Results
97-26-02: This amendment adopts a new airworthiness directive (AD) that is applicable to Eurocopter Deutschland GmbH (ECD) (Eurocopter Deutschland) Model BO-105A, BO-105C, BO- 105S, BO-105LS A-1, and BO-105LS A-3 helicopters; and Eurocopter Canada Ltd. Model BO-105LS A-3 helicopters. This action requires visual inspections for cracks in the ribbed area of the main rotor mast flange (flange). This amendment is prompted by a report of an operator discovering a crack in the flange after experiencing in-flight vibrations. The actions specified in this AD are intended to detect cracks in the flange, which, if not detected, could result in failure of the flange and subsequent loss of control of the helicopter.
97-26-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Boeing Model 737-100, -200, -300, -400, and -500 series airplanes, that requires repetitive inspections to detect galling on the input shaft and bearing of the standby rudder power control unit (PCU), and replacement of the standby rudder actuator with a serviceable actuator, if necessary. This amendment also requires eventual replacement of the input bearing of the standby PCU with an improved bearing, which constitutes terminating action for the inspections to detect galling. This amendment is prompted by a review of the design of the flight control systems on Model 737 series airplanes. The actions specified by this AD are intended to prevent galling on the input shaft and bearing of the standby PCU, which could result in uncommanded movement of the rudder or increased pedal forces. These conditions, if not corrected, could result in reduced controllability of the airplane.
89-04-01 R3: 89-04-01 R3 SIKORSKY AIRCRAFT: Amendment 39-6131, as revised by Amendment 39- 6279 and 39-6340, is further revised by Amendment 39-6451. Docket No. 88-ASW-56. Applicability: Model S-61N and S-61NM helicopters, certificated in any category. Compliance: Required as indicated, unless already accomplished. To preclude possible fracture of the sponson truss assembly components, accomplish the following: (a) Within the next 100 hours' time in service after the effective date of this AD, conduct a hardness test of each welded lug of sponson truss tube assemblies, Part Number (P/N) S6125-51212-4 and 61250-51233-042, aft lower truss tube assembly-left side; S6125- 51212-5 and 61250-51233-043, aft lower truss tube assembly-right side; S6125-51214-3 and 61250-51235-041, forward upper truss assembly-left and right side; and S6125-51214-4 and 61250-51235-042, aft upper truss tube assembly-left and right side as follows: NOTE: Tube assemblies with single piece end fittings do not require a hardness check of the lug. (1) Remove paint from lugs using a clean cloth dampened with suitable paint remover. NOTE: Paint remover in compliance with MIL-R-81294 is acceptable. (2) Rinse lugs with fresh water and wipe dry. (3) Using a portable calibrated Rockwell hardness tester such as Wilson Model M-51, or FAA-approved equivalent, conduct a hardness test on any exposed area of the lug that is inboard of the edge and away from the weld, using the Rockwell C-scale. Repeat the test three times. Maintain a distance of at least three penetration diameters between tester indentations. NOTE: Removal of the truss component from the helicopter for the hardness test is not required; however, removal is recommended to improve consistency of hardness readings obtained. (4) If the variation of hardness readings exceeds a three-point spread on the Rockwell C-scale, repeat the test until any three readings have a variation within a three-point spread. (5) Determine the average of these three readings and mark the average value of three Rockwell C-scale hardness readings on the welded lug to indicate the test has been done. Use a vibro-peen pencil or equivalent FAA-approved marking method. (6) Dispose of the components as follows: (i) If the lug average hardness is less than Rockwell C34, replace