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94-24-01: This amendment supersedes Airworthiness Directive (AD) 87-04-04 and AD 89-16-02, which currently require the following on Jetstream Aircraft Limited (JAL) HP 137 Mk1, Jetstream series 200, and Jetstream Models 3101 and 3201 airplanes: repetitively inspecting the universal joints and universal rivets, replacing any damaged part, and limiting the in-service life of the torque tube shaft assembly. JAL has introduced an improved flap torque tube shaft assembly that includes universal joints that are not life limited. The Federal Aviation Administration's policy on aging commuter-class aircraft is to eliminate or, in certain instances, reduce the number of certain repetitive short-interval inspections when improved parts or modifications are available. This action requires installing this improved flap torque shaft assembly in place of the repetitive inspection and life limit requirements of the two existing AD's. The actions specified by this AD are intended to prevent failure of the flap torque assembly, which could result in asymmetric flap deployment and loss of control of the airplane.
85-22-05: 85-22-05 BEECH: Amendment 39-5146. Applies to Beech airplanes listed in Table I below, certificated in any category, upon accrual of five years in service. This AD does not apply to those airplanes in which bolts and nuts made of Inconel have been installed in the wing attachment joints that are specified in Table I below. 85-22-05 TABLE I BEECH MODEL (MILITARY MODEL) SERIAL NUMBER JOINTS (1) INSTRUCTION (2) 34C GP-1 and up LF,UF,UR,LR T-34C-1-0083 T-34C GL-1 and up LF,UF,UR,LR T-34C-0158R2 T-34C-1 GM-1 and up LF,UF,UR,LR T-34C-1-0083 50 H-1 thru H-11 LF* P/N 98-39006 50 (L-23A,U-8A) LH-1 thru LH-55 LF* See Note 3a B50 CH-12 thru CH-110 LF* P/N 98-39006 B50 (L-23B) LH-56 thru LH-95 LF* See Note 3b C50 CH-111 thru CH-360 LF* P/N 98-39006 D50 (L-23E, L-23G) DH-1 thru DH-154 LF* P/N 98-39006 D50A, D50B, D50C DH-155 thru DH-300 LF* P/N 98-39006 D50E, D50E-5990 DH-301 thru DH-347 LF* P/N 98-39006 E50 EH-1 thru EH-70 LF* P/N 98-39006 E50 (L-23D, U-8D) LH-96 and up LF* See Note 3b E50 (RL-23D, RU-8D) RLH-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) LHC-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) LHD-1 and up LF* See Note 3b E50 (RL-23D, RU-8D) RLHE-1 and RLHE-2 LF* See Note 3b E50 (RL-23D, RU-8D) LHE-3 and up LF* See Note 3b F50 FH-71 thru FH-96 LF* P/N 98-39006 G50 GH-94 thru GH-119 LF* P/N 98-39006 H50 HH-120 thru HH-149 LF* P/N 98-39006 J50 JH-150 thru JH-176 LF* P/N 98-39006 60, A60, B60 P-4 and up LF, UF, LR P/N 60-590001- 25A13 65 LC-1 thru LC-239 LF* P/N 98-39006 65 (L-23F, U-8F) L-1 thru L-6 LF* See Note 3b 65 (L-23F, U-8F) LF-7 and up LF* See Note 3b A65, A65-8200 LC-240 thru LC-335 LF* P/N 98-39006 65-80, 65-A80, 65-A80-8800 LD-1 thru LD-269 LF* P/N 98-39006 65-B80 LD-270 and up LF* P/N 98-39006 65-88 LP-1 thru LP-47 LF* P/N 98-39006 65-A90-1 (U-21A, RU-21A, RU-21D, JU-21A, U-21G, RU-21H) LM-1 and up LF* See Note 3c 65-A90-2 (RU-21B) LS-1 and up LF* See Note 3c 65-A90-3 (RU-21C) LT-1 and up LF* See Note 3c 65-A90-4 (RU-21E, RU-21H) LU-1 and up LF* See Note 3c 70 LB-1 thru LB-35 LF* P/N 98-39006 65-90, 65-A90, B90, C90 LJ-1 thru LJ-993 LJ-995 thru LJ-1007 LJ-1009 thru LJ-1034 LJ-1037 & LJ-1039 thru LJ-1044 LF* LF* LF* LF* LF* P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 E90 LW-1 thru LW-347 LF* P/N 98-39006 F90 LA-2 thru LA-90 LA-92 thru LA-156 LA-158 thru LA-169 LA-171 thru LA-173 LA-175 thru LA-182 LA-185, LA-187 LA-189 thru LA-191 LA-193 thru LA-196 and LA-199 LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR LF,UR,LR P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 H-90 (T-44A) LL-1 thru LL-18 LL-20 thru LL-40 LL-42 thru LL-48 LL-50 thru LL-61 LF LF LF LF T-44A-0049R1 T-44A-0049R1 T-44A-0049R1 T-44A-0049R1 99, 99A, 99A (FACH) U-1 thru U-49 and LF* P/N 98-39006 A99, A99A, & B99 U-51 thru U-164 LF* P/N 98-39006 C99 U-50 and U-165 thru U-179 U-181 thru U-184 U-186 thru U-192 U-194 thru U-196 LF,UF LF,UF LF,UF LF,UF LF,UF P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 100, A100 & A100A B-1 thru B-247 LF* P/N 98-39006 B100 BE-2 thru BE-131, and BE-135 LF* LF* P/N 98-39006 P/N 98-39006 A100-1 BB-2 thru BB-342 