Results
63-07-03: 63-07-03 PIPER: Amdt. 551 Part 507 Federal Register April 4, 1963. Applies to All Models PA-23 and PA-23-160 Aircraft. Compliance required as indicated. Cracks and loose or working rivets have been found in the attachment area of the rudder trim tab horn, P/N 17059-00, to the rudder trim tab skin, P/N 17068-27. Cracked rudder trim tab ribs P/N 17068-17 also have been found. To preclude the separation of the trim tab horn from the trim tab, accomplish the following: (a) On aircraft having 690 or more hours time in service on the effective date of this AD, within the next 10 hours' time in service unless already accomplished within the last 90 hours' time in service and within each 100 hours' time in service thereafter from the last inspection, accomplish the inspection specified in (c). (b) On aircraft having less than 690 hours' time in service on the effective date of this AD, accomplish the inspection specified in (c) between 600 and 700 hours' time in service and within each 100 hours' time in service thereafter from the last inspection. (c) Visually inspect with at least a 10-power magnifying glass the rudder trim tab horn attach rivets for looseness or evidence of rivet working and the adjacent tab skin for cracks. (d) If skin cracks, loose or working rivets are found, visually inspect the rudder trim tab support rib, P/N 17068-17 for cracks. (e) If any of the defects mentioned herein are found, repair in accordance with Civil Air Regulations Part 18 or replace the affected part, or parts, with an airworthy assembly prior to further flight. (Piper's Periodic Inspection Report pertains to this same subject.) This directive effective April 15, 1963.
2005-11-05: The FAA adopts a new airworthiness directive (AD) for all airplanes equipped with Precise Flight, Inc. (Precise Flight) Models SVS I and SVS IA standby vacuum systems (SVS) installed under certain supplemental type certificates or through field approval. This AD requires you to replace the airplane flight manual supplement (AFMS) in the airplane flight manual with the appropriate revision and install placards as defined in the AFMS, upgrade the Model SVS I or SVS IA SVS to the Model VI SVS, and add the instructions for continued airworthiness (ICA) to the maintenance schedule for the aircraft. This AD results from several reports of failed shuttle control valves of the standby vacuum system (SVS) and one report of an airplane crash with a fatality in which improper use of the SVS was a factor. We are issuing this AD to correct problems with the SVS before failure or malfunction during instrument flight rules (IFR) flight that can lead to pilot disorientation and loss of control of the aircraft.
2005-11-11: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier Model DHC-8-400, -401, and -402 series airplanes. This AD requires repetitive inspections to detect discrepancies of the attachment fittings of the outboard flap front spar at flap track Number 4 and Number 5 locations, and corrective actions if necessary. This AD also requires eventual replacement of the attachment fittings as terminating action for the repetitive inspections. This AD is prompted by the discovery of several airplanes that have loose flap front spar attachment fittings at flap track Number 4 and Number 5 locations. We are issuing this AD to prevent the attachment fittings from becoming detached, and consequent loss of control of the airplane.
71-17-08: 71-17-08 CESSNA: Amdt. 39-1273. Applies to Model 337 Airplanes. Compliance. To minimize the possible adverse affects of undetected rear engine stoppage during take-off, within the next 10 hours' time in service after the effective date of this AD, the following modifications must be accomplished and the following operational procedures will be applicable: A) Readjust the rear engine idle RPM from 575-625 to 625-675 and readjust idle mixture accordingly. B) Taxi primarily by use of the rear engine. C) Initiate all take-offs by advancing the throttle on the rear engine to a point where the normal functioning of the rear engine has been established before advancing the throttle of the front engine. D) Install a permanent placard to the right of the tachometer instrument to read as follows: TAXI AND TAKE-OFF LEAD WITH REAR ENGINE POWER CHECK RPM AND FUEL FLOW NOTE: The operator may make and install a temporary placard using minimum 1/8 inch highletters until the permanent placard is obtained from the manufacturer and properly installed. Cessna Service Letter ME 71-21 refers to the subject matter of this AD. This amendment becomes effective August 24, 1971.
