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2006-13-11: We are adopting a new airworthiness directive (AD) that supersedes AD 2002-21-08, which applies to certain Pilatus Aircraft Ltd (Pilatus) Model PC-6 airplanes. AD 2002-21-08 currently requires you to inspect the aileron assembly for correct configuration and modify as necessary. Since we issued AD 2002-21-08, the FAA determined the action should also apply to all the models of the PC-6 airplanes listed in the type certificate data sheet of Type Certificate (TC) No. 7A15 that were produced in the United States through a licensing agreement between Pilatus and Fairchild Republic Company (also identified as Fairchild Industries, Fairchild Heli Porter, or Fairchild-Hiller Corporation). In addition, the intent of the applicability of AD 2002-21-08 was to apply to all the affected serial numbers of the airplane models listed in TC No. 7A15. This AD retains all the actions of AD 2002-21-08, adds those Fairchild Republic Company airplanes to the applicability of this AD, and lists theindividual specific airplane models. We are issuing this AD to correct improper aileron assembly configuration, which could result in failure of the aileron mass balance weight. Such failure could lead to loss of control of the airplane.
2006-10-21: The FAA is adopting a new airworthiness directive (AD) for certain Lycoming Engines (formerly Textron Lycoming) 360 and 540 series reciprocating engines with ECi connecting rods, part number (P/N) AEL11750, installed. The Airmotive Engineering Corp, Division of Engine Components Incorporated (ECi), holds the Parts Manufacturer Approval (PMA) for the affected parts, and markets the parts as ECi parts. This AD requires replacing certain lot and serial numbered connecting rods, P/N AEL11750, having forging part number AEL11488. This AD would also prohibit installing certain ECi connecting rods, P/N AEL11750, into any Lycoming 360 or 540 series reciprocating engines. This AD results from reports of connecting rods with excessive variation in circularity of the journal bores. We are issuing this AD to prevent fatigue failure of the connecting rod and a possible uncommanded shutdown of the engine.
48-04-02: 48-04-02 AERONCA: Applies to 7AC Serial Numbers 7AC-1 Through 7AC-7129; 11BC Serial Numbers 11BC-1 Through 11BC-173; and 11AC Serial Numbers 11AC-1 and Up. \n\n\tCompliance required by March 1, 1948. \n\n\tInspect the wing leading edge for buckled nose ribs or loose PK screws by pressing leading edge skin with hand to nose ribs. If the skin can be depressed beyond the normal wing contour, other than the extreme nose radius, indicated by section A-A in Figure 1, the fabric should be cut open on the bottom surface just forward of the front spar for thorough inspection. Item No. 1 below should be accomplished on all wings whether damage has occurred or not, whereas item No. 2 pertains only to damaged ribs found in the above inspection. Repair need not be made if buckling is confined to area forward of section A-A of Figure 1. \n\n\n\n\t1.\tTo help prevent further failures of the nose ribs five additional No. 4 x 1/4 PK screws or, as an alternate, Cherry CR163-4-2 rivets or equivalentare to be installed in all nose ribs. Four PK screws or rivets are to be installed on the top surface and one PK screw or rivet is to be installed in the bottom as shown in Figure 1. Apply dope liberally under the PK screw head before tightening. It is not necessary to remove the fabric to accomplish this modification. \n\n\t2.\tDamaged ribs should be cut away at the top and bottom of spar. The new nose ribs are installed by means of two gussets on the side of the ribs as shown in Figure 1. Factory kits, Aeronca P/N 5-185-2 and 5-190-2, are to be used. \n\n\t(Aeronca Helps and Hints No. 17 with three supplements thereto covers this same subject.)
2006-13-01: The FAA is superseding an existing airworthiness directive (AD), which applies to certain Boeing Model 727-200 series airplanes. That AD currently requires initial and repetitive inspections for cracks in the forward frame of the No. 3 cargo door cutout; and corrective actions, if necessary. The existing AD also provides for an optional structural modification, which terminates the repetitive inspections. This new AD reduces the compliance time for the initial inspections and adds an optional method of inspection for both the initial and repetitive inspections. This AD also adds initial and repetitive inspections of an additional area, and repair if necessary. Additionally, this AD clarifies that the previously optional structural modification is now required by other rulemaking. This AD results from additional reports of cracking in the forward frame of the No. 3 cargo door cutout. We are issuing this AD to detect and correct cracking of the forward frame and fuselage skin ofthe No. 3 cargo door cutout, which could result in failure of the frame and skin, and consequent rapid decompression of the airplane.
