Results
49-40-01: 49-40-01 LUSCOMBE: Applies to All Model 11A Aircraft. Compliance required on or before the next periodic inspection but not later than December 1, 1949. To preclude the possibility of the elevator trim tab actuating horn becoming disconnected from the trim tab, with consequent serious vibration of the horizontal tail surfaces, it is necessary to rework the attachment of the trim tab horn by adding more rigidity to the attachment. This rework can be accomplished by fabricating two blocks from solid 24ST aluminum alloy that will fit inside the inboard end of the trim tab, one located at the extreme inboard end to which the steel trim tab horn attaches and other one located diagonally chordwise inside the trim tab, with the forward end located approximately 2 1/2 inches and the aft end approximately 1 inch from the inboard end of the trim tab. These blocks, which actually are equivalent to solid ribs, should be approximately 3/8-inch wide and shaped in elevation to fit the inside contour of the trim tab. The attachment of these ribs should be effected by four AN 456AD4 rivets in each, drilled on assembly, with the rivets driven through both upper and bottom skins of the trim tab. The trim tab horn should be attached to the trim tab through their regular attaching holes, riveting the horn with two AN 456AD4 rivets to the chordwise end of the inboard revised solid rib and the two remaining holes as originally attached with two AN 456AD4 rivets. The aluminum alloy blocks or ribs should be finished with a protective coating of zinc chromate prior to assembly of the trim tab. An equivalent modification to that described above and in Luscombe Service Bulletin is acceptable. (Luscombe Service Bulletin No. 1-1149, dated January 25, 1949, covers this same subject.)
64-27-01: 64-27-01 BEECH: Amdt. 39-6 Part 39 (New) Federal Register December 1, 1964. Applies to Models P35 and S35 Aircraft, Serial Numbers D-6842 through D-7630, except Serial Numbers D-7610, D-7615, D-7617, D-7620, and D-7625 through D-7628. Compliance required within 25 hours' time in service after the effective date of this AD, unless already accomplished. To prevent the elevator control column stop fitting from binding between overtrimmed stops or going past the stops and jamming the control column in the full aft (up elevator) position, accomplish the following. (a) Remove the large inspection plate on the left firewall. With the control column in the most aft position, mark the contact area on the fixed stops, using the control column fitting stop as a guide. The combined contact area width on the two stops measured in a horizontal plane shall be at least 3/16 inch. (b) Fabricate and install a new elevator control fixed stop on aircraft having less than 3/16 inchcombined contact area width on the two stops measured in a horizontal plane in accordance with Beech Service Bulletin No. 64-19, or an FAA-approved equivalent. (Beech Service Bulletin No. 64-19, covers this same subject.) This directive effective December 7, 1964.
64-26-03: 64-26-03 GRUMMAN: Amdt. 39-4 Part 39 (New) Federal Register November 25, 1964. Applies to Models TBM and TBF Series Aircraft. Compliance required as indicated. As a result of loose and sheared rivets found on the elevator push rod assembly, P/N 21399, accomplish the following: (a) Within 50 hours' time in service after the effective date of this AD, unless already accomplished within the last 50 hours' time in service, and thereafter within every 100 hours' time in service from the last inspection, inspect the four rivets on the elevator push-pull rod assembly which retain the clevis and bearing terminal fittings, P/N's G76-10 and 21356, to the tube, P/N 21399-1, and determine if any are loose or sheared. (The rod assembly, which connects the elevator horn to the sector, is located at the airplane centerline at approximately fuselage Station 326.) (b) If any rivets are determined to be loose or sheared, accomplish the following modification before further flight, except that one flight may be made in accordance with the provisions of CAR 1.76 for the purpose of obtaining these repairs: (1) Remove the four G10-D3-102 rivets. If the rivet holes in the tube or terminal fittings are elongated beyond the maximum diameter for the replacement rivet, replace that part with a new part of the same part number, or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (2) Plug both hollow terminal fitting shanks with an aluminum alloy plug, or an FAA approved equivalent. The plug must be of such diameter so that it may be inserted into the terminal fitting shank creating an interference fit of 0.0005 inch, plus 0.000 inch, minus 0.0005 inch. The plug length must be long enough to fill the area occupied by the rivets. (3) Drill all four holes, including the plugs, to accommodate MS 20470AD8 rivets, or an FAA approved equivalent. (4) Install all four MS 20470AD8 replacement rivets, or an FAA approved equivalent. (c) The 100-hour repetitive inspections of (a) need not be made on aircraft which already incorporate aluminum alloy terminal plugs and four MS 20470AD8 rivets, or an FAA approved equivalent, and may be discontinued on aircraft that are modified in accordance with (b)(1) through (b)(4) inclusive, or an FAA approved equivalent. This directive effective December 25, 1964.
