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72-10-01: 72-10-01 HAWKER SIDDELEY AVIATION: Amendment 39-1441. Applies to Hawker Siddeley de Havilland Model DH-104 "Dove" airplanes. Compliance is required as indicated. To detect cracks in the upper arm of the rudder control pedal reversal lever, P/N 4CF.767A, accomplish the following: (a) Within the next 200 hours' time in service after the effective date of this AD, unless already accomplished within the last 2,200 hours' time in service, and thereafter at intervals not to exceed 2,400 hours' time in service from the last inspection, inspect the upper arm of the rudder control pedal reversal lever, P/N 4CF.767A, for cracks, using a dye penetrant method, in accordance with Hawker Siddeley Technical News Sheet Series: CT(104) No. 219, Issue 1, dated October 19, 1970, or later ARB-approved issue or FAA-approved equivalent. If no cracks are found, visually inspect the upper radiused surface of the upper arm of the rudder control reversal lever for marks or other evidence of mechanical damage. (b) If a crack is found during an inspection required by paragraph (a), before further flight replace the rudder control pedal reversal lever, P/N 4CF.767A, with a new part of the same part number and continue to inspect in accordance with paragraph (a). (c) If marks or other evidence of mechanical damage are found during an inspection required by paragraph (a), before further flight smoothly blend out the marks and renew the protective treatment, using selenious acid treatment, or an FAA-approved equivalent, and repaint the part. After refitting the part, continue to inspect in accordance with paragraph (a). This amendment is effective May 31, 1972.
2003-24-01: This amendment adopts a new airworthiness directive (AD) for the specified model helicopters modified with a Helicopter Technology Company, LLC, Supplemental Type Certificate (STC) No. SR09172RC, SR09074RC, or SR09184RC. This action requires recording on the component history card or equivalent record the number of torque events (TEs) on each main rotor blade (blade). When a blade accumulates 13,720 TEs and 750 hours time-in-service (TIS), the AD requires inspecting both surfaces of the blade for a crack at specified intervals. If a crack is found, the AD also requires replacing the blade with an airworthy blade. Also, the AD establishes life limits for certain part- numbered blades. This proposal is prompted by several reports, including a recent report dated July 24, 2003, of blade cracks due to a high number of TEs per hour. The actions specified in this AD are intended to prevent fatigue cracking of the blade, blade failure, and subsequent loss of control of the helicopter.
70-14-07: 70-14-07 TELEDYNE CONTINENTAL: Amendment 39-1028 as amended by Amendment 39-1092 is further amended by Amendment 39-2014. Applies to Teledyne Continental Models IO-360-A, -C, -D; IO-520-A, -B, -C, -D, -E, -F, -J, -K; IO-470-C, -D, -E, -F, -H, -K, -L, -M, -N, -S, -J, -U, -V, -VO; TSIO-470-B, -C and -D engines. NOTE: Compliance with AD 70-14-07 on the Model IO-470-J, -U, -V, and -VO engines added to the applicability statement commences on the effective date of amendment 39-1092. Compliance: Required as indicated unless already accomplished. To prevent loss of the fuel injection pump adjustable bypass needle with subsequent power failure accomplish the following or any equivalent procedure approved by Chief, Engineering and Manufacturing Branch, FAA, Central Region, Kansas City, Missouri: A) Within 25 hours' time in service after the effective date of this AD, visually inspect the fuel injection pump below the fuel inlet elbow for the presence of a hexagonal head brass plug containing a stainless steel adjusting needle with a screwdriver slot. If the pump does not have this feature, or if the adjusting needle is retained in the brass plug by an internal spring circular clip, no further action is required. If it does, and the slotted head of the needle is flush or below the face of the plug, secure the needle in its present position by cleaning the plug and needle with gasoline or carbon tetrachloride and applying LePages Epoxy or Loctite #2508 epoxy cement or equivalent over the slotted head of the needle and face of the plug. If the slotted head of the needle extends outward beyond the face of the plug, accomplish Paragraph B before further flight. B) At next engine or fuel injection pump overhaul, fuel injection pump adjustment or as indicated in Paragraph A, replace the existing fuel injection bypass needle with P/N 637766 or 637767 needle as appropriate in accordance with the following: 1. Before removing the needle start the engine and adjust engine to obtain full throttle and maximum RPM, then record fuel flow or pressures and RPM for future reference. Stop the engine and record the number of turns required to bottom the needle. 2. Replace the present needle with P/N 637766 needle having 8-32 threads or P/N 637767 needle having 10-32 threads as applicable. Use a new "O" ring, P/N AN123957, with the new needle. 3. Bottom the new needle and back out to the previous setting, start the engine, and stabilize power at full throttle and maximum RPM, then adjust the needle to obtain the previously recorded full throttle, maximum RPM, fuel pressure or flow. 4. Safety wire the needle through the drilled hole in the needle shank and the unused hole in the brass plug, or around the pressure relief valve adjusting screw in the cover at the back of the pump. Exercise care to avoid disturbing the needle setting. Teledyne Continental Motors Service Bulletin M70-10, Revision 1, dated June 25, 1970, or later FAA- approved revision, refers to this subject. NOTE: The presently FAA-approved manufacturer's fuel pump adjustment procedures contained in his Overhaul and Parts Catalog for Fuel Injection Systems, Form X-30091 may be substituted for the adjustment instructions contained in Paragraphs B.1., 2. and 3. Amendment 39-1028 became effective July 16, 1970. Amendment 39-1092 became effective October 13, 1970. This Amendment 39-2014 becomes effective November 22, 1974.
