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62-25-01: 62-25-01 BOEING: Amdt. 509 Part 507 Federal Register November 21, 1962. Applies to Models 707 and 720 Series Aircraft. \n\n\tCompliance with certain service bulletins pertaining to the Model 707 and Model 720 flight control systems is considered necessary to provide for significant improvement in the safety and reliability of operation of those aircraft models. Accordingly, the aircraft models listed below shall be inspected and/or modified within the compliance times and in accordance either with the service bulletins as indicated or with equivalent methods approved by the Chief, Engineering and Manufacturing Branch, FAA Western Region. Airplanes modified in accordance with later FAA approved revisions of the service bulletins listed below will be considered to have complied with the appropriate provisions of this AD. \n\n\t(a) Compliance required within the next 400 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.1. Stabilizer trim actuator auxiliary brake retaining nut. \n707 & 720\n984 \n\n\t(b) Compliance required within the next 650 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.\n1. Rudder pedal push-rod attachment. \n707\n337 \n2. Inboard aileron tab nose weight attachment screws. \n707\n860 \n3. Spoiler and emergency flap switch placard installation. \n707 & 720\n1524 \n\n\tBAC P/N 10-60424-621 (Type I) and P/N 10-60424-184 (Type I) are approved equivalents. \n\n\t(c) Compliance required within the next 2,700 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No.\n1. Bearing retainer installation for center and inboard hinges for inboard aileron tab.\n707\n307 (R-1) and 307 (R-1)A.\n2. Guard installation for chain in stabilizer trim unit.\n707\n655 and 655B\n3. Flap drive torque tube guard installation in wheel well area.\n707\n680 and 680A\n4. Rudder controlinput stop modification and directional bushing replacement.\n707\n735 (R-1)\n5. Replacement of stabilizer trim actuator.\n707 & 720\n889\n6. Flap takeoff warning switch relocation.\n707 & 720\n1016 (R-1) and 1016 (R-1)C\n7. Stabilizer trim actuator motor replacement.\n707\n1247\n8. Emergency flap switch installation.\n707 & 720\n1251\n9. Rudder power con- trol unit replacement, "Extension Sleeve Revision".\n707 & 720\n1479 (R-1) Part I only\n10. Rudder pressure control valve bypass installation.\n707 & 720\n1482 (R-1).\n11. Rudder control centering spring cable modification.\n707 & 720\n1625\n12. Rudder control centering spring cable guard installation.\n707 & 720\n1680 & 1680A\n \n\n\t(d) Compliance required within the next 3,500 hours time in service following the effective date of the AD. \n\n\nModification\nModel\nService Bulletin No. \n1. Inboard aileron centering spring cartridge. \n707 & 720\n1344 \n2. Control wheel stabilizer trim switch installation. \n707 & 720\n1410 and 1410B\n3. Replacement of rudder hydraulic system solenoid valve. \n707 & 720\n1490 (R-1) \n\n\tBAC P/N's 10-3265-6 and 10-60705-1 are approved equivalents. \n\n\t(e) Compliance required within the next 5,000 hours' time in service following the effective date of this AD: \n\n\nModification\nModel\nService Bulletin No. \n1. Stabilizer trim actuator brake unlock gear ball bearing adapter addition. \n707 & 720\n1128 & 1128A \n2. Stabilizer trim actuator brake pawl spring. \n707 & 720\n1237\n3. Outboard spoiler shutoff valve consolidation.\n707 & 720\n1336 (R-1) and 1336 (R-1)B. \n4. Replacement of spoiler hydraulic system shutoff valve. \n707 & 720\n1484\n\n\tThis directive effective December 20, 1962. \n\n\tRevised December 24, 1963.