LF,UF P/N 98-39006 (RU-21J), 200, BB344 thru BB-983 LF,UF P/N 98-39006 and B200 BB-985 thru BB-1038 LF,UF P/N 98-39006 B200 BB-1040 thru BB-1045 BB-1047 thru BB-1049 BB-1053 thru BB-1078 & BB-1080 LF,UF LF,UF LF,UF LF,UF P/N 98-39006 P/N 98-39006 P/N 98-39006 P/N 98-39006 200C, B200C BL-1 thru BL-51 LF,UF P/N 98-39006 (C-12F) BL-53, BL-55 LF,UF P/N 98-39006 200 CT BN-1 LF,UF P/N 98-39006 200 T BT-1 thru BT-22 LF,UF P/N 98-39006 A200 (C-12A, C-12C)BC-1 thru BC-75 BD-1 thru BD-30 LF,UF LF,UF P/N 98-39006 P/N 98-39006 A200C (UC-12B) BJ-1 thru BJ-47 LF,UF P/N 98-39006 A200CT (C-12D,FWC-12D) BP-1 thru BP-27 LF,UF P/N 98-39006 (*-See Note 4) NOTE 1: Wing attachment joints, on left and right sides of each airplane, are abbreviated as: LF=lower forward, UF=upper forward, UR=upper rear, LR=lower rear. NOTE 2: T-34C-1-0083-1, T-34C-0158 Rev. 2, and T-44A-0049 Rev. 1 are Beech Service Instructions. Cited Beech Manuals, and their earliest applicable revision dates are: PART NUMBER NAME DATE 60-590001-25 Maintenance Manual June 13, 1984 98-39006 Structural Inspection and Repair Manual December 20, 1984 NOTE 3: Apply the following portions of P/N 98-39006 manual even though applicability to military models is not shown within the P/N 98-39006 manual: NOTE MANUAL SECTION REFERENCE FIGURE BOLT P/N NUT P/N 3a 57-10-00 209 NAS495-14-27 EB-144 3b 57-11-00210 MS20014-29 EB-144 3c 57-13-00 212 LWB-14-32 FN22-1414 Compliance: Required initially, upon accrual of five years after first airworthiness certification or within 60 days after the effective date of this AD (whichever is later), unless already accomplished, and thereafter at intervals which do not exceed five years. To assure structural integrity of attachments of outer wing panels to the wing center section, use procedures in instructions identified in Table I of this AD to accomplish the following at each wing attachment joint that is specified for a particular airplane by Table I of this AD: (a) Remove each steel nut and each steel tension bolt. Use visual and magnetic particle methods to inspect the bolt and nut for cracking and corrosion in parent steel, and replace each bolt and nut found cracked or corroded. NOTE 4: In lower forward joints that are asterisked in Table I of this AD, while bolts are removed for accomplishment of Paragraph (a), above, it is recommended that inboard and outboard fittings be inspected, by a fluorescent penetrant method, for fatigue cracks in washer face areas of the fittings. For some of the asterisked joints, inspections of fittings are required by other AD's, but inspections of fittings are not required by this AD. (b) During reassembly of each joint, coat the bolt, nut, and adjacent parts with MIL- C-16173 Grade 2 corrosion preventative compound. (c) Within the next 150 hours of flight time, check joint tightness, and tighten as necessary. (d) Inject MIL-C-16173 Grade 2 corrosion preventative compound into a lubrication fitting on each barrel nut, (wherever a barrel nut is used) when joint tightness is checked per Paragraph (c), above, and thereafter at intervals which do not exceed one year. (e) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (f) For destructive or nondestructive examination and any specified return, each nut and each bolt that is replaced in response to this AD must be identified with the related years in service, joint, and airplane serial number and sent to FAA/AVN-112, Room 203, Aviation Records Building, Mike Monroney Aeronautical Center, 6500 South MacArthur Boulevard, Post Office Box 26460, Oklahoma City, Oklahoma 73125. Parts so sent will be destroyed if return to a specified address is not requested. Reporting requirements approved by OMB pursuant to clearance No. 2120 0056. (g) An equivalent means of compliance with this AD may be used if approved by the Manager, FAA, Wichita Aircraft Certification Office, Room 100, 1801 Airport Road, Wichita, Kansas 67209; telephone (316) 946-4400. All persons affected by this directive may obtain copies of the documents referred to herein upon request to Beechcraft Aero and Aviation Centers; Beech Aircraft Corporation, 9709 East Central, Post Office Box 85, Wichita, Kansas 67201, or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment becomes effective on November 12, 1985.
