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81-23-01 R1: 81-23-01 R1 BEECH: Amendment 39-4289. Applies to the following model airplanes regardless of the category or categories of airworthiness certification: MODELS SERIAL NUMBER (S/N)* 65, A65 & A65-8200 LC-181 through LC-335 70 LB-1 through LB-35 65-A80, 65-A80-8800 & 65-B80 LD-151 through LD-511 and LD-34, LD-46, LD-119 65-A88, 65-88 LP-1 through LP-54 65-90, 65-A90, B90 & C-90 LJ-1 through LJ-929 E90 LW-1 through LW-342 99, 99A, B99 U-1 through U-164 100 & A100 B-1 through B-247 B100 BE-1 through BE-102, and BE-104 Military: L23F** ** LF-7 through LF-76** 65-A90-1 LM-1 through LM-144 65-A90-2 LS-1, -2, -3 65-A90-3 LT-1, -2 65-A90-4 LU-1 through LU-16 NU-8F LG-1 *Except that airplanes which have installed BEECHCRAFT Kit No. 90-4077-1 S, BEECHCRAFT Kit No. 99-4023-1 S, or Aviadesign Supplemental Type Certificate SA1178CE or SA1583CE are not affected by this AD. **Except that Model L23F airplanes which do not have a preload indicating washer assembly (i.e., one with radial holes in a center ring) are not affected by this AD. COMPLIANCE: Required as indicated, unless already accomplished. In order to assure integrity of bolts and nuts at the lower forward attachments of outer wing panels to the wing center section, accomplish the following: A) Prior to next flight, accomplish all of the following: 1. Remove all bolts, washers, and nuts from each lower forward wing attachment and thoroughly clean each removed part. Throughout all action required by this AD: a. Use procedures in the applicable Beech Maintenance Manual except where other procedures are specified by this AD, b. Unless different instruction from Beech Aircraft Corporation is obtained and followed, reposition wing, as necessary, to remove or reinstall bolt by hand without using any tool, c. Keep parts of each preload indicating washer assembly together so that parts of one assembly cannot be intermingled with parts of another assembly, d. Clean each removed part with naphtha or methyl ethyl ketone (MEK) using a bristle brush, and repeat this cleaning as necessary prior to each subsequently specified action until lubricant is applied, and e. Accomplish all of the specified actions on both (i.e., left and right) sides of the airplane. 2. Visually inspect each bolt and nut for reddish rust. Do not classify copper residue over cadmium plating as rust. For a bolt, rust is acceptable only on the end (including not more than one thread) farthest from the head and within counterboard recess between wrench serrations of the bolt head. For compliance with Paragraph A)6 and C), below, classify a bolt as rusted if rust is found elsewhere. Classify a nut as rusted if rust is found anywhere. 3. Visually inspect each bolt and nut for a pit or crack in steel (not cadmium or copper plating) material. Use 10X or stronger magnifying glass. For each bolt, pay particular attention to the fillet and shank, including threads. For each nut, pay particular attention to the chamfer (that faces the bolt head when installed) and perceptible threads adjacent to this chamfer. (Refer to Paragraphs A)6 and C) below.) 4. Bake each bolt and nut continuously for 23 hours at 350 degrees to 400 degrees Fahrenheit and cool in still air. 5. After accomplishment of Paragraph A)4, above, use a magnetic particle method of Advisory Circular AC43.13-1A to inspect each bolt and nut for a crack, paying particular attention to locations specified in Paragraph A)3, above. For each bolt, use a fluorescent particle method with 5250 to 6750 ampereturns in a coil to produce longitudinal magnetization in each bolt. (6,000 ampereturns means 2,000 amperes in a 3-turn coil or 1,000 amperes in a 6-turn coil, etc.) For each nut, use any magnetic particule method with 500 to 700 amperes through a central conductor of at least 0.6-inch diameter through two nuts to produce circular magnetization. Demagnetize each bolt and nut after the above inspection. 6. Replace each rusted, pitted, and/or cracked nut and bolt with a new Part Number (P/N) as follows: a. If new preload indicating (PLI) washer assembly is to be used in accordance with Paragraph A)9, below, nut P/N is 72789-1414, 72789M-1414, FN22-1414, or FN22M-1414. ("M" in P/N denotes black coating. All eligible nuts have a locking feature which necessitates use of a wrench for full engagement with bolt.) b. If a used PLI washer assembly is reinstalled in accordance with Paragraph A)9, below, nut P/N is 72789-1414 or FN22-1414. c. Bolt is P/N LWB 22-14-XX or VEP 220121-14-XX where XX is 31 for airplanes with S/N LD-34, LD-46, LD-119, and LJ-1 through LJ-67, and XX is 32 for all other airplanes affected by this AD. Replace preload indicating washer with new P/N 61475-14-43.5 assembly (not any other P/N) if this assembly is available. Obtain new parts only from BeechService Centers or Beech Aircraft Corporation. (Neither baking nor field inspection of new parts is necessary.) Do not replate any part. 7. Clean the bore and recessed washer seat area of the outboard and inboard wing fittings with naphtha or methyl ethyl ketone (MEK). Visually inspect these areas for corrosion, burrs, gouges and coining. If any defect is found, contact Beech Aircraft Service Department, 9709 East Central, Wichita, Kansas 67201; telephone (316) 681-7261, 7278, or 7352, for rework disposition. Also, if any defect is found, treat the bore and recessed washer seat areas of the inboard and outboard wing fittings with Alodine 1200, 1200S, or 1201. Allow the alodine coating to dry for 5 minutes. Wash the coating with water and blow dry with air without wiping. Paint treated washer seat areas with zinc chromate primer (obtain locally) and allow primer to dry. 8. Coat the inspected areas of the wing fittings, all of each bolt, all of each nut, and all of each preload indicating washer assembly with either clean MIL-C-16173, Grade 2 corrosion preventative compound or clean General Electric G322L Versilube Silicone Lubricant. 9. Install removed or new parts using standard procedures except as follows: a. Preload indicating (PLI) washer assembly may be reused with P/N 72789-1414 and/or P/N FN22-1414 nuts, only. b. Ascertain that a radius of the adjacent washer is next to the fillet under the bolt head and next to the outer edge of the recess in each wing fitting. Position wing as necessary to allow bolt to slide into fitting without use of any tool. c. Tighten the joint by rotating the nut (do not turn the bolt). Use standard procedure if new PLI washer assembly is installed. If used PLI washer assembly is reinstalled, make necessary correction for any torque wrench adapter and apply 3250 to 3400 inch-pounds torque, but install new PLI washer assembly if center ring of the used assembly turns after 3400 inch-pounds torque is applied. Do not allow wrench to bear against fitting. d. Coat entire portion of bolt that projects beyond nut, using a material that is specified in Paragraph A)8, above. e. Make aircraft maintenance record entry showing work accomplished, especially procedure used for tightening nut, and whether new or used PLI washer was installed. Indicating washer assembly with either clean MIL-C-16173, Grade 2 corrosion preventative compound or clean General Electric G322L Versilube Silicone Lubricant. B) Between 90 and 110 hours time-in-service after accomplishment of action specified by Paragraph A) of this AD, check nut tightness, using the same procedure that was used for accomplishment of Paragraph A)9c, above. C) Within 3 days after replacing a part in accordance with Paragraph A)6, above, or noting a defect when complying with this AD, submit a written report to the Federal Aviation Administration via an FAA M or D Report (FAA Form 8330-2) or a letter to the office specified in Paragraph E), below and send the replaced part(s) to Beech Aircraft Corporation. In the submitted report, please advise date of last previous bolt removal. D) A special flight permit in accordance with Federal Aviation Regulation 21.197 for flight to the nearest base is permitted in order to accomplish Paragraph A) of this AD. The nearest FAA Flight Standards District Office may be contacted to obtain a telegraphic special flight permit. E) Any equivalent method of compliance with this AD must be approved by the Chief, Aircraft Certification Program, Federal Aviation Administration, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209; Telephone (316) 269-7000, 7001, or 7002. This amendment becomes effective on January 4, 1982, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated October 31, 1981.