92-06-17: 92-06-17 ROBINSON HELICOPTER COMPANY (RHC): Amendment 39-8293. Docket No. 92-ASW-14. Final Rule of a Priority Letter AD. Applicability: All Model R22 series helicopters, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent failure of NAS1304-16 AF bolts, resulting in loss of helicopter control, accomplish the following: (a) Within the next 10 hours' time in service or before March 30, 1992, whichever comes first, visually inspect the helicopters specified in (1), at the inspection areas or locations specified in (2), to determine the identification of the NAS1304-16 bolts. (1) The helicopters affected are-- (i) R22 Serial Numbers (S/N) 1880 through 2060 and S/N 2073; (ii) All R22 helicopters regardless of S/N overhauled or repaired at Robinson Helicopter Company between July 9, 1991, and March 1, 1992; and (iii) All R22 helicopters regardless of S/N for which maintenance was performedafter July 9, 1991, in the inspection locations specified in paragraph (a)(2). (2) The inspection locations are-- (i) The tail rotor blade control assembly at the aft end of the tail cone, including tail rotor controls connecting the rotor blade pitch link to the rotor pitch control cross head (slider) arms, and the rotor pitch link to the rotor blade attachment; (ii) The lower aft corners of the cabin, both left-hand and right-hand sides, where the attachment joins the cabin to the welded frame assembly; and (iii) The region above the swashplate attaching the counter weights (balance weights) to the swashplate assembly (also described as the main rotor balance weights attachment to the Chord Arm Yoke). NOTE: Further details of the installations are contained in Robinson Model R22 Illustrated Parts Catalog (IPC). (b) Before further flight, remove those NAS1304-16 bolts bearing the identification letters "AF" on the head of the bolt and replace with serviceable NAS1304-16 bolts with head marking other than "AF" or with NAS6604-16 bolts. (c) After the effective date of this AD, NAS1304-16 bolts identified with the letters "AF" shall no longer be installed as a replacement part in any application on these helicopters. (d) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office, 3229 E. Spring Street, Long Beach, California 90806-2425. The request shall be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles Aircraft Certification Office. (e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the helicopter to a location where the requirements of this AD can be accomplished. (f) Copies of the applicable service information may be obtained from Robinson Helicopter Company, 24747 Crenshaw Blvd, Torrance, CA 90505. This information may be examined at the FAA, Office of the Assistant Chief Counsel, Rules Docket, 4400 Blue Mound Road, Bldg. 3B, Room 158, Fort Worth, Texas. (g) This amendment becomes effective July 24, 1992, to all persons except those persons to whom it was made immediately effective by Priority Letter AD 92-06-17, issued March 17, 1992, which contained the requirements of this amendment.
2016-12-01: We are adopting a new airworthiness directive (AD) for PILATUS AIRCRAFT LTD. Models PC-12, PC-12/45, PC-12/47, and PC-12/47E airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as incorrect installation instructions of the torlon plates in the airplane maintenance manual resulting in the incorrect installation of the torlon plates in the forward wing-to-fuselage attachment. We are issuing this AD to require actions to address the unsafe condition on these products.
