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2016-22-02: We are adopting a new airworthiness directive (AD) for Embraer S.A. Models EMB-500 and EMB-505 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 incorrect installation of passenger seat attachment fittings. We are issuing this AD to correct the unsafe condition on these products.
53-09-03: 53-09-03 HILLER: Applies to All UH-12, UH-12A, UH-12B, HTE-1, HTE-2, H-23A and H-23B Model Helicopters. Compliance required as soon as possible but no later than the next 25-hour inspection and as indicated in part C below. The following has been found necessary to prevent fatigue failure of the clevis head on the outboard tension torsion bar pins, P/N 51414-1: A. Inspect P/N 51414-1 pins to determine the fillet radius under the clevis head. If radius is less than 0.030 inch the pin must be scrapped and replaced with a pin having at least 0.030-inch radius before further flight. B. If radius is 0.030 or greater and the pin has less than 500 hours total time it may be reinstalled. If a featheredge is found at the base of the clevis it should be carefully removed. C. All 51414-1 pins must be replaced when they have accumulated a total of 500 hours of flight time. (Hiller Service Bulletin No. 36 covers this procedure.)
2016-22-07: We are superseding Airworthiness Directive (AD) 75-26-05 for Bell Helicopter Textron (Bell) Model 204B, 205A-1 and 212 helicopters. AD 75-26-05 required removing and visually inspecting each main rotor (M/R) blade and, depending on the inspection's outcome, repairing or replacing the M/R blades. This new AD requires more frequent inspections of certain M/R blades and applies to Model 205A helicopters. This AD does not require that helicopter blades be removed to conduct the initial visual inspections. We are issuing this AD to detect a crack and prevent failure of an M/R blade and subsequent loss of helicopter control.
78-26-04: 78-26-04 HUGHES HELICOPTERS: Amendment 39-3374. Applies to Hughes Model 269C, Serial Number 0004 through 0716, and Models 269A, 269A-1, and 269B certificated in all categories, equipped with LTS Tail Rotor Assembly P/N 269A6034, including Military Model TH-55A equipped with LTS Tail Rotor Assembly P/N 269A6034, equipped with Tail Rotor Hub Assembly Part No. 369A1725 or Tail Rotor Hub Assembly Part No. 369A1725-5, Serial No. 001 through 862. Also, applies to Hughes Model 369H, 369HM, 369HS, 369HE and 369D, Serial Numbers 0001D through 0324D and 0331D through 0333D certificated in all categories, equipped with Tail Rotor Hub Assembly Part No. 369A1725 or Tail Rotor Hub Assembly Part No. 369A1725-5, Serial No. 001 through 862 or Tail Rotor Hub Assembly Part No. 369A1725- 501. Compliance required as indicated. To prevent failure of Tail Rotor Hub Assembly, Hughes P/N 369A1725, 369A1725-5, Serial No. 001 through 862 and 369A1725-501, which could result in a loss of the rotorcraft, accomplish the following: (a) Within 100 hours time in service or six months after the effective date of this AD, whichever occurs sooner, unless already accomplished, inspect the tail rotor hub for cracks, corrosion or other damage and rework the tail rotor hub in accordance with Part II - Hub Inspection and Rework Paragraphs a through h of the following FAA approved Hughes Service Information Notices: (1) For Hughes Model 269 Series Helicopter use Hughes Service Information Notice No. N-153, dated September 1, 1978. (2) For Hughes Model 369H Series Helicopter use Hughes Service Information Notice No. HN-128.1, dated December 8, 1978. (3) For Hughes Model 369D Series Helicopter use Hughes Service Information Notice No. DN-27.1, dated December 8, 1978. NOTE: Hughes Service Information Notices of the above listed dates are the only versions of S.I.N.s suitable for compliance with this AD. (b) Remove all cracked hubs from service prior to further flight and replace with like serviceable part. If replacement part has part number suffix identification SP, no further action is required by this AD. If replaced with part number hub affected by this AD, revert to the inspection requirements of Paragraph (c). (c) Reinspect the tail rotor hub with a 10 power magnifying glass at each annual inspection or 300 additional hours time in service since the last inspection, whichever occurs earlier. If cracks, pits, corrosion or other damage is found, inspect and rework, per the instructions of Paragraphs (a) and (b). (d) Equivalent inspection procedures and repairs may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. This amendment becomes effective December 27, 1978.
