Results
61-06-05: 61-06-05 DOUGLAS: Amdt. 270 Part 507 Federal Register March 23, 1961. Applies to All DC-8 Series Aircraft. \n\n\tCompliance required within the next 20 hours' time in service, unless already accomplished in accordance with Douglas Alert Bulletin dated February 24, 1961. \n\n\tAs a result of reported instances of loosened rudder boost piston end assembly locknut, the following must be accomplished. \n\n\tRefer to DC-8 Overhaul Manual, Chapter 27-19-2, page 18. Visually inspect for loose locknut P/N NAS 509-16 and proper installation and safetying to locknut of lockwasher P/N NAS 513-16. Insure that lockwasher is not installed backwards and that tang on lockwasher is engaged in notch in piston end and that piston has not rotated out of its proper position. If any discrepancies are found, they must be corrected in accordance with the DC-8 Overhaul Manual and Douglas Alert Bulletin dated February 24, 1961, prior to further flight. \n\n\tThis directive becomes effective upon publicationin the Federal Register for all person except those to whom it was made effective immediately by individual telegrams dated March 8, 1961.
83-12-02: 83-12-02 PILATUS BRITTEN-NORMAN LTD.: Amendment 39-4668. Applies to Model BN-2A MK III Trislander Series airplanes (all serial numbers), not incorporating Britten- Norman Mod. NB/M/908, certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent the structural failure of the rudder drive lever support assembly and loss of rudder control, within the next 100 hours time-in-service after the effective date of this AD, accomplish the following: a) Replace the rudder drive lever support assembly attachment rivets with bolts in accordance with the instructions in "Part I" of the "Rectification" section of Britten-Norman Service Bulletin No. BN-2/SB.102, Issue 1, dated May 12, 1977 (hereinafter referred to as the SB). b) Visually inspect the corners of the two rudder lever mounting channels for cracks in accordance with the "Inspection" section of the SB. 1) If no evidence of cracks in the lever mounting channels is found, repeat the inspection for cracks specified in paragraph b) of this AD at intervals not exceeding 100 hours time-in-service. 2) If cracks less than 0.25 inches in length are found, within the next 25 hours time-in-service, stop drill ends of cracks and install the modification described in "Part 2" of the "Rectification" section of the SB. 3) If cracks greater than 0.25 inches in length are found, before further flight, stop drill end of cracks and install the modification described in "Part 2" of the "Rectification" section of the SB. c) The intervals between the repetitive inspections required by this AD may be adjusted up to 10 percent of the specified interval to allow accomplishing these inspections concurrent with other scheduled maintenance of the airplane. d) The repetitive inspections required by paragraph b) of this AD may be discontinued upon installation of the modification as described in "Part 2" of the "Rectification" section of the SB. e)Aircraft may be flown in accordance with FAR 21.197 to a location where this AD can be accomplished. f) An equivalent method of compliance with this AD, if used, must be approved by the Manager, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. This amendment becomes effective on July 27, 1983.
2017-22-12: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 757-200, -200PF, and -200CB series airplanes. This AD was prompted by reports of slats disbonding on airplanes on which the terminating actions of AD 2005-07-08 had been performed. We have also received reports of slats disbonding on airplanes outside of the applicability of AD 90-23-06, AD 91-22-51, and AD 2005-07-08, which also addressed slat disbonding. This AD requires determining the type of trailing edge slat wedges of the leading edge slats, repetitive inspections for disbonding on certain trailing edge slat wedges, and corrective actions if necessary. This AD also provides an optional terminating action for the repetitive inspections. We are issuing this AD to address the unsafe condition on these products.
