Results
70-12-03: 70-12-03 FAIRCHILD-HILLER: Amdt. 39-996. Applies to F-27 and FH-227 type airplanes certificated in all categories. To assure that the outboard flaps are contained in the event of over-travel, by the addition of positive stops to the screwjacks, accomplish the following within the next 250 hours in service after the effective date of this AD, unless already accomplished. (a) Comply with the applicable Fairchild Hiller Service Bulletin, No. F-27-27-72 dated January 16, 1970, or No. FH-227-27-30 dated January 16, 1970, or later revision or equivalent method both approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. (b) Upon request with substantiating data submitted through an FAA Maintenance Inspector, the compliance time specified in this AD may be increased by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. This amendment is effective June 19, 1970.
75-03-04: 75-03-04 FAIRCHILD (HILLER): Amendment 39-2071 as amended by Amendment 39-2251. Applies to model 1100 and FH1100 type helicopters certificated in all categories. To detect cracks in the tail fin spar channel, P/N 24-62030-7 or P/N 24-62030-43 in the area of the tail rotor gear box mount, P/N 24-62006-3, accomplish the following inspection or an equivalent inspection approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, within the next five hours in service after the effective date of this AD unless already accomplished within the last 95 hours in service and at every 100 hours in service thereafter: 1. Remove tail rotor gear box fairing and fin leading edge cover. 2. Clean the tail fin spar in an area one inch in diameter around the left and right forward attachment bolts (two of ten to which the tail rotor gear box mount is attached to the spar) and the spar surface between these two attachments and forward, for a distance of 1 1/2 inches with metachlor or equivalent grease and oil remover by light scrubbing with a stiff bristle brush. 3. Inspect the cleaned area for cracks with at least a ten power magnifying glass by looking through the front end of the tail rotor gear box mount fitting. 4. If a crack is found, replace with an uncracked fin assembly that has been inspected in accordance with the above procedure or alter fin in accordance with an alteration approved by the Chief, Engineering and Manufacturing Branch, Eastern Region before further flight. Amendment 39-2071 was effective January 27, 1975, and was effective for all recipients of the airmail letter of December 9, 1974, upon receipt. This amendment 39-2251 is effective July 8, 1975.
53-15-02: 53-15-02 LOCKHEED: Applies to All Models 049, 149, 649 and 749 Series Aircraft. Compliance required as indicated. At the first arrival at the main base, unless already accomplished, inspect for cracks in the forward flange of the lower front spar cap at wing Station 326, left and right, with particular reference to spar cap joggle areas using dye penetrant inspection method or equivalent. 1. If crack is found in the forward flange and does not extend into the vertical leg, stop drill the crack unless it terminates in a rivet hole and make permanent repair or install the serviceable repair in accordance with LAC Drawing 325800. When serviceable repair is used, a visual inspection must be conducted at periodic intervals not to exceed 50 hours with dye penetrant inspection or equivalent method to be used at periods not to exceed every 200 hours until incorporation of the reinforcement per LAC Drawing 325667, change A or equivalent. 2. If crack in the front spar flange extends into the vertical leg, remove tank sealant as necessary for skin and web inspection using dye penetrant inspection method or equivalent. Reinforcement per LAC Drawing 325667, change A or equivalent is necessary before resuming commercial operation and normal inspection procedures. 3. If no cracks are found, reinspect using dye penetrant inspection method or equivalent on all aircraft with 10,000 hours or more total flight time, at intervals not to exceed 200 hours, and on all other aircraft at each major airframe inspection period until such time as reinforcement per LAC Drawing 325667, change A or equivalent is accomplished.
73-25-03: 73-25-03 HILLER AVIATION: Amdt. 39-1752. Applies to Hiller UH-12D helicopters certificated in all categories. Compliance required prior to further flight for all UH-12D helicopters which have been converted from the military version (H-23D) before the effective date of this AD, and at the time of conversion for those helicopters which are converted to the UH-12D after the effective date of this AD. To insure safe service life for the finite life components of the Hiller Model UH-12D Helicopters, accomplish the following: Replace the finite life components listed in Hiller Aviation's UH-12D Inspection Guide, Airworthiness Limitations Section, dated November 5, 1973, at the times specified therein with new or serviceable parts. NOTE: A copy of the finite life components list can be obtained from Hiller Aviation, 2075 West Scranton Avenue, Porterville, California, 93257, or from the FAA, Aircraft Engineering Division, P. O. Box 92007, World Way Postal Center, Los Angeles, California 90009. This amendment becomes effective January 10, 1974.
