliance required as indicated. As a result of cracks discovered in Lockheed 749A wing skin and stringers, the following inspections shall be accomplished on the various model aircraft as indicated, and if any cracks are discovered, they must be repaired prior to further operation. Any FAA/LAC approved repair may be used. Inspect and reinspect for cracks in the lower wing skin and stringers, left and right, from wing Station 125 through Station 215 between the front and rear beams. Inspections to be conducted at the following specified times and intervals using X-ray and visual, or visual means. The X-ray inspection method is recommended if equipment is available, since cracks under the stringers would be detected. I. For Models 649, 649A, 749, 749A and 1049-54: A. The first inspection should be performed before 20,500 hours have been accumulated on the aircraft. For aircraft on which inspections of STA 191 through 206 have already been made in accordance with AD 56-03-01, initial inspections of additional indicated areas need not be earlier than and may be correlated with reinspections required by B(1), B(2), B(3), and B(4). B. Reinspections must be accomplished in accordance with one of the following programs: (1) X-ray at 2,500 hours (maximum) intervals without opening the fuel tanks following the recommendations and technique outlined on Lockheed Sketch No. 101057 or a FAA/LAC approved equivalent. In addition to the X-ray inspection at this time, the bottom side of the wing skin must be visually inspected from front to rear beam beneath the nacelle to wing fillets on the inboard and outboard sides of No. 2 and No. 3 nacelles. This necessitates opening the kidney plate inspection holes in these fillets and/or removal of the tail cone assembly. See Lockheed Sketch No. 101057 for location of cracks which have previously been discovered; or (2) X-ray at 3,200 hours (maximum) intervals by opening the fuel tanks and following the technique outlined on Lockheed Sketches No. 101057 and No. 101058, or FAA/LAC approved equivalent. In addition to the X-ray inspections at 3,200 hours, aircraft with over 20,000 hours must be visually inspected at 200-hour (maximum) intervals as follows: Inspect the bottom side of lower wing skin, for leaks resulting from cracks, from front beam to rear beam between W.S. 125 and W.S. 145 and between W.S. 191 and W.S. 215. This necessitates opening the kidney plate inspection holes in the nacelle to wing fillets on the inboard and outboard sides and/or removal of the tail cone assembly of nacelles No. 2 and No. 3. This area should be given special attention. If leaks are discovered and cracks suspected, tanks must be opened and stripped of sealant to visually inspect upper side of skin. Inspect the upper side of lower wing skin for cracks in the dry area from front beam to rear beam between W.S. 145 and W.S. 191. See Lockheed Sketches No. 101057 and No. 101058 for location of cracks which have been previously discovered; or (3) When X-ray equipment is not available, a visual inspection must be made at 800-hour (maximum) intervals after opening the fuel tanks and removing the sealant from the designated areas. (It should be noted that cracks under stringers cannot be detected by the visual inspection method); or (4) X-ray at 2,800 hours (maximum) intervals by opening the fuel tanks and following the technique outlined on Lockheed Sketches No. 101057 and No. 101058, or FAA approved equivalent. In addition to the X-ray inspections at 2,800 hours, visually inspect at 350-hour (maximum) intervals as follows: Inspect the bottom side of lower wing skin, for leaks resulting from cracks, from front beam to rear beam between W.S. 125 and W.S. 145 and between W.S. 191 and W.S. 215. This necessitates opening the kidney plate inspection holes in the nacelle to wing fillets on the inboard and outboard sides and/or removal of the tail cone assembly of nacelles No. 2 and No. 3. This area should be given special attention. If leaks are discovered and cracks suspected, tanks must be opened and stripped of sealant to visually inspect upper side of skin. Inspect the upper side of lower wing skin for cracks in the dry area from front beam to rear beam between W.S. 145 and W.S. 191. See Lockheed Sketches No. 101057 and No. 101058 for location of cracks which have been discovered previously. C. The reinspections required as per paragraphs B(1), B(2), B(3), or B(4) may be discontinued when permanent reinforcement per Lockheed Drawing No. 550236 has been accomplished, except that: In the area from W.S. 125 to W.S. 191 where the size and kind of material remains unchanged (i.e., the old material is merely replaced with new) the reinspection program noted above must be reinstated not later than 20,000 hours after rework. D. Lockheed Drawing Nos. 11755, 490668, 492806, and 493312, describe approved permanent repairs for individually affected areas in which cracks have been previously discovered. Reinspections in the area between W.S. 191 and W.S. 215 may be discontinued if permanent repair is made per Lockheed Drawing No. 11755. The reinspection program must be reinstated not later than 20,000 hours after rework is accomplished in the individually affected areas per drawing numbers 490668, 492806, or 493312. E. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. II. For Models 49-46 and 149: A. The first inspection should be performed before 25,500 hours have been accumulated on the aircraft. For aircraft on which inspections of STA 191 through 206 have already been made in accordance with AD 56-03-01, initial inspections of additional indicated areas need not be earlier than and may be correlated with reinspection required by B(1), B(2), B(3), and B(4). B. Same as paragraph I.B. 1. Same as paragraph I.B.(1). 2. Same as paragraph I.B.(2) (except substituted 25,000 hours for 20,000 hours). 3. Same as paragraph I.B.(3). 4. Same as paragraph I.B.(4). C. The reinspections required as per paragraphs B(1), B(2), B(3), or B(4) may be discontinued when permanent reinforcement per Lockheed Drawing No. 550236 has been accomplished, except that: In the area from W.S. 125 to W.S. 191 where the size and kind of material remains unchanged (i.e., the old material is merely replaced with new) the reinspection program noted above must be reinstated not later than 25,000 hours after rework. D. Lockheed Dwg. Nos. 11755, 490191, 490668, 492806, and 493312 describe approved permanent repairs for individually affected areas in which cracks have been previously discovered. Reinspections in the area between W.S. 191 and W.S. 215 may be discontinued if permanent repair is made per Lockheed Drawing No. 11755 or 490191. The reinspection program must be reinstated not later than 25,000 hours after rework is accomplished in the individually affected areas per drawing Nos. 490668, 492806, or 493312. E. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. This supersedes AD 56-03-01. Revised September 26, 1963.