84-02-05: 84-02-05 BOEING: Amendment 39-4798. Applies to Boeing Model 747 series airplanes certificated in all categories. Compliance required as indicated, unless already accomplished. \n\n\tA.\tTo clarify the operation of the anti-icing system, emphasize the need to maintain the specified minimum N1 engine rpm during icing conditions, and expand the definition of icing conditions, accomplish the following: Within 120 days from the effective date of this AD, unless already accomplished, revise the FAA approved Airplane Flight Manual (AFM) CERTIFICATE LIMITATIONS SECTION by adding: \n\n\t\t\t\t"ENGINE ANTI-ICE SYSTEM \n\n\tWhen penetrating or operating in Icing Conditions, maintain a minimum of 50 percent N1 rpm at 10,000 feet and above, and 45 percent N1 rpm for Pratt & Whitney JT9D and General Electric CF6 engines, and 42 percent N1 for Rolls Royce RB211 engines, below 10,000 feet altitude, except as required for landing. \n\n\tNacelle anti-ice must be ON during all ground and flight operationswhen icing conditions exist or are anticipated, except during climb and cruise when the temperature is below -40 degrees C SAT. Nacelle anti-ice must be ON prior to and during descent in all icing conditions, including temperatures below -40 degrees C SAT. \n\n\t\tNOTE: Icing Conditions - Icing Conditions exist when the OAT on the ground and for takeoff, or TAT inflight is 10 degrees C or below and visible moisture in any form is present (such as clouds, fog with visibility of one mile or less, rain, snow, sleet, and ice crystals). \n\n\tIcing conditions also exist when the OAT on the ground and for takeoff is 10 degrees C or below when operating on ramps, taxiways or runways where surface snow, ice, standing water, or slush may be ingested by the engines or freeze on engines, nacelles or engine sensor probes." \n\n\tB.\tTo alert the flight crew of engine operation at a lower N1 than required for icing condition, install a LOW N1 rpm caution indication system as follows: \n\n\t\tWithin 24 months from the effective date of this AD, unless already accomplished, provide "LOW N1" indication that will alert the flight crew that the nacelle anti-ice is "ON and N1 is less than 45 percent N1 (42 percent N1 for RB211 engines) below 10,000 feet, and is less than 50 percent N1 above 10,000 feet altitude. \n\n\t\tNOTE: The LOW N1 indication may be provided by incorporating Boeing Service Bulletin S/B 747-77-2060 for the JT9D Pratt & Whitney powered airplanes and S/B 747-77-2063 for General Electric CF6 and Rolls Royce RB211 powered airplanes. \n\n\t\tBoth service bulletins have been approved by the FAA and were released on February 14, 1983. The service bulletins may be obtained from the Boeing Company at the following address: The Boeing Company, P.O. Box 3707, Seattle, Washington 98124. \n\n\tC.\tAlternate means of compliance with the AD which provide an equivalent level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, 9010 East MarginalWay South, Seattle, Washington. \n\n\tD.\tA special flight permit may be issued in accordance with FAR 21.197 and 21.199 for the purpose of flying the aircraft which has exceeded the compliance period to a maintenance facility where the modification can be performed. \n\n\tThis amendment becomes effective on March 2, 1984.