the component with a serviceable part before further flight. (ii) If the lug average hardness is in the proper range, Rockwell C39 to C43, the component is serviceable provided the component is inspected for cracks in accordance with the requirements of paragraph (b) of this AD. (iii) If the lug average hardness is a range of Rockwell C34 to less than C39 or above C43, the part is serviceable provided the component is inspected for cracks as prescribed in paragraph (b)(1) through (9) of this AD at intervals not to exceed 500 landings and removed from service not later than June 30, 1989.(7) Repaint cleared areas with primer coating and paint, and reinstall jacking pad if removed for lug inspection. NOTE: Sikorsky Aircraft Alert Service Bulletin No. 61B25-15, Part IIA, pertains to the hardness test required by this AD. (b) Prior to the accumulation of 1,000 landings after the effective date of this AD, and thereafter at intervals not to exceed those landing intervals stated in Table 1, inspect the sponson truss tube assemblies for cracks in the locations noted in the table as follows: TABLE 1 INSPECTION SCHEDULE AND LOCATIONS Paragraph (b) Inspection Interval, Number of Landings Sponson Truss Tube Assembly P/N Inspection Locations 2500 S6125-51212-1 61250-51233-041 Inboard & outboard tube-to-fitting welds and clevis. Two welded manufacturing holes S6125-51212-4, -5 S6125-51214-3, -4 61250-51233-042, -043 61250-51235-041, -042 Inboard & outboard tube- to-fitting welds Two welded manufacturing holesS6125-51213-1, -041 Inboard tube-to-fitting weld 61250-51234-041 Welded manufacturing hole 4700 S6125-51212-4, -5 Lug-to-fitting weld S6125-51214-3, -4 Lug hole 61250-51233-042, -043 Lug hole 61250-51235-041, -042 Lug-to-fitting weld (if applicable) S6125-51217-1, -041 Clevis lug hole (1) Raise the helicopter using jacks to unload the trusses. NOTE: Jacking is described in Sikorsky Maintenance Manual, SA 4045-80, Section 7-2- 0. (2) Remove the left and right upper and lower truss tube assemblies, and diagonal brace assemblies from the helicopter sponsons. NOTE: Sponson removal is described in Sikorsky Maintenance Manual, SA 4045-80, Section 32-10-1. (3) Remove paint from lugs, fittings, and welded manufacturing holes indicated in Table 1 using a clean cloth dampened with suitable paint remover (ref. paragraph (a)(1) NOTE.) Rinse with fresh water and dry. (4) Inspect the cleanedareas identified in Table 1, using a fluorescent penetrant or equivalent inspection method. (5) If cracks are found, replace the component with a serviceable part prior to further flight. (6) If no cracks are found, repaint cleaned areas with suitable primer coating and paint. (7) Reassemble and install the sponson. (8) Seal pockets, joints, rivets, bolts, nuts, and inspection holes of all tube and brace assemblies with brush-type sealing compound such as PR-1440, Class A, Products Research and Chemical Corporation, or Pro-Seal 890, Class A, Essex Chemical Co., or FAA-approved equivalent, after reassembly of landing gear. (9) Remove jacks. NOTE: Sikorsky Alert Service Bulletin No. 61B25-15, Part III A (3), (4), (6), (7), (8), and (9) pertains to the inspections required by this AD. An alternate method of compliance or adjustment of the compliance schedule which provides an equivalent level of safety, may be used if approved by the Manager, Boston Aircraft Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone (617) 273-7118. Substantiating data compiled by an owner or operator must be submitted through the cognizant FAA Aviation Safety Inspector to support and recommend the request for compliance time adjustment. This amendment further amends Amendment 39-6131 (54 FR 6512; February 13, 1989), AD 89-04-01, as amended by Amendment 39-6279, AD 89-04-01 R1, (54FR 31505; July 31, 1989) and by Amendment 39-6340, AD 89-04-01 R2, (54 FR 40639; October 3, 1989). This amendment (39-6451, AD 89-04-01 R3) becomes effective on February 5, 1990.