2017-18-17: We are superseding Airworthiness Directive (AD) 2004-23-20, which applied to certain Airbus Model A300, A300 B4-600, and A300 B4- 600R series airplanes; and Model A300 F4-605R and A300 C4-605R Variant F airplanes. AD 2004-23-20 required, for certain airplanes, repetitive inspections for cracking around certain attachment holes, installation of new fasteners for certain airplanes, and follow-on corrective actions if necessary. AD 2004-23-20 also required modifying certain fuselage frames, which terminated certain repetitive inspections. This new AD reduces certain compliance times, expands the applicability, and requires an additional repair on certain modified airplanes. This AD was prompted by a report indicating that the material used to manufacture the upper frame feet was changed and negatively affected the fatigue life of the frame feet. We are issuing this AD to address the unsafe condition on these products.
2017-19-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 727 airplanes. This AD was prompted by analysis of the cam support assemblies of the main cargo door (MCD) that indicated the repetitive high frequency eddy current (HFEC) inspections required by the existing maintenance program are not adequate to detect cracks before two adjacent cam support assemblies of the MCD could fail. This AD requires repetitive ultrasonic inspections for cracking of the cam support assemblies of the MCD and replacement if necessary. We are issuing this AD to address the unsafe condition on these products.
90-15-04: 90-15-04 BRITISH AEROSPACE: Amendment 39-6652. Docket No. 90-NM-41-AD. Applicability: Model BAC 1-11 200 and 400 series airplanes, pre-modification PM5384, certificated in any category. Compliance: Required within 2,400 hours time-in-service or two years after the effective date of this AD, whichever occurs first, unless previously accomplished within the past 2,400 hours time-in-service or within the past two years; and thereafter at intervals not to exceed 4,800 hours time-in-service or four years, whichever occurs first. To prevent tailplane trim gearbox oil from being contaminated with water, accomplish the following: A. Remove the tailplane trim gearbox from the airplane, drain the oil, flush and refill with clean oil, and replace the filler plug and wire lock, in accordance with paragraph 2.2 of British Aerospace Alert Service Bulletin 27-A-PM5384, Issue 1, dated July 24, 1989. Reinstall the gearbox in the airplane and test in accordance with Maintenance Manual Chapter 27-40. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM-113, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Standardization Branch, ANM-113. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041-0414. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6652, AD 90-15-04) becomes effective on August 14, 1990.
2001-17-26 R1: This document corrects and clarifies information in an existing airworthiness directive (AD) that applies to certain Raytheon Model DH.125, HS.125, BH.125, and BAe.125 (U-125 and C-29A) series airplanes; Model Hawker 800, Hawker 800 (U-125A), Hawker 800XP, and Hawker 1000 airplanes. That AD currently requires an inspection for cracking or corrosion of the cylinder head lugs of the main landing gear actuator and follow-on/corrective actions. This document corrects and clarifies the affected airplane serial numbers. This correction is necessary to ensure that operators do not misinterpret which airplanes are subject to the requirements of this AD. The incorporation by reference of certain publications listed in the regulations was approved previously by the Director of the Federal Register as of October 3, 2001 (66 FR 45575, August 29, 2001).
85-25-03: 85-25-03 SIKORSKY AIRCRAFT: Amendment 39-5172. Applies to Model S-64E helicopters, certificated in any category. Compliance is required as indicated, unless already accomplished. To prevent operation with a cracked main rotor head torque tube inner bracket, accomplish the following: (a) Prior to the first flight of each day, after the effective date of this AD, visually inspect with a 10-power or higher magnifying glass the main rotor head torque tube inner bracket assembly, Part Number S1510-21332-0, for cracks and/or corrosion in accordance with Section 2, Paragraph A, of Sikorsky Alert Service Bulletin (ASB) No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (b) If the torque tube inner bracket assembly is cracked, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (c) If the torque tube inner bracket assembly is corroded, determine the extent and limits of the corrosion prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB-No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. If the extent or limits of the corrosion are exceeded, replace with a serviceable torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph A, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. Otherwise, rework the torque tube inner bracket assembly prior to further flight in accordance with Section 2, Paragraph B, of Sikorsky ASB No. 64B10-4A, dated July 17, 1985, or an FAA-approved equivalent. (d) Aircraft may be ferried in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the AD can be accomplished. (e) Upon request, an alternative means of compliance with the requirements of this AD which provide an equivalent level of safety maybe used when approved by the Manager, Boston Aircraft Certification Office, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone (617) 273-7112. Sikorsky ASB No. 64B10-4A, dated July 17, 1985, identified and described in this directive is 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 this document from the manufacturer may obtain copies upon request to Sikorsky Aircraft Division, United Technologies Corporation, North Main Street, Stratford, Connecticut 06601. These documents may also be examined at the Office of the Regional Counsel, FAA, Southwest Region, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment becomes effective December 30, 1985.