98-13-05: This amendment adopts a new airworthiness directive (AD) that applies to certain Alexander Schleicher Segelflugzeugbau (Alexander Schleicher) Model AS-K13 sailplanes. This AD requires inspecting the main spar fitting for excessive tolerance, traces, movement, etc., and repairing the main spar fitting if any of the above conditions exist. 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 failure of the main spar caused by excessive movement of the main spar fitting, which could result in loss of control of the sailplane.
2012-22-15: We are superseding an existing airworthiness directive (AD) for all Fokker Services B.V. Model F.28 Mark 0070 and 0100 airplanes. That AD currently requires revising the airworthiness limitations section (ALS) of the instructions for continued airworthiness for certain airplanes, and the FAA-approved maintenance program for certain other airplanes, to incorporate new limitations. This new AD requires revising the maintenance program to incorporate the limitations, tasks, thresholds, and intervals specified in certain revised Fokker maintenance review board (MRB) documents. This AD was prompted by a revised Fokker 70/100 MRB document with revised limitations, tasks, thresholds, and intervals. We are issuing this AD to reduce the potential of structural failures or of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors, could result in fuel tank explosions and consequent loss of the airplane.
84-17-02: 84-17-02 GATES LEARJET: Amendment 39-4902. Applies to Gates Learjet Model 24 series airplanes certificated in all categories. Compliance required within eighteen (18) months after effective date unless previously accomplished or prior compliance with AD 81-16-08, Amendment 39-4546. Accomplish the requirements of this AD at an FAA approved repair station. The modification and inspection of the horizontal stabilizer trim actuator must be performed by an FAA certificated repair agency authorized to overhaul and test the Gates Learjet Horizontal Stabilizer Trim Actuator. The necessary shop equipment or the equivalent, as referenced in Learjet Repair Manual Number 1711-9, is listed in Attachment I hereto. A. Modify Learjet Model 24 airplane flight control systems, stall warning, and control wheel by incorporating the airplane modification kit listed in Table I. Table I MODIFICATION KIT MODEL AMK 82-5 24, 24A AMK 82-1 24B, 24B-A AMK 81-18 24D, 24D-A AMK81-13 24E, 24F, 24F-A NOTE: Modification of JET Autopilot Controllers and Computers as required by the instructions in the above kits must be performed in an FAA approved facility for maintenance of the JET FC-110 autopilot. B. Required Airplane Maintenance Record entry must be accomplished by the facility performing its portion of the AD as prescribed in paragraph A. of this AD. C. After airplane modification, verify that the Airplane Flight Manual (AFM) contains the revision/change listed in Table II, below, or later FAA approved revision/change. Upon completion of the modifications required by paragraph A. of this AD, the more restrictive paragraphs A.2., A.5., and A.6. of AD 80-19-11 are no longer applicable. Table II Gates Learjet Airplane Flight Manual/Supplement Revision/Change Modification kit 1. 24 AFM, FM- 004 Revision 32 AMK82-5 2. 24 RAS AFM, W0159 Revision 4 AMK82-5 3. 24A AFM, FM-005 Revision 18 AMK82-5 4. 24E CR 736 AFM, FM-008 Revision 7 AMK82-5 5. FC-110 Autopilot, W1037 New AMK82-5 6. 24B AFM, FM-006 Revision 21 AMK82-1 7. 24B-A AFM, FM-007 Revision 21 to 24B AMK82-1 8. 24B RAS AFM, W0157 Revision 3 AMK82-1 9. FC-110 Autopilot, W1041 New AMK82-1 10. 24D AFM, FM-009 Change 18 AMK81-18 11. 24D-A AFM, FM-010 Change 18 to 24D AMK81-18 12. 24D RAS AFM, W-0119 Change 3 AMK81-18 13. FC-110 Autopilot, W1030 New AMK81-18 14. 24E AFM, FM-011 Change 9 AMK81-13 15. 24F AFM, FM-012 Change 7 AMK81-13 16. FC-110 Autopilot, W1018 New AMK81-13 D. Prior to accomplishing the modification required by paragraph A. of this AD, contact the Wichita Aircraft Certification Office, FAA, Central Region, telephone (316) 269- 7008, if any modification or alteration has been performed on the affected airplane systems, for further instruction relative to the compatibility of the modification and this AD. E. Alternate methods of compliance with this AD may be used if they are approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. This amendment becomes effective October 5, 1984. ATTACHMENT I The stabilizer actuator test stand (P/N ST-00463) is used to functionally test the stabilizer actuator after overhaul. The physical structure of the test stand must be capable of withstanding a minimum load of 2500 lbs. without any bending or deformation. The stabilizer actuator is vertically mounted on the test stand with one end stationary and the other end movable through a hydraulic actuator. The test stand consists of the following components: a. Hydraulic Actuator - The hydraulic actuator is capable of applying a regulated load of 0 to 2500 lbs. on the stabilizer actuator during the entire extend or retract cycles. b. Hydraulic Pressure Regulator - The pressure regulator is used to select hydraulic pressures applied to the stabilizer actuator during the functional test. c. Hydraulic Pressure Gauge - The hydraulic pressure gauge is used to monitor hydraulic pressure applied to the stabilizer actuator. The gauge must be certified at least monthly. d. Digital Position Readout - The digital position readout indicator is used to monitor the travel of the stabilizer actuator. Signals to the indicator are picked up from a rigid mounted linear potentiometer and movable wiper attached to the hydraulic actuator. The digital readout is accurate to 1/1000th of an inch. e. Linear Scale - A linear scale, graduated in 100th of an inch, is permanently mounted on the test stand to verify the digital readout. A tool of known length is used to verify the linear scale and digital readout before the stabilizer actuator functional test is performed. The tool length must be certified at least yearly. f. Lapse Timer - A lapse timer is coupled to the control switches and the stabilizer actuator to monitor travel time during the extend and retract cycles. The lapse timer must measure seconds to beaccurate to 1/100th of a second. g. Trim Controller - The trim controller is used to simulate two-speed input to the stabilizer actuator primary motor. The trim controller part number is EM 2079-6. h. Pre-Select Timer - The pre-select timer is used to check stabilizer actuator travel vs. time, voltage, and amperage inputs in accordance with the functional test. i. Power Supply - The power supply is variable through 0-30 volts DC and 0-30 amperes DC. j. DC Voltmeter - The DC voltmeter must be capable of measuring 0-30 volts DC and must be certified at least yearly. The voltmeter is used to monitor the voltage inputs to the stabilizer actuator in accordance with the functional test. k. DC Ammeter - The DC ammeter must be capable of measuring 0-30 amperes DC and must be certified at least yearly. The ammeter is used to monitor the amperes inputs to the stabilizer actuator in accordance with the functional test. l. Millivolt Meter - The millivolt meter is used to monitor the stabilizer actuator linear potentiometer for a smooth and steady signal output. The meter is 0-50 volts graduated in 100 mv increments. m. Switches - Necessary switches installed to operate the stabilizer actuator primary and secondary motors to extend or retract. n. A digital or Simpson 260 meter, not a part of the test stand, is used to verify the resistance of the stabilizer actuator linear potentiometer. The digital or Simpson 260 meter must be certified at least every 90 working days.
