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78-13-02: 78-13-02 BRITISH AEROSPACE: Amendment 39-3245. Applies to Hawker Siddeley Model BH/HS-125 Series 600A and 700A airplanes, certificated in all categories, that have either the RCA AVQ21 or Primus 40 weather radar systems installed. Compliance required as indicated. To prevent failure of the restraint provisions for the weather radar receiver/transmitter, accomplish the following: (a) Within 10 hours time in service after the effective date of this AD, unless already accomplished within the last 40 hours time in service, and thereafter at intervals not to exceed 50 hours time in service from the last inspection, inspect the support brackets for cracks and the "T" bolts for failure of the brazing in accordance with section 2, "Accomplishment Instructions" of British Aerospace Alert Service Bulletin 34-A134, dated April 1, 1978, or an FAA approved equivalent. (b) If a crack in a support bracket or a failure of the brazing of the "T" bolt is found during an inspection required by paragraph (a) of this AD, before further flight, except that the airplane may be flown in accordance with FAR 21.197 and 21.199 to a base where the replacement can be accomplished, replace the mounting tray with a serviceable part of the same part number, or repair the existing mounting tray in accordance with an FAA approved repair scheme and continue to inspect in accordance with paragraph (a) of this AD, or replace the mounting tray in accordance with paragraph (c) of this AD. (c) The inspections required by this AD may be discontinued upon replacement of the mounting tray with an improved standard tray, P/N 1719353-501 (Rev. E), in accordance with British Aerospace Modification 258171 or an FAA approved equivalent. This amendment becomes effective July 5, 1978.
89-25-10: 89-25-10 BEECH: Amendment 39-6409. Applicability: Models 65-90 and 65-A90 (Serial Number (S/N) LJ-1 thru LJ-317); 65-A90-1, 65-A90-2, 65-A90-3, 65-A90-4, B90, C90 (all S/N); C90A (S/N LJ-1063 thru LJ-1087, except LJ-1085); E90, 100, A100 and B100 (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, whichever occurs later, unless previously accomplished per AD 87-23-09, Amendment No. 39-5765, or AD 70-25-04, Amendment No. 39-1332, 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 sectionand 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), Part Number 98-39006, Revision A4, dated May 1, 1987. The inspection must be performed by personnel specifically trained by Beech Aircraft Corporation. 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 (g) 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) SA1178CE or SA1583CE for proper tension and condition in accordance with Aviadesign Engineering Order E.O. B-8001, Issue 3, dated May30, 1985. Correct any discrepancy prior to further flight. For airplanes so equipped and inspected, 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 Beech SIRM or other FAA approved instructions provided by Beech Aircraft Corporation. (d) Within one week after completion of any inspection required by paragraph (a) or (b) of this AD, complete the reporting form included with this AD as Figure 1 and mail it to the address shown thereon (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056). (e) The initial and repetitive inspections specified in this AD are no longer required when the airplane is modified by Beech Wing Modification Kit No. 90-4077-1S or 100-4007-1S. (f) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. (g) 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, 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. ThisAD supersedes AD 87-23-09, Amendment 39-5765, and AD 70-25-04, Amendment 39-1332. This amendment (39-6409, AD 89-25-10) becomes effective on January 4, 1990. REPORTING FORM - 89-25-10 Airplane Model No. ___________________________________________________ Airplane Serial No.____________________________________________________ Date of inspection per this AD____________________________________________ Airframe total hours time-in-service________________________________________ Were any fatigue cracks found? No Yes __ If "Yes" was checked above, complete the following: Location of crack_____________________________________________________ Was crack removable by reaming or grinding? No Yes __ Additional Comments __________________________________________________ Mailing Address: FAA, Wichita ACO Airframe Branch, Room 100 1801 Airport Road Wichita, KS 67209 FIGURE 1 - 89-25-10
76-07-04: 76-07-04 HAWKER SIDDELEY AVIATION LTD: Amendment 39-2563. Applies to de Havilland Model DH-114 "Heron" airplanes certificated in all categories which have not been altered in accordance with Heron Modification 1612. Compliance is required as indicated. To detect cracks in the nose landing gear inner casing, and prevent the possible collapse of the nose landing gear upon landing, accomplish the following: (a) Within the next 50 hours time in service after the effective date of this AD, unless already accomplished within the preceding 600 hours time in service, and thereafter, at intervals not to exceed 600 hours time in service from the last inspection, inspect the nose landing gear inner casing for cracks in accordance with paragraphs 3.1 and 3.2 of section 3 entitled "Inspection" of Hawker Siddeley Aviation Ltd., Technical News Sheet No. U.17, Issue 1, dated September 17, 1973, or an FAA-approved equivalent. (b) If any cracks are found during an inspection required by paragraph (a) of this AD, before further flight, except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed, replace the inner casing with a new part of the same part number or a serviceable used part of the same part number that has been inspected and found to be free of cracks in accordance with the inspection prescribed in paragraph (a) of this AD. Continue to inspect the replacement casing for cracks in accordance with Hawker Siddeley Aviation Ltd., Technical News Sheet, No. U.17, Issue 1, dated September 17, 1973, or an FAA-approved equivalent at intervals not to exceed 600 hours time in service from replacement. This amendment becomes effective April 12, 1976.
77-14-04: 77-14-04 HAWKER SIDDELEY AVIATION, LTD.: Amendment 39-2952. Applies to DH-104 "Dove" and DH-114 "Heron" airplanes. Compliance is required within the next 500 hours time in service after the effective date of this AD unless already accomplished. To prevent the possibility that a loss of generated electrical power would be undetected by the flight crew, accomplish the following: (a) Alter the electrical system to incorporate a bus bar low voltage sensing unit, a bus bar low voltage warning light, and an essential service switch, designed and installed in accordance with paragraphs 5 and 6 of Hawker Siddeley Aviation, Ltd., Technical News Sheet, Series: Heron (114), No. N. 6., Issue 3 (for DH-114 "Heron") and CT104, No. 227, Issue 3 (for DH-104 "Dove"), both dated July 3, 1972, as amended to November 20, 1972, or FAA-approved equivalent of either. (b) Amend the "Normal and Emergency Procedures", Part B, of the "Operating Procedures" section, Section II, of the applicable Airplane Flight Manual by adding the electrical system operation information contained in paragraphs 7 and 8 and Figure 1 of the applicable Technical News Sheet, referred to in paragraph (a) of this AD, or an FAA-approved equivalent. (c) Check the condition of the electrical distribution and generator system in accordance with paragraph 6 of the applicable Technical News Sheet, referred to in paragraph (a) of this AD, or an FAA-approved equivalent, and repair, as necessary. The checks required by this paragraph may be performed by persons authorized to perform preventive maintenance under FAR 43. This amendment becomes effective August 5, 1977.