91-05-21: 91-05-21 GENERAL ELECTRIC COMPANY: Amendment 39-6900. Docket No. 90-ANE- 28. Applicability: General Electric Company (GE) CF6-80C2A5 and CF6-80C2B6 engines, Serial Numbers (S/N) 690-101 through 690-369, and S/N 695-101 through 695-423; and CF6- 80C2B6F and CF6-80C2D1F engines, S/N 702-101 through 702-470, and S/N 703-101 through 703-136, which do not incorporate the increased shroud cooling design features of paragraph (b) of this AD, installed on, but not limited to, Airbus A300, Boeing 767, and McDonnell Douglas MD-11 aircraft. Compliance: Required as indicated, unless previously accomplished. To prevent high pressure turbine (HPT) failure and possible aircraft damage, accomplish the following: (a) Borescope inspect engines in accordance with sections 2.B., 2.C., and 2.D of the Accomplishment Instructions in GE CF6-80C2 Service Bulletin (SB) 72-473, Revision 1, dated September 21, 1990, unless previously accomplished, according to the following schedule based upon cycles since new (CSN) on the effective date of this AD: (1) Inspect within 10 cycles in service (CIS) after the effective date of this AD or prior to accumulating 520 CSN, whichever occurs later, for CF6-80C2A5 and CF6- 80C2B6 engines, S/N 690-101 through 690-369, and S/N 695-101 through 695-350; and CF6- 80B6F engines, S/N 702-101 through 702-315, and S/N 702-317 through 702-321. (2) Inspect within 10 CIS after the effective date of this AD or prior to accumulating 1,250 CSN, whichever occurs later, for CF6-80C2A5 and CF6-80C2B6 engines, S/N 695-351 through 695-423; and CF6-80C2B6F and CF6-80C2D1F engines, S/N 702-316, 702-322 through 702-470, and S/N 703-101 through 703-136. (3) Remove from service or reinspect in accordance with the following: (i) Remove from service prior to further flight, engines with at least one Category 4 shroud. (ii) Remove from service within 25 hours time in service (TIS) since last inspection (SLI), engines with no Category 4 shrouds, but at least one Category 3 shroud. (iii) Borescope reinspect at intervals not to exceed 125 hours TIS SLI, engines with no Category 3 or 4 shrouds, but at least one Category 2 shroud. (iv) Borescope reinspect at intervals not to exceed 300 hours TIS SLI, engines with no Category 2,3, or 4 shrouds, but at least one Category 1 shroud. (v) Borescope reinspect at intervals not to exceed 520 CIS SLI, engines with no Category 1, 2, 3, or 4 shrouds. (b) Replace the HPT stator stage one shroud support assemblies, Part Numbers (P/N) 9381M61G06 and 9381M61G07; the HPT stator support hanger assemblies, P/N 9397M73G05 and 9397M73G06; and the HPT stage one shrouds, P/N 1333M75P05, 1333M75P06, 1333M75P07, 1333M75P08, 1333M75P09, and 1333M75P10 in accordance with the Accomplishment Instructions in GE CF6-80C2 SB 72-474, Revision 1, dated December 11, 1990, at the next HPT module exposure after the effective date of this AD, but prior to December 31, 1994. (c) For the purpose of this AD, HPT module exposure is defined as the separation of the HPT stator support case from the compressor rear frame. (d) For the purpose of this AD, the shroud Categories are defined in GE CF6-80C2 SB 72-473, Revision 1, dated September 21, 1990. (e) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (f) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. The borescope inspections and installation of improved HPT hardware shall be done in accordance with the following documents: Document Page Revision Date CF6-80C2SB 72-473 5, 6, 7, Original 7/3/90 10-23 1, 2, 3, Rev. 1 9/21/90 4, 8, and 9 CF6-80C2 SB 72-474 1-24 Rev. 1 12/11/90 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 General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the Office of the Assistant Chief Counsel, FAA, New England Region, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. This amendment (39-6900, AD 91-05-21) becomes effective on March 27, 1991.