2022-08-07: The FAA is adopting a new airworthiness directive (AD) for all Embraer S.A. Model ERJ 170-100 STD, -100 LR, -100 SU, and -100 SE; ERJ 170-200 STD, -200 LR, -200 SU; ERJ 190-100 STD, -100 LR, -100 IGW, and -100 ECJ; and ERJ 190-200 STD, -200 LR, and -200 IGW airplanes. This AD was prompted by a report of the failure of the inner pane of certain passenger windows to meet maximum operating pressure and lack of fail- safe design. This AD requires determining if certain NORDAM passenger windows are installed, and performing corrective actions if any affected part is installed. This AD also prohibits the installation of affected parts. The FAA is issuing this AD to address the unsafe condition on these products.
2010-11-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated 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: During ERJ 170 airplane full scale fatigue test, cracks were found in some structural components of the airplane. Analysis of these cracks resulted in modifications on the airplane Airworthiness Limitation Items (ALI), to include new inspections tasks or modification of existing ones and its respective thresholds and intervals. Failure to inspect these components according to the new tasks, thresholds and intervals, could prevent a timely detection of fatigue cracks. Undetected fatigue cracks in these areas could adversely affect the structural integrity of these airplanes. * * * * * We are issuing this AD to require actions to correct the unsafecondition on these products.
68-17-08: 68-17-08\tBOEING: Amendment 39-640 as amended by Amendments 39-670, 39-852, 39- 891, and 39-923 is further amended by Amendment 39-1420. Applies to all Boeing 707/720 series aircraft listed in Boeing Service Bulletin 1995 (Revision 5) dated September 28, 1967 or later FAA approved revision. \n\tCompliance required as indicated. \n\tTo detect cracking and prevent failure of the lower wing skin at front spar station 392, accomplish the following: \n\t(a)\tInspect the lower wing skin of aircraft which have not been repaired by installation of the small repair doubler (identified on Page 25, Boeing Service Bulletin 1995, Revision 5), for cracks emanating from the two outboard fasteners of the splice plate tab as noted in Figure 1 of Boeing Service Bulletin 1995, (Revision 5 or later FAA approved revisions) by the use of either a dye penetrant or an eddy current inspection technique, or an equivalent inspection method approved by the Chief, Aircraft Engineering Division, FAA Western Region, atthe times specified in (h), (i), (j) or (k) as appropriate and, if cracks are found, repair prior to further flight per (f) or (g). \n\t(b)\tInspect the lower wing skin of aircraft which have been repaired by installation of the small repair doubler in accordance with Boeing Service Bulletin 1995, within 1600 hours (for 720 Series) or 2000 hours (for 707 Series) after installation or within the next 400 hours (for 720 Series) or 600 hours (for 707 Series) time in service after the effective date of this AD, unless inspected within the previous 1200 (for 720 Series) or 1400 (for 707 Series) hours time in service and at intervals thereafter not to exceed 1600 (for 720 Series) or 2000 (for 707 Series) hours time in service, per (e). \n\t(c)\tInspect the lower wing skin for cracks emanating from the attachments of the front spar support fitting as noted in Figure 1 of Boeing Service Bulletin 1995 (Revision 7) dated 20 August 1969 or later FAA-approved revisions, at the times specified in (h),(i), (j), or (k) as appropriate, and, if cracks are found, repair prior to further flights per (g). The initial inspection must be accomplished either by means of a dye penetrant technique or in accordance with the eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region. The eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revision, or an equivalent FAA-approved inspection technique, must be used for all inspections thereafter. \n\t(d)\tOn those aircraft which have not had the drag fitting trimmed and the fairing attach angle modified in accordance with Boeing Service Bulletin 1995 (Revision 7) or later FAA- approved revisions, within the next 400 hours (for 720 Series) or 600 hours (for 707 Series) time in service after the effective date of this AD and thereafter at intervals not to exceed 800 hours (for 720 Series) or 1,200 hours (for 707 Series) time in service, inspect for cracks in the lower wing skin, emanating from the forward fastener for the drag fitting and from the fasteners for the fairing attach angle as noted in Figure 1 of Boeing Service Bulletin 1995 (Revision 7) or later FAA-approved revisions, at the threshold times as specified in (h), (i), (j), or (k) as appropriate. The initial inspection must be accomplished either by means of a dye penetrant technique or by use of eddy current inspection techniques described in S.B. 1995 (Revision 7), or later FAA- approved revisions, or an equivalent inspection approved by the Chief, Aircraft Engineering Division, FAA Western Region. The eddy current inspection technique described by S.B. 1995 (Revision 7) or later FAA-approved revision, must be used for all inspections thereafter. If cracks are found around the fairing attach angle or emanating aft from the drag fitting fastener, rework the drag fitting, double, andskin prior to further flight in accordance with (g). If cracks are found emanating forward from the drag fitting fastener, rework the drag fittings, doubler and skin, prior to further flight in accordance with Boeing Service Bulletin 1995 (Revision 7) or later FAA-approved revision, or in accordance with an equivalent rework or modification approved by the Chief, Aircraft Engineering Division, FAA Western Region. \n\t(e)\tInspect the lower wing skin covered by the small repair doubler for cracks by use of the x-ray inspection techniques noted in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) or an equivalent inspection technique approved by the Chief, Aircraft Engineering Division, FAA Western Region. Repeat inspections at intervals not to exceed 1600 hours (for 720 Series) or 2000 hours (for 707 Series) time in service. If crack growth is found, repair prior to further flight in accordance with (g). \n\t(f)\tIf the cracks fall within the crack length limitsoutlined in the paragraph titled "Installation of the small repair doubler," (Part II, Boeing Service Bulletin 1995, Revision 5 or later FAA-approved revisions) repair in accordance with that section of the bulletin or later FAA- approved revisions. Within 1600 hours (for 720 Series) or 2000 hours (for 707 Series) after installation of the doubler, inspect in accordance with (e). \n\t(g)\tUpon completion of any of the following modifications, the inspections required by this AD may be discontinued: \n\t\t(1)\tInstallation of 65-56257-1-2 or 65-57788-1-2 doublers as appropriate, per Boeing Service Bulletin 1995 (Revision 5 or later FAA-approved revisions). \n\t\t(2)\tBoeing Service Bulletin 2484. \n\t\t(3)\tBoeing Service Bulletin 2487. \n\t\t(4)\tInstallation of the 720 Wing Structural Improvement Program (Per Boeing Document 65-12700) accomplish at Boeing's Wichita facility. \n\t\t(5)\tAn equivalent installation approved by the Chief, Aircraft Engineering Division, Western Region, FAA. \n\t(h)\tFor those airplanes listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part I, and