78-03-05: 78-03-05 MITSUBISHI HEAVY INDUSTRIES, LTD: Amendment 39-3137. Applies to models MU-2B, MU-2B-10, MU-2B-15, MU-2B-20, MU-2B-25 and MU-2B-26 airplanes with serial numbers up through and including 347 except 313 and 321 and models MU-2B-30, MU-2B-35, and MU-2B-36 airplanes with serial numbers up through and including 696 except 652 and 661. NOTE: This AD is not applicable to MU-2B series airplanes having serial numbers with the suffix "SA." Compliance is required as indicated. To prevent failure of the cowling latches between the engine nacelle upper door and side doors, subsequent separation of the upper cowling panel, and possible loss of control of the airplane, accomplish the following: Within the next 25 hours time in service after the effective date of this AD, unless already accomplished, replace the cowling latch links between the engine nacelle upper door and side doors in accordance with the instructions contained in Mitsubishi MU-2 Service Bulletin No. 171A datedJuly 14, 1975, as supplemented by Mitsubishi MU-2 Service Bulletin No. 180 dated August 26, 1977, or Mitsubishi Service Bulletin No. 180A dated November 17, 1977, or an FAA- approved equivalent, approved by the Chief, Engineering and Manufacturing District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. This supersedes amendment 39-2695, (41 FR 34009), AD 76-16-05. This amendment becomes effective February 23, 1978.
76-18-12: 76-18-12 GRUMMAN AMERICAN AVIATION CORPORATION: Amendment 39- 2721. Applies to Model G-1159 airplanes certificated in all categories, serial numbers 1 through 154 and 775. Compliance required within the next 100 hours time in service after the effective date of this AD, unless already accomplished. To detect loose terminal connections at the generator terminal boards and to prevent the loosening of these connections, accomplish the following: 1. Modify each engine electrical junction box to provide an access hole for inspecting the two generator terminal boards and the eight associated wiring connections. Grumman American Aircraft Service Change No. 203 Amendments 1 and 2 provide the information for accomplishing this modification. 2. If the inspection reveals that all generator terminals are secure, i.e., that all lock washers are compressed and torque stripes not broken, the connections are satisfactory. 3. Reinspect in accordance with A.S.C. No. 203 Amendments 1 and 2 in intervals of 300 hours time in service until the basic A.S.C. No. 203 has been accomplished. 4. If the lock washer on any terminal is found not to be compressed, or evidence of arcing at any connection is noted, the affected engine shall be removed and the basic A.S.C. No. 203 accomplished to the corresponding engine junction box. 5. The reinspection procedure must continue for the remaining engine junction box until the basic A.S.C. No. 203 has been accomplished. This service change requires the removal of a plain nut, a plain washer, and lockwasher from each generator terminal, and replacing them with a self-locking nut and a plain washer. 6. Compliance with the basic A.S.C. No. 203 must be accomplished at the next engine removal, if not done prior to that time. This Airworthiness Directive may be accomplished by any other means approved by the Chief, Engineering and Manufacturing Branch of the Southern Region, Atlanta, Georgia. This amendmentbecomes effective September 22, 1976.
77-14-01: 77-14-01 AGUSTA: Amendment 39-2949. Applies to Model A-109A helicopters equipped with main rotor hub flap hinge bearings P/N SJ7355/IR7355. Compliance is required as indicated, unless already accomplished. To prevent possible improper operation of the main rotor flaps due to premature failure of any one of the main rotor hub flap hinge bearings, accomplish the following: (a) For helicopters with serial numbers up to and including S/N 7120, except S/N 7119 - (1) Within the next 10 hours time in service after the effective date of this AD, replace the main rotor hub flap hinge bearings P/N SJ7355/IR7355 in accordance with Part I of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent; and (2) Within 100 hours time in service after complying with paragraph (a)(1) of this AD, and thereafter at intervals not to exceed 100 hours time in service, or any time abnormal oil leaks occur from flap hinges, perform inspections and replace bearings, as necessary, in accordance with Part II of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent. (b) For helicopters with serial numbers S/N 7119, S/N 7121 and up, within the next 25 hours time in service after the effective date of this AD, except for those which have been inspected within the previous 75 hours, and thereafter at intervals not to exceed 100 hours time in service or any time abnormal oil leaks occur from flap hinges, perform inspections and replace bearings, as necessary, in accordance with Part II of Agusta Technical Bulletin No. 109-3, dated May 4, 1977, or an FAA-approved equivalent. (c) Equivalent methods of complying with this AD must be approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region, c/o American Embassy, APO New York, N.Y. 09667. This amendment becomes effective July 15, 1977.
76-16-01: 76-16-01 AIRESEARCH MANUFACTURING COMPANY OF ARIZONA: Amendment 39-2683. Applies to the following groups of Model TPE331 series engines: Group A: TPE331-1-101B, S/N 93058 through 93061; TPE331-1-151A, S/N 92249 and 92336 through 92354; TPE331-1-151K, S/N 26001 through 26014; TPE331-1-151G, S/N 91193 through 91198; TPE331- 2-201A, S/N 90218 through 90278, TPE331-3U-303G or TPE331-3UW-303G, S/N 03108, 03109, and 03112 through 03180 and 05031 through 05042; TPE331-3U-307G, S/N 03001 and 03009; TPE331-5-251C, S/N 22006 through 22057; TPE331-5-251K, S/N 06113, 06190 through 06442 and 06444 through 06454; TPE331-6-251M or TPE3310-6-252M, S/N 20144 and 20182 through 20467 and 20469 through 20533; TPE331-6-252B, S/N 27001 and 27002, and any Model TPE331-1, -2, -3, -5, or -6 engines which have been modified in accordance with AiResearch Service Bulletin TPE331-72-0064 dated February 1, 1974 or subsequent revisions. Group B: TPE331-1-101B, S/N 93062, 93063; TPE331-1-151A, S/N 92355 through 92357; TPE331- 1-151K, S/N 26015 through 26023; TPE331-2-201A, S/N 90279 through 90296; TPE331-3U-303G or TPE331-3UW-303G, S/N 03181 through 03197, 05043 through 05052; TPE-5-251C, S/N 22058 through 22119; TPE331-5-251K, S/N 06443, 06455 through 06556; TPE331-6-251M or TPE331-6-252M, S/N 20534 through 20577. Group C: TPE331-1-101F, S/N 98017 through 98019; TPE331-1-151K, S/N 91194 and 91195; TPE331-2- 201A, S/N 90297 through 90303; TPE331-3U-303G or TPE331-3UW-303G, S/N 05053 through 05058, S/N 05016, S/N 03001, S/N 03010, S/N 03011, and S/N 03198 through 03205; TPE331-5-251K, S/N 06557 through 06576; TPE331-5-251C, S/N 22120 through 22139, TPE331-5-251M, S/N 28001 through 28005; TPE331-6-251M, S/N 20468, S/N 20017, and S/N 20018; TPE331-6-252B, S/N 27003 and S/N 27004, and any Model TPE331-1, -2, -3, -5, or -6 engines which have been modified in accordance with AiResearch Service Bulletin TPE331-72-0064, dated February 1, 1974, or subsequent revisions. NOTE: Operators of airplanes incorporating the engines affected by this airworthiness directive are advised to examine the applicability section, Configuration Identification, Table 1, or AiResearch Service Bulletin TPE331-72-0092, Revision 5, dated June 8, 1976, or later FAA-approved revision, to identify the extent of work required by this airworthiness directive which may have been previously performed. Compliance required as indicated. To detect and prevent fatigue failures of the high speed pinion (HSP) bearing oil transfer tube, supply tube, and lubricating