75-09-03: 75-09-03 BOEING: Amendment 39-2175. Applies to all Boeing Model 707-100 and 720 series airplanes noted in Boeing Alert Service Bulletin 3220 certificated in all categories. Compliance required as indicated. \n\tA.\tTo detect wing center section lower skin cracks, within the next 100 flights after the effective date of this AD, for airplanes with 20,000 or more flights, visually inspect the wing center section lower skin for cracks at the air conditioning duct anchor channel just outboard of RBL 12.78 beam on stringer 17 and at the air cycle machine support clip just outboard of LBL 12.78 beam on stringer 18. Particular attention should be given to the wing skin at the forward and aft edges of the two support fittings. \n\tB.\tSkins found cracking are to be repaired in accordance with the Boeing structural repair manual or in a manner approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tC.\tThis directive will be amended to require repetitive inspections when such have been determined. \n\tD.\tFor the purpose of complying with the Airworthiness Directive, subject to acceptance by the assigned FAA maintenance inspector, the number of flights may be determined by dividing each airplane's hours' time in-service by the operator's fleet average time from takeoff to landing for the airplane type. One flight is defined as one takeoff and landing. \n\tThe 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). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. The documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tThis amendment becomes effective April 16, 1975.
75-21-03: 75-21-03 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE (formerly Sud Aviation). Amendment 39-2381. Applies to Alouette III Model SE.3160, SA.316B, SA.316C, and SA.319B helicopters incorporating tail boom P/N 3160S.23.11.000.9 to 3160S.23.11.000.13, inclusive, certificated in all categories. Compliance is required as indicated, unless already accomplished in accordance with Alouette Service Bulletin No. 53.31. To prevent failure of the tail boom, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 50 hours' time in service since the last check, until reinforced in accordance with paragraph (d) of this AD, check the upper section of the tailboom at Frame 8 for loose rivets. (b) If more than one rivet is found loose, comply with paragraph (d) of this AD before further flight, except that the helicopter may be flown in accordance with FAR Sections 21.197 and 21.199 to abase where the work can be performed. (c) If no more than one rivet is found loose, within the next 400 hours' time in service after the effective date of this AD, comply with paragraph (d) of this AD. (d) Where required by paragraph (b) or (c) of this AD, reinforce the tail boom attachment at Frame 8 in accordance with subparagraph 1C of Alouette Service Bulletin No. 53.31, as revised April 29, 1974, or equivalent approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa, and Middle East Region, c/o American Embassy, A.P.O. New York, N.Y. 09667. The checks required by paragraph (a) of this AD may be performed by the pilot. This amendment becomes effective October 20, 1975.
74-22-07: 74-22-07 AVCO LYCOMING: Amendment 39-1997. Applies to all Avco Lycoming T5313A turboshaft engines. Compliance required, unless already accomplished, prior to the accumulation of 100 hours in service after the effective date of this AD. To prevent possible compressor disk or centrifugal impeller assembly failures leading to partial or complete loss of engine power, remove and replace the first, second, third, fourth, and fifth stage compressor rotor disks and centrifugal impeller assemblies in accordance with the table below. Remove P/N 1st Stage Compressor 1-100-700-05 Rotor Disk Assembly Socket Head Cap Screw 1-100-214-06 2nd Stage Compressor 1-100-710-05 Rotor Disk Assembly 1-100-710-09 1-100-710-10 3rd Stage Compressor 1-100-230-06 Rotor Disk Assembly 1-100-230-07 4th Stage Compressor 1-100-240-08 Rotor Disk Assembly 1-100-240-10 1-100-240-15 5th Stage Compressor 1-100-450-09 Rotor Disk Assembly 1st Stage Compressor 1-100-201-05 Rotor Spacer 2nd Stage Compressor 1-100-202-03 Rotor Spacer 3rd Stage Compressor 1-100-203-03 Rotor Spacer 4th Stage Compressor 1-100-204-06 Rotor Spacer 5th Stage Compressor 1-100-416-07 Rotor Spacer Sleeve Seal 1-100-482-01 Centrifugal Compressor 1-100-440-07 Impeller Assembly Compressor Rotor 1-100-207-05 Sleeve Compressor Rear Shaft 1-100-137-04 Install P/N Front Shaft Assembly 1-101-080-01 Socket Head Cap Screw 1-100-506-02 Compressor Rotor 1-101-090-02 Subassembly Centrifugal Compressor 1-100-078-08 Impeller Assembly Compressor Rear Shaft 1-100-501-01 Lock Cup Washer 1-100-148-01 Double Hexagon Extended 1-100-502-02 Washer Head Bolt Upon submission of substantiating data through an FAA Maintenance Inspector, the Chief, Engineering and Manufacturing Branch, FAA, New England Region, may adjust the compliance time. NOTE: (Avco Lycoming Service BulletinNumber 0037 pertains to this subject.) This amendment is effective October 31, 1974.