75-08-09: 75-08-09 LYCOMING: Amendment 39-2155 as amended by Amendments 39-2260 and 39-2470 is further amended by Amendment 39-3013. Applies to O-235, O-290, O-320, IO-320, LIO-320, O-360, IO-360, HO-360, HIO-360, VO-360-A1A, VO-360-A1B, VO-360-B1A, IVO-360-A1A, LIO-360, TIO-360, AIO-360, AEIO-360, O-540 and IO-540 series Lycoming engines listed below and the same series engines overhauled/remanufactured by Lycoming between December 18, 1972, and December 10, 1974, and the same series engines overhauled after December 10, 1972, at facilities other than the manufacturer, in which the provisions of Lycoming Service Instructions No. 1272 have been incorporated. Applicable Applicable Excepted Models Serial Nos. Serial Nos. O-235 series L-11268-15 thru L-12098-15 L-12099-15,12101-15 and up and L-12100-15 O-290 series Any engine modified Any engine not modified in accordance with Lycoming in accordance with Service Instruction No. 1272. Lycoming S.I.No. 1272 O-320 series L-33329-27A thru L-41054-27A L-41055-27A and up O-320-E2D Series: L-41029-27A and up; O-320-E3D series: L-41017-27A, L-41021-27A and up. O-320-B and -D L-6809-39A thru L-6971-39A L-6972-39A and up IO-320 series L-4953-55A thru 5270-55A L-5271-55A and up LIO-320 series L-292-66A thru L-296-66A L-297-66A and up O-360, HO-360- L-17440-36A thru L-19846-36A L-19817-36A, L-19818-36A, B1A, B1B Series and L-17427 L-19847-36A and up VO-360, IVO-360 Any engine modified in Any engine not modified in Series accordance with Lycoming accordance with S.I. 1272 Lycoming S.I. 1272 HIO-360-A1A, -B1A L-10179-51A thru L-13351-51A L-12557-51A, L-12727-51A, L-12853-51A, L-12890-51A, L-13513-51A and up HIO-360-D1A L-10179-51A thru L-13512-51A L-12892-51A thru L-12894-51A, L-12919-51A, L-12966-51A thru L-12968-51A, L-12979-51A, L-13034-51A thru L-13040-51A, L-13124-51A thru L-13128-51A, L-13170-51A thru L-13174-51A, L-13257-51A thru L-13262-51A, L-13280-51A thru L-13283-51A, L-13513-51A and up HIO-360-C1A L-10179-51A thru L-13372-51A L-11578-51A, L-12193-51A, L-12445-51A, L-12763-51A, L-12845-51A, L-12847-51A thru L-12849-51A, L-12895-51A, L-12897-51A, L-12898-51A, L-12911-51A, L-12912-51A, L-12914-51A thru L-12916-51A, L-12918-51A, L-12969-51A thru L-12972-51A, L-13041-51A, L-13042-51A, L-13119-51A thru L-13123-51A, L-13142-51A thru L-13148-51A, L-13271-51A thru L-13275-51A, L-13373-51A and up HIO-360-C1B L-10179-51A thru L-13551-51A L-13352-51A and up AEIO-360 series L-10179-51A thru L-13616-51A L-13617-51A and up IO-360 series L-10146-51A thru L-13540-51A L-13541-51A and up IO-360-A1B6D L-10115-51A thru L-13529-51A L-13530-51A and up AIO-360 series L-171-63A thru L-208-63A L-209-63A and up LIO-360 series L-634-67A thru L-1059-67A L-1060-67A and up TIO-360 series L-116-64A thru L-145-64A L-146-64A and up O-540 series L-15327-40A thru L-17105-40A L-17098-40A, L-17103-40A, except O-540- L-17106-40A and up H1A5D, -H1B5D, -H2A5D, -H2B5D series IO-540 series L-10536-48 thru L-12896-48 L-10623-48, L-10624-48, except IO-540- L-10813-48, L-10814-48 K1A5D, -K1B5D, L-11246-48, L-11247-48, -K1E5D, -K1F5D L-11266-48, L-11267-48, -M2A5D, -P1A5 L-12144-48 thru L-12147-48,-S1A5, -T4A5D series L-12231-48, L-12287-48 thru L-12298-48, L-12371-48 thru L-12378-48, L-12463-48, L-12464-48, L-12636-48, L-12637-48, L-12684-48, L-12685-48, L-12711-48 thru L-12713-48, L-12726-48 thru L-12729-48, L-12734-48 thru L-12739-48, L-12744-48 thru L-12753-48, L-12806-48, L-12821-48 thru L-12823-48, L-12840-48 thru L-12844-48, L-12859-48 thru L-12868-48, L-12888-48, L-12897-48 and up Also applies to the same models and series engines overhauled/remanufactured by Lycoming between December 18, 1972 and December 10, 1974 and to any other engine in which the provisions of Lycoming Service Instruction No. 1272 have been incorporated. Compliance required as indicated. 1. For the Lycoming O-360-C2D, HO-360, HIO-360, VO-360 and IVO-360 series engines, compliance is required within the next 10 hours in service after the effective date of this AD or before the engines have accumulated 400 hours in service, whichever occurs later, unless already accomplished. 2. For the O-235, O-290, O-320, IO-320, LIO-320, O-360, IO-360, AEIO-360, AIO-360, LIO-360, TIO-360, O-540 and IO-540 series engines compliance is required within the next 50 hours in service after the effective date of this AD or before the engines have accumulated 400 hours in service, whichever occurs later, unless already accomplished. To prevent oil pump failures, inspect, replace and assemble the oil pump drive shaft and drive impeller in accordance with the inspection and procedure paragraphs of Lycoming Service Bulletin No. 381B or No. 385C or later revision approved by Chief, Engineering and Manufacturing Branch, FAA, Eastern Region. Upon submission of substantiating data thru an FAA maintenance inspector, the Chief, Engineering & Manufacturing Branch, Eastern Region, may adjust the compliance time specified in this A.D. The manufacturer's inspections and replacement 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 Avco Lycoming Division Service Department, Williamsport, Pennsylvania 17701. These documents may also be examined at the Engineering and Manufacturing Branch, Federal Aviation Administration, Eastern Region, Federal Building, John F. Kennedy International Airport, Jamaica, New York 11430. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its Eastern Region Headquarters. Amendment 39-2155 was effective April 9, 1975. Amendment 39-2260 was effective July 16, 1975. Amendment 39-2470 was effective December 24, 1975. This amendment 39-3013 is effective August 18, 1977.
