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98-24-08: This amendment adopts a new airworthiness directive (AD) that applies to all Burkhart Grob Luft-und Raumfahrt (Grob) Models G115, G115A, G115B, G115C, G115C2, G115D, and G115D2 airplanes. This AD requires inspecting the area of the elevator trim tab hinges for cracks and a secure fit, and repairing any elevator trim tab hinges with cracks or where a proper secure fit is not found. 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 prevent structural damage of the trim tab hinges caused by cracks, which could result in trim tab failure with consequent loss of control of the airplane.
77-01-08: 77-01-08 ROCKWELL INTERNATIONAL: Amendment 39-2805. Applies to Model 112 airplanes (Serial Numbers 3 through 470) certificated in all categories. Compliance required as indicated. To prevent engine failure due to loss of engine oil from failure of the engine oil pressure tube assembly, accomplish the following: a. As a part of each preflight check, after the effective date of this airworthiness directive, conduct a visual check of the P/N 46320-1 copper oil pressure tube assembly for cracks and oil leakage, particularly in the areas where the copper tube and its fittings attach. If any cracks or leakage are detected, remove and replace such assembly as described in paragraph c below before further flight. b. Within the next ten hours' time in service after the effective date of this airworthiness directive and thereafter at intervals not to exceed ten hours' time in service, accomplish the following: (1) Remove the upper engine cowling. (2) Clean the copper oil pressure tube assembly, P/N 46320-1, with an oil soluble solvent. (3) Start and operate the engine until it is warm enough to respond smoothly to throttle changes (monitor oil temperature and cylinder head temperature gauges to maintain temperatures within limits), then stop the engine. (4) Check the copper oil pressure tube assembly, P/N 46320-1, for cracks, chaffing, or any signs of oil leakage. (5) If any of the conditions in step 4 above are detected, remove and replace such assembly as described in paragraph c below before further flight. (6) Part I(1)(b) of Rockwell International Service Bulletin No. SB-112-46 dated November 2, 1976, refers to this same subject. A pilot may perform the requirements of items a and b above pursuant to the provisions of FAR 43.3(h). NOTE: For the requirements regarding the listing of compliance and method of compliance with paragraphs a and b of this AD in the aircraft permanent maintenance record, see FAR 91.173. c. Within 50 hours' time in service after the effective date of this airworthiness directive, remove the P/N 46320-1 copper oil pressure tube assembly, the P/N 46080-1 elbow, the AN919-0 reducer, and the P/N 46136-47 hose assembly and replace them with P/N 46136-63 hose assembly and P/N 46080-3 elbow in accordance with Part II of Rockwell Commander Service Bulletin No. SB-112-46 dated November 2, 1976, or later FAA DOA approved revision. Previous compliance with Rockwell Commander Service Letter No. SL-112-26, June 3, 1976, satisfies this airworthiness directive. d. The checks required by paragraphs a and b of this airworthiness directive are no longer required after paragraph c of this airworthiness directive is accomplished. e. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Customer Service Department, Avenue, Bethany, Oklahoma 73008. 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 Region Office in Fort Worth, Texas. f. Any alternate equivalent method of compliance with this airworthiness directive must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration. This amendment becomes effective on January 14, 1977.
53-20-01: 53-20-01 MARTIN: Applies to All Models 202 and 202A Airplanes With Hamilton Standard Reversing Propeller Installations In Which the Reverse Feature is Not Used, and In Which the Normal Reverse Pitch Stop Ring Has Been Relocated to Act Essentially as a Low Pitch Stop. Compliance required as indicated. With the reverse pitch stop ring relocated to act as a low pitch stop, and with inadvertent energization of the reverse solenoid valve, the stop ring will fail under the extreme loads resulting from high oil pressure surge plus the high blade twisting moments attendant to rapid pitch change toward low pitch. This failure will result in the jamming of the pitch changing mechanism, or an inadvertent propeller reversal. To preclude any hazardous incidents, modification of the internal mechanism of the propeller dome, or modification of the propeller control system is necessary. Accomplish items I, II, and III. I. Comply with AD 52-15-02. A. Item I of AD 52-15-02 isto be accomplished by means of a progressive modification program to be submitted to and approved by the FAA. The program shall begin no later than November 1, 1953, and shall be completed no later than August 1, 1954. B. Item III shall be instituted when reversing is reactivated. II. Modify the internal mechanism of the propeller dome. The modifications include removal of the normal low pitch stops, and incorporation of a dump valve which opens just above the normal low pitch position to maintain oil pressure of 50 to 100 p.s.i. on the decrease pitch side of the piston. A. Replace the present oil transfer housing 70451 or 