having less than 6000 (for 720 Series) or less than 10,000 (for 707 Series) hours time in service on the effective date of this AD, prior to the accumulation of 6800 (for 720 Series) or 11,200 (for 707 Series) hours time in service, respectively, and thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service from the last inspection. \n\t(i)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part I, and having 6000 or more (in the case of 720 Series aircraft) or 10,000 or more (in the case of 707 Series aircraft) hours time in service on the effective date of this AD, within the next 400 (for 720 Series) or 600 (for 707 Series) hours time in service, unless accomplished within the last 400 (for 720 Series) or 600 (for 707 Series) hours time in service, and at intervals thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service. \n\t(j)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part II, and having less than 10,000 (in the case of 720 Series) or less than 15,000 (in the case of 707 Series) hours time in service on the effective date of this AD, prior to the accumulation of 10,800 or 16,200 hours time in service, respectively, and thereafter at intervals not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service from the last inspection. \n\t(k)\tFor aircraft listed in Boeing Service Bulletin 1995 (Revision 5 or later FAA- approved revisions) Part II, and having 10,000 or more (in the case of the 720 Series) or 15,000 or more (in the case of the 707 Series) hours time in service on the effective date of this AD, within the next 400 (for 720 Series) or 600 (for 707 Series) hours time in service unless accomplished within the last 400 (for 720 Series) or 600 (for 707 Series) hours time in service, and at intervals thereafter not to exceed 800 (for 720 Series) or 1200 (for 707 Series) hours time in service. \n\t(l)\tAirplanes having cracks which require rework under this AD may be flown in accordance with FAR 21.197 with the concurrence of Chief, Aircraft Engineering Division, FAA Western Region, to a base where the rework can be accomplished. \n\t(m)\tUpon request of an operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator. \n\tAmendment 39-640 became effective September 23, 1968. \n\tAmendment 39-670 became effective October 18, 1968. \n\tAmendment 39-852 became effective November 1, 1969. \n\tAmendment 39-891 became effective December 16, 1969. \n\tAmendment 39-923became effective January 17, 1970. \n\tThis Amendment 39-1420 becomes effective April 1, 1972.
98-13-01: This amendment adopts a new airworthiness directive (AD), applicable to certain Aerospatiale Model ATR42 and ATR72 series airplanes, that requires replacement of the left longitudinal net of the forward cargo compartment with a new reinforced net. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent blockage of the access door, which could restrict access for crewmembers between the flight deck and the passenger compartment during normal operations or an emergency evacuation.
90-25-12: 90-25-12 BOEING: Amendment 39-6821. Docket No. 90-NM-119-AD. \n\n\tApplicability: Model 737 series airplanes, listed in Boeing Service Bulletin 737-35- 1033, dated March 15, 1990, certificated in any category. \n\n\tCompliance: Required within 3,000 hours time-in-service after the effective date of this AD, unless previously accomplished. \n\n\tTo prevent fire caused by the chafing of wire bundles on crew oxygen system tubing, accomplish the following: \n\n\tA.\tInspect the clearances between the crew oxygen system tubing, the auxiliary power unit (APU) power feeder wire bundle, and the horizontal stabilizer trim control cables, located below the control cabin floor, in accordance with Boeing Service Bulletin 737-35-1033, dated March 15, 1990. \n\n\t\t1.\tIf there is inadequate clearance or damage has occurred, prior to further flight, repair the damage, replace the oxygen system tubing with modified tubing, and perform a leak check, in accordance with the Service Bulletin. \n\n\t\t2.\tIfclearance is inadequate between the crew oxygen tubing and the wire bundle only, and no damage has occurred, install floating loop clamps and spacers to obtain sufficient clearance between the tubing and the wire bundle, in accordance with the Service Bulletin. \n\n\tB.\tAn 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, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Seattle ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Seattle ACO. \n\n\tC.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington. \n\n\tThis amendment (39-6821, AD 90-25-12) becomes effective on January 7, 1991.
85-01-02 R1: 85-01-02 R1 McDONNELL DOUGLAS: Amendment 39-4978 as amended by Amendment 39-5241. Applies to Model DC-9 and C-9 (Military) series airplanes, Fuselage Numbers 1 through 1087, certificated in any category. \n\n\tCompliance required as indicated, unless previously accomplished. \n\n\tTo detect fatigue cracks and prevent structural failure of the fuselage cabin aft ventral pressure bulkhead, P/N 5910130 (and rework drawing numbers 5924597 and 5924350), accomplish the following: \n\n\tA.\tExcept for those airplanes that are subject to paragraph K., below, for airplanes with 15,000 or more landings on or after the effective date of this AD, within the next 700 additional landings or, for airplanes which have been inspected in accordance with AD 80-10-03, at the next inspection scheduled in accordance with AD 80-10-03 after the effective date of this AD, conduct visual and x-ray inspections in accordance with the instructions contained in Service Sketch 2934C, of McDonnell Douglas Service Bulletin 53-137, Revision 3, dated August 4, 1983. \n\n\tNOTE: McDonnell Douglas Service Bulletin 53-137, Revision 3, dated August 4, 1983, is herein referred to as SB 53-137, R3. \n\n\tB.\tIf no cracks are found in the bulkhead web and doublers, in the areas of the doorjamb upper LH and RH corners, identified in Figure 1 of Service Sketch 2934C of SB 53-137, R3, for those corner areas: \n\n\t\t1.\tAt intervals not to exceed 2,000 landings from last inspection, repeat the visual inspections required by paragraph A., above; and \n\n\t\t2.\tAt intervals not to exceed 4,000 landings from last inspection, repeat the x-ray inspections required by paragraph A., above. \n\n\t\t3.\tRepetitive inspections must be continued until terminating action in accordance with paragraph K., below, has been accomplished. \n\n\tC.\tIf cracks are found in the bulkhead web and doublers, for both Group I and II airplanes (as defined in SB 53-137, R3) accomplish the following: \n\n\t\t1.\tFor cracks found in the upper left and/or right bend radius, with no radial cracks, modify and inspect upper corners of bulkhead in accordance with paragraph 2.C. (Accomplishment Instructions) of SB 53-137, R3. \n\n\t\t2.\tFor new cracks found in the upper left bend radius, with no radial cracks and with Condition II modification installed, modify and inspect corners of bulkhead per paragraph 2.D. of SB 53-137, R3. \n\n\t\t3.\tFor new cracks found in the upper right bend radius, with no radial cracks and with Condition I modification installed, modify and inspect upper corners of bulkhead in accordance with paragraph 2.E. of SB 53-137, R3. \n\n\t\t4.\tFor radial crack(s) found in any or all members of upper right corner, with or without bend radius cracks, modify and inspect corners of bulkhead in accordance with paragraph 2.I. of SB 53-137, R3. \n\n\t\t5.