Results
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.
69-21-03: 69-21-03 ALLISON GAS TURBINE DIVISION Amendment 39-862. Applies to Models 501-D13 and 501-D22 Series Engines. Compliance: Required as indicated. To insure adequate life limit margin for Allison P/Ns 6738651, 6788581, 6789681, 6805651, 6805701, 6844081, and 6829661 first stage compressor wheels with balance weight pin holes in the front lip of the outer rim, accomplish the following: (A) Wheels which have 16,000 cycles or less on September 30, 1969, shall be removed from service prior to reaching 19,000 cycles or, unless already accomplished, must be inspected in accordance with Allison Commercial Service Letter CSL-120 Supplement dated October 2, 1969, or Allison Commercial Service Letter 501-D22 CSL-1001 Supplement dated October 2, 1969, or any other method approved by the Chief, Engineering and Manufacturing Branch, Central Region. (B) Wheels which exceeded 16,000 cycles on September 30, 1969, must be removed from service in accordance with Inspection/Removal Chart contained on page 3 of Allison Commercial Service Letter CSL-120 Supplement dated October 2, 1969, or contained on page 3 of Allison Commercial Service Letter 501-D22 CSL- 1001 Supplement dated October 2, 1969, unless inspected in accordance with said Allison Commercial Letters. (C) No first stage compressor wheels shall remain in service beyond 25,100 cycles. (D) For the purposes of this airworthiness directive, a cycle is defined as one takeoff. This amendment becomes effective October 21, 1969.
2006-08-12: This amendment supersedes an existing airworthiness directive (AD) for the MD Helicopters, Inc. (MDHI) Model 600N helicopters, that currently requires inspecting both upper tailboom attachment fittings, nut plates and both angles for a crack or thread damage, and repairing or replacing any cracked or damaged part. That AD also requires replacing the upper right tailboom attachment bolt with a new attachment bolt, and if the upper right attachment bolt is broken, replacing the three remaining attachment bolts with airworthy bolts. Adding a washer to each bolt and modifying both access covers is also required. Thereafter, inspecting the upper tailboom attachments and repairing or replacing any cracked part is required by that AD. That AD was prompted by the discovery of a cracked attachment bolt on a helicopter. This AD requires those same actions, plus installing additional inspection holes in the aft fuselage skin panels and inspecting the upper and lower tailboom attachmentfittings, the upper longerons, and the angles and nut plates for cracks. It also requires, within a specified time, replacing the upper right tailboom attachment fitting, painting the inspection area, and replacing existing nut plates. Additionally, it requires inspecting the attachment bolts for any damage or wear. This amendment is prompted by an accident involving a Model 600N helicopter. The actions specified by this AD are intended to prevent failure of the tailboom attachment fittings, separation of the tailboom from the helicopter, and subsequent loss of control of the helicopter.
65-13-05: 65-13-05 LOCKHEED: Amdt. 39-87 Part 39 Federal Register June 22, 1965. Applies to Model 1329 Aircraft Serial Numbers 5001 through 5058. Compliance required before further flight except that one flight may be made in accordance with the provisions of FAR 21.197 for the purpose of obtaining inspection and rework. To prevent loss of elevator inboard counterweights accomplish the following: (a) Remove empennage trailing edge fairing installation P/N JE114 to gain access to inboard elevator counterweights and their attach bolts P/N AN 174H10A. (b) Remove aft outboard and two inboard attach bolts one at a time and measure shank length. Replace bolts with AN 174H10A bolt if shank length is less than 1.016 inch. Torque to 70-100 inch-pounds. (c) Back off forward outboard attach bolt eight turns. If bolt is still engaged, retorque to 70-100 inch-pounds. If bolt is not engaged at eight turns, but is engaged at least four turns, retorque to 70-100 inch-pounds. Replace any bolt having less than a four turn engagement with a AN 174H10A bolt. This can be accomplished by disconnecting the vertical push-pull rod from elevator horn on side being worked. Then remove torque tube flange bolt connecting the elevator. This will free elevator assembly so that counterweight can be rotated sufficiently for bolt access. Upon completion reassemble, torque and safety wire bolts. (d) Inspect two aft adjustment weight attach bolts, if installed, for proper length. Correct length is at least one thread extending through barrel nut. Replace as required with AN 174H()A bolt to obtain proper length. Torque to 70-100 inch-pounds. (e) Safety wire in pairs with MS20995C40 or equivalent, the attach bolt head of the two outboard, the two inboard and the two aft adjustment weight attach bolts, if installed. (f) Replace empennage trailing edge fairing. (g) At the next 100-hour inspection, remove and replace any forward outboard attach bolt found by inspection of (c) to have less than eight turn engagement, with AN 174H10A bolt torqued to 70-100 inch-pounds and safety wired in accordance with (e). (Lockheed Service Bulletin 329-206 covers this same subject.) NOTE: It is requested that the results of this inspection be reported to the FAA, Engineering and Manufacturing Branch, Southern Region, P.O. Box 20636, Atlanta, Georgia 30320. This directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated June 13, 1965.