50-15-01: 50-15-01 GRUMMAN: Applies to All Model G-21A Aircraft (Converted JRF-5, JRF- 6B) Equipped With Reverse Direction Mixture Controls. Compliance required not later than next 25-hour inspection. To conform with conventional mixture control operation ("forward" for full rich position) on aircraft equipped with Bendix NAR9B carburetors with manual mixture control, rotate the position of the mixture bellcranks 180 degrees on the carburetors and reverse the tooth segments on the cockpit control end for end. Revise the cockpit control placard accordingly. On aircraft equipped with Bendix NAR9C2 carburetors with automatic mixture control, the cockpit quadrant is already arranged in the correct sense and requires no revision. It should be noted that an additional Manual Lean position is provided forward of Full Rich and caution must be exercised to prevent inadvertently positioning the control incorrectly if the Manual Lean sector of the quadrant is retained. This supersedes AD 48-14-02.
47-42-01: 47-42-01\tDOUGLAS: (Was Mandatory Note 20 of AD-762-7.) Applies to DC-4 and C-54 Aircraft. \n\nTo be accomplished not later than April 1, 1948. \n\nTo prevent the possibility of the gust lock control becoming engaged in flight or during taxiing, a latch assembly must be installed to safety the control handle in the gust lock "OFF" position. Early aircraft incorporated a short gust lock control handle. In later aircraft, the control handle design was changed and the length of the handle increased to provide more leverage. On aircraft incorporating the short gust lock control handle, latch assembly, P/N 3356892, must be installed. In aircraft incorporating the new and longer handle, latch assembly, P/N 4356957, must be installed and the gust lock handle link assembly, P/N 4248396, must be reworked by removing and replacing the spring, P/N 2356732 (or 1248420), and plunger, P/N 1248421, with new bolt P/N 1356885. \n\nIn addition to the above, the elevator and rudder gust lock in the tail section and the aileron gust lock in the fuselage center section must be reworked by removing shaft, P/N 1165889, and replacing with new piston, P/N 2356840. After completing the rework, care must be exercised improperly rigging the gust lock control system. \n\n(NOTE: Some operators have obtained approval of a gust lock latch of their own design. In such cases, the Douglas designed latch need not be installed, however, the remainder of the rework described above must be accomplished.) \n\n(Douglas Service Bulletin DC-4 No. 79 covers this same subject.)
64-19-03: 64-19-03 DOUGLAS: Amdt. 784 Part 507 Federal Register August 7, 1964. Applies to Model DC-8 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tTo assure flap system reliability and to eliminate difficulties which could result in the inadvertent retraction of the flaps during critical portions of the flight regime, accomplish the following: \n\n\t(a) Modify each aircraft within 6,000 hours' time in service after the effective date of this AD to incorporate a flap lockout system per Douglas DC-8 Service Bulletin No. 27-132, Revision No. 2, dated December 12, 1963, or an FAA Western Region, Aircraft Engineering Division approved equivalent modification. When the flap system is modified as indicated, the hose inspection and replacement provisions of this AD may be discontinued. \n\n\t(b) Until the modification required by (a) is incorporated, inspect the flap system hoses and remove the hoses from service as follows: \n\n\t\t(1) Within 900 hours' hose time in service since the last inspection or 90 days after the effective date of this AD, whichever occurs first, inspect all flaps hoses bearing Douglas basic P/N 5716378-4 for any evidence of cracking, splitting, abrasion, or other damage to the covering. Reinspect at intervals of 900 hours hose time in service or 90 days, whichever occurs first, until the modification in (a) is accomplished. Remove from service any hose on which only the covering is found to be cracked or abraded and the hose itself is found to be undamaged prior to the next 300 hours hose time in service or 30 days, whichever occurs first. Remove from service any hose exhibiting damage other than cracked or abraded covering before further flight. Remove from service undamaged hoses bearing Douglas basic P/N 5716378-4 prior to 1,500 hours total hose time in service. The inspections and the removal from service requirements of this paragraph apply also to new hoses installed as replacements pursuant to this paragraph. \n\n\t\t(2) Priorto 1,800 hours total hose time in service, remove from service all hoses bearing Douglas basic P/N S5773937-4 (same length code) or Aeroquip basic P/N 611049-4 (same length code). \n\n\t\t(3) Remove from service all hoses bearing Douglas basic P/N S5776432-4 (same length code) or Aeroquip basic P/N's 677219-4, 677220-4 (same length code) as follows: \n\n\t\t(i) Remove from service hoses with less than 1,900 hours total hose time in service on the effective date of this AD prior to 2,000 hours total hose time in service. \n\n\t\t(ii) Remove from service hoses with 1,900 or more hours total hose time in service on the effective date of this AD within the next 100 hours hose time in service. \n\n\t\t(iii) Remove from service hoses installed as replacements under this paragraph prior to 2,000 hours hose time in service. \n\n\t\t(4) Remove from service all hoses bearing Douglas basic P/N S5778051 (same length code) or Resistoflex basic P/N's R23718-4, R23708-4 (same length code) as follows:(i) Remove from service hoses with less than 2,300 hours total hose time in service on the effective date of this AD prior to 2,400 hours total hose time in service. \n\n\t\t(ii) Remove from service hoses with 2,300 or more hours total hose time in service on the effective date of this AD within the next 100 hours hose time in service. \n\n\t\t(iii) Remove from service hoses installed as replacements under this paragraph prior to 2,000 hours hose time in service. \n\n\t\t(5) Prior to 2,000 hours total hose time in service, remove from service all hoses bearing Aeroquip basic P/N's 677233-4, 677235-4 (same length code) or Resistoflex basic P/N's R24717-4, R24718-4, R24719-4 (same length code). \n\n\t(c) Do not install green or black flap actuating cylinder hoses dated prior to 1962. \n\n\t(Douglas DC-8 Service Bulletins No. 27-113, Reissue No. 1 dated November 14, 1962, No. A27-146, Reissue No. 2, dated December 27, 1963, and No. 27-132, Revision No. 2 dated December 12, 1963, pertain tothis same subject.) \n\n\tThis supersedes AD 62-20-01. \n\n\tThis directive effective September 7, 1964. \n\n\tRevised October 13, 1964.