75-09-04 R1: 75-09-04 R1 BOEING: Amendment 39-2174 as amended by Amendment 39-4926. Applies to Boeing Model 727 series airplanes, certificated in all categories, listed in Boeing Service Bulletin 727-55-62, or later FAA approved revisions (line numbers 1 through 641, inclusive). Compliance required as indicated. \n\n\tTo detect cracks in the horizontal stabilizer rear spar center section fitting, accomplish the following: \n\n\tA.\tWithin the next 750 flight hours after the effective date of this AD, unless accomplished within the last 2250 flight hours, inspect the horizontal stabilizer rear spar center section fitting in accordance with Paragraph III of Boeing Service Bulletin 727-55-62, or later FAA approved revisions or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t1.\tIf no cracks are found in the fitting, repeat the inspections at intervals not to exceed 3000 flight hours, until replaced per Paragraph B. \n\n\t\t2.\tIf a crackis found at the upper or lower flanges of the fitting and is within the allowable limits specified in Figure 2 of Boeing Service Bulletin 727-55-62, or later FAA approved revisions, prior to further flight, stop drill the crack per the service bulletin. Inspect the stop drilled areas for crack growth at 1500 flight hour intervals, until repaired or replaced in accordance with Boeing Service Bulletin 727-55-62, or later FAA approved revisions, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\n\t\t3.\tIf a crack is found in any of the hinge lugs and is within the allowable limits specified in Boeing Service Bulletin 727-55-62, or later FAA approved revisions, before further flight, repair the fitting in accordance with Figure 3 of the service bulletin, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. Inspect the repaired fitting thereafter at intervals not to exceed 1500 flight hours. \n\n\t\t4.\tIf a crack at any location is beyond the allowable limits specified in Boeing Service Bulletin 727-55-62, or later FAA approved revisions, before further flight, repair in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, or replace the fitting with a new improved fitting per Paragraph B. \n\n\tB.\tAs terminating action for this AD, replace the horizontal stabilizer rear spar center section fitting with a new improved 7075-T73 aluminum allow fitting. \n\n\tC.\tAirplanes having cracked horizontal stabilizer rear spar center section fittings which require replacement under this AD may be flown in accordance with FAR 21.197 to a base where the replacement can be accomplished. \n\n\tD.\tUpon request of an operator, an FAA Maintenance Inspector, subject to the prior approval of the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, may adjust the inspection intervals in this AD, if the request contains substantiating data to justify the increase for that operator. \n\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\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. The documents may also be examined at FAA Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington. \n\n\tAmendment 39-2174 became effective April 29, 1975. \n\n\tThis Amendment 39-4926 becomes effective January 22, 1985.
90-23-12: 90-23-12 BOEING: Amendment 39-6799. Docket No. 90-NM-134-AD. \n\n\tApplicability: Model 737-300 and 737-400 series airplanes, listed in Boeing Service Bulletin 737-49-1071, dated May 10, 1990, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent auxiliary power unit (APU) rotor failure resulting from an undetected EGT overtemperature condition, accomplish the following: \n\n\tA.\tFor Model 737-400 series airplanes: Within 1,000 hours time-in-service after May 29, 1990 (the effective date of Amendment 39-6583, AD 90-09-05), modify the APU instrumentation wiring in a manner that will assure continuous flight-compartment APU exhaust gas temperature (EGT) indication following a shutdown. The modification must be accomplished in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Transport Airplane Directorate; or in accordance with Boeing Service Bulletin 737-49-1071, dated May 10, 1990. \n\n\tB.\tFor Model 737-300 series airplanes: Within 1,000 hours time-in-service after the effective date of this amendment, modify the APU instrumentation wiring in accordance with Boeing Service Bulletin 737-49-1071, dated May 10, 1990. \n\n\tC.\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\tD.\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 mayobtain 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\tAirworthiness Directive 90-23-12 supersedes AD 90-09-05, Amendment 39-6583. \n\tThis amendment (39-6799, AD 90-23-12) becomes effective on December 11, 1990.
59-12-05: 59-12-05 COLONIAL: Applies to Models C-1 and C-2 Aircraft Serial Numbers 1 Through 132. Compliance required as indicated. Due to a recent incident where the plastic lock for the control surface hinge pin cracked, thus making it possible for the hinge pin to work out, the following inspection and replacement of all plastic locks is required. Prior to next flight inspect the control surface hinge pin locks. (1) If made of metal, no further action necessary. (2) If made of plastic material inspect for cracks. Parts found cracked must be replaced with locks fabricated of 0.025 2024-T3 aluminum alloy material or equivalent before further operation. (3) All plastic locks must be replaced within the next 10 hours of operation with metal locks fabricated of 0.025 2024-T3 aluminum alloy material or equivalent. (Colonial Service Bulletin No. 15 covers this same subject.)