53-09-01: 53-09-01 PRATT & WHITNEY: Applies to R-2000 Series Engines Overhauled by San Antonio Air Depot Between July 1, 1952, and April 9, 1953. Compliance required prior to further carriage of passengers or cargo in aircraft with such engines installed. Aircraft may be ferried to base where inspection is to be conducted. Several operators of C-54 aircraft utilizing military overhauled engines, have experienced failures of link (knuckle) pins in the subject engines due to improper overhaul, inspection, or assembly procedures. The failures have occurred in comparatively low-time engines, and cracked pins have been found in engines with zero TSO. To preclude the possibility of further failures of this nature, engines falling within the category as noted must be dissassembled and the link pins magnetically inspected, eliminating any found with cracks, then reassembled in accordance with manufacturer's instructions, prior to further flight. Date of overhaul and identification ofoverhaul base stamped on exterior surface of engine nose housings.
77-10-01: 77-10-01 MCDONNELL DOUGLAS: Amendment 39-2895. Applies to DC-4 and C-54 series airplanes, certificated in all categories, using Hamilton Standard propellers. \n\n\tCompliance required by the first engine change after the effective date of this AD or October 1, 1977, whichever occurs earlier, unless previously accomplished in accordance with AD 55-15-03 as amended by Amendment 55-24 to Part 507. \n\n\tTo increase fire resistance integrity of the propeller feathering system against damage by a powerplant fire, all flexible hose components of propeller feathering lines forward of the firewall must be replaced with lines and fittings which will meet the fire resistance requirements of the hose assemblies specified in (a) through (f) herein. However, if the feathering lines in Zone I include a section of steel tubing, flexible hose assemblies located forward of the cylinders and connecting to the governor are not affected by this directive. \n\n\tThe following flexible hose assemblies are acceptable for use in this application: \n\n\t(a)\tResistoflex SSFR-3800-10 hose assemblies. \n\n\t(b)\tAeroquip 680-10S hose assemblies with Aeroquip 304 protective sleeves over end fittings (Aeroquip Assembly P/N 304000). \n\n\t(c)\tAeroquip 309009 hose assemblies. \n\n\t(d)\tAeroquip 309009-8S hose assemblies (where feathering system requires this size). \n\n\t(e)\tAeroquip 634000-8 or -10, as applicable, hose assemblies. \n\n\t(f)\tHose assemblies that fully comply with FAR 37.140 (TSO-C42) and have a pressure rating equal to or greater than that of the propeller feathering system installed on the airplane. \n\n\t(g)\tEquivalent hose assemblies or other means of compliance may be used when approved by the Chief Aircraft Engineering Division, FAA Western Region. \n\n\tSpecial flight permits may be issued in accordance with FAR's 21.97 and 21.199 to operate the airplane to a base for the accomplishment of this AD. \n\n\tThis supersedes AD 55-15-03, as amended by Amendment 55-24 to Part 507. \n\n\tThis Amendment becomes effective June 16, 1977.
96-16-03: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320-200 series airplanes, that requires modification of the shock absorber sub-assembly of the main landing gear (MLG). This amendment is prompted by reports of internal damage to the shock absorber sub-assembly due to loose screws in the upper bearing dowels. The actions specified by this AD are intended to prevent such damage, which could result in the overextension of the shock absorber and failure of the torque link. This situation may lead to the inability of the MLG to retract and subsequent collapse of the MLG.