78-07-01: 78-07-01 CESSNA: Amendment 39-3163. Applies to TP206 (Serial Numbers P206-0191 thru P20600647), TU206 (Serial Numbers U206-0487 thru U20603693), T207 (Serial Numbers 20700001 thru 20700378) and T210 (Serial Numbers T210-0001 thru T210-0454 and Serial Numbers 21059200 thru 21061758) series airplanes certificated in all categories. To preclude engine oil pump failure due to contamination by the turbocharger thrust bearing anti-rotation pins and failure of the turbocharger shaft, within the next 25 hours' time-in-service after the effective date of this AD, accomplish the following: A. Check the turbocharger nameplate or aircraft permanent maintenance records to determine if the turbocharger serial number is prefixed by any of the following letter combinations: EF EFR FA FAR FH FHR EG EGR FB FBR FI FIR EH EHR FC FCR FJ FJR EI EIR FD FDR FK FKR EJ EJR FE FER FL FLR EK EKR FF FFR GA GAR EL ELR FG FGR GB GBR B. If the serial number on the turbocharger nameplate is not prefixed by any of the letter combinations set forth in Paragraph A, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. C. If the serial number on the turbocharger nameplate is prefixed by any of the letter combinations set forth in Paragraph A, check the aircraft permanent maintenance records to determine whether, when complying with AD 77-06-02, the turbocharger center housing was replaced by a mechanic or repair agency in accordance with Cessna Service Kit SK 210-75, dated February 24, 1977 (Reference Cessna Service Letter SE 77-3, Supplement #2 dated February 24, 1977) or by the turbocharger manufacturer (AiResearch). D. If the turbocharger center housing was replaced by AiResearch, make an entry in the aircraft permanent maintenance records indicating this finding and no further action is required. E. If the turbocharger center housing was replaced by a mechanic or repair agency using instructions in Cessna Service Kit SK 210-75, visually inspect the turbocharger for signs of damage and proper compressor wheel attachment in accordance with Cessna Service Kit SK 210-78 dated November 15, 1977, or later revision (Ref. Cessna Service Letter SE 77-42 dated December 2, 1977, or later revisions) for damage which may have resulted from incomplete compressor wheel locknut torquing procedures prescribed in Cessna Service Kit SK 210-75. (1) If visual signs of damage are evident, return the turbocharger to AiResearch in accordance with Cessna Service Kit SK 210-78. (2) If no visual signs of damage are present but the compressor wheel attachment does not meet the criteria set forth in Cessna Service Kit SK 210-78, conduct additional inspections prescribed therein. Units found acceptable as a result of this inspection may be returned to service after reassembly per this kit. Return unacceptable units to AiResearch in accordance with instructions in Cessna Service Kit SK 210-78. (3) If no visual signs of damage are found and the compressor wheel attachment meets the criteria set forth in Cessna Service Kit SK 210-78, reassemble and identify the turbocharger in accordance with Cessna Service Kit SK 210-78. F. If the turbocharger center housing has not been replaced in accordance with AD 77-06-02, replace the turbocharger center housing in accordance with Cessna Service Kit SK 210-75B dated October 27, 1977, or later revision incorporating a compressor wheel seating procedure. (Ref. Cessna Service Letter SE 77-42.) G. Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. H. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment supersedes Amendment 39-2853 (42 FR 15894), AD 77-06-02. This amendment becomes effective April 6, 1978.
2001-26-53: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 2001-26-53, which was sent previously to all known U.S. owners and operators of Eurocopter France (ECF) Model AS350B, B1, B2, B3, BA, D, and AS355E helicopters by individual letters. This AD requires, before further flight, removing certain serial-numbered servocontrols. This AD is prompted by a report of manufacturing defects in a batch of main servocontrol rods. The actions specified by this AD are intended to prevent failure of a main servocontrol in the flight control system and subsequent loss of control of the helicopter.
2017-18-18: We are adopting a new airworthiness directive (AD) for all Airbus Model A350-941 airplanes. This AD requires repetitive on-ground power cycles to reset the internal timer. This AD was prompted by the in-service loss of communication between some avionics systems and the avionics network. We are issuing this AD to address the unsafe condition on these products.
89-25-08: 89-25-08 BEECH: Amendment 39-6410. Applicability: Models 65 (Serial Numbers (S/N) L-1, L-2, L-6, LF-7 through LF-76, and LC-1 through LC-180); 65-80 and 65-A80 (S/N LD-1 through LD-244); 65-A80 (S/N LD-245 through LD-269) when Beech Modification Kit No. 80-4004-1 or -3 is installed; and 65-B80 (all S/N) airplanes certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To detect possible fatigue cracking of the wing main spar lower cap and associated structure, accomplish the following: (a) Within the next 200 hours time-in-service (TIS) after the effective date of this AD, or upon accumulating 3000 hours TIS on Models 65-80 and 65-A80 airplanes, or upon accumulating 5000 hours TIS on Models 65 and 65-B80 airplanes, whichever occurs later, unless previously accomplished per AD 70-25-01, Amendment 39-1609, and thereafter at intervals not to exceed 1000 hours TIS (except as provided in paragraph (b) below) after the initial inspection, inspect the wing lower forward spar attach fittings, center section and outboard wing spar caps adjacent to the attach fittings by visual, fluorescent penetrant and eddy current methods as specified in the applicable section of Beech Structural Inspection and Repair Manual (SIRM), P/N 98-39006, Revision A4, dated May 1, 1987. NOTE 1: Beech offers a two-day training course free of charge to qualified personnel who have prior knowledge of eddy current inspection techniques. A listing of Beech Corporate maintenance facilities may be obtained from the sources contained in paragraph (e) of this AD. A listing of other facilities employing qualified inspectors is not available. (b) At each inspection required by paragraph (a) above, inspect any reinforcing strap installed per Supplemental Type Certificate (STC) SA1583CE for proper tension and condition in accordance with Aviadesign Engineering Order E.O. B-8001, Issue 3, dated May 30,1985. Correct any discrepancy prior to further flight. For airplanes equipped with STC SA1583CE and inspected in accordance with paragraph (a) above, the repetitive inspection interval of 1000 hours TIS in paragraph (a) above may be extended to 3000 hours TIS. (c) If any crack is found in a main spar lower cap or fitting, prior to further flight repair or replace the defective part using the instructions and limitations specified in the SIRM or other FAA approved instructions provided by Beech Aircraft Corporation. (d) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (e) An alternate method of compliance or adjustment of the initial or repetitive compliance times which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209; Telephone (316) 946-4400. NOTE 2: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the documents referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or Western Aircraft Maintenance, 4444 Aeronca Street, Boise, Idaho 83705; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This AD supersedes AD 70-25-01, Amendment 39-1609. This amendment (39-6410, AD 89-25-08) becomes effective on January 4, 1990.