2010-10-10: We are adopting a new airworthiness directive (AD) for certain Hawker Beechcraft Corporation Model 390 airplanes. This AD requires you to inspect the essential bus lightning strike protection for proper installation of metal oxide varistor (MOV) and spark gap wiring. This AD also requires you to rework the wiring as necessary to achieve the required lightning strike/surge protection. This AD results from a report that the wires to the MOV and spark gap were swapped. We are issuing this AD to detect and correct improper installation of the MOV and spark gap wiring, which could result in overload of the MOV in a lightning strike and allow electrical energy to continue to the essential bus and disable equipment that receives power from the essential bus. The disabled equipment could include the autopilot, anti-skid system, hydraulic indicator, spoiler system, pilot primary flight display, audio panel, or the 1 air data computer. This failure could lead to a significant increase in pilot workload during adverse operating conditions.
91-13-06: 91-13-06 McDONNELL DOUGLAS: Amendment 39-7037. Docket No. 91-NM-03-AD. Supersedes AD 90-03-17. \n\n\tApplicability: Model DC-9 series, Model DC-9-80 series, C-9 (Military), and Model MD-88 airplanes, equipped with Westinghouse bus control unit (BCU) Part Number 947F946-2, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent the loss of generator electrical power, accomplish the following: \n\n\tA.\tWithin 30 days after February 15, 1990 (the effective date of AD 90-03-17, Amendment 39-6496), add the following to the LIMITATIONS section of the approved Airplane Flight Manual (AFM). This may be accomplished by inserting a copy of this AD into the AFM. \n\n\t\t1.\tOn takeoff with both engine generators and APU generator operating and APU bus switches selected to the ON position, the AC BUS X-TIE switch must be placed in the OPEN position. At or above 10,000 feet MSL with both engine generators operating, the APU may be shut down and the AC BUS X-TIE switch placed in the AUTO position. \n\n\t\tNOTE: In the event of an in-flight failure of an engine generator that results in the APU generator powering an AC bus, de-activate all galley power and place the other APU BUS switch in the ON position. \n\n\t\t2.\tOn takeoff with the APU generator inoperative, or an engine generator inoperative, dispatch is permitted in accordance with the present MMEL conditions except that the AC BUS X-TIE switch must be in the OPEN position. Verify that all transformer rectifiers (TR's) are operating and place the DC BUS X-TIE switch in the CLOSE position. Takeoff minimums are restricted to ceiling 1,000 foot and visibility 3 miles. The Captain must make the takeoff with his instrument incandescent flood lights adjusted to a level which would adequately compensate for the subsequent potential loss of his instrument integral lights. At or above 10,000 feet MSL, place the DC BUS X-TIE switch in the OPEN position and the AC BUSX-TIE switch in the AUTO position. \n\n\t\tNOTE: In the event of an in-flight failure of an engine generator following dispatching with the APU generator powering an AC bus, de-activate all galley power and place the other APU BUS switch in the ON position. \n\n\t\t3.\tWhen operating with the AC BUS X-TIE switch in the AUTO position, if rapid cycling of the AC cross-tie relay occurs, manifested by a buzzing/chattering sound from the electrical power center and any combination of random circuit breaker trips, inappropriate aural warning messages, loss of some flight instruments, and/or flashing cockpit annunciators, place the AC BUS X-TIE switch to the OPEN position. If a generator trips off-line, it may be reset only once. If the engine generator fault cannot be cleared, the APU should be utilized, if available. \n\n\t\t4.\tPrior to the approach with both engine generators operating, start the APU and place both APU BUS switches to the ON position. Place the AC BUS X-TIE switch to the OPENposition after APU electrical power becomes available. \n\n\t\t5.