Results
2000-04-20: This amendment adopts a new airworthiness directive (AD), applicable to Bell Helicopter Textron Canada (BHTC) Model 407 helicopters, that requires replacing a certain hydraulic relief valve (valve) with a different valve. This amendment is prompted by the discovery of a manufacturing defect in a valve. The actions specified by this AD are intended to prevent intermittent loss of hydraulic pressure to the flight controls and subsequent loss of control of the helicopter
46-46-01: 46-46-01 ERCO: (Was Mandatory Note 9 of AD-718-6.) Applies Only to 415-C, -CD and -D Aircraft Serial Numbers 2623 to 2994, Inclusive. Compliance required prior to January 1, 1947. Install a new fuselage gas tank overflow line (Erco P/N 415-48162) and replace the imperial brass compression sleeve No. 60F with a rubber washer No. A549, Kohler Co. of A-64-3, Hayes Industries, Inc. (Erco Service Department Bulletin No. 15, dated August 24, 1946, covers this same subject.)
2014-20-13: We are adopting a new airworthiness directive (AD) for Pacific Aerospace Limited Model 750XL airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued 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 failure of the fin forward pickup due to possible fatigue cracks. We are issuing this AD to require actions to address the unsafe condition on these products.
2014-20-15: We are superseding airworthiness directive (AD) 2012-02-13 for certain Airbus Helicopters, Inc. (Airbus Helicopters) Model EC130B4 helicopters. AD 2012-02-13 required inspecting certain areas of the tailboom/Fenestron junction frame (junction frame) for a crack. This AD retains the requirements of AD 2012-02-13, expands the inspection area of the junction frame, and reduces the repetitive inspection interval. These actions are intended to detect a crack in the junction frame, which could result in detachment of the Fenestron and subsequent loss of control of the helicopter.
2000-04-17: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Boeing Model 747-100, -200, and -300 series airplanes. This action requires repetitive inspections to detect fatigue cracking in the upper deck floor beams located at certain body stations, and repair, if necessary. This amendment is prompted by a report by the manufacturer that, during a fatigue test, the upper chord and web of the upper deck floor beams located at body stations (BS) 340 and 360 were found severed at approximately 34,000 total flight cycles. Another report by an operator indicated that a severed upper chord and web were found in the upper deck floor beam at BS 380 at approximately 33,000 total flight cycles. In addition, cracking was found at multiple fastener hole locations. The actions specified in this AD are intended to prevent failure of the upper deck floor beams at certain body stations due to fatigue cracking, which could result in rapid decompression and consequent reduced controllability of the airplane.
2000-04-18: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 757 series airplanes, that requires replacement of transmission assemblies for the trailing edge flaps with modified transmission assemblies. This amendment is prompted by reports of broken bolts that attach the transmission assemblies for the trailing edge flaps. The actions specified by this AD are intended to prevent damage to the flap system, adjacent system, or structural components; and excessive skew of the trailing edge flap; which could result in reduced controllability of the airplane.
93-05-12: 93-05-12 AIRBUS INDUSTRIE: Amendment 39-8516. Docket 92-NM-214-AD. Applicability: Model A320 series airplanes; serial numbers 002 through 071, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the wing, accomplish the following: (a) Prior to the accumulation of 13,000 total landings, or within 1,000 landings after the effective date of this AD, whichever occurs later: Perform a detailed visual inspection to detect cracks in the left- and right-hand sides of the wing rear spar between ribs 1 and 2, in accordance with Airbus Industrie Service Bulletin A320-57-1020, dated September 5, 1991. (1) If any crack is found, prior to further flight, repair in accordance with a method approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. (2) If no cracks are found, repeat the inspection required by paragraph (a) of this AD thereafter at intervals not to exceed 3,000 landings. (b) Within 3 years after the effective date of this AD, modify the outer wing rear spar forward face between ribs 1 and 2, in accordance with Airbus Industrie Service Bulletin A320-57-1021, dated September 5, 1991. (c) Accomplishment of the modification required by paragraph (b) of this AD constitutes terminating action for the repetitive inspections required by paragraph (a) of this AD. (d) 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, ANM-113, 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, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (e) 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. (f) The inspection shall be done in accordance with Airbus Industrie Service Bulletin A320-57-1020, dated September 5, 1991. The modification shall be done in accordance with Airbus Industrie Service Bulletin A320-57-1021, dated September 5, 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 Airbus Industrie, Airbus Support Division, Avenue Didier Daurat, 31700 Blagnac, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (g) This amendment becomes effective on May5, 1993.