2006-09-03: The FAA is adopting a new airworthiness directive (AD) for all Boeing Model 727, 727C, 727-100, and 727-100C series airplanes. This AD requires repetitive inspections for cracks in the body skin and bear strap at the upper and lower hinge cutouts of the mid-cabin galley doorway, along the upper fastener row of the stringer 14R lap splice, and in the doorstop fitting adjacent to the upper hinge cutout; and corrective action if necessary. This AD also provides for optional terminating action for certain inspections. This AD results from reports of skin and bear strap cracking at the upper and lower hinge cutout and along the upper fastener row of the stringer 14R lap splice, and cracking in the doorstop fitting adjacent to the upper hinge cutout. There are also reports of cracking on airplanes previously modified in production to resist such cracking. We are issuing this AD to find and fix fatigue cracking of the fuselage, which could result in reduced structural integrity and consequent rapid decompression of the airplane.
69-26-02: 69-26-02 HAWKER SIDDELEY AVIATION, LIMITED: Amdt. 39-894. Applies to Heron Model D.H. 114 Airplanes. Compliance is required as indicated. To reduce the possibility of fatigue cracks developing in the engine mount pick-up wing straps located at the top outboard position of the left and right inboard engine-to-wing attachment structure, accomplish the following: (a) Within 50 hours' time in service after the effective date of this AD, unless already accomplished within the last 250 hours' time in service, remove the top cowling panel over the oil tank on the right and left inboard engine installations and visually inspect the upper outboard engine mounting pick-up wing straps for fatigue cracks using a dye penetrant method and in accordance with Hawker Siddeley Technical News Sheet, Series Heron 114, No. W.15, Issue 1, dated 27 October 1969, or an FAA-approved equivalent. (b) Within the next 300 hours' time in service after accomplishing the inspection required byparagraph (a), accomplish the following: (1) Remove and discard the two aft bolts and nuts which attach the upper outboard engine mount pick-up straps of the left and right inboard engine installation structure to the engine mounting pick-up fitting. (2) Accomplish the inspection required in paragraph (a). (3) Replace the bolts and nuts removed in accordance with subparagraph (1) with new bolts, P/N A.25/4E, and new nuts, P/N A.16Y/ET in accordance with Hawker Siddeley Technical News Sheet, Series Heron (114), No. W.15, Issue 1, dated 27 October 1969, or an FAA-approved equivalent. (c) Within the next 600 hours' time in service after compliance with paragraph (b) and thereafter at intervals not to exceed 600 hours' time in service since the last inspection, accomplish the inspection required by paragraph (a). (d) If cracks are found during any of the inspections required by paragraphs (a), (b), and (c) replace the cracked wing strap with a new strap in accordance with Hawker Siddeley Technical News Sheet, Series Heron (114), No. W.15, Issue 1, dated 27 October 1969, or an FAA-approved equivalent. This amendment becomes effective December 18, 1969.
2006-12-23: The FAA is superseding an existing airworthiness directive (AD) that applies to certain Boeing Model 737-100, -200, -200C, -300, - 400, and -500 series airplanes. The existing AD currently requires initial and repetitive inspections of the elevator tab assembly to find any damage or discrepancy; and corrective actions if necessary. This new AD adds certain new inspections and removes certain existing inspections. This AD results from additional reports of airframe vibrations of the elevator tab during flight on airplanes inspected per the existing AD; subsequently, considerable damage was done to the elevator tab, elevator, and horizontal stabilizer. In several incidents, a portion of the elevator tab separated from the airplane. We are issuing this AD to prevent excessive in-flight vibrations of the elevator tab, which could lead to loss of the elevator tab and consequent loss of controllability of the airplane. \n\nDATES: This AD becomes effective July 3, 2006. \n\n\tThe Director of the Federal Register approved the incorporation by reference of a certain publication listed in the AD as of July 3, 2006. \n\n\tOn February 19, 2002 (67 FR 1603, January 14, 2002), the Director of the Federal Register approved the incorporation by reference of Boeing Service Bulletin 737-55A1070, Revision 1, including appendices A, B, and C, dated May 10, 2001. \n\n\tWe must receive any comments on this AD by August 15, 2006.
73-20-04: 73-20-04 CESSNA: Amendment 39-1726. Applies to Model R172E (T-41B) Serial Numbers R172-0001 through R172-0256) airplanes. Compliance: Within 50 hours' time in service after the effective date of this AD, unless already accomplished. To prevent complete loss of engine power when using the fuel boost pump, accomplish the following: Modify the fuel boost pump electrical circuit by installing a new boost pump switch, electrical resistors, and placard in accordance with Cessna Service Letter No. SE73-24, dated August 24, 1973, and Service Kit SK172-43, or later FAA-approved revisions, or any equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective October 3, 1973.