61-06-03: 61-06-03 DOUGLAS: Amdt. 264 Part 507 Federal Register March 11, 1961. Applies to All DC-7 and DC-7B Aircraft, Fuselage Numbers 1 to 720 Inclusive, Having in Excess of 8,000 Hours' Time in Service. \n\n\tCompliance required as indicated. \n\n\tThere have been thirteen (13) reported cases of upper front spar cap cracking on DC-7 Series aircraft. Cracking usually occurs in spar cap tangs in the area of the Station 60 attachments and progresses chordwise. In addition, service experience has shown that the temporary repair of the above difficulty per Douglas Rework Drawing 5611387 does not have the service life originally anticipated. As a result of this service experience, the upper and lower, front and center spar caps in the area of wing Station 60, with special attention to the spar cap tangs between wing Stations 55 and 65, must be visually or radiographically inspected for cracks as follows: \n\n\t(a) The upper and lower, front and center spar caps must be inspected within the next 450 hours' time in service unless already accomplished. Aircraft inspected prior to issuance of this AD must also comply with the repetitive inspections, rework and/or repairs specified in (b), (c), (d), and (e). \n\n\t(b) The upper front and center spar caps must be reinspected at intervals not to exceed 1,600 hours' time in service. \n\n\t(c) The lower front and center spar caps must be reinspected at intervals not to exceed 3,250 hours' time in service. \n\n\t(d) If cracks are found, FAA approved permanent rework or temporary repair of the spar cap is required prior to further flight except ferry flight in accordance with provisions of CAR 1.76. Temporary repairs may be made per Douglas Rework Drawing 5611387, or FAA approved equivalent, providing crack limitations as established on this drawing have not been exceeded. Douglas DC-7 Service Bulletin No. 167 revised June 3, 1960, contains an FAA approved permanent rework consisting of Kit X plus the appropriate kits from the following list as indicated on Page 4 of the service bulletin: A, B, C, D, E, S, U, H, J, K, L. \n\n\t(e) Aircraft incorporating a temporary repair must be reinspected at intervals not to exceed 750 hours' time in service pending the accomplishment of an FAA approved permanent rework. Permanent rework must be accomplished within 4,200 hours' time in service after incorporating the temporary repair. \n\n\t(f) The special inspections specified in this AD are no longer required after an FAA approved permanent rework is accomplished. \n\n\t(Douglas DC-7 Service Bulletin No. 167 revised June 3, 1960, covers this subject.) \n\n\tThis directive effective April 11, 1961. \n\n\tRevised May 6, 1961.
2017-23-09: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model CL-600-2A12 (CL-601 Variant) and CL-600-2B16 (CL-601-3A, CL-601-3R, and CL-604 Variants) airplanes. This AD was prompted by a determination that the bushing holes on the engine mount rib might not conform to the engineering drawings and that certain inspections of the engine mount rib must be included in the airworthiness limitations section (ALS) of the Instructions for Continued Airworthiness (ICA). This AD requires revising the maintenance or inspection program to incorporate certain airworthiness limitation items (ALIs). We are issuing this AD to address the unsafe condition on these products.