\tPrior to the approach with only two generators operating, place the AC BUS X-TIE switch in the OPEN position. Landing minimums are restricted to Category I and the Captain must make the approach with his instrument incandescent flood lights adjusted to a level which would adequately compensate for the subsequent potential loss of his instrument integral lights. \n\n\t\t6.\tIn the event of an in-flight failure that results in an AC bus not powered, place the DC BUS X-TIE switch in the CLOSE position. \n\n\t\t7.\tIn the event of an in-flight failure that results in both AC BUSES being powered by only one generator, the landing minimums are restricted to ceiling 1,000 feet and 3 miles visibility. The Captain must make the approach and landing. \n\n\t\t8.\tAutoland is permitted with two engine generators operating and APU generator operating with both APU BUS switches in the ON position and the AC BUS X-TIE switch in the OPEN position. Reconfirm APU generator availability after "AUT LND/AUT LND" is indicated on the Flight Mode Annunciator (FMA). An autoland approach must be discontinued following a failure of an engine generator. \n\n\tB.\tWithin 2 years after the effective date of this AD, replace the Westinghouse Bus Control Unit, Part Number 947F946-2 with the Westinghouse Bus Control Unit, Part Number 947F946-3, in accordance with McDonnell Douglas DC-9 Service Bulletin 24-119 dated January 24, 1990. The limitations on the electrical operating procedures and restricted operation of the automatic landing system required by paragraph A. of this AD may be removed from the AFM when the modified Westinghouse BCU is installed. \n\n\tC.\tAn alternate means of compliance and adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be submitted directly to the Manager, Los Angeles ACO, and a copy sent to the cognizant FAA Principal Inspector (PI). The PI will then forward comments or concurrence to the Los Angeles ACO. \n\n\tD.\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 requirement 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 McDonnell Douglas Corporation, 3855 Lakewood Boulevard, Long Beach, California 90846, Attention: Business Unit Manager, Technical Publications, C1-HCW (54-60). These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington, or the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California. \n\n\tThis amendment supersedes Amendment 39-6496, AD 90-03-17. \n\n\tThis amendment(39-7037, AD 91-13-06) becomes effective on July 15, 1991.
67-05-03: 67-05-03 MCDONNELL DOUGLAS: Amend. 39-347 as amended by amendment 39-1160. Applies to Model DC-8 Series (except DC-8-61) and DC-8F Series Airplanes as Indicated Herein. \n\n\tCompliance required as indicated. \n\n\tNumerous reports have been received concerning cracks in the inboard and outboard pylons and pylon stub wing structure. These cracks have been found in the skins, bulkheads, and spar caps. The cracks have varied in length, and in some cases, the parts have cracked completely through. To prevent further failure of this nature, accomplish the following, or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region: \n\n\t(a)\tFor all inboard and outboard pylons that have accumulated a total of 6,000 or more hours' time in service at the effective date of this AD, unless already accomplished within the last 2,000 hours' time in service, and for all inboard and outboard pylons accumulating a total of 6,000 hours' time in service after the effective date of this AD, within the next 1,000 hours' time in service, conduct a visual inspection as follows: \n\n\t\t(1)\tInspect outboard pylons in the area of the pylon stub wing lower skin and doubler for cracks in the cutout for the leading edge lower slot door on airplanes with Serial Numbers 45256-45272, 45274-45277, 45282, 45284-45289, 45292-45300, 45304-45306, 45376=45382, 45384-45390, 45392, 45393, 45408-45413, 45418, 45419, 45421, 45425-45427, 45429-45431, 45433-45437, 45445, 45565-45567, 45569, 45570, 45596, 45597, 45602, 45603, 45605, 45606, 45609-45612, 45626, 45627, 45638, 45645, 45646, 45649, 45650, 45672, 45673, 45687-45690, 45806-45808, 45815, 45877. Refer to Service Sketch No. 754 of Douglas Engineering Service Letter C1-78-2016/DBA, October 18, 1966, or later revisions for location of the cracks. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. If cracks arefound during this inspection or any reinspection, the defective parts must be replaced before further flight with uncracked parts, or parts reworked in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended as a redesign to prevent further cracking of these areas and approved as such, the repetitive inspection requirements may be discontinued. \n\n\t\t(2)\tInspect at Station YOP 229 the outboard pylon inboard skin in the vicinity of the slot leading edge on airplanes with Serial Numbers 45252-45272, 45274-45283, 45289, 45291-45300, 45304-45306, 45376-45382, 45384-45393, 45418, 45419, 45421-45427, 45429-45431, 45433-45437, 45442-45445, 45565-45567, 45569, 45570, 45588-45597, 45602, 45603, 45605, 45609-45612, 45626, 45627, 45638, 45645, 45646, 45649, 45650. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during the inspection or any reinspection must be reworked before further flight in accordance with Douglas Service Bulletins Nos. 54-30, Kits J and K, dated November 27, 1962, 54-31, Kits L and M dated November 30, 1962, or 54-32, Kits J and K, dated December 28, 1962, or later FAA-approved revisions, whichever is applicable to the serial number of the airplanes listed therein. After rework in accordance with this AD, the repetitive inspection requirement may be discontinued. \n\n\t\t(3)\tInspect at YOP Station 252 to YOP Station 262 the aft inboard flex panel on the outboard pylons on airplanes with Serial Numbers 45252-45263, 45274-45276, 45278-45283, 45289, 45291-45297, 45376-45379, 45384-45387, 45391, 45392, 45416, 45418, 45419, 45422-45427, 45429, 45442-45445, 45567, 45569, 45588-45595, 45598-45600, 45602, 45603. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during the inspection or any reinspection must be repaired in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region, before further flight or replaced before further flight in accordance with Douglas Service Bulletins Nos. 54-10 Kit F, dated August 17, 1960, 54-14, Kit B, dated November 3, 1960, or later FAA approved revisions, whichever is applicable to the serial number of the airplane listed therein. After repair or replacement in accordance with this AD, the repetitive inspection requirement may be discontinued. \n\n\t\t(4)\tInspect the inboard side of the inboard pylon in the area of the D-duct spar angle and the pylon side skin on airplanes with Serial Numbers 45253-45272, 45274-45283, 45289, 45291-45300, 45304-45306, 45376-45382, 45384-45393, 45418, 45419, 45421-45427, 45429-45431, 45433-45437, 45442-45445, 45565-45567, 45569, 45570, 45588-45597, 45602, 45603, 45605,45606, 45609-45612, 45626, 45627, 45638. Refer to Service Sketch No. 756, Douglas Service Engineering Letter C1-78-2106/DBA, October 18, 1966, or later revisions for location of cracks. If no cracks are found during this inspection, repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. Parts found cracked during this inspection or any reinspection, must be replaced before further flight with uncracked parts or parts reworked in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended as a redesign to prevent further cracking of these areas, and approved as such, the repetitive inspections required may be discontinued. \n\n\t(b)\tOn all airplanes having Serial Numbers 45252-45272, 45274-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45427, 45429-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45614, 45616-45630, 45632-45638, 45640-45651, 45653, 45655-45663, 45665-45673, 45676, 45684, conduct an inspection for cracks on all upper inboard and outboard spar caps of the outboard pylons in the area between Sta. YOP 214 and Sta. YOP 255, and repair or replace if necessary as follows: \n\n\t\t(1)\tUnless already accomplished within the last 475 hours' time in service within the next 25 hours' time in service after the effective date of this AD, visually inspect the outboard pylons that have accumulated a total of 8,000 hours' time in service as of the effective date of this