74-05-02: 74-05-02 CANADAIR: Amdt. 39-1790. Applies to CL-44-D4 and CL-44-J airplanes certificated in all categories. Compliance required as indicated. To prevent cracks in the main landing gear uplock actuator cylinder on the Canadair Models CL-44-D4 and CL-44-J aircraft, resulting in the inability to extend the landing gear, accomplish the following: 1. Prior to accumulation of 3,500 hours time in service on the main landing gear uplock actuator cylinders, modify the aircraft in accordance with Canadair Service Bulletin No. CL-44-D4-381, revised September 28, 1968, observing the shimming requirements of Canadair Service Information Circular No. 317-CL-44-D4, dated August 28, 1964, or in accordance with an equivalent procedure approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. 2. For those aircraft modified as described in paragraph 1, the following apply: a. An uplock actuator cylinder installed in a primary uplock actuator must be replaced prior to the accumulation of 8,000 hours in service in primary uplock actuators. Service time accumulated on uplock cylinders prior to accomplishment of Canadair Service Bulletin No. CL-44-D4-381, must be counted as part of the aforementioned 8,000 hours. b. Cylinders removed from primary uplock actuators including unmodified, single cylinder uplock systems, may be used in emergency uplock actuators provided that: (a) Cylinders are subjected to a dye-penetrant inspection, and are found to be free from cracks. (b) Satisfactory dye-checked cylinders are reidentified to distinguish them from new cylinders. (c) Cylinders do not exceed 16,000 hours total combined service time in primary and emergency systems. c. A new actuator cylinder installed in an emergency uplock system has an unrestricted service life. This AD supersedes AD 65-04-04. This amendment is effective February 26, 1974.
75-05-13: 75-05-13 GENERAL ELECTRIC: Amendment 39-2114 as amended by Amendment 39-2524. Applies to all General Electric CF700-2C, CF700-2D, and CF700-2D-2 turbofan engines having compound turbine and fan blade, Part Number 6002T57P01, with the following serial numbers: 001CY through 9994CY 001DN through 999DN 001CZ through 1038CZ 001DO through 999DO 001DA through 999DA 001DP through 999DP 001DB through 999DB 001DQ through 999DQ 001DC through 999DC 001DR through 999DR 001DE through 999DE 001DS through 999DS 001DF through 999DF 001DT through 999DT 001DG through 999DG 001DU through 999DU 001DH through 999DH 001DV through 999DV 001DW through 999DW 001DJ through 999DJ 001DX through 999DX 001DK through 999DK 001DY through 999DY 001DL through 999DL 001DI through 999DI 001DZ through 999DZ 001DM through 999DM Compliance required as indicated unless already accomplished. To preclude possible failures of compound turbine and fan blades, perform an eddy current inspection as specified in the Accomplishment Instructions section of General Electric Alert Service Bulletin No. (CF700) A72-124, or (CF700) A72-125 or later FAA approved revisions or a fluorescent-penetrant inspection as specified in General Electric Overhaul Manual SEI-133 Section 72-72-1, or equivalent inspection method approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region, in accordance with the following schedule: a. Compound turbine and fan blades with 400 or more hours total time in service on the effective date of this AD: Inspect within the next 30 hours time in service and inspect every 400 hours time in service thereafter. b. Compound turbine and fan blades with less than 400 hours total time in service on the effective date of this AD: Inspect prior to the accumulation of 430 hours total time in service and inspect every 400 hours time in service thereafter . Replace all compound turbine and fan blades, confirmed as cracked, prior to further flight. Upon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, FAA, New England Region, may adjust the inspection interval specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. Amendment 39-2114 was effective March 12, 1975. This amendment 39-2524 becomes effective March 3, 1976.
