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2008-23-13: We are adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation (HBC) Model 390 airplanes. This AD requires you to remove the current preformed packing, elbow fitting, and jam nut from the left and right hydraulic pump pressure output port and replace with new parts. This AD also requires you to install a hydraulic pump case drain check valve. This AD results from nine occurrences of hydraulic fluid leaking from the engine hydraulic pump output fitting as a result of an improperly installed elbow connecting the output port to the pulse dampener hose. We are issuing this AD to prevent hydraulic fluid leaks from the left and right hydraulic fluid pump and to prevent the flow of hydraulic fluid into the engine compartment. The loss of hydraulic fluid can result in loss of airplane hydraulic system pressure and the consequent loss of hydraulic system functions including gear extension/retraction, spoiler functions, and anti-skid braking system actuation. The inability of the hydraulic installation to isolate flow of hydraulic fluid could result in a hazardous amount of flammable fluid in the corresponding engine compartment. These conditions, if not corrected, could result in loss of system functions and/or fire in the engine compartment.
94-08-11: This amendment adopts a new airworthiness directive (AD), applicable to certain Jetstream Model ATP airplanes, that requires replacement of certain circuit breakers on the left- and right-hand AC generator panel assemblies with new circuit breakers. This amendment is prompted by reports of failures due to localized overheating of the electric power circuit for the air conditioning recirculation fan in the environmental control system (ECS). The actions specified by this AD are intended to prevent the recirculation fan circuit from overheating, which could lead to smoke and/or flame in the fuselage.
65-09-02: 65-09-02 DOUGLAS: Amdt. 39-56 Part 39 Federal Register April 21, 1965. Applies to Model DC-6 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tDue to the failure of a main gear torque link, P/N 8065-46, the wheels and axle of the main gear rotated 180 degrees. To correct this condition, accomplish the following on torque links, Douglas P/N 8065-46, which have not been reworked in accordance with Douglas Service Engineering Letters C1-78-133/DJW dated January 28, 1964, C1-78-977/DJW dated June 22, 1964, or C1-78-1311/DJW dated August 11, 1964, and attached service sketches: \n\n\t(a)\tWithin 600 hours' time in service after the effective date of this AD for torque links which have been installed on aircraft for 20,000 or more hours' time in service on the effective date of this AD, and prior to the accumulation of 20,600 hours' time in service for torque links which have been installed on aircraft for less than 20,000 hours' time in service on the effective date of this AD - \n\n\t\t(1)\tInspect for cracks in the area of the vertical webs adjacent to the trunnion holes, using a 10-power glass or an equivalent method approved by the Aircraft Engineering Division, FAA Western Region. \n\n\t\t(2)\tTorque links with cracks greater than 0.170 inch in depth shall be replaced before further flight with torque links reworked in accordance with Douglas Service Engineering Letters C1-78-133/DJW dated January 28, 1964, C1-78-977/DJW dated June 22, 1964, or C1-78-1311/DJW dated August 11, 1964. \n\n\t\t(3)\tTorque links with cracks 0.170 inch in depth or less shall be: \n\n\t\t(i)\tReworked before further flight in accordance with Douglas Service Engineering Letter C1-78-1311/DJW dated August 11, 1964, and Douglas Service Sketch No. 608-A attached thereto, or an equivalent method approved by the Aircraft Engineering Division, FAA Western Region; or \n\n\t\t(ii)\tReplaced before further flight with torque links reworked in accordance with Douglas Service Engineering Letters C1-78-133/DJW dated January 28, 1964, C1-78-977/DJW dated June 22, 1964, or C1-78-1311/DJW dated August 11, 1964. \n\n\t\t(4)\tIf no cracks are found, repeat the inspection described in subparagraph (1) at intervals not to exceed 600 hours' time in Service from the last inspection. \n\n\t(b)\tThe repetitive inspections required by (a)(4) may be discontinued on torque links reworked in accordance with (a)(3). \n\n\t(c)\tOperators who have not kept records of hours' time in service on individual torque links shall substitute airplane hours' time in service in lieu thereof. \n\n\t(d)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, 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 Service Engineering Letter C1-78-1311/DJW dated August 11, 1964, and Service Sketch No. 608-A attached thereto, cover this same subject.) \n\n\tThis directive effective May 21, 1965.