adapter, and to detect, correct, and prevent loosening of the HSP bearing carrier bolts, accomplish the following: (a) For engines in Group A, above, within the next 100 hours time in service after December 30, 1974 (the effective date of Amendment 39-2054, AD 74-26-11), unless already accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, dated December 9, 1974, or later FAA-approved revision: (1) Replace the two high speed pinion bearing carrier bolts, P/N MS21279-07, with two bolts, P/N MS9489-07, and lockplate, P/N 3101483-1, (2) Inspect to insure proper torque on bolt, P/N MS21279-10, securing lube nozzle, P/N 3101209, and (3) If aluminum lube adapter, P/N 3101210, is not replaced as described in paragraph (d) of this directive, inspect to insure proper torque on bolt, P/N MS21297-07, securing this lube adapter. (b) For engines in Group A and Group B, above, which incorporate the oil transfer tube, P/N 3101187-1, within the next 100 hours time in service after December 30, 1974, unless already accomplished within the last 100 hours time in service prior to December 30, 1974, and thereafter at intervals not to exceed 200 hours time in service from last inspection, until the modifications of paragraph (c) are accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, dated December 9, 1974, or later FAA-approved revision: (1) Inspect the integral support bracket on oil transfer tube, P/N 3101187-1, for cracks or separation, and (2) If the oil transfer tube bracket is cracked or separated, before further flight, replace both the oil transfer tube and oil supply tube with new tubes, oil transfer tube, P/N 3101475-1, and oil supply tube, P/N's 3101473-1 or 3101473-2, and associated hardware. (c) For engines in Group A and Group B, above, which incorporate the oil transfer tube, P/N 3101187-1, before accumulating a total of 1800 hours time in service after December 30, 1974, unless already accomplished, replace the oil transfer tube, P/N 3101187-1, with one of the following serviceable oil transfer tubes in accordance with AiResearch Service Bulletin TPE331-72-0092, Revision 1, dated, January 1, 1975, or later FAA-approved revision: (1) P/N 3101187-2 which has been heat-treated and reidentified as P/N 3101187-3, (2) P/N 3101187-3, or (3) P/N3101475-1 (d) For engines in Group A, Group B, and Group C, above, within the next 200 hours time in service after the effective date of this AD, unless already accomplished, accomplish the following in accordance with AiResearch Service Bulletin TPE331-72-0092, Revision 5, dated June 8, 1976, or later FAA-approved revision: (1) Replace the aluminum main journal bearing lubricating adapter, P/N's 3101210-1 or 3101210- 2, with a new steel lubricating adapter, P/N's 3101474-1 or 3101474-2, and (2) If two carrier bolts, P/N MS9489-07, and lockplate, P/N 3101483-1, are not replaced, inspect these carrier bolts manually for any indication of looseness. If loose bolts are found, before further flight install a new lockplate, P/N 3101483-1, and retorque the two carrier bolts, P/N MS9489-07 (e) Equivalent procedures, inspections, modifications, or parts may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. (f) Aircraft may be flown to a base for accomplishment of inspections and modifications required by this AD per FAR's 21.197 and 21.199. This supersedes Amendment 39-2054 (39 F.R. 4439), AD 74-26-11, as amended by Amendments 39-2092 (40 F.R. 6771), 39-2214 (40 F.R. 22126), 39-2254 (40 F.R. 28605), 39-2367 (40 F.R. 42740), and 39-2382 (40 F.R. 48499). This amendment becomes effective August 6, 1976.
77-07-10: 77-07-10 BRITTEN NORMAN, LTD.: Amendment 39-2869. Applies to Models BN-2A Islander and BN-2A Mk III Trislander airplanes, all series, certificated in all categories, except those airplanes incorporating Britten Norman Modification NB/M878. Compliance is required as indicated, unless already accomplished. To detect internal corrosion and prevent possible failure of the aileron mass balance support arm, accomplish the following: (a) For ailerons modified in accordance with Britten-Norman, Ltd. Modification NB/M/336, within the next 25 hours time in service after the effective date of this AD or prior to 3 years since new, whichever occurs later, accomplish the following: (1) Gain access to and inspect the interior surface of the aileron mass balance support arm in accordance with steps 1 thru 4 of Part 1 and Figure 1 of Britten-Norman, Ltd. Service Bulletin BN-2/SB.98, Issue 1, dated October 27, 1976 (hereinafter S.B. BN- 2/SB.98), or an FAA-approved equivalent.(2) If no corrosion or acceptable corrosion (as defined in Table 1 of Figure 1 of S.B. BN-2/SB.98) is found as a result of the inspection required by paragraph (a)(1) of this AD, clean and protect the interior surface of the support arm and the shank of the bobweight in accordance with subpart (a) of step 5, and reassemble in accordance with steps 6 through 9, of Part 1 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (3) If marginal corrosion (as defined in Table 2 of Figure 1 of S.B. BN- 2/SB.98) is found as a result of the inspection required by paragraph (a)(1) of this AD, before further flight, clean, protect, and reassemble the support arm, in accordance with paragraph (a)(2) of this AD. Thereafter, within the next 300 hours time in service or 3 months, whichever occurs sooner, replace the support arm with a new part of the same part number or an FAA-approved equivalent, and reassemble in accordance with paragraph (a)(2) of this AD. (4) If, during the inspection required by paragraph (a)(1) of this AD, corrosion is found beyond acceptable limits (as defined in Table 3 of Figure 1 of S.B. BN- 2/SB.98), before further flight, replace the support arm with a new part of the same part number, or an FAA-approved equivalent, and reassemble in accordance with paragraph (a)(2) of this AD. (b) For ailerons not modified in accordance with Britten-Norman, Ltd. Modification NB/M/336, within the next 25 hours time in service after the effective date of this AD or prior to 3 years since new, whichever occurs later, accomplish the following: (1) Gain access to and inspect the interior surface of the mass balance support arm in accordance with steps 1 thru 4 of Part 2 and Figure 2 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (2) If no corrosion or acceptable corrosion (as defined in Table 1 of Figure 1 of S.B. BN-2/SB.98) is found as the result of the inspection required by paragraph (b)(1) of this AD, clean and protect the interiorsurface of the support arm and the shank of the bobweight in accordance with subpart (a) of Step 5, Part 1, and repair the support arm in accordance with Figure 2 and reassemble in accordance with steps 6 thru 9, of Part 2 of S.B. BN-2/SB.98, or an FAA-approved equivalent. (3) If marginal corrosion (as defined in Table 2 of Figure 1 of S.B. BN- 2/SB.98) is found as a result of the inspection required by paragraph (b)(1) of this AD, before further flight, clean, protect, repair, and reassemble the support arm and shank of the bobweight in accordance with paragraph (b)(2) of this AD. Thereafter within the next 300 hours time in service or 3 months, whichever occurs sooner, replace the support arm in accordance with the repair and reassembly instructions specified in Part 3 and Figure 3, of S.B. BN-2/SB.98, or an FAA-approved equivalent. (4) If, during the inspection required by paragraph (b)(1) of this AD, corrosion is found beyond acceptable limits (as defined in Table 3 of Figure 1 of S.B. BN- 2/SB.98), before further flight, replace the support arm in accordance with Part 3 and Figure 3 of S.B. BN-2/SB.98, or an FAA-approved equivalent. This amendment becomes effective April 18, 1977.