90-15-09: 90-15-09 AIRBUS INDUSTRIE: Amendment 39-6658. Docket No. 90-NM-20-AD. Applicability: Models A300, A310, and A300-600 series airplanes, equipped with BF Goodrich, Aircraft Evacuation Systems, Slide/Raft, Part Number (P/N) 7A1300-() or 7A1359-(), certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent improper slide/raft deployment, accomplish the following: A. Within 6 months after the effective date of this AD, accomplish the modification of the evacuation slide/rafts in accordance with Section 2, Accomplishment Instructions, of BF Goodrich Service Bulletin 7A13000/7A1359-25-227, Revision 1, dated January 5, 1990. B. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. NOTE: 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. C. Special flight permits may be issued in accordance with FAR Sections 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to BF Goodrich, Aircraft Evacuation Systems, 3414 South 5th Street, Phoenix, Arizona 85040. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California. This amendment (39-6658, AD 90-15-09) becomes effective August 20, 1990.
77-13-22: 77-13-22\tTELEDYNE CONTINENTAL MOTORS: Amendment 39-2947 as amended by Amendment 39- 3188. Applies to the following engine models: \n\n\tIO-520-A, -B, -BA, -C, -D, -E, -F, -J, -K, -L, and -M which do not incorporate the identification markings in Table 1. \n\tTSIO-520-B, -C, -D, -E, -G, -H, -J, -K, -L, and -N which do not incorporate the identification markings in Table 1. \n\tGTSIO-520-C, -D, -F, and -H which do not incorporate the identification markings in Table 1. \n\tNOTE: This A.D. does not apply to engines bearing the following serial numbers: \n\n\t\t\tNEW\t\t\tTCM REBUILT \n\tIO-520-A\t550024 & up\t\t112354-R & up \n\t-B, -BA\t\t562678 "\t\t122539-R " \n\t-C\t\t561476 "\t\t172903-R " \n\t-D\t\t563853 "\t\t174412-R " \n\t-E\t\t556319 "\t\t215510-R " \n\t-F\t\t564398 "\t\t195707-R " \n\t-J\t\t558005 "\t\t216505-R " \n\t-K\t\t557306 "\t\t224007-R " \n\t-L\t\t554919 "\t\t220537-R " \n\t-M\t\t565039 "\t\t227309-R " \n\n\tTSIO-520-B\t500457 & up\t\t176235-R & up \n\t-C\t\t509518 "\t\t178143-R " \n\t-D\t\t505004 "\t\t180043-R " \n\t-E\t\t510125 "\t\t182692-R " \n\t-G\t\t507057 "\t\t216007-R " \n\t-H\t\t506853 "\t\t217031-R " \n\t-J\t\t503582 "\t\t218621-R " \n\t-K\t\t504311 "\t\t224529-R " \n\t-L\t\t508311 "\t\t227611-R " \n\t-N\t\t509516 "\t\tNo Rebuilt \n\n\tGTSIO-520-C\t602221 & up\t\t155418-R & up \n\t-D\t\t601051 "\t\t219249-R " \n\t-F\t\t603112 "\t\t224227-R " \n\t-H\t\t600915 "\t\t218260-R " \n\n\nTABLE 1 \n\n\nThe improved design crankcase may be identified by the following table. This A.D. does not apply to engines so identified. \n\n\n\tTo prevent the possibility of undetected crankcase cracks progressing into a crankcase failure, accomplish the following: \n\t(a)\tWithin the next 50 hours time in service after the effective date of this A.D., visually inspect the engine crankcase for evidence of cracks. Particular attention during inspection should be given to the upper rear stud area of the number 2 and number 6 cylinders, the magneto pads and in the sections between the cylinder mounting bases. \n\t\t(1)\tIf no cracks are found, reinspecteach 100 hours time in service thereafter, unless replaced by a crankcase described in Table 1. \n\t\t(2)\tIf cracks are found, proceed to paragraph (c). \n\t(b)\tCritical (shaded) and non-critical (non-shaded) areas are illustrated in Figures 1, 2, and 3 below. \n\n\n\n\t(c)\tIf any crack is observed in the non-critical (non-shaded) area that exceeds two (2) inches in total length or is leaking oil, regardless of length, or if any cracks are observed in the critical (shaded) area, the engine crankcase must be replaced prior to further flight with a serviceable crankcase, except that the engine may be operated on an aircraft which is flown in accordance with FAR's 21.197 and 21.199 to a base where replacement may be accomplished. \n\t\t(1)\tIf a crack of two (2) inches or less in length is observed in any of the non-critical (non-shaded) areas, the crack extremity may be stop-drilled using a 3/32" drill bit (.0938"). Prior to drilling, identify the end of the crack using the dye penetrant method and locate the hole one-eighth (1/8) inch from the crack extremity. Do not stop-drill the end of any crack which would require the hole to intersect with a cylinder hold-down stud, require the hole to be placed in a heavy fillet area such as that which immediately surrounds the cylinder mount pad, or which would require the hole to