77828 with engine shaft extension assembly 70300. B. Remove the present low pitch stop lever assembly 71042, 71676 or 76149 from the propeller entirely. C. Replace the present piston sleeve 68425 with a new piston sleeve 72259 and snap ring 67698. Since the piston sleeve is pressed into the piston, the internal diameter of the sleevemust be machined after assembly. D. Purchase Orders for the above parts must be submitted to Hamilton Standard by November 1, 1953. E. Hamilton Standard Service Bulletin No. 264 covers this same subject. F. Compliance required prior to November 1, 1954. III. Modify the propeller control system. A. At the A or unfeathering relays in the propeller relay control box, disconnect the wires that run through the nose junction box and into the reversing control boxes (Hamilton Standard P/N 72400) to connect the A relays with the throttle microswitches. (1) Either physically remove the full run of these wires from the A relay terminal to the nose junction box terminal strip, or (2) Physically remove the portion from the A relays to the pin connector in the propeller control relay box, and disconnect these wires from the mating pin in the external portion of the pin connector. Insulate the exposed ends of these wires, and secure them from movement in such manner as to preclude their becoming grounded or contacting any terminals. B. Modify the reverse solenoid circuit wiring from the reverse solenoids to the terminal on the A relays by providing continuous wires physically isolated from all other circuits. C. Comply with item I.C. of AD 52-15-02. D. Revise the reversing solenoid circuits to comply with Hamilton Standard Service Bulletin No. 236, Ref. 955. E. Provide positive mechanical stops for both throttle levers so they cannot be moved into the reverse range. (1) Item 2, of NWA Mechanical Order No. 479, dated November 21, 1950, is considered acceptable to accomplish this. (2) Comply with items 3 and 4 of NWA Mechanical Order No. 479, dated November 21, 1950. F. Items III.A, B, C, D, and E of AD 53-20-01 are to be accomplished by April 1, 1954. G. No later than November 1, 1953, all operating instructions regarding unfeathering procedures shall specify that the following practicesare to be observed, and shall indicate that the reason is to guard against jamming the pitch changing mechanism or possible inadvertent reversal during the unfeathering operation: (1) If unfeathering is being accomplished at night the wing illumination lights or landing lights are to be used to permit observation of propeller operation. (2) The propeller is to be watched during unfeathering and the button is to be released when rotation starts. (This should normally be in 1 or 2 seconds.) (3) The tachometer is not to be used as a guide for determining when unfeathering is to be terminated.
2013-11-01: We are adopting a new airworthiness directive (AD) for Iniziative Industriali Italiane S.p.A. Models Sky Arrow 650 TC, Sky Arrow 650 TCN, Sky Arrow 650TCS, and Sky Arrow 650TCNS airplanes. 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 improper installation of the spherical bearing on the central hinge lever and a crack on the weld length of the horizontal tail/elevator plane hinge assembly. We are issuing this AD to require actions to address the unsafe condition on these products.
74-07-03: 74-07-03 DOWTY ROTOL: Amendment 39-1803 as amended by Amendment 39-2110. Applies to Dowty Rotol type (c)R 209/4-40-4.5/2, (c)R 245/4-40-4.5/13, and (c)R 259/4-40- 4.5/17 propellers installed on, but not necessarily limited to, Nihon Model YS-11 and YS-11A Series airplanes and Convair Models 600(240D), 640(340D), and 640(440D) Series airplanes equipped with Rolls-Royce Dart Model 542 Series engines. Compliance is required as indicated. To prevent possible propeller failure resulting from cracking of full width case hardened rollers in the bottom (C.F.) race of the propeller blade bearings, accomplish the following: (a) For propellers having blade bearing assemblies that incorporate Modification No. (c) VP2416 (SB61-509) or Modification No. (c) VP2677 (SB61-709) having sets of rollers P/N's 601026724 or 601026940, comply with paragraphs (b) and (c). (1) Before further flight, after each report of significant propeller induced vibration in flight, except that theairplane may be flown in accordance with FAR Section 21.197 to a base where the repair can be performed; and (2) If initial compliance is not required by paragraph (a)(1), within the next 600 hours' time in service after the effective date of this AD or before the accumulation of 2,000 hours' time in service on blade bearing bottom (C.F.) race rollers, whichever occurs later. (b) Replace sets of rollers specified in paragraph (a) in accordance with Dowty Rotol Service Bulletin No. 61-542-8, Revision 2, dated December 20, 1972, for type (c)R 209/4-40- 4.5/2 propellers; Dowty Rotol Service Bulletin No. 61-542-9, dated June 21, 1973, for type (c)R 245/4-40-4.5/13 and (c)R 259/4-40-4.5/17 propellers; (1) With new parts of the same part number and thereafter continue to replace sets of rollers specified in paragraph (a) in accordance with subparagraph (a)(1) and at intervals not to exceed 2,000 hours' time in service on blade bearing bottom (C.F.) race rollers, and comply with paragraph (c) at each replacement; or (2) With through