62-26-06: 62-26-06 PIPER: Amdt. 516 Part 507 Federal Register December 8, 1962. Applies to All Models PA-28-150 and PA-28-160 Airplanes With Serial Numbers 28-03, 28-1 through 28-671, 28-678, 28-679, 28-681, 28-685, 28-690, 28-691, 28-693, 28-697, 28-706, and 28-707, Except Those With Exhaust Assembly P/N 63726 Installed. Compliance required within the next 10 hours' time in service after the effective date of this AD, and each 50 hours' time in service thereafter. There have been cracks found in the exhaust system on Piper PA-28-150 and PA-28-160 airplanes. As this condition is likely to occur in other airplanes of the same type design, accomplish the following: (a) Remove carburetor heat muff shroud and conduct a close visual inspection of the complete exhaust system piping for cracks in all welded joints and tubing bends. Pay particular attention to the area near the junction of the cylinder stacks and main exhaust manifold. (b) If cracks are found, repair by gaswelding. (The exhaust pipe is AISI 321 or 347 corrosion resistant steel.) (Piper Service Letter 389 covers this same subject.) This directive effective December 14, 1962.
62-23-03: 62-23-03 DOUGLAS: Amdt. 500 Part 507 Federal Register October 30, 1962. Applies to DC-8 Aircraft Serial Numbers 45252 to 45289 Inclusive, 45291 to 45306 Inclusive, 45376 to 45393 Inclusive, 45408 to 45413 Inclusive, 45416 to 45431 Inclusive, 45433 to 45437 Inclusive, 45442 to 45445 Inclusive, 45526, 45565 to 45570 Inclusive, 45588 to 45606 Inclusive, 45609 to 45614 Inclusive, 45616 to 45629 Inclusive, 45636, and 45638. \n\n\tCompliance required as indicated. \n\n\tTo prevent failure of the main landing gear actuating cylinder upper attach bracket accomplish the following: \n\n\t(a)\tWithin 440 hours' time in service after the effective date of this AD, unless already accomplished, on attach brackets, P/N's 5641950-1, -2, -501 or -502 having 4,000 or more hours' time in service and, prior to the accumulation of 4,440 hours' time in service on brackets having less than 4,000 hours' time in service, conduct a hardness check on the attach bracket following the procedures described in theaccomplishment instructions, paragraphs B (4), B(5) and C of Douglas Alert Bulletin No. A32-76, Reissue No. 1, Revision No. 1, dated April 5, 1962, or later, or an FAA approved equivalent. \n\n\tNOTE: To gain access to the bracket it is necessary to remove the hydraulic lines from the rework area, remove swivel glands as applicable, remove the retract cylinder upper bolt and swing the cylinder down out of the way. \n\n\t(b)\tBrackets testing within Rockwell C39.0 (175,000 p.s.i.) and C43.0 (200,000 p.s.i.) range are acceptable for further use and may be continued in service provided they are not otherwise defective. These brackets shall be identified by applying a 1/4 to 1/2-inch stripe of Cat-A-Lac, yellow No. 443-3-129 enamel or equivalent on the bracket. The identification stripe should be located in a place on the bracket so that it is clearly visible without having to remove the retract cylinder to observe the stripe. \n\n\t(c)\tBrackets testing outside the C39.0 to C43.0 heat-treatrange shall be inspected for cracks in the area of the junction of the grease fitting hole and the actuating cylinder attach hole using a dye penetrant or equivalent inspection method. The bushing, Douglas P/N 2641952 must be removed to conduct this inspection. \n\n\t\t(1)\tAny of these brackets found to be free of cracks during inspection, and not otherwise defective, may be continued in service provided the inspection is repeated at intervals not exceeding 350 hours' time in service thereafter. \n\n\t\t(2)\tReplace cracked parts prior to further flight with an appropriate part (Douglas P/N 5641950-1 or -2, 5641950-501 or -502) falling within the C39.0 to C43.0 heat-treat limits and marked with the yellow stripe for identification, with new P/N's 5774066-501 or -502, or with an FAA approved equivalent. \n\n\t(d)\tUpon 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. A32-76, Reissue No. 1, Revision No. 1, dated April 5, 1962.) \n\n\tThis directive effective November 29, 1962.