2002-01-12: This amendment adopts a new airworthiness directive (AD), that is applicable to General Electric Company (GE) GE90 series turbofan engines. This amendment requires removing from service high pressure turbine (HPT) interstage seals, identified by GE as the pre-life-improved rotor (pre-LIR) configuration, and installing a new design, identified by GE as the life improved rotor (LIR) configuration seal. This amendment also requires a new lower life limit for the LIR configuration seal. This amendment is prompted by an uncontained engine failure which occurred during a factory development engine ground test. The actions specified by this AD are intended to prevent failure of the HPT interstage seal that could result in an uncontained engine failure and damage to the airplane.
87-11-05 R4: 87-11-05 BOEING VERTOL COMPANY (VERTOL) AND KAWASAKI HEAVY INDUSTRIES, LTD.: Amendment 39-5591. Applies to Model 107-II helicopters and Kawasaki Model KV107-II and IIA helicopters, certificated in any category. Compliance is required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. (a) To prevent hazards in flight associated with the fatigue failure of life limited components, the following retirement lives are imposed: LIFE LIMIT COMPONENT PART PART NUMBER HOURS APPENDIX Alum Synch 107D3141 9,940 Shaft Assembly 107D3341 9,940 Steel Synch 107D3140 7,850 Shaft Assembly 107D3340 7,850 A15D3840-1 24,750 Synchronizing 107D3154-2 500 No. 1 Shaft Splined 107D3154-3, -4 3,000 Adapter 107D3144-1 14,615 Mix Box 107D2066-10 11,075 Collector Gear thru -28 Fwd Rotor Shaft A02D1259-3 2,100 and Carrier Assy A02D1269-1 2,100A07D1269-1 2,100 A02D1259-3SP 13,580 A02D1269-1SP 13,580 A07D1269-1SP 13,580 A02D1269-2 13,580 Quill Shaft 107D2067-1, 5,620 No. 2 -3, -5 Aft Transmission 107D2419 or 8,700 Planet Carrier A02D2419 Aft Rotor Drive Shaft upper extension A02D3147 7,200 Aft Rotor Shaft 107D3171-2, -3, -4, -5 15,860 Lower Section A2D3171-3, -4, -6 15,860 Aft Rotor Shaft 107D3151 19,140 Center Section A02D3151 19,140 Fwd Hub 107R2550 2,500/1,200 No. 3 Aft Hub 107R2550 10,460/3,000 Fwd and Aft 107R2551-1 or 6,610 Connecting Link A02R2551-8 Fwd and Aft Hub A02R2589-3 25,140 Horizontal Hinge Pin 107R2589-3 25,140 Fwd and Aft Lag Damper Piston 107R2601-1 4,670 Rod End APPENDICES: No. 1. Synchronizing shaft splined adapter, Part Number (P/N) 107D3154-2, shall be repetitively inspected as specified in Boeing Vertol (BV) Service Bulletin 107-116 (R-1), Revision B, dated February 21, 1983, or FAA-approved equivalent. P/N 107D3154-3 and -4 require repetitive visual and dye check inspections as specified in the BV107-6 Maintenance Schedule, Section 2, dated February 7, 1986, and Service Bulletin 107-116 (R-1)B. No. 2. Quill shaft, P/N 107D2067-1, has a life limit of 120 hours unless the modifications specified in paragraph (e) of AD 63-24-04 revised October 15, 1964, Amendment 648 of Part 507 (28 FR 12614) as amended by Amendment 821 of Part 507 (29 FR 14169) are accomplished. (Also, reference Boeing Service Bulletins 107-113, Revision A, dated November 22, 1963, and 107-182, Revision B, dated July 26, 1965.) No. 3. The life limit for the forward rotor hub spline is 2,500 hours. After 2,500 hours of service, the hub is to be magnetic particle inspected. If found to be free of cracks and not having a wear step on the profile face of the spline in excess of 0.002 inch, the hub may be inverted and installed on the aft rotor head for an additional 3,000 hours of use. On completion of the additional 3,000 hours, the hub is to be retired from service. The life limit for the aft rotor hub spline is 10,460 hours. After 10,460 hours of use, the hub is to be magnetic particle inspected. If found to be free of cracks and not having a wear step on the profile face of the spline in excess of 0.002 inch, the hub may be inverted and installed on the forward rotor head for an additional 1,200 hours of use. On completion of the additional 1,200 hours, the hub is to be retired from service. (An acceptable procedure for measuring the 0.002-inch step wear is contained in Boeing Vertol Overhaul Manual 107-5, Chapter 60-20-1.) Inverting and switching of main rotor hubs are permitted to the life limit stated above any number of times provided accurate records are maintained of the total hours the hubs are installed on an aft head and on a forward head. (b) An alternate method of compliance or adjustment of the compliance time, which provides an equivalent level of safety, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region, 181 South Franklin Avenue, Room 202, Valley Stream, New York 11581. (c) Aircraft may be ferried to a base in accordance with FAR Sections 21.197 and 21.199 where compliance can be accomplished. The repair and inspection procedures in Boeing Vertol (BV) Service Bulletin (SB) 107- 113, Revision A, dated November 22, 1963; BV SB 107-116 (R-1), Revision B, dated February 21, 1983; BV SB 