2008-22-18: We are adopting a new airworthiness directive (AD) for certain Cessna Aircraft Company (Cessna) 150 series airplanes with the BRS-150 Parachute System installed via Supplemental Type Certificate (STC) SA64CH. This AD requires you to replace the pick-up collar support and nylon screws for the BRS-150 Parachute System. This AD results from notification by Ballistic Recovery Systems, Inc. (BRS) that the pick-up collar assembly may prematurely move off the launch tube and adversely affect rocket trajectory during deployment. We are issuing this AD to prevent premature separation of the collar. This condition could result in the parachute failing to successfully deploy.
49-16-01: 49-16-01 GRUMMAN: Applies to Model G-21A Aircraft Serial Numbers B-34, B-35, B- 38 Through B-42, B-45 Through B-51, B-53, B-54, B-55, B-57 Through B-61, B-63, B-64, B-65, B-67, B-68, B-70, B-71, B-74, B-76, B-77, B-82, B-83, B-85 Through B-90, B-92, B-96 Through B-99, B-101, B-106, B-107, B-111, B-116, B-118, B-119, B-120, B-124, B-125, B-127 Through B-134, B-137 Through B-141, B-143, B-144, and B-145. Compliance required as indicated. By June 1, 1949, inspect the fuel tank baffles at wing Stations 42, 54, and 75 through the handholds in bottom of integral fuel tanks. If baffles are found riveted to angle stiffness, no further action is required. If baffle stiffeners are attached by spotwelds, inspect for cracks. Airplane may continue in service, if no cracks are found in baffles, providing inspection is repeated each 100 hours. If cracking is not extensive and no spotwelds are broken from ribs, the airplane may be operated if inspected each 50 hours. Extensively cracked baffles should be repaired by replacing spotwelded baffles with riveted baffles. For further details, contact Grumman Aircraft Engineering Corporation, Bethpage, N.Y.
71-26-01: 71-26-01 BOEING: Amdt. 39-1362 as amended by Amendment 39-1387. Applies to Model 727 series airplanes listed in Boeing Service Bulletin 32-196, dated 16 September 1971, and Revision 1, dated 24 November 1971, or later FAA approved revisions, incorporating main landing gear actuator beam support link shaft P/N 69-19167-1 and -2. \n\n\tCompliance required as indicated: \n\tTo detect cracks in the main landing gear actuator beam support link shaft, accomplish the following: \n\n\t(a)\tFor all shafts which have accumulated 12,000 or more landing cycles on or after 25 January 1972, inspect the shaft within the next 1000 landings after 25 January 1972, unless already accomplished within the last 1000 landings, and thereafter at intervals not to exceed 2000 landings since the last inspection, per (b) below, until the shaft is replaced or reworked per (c) and (d) below. \n\t(b)\tInspect the shaft in accordance with Boeing Service Bulletin 32-196, Revision 1, dated 24 November 1971, or later FAA approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA, Western Region. If evidence of a crack is found, replace the shaft, prior to further flight, with shaft P/N 69-19167-3 or with a shaft that (1) has accumulated less than 12,000 landing cycles, or (2) has been previously inspected per this AD, and found to be uncracked, or (3) has been reworked per (c) below. \n\t(c)\tRework or replace shafts per Boeing Service Bulletin 32-196, Revision 1, dated 24 November 1971, or later FAA approved revisions, or an equivalent rework approved by the Chief, Aircraft Engineering Division, FAA, Western Region. \n\t(d)\tWithin 16,000 landings after such rework per (c) above, replace all shafts with acceptable shafts as identified in (b) above. Identify the shafts replaced so as to prevent inadvertent return to service. \n\t(e)\tFor the purpose of this AD, when conclusive records are not available to show the number of landings accumulated by a particular shaft, the number of landings may be computed by dividing the airplane time in service since the shaft was installed in the airplane by the operator's fleet average time per flight for his model 727 airplanes. \n\t(f)\tInspections prescribed by this AD do not apply to new replacement shafts P/N 69- 19167-3 installed on Boeing 727 aircraft. \n\n\tAmendment 39-1362 became effective January 25, 1972.\n\tThis Amendment 39-1387 becomes effective February 4, 1972.