85-04-01 R1: 85-04-01 R1 FAIRCHILD: Amendment 39-5005 as amended by Amendment 39-5138. Applies to Models SA226-T, SA226-T(B), SA226-AT (all serial numbers) and Model SA226-TC (all serial numbers below S/N TC398) airplanes certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent cockpit fires and hydraulic failures by increasing the fatigue resistance of certain hydraulic lines and replacing nonmetallic oxygen lines with metal lines, accomplish the following: (a) Within the next 25 hours time-in-service, after the effective date of this AD on airplanes not previously inspected per AD 83-19-02 or within 200 hours time-in-service after the last inspection per AD 83-19-02 and within each 200 hours time-in-service thereafter: (1) Visually inspect the hydraulic and oxygen lines for leakage in the vicinity of the side panel and behind the instrument panel on both sides of the aircraft. Apply maximum pilot effort to the brakepedals while inspecting the brake lines. (2) Visually inspect, in the forward-pressure bulkhead area, for hydraulic fluid contamination from the brake reservoir vent. (3) Before further flight, clean or replace any hydraulic fluid-contaminated structure, material or equipment and replace any lines or tubing which leak or have stress cracks which could cause future leaks found during inspection per paragraphs (a)(1) and (2). NOTE: Follow FAA Advisory Circular 43.13-1A, Chapter 10, paragraph 393, and Chapter 8, paragraph 363, when accomplishing these inspections and corrective action required by paragraphs (a)(1), (2), and (3). (4) Visually inspect the electrical wires in the vicinity of the cockpit side panels and behind the instrument panel on both sides of the aircraft for contact or inadequate clearance between the wires and adjacent components, especially hydraulic and oxygen lines. Determine that the wire bundles near the rudder pedals have adequateseparation with the pedals in their extreme positions. Prior to further flight, add additional supports or reroute, as necessary, to prevent wire contact or chafing which may damage the wire insulation, and clean any contamination from the bundles. NOTE: Follow FAA Advisory Circular 43.13-1A, Chapter 11, when accomplishing these inspections and corrective action. (b) Within the next 25 hours time-in-service after the effective date of this AD unless already accomplished: (1) Inspect wires and wire terminations within and below the generator control junction box (J-box) and install phenolic insulator on side of J-box and spiral wrap on wires in accordance with Fairchild Service Bulletin 24-021 dated March 21, 1983. (2) On Aircraft Models SA226-T (S/Ns T201 through T287); SA226-AT (S/Ns AT001 through AT066); and SA226-TC (S/Ns TC201 through TC247) in which MIL-H- 5606 hydraulic fluid is used, drain and purge the main hydraulic and brake system reservoirs and refill these reservoirs with MIL-H-83282 fluid. Change the placards on both reservoirs to specify MIL-H-83282 fluid. (c) On or before December 31, 1985, modify Model SA226 airplanes in accordance with the following: (1) On Models SA226-T (S/N's T201 through T275, T277 through T291), SA226-T(B) (S/N's TB276, TB292 through TB417), SA226-AT (S/N's AT001 through AT074), and SA226-TC (S/N's TC201 through TC397) airplanes, modify the hydraulic system in accordance with Fairchild S/B 226-29-005 revised July 19, 1985. NOTE: Inspect condition of cushion clamps when changing hydraulic lines and change as necessary. (2) On Models SA226-T (S/N's T249 through T275, T227 through T291), SA226-T(B) (S/N's TB276, TB292 through TB417), SA226-AT (S/N's AT001 through AT074), and SA226-TC (S/N's TC201 through TC397) airplanes, modify the crew oxygen system in accordance with Fairchild S/B 226-35-003 revised July 19, 1985. (d) On airplanes modified in accordance with paragraph (c) of this AD. (1) If equipped with an anti-skid brake system, compliance with inspections required by paragraphs (a) (1), (2), and (4) of this AD is not required. (2) If not equipped with an anti-skid brake system, compliance with inspections required by paragraph (a) (1) and (4) of this AD is not required. (e) Airplanes may be flown in accordance with FAR 21.197 to a location where this AD may be accomplished. (f) The intervals between repetitive inspections required by this AD may be adjusted up to 10 percent of the specified interval to allow accomplishing these inspections concurrent with other scheduled maintenance of the airplane. (g) An equivalent method of compliance with this AD may be used when approved by the Manager, Aircraft Certification Division, ASW-100, Southwest Regional Office, FAA, Fort Worth, Texas 76101; Telephone (817) 624-4911, Extension 511. This AD supersedes AD 83-19-02. Amendment 39-5005 became effective on March 30, 1985. This Amendment 39-5138 becomes effective October 28, 1985.
2012-08-01: We are adopting a new airworthiness directive (AD) for Sikorsky Aircraft Corporation (Sikorsky) Model S-92A helicopters. This AD was prompted by the manufacturer's analysis of engine data that revealed the data was inaccurate in dealing with available above specification engine power margin. This AD requires revising the Operating Limitations section of the Sikorsky Model S-92A Rotorcraft Flight Manual (RFM). The actions are intended to prevent the use of inaccurate engine performance data in calculating maximum gross weight by revising the Operating Limitations section of the RFM.