107-182, Revision B, dated July 26, 1965, and BV 107-6 Maintenance Schedule, Section 2, temporary Revision 31, dated February 7, 1986, incorporated by reference in this directive were approved by the Director of the FEDERAL REGISTER pursuant to 5 U.S.C. 552(a)(1). Copies may be obtained from Boeing Vertol Company, Boeing Center, P.O. Box 16858, Philadelphia, Pennsylvania 19142. These documents may be examined at the Office of the Regional Counsel, FAA, Southwest Region 4400 Blue Mound Road, Fort Worth, Texas 76106 or the Office of the Federal Register, 1100 L Street, NW., Room 8401, Washington, D.C. This amendment supersedes the following: Amendment 656 of Part 507 (28 FR 13931) as amended by Amendment 39-107 (30 FR 8963), AD 63-26-04; Amendment 39-2993 (42 FR 38803), AD 77-16-03; Amendment 39-3746 (45 FR 25050), AD 80-08-11; Amendment 39-4522 (47 FR 57484), AD 82-27-06; and Amendment 39-4757 (48 FR 50069), AD 83-22-02. This amendment, 39-5591, becomes effective May 14, 1987.
70-26-02: 70-26-02 WOODWARD: Amdt. 39-1129. Applies to Woodward propeller governors of the following listed models having serial numbers below 992601 which were manufactured prior to 1970 used on single, reciprocating engine aircraft: Woodward Governor Models 210452, A210452, B210452, C210452, D210452, E210452, F210452, G210452, H210452, J210452, K210452, L210452, M210452, P210452, 210453, 210458, 210460, B210460, 210462, A210462, 210472, and C210472. Date of manufacture can be determined from a decal attached to the governor body which shows the quarter and the year. Example: "1Q70" indicates first quarter 1970. Compliance: Required within the next 50 hours' time in service after the effective date of this AD, unless previously accomplished. To prevent loss of propeller control accomplish the following or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. A. Inspect the propeller governor lever arm for security andengagement on the speed control shaft as follows: 1. Inspect axial security by applying, alternately, in both directions, a manual force of 5 to 10 pounds to the lever arm directly in line with the axis of the shaft. Do not mistake end play of the shaft in the governor cover for a loose lever. 2. Inspect rotational security by observing the arm and shaft while the cockpit propeller control is moved from full increase to full decrease and back to full increase RPM positions. 3. Inspect axial location of lever arm and offset lever arm extension on the shaft. On those installations which use an offset extension which bolts to the outboard face of the lever arm and has an alignment hole for locating on the shaft, the shaft must protrude through the full thickness of the extension. When no extension is used the shaft must protrude beyond the lever arm by at least .050 inch. B. If the inspections in accordance with Paragraphs A1 and A2 disclose movement of the leverarm relative to the shaft or if the location of the arm or extension do not meet the limits defined in Paragraph A3 proceed as follows: 1. Remove arm from shaft and inspect serrations on both parts for wear and damage. Before removing arm, provisions, such as match-marking, should be made to assure reinstallation in the same circumferential location of the shaft. Later design shafts have a retaining ring and groove at the end of the shaft serrations to provide positive retention of the lever arm. To remove the arm from these shafts move the arm toward the governor cover until the retainer is exposed, then remove retainer. 2. If the serrations are damaged or excessively worn, replace the governor with a serviceable unit. 3. If the serrations are in satisfactory condition replace the lever arm on the shaft in its original circumferential location. If retainer ring was removed pursuant to Paragraph B1 reinstall it. Position axially on shaft to maintain .020 to .045 inch clearance between bottom side of lever arm and the top of governor cover at the maximum RPM setting. Torque the clamping screw in the lever arm to 33 to 38 inch pounds. (This value is specified in Woodward Overhaul Bulletin 33017A.) Recheck security per Paragraph A and if tight, safety the clamping screw with AMS 5685 .024-.026 wire or equivalent, taking care that the wire will not interfere with the aircraft manufacturer's lever arm extension. 