70-05-03: 70-05-03 BRITISH AIRCRAFT CORPORATION: Amdt. 39-947. Applies to model BAC 1- 11, 400 series airplanes. Compliance required as indicated. To prevent fatigue failure of the main landing gear, inner door jack attachment, saddle bracket structure, located on the keel in the main landing gear bay at station 575, accomplish the following: (a) For airplanes which have not incorporated British Aircraft Corporation Modification PM2510 or PM3082 on the effective date of this AD, comply with the following: (1) Visually inspect all components of the main landing gear, inner door jack attachment, saddle bracket structure assembly which projects outboard on either side of keel structure for cracks or signs of damage in accordance with BAC 1-11 Alert Service Bulletin 53-A- PM 2510, Issue 5, dated May 16, 1969, or later ARB-approved issue, or an FAA-approved equivalent at the following intervals: (i) For airplanes with less than 2,000 landings on the effective date of thisAD, inspect prior to the accumulation of 2,350 landings, and thereafter at intervals not to exceed 350 landings from the last inspection until the accumulation of 3,500 landings and thereafter at intervals not to exceed 50 landings from the last inspection. (ii) For airplanes with from 2,000 to 3,500 landings on the effective date of this AD, unless already accomplished within the last 350 landings, inspect within the next 350 landings after the effective date of this AD and thereafter at intervals not to exceed 350 landings from the last inspection until the accumulation of 3,500 landings, and thereafter at intervals not to exceed 50 landings from the last inspection. (iii) For airplanes with 3,500 or more landings on the effective date of this AD, unless already accomplished within the last 50 landings, inspect within the next 50 landings after the effective date of this AD and thereafter at intervals not to exceed 50 landings from the last inspection. (2) Within thenext 50 landings after the effective date of this AD or before the accumulation of 5,000 landings, whichever occurs later after the effective date of this AD, unless already accomplished, and thereafter at intervals not to exceed 5,000 landings from the last replacement, replace the upper web angles P/N AK27-1639 with serviceable web angles of the same part number. Compliance with this subparagraph may be discontinued after the modifications specified in either subparagraph (3)(ii), (3)(iii) or (3)(iv) have been accomplished. (3) If cracks or damage are found during the inspections required by paragraph (a) before further flight accomplish one of the following: (i) Replace cracked or damaged components with serviceable components of the same part number. (ii) Modify the saddle bracket structure in accordance with British Aircraft Corporation Model BAC 1-11 Service Bulletin No. 53-PM-2510, Revision 4, dated October 30, 1967, or later ARB-approved issues, or an FAA-approved equivalent, and replace all cracked or damaged components not replaced by Modification PM2510 with new components of the same part number. (iii) Accomplish the modification and replacement required by subparagraph (ii) of this paragraph and in addition modify the saddle bracket in accordance with British Aircraft Corporation Model BAC 1-11 Service Bulletin No. 53-PM 3082, Revision 3, dated December 2, 1968, or later ARB-approved issues, or an FAA-approved equivalent. (iv) Accomplish the modifications and replacement required by subparagraph (ii) an (iii) of this paragraph and in addition modify the saddle bracket structure in accordance with British Aircraft Corporation Model BAC 1-11 Service Bulletin No. 53-PM 3871, dated March 31, 1969, or later ARB-approved issues, or an FAA-approved equivalent. (4) The inspections required by subparagraph (1) of this paragraph may be discontinued after the modifications specified in either subparagraph (3)(ii), (3)(iii), or(3)(iv) of this paragraph have been accomplished. (b) For airplanes which have incorporated BAC Modification PM2510 within the next 100 landings or before the accumulation of 14,000 landings after Modification PM 2510 was accomplished, whichever occurs later after the effective date of this AD, and thereafter at intervals not to exceed 14,000 landings since the last replacement, replace upper web angles P/N AK27- 10133 with new web angles of the same part number. (c) For airplanes which have incorporated the modifications required by paragraph (a)(3)(iii), within the next 100 landings or before the accumulation of 20,000 landings after Modification PM 3082 was installed, whichever occurs later after the effective date of this AD, and thereafter at intervals not to exceed 20,000 landings since the last replacement, replace the upper web angles P/N AK27-10133 with new web angles of the same part number. (d) For the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing each airplane's hours' time in service by the operators' fleet average time from takeoff to landing for the airplane type. This supersedes Amendment 39-471 (32 F.R. 12711), AD 67-25-3. This amendment becomes effective March 23, 1970.
79-06-04: 79-06-04 MOONEY: Amendment 39-3438. Applies to models M20F and M20J, registration numbers: N3207F N201A N9545M N201WX N201EU N201MF N201TN N201MR N201NW N201WT N201TS N201KJ HB-DWK N201YD N201YB N201WE N201JF OE-DHM N201TW N201ND N201MS N201GJ N201WM N201MW N201SE Compliance is required within the next 25 hours of flight time after the effective date of this Airworthiness Directive (unless already accomplished). To prevent elevator control lock, accomplish the following: 1. Inspect the pitch servo installation in accordance with Edo-Aire Mitchell Service Bulletin MB-14 dated 3/30/78 and assure servo stop bracket part number 7B1320 is installed and adjusted in accordance with SB MB-14, paragraph 5. 2. If servo stop bracket is not installed, accomplish the installation in accordance with Service Bulletin MB-14. This amendment becomes effective March 26, 1979.
