Results
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.
2016-01-17: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-2C10 (Regional Jet Series 700, 701, & 702) airplanes. This AD was prompted by reports of cracked forward door members of the inboard main landing gear (MLG) doors. This AD requires repetitive inspections of the inboard MLG doors, repairs if necessary, and replacement of the inboard MLG doors. This AD also provides optional terminating action for the door replacement. We are issuing this AD to prevent loss of an MLG door during flight, which could result in damage to the airplane.
2010-18-10: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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 events have been reported where insulation material was found to be fouling pulleys in the aileron interconnect circuit in the cabin roof area. * * * Interference between the cable and the insulation bag causes the material to be drawn into the gap between the pulley and the pulley guard. This condition, if not detected and corrected, could lead to restricted aileron movement and consequently, reduced control of the aeroplane. * * * * * We are issuing this AD to require actions to correct the unsafe condition on these products.
2005-10-16: The FAA is adopting a new airworthiness directive (AD) for General Electric (GE) CF6-80E1 series turbofan engines that have an electronic control unit (ECU) with software version E.1.M. or earlier installed. This AD requires installing improved software for the ECU. This AD results from an uncommanded engine acceleration event caused by a failure of the ECU digital interface unit (DIU). We are issuing this AD to prevent an undetected failure of the ECU DIU, which could result in uncommanded acceleration to the overspeed limit without response to throttle commands. The airplane could then experience asymmetric thrust.
2016-01-15: We are adopting a new airworthiness directive (AD) for Agusta S.p.A. (Agusta) Model AB139 and AW139 helicopters. This AD requires visually inspecting certain subfloor frames for a crack. This AD was prompted by reports of cracks on in-service helicopters. The actions of this AD are intended to detect or prevent a crack in the subfloor frame, which could result in failure of the pilot and co-pilot pedal support frame and subsequent loss of control of the helicopter.
2005-10-17: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 777-200 and -300 series airplanes, that requires modification of the bolt holes of the lower side of the body splice t-chord common to the paddle fittings of the lower wing panel. The modification includes performing a high frequency eddy current inspection of the fastener holes for cracks, repairing the hole if necessary, and replacing the fasteners with new inconel bolts. This action is necessary to prevent fatigue cracks in the lower t-chord at the bolt holes common to the paddle fittings that could result in fractures of one or more of the t-chord segments, which could lead to detachment of the lower wing panel and consequent loss of the wing. This action is intended to address the identified unsafe condition.
73-26-06: 73-26-06 BOEING: Amdt. 39-1762. Applies to all Model 727 airplanes, certificated in all categories, listed in Boeing Service Bulletin 727-32-203, Revision 3, dated November 21, 1973, or later FAA approved revisions. \n\tA.\tFor reworked torsion shafts which were not plated in a manner to provide corrosion protection as specified in Boeing Service Bulletin 727-32-203, Revision 3, dated November 21, 1973, within the next 750 landings after the effective date of this AD, unless already accomplished within the last 750 landings, and thereafter at intervals not to exceed 1500 landings since the last such inspection, inspect the shafts for cracks in accordance with such service bulletin or later FAA approved revisions. Shafts found cracked are to be replaced prior to further flight with one of the following, any one of which is terminating action. \n\t\t1)\tA shaft P/N 65-18949-1 which has never been reworked; or \n\t\t2)\tA shaft P/N 65-18949-1 which has been reworked in accordance with BoeingService Bulletin 727-32-203, Revision 3, dated November 21, 1973, or later FAA approved revisions; or \n\t\t3)\tA shaft (P/N 65-18949-1) which has been reworked in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region; or \n\t\t4)\tA newly designed shaft as specified in and installed in accordance with Boeing Service Bulletin 727-32-205, dated March 14, 1973, or later FAA approved revisions. \n\tB.\tPrior to accumulation of 8,000 landings on a reworked torsion shaft which was not plated in a manner to provide corrosion protection, as specified in Boeing Service Bulletin 727- 32-203, Revision 3, dated November 21, 1973, or within the next 3,000 landings time in service after the effective date of this AD, whichever occurs later, replace the shaft with one of those described in A(1), A(2), A(3), or A(4). \n\tC.\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 oflandings 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\tD.\tUpon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering & Manufacturing Branch, FAA Northwest 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 that operator. \n\n\tThis Amendment becomes effective upon publication.