64-15-03: 64-15-03 MCDONNELL DOUGLAS: Amendment 757 Part 507 Federal Register July 7, 1964 as amended by Amendment 801 (29 F.R. 12068), is further amended by Amendment 39-1287. Applies to Model DC-8 Series Aircraft (Except Model DC-8F). \n\n\tCompliance required as indicated. \n\n\tRecent fatigue cycle testing of some flap system components has shown a need to establish new service life limits. Accordingly, the following flap system components must be retired in accordance with applicable schedules specified herein: \n\n\t(a) Flap cylinder rod end bearing P/N 4648686-501, Station X(subscript w) = 97.906 and Station X(subscript F) #219.498. \n\n\t\t(1) Part Number 4648686-501 with more that 12,000 hours' time in service or more than 4,300 landings on the effective date of this AD shall be retired from service within 500 hours' time in service after the effective date of this AD. \n\n\t\t(2) Part Number 4648686-501 with 12,000 or less hours' time in service and 4,300 or less landings on the effective date of this AD, and parts installed subsequent to the effective date of this AD shall be retired prior to the accumulation of 12,500 hours' time in service or 4,500 landings, whichever occurs first. \n\n\t(b) Flap cylinder rod end bearing P/N 4648686-503, Station X(subscript W) = 97.906 and Station X(subscript F) = 219.498. \n\n\t\t(1) Part Number 4648686-503 with more than 13,500 hours' time in service or more than 4,400 landings on the effective date of this AD shall be retired from service within 500 hours' time in service after the effective date of this AD. \n\n\t\t(2) Part Number 4648686-503 with 13,500 or less hours' time in service and 4,400 or less landings on the effective date of this AD, and parts installed subsequent to the effective date of this AD shall be retired from service prior to the accumulation of 14,000 hours' time in service or 5,000 landings, whichever occurs first. \n\n\t(c) Flap cylinder rod end bolt P/N 2645104, Station X(subscript W) = 97.906, and Station X(subscript F) = 219.498 and 339.723. \n\n\t\t(1) Part Number 2645104 with more than 27,500 hours' time in service or more than 9,800 landings on the effective date of this AD shall be retired from service within 500 hours' time in service after the effective date of this AD. \n\n\t\t(2) Part Number 2645104 with 27,500 or less hours' time in service and 9,800 or less landings on the effective date of this AD, and parts installed subsequent to the effective date of this AD shall be retired prior to the accumulation of 28,000 hours' time in service or 10,000 landings, whichever occurs first. \n\n\t(d) Compensating flap actuating link P/N's 3648269, 3648268, 3648270, Station X(subscript W) = 97.906, Station X(subscript F) = 219.488, Station X(subscript F) = 399.723, respectively. \n\n\t\t(1) Part Numbers 3648268, 3648269, 3648270 with no bushings installed in the 0.9995/1.005 inch-diameter holes, and with more than 11,800 hours' time in service or more than 4,200 landings on the effective date of this amendment, shall be inspected and reworked as specified in (d)(5) within the next 500 hours' time in service, except as provided in (e)(1). \n\n\t\t(2) Part Numbers 3648268, 3648269, 3648270 with no bushings installed in the 0.9995/1.005 inch-diameter holes and with 11,800 or less hours' time in service and 4,200 or less landings on the effective date of this amendment and parts installed subsequent to the effective date of this amendment shall be inspected and reworked as specified in (d)(5), prior to the accumulation of 12,800 hours' time in service or 4,400 landings, whichever occurs first, except as provided in (e)(1). \n\n\t\t(3) Part Numbers 3648268, 3648269, 3648270 with bushings installed in the 0.9995/1.005 inch-diameter holes and having more than 4,500 hours' time in service or more than 1,600 landings on the effective date of this amendment shall be inspected and reworked as specified in (d)(5) within 500 hours' time in service after the effective date of this amendment. \n\n\t\t(4) Part Numbers 3648268, 3648269, 3648270 with bushings installed in the 0.9995/1.005 inch-diameter holes and with 4,500 or less hours' time in service and 1,600 or less landings on the effective date of this amendment shall be inspected and reworked as specified in (d)(5) prior to the accumulation of 5,000 hours' time in service or 1,780 landings, whichever occurs first. \n\n\t\t(5) Parts which have not been reworked in accordance with Douglas Service Bulletin No. 27-144, dated June 7, 1963, as of the effective date of this amendment shall be inspected and reworked in accordance with the accomplishment instructions of Douglas Service Bulletin No. 27-144, Reissue No. 1, dated August 3, 1964, or an FAA-approved