74-11-05: 74-11-05 GRUMMAN AMERICAN AVIATION CORPORATION: Amendment 39- 1849. Applies to Grumman G-1159, Serial Number 1 through 110 and 775 except 91 and 95, certificated in all categories. Compliance required within the next 100 hours time in service after the effective date of this AD, unless already accomplished. To provide dual electrical fire extinguisher squib circuits, accomplish the following or an equivalent rework approved by the Chief, Engineering and Manufacturing Branch, Southern Region. Remove the left and right fire extinguisher switches MS24523-21 and install MS24524- 21 switches in their place. Wire switches in accordance with Grumman Gulfstream II Aircraft Service Change 121. Replace pilot's circuit breaker panel nameplate 1159AV2042 with nameplate 1159SB20082-11. This rework is outlined in detail in the Grumman Gulfstream II Aircraft Service Change 121. This amendment is effective May 24, 1974.
62-12-04: 62-12-04 DOUGLAS: Amdt. 439 Part 507 Federal Register May 17, 1962. Applies to All DC-8 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tA number of wing flap actuating cylinder rod-end bearings have cracked and in one case there was complete failure. To preclude failure of the rod-end bearings in cylinders located at wing Stations 98 and 219, accomplish the following: \n\n\t(a) Assemblies having 4,650 hours' or more time in service as of the effective date of this AD shall be inspected in accordance with (b) within the next 350 hours' time in service after the effective date of this AD and thereafter at intervals not exceeding 350 hours' time in service. Assemblies having less than 4,650 hours' time in service as of the effective date of this AD shall be inspected in accordance with (b) prior to the accumulation of 5,000 hours' time in service and thereafter at intervals not exceeding 350 hours' time in service. \n\n\t(b) Visually inspect rod-end bearings P/N 4648686 in the area of the bearing case adjacent to the bearing dust shield using at least a 10-power glass, for evidence of cracks. Caution: Do not remove rod-end bearing dust shield or magnetically inspect the rod-ends. \n\n\t(c) If cracks are found, replace the defective part. If no cracks are found, the 350 hour periodic inspection, specified in (b), shall be continued. \n\n\t(d) When replacement assemblies P/N 4648686 are used, the inspection prescribed in (b) shall be accomplished prior to the accumulation of 5,000 hours' time in service and thereafter at intervals not exceeding 350 hours' time in service. \n\n\t(e) When assembly P/N 4648686 is replaced by Douglas P/N 4648686-501 (Shafter P/N YD-200A) or -503 (Shafter P/N YD-200B), or FAA approved equivalent, the inspections called for in this AD may be discontinued. \n\n\t(f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering & 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 Service Bulletin No. 27-127 reissue No. 2 dated January 28, 1964, covers this same subject.) \n\n\tThis directive effective May 17, 1962. \n\n\tRevised August 12, 1964.
2005-03-14: This amendment supersedes an existing airworthiness directive (AD) that applies to all Airbus Model A300 B2 and B4 series airplanes. The existing AD currently requires determining the part and amendment number of the variable lever arm (VLA) of the rudder control system to verify the parts were installed using the correct standard, and corrective actions if necessary. For certain VLAs, this new AD requires repetitive inspections of the VLA and corrective action if necessary. This new AD also provides a terminating action for the repetitive inspections. Furthermore, this new AD reduces the applicability of affected airplanes. The actions specified by this AD are intended to prevent failure of both spring boxes of certain VLAs due to corrosion damage, which could result in loss of rudder control and consequent reduced controllability of the airplane. This action is intended to address the identified unsafe condition.
72-12-05: 72-12-05 HAWKER SIDDELEY AVIATION LTD.: Amdt. 39-1458. Applies to Hawker Siddeley deHavilland Model DH-114 "Heron" airplanes which do not have Modification 1161 incorporated on all of the main landing gear and nose landing gear operating jacks. Compliance required as indicated. To prevent the possible failure of the main or nose landing gear operating jacks accomplish the following: (a) Within the next 50 landings after the effective date of this AD, or before accumulating a total of 5,000 landings on a piston, P/N AHO 19742, in an operating jack assembly, whichever occurs later (1) Replace the affected operating jack assembly with a serviceable jack assembly which has been modified to incorporate piston, P/N ACO 23389, (Modification 1161) in accordance with de Havilland Aircraft Service, Modification News Sheet No. Heron 1161, dated March 17, 1958, or an FAA-approved equivalent, or (2) Replace the piston in each affected operating jack assembly with a newpiston of the same part number and thereafter continue to replace P/N AHO 19742 pistons at intervals not to exceed 5,000 landings. (b) Operators who have not kept records of the number of landings on individual pistons, P/N AHO 19742, shall substitute the total number of airplane landings accumulated in lieu thereof. (de Havilland Aircraft Service, Technical News Sheet, Series: Heron (114) No. S.2 dated March 1, 1959, covers this same subject.) This amendment is effective upon publication in the Federal Register as to all persons except those persons to whom it was made immediately effective upon receipt of the airmail letter dated April 20, 1972, which contained this amendment.