2014-20-10: We are superseding Airworthiness Directive (AD) 2013-11-14 for certain The Boeing Company Model 777-200 and -300 series airplanes. AD 2013-11-14 required repetitive general visual inspections of the strut forward dry bay for the presence of hydraulic fluid, and related investigative and corrective actions (including checking drain lines for blockage due to hydraulic fluid coking, and cleaning or replacing drain lines to allow drainage) if necessary. This AD adds airplanes to the applicability. This AD was prompted by reports of hydraulic fluid contamination (including contamination caused by hydraulic fluid in its liquid, vapor, and/or solid (coked) form) found in the strut forward dry bay. We are issuing this AD to detect and correct hydraulic fluid contamination of the strut forward dry bay, which could result in hydrogen embrittlement of the titanium forward engine mount bulkhead fittings, and consequent inability of the fittings to carry engine loads, resulting in engine separation. Hydrogen embrittlement also could cause a through-crack formation across the fittings through which an engine fire could breach into the strut, resulting in an uncontained strut fire.
2000-04-13: This amendment adopts a new airworthiness directive (AD), applicable to certain Aerospatiale Model ATR72 series airplanes, that requires initial and repetitive inspections to detect fatigue cracking in certain areas of the fuselage, 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 fatigue cracking of the fuselage and the passenger and service doors, which could result in reduced structural integrity of the airplane.
70-13-02: 70-13-02 GENERAL DYNAMICS: Amdt. 39-1016. Applies to all Model 340 and 440 Series airplanes including those modified in accordance with STC SA4-1100 or STC SA1096WE. Compliance required as indicated. To detect cracks in the main windshield lower longeron splice channel, and prevent possible failure of the windshield support structure and loss of pressurization, accomplish the following: (a) Within the next 150 hours' time in service after the effective date of this AD or before the accumulation of 10,000 hours' time in service, whichever occurs later, unless already accomplished within the last 1850 hours' time in service, inspect the longeron splice channel P/N 340-3110113-15 for cracks in accordance with General Dynamics 640(340D) Service Bulletin No. 53-3, dated April 24, 1970, or later FAA approved revision, or an equivalent inspection procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region. (b) If no cracks are found, repeat the inspection of paragraph (a) above at intervals not to exceed 2000 hours' time in service. (c) If cracks are found, either: (1) Replace the longeron splice channel P/N 340-3110113-15 before further flight, and perform the inspection per (d) below, or (2) (i) Install on the instrument panel in full view of both the pilot and copilot, a placard limiting operation to a maximum of 18,000 feet above sea level. (ii) Change the cabin pressure relief valve nominal setting to 3 psi. (iii) Visually inspect the adjacent areas per S.B. 53-3 at intervals not to exceed 150 hours' time in service, until the channel can be replaced. A cracked channel must be replaced with a new channel within 1000 hours' time in service from the initial discovery of the crack. Channel P/N 340-3110113-15 must be replaced before further flight if any cracking of the adjacent structure is found during these 150 hour interval inspections. (iv) The cabin pressure relief valve setting may be returned to normal and the placard removed from the instrument panel when the cracked channel is replaced. Compliance with the inspection intervals of (d) below is required after replacing the channel. (d) Normal inspection intervals may be resumed for a period of 10,000 hours' time in service when a cracked channel P/N 340-3110113-15 is replaced with a new channel. Before the accumulation of 10,000 hours' time in service after replacement of the channel and thereafter at intervals not to exceed 2000 hours' time in service, inspect the channel for cracks in accordance with General Dynamics 640(340D) Service Bulletin No. 53-3, dated April 24, 1970, or later FAA approved revision, or an equivalent inspection procedure approved by the Chief, Aircraft Engineering Division, FAA Western Region. If cracks are found, comply with Paragraph (c) above. (e) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Regional Director, 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 that operator. This amendment becomes effective July 28, 1970.
77-15-12: 77-15-12 DETROIT DIESEL ALLISON: Amendment 39-2977 as amended by Amendment 39-3037. Applies to 250-B15 and B17 and 250-C18 and C20 series engines equipped with power turbine outer coupling nuts Part Number (P/N) 6846278 as installed in aircraft certificated in all categories. NOTE THAT THIS AMENDED VERSION OF AD 77-15-12 REQUIRES THAT ALL POWER TURBINE COUPLING NUTS, BOTH COATED AND UNCOATED, BE INSPECTED ON A REPETITIVE BASIS. Compliance schedule - Compliance required as follows unless previously accomplished. (a) For engines with turbine sections having more than 400 hours or one calendar year time in service since installation in an aircraft, initial compliance must be accomplished within the next 100 hours or 90 days, whichever occurs first, from the effective date of the original amendment of July 27, 1977. (b) For engines with turbine sections having less than 400 hours and less than one calendar year time in service since installation in an aircraft, initial compliance must be accomplished within 500 hours or one year time in service since installation in an aircraft, whichever occurs first. (c) For each turbine section incorporating an acceptable uncoated or coated coupling nut, as defined by the below referenced service letters, repetitive compliance is required at intervals not to exceed 500 hours or one calendar year time since last inspection, whichever occurs first. To preclude possible engine failure resulting from power turbine coupling nut failure: Inspect the power turbine outer coupling nut P/N 6846278 in accordance with Detroit Diesel Allison Commercial Service Letter Numbers 88 for the C18 series, 1060 for the C20 series, 25 for the B15 series, and 1030 for the B17 series engines. All Service Letters are Revision 3 or later FAA approved revisions. If an unacceptable extent of corrosion, as defined by the Service Letter, is found, the nut shall be replaced before further flight except that the aircraft may be flown in accordance with FAR 21.197 to a base where the repair can be performed. Amendment 39-2977 was effective July 27, 1977. This amendment 39-3037 is effective September 23, 1977, and was effective immediately for all recipients of airmail letters dated August 19, 1977 which contained this amendment.