equivalent, within the time limits specified in (d), (1), (2), (3), or (4), as appropriate. \n\n\t(e) Service life limits for reworked parts are as follows: \n\n\t\t(1) Parts which have been reworked in accordance with Douglas Service Bulletin No. 27-144, datedJune 7, 1963, as of the effective date of this amendment may be continued in service not to exceed 10,000 landings or an additional 28,000 hours' time in service from the time of rework of the 0.250/0.25A diameter holes, whichever occurs first, and then must be retired from service. \n\n\t\t(2) Parts with the bushings installed as noted in (d)(3) and (d)(4) may be continued in service not to exceed an additional 3,570 landings or 10,000 hours' time in service, whichever occurs first, from the time of rework of the 0.250/0.25A diameter holes and then must be retired from service. \n\n\t\t(3) Parts without bushings installed during rework of the 0.250/0.25A diameter holes, may be continued in service not to exceed an additional 10,000 landings or 28,000 hours' time in service, whichever occurs first, from the time of rework, and then must be retired from service. \n\n\tNOTE: Due to the improved orientation of the 0.250/0.254 inch diameter lock pin hole, compensating flap actuating link P/N's 3648268, 3648269, and 3648270 which have had that hole drilled in accordance with McDonnell Douglas Assembly Drawing Nos. 4717382 (Change letter B or later), 4717381 (change letter A or later), and 4717383 (change letter B or later), respectively, are not subject to the service life limits of Paragraph (e) above. \n\n\t(f) Parts with evidence of cracks remaining after the initial chamfer is specified in Douglas Service Bulletin No. 27-144, Reissue No. 1, dated August 3, 1964, are not eligible for further use. \n\n\t (Douglas Service Bulletins No. 27-127, Revision No. 2 dated January 28, 1964, and No. 27-144 Reissue No. 1 dated August 3, 1964, cover this same subject.) \n\n\tThis directive became effective July 7, 1964. \n\n\tRevised August 25, 1964. \n\n\tThis Amendment 39-1287 becomes effective September 11, 1971.
44-20-02: 44-20-02 BOEING: (Was Service Note 1 of AD 719-1 and Service Note 1 of AD-726-1.) Applies to 307 Series Aircraft. \n\tInspect by visual means all square aluminum alloy 24SRT tubing for cracks in the following locations: wing spars, front spar fuselage bulkhead, rear spar fuselage bulkhead, and fin and stabilizer attachment bulkheads. These inspections shall be conducted at intervals specified and in the following manner: \n\tA.\tSA-307B. In the inspection of 24SRT members in this model airplane, it is recommended that the visual inspection procedure outlined for Boeing Model 314 and A-314 under AD 45-04-01 be followed. If defects are located, they shall be reported to the FAA for evaluation. Past experience has shown that once cracking starts, it may progress at a rapid rate, thus requiring closer inspections and corrective action. It shall also be the operator's responsibility to keep a record of all cracks on this model airplane. This record shall be revised periodically to showthe status of existing cracks and to record newly developed cracks. Copies of the original report and all revised pages should be submitted to the FAA for examination. \n\t\t1.\tINSPECTIONS OF READILY ACCESSIBLE AREAS. These inspections shall be conducted at intervals not to exceed 150 hours of operation or 90 days, whichever occurs first. This inspection is intended to cover only those portions of 24SRT tubing that are accessible to visual inspection through available inspection panels, removal of gap strips and the openings in the nacelles. \n\t\t2.\tDETAILED INSPECTIONS. These inspections will be conducted annually or at engine overhaul periods, whichever occurs first. This inspection is required of all 24SRT tubing visible through all available inspection panels, removal of gap strips, leading edges, wing tips, stress plates and fuel tanks. The use of at least a 10-power glass will be required. To more thoroughly cover the wing area, it will be necessary for a man to crawl outboardin the wings as far as possible. \n\t\t3.\tX-RAY INSPECTION. This type of inspection is required annually. Inspect by x- ray all inaccessible portions of the 24SRT spar chord members for their entire length. This inspection may coincide with annual inspection noted under 2. \n\tB.\tSA-307B-1. At intervals not to exceed 850 hours of operation or 120 days, whichever occurs first. If defects are located, they shall be repaired in a manner satisfactory to the FAA. \n\tC.\tS-307. At intervals not to exceed 700 hours of operation or 120 days, whichever occurs first. If defects are located, they shall be repaired in a manner satisfactory to the FAA.