61-05-02: 61-05-02 DOUGLAS: Amdt. 259 Part 507 Federal Register March 3, 1961, as amended by amendment 39-3354. Applies to All Models DC-6 Series Prior to Serial Number 44888 and DC-7 Series Up to Serial Number 44872 Having 32,000 or More Hours' Time in Service Including Military Models. \n\n\tCompliance required as indicated. \n\n\tInstances have been reported of loss of emergency exit doors during pressurized flight due to failure of the bottom hinge fitting with subsequent failures of the top hinge fitting. To preclude further occurrences, the following shall be accomplished on the emergency exit door hinges: \n\n\t(a) Unless already accomplished within the last 350 hours' time in service, emergency exit door hinges with less than a 1-inch door skin recess radius shall be inspected for cracks in the radius using dye penetrant, or equivalent, within the next 50 hours' time in service and repeated within each 400 hours' time in service thereafter until hinges are replaced as indicated in (b). If cracks are found, the hinges must be replaced prior to further pressurized flight. \n\n\t(b) Aircraft with hinges having less than a 1-inch door skin recess radius must have the hinges replaced within the next 2,000 hours' time in service. Parts not replaced at the end of 2,000 hours' time in service can be kept in service an additional 600 hours providing inspections per (a) are made each 200 hours' time in service. \n\n\t(c) All replacement hinges must have at least a 1-inch door skin recess radius and hinges having a 3/32-inch radius may not be reworked to the 1-inch radius. \n\n\t(d) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\n\t(e) Aircraft modified for unpressurized operations by a modification approved by the Chief, Aircraft Engineering Division, FAA Western Region, and in which the initial inspections required by paragraphs (a), (b), (c) or (d) of this AD have been accomplished, may substitute the following in lieu of the use of a dye penetrant in accomplishing the inspections required by paragraphs (a), (b) (c), or (d): \n\n\t\t1. At intervals not to exceed 200 hours' time in service, perform a visual inspection of all emergency exit door hinges and door skin recesses for cracks in accordance with an FAA approved maintenance program. \n\n\t\t2. This adjusted inspection requirement does not void the requirement to replace the hinges in accordance with paragraph (b). \n\n\t\t3. Restoration of the cabin pressurization system to operational status negates the relief provided by paragraph (e). \n\n\t(f) Equivalent inspection procedures and repairs may be used when approved by the Chief, Aircraft Engineering Division,FAA Western Region. \n\n\t(g) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes without differential pressure, to a base for the accomplishment of inspections required by this AD. \n\n\t(Douglas Service Bulletins DC-6 No. 848; second reissue dated January 4, 1961, and DC-7 No. 393 dated January 3, 1961, cover the above subject.) \n\n\tAmendment 259 supersedes AD 54-25-01. \n\n\tThis directive was effective March 15, 1961. \n\n\tRevised January 18, 1962. \n\n\tThis amendment 39-3354 becomes effective December 1, 1978.
79-16-04: 79-16-04 SL INDUSTRIES (CALLAIR): Amendment 39-3523. Applies to Models B-1 and B-1A Airplanes, all Serial Numbers, certificated in all categories with over 500 hours' time in service. Airworthiness Docket No. 79-ASW-4. Compliance required within the next 50 hours after the effective date of this AD, unless already accomplished within the last 50 hours' time in service and thereafter at intervals not to exceed 100 hours' time in service from the last inspection. To prevent failure of the wing rear spar at the wing-fuselage attachment fitting (P/N 160049) due to cracking, accomplish the following: (a) Inspect both rear wing spars by the following procedures: (1) Remove the wing-fuselage fairings covering the rear spar attach bolt. (2) Hoist the wing and jack the landing gear to relieve the load on the rear spar attach bolt and remove the rear spar attach bolt. (3) Remove the P/N 16049-1 and -2 spar fittings by removal of (6) AN 4 bolts attaching fittings to the spar web. (4) Visually and with dye penetrant methods, inspect the spar web in the area of the P/N 16049-1 and -2 fittings for cracks. Give particular attention to the upper and lower fasteners in the pattern of (6) fasteners attaching the P/N 16049-1 and -2 fittings to the spar web. Inspect these holes for elongation beyond .254 inches. (5) Visually inspect the inboard 24 inches of the spar web for corrosion. (b) Repair any discrepancies found during the inspections of paragraph (a) as follows: (1) For cracks or elongated fastener holes in the spar web, replace the damaged spar with a new spar. (2) For corrosion on the spar web, repair as follows: (i) Remove corrosion in accordance with AC 43.13-1A dated 1972. (ii) If corrosion is less than 10 percent of spar web thickness locally and reduces the spar web cross-section area by less than 2 percent, treat the corroded spar web area with two coats of zinc chromate primer in accordance with MIL-P-6889 or MIL-P-8585A. (iii) If corrosion exceeds 10 percent of local spar web thickness or 2 percent of spar web cross-section area, replace the spar or repair with a repair approved by the Chief, Engineering and Manufacturing Branch, FAA, Southwest Region. (iv) Reinstall the 16049-1 and -2 spar fittings with the fasteners removed as specified in paragraph (a)(3). (v) Reinstall the wing fairings removed during paragraph (a) inspections. NOTE: (SL Industries Service Bulletin A-25 covers this same subject.) (c) Modification of the rear spar by installation of the new wing-fuselage attach fittings of SL Industries Service Bulletin A-25A (or equivalent modifications approved by the Chief, Engineering and Manufacturing Branch, FAA, Southwest Region) will remove the requirement for the 100-hour repetitive inspections. (d) Aircraft may be flown in accordance with FAR 21.197 to a base where the requirements of this AD can be accomplished.This amendment becomes effective September 7, 1979.