90-08-17: 90-08-17 BEECH: Amendment 39-6564. Docket No. 89-CE-39-AD. Applicability: Models 65-90, 65-A90, B90, C90 and C90A (Serial Numbers (S/N) LJ-1 through LJ-1222); E90, F90, H90, 100, A100 and B100 (all S/N); 200 and B200 (S/N BB-2 through BB-1344); 200C, B200C, 200CT, B200CT, 200T, B200T, A200, A200C and A200CT (all S/N); and 300 (S/N FA-1 through FA-204 and FF-1 through FF-19) airplanes certificated in any category. Compliance: Required within the next 100 hours time-in-service after the effective date of this AD, unless already accomplished. To prevent water accumulation in the aft fuselage which can freeze and restrict control movement, accomplish the following in accordance with the instructions in Beech Service Bulletin Number 2312, dated December 1989: (a) Inspect the aft fuselage moisture drain system. (1) If the drain system is unobstructed and the drain openings are not undersized, no further action is required. (2) If any obstruction of thedrain system or any undersized opening is found, prior to further flight remove the obstruction or enlarge the opening as required; and (3) Report, in writing, any defects found to the Manager, Wichita Aircraft Certification Office (address below) within 7 days of the inspection. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056.) (b) Airplanes may be flown in accordance with FAR 21.197 to a location where the AD may be accomplished. (c) An alternate method of compliance or adjustment of the compliance time 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. NOTE: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. All persons affected by this directive may obtain copies of the document referred to herein upon request to the Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; Telephone (316) 681-7111; or may examine this document at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment (39-6564, AD 90-08-17) becomes effective on May 7, 1990.
92-10-07: 92-10-07 ROLLS-ROYCE PLC: Amendment 39-8241. Docket No. 87-ANE-05. Applicability: Rolls-Royce plc (R-R) Viper Mk. 521, Mk. 522, and Mk. 601 turbojet engines installed on but not limited to Hawker-Siddeley HS125 and HS125-600 aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent failure of high pressure compressor (HPC) rotor disks, causing an engine failure, accomplish the following: (a) Remove from service HPC stage 5 and 6 rotor disks with the following part numbers (includes all Modification Standards): Viper Mk. 521: V.20492, V.20494, V.27430, V.27433, V.39638, V.39639, V.39642, V.39643, V.39677, V.39678, V.39681, V.39682, V.42057, V.42059, V.46180, V.46181, V.46182, V.46204, V.46205, and V.46206; Viper Mk. 522: V.20492, V.20494, V.27430, V.27433, V.39638, V.39639, V.39642, V.39643, V.39677, V.39678, V.39681, V.39682, V.42057, V.42059, V.46180, V.46181, V.46182, V.46204, V.46205, and V.46206; Viper Mk. 601: V.42057, V42059, V.44768,V.44769, V.44780, V.44781, V.46182, and V.46206, and replace with a serviceable part, as follows: (1) HPC stage 5 and 6 rotor disks that have accumulated 12,500 or more cycles since new on the effective date of this AD, within the next 100 cycles in service. (2) HPC stage 5 and 6 rotor disks that have accumulated less than 12,500 cycles since new on the effective date of this AD, at or prior to accumulating 12,600 cycles since new. NOTES: (1) Further information pertaining to disk life limits can be obtained in Chapter 5 of the R-R Viper Mk. 521, Mk. 522, and Mk. 601 Maintenance and Shop Manuals. (2) Reference R-R plc Service Bulletins Number 72-A154 for Viper Mk. 601, and Number 72-A372 for Viper Mk. 521 and Viper Mk. 522 engines. (b) 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. (c) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety, may be used if approved by the Manager, Engine Certification Office, FAA, Engine and Propeller Directorate. The request should be forwarded through an 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. (d) This amendment becomes effective on September 21, 1992.
93-05-04: 93-05-04 AIRBUS INDUSTRIE: Amendment 39-8509. Docket 92-NM-112-AD. Applicability: Model A300 B2-1C, B2-203, B2K-3C, B4-2C, B4-103, and B4-203 series airplanes; equipped with General Electric (GE) engines on which engine Pylon Modification No. 2434, as described in Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, has not been installed; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the engine pylon, accomplish the following: (a) Prior to the accumulation of 22,400 total hours time-in-service, or within 1,000 hours time-in-service after the effective date of this AD, whichever occurs later, perform an initial visual and eddy current inspection to detect wear or cracks of the inner doubler on the pylon side panel around the fire extinguisher access doors, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992. If cracks are detected, prior to further flight, accomplish the requirements of either paragraph (a)(1) or (a)(2) of this AD, as applicable: (1) If cracks less than 5 mm (0.197 inch) are detected, prior to further flight, repair in an accordance with a method approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. (2) If cracks greater than or equal to 5 mm (0.019 inch) are detected, accomplish either paragraph (a)(2)(i) or (a)(2)(ii), as applicable: (i) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. (ii) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD; and repeat those inspections thereafter in accordance with paragraphs (b), (c), (d), or (e) of this AD, as applicable. (b) For airplanes having Configuration No. 1.A.: Repeat the visual and eddy current inspections to detect doubler wear and cracks, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,400 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,400 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 1,800 hours time-in-service. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Accomplish either paragraph (b)(5)(i), (b)(5)(ii), or (b)(5)(iii) of this AD. (i) Repeat the visual and eddy current inspection at intervals not to exceed 350 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (b) of this AD. (6) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 6 mm (0.236 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (c) For airplanes having Configuration No. 1.B: Repeat the visual and eddy current inspections for doubler wear and cracks, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detectedby any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,400 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,400 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Accomplish either paragraph (c)(4)(i), (c)(4)(ii), or (c)(4)(iii) of this AD. (i) Repeat the visual and eddy current inspection at intervals not to exceed 300 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after the replacement date of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (c) of this AD. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 4 mm (0.157 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (d) For airplanes having Configuration No. 1.A. on which the modification described in Airbus Industrie Service Bulletin No. A300-54-008 (door with 4 latches) has been accomplished; and for airplanes having Configuration No. 2.A.: Repeat the visual and eddy current inspections for doubler wear and cracks at zones D1, D2, D3, and D4, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is equal to or greater than to 0.1 mm (0.004 inch) and less than or equal to 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 5,800 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 5,800 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,300 hours time-in-service. (5) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Repeat the visual and eddy current inspection at intervals not to exceed 2,300 hours time-in-service. (6) For airplanes on which the measured doublerwear, as detected by any visual inspection required by this AD, is greater than 6 mm (0.236 inch) and less than 8 mm (0.315 inch): Accomplish either paragraph (d)(6)(i), (d)(6)(ii), or (d)(6)(iii) of this AD. (i) Repeat the visual and eddy current inspections at intervals not to exceed 450 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (d) of this AD. (7) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 8 mm (0.315 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (e) For airplanes having Configuration No. 1.B. on which the modification described in Airbus Industrie Service Bulletin No. A300-54-008 (door with 4 latches) has been accomplished; and for airplanes having Configuration No. 2.B.: Repeat the visual and eddy current inspections for doubler wear and cracks at zones D1, D2, D3, and D4, as required by paragraph (a) of this AD, in accordance with Airbus Industrie Service Bulletin A300-54-070, Revision 1, dated March 17, 1992, as follows: (1) No further action is necessary for the following airplanes: (i) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.5 mm (0.019 inch). (ii) Airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, and the measured doubler wear, as detected by any visual inspection required by this AD, is less than 0.1 mm (0.004 inch). (2) For airplanes on which the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished, if the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.1 mm (0.004 inch) and less than 0.5 mm (0.019 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,500 hours time-in-service. NOTE: Subsequent action to be taken (repetitive inspections or inner doubler replacement) will depend upon the "worst finding," defined as the area with the largest amount of measured doubler wear. (3) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 0.5 mm (0.019 inch) and less than 2 mm (0.078 inch): Repeat the visual and eddy current inspections at intervals not to exceed 4,500 hours time-in-service. (4) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 2 mm (0.078 inch) and less than 4 mm (0.157 inch): Repeat the visual and eddy current inspections at intervals not to exceed 3,300 hours time-in-service. (5) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 4 mm (0.157 inch) and less than 6 mm (0.236 inch): Repeat the visual and eddy current inspections at intervals not to exceed 1,800 hours time-in-service. (6) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than 6 mm (0.236 inch) and less than 8 mm (0.315 inch): Accomplish either paragraph (e)(6)(i), (e)(6)(ii), or (e)(6)(iii) of this AD. (i) Repeat the visual and eddy current inspections at intervals not to exceed 400 hours time-in-service. Or (ii) If the modification described in Airbus Industrie Service Bulletin A300-54-046 (screwed doors) has been accomplished: Prior to further flight, replace the doubler. Replacement of the doubler constitutes terminating action for the inspections required by this AD. Or (iii) If the modification described in Airbus Industrie Service Bulletin No. A300-54-046 (screwed doors) has not been accomplished: Prior to further flight, replace the doubler. Prior to the accumulation of 22,400 hours time-in-service after replacement of the doubler, perform the visual and eddy current inspections in accordance with paragraph (a) of this AD, and repeat those inspections thereafter in accordance with paragraph (e) of this AD. (7) For airplanes on which the measured doubler wear, as detected by any visual inspection required by this AD, is greater than or equal to 8 mm (0.315 inch): Prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (f) Accomplishment of the replacement of the inner doubler on the pylon side panel around the fire extinguisher access doors; and installation of the modification (screwed doors) in accordance with Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982, and Change Notice, dated July 8, 1985; constitutes terminating action for the visual and eddy current inspections required by this AD. (g) 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. (h) 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. (i) The modification shall be done in accordance with Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982; and Change Notice, dated July 8, 1985, for Airbus Industrie Service Bulletin No. A300-54-046, dated June 24, 1982. The inspections shall be done in accordance Airbus Industrie Service Bulletin No. A300-54-070, Revision 1, dated March 17, 1992, which contains the following list of effective pages: Page Number Revision Level Shown on Page Date Shown on Page 1-10, 31-32 1 March 17, 1992 11-30, 33-34 Original February 6, 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 Airbus Industrie, 1 Rond Point Maurice Bellonte, 31707 Blagnac Cedex, 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. (j) This amendment becomes effective on April 29, 1993.
90-05-05: 90-05-05 BOEING: Amendment 39-6516. Docket No. 89-NM-166-AD. \n\n\tApplicability: Model 767 series airplanes equipped with Kidde engine fire and overheat detection systems, as listed in Boeing Service Bulletin 767-26-0037, Revision 1, dated June 1, 1989, certificated in any category. \n\n\tCompliance: Required within 6 months after the effective date of the AD, unless previously accomplished. \n\n\tTo prevent false engine fire and overheat warnings, which could result in unnecessary engine in-flight shutdowns and airplane diversions that unduly jeopardize continued safe operation of the airplane, accomplish the following: \n\n\tA.\tModify the engine fire and overheat detection system on each engine in accordance with Boeing Service Bulletin 767-26-0037, Revision 1, dated June 1, 1989. \n\n\tB.\tAn 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: 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, Seattle Aircraft Certification Office.\n \n\tC.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment (39-6516, AD 90-05-05) becomes effective on March 27, 1990.
90-26-08: 90-26-08 BOEING: Amendment 39-6831. Docket No. 90-NM-252-AD. \n\n\tApplicability: Model 747-400 series airplanes, listed in Boeing Alert Service Bulletin 747-26A2171, dated October 4, 1990, certificated in any category. \n\n\tCompliance: Required within 45 days after the effective date of this airworthiness directive, unless previously accomplished. \n\n\tTo prevent the failure of the fire detection warning system to annunciate a fire, accomplish the following: \n\n\tA.\tModify the wiring of the aural and master warning indication systems of the fire detection system, in accordance with Boeing Alert Service Bulletin 747-26A2171, dated October 4, 1990. \n\n\tB.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tC.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. This information may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tThis amendment (39-6831, AD 90-26-08) becomes effective on December 26, 1990.
90-15-05: 90-15-05 BOEING: Amendment 39-6654. Docket No. 90-NM-24-AD. \n\n\tApplicability: Model 767 series airplanes, identified in paragraphs A. and B., below, certificated in any category. \n\n\tCompliance: Required within the next 20 months after the effective date of this AD, unless previously accomplished. \n\n\tTo provide satisfactory reliability of the evacuation system, accomplish the following: \n\n\tA.\tFor airplanes identified in Boeing Service Bulletin 767-25-0120, dated December 14, 1989: Modify the off-wing evacuation system (compartment door closed proximity sensor installation) in accordance with that service bulletin. \n\n\tB.\tFor airplanes identified in Boeing Alert Service Bulletin 767-25A0131, Revision 1, dated November 9, 1989: Modify the off-wing evacuation system by replacing the escape system actuator in accordance with Boeing Alert Service Bulletin 767-25A0131, Revision 1, dated November 9, 1989, or Boeing Service Bulletin 767-25-0011, Revision 2, dated October 12, 1989 or Revision 3, dated December 14, 1989.\n \n\tC.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant Principal Inspector (PMI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tE.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport AirplaneDirectorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThe amendment (39-6654, AD 90-15-05) becomes effective on August 20, 1990.