hardened sets of rollers which incorporate Modification No. (c) VP2762 (SB61-771) or Modification No. (c) VP2814 (SB61-795). (c) At each set of roller replacement required by paragraphs (a) and (b), determine the number of broken rollers and the preload in each bearing assembly in accordance with Dowty Rotol Service Bulletin No. 61-542-8, Revision 2, dated December 20, 1972 for type (c)R 209/4- 40-4.5/2 propellers; Dowty Rotol Service Bulletin No. 61-542-9, dated June 21, 1973, for type (c)R 245/4-40-4.5/13 and (c)R 259/4-40-4.5/17 propellers; or an FAA-approved equivalent. If ten or more rollers are found to be broken or if the preload is found to be less than .0035 inches, before further flight, remove the associated propeller blade, blade retaining bolt, and bearing assembly from service, mark them in a manner that will prevent their further use, and replace them with serviceable parts of the same part numberor FAA-approved equivalents. (d) The replacement of sets of rollers required by paragraphs (a) and (b) and the inspections required by paragraph (c) may be discontinued when through hardened sets of rollers which incorporate Modification No. (c) VP2762 (SB-771) or Modification No. (c) VP2814 (SB61-795) are installed in accordance with Dowty Rotol Service Bulletin No. 61-542-8, Revision 2, dated December 20, 1972, for type (c)R 209/4-40-4.5/2 propellers; Dowty Rotol Service Bulletin No. 61-542-9, dated June 21, 1973, for type (c)R 245/4-40-4.5/13 and (c)R 259/4-40-4.5/17 propellers; or an FAA-approved equivalent. Amendment 39-1803 became effective April 19, 1974. This amendment 39-2110 is effective March 27, 1975.
98-24-06: This amendment adopts a new airworthiness directive (AD) that applies to all Dornier-Werke G.m.b.H. (Dornier) Model Do 27 Q-6 airplanes. This AD requires repetitively inspecting the rivets that attach the forward stabilizer attach fitting to the airplane fuselage for looseness, and replacing any loose rivets. 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 prevent the stabilizer from detaching at the forward stabilizer attach flanges because of loose rivets, which could result in reduced or loss of control of the airplane.
98-24-13: This amendment supersedes an existing airworthiness directive (AD), applicable to Eurocopter Model MBB-BK117 A-1, A-3, A-4, B-1, B-2, and C-1 helicopters, that currently requires initial and repetitive inspections of both surfaces of the tail boom vertical fin (vertical fin) spar, the skin, and the left-hand and right-hand frame sheets for cracks or loose rivets. That AD also requires repairing certain cracks, if found, and repairing and reporting those loose rivets and certain other cracks, if found. This amendment requires the same inspections, repairs, and reporting as the existing AD, but changes the reference to the service bulletin and prohibits the use of blind rivets for the vertical fin spar repair. This amendment is prompted by an accident that occurred on April 15, 1997, resulting in one fatality. The actions specified by this AD are intended to prevent failure of the vertical fin and subsequent loss of control of the helicopter.
89-02-08 R1: 89-02-08 R1 CASA: Amendment 39-6111 as revised by Amendment 39-6280. Applicability: All CASA Model C-212 series airplanes, certificated in any category. Compliance: Required as indicated below, unless previously accomplished. To ensure continuing structural integrity of the wing flap control system, accomplish the following: A. Within six months after February 17, 1989 (the effective date of AD 89- 02-08, Amendment 39-6111), incorporate a revision into the FAA-approved maintenance inspection program that will provide for inspection of the wing flap control system in accordance with CASA Document COM. 212-206, Revision 1, dated May 20, 1988. The non-destructive inspection techniques set forth in the CASA C-212 non-destructive procedures (27-50-01 through 27-50-05) provide acceptable methods for accomplishing the inspections required by this AD. All inspection results, positive or negative, must be reported to CASA Product Support, in accordance with instructions in the CASA Flap Control System Inspection Document. This Supplemental Structural Inspection (SSI) is to be repeated at intervals not to exceed 4,000 landings. B. Prior to the accumulation of 4,000 landings, or within six months after the effective date of this amendment, whichever occurs later, inspect the wing flap control system in accordance with CASA Document COM. 212-206, Revision 1, dated May 20, 1988. C. Cracked structure or damaged components detected during the inspections required by paragraphs A. and B., above, must be replaced prior to further flight, in accordance with CASA Document COM. 212-206, Revision 1, dated May 20, 1988. D. 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, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI),who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. E. 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 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 Construcciones Aeronauticas S.A. (CASA), Getafe, Madrid, Spain. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This AD revises AD 89-02-08 (Amendment 39-6111) which became effective on February 17, 1989. This amendment (39-6280, AD 89-02-08 R1) becomes effective on September 5, 1989.