2005-11-01: The FAA is adopting a new airworthiness directive (AD) for Turbomeca S.A. Arrius 1A turboshaft engines. This AD requires initial and repetitive testing of the Free Turbine Overspeed Protection System. This AD results from an investigation into the Digital Electronic Control Unit (DECU) that revealed a malfunction of the Free Turbine Overspeed Protection System. This malfunction can exist despite the DECU passing all functional tests specified in the Engine Maintenance Manual. We are issuing this AD to prevent uncontained engine failure if a free turbine overspeed occurs.
75-09-08: 75-09-08 MOONEY: Amendment 39-2181. Applies to Mooney Models M20C, M20D, M20E, M20F, and M20G airplanes with the following serial numbers having accumulated 400 or more hours' time in service: M20C - 1852, 1940 thru 3466, 670001 thru 700091, 20-0001 thru 20-1156. M20D - 101 thru 260. M20E - 101 thru 1308, 670001 thru 700061, 21-0001 thru 21-1165. M20F - 660003 thru 700072, 22-0001 thru 22-1207. M20G - 680001 thru 700006. Compliance required within the next 50-hours' time in service after the effective date of this AD or before the accumulation of 450-hours' time in service, whichever occurs later, unless already accomplished within the last 50-hours' time in service, and thereafter at intervals not to exceed 100-hours' time in service from the last inspection. a. To detect cracked engine mount members, inspect the mount in accordance with instructions in Mooney Service Bulletin No. M20-192 dated April 7, 1975, or later FAA approved revision, or an FAA approved equivalent method. b. If cracks are detected, repair or replace in accordance with instructions in Mooney Service Bulletin No. M20-192 dated April 7, 1975, or later FAA approved revision or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Southwest Region, Fort Worth, Texas. c. The 100-hour repetitive inspection may be discontinued upon repair of a cracked or uncracked mount, or upon installation of a new mount in accordance with Mooney Service Bulletin No. M20-192 dated April 7, 1975, or later FAA approved revision or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, Southwest Region, Federal Aviation Administration, Fort Worth, Texas. This amendment becomes effective April 30, 1975.
2000-06-13 R1: This amendment revises an existing airworthiness directive (AD), applicable to certain Boeing Model 737-200, -200C, -300, and -400 series airplanes, that currently requires repetitive visual and high frequency eddy current (HFEC) inspections to detect cracking of the corners of the door frame and the cross beams of the aft cargo door, and corrective actions, if necessary. That amendment also mandates accomplishment of a modification to the aft cargo door, which would terminate the repetitive inspection requirements. This amendment revises the compliance time for the terminating modification. The actions specified by this AD are intended to prevent fatigue cracking of the corners of the doorframe and the crossbeams of the aft cargo door, which could result in rapid depressurization of the airplane. \n\n\tThe incorporation by reference of Boeing Alert Service Bulletin 737-52A1079, Revision 6, dated November 18, 1999, as listed in the regulations, was approved previously by the Director of the Federal Register as of May 9, 2000 (65 FR 17583, April 4, 2000).\n\n\n\tThe incorporation by reference of Boeing Service Bulletin 737-52-1079, Revision 5, dated May 16, 1996, as listed in the regulations, was approved previously by the Director of the Federal Register as of December 24, 1998 (63 FR 67769, December 9, 1998).
73-26-05: 73-26-05 SCOTT AVIATION DIVISION of ATO INC: Amdt. 39-1756. Applies to chemical oxygen generators, Scott Part Numbers 801386-01, -02, and -03, and 801462-01, -02, and -03. Compliance required within the next six months after the effective date of this AD, unless already accomplished. To detect generators showing evidence of corrosion, and to replace corroded units, accomplish Scott Aviation Service Bulletin 35-38, dated February 1, 1973, and Revision Number 1, dated May 11, 1973, or an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective December 20, 1973.