79-19-07 R1: 79-19-07 R1 COSTRUZIONI AERONAUTICHE GIOVANNI AGUSTA: Amendment 39-3559 as amended by Amendment 39-4430. Applies to Model A109A helicopters equipped with 90 degree gearbox attach sleeve, P/N 109-0435-29-3, certificated in all categories. Compliance is required as indicated, unless already accomplished. To prevent possible tail rotor gearbox mounting failure, within the next 50 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 100 hours' time in service since the previous inspection, accomplish the following in accordance with Agusta Bollettino Tecnico No. 109-10, Revision A, dated December 17, 1980 (hereinafter referred to as the service bulletin), or an FAA-approved equivalent. (a) Inspect the tail rotor gear box attach nutplates for condition and security. Replace any defective nutplates that are found. (b) Inspect the tail rotor gear box attach sleeve and shim, P/N 109-0435-29-3 and P/N 109-0372-18- 5, respectively, for cracks, fretting, nicks, and wear. (c) If a cracked sleeve or shim is found during the inspection required by paragraph (b) of this AD, replace the defective part with a serviceable part of the same part number or, in the case of sleeves, by P/N 109- 0435-29-5 or P/N 109-0435-31-1. (d) If a sleeve is found to have wear, fretting, or nicks, during the inspection required by paragraph (b) of this AD, and if: (1) They are not more than 0.2 mm (0.008 in.) in depth, remove the defect; and (2) If they are more than 0.2 millimeters (0.008 inch) in depth, replace the sleeve with sleeve of the same part number, or P/N 109-0435-29-5 or P/N 109-0435-31-1. (e) If a shim is found to have wear, fretting, or nicks during the inspection required by paragraph (b) of this AD, replace the shim with a new part of the same number. (f) Inspect and, as necessary, correct the alignment and coaxiality between the sleeve and helicopter tail boom, and the sleeve-to-shim flatness fit, in accordance with paragraph 6 of the Service Bulletin or an FAA- approved equivalent. (g) For purposes of this AD, an FAA-approved equivalent must be approved by the Chief, Aircraft Certification Division, Southwest Region, Federal Aviation Administration, P.O. Box 1689, Fort Worth, Texas 76101, or by the Chief, Aircraft Certification Staff, Federal Aviation Administration, Europe, Africa, and Middle East Office, c/o American Embassy, Brussels, Belgium. Amendment 39-3559 became effective October 10, 1979. This Amendment 39-4430 becomes effective August 16, 1982.
2017-23-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-200, -200C, -300, -400, and -500 series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the fuselage crown skin panels are subject to widespread fatigue damage (WFD). This AD requires repetitive inspections, replacement, and applicable on-condition actions for certain fuselage crown skin panels. We are issuing this AD to address the unsafe condition on these products.