48-17-01: 48-17-01 DOUGLAS: Applies to All DC3 Series Aircraft As Specified by Civil Air Regulations Amendment 41-3, 41-18, 42-2, 42-8, 61-2, and 61-16. \n\n\tTo be accomplished not later than the dates specified in the above amendments and any subsequent regulations effecting these compliance dates. \n\n\tAll air carrier aircraft must be modified to comply with the fire prevention requirements as outlined in CAR Amendments 41-3, 41-18, 42-2, 42-8, 61-2, and 61-16. The modification outlined in the following listed Douglas Service Bulletins are required for compliance with these amendments. Other modifications shown to be equivalent to those covered by the Service Bulletins will also be acceptable. \n\n\tDC3 No. 250, "Installation of Fire Detector in Engine Accessory Section and Smoke Detector in Aft Cargo Compartment"; DC3 No. 252, "Rear Baggage Compartment Access Door and Vent"; DC3 No. 258, "Elimination of Holes in Firewall, Addition of Control Cable Seals, Replacement of Dural Plates and Fittings With Steel Plates and Fittings. Replacement of Fluid Carrying Lines Forward of Firewall With Steel or Fire Resistant Flexible Hoses'; DC3 No. 259, "Installation of Shut-Off Valves on Lines Carrying Combustible Fluids Into the Engine Accessory Section". (Installation of additional fuel valves listed on Page 2 of this Bulletin is recommended but is not mandatory.) \n\n\tNOTE: It will be noted that Service Bulletins DC3 No. 258 and No. 259 apply to all DC3C and DC3D (C-47 and C-117) Series airplanes only with P&W R-1803 engines. Since there are various differences in early DC3 powerplant installations with P&W S1C3-G engines and Wright GR-1820 engines, it will be the operator's responsibility to use these two Bulletins as a guide and develop the fire prevention items for other DC3 Series airplanes accordingly. \n\n\tIn addition to the above, it will be necessary to ascertain that all interior materials and finishes comply with applicable sections of CAR 4b. Safety Regulation Release 259 outlines acceptable procedures for complying with these particular requirements.
64-08-01: 64-08-01 DOUGLAS: Amdt. 710 Part 507 Federal Register April 2, 1964. Applies to All Models DC-8 and DC-8F Series Aircraft.\n\n\tCompliance required as indicated.\n\n\tIt has been found that excessive torquing of pilot and static line fittings of the pitch trim compensator and the use of a nonreinforced hose can result in the twisting or bending of the pitot and/or static lines and the possible loss of required instruments. To correct this condition, accomplish either (a) or (b) as follows:\n\n\t(a)\tWithin 300 hours time in service after the effective date of this AD, modify the pitch trim compensator system in accordance with the provisions of paragraph (c).\n\n\t(b)\tAccomplish the provisions specified in (1), (2) and (3) and in addition, within 1,500 hours' time in service after the effective date of this AD modify the pitch trim compensator in accordance with the provisions of paragraph (c).\n\n\t\t(1)\tWithin 300 hours time in service after the effective date of this AD, unless already accomplished, visually inspect all aircraft for any evidence of bending or twisting of the pitot and static lines associated with the pitch trim compensator in accordance with the Douglas "Pitot Static System Inspection-DC-8" Service Letter CL-78-1966/DEG8-34-15-0 dated November 15, 1963.\n\n\t\t(2)\tReplace any damaged lines before further flight.\n\n\t\t(3)\tEach time the pitot and static line fittings associated with the pitch trim compensator are removed, loosened or retightened, visually inspect the pitot-static lines associated with the pitch trim compensator for any evidence of bending or twisting in accordance with the Douglas "Pitot Static System Inspection - DC-8" Service Letter CL-78-1966/DEG 8-34-15-0 dated November 15, 1963.\n\n\t(c)\tModify the pitch trim compensator system by replacing the pitot and static lines with reinforced lines and securing the pitot and static line fittings in accordance with paragraph 2, Accomplishment Instructions of the Douglas DC-8 Service BulletinNo. 34-51 dated January 9, 1964, or an FAA Western Region, Engineering and Manufacturing Branch approved equivalent.\n\n\t(Douglas DC-8 Service Bulletin No. 34-51 dated January 9, 1964, and Douglas DC-8 Service Letter CL-78-1966/DEG 8-34-15-0 dated November 15, 1963, pertain to this AD.)\n\n\tThis directive effective April 2, 1964.
2002-14-12: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model MD-11 and -11F airplanes, that requires an inspection of the wiring in the fuel control panel of the wings for chafing damage and for proper routing of the wiring; and corrective action(s), if necessary. This action is necessary to prevent chafing of the wiring in a cutout area in the wing fuel control panel due to improperly routed wiring, which could result in electrical arcing in an abnormal fuel vapor zone and consequent possible ignition of the fuel vapor. This action is intended to address the identified unsafe condition.
2008-22-16: The FAA is adopting a new airworthiness directive (AD) for certain GE CT58 series turboshaft engines. This AD requires recalculating the lives of certain part numbered compressor spools using a new repetitive heavy lift (RHL) multiplying factor. This AD results from reports of cracks originating from the inner faces of the locking screw holes in the compressor spool. We are issuing this AD to prevent cracks due to RHL missions. Cracks could result in an uncontained rotor burst and damage to, or loss of, the helicopter and serious injuries to any person onboard.