2009-25-10: This amendment adopts a new airworthiness directive (AD) for the Sikorsky Model S-92A helicopters. This action requires a one-time visual inspection of the main gearbox (MGB) lube system filter assembly for oil filter damage. This action also requires if either the primary or secondary oil filter is damaged, replacing both filters, all packings, and the studs before further flight. This AD also requires replacing the oil filter bowl within 30 days after replacing a damaged filter and a daily leak inspection for an oil leak (no oil leaks allowed) during that 30-day interim period. This amendment is prompted by three reports of damaged oil filters or packings resulting from installing the filter assembly with an oversized packing possibly because of incorrect part numbers in the maintenance manual. Based on a previous accident investigation, failure of the oil filter bowl or mounting studs can result in sudden and complete loss of oil from the MGB. The actions specified in thisAD are intended to prevent complete loss of oil from the MGB, failure of the MGB, and subsequent loss of control of the helicopter.
90-04-09: 90-04-09 BRITISH AEROSPACE: Amendment 39-6511. Docket No. 89-NM-216-AD. Applicability: Model BAe 146-200A and -300A series airplanes, as listed in British Aerospace Service Bulletin 53-84-00737D, Revision 1, dated August 22, 1989, certificated in any category. Compliance: Required prior to the accumulation of 3,000 landings since new, or within 30 days after the effective date of this AD, whichever occurs later, unless previously accomplished. To prevent reduced structural integrity of the fuselage, accomplish the following: A. Modify the fuselage rear section by adding eight rivets to the Stringer 21P end termination area, in accordance with British Aerospace Service Bulletin 53-84-00737D, Revision 1, dated August 22, 1989. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base 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, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. 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 (39-6511, AD 90-04-09) becomes effective on March 19, 1990.
71-09-04: 71-09-04 BOEING: Amdt. 39-1199. Applies to Model 727 airplanes listed in Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revisions. \n\tCompliance required as indicated. \n\tTo detect cracking of the main landing gear actuator beam, accomplish the following: \n\tA.\tUnless already accomplished within the last 300 hours' time in service preceding the effective date of this AD, within the next 300 hours' time in service after the effective date of this AD or prior to the accumulation of 5300 hours' time in service, whichever occurs later, accomplish one of the following: \n\t\t1.\tVisually inspect the main landing gear actuator beam for any evidence of cracking in accordance with Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revision, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat the visual inspection at intervals not to exceed 600 hours' time in service, or \n\t\t2.\tUltrasonically inspect the main landing gear actuator beam for cracks in accordance with Boeing Service Bulletin 32-188, dated March 29, 1971, or later FAA approved revision, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat the ultrasonic inspection at intervals not to exceed 1500 hours' time in service. \n\tB.\tIf cracks are found replace the beam with a serviceable beam. \n\tC.\tWhere records maintained by the operator are such as will permit a clear determination of the number of hours' time in service accumulated by the main landing gear actuator beam, P/N 65-17658-11, installed on the airplane, the inspection times prescribed by this AD may be applied to the beam rather than to the airplane. \n\tD.\tInspections prescribed by this AD do not apply to new replacement beams, P/N 65- 57153-3, or to P/N 65-17658-11 beam until 5300 hours' time in service is reached. \n\tE.\tAirplanes having cracked main landing gear actuator beams which require replacement under this AD may, in accordance with FAR 21.197, be flown with the landing gear extended to a base where the replacement can be accomplished. \n\tThis AD becomes effective May 28, 1971.
90-23-17: 90-23-17 BRITISH AEROSPACE: Amendment 39-6800. Docket No. 90-NM-122-AD. Applicability: Model DH.125-1A series airplanes, equipped with Hawker Siddeley Dynamics Air Conditioning System and Rolls Royce Viper Engines, certified in any category. Compliance: Required as indicated, unless previously accomplished. To prevent chafing between the air conditioning duct and the rear pressure bulkhead, and subsequent rapid decompression of the airplane, accomplish the following: A. Within 30 days after the effective date of this AD, perform a detailed visual inspection for chafing on the aft face of the rear pressure bulkhead, in accordance with the Accomplishment Instructions of British Aerospace Service Bulletin 53-71, dated November 1, 1989. B. If defects are found, prior to further flight, perform a dye penetrant inspection to detect cracks in the vicinity of the affected area; and perform a dial test indicator measurement to determine the depth of damage in therear pressure bulkhead, in accordance with British Aerospace Service Bulletin 53-71, dated November 1,1989. 1. If the damage to the rear pressure bulkhead is less than 0.003 inch deep, prior to further flight, carefully blend out, polish, and then restore protective treatment in accordance with the service bulletin. 2. If the damage to the rear pressure bulkhead is greater than 0.003 but less than 0.010 inch deep, within 100 landings, repair in accordance with Appendix B of the Service Bulletin. 3. If the damage to the rear pressure bulkhead is greater than 0.010 inch deep, prior to further flight, repair in accordance with Appendix B of the Service Bulletin. C. Within 30 days after the effective date of this AD, adjust the clearance between the air conditioning duct clamp and the rear pressure bulkhead so there is at least a 3/4-inch clearance. This can be accomplished by rotating and adjusting the duct position. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Standardization Branch, ANM-113, and a copy sent to the cognizant Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Standardization Branch, ANM-113. E. 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, 1601 Lind Avenue, S.W., Renton, Washington. This amendment (39-6800, AD 90-23-17) becomes effective on December 11, 1990.