76-16-08: 76-16-08 PITTS AVIATION ENTERPRISES, INC.: Amendment 39-2692. Applies to Model S-1S, Serial Numbers 1-0001 through 1-0043, and Model S-2A, Serial Numbers 2001 through 2122, airplanes certificated in all categories. Compliance required as indicated, unless already accomplished: To prevent partial loss of rudder control due to control cable slippage resulting from improper swaging of the Nicopress sleeves, accomplish the following: A. Within 10 hours' time in service after the effective date of this airworthiness directive or by September 13, 1976, whichever occurs first, inspect the 3 swaged grooves of the Nicopress part number 18-3-M sleeves in the rudder control system to assure that the major axis diameter is between .348 and .353 inches. The Nicopress "go" gage for this sleeve is part number 64-CGMP and has an opening of .353 inches. The correct opening for checking these sleeves is marked "M." The Model S-2A airplane has a total of 8 sleeves to be checked and the Model S-1S has 4. If improperly swaged sleeves are found, before further flight, reswage to the proper limits or replace as required. B. Until compliance with paragraph A is accomplished, acrobatic flight (including but not limited to, maneuvers delineated in the respective FAA Approved Airplane Flight Manuals) is prohibited and prior to further flight a placard must be installed on the instrument panel in full view of the pilot which reads: "ACROBATIC FLIGHT PROHIBITED." C. Upon compliance with paragraph A, the placard required by paragraph B may be removed. This amendment becomes effective August 20, 1976.
90-10-07: 90-10-07 BOEING: Amendment 39-6592. Docket No. 89-NM-102-AD. \n\n\tApplicability: Model 767 series airplanes, line numbers 1 through 243, 257, 262, 263, 269, 273 through 275, 282, 285, 289, and 291, certificated in any category. \n\n\tCompliance: Required within the next 24 months after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent uncommanded extension of three flight spoilers on one wing, due to a failure of a spoiler wheel command unit, accomplish the following: \n\n\tA.\tFor Group 1 airplanes (as listed in the Boeing service bulletin): Replace both spoiler wheel command units in accordance with Boeing Service Bulletin 767-27-0085, Revision 1, dated November 30, 1989. \n\n\tB.\tFor Group 2 airplanes (as listed in the Boeing service bulletin): Replace the left side spoiler wheel command unit in accordance with Boeing Service Bulletin 767-27-0085, Revision 1, dated November 30, 1989. \n\n\tC.\tAn 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, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tNOTE: 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, Seattle Aircraft Certification Office. \n\n\tC.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplanes, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Seattle Aircraft Certification Office, 9010 East Marginal WaySouth, Seattle, Washington. \n\n\tThis amendment (39-6592, AD 90-10-07) becomes effective on June 12, 1990.
77-14-13: 77-14-13 HAWKER SIDDELEY AVIATION, LTD: Amendment 39-2958. Applies to Model BH-125 Series 600A airplanes, S/N's 25/6001-6004, 6007, 6009-6011, 6013, 6014, 6016, 6018, 6020, 6022-6026, 6032, 6034, 6038, 6040, 6044, 6046, certificated in all categories. Compliance is required within the next 300 hours time in service after the effective date of this AD, unless already accomplished. To prevent possible unwanted rolling of the airplane when operating at buffet onset, add vortex generators to the leading edge of each wing by incorporating Hawker Siddeley Aviation, Ltd., Modification No. 252442 in accordance with Section 2 entitled "Accomplishment Instructions" of Hawker Siddeley Aviation, Ltd., Service Bulletin 57-48-(2442), dated June 25, 1975, including Revision 1, dated July 23, 1975, or an FAA-approved equivalent. This amendment becomes effective August 8, 1977.
92-07-06: 92-07-06 BRITISH AEROSPACE, REGIONAL AIRCRAFT LIMITED: Amendment 39-8202. Docket No. 91-CE-72-AD. Supersedes AD 83-07-01, Amendment 39-4598. Applicability: Beagle B121 Pup series 1, 2, and 3 (all serial numbers) airplanes without Modification No. BE.214 incorporated on both mainplanes, certificated in any category. NOTE 1: Modification No. BE.214 is a replacement of the wing/fuselage joint plate that is equivalent to the requirements of this AD action. Compliance: Required upon the accumulation of 1,300 hours time-in-service (TIS) or within the next 100 hours TIS after the effective date of this AD, whichever occurs later, unless already accomplished. NOTE 2: The requirements of this AD may have been accomplished in accordance with superseded AD 83-07-01, Amendment 39-4598. To prevent failure of each wing spar, which could result in loss of control of the airplane, accomplish the following: (a) Inspect each wing spar web in the area immediately outboard of the root rib by accomplishing paragraphs 3.2.2 through 3.2.4 of paragraph 3. ACTION in BAe Pup Mandatory Service Bulletin (SB) B121/79, Revision 1, dated February 15, 1991. (1) If found cracked, prior to further flight, obtain an FAA-approved repair scheme from the manufacturer through the Brussels Aircraft Certification Office at the address specified in paragraph (d) of this AD, and incorporate this repair scheme. (2) If cracks are not found, prior to further flight, modify each wing spar web by incorporating Repair Scheme BE.03.10169 in accordance with Drawing No. BE.03.10169, and accomplish paragraphs 3.2.6 and 3.2.7 of paragraph 3. ACTION in BAe Pup Mandatory SB B121/79, Revision 1, dated February 15, 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 compliancetime that provides an equivalent level of safety may be approved by the Manager, Brussels Aircraft Certification Office, FAA, Europe, Africa, and Middle East Office, c/o American Embassy, B-1000 Brussels, Belgium. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. (e) The inspection and modification required by this AD shall be done in accordance with British Aerospace Pup Mandatory Service Bulletin (SB) B121/79, Revision 1, dated February 15, 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 British Aerospace, Regional Aircraft, Ltd., Manager Product Support, Prestwick Airport, Ayrshire, KA9 2RW Scotland. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, KansasCity, Missouri, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. (f) This amendment (39-8202) supersedes AD 83-07-01, Amendment 39-4598. (g) This amendment (39-8202, AD 92-07-06) becomes effective on April 30, 1992.