2005-11-02: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 767-200, -300, and -300F series airplanes. That AD currently requires inspections for fatigue cracking of the horizontal stabilizer pivot bulkhead, and repetitive inspections or other follow-on actions. That action also provides a permanent repair, which is optional for airplanes with no cracks, and, if accomplished, ends the repetitive inspections. For airplanes on which the permanent repair is not installed, this new AD requires repetitive inspections of the same and additional inspection locations at new inspection intervals; a one-time torque test; and related investigative and corrective actions. For airplanes on which the permanent repair is installed, this new AD would require repetitive inspections of the repaired area and, if necessary, corrective action. This AD is prompted by reports of loose tension bolts and crack indications in the fuselage skin. We are issuingthis AD to find and fix fatigue cracking of the horizontal stabilizer pivot bulkhead and adjacent structure, which could result in loss of the horizontal stabilizer. \n\nDATES: This AD becomes effective July 1, 2005. \n\n\tThe incorporation by reference of Boeing Alert Service Bulletin 767-53A0078, Revision 3, dated November 15, 2001; and Boeing Alert Service Bulletin 767-53A0078, Revision 4, dated September 26, 2002, as listed in the AD is approved by the Director of the Federal Register as of July 1, 2005. \n\n\tOn May 24, 2001 (66 FR 23538, May 9, 2001), the Director of the Federal Register approved the incorporation by reference of Boeing Service Bulletin 767-53-0078, Revision 2, dated April 19, 2001.
62-02-05: 62-02-05 PIPER: Amdt. 386 Part 507 Federal Register January 18, 1962. Applies to Models PA-18 and PA-18A Series aircraft Serial Numbers 18-1 through 18-7768 inclusive. Compliance required within the next 25 hours' time in service after the effective date of this AD unless already accomplished. Several failures have been reported of the rudder cable attachment fitting lug, P/N 40831, at the weld on rudder pedals, P/N's 40842-04 and 40842-05. Accordingly, the following shall be accomplished: Visually inspect the weld that attaches the rudder cable lug, P/N 40831, to the foot bar tube on both left and right rudder pedals for evidence of separation, cracks in the weld, insufficient length of weld, or excessive wear of weld. A weld of sufficient length must cover at least the top half perimeter of the tube (minimum). If the weld is separated, cracked, excessively worn, or of insufficient length, replace the rudder pedal assembly prior to further flight. (Piper Service Bulletin No. 207A, dated May 10, 1962, covers this subject.) This directive effective January 23, 1962. Revised June 20, 1962.
61-24-02: 61-24-02 DOUGLAS: Amdt. 370 Part 507 Federal Register November 23, 1961. Applies to All DC-8 Aircraft, Serial Nos. 45252-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45419, 45421-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45606, 45609-45614, 45616-45628, 45636. \n\n\tCompliance required within 1,000 hours' time in service after November 23, 1961, but in no event later than February 28, 1962, except as provided in paragraph (e). \n\n\tAs a result of numerous recent incidents associated with hydraulic system failures or malfunctions, the aircraft shall be modified to incorporate the following: \n\n\t(a) A dual source of hydraulic power for the actuation of all wing spoilers during landing roll, in accordance with Douglas DC-8 Service Bulletin 29-39 or FAA approved equivalent. \n\n\t(b) Increased brake accumulator capacity, in accordance with Douglas DC-8 Service Bulletin 29-41 or FAA approved equivalent. \n\n\t(c) A dual source of hydraulic power for the rudderpower system, in accordance with Douglas DC-8 Service Bulletin 27-117 or FAA approved equivalent. \n\n\t(d) A source of power to actuate the nose wheel steering system when the airplane hydraulic system is being operated with the hydraulic system selector handle in the "main gear down lock and flaps" position, and additional hydraulic fluid reserve capacity and related changes in the fluid reservoir quantity indicating system, in accordance with Douglas DC-8 Service Bulletin 32-73 or FAA approved equivalent. \n\n\t(e) To reduce to acceptable levels the hydraulic system pressure fluctuations induced by the auxiliary hydraulic pump: \n\n\t\t(1) By no later than February 28, 1962, bypass the surge damper which may have been installed in accordance with Douglas DC-8 Service Bulletin 29-35. The bypass method shall be