52-01-05: 52-01-05 BELL: Applies to All Models 47B and 47B3 Helicopters and to Model 47D Helicopters Serial Numbers 1 to 110, Inclusive. Compliance required at next 300-hour overhaul, but not later than March 1, 1952. To provide a tail rotor blade grip retaining bolt of increased service life, replace each existing 47-641-026-1 bolt with the 47-641-052-1 bolt, and torque to 160-190 inch-pounds. Also, replace the 47-641-036-1 flat washer type micarta seal in the blade yoke on Models 47B and 47B3 helicopters with the 47-641-042-1 cup seal type. (Bell Service Bulletin No. 80 covers this same subject.)
78-07-04: 78-07-04 AVCO LYCOMING: Amendment 39-3166. Applies to Avco Lycoming T5508D model turboshaft engines. Compliance is required by May 1, 1978. To prevent fatigue failure of the engine accessory inner bevel gear leading to engine stoppage, remove inner bevel gear P/N 2-070-005-02 and replace with inner bevel gear assembly P/N 2-070-030-01, in accordance with Avco Lycoming Service Bulletin No. 5508-0013, dated March 15, 1978, or later revision approved by the Chief, Engineering and Manufacturing Branch, FAA, New England Region. The manufacturer's service bulletin identified and described in this directive is 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 Avco Lycoming Division, 550 South Main Street, Stratford, Connecticut 06497. These documents may also be examined at Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803, and FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. This amendment becomes effective March 30, 1978.
76-12-10: 76-12-10 BEECH: Amendment 39-2644. Applies to Model 200 (Serial Numbers BB-3 thru BB-41) airplanes. Compliance: Required as indicated, unless already accomplished. To prevent flight into icing conditions with inoperative wing de-ice boots, accomplish the following: A) Within 50 hours' time in service after the effective date of this AD: 1. Remove the access plates on the outboard side of the nacelles, forward of the main spar, and visually inspect the wing de-ice tubes to determine if they are high temperature supply tubes which are identified by their black color. 2. If they are black high temperature supply tubes no further action is required. 3. If they are not black high temperature supply tubes, either install a placard on the instrument panel in full view of the pilot which reads: "DO NOT OPERATE IN ICING CONDITIONS" and operate the aircraft in accordance with this limitation or comply with Paragraph B. B) On or before October 1, 1976,install new P/N 130936N8D1200 black high temperature supply tubes in the right and left hand wings outboard of the engine nacelles. When this has been accomplished the placard and operating instructions required in Paragraph A are no longer applicable. C) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Procedures for inspecting and replacing the wing de-ice boot supply tubes are provided in Beechcraft Service Instructions 0822-193 or later approved revisions. This amendment becomes effective June 23, 1976.
76-12-11: 76-12-11 LAKE AIRCRAFT, DIVISION of CONSOLIDATED AERONAUTICS, INC: Amendment 39-2646. Applies to all Model LA-4-200 airplanes equipped with the AC GF- 416 fuel filter. Compliance required as specified below unless already accomplished. 1. To prevent engine power loss resulting from fuel injector contamination caused by filter housing corrosion, remove the AC GF-416 fuel filter from aircraft, cut open the housing and inspect the inside for corrosion in accordance with the following schedule: a. Inspect the AC GF-416 filters installed in aircraft 60 or more days, within the next 30 days. b. Inspect AC GF-416 filters installed in aircraft less than 60 days, prior to the accumulation of 90 days on aircraft. Thereafter, c. Inspect AC GF-416 replacement filters prior to the accumulation of 90 days on aircraft. If corrosion is observed, inspect the Bendix Fuel Injection Unit in accordance with Paragraphs 2 a, b, and c, prior to further flight. 2. If the aircraft has had at least one fuel filter changed since the installation of a new or overhauled fuel injection unit before the effective date of this AD, perform the following inspection of the Model RSA-5AD1 Bendix Fuel Injection Unit, P/N 252054-4, within the next 30 calendar days: a. Remove and inspect fuel inlet strainer. b. Inspect housing area enclosing strainer. c. Remove and inspect each of the four injector nozzles P/N 2524107, including the shield P/N 2520418 and screen P/N 367687. If corrosion contamination is found in the fuel injector nozzle assembly, strainer or housing area, replace or overhaul the fuel injection unit and the flow divider P/N 2524057-2 prior to further flight. Reference: Lake Aircraft Division Service Bulletin B-57. Equivalent methods of compliance may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803.This amendment becomes effective June 29, 1976.