2008-22-11: The FAA is adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation (HBC) Model 390 airplanes. This AD requires you to inspect hydraulic pump pressure output hose assemblies to determine if they are from the affected lots, inspect for hydraulic fluid leaks if the hose assemblies are from the affected lots, and replace all affected hose assemblies. This AD results from reports of hydraulic leaks from the hydraulic pump pressure output hose assemblies. We are issuing this AD to prevent leakage of hydraulic fluid from the pump output hose within the engine compartment, which could result in an in-flight fire.
95-10-09: This amendment adopts a new airworthiness directive (AD), applicable to Sikorsky Aircraft Model S-58 and S-58T series helicopters, that requires the removal and replacement of the transmission main gear box ring gear (ring gear) within certain time intervals, and establishes a retirement life for the ring gear. This amendment is prompted by reports of failures of the ring gear due to slow-growth fatigue cracks. The actions specified by this AD are intended to prevent failure of the ring gear, failure of the main transmission, and subsequent loss of control of the helicopter.
48-41-01: 48-41-01 CONVAIR: Applies to All Model 240 Aircraft. Compliance required as indicated. I. Inspect wing bulkhead flanges and stringers at their intersections in the fuel tank area for cracks and repair as necessary at each No. 2 inspection (or equivalent periodic inspection approximating 100 hours) until permanent repairs and rework are accomplished. II. Complete rework in accordance with CVAC Service Bulletin No. 240-166A dated September 27, 1948, or equivalent should be accomplished not later than the next engine change.
47-21-14: 47-21-14 REPUBLIC: (Was Mandatory Note 4 of AD-769-2.) Applies only to Model RC-3 Aircraft Serial Numbers 5 to 500, Inclusive. Compliance required at the next 25-hour inspection or by August 1, 1947, whichever occurs first. To prevent fouling of the lower elevator cable on the elevator balance weight in the tail boom, incorporate Republic SK-17-14052-2 in the lower elevator control cable system. This elevator control cable guide is installed on the cross channel, in the tail boom, with the existing bolts holding the two inboard rudder pulley brackets. (Republic Service Bulletin No. 14 dated March 31, 1947, covers the same subject.)
63-27-02: 63-27-02 GRUMMAN: Amdt. 655 Part 507 Federal Register December 19, 1963. Applies to All Model G-21 Series Aircraft. Compliance required as indicated. As a result of cracks found on the rod end fitting, P/N 12727-7 located at the aft end of the rudder control push rod assembly, P/N 12727-1, accomplish the following within 100 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 100 hours' time in service from the last inspection. (a) Remove the rudder control push rod assembly, P/N 12727-1. This assembly consists of a tube with a fitting, P/N 12727-7, attached at each end. The length of the assembly is 20 1/2 inches from fitting center to fitting center. The rod assembly, used in conjunction with the left rudder pedals, is located below the pilot's compartment floor, the forward end approximately 2 1/2 inches below, and the aft end approximately 12 inches below . Laterally, the rod assembly is located approximately9 1/2 inches to the right of the aircraft's centerline. (b) Clean both rod end fittings, removing all grease and dirt. (c) Inspect both rod end fittings for cracks using a dye penetrant in conjunction with at least a 10-power magnifying glass, or an FAA approved equivalent inspection. (d) If a crack is found, that part shall be replaced in accordance with Grumman Drawing No. 12727 with a new part having the same part number, or an FAA approved equivalent before further flight. (e) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Southern Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. NOTE: It should be ascertained that the proper rudder control stop bolt, Grumman P/N G19-4-11, or anFAA approved equivalent bolt (having a smooth head) is presently installed in the pedal stop block (P/N 12779) which is part of the rudder and elevator torque shaft support assembly (P/N 12722) at Hull Station 11. This directive effective January 20, 1964.
2008-22-01: We are adopting a new airworthiness directive (AD) for various transport category airplanes. This AD requires deactivation of PATS Aircraft, LLC, auxiliary fuel tanks. This AD results from fuel system reviews conducted by the manufacturer, which identified unsafe conditions for which the manufacturer has not provided corrective actions. We are issuing this AD to prevent the potential of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane.
91-24-06: 91-24-06 SAAB-SCANIA: Amendment 39-8092. Docket No. 91-NM-132-AD. Applicability: Model SF-340A series airplanes, Serial Numbers 004 through 159; and Model SAAB 340B series airplanes, Serial Numbers 160 through 200; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent the cowling doors from opening during flight, accomplish the following: (a) Within 90 days after the effective date of this AD, inspect and measure the Avibank latches on the nacelle forward cowling doors 7271110- 501/601, in accordance with SAAB Service Bulletin 340-71-035, Revision 1, dated December 18, 1990. (1) If the measurement is within the limits specified in Figure 2 of the service bulletin, no further action is required. (2) If the measurement is outside the limits specified in Figure 2 of the service bulletin, prior to further flight, install new latch triggers in accordance with paragraph 2.C. of the service bulletin.(b) An 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, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may 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. (d) The inspection and replacement requirements shall be done in accordance with SAAB Service Bulletin 340-71-035, Revision 1, dated December 18, 1990. 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 SAAB-Scania AB, Product Support, S-581.88, Linkoping, Sweden. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-8092, AD 91-24-06) becomes effective on January 7, 1992.