\tFor radial and/or bend radius crack(s), found in upper left and/or right corner(s) or portion of upper left and/or right bulkhead skin which exceeds the repairable limits of Conditions I through XVII, XX, or XXII through XXXIV (any noted conditions), or at operator's option for replacement of cracked parts that are within repairable limits, in lieu of repair modification, replace cracked parts in accordance with paragraph 2.J. of SB 53-137, R3. Reinstate repetitive inspections of the upper corners within 15,000 landings and at intervals noted in paragraphs B.1., B.2., and B.3., above. \n\n\t\t6.\tFor radial crack(s) found in any or all members of lower left and/or right corner radius (except P/N 5910130-9/-11 side channel), with or without cracks in the lower inboard tab portion of the P/N 5910130-3/-5 bulkhead skin, modify and inspect corners of bulkhead in accordance with paragraph 2.L. of SB 53-137, R3. \n\n\t\t7.\tFor crack(s) found in any or all members which extends below the lower edge of the P/N 5910130-37 channel forward vertical leg; replace cracked parts, and inspect corners in accordance with paragraph 2.N. of SB 53-137, R3. Reinstate repetitive inspectionswithin 15,000 landings and at intervals noted in paragraphs B.1., B.2., and B.3., above. \n\n\t\t8.\tFor radial crack(s) found in P/N 5910130-3/-5 side bulkhead skin at one or more of the five lower inboard fasteners through the P/N 5910130-9/-11 side channel, install doublers and inspect lower corners in accordance with paragraph 2.0. of SB 53-137, R3. \n\n\t\t9.\tFor new radial crack(s) found in P/N 5910130-3/-5 side bulkhead skin at one or more of the five inboard fasteners through the P/N 5910130-9/-11 side channel, and preventive corner repair modification in accordance with Condition XIII has been installed, modify and inspect in accordance with paragraph 2.P. of SB 53-137, R3. \n\n\t\t10.\tFor new radial crack(s) found in any or all members of upper right corner, with or without bend radius cracks in P/N 5910130-13 doubler, and with Condition I modification installed, modify and inspect upper corners in accordance with paragraph 2.Y. of SB 53-137, R3. \n\n\tD.\tIf no crack(s) are found inthe lower corners of the bulkhead web and doublers, for both Groups I and II airplanes (as defined in SB 53-137, R3); install preventive repair modification and repetitively inspect lower corners at 15,000 landing intervals in accordance with paragraph 2.M. of SB 53-137, R3. If preventive repair is not installed, repetitively inspect at intervals in accordance with paragraph B., above. \n\n\tE.\tIf cracks are found in the bulkhead web and doublers in Group I airplanes, accomplish the following: \n\n\t\t1.\tFor radial crack(s) found in any or all members of upper left or both upper left/right corners, with or without bend radius cracks, modify and inspect corners of bulkhead in accordance with paragraph 2.F. of SB 53-137, R3. \n\n\t\t2.\tFor new radial crack(s) found in any or all members of upper right corners, with or without bend radius cracks, and with Condition XXII modification installed, modify and inspect corners of bulkhead in accordance with paragraph 2.Q. of SB 53-137, R3. \n\n\t\t3.For radial crack(s) found in P/N 5910130-13 doubler at upper left and/or right corner running along heel line of P/N 5910130-95 left and/or -96 right door stop angle, with or without radial cracks in any or all member of upper left and/or right corners, and with or without crack(s) in upper left and/or right bend radius of the forward flange of P/N 5910130-13 doubler, modify and inspect upper corners of bulkhead in accordance with paragraph 2.T. of SB 53-137, R3. \n\n\t\t4.\tFor new radial crack(s) found in any or all members of upper left corner, with or without bend radius cracks, and with Condition V modification installed, modify and inspect upper corners of bulkhead in accordance with paragraph 2.V. of SB 53-137, R3. \n\n\t\t5.\tFor new radial crack(s) found in any or all members of upper left corner, with or without bend radius cracks in P/N 5910130-13 doubler, and with Condition II modification installed, modify and inspect upper corners in accordance with paragraph 2.W. of SB 53-137,R3. \n\n\tF.\tIf no crack(s) are found in the bulkhead web and doublers in Group I airplanes, the interim crack preventative repair modification may be installed and the upper corners repetitively inspected at 15,000 landing intervals in accordance with paragraph 2.G. of SB 53-137, R3. If preventive repair modification is not installed, repetitively inspect at intervals specified in accordance with paragraph B., above. \n\n\tG.\tIf cracks are found in the bulkhead web and doublers for Group II airplanes, accomplish the following: \n\n\t\t1.\tFor radial crack(s) found in any or all members of upper left or both upper left/right corners, with or without bend radius cracks, modify and inspect corners of bulkhead in accordance with paragraph 2.H. of SB 53-137, R3. \n\n\t\t2.\tFor new radial crack(s) found in any or all members of upper right corner, with or without bend radius cracks, and with Condition IV modification installed, accomplish paragraph 2.R. of SB 53-137, R3. \n\n\t\t3.\tFor new radialcrack(s) found in any or all members of upper left corner, with or without bend radius cracks, and with Condition V modification installed, accomplish paragraph 2.S. of SB 53-137, R3. \n\n\t\t4.\tFor radial crack(s) found in P/N 5910130-13 doubler, at upper left and/or right corner running along heel line of P/N 5910130-95/-96, side door stop angle, with or without radial cracks in any or all members of upper left and/or right corners, and with or without crack(s) in upper left and/or right bend radius of forward flange of P/N 5910130-13, doubler, modify and inspect corners of bulkhead in accordance with paragraph 2.U. of SB 53-137, R3. \n\n\t\t5.\tFor radial crack(s) found in any or all members of upper left corner, with or without bend radius cracks in 5910130-13 doubler, and with Condition II modification installed, modify and inspect upper corners in accordance with paragraph 2.X. of SB 53-137, R3. \n\n\tH.\tIf no crack(s) are found in the bulkhead web and doublers in Group II airplanes, the interim crack preventive modification may be installed and the upper corners repetitively inspected at 15,000 landing interval in accordance with paragraph 2.K. of SB 53-137, R3. If preventive repair modification is not installed, repetitively inspect at intervals specified in accordance with paragraph B., above. \n\n\tI.\tMcDonnell Douglas DC-9 Service Bulletin A53-144, or later FAA approved revisions, must be accomplished within 18 months after the effective date of this AD. \n\n\tJ.\tFor aircraft modified per DC-9 Service Bulletin 53-139 (basic), or production equivalent, accomplish McDonnell Douglas DC-9 Service Bulletin 53-157 within 18 months after effective date of this AD. \n\n\tK.\tFor airplanes previously modified in accordance with DC-9 Service Bulletin 53-139 (Original Issue or Revision 1), or production equivalent, accomplish modification of the pressure bulkhead in accordance with Part 2 of the Accomplishment Instructions of DC-9 Service Bulletin 53-165 prior to the accumulation of 15,000 cycles after accomplishment of the modification or aircraft delivery (as applicable) or prior to the accumulation of the 5,400 landings after the effective date of Amendment 39-5241, whichever occurs later. Accomplishment of the modification referenced above constitutes terminating action for the requirements of this AD. \n\n\tL.