90-02-04: 90-02-04 BOEING OF CANADA, LTD., DE HAVILLAND DIVISION: Amendment 39- 6467. Docket No. 89-NM-207-AD. Applicability: Model DHC-8 series airplanes, serial numbers 3 through 144 inclusive; certificated in any category; equipped with main landing gear uplock actuator, Part Numbers 10800-103, -105, -107, -109, -111; and which do not have Modification 8/1098 incorporated. Compliance: Required as indicated, unless previously accomplished. To prevent failure of the main landing gear (MLG) to extend, accomplish the following: A. Within 50 hours time-in-service after the effective date of this AD, inspect the MLG uplock installation in accordance with the compliance section of de Havilland Service Bulletin No. 8-32-79, Revision A, dated February 3, 1989. If the sensor bracket is under flush, and there are no spacers fitted, prior to further flight, rework the actuator in accordance with the service bulletin. B. Whenever the proximity sensor and/or the proximity sensor mounting bracket on an uplock actuator has been reset or replaced, prior to further flight, perform the inspection and any necessary repair as required by paragraph A., above. C. Within one year after the effective date of this AD, rework and reidentify all uplock actuators in accordance with DOWTY Alert Service Bulletin (part of de Havilland Service Bulletin No. 8-32-79) paragraph 2.E., "Rework" instructions. After rework is accomplished, the inspections required in paragraphs A. and B., above, may be terminated. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, New York Aircraft Certification Office. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing of Canada, Ltd., de Havilland Division, Garrett Boulevard, Downsview, Ontario M3K 1Y5, Canada. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. This amendment (39-6467, AD 90-02-04) becomes effective on February 12, 1990.
92-15-08: 92-15-08 FOKKER: Amendment 39-8302. Docket No. 92-NM-47-AD. Applicability: Model F28 Mark 0100 series airplanes; serial numbers 11244 through 11308, inclusive; 11310, and 11312 through 11314, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent fatigue damage and subsequent reduced structural capability of the horizontal stabilizer attachment, accomplish the following: (a) Within 7 months after the effective date of this AD, remove the normal maximum (second) detent for the reverse thrust control and install a modified detent, in accordance with Fokker Service Bulletin SBF100-76-008, dated May 8, 1991. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The removal and installment shall be done in accordance with Fokker Service Bulletin SBF100-76-008, dated May 8, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW.,Room 8401, Washington, DC. (e) This amendment becomes effective on September 8, 1992.
77-09-07: 77-09-07 DOWTY ROTOL: Amendment 39-2890. Applies to Dowty Rotol Type (c)R175/4-30-4/13E and (c)R212/4-30-4/22 propellers having hub and driving center, P/N 601023166, 601023223, or 601023227 modified to Dowty Rotol Mod. No. (c)VP2381 standard. These propellers are installed on, but not necessarily limited to, Fokker Model F-27 Mks. 100, 300, and 700 Series and Hawker Siddeley Model 748 Series 2A airplanes. Compliance is required at the next propeller overhaul or within the next 4700 hours propeller time in service after the effective date of this AD, whichever occurs first, unless already accomplished. To prevent cracking and possible failure of the hub driving center due to improper machining, accomplish the following: (a) For Type (c)R175/4-30-4/13E propellers, that incorporate the following hub and driving center assemblies, rework the hub driving center to salvage scheme No. 640144020 in accordance with Paragraph 2A of Dowty Rotol Service Bulletin 61-857, datedFebruary 3, 1975, or an equivalent approved in accordance with Paragraph (c) of this AD: Hub and Driving Center Assembly Serial Numbers 175/58/67 175/58/72 175/58/91 175/58/146 175/59/184 (b) For Type (c)R175/4-30-4/13E propellers that incorporate hub and driving center assemblies having serial numbers not listed in paragraph (a) of this AD and Type (c)R212/4-30- 4/22 propellers - (1) Inspect the hub driving center to verify correct machining in accordance with Paragraph 2B of Dowty Rotol Service Bulletin 61-857 dated February 3, 1975, or an equivalent approved in accordance with paragraph (c) of this AD; and (2) If the hub driving center is found to have been incorrectly machined, rework it to salvage scheme No. 640144020 in accordance with Paragraph 2A of Dowty Rotol Service Bulletin 61-857, dated February 3, 1975, or an equivalent approved in accordance with paragraph (c) of this AD. (c) The equivalent means of compliance specified inparagraphs (a) and (b) of this AD must be approved by the Chief, Aircraft Certification Staff, FAA, Europe, Africa and Middle East Region c/o American Embassy, APO New York, N.Y. 09667. This amendment becomes effective May 19, 1977.