97-25-09: This amendment adopts a new airworthiness directive (AD) that is applicable to Allison Engine Company Model 250-C40B turboshaft engines. This action requires installation of a placard requiring pilots to record torque level and time in service operating above 86% engine torque until the defective parts have been replaced, no later than December 31, 2000, or when certain maintenance actions are accomplished, or when certain operational restrictions are exceeded, whichever occurs earliest. This amendment is prompted by a report from Allison Engine Company of a manufacturing defect in certain helical power takeoff gearshaft assemblies, identified by serial numbers. The actions specified in this AD are intended to prevent fatigue failure of the helical power takeoff gearshaft assembly, which could result in a loss of engine power and inflight engine shutdown.
78-08-08: 78-08-08 BELL: Amendment 39-3189. Applies to Bell Models 214B and 214B-1 helicopters, S/N 28001 through 28024 and S/N 28032 through 28036, certificated in all categories. Compliance required as indicated. To detect possible cracks in the tail fin forward spar, to improve the strength and fatigue resistance of the tail fin and boom and to preclude possible failure of the fin, accomplish the following repetitive inspection and modification. (a) Before the first flight of each day after the effective date of this AD, conduct the following inspection of the tail fin forward spar, left side cap angle, and the spar web until the modification in paragraph (b) is accomplished: (1) Remove the 42 degree gear box cover from the tail boom. (2) Remove the paint finish and clean the spar left side cap angle and left forward side of the spar web in the area surrounding the rivets where the fin intersects the tail boom, using a cloth and methyl-ethyl ketone or equivalent.(3) Inspect the cap angle and web in the clear area of the spar for cracks using a three-power or higher magnifying glass and a light, or a dye penetrant or equivalent inspection method. (4) If no cracks are found, install the gear box cover. (5) If cracks are found, remove the tail boom before further flight and replace with a tail boom and fin modified in accordance with the data specified in paragraph (b) of this AD. (b) On or before July 30, 1978, modify the tail boom and fin in accordance with Bell Helicopter Textron Drawing No. 214-961-151 or in accordance with an equivalent means approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southwest Region. (c) The manufacturer's instructions identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Textron, P.O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. (Bell Helicopter Textron Service Bulletin No. 214-77-8 pertains to inspections and Service Bulletin No. 214-78-1 pertains to modifications.) This amendment becomes effective April 21, 1978.
64-10-03: 64-10-03 DAVIS AIRCRAFT PRODUCTS: Amdt. 39-856. Applies to All Aircraft Equipped With Davis Aircraft P/N FDC-2700 Series Safety Belts Manufactured Prior to September 1, 1963. Compliance required within 125 hours' time in service after the effective date of this AD. As the result of the investigation of reports of the failure of Davis Aircraft P/N FDC-2700 Series safety belts to latch properly, it has been determined that binding of the latching mechanism has resulted from the loss of lubricant during belt cleaning processes and also from certain manufacturing and assembly errors. (a) Inspect each belt assembly as follows: (1) Determine whether the buckle cover P/N FD-2674 is fully closed. If the buckle cover is fully opened and does not automatically snap back into the completely closed position when released, it may be assumed that the spring is damaged. (2) Slowly raise the buckle cover taking careful note of any tendency of binding of the latching mechanism components. The buckle cover P/N FD-2674 and release latch P/N FD-2668 must be free to snap into the fully closed position when the buckle cover is opened and released. (b) If any of the deficiencies specified in (a) are found, disassemble and further inspect the components to determine whether they meet the following requirements: (1) The tangs or straight ends of the coil spring P/N FD-3007 shall be 9/32 plus 0- 1/16 inch and be straight throughout this length. (A bent tang will result in binding of the latching mechanism.) (2) The release latch P/N FD-2668, shall be counterbored in one end to receive spring P/N FD-3007. The counterbored depth shall be 15/32 plus or minus 1/32 inch. (c) Replace any components found to be defective under the inspection required in paragraph (b). (d) Clean latch components and relubricate hexagonal headed hinge bolt and spring of the latching mechanism as required using Alemite No. 33 lubricant or equivalent.(e) After reassembly inspect the spring to determine that it is seated in the release latch retaining groove and apply spring tension by rotating the hexagonal headed hinge bolt from the unloaded position through two to three flats of the hexagonal head. Secure the hinge bolt in the hexagonal cutout in the buckle frame. Effective June 5, 1964. Revised October 8, 1969.