73-20-02: 73-20-02 BOEING: Amendment 39-1722 as amended by Amendment 39-2763 is further amended by amendment 39-3393. Applies to all B-17F & G airplanes certificated in all categories. Compliance required within the next 25 hours' time-in-service or within 6 months after the effective date of this AD, unless already accomplished within the last 25 hours' time-in-service or 6 months, and thereafter at intervals not to exceed 50 hours' time-in-service or 12 months, whichever comes first, (from the last inspection). To detect cracking in the front spar lower cap center section, P/N 753424-2, accomplish the following: (a) Remove the most inboard bolt from the 8 bolt pattern attaching the front spar center section lower chord P/N 753424-2 to the terminal plates, P/N 56-3852-500 (aft) and 46-3852 (forward), left and right-hand sides. The bolt is approximately 8 inches inboard of the inboard jack pad. Using eddy current inspection procedures, or borescope methods in conjunction with dyepenetrants, inspect the front spar lower chord center section for cracks around the bolt hole in both the forward and aft wall of the tube. Particular attention should be given to the top and bottom portion of the tube. Removal of the bolt may necessitate installation of an access panel in the wing fillet fairing just forward of the front spar. The access panels may be installed using the procedures of FAA Advisory Circular 43.13-2. (b) If cracks are found, replace the spar cap with a serviceable part of the same part number, or repair in accordance with Army T.O. No. 01-20E-3 or other method approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. (c) After repairs per (b) have been accomplished, reinspect in accordance with (a) at intervals not to exceed 150 hours time-in-service or every 12 months, whichever comes first. (d) The bolt holes described in paragraph (a) above may be reamed .063 inch oversize for a close tolerance oversize bolt, provided no cracks are detected when the chords are inspected in accordance with paragraph (a) above using the eddy current inspection methods. Any holes reworked with the oversize bolts must be reinspected in accordance with (a) above within 1500 flight hours after such rework. Upon accumulation of 1500 flight hours on the reworked holes, the repeat inspection interval reverts to the interval specified in (c) above. (e) Any new replacement beam chords must be inspected within 2500 flight hours after installation and thereafter at intervals specified in paragraph (c) above. Amendment 39-1722 became effective upon publication in the Federal Register. Amendment 39-2763 became effective November 17, 1976. This amendment 39-3393 becomes effective January 26, 1979.
53-24-01: 53-24-01 GRUMMAN AIRCRAFT: Amendment 39-1264. Applies to all G-21A (Converted OA-9, JRF-1 through JRF-5 Under TC 654) Aircraft. Compliance required as indicated. 1. Within the next 50 hours in service after the effective date of the revision to this airworthiness directive, unless already accomplished within the last 50 hours' in service and at intervals thereafter not to exceed 50 hours' in service, accomplish Paragraph (3) for the stabilizer forward attachment fittings (P/N 12548 and P/N 12093). 2. Within the next 100 hours' in service after the effective date of this airworthiness directive, unless already accomplished within the last 100 hours in service and at intervals thereafter not to exceed 100 hours in service, accomplish paragraph (3) for the upper terminal (P/N 12561-1) of stabilizer strut (P/N 12560). 3. Inspect for cracks extending radially from the outside edge of the ears to the inside of the hole into which the shoulder bushings are pressed.Cracked parts must be replaced before further flight with parts inspected in accordance with this Directive or with equivalent parts approved by the Chief, Engineering and Manufacturing Branch, FAA, Southern Region. 4. Upon request with substantiating data submitted through an FAA Maintenance Inspector, the compliance times specified in this airworthiness directive may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. 5. (Grumman Aircraft Engineering Corporation Service Bulletin No. 21 dated March 17, 1948, and Customer Bulletin No. 5 dated October 30, 1953 cover this same subject). AD 53-24-01 which was to have been accomplished by January 15, 1954 supersedes AD 48-18-01. This amendment 39-1264 is effective August 17, 1971.