4. Assure security of aircraft linkage to governor. If any aircraft linkage settings were changed as a result of work performed above, check rigging in accordance with the aircraft manufacturer's instructions. Woodward FAA-approved Service Bulletin No. 33534 or later FAA-approved revisions pertain to this subject. NOTE: The above listed governors may be installed on the following single, reciprocating engine aircraft but this listing is not all inclusive: BEECH Models E33, F33, E33A, E33C, F33A, F33C, 35-33, 35-A33,35-B33, 35-C33, 35-C33A, H35, J35, K35, M35, N35, P35, S35, V35, V35A, V35B, 36 and A36 airplanes. BELLANCA Models 14-19-3A, 17-30, 17-30A, 17-31, 17-31A, 17-31TC, 17-31ATC airplanes. CESSNA Models 180, 180A, 180B, 180C, 180D, 180E, 180F, 180G, 180H, 182, 182A, 182B, 182C, 182D, 182E, 182F, 182G, 182H, 182J, 182K, 182L, 182M, 182N, 185, 185A, 185B, 185C, 185D, 185E, A185E, 188, A188, 188A, A188A, 206, U206, P206, U206A, P206A, P206B, TU206A, TU206B, TP206A, TP206B, U206B, P206C, TP206C, P206D, TP206D, P206E, TP206E, U206C, TU206C, U206D, TU206D, U206E, TU206E, 207, T207, 210B, 210C, 210-5(205), 210-5A(205A), 210D, 210E, T210F, 210F, T210G, T210H, 210G, 210H, T210J, 210J, 210K and T210K airplanes. MAULE Models M-4-210, M-4-210C, M-4-210S, M-4-210T, M-4-220, M-4-220C, M-4-220S, M-4-220T and M-4-180 airplanes. MOONEY Models M20C and M20D airplanes. NAVION H Model airplanes. This amendment becomes effective December 27, 1970.
2025-09-13: The FAA is adopting a new airworthiness directive (AD) for Airbus Helicopters Model AS350B, AS350BA, AS350B1, AS350B2, AS350B3, AS350D, AS355E, AS355F, AS355F1, AS355F2, AS355N, and AS355NP helicopters. This AD was prompted by a report of a sliding door that was locked in the open position detaching from the helicopter during flight. This AD requires modifying certain upper rail rollers, installing a label on each sliding door, and prohibits installing affected upper rail rollers or a door having an affected upper rail roller. These actions are 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.
2017-18-11: We are adopting a new airworthiness directive (AD) for certain Textron Aviation Inc. Model 390 airplanes (type certificate previously held by Beechcraft Corporation). This AD was prompted by reports of hydraulic fluid loss from the engine driven pumps (EDPs) on three different airplanes. This AD requires an inspection to determine if an affected EDP is installed with replacement as necessary. We are issuing this AD to address the unsafe condition on these products.
2011-01-08: We are superseding an existing airworthiness directive (AD) that applies to the products listed above. This AD results 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 describes the unsafe condition as: Two reports have been received where, during inspection of the vertical stabilizer of F28 Mark 0100 aeroplanes, one of the bolts that connect the horizontal stabilizer control unit actuator with the dog-links was found broken (one on the nut side & one on the head side). In both occasions, the bolt shaft was still present in the connection and therefore the horizontal stabilizer function was not affected. If a single dog-link connection fails, the complete stabilizer load is taken up by the remaining dog-link connection. * * * To address and correct this unsafe condition EASA [European Aviation Safety Agency] issued AD 2007-0287[corresponding FAA AD 2008-22-14] that required a one-time inspection of the affected bolts, * * * and replacement of failed bolts with serviceable parts. EASA AD 2007-0287 also required the installation of a tie wrap through the lower bolts of the horizontal stabilizer control unit, to keep the bolt in place in the event of a bolt head failure. Recent examination revealed that the bolts failed due to stress corrosion, attributed to excessive bolt torque. Investigation of the recently failed bolts showed that the modification as required by AD 2007-0287 is not adequate. * * * * * Loss of horizontal stabilizer function could result in partial loss of control of the airplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
2011-01-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results 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 describes the unsafe condition as: During a routine maintenance check on an A300-600 aeroplane, the operator found the pitch uncoupling unit installed at an incorrect location. The pitch uncoupling unit was inverted with the rod assembly. After a complete inspection of all A300-600 aeroplanes of its fleet, the operator identified the same incorrect installation on another aeroplane. * * * * * This condition, if not detected and corrected, in combination with particular failure modes, could lead to loss of control of the aeroplane during the takeoff phase. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
2023-03-05: The FAA is superseding Airworthiness Directive (AD) 2022-09- 06, which applied to certain Airbus SAS Model A350-941 and -1041 airplanes. AD 2022-09-06 required revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations. This AD was prompted by a determination that new or more restrictive airworthiness limitations are necessary. This AD continues to require the actions in AD 2022-09- 06 and requires revising the existing maintenance or inspection program, as applicable, to incorporate additional new or more restrictive airworthiness limitations, as specified in a European Union Aviation Safety Agency (EASA) AD. The FAA is issuing this AD to address the unsafe condition on these products.