2012-06-20: We are adopting a new airworthiness directive (AD) for certain Fokker Services B.V. Model F.28 Mark 0070 and 0100 airplanes. This AD was prompted by a report that the fuel crossfeed valves cannot be controlled when only emergency electrical power is available, that an unwanted configuration of the indication logic for the fuel fire shutoff valve was introduced during production, and that current fuel crossfeed indications are based on selection by the flightcrew instead of actual position of the crossfeed valve actuators. This AD requires modifying the crossfeed valve control and power supply, the crossfeed indication logic and power supply, and the indication logic for the fuel fire shutoff valve; modifying the overhead panel; and for certain airplanes, modifying the transfer logic of the center wing fuel tank. We are issuing this AD to prevent failure of an in-flight engine re- light following a double engine flame-out event, which could result in loss of the airplane.
2010-09-02: We are adopting a new airworthiness directive (AD) for the products listed above. 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: BAE Systems have received three reports of uncommanded flap extensions affecting different Jetstream 31 aeroplanes. In one instance, the aeroplane exceeded the airspeed limit allowed for the uncommanded flap configuration, resulting in damage to the wing trailing edge. Following investigation, it was considered that a loss of electrical signal to the `up' solenoid of the flap selector valve had occurred and, combined with the normal internal leakage in the hydraulic system, resulted in hydraulic pressure being supplied to the `down' side of the flap hydraulic jack. The loss of signal could have been intermittent, and the evidence strongly implicated oxide debris contamination of the flap selector switch contacts. This condition, if not corrected, could lead to further cases of damage to the aeroplane due to airspeed limit exceedance, possibly resulting in asymmetric flap deployment, which could lead to loss of control of the aeroplane. We are issuing this AD to require actions to correct the unsafe condition on these products.
97-25-05: This amendment adopts a new airworthiness directive (AD), applicable to Robinson Helicopter Company (Robinson) Model R22 helicopters with a Lycoming 0-360-J2A engine installation. This AD requires replacing the carburetor and carburetor air temperature (CAT) gage with an improved carburetor that does not require manual leaning of the fuel/air mixture during flight, and a remarked CAT gage; and revising the Rotorcraft Flight Manual to remove the reference to leaning the engine. This amendment is prompted by a report from the Civil Aviation Authority of Great Britain that cautioned that the mixture control could inadvertently be placed in the idle cutoff position during in-flight manual leaning of the fuel/air mixture in the carburetor of the Lycoming 0-360-J2A engine. The actions specified by this AD are intended to prevent inadvertent placement of the mixture control to the idle cutoff position during in-flight leaning of the engine, which could result in an engine shutdown and subsequent loss of control of the helicopter.
86-05-08: 86-05-08 FOKKER B.V.: Amendment 39-5243. Applies to Model F28 airplanes, certificated in any category. 11003 to 11189 inclusive 11190 to 11192 inclusive (RH side only) Serial Numbers: 11991 and 11992 To prevent failures of the wing center section, accomplish the following, unless already accomplished, within the next 60 days after the effective date of this AD, or before the airplane reaches six years of age (from date of delivery), whichever occurs later: A. Inspect the center wing rear spar end fittings, and repair if cracks are found, in accordance with Fokker Service Bulletin F28/57-73, dated June 18, 1984. B. Repeat the inspection and repairs required by paragraph A., above, at intervals not to exceed one year. C. Alternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. All persons affected by this directive who have not already received the appropriate service document from the manufacturer may obtain copies upon request to the Manager, Maintenance and Engineering, Fokker B.V., Product Support, P.O. Box 7600, 11172J Schiphol Oost, The Netherlands. This document may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective April 7, 1986.
2021-13-06: The FAA is adopting a new airworthiness directive (AD) for certain Airbus SAS Model A350-941 and -1041 airplanes. This AD was prompted by a determination 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 two European Union Aviation Safety Agency (EASA) ADs, which are incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2012-06-15: We are adopting a new airworthiness directive (AD) for DG Flugzeugbau GmbH Models DG-500 Elan Orion, DG-500 Elan Trainer, DG-500/ 20 Elan, and DG-500/22 Elan sailplanes and Models DG-500M and DG-500MB powered sailplanes. 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 re-installation of the rear cockpit securing rope for the headrest of the rear seat during maintenance, which could cause the rear seat to interfere with the control stick of the sailplane. We are issuing this AD to require actions to address the unsafe condition on these products.