2014-20-14: We are superseding Airworthiness Directive (AD) 2014-04-03 for all Pacific Aerospace Limited Model 750XL airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued 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 broken control column attachment bolts failing in service. We are issuing this AD to require actions to address the unsafe condition on these products.
75-11-09: 75-11-09 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-2219. Applies to Model L-1011-385-1 Series Airplanes, certificated in all categories, incorporating Lockheed counterbalances, P/N 671827-107, -109, -113. To prevent possible failure of Belleville springs resulting in inoperative counterbalances of the Type I. emergency door actuation system, which could cause the most rearward Type I. left and right doors to be inoperative in the emergency mode, accomplish the following: Compliance required as indicated. (a) For airplanes incorporating P/N 671827-113 counterbalances: (1) Within 300 additional flight hours after the effective date of this AD, unless already accomplished, reduce the spring prewind per accomplishment instructions of Lockheed-California Company Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions. (2) Within 300 additional flight hours after the effective date of this AD, unless previously accomplishedwithin the last l,300 flight hours prior to the effective date of this AD, and at intervals not to exceed 1,600 hours thereafter, until (3), below is accomplished, perform the operational checks on the door emergency mode for proper functioning of the counterbalance for the most rearward Type I. left and right doors, per the accomplishment instructions of Lockheed-California Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions. (i) If the operational checks are not satisfactory, replace the defective counterbalances, P/N 671827-113, with either a P/N 671827-107, or -109 counterbalance per Lockheed-California Company Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions, or a counterbalance P/N 671827-115 or -117 per Lockheed-California Service Bulletin 093-52-076, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions. Prior to further flight perform the operational checks toensure that the operation of the doors with the new counterbalances is satisfactory. (ii) Mark the defective P/N 671827-113 counterbalances in a conspicuous manner to prevent inadvertent return to service. (3) Within 8000 hours additional flight time after the effective date of this AD, unless already accomplished, replace all counterbalances P/N 671827-113 with either counterbalances P/N 671827-115 or -117, per Service Bulletin 093-52-076, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions or with counterbalances P/N 671827-107 or - 109 per Lockheed-California Company Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions. Prior to further flight perform the operational checks to ensure that the operation of the doors with the new counterbalances is satisfactory. If satisfactory operation of the doors exist no further inspections are required per this AD. (b) For airplanes incorporating counterbalances P/N671827-107 and -109: (1) Within 300 additional flight hours after the effective date of this AD, unless already accomplished, perform the operational checks of the door emergency mode for the most rearward Type I. left and right doors, per the accomplishment instructions of the Lockheed- California Company Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975, or later FAA-approved revisions. (i) If the operational checks are satisfactory then no further checks are required per this AD. (ii) If the operational checks are not satisfactory, replace the defective counterbalances, P/N 671827-107 or -109, with either a new counterbalance, P/N 671827-107 or - 109, per Lockheed-California Company Service Bulletin 093-52-072, Revision Number 1, dated May 8, 1975 or later FAA-approved revisions, or with counterbalances, P/N 671827-115 or -117, per Lockheed-California Company Service Bulletin 093-52-076, Revision Number 1, dated May 8, 1975 or later FAA-approved revisions. Perform the doors operational checks to ensure that the operation of the doors with the new counterbalance is satisfactory. If satisfactory operation of the doors exist no further inspections are required per this AD. (iii) Mark the defective P/N 671827-107 or -109 counterbalances in a conspicuous manner to prevent inadvertent return to service. (c) Equivalent inspections may be approved by the Chief, Aircraft Engineering Division, FAA Western Region. (d) Airplanes may be flown to a base where the inspections and modifications can be performed per FAR's 21.197 and 21.199. This amendment becomes effective May 30, 1975.
48-13-06: 48-13-06 BELLANCA: Applies to Models 14-13, 14-13-2 Serial Numbers 1060 Through 1310. Compliance required after each 25 hours of operation. To prevent failure of the four engine-cowl-support brackets, Bellanca P/N 9892-13, mounted on the firewall and possible cowl loss in flight, the brackets should be closely examined for cracks. If cracks are noted, heavier brackets available from the factory should be installed, in which case inspection is no longer required. (Bellanca Service Bulletin No. 16 dated December 8, 1947, covers this same subject.)
