Results
64-28-02: 64-28-02 HUGHES: Amendment 39-617. Applies to Models 269A, 269A-1, 269A-2 and 269B helicopters. Compliance required as indicated. Due to inadequate lubrication of the main rotor blade flapping hinge bearings, binding has occurred, resulting in overstressing and failure of one main rotor blade. To prevent main rotor blade overstressing and failure, accomplish the following: (a) For helicopters having less than 100 hours' time in service on the effective date of this AD inspect the main rotor blade flapping hinge bearings for the condition indicated in (d), prior to the accumulation of 110 hours' time in service, unless already accomplished, and thereafter at periods not to exceed 400 hours' time in service from the last inspection. (b) For helicopters having 100 hours' or more time in service on the effective date of this AD inspect the main rotor blade flapping hinge bearings for the condition indicated in (d), within the next 10 hours' time in service, unless already accomplished within the last 390 hours' time in service, and thereafter at periods not to exceed 400 hours' time in service from the last inspection. (c) Within 10 hours' time in service, after the effective date of this AD, unless already accomplished within the last 15 hours' time in service, and thereafter at periods not to exceed 25 hours' time in service, apply MIL-G-25537 grease through the grease fittings to all main rotor blade flapping hinge bearings and inspect to ascertain that a thorough purging of the bearings has been achieved. If new grease does not exude from all six bearings, inspect the bearings at the location where the grease does not exude, for the condition indicated in (d), before further flight. (d) If bearings are found to be worn or damaged in excess of the limits specified in Hughes Service Information Notice No. 2A-39.1, 2A-1-06.1 or 2B-07-.1, remove from service before further flight. (e) If the inner race of any bearing is found to be brinelled or worn in excess of the 0.002 inch limit specified in Hughes Service Information Notice No. 2A-39.1, 2A-1-06.1 or 2B- 07.1, remove the corresponding main rotor blade from service before further flight, and conspicuously mark it to prevent inadvertent return to service. (f) If the inner race of any bearing is brinelled or worn, but not in excess of the 0.002 inch limit specified in Hughes Service Information Notice No. 2A-39.1, 2A-1-06.1 or 2B- 07.1, inspect the corresponding main rotor blade for skin cracks in the exposed areas adjacent to the edges of the blade root fitting beside the outboard bolt. Inspect both upper and lower surfaces of the blade, using a 4- to 6-power magnifying glass. If any cracks are found remove the blade from service before further flight. (g) Starting with the effective date of this AD, conduct initial and repetitive inspections of all main rotor blades in a manner and at periods specified in Hughes Service Information NoticesNo. 2A-38 dated September 9, 1964, No. 2A-1-05 dated September 14, 1964, and No. 2B-06 dated September 16, 1964, or later revisions approved by FAA Western Region Aircraft Engineering Division. (h) The following Hughes Service Information Notices or later FAA approved revisions are approved alternatives to those specified in (d), (e), (f) and (g) above: Notice No. N-10.1 dated December 14, 1967, is an approved alternate to Notices 2A-38, 2A-1-05 and 2B-06. Notice No. N-47 dated April 18, 1968, is an approved alternate to Notices 2A-39.1, 2A- 1-06.1 and 2B-07-.1. This amendment becomes effective on July 6, 1968.
2017-20-13: We are adopting a new airworthiness directive (AD) for PIAGGIO AERO INDUSTRIES S.p.A. Model P-180 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as disbonding of the upper and lower metal skin from the honeycomb core on the elevator assembly and other flight control surfaces. We are issuing this AD to require actions to address the unsafe condition on these products.
60-13-01: 60-13-01 DOUGLAS: Amdt. 173 Part 507 Federal Register June 24, 1960. Applies to All DC-8 Series Aircraft. \n\n\tCompliance required as indicated. \n\n\tSeveral cases of fuel dump chute oscillation have occurred during extension and retraction of the chutes. This has caused cracking of the chutes which may create a hazardous condition by allowing fuel impingement on the wing and flap when dumping fuel. Oscillation of dump chutes occurs at speeds in excess of 200 knots indicated airspeed. To prevent possible fuel dump chute cracking, the following action is required. \n\n\t(a) No later than 10 days following publication of this airworthiness directive in the Federal Register, the maximum speed for fuel dump chute operation and fuel dumping is restricted to not more than 200 knots indicated airspeed and the following placard shall be posted in full view of the pilot: \n\n\t"Do not exceed 200 KIAS when operating dump chute and dumping fuel." \n\n\tThe limitations section of the FAA approved airplane flight manual is hereby amended to incorporate this limitation. \n\n\t(b) Not later than the next periodic inspection ascertain that the dump cable control system is properly rigged in accordance with Douglas drawing No. 7651290 BS change. \n\n\t(c) The 200 knots indicated airspeed restriction may be removed from those aircraft which have been reworked in accordance with Douglas Service Bulletin 28-16, "Fuel-Fuel Dump System and Controls". \n\n\t(Douglas telegraphic alert Service Bulletin No. A28-16 covers the above speed limitation.) \n\n\tRevised January 25, 1961.