2013-11-03: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc. Model CL-215-1A10 and CL-215-6B11 (CL-215T Variant) airplanes. This AD requires repetitive detailed inspections for cracking of the left-hand (LH) and right-hand (RH) wing lower skin, and repair if necessary. This AD also provides terminating action for the repetitive detailed inspections. This AD was prompted by reports of a fractured wing lower rear spar cap and reinforcing strap. We are issuing this AD to detect and correct cracked wing structure, which could result in failure of the wing.
2013-11-02: We are adopting a new airworthiness directive (AD) for Aircraft Industries a.s. Model L-420 airplanes. 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 in-flight engine flame out occurred at take-off with water injection after reduction of engine power. We are issuing this AD to require actions to address the unsafe condition on these products.
52-13-02: 52-13-02 LOCKHEED: Applies to All Constellation (49 Series) Airplanes With Hamilton Standard Reversing Propellers. Items I, II and III 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 November 1, 1953. I. In 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 specific features to be considered in isolation of the circuits. Airplanes which have other features not specifically referred to in this list shall be treated in an equivalent manner: A. The multiple pin connector assembly at the reverse coordinating relay panel must comply with item 2 of attachment A. B. Protect the following exposed terminals as specified in item 3 of attachment A: (1) Exposed terminal on "A" relay in reverse coordinating relay panel; (2) Exposed terminals at throttle reversing switches (not required if item IIA(1) is installed). C. Modify Hamilton Standard relay box, where used, to shield the reversing solenoid circuit relay contacts, etc., 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. D. Reversing solenoid circuit wiring: Modify in accordance with item 4 of attachment A. II. Other circuit modifications: A. All airplanes shall be modified in one of the following ways: (1) Install an additional switch in the reversing solenoid circuit which will prevent the application of power to the circuit until the switch has been closed by operation of the landing gear actuated throttle reversing lock system, or (2) Comply with Hamilton Standard Service Bulletin No. 221. III. Pedestal design (same compliance date as item I): A. Reversing throttle switch and lock bar assembly: Modify the assembly by adding stop pin and plate for the lock bar and cutting clearance notches on No. 1 and No. 4 switch triggers. IV. Maintenance practices: A. At each nearest scheduled service to 350 hours: (1) Inspect the points specified in item IB and any other critical points in the systems where two or more solenoid wires run together. These inspections may be discontinued if the modifications made to the system are of the type described in item 1(a) or 1(b) of attachment A. (2) 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 operation 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. (3) If item IIA(1) is installed, perform an electrical check of the switch to assure that it opens when the throttle lock bar is in the locked position. B. At any time that an electrical fault occurs in a circuit which is carried in the same bundles or the same conduit 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 be replaced as necessary. V. Operating instructions: Comply with item 5 of Attachment A. VI. (Note: Propeller governor design changes which are under development and whose purpose is to provide a high pressure hydraulic circuit bypass to safeguard against inadvertent reversing and to provide ability to feather even when the reversing solenoid is energized, are still under consideration and may be the subject of a future Directive.) ATTACHMENT A Criteria for isolation of reversing circuits at terminal points and connections: 1. Terminal Strips. The following methods of isolation can be used: (a) Elimination of connections at terminal strips by using continuous wiring, (b) Providing separate, covered terminal strip for reversing lead connections, (c) Isolating the solenoid lead stud, terminals and associated hardware from all other nearby studs or terminals by enclosing these components in an insulating cover which is so designed orsecured to the wiring that the wire will stay in place in case of breakage at the terminal or so that the broken wire and terminal will remain insulated by the cover from contact with other circuits if the wire comes off its terminal. The nature of the cover design or provisions for its attachment must be such that its installation will not be overlooked during maintenance, (d) Removing or grounding studs adjacent to solenoid valve lead stud and securing all adjacent wiring and the reversing solenoid lead to prevent contact of broken leads with reversing solenoid terminal or contact of broken reversing solenoid lead with other terminals. If the adjacent studs are grounded, rather than removed, the studs must be identified distinctively so that they will not inadvertently be used for the attachment of wires serving other circuits. 