99-06-14: This amendment adopts a new airworthiness directive (AD), applicable to all Dornier Model 328-100 series airplanes, that requires one-time visual inspections of the elevator trim system for paint contamination on the actuator pistons and to determine the moisture level of the moisture indicator; verification of the installation and condition of the gasket of the flex drive; and corrective actions, if necessary. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent failure of the elevator trim system due to paint/moisture contamination, and consequent reduced controllability of the airplane.
59-11-03: 59-11-03 MOONEY: Applies to Mark 20A Aircraft Through Serial Number 1534. Compliance required as indicated. To guard against the possibility of carbon monoxide entering the cabin, the following inspection and/or repair is required: (1) Before next flight, visually inspect the exhaust pipes for cracking around the heater jacket bulkheads. Repair or replace as required. No welding is permitted inside the muff or at heater jacket bulkheads. (2) Within next 25 hours of flight and periodically every 50 hours thereafter, remove the heater muff jacket and inspect the complete exhaust system for cracking. Repair or redlace defective parts as required. The 50-hour inspection is not required if exhaust heater muff assembly P/N 6363 is installed. (Mooney Service Letter 20-38 concerns this subject.)
55-20-03: 55-20-03 MARTIN: Applies to All Models 202, 202A, and 404 Aircraft. Compliance required as indicated. Several cases of nose gear steering shaft failures have occurred at the machined splines, due to torsional fatigue. Accordingly, the following inspections using dye penetrant, magnetic particle or vapor blast, are required to check for the presence of cracks. 1. New type shafts, Menasco P/N 526681, installed on all 202 and 202A aircraft, and on all 404 aircraft incorporating shimmy dampeners, must be inspected every 2,500 hours. On 404 aircraft not incorporating shimmy dampeners, the inspection must be conducted every 1,000 hours. Cracked shafts should be removed from service pending instructions from Martin. 2. Original type shafts, Menasco P/N 511681, which have never cracked, may be continued in service subject to the same conditions and inspections as the new type in item 1 providing the 1.628+0.005-inch relief cut is added. This is accomplished by grinding the serration run out circumferentially to a relief diameter of 1.628 inches starting 5/8 inch from upper shoulder, with 1/16- inch corner radii. Cracked shafts may be ground down to a minimum diameter of 1.530 inches to remove cracks. If cracks are removed, the shaft may be returned to service, but must be reinspected as required in the following paragraph 3. 3. All original type shafts which have been ground to remove cracks must be inspected at 325-hour intervals. Shafts may be ground down to a 1.530-inch minimum diameter to remove cracks. If cracks are removed, the shafts may be returned to service, continuing this inspection. If cracks are not removed at the 1.530 diameter, the shaft must be replaced.