61-20-03: 61-20-03 RUSSELL MANUFACTURING COMPANY SAFETY BELT: Amdt. 337 Part 507 Federal Register September 20, 1961. Applies to All Aircraft Equipped With Rusco Model RM23 P/N SB1709 Safety Belts Manufactured From October 1959, through June 1960, and Equipped With D-832B Buckles. Compliance required with the next 25 hours' time in service after the effective date of this directive. Service difficulties with Rusco Model RM23, P/N SB1709, safety belt assemblies manufactured from October 1959, through June 1960, have been reported wherein the buckle cam does not securely grip the webbing thereby causing the assembly to slip. In order to determine the airworthiness of these belts, a minimum tensile load of 100 pounds must be applied to the belt assembly. If the webbing slips through the buckle as a result of this test, the belt must be considered unairworthy and replaced with an airworthy belt assembly conforming to TSO-C22 standards prior to further flight. (Russell Manufacturing Company Bulletin No. 61-5 covers this same subject.) This directive effective October 20, 1961.
74-04-01: 74-04-01 CESSNA: Amdt. 39-1783. Applies to Model 172 (Serial Numbers 17260759 through 17261495, F17200905 through F-17201034 and FR17200351 through FR17200440) series airplanes. Compliance: Required as indicated, unless already accomplished. To detect cracks in the aft fuselage bulkhead assembly and provide means for repair of cracked assemblies, accomplish the following: A) Within the next 100 hours' time in service after the effective date of this AD, visually inspect the aft fuselage bulkhead assembly (P/N 0512157-7) for cracks in accordance with Cessna Service Letter SE73-37 dated November 16, 1973, or later FAA-approved revisions. B) If as a result of any inspection required herein, cracks are found, prior to further flight, repair the bulkhead assembly in accordance with Cessna Service Kit SK 172-44. C) Any alternate methods of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Thisamendment becomes effective February 12, 1974.
2005-10-21: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 747 airplanes. That AD currently requires inspections to detect cracks in the front spar pressure bulkhead chord, and repair if necessary. This new AD requires repetitive high frequency eddy current (HFEC) inspections of the body station (BS) 1000 bulkhead chord for cracks, repetitive detailed inspections of the bathtub fittings, if installed, for cracks, and corrective action if necessary. Initiation of the new inspections ends the inspections of the existing AD. This AD also revises the applicability of the existing AD to include additional airplanes. This AD is prompted by reports of cracks in the BS 1000 bulkhead chord. We are issuing this AD to detect and correct fatigue cracks in the BS 1000 bulkhead chord, which, if not repaired before they reach critical length, could result in the failure of the adjacent structure and skin and lead to in-flight depressurization of the airplane. \n\nDATES: This AD becomes effective June 23, 2005. \n\n\tThe incorporation by reference of Boeing Alert Service Bulletin 747-53A2471, dated March 27, 2003; and Boeing Service Bulletin 747-53- 2064, Revision 4, including Addendum, dated September 23, 1983; as listed in the regulations; is approved by the Director of the Federal Register as of June 23, 2005.