2008-23-04: 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 affecting only RB211 Trent 500 series turbofan engines that have not incorporated Rolls-Royce plc (RR) Service Bulletin (SB) No. RB.211-72-D733, dated August 21, 2002, or Revision 1 of that SB, dated March 6, 2008, as follows: The intermediate-pressure (IP) turbine blade shrouds of the RB211 Trent 500 series engines feature closure welds (dust caps). Development engine testing has revealed the potential for dust caps to crack, lift and release. The latter may potentially allow hot annulus gas to be ingested down the core passages of IP turbine blades. Radial inflow of annulus gas into the IP disc rim region could cause local heating of the disc firtree, resulting in creep of the disc material. Failure of the disc rim in creep could simultaneously release two blades and a disc post. Failure to this extent could be beyond the containment capabilities of the casing. Consequently, release of the dust caps would constitute a potentially unsafe condition. This AD requires actions that are intended to address the unsafe condition described in the MCAI, which could result in uncontained release of IP turbine blades and disc posts, resulting in damage to the airplane.
66-01-02: 66-01-02 GRUMMAN: Amdt. 39-174 Part 39 Federal Register December 28, 1965. Applies to Model G-164 Series Airplanes Equipped with J-5404/MA96K, SR-5404R/MA96K or SR-5404/MA96K Propellers on Continental W-670 (R-670) Series, Gulf Coast W-670-240, and Jacobs R-755 and L-4M Series Engines. Compliance required within 25 hours' time in service after the effective date of this AD, unless already accomplished. To prevent further instances of loose propeller hubs, accomplish the following: (a) Remove propeller and check for matched sets of cone and spacer. The AN 5008- 20 rear cone (0.969 inch) must be used with the AN 5010-20 spacer (1.00 inch) and a No. 525-B rear cone (0.875 inch) must be mated with a No. 30091 spacer (1.105 inch). The overall average length of these sets must be 1.969 inch and l.980 inch respectively. An exception to these requirements is the Gulf Coast W-670-240 engine modified after 3/10/64 wherein the spacer has been omitted from the propeller installationdue to the 1-inch shorter propeller shaft. (b) Determine that the propeller nut does not bottom on the crankshaft threads by comparing the number of total turns with propeller removed to the number of turns with propeller installed and properly torqued. To correctly seat the propeller hub on the crankshaft there should be at least one full turn less with the propeller installed. If a proper fit cannot be obtained, new matched sets must be installed. If this procedure indicates the need for a spacer on the Gulf Coast W-670-240 installation, install Grumman spacer P/N A1600-99 (1/8 inch long). (Grumman Service Bulletin No. 32 dated September 9, 1965, and Service Bulletin No. 32A dated October 21, 1965, cover this same subject.) This directive effective December 28, 1965.
