Results
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.)
67-03-03: 67-03-03 HOWARD AND LOCKHEED: Amdt. 39-339 Part 39 Federal Register January 18, 1967. Applies to All Howard Model 500 Series Airplanes and Lockheed Model PV-1 Series Airplanes, Serial Numbers 5275, 5336, 5892, 5492, 5331, 5599, 5272, 5372, 5283, 4385, 4401, 6642, 5373, 5500, 5554, 5560, 5494, 5280, 5700, 5598, 5696, 5894, 2020, 5591, 5267, 5702, 5694, 5887, 5891, 5705, 5371, 5489, 5492, 5274, and 5269. Compliance required as indicated. To detect cracked elevator torque tube collars, P/N 5-408380-1, unless already accomplished, within the next 10 hours' time in service after the effective date of this AD, and thereafter at intervals not to exceed 100 hours' time in service from the date of the last inspection, inspect the elevator torque tube collars P/N 5-408380-1 for cracks in the radius using dye penetrant or an FAA-approved equivalent. Before further flight, replace parts found cracked in accordance with Dee Howard Company, San Antonio, Texas, Drawing Number 13-0766-008 or later FAA Engineering and Manufacturing Branch approved equivalent. The inspection provisions of this AD may be discontinued when P/N 5-408380-1 is replaced in accordance with this AD. This directive effective January 23, 1967.
56-16-03: 56-16-03 SIKORSKY: Applies to All Model S-55 Helicopters With S14-10-2000-1, -2, - 3, -4, and -5 Main Rotor Blades Bearing Serial Numbers 100-3700 Inclusive, and All Model S-51 Helicopters With S10-10-2065 Series Main Rotor Blades Bearing Serial Numbers 100-3700 Inclusive. Compliance required as indicated. Main rotor blades of the S10-10-2065 Series on the S-51 and of the S14-10-2000-1, -2, - 3, -4 and -5 Series on the S-55 have been found with fine cuts on the spar beneath the edges of the pockets. These cuts are believed to be a result of repairs to blade pockets and constitute serious stress raisers which may materially affect the service life of the blades. Inspections are to be accomplished as indicated below. The necessary preparations for inspections, the inspections, the finishing of serviceable blades and the rework of salvable blades are to be in accordance with the procedures recommended in Sikorsky Service Information Circular No. 676 for the S-51 and 1410-669 for the S-55. 1. All blades with less than 500 hours service since manufacture or complete rebuilding prior to May 1953, should be inspected as soon as practicable but prior to the accumulation of 500 hours. 2. All blades with 500 hours or more service since manufacture or complete rebuilding prior to May 1953, should be inspected as soon as practicable but not later than August 15, 1956. 3. Blades found to have no cuts may be returned to service after refinishing. 4. S-55 blades with less than 1,200 hours service and S-51 blades with less than 1,000 hours service and found to have cuts on the top or bottom may be salvaged. 5. Blades found to have cuts on the back are to be returned to the manufacturer for evaluation. 6. S-55 blades with 1,200 hours or more service and S-51 blades with 1,000 hours or more service and found to have cuts on the top, bottom or back are to be returned to the manufacturer for evaluation.
56-16-02: 56-16-02 LOCKHEED: Applies to All Models 49, 149, 649, 749 and 1049 Series Aircraft Equipped With Bendix Eclipse-Pioneer Type PB-10 Series Automatic Pilot Installations. Compliance required as indicated. Numerous instances of fatigue failures have been reported in the side webs of the Bendix Eclipse-Pioneer Type DQ-15 Series servo disconnect mounting brackets in the primary flight control system. Since complete failure could result in disabling the flight control system, the following must be accomplished as indicated: 1. The rudder and elevator servo disconnect mounting bracket installations in the primary flight control system must be inspected by dye penetrant or equivalent method, at 125-hour intervals until item 2 is accomplished. Defective servo mounting brackets must be discarded. 2. To be accomplished as soon as possible but not to exceed next overhaul period after parts become available. The rudder and elevator servo must be reinforced by installation of a secondary servo support bracket to provide additional support and vibration dampening. (Bendix Eclipse-Pioneer Service Bulletin No. 858 and Lockheed Service Bulletins Nos. 49-864 and 1049-2830 also cover this subject.)
69-05-06: 69-05-06 SCHLEICHER: Amdt. 39-733. Applies to Schleicher Model Ka6E gliders, Serial Nos. 1 through 4232 except Serial No. 4226. Compliance required within the next 100 hours' time in service after the effective date of this AD, unless already accomplished. To prevent the failure of the thermos-bottle mounting brackets, install new mounting clamps with rubber insert, in accordance with Schleicher Technical Note No. 17, dated September 10, 1968, or later LBA-approved issue or an FAA-approved equivalent. This amendment becomes effective April 6, 1969.
2000-15-08: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 747 series airplanes, that currently requires repetitive inspections for damage or cracking of the aft pressure bulkhead, and cracking of the bulkhead web-to-Y-ring lap joint area and the upper segment of the bulkhead web. That AD also requires certain follow-on actions, if necessary. This amendment requires that a currently required one-time inspection to detect cracking of the upper segment of the bulkhead web be accomplished repetitively, and adds additional repetitive inspections to detect cracking of the upper and lower segments of the aft bulkhead web. The actions specified by this AD are intended to detect and correct fatigue cracking of the bulkhead web, which could result in rapid depressurization of the airplane, and consequent reduced controllability of the airplane.