60-25-02: 60-25-02 LOCKHEED: Amdt. 229 Part 507 Federal Register December 2, 1960. Applies to All Model 188 Series Aircraft. Compliance required as indicated. As a result of reported cracked aileron counterweight attaching angles, the following is required: (a) At intervals not to exceed 300 hours' time in service, inspect for cracks in the angles which attach the aileron balance-weight brackets to the aileron spar at aileron Stations 13.85, 30.95, 48.01, 82.16, 116.31, 150.46 and 184.61. Cracked angles must be replaced before further flight. (b) The replacement of angles found to be cracked must be made in accordance with either (1) or (2). (1) Original angles may be replaced with parts having the same dimensions as the original parts but fabricated from 0.090-inch thickness AISI 8630 or AISI 4130 steel heat- treated to 150,000-170,000 p.s.i. ultimate tensile strength, cadmium plated and baked. Lockheed Process Specifications Nos. 522, 491, and 170 or FAA approved equivalent must be used for the respective processes indicated. (2) Original angles may be replaced in accordance with the instructions in the kits outlined in Lockheed Service Bulletin 88/SB-451 or FAA approved equivalent. This service bulletin also contains detailed instructions for accomplishing this modification. (c) After replacement of the original parts in accordance with the instructions in (b), subsequent inspections of replaced parts may then be made at normal inspection periods. This supersedes AD 60-11-04. This directive effective December 2, 1960.
80-03-06: 80-03-06 GENERAL ELECTRIC: Amendment 39-3683. Applies to General Electric CF6-6 series engines, Stage 2 HPT disks, Part Numbers 9083M48P01, 9084M52P02, 9084M52P05, 9137M15P01, and 9137M15P02. Compliance is required as indicated. To preclude disk failure due to fatigue, these disks are to be removed from service as follows: (a) All disks, except those noted in (c) below, which have more than 8400 cycles on the effective date of this AD are to be removed from service within 100 cycles. (b) All disks, except those noted in (c) below, which have less than 8400 cycles are to be removed before they accrue 8500 cycles. (c) Those disks, P/N 9137M15P02, with Serial Numbers MP036073, MP030786, MP030797, MP035005, and RPO24722 because of their refurbishment schedule shall be removed from service before they accrue 8800 cycles. (d) Other disks than those identified in (c) above may be allowed additional cycles on an individual basis when approved by the Chief, Engineering and Manufacturing Branch, FAA, Great Lakes Region. CF6-6 Service Bulletin 72-761 dated November 28, 1979 also applies to this subject. This amendment becomes effective on February 8, 1980.
64-14-05: 64-14-05 LOCKHEED: Amdt. 738 Part 507 Federal Register May 26, 1964. Applies to All Models 188A and 188C Series Aircraft on Which The Main Landing Gear Actuator Support Fitting, Lockheed P/N 800618, Has Accumulated More Than 5,000 Hours' Time in Service. Compliance required as indicated. To eliminate the possibility of jamming the main landing gear, accomplish the following: (a) For operators maintaining records of landings, within 225 landings after the effective date of this AD unless already accomplished within the last 775 landings, and thereafter at intervals not to exceed 1,000 landings, comply with (c). If a fitting is replaced in accordance with (d), the next inspection in accordance with (c) shall be accomplished within 5,225 landings from that time and at periodic intervals thereafter not to exceed 1,000 landings. Where past records of landings are unavailable, the number of landings prior to the effective date of this AD may be obtained by substituting one landing for each hour of time in service prior to the effective date of this AD. (b) For operators not maintaining records of landings, within 225 hours' time in service after the effective date of this AD, unless already accomplished within the last 775 hours' time in service and thereafter at intervals not to exceed 1,000 hours' time in service, comply with (c). If a fitting is replaced in accordance with (d), the next inspection in accordance with (c) shall be accomplished within 5,225 hours' time in service from that time and at periodic intervals thereafter not to exceed 1,000 hours' time in service. (c) Inspect the aft end of the main landing gear actuator support fittings for fatigue cracks in accordance with Lockheed 88 Alert Service Bulletin No. 599A, Section 2.A through 2.C or an FAA approved equivalent. (d) Replace cracked actuator support fittings detected during the accomplishment of (c) before further flight in accordance with Section 2.E(1) through 2.E(10) of Lockheed 88 Alert Service Bulletin No. 599A and Section 2 of Lockheed Service Bulletin No. 538 or equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region. (e) Spar cap cracks and plank cracks detected during compliance with Section 2.E(4) of 88/SB-599A shall be repaired in accordance with the Lockheed Structural Repair Manual or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight, except that the aircraft may be ferried in accordance with the provisions of CAR 1.76 to the base at which the repairs are to be accomplished. (f) The periodic inspections described in (a) and (b) may be discontinued when P/N 800618-3 is replaced by P/N 841275-101 and P/N 800618-4 is replaced by P/N 841275-102. (g) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, 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. (Lockheed Alert Service Bulletin Nos. 88/SB-599A and 88/SB-538 cover this same subject.) This directive effective June 26, 1964.