65-05-02: 65-05-02 DOUGLAS: Amdt. 39-39 Part 39 Federal Register February 25, 1965. Applies to Model DC-8 Series Aircraft Equipped with the Standard Wing Leading Edge Configuration, Except Aircraft with Serial Numbers 45687 to 45690 inclusive. \n\n\tCompliance required as indicated. \n\n\tAs a result of failures of the upper outboard spar cap structure of the inboard pylons, accomplish the following: \n\n\t(a)\tWithin the next 350 hours' time in service unless already accomplished, within the last 650 hours' time in service, inspect in accordance with (b) all inboard pylons having 8,000 or more hours' time in service as of the effective date of this AD. Prior to the accumulation of 8,350 hours' time in service unless already accomplished within the last 1,000 hours' time in service prior to the effective date of this AD, inspect in accordance with (b) all inboard pylons having less than 8,000 hours' time in service as of the effective date of this AD. \n\n\t(b)\tGain access to the area to be inspected by removing outboard access door numbers 213 and 314. Inspect the outboard spar cap, P/N 5640144-1, of No. 2 inboard pylon and the outboard spar cap, P/N 5640144-2, of No. 3 pylon for evidence of cracks in the area of Stations Yip = 225.000 to Yip = 233.000. Use close visual or dye penetrant inspection methods. \n\n\t(c)\tIf cracks are found, rework in accordance with paragraph 2 of Douglas Service Bulletin No. 54-34, Reissue No. 1, dated July 6, 1964, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight. \n\n\t(d)\tIf no cracks are found, reinspect parts as outlined in (b) at periods not to exceed 1,000 hours' time in service from the last inspection until reworked in accordance with paragraph 2 of Douglas Service Bulletin No. 54-34, Reissue No. 1, dated July 6, 1964, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. The repetitive inspection may be discontinued onall parts reworked in accordance with (c) or (d). \n\n\t(e)\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, 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 DC-8 Service Bulletin No. 54-34, Reissue No. 1, dated July 6, 1964, covers this same subject.) \n\n\tThis directive effective March 26, 1965.
2008-22-06: We are adopting a new airworthiness directive (AD) for certain Boeing Model 767-200 and -300 series airplanes. This AD requires replacing the wire segments of the four Fuel Quantity Indicating System (FQIS) wire bundles with new, improved wire segments. This AD results from operator inspections of the FQIS wire bundles that revealed corrosion at the connections between the ground wire and shield of each of the four FQIS wire bundles. We are issuing this AD to prevent this corrosion, which could reduce system protection of the lightning shield and result in loss of the electrical grounding between the lightning shield and the airplane structure. This condition, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane.
48-50-04: 48-50-04 CONVAIR: Applies to All Model 240 Aircraft. Compliance required by May 1, 1949. 1. Install steel nose gear upper centering cam, Bendix P/N 157627 in lieu of bronze cam. 2. Install main nose landing gear shock strut bearing and packing nut lock pins, in six places, Bendix P/N 54201 in lieu of lock rings. 3. Rework nose gear steering mechanism. 4. Install nose gear centering guides in nose wheel well. (Consolidated-Vultee Aircraft Corp. Service Bulletin Nos. 240-104A, -161, -162A, -167, and -201; CVAC Service Information Letter No. 310, and Bendix Service Bulletin No. L. G. 504, also cover these same subjects.)
47-20-06: 47-20-06 ERCO: (Was Mandatory Note 14 of AD-718-6.) Applies Only to 415-C, -CD and -D Aircraft Serial Numbers 113-3784, Inclusive, Except the Following Which Have New Design Incorporated: 3719, 3720, 3723, 3724, 3726, 3729, 3732, 3735, 3738, 3741, 3742, 3744, 3745, 3747, 3750, 3753, 3756, 3759, 3762, 3764, 3765, 3767, 3768, 3771, 3774, 3777, 3780, 3783. Compliance required not later than next 100-hour inspection unless visual inspection indicated immediate repair is required. Flexing of the lower aileron skin has resulted in fatigue cracks in the beam in the balance weight area. Inspect the beam and lower aileron skin carefully for cracks and drill relief holes at the ends of all cracks. Then add reinforcement plates (Erco P/N 415-16039-5 and -6) to the front face of the aileron beam and lower surface of the lower aileron skin, respectively, following the procedure outlined in Ercoupe Service Department Bulletin No. 20. (Blind, Type A, AN 450-4-10 may be used in lieu of DupontExplosive DR134A-8 and DR134A-10 rivets, respectively.) Use new longer AN 526C632-7 truss head screws to reinstall the balance weight. Check the aileron rigging and the aileron bellcrank pushrod for freedom from binding in the rod end under full aileron travel before again placing the airplane in operation. (Ercoupe Service Department Bulletin No. 20 dated February 17, 1947, covers this subject in greater detail.)