61-03-01: 61-03-01 BELL: Amdt. 249 Part 507 Federal Register February 7, 1961. Applies to All Model 47J-2 Helicopters Except Those Equipped With Lycoming VO-540-B1B3 Engines (Third Order Torsional Dampers). Compliance required as indicated. To preclude failures of the pinion gear on the inboard end of the P/N 47-620-539-1 cooling fan drive assembly which are associated with a poor tooth mesh condition, the following inspections and corrective action must be accomplished no later than the next 10 hours of flight time after the effective date of this directive, and by the completion of the 25 hour, 50 hour, 100 hour, and 300 hour flight time periods subsequent to the accomplishment of the initial inspection. Inspections accomplished in accordance with Bell Mandatory Service Bulletin 130SB prior to the effective date of this directive need not be repeated and only those remaining inspections necessary to complete the series are required. (a) Remove cooling fan drive assembly, P/N 47-620-539-1, from transmission in accordance with instructions contained in Bell Model 47J-2 Maintenance and Overhaul Instructions, and inspect driven side of pinion gear teeth (P/N 47-620-530-1). (b) Replace gear if conditions defined by Figure 2 of Bell Service Bulletin 130SB are found. (c) If replacement of the pinion gear is required, further disassembly as necessary and inspection of the cooling fan driving gear (P/N 47-620-207-1) must be conducted. If conditions defined by Figure 4 of Bell Service Bulletin 130SB are found, this gear must also be replaced. (d) If replacement of either gear is required, back lash and gear pattern must be established in accordance with instructions contained in Maintenance and Overhaul Instructions. (e) Reassemble cooling fan drive assembly in accordance with Maintenance and Overhaul Instructions except that nut, P/N 47-620-565-1 shall be torqued to 960 inch-pounds and fan pulley bolt AN 6H5A shall be torqued to 300 inch-pounds. (f) If either or both gears are replaced, perform the above inspections at 10, 25, 50, 100 and 300 hours of flight time following gear replacement. This procedure must be reconducted until a satisfactory wear pattern on the pinion gear and drive gear is maintained through 300 hours of flight time, after which inspections in accordance with this directive may be discontinued. (Bell Mandatory Service Bulletin 130SB covers this same subject.) This directive effective February 22, 1961. Revised May 13, 1961. Revised November 5, 1966.
2017-22-11: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model CL-600-2B16 (CL-604 Variant) airplanes. This AD was prompted by reports of in-service incidents regarding the loss of all air data system information provided to the flightcrew. This AD requires revising the airplane flight manual (AFM) to provide ``Unreliable Airspeed'' procedures to the flightcrew to stabilize the airplane's airspeed and attitude for continued safe flight and landing. We are issuing this AD to address the unsafe condition on these products.
60-21-01: 60-21-01 ALLISON: Amdt. 208 Part 507 Federal Register October 8, 1960. Applies to All Model 501-D13 Series Engines. Compliance required at next overhaul of engine, power section or torquemeter, whichever occurs first, after the effective date of this AD. Several cases of rubbing of the torquemeter housing by the torquemeter reference shaft have resulted in complete separation of the housing into two sections. To preclude such failures, a mid-bearing torquemeter assembly, P/N 6823900, identified by a 1/2-inch by 2 1/2-inch blue stripe on the forward bevel of the housing shall be installed. (Allison Commercial Engine Bulletin No. 72-113 covers the same subject.) This directive effective November 9, 1960.