68-20-08: 68-20-08 SIKORSKY: Amdt. 39-633 as amended by Amendment 39-665 is further amended by Amendment 39-1141. Applies to S-61A, S-61L and S-61N Type Helicopters. To assure proper adjustment of the Automatic Flight Control System (AFCS) accomplish the following: A. Prior to next AFCS flight inspect all AFCS electrical connectors for electrical integrity, pin alignment and perform a functional test of the AFCS system including amplifier and amplifier rack connectors Cannon Part No. DPJ-98-33S-B or DPJM-98-33S-B and box mating connectors Part No. DPJ-98-34P-A. B. Prior to the first AFCS flight each day accomplish the following checks: 1. A pitch stick sensor check on the ground with electrical and hydraulic power "ON" as follows (a) engage the AFCS (b) place C.G. trim knob to center of its travel (c) select "P" on null indicator switch (d) move pitch cyclic stick to center the null indicator (e) verify that forward stick motion deflects the needle to the right. 2. A roll stick sensor check on the ground with electrical and hydraulic power "ON" as follows: (a) Engage AFCS (B) select "R" on null indicator (c) move roll cyclic stick to center the null indicator (d) verify that left stick motion deflects the needle to the left. 3. A pitch and roll AFCS hard-over check on the ground in accordance with Sikorsky Service Bulletin No. 61B65-1 Paragraphs B-1 thru B-4 dated April 12, 1967, or later FAA approved revision or alternate check approved by the Chief, Engineering & Manufacturing Branch, Eastern Region. 4. Yaw Hard-over check on the ground as follows: (a) position null indicator switch to position "Y" and position pedals to full right position (b) position yaw channel disengage switch to "ON" and yaw hard-over switch to "LEFT". Verify that pointer on null indicator also moves hard-over and pedals start motion to the left (c) verify that it is possible to push right pedal to override hard-over condition (d) position hard-over switch to "RIGHT" and pedals to the full left position. Verify that the pointer on null indicator moves hard-over and pedals start motion to the right. (e) Verify that it is possible to push left pedal to override hard- over condition. C. After the effective date of this AD as amended and prior to the next AFCS Flight, if not already accomplished, with AFCS "OFF" measure the adjustment of the auxiliary servo linkage, by placing dial indicators at the output of the servo and at the bolt located in the sloppy link to ascertain that the adjustment of the linkage is .010" or less in the "AUX-ON" and "AUX- OFF" positions. Perform stick jump test to further verify these measurements. Perform these checks in the roll attitude, pitch and yaw functions. D. Within the next 1200 hours time in service and thereafter at 1200 hour intervals of time in service, accomplish the following: 1. Perform AFCS hard-over check according to Sikorsky Service Bulletin No. 61B65-1 dated April 12, 1967 paragraph B-5, or alternate check approved by the Chief, Engineering and Manufacturing Branch, Eastern Region. E. Compliance with this AD is not required if the AFCS installation is disconnected. Amendment 39-633 was effective August 17, 1968. Amendment 39-665 was effective October 10, 1968. This Amendment, 39-1141, is effective January 14, 1971.
86-20-01: 86-20-01 MITSUBISHI: Amendment 39-5428. Applies to Models MU-2B, -10, -15, -20, -25, -26, -26A, -30, -35, -36, -36A, -40, and -60 (Serial Numbers 1 through 753 inclusive, with or without the SA suffix) airplanes certificated in any category. NOTE: The serial numbers of airplanes manufactured in the United States by MAI under Type Certificate A10SW are suffixed by "SA". The serial numbers of airplanes manufactured in Japan by MHI under Type Certificate A2PC have no suffix. Compliance required as indicated unless already accomplished. To assure anti-ice capability of pitot system and proper electric load distribution of anti-ice/deice circuits, accomplish the following: (a) Before next flight after the effective date of this AD: (1) Modify the following in accordance with Mitsubishi Heavy Industries (MHI) Ltd., Service Recommendation (SR) 053, Revision A, dated October 23, 1984, or Mitsubishi Aircraft International, (MAI) Inc. SR 020/34-005, Revision B, dated May 24, 1985, as applicable: (i) The electrical circuitry of the pitot tube and anti-ice/deice systems and, (ii) The pitot tube system by installing higher heat capacity pitot tube, P/N PH1100. (2) As an alternate means of compliance: (i) Before further flight: (A) Fabricate and install a temporary placard(s) in full view of the pilot, using letters of minimum 0.10 inch in height which state: (I) "FLIGHT IN KNOWN ICING CONDITIONS IS PROHIBITED". (II) "TURN PITOT HEAT ON DURING FLIGHT IN VISIBLE MOISTURE". (III) "Pilot and copilot airspeed indicators may display erroneous data after: (a) Flight in visible moisture; (b) Outside storage in rain without pitot covers; or (c) Washing of airplane. Refer to AFM for corrective action", and (B) On the "TYPES OF OPERATION" placard located in the cockpit delete, using opaque tape, the words "ICING CONDITIONS", and (C) Add the following information to the "LIMITATIONS" section of the FAA Approved Airplane Flight Manual (AFM) which supersedes any other AFM information which may be contradictory: (I) "FLIGHT in known icing conditions is PROHIBITED", and (II) "TURN PITOT HEAD HEATER ON DURING FLIGHT IN VISIBLE MOISTURE", and (III) "The pilot and copilot airspeed indicator may display erroneous data after any: (a) Flight in visible moisture, or (b) Period of outside storage