94-09-08: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Allied-Signal Inc., Garrett Engine Division, TPE331 series turboprop and Model TSE331-3U turboshaft engines. This action requires inspection of certain third stage turbine stator assemblies, and replacement, if necessary, with serviceable assemblies. This amendment is prompted by reports of six third stage turbine stator assemblies assembled with inner seal supports made of incorrect material that results in a significantly reduced cyclic life. The actions specified by this AD are intended to prevent an uncontained failure of the third stage turbine wheel.
94-19-03: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F28 Mark 0100 series airplanes, that requires replacement of the autopilot disconnect switches with modified units. This amendment is prompted by several incidents in which the flight crew did not depress both halves of the autopilot disconnect switch during the LAND 2 or LAND 3 approach and, as a result, one autopilot remained engaged. This condition resulted in unanticipated movements of the stabilizer trim and higher than anticipated control forces of the flight controls. The actions specified by this AD are intended to prevent the flight crew from inadvertently disconnecting only one autopilot when both autopilots are engaged, which could result in unanticipated control surface movements.
92-12-05: 92-12-05 TEXTRON LYCOMING: Amendment 39-8265. Docket No. 91-ANE-48. Applicability: Textron Lycoming O-320 series, IO-320 series, LIO-320 series, AIO-320 series, AEIO-320 series, O-340 series, O-360 series, LO-360 series, HO-360 series, VO-360 series, IVO-360 series, IO-360 series, AIO-360 series, HIO-360 series, LHIO-360 series, LIO-360 series, AEIO-360 series, TO-360 series, LTO-360 series, TIO-360 series, O-480 series, GSO-480 series, IGSO-480 series, IGO-480 series, GO-480 series, O-540 series (excluding models O-540-J and O-540-L), VO-540 series, IO-540 series (excluding model IO-540-W), HIO-540 series, AEIO-540 series, IGSO-540 series, IGO-540 series, TVO-540 series, TIVO-540 series, IVO-540 series, TIO-540 series, LTIO-540 series, TIO-541 series, TIGO-541 series, and IO-720 series opposed piston engines; and in addition those engine models and serial numbers listed in Textron Lycoming Service Bulletin (SB) No. 501, Revision B, dated November 15, 1991; installed on but notlimited to Cessna 172 and Piper PA-28 aircraft. Compliance: Required as indicated, unless accomplished previously. To prevent piston pin failure, or piston release, and engine failure, accomplish the following: (a) For engines with serial numbers listed in Textron Lycoming SB No. 501, Revision B, dated November 15, 1991, with more than 75 hours time in service (TIS) since new, since remanufacture, or since factory overhaul on the effective date of this AD, remove all piston pins, Part Number (P/N) LW-14077, within 25 hours TIS after the effective date of this AD, and replace with serviceable parts. (b) For engines with serial numbers listed in Textron Lycoming SB No. 501, Revision B, dated November 15, 1991, with 75 hours or less TIS since new, since remanufacture, or since factory overhaul on the effective date of this AD, remove all piston pins, P/N LW-14077, within 100 hours TIS since new, since remanufacture, or since factory overhaul and replace with serviceable parts. (c) For engines not listed in Textron Lycoming SB No. 501, Revision B, dated November 15, 1991, accomplish the following: (1) Within 15 days after the effective date of this AD, conduct a search and review of maintenance and purchase records to determine if piston pin, P/N LW-14077, had been purchased from Textron Lycoming or a Textron Lycoming distributor from June 18, 1991, through August 5, 1991. (2) For installed piston pins, P/N LW-14077, purchased from Textron Lycoming or a Textron Lycoming distributor from June 18, 1991, through August 5, 1991, accomplish the following: (i) For engines with more than 75 hours TIS since piston pin installation on the effective date of this AD, remove all piston pins, P/N LW-14077, purchased from Textron Lycoming or a Textron Lycoming distributor from June 18, 1991, through August 5, 1991, within 25 hours TIS after the effective date of this AD, and replace with serviceable parts. (ii) For engines with 75 hours or less TIS since piston pin installation on the effective date of this AD, remove all piston pins, P/N LW-14077, purchased from Textron Lycoming or a Textron Lycoming distributor from June 18, 1991, through August 5, 1991, within 100 hours TIS since piston pin installation and replace with serviceable parts. (d) Piston pins, P/N LW-14077, purchased from Textron Lycoming or a Textron Lycoming distributor from June 18, 1991, through August 5, 1991, that are not installed in engines are considered unairworthy and shall not be placed in service. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, New York Aircraft Certification Office, Engine and Propeller Directorate. The request should be forwarded through an FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, New York Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this airworthiness directive, if any, may be obtained from the New York Aircraft Certification Office. (f) 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. (g) The identification of certain engines to which this AD is applicable shall be done in accordance with the following Textron Lycoming service document: DOCUMENT NO. PAGES REVISION DATE SB No. 501 1-3 Rev. B Nov. 15, 1991 Total pages: 3 This incorporation 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 Textron Lycoming, 652 Oliver Street, Williamsport, PA 17701. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, Room 311, 12 New England Executive Park, Burlington, Massachusetts; or at the Office of the Federal Register, 1100 L Street, NW., Room 8401, Washington, D.C. (h) This amendment becomes effective on July 10, 1992.
95-07-06: This amendment adopts a new airworthiness directive (AD), applicable to all British Aerospace Model BAC 1-11-200 and -400 series airplanes, that requires inspections of the bearings of the aileron control system, and correction of discrepancies. This amendment is prompted by a report indicating that an operator experienced difficulties wherein considerable pressure was required to manually input roll control due to seized bearings in the aileron control system. The actions specified by this AD are intended to prevent such seizure of bearings, which could reduce the pilot's ability to initiate roll control during critical phases of flight.