2017-20-12: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 737-100, -200, and -200C series airplanes. This AD was prompted by a report of incidents involving fatigue cracking in transport category airplanes that are approaching or have exceeded their design service objective and a structural reevaluation that was conducted by the manufacturer. This AD requires revising the maintenance or inspection program, as applicable, to add supplemental inspections. This AD also requires inspections to detect cracks in each structural significant item (SSI), and repair of any cracked structure. We are issuing this AD to address the unsafe condition on these products.
2017-20-04: We are adopting a new airworthiness directive (AD) for certain Airbus Model A300 B4-600, B4-600R, and F4-600R series airplanes, Model A300 C4-605R Variant F airplanes (collectively called Model A300-600 series airplanes), and Model A310 series airplanes. This AD was prompted by reports of unreliable airspeed indications that were caused by pitot heater resistance shorted to ground. This AD requires replacement of certain parts. We are issuing this AD to address the unsafe condition on these products.
60-09-04: 60-09-04 VICKERS: Amdt. 135 Part 507 Federal Register April 26, 1960. Applies to All Viscount Models 745D and 810 Series Aircraft. Compliance required as indicated. Due to cracks found in the flanges of the main undercarriage cross shaft bearing channels on the inboard and outboard sides of the rib at Station 131 on both wings, the following shall be accomplished: (a) Aircraft which have accumulated 4,000 flights, or more, must be visually inspected, at intervals not to exceed 160 hours' time in service beginning with the effective date of this AD, for cracks in the bearing channels (P/N 70103-1515ND and -1516ND for 700 Series aircraft; P/N 81003- 1061ND and -1062ND for 810 Series aircraft) on the inboard side of the rib in the wheel bay area. Cracked bearing channels must be either repaired and reinforced in accordance with Vickers Mod. D.2866, Parts (A) and (C) for Model 745D's, FG.1513 Part (A) and (C) for Model 810's or replaced and reinforced in accordance with Mod. D.2866 Part (A) and FG.1513 Part (A). Details of crack limits on repairable channels are specified in the respective Modification Leaflets. (b) If wheel bay channels are found cracked, the corresponding channels on the outboard side of the rib in the tank bay area must be either repaired and reinforced or replaced and reinforced at the same time as the inboard components. (c) Not later than April 1, 1961, aircraft which have accumulated 4,000 flights or more must incorporate reinforced inboard and outboard bearing channels at rib Station 131 on both right and left wings in accordance with Vickers Mod. Bulletin D.2866 Part (a) for Model 745D or FG.1513 Part (A) for Model 810 Series, except that aircraft may remain in service on or after April 1, 1961, without the incorporation of the reinforcements provided the following is accomplished in addition to the provisions of (a). (1) Aircraft must be inspected within the next 1,500 flights after April 1, 1961, andevery 1,000 flights thereafter, for cracks in the bearing channels (P/N 70103-1515ND and -1516ND for 700 Series aircraft; P/N 81003-1061ND and -1062ND for 810 Series aircraft) on the outboard side of the rib in the tank bay area. Cracked bearing channels must be repaired and reinforced per (a). (2) If tank bay channels are found cracked, the corresponding channels on the inboard side of the rib in the wheel bay area must be either repaired and reinforced or replaced and reinforced at the same time as the outboard components. (d) The inspections in (a) and (c)(1) are not required after incorporation of the reinforcements. (Vickers-Armstrongs PTL 211 and Mod. D.2866 for 700 Series aircraft and PTL 77 and Mod. FG. 1513 for 800/810 Series aircraft cover this subject.) Revised February 28, 1961.