\tFor airplanes not previously modified in accordance with DC-9 Service Bulletin 53-139 (Original Issue and Revision 1), or production equivalent, accomplishment of modifications in accordance with DC-9 Service Bulletin 53-166, Revision 1, dated January 31, 1983, or later FAA-approved revision, constitutes terminating action for the requirements of this AD. \n\n\tM.\tPrevious accomplishment of any rework or inspection(s), or portion(s) thereof, which is outlined in McDonnell Douglas DC-9 Service Bulletin 53-137, Revision 1, dated December 6, 1979, required by AD 80-10-03, Amendment 39-3769, effective May 15, 1980, which is also outlined in McDonnellDouglas DC-9 Service Bulletin 53-137, Revision 2, or SB 53-137, R3, provided for in AD 80-10-03, may be considered an equivalent to that requirement of this AD. \n\n\tN.\tThe inspections and modifications required by this AD need not be accomplished if, after the effective date of this AD, the aircraft is operated without cabin pressurization and a placard is installed in the cockpit, in full view of the pilots, stating: "OPERATION WITH CABIN PRESSURIZATION IS PROHIBITED." \n\n\tO.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of modifications required by this AD. \n\n\tP.\tUpon the request of an operator, an FAA Maintenance Inspector, subject to prior approval by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region, may adjust the repetitive inspection intervals of the operator if the request contains substantiating data to justify the increase for that operator. \n\n\tQ.Alternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). These documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or the Los Angeles Aircraft Certification Office, 4344 Donald Douglas Drive, Long Beach, California. \n\n\tThis supersedes AD 80-10-03, Amendment 39-3769 (45 FR 31052; May 15, 1980). \n\n\tAmendment 39-4978 became effective February 14, 1985. \n\n\tThis Amendment 39-5241 becomes effective March 31, 1986.
2012-22-13: We are adopting a new airworthiness directive (AD) for certain Sikorsky Aircraft Corporation (Sikorsky) Model S-76C helicopters. This AD requires installing an improved throttle stop and a wider trigger on the engine control levers (ECL). This AD was prompted by a bird-strike to the windshield that resulted in unintended movement of the engine control levers from the forward position and towards the flight-idle position, which reduced power on both engines. These actions are intended to prevent unintended movement of the ECLs, resulting in main rotor speed decay and subsequent loss of control of the aircraft.
2012-22-14: We are superseding an existing Emergency airworthiness directive (AD) for Sikorsky Aircraft Corporation (Sikorsky) Model S-70, S-70A, S-70C, S-70C(M), and S-70C(M1) helicopters with a certain part- numbered intermediate gearbox (IGB). The existing Emergency AD requires a one-time inspection of the internal oil passages of the IGB for an obstruction. That Emergency AD was prompted by an accident that resulted from blockage of oil in the IGB by a plug that was inadvertently left in the IGB during the coating of the IGB housing. We are issuing this supersedure to that Emergency AD to include two additional part numbers of affected IGBs and identify a specific date since new or overhaul of the affected IGBs. The actions specified by this AD are intended to detect a plug in the IGB and prevent overheating and seizing of the IGB, failure of the tail rotor drive output shaft, loss of tail rotor drive, and subsequent loss of control of the helicopter.
98-12-29: This amendment adopts a new airworthiness directive (AD), applicable to Lucas Air Equipment electric hoists (hoists) installed on, but not limited to, all models of Eurocopter France SA-360 and SA-365 helicopters that requires visually inspecting the cable for damage before the next hoist operation, blanking (plugging) the electronic control box upper vent, and performing an end-of-travel procedure before each hoist operation. This amendment is prompted by several incidents of cable failures caused by dynamic overload on the winding-up limit due to uncontrolled excessive speed of the cable, which is normally regulated by the automatic speed-reducing mechanism or the operator. The actions specified by this AD are intended to prevent breaking of the cable, which could become entangled with a main rotor or tail rotor blade, and result in damage or separation of a rotor blade, and subsequent loss of control of the helicopter.
87-06-11: 87-06-11 BOEING: Amendment 39-5576. Applies to Model 747 convertible and freighter series airplanes listed in Boeing Alert Service Bulletin 747-53A2273, dated July 29, 1986, certificated in any category. \n\n\tTo prevent inadvertent in-flight movement of main deck cargo on freighter or convertible airplanes, accomplish the following, unless already accomplished: \n\n\tA.\tWithin 10 landings after the effective date of this AD, or immediately after the replacement or reinstallation of the restraint hardware, unless the requirements of paragraph B., for removal of placards, decals, or stencils, are accomplished, install suitable placards, decals, or stencils at all restraints along both left and right buttock lines (BL) 98.2, 9.8, and 1.8, between body stations (BS) 980 and 1500, that state the following: \n\n\t\t"This restraint is inoperative. Cargo must be tied down per FAA-approved procedures." \n\n\tB.\tPlacards, decals, or stencils installed in accordance with the provisions of paragraph A., above, may be removed at each location where a determination is made in accordance with Boeing Service Bulletin 747-53A2273, dated July 29, 1986, or later FAA- approved revisions, that the restraint lip overlap meets or exceeds the minimum value specified therein, or if terminating action defined in paragraph C., below, is accomplished. \n\n\tC.\tTerminating action for this amendment consists of the inspection of the seat track alignment and, if necessary, modification of the floor beams between body station 980 and 1480 in accordance with Section III., Part II of the Boeing Service Bulletin 747-53A2273, dated July 29, 1986, or later FAA-approved revisions. \n\n\tD.\tAn alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\tE.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received the applicable service bulletin from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124-2207. This information 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. \n\n\tThis Amendment becomes effective April 13, 1987.
2012-22-16: We are adopting a new airworthiness directive (AD) for all Pratt & Whitney Division PW4050, PW4052, PW4056, PW4060, PW4060A, PW4060C, PW4062, PW4062A, PW4152, PW4156, PW4156A, PW4158, PW4160, PW4460, PW4462, and PW4650 turbofan engines, including models with any dash number suffix. This AD was prompted by 16 reports of damaged or failed 3rd stage low-pressure turbine (LPT) duct segments. This AD requires removing from service certain part numbers (P/Ns) of 3rd stage LPT duct segments. We are issuing this AD to prevent failure of the 3rd stage LPT duct segments, which could lead to LPT rotor damage, uncontained engine failure, and damage to the airplane.