2. Multiple Pin Connector Assemblies. The following methods of isolation can be used: (a) Elimination of pin connectors by using continuous wiring, (b) Providing separate pin connectors for each reversing solenoid circuit, (c) Deactivating all pins adjacent to the one carrying the reversing solenoid circuit. These pins are to be retained in the connector but identified distinctively so that they will not be used inadvertently. When distinctively identified, these pins may also be used for circuits which cannot supply sufficient energy to actuate the reversing solenoid or circuits which are energized only when reversing is desired. At the points where wires are attached to the connector pins, all exposed metal parts are to be protected with insulating covers so secured that contact between circuits cannot occur in case of failure at the connection or in case foreign material is left in the connector assembly. 3. Exposed Terminals on Relays and Switches. Protect these terminals in either one of the following ways: (a) As specified in item 1(c) for terminal strips, or (b) If the terminal is a type which cannot be protected as specified above, cover all exposed metal components with insulating material and secure all wires so that no wire can touch another terminal if the wire breaks or falls off its own terminal. Install insulating barriers as necessary to prevent inadvertent contact between broken or loose wires and other terminals. 4. Reversing Solenoid Circuit Wiring. Modify in one of the following ways: (a) Physically isolate the wiring from all other circuits. (b) If the wiring is run in bundles with other wires, a shielded wire is to be used. The shielding shall be grounded at both ends and a protective cover shall also be provided over the shielding. The shielding shall be carried as close as possible to the terminal points. 5. Operating Instructions. Not later than August 1, 1952, all operating instructions regarding unfeathering procedures shall specify that the following practices are to be observed and shall indicate that the reason is to guard against inadvertent reversing during the unfeathering operation: (a) If unfeathering is being accomplished at night the wing illumination lights or landing lights are to be used to permit observation of propeller operation. (b) The propeller is to be watched during unfeathering and the button is to be released when rotation starts. (This should normally be in 1 or 2 seconds.) (c) The tachometer is not to be used as a guide for determining when unfeathering is to be terminated.
98-23-14: This amendment adopts a new airworthiness directive (AD) that applies to certain Industrie Aeronautiche e Meccaniche (I.A.M.) Model Piaggio P-180 airplanes. This AD requires inspecting the elevator and aileron control retaining pins for proper installation and damage, and replacing any improperly installed or damaged pins. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for Italy. The actions specified by this AD are intended to prevent the retaining pins from interfering with the flight control elements, which could result in loss of the cable retaining function with consequent loss of control of the airplane.
98-24-02: This amendment adopts a new airworthiness directive (AD) that is applicable to certain McDonnell Douglas Model MD-11 series airplanes. This action requires a one-time inspection to identify the part numbers of two dimmer controls for the overhead instrument panel light and circuit breaker lightplate located in the flight compartment. For airplanes on which a dimmer control having an incorrect part number is installed, this action also requires replacing the dimmer control with a new part; modifying and reinstalling the existing dimmer control; or reinstalling a dimmer control following modification of the part by the part manufacturer. This amendment is prompted by reports of smoke emitting from the overhead panels in the cockpit area. The actions specified in this AD are intended to prevent an electrical failure in the overhead dimmer control due to overheating of a printed circuit board capacitor in the dimmer control, which could result in rupture of the capacitor andsmoke in the flight compartment.
93-11-08: 93-11-08 DE HAVILLAND, INC.: Amendment 39-8598. Docket 92-NM-69-AD. Applicability: Model DHC-7 series airplanes; serial numbers 1 through 30, inclusive; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent structural failure, accomplish the following: (a) Within 6 months after the effective date of this AD, perform a one-time inspection of the lower two securing rivets on the left- and right-hand ground spoiler actuator attachment brackets to ensure the integrity of these rivets and to detect signs of fuel leakage, in accordance with de Havilland Service Bulletin 7-57-12, dated January 15, 1982. (1) If any loose or damaged rivet is found, or if there is evidence of fuel leakage, prior to further flight, modify the ground spoiler actuator attachment brackets (Modification No. 7/1827), in accordance with the service bulletin. (2) If no loose or damaged rivet and no evidence of fuel leakage is found, within 6 months after accomplishing the inspection required by paragraph (a) of this AD, modify the ground spoiler actuator attachment brackets (Modification No. 7/1827), in accordance with the service bulletin. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, New York Aircraft Certification Office (ACO), ANE-170, FAA, Engine and Propeller Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, New York ACO. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the New York ACO. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The inspection and modification shall be done in accordance with de Havilland Service Bulletin 7-57-12, dated January 15, 1982. 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 de Havilland, Inc., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Engine and Propeller Directorate, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on July 22, 1993.