AD; or \n\n\t\t(2)\tFor outboard pylons that have not accumulated a total of 8,000 hours' time in service as of the effective date of this AD, visually inspect them within 25 hours' time in service after 8,000 hours' time in service is reached. \n\n\t\t(3)\tIf no cracks are found during the inspections required by paragraphs (b)(1) or (b)(2), the inspections required therein must be repeated at intervals not to exceed 500 hours' time in service from the date of the last inspection. If cracks are found during this inspection or any reinspection before further flight, replace or repair in accordance with Douglas Service Bulletins Nos. 54-33, dated March 15, 1963, 54-35, dated April 9, 1965, or later FAA-approved revisions, whichever is applicable to the affected airplane. Upon replacement or repair as provided for in this AD, the 500 hours' repetitive inspection may be discontinued. \n\n\t\t(4)\tThe inspections required in Paragraph (b)(1) and (b)(2) do not apply to airplanes modified in accordance with McDonnell Douglas Service Bulletin Nos. 54-33, dated March 15, 1963; 54-35, dated April 9, 1965; or later FAA-approved revisions. Each service bulletin contains the airplane serial numbers to which it is applicable. \n\n\tHowever, all airplanes modified in accordance with S.B. Nos. 54-33 or 54-35 must be reinspected per Paragraph (b) above within the next 1500 hours' time in service after the effective date of this AD amendment, unless already accomplished within the last 1500 hours' time in service, and thereafter at intervals not to exceed 3000 hours' time in service. Upon installation of an improved spar kit per McDonnell Douglas S.B. 54-57, Revision 1, dated December 9, 1969, or later FAA approved revision, these repetitive inspections may be discontinued. \n\n\t(c)\tOn all airplanes with Serial Number 45252-45272, 45274-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45427, 45429-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45614, 45616-45630, 45632-45638, 45640-45651, 45653, 45655-45694, 45750, 45752-45760, 45762-45769, 45800, 45809, 45814-45821, 45824, 45850-45862, 45877-45882, 45886, 45916, conduct an inspection on the cant bulkhead fittings in all outboard pylons, and on the cant bulkhead fittings in all inboard pylons for the referenced serial numbers, with the addition of Serial Number 45883, in the area shown on Service Sketch No. 669, Douglas Engineering Service Letter C1-78-2016/DBA, October 18, 1966, or later FAA-approved revision, and rework or replace, if necessary, as follows - \n\n\t\t(1)\tUnless already accomplished within the last 2,000 hours' time in service, within the next 1,000 hours' time in service after the effective date of this AD, visually inspect the inboard and outboard pylons that have accumulated 10,000 or more hours' time in service as of the effective date of this AD; or \n\n\t\t(2)\tFor those inboard and outboard pylons referenced in paragraph (c) that have not accumulated a total of 10,000 hours' time in service as of the effective date of this AD, visually inspect them within 1,000 hours' time in service after 10,000 hours' time in service is reached. \n\n\t\t(3)\tIf no cracks are found during the inspection required by paragraphs (c)(1) and (2), repeat the inspection at intervals not to exceed 3,000 hours' time in service from the date of the last inspection. If cracks are found during inspection or reinspectionbefore further flight, replace the defective parts with uncracked parts, or rework in a manner approved by the Chief, Aircraft Engineering Division, FAA Western Region. When a rework is accomplished, or a replacement part is installed that is specifically intended to prevent further cracking of the structure and is approved as such, the repetitive inspection requirement may be discontinued. \n\n\t(d)\tUpon request of the 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 the operator. \n\n\t(Douglas Service Engineering Letter C1-78-2016/DBA dated October 18, 1966, or later revisions, covers this subject.) \n\n\tAmendment 39-347 effective February 4, 1967. \n\n\tThis Amendment 39-1160 becomes effectiveApril 4, 1971.