81-14-05: 81-14-05 KAWASAKI HEAVY INDUSTRIES, LTD.: Amendment 39-4149. Applies to Model KV 107-II and KV 107-IIA helicopters, certificated in all categories. Compliance required as indicated, unless already accomplished. To prevent fatigue failure of the rotor pitch housing, accomplish the following: (a) For aft rotor pitch bearing housings, P/N's 107R2553-8, -10, -14, and -16: (1) Prior to the accumulation of 1,000 hours time in service, or within the next 50 hours time in service after the effective date of this AD, whichever occurs later, install crack detector wires on pitch bearing housings in accordance with Part I, "Installation Procedure," of Kawasaki Service Bulletin No. KSB-V107-615, dated September 10, 1980 (hereinafter referred to as the Service Bulletin), or an FAA-approved equivalent. (2) Within the next 50 hours time in service after the effective date of this AD, inspect the lug area of the pitch bearing housings for cracks in accordance with PartII, "Inspection Procedure," of the Service Bulletin, or an FAA-approved equivalent, and continue to inspect at intervals not to exceed 25 hours time in service. (b) For forward rotor pitch bearing housings, P/N's 107R2553-7, -9, -13, and -15: (1) Prior to the accumulation of 2,000 hours time in service, or within the next 100 hours time in service after the effective date of this AD, whichever occurs later, install crack detector wires on pitch bearing housings in accordance with Part I, "Installation Procedure," of the Service Bulletin, or an FAA-approved equivalent. (2) Within the next 100 hours time in service after the effective date of this AD, inspect the lug area of the pitch bearing housings for cracks in accordance with Part II, "Inspection Procedure," of the Service Bulletin, or an FAA-approved equivalent, and continue to inspect at intervals not to exceed 50 hours time in service. (c) Conduct a visual inspection for cracks in the lug area of blade sockets, P/N's 42R1043-11, -12, -13, and -14, at intervals not to exceed 50 hours time in service. This may be accomplished without disassembly from the helicopter. (d) If any cracks are found as a result of the inspections required by paragraphs (a), (b), and (c) of this AD, before further flight, replace with a serviceable part of the same part number, or an FAA-approved equivalent, and continue to inspect in accordance with this AD. (e) Retire from service all rotor pitch bearing housings, P/Ns 107R2553-7, -8, -9, -10, -13, -14, -15, and -16, prior to the accumulation of 5,000 hours time in service. (f) If an equivalent means of compliance is used in complying with this AD, that equivalent means must be approved by the Chief, Airworthiness District Office, FAA, Pacific-Asia Region, Honolulu, Hawaii. (g) Upon request of an operator and submission of substantiating data, the Chief, Airworthiness District Office, Pacific-Asia Region, may upon recommendation of the cognizant FAA aviation safety inspector adjust the compliance time specified in this AD. 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 Kawasaki Heavy Industries, Ltd., Aircraft Division, Kawasakicho, Kakamigahara, Gifu Prefecture, Japan. The documents may also be examined at the FAA, Pacific-Asia Region, Airworthiness District Office, 300 Ala Moana Blvd., Room 7321, Honolulu, Hawaii 96850, and Rules Docket, Room 916, 800 Independence Avenue, SW., Washington, DC 20591. This amendment becomes effective July 9, 1981.
60-05-03: 60-05-03 PIPER: Amdt. 108 Part 507 Federal Register March 2, 1960. Applies to PA-22 "150" and PA-22 "160" Aircraft Serial Numbers 22-3218, 22-3387, 22-3388 to 22-7049 Inclusive, and 22-7054. Compliance required by April 1, 1960. Install safety belt extension, P/N 14920-02 or equivalent, on the front seat belt in order to eliminate deterioration due to heat from the rear seat heater outlet and chafing where the web attaches to the attaching lug. P/N 14920-02 has the same geometric design as P/N 14920-0 being replaced, except that P/N 14920-02 is one inch longer measuring 3.5 inches between bolt centerlines. (Piper Service Bulletin No. 184 covers this same subject.) This supersedes AD 57-17-02. Revised November 2, 1960.
2017-22-02: We are adopting a new airworthiness directive (AD) for certain Ipeco Holdings Ltd. (Ipeco) pilot and co-pilot seats. This AD requires modification and reidentification of the affected seats. This AD was prompted by reports of unexpected movement of pilot and co-pilot seats on takeoff and landing. We are issuing this AD to address the unsafe condition on these products.