2013-09-10: We are superseding an existing airworthiness directive (AD) for all The Boeing Company Model 737-100, -200, -200C, -300, -400, and -500 series airplanes. That AD currently requires repetitive inspections to detect cracking of the lower corners of the door frame and cross beam of the forward cargo door, and corrective actions if necessary. That AD also requires eventual modification of the outboard radius of the lower corners of the door frame and reinforcement of the cross beam of the forward cargo door, which terminates the existing repetitive inspections. This new AD revises the compliance times for the preventive modification; adds certain inspections for cracks in the number 5 cross beam of the forward cargo door; and adds inspections of the number 4 cross beam if cracks are found in the number 5 cross beam, and corrective actions if necessary. For certain airplanes, this \n\n((Page 31390)) \n\nnew AD also adds a one-time inspection for airplanes previously modified or repaired, and a one-time inspection of the reinforcement angle for excessive shimming or fastener pull-up, and corrective actions if necessary. This AD was prompted by additional reports of fatigue cracking in the radius of the lower frames and in the lower number 5 cross beam of the forward cargo door. We are issuing this AD to prevent fatigue cracking of the lower corners of the door frame and number 5 cross beam of the forward cargo door, which could result in rapid depressurization of the airplane.
98-23-08: This amendment adopts a new airworthiness directive (AD), applicable to certain Pratt & Whitney (PW) PW4000 series turbofan engines not incorporating modifications described in certain PW service bulletins listed in the applicability section, that requires high pressure compressor (HPC) blade tip grinding of the rotor assembly, installation of aluminum oxide coated HPC blade tips in stages 9 through 12, modification of HPC 8th through 14th stage stators, incorporation of 1st stage high pressure turbine (HPT) vanes with increased airflow area which also requires additional HPT hardware modifications, and incorporation of HPC 13th-15th stage zirconium oxide blade tips. This amendment is prompted by reports of HPC surge caused by excessive HPC rear stage rotor-to-case clearance. The actions specified by this AD are intended to prevent HPC surge, which can result in engine power loss at a critical phase of flight such as takeoff or climb. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of January 12, 1999.
91-07-17: 91-07-17 GENERAL ELECTRIC COMPANY: Amendment 39-6954. Docket No. 90-ANE-23. Applicability: General Electric Company (GE) CT58-140-1 turboshaft engines overhauled by H&S Aviation, Division 3 (Formerly Hants and Sussex), Portsmouth, England, between March 14, 1989, and August 21, 1989, installed on, but not limited to, Sikorsky S-61 and Boeing V107 aircraft. The affected parts are identified by part number (P/N) and serial number (S/N) and are installed in engines with S/N, as follows: Engine S/N Part Name P/N Part S/N 295235 Stage 1 forward cooling plate 37C30005P101 BJWTMS 7732 Stage 1 aft cooling plate 3002T25P01 BJWTMS 6649 Stage 2 forward cooling plate 000T88P02 ASVA 3663 Stage 2 aft cooling plate 3002T27P01 ASVA 0125 Power Turbine (PT) wheel & shaft 5002T30P01 GAT 59510 295206 Stage 1 turbine wheel 4002T17P02 GATBK 483 Stage 2 turbine wheel 4002T96P02 GATFE 352 280326 Stage 2 turbine wheel 4002T96P02 GATL 2835 PT wheel and shaft 5002T30P01 GAT 59722 280218 PT wheel and shaft 5002T30P01 GAT 59692 280294 Stage 1 forward cooling plate 37C300055P101 BJWTMS 7663 Stage 1 aft cooling plate 3002T25P01 5778 5778 Stage 2 forward cooling plate 3000T88P02 BJWTMS 5951 Stage 2 aft cooling plate 3002T27P01 BJWTMS 7356 Stage 2 turbine wheel 4002T96P02 GATEL 092 Compliance: Required as indicated, unless already accomplished. To prevent failure of suspect gas generator (GG) turbine rotor parts and power turbine (PT) wheels, which may have a reduced low cycle fatigue life due to inadvertent shotpeening which could result in an uncontained engine failure, remove and inspect all surfaces of GG turbine rotor parts and PT wheels identified above by 20x microscope as follows: (a) Part S/N BJWTMS 7732, BJWTMS 6649, ASVA 3663, ASVA 0125, and GAT 59510, prior to accumulating 5,853 cycles since new (CSN). (b) Part S/N GATBK 483, and GATFE 352, prior to accumulating 5,249 CSN. (c) Part S/N GATL 2835, prior to accumulating 10,792 CSN. (d) Part S/N GAT 59722, prior to accumulating 10,856 CSN. (e) Part S/N GAT 59692, prior to accumulating 9,660 CSN. (f) Part S/N BJWTMS 7663, BJWTMS 5778, BJWTMS 5951, BJWTMS 7356, and GATEL 092, prior to accumulating 5,696 CSN. (g) Remove from service parts found with evidence of shotpeening and replace with a serviceable part. Reidentify parts found with no evidence of shotpeening prior to return to service. NOTE: GE Alert Service Bulletin CT58 A72-173 (CEB-275), contains information in reference to paragraphs (a) through (g) above. (h) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished. (i) Upon submission of substantiating data by an owner or operator through an FAA Airworthiness Inspector, an alternate method of compliance with the requirements of this AD or adjustments to the compliance schedule specified in this AD may be approved by the Manager, Engine Certification Office, ANE-140, Engine and Propeller Directorate, Aircraft Certification Service, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803-5299. All persons affected by this directive who have not already received the appropriate alert service bulletin from the manufacturer may obtain copies upon request to GE Aircraft Engines, 1000 Western Avenue, Lynn, Massachusetts 01910. These documents may be examined at the FAA, New England Region, Office of the Assistant Chief Counsel, Room 311, 12 New England Executive Park, Burlington, Massachusetts. This amendment (39-6954, AD 91-07-17) becomes effective on April 26, 1991.