97-25-04: This amendment adopts a new airworthiness directive (AD) that applies to all Cessna Aircraft Company (Cessna) Models 208, 208A, 208B, 425, and 441 airplanes. This AD requires amending the Limitations Section of the airplane flight manual (AFM) to prohibit the positioning of the power levers below the flight idle stop while the airplane is in flight. This AFM amendment will include a statement of consequences if the limitation is not followed. This AD results from numerous incidents and five documented accidents involving airplanes equipped with turboprop engines where the propeller beta was improperly utilized during flight. The actions specified by this AD are intended to prevent loss of airplane control or engine overspeed with consequent loss of engine power caused by the power levers being positioned below the flight idle stop while the airplane is in flight.
2012-07-07: We are adopting a new airworthiness directive (AD) for all The Boeing Company Model 747 airplanes. This AD was prompted by reports of fractured latch pins found in service; investigation revealed that the cracking and subsequent fracture were initiated by fatigue and propagated by a combination of fatigue and stress corrosion. This AD requires repetitive general visual inspections for broken or missing latch pins of the lower sills of the forward and aft lower lobe cargo doors; repetitive detailed inspections for cracking of the latch pins; and corrective actions if necessary. We are issuing this AD to detect and correct fractured or broken latch pins, which could result in a forward or aft lower lobe cargo door opening and detaching during flight, and consequent rapid decompression of the airplane.
78-21-05: 78-21-05 SIKORSKY AIRCRAFT: Amendment 39-3322. Applies to S-58BT, S-58DT, S-58ET, S-58FT, S-58HT, and S-58JT helicopters certificated in all categories. To prevent possible tail rotor instability accomplish the following: Within 200 hours time in service or three months, whichever comes first, after the effective date of this AD, unless already accomplished, complete the following: (a) Install improved abrasion strips, Sikorsky Modification Kit 58070-10008-011, on tail rotor blades, P/Ns S1615-30100-3, -4, -7, -10, -11, -15, -16, -041, -042, and -043. This modification is covered by Sikorsky Service Bulletin No. 58B15-18. (b) Install Sikorsky control rod assembly, P/N 58400-64010-101. This modification is covered by Sikorsky Service Bulletin No. 58B40-5. This amendment becomes effective November 30, 1978.
97-24-14: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires inspection of the two-way check valve on the engine fire extinguishing system for discrepancies, and corrective action, if necessary. This amendment is prompted by issuance of mandatory continued airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent discrepancies of the check valve, which could result in improper functioning of the engine fire extinguishing system.
91-07-05: 91-07-05 GULFSTREAM: Amendment 39-6942. Docket No. 90-NM-187-AD. Applicability: Model G-IV series airplanes, Serial Numbers 1060 through 1089, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent hazardous operation of the Avionics Standard Communication Bus (ASCB) during turbulent weather conditions (Instrument Meteorological Conditions), accomplish the following: A. Within 180 days after the effective date of this AD, perform a detailed integrity test of the ASCB, in accordance with Gulfstream Aerospace Report No. GIV-GER-276, "ASCB Databus Cable, Coupler, and Connector Integrity Test: Phase II Incorporation," dated April 2, 1990. If defective ASCB connectors are found, prior to further flight, repair or replace all defective connectors in accordance with Gulfstream Aerospace Report No. GIV-GER-276, dated April 2, 1990. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Atlanta Aircraft Certification Office (ACO), FAA, Small Airplane Directorate. NOTE: The request should be submitted directly to the Manager, Atlanta ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Manager, Atlanta ACO. C. 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 Gulfstream Aerospace Corporation, P.O. Box 2206, Savannah, Georgia 31402-2206. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington, or at the FAA, Small Airplane Directorate, Atlanta Aircraft Certification Office, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia. This amendment (39-6942, AD 91-07-05) becomes effective on April 22, 1991.
2012-02-16: We are superseding an existing airworthiness directive (AD) for all The Boeing Company Model 747-100, 747-100B, 747-100B SUD, 747- 200B, 747-200C, 747-200F, 747-300, 747-400, 747-400D, 747-400F, 747SR, and 747SP series airplanes. That AD currently requires an inspection of the No. 2 and No. 3 windows on the left and right sides of the airplane to determine their part numbers, related investigative and corrective actions if necessary, and repetitive inspections of single pane windows. This new AD requires installing dual pane No. 2 and No. 3 windows. This new AD also removes certain airplanes from the applicability. This AD was prompted by loss of a No. 3 window in flight, which could result in consequent rapid loss of cabin pressure. Loss of the window could also result in crew communication difficulties or incapacitation of the crew. We are issuing this AD to correct the unsafe condition on these products.