in accordance with steps 1 through 6 of Addendum Number 1 to Service Bulletin 29-35 dated February 16, 1962. \n\n\t\t(2) As soon as necessary parts are available but in no eventlater than April 30, 1962, install surge damper in accordance with Douglas Service Bulletin 29-35, Reissue Number 1 dated November 8, 1961, as amended by Addendum Number 1 of Service Bulletin 29-35 dated February 16, 1962. \n\n\t(f) Dampers to reduce to an acceptable level the surge pressure induced in the hydraulic system when the aileron and rudder power systems are activated, in accordance with Douglas DC-8 Service Bulletin 27-109, revision No. 1 to reissue No. 1 dated November 15, 1961, or later FAA approved version. \n\n\t(g) The FAA approval of the equivalent methods of compliance with the modifications required by paragraphs (a) through (d) shall be obtained through the Chief, Engineering and Manufacturing Branch, FAA Western Region. \n\n\tThis directive effective November 23, 1961. \n\n\tRevised March 14, 1962, for all persons except those to whom it was made effective immediately by telegram dated February 21, 1962.
75-18-08: 75-18-08 MCDONNELL DOUGLAS: Amendment 39-2351 as amended by Amendment 39-2411. Applies to DC-10-10, -10F, -30, -30F and -40 series airplanes, certificated in all categories. \n\n\tTo prevent possible airplane misalignment during ILS approaches without readily available failure indication, accomplish the following: \n\n\t(A)\tUnless already accomplished or unless one of the relieving provisions, specified below, has been accomplished, within four calendar days after receipt of this telegram the airplane is limited to Category I minima and the following placard or its equivalent must be placed in plain view of the pilots: AIRCRAFT LIMITED TO CATEGORY I APPROACH MINIMA. \n\n\t(B)\tOperators shall, by the most immediate and practicable means, notify flight crews of the foregoing. \n\n\t(C)\tThe Category I limitation does not apply to an airplane after accomplishing one of the following: \n\n\t\t(1)\tAccomplish modification defined in McDonnell Douglas Service Bulletin 34-66, dated July 23, 1975, or later FAA-approved revisions; or \n\n\t\t(2)\tAccomplish a modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\t(D)\tPrior to August 6, 1976, accomplish McDonnell Douglas Service Bulletin 34-66, dated July 23, 1975, or later FAA-approved revisions, or an equivalent modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\n\tAmendment 39-2351 was effective September 2, 1975, for all persons except those to whom it was made effective immediately by telegrams dated August 6, 1975. \n\n\tThis amendment 39-2411 is effective November 10, 1975 for all persons except those to whom it was made effective by telegrams, dated October 10, 1975, which contained this amendment.
2005-11-04: The FAA is adopting a new airworthiness directive (AD) for certain Bombardier Model CL-600-1A11 (CL-600), CL-600-2A12 (CL-601), and CL-600-2B16 (CL-601-3A, CL-601-3R, and CL-604) airplanes modified by STC SA4900SW. This AD requires revising the airplane flight manual (AFM) to require repetitive visual checks of the microphone jack assemblies on both control columns to detect damage that may interfere with movement of the control column. This AD also requires modification of the microphone jack assembly, related investigative actions, and corrective actions if necessary, which allows the AFM revision to be removed from the AFM. This AD is prompted by a report of a rejected take-off and subsequent runway overrun due to restricted movement of the co-pilot's control column, which resulted in collapse of the nose landing gear and consequent damage of the forward fuselage. We are issuing this AD to prevent a damaged microphone jack assembly from interfering with movement of the control column, which could result in loss of control of the airplane.
97-11-11: This amendment adopts a new airworthiness directive (AD) that applies to Aerospace Technologies of Australia Pty Ltd. (ASTA) Models N22B, N22S, and N24A airplanes. This action requires repetitively inspecting the horizontal stabilizer upper and lower skin, intercostal angles, and the horizontal stabilizer trailing edge channel for cracks; and repairing any crack or replacing any cracked parts, as applicable. This AD results from numerous reports of cracking in these horizontal stabilizer areas on the affected airplanes. The actions specified by this AD are intended to prevent structural failure of the horizontal stabilizer caused by fatigue cracks, which could result in loss of control of the airplane.