98-23-15: This amendment adopts a new airworthiness directive (AD), applicable to certain Raytheon Model Hawker 800XP series airplanes, that requires replacement of the fuel feed hose assemblies of the auxiliary power unit (APU) with new hose assemblies. This amendment is prompted by a report of the collapse of the inner casing of the fuel feed hose that supplies fuel to the APU. The actions specified by this AD are intended to prevent failure of the fuel feed hose assemblies, which could result in fuel leakage and consequent risk of fire in the aft equipment bay.
98-23-12: This amendment supersedes an existing airworthiness directive (AD), applicable to all de Havilland Model DHC-7 series airplanes, that currently requires certain structural inspections, and repair, if necessary. This amendment requires an additional structural inspection. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to detect and correct fatigue cracking in certain significant structural areas, which could reduce the structural integrity of these airplanes.
77-13-14: 77-13-14 HAWKER SIDDELEY AVIATION, LIMITED: Amendment 39-2939. Applies to Model DH-114 airplanes, certificated in all categories, except those modified in accordance with either Hawker Siddeley Modification 1102 or 1103. Compliance is required as indicated, unless already accomplished. To detect corrosion and cracking of the drag brace fittings at the center section spar lower cap, and prevent possible structural failure of the wing, accomplish the following: (a) Within the next 50 hours time-in-service after the effective date of this AD, inspect the drag brace fittings, P/N 4 FS.1797(L.H.) and 4 FS.1798(R.H.), for corrosion and cracks in accordance with paragraph 4, entitled "Inspection of Fittings," of Hawker Siddeley Aviation, Ltd., Technical News Sheet, Series 114, No. F.18, dated April 9, 1973, or an FAA- approved equivalent. (b) If a crack or evidence of severe corrosion is found in either drag brace fitting during the inspection specified in paragraph (a)of this AD, before further flight, replace the drag brace fitting with an improved fitting, P/N 14FS.6069 (L.H.) or P/N 14FS. 6070 (R.H.), in accordance with the instructions set forth in paragraph 5.2 of Hawker Siddeley Aviation, Ltd., Technical News Sheet, Series 114, No. F.18, dated April 9, 1973, or an FAA-approved equivalent. (c) "Severe corrosion" as used in this AD is internal corrosion that has progressed to a point where there is evidence of corrosion breaking through the exterior surface of the fitting. "Surface corrosion" as used in this AD means corrosion limited to surface pitting. NOTE: The Hawker Siddeley Aviation, Ltd., Model DH-114 Maintenance and Repair Manual further defines the terms "severe corrosion" and "surface corrosion." (d) If, during an inspection specified in paragraph (a) of this AD, no crack is found and evidence of corrosion is found but it is limited to surface corrosion and does not constitute the severe corrosion specified in paragraph (b) of this AD, before further flight, remove the corrosion and protect the surface from further corrosion in accordance with FAR 43.13, and thereafter reinspect the drag brace fittings in accordance with the inspection requirements of paragraph (a) of this AD at intervals not to exceed 100 hours time-in-service from the last inspection. NOTE: Advisory Circular AC 43.13-1A, "Acceptable Methods, Techniques and Practices - Aircraft Inspection and Repair," contains information relating to corrosion protection and removal. (e) If, during an inspection specified in paragraph (a) of this AD, no crack is found in the fittings, and the fittings are free of evidence of any corrosion, reinspect the fittings in accordance with the inspection requirements of paragraph (a) of this AD at intervals not to exceed 100 hours time-in-service from the last inspection. (f) The repetitive inspections required by paragraphs (d) and (e) of this AD may be terminated upon the installation of the improved fittings, P/Ns 14FS. 6069 (L.H.) and 14FS. 6070 (R.H.), in accordance with the instructions set forth in paragraph 5.2 of Hawker Siddeley Aviation, Ltd., Technical News Sheet, Series 114, No. F.18, dated April 9, 1973, or an FAA- approved equivalent. This amendment becomes effective July 7, 1977.