90-17-02: 90-17-02 ENSTROM HELICOPTER CORPORATION: Amendment 39-6688. Docket No. 90-ASW-28. \n\tApplicability: Enstrom Model F-28, F-28A, F-28C, F-28C-2, F-28F, 280, 280C, 280F, and 280FX series helicopters certificated in any category. \n\n\tCompliance: Required as indicated, unless already accomplished. \n\n\tTo prevent the loss of tail rotor thrust and directional control that could result in substantial damage or loss of the helicopter, accomplish the following: \n\n\t(a)\tWithin the next 5 hours' time in service after the effective date of this AD, inspect the tail rotor drive shaft couplings. Enter the part number of the two drive shaft couplings installed, the number of hours' time in service on each coupling, and the date in the log book. \n\n\tNOTE: Enstrom helicopters use one of the following three coupling designs: \n\n\t\t(1)\tsplined coupling P/N 28-13609-1, (2) five-plate flex pack coupling (Dana Corp. Element No. A005-1991); or (3) seven-plate flex pack coupling, (Dana Corp. Element No. A005-1992). \n\n\t(b)\tBefore further flight, remove any five-plate flex-pack couplings, Dana Corp. P/N A005-1991, found during the inspection. After removal: \n\n\t\t(1)\tInspect the condition of the coupling hub flanges, Enstrom P/N 28-13613-1 and 28-13614-1, to which the couplings mount. The flanges must be flat within 0.010-inch, and the bolt holes must not have any evidence of elongation. Replace any flange deformed beyond these limits with an airworthy part before further flight; and \n\n\t\t(2)\tInstall the seven-plate flex-pack coupling, Enstrom Kit No. 28-01041-1. \n\n\tNOTE: Enstrom Kit No. 28-01041-1 includes all of the necessary parts to replace a five-plate coupling with a seven-plate coupling. Two kits per rotorcraft are required. The replacement beveled washers for the seven-plate flex-pack coupling are 0.010-inch thinner than beveled washers for the five-plate flex-pack coupling. Consequently, a direct exchange should not require any additional shimming. \n\n\t(c)\tWithin the next 100 hours' time in service and at 100-hour intervals thereafter, inspect each seven-plate flex-pack coupling, Dana Corp. P/N A005-1992 or Enstrom Kit No. 28- 01041-1, for compliance with deformation and airworthiness limits as follows: \n\n\t\t(1)\tLocate the tension and compression sides by comparing the couplings to Figure 1. \n\n\t\t(2)\tInspect the couplings for flex pack distortion in the shape of a bow. Compare any distortion to the acceptable limits shown in Figure 2. \n\n\t\t(3)\tInspect the couplings for flex pack distortion in the shape of an offset bend. The acceptable limits shown in Figure 3 are as follows: \n\n\t\t\t(i)\tThe offset on one set of plates must not exceed 0.015 inches. \n\n\t\t\t(ii)\tThe maximum difference in offset from side to side (one pair of bolts to the other pair) must be less than 0.007 inches. \n\n\t\t\t(iii)\tThe maximum allowable shim thickness is 0.072 inches (not including bevel washers). \n\n\tNOTE: The replacement of the five plate coupling with the seven-plate coupling requires a thinner set of beveled washers which are included in Enstrom Kit Number 28-01041-1. \n\n\t\t(4)\tReplace any seven-plate coupling, that either-- \n\n\t\t\t(i)\tExceeds the airworthiness limits specified by paragraph (c); or, \n\n\t\t\t(ii)\tHas accumulated 1,200 hours' time in service. \n\n\n\n\n\t\t\t AD 90-17-02 FIGURE 1 \n\n\t\t\tTension Sides Versus Compression Sides \n\t\t\tTail Rotor Drive Shaft Coupling Installation. \n\t\t\tRear coupling shown, forward coupling similar. \n\t\t\tNOTE: Tie-wraps eliminated for clarity. \n\n\t(d)\tAfter removal of either seven-plate flex-pack coupling, inspect and replace, as necessary, the coupling hub flanges in accordance with paragraph (b)(1) of this AD. \n\n\t(e)\tIn accordance with FAR Sections 21.197 and 21.199, flight is permitted to a base where the requirements of this AD may be accomplished. \n\n\t(f)\tAn alternate method of compliance or adjustment of the compliance time, which provides an equivalent levelof safety, may be used if approved by the Manager, Chicago Aircraft Certification Office, ACE-115C, FAA, 2300 East Devon Avenue, Des Plaines, Illinois 60018. \n\n\tThis amendment (39-6688, AD 90-17-02) becomes effective on September 7, 1990. \n\n\n\nFigure 2 - Acceptable Limits of Bowed Flex Packs. \nNOTE: Tie-wraps eliminated for clarity. \nAD 90-17-02 \n\n\n\t\n\t\t\tFigure 3 - Acceptable Limits of Flex Packs with Offset Bends. \n NOTE: Tie-wraps eliminated for clarity.\n AD-90-17-02
92-11-10: 92-11-10 MCDONNELL DOUGLAS: Amendment 39-8260. Docket No. 91-NM-232-AD. Supersedes AD 91-13-09, Amendment 39-7040. \n\n\tApplicability: Model DC-9-20, -30, -40, -50, and C-9 (Military) series airplanes; which correspond to factory serial numbers listed in McDonnell Douglas Service Bulletin 27-196, Revision 1, dated September 28, 1984, or Revision 2, dated December 17, 1990; McDonnell Douglas Service Bulletin 27-250, dated August 29, 1984, or Revision 1, dated October 18, 1984, or Revision 2, dated January 3, 1990; and McDonnell Douglas Alert Service Bulletin A27-250, Revision 3, dated May 15, 1991; certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\tTo detect cracks and prevent failure of the slat drive mechanism and its interrelated structure, accomplish the following: \n\n\t(a)\tWithin 20 days or 135 landings, whichever occurs first after January 10, 1985 (the effective date of AD 84-24-03, Amendment 39-4956), inspect the left and right actuator slat drive mechanism in accordance with McDonnell Douglas Service Bulletin 27-196, Revision 1, dated September 28, 1984, or Revision 2, dated December 17, 1990. \n\n\t\t(1)\tIf no cracks are found, no further inspection is required. \n\n\t\t(2)\tIf a crack is found, and the crack is less than one inch in length, continue to inspect the actuator slat drive shaft at intervals not to exceed 1,600 landings in accordance with the service bulletin. \n\n\t\t(3)\tIf a crack is found, and the crack is one inch or greater in length, prior to further flight, replace the actuator slat drive shaft in accordance with Condition II specified in the service bulletin. Such replacement constitutes terminating action for the inspections required by paragraph (a)(2) of this AD. \n\n\t(b)\tWithin 20 days or 135 landings, whichever occurs first after January 10, 1985, inspect the forward slat drive drums' bellcrank shafts that have accumulated 4,000 or more landings since new or last overhaul.(1)\tIf no cracks are detected, continue to inspect the slat drive drum bellcrank shaft for cracks at intervals not to exceed 1,500 landings as shown in Figure 1 of McDonnell Douglas Service Bulletin 27-250, dated August 29, 1984; Revision 1, dated October 18, 1984; or Revision 2, dated January 3, 1990. \n\n\t\t(2)\tIf cracks are found, prior to further flight, replace the slat drive drum bellcrank shaft in accordance with Condition II of the service bulletin. Such replacement constitutes terminating action for the repetitive inspection requirements of paragraph (b)(1) of this AD. \n\n\t(c)\tWithin 500 landings after July 8, 1991 (the effective date of AD 91-13-09, Amendment 39-7040), or at the next scheduled inspection in accordance with paragraph (b) of this AD, whichever occurs earlier, inspect the slat drive drums' bellcrank shafts for cracks, in accordance with McDonnell Douglas Alert Service Bulletin A27-250, Revision 3, dated May 15, 1991. This inspection constitutes terminating action for the repetitive inspection requirements of paragraph (b)(1) of this AD. \n\n\t\t(1)\tIf no cracks are found, repeat the inspection of the slat drive drum bellcrank shaft for cracks at intervals not to exceed 750 landings, in accordance with the alert service bulletin. \n\n\t\t(2)\tIf cracks are found, prior to further flight, replace the slat drive drum bellcrank shaft with a new drum shaft, P/N 5920212-505, in accordance with Condition II specified in the alert service bulletin. Such replacement constitutes terminating action for the repetitive inspection requirements of paragraph (c)(1) of this AD. \n\n\t(d)\tIf cracks are found in locations in the slat drive shaft(s) other than those specified in McDonnell Douglas Service Bulletin 27-196, Revision 1 or 