97-24-05: This amendment adopts a new airworthiness directive (AD) that applies to all Aerospace Technologies of Australia Pty Ltd. (ASTA) Models N22B, N22S, and N24A airplanes. This AD requires repetitively inspecting the aft wing break connectors for arcing damage, deposits between contacts, and looseness of contacts; and removing deposits between contacts, tightening any loose contacts, and replacing any aft wing break connectors with arcing damage. This AD results from several reports of uncommanded flap extensions and displays of incorrect stall warning indications on the affected airplanes. The actions specified by this AD are intended to prevent contamination in the aft wing break connectors, which could result in uncommanded flap extensions and incorrect stall warning indications with consequent loss of airplane control.
84-15-02: 84-15-02 DOUGLAS AIRCRAFT COMPANY: Amendment 39-5031. Applies to Douglas Aircraft Model A-26/B26 series airplanes certificated in all categories. Compliance schedule as prescribed in the body of the AD, unless previously accomplished. To prevent loss of wing structural integrity due to failure of a lower spar cap, accomplish the following: A. Within ten (10) hours time in service after the effective date of this AD, perform a visual dye penetrant inspection of the lower forward and aft spar cap, outboard and inboard of each nacelle in the area where the nacelle fairing upper edge runs across the lower spar cap surface. Trim the fairing edge if necessary, to ensure that the fairing edge is smooth and that there is a minimum of one-sixteenth inch (1/16") clearance between the fairing and spar cap surface. B. If no cracking or fretting of the spar cap is detected, repeat the inspection for cracks, surface clearance, and condition in accordance with Paragraph A. of this AD at intervals not to exceed 500 hours time in service. C. If any evidence of cracking or fretting is found in the spar caps, polish out to a machine finish not to exceed approximately 125 microinches on both sides of the damaged area to a maximum depth of 0.030 inches and repeat the inspection of Paragraph A., above. Continue to inspect in accordance with Paragraph A. at intervals not to exceed 30 hours time in service. D. If cracking or fretting in excess of 0.030 inches in a spar cap is detected, repair in a manner approved by the Manager, Western Aircraft Certification Office, FAA, Northwest Mountain Region, Hawthorne, California. E. For those aircraft which have been modified to incorporate a steel or titanium lower front spar cap strap (in the area where the nacelle fairing upper edge is in contact with the lower wing surface) in accordance with AD 64-12-03, the requirements of Paragraph A. of this AD are applicable only to the lower aft spar cap, outboardand inboard of each nacelle. F. Within 72 hours after performing the inspections required by Paragraph A., above, report the results of the inspections to the Manager, Western Aircraft Certification Office, ANM-170W, FAA, Northwest Mountain Region, 15000 Aviation Blvd., Hawthorne, California. Mailing address: P.O. Box 92007, Worldway Postal Center, Los Angeles, California 90009. The reports should cite the airplane registration and serial number, crack location and extent of damage, total airplane operating hours, and time since last inspection. G. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections required by this AD. H. Alternate inspections, modifications, or other actions which provide an acceptable level of safety may be used when approved by the Manager, Western Aircraft Certification Office, FAA, Northwest Mountain Region, Hawthorne, California. This amendment becomes effective April 16, 1985, and it was effective earlier to all recipients of priority mail AD 84-15-02, issued August 1, 1984.