77-20-06: 77-20-06 PILATUS AIRCRAFT, LTD. AND FAIRCHILD HILLER: Amendment 39-3050. Applies to Model PC-6 airplanes (all variants) manufactured by Pilatus Aircraft, Ltd., up through S/N 724 and to Model PC- 6 airplanes (all variants) manufactured by Fairchild Hiller, S/N's 2001 through 2047, certificated in all categories. Compliance is required within the next 25 hours time in service after the effective date of this AD, unless already accomplished within the last 75 hours time in service, and thereafter at intervals not to exceed 100 hours time in service from the last inspection, until the conditions of paragraph (c) are met. (a) To prevent a hazardous degree of corrosion from developing inside the wing struts, accomplish the following: (1) Visually inspect the internal surface of each wing strut for corrosion in accordance with paragraph 2.1 of Pilatus Aircraft Ltd., Service Bulletin No. 105, dated May 1971, (hereinafter S.B. No. 105) for Pilatus manufactured airplanes or paragraph2A of Fairchild Hiller Service Bulletin PC6-57-3, dated July 15, 1971 (hereinafter S.B. PC6-57-3), for Fairchild Hiller airplanes, or an FAA-approved equivalent. (2) If only light corrosion (corrosion which has not caused surface blistering) is found during an inspection required by paragraph (a)(1) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, remove the corrosion from, and apply an anticorrosive treatment to, the inside of the wing strut in accordance with paragraph 2.3 of S.B. No. 105 or paragraph 2D of S.B. PC6-57-3, as applicable, or an FAA-approved equivalent. (3) If corrosion is found during an inspection required by parargraph (a)(1) of this AD which has resulted in exceeding the limits prescribed in paragraph (a)(2), within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, replace the wing strut with a serviceable strut of the same part number that has had anticorrosive treatment applied to the inside surface in accordance with paragraph 2.3 of S.B. No. 105 or paragraph 2D of S.B. PC6-57-3, as applicable, or an FAA- approved equivalent. (b) For Pilatus Aircraft, Ltd., Model PC-6 airplanes, S/N's 338 through 701, to prevent a hazardous degree of corrosion from developing on the wing strut attachment brackets, acocmplish the following: (1) Visually inspect each wing strut attachment bracket for corrosion in accordance with paragraph 2.1 of Pilatus Aircraft, Ltd., Service Bulletin No. 93, dated June 1969, (hereinafter S.B. NO. 93) or an FAA-approved equivalent. (2) If only light corrosion (corrosion which has caused 2% to 10% reduction in cross-section per paragraph 2.2 of S.B. No. 93) is found during an inspection required by paragraph (b)(1) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, remove the corrosion and apply an anti-corrosive treatment to the wing strut attachment bracket in accordance with paragraph 2.4 of S.B. No. 93, and reinstall the bracket in accordance with paragraph 2.5.1 of S.B. No. 93 or an FAA-approved equivalent. (3) If corrosion is found during an inspection required by paragraph (b)(1) of this AD which has resulted in exceeding the limits prescribed in paragraph (b)(2) of this AD, within the next 100 hours time in service or within the next 60 days after finding the corrosion, whichever occurs sooner, replace the wing strut attachment bracket with a new wing strut attachment bracket (P/N 111.35.06.055 (left) or 111.35.06.056 (right)) in accordance with paragraph 2.5.2 of S.B. No. 93 or an FAA-approved equivalent. (c) The inspections required by paragraph (a)(1) or (b)(1) of this AD may be discontinued, in accordance with the following: (1) Inspection of the wing strut when the strut has had light corrosion removed and has had the anticorrosive treatment in accordance with paragraph (a)(2) or when the strut has been replaced in accordance with paragraph (a)(3) of this AD. (2) Inspection of the wing strut attachment bracket when the bracket has had light corrosion removed and has had the anticorrosive treatment in accordance with paragraph (b)(2), or when the bracket has been replaced in accordance with paragraph (b)(3) of this AD. This amendment becomes effective November 3, 1977.