75-12-11: 75-12-11 LOCKHEED: Amendment 39-2234 as amended by Amendment 39-2349. Applies to L-1011-385-1 Series airplanes, certificated in all categories with Collins FCS-110 autopilot installed. To prevent possible unwanted pitch-up while in the autoland mode, accomplish the following: (1) Effective 48 hours after receipt of this telegram, the following operating limitation applies: 'Autopilot command mode use prohibited below 100 feet AGL', and a placard must be installed in plain view of the pilots stating: 'AUTOPILOT CMD MODE USE PROHIBITED BELOW 100' AGL'. (2) Operators shall, by the most immediate and practicable means, notify flight crews of the foregoing. (3) When an operator's entire fleet, including spares, incorporates the auto flight pitch computer modification described in Lockheed Service Bulletin 093-22-080, dated August 11, 1975, or later FAA-approved revisions, remove the operating limitation and placard. Amendment 39-2234 was effective 6-12-75 for allpersons except those to whom it was made effective immediately by telegrams dated May 14, 1975. This amendment 39-2349 is effective September 2, 1975.
2002-22-04: This amendment adopts a new airworthiness directive (AD) that applies to certain Stemme GmbH & Co. KG (Stemme) Model S10-VT sailplanes. This AD requires you to modify the engine compartment fuel and oil system and firewall. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Germany. The actions specified by this AD are intended to reduce the potential for a fire to ignite in the engine compartment and increase the containment of an engine fire in the engine compartment. A fire in the engine compartment could lead to loss of control of the sailplane.
2009-05-06: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: \n\n\tThere is a possibility that during a go around procedure with a flap system failed the stall warning and the stick pusher triggering angles are anticipated reducing the margin between the real angle of attack and the stick pusher triggering angle. If the stick pusher is activated at a low altitude the pilot may be not able to recover the airplane control. Since this condition affects flight safety, an immediate corrective action is required. Thus, sufficient reason exists to request compliance with this EAD in the indicated time limit without prior notice. \n\nThis AD requires actions that are intended to address the unsafe condition described in the MCAI.
92-07-02: 92-07-02 BOEING: Amendment 39-8198. Docket No. 91-NM-186-AD. Supersedes AD 90- 08-08, Amendment 39-6572. \n\n\tApplicability: Model 737-300 series airplanes, listed in Boeing Service Bulletin 737-54- 1028, Revision 1, dated July 11, 1991, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo ensure the integrity of the engines' nacelle strut firewall seal, accomplish the following: \n\n\t(a)\tFor airplanes listed in Boeing Service Bulletin 737-54-1028, dated August 17, 1989: At the next scheduled engine removal, or within 8,000 flight hours after May 14, 1990 (the effective date of AD 90-08-08, Amendment 39-6572), which ever occurs sooner, inspect the engines' nacelle strut door assemblies for proper application of firewall sealant in accordance with Boeing Service Bulletin 737-54-1028, dated August 17, 1989, or Revision 1, dated July 11, 1991. The door assemblies are located between nacelle station 200.00 and 235.00 and attached to the underside of the strut and spar web at approximately nacelle waterline 132.00. If there are gaps, holes, or voids in the firewall sealant, apply sealant prior to further flight, in accordance with the service bulletin. \n\n\t(b)\tFor airplanes listed in Boeing Service Bulletin 737-54-1028, Revision 1, dated July 11, 1991, that are not subject to paragraph (a) of this AD: At next scheduled engine removal or within 8,000 flight hours after the effective date of this AD, which ever occurs sooner, inspect the engines' nacelle strut door assemblies for proper application of firewall sealant in accordance with Boeing Service Bulletin 737-54-1028, Revision 1, dated July 11, 1991. The door assemblies are located between nacelle station 200.00 and 235.00 and attached to the underside of the strut and spar web at approximately nacelle waterline 132.00. If there are gaps, holes, or voids in the firewall sealant, apply sealant prior to further flight, in accordance with the previously described service bulletin. \n\n\t(c)\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Seattle ACO. \n\n\t(d)\tSpecial 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. \n\n\t(e)\tThe inspections shall be done in accordance with Boeing Service Bulletin 737- 54-1028, Revision 1, dated July 11, 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 Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, DC. \n\n\t(f)\tThis amendment becomes effective April 28, 1992.
2024-17-03: The FAA is superseding Airworthiness Directive (AD) 2019-24- 16, which applied to certain Embraer S.A. Model ERJ 190-100 STD, -100 LR, -100 IGW, and -100 ECJ airplanes; and Model ERJ 190-200 STD, -200 LR, and -200 IGW airplanes. AD 2019-24-16 required revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations. Since the FAA issued AD 2019-24-16, the FAA has determined that new or more restrictive airworthiness limitations are necessary. This AD requires revising the existing maintenance or inspection program, as applicable, to incorporate new or more restrictive airworthiness limitations, as specified in an Ag[ecirc]ncia Nacional de Avia[ccedil][atilde]o Civil (ANAC) AD, which is incorporated by reference (IBR). The FAA is issuing this AD to address the unsafe condition on these products.