2; and McDonnell Douglas Alert Service Bulletin A27-250, Revision 3; prior to further flight, replace or rework the cracked component(s) in a manner approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Transport Airplane Directorate. \n\n\t(e)\tReplacement of both the actuator slat drive mechanism and the slat drive drum bellcrank shaft in accordance with Condition II of the following service bulletins, as applicable, constitutes terminating action for the requirements of this AD: \n\t\n\nMcDonnell Douglas \nService Bulletin Number\nRevision Level\n\tDate \n27-196\t\t\nRevision 1\nSeptember 28, 1984 \n\nRevision 2\t\nDecember 17, 1990 \n\n\n\n27-250\t\nOriginal\t\nAugust 29, 1984 \n\nRevision 1\t\nOctober 18, 1984 \n\nRevision 2\t\nJanuary 3, 1990 \n\n\n\nA27-250\t\t\nRevision 3\nMay 15, 1991 \n\t\n\t(f)\tWithin 18 months or 2,250 flight cycles after the effective date of this AD, whichever occurs later, replace the slat drive drum bellcrank shaft with a new drum shaft, P/N 5920212-505, in accordance with Condition II specified in McDonnell Douglas Alert Service Bulletin 27-250, Revision 3, dated May 15, 1991. Such replacement constitutes terminating action for the repetitive inspection requirements of paragraphs (b)(1) and (c)(1) of this AD. \n\n\t(g)\tSpecial 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. \n\n\t(h)\tAn alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles ACO. \n\n\tNOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the Los Angeles ACO. \n\n\t(i)\tUpon the request of an operator, an FAA Maintenance Inspector, subject to prior approval by the Manager, Los Angeles Aircraft Certification Office, FAA, Transport Airplane Directorate, may adjust the inspection times specified in this AD to permit compliance at an established inspection period of that operator if the request contains substantiating data to justify the change for that operator. \n\n\t(j)\tThe inspection and replacement requirements shall be done in accordance with McDonnell Douglas Service Bulletin 27-196, Revision 1, dated September 28, 1984, or Revision 2, dated December 17, 1990; McDonnell Douglas Service Bulletin 27-250, dated August 29, 1984, or Revision 1, dated October 18, 1984, or Revision 2, dated January 3, 1990; and McDonnell Douglas Alert Service Bulletin A27-250, Revision 3, dated May 15, 1991. This incorporation by reference was previously approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51, as of July 8, 1991 (56 FR 28479, June 21, 1991). Copies may be obtained from McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Business Unit Manager of TechnicalPublications - Technical Administrative Support, C1-L5B (54-60). Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at FAA, Transport Airplane Directorate, Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California 90806-2425; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. \n\n\t(k)\tThis amendment becomes effective on July 23, 1992.
77-20-02: 77-20-02 ENSTROM: Amendment 39-3046. Applies to Enstrom Models F28, F28A, F28C, 280 and 280C helicopters certificated in all categories with a 3.3 inch tail rotor blade with spindle P/N 28-15202. Within the next 50 hours time in service or six (6) months calendar time, whichever occurs first, accomplish the following unless already complied with: 1. Replace tail rotor spindle P/N 28-15202 with a new spindle P/N 28-15202-13 or a manufacturer-reworked spindle which may be identified by an eight (8) character serial number acid-etched on the outside of the spindle center section. Reference should be made to Enstrom Maintenance Manual Section 15 for disassembly and reassembly instructions. Special care should be taken to have the three bearing stack positioned with the closed side toward the hub and the retaining nut properly torqued to 35-40 foot-pounds and safetied. 2. Install strike tabs P/N 28-17308 extending forward on the blade leading edge under the outer balancingscrew. This supersedes Amendment 39-2839 (42 FR 10841) AD 77-04-04. This amendment is effective October 6, 1977, and was effective for all recipients of the airmail letter dated August 8, 1977, as amended by airmail letter dated August 22, 1977, upon receipt thereof.
96-22-05: This amendment adopts a new airworthiness directive (AD), applicable to certain British Aerospace Model BAe 146 series airplanes and Model Avro 146-RJ series airplanes, that requires a one-time inspection of terminal block "D" to ensure that a two-way link is installed, and installation of a new link, if necessary. This amendment is prompted by a report indicating that a two-way link that should be installed on direct current (DC) panel No. 1 may be missing from certain airplanes. The actions specified by this AD are intended to ensure that a two-way link is installed. If the link is not installed, it could result in loss of the emergency electrical system and, consequently, increased pilot workload and possible reduced controllability of the airplane.
78-11-10: 78-11-10 SHORT BROTHERS LIMITED: Amendment 39-3218. Applies to Model SD3-30 airplanes, certificated in all categories. Compliance required as indicated. To detect cracking and prevent possible failure of the main landing gear downlock stop, P/N 17780-1 or -3, which could result in collapse of the gear, accomplish the following: (a) Prior to accumulating a total of 2000 landings or within the next 50 landings after the effective date of this AD, whichever occurs later, unless already accomplished within the last 450 landings, and thereafter at intervals not to exceed 500 landings, inspect the downlock stop for cracks using a dye penetrant method of inspection in accordance with paragraph 7, "Accomplishment Instructions", and figure 1 of the Menasco Manufacturing Service Bulletin 32- 18, Revision 1, dated October 13, 1977, or an FAA approved equivalent. (b) If a crack is found during an inspection required by paragraph (a) of this AD, before further flight, exceptthat the airplane may be flown in accordance with FAR 21.197 and 21.199 to a base where the repair can be performed - (1) Replace the downlock stop with a new part of the same part number and continue to inspect in accordance with paragraph (a) of this AD; or (2) Replace the stop as specified in paragraph (c) of this AD. (c) The inspections required by this AD may be terminated upon replacement of the downlock stop, P/N 17780-1 or -3, with a stop, P/N 17780-5, in accordance with paragraph 10, "Accomplishment Instructions", of Menasco Manufacturing Service Bulletin 32-7, Revision 1, dated October 12, 1977, or an FAA approved equivalent. NOTE: Short Brothers Limited Service Bulletins SD3-32-25 and SD3-32-18 cover this same subject. This amendment becomes effective June 6, 1978.
77-02-08: 77-02-08 NARCO: Amendment 39-2818. Applies to all Narco Model ELT-10 77 Emergency Locator Transmitter batteries, with a battery replacement date prior to November 1977. Compliance required as indicated. As a result of battery corrosion which results in battery lead breakage and the resulting failure of the ELT to operate, accomplish the following: (a) Within the next 90 days and thereafter, at each additional 90-day interval after the effective date of this Airworthiness Directive, accomplish the following: 1. Remove the control head from battery pack by removing four screws. 2. Check for battery leakage, which will appear as a white residue which normally occurs first around battery leads. 3. Check for corrosion at the battery terminals, printed circuit board and components. 4. Check for secure battery leads. 5. If no corrosion is found, determine ELT transmits properly in accordance with FAA Advisory Circular 91-44 or 20-81, and reinstall in accordance with paragraph 8 of the Narco Owner's Manual or approved alternate installation procedures. 6. If corrosion is found, replace battery (Narco P/N 57674-0001) with one which has a replacement date of November 1977, or later. 7. The checks required by this Airworthiness Directive may be conducted by the pilot or aircraft owner. (Narco Service Bulletin ELT-10 No. 6, dated August 2, 1976, covers this same subject.) This amendment is effective February 1, 1977.

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