86-07-02 R1: 86-07-02 R1 PILATUS BRITTEN-NORMAN LTD: Amendment 39-10171; Docket No. 86-CE-23-AD. Revises AD 86-07-02, Amendment 39-5382. Applicability: Models BN2A MK. 111, BN2A MK. 111-2, and BN2A MK. 111-3 airplanes (all serial numbers), certificated in any category. NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (f) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. Compliance: Required prior to further flight after the effective date of this AD (see NOTE 2) or within 100 hours time-in-service (TIS) after the last inspection accomplished in accordance with AD 86-07-02, whichever occurs later, and thereafter at intervals not to exceed 100 hours TIS. NOTE 2: The "prior to further flight after the effective date of this AD" compliance time was the original initial compliance time of AD 86-07-02, and is being retained to provide credit and continuity for already-accomplished and future inspections. To prevent failure of the main landing gear caused by cracks in the torque link assembly area, which could lead to loss of control of the airplane during landing operations, accomplish the following: (a) Inspect the junction of the torque link lug and upper case for cracks (using a 10- power magnifying glass or by dye penetrant methods) in accordance with Fairey Hydraulics Limited Service Bulletin (SB) 32-7, Issue 3, dated January 30, 1990; or Fairey Hydraulics SB 32- 10, Issue 2, dated November 10, 1992, as applicable. Pilatus Britten-Norman SB BN-2/SB. 173, Issue 3, dated November 16, 1990, references Fairey Hydraulic Limited SB 32-7; and Pilatus Britten-Norman SB BN-2/SB.209, Issue 1, dated November 30, 1992, references Fairey Hydraulic Limited SB 32-10. (b) If cracked parts are found during any of the inspections required by this AD, prior to further flight, replace the cracked parts with airworthy parts in accordance with the applicable maintenance manual. (c) If the landing gear is replaced, only equal pairs of the same manufacturer are approved as replacement parts. Mixing of different manufacturer landing gears is not authorized. (d) The intervals between the repetitive inspections required by this AD may be adjusted up to 10 percent of the specified interval to allow accomplishing these actions along with other scheduled maintenance on the airplane. (e) Special flight permits maybe issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the inspection requirements of this AD can be accomplished. (f) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Small Airplane Directorate, 1201 Walnut, suite 900, Kansas City, Missouri 64106. (1) The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Small Airplane Directorate. (2) Alternative methods of compliance approved for AD 86-07-02 are considered approved as alternative methods of compliance for this AD. NOTE 3: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Small Airplane Directorate. (g) The inspections required by this AD shall be done in accordance with Fairey Hydraulics Limited Service Bulletin 32-7, Issue 3, dated January 30, 1990, or Fairey Hydraulics Service Bulletin 32-10, Issue 2, dated November 10, 1992, as applicable. 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 Fairey Hydraulics Limited, Claverham, Bristol, England; or Pilatus Britten-Norman Limited, Bembridge, Isle of Wight, United Kingdom PO35 5PR. Copies may be inspected at the FAA, Central Region, Office of the Regional 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. (h) This amendment (39-10171) revises AD 86-07-02, Amendment 39-5382. (i) This amendment becomes effective on November 28, 1997.
86-07-04: 86-07-04 SHORT BROTHERS, LTD.: Amendment 39-5272. Applies to Models SD3- 30 and SD3-60 airplanes as listed in Short Brothers, Ltd. Service Bulletins SD3-27-29, dated April 1985 (for SD3-30 airplanes), and SD360-27-06, dated April 1985 (for SD3-60 airplanes), certificated in any category. To prevent the loss of elevator control, accomplish the following within the next 90 days after the effective date of this AD, unless already accomplished: 1. Modify the elevator torque tube assembly in accordance with Short Brothers, Ltd. Service Bulletin SD3-27-29, dated April 1985 (for SD3-30 airplanes), or SD360-27-06, dated April 1985 (for SD3-60 airplanes). 2. Alternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. 3. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment ofinspections and/or modifications required by this AD. All persons affected by this directive who have not already received the appropriate service document from the manufacturer may obtain copies upon request to Shorts Aircraft, 1725 Jefferson Davis Highway, Suite 510, Arlington, Virginia 22202. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective May 8, 1986.
2012-07-04: We are adopting a new airworthiness directive (AD) for certain Cessna Aircraft Company Model 680 airplanes. This AD was prompted by a false cross-feed command to the right-hand fuel control card, due to the cross-feed inputs on the left- and right-hand fuel control cards being connected together and causing an imbalance of fuel between the left and right wing tanks. This AD requires adding diodes to the fuel cross-feed wiring, and revising the airplane flight manual to include procedures to use when the left or right generator is selected OFF. We are issuing this AD to prevent lateral imbalance of the airplane, resulting from uncontrolled fuel cross-feed, which can be corrected by deflecting the aileron trim; deflecting the aileron trim increases the pilot's workload and could exceed the airplane's limitation in a short period of time, resulting in reduced controllability of the airplane.