98-06-01: This amendment adopts a new airworthiness directive (AD), applicable to all Dassault Model Mystere-Falcon 50 series airplanes, that requires a one-time inspection of the clearances around the wiring harnesses of the right-hand electrical cabinet, and readjustment of the clearances, if necessary. This amendment will also require installation of protective strips on the wiring harnesses and equipment supports. This amendment is prompted by issuance of mandatory continued airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent interference between the wiring harnesses and adjacent equipment, support brackets, and structural elements, which could cause an electrical short circuit resulting in fire, and consequent loss of electrical power to essential flight systems.
2006-12-09: The FAA is superseding an existing airworthiness directive (AD), which applies to certain BAE Systems (Operations) Limited Model BAe 146 and Avro 146-RJ airplanes. That AD currently requires one-time inspections of the inner webs and flanges at frames 15, 18, 41, and 43 for evidence of corrosion or cracking; and corrective actions if necessary. This new AD instead requires new repetitive inspections and expands the area to be inspected. This new AD also expands the applicability and provides an optional action that would extend the repetitive inspection interval. This AD results from a report indicating that in some cases the inspections required by the existing AD revealed no damage, yet frame corrosion and cracking were later found during scheduled maintenance in the two forward fuselage frames 15 and 18. We are issuing this AD to prevent reduced structural integrity of the airplane.
73-16-01: 73-16-01 SIKORSKY: Amendment 39-1694. Pursuant to the authority of the Federal Aviation Act of 1958, as amended, delegated to me by the Administrator, the following airworthiness directive applicable to operators of Sikorsky S-55 series helicopters modified to incorporate AiResearch TSE331-3U-303N engines installed in accordance with Supplemental Type Certificate No. SH125-WE or SH127WE was effective immediately upon receipt of the telegram dated July 17, 1973. This directive is necessary because of the possibility of engine overspeed and resultant third stage turbine failure. Except as provided by FAR 21.197, the following is required prior to further flight: (a) Install a fuel bypass system in accordance with AiResearch Service Bulletin No. TSE331-73-5004 and Aviation Specialties Service Bulletin No. AS55-01-1, both dated July 16, 1973, or later FAA-Approved revisions thereto. (b) Incorporate the FAA-Approved Aviation Specialties Rotorcraft Flight Manual Revision3 dated July 17, 1973. (c) Equivalent modifications and rotorcraft flight manual revisions may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. This amendment is effective July 31, 1973, for all persons except those to whom it was made effective immediately by telegram dated July 17, 1973.
2006-12-04: The FAA is adopting a new airworthiness directive (AD) for certain Viking Air Limited Model DHC-7 airplanes. This AD requires revising the FAA-approved Airworthiness Limitations section of the airplane maintenance manual to prohibit operation of the airplane past its designed life limit for the primary structure, which is 80,000 total flight cycles. This AD also requires contacting the FAA for approval of analysis that substantiates that the airplane is safe to continue operation beyond the designed life limit. This AD results from a report that the designed life limit for the primary structure for the affected airplanes is 80,000 total flight cycles. We are issuing this AD to prevent continued operation of an airplane beyond its designed life limit for the primary structure, which could result in reduced structural integrity of the airplane.
71-02-03: 71-02-03 GENERAL DYNAMICS: Amdt. 39-1145 as amended by Amendment 39- 1165. Applies to Model 340, 440 and C-131E Airplanes including those airplanes converted to turbopropeller power, certificated in all categories. Compliance required within the next 50 hours time in service after the effective date of this AD unless already accomplished within the last 200 hours time in service. To prevent failures of the left and right main landing gears, accomplish the following: (a) Inspect the entire outer surface of the main landing gear cylinders (P/N 528002 or P/N 528402), including the fulcrum arms, for cracks using magnetic particle or dye penetrant methods, or an equivalent method approved by the Chief, Aircraft Engineering Division, FAA Western Region. (b) If cracks are found, before further flight either rework the cylinder in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region or replace the cylinder with a cylinder which has been inspected per (a) above and found free of cracks. (c) If cylinders are reworked in accordance with (b) above, accomplish the following before further flight: 1. Identify and record the cylinders and areas reworked. 2. Limit the use of reworked cylinders to those aircraft operating below or at a maximum of 55,000 lbs. take-off gross weight. If any cylinder reworked per (b) above is installed in any aircraft converted to turbopropeller power in accordance with STC SA4-1100 (known as Model 580), install a placard in only those aircraft previously approved for operating weights above 55,000 pounds, and in full view of the pilot, which reads as follows: "Maximum take-off gross weight 55,000 lbs." 3. Repeat (a) above at intervals of 1,000 hours' time in service from the last inspection. (d) If no cracks are found as a result of the inspection required by (a) above and it is definitely determined that a cylinder has less than 14,000 hours' time in service, repeat (a) above, before 15,000 hours' time in service have been accumulated. (e) If no cracks are found as a result of the inspection required by (a) or (d) above, and a cylinder is considered to have more than 14,000 hours' time in service, repeat (a) above at intervals of 1,000 hours time in service from the last inspection. Amendment 39-1145 effective January 21, 1971. This amendment 39-1165 becomes effective March 9, 1971.