80-16-02: 80-16-02 COSTRUZIONI AERONAUTICHE GIOVANNI AGUSTA: Amendment 39- 3860. Applies to Model A109A series helicopters, certificated in all categories, which have nose landing gear mount assembly, P/N 1808GR90 (Agusta P/N 109-0500-12-1), manufactured from forging lot SS5211, installed. Compliance required as indicated, unless already accomplished. To prevent failure of the nose landing gear, inspect the mount assembly, P/N 1808GR90 (Agusta P/N 109-0500-12-1), as follows: (a) Within 25 hours time in service after the effective date of this AD, inspect the nose landing gear strut mount, P/N 1808GR90 (Agusta P/N 109-0500-12-1), to determine the manufacturer's lot number as etched on the mount assembly below the shimmy control group. (b) If, as a result of the inspection required by paragraph (a) of this AD, the lot number is found to be other than SS5211, apply an identifying mark to the mount in accordance with "ACCOMPLISHMENT INSTRUCTIONS," Part I, paragraph 1.a., of Costruzioni Aeronautiche Giovanni Agusta Service Bulletin No. 109-14, dated February 12, 1979, (hereinafter referred to as the Service Bulletin), or an FAA-approved equivalent, and return to service. (c) If, as a result of the inspection required in paragraph (a) of this AD the lot number is found to be SS5211 or the lot number is not legible, or is non-existent, before further flight, except as provided in paragraph (j) of this AD, using a mirror and a light source, visually inspect the upper portion of the strut mount for cracking or evidence of cracking. NOTE: During the inspection required by paragraph (c) of this AD, particular attention should be directed to the actuating cylinder attach pin area and the mount horns attach area to the fuselage structure in accordance with the Service Bulletin, or an FAA-approved equivalent. (d) If, as a result of the inspection required in paragraph (c) of this AD, no cracking or evidence of cracking is found, return the helicopterto service and repeat the inspection in paragraph (c) of this AD before the first flight of each day, until paragraph (g) is accomplished. (e) If, as a result of the inspection required in paragraph (c) or (d) of this AD, cracking of the paint or other evidence of cracking is found, inspect the nose landing gear mount assembly as described in paragraph (c) of this AD using the flourescent penetrant method. (f) If, as a result of the inspection in paragraph (e) of this AD, no cracks are found, return the helicopter to service and repeat the inspection in paragraph (c) before the first flight of each day, until paragraph (g) is accomplished. (g) If, as a result of the inspection required in paragraphs (c), (d), (e) or (f) of this AD, cracking is found, before further flight, except as provided in paragraph (j) of this AD: (1) Accomplish paragraph (h) of this AD, and (2) Replace the nose landing gear strut mount assembly, P/N 1808GR90 (Agusta P/N 109-0500-12-1), with a new part of the same P/N with lot number other than SS5211, in accordance with "ACCOMPLISHMENT INSTRUCTIONS," PART II, paragraph 3 through 6, of the Service Bulletin, or an FAA-approved equivalent. (h) If, as a result of the inspections required in paragraph (c), (d), and (e) or (f) of this AD, removal of the nose landing gear is necessary, before reinstallation of the nose landing gear, visually inspect the cylinder aft fitting assembly, P/N 109-0501-19-5, and related attach components for cracks, warpage, elongation of holes and fretting in accordance with "ACCOMPLISHMENT INSTRUCTIONS," PART II, of the Service Bulletin, or an FAA- approved equivalent. Replace damaged components as necessary. (i) When paragraphs (g) has been accomplished, no further inspections are required by this AD. (j) The helicopter may be flown in accordance with FAR 21.197 and FAR 21.199 to a base where the inspection or replacement can be performed. (k) For the purpose of complyingwith this AD, an FAA-approved equivalent may be approved by the Chief, Aircraft Certification Staff, AEU-100, Europe, Africa and Middle East Office, Federal Aviation Administration, c/o American Embassy, Brussels, Belgium. This amendment becomes effective August 11, 1980.