91-15-04: 91-15-04 MCCAULEY ACCESSORY DIVISION, CESSNA AIRCRAFT COMPANY: Amendment 39-7067. Docket No. 91-ANE-22. Applicability: McCauley Model ( )2( )34C( )-( ) Series two bladed constant speed propellers with threaded retention hubs, including those with feathering capabilities listed as follows: Affected Propeller Hub Models CONSTANT SPEED FEATHERING 2D34C8-( ) D2AF34C30-( ) 2D34C9-( ) 2AF34C55-( ) 2D34C53-( ) D2AF34C56-( ) B2D34C53-( ) D2AF34C61-( ) D2A34C58-( ) D2AF34C65-( ) F2A34C58-( ) D2AF34C81-( ) 2A34C66-( ) E2A34C70-( ) E2A34C73-( ) D2A34C78-( ) D2A34C98-( ) The parentheses used in the above list indicate the presence or absence of an additional letter(s) which vary the basic hub model designation. These letter(s) define minor changes that do not affect interchangeability or eligibility, and therefore, this AD still applies regardless of whether these letters are present or absent on the hub model designation. The abovelisted McCauley propeller hubs are found on, but not limited to, the following aircraft certificated in any category: Beech A23-24, A24, A24R, 58, 58A; 95-55, -A55, -B55,-B55A, -B55B, -C55, -C55A; D55, D55A, E55, E55A. Bellanca 17-30, 17-30A Cessna 180, 182H, 185, 185A thru D, A185E, A185F, 188, 188A, 188B, A188, A188A, A188B, 206, P206, P206A thru E, TP206A thru E, TU206A thru G, U206, U206A thru G, 207, T207, 210, 210A thru H, 210J thru L, 210-5, 210-5A, T210F thru H, T210J thru L, 305B, 305E, 310J, E310J, 310K, 310L, 310N, 336, 337, 337A thru F, M337B, T337B thru F. Fuji FA-200-180 Interceptor (AeroCommander/Meyers) 200A thru C Mooney M20C, M20D, M20G Navion A, B, D thru H Procaer F15/C Reims F337E, F337F, FT337E, FT337F Transavia PL-12/T-300 Windecker AC-7 Compliance: Required as indicated, unless previously accomplished. To prevent possible blade separation, which could result in the loss of the engine and subsequent loss of aircraft control, accomplish the following in accordance with the compliance schedule as indicated: PRIOR PROPELLER UTILIZATION (Hours/calendar months given as time-in-service) COMPLIANCE SCHEDULE OF PROPELLER INSPECTION AND MODIFICATION Greater than 900 hours, or 59 calendar months since last overhaul/penetrant inspection or installed new, or prior time-in-service unknown. Within the next 100 hours, or at the next annual inspection, or within 12 calendar months after the effective date of this AD, whichever occurs first. Less than or equal to both 900 hours and 59 calendar months since last overhaul/penetrant inspection or installed new. Prior to the accumulation of 1000 hours or 60 calendar months since last overhaul/penetrant inspection, or installed new, whichever occurs first. (a) For propellers which have incorporated a hub containing oil with red dye and have been designated at initial production as a hub model number listed in the Appendix to this AD, or prior manufactured propellers whose hubs have been modified to contain oil with a red dye and reidentified as a hub model number listed in the Appendix to this AD, compliance is required only with paragraphs (f) and (h) of this AD. (b) Perform propeller disassembly in accordance with the procedures specified for the affected hub model number listed in Paragraph 1 on page 4 of McCauley Service Bulletin (SB) 184, dated March 15, 1991. (c) Penetrant inspect the propeller assembly for cracks in the propeller blade threaded retention area, the hub blade socket threads, the retention nut threads, and the ferrule threads in accordance with the procedures specified for the affected hub model number listed in Paragraph 2 on page 5 of McCauley SB 184, dated March 15, 1991. (d) Remove from service, prior to further flight, propeller assemblies which exhibit cracks and replace with a serviceable unit, modified in accordance with paragraph (e) of this AD, or with an equivalent initial production propeller which has incorporated a hub with oil containing red dye. (e) Modify the affected propeller hub assembly to contain oil with a red dye and reidentify in accordance with the procedures specified for the affected hub model number listed in Paragraph 3 on page 6 of McCauley SB 184, dated March 15, 1991. NOTE: The modification of the propeller hub assembly to contain oil with a red dye provides an "on-condition" (in-service) means of early crack detection to prevent blade separation and also improves lubrication and corrosion protection. The oil will add approximately 2.8 lbs. to the weight of the propeller assembly. (f) If leakage of oil containing red dye is detected in service (whether during flight or while on the ground), determine prior to further flight, the source of leakage in accordance with the procedures specified for the affected hub model number listed in Paragraph 4 on page 7 of McCauley SB 184, dated March 15, 1991. If the inspection reveals a crack,compliance with Paragraph (d) of this AD is required. (g) The "calendar month" compliance times stated in this AD allow the performance of the required action prior to the last day of the month in which compliance is required. NOTE: For example, a required inspection and modification 60 months from last overhaul/penetrant inspection that was performed on December 15, 1986, would allow the penetrant inspection and modification to be performed no later than December 31, 1991. (h) Report in writing any cracks found during inspections accomplished in accordance with paragraphs (c) or (f) of this AD to the Manager, Chicago Aircraft Certification Office, within ten (10) days of the inspection. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (P.L. 96-511) and has been assigned OMB Control Number 2120-0056. (i) Aircraft may be ferried in accordance with the provisions of Federal Aviation Regulations (FAR) 21.197 and 21.199 to a base where the AD can be accomplished. (j) Upon submission of substantiating data by an owner or operator through an FAA Inspector (maintenance, avionics, or operations, as appropriate) an alternate method of compliance with the requirements of this AD or adjustments to the compliance times specified in this AD may be approved by the Manager, Chicago Aircraft Certification Office, Small Airplane Certification Directorate, Aircraft Certification Service, FAA, 2300 East Devon Avenue, Des Plaines, Illinois 60018. The disassembly, inspection, and modification shall be done in accordance with the procedures listed in McCauley SB 184, dated March 15, 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 McCauley Accessory Division, The Cessna Aircraft Company, 3535 McCauley Drive, Vandalia, Ohio 45377. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, 12 New England Executive Park, Room 311, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, DC. Airworthiness Directive 91-15-04, Amendment 39-7067, supersedes, AD 77-17-09, Amendment 39-3020, AD 77-20-03, Amendment 39-3044, AD 77-23-01, Amendment 39-3073, AD 77-24-04, Amendment 39-3086, AD 78-20-01, Amendment 39-3304. This amendment (39-7067, AD 91-15-04) becomes effective on August 7, 1991. AD 91-15-04 APPENDIX OIL-FILLED PROPELLER HUB COMPLIANCE INDICATOR TABLE Propeller Hub Model Compliance Indicator Propeller Hub Model Compliance Indicator + 2D34C8 2D34C8-()P and/or oil-fill plug in side of hub F2A34C58 F2A34C58-()0 and/or oil-fill plug in side of hub 2D34C9 2D34C9-()P and/or oil-fill plug in side of hub D2AF34C61 D2AF34C61-()0 and/or oil-fill plug in side of hub D2AF34C30 D2AF34C30-()P and/or oil-fill plug in side of hub D2AF34C65 D2AF34C65-()0 and/or oil-fill plug in side of hub B2D34C53 B2D34C53-()0 and/or oil-fill plug in side of hub 2A34C66 2A34C66-()P and/or oil-fill plug in side of hub 2D34C53 2D34C53-()0 and/or oil-fill plug in side of hub E2A34C70 E2A34C70-()P and/or oil-fill plug in side of hub 2AF34C55 2AF34C55-()0 and/or oil-fill plug in side of hub E2A34C73 E2A34C73-()P and/or oil-fill plug in side of hub D2AF34C56 D2AF34C56-()0 and/or oil-fill plug in side of hub D2A34C78 D2A34C78-()P and/or oil-fill plug in side of hub D2A34C58 D2A34C58-()0 and/or oil-fill plug in side of hub D2AF34C81 D2AF34C81-()0 and/or oil-fill plug in side of hub D2A34C98 D2A34C98-()0 and/or oil-fill plug in side of hub +Propeller models are listed in numerical sequence following the letter C in the model designation.
71-07-04: 71-07-04 CESSNA: Amdt. 39-1185. Applies to Models 401 (Serial Numbers 401-0041 and up), 401A (all Serial Numbers), 401B (Serial Numbers 401B0001 thru 401B0052), 402 (Serial Numbers 402-0041 and up), 402A (all Serial Numbers), 402B (Serial Numbers 402B0001 thru 402B0030), 411 and 411A (Serial Numbers 411-0235 and up) airplanes. Compliance: Required as indicated, unless already accomplished. To assure safe operation of the emergency exit window, accomplish the following: Within 50 hours' time in service after the effective date of this AD, modify the emergency exit window in accordance with Cessna Service Letter No. ME70-30 dated August 21, 1970, and Cessna Service Kit SK402-23 dated June 15, 1970, or by any equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. NOTE: The installation of Cessna Service Kit SK402-23 does not require pulling the emergency release handle. If handle is inadvertently pulled, assure that existing emergency exit window retention pins are properly installed. This amendment becomes effective April 6, 1971.
2024-17-01: The FAA is superseding Airworthiness Directive (AD) 2021-11-17 and AD 2021-11-22, which applied to all Airbus Helicopters Deutschland GmbH (AHD) Model EC135P1, EC135P2, EC135P2+, EC135P3, EC135T1, EC135T2, EC135T2+, EC135T3 and EC635T2 helicopters. AD 2021-11-17 required a one-time visual inspection of certain part-numbered main rotor actuators (MRAs). AD 2021-11-22 required revising the life limits of certain parts and removing each part that had reached its life limit. Since the FAA issued those ADs, it was determined that repetitive inspections of the MRAs are necessary, new and more restrictive tasks and limitations have been issued, and that it is necessary to expand the applicability. This AD continues to require the actions required by AD 2021-11-17 and AD 2021-11-22, except this AD requires changing the one-time MRA inspection to a repetitive inspection and incorporating other new and more restrictive tasks and limitations by revising the airworthiness limitations section (ALS) of the existing helicopter maintenance manual or instructions for continued airworthiness and the existing approved maintenance or inspection program, as applicable. This AD also expands the applicability by adding Model EC635T2+ helicopters. These actions are specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2016-21-03: We are adopting a new airworthiness directive (AD) for Airbus Helicopters Model MBB-BK 117 C-2 helicopters with certain duplex trim actuators installed. This AD requires repetitively inspecting the lateral and longitudinal trim actuator output levers for correct torque of the nuts. This AD was prompted by a design review that indicated the attachment screws can become loose under certain circumstances. These actions are intended to prevent the loss of an attachment screw, which could result in movement of the output lever in an axial direction, contact of a bolt connecting the control rod to an output lever with the actuator housing, and subsequent loss of helicopter control.
2004-09-11: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 767-200, -300, and -300F series airplanes, that requires performing, for both main landing gear (MLG), gap measurements of the upper and lower joint gaps; an ultrasonic inspection of the outer cylinder of the MLG for cracks between the downlock fitting attach lugs; and follow-on and corrective actions if necessary. This action is necessary to detect and correct cracks in the outer cylinder of the MLG, which could result in collapsed MLG and consequent reduced controllability of the airplane during takeoff and landing. This action is intended to address the identified unsafe condition.