75-07-08: 75-07-08 SIKORSKY AIRCRAFT: Amendment 39-2138. Applies to Models S-58 and S-58T series helicopters certificted in all categories, including Military Type HSS-1, HSS-1F, HSS-1N, HUS-1, HUS-1A, HUS-1AN, HUS-1G, HUS-1Z, H-34A, H-34C, H-34J, CH-34A, CH- 34C, HH-34F, SH-34G, SH-34H, SH-34J, UH-34D, UH-34E, UH-34G, UH-34J, VH-34C, VH- 34D, equipped with S1610-31100 series rotary rudder assemblies and S1610-34200 series counterweight assembly. Compliance required as indicated. To prevent excessive stresses and possible structural failures in the tail rotor blades, tail rotor gear box assembly, and tail pylon area, from tail rotor unbalance, accomplish the following: a. Within the next 10 hours time in service after the effective date of this AD unless already accomplished, lubricate and inspect the rotary assembly, P/N S1610-31100 series and counterweight assembly, P/N S1610-34200 series, for balance in accordance with Section 2, Paragraph A of Sikorsky Service Bulletin No. 58B15-14A, dated December 19, 1974, or later FAA approved revision. b. Thereafter, lubricate and inspect at intervals not to exceed 50 hours time in service from the last inspection, or earlier if unbalance is suspected, in accordance with Section 2, Paragraph B of Sikorsky Service Bulletin No. 58B15-14A or later FAA approved revision. c. If after two consecutive 50 hour repetitive inspections as specified in Paragraph b., the rotor balance remains unchanged, and within limits, these inspections can be extended to a 100 hour interval. These 100 hour intervals can be maintained if all subsequent 100 hour inspections show that rotor balance remains unchanged. d. Rotary rudder assemblies, P/N S1610-31100 series and counterweight assembly P/N S1610-34200 series which are not balanced as specified in Paragraphs a., b., and c must be either rebalanced or removed for corrective action in accordance with Section 2, Paragraphs A and B of Sikorsky Service Bulletin No. 58B15-14A or laterFAA approved revision, before further flight. e. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Sikorsky Aircraft, Division of United Aircraft Corporation, Stratford, Connecticut 06602. These documents may also be examined at the Office of the Regional Counsel, New England Region, Federal Aviation Administration, 12 New England Executive Park, Burlington, Massachusetts 01803. This amendment becomes effective April 8, 1975.
61-16-05: 61-16-05 PIPER: Amdt. 311 Part 507 Federal Register July 26, 1961. Applies to PA-18/150, PA-18A/125, PA-18A/135, PA-18A/150, PA-18S/150, PA-18AS/150: Serial Numbers 18-1667, 18-1887, 18-4260, 18-4860, 18-6127, 18-6289, 18-6301, 18-6332, 18-6375, 18-6398, 18-6418, 18-6424, 18-6434, 18-6437, 18-6438, 18-6444, 18-6463, 18-6466, 18-6501, 18-6540, 18-6604, 18-6606, 18-6609, 18-6616, 18-6657, 18-6668, 18-6679, 18-6680, 18-6681, 18-6682, 18-6771, 18-6855, 18-6914, 18-6924, 18-6981, 18-6991, 18-6992, 18-6993, 18-7018, 18-7025, 18-7037, 18-7043, 18-7047, 18-7058, 18-7071, 18-7072, 18-7075, 18-7078, 18-7093. PA- 22/150, PA-22S/150, PA-22/160, PA-22S/160: Serial Numbers 22-6116, 22-6167, 22-6359, 22-6421, 22-6466, 22-6550, 22-6704, 22-6758, 22-6883. Compliance required within 10 hours' time in service after the effective date of this AD unless already accomplished. To preclude loss of control of the airplane as a result of failed nicopress sleeves in the control system, the following must be accomplished: (a) In PA-18 Series aircraft, aileron, lower elevator, and flap flexible stainless steel cable assemblies, P/N's 12794-03, 13271-02, 13745-02, 40123-44, 40123-86, 40123-87, and 10870-12 shall be replaced with respective assemblies P/N's 12794-00, 13271-00, 13745-00, 40123-03, 40123-77, 40123-76, and 10870-08, or respective assemblies P/N's 12794-04, 13271- 03, 13745-03, 14300-19, 14300-27, 14300-28, and 10870-17. (b) In PA-22 Series aircraft, aileron, lower elevator, flap and rudder cable assemblies P/N's 11527-03, 12515-04, 13108-03, 13109-13, 13109-15, 40123-83, 40123-84, and 40123-94 shall be replaced with respective cable assemblies P/N's 11527-02, 12515-03, 13108-02, 13109- 10, 13109-12, 40123-68, 40123-69, and 40123-93; or respective cable assemblies P/N's 11527- 04, 12515-06, 13108-11, 13109-18, 13109-19, 14300-24, 14300-25, and 14300-29. Standard landplane galvanized cables are satisfactory for continued or replacement use. (PiperService Bulletin Number 181 dated November 5, 1959, and Piper Service Letter No. 355, dated February 23, 1961, cover this same subject.) This supersedes AD 59-26-06. This directive effective July 26, 1961.