58-08-06: 58-08-06 VICKERS: Applies to all Viscount 700 Series Aircraft. Compliance required as indicated. As a result of further cracks found in main chassis ram foot fittings, Vickers-Armstrong (Aircraft) Limited has recommended revised inspections which the British Air Registration Board considers mandatory. When each main chassis ram foot fitting has completed 1,500 landings, it should be inspected as follows: 1. A visual inspection is to be made within every 135 flying hours for the possible presence of cracks in the external surface of each ram foot fitting, particularly in the area of the base and sides of the ram socket for a distance of approximately 2 inches vertically from the base; 2. Inspect within the next 600 flying hours and thereafter within each subsequent 3,000 flying hours as follows: Remove the ram foot from each main undercarriage assembly and inspect for the possible presence of cracks, both inside and outside of the ram socket bore. The examination should be carried out using an approved method of crack detection and particular attention should be given to the radius at the bottom of the ram socket bore joining the bottom flange and the wall of the bore; 3. Any fittings found cracked should be replaced by new parts; 4. After compliance with Vickers-Armstrongs Modification D. 2695, the inspection outlined above may be discontinued. After January 31, 1959, all ram foot fittings exceeding 1,500 landings must incorporate Modification D.2695 or be replaced. The FAA concurs with this action and considers compliance therewith mandatory. (Vickers-Armstrongs PTL No. 175, Issue 2, and Modification D.2695, covers this subject.) This supersedes AD 57-26-01.
68-07-01: 68-07-01 GENERAL DYNAMICS: Amendment 39-568. Applies to Models 340, 440 and C-131E airplanes including those using turbo propeller power. Compliance required as indicated. Due to numerous reports of cracks developing in the Pilot and Copilot Direct Vision Window Frame Casting, PNs 340-3110314-9, -10, -13 and -14 (hereinafter referred to as the Casting) which affects the structural integrity of the Casting and which, in some cases, has caused loss of cabin pressurization, accomplish the following: (a) Inspect each Casting with 4500 or more hours' time in service on the effective date of this AD for cracks in accordance with Paragraph (c) within the next 250 hours' time in service after the effective date of this AD and thereafter at intervals not to exceed 250 hours' time in service from the last inspection. (b) Inspect each Casting with less than 4500 hours' time in service on the effective date of this AD for cracks in accordance with Paragraph (c) prior to the accumulation of 4750 hours' time in service and thereafter at intervals not to exceed 250 hours' time in service from the last inspection. (c) Inspect all visible areas of the main body of each Casting for cracks by means of either a visual inspection with the aid of an eight power glass, a dye penetrant inspection method, an eddy current inspection method, or by a method approved by the Chief, Aircraft Engineering Division, FAA Western Region. NOTE: In performing the inspection specified in Paragraph (c), special attention should be given to the lower left and right hand corner of the Casting. NOTE: For purposes of complying with this AD, the main body of the Casting includes only that part of the Casting which outlines the Direct View Window and does not include the attach flanges. (d) If a crack or cracks are found in a Casting comply with subparagraphs (1) and (2) of this Paragraph as appropriate: (1) If the cracked Casting is completely severed at any point, replace the affected Casting with a new part, P/N 340-3110314-9 or -13 (left hand side) or P/N 340- 3110314-10 or -14 (right hand side), prior to further flight after discovery of the crack (except that the airplane may be flown at a cabin pressure differential of zero p.s.i. in accordance with FAR 21.197 to a base where the replacement can be accomplished); and (2) If the cracked Casting is not completely severed at any point replace the affected Casting with a new part, P/N 340-3110314-9 or -13 (left hand side) or P/N 340- 3110314-10 or -14 (right hand side) within 250 hours' time in service after discovery of the initial crack except that until such time as the affected Casting is replaced in accordance with this Subparagraph: (i) The affected airplane must be operated at a cabin pressure differential of zero p.s.i.; and (ii) Prior to the initial takeoff after discovery of the initial crack, an operating limitation in the form of a placard must be installed inthe affected airplane in clear view of the pilot stating: "Operation Limitation. Pressurized Flight Prohibited." (e) Operators who have not kept records of hours' time in service of individual Castings shall substitute hours' time in service of the airplane in lieu thereof. (f) Upon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Regional Director, FAA Western Region, may adjust the compliance times herein if the request contains substantiating data to justify the increase for that operator. This amendment becomes effective on April 29, 1968.