2013-09-09: We are superseding an existing airworthiness directive (AD) for all Slingsby Sailplanes Ltd. Models Dart T.51, Dart T.51/17, and Dart T.51/17R sailplanes equipped with aluminum alloy spar booms. 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 an incident of glue joint failure on a starboard wing caused by water entering the area of the airbrake box that resulted in delamination and corrosion in the area of the aluminum alloy spar booms and the wing attach fittings. We are issuing this AD to require actions to address the unsafe condition on these products.
98-23-13: This amendment supersedes an existing airworthiness directive (AD), applicable to all British Aerospace Model Viscount 700, 800, and 810 series airplanes, that currently requires repetitive inspections to detect cracks and corrosion in the inboard and outboard engine nacelle structures on the wings; replacement of any cracked fittings and mating struts; and treatment or replacement of any corroded fittings or struts. This amendment requires repetitive inspections to detect cracking or corrosion of the eye end fittings of the outboard engine lower support or of the bore of the taper pin holes, and repair, if necessary. This amendment also limits the applicability of the existing AD. This amendment is prompted by reports of cracked and separated lower eye end fittings. The actions specified by this AD are intended to detect and correct cracking of the eye end fittings of the outboard engine lower support, which could result in reduced structural integrity of the engine nacelle support structures.
98-22-16: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 98-22-16 which was sent previously to all known U.S. owners and operators of RHC Model R44 helicopters by individual letters. This amendment supersedes AD 98-12-19, issued August 5, 1998, applicable to RHC Model R44 helicopters, that currently requires main rotor blade inspections and replacement if a crack is found. This amendment requires the same inspections as AD 98-12-19, but mandates replacement of all the affected main rotor blades prior to further flight after November 15, 1998. This amendment is prompted by an incident in which a crack was discovered in a main rotor blade. The actions specified by this AD are intended to prevent failure of a main rotor blade and subsequent loss of control of the helicopter.
98-23-05: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Boeing Model 767 series airplanes, that currently requires an inspection to detect damage of the wire bundles in the left side of the flight compartment in the vicinity of the stowage box for the captain's oxygen mask, and repair, if necessary; a continuity check on repaired wires; installation of sleeving over the wire bundles; and rerouting of the wire bundles. This amendment requires modifications of the captain's and first officer's consoles in the flight compartment to ensure adequate clearance between oxygen equipment and adjacent wire bundles. This amendment is prompted by reports indicating that chafed wiring and wire insulation wear occurred in the vicinity of the stowage box for the captain's oxygen mask due to interference between oxygen line fittings and adjacent wire bundles. The actions specified by this AD are intended to prevent such chafing and inadequate clearance,which could result in electrical arcing and consequent oxygen leakage in the vicinity of the stowage box; these conditions, if not corrected, could result in a fire in the flight compartment.
91-15-01: 91-15-01 AEROSPATIALE: Amendment 39-7064. Docket No. 91-NM-123-AD. Applicability: Model ATR42-200 and -300 series airplanes, serial numbers 003 to 208, 213, 214, 218, 221, 225, 226, and 228; and Model ATR72-100 and -200 series airplanes, serial numbers 126 to 189, 195, 198, and 210; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent failure of the rudder pedal connection rod and subsequent reduced controllability of the airplane, accomplish the following: (a) Within 10 days after the effective date of this AD, perform a visual inspection of the rudder pedal connection rod (captain side) to determine the rod vendor, in accordance with Aerospatiale Service Bulletin ATR42-27-0052 (for Model ATR42 series airplanes), Revision 1, dated April 4, 1991; or Aerospatiale Service Bulletin ATR72-27-1015 (for Model ATR72 series airplanes), Revision 1, dated April 4, 1991; as applicable. (1) If the rod is manufactured by TAC, no further action is required. (2) If the rod is manufactured by SARMA, vendor P/N 14132B, prior to further flight, replace the rod with a TAC rod; or a SARMA rod, vendor P/N 14132-C; in accordance with the applicable service bulletin. (b) As of the effective date of this AD, no SARMA rudder pedal connection rod, P/N 14132B, shall be installed on any airplane. (c) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113. (d) 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. (e) The inspection and replacement requirements shall be done in accordance with Aerospatiale Service Bulletin ATR42-27-0052 (for Model ATR42 series airplanes), Revision 1, dated April 4, 1991; or Aerospatiale Service Bulletin ATR72-27-1015 (for Model ATR72 series airplanes), Revision 1, dated April 4, 1991; which include the following list of affected pages: Service Bulletin Page Number Revision Level Date 1, 3, 4, 7 1 April 4, 1991 ATR42-27-0052 2, 5-6, 8 through 17 (Original) March 7, 1991 1, 2, 5 1 April 4, 1991 ATR72-27-1015 3, 4, 6 through 15 (Original) March 7, 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 Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-7064, AD 91-15-01) becomes effective on July 17, 1991.