85-11-03: 85-11-03 BOEING: Amendment 39-5067. Applies to all Model 757 series airplanes equipped with Air Cruisers evacuation slides, part numbers (P/N) as specified in Boeing Service Bulletin 757-25-0040 dated December 21, 1984. To assure slides do not become unsafe due to porosity, accomplish the following, unless already accomplished. \n\n\tA.\tAccomplish inspection procedures in accordance with the service bulletin or later FAA-approved revisions, as follows: \n\n\t\t1.\tFor slides manufactured prior to six months before the effective date of this AD, accomplish the inspection within the next 12 months. \n\n\t\t2.\tFor all other slides, accomplish the inspection within 18 months after the date of manufacture.\n \n\t\t3.\tSlides which do not meet the limitations set forth in the service bulletin must be replaced with a serviceable slide prior to further flight. \n\n\tB.\tRepeat the inspection procedures of paragraph A., above, within 18 months after the last leak check inspection performed in accordance with the service bulletin, or later FAA- approved revisions. \n\n\tC.\tAlternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office. \n\n\tD.\tUpon request of an operator, an FAA Principal Maintenance Inspector, subject to prior approval by the Manager, Seattle Aircraft Certification Office, may adjust the compliance times if the request contains substantiating data to justify the request. \n\n\tE.\tAircraft may be ferried to a base for maintenance in accordance with Sections 21.197 and 21.199 of the Federal Aviation Regulations. \n\n\tThis amendment becomes effective June 28, 1985.
2017-22-05: We are superseding Airworthiness Directive (AD) 2013-15-03 for Eurocopter France Model AS350B, AS350BA, AS350B1, AS350B2, AS350B3, AS350C, AS350D, and AS350D1 helicopters. AD 2013-15-03 required inspecting the hydraulic pump drive pulley bearing (bearing) for leaks, rust, overheating, and condition. This new AD adds a requirement to grease the bearing and inspect for bronze particles in the grease, and changes the inspection and inspection intervals of the bearing until it is replaced with an improved bearing. This AD was prompted by additional reports of hydraulic pump drive belt failure caused by bearing seizures. The actions of this AD are intended to prevent an unsafe condition on these products.
2017-22-04: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-200, -200C, -300, -400, and -500 series airplanes. This AD was prompted by reports of skin doublers that disbonded from their skin panels. This AD requires repetitive inspections of fuselage skin panels, and applicable on-condition actions. We are issuing this AD to address the unsafe condition on these products.
60-20-01: 60-20-01 AERO COMMANDER: Amdt. 204 Part 507 Federal Register September 28, 1960. Applies to Models 680-E and 720, Serial Numbers 501, 623 Through 873 Except 820, 850, 860, 867 and 872. Compliance required within the next 100 hours' time in service after the effective date of this amendment. The manufacturer's inspection has determined a nonconformity with the approved design data and it is possible that aircraft in service may have the following nonconformity: AN 426AD-5 rivets have been installed instead of 3/16-inch huckbolts in the lower surface of the wing at the rear spar between wing station 54 and the inboard nacelle attach angle on both the left and right wings. (a) Inspection. Inspect the lower wing at rear spar between wing station 54 and the inboard nacelle attach angles on both the left and right wings to determine whether 3/16-inch huckbolts or AN 426AD-5 rivets have been installed. If the AN 426AD-5 rivets are installed, the wing shall be reworked as outlined in paragraph (b). (b) Rework. Remove flaps and wing trailing edge closeout skins on both left and right wings. Drill out the AN 426AD-5 rivets and replace with AN 426AD-6 rivets. These rivet heads will protrude below the wing surface by approximately 0.030 inch. Do not overdrive the rivets in an attempt to sink them completely. Measure the distance between the rivet which passes through the wing skin and rear spar cap at wing station 54 and the screw which passes through the inboard nacelle attach angle. This distance should be approximately 4.5 inches and should contain six rivets (0.75 inch on center) and the screw. If only five rivets exists in this area, a brazier head rivet (AN 456AD-6) must be added between the nacelle attach angle screw and the next rivet inboard. If sufficient space does not exist to permit minimum rivet to rivet spacing of three rivet diameters, contact the Service Department, Aero Design & Engineering Company for approved repairinstructions. Replace flaps and left and right wing trailing edge skins. (Aero Design Service Bulletin No. 62 covers this same subject.) This directive effective October 28, 1960.