75-08-04: 75-08-04 LOCKHEED: Amendment 39-2153. Applies to L-1011-385-1 airplanes certificated in all categories. Compliance required as indicated. To prevent a possible fire hazard in the C-1 or C-1A, C-2 and C-3 cargo compartment: (a) Within the next 300 flight hours, perform the following, unless a modification acceptable per (b), below, has been accomplished: (1) Deactivate the cargo door compartment manual light switch in accordance with Lockheed Service Bulletin 093-33-052, dated March 5, 1975, or later FAA-approved revisions; or an equivalent deactivation modification procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region. (2) Check the functioning of the door switch which controls the cargo compartment light per Lockheed Service Bulletin 093-33-052, dated March 5, 1975, or later FAA- approved revisions. (b) Within 3000 hours time in service and each 3000 hours thereafter, recheck the door switch per item (a)(2), above. These tests may be discontinued after installation of a modification to the cargo door light, acceptable to the Chief, Aircraft Engineering Division, FAA Western Region. The approved data must reference this AD, and the approval date. (c) After incorporation of an acceptable modification, per (b) above, the manual switch may be reactivated. (d) An airplane may be flown to a base for the performance of the work required by this AD per FAR's 21.197 and 21.199. This amendment becomes effective April 9, 1975.
2005-10-23: The FAA adopts a new airworthiness directive (AD) for all DG Flugzeugbau GmbH Model DG-500MB sailplanes equipped with a Solo engine and Glaser-Dirks Flugzeugbau GmbH Model DG-800B sailplanes equipped with a Solo engine. This AD requires you to inspect the propeller for damage, specifically foam core separation, and replace any damaged propeller. This AD results from mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. We are issuing this AD to detect and correct damage to the propeller, which could result in failure of the propeller to perform properly. This failure could lead to reduced or loss of control of the sailplane.
97-21-06: 97-21-06 CASA: Amendment 39-10159. Docket 96-NM-137-AD. Applicability: CASA Model CN-235 airplanes; as listed in CASA Service Bulletin SB-235-27-05, Revision 1, dated September 29, 1993 (non-military airplanes), and CASA Service Bulletin SB-235-27-05M, Revision 2, dated January 25, 1996 (military airplanes); certificated in any category. NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been otherwise modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (e) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. Compliance: Required as indicated, unless accomplished previously. To prevent loss of control of the elevator and/or rudder, due to failure of the elevator and/or rudder assemblies as a result of stress corrosion cracking in the torsion tubes and fittings, accomplish the following: NOTE 2: Actions required by this AD that were accomplished previous to the effective date of this AD, and in accordance with earlier versions of the specified CASA service bulletins, are considered acceptable for compliance with the applicable requirements of this AD. (a) At the applicable time specified in either paragraph (a)(1) or (a)(2) of this AD, conduct a visual inspection of the torsion (torsion) tubes on the elevator and rudder assemblies to detect stress corrosion cracking, in accordance with CASA Service Bulletin SB-235-27-05, Revision 1, dated September 29, 1993 (for non-military airplanes) or CASA Service Bulletin SB-235-27-05M, Revision 2, dated January 25, 1996 (for military airplanes), as applicable. (1) For airplanes that have accumulated more than 600 total hours time-in-service, or more than 1,000 total landings, as of the effective date of this AD: Conduct the inspection required by paragraph (a) of this AD prior to the accumulation of 50 hours time-in-service, or 100 landings, or within 3 months, after the effective date of this AD, whichever occurs first. (2) For all other airplanes: Conduct the inspection required by paragraph (a) of this AD prior to the accumulation of 600 total hours time-in-service, or 1,000 total landings, or within 6 months, after the effective date of this AD, whichever occurs first. (b) If no cracking is detected during the inspection required by paragraph (a) of this AD, repeat that inspection at intervals not to exceed 600 hours time-in-service, or 1,000 landings, or 6 months, whichever occurs first.