in rain with no pitot covers installed, or (c) Washing of airplane with no pitot covers installed. If erroneous airspeed indication(s) has (have) been observed, corrective action is required prior to next flight by draining the affected pitot line(s) and performing the 'OPERATIONAL CHECK OF PITOT LINE' in accordance with the applicable Mitsubishi MU-2 maintenance manual." (ii) Within the next 100 hours time-in-service after the effective date of this AD, modify the electrical circuitry of the pitot tube and anti-ice/deice systems in accordance with the applicable service information as follows: A. Mitsubishi Heavy Industries International (MAI) Ltd., Service Recommendation (SR) 053, Revision A, dated October 23, 1984, or B. Mitsubishi Aircraft International (MAI), Inc. SR 020/34-005, Revision B, dated May 24, 1985. (iii) Replacement of the existing pitot tube(s) with the high heat producing pitot tube(s) in accordance with paragraph (a)(1)(ii) of this AD may be delayed until September 1, 1988, if compliance with paragraph (a)(2)(i) of this AD is accomplished prior to further flight. (b) Insertion of a copy of this AD in the "LIMITATIONS" section of the AFM satisfies the requirements of paragraph (a)(2)(i)(C) of this AD. (c) The requirements of paragraphs (a)(2)(i)(A), (a)(2)(i)(B) and (b) of this AD may be accomplished by the holder of a pilot certificate issued under Part 61 of the Federal Aviation Regulations on any airplane owned or operatedby him. The person accomplishing these actions must make the appropriate aircraft maintenance record entry as prescribed by FAR 91.173. (d) Remove the temporary placard(s) and AFM textual addition required by paragraph (a)(2)(i) of this AD when the requirements of paragraph (a)(1) of this AD are accomplished. (e) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (f) An equivalent method of compliance with this AD may be used on the MHI airplanes, if approved by the Manager, Western Aircraft Certification Office, ANM-170W, Federal Aviation Administration, P.O. Box 92007, Worldway Postal Center, Los Angeles, California, 90009-2007, and on the MAI airplanes, if approved by the Manager, Wichita Aircraft Certification Office, ACE-115W, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas, 67209. All persons affected by this directive may obtain copies of the documents referred to herein upon request toBeech Aircraft Corporation (Licensee for Mitsubishi), 9709 East Central, P.O. Box 85, Wichita, Kansas 67201, or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106. This amendment supersedes Amendment 39-5006 (50 FR 8321) as amended by Amendment 39-5328 (51 FR 21515), AD 85-04-03R1. This amendment becomes effective October 6, 1986.
2000-04-15: This amendment adopts a new airworthiness directive (AD) applicable to Bell Helicopter Textron Canada (BHTC) Model 222, 222B, 222U, and 230 helicopters. This action requires inspecting the swashplate assembly drive pin (drive pin) for damage or looseness, torque testing to determine if the interference fit between the drive pin and rotating ring (ring) is adequate, and replacing any unairworthy drive pin. This amendment is prompted by an accident investigation that revealed fatigue failure of a drive pin. The actions specified in this AD are intended to prevent fatigue failure of a drive pin and subsequent loss of control of the helicopter.
93-03-02: 93-03-02 BEECH AIRCRAFT CORPORATION: Amendment 39-8501; Docket No. 93-CE-09-AD. Applicability: The following model and serial number airplanes equipped with a Bendix/King KLN-88 Loran receiver system, certificated in any category: Models Serial Numbers F33A CE-1620, CE-1640, CE-1655, CE-1659, CE-1662, CE-1680, CE-1682, and CE-1693 A36 E-2563, E-2564, E-2575, E-2582, E-2599, E-2604, E-2614, E-2631, E-2633, E-2643, E-2645, E-2650, E-2660, E-2664, E-2673, E-2675, E-2680, E-2681, E-2688, E-2690, E-2693, E-2695, E-2698, E-2711, E-2727, E-2728, E-2737, E-2738, E-2743, E-2746, and E-2752. B36TC EA-515, EA-516, EA-519, EA-522, EA-524, and EA-529. 58 TH-1581, TH-1588, TH-1592, TH-1597, TH-1598, TH-1599, TH-1600, TH-1602, TH-1612, TH-1616, TH-1621, TH-1623, TH-1627, TH-1628, TH-1632, TH-1633, TH-1640, TH-1642, TH-1646, TH-1650, TH-1651, TH-1653, TH-1661, TH-1663, TH-1668, TH-1672, and TH-1673. Compliance: Required within the next 50 hourstime-in-service after the effective date of this AD, unless already accomplished. To prevent the display of a Loran lateral deviation signal when the instrument landing system localizer frequency is selected, which could result in navigational errors, accomplish the following: (a) Modify the wiring configuration of the switching system between the Bendix/King KLN-88 Loran receiver system and the instrument landing system localizer in accordance with the ACCOMPLISHMENT INSTRUCTIONS section of Beech Service Bulletin No. 2496, dated January 1993. (b) 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. (c) An alternative method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209. The request shall be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and send it to the Manager, Wichita Aircraft Certification Office. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Wichita Aircraft Certification Office. (d) The modification required by this AD shall be done in accordance with Beech Service Bulletin No. 2496, dated January 1993. 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 the Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201-0085. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment (39-8501) becomes effective on March 12, 1993.