67-29-06: 67-29-06 SIKORSKY: Amdt. 39-502 Part 39 Federal Register November 2, 1967. Applies to all S-51 type Helicopters. Compliance required as indicated unless already accomplished. To preclude the possibility of loss of tail rotor power due to failure of sleeve and tube assembly, P/N S535180, accomplish the following: (a) Immediately upon the effective date of this AD and every 50 aircraft hours' time in service thereafter, perform the following torsional check on sleeve and tube assembly P/N S535180 as follows: 1. Apply rotor brake. 2. Induce a torsional force in both clockwise and counter-clockwise directions to tail rotor drive shafting by hand applied light rotational pressure on tail rotor blades. 3. Inspect shaft ends for evidence of movement between end fittings, rivets and tube. 4. If evidence of movement is observed, replace sleeve and tube assembly P/N S535180 with like new or serviceable part. (b) Following the effective date of this AD daily inspect sleeve and tube assembly P/N S535180 for loose rivets as evidenced by the presence of black fretting residue around rivet heads, or actual movement of the rivets, and the presence of cracks in the areas where the tube and sleeves (adapters) are joined. If cracks are found replace sleeve and tube assembly P/N S535180 with like new or serviceable part. Loose rivets may be replaced by removing the existing rivets, drilling No. 2 (.221) diameter and installing new rivets P/N MS20470-B7-9. Alternate oversize rivets P/N MS20470-B8-10 may be used if holes are drilled to No. F. (.257) diameter. (c) Upon request, with substantiating data submitted through an FAA maintenance inspector, compliance times may be increased by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (Sikorsky Aircraft telegram to all operators dated September 5, 1967 covers this same subject.) This amendment effective October 31, 1967.
55-25-02: 55-25-02 de HAVILLAND: Applies to All Model 104 "Dove" Aircraft. Compliance required as indicated. Cases have occurred recently where cracks have been found in the left-hand front fin attachment brackets, P/N 4FS.1749 (Pre. Dove Model 7, "Individual fin attachment and rudder control pulley bracket"), and P/N 4FS.6781 (Dove Mod. 7, "To introduce single casting for front fin attachments and control pulley brackets"). The cracks generally emanate from the top rivet hole in the left row and pass through the flange. The de Havilland Service strongly recommends inspection of the fin attachment brackets at an early date with which the FAA concurs and considers mandatory. Inspect both front fin attachment brackets for cracks, using a magnifying glass after removing the paint, as soon as practical, but not later than the next 25 hours operation unless already accomplished, and thereafter at each check II (approximately 100-hour periods). Access can be made by enteringthrough bulkhead No. 5. Should any cracks be found, install new front fin attachment and pulley brackets, P/N 4FS.9165 L. H. and 4FS.9166 R. H. (Ref. Dove Modification 903) and secure to bulkhead No. 6 using 2BA bolts and nuts or equivalent in the top six holes. Rivets are used in the other positions. Repetitive inspection may be discontinued when the new front fin attachment and pulley brackets per Modification 903 are installed. (de Havilland Technical News Sheet CT (104) No. 112, Issue 2, dated September 1, 1954, covers this same subject.)
73-09-03: 73-09-03 SOCIETE NATIONALE INDUSTRIELLE AEROSPATIALE: (S.N.I.A.S.) - formerly SUD Aviation). Amendment 39-1626. Applies to Lama Model SA-315B and Alouette Models SA-316B, SA-316C, and SA- 319B helicopters having main gear box P/N's 319A.62.00.000.1 and .2, serial numbers up to and including 2,000, installed. Compliance required as follows, unless already accomplished: (1) For helicopters that have accumulated 500 or more hours' time in service on the main rotor gear box since new or since overhauled, compliance is required before further flight, except that the aircraft may be flown in accordance with FAR 21.197 to a base where the work can be performed. (2) For all other helicopters, compliance is required before the accumulation of 500 hours' time in service on a main rotor gear box since new or since overhauled. To prevent the possible failure of a main rotor gear box because of defective planetary gears, remove and disassemble gear box P/N's 319A.62.00.000.1 or .2 and inspect the planet pinions of the second stage planetary gear, P/N 3160S.62.05.208, for cracks and evidence of overheat, due to grinding, in accordance with the inspection procedures specified in Aerospatiale Service Letter No. 137-01-72, dated November 30, 1972, or an FAA-approved equivalent and - (a) For Lama Model SA-315B helicopters, Aerospatiale Service Bulletin No. 05-02, dated October 27, 1972, as amended on November 30, 1972, or an FAA-approved equivalent; and (b) For Alouette Models SA-316B, SA-316C and SA-319B helicopters, Aerospatiale Service Bulletin No. 05-47, dated October 27, 1972, or an FAA-approved equivalent. (c) If cracks or evidence of overheat are found in the planet pinions of the second stage planetary gear, P/N 3160S.62.05.208, replace with serviceable planet pinions of the same part number. 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 airmail letter dated March 23, 1973, which contained this amendment.