63-15-06: 63-15-06 PIAGGIO: Amdt. 591 Part 507 Federal Register July 24, 1963. Applies to Model P.166 Aircraft, Serial Numbers 1, through 403. Compliance required within 25 hours' time in service after the effective date of this AD. To preclude failure of the elevator trim tab control system and lever, P/N 5069, because of unsound welding seams, accomplish the following: (a) Inspect right and left levers, P/N's 5069.03 and 5069.02, respectively, for cracks or unsatisfactory machining in accordance with Piaggio Service Bulletin No. 166-30 dated February 1, 1963. (b) If any defects specified in the service bulletin are found, before further flight replace the lever with a lever inspected and found to have no defects. This directive effective August 23, 1963.
2017-21-08: We are adopting a new airworthiness directive (AD) for all Airbus Model A310 series airplanes. This AD was prompted by a revision of certain airworthiness limitation items (ALI) documents, which require more restrictive maintenance requirements and airworthiness limitations. This AD requires revising the maintenance or inspection program to incorporate the maintenance requirements and airworthiness limitations. We are issuing this AD to address the unsafe condition on these products.
2017-21-04: We are adopting a new airworthiness directive (AD) for certain Gulfstream Aerospace LP Model Gulfstream G150 airplanes. This AD was prompted by a report indicating that the main entrance door (MED) opened during flight, and by the determination that the ``CABIN DOOR UNLOCK'' crew alerting system (CAS) message may extinguish before the handle latch pin is fully engaged. This AD requires accomplishing an updated rigging procedure for the adjustment of the MED microswitch. We are issuing this AD to address the unsafe condition on these products.
2017-21-03: We are adopting a new airworthiness directive (AD) for certain Gulfstream Aerospace LP Model Gulfstream 100, Astra SPX, and 1125 Westwind Astra airplanes. This AD was prompted by a report indicating that the main entrance door (MED) opened during flight, and by the determination that the ``CABIN DOOR UNLOCK'' crew alerting system (CAS) message may extinguish before the handle latch pin is fully engaged. This AD requires accomplishing an updated rigging procedure for the adjustment of the MED microswitch. We are issuing this AD to address the unsafe condition on these products.
60-16-01: 60-16-01 BEECH: Amdt. 186 Part 507 Federal Register August 4, 1960. Applies to All Model C45G, C45H, TC45G, and TC45H Airplanes Which Have Been Converted From Military Status to Civil Certification. Compliance required not later than October 1, 1960. The emergency position switch of the electrical turn and bank indicator that bypasses the master switch arrangement contrary to CAR 3.688, must be removed. The live wire connected to the switch must be disconnected at the battery terminal and either removed from the airplane or carefully insulated and secured. Passenger seats (P/N 734-183302) which partially block the emergency exit must be removed, relocated, or reversed to provide a clear and unobstructed opening as required by CAR 3.387. Two configurations of seat P/N 734-183302 were delivered to the military only one of which has been structurally substantiated for aft facing mounting. This seat can be identified by the triangular shaped closed rear leg formed from 2 sheets of 0.040 alal with a long stiffening bead on the outer face of the leg. FAA approval must be obtained for any modification of the seating arrangement, other than removing or reversing (if applicable) the obstructing seat.
61-13-01: 61-13-01 CONVAIR: Amdt. 297 Part 507 Federal Register June 20, 1961. Applies to All Model 22 (880) Aircraft. Instances of fire due to overheating of Bussman 60 and 70 amp type ACO and ACY limiters in the freon compressor and recirculation fan motor circuits have occurred. To preclude fires of this type, the following modifications must be accomplished: Unless already accomplished, compliance with items (a) and (b) is required within the next 130 hours' time in service: (a) Replace the Bussman 60 and 70 amp type ACO and ACY limiters and their holders which are located in the AC power distribution box with type AHB limiters and their holders. (b) Replace the nameplates adjacent to the 60 and 70 amp limiters with similar nameplates made of a fire resistant material such as impregnated fiberglass. (Convair Service Bulletin No. 24-42 covers this same subject.) This directive effective June 20, 1961.