93-14-17: 93-14-17 AEROSPATIALE: Amendment 39-8641. Docket 93-NM-15-AD. Applicability: Model ATR72-100 and -200 series airplanes on which Modification 3196 has not been installed; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the wing spar fittings, accomplish the following: (a) Within 12 months after the effective date of this AD, perform a one-time detailed visual inspection of the fastener holes on the front and rear wing spar fittings to ensure that spotfacing of the fastener holes has been accomplished, in accordance with Aerospatiale Service Bulletin ATR72-57-1008, dated November 19, 1992. (1) If spotfacing of the fastener holes has been accomplished, no further action is required by this AD. (2) If spotfacing of the fastener holes has not been accomplished, prior to further flight, perform a one-time general visual inspection of the fastener holes for peening or cracks, in accordance with the service bulletin. (i) If no peening or crack is found, prior to further flight, install a shim and replace existing nuts with self-aligning nuts, in accordance with the service bulletin. (ii) If any peening or crack is found, prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (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 inspections, installation, replacement, and repair shall be done in accordance with Aerospatiale Service Bulletin ATR72-57-1008, dated November 19, 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 Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on September 16, 1993.
2017-18-14: We are superseding Airworthiness Directive (AD) 2015-02-22 for certain Rolls-Royce Corporation (RRC) model 250 turboprop and turboshaft engines. AD 2015-02-22 required repetitive visual inspections and fluorescent-penetrant inspection (FPIs) on certain 3rd- stage and 4th-stage turbine wheels for cracks in the turbine wheel blades. This AD requires repetitive visual inspections and FPIs of 3rd- stage turbine wheels while removing from service 4th-stage turbine wheels. We are also revising the applicability to remove all RRC turboprop engines and add additional turboshaft engines. This AD was prompted by our finding that it is necessary to remove the 4th-stage wheels at the next inspection. We are issuing this AD to address the unsafe condition on these products.
2002-01-10: This amendment adopts a new airworthiness directive (AD) that applies to certain Raytheon Aircraft Company (Raytheon) Beech Models 65-90, 65-A90, 65-A90-1, 65-A90-4, B90, C90, C90A, E90, and H-90 airplanes. This AD requires you to repetitively inspect the main landing gear upper torque knees and lower torque knees for evidence of fatigue cracks; and replace any torque knee with evidence of fatigue cracks. This AD is the result of reports of many incidents of main landing gear torque knees cracking or breaking on the above-referenced airplanes. The actions specified by this AD are intended to detect and replace cracked main landing gear torque knees, which could result in failure of the main landing gear and consequent loss of control of the airplane during takeoff, landing, or other ground operations.
89-22-01 R1 L: 89-22-01 R1 BELL HELICOPTER TEXTRON, INC. (BHTI): Priority Letter AD 89-22-01 as revised by Priority Letter AD 89-22-01 R1. Docket No. 89-ASW-54. Applicability: Model 206A, 206B, 206L, 206L-1, and 206L-3 helicopters, certificated in any category, with main rotor blades part numbers (P/N) 206-015-001-001, 206-015-001-103, 206-015-001-105, or 206-010-200-033 installed. Compliance: Required before further flight, unless already accomplished. To prevent failure and separation of the main rotor blades and subsequent loss of the helicopter, accomplish the following: (a) Visually inspect the Model 206A and 206B main rotor blades and determine if one of the following S/N blades is installed: TAC-0089, TAC-0542, TAC-0607, TAC-0614, TAC-0624, TAC-1643, TAC-1749, TAC-1776, TAC-1831, TAC-1911, TAC-1922, TAC-2399, TAC-2768, TAC-5742, TKK-9794, TKK-9883, or TKK-9933. If any one of these main rotor blades is installed, remove and replace with a serviceable part prior tofurther flight. (b) Visually inspect the Model 206L, 206L-1, and 206L-3 main rotor blades and determine if one of the following S/N blades is installed: T-92, T-245, T-417, TLY-0075, TLY- 0095, TLY-0764, TLY-0770, TLY-0973, TLY-1438, TLY-1619, TLY-1653, TLY-1697, TLY- 1766, TLY-1801, TLY-1858, TLY-1953, TLY-1984, TLY-2031, TLY-2039, TLY-2064, TLY- 2081, TLY-2148, TLY-2335, TLY-2337, TLY-2549, TLY-2603, TLY-2604, TLY-2625, TLY- 2633, TLY-2648, TLY-2745, TLY-2786, TLY-2951, or TLY-2954. If any one of these main rotor blades is installed, remove and replace with a serviceable part prior to further flight. NOTE: The serial number may be found on the Bell Helicopter data plate located on top of the blade at the root end and is also marked on the root end of the lower grip plate in the 1.5 inch radius. (c) If the serial number of the main rotor blade matches one listed in paragraph (a) or (b) of this AD, report the registration number and serial number of the affected helicopter and provide a copy of the parts tag with which the part was delivered, if available. This report is to be made to the Manager, Rotorcraft Certification Office, Southwest Region, Federal Aviation Administration, Fort Worth, Texas 76193-0170, telephone (817) 624-5170, within 10 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056.) (d) An alternate method of compliance which provides an equivalent level of safety, may be used if approved by the Manager, Rotorcraft Certification Office, Southwest Region, Federal Aviation Administration, Fort Worth, Texas 76193-0170, telephone (817) 624-5170. This Priority Letter AD (89-22-01 R1) issued on November 21, 1989, revises Priority Letter AD 89-22-01, issued on October 18, 1989, and becomes effective immediately upon receipt.