66-03-01: 66-03-01 BRISTOL SIDDELEY: Amdt. 39-186 Part 39 Federal Register January 28, 1966. Applies to de Havilland Model Gipsy Queen Series 30 Engines With Crankshafts With Modification 2036 or 2162 Installed. Compliance required as indicated. To prevent further failures of the crankshaft front end that could result in loss of the propeller, accomplish the following: (a) Unless already accomplished within the last 375 hours' time in service, inspect crankshaft for cracks within the next 25 hours' time in service after the effective date of this AD in accordance with Bristol Siddeley Technical News Sheet (TNS) GQ30 No. 70, dated November 16, 1965, or later ARB-approved issue, or an FAA- approved equivalent, and thereafter at intervals not to exceed 400 hours' time in service from the last inspection until modified in accordance with (c). (b) Replace cracked crankshafts before further flight with an uncracked crankshaft. (c) Incorporate Modification 2925 (which introduces shot peening underneath the chrome to improve the fatigue strength) on crankshaft within the next 1,600 hours' time in service after the effective date of this AD unless already accomplished. This directive effective January 28, 1966.
61-14-02: 61-14-02 BELL: Amdt. 300 Part 507 Federal Register July 1, 1961. Applies to All Models 47J, 47J-2, 47G-3, and 47G-2A Helicopters With Float Gear Kit No. 47-706-661 installed. Compliance required as indicated. Several cases of failures have been reported of the forward cross tube failing near the left attaching point (cross tube to centerframe) through drilled holes which have been welded closed and filed smooth. To remove defective tubes and preclude the possibility of float cross tube No. 47-512-111-9 failing during landing, the following one-time inspection is required: (a) Within the next 10 hours' time in service after the effective date, remove paint from around the P/N 47-512-111-9 cross tube for an area approximately 2 inches on each side of the stop plates, P/N 47-500-024-1 and visually inspect for any indication of welded holes in this area. (b) If no evidence of welds is found, prime and repaint as necessary. (c) If any indication of a weldexists, replace the cross tube with an FAA approved part within the next 5 hours' time in service following the above inspection. (Bell Service Bulletin No. 132 dated May 23, 1961 covers this subject.) This directive effective July 12, 1961.
2005-10-18: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 747 series airplanes. That AD currently requires a one-time inspection to determine the material type of the stop support fittings of the main entry doors (MEDs). That AD also currently requires repetitive detailed inspections to detect cracks of certain stop support fittings of the MEDs, and replacement of any cracked stop support fitting with a certain new stop support fitting. This new AD adds new inspections, and replacement if necessary, of the stop support fittings of MED 3, and adds airplanes to the applicability. This AD is prompted by reports of MED 3 having certain stop support fittings that are susceptible to stress corrosion cracking. We are issuing this AD to detect and correct stress corrosion cracking of the stop support fittings of the MEDs, which could result in damage to the adjacent forward edge frame of the door and consequent loss of a MED and rapid decompression of the airplane. \n\nDATES: This AD becomes effective June 23, 2005. \n\n\tThe incorporation by reference of Boeing Service Bulletin 747-53- 2358, Revision 1, dated April 19, 2001; and Boeing Special Attention Service Bulletin 747-53-2485, dated January 8, 2004; as listed in the AD, is approved by the Director of the Federal Register as of June 23, 2005. \n\n\tOn January 25, 1999 (63 FR 70316, December 21, 1998), the Director of the Federal Register approved the incorporation by reference of Boeing Service Bulletin 747-53-2358, dated August 26, 1993.