52-28-05: 52-28-05 DOUGLAS: Applies to All Model DC-6 Series Aircraft (Fuselage Numbers 1 to 213 Inclusive). \n\n\tCompliance required as indicated. \n\n\tThe following inspections and rework pertain to the center wing lower surface access hole structure at Station 149. \n\n\t1.\tInspection. \n\n\t\t(a)\tConduct following inspection as soon as practical but not later than the next 50 hours operation unless already accomplished and continuing thereafter at regular periodic inspection intervals nearest to 500 hours from the time of initial inspection until permanent repair is made as outlines in 2(b). Using at least an 8-power magnifying glass and/or dye check method or equivalent, make inspections for cracks in the lower wing skin and doubler at the aft access hole paying particular attention to the corner areas. Alternate inspection procedures which will provide equivalent safety may be approved. If cracks are found, make repairs as indicated in item 2 before the next schedules flight. \n\n\t\t(b)Periodic visual inspection must be continued at the most frequently established inspection period between 15 and 35 flying hours for airplanes reworked, as per item 2(a) until the rework of item 2(b) is accomplished. If a crack is found beyond the stop drill hole prior to the replacement period as indicated in item 2(c) make repair as per item 2(b) before the next scheduled flight. \n\n\t2.\tRepair. \n\n\t\t(a)\tIf cracks are found in either the lower wing skin or doubler less than 1-inch long, stop drill using a 1/4-inch drill or 3/8-inch drill hole if space permits. The combined length of the crack and drill hole should not exceed 1 1/4 inches in total length. \n\n\t\t(b)\tIf cracks are found in either the lower wing skin or doubler greater than 1-inch long, or if the cracks extend under the adjacent angle which cannot be visually inspected, incorporate the rework on Douglas Drawing No. 5400661 before the next scheduled flight. In cases where only one corner of the access hole is cracked, Douglas approved interim repair may be used subject to replacement with permanent rework, per Drawing No. 5400661, within a period not to exceed 1,500 hours from time interim repair is made. \n\n\t\t(c)\tThe rework of item 2(a) must be replaced with the rework reinforcement of item 2(b) within 3,000 flying hours from time rework of item 2(a) is accomplished. \n\n\t(Douglas Service Letter No. 130, dated July 10, 1952, also covers this same subject.)
66-28-05: 66-28-05 PILATUS: Amdt. 39-306, Part 39, Federal Register November 9, 1966. Applies to Model PC-6 Series Airplanes. Compliance required as indicated. To prevent failure of the rudder pedal support, accomplish the following: (a) Within the next 50 hours' time in service 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, until modified in accordance with (b)(3), visually inspect the guide tube welding seams of rudder pedal support, P/N 6232.196 for cracks, using a lamp and mirror. (b) If a crack is found during an inspection required by (a), before further flight, accomplish one of the following or an FAA-approved equivalent - (1) Repair the part in an FAA-approved manner; (2) Replace the part with an unmodified part of the same part number; or, (3) Replace the part with one modified or one repaired and reinforcedin accordance with Swiss Office Federal de l'Air-approved Pilatus Service Bulletin No. 65. This directive effective November 19, 1966.
57-12-04: 57-12-04 BELL: Part A Below Applies to the Following Model 47 Helicopters Having Metal Tail Rotor Blades: 47B, 47B3, 47D, 47D1, 47G, 47G2, 47H1, and 47J. Part B Below Applies to the Following 47 Helicopters Having Metal Tail Rotor Blades P/N 47-642-102-5; 47B, 47B3, 47D, 47D1, 47G, 47G2, and 47H1. Compliance required as soon as possible but not later than August 1, 1957. Part A. Due to the possibility of excessive play in the metal tail rotor blade and hub assembly and the pitch control mechanism which can result in blade flutter, the inspection as required in Part A of Bell Mandatory Service Bulletin No. 121SB, dated April 2, 1957, must be accomplished. Part B. Metal tail rotor blades, P/N 47-642-102-5 should be inspected for proper thickness at blade station 14.00. This thickness should be a minimum of 0.750 inch at the thickest part of the blade. Blades measuring less than 0.750 inch are required to be removed and replaced with acceptable blades. (Part B ofBell Mandatory Service Bulletin No. 121SB dated April 2, 1957, covers this same subject.)