98-12-26: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F.28 Mark 1000, 2000, 3000, and 4000 series airplanes, that requires a one-time inspection to determine the torque values of the coupling fitting attachment bolts at fuselage station 10790, and corrective action, if necessary. This amendment is prompted by issuance of mandatory continuing airworthiness information by a foreign civil airworthiness authority. The actions specified by this AD are intended to prevent loss of the coupling fitting attachment bolts between the center wing section and the fuselage, and consequent reduced structural integrity of the airplane.
83-22-08: 83-22-08 B.F. GOODRICH: Amendment 39-4765. Applies to B.F. Goodrich Emergency Evacuation Slide/Rafts P/Ns 7A1340 series, 7A1342 series, 7A1371 series, and 7A1373 series. These slide/rafts are installed on Boeing Model 747-100 and 747-200B airplanes in accordance with Supplemental Type Certificate (STC) SA574GL, and on Boeing Model 747-100B, 747SR, and 747-300 airplanes in accordance with STC SA575GL. Compliance required as indicated below. To assure proper functioning of B.F. Goodrich Slide/Rafts, accomplish the following unless previously accomplished in accordance with the procedures in B.F. Goodrich Service Bulletin 25-081, Revision 1, dated September 27, 1983. \n\n\tA.\tFor B.F. Goodrich Slide/Rafts, P/Ns 7A1340 series or 7A1371 series S/Ns G001 thru G289, and P/Ns 7A1342 series or 7A1373 series, S/Ns G001 thru G087: Within 20 calendar days after the effective date of this AD perform the maintenance procedures contained in B. F. Goodrich Service Bulletin 25-081, Revision 1, dated September 27, 1983, or subsequent FAA approved revisions, on installed slide/rafts or on slide/rafts prior to installation. Repeat inspections per paragraph C., below. \n\n\tB.\tFor all B.F. Goodrich Slide/Rafts, P/N's 7A1340 series, 7A1342 series, 7A1371 series and 7A1373 series, inspect all serial numbers installed prior to September 1, 1983, and not inspected per paragraph A., above, in accordance with B.F. Goodrich Service Bulletin 25-081, Rev. 1, dated September 27, 1983, or subsequent FAA approved revision, within 120 days of the effective date of this AD. Slides installed subsequent to August 31, 1983, must be inspected in accordance with B.F. Goodrich Service Bulletin 25-081, Revision 1, dated September 27, 1983, or subsequent FAA approved revision, within 180 days of last leak check inspection. (Reference Service Bulletin Section 2B.) Repeat inspections per Paragraph C., below. \n\n\tC.\tTo prevent undetected deterioration repeat the inspection of all installed B.F. Goodrich Slide/Rafts, P/Ns 7A1340 series, 7A1342 series, 7A1371 series and 7A1373 series, in accordance with Section 2B of B.F. Goodrich Service Bulletin 25-081, Rev. 1, dated September 27, 1983, or subsequent FAA approved revision, within 180 days of last inspection or within 180 days of installation. \n\n\tD.\tAlternate means of compliance which provide an equivalent level of safety may be used when approved by the Manager, Chicago Aircraft Certification Office, FAA, Central Region. \n\n\tE.\tUpon request of operator, an FAA Principal Maintenance Inspector, subject to prior approval by Manager, Chicago Aircraft Certification Office, FAA, Central Region, can adjust the compliance times if the request contains substantiating data to justify the increase for the operator. \n\n\tF.\tSpecial 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. \n\n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to B.F. Goodrich Company, Attn: Mr. Earl Lucas, Dept. 1809, Bldg. 17F, 500 South Main Street, Akron, Ohio 44318. \n\n\tThese documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington. \n\n\tThis amendment becomes effective November 15, 1983.
86-01-51 R1: 86-01-51 R1 BOEING: Amendment 39-5269. Applies to all Model 747 series airplanes equipped with General Electric CF6 engines, certificated in any category. To prevent accumulation of flammable fluid in the engine strut, which constitutes a fire hazard, accomplish the following, unless already accomplished: \n\n\t1.\tWithin the next 72 hours time-in-service, inspect and, if necessary, clear the engine strut drains in accordance with the procedures specified in Chapter 71-71-00 of Boeing Maintenance Manual for the Boeing Model 747 airplanes equipped with General Electric CF6 engines. Repeat these inspections at intervals not to exceed 1,000 hours time-in-service. Report results of inspections to FAA, Northwest Mountain Region, Seattle Aircraft Certification Office, Propulsion Branch, ATTN: Kanji Patel, ANM-140S, 17900 Pacific Highway South, C-68966, Seattle, Washington 98168. \n\n\t2.\tThe repetitive inspections required by paragraph 1., above, may be terminated upon the accomplishment of the following: \n\n\t\tA.\tFor Group I and II airplanes defined in Boeing Service Bulletin 747-71-2146, dated November 24, 1978, accomplish the replacement of the inboard and outboard pylon drain lines only, as shown in Figures 2 and 3, respectively, of that service bulletin or later FAA-approved revisions. \n\n\t\tNOTE: The portion of this service bulletin pertaining to "Strut Pylon Drain Line Engine Tube Replacement," Figure 1, has been superseded by Boeing Service Bulletin 747-71-2155, Revision 3, dated September 28, 1984. \n\n\t\tB.\tFor Group I airplanes, as defined in Boeing Service Bulletin 747-71-2155, Revision 3, dated September 28, 1984, accomplish replacement of the engine strut aft drain and incorporate the bracket/heat shield, in accordance with Work Package I described in that service bulletin, or later FAA-approved revisions; and install forward drain fitting, drain tube, and fluid dam in accordance with Work Package II described in that same service bulletin, or later FAA-approved revisions. \n\n\t\tC.\tFor Group II airplanes, as defined in Boeing Service Bulletin 747-75-2155, Revision 3, dated September 28, 1984, install forward drain fitting, drain tube, and fluid dam in accordance with Work Package II described in that service bulletin, or later FAA-approved revisions. \n\n\t\tD.\tFor Group III airplanes, as defined in Boeing Service Bulletin 747-75-2155, Revision 3, dated September 28, 1984, accomplish replacement of the engine strut aft drain and incorporate the bracket/heat shield, in accordance with Work Package I described in that service bulletin, or later FAA-approved revisions. \n\n\t\tE.\tFor Group I, II, and IV airplanes, as defined in Boeing Service Bulletin 747-75- 2155, Revision 3, dated September 28, 1984, accomplish Work Package III, in accordance with that service bulletin, or later FAA-approved revisions, if the thrust reverser follow-up cable support bracket is found to interfere with the fluid dam. \n\n\t3.\tAlternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA Northwest Mountain Region. \n\n\t4.\tSpecial 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 inspections and, if necessary, cleaning of the engine strut drain lines. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. 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. \n\n\tThis amendment becomes effective April 18, 1986, as to all persons, except those persons to whom it was made immediately effective by telegraphic AD T86-01-51, issued on January 9, 1986.