2008-22-08: We are adopting a new airworthiness directive (AD) for certain Boeing Model 757-200 and 757-300 series airplanes. This AD requires installing a bonding jumper between a ground and the clamp on the tube of the forward and aft gray water composite drain masts. For certain airplanes, this AD requires inspecting existing aft bonding jumper assemblies that might be too short, repair if necessary, and replacing the bonding jumper assembly with a new, longer bonding jumper assembly if necessary. This AD results from a report of charred insulation blankets and burned wires around the forward gray water composite drain mast found during an inspection of the forward cargo compartment on a Model 767-300F airplane. We are issuing this AD to prevent a fire near a composite drain mast and possible disruption of the electrical power system due to a lightning strike on a composite drain mast, which could result in the loss of several functions essential for safe flight.
47-43-02: 47-43-02 CESSNA: (Was Mandatory Note 13 of AD-768-5.) Applies to 120 and 140 Aircraft Serial Numbers 8001 to 8480, Inclusive. Compliance required prior to January 1, 1948. To eliminate the possibility of confusion in the operation of the fuel selector valve, remove the embossed pointer from the selector valve handle and ascertain that the selector valve handle and ascertain that the selector valve handle is installed so that the handle indicates correctly the position of the selector valve as shown by the valve placard.
72-25-02: 72-25-02 ROLLS ROYCE (1971) LIMITED: Amendment 39-1567 as amended by Amendment 39-1624. Applies to Rolls Royce Continental engines that have oil pump drive gear P/N's 22354/RR or 23403/RR installed. Those engines may include Model RR C90, except S/N's 11R021 and subsequent; Model RR O-200 engines, except S/N's 23R590 through 23R600 and 23R638 and subsequent and Model RR O-300 engines, except S/N's 31R162 and subsequent. Compliance required as indicated. To prevent the loss of oil pressure due to excessive wear of oil pump drive gear P/N's 22354/RR or 23403/RR accomplish the following: (a) Within the next 25 hours' time in service after the effective date of this AD, unless already accomplished within the last 25 hours' time in service before the effective date of this AD, and thereafter at the intervals specified in paragraph (b), determine the amount of back lash (angular movement) in the oil pump drive gear in accordance with Rolls Royce Service Bulletin No. T-200, dated November 26, 1971, or an FAA-approved equivalent. (b) If the amount of back lash (angular movement) determined in accordance with paragraph (a) during an inspection required by paragraph (a) or this paragraph is - (1) Six degrees or less, continue to inspect in accordance with paragraph (a) at intervals not to exceed 300 hours' time in service since the last inspection; (2) More than 6 degrees but not more than 10 degrees, continue to inspect in accordance with paragraph (a) at intervals not to exceed 100 hours' time in service since the last inspection; (3) More than 10 degrees but not more than 14 degrees, continue to inspect in accordance with paragraph (a) at intervals not to exceed 25 hours' time in service since the last inspection; (4) More than 14 degrees, before further flight, replace oil pump drive gear P/N's 22354/RR or 23403/RR, as applicable, with - (i) An improved oil pump drive gear as specified in Rolls Royce Service Bulletin No. T-200, dated November 26, 1971, or an FAA-approved equivalent; or (ii) An oil pump drive gear of the same part number with 14 degrees or less of back lash (angular movement), determined in accordance with paragraph (a), and continue to comply with paragraph (b), as applicable. (c) The repetitive inspections required by subparagraphs (b)(1), (b)(2), (b)(3), and (b)(4)(ii) may be discontinued when oil pump drive gear P/N's 22354/RR or 23403/RR are replaced with improved oil pump drive gear as specified in Rolls Royce Service Bulletin No. T-200, dated November 26, 1971, or an FAA-approved equivalent. NOTE: Copies of Rolls Royce Service Bulletin No. T-200, dated November 26, 1971, may be obtained by request to the following address: Teledyne Continental Motors Attention: Circulation Department P.O. Box 90 Mobile, Alabama 36601 Amendment 39-1567 became effective December 4, 1972. This amendment 39-1624 becomes effective April 23, 1973.
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.