Results
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.
2017-19-19: We are adopting a new airworthiness directive (AD) for certain Rolls-Royce plc (RR) Trent XWB-75, Trent XWB-79, Trent XWB-79B, and Trent XWB-84, turbofan engines. This AD requires replacement of the low-pressure compressor (LPC) case support inboard pins. This AD was prompted by LPC case support inboard pins that may have reduced integrity due to incorrect heat treatment. We are issuing this AD to correct the unsafe condition on these products.
78-24-06 R1: 78-24-06 R1 BELL: Amendment 39-3358 as amended by Amendment 39-4032. Applies to Bell Models 206L and 206L-1 helicopters, certificated in all categories (Airworthiness Docket No. 78-ASW-53). Compliance required as indicated, unless previously accomplished. To prevent possible failure of the horizontal stabilizer, P/N 206-023-119, all dash numbers, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, modify the left and right upper stabilizer supports, P/N 206-023-100-009 and -010, respectively, in accordance with Bell Helicopter Textron Service Bulletin 206L-78-3 dated October 23, 1978, or Bell Helicopter Textron Alert Service Bulletin 206L-80-16 dated November 17, 1980, or FAA approved equivalent, so that the critical area can be checked. (b) Before the first flight of each day after compliance with paragraph (a), visually check the stabilizer skin area exposed by the cutouts, in the upper stabilizer supports, forcracks. (1) If a crack is found, remove and replace the horizontal stabilizer before further flight. (2) If no cracks are found, continue the repetitive check specified above. (c) The checks required by this AD may be performed by the pilot. NOTE: For the requirements regarding the listing of compliance and method of compliance with this AD in the aircraft's permanent maintenance record, see FAR 91.173. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Textron, P.O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. Amendment 39-3358 became effective December 5, 1978. This amendment 39-4032 becomes effective May 11, 1981.
2017-19-26: We are superseding Airworthiness Directive (AD) 2008-12-04, which applied to certain The Boeing Company Model 737-600, -700, -700C, -800, and -900 series airplanes. AD 2008-12-04 required various repetitive inspections to detect cracks along the chem-milled steps of the fuselage skin, and to detect missing or loose fasteners in the area of a certain preventive modification or repairs; replacement of the time-limited repair with a permanent repair, if applicable; and applicable corrective actions which would end certain repetitive inspections. This AD reduces the post-modification inspection compliance times, limits installation of the preventive modification to airplanes with fewer than 30,000 total flight cycles, and adds repetitive inspections for modified airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) that indicated that the upper skin panel at the chem-milled step above the lap joint is subject to widespread fatigue damage (WFD) if themodification was installed after 30,000 total flight cycles. We are issuing this AD to address the unsafe condition on these products.
58-22-03: 58-22-03 PIPER: Applies to All Model PA-23 Aircraft Equipped With Goodrich G-3-787 Main Wheel Assemblies. Compliance required as indicated. Failures of the Goodrich G-3-787 wheel assembly are being reported. These wheels may be continued in service subject to inspection as specified below, or replaced with Goodrich G-3-880 wheel assembly or Cleveland Aircraft Products wheel assembly Model 3060 or 3080A and brake assembly Model 3000-500 or another equivalent approved type wheel and brake combination. 1. Remove wheels and tires and inspect wheels at each one hundred hours of operation or at each tire change, whichever occurs first. 2. Inspect the flange area of both wheel halves by means of Dy-Check or Zyglo, whichever is available. Cracks may appear on either the inside or outside surface of the flange. 3. If cracks are present in the flange area, remove the defective wheel half from service and replace as indicated above. 4. To detect possible flange failures during preflight inspection, look for outward deformation of flanges. A wheel with a flange failure will appear to wobble when rotated. (Goodrich Service Bulletin No. 102, dated July 25, 1956, and Piper Service Letter No. 291, dated June 5, 1957, covers this same subject.)