53-24-02: 53-24-02 HILLER: Applies to All Model UH-12, UH-12A and UH-12B Helicopters. Compliance required as indicated below. There have been several recent failures of the tail rotor pitch change rod on UH-12B helicopters, resulting in loss of directional control. This pitch change rod, P/N 25036 or 25009, must be reinforced and/or inspected as follows on all UH-12 Series helicopters. (1) Prior to the next flight and thereafter at periods not to exceed 10 hours operating time, remove the tail rotor pitch change arm and inspect the pitch change rod, P/N 25036 or 25009, in the shoulder area by dye penetrant method. If cracks are found the rod must be removed from service. If no cracks are found, the pitch change rod is considered satisfactory for an additional 10 hours operating time. On reassembling the parts determine that chamfer of hole in arm does not interfere with radius on rod. If interference is noted, enlarge chamfer on arm to clear. Reinstall arm on pitch change rod, making sure woodruff key or bolt properly locates arm and that arm bottoms on shoulder. Install washer and nut on rod and torque to 50-75 inch-pounds (UH-12 and -12A Models) or 100-150 inch-pounds (UH-12B Models). This periodic inspection must be continued until such time as the rod is reinforced as set forth in part (2). (2) When the pitch change rod is reinforced by the addition of bracket, Hiller P/N 25104, the 10-hour inspections may be discontinued. Installation of this bracket is described in Hiller Service Bulletin No. 41. The service life of this installation is limited to 300 hours from time of modification. This bracket, P/N 25104, should be installed as soon as possible but not later than January 31, 1954.
2014-20-16: We are adopting a new airworthiness directive (AD) for Brantly International, Inc. (Brantly) Model B-2, Model B-2A, and Model B-2B helicopters with certain main rotor blades. This AD requires inspecting each main rotor (M/R) blade for a crack or delamination and removing the blade if a crack exists or if the delamination exceeds certain thresholds. This AD was prompted by multiple reports of M/R blade cracks and an incident in which a crack that originated near the M/R blade trailing edge resulted in the loss of a large section of the M/R blade. The actions of this AD are intended to prevent loss of the M/R blade and subsequent loss of control of the helicopter.
91-14-02: 91-14-02 SAAB-SCANIA: Amendment 39-7044. Docket No. 91-NM-28-AD. Supersedes AD 90-16-02. Applicability: Model SF-340A series airplanes, Serial Numbers 031 through 159; and SAAB 340B series airplanes, Serial Numbers 160 through 186; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent inhibited airplane rudder control due to cracking in the vertical stabilizer top closure rib, accomplish the following: A. Prior to the accumulation of 500 hours time-in-service since new or within 100 hours time- in-service after August 30, 1990 (the effective date of Amendment 39-6676, AD 90-16-02), whichever occurs later, inspect the vertical stabilizer top closure rib for evidence of cracking, in accordance with SAAB Service Bulletin 340-55-022, Revision 1, dated February 27, 1990. B. If no evidence of cracking is found, reinspect the vertical stabilizer top closure rib for cracking at intervals not to exceed 500 flight hours time-in-service. C. If cracking is found, prior to further flight, stop drill the ends of the cracks, blend, clean, and apply aluminum tape, as specified in SAAB Service Bulletin 340-55-022, Revision 1, dated February 27, 1990. Reinspect for additional cracking and the condition of the aluminum tape at intervals not to exceed 100 hours time-in-service. D. Within one year after the effective date of this amendment, either replace the rib with a new thicker rib with a larger radius, or reinforce the rib and replace the attachment angle, in accordance with SAAB Service Bulletin 340-55-023, dated October 1, 1990. Accomplishment of either of these modifications constitutes terminating action for repetitive inspections required by paragraphs B. and C. of this AD. E. An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. F. 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 SAAB-Scania AB, Product Support, S-581.88, Linkoping, Sweden. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. Airworthiness Directive 39-7044 supersedes AD 90-16-02, Amendment 39-6676 which superseded AD 88-13-10, Amendment 39-5963. This amendment (39-7044, AD 91-14-02) becomes effective on July 29, 1991.