60-10-08: 60-10-08 PIPER: Amdt. 149 Part 507 Federal Register May 13, 1960. Applies to All PA- 22, PA-20, PA-18 Airplanes Equipped With Two Wing Tanks. Compliance required prior to July 15, 1960, and every 100 hours' time in service thereafter. Several accidents have occurred involving engine fuel starvation attributed to a lack of detent action in the fuel selector valve (P/N 11383), causing the pilot to position the selector improperly. If the detent pin in the valve shaft is improperly centered or if the spring retaining washer is installed upside down, the pin will not engage the slotted detent washer. Therefore, the fuel selector valve in the above listed models must be thoroughly cycled to determine whether or not detent engagement is positive. There should be four distinct detents in one complete cycle. If detent engagement is not positive, the valve must be replaced prior to further flight. Also, determine if the position of the fuel valve handle at detent engagement coincides with the proper markings on the indicator plate. If the handle does not coincide with the markings, the plate must be repositioned accordingly. (Piper Service Bulletin No. 141 covers this subject.)
2017-20-07: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model DHC-8-400 series airplanes. This AD was prompted by the failure of the fire control amplifier (FCA), which was likely caused by an electrical short in a discharged squib for a fire extinguishing bottle. This AD requires replacing certain circuit breakers. We are issuing this AD to address the unsafe condition on these products.
60-04-01: 60-04-01 BENDIX: Amendment 107 Part 507 Federal Register February 20, 1960. Applies to All Aircraft Public Address Speaker Systems Using the Bendix MI-51B Amspeaker, Serial Numbers 1001 To 1201, Which Have Carbon Resistors in Parallel With or in Lieu of Wire Wound Resistors R-103 and R-106. (Serial Numbers 1001 To 1051 May Have Been So Modified In Service.) Compliance required by June 1, 1960. Failures have occurred where the speaker cone of the MI-51B was destroyed by fire as a result of overheating of these carbon resistors, thus creating a possible hazardous condition. Due to the seriousness of the fire hazard associated with these failures, any carbon resistors paralleled with or substituted for resistors R-103 and R-106 shall be removed in accordance with either of the following methods: (a) Method No. 1. Clip out 3.9 ohm carbon resistors connected across wire wound resistors R-103 and R-106. (Removal of these resistors will reduce the audio output by approximately 10 percent.) (b) Method No. 2. Replace each of the parallel networks composed of 3.9 ohm carbon resistors in parallel with wire wound resistor R-103 and R-106, with a 0.75 ohm wire wound resistor (plus-minus 5 percent, 1/2 W). (Bendix Service Bulletin No. M-273 dated July 8, 1959 covers the same subject.)
2017-20-01: We are adopting a new airworthiness directive (AD) for all Honeywell International Inc. (Honeywell) TFE731-20 and TFE731-40 turbofan engines. This AD was prompted by two fan disks found with a manufacturing-caused flaw. This AD requires removing affected fan disks and replacing fan disks with a part eligible for installation. We are issuing this AD to address the unsafe condition on these products.
63-18-05: 63-18-05 ROLLS-ROYCE: Amdt. 606 Part 507 Federal Register August 22, 1963. Applies to All Tyne 512 and 515 Engines. Compliance required as indicated. To prevent failure of the high pressure cooling air manifold accomplish the following: (a) If not already accomplished within the past 475 hours' time in service, within the next 25 hours' time in service after the effective date of this AD, and thereafter at intervals not exceeding 500 hours' time in service from the last inspection, inspect the joints adjacent to the flanged connections of the high pressure cooling air manifold for cracks. Cracks across the tack welds are acceptable, however if cracks are found in the parent metal of the flange or pipe connection replace the complete high pressure cooling air manifold assembly before further operation. (b) Upon incorporation of an FAA approved modification or compliance with Section 4(b) of Rolls-Royce Alert Service Bulletin No. Ty A.72-438 revised June 24, 1963, the special inspections in (a) can be discontinued. (c) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, International Engineering and Manufacturing Branch, FAA International Division, 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. (Rolls-Royce Alert Service Bulletin No. Ty A.72-438 revised June 24, 1963, covers this same subject.) This directive effective September 21, 1963.