60-21-02: 60-21-02 LOCKHEED: Amdt. 207 Part 507 Federal Register October 7, 1960. Applies to All Model 18 Aircraft Which Have Been Converted to the "Learstar" Configuration. Compliance required as indicated. Flight tests have disclosed that excessive temperature stratification at the carburetor air screen exists during operation with partial carburetor air preheat. This stratification results in erroneous readings on the cockpit carburetor air temperature gage and may, under icing conditions, cause ice formation on the cold portions of the air screen and in the carburetor. While full preheat is available if needed for ice elimination, only partial preheat should be used for continuous operation under certain temperature conditions in order to avoid exceeding the engine manufacturer's carburetor air temperature limit of 38 degrees C. The following action is required: (a) Effective November 15, 1960, Learstar aircraft shall be restricted against operation in known icing conditions until modifications to the air preheat system covered in paragraph (b) are accomplished. The following placard shall be posted in full view of the pilot: "OPERATION INTO KNOWN ICING CONDITIONS PROHIBITED." The limitations section of the FAA approved Airplane Flight Manual is hereby amended to incorporate this limitation. (b) The operating restriction into known icing conditions shall continue until modifications are accomplished to the carburetor air preheat system which will result in conservative C.A.T. indications for the prevention of ice formation on the carburetor screen and engine induction system at all preheat positions. Such modifications shall also permit operation with preheat under varying power and ambient temperature conditions without resulting in excessive C.A.T. A satisfactory modification to meet the requirements is covered in FAA approved PacAero Engineering Corporation Service Bulletin No. 14, dated August 26, 1960. An FAA approved airplane flight manualrevision setting forth recommended procedures for safe operation of the system will be supplied by Pac Aero with the modification kit. Any deviations from the modifications or procedures set forth in the service bulletin and airplane flight manual revision must be approved by FAA, Region Four, Engineering and Manufacturing Branch, Los Angeles, California. (c) Upon compliance with paragraph (b), the operating restriction set forth in paragraph (a) is cancelled.
79-17-01: 79-17-01 ROCKWELL INTERNATIONAL: Amendment 39-3525. Applies to Rockwell NA-265-60 airplanes which have been modified by Raisbeck Group Supplemental Type Certificate SA687NW. To prevent flutter caused by the accumulation of undrained water in the elevators, accomplish the following, unless already accomplished, within 30 days or 30 flight hours, whichever occurs first: A. Modify the elevators to provide water drainage provisions in accordance with Raisbeck Service Bulletin No. 8, dated July 13, 1979, or later FAA approved revisions. B. Using the procedure specified in Raisbeck Service Bulletin No. 8, dated July 13, 1979, or later FAA approved revisions, determine the static balance of the elevators and if required, rebalance them in accordance with instructions in that service bulletin. C. Equivalent modifications may be approved by the Chief, Engineering and Manufacturing Branch, Northwest Region. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to The Raisbeck Group, 7777 Perimeter Road South, Boeing Field International, Seattle, Washington 98108. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective August 15, 1979.
2017-19-17: We are superseding Airworthiness Directive (AD) 2016-17-02, which applied to certain Dassault Aviation Model FALCON 900EX and FALCON 2000EX airplanes. AD 2016-17-02 required revising the airplane flight manual (AFM) to include procedures to follow when an airplane is operating in icing conditions. AD 2016-17-02 also provided optional actions after which the AFM revision may be removed from the AFM. Since we issued AD 2016-17-02, we have determined additional actions are necessary to address the identified unsafe condition. This new AD retains the requirement of AD 2016-17-02, and also requires a detailed inspection of the wing anti-ice system ducting (anti-ice pipes) for the presence of a diaphragm, and replacement of ducting or re- identification of the ducting part marking. We are issuing this AD to address the unsafe condition on these products.
58-01-06: 58-01-06 PIPER: Applies to Model PA-23 Aircraft, Serial Numbers 23-1 to 23-1219 Inclusive. Compliance required as indicated. Due to the installation of the front stabilizer-to-fuselage attachment fitting P/N 17093-00, on additional aircraft to those covered by AD 57-13-09 and since special inspections are not required when the redesigned fitting P/N 17093-03 is installed, this supersedes the portions of AD 57-13-09 concerning this fitting and revision issued on Card No. 57-22. Inspect visually for cracks, the front stabilizer fitting, P/N 17093-00 every 100 hours until replaced with the redesigned fitting P/N 17093-03. Fittings found cracked must be replaced. (Piper Service Bulletin No. 160 dated October 7, 1957, covers the same subject.)
58-01-07: 58-01-07 PIPER: Applies to All J-3 Series and J-5 Series Aircraft. Compliance required by February 1, 1958. To preclude the possibility of failures of the fork end of the turnbuckles in the control system, the following inspection and rework is necessary. Failures of the fork end of the turnbuckles have occurred in the area covered by the safety wire. This results from binding caused by the attaching bolt being drawn up too tightly on the fork end of the turnbuckle. Inspect the turnbuckle to horn attachment at the elevators, rudder and ailerons to determine that an AN 23-12 clevis bolt is installed with one AN 960-10 washer under the nut. This assembly should swivel freely.