2000-03-09: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Cessna Model 560 series airplanes, that currently requires revising the FAA-approved Airplane Flight Manual (AFM) to provide the flightcrew with limitations, operational procedures, and performance information to be used during approach and landing when residual ice is present or can be expected. This amendment is prompted by reports indicating that, while operating in icing conditions or when ice is on the wings, some of these airplanes have experienced uncommanded roll at (or slightly higher than) the speed at which the stall warning system is activated. This amendment requires revising the AFM and revises the applicability of the existing AD. This amendment also requires modification of the stall warning system of the angle-of-attack computer. The actions specified by this AD are intended to prevent uncommanded roll of the airplane during approach and landing when residual ice ispresent or can be expected.
2014-20-01: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-600-2B16 (CL-601-3A, CL-601-3R, and CL-604 Variants) airplanes. This AD requires repetitive inspections for any fuel leak in the right-hand landing lights compartment, and related investigative and corrective actions if necessary. This AD also provides for an optional replacement of the connector of the fuel boost pump canister of the auxiliary power unit (APU), which terminates the repetitive inspections. This AD was prompted by a report of fuel leaks in the connector cavity of the APU fuel boost pump canister and at the electrical conduit connection of the APU fuel boost pump in the right- hand landing lights compartment. We are issuing this AD to detect and correct fuel leaks in the right-hand landing lights compartment, which, in combination with the heat generated by the taxi lights and landing lights on the ground reaching the auto-ignition temperature of the fuel, could result in ignition of any fuel or fumes present in the right-hand landing lights compartment.
2004-14-06: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Airbus Model A310 series airplanes, that currently requires repetitive inspections of the fuselage skin to detect corrosion or fatigue cracking around and under the chafing plates of the wing root; and corrective actions, if necessary. That AD also provides an optional terminating action for the repetitive inspections. This amendment reinstates repetitive inspections in certain areas where corrosion was detected and reworked as required by the existing AD. The actions specified by this AD are intended to detect and correct fatigue cracks and corrosion around and under the chafing plates of the wing root, which could result in reduced structural integrity of the airplane. This action is intended to address the identified unsafe condition. DATES: Effective August 13, 2004. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of August 13, 2004. The incorporation by reference of certain other publications, as listed in the regulations, was approved previously by the Director of the Federal Register as of June 3, 1998 (63 FR 23377, April 29, 1998).
83-17-05: 83-17-05 BEECH: Amendment 39-4715. Applies to Model E33C (Serial Numbers CJ-1 through CJ-25) and Model F33C (Serial Numbers CJ-26 through CJ-155) airplanes certificated in acrobatic and utility categories. \n\n\tCOMPLIANCE: Required as indicated, unless previously accomplished. \n\n\tTo prevent maneuvers which may develop into an unrecoverable spin, within the next 100 hours time-in-service after the effective date of this AD, accomplish paragraph a) or b) below. \n\n\ta)\tInstall Beech Spin Improvement Kit 33-4002-3 (E33C Serial Numbers CJ-1 through CJ-25 airplanes and F33C Serial Numbers CJ-26 through CJ-38 and CJ-40 through CJ-51 airplanes) or 33-4002-1 (F33C Serial Number CJ-39 and Serial Numbers CJ-52 through CJ-155 airplanes) in accordance with instructions contained in Beechcraft Class I Service Instruction No. 1249. \n\n\tb)\tRemove approval for operation of the airplane in acrobatic category in accordance with the following: \n\n\t\t1.\tPlace a copy of this AD in the limitations section of the Pilot's Operating Manual and FAA Approved Airplane Flight Manual. \n\n\t\t2.\tRemove the Airplane Flight Manual Supplement (Beech Part Number 33-590006-11, -13, or -17) pertaining to operation in the Acrobatic Category. \n\n\t\t3.