2004-09-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB SF340A and SAAB 340B series airplanes, that requires relocating the most outboard latch in the right-hand leading edge of the refueling panel, and sealing of the original latch-mounting cutout. This action is necessary to prevent wear of the signal conditioner wiring harness behind the refueling panel, which could result in a short circuit and consequent smoke or fire behind the refueling panel. This action is intended to address the identified unsafe condition.
2002-14-17: This amendment supersedes an existing airworthiness directive (AD), applicable to Eurocopter Deutschland GmbH (ECD) (Eurocopter) Model EC135 helicopters, that currently requires initial and repetitive visual inspections and a one-time dye-penetrant inspection of a certain main rotor hub shaft (shaft) for cracks, and replacement of any cracked shafts. This amendment requires the same actions as the existing AD, but corrects the shaft part number (P/N) in the current AD, includes additional P/N's, increases the area of inspection from a 40mm area to a 50mm area of the shaft, and provides an option for using either a visual or dye-penetrant inspection to satisfy the repetitive inspection requirement. This amendment is prompted by the need to correct the shaft part number and increase the area of inspection, as well as add additional affected shaft P/N's. The actions specified by this AD are intended to detect fatigue cracks in the shaft that could lead to shaft failure and subsequent loss of control of the helicopter.
73-12-03 R1: 73-12-03 R1 ROLLS ROYCE LTD: Amendment 39-1652 as amended by Amendment 39-4277. Applies to Rolls-Royce DART engine models 542-10, 542-10J, and 542-10K which do not have Rolls-Royce, Ltd., Modification No. 1681 and Modification No. 1768 incorporated, as installed on, but not limited to, NAMC YS-11 airplanes. Compliance is required as indicated. To detect engine mounting foot cracks, and engine mounting foot stud fractures, looseness, and loose nuts, accomplish the following: (a) Within the next 400 hours time in service after the effective date of this AD, or within 400 hours time in service since the last inspection, whichever occurs sooner, and thereafter at intervals not to exceed 400 hours time in service since the last inspection, inspect engine top mounting feet which do not incorporate Modification 1681 and side mounting feet which do not incorporate Modification 1768 for cracks and the mounting feet studs for fracture, looseness, and loose nuts, in accordancewith Rolls-Royce Service Bulletin Da72-384, Revision 2, dated September 1979. (b) If any engine mounting feet are found cracked or any engine mounting feet studs are found to have loose nuts, or to be loose or fractured during an inspection required by paragraph (a), before further flight, repair in accordance with Rolls Royce Service Bulletin Da72-384 Revision 2, dated September 1979 or an FAA- approved equivalent. NOTE: Rolls-Royce DART Service Bulletins Da72-411 and Da72-444 concern Modification No. 1681 and Modification No. 1768, respectively. Amendment 39-1652 became effective July 1, 1973. This amendment 39-4277 becomes effective December 10, 1981.
2016-20-02: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-300, -400, and -500 series airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) indicating that the aft pressure bulkhead is subject to widespread fatigue damage (WFD). This AD requires repetitive inspections of the aft pressure bulkhead web for any cracking, incorrectly drilled fastener holes, and elongated fastener holes; and related investigative and corrective actions, if necessary. We are issuing this AD to detect and correct fatigue cracking of the aft pressure bulkhead web at the ''Y''-chord, which could result in reduced structural integrity of the airplane and rapid decompression of the fuselage.
59-08-02: 59-08-02\tBOEING: Applies to all Model 707-100 Series Aircraft with Bendix Flux Gate Compass system. \n\tCompliance required as indicated. \n\tReports have indicated that excessive indicator errors can be introduced in the Bendix remote compass system when it is in the slaving mode and when the aircraft is exposed to wing oscillations resulting from rough air, spoiler operation, etc. As an interim safety measure pending further investigation and development by the manufacturer of an improved Bendix flux valve installation, the following aircraft operating limitation is required for all aircraft incorporating the Bendix system: \n\tWhenever heading information is required by the pilots, the two Bendix systems (two RMI and two CDI indicators) should be frequently checked against each other and against the magnetic standby compass to detect obvious errors in indication. This information must be included, within the next 10 hours' flight time, on a placard adjacent to the RMI and CDI indicators until an approved Airplane Flight Manual revision covering this subject is available. The foregoing limitation does not apply to aircraft in which the magnetic sensing device (flux gate) is installed in the fuselage body by the manufacturer or installed in accordance with Boeing 707 Service Bulletin No. 587 "Bendix Compass System Flux Gate Relocation," which describes a satisfactory and approved means of mounting the flux valves.