2012-21-15: We are adopting a new airworthiness directive (AD) for all Airbus Model A300 B4-600, B4-600R, and F4-600R series airplanes, and Model A300 C4-605R Variant F airplanes (collectively called A300-600 series airplanes); and Model A310 series airplanes. This AD was prompted by events of excessive rudder pedal inputs and consequent high loads on the vertical stabilizer on several airplanes. This AD requires either incorporating a design change to the rudder control system and/ or other systems, or installing a stop rudder inputs warning (SRIW) modification. We are issuing this AD to prevent loads on the vertical stabilizer that exceed ultimate design loads, which could cause failure of the vertical stabilizer and consequent reduced controllability of the airplane.
98-12-15: This amendment adopts a new airworthiness directive (AD), applicable to Eurocopter France Model AS 332C, L, L1, and L2 helicopters that requires visually inspecting the intermediate gearbox-to-structure attachment stirrup (stirrup) front tabs for cracks, and if a crack is discovered, removing the intermediate gearbox and replacing it with an airworthy intermediate gearbox; and inspecting for the conformity of the attachment parts. This amendment is prompted by five reports of failure of the two stirrup tabs. The actions specified by this AD are intended to prevent failure of the intermediate gearbox stirrup front tabs, loss of anti-torque drive, and subsequent loss of control of the helicopter.
98-12-16: This amendment adopts a new airworthiness directive (AD), applicable to Eurocopter France Model SA 330F, G, and J helicopters that requires visually inspecting the intermediate gearbox (IGB) fairing safety stop (safety stop) for cracks, crazing, or edge wear, and if a crack, crazing, or edge wear exceeds the established limits, replacing the safety stop; and, inspecting to ensure that the inclined drive shaft fairing hinge pin is properly locked. A terminating action is provided in the AD by installing an additional safety stop on the IGB fairing. This amendment is prompted by one report of an accident involving the loss of the inclined drive shaft fairing. The actions specified by this AD are intended to prevent loss of the inclined drive shaft fairing, impact with the tail rotor, and subsequent loss of control of the helicopter.
98-12-10: This amendment supersedes Airworthiness Directive (AD) 93-10-11, which currently requires installing an inspection opening in the wing, repetitively inspecting the upper wing spar cap for cracks, and repairing any cracks on all Avions Mudry et Cie (Avions) Model CAP 10B airplanes. This AD will retain the same actions already required by AD 93-10-11, and will add inspecting, and repairing if necessary, the lower surface of the wing spar. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for France. The actions specified by this AD are intended to prevent structural cracks in the wing spar, which could lead to loss of a wing and loss of control of the airplane.
79-09-03: 79-09-03 BOEING: Amendment 39-3454. Applies to all Model 747 series airplanes certificated in all categories. Compliance required within 500 hours time-in-service after the effective date of this AD unless already accomplished. \n\tTo prevent takeoff with a takeoff aural warning system that would not indicate an unsafe takeoff condition, accomplish the following: \n\tPerform a one-time check of the aural warning system and, as necessary, adjust and re- check the aural warning activation switch, the auto-brake deactivation and auto-speedbrake retraction switches in accordance with Boeing Alert Service Bulletin 747-31A2056, Rev. 1, dated April 6, 1979, or later FAA approved revisions, or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA, Northwest Region. \n\tThe 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). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis amendment becomes effective May 10, 1979.
81-19-05: 81-19-05 McDONNELL DOUGLAS: Amendment 39-4216. Applies to McDonnell Douglas Model DC-9-80 series airplanes, certificated in all categories with thrust reverser barrel assemblies P/N 5938033-1 installed. Compliance required as indicated unless already accomplished. To prevent the separation of the thrust reverser from the airplane as a result of a failure of the brazed joint at the thrust reverser attach flange, accomplish the following:\n\n\tA.\tWithin three days from the effective date of this AD, and at intervals of 24 hours calendar time thereafter, conduct a visual inspection, with a 3 to 5 power magnifying glass, of the exhaust nozzle barrel assembly exterior skin P/N 5938033-5 at attach flange P/N 5938033-37 joint for buckling, obvious separation, or circumferential cracks especially in the zone 30 degrees above and below the thrust reverser stang fairings on each side. If buckling separation or cracks are found remove and replace barrel assembly prior to further flight. Chapter 78-30-01 of the maintenance manual contains additional information on this subject.\n\n\tB.\tWithin three days from the effective date of this AD inspect all thrust reversers to identify each thrust reverser barrel serial number. Remove from service reverser assemblies identified by the following barrel serial numbers.\n\n\tAccomplishment instructions for this inspection are contained in McDonnell Douglas Alert Service Bulletin A78-50 Revision 1, dated July 20, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, Northwest Region.\n\n\tSerial Numbers D3-0005, D3-0008, D3-0010 thru D3-0012, D3-0024, D3-0025, D3-0043, D3-0047, D3-0048, D3-0057, D3-0058, D3-0060, D3-0063, D3-0064, D3-0069, D3-0084, D3-0091, D3-0093, D3-0095, D3-0100 thru D3-0105, D3-0108, D3-0112, D3-0113, D3-0115, D3-0119, D3-0129A, D3-0132, D3-0137, D3-0156, D3-0162, D3-0165, and D3-0172. \n\n\tNOTE:\n\n\tTypical example of the number found on the reverser barrel assembly is "5938033-1-01-D3-0105." This number is located approximately six inches aft of the upper hydraulic thrust reverser latch and is etched on the exterior surface of the thrust reverser exhaust duct barrel assembly near top centerline. 5938033-1 is the part number and will be the same on each unit. The last six digits (D3-0105) constitute the listed serial number.\n\n\tC.\tWithin 120 days from the effective date of the AD modify all P/N 5938033-1 reverser barrel assemblies by welding the engine attach flange - outer barrel skin joint in accordance with the accomplishment instructions of McDonnell Douglas Service Bulletin 78-51 dated August 25, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, Northwest Region. Upon completion of this modification reidentify the reverser barrel assembly as P/N 5938033-501 and the thrust reverser assembly as P/N 5938050-501 in accordance with paragraph 2.B. (6) and (9) of the accomplishment instructions of the bulletin. This modification constitutes terminating action for all requirements of this AD. All reverser barrel assemblies removed from service per paragraph B above may be modified in accordance with this paragraph and returned to service.\n\n\tD.\tSpecial flight permits may be issued in accordance with FAR's 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD.\n\n\tE.\tAlternative means of compliance or other actions which provide an equivalent level of safety may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Region. \tThe 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).\n\n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Director, Publications and Training, C1-750 (54-60). These documents also may be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108, or 4344 Donald Douglas Drive, Long Beach, California 90808.\n\n\tThis amendment becomes effective September 28, 1981.
98-12-14: This amendment adopts a new airworthiness directive (AD) that applies to certain AERMACCHI S.p.A. (AERMACCHI) S.205 series and Models S.208 and S.208A airplanes. This AD requires inspecting the flap cable pulley bracket for correct alignment and correcting any misalignment; inspecting the flap control cable for wear (nicks, cuts, frays, etc.), and replacing the flap control pulley bracket and flap control cable if worn. 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 flap control failure, which could result in loss of control of the airplane.