(c) If any cracking is detected during the inspection required by paragraph (a) of this AD, prior to further flight, accomplish either paragraph (c)(1) or (c)(2) of this AD. (1) Replace cracked parts with new parts of the original design, in accordance with the service bulletin. After replacement, repeat the visual inspection required by paragraph (a) of this AD at intervals not to exceed 600 hours time-in-service, or 1,000 landings, or 6 months, whichever occurs first. OR (2) Replace cracked parts with newly-designed parts, in accordance with CASA Service Bulletin SB-235-27-05, Revision 1, dated September 29, 1993 (for non-military airplanes); or CASA Service Bulletin SB-235-27-05M, Revision 2, dated January 25, 1996 (for military airplanes); as applicable. This replacement constitutes terminating action for the repetitive visual inspections of that part required by paragraph (b) of this AD. (d) Within 2 years after the effective date of this AD, replace all original design parts comprising the torsion tube assemblies on the elevator and rudder assemblies with newly-designed parts, in accordance with CASA Service Bulletin SB-235-27-05, Revision 1, dated September 29, 1993 (for non-military airplanes); or CASA Service Bulletin SB-235-27-05M, Revision 2, dated January 25, 1996 (for military airplanes); as applicable. This action constitutes terminating action for the inspection requirements of this AD. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE 3: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (f) Special flight permits may be issued in accordance with sections 21.197 and P21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. (g) The actions shall be done in accordance with CASA Service Bulletin SB-235-27-05, Revision 1, dated September 29, 1993; and CASA Service Bulletin SB-235-27-05M, Revision 2, dated January 25, 1996; as applicable; which contain the specified effective pages: Service Bulletin Referenced and Date Page Number Revision Level Shown on Page Date Shown on Page SB-235-27-05, Revision 1, September 29, 1993 1, 2 3-23 1 Original September 29, 1993 February 5, 1993 SB-235-27-05M, Revision 2, January 25, 1996 1 2-23 2 Original January 25, 1996 October 28, 1991 This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from Construcciones Aeronauticas, S.A., Getafe, Madrid, Spain. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. NOTE 4: The subject of this AD is addressed in Spanish airworthiness directive 06/94, dated August 1994. (h) This amendment becomes effective on November 24, 1997.
61-21-02: 61-21-02 DOUGLAS: Amdt. 350 Part 507 Federal Register October 20, 1961. Applies to All DC-8 Aircraft. \n\n\tCompliance required within the next 10 hours' time in service. \n\n\tAs an interim safety measure pending development of modifications to the throttle control system and reversing mechanism the following procedures shall be followed: \n\n\tReverse thrust after landing shall not be used until: \n\n\t(a)\tThe main and nose gear of the aircraft are firmly on the ground, \n\n\t(b)\tthe blue ejector lights are on steady, \n\n\t(c)\tthe reverse power lever is positioned to reverse detent, and \n\n\t(d)\tthe amber reverser lights are on steady. \n\n\tIf the amber reverser lights are on steady, additional symmetrical reverse may be used on the inboard engines and thereafter reverse power may be used on the outboard engines if required. \n\n\tThis directive effective for all persons except those to whom it was made effective immediately by telegram dated October 6, 1961.
69-26-08: 69-26-08 LOCKHEED: L-188 Series Airplanes. Amdt. 39-904. Applies to Lockheed L-188 airplanes listed in Lockheed Service Bulletin 88/5B-672, dated 10 November 1969, or later FAA approved revisions, and to any other L-188 airplane which has the Lockheed takeoff warning system installed. Compliance required as indicated. To provide for arming the aural takeoff warning system at lower ambient temperatures, and to provide interim placard information to the flight crew on the present ambient temperature at which the system will not be armed, accomplish the following: (a) Within 50 hours time in service after the effective date of this AD install a placard in the cockpit in full view of the pilot to read: "Takeoff aural warning system will not be armed when setting takeoff power at ambient temperatures below 23 degrees F" This placard may be removed when Item (b) is accomplished. (b) Within 3000 hours time in service, unless previously accomplished modify the arming switches installation in the takeoff warning system to provide for arming the switches at about 68 degrees-70 degrees power lever advancement in accordance with Lockheed Service Bulletin 88/SB-672 dated November 10, 1969, or later FAA approved revisions, or an equivalent modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. This amendment becomes effective on January 10, 1970.