70-02-10: 70-02-10 SCHEMPP-HIRTH K.G.: Amdt. 39-926. Applies to Schempp-Hirth Cirrus sailplanes having serial numbers 19, 46, 47, 50, 53, 55, 57 through 69, 71 through 75, 77, and 79 through 82. Compliance is required as indicated. To prevent failure of elevator, aileron or airbrake flight control systems, before further flight after the effective date of this AD, unless already accomplished, accomplish the following: Inspect askubal rod ends located in wings (4), fuselage (2), and fin (1) in accordance with Schempp-Hirth K.G. Technical Note No. 1/69 dated 20 October 1969, or an FAA-approved equivalent, to ensure that correct rod ends are installed in the elevator, aileron, and airbrake flight control systems. Incorrect rod ends may be recognized by thread of shanks going up to neck of bearing housing. If incorrect rod ends are installed, before further flight, replace them with correct rod ends in accordance with Schempp-Hirth K.G. Technical Note No. 1/69 dated 20 October 1969, or an FAA-approved equivalent. This amendment is effective upon publication in the Federal Register as to all persons except those persons to whom it was made immediately effective by the telegram dated 12 December 1969, which contained this amendment.
71-03-01: 71-03-01 RATIER-FIGEAC: Amdt. 39-1150. Applies to Ratier-Figeac FH.146 propellers installed on, but not necessarily limited to, Nord Aviation NORD 262A-12 airplanes. To prevent fatigue cracks and possible failures in service, the propeller components listed in Bulletin Ratier-Figeac Service No. 61-107, dated August 19, 1970, must be removed from service in accordance with the hours' time in service life limits specified in that Bulletin, or later SGAC-approved revision, or an equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region. This amendment becomes effective February 22, 1971.
2006-07-15: The FAA is adopting a new airworthiness directive (AD) to supersede AD 2003-07-01, which applies to certain Thrush Aircraft, Inc. Model 600 S2D and S2R (S-2R) series airplanes (type certificate previously held by Quality Aerospace, Inc. and Ayres Corporation). AD 2003-07-01 currently requires you to repetitively inspect the 1/4- inch and 5/16-inch bolt hole areas on the lower wing spar caps for fatigue cracking; replace or repair any lower wing spar cap where fatigue cracking is found; and report any fatigue cracking found. This AD is the result of the analysis of data from 112 cracks found in the last 8 years on similar design Model 600 S2D and S2R (S-2R) series airplanes, and FAA's determination that an immediate initial inspection and more frequent repetitive inspections are necessary to address the unsafe condition for certain airplanes. Consequently, this AD would require you to increase the frequency of the repetitive inspections on Groups 1, 2, 3, and 6 airplanes; and decrease the hours time-in-service (TIS) for the initial inspection on Group 2 airplanes. We are issuing this AD to prevent lower wing spar cap failure caused by undetected fatigue cracks. Such failure could result in loss of a wing with consequent loss of airplane control.