47-43-08: 47-43-08 CESSNA: (Was Service Note 6 of AD-768-5.) Applies to All 120 and 140 Aircraft Equipped With Beech R003 Propeller Having R003-201 Blades and Continental C-85 Series Engine. Compliance required prior to January 1, 1948, and thereafter upon completion of each 25 hours of operation. Remove the R003-201 propeller blades and visually inspect the propeller blade retainer ferrule for cracks at the fillet joining the cylindrical outer surface of the ferrule with the retaining face of the flange. Particular caution should be exercised not to injure or contaminate the thrust bearing which must be pressed away from the flange for the inspection. The propeller manufacturer's assembly and service instructions are to be followed during disassembly and reassembly of the propeller. If any indication of a crack is found, both blades should be replaced with the R003-225 blades. The 25-hour inspection may be discontinued if R003-225 blades are installed. The R003-225 blades are sufficiently similar to the R003-201 blades to be considered aerodynamically interchangeable in the same diameter without a flight test. (Beech Aircraft Co. propeller Service Letter No. 1 covers this same subject.)
63-20-03: 63-20-03 DOUGLAS: Amdt. 617 Part 507 Federal Register September 14, 1963. Applies to All DC-6, DC-6A, and DC-6B Aircraft, Except Serial Number 44430 (Fuselage No. 500) and Subsequent. \n\n\tCompliance required as indicated. \n\n\tThere have been several instances of cracks causing failure of the lower center spar caps at Station 121, as well as cracking of the wing skin in the same area. Accordingly, the following shall be accomplished: \n\n\t(a) Within 100 hours' time in service after the effective date of this AD, unless already accomplished within the last 400 hours' time in service, visually or X-ray inspect for cracks in the lower center spar cap and the surrounding wing skin area from the inboard side of Numbers 2 and 3 engine nacelles inboard to Station 114.500 and from the outboard side of Number 2 and 3 engine nacelles outboard to Station 184.000. (For X-ray inspection see X-ray procedures and information as described in Figure 2 in Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962.) Pay particular attention to the area around the end attachments through the splice fittings. Reinspect at intervals not to exceed 500 hours' time in service from the last inspection. \n\n\t(b) If cracks are found in the lower spar cap, replace the part or repair it in accordance with a method approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region, before further flight, except for a ferry flight in accordance with CAR 1.76. When the new spar cap is installed, the original splice may be reworked or replaced with a redesigned splice. The rework instructions and redesign data are described in Accomplishment Instructions, Parts II and III, respectively, of Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962. Also, the splice may be reworked or replaced with a new splice in accordance with a method approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region. If cracks arefound in the surrounding skin area as set forth in (a), the skins shall be replaced or reworked in accordance with the manufacturer's instructions as authorized in Part I, paragraph (2), of Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962, or a method approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region. \n\n\t(c) The repetitive inspection specified in (a) may be temporarily discontinued for a period not to exceed 4,000 hours' time in service on those aircraft on which the temporary rework, described in Accomplishment Instructions, Part II of Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962, is accomplished. The 4,000 hour temporary discontinuance period will be computed as starting at the time of the temporary rework accomplishment. If the preventive rework as described in (d) is not accomplished prior to the end of the 4,000 hours' time in service period, the repetitive inspection of (a) must bereinstituted and the first reinspection accomplished prior to the expiration of the 4,000 hour period. \n\n\t(d) The special inspections described in (a), (b), and (c) may be discontinued when a specific area as described in (a) has been reworked with the preventive rework as outlined in Accomplishment Instructions, Part III of Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962. \n\n\t(e) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(Douglas Alert Service Bulletin No. A-849, Reissue No. 1, dated October 1, 1962, covers the same subject.) \n\n\tThis directive effective October 15, 1963. \n\n\tRevised November 21, 1963.
2008-21-07: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Three in-service propellers have been found to have blades which have lost the bonded metallic leading edge guard. If the leading edge guard comes off as the propeller turns, it could cause secondary damage to aircraft or injury to personnel. For the reasons described above, EASA issued Emergency AD 2007-0223-E to require repetitive inspections of the blade Leading Edge (L/E) guards for correct bonding until they accumulate more than 1,200 flight hours (FH) time in service. This AD requires actions that are intended to address the unsafe condition described in the MCAI, which could result in the loss of the bonded metallic leading edge guard, and could result in damage to the airplane or injury to personnel.