80-17-03: 80-17-03 MOONEY: Amendment 39-3868. Applies to Model M20J airplanes Serial Numbers 24-0084, 24-0378 through 24-0906, 24-0908 through 24-0925, 24-0927 through 24- 0942, and 24-0946 (Airworthiness Directive Docket No. 80-ASW-3). \n\n\tCOMPLIANCE: Required as indicated unless already accomplished. To prevent total loss of engine oil, accomplish the following: \n\n\t(a)\tBefore further flight, unless already accomplished, remove the top cowl and inspect the positioning of the Rochester Model 3060-18 oil pressure transducer and the condition of the two 45-degree elbows as follows: \n\n\t\t(1)\tClean the oil pressure transducer fitting and the two 45-degree fittings with an oil soluble solvent. \n\n\t\t(2)\tInspect the Rochester Model 3060-18 oil pressure transducer for any evidence of contact with the engine mount. If damage is present, remove and replace it before verifying the clearance outlined in paragraph (a)(3). \n\n\t\t(3)\tVerify a minimum of 0.40 inch clearance between the Rochester Model 3060-18 oil pressure transducer body and the upper right hand engine mount ring and between the Rochester Model 3060-18 oil pressure transducer and the two 45-degree elbow fittings and the engine mount tube. If any clearance is less than 0.40 inch, rotate the oil pressure transducer and its fittings to obtain this minimum clearance. \n\n\tNOTE: Measure the 0.40 inch in an arc perpendicular to the crankshaft centerline. When this clearance exists the oil pressure transducer is almost contacting the vacuum pump body. \n\n\t\t(4)\tStart and operate the engine until it is warm enough to respond smoothly to throttle changes (monitor oil temperature and cylinder head temperature gauges to maintain temperatures within limits), then stop the engine. \n\n\t\t(5)\tInspect the oil pressure transducer and its fittings for any signs of oil leakage. \n\n\t\t(6)\tIf any signs of oil leakage are detected, comply with paragraph (b) of this AD before further flight. \n\n\t(b)\tWithin the next 25 hours' time in service after the effective date of this AD, unless already accomplished, modify the mounting provisions of the oil pressure transducer as follows: \n\n\t\t(1)\tRemove the top cowl. \n\n\t\t(2)\tDisconnect the wire from the oil pressure transducer, P/N 3060-18, and remove the transducer from the 45-degree fitting; retain for reinstallation. \n\n\t\t(3)\tRemove the 45-degree fitting which was attached to the transducer. Leave the 45-degree fitting installed in the engine case port. \n\n\t\t(4)\tConnect an AN 816-3 adapter to this 45-degree fitting. \n\n\t\t(5)\tConnect the flex hose, P/N S94B90145, to this adapter and the other end, 1/8 inch pipe thread, to the transducer; use Tite Seal on all pipe threads. Position hose and fittings to obtain a clearance of .20 inch between hose socket and vacuum pump housing. \n\n\t\t(6)\tRoute the transducer and hose under the upper right hand engine mount tubes and secure the upper outboard tube with an AN742D25 and a MS21919DG8 clamp and an AN3-5A bolt, two AN960-10 washers and an AN363-1032 nut; the smaller clamp, MS21919DG8, clamps around the engine mount tube and the larger clamp, AN742D25, clamps around the transducer body as shown in Figure 1. \n\n\t\t(7)\tConnect the ground wire, Wire No. 21DH04C20, under the bolt head holding the clamps together and to the landing light ground located on the firewall. \n\n\t\t(8)\tSecure the hose to the engine mount with TY-RAP, MS3367-1 ensuring clearance between adjacent components. \n\n\t\t(9)\tReconnect the oil pressure gauge wire to the transducer connection post. \n\n\t\t(10)\tRun the engine and check for oil leaks at the fittings connecting the flex hose and the engine and the flex hose and the oil pressure transducer. If leaks are noted, correct this situation before proceeding. Check the operation of the oil pressure gauge. \n\n\tNOTE: Secure bottom cowling or remove prior to engine run. \n\n\t\t(11)\tReplace cowling and secure all connections. \n\n\tNOTE: Mooney Service Bulletin No. M20-221, "Oil PressureTransducer Installation Modification," pertains to this same subject. \n\n\t(c)\tSpecial flight permits may be issued in accordance with FAR 21.197 and FAR 21.199 to fly airplanes to a base where this AD can be accomplished. \n\n\t(d)\tAny alternate equivalent method of compliance with this airworthiness directive must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration. \n\n\tThis amendment supersedes Amendment 39-3725 (45 FR 20780) AD 80-07-12. \n\n\tThis amendment becomes effective August 18, 1980. \n\n\n\n\nAD 80-17-03 \nFIGURE 1