82-04-02: 82-04-02 MCDONNELL DOUGLAS: Amendment 39-4317. Applies to all McDonnell Douglas Model DC-9 and C-9 series airplanes, certificated in all categories with rudder pedal arm P/N 3616012 installed with more than 13,500 hours time in service. (Note: Time in service on the rudder pedal arm may be used, if the operator has records to substantiate it.) Compliance required as prescribed herein. To detect fatigue cracking and possible structural failure of the rudder pedal arms, P/N 3616012, accomplish the following, unless already accomplished: \n\n\tA.\tWithin the next 2,000 landings or six months after the effective date of this AD, whichever occurs first, perform ultrasonic and dye penetrant inspections on rudder pedal arm assemblies, P/N 3616012, as outlined in Service Sketch 3251 and Accomplishment Instructions of McDonnell Douglas DC-9 Service Bulletin 27-209 dated May 29, 1981, or later revisions approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region. \n\n\tB.\tIf no cracks are found, replace the rudder pedal arms with new P/N 3953505 aluminum rudder pedal arm assemblies or retain the 3616012 parts and repeat ultrasonic and dye penetrant inspections at intervals not to exceed 4,000 landings or one year, whichever occurs first. Replacement with aluminum rudder pedal arm assemblies constitutes terminating action for this AD. \n\n\tC.\tIf cracks are found, prior to further flight replace the rudder pedal arms with: \n\n\t\t1.\tNew P/N 3953505 aluminum rudder pedal arm assemblies and thereby terminate the repetitive inspection requirements of this AD, or \n\n\t\t2.\tReplace with new P/N 3616012 magnesium rudder pedal arm assemblies, and repeat inspections specified in paragraph B above. \n\n\tD.\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\tE.\tFor the purposes of complying with this AD, subject to acceptance by the assigned FAA Maintenance Inspector, the number of landings may be determined by dividing each airplane's number of hours time in service by the operator's fleet average time from takeoff to landing. \n\n\tF.\tUpon the request of an operator, an FAA Maintenance Inspector, subject to prior approval by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region, may adjust the inspection times specified in this AD to permit compliance at an established inspection period of that operator if the request contains substantiating data to justify the change for that operator. \n\n\tG.\tAlternative means of compliance with this AD which provide an equivalent level of safety may be used when approved by the Chief, Los Angeles Area Aircraft Certification Office, FAA Northwest Mountain Region. \n\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).All persons affected by this proposal 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 the FAA Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington 98108, or 4344 Donald Douglas Drive, Long Beach, California 90808. \n\n\tThis airworthiness directive becomes effective March 21, 1982.
80-09-05: 80-09-05 BOEING: Amendment 39-3760. Applies to all 737-100 and -200 series airplanes certificated at takeoff weights in excess of 97,800 pounds and containing the horizontal stabilizer trim actuator P/N 10-61326-4 or P/N 10-61326-5. \n\n\tNOTE: The 737 airplanes from line 482 and on were delivered with horizontal stabilizer trim actuator P/N 10-61326-6. This AD will apply to those 737 airplanes line number 482 and on if P/N 10-61326-4 or -5 have been exchanged for P/N 10-61326-6.\n \n\tCompliance is required as follows: To assure sufficient horizontal stabilizer trim capability, accomplish either A, B or C below within the next six (6) months after the effective date of this AD, unless already accomplished. \n\n\tA.\tFor horizontal stabilizer trim actuators having the P/N 10-61326-4 or P/N 10-61326- 5, test the actuator stall torque in accordance with Boeing Service Bulletin No. 737-27-1101, dated February 1, 1980. The actuators found to have less than 350 inch pounds of torque must be replaced with a serviceable actuator P/N 10-61326-4, -5, or -6. Thereafter, for the actuators P/N 10-61326-4 and -5, conduct repetitive torque test per the "Actual Stall Torque/Maximum Test Interval" chart, Figure 2, of that service bulletin at intervals not to exceed the maximum test intervals (hours) indicated by the curve on the chart. \n\n\tB.\tReplace the horizontal stabilizer trim actuator, Boeing P/N 10-61326-4 or -5 with stabilizer trim actuator, Boeing P/N 10-63126-6 in accordance with Boeing Service Bulletin No. 737-27-1101, dated February 1, 1980. Replacing the stabilizer trim actuator, Boeing P/N 10- 61326-4 or -5, with Boeing P/N 10-61326-6 is the terminating requirement for this AD. \n\n\tC.\tPerform an equivalent inspection and/or installation approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereofpursuant 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 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 supersedes AD 70-25-10.\n \n\tThis amendment becomes effective May 6, 1980.