52-14-01: 52-14-01 DOUGLAS: Applies to All DC-6, DC-6A and DC-6B Airplanes With Hamilton Standard Propellers Except as Otherwise Indicated. \n\n\tItems I and II are to be accomplished by means of a progressive modification program to be submitted to and approved by the FAA. This program shall begin no later than August 1, 1952, and shall be completed no later than August 1, 1953. \n\n\tI.\tIn order to prevent inadvertent actuation of the propeller reversing solenoid valves, protect the reversing solenoid circuits from all other electrical circuits and protect the reversing solenoid circuits from each other. This is to be accomplished in accordance with attachment A and the following instructions which pertain to some of the specific features to be considered in isolation of the reversing circuits from other circuits. Other features which are not specifically referred to in this list shall be treated in an equivalent manner: \n\n\t\tA.\tModify the following multiple pin connector assemblies as specified in item 2 of attachment A (See AD 52-13-02 Lockheed for Attachment A): \n\n\t\t\t1.\tFirewall connector (if the reversing solenoid lead has not already been removed). \n\n\t\t\t2.\tConnector at the front of the control pedestal. \n\n\t\t\t3.\tConnector at Hamilton Standard relay box (if used). \n\n\t\tB.\tModify the following terminal strips as specified in item 1 of Attachment A: \n\n\t\t\t1.\tFirewall junction box terminal strip (if used). \n\n\t\t\t2.\tTerminal strip at synchronizer compartment (if used). \n\n\t\t\t3.\tTerminal strip within propeller control box located behind pilot's seat. \n\n\t\tC.\tProtect the following exposed terminals as specified in item 1(c) of Attachment A: \n\n\t\t\t1.\tExposed terminals at secondary throttle lock relays located behind pilot's seat. \n\n\t\t\t2.\tExternal A2 and A3 terminals on "C" relays in propeller control box behind pilot's seat. \n\n\t\tD.\tHamilton Standard reversing relay box (if used): Reversing solenoid circuit relay contacts, etc., to be shielded from all other circuits which are energized at any time except when reversing is desired. Reversing relay boxes which have separate pin connectors for the reversing solenoid wire and the remaining circuits shall be so installed that it will not be possible inadvertently to interchange any connectors on any two relay boxes. \n\n\t\tE.\tPressure seal disconnect: Modify in one of the following ways: \n\n\t\t\t1.\tBypass the pressure seal by using continuous wiring. \n\n\t\t\t2.\tProvide a separate connector for the reversing solenoid lead. \n\n\t\t\t3.\tPins adjacent to the reversing solenoid pins shall be deactivated or used only in circuits which cannot provide sufficient energy to activate the reversing solenoid or circuits which are energized only when reversing is desired. Also, provide an insulating shield for the reversing solenoid pins on both sides of the pressure seal. The reversing solenoid wire shall be secured to this shield or in some equivalent manner to prevent the wire from falling free in casethe terminal pin comes out of the socket. \n\n\t\t\t4.\tIsolate the solenoid valve lead insert on both sides of the seal by covering both the socket and the wire terminal with insulation which will cover all exposed metal parts when the wire is in place, when the wire terminal has come out of the socket, or when the wire is broken at any point up to the point at which it is secured in the bundle. The nature of the insulation or the provisions for securing it in place must be such that its installation will not be overlooked during maintenance. \n\n\t\tF.\tReversing solenoid circuit wiring: Modify in accordance with item 4 of attachment A. \n\n\t\tG.\tOther circuit modifications: \n\n\t\t\t1.\tAll airplanes with fuselage numbers below 233 shall be modified to comply with Hamilton Standard Service Bulletin No. 221. \n\n\tII.\tReverse solenoid lock assembly: \n\n\t\tA.\tComply with Douglas Service Bulletin DC-6, NO. 356, dated March 9, 1949, to prevent excessive deflection of lock assembly components.B.\tThe "Reverse operable" warning device shall be clearly visible when the lock is open just a sufficient amount to permit pulling the throttles into the reverse regime. \n\n\tIII.\tMaintenance practices (to be instituted not later than August 1, 1952): \n\n\t\tA.\tAt each nearest scheduled service to 350 hours: \n\n\t\t\t1.\tInspect all points specified in items IB, IC, and IE. These inspections may be discontinued if the modifications made to the system are of the type described in item E1 or E2; item 1A or 1B of attachment A and item 2A or 2B of attachment A. \n\n\t\tB.\tAt any time that an electrical fault occurs in a circuit which is carried in the same bundles or the same conduits as the reversing solenoid circuit, representative terminal points in the faulty circuit are to be inspected to determine whether any damage may have occurred within the bundles or conduit. If there is evidence of possible damage, all the wiring involved is to be removed and inspected. Damaged wiring is to bereplaced as necessary. \n\n\t\tC.\tAt each nearest scheduled service to 350 hours, perform an electrical check of the reverse safety switches in the pedestal assembly to assure that the switch is open when the throttles are moved forward out of the reverse position, unless it is shown that failure of any of the reverse safety switches to open will be clearly apparent to the flight crew by reason of improper operations of the propeller control system. Because of the many technical considerations involved, analyses showing that the objective of this revision has been accomplished should be referred to the FAA for engineering evaluation and approval. \n\n\t\tD.\tConduct the mechanical functional test specified in AD 50-16-01 at each nearest scheduled service to 350 hours. \n\n\tIV.\tOperating instructions: Comply with item 5 of attachment A. \n\n\tV.\t(NOTE: Propeller governor design changes, which are under development and whose purpose is to provide a high pressure hydraulic circuit bypass tosafeguard against inadvertent reversing and ability to feather even when the reversing solenoid is energized, are still under consideration and may be the subject of a future directive.)
2024-17-06: The FAA is adopting a new airworthiness directive (AD) for all BAE Systems (Operations) Limited Model BAe 146 and Avro 146-RJ series airplanes. This AD was prompted by a report of cracking on the radius of the rib 0 forward longeron at a certain frame. This AD requires a one-time inspection for defects of the radius, and repair if necessary. The FAA is issuing this AD to address the unsafe condition on these products.
2024-16-12: The FAA is adopting a new airworthiness directive (AD) for certain Airbus SAS Model A330-243, -302, -343, and -941 airplanes. This AD was prompted by a determination that a certain aft bulkhead cover panel may have been made with a non-conforming material. This AD requires replacing the aft bulkhead cover panel and prohibits the installation of affected parts, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.