55-25-04: 55-25-04 SIKORSKY: Applies to All Model S-55 Helicopters. Compliance required as indicated. A recent failure of the tail cone skin in the area where the tail cone is spliced to the fuselage has been reported in a Model S-55 helicopter. Since this is the means of attaching the tail cone to the fuselage and to preclude further failures, the following inspections and reinforcements for the splice area are considered necessary. 1. Tail cones with less than 1,500 hours, add reinforcements at the next major inspection or within the next 150 hours. 2. Tail cones with 1,500 hours to 2,000 hours, inspect the skin for cracks at the next major inspection or within the next 150 hours. If no cracks are found, add reinforcements. If cracks are found, replace the skin and add reinforcements. 3. Tail cones with 2,000 hours or more, inspect the skin for cracks at the next intermediate inspection or within the next 50 hours. If no cracks are found, add reinforcements. Ifcracks are found, replace the skin and add reinforcements. 4. The inspection procedures to be followed, the reinforcements to be added and the method of attaching the reinforcements are explained in detail in Sikorsky Information Circular No. 1420-632.
61-12-01: 61-12-01 DOUGLAS: Amdt. 293 Part 507 Federal Register June 3, 1961. Applies to All Model DC-8 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tAn instance has been reported wherein the left wing flight spoilers extended during takeoff resulting in a left wing heavy condition. The loss of spoiler control was attributed to a sheared rivet in rod assembly P/N 4719278 of the aileron shift mechanism in the left wheel well. To preclude further difficulties with this part which can result in a hazardous condition, the following inspection, operational checks or rework must be accomplished: \n\n\t(a)\tPrior to each flight, conduct an operational check of the outboard spoiler control system, observing the pressure drop in the spoiler hydraulic system during the check. During rapid rotation of the aileron control wheel a pressure drop of approximately 200 p.s.i. is normal. If a drop in hydraulic pressure does not occur, inspect the spoiler system for a sheared rivet in rod assembly P/N 4719278 of the spoiler shifting mechanism. If the rivet is not sheared a more thorough examination of the spoiler system should be made to determine and correct the cause of the malfunction since operation of the system should result in some pressure drop. In addition, after any maintenance work on the outboard spoilers or outboard spoiler control system and prior to return of the aircraft to service, the spoiler system should be thoroughly checked to ascertain that the rivet in rod assembly P/N 4719278 is not sheared and that the system operates normally. If sheared, the rivet must be replaced or "Rod Assembly-Lateral Control Spoiler" P/N 4719278-1 and "Spring" P/N 4771197-1B installed per Douglas Drawing No. 5718924 "J" change, or FAA approved equivalent, prior to next flight. Pilots and flight engineers, in addition to authorized maintenance personnel, are authorized to perform the operational check. \n\n\t(b)\tThe operational check and inspection described in (a) may bediscontinued when rework per Douglas Drawing No. 5718924 "J" change, or FAA approved equivalent, is accomplished. \n\n\t(Douglas Service Bulletin No. A27-104 dated March 27, 1961, covers this subject.) \n\n\tThis directive effective June 9, 1961.
2005-10-19: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 747-100, 747-100B, 747-200B, 747-300, 747-400, 747-400D, 747SR, and 747SP series airplanes. This AD requires replacing or modifying the control panels for the galley cart lift and modifying related electrical cable assemblies, as applicable. This AD is prompted by reports of injuries to catering personnel and flight attendants who were loading or unloading galley carts on one deck when, due to a disabled or malfunctioning safety interlock door switch, the galley cart lift unexpectedly moved when it was activated from the control panel on the other deck. We are issuing this AD to ensure that the galley cart lift can be sent only from the deck on which it is in use, which will prevent unexpected movement of the cart lift that could result in possible injury to catering personnel or flight attendants.
54-02-01: 54-02-01 LYCOMING: Applies to All Model O-290-D2 Engines. Compliance required by February 1, 1954, and at each subsequent 100-hour inspection. In order to preclude the possibility of serious engine damage from incipient detonation under certain atmospheric and altitude conditions, the magneto timing of all Lycoming O-290-D2 engines must be set at 18 degrees BTC. To insure that the magneto timing does not change substantially from this setting, it should be checked and reset, if necessary, at each subsequent 100-hour inspection. (Lycoming Service Bulletin No. 169 covers this same subject, but recommends that the timing be checked every 50 hours.)