67-27-02: 67-27-02 AVIONS MARCEL DASSAULT: Amdt. 39-478, Part 39, Federal Register September 9, 1967. Applies to Fan Jet Falcon Airplanes Serial Numbers 1 thru 89, except Serial Numbers 73, 78, 82, 85 and 87. Compliance required as indicated. To detect and prevent corrosion of the wing to fuselage recess and the wing to fuselage attachment bolts accomplish the following, unless already accomplished: (a) For all airplanes, within the next 200 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 200 hours' time in service from the last inspection, visually inspect the recesses of the wing fuselage junction for signs of corrosion, in accordance with Dassault Service Bulletin No. 282 (57-14), dated April 12, 1967, or later SGAC- approved or FAA-approved revision, or in accordance with an FAA-approved equivalent. (b) For airplanes without Dassault Modifications M1014A and M1014C and with more than 400 hours' time in service on the effective date of this AD, within the next 200 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 600 hours' time in service from the last inspection, inspect the wing to fuselage attachment bolts for signs of corrosion, in accordance with Dassault Service Bulletin No. 282 (57-14), dated April 12, 1967, or later SGAC-approved or FAA-approved revision, or in accordance with an FAA- approved equivalent. (c) For airplanes without Dassault Modifications, M1014A and M1014C and with less than 400 hours' time in service on the effective date of this AD, prior to the accumulation of 600 hours' time in service, and thereafter at intervals not to exceed 600 hours' time in service from the last inspection, inspect the wing to fuselage attachment bolts for signs of corrosion, in accordance with Dassault Service Bulletin No. 282 (57-14), dated April 12, 1967, or later SGAC- approved or FAA-approved revision, or in accordance with an FAA-approved equivalent. (d) If corrosion is found when conducting the inspections required by paragraphs (a), (b), or (c), within the next 200 hours' time in service, comply with paragraph (f). (e) If no corrosion is found during the inspections required by paragraphs (a), (b), or (c), incorporate the modifications specified in paragraph (f) prior to the accumulation of 1200 hours' time in service from the effective date of this AD, but in any event not later than December 31, 1968. (f) Incorporate Dassault Modifications M1014C and M1057, in accordance with Dassault Service Bulletins No. 234, revision 1, dated April 12, 1967, and No. 259, dated April 12, 1967, or later SGAC-approved or FAA-approved revisions, or an FAA-approved equivalent. (g) The repetitive inspections, required by paragraphs (a), (b), and (c), may be discontinued after the incorporation of Dassault Modifications M1014A, M1014C and M1057. This amendment effective October 9, 1967.
70-03-05: 70-03-05 BEECH: Amdt. 39-935 as amended by Amendment 39-1030 is further amended by Amendment 39-1472. Applies to Models H35, equipped with Continental O-470-G- CI engines, J35, K35, M35, N35, P35, S35, S35TC, V35, V35-TC, V35A, and V35A-TC, Serial Numbers D5062, D5331 through D9068; Models 35-33, 35-A33, 35-B33, 35-C33 and E33, Serial Numbers CD-1 through CD-1234; Models 35-C33A and E33A, Serial Numbers CE-1 through CE-289; Model F33C, Serial Numbers CJ-26 and up; Model E33C, Serial Numbers CJ-1 and up; Model 36, Serial Numbers E-1 through E-184 airplanes. Compliance: Required as indicated, unless already accomplished. A) Effective immediately, turning type takeoffs and a takeoff immediately following a fast taxi turn are prohibited. Avoid prolonged slips (20 seconds or more) with fuel tanks less than half full. B) Within 20 hours' time in service after the effective date of this AD, install a permanent type placard on the instrument panel in clear view of the pilot utilizing a minimum of 1/8 inch high letters, or at any equivalent location approved by an FAA Flight Standards Inspector, with the following