60-02-05: 60-02-05 DOUGLAS: Amdt. 84 Part 507 Federal Register January 16, 1960. Applies to the Following Aircraft: DC-6 Serial Numbers 42878, 43030 To 43033 Inclusive, 43136, 43148 To 43151 Inclusive, 43212 To 43214 Inclusive, and 43216 To 43218 Inclusive. \n\n\tCompliance required as indicated. \n\n\tTo detect cracking of the lower front and center spar cap tangs at intersection with lower fuselage attach angle the following must be accomplished on affected DC-6 aircraft having in excess of 16,000 hours service time. \n\n\t(a)\tInspect lower front spar cap at the nearest maintenance inspection period to 200 hours service time unless similar inspection has been conducted within the last 1,250 hours service time. \n\n\t(b)\tInspect lower front and center spar caps at maintenance inspection period nearest to each succeeding 1,250 hours service time. \n\n\t\t(1)\tAt the first 1,250-hour inspection period, the holes located in aft tang of front spar lower cap and fuselage attach angle should be enlarged and new attachments installed. (Kit "A" of Douglas SB A-845 or equivalent.) \n\n\t\t(2)\tAt next regularly scheduled overhaul period, the holes located in forward tang of front spar lower cap should be enlarged and new attachments installed. (Kit "A" of Douglas SB A-845 or equivalent.) \n\n\t(c)\tIf spar cap cracks are found, temporary rework per Drawing No. 3645935 (Kit "B"), or permanent rework per Drawing No. 5765079 (Kit "C") or equivalent, must be accomplished. If temporary rework is installed, inspection must be repeated at 1,250-hour intervals for a maximum of 3,200 hours service time, at which time permanent rework per Drawing No. 5765079 (Kit "C"), or equivalent, must be accomplished. \n\n\t(d)\tAll aircraft must have permanent rework per Drawing No. 5765079 (Kit "C"), or equivalent, accomplished within next 6,400 hours service time. \n\n\t(e)\tAfter installation of permanent rework per Kit "C", or equivalent, operators may revert to normal repetitive inspection periods not to exceed 3,200 hours service time. \n\n\t(Douglas Service Bulletin DC-6 No. A-845 dated July 31, 1959, covers this same subject.)
2017-19-16: We are adopting a new airworthiness directive (AD) for certain Rolls-Royce plc (RR) RB211 Trent 553-61, Trent 553A2-61, Trent 556-61, Trent 556A2-61, Trent 556B-61, Trent 556B2-61, Trent 560-61, and Trent 560A2-61 turbofan engines. This AD requires replacement of the low- pressure compressor (LPC) case A-frame hollow locating pins. This AD was prompted by LPC case A-frame hollow locating pins that may have reduced integrity due to incorrect heat treatment. We are issuing this AD to correct the unsafe condition on these products.
81-02-09: 81-02-09 BOEING: Amendment 39-4024. Applies to all Boeing Model 727-100 and 727-100C series airplanes equipped with the aft airstair emergency extension system, except all-cargo configurations and those airplanes where the aft stair emergency extension system has been deactivated.\n \n\tCompliance is required as indicated. Accomplish the following:\n \n\tA.\tPrior to June 1, 1981, replace or modify the aft airstair emergency extension control handle in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. (Note: Accomplishment of Boeing Service Bulletin 727-52-120 dated March 21, 1980, or later FAA approved revisions has been approved as a means of compliance with the requirements of this AD.) \n\n\tB.\tUpon request of the operator, an FAA aviation safety inspector, subject to prior approval by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, may adjust the compliance date if the request contains substantiating data to justify the change. \n\n\tThis amendment becomes effective April 1, 1981.
2017-19-27: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model DHC-8-401 and -402 airplanes. This AD was prompted by the discovery of cracking on two test spoiler power control unit (PCU) manifolds during testing by the manufacturer. This AD requires replacement of affected spoiler PCUs. We are issuing this AD to address the unsafe condition on these products.
59-26-02: 59-26-02 PIPER: Applies to PA-24 and PA-24 "250" Airplanes Serial Numbers 24-1 To 24-1373 Inclusive. Compliance required by January 15, 1960. To prevent clogging, the two fuel cell vent tubes which are located under the wings shall be modified in the following manner: Measure a distance of 1/2-inch down from the bottom of the wing skin along the forward side of each protruding vent tube. At this point, cut the tube off at a 45-degree angle to the bottom skin so that the end of the tube remains square. (Piper Immediate Action Service Bulletin No. 180 covers this subject.)