\tCut both electrical leads to the acrobatic category fuel boost pump P/N 4140-00-191 (E33C Serial Numbers CJ-1 through CJ-25 and F33C Serial Numbers CJ-26 through CJ-148 airplanes) or P/N 1297-00-1 (F33C Serial Numbers CJ-149 through CJ-155 (airplanes) at a point near the pump motor housing. Cap and stow these leads per AC 43-13. \n\n\t\t4.\tReplace existing 3-position fuel boost pump switch P/N 35-380053-27 with 2-position switch P/N 35-380053-21, restoring electrical connections for correct operation of the auxiliary fuel boost pump P/N 4140-00-39 (E33C Serial Numbers CJ-1 through CJ-25 and F33C Serial Numbers CJ-25 through CJ-148 airplanes) or P/N 1296-00-1 (F33C Serial Numbers CJ-149 through CJ-155 airplanes). Cap and stow the unused electrical leadper AC 43-13. Secure knob P/N 96-384050-9 to the installed switch with LOCTITE 222 or equivalent bonding agent. \n\n\t\t5.\tModify the auxiliary fuel pump operation placard (including switch position placarding) to read as follows: \n\n\n\n\n\n\n\t\t6.\tObliterate the words "Remove door hold open rod prior to operation in Acrobatic Category" from the placard on the cabin door side panel. \n\n\t\t7.\tRemove the following placards on the left hand sidewall: \n\n\nDURING ACROBATIC CATEGORY \nOPERATION OCCUPANCY \nLIMITED TO PILOT'S OR PILOT'S \nAND COPILOT'S SEAT \n\n\t\tand, if installed, \n\nREMOVE THIRD & FOURTH \nPASSENGER SEATS \nPRIOR TO OPERATION \nIN ACROBATIC CATEGORY \n\n\t\t8.\tObliterate from the airplane operation limitations placards on the left side panel the heading "Acrobatic Category Airplane" and all portions of the placard under that heading. \n\n\t\t9.\tObliterate the word "AEROBATIC" located below the pilot's and co-pilot's window on the exterior of the airplane. \n\n\t\t10.\tOn ModelE33C (Serial Numbers CJ-1 through CJ-25) airplanes, obliterate from P/N 33-590003-7 Pilot's Operating Handbook and FAA Approved Airplane Flight Manual: \n\n\t\t\ti)\tOn the cover page, the words "ACROBATIC See Flight Manual Supplement" after "E33C." \n\n\t\t\tii)\tOn page 9-3, "33-590006-17 Acrobatic Bonanza E33C." \n\n\t\t11.\tOn Model F33C (Serial Numbers CJ-26 through CJ-128) airplanes: \n\n\t\t\ti)\tIn P/N 33-590009-9 Pilot's Operating Manual and FAA Approved Airplane Flight Manual: \n\n\t\t\t\tA)\tObliterate the following: \n\n\t\t\t\t\t1)\tOn Page 5-8, all information pertaining to F33C Acrobatic Category limits. \n\n\t\t\t\t\t2)\tOn Page 8-1, "Pilot's Operating Manual Acrobatic Supplement P/N 33-590006-9." \n\n\t\t\t\t\t3)\tOn Page 8-3, "33-590006-11 Acrobatic Supplement for the F33C 11/3/72." \n\n\t\t\t\tB)\tRemove P/N 33-590006-9 Pilot's Operating Manual Supplement or \n\n\t\t12.\tIn P/N 33-590009-15A1 Pilot's Operating Handbook and FAA Approved Airplane Flight Manual: \n\n\t\t\ti)\tObliterate the following: \n\n\t\t\t\tA)\tOn the cover page, the word "Acrobatic" after F33C and the words "See Flight Manual Supplement" after CJ-128. \n\n\t\t\t\tB)\tOn page 6-1: \n\n\t\t\t\t\t1)\t"Sample Weight and Balance Load Form (Acrobatic Category) - 6-17." \n\n\t\t\t\t\t2)\t"Weight and Balance Loading Form (Acrobatic Category) - 6-18." \n\n\t\t\t\tC)\tOn page 6-12, all information pertaining to F33C Acrobatic Category limits. \n\n\t\t\t\tD)\tOn page 9-3, "33-590006-11 Acrobatic Supplement for the F33C, Rev. No. 2/11/77." \n\n\t\t\t\t\tii)\tRemove pages 6-17 and 6-18. \n\n\t\t13.\tOn Model F33C (Serial Numbers CJ-129 through CJ-155) airplanes, obliterate from P/N 33-590009-13A3 Pilot's Operating Handbook and FAA Approved Airplane Flight Manual the following: \n\n\t\t\ti)\tOn page 6-10, all information pertaining to F33C Acrobatic Category limits. \n\n\t\t\tii)\tOn cover page: \n\n\t\t\t\tThe word "Acrobatic" before F33C and the words "(See Flight Manual Supplement)" after "CJ-129 and after." \n\n\t\t\tiii)\tOn page 9-3, "33-590006-13 Acrobatic Supplement for the F33C Rev. No. 2/9/78." \n\n\t\t14.\tRemove the existing Utility/Acrobatic Category Airworthiness Certificate and replace it with a new Utility Category Airworthiness Certificate as provided in Paragraph b 15 of this AD. \n\n\t\t15.\tObtain a new Utility Category Airworthiness Certificate from any FAA General Aviation District Office or Flight Standards District Office by presenting a completed FAA Form 8130-6, Application for Airworthiness Certificate, together with the removed Utility/Acrobatic Category Airworthiness Certificate, citing compliance with this AD as reason for the replacement. \n\n\tc)\tAn equivalent method of compliance with this AD may be used if approved by the Manager, Wichita Aircraft Certification Office, FAA, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 269-7000. \n\t\n\tThis amendment becomes effective October 1, 1983.