wording: "TURNING TYPE TAKEOFFS, AND TAKEOFF IMMEDIATELY FOLLOWING FAST TAXI TURN PROHIBITED. AVOID PROLONGED SLIPS (20 SECONDS OR MORE) WITH FUEL TANKS LESS THAN HALF FULL." NOTE: The operator/owner may make and install the placard. C) Within 20 hours' time in service after the effective date of this AD, on Beech Models V35A (Serial Numbers D-8828 through D-9068), E33 (Serial Numbers CD-1181 through CD-1234), E33A (Serial Numbers CE-227 through CE-289) and 36 (Serial Numbers E-1 through E-184) airplanes and those Beech Model airplanes previously modified in accordance with Beech Service Instruction 0133-286 obliterate or remove Beech placard, P/N 33-924017, located either on the Fuel Selector Valve Cover Plate or on the Top Center of the floating instrument panel, which reads: "CAUTION - TO PREVENT FUEL FLOW INTERRUPTIONS DUE TO GRAVITY OR CENTRIFUGAL FORCE, SELECT THE HIGH-WING TANK IN SLIPS AND INSIDE TANK DURING TURNING TAKEOFFS." Naphtha will remove the instrument panel placard. (Beech Service Instruction 0133-286 has been cancelled.) D) Within 20 hours' time in service after the effective date of this AD, revise the Airplane Flight Manual, P/N 33-590004-1, dated September 28, 1968, on the Beech Model E33 airplanes and the Airplane Flight Manual, P/N 35-590116-3, dated September 20, 1968, on the Beech Model V35A airplanes as follows: In Section I, Limitations, under Item I obliterate the "required placard" paragraph which states, "CAUTION - TO PREVENT FUEL FLOW INTERRUPTIONS DUE TO GRAVITY OR CENTRIFUGAL FORCE, SELECT THE HIGH- WING TANK IN SLIPS AND INSIDE TANK DURING TURNING TAKEOFFS", and in its place insert a new paragraph with the words specified on the placard required by Paragraph B of the AD. Accomplish this insertion by affixing a typewritten or printed insert over the existingparagraph. Note: This insert may be made and installed by the operator/owner. E) (1) Beech Models K35, M35, N35, P35, S35, S35-TC, V35, V35-TC, V35A, V35A-TC, 35-33, 35-A33, 35-B33, 35-C33, 35-C33A, E33, E33A, E33C, F33C and 36 airplanes with fuel cell baffles installed in both wings in accordance with Beech Service Instructions 0459- 281 (Beech Kit Nos. 35-9009-1S, 35-9009-2S, 35-9009-3S or 35-9009-4S) or 365-281, Rev. 1, (Beech Kit Nos. 35-9009S or 35-9009-5S) or later revisions, or fuel reservoirs installed in both wings per Beech Kit 35-9012 or a fuel reservoir in one wing and a baffled cell in the other are exempt from compliance with the turning takeoff and 20 second side slip limitations of this AD. (2) On Models 35-C33A, E33A, E33C, F33C and 36 airplanes which have complied with Paragraph E(1) install a placard on the instrument panel in full view of the pilot with the wording, "MAXIMUM SIDESLIP DURATION 30 SECONDS", and operate the airplane accordingly.(3) On all other model airplanes listed in Paragraph E(1) (except those listed in Paragraph E(2), which have complied with Paragraph E(1), operate the airplane in accordance with the limitations set forth in Airplane Flight Manual Supplement P/N 35-590118-15 dated February 11, 1972, or later revision. F) Beech Models H35 (equipped with Continental O-470-G-CI engines) and J35 airplanes which have complied with Beech Service Instruction No. 0459-281 are exempt from compliance with this AD. Amendment 39-935 became effective February 5, 1970. Amendment 39-1030 became effective July 18, 1970. This Amendment 39-1472 becomes effective June 30, 1972.
71-20-03: 71-20-03 SUD AVIATION: Amendment 39-1296. Applies to Sud Model SE.210, MK VI-R, "Caravelle" airplanes. Within the next 500 hours' time in service after the effective date of this AD, unless already accomplished, incorporate S.A. Modification 1592 by installing a holding relay in the elevator servodyne jamming warning circuit in accordance with Sud Service-Caravelle Bulletin No. 27-218 at Revision 4, dated March 27, 1970, or an FAA-approved equivalent. This amendment becomes effective October 18, 1971.