2017-19-15: We are adopting a new airworthiness directive (AD) for certain Technify Motors GmbH TAE 125-02 reciprocating engines. This AD requires replacement of the clutch with a dual mass flywheel. This AD was prompted by a loss of engine power in flight caused by oil leaking from the gearbox radial shaft sealing ring that contaminated the clutch. We are issuing this AD to correct the unsafe condition on these products.
59-25-05: 59-25-05 FORNEY (ERCOUPE): Applies to All (Ercoupe) Forney Aircraft With Serial Numbers Up to 3,335 Inclusive. Compliance required by December 31, 1959, and thereafter every 100 hours of operation or periodic inspection, whichever occurs first. Fatigue failures have continued to occur in the rudder main rib where the control horn is attached after installation of reinforcement plates. Therefore, it is required that a visual inspection be made of the area around the rudder control horn for excessive deflection of the horn, canning of rudder skin, or any other unusual peculiarity which would indicate main rudder rib damage. If damage is evident, rudder rib Erco P/N 415-240 12 L/R must be replaced with Forney P/N F-24015 L/R or equivalent. This inspection may be discontinued when the heavier gage rib is installed. (Forney Service Bulletin No. 105 covers this subject.) This supersedes AD 47-20-07.
59-13-02: 59-13-02 PIPER: Applies to Models PA-24 and PA-24 "250" Aircraft Serial Numbers 24-1 to 24-978 Inclusive and 24-980. Compliance required within the next 100 hours of operation or by October 1, 1959, whichever occurs first. Service experience indicates that cracks have developed in the aileron balance weight attachment bulkheads. These bulkheads are riveted to the front spar of the aileron and are the supports to which the balance weight arm is attached. To reduce the probability of failure of the aileron balance weight arm attachment install reinforced bulkheads on both ailerons except on Serial Number 24-980 replace the balance weight attachment bulkhead on right aileron only. (Piper Service Bulletin No. 173 also covers this subject and states "Service Kit, Part Number 734-233, is available from your nearest Piper distributor or dealer free of charge if the airframe serial number is included on the purchase order.")
79-19-01 R2: 79-19-01 R2 BOEING: Amendment 39-3556 as amended by Amendment 39-4087 is further amended by Amendment 39-4486. Applies to all Boeing 720/720B, 707-300, 707-400, 707- 300B, and 707-300C series airplanes. \n\tA.\tAfter the effective date of this amendment, perform a low frequency eddy current inspection for cracks in the wing lower surface splice stringers in accordance with Boeing Service Bulletin 3226, Rev. 5, dated November 15, 1981, or later FAA approved revisions, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. Inspections are to be made at the threshold times, within the prescribed initial interval and at repetitive intervals shown below: \n\tAt the inboard nacelle strut drag brace, the affected lower skin area covered by the fairing may be visually inspected for cracks and evidence of fuel leakage. If crack indications are noted in stringers, or skin cracks or fuel leakage are found at the diagonal brace fairing area, tank entry and inspection by high frequency eddy current of the wing splice stringers is required. \n\n\nAirplane\nThreshold\nInitial Inspection within\n\nRepetitive Interval \n\n\n\nunless accomplished \nwithin the last \n\n720/720B\n14,000 ldgs\n715 ldgs\n715 ldgs\n1,430 ldgs\n707-300/400\n21,000 ldgs\n1,675 ldgs\n1,675 ldgs\n3,350 ldgs\n707-300B\n19,000 ldgs\n1,425 ldgs\n1,425 ldgs\n2,850 ldgs\n707-300C\n17,000 ldgs\n725 ldgs\t\n725 ldgs\n1,450 ldgs \n707-300C\n\n17,000 ldgs\n1,425 ldgs\n1,425 ldgs\n2,850 ldgs \n\n(passenger only) \n\n\n\n\n\t\t\t\t\t\t\t\t\t\t \n\tB.\tIf cracks are found, repair prior to further revenue flight in accordance with a method approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\tC.\tFor the purpose of complying with this AD and subject to acceptance by the assigned FAA Maintenance Inspector, the number of landings may be determined by dividing each airplane's time-in-service by the operator's fleet average from takeoffto landing for the airplane type. \n\tD.\tUpon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region may adjust the inspection interval if the request contains substantiating data to justify the increase for that operator. \n\tE.\tAirplanes with cracked splice stringers may be flown in accordance with FAR 21.197 to a base where repairs can be performed. \n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at the FAA, Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tAmendment 39-3556 became effective September 18, 1979. \n\tAmendment 39-4087 became effective May 17, 1981. \n\tThis Amendment 39-4486 becomes effective November 15, 1982.