2006-10-16: The FAA is superseding two existing airworthiness directives (ADs); one AD is applicable to all Boeing Model 747 airplanes and the other AD is applicable to certain Boeing Model 747 airplanes. The first AD currently requires repetitive inspections for cracking of the upper skin of the horizontal stabilizer center section and the rear spar upper chord, and repair if necessary. The other AD currently requires repetitive inspections for cracking of the upper skin of the outboard and center sections of the horizontal stabilizer and the rear spar structure, hinge fittings, terminal fittings, and splice plates; and repair if necessary. This new AD adds, for certain airplanes, repetitive inspections for cracking of the outboard and center sections of the horizontal stabilizer and repair if necessary. For certain other airplanes, this new AD adds a detailed inspection to determine the type of fasteners, related investigative actions, and repair if necessary. This new AD also revises the compliance times for certain inspections and adds alternative inspections for cracking of the upper skin of the center section and rear spar upper chord. This AD results from reports of cracking in the outboard and center section of the aft upper skin of the horizontal stabilizer, the rear spar chord, rear spar web, terminal fittings, and splice plates; and a report of fractured and cracked steel fasteners. We are issuing this AD to detect and correct this cracking, which could lead to reduced structural capability of the outboard and center sections of the horizontal stabilizer and could result in loss of control of the airplane. \n\nDATES: This AD becomes effective June 21, 2006. \n\n\tOn July 15, 2003 (68 FR 38583, June 30, 2003), the Director of the Federal Register approved the incorporation by reference of Boeing Alert Service Bulletin 747-55A2050, Revision 1, dated May 1, 2003. \n\n\tOn April 3, 2002 (67 FR 12464, March 19, 2002), the Director of the Federal Register approved the incorporation by reference of Boeing Alert Service Bulletin 747-55A2050, dated February 28, 2002.
98-06-20: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Airbus Model A320 series airplanes, that currently requires inspections to detect cracking of certain floor beams and side box-beams, and repair of cracks; and modification of the pressure floor. That AD was prompted by results of a full-scale fatigue test. This amendment adds a one-time inspection to verify proper clearance between the fasteners of the reinforcement bracket and the bellcrank of the free-fall extension system of the main landing gear (MLG) and its associated tie rod attachment nut. This amendment also adds a requirement for a new improved modification of the pressure floor. The actions specified by this AD are intended to prevent reduced structural integrity of the fuselage, restricted operation of the MLG free-fall system and, consequently, reduced ability to use the MLG during an emergency.
70-21-05: 70-21-05 HAWKER SIDDELEY AVIATION, LIMITED: Amendment 39-1090. Applies to deHavilland Model DH.114 "Heron" airplanes. To prevent failure of the flap datum hinge assemblies, unless already accomplished, accomplish the following within the next 3,000 hours' time in service after the effective date of this AD, or by March 31, 1971, whichever occurs first: (a) Inspect the wall thickness of the bearing housing recess of both the right wing and left wing flap datum hinge links in accordance with Hawker Siddeley Aviation, Limited, Technical News Sheet Heron (114) No. CF.14 Issue 1, June 15, 1970, or later ARB-approved issue or an FAA-approved equivalent. If the wall thickness is found to be less than 0.17 inches, replace the flap datum hinge link with a serviceable link of Modification 837 standard. (b) Incorporate Modification 837 by replacing the flap datum hinge assemblies P/N 14WF.16A(R.H.) and P/N 4WF.15A(L.H.) with assemblies P/N 14WF.456A(R.H.) and P/N 14WF.455A(L.H.)in accordance with Hawker Siddeley Aviation, Limited, Modification News Sheet, Modification No. Heron 837, dated June 15, 1956, or later ARB-approved issue or an FAA-approved equivalent. This amendment becomes effective November 5, 1970.
2006-10-04: The FAA is adopting a new airworthiness directive (AD) for certain Boeing Model 747-200B, 747-200C, 747-200F, 747-300, 747-400, and 747SP series airplanes. This AD requires doing a detailed inspection of the left and right longeron extension fittings, and corrective action if necessary. This AD results from cracking found in the longeron extension fitting at body station 1480 due to accidental damage during production. We are issuing this AD to detect and correct cracking in the longeron extension fitting, which could result in rapid decompression of the airplane and possible in-flight breakup of the airplane fuselage.
68-02-02: 68-02-02 BRITISH AIRCRAFT: Amendment 39-543. Applies to BAC 1-11 200 and 400 Series airplanes. Compliance required as indicated, unless already accomplished. To prevent possible inadvertent "stick push" during takeoff resulting from oleo relay failure, accomplish the following: Within the next 1,000 hours' time in service after the effective date of this AD, modify the stall protection system oleo relays by removing the links strapping the oleo relays and connecting the oleo relays in a series configuration, in accordance with British Aircraft Corporation BAC 1-11 Service Bulletin 34-PM 2784, dated August 28, 1967, or later ARB- approved issue, or an FAA-approved equivalent. This amendment becomes effective February 16, 1968.