be placed in an upper backbone bolt boss. (Cracks in these areas must be scribe marked at their extremities and reinspected at the intervals stated in (c)(2) to check for progression.) Stop drill only those cracks which terminate in the crankcase wall sections. Coat the bit with beeswax or modeling clay to retain chips, and operate the drill at the slowest possible speed. Additionally, pressurization of the crankcase with vacuum cleaner outlet air or regulated shop air (not to exceed 5 psi) may be used to help prevent drilling chips from entering the crankcase. Thoroughly clean the immediate area surrounding the newly drilled hole. Fill the hole with a small amount of freshly mixed epoxy adhesive such as Locktite brand, "Fast Cure" Epoxy Adhesive, Locktite P/N 44581 or 3M, "Scotch Weld" brand Structural Adhesive, 3M P/N 1838 B/A, or equivalent, being careful not to allow any epoxy to enter the interior of the crankcase. See Figures 4, 5, and 6 for additional details. \n\n\n\n\n\n\t(2)\tReinspect for crack progression and additional cracks within the next fifty (50) hours time in service. If no crack progression has taken place or no additional cracks are found, continue to reinspect at intervals not to exceed 100 hours time in service, until replaced with a crankcase described in Table 1. If further crack progression is noted, or additional cracks are found, repeat Paragraph (c). \n\t(d)\tIn addition to the above inspections, accomplish the following on the Models IO-520-B, -BA, -C, and -M, and TSIO-520-B, -D, -E, -J, -K, -L, and -N engines. \n\t\t(1)\tWithin the next fifty (50) hours time in service after the effective date of thisA.D., inspect the crankcase immediately above the backbone bolts to determine if the crankcase is of the design which incorporates raised backbone bolt bosses (bumps). \n\t\t\t(i)\tIf the crankcase incorporates the raised backbone bolt bosses, no further action is required. \n\t\t\t(ii)\tIf the crankcase does not incorporate the raised backbone bolt bosses, install Teledyne Continental Motors backbone bolt kit Number EQ 6541 in lieu of the current backbone bolts in accordance with steps (iii) and (iv) below. \n\t\t\t(iii)\tRemove backbone bolts three (3) through ten (10) inclusive (counting from front to rear). CAUTION: Remove no more than two (2) bolts at a time and install replacement bolts prior to removing the next two (2) bolts, etc. \n\t\t\t(iv)\tInstall the longer bolts and thicker washers as shown in the instruction drawing included in kit Part Number EQ 6541. Torque to 160 to 180 in.-lbs. \n\tNOTE: Cracks found in the upper backbone boss area are acceptable; however, if cracks extend downinto the crankcase wall, follow procedure outlined in paragraph (c). \n\t(e)\tMake a log book entry indicating compliance with applicable portions of this A.D. Include backbone kit installation, if applicable, and engine time in service. \n\t(f)\tUpon request of the operator, an FAA maintenance inspector may adjust the repetitive inspection interval specified in paragraph (a)(1) of this A.D. up to a maximum time between inspections of 125 hours to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. The compliance time for inspection of engines with existing crankcase cracks required by paragraph (c)(2) of this A.D. may not be adjusted. \n\tAlternate methods of compliance must be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, Southern Region. \n\tTCM Service Bulletin M77-14 pertains to this subject. \n\tAmendment 39-2947 became effective July22, 1977. \n\tThis Amendment 39-3188 becomes effective immediately upon publication in the Federal Register.
65-14-07: 65-14-07 VICKERS: Amdt. 39-93 Part 39 Federal Register June 26, 1965. Applies to Viscount Model 810 Series Aircraft. Compliance required as indicated. To prevent further failures of entrance doors during pressurized flight: (a) Within the next 250 landings after the effective date of this AD, accomplish the following: (1) Visually inspect the door locking mechanism of each entrance door in accordance with paragraph B.1 of Vickers-Armstrongs Preliminary Technical Leaflet No. 112, Issue 2. Repair any found defective before further flight. (2) Inspect for position accuracy the check markings applied to each claw and fuselage aperture. Where no such markings exist they must be applied in accordance with Figure 2 of PTL No. 112, Issue 2, or an FAA-approved equivalent. (b) Within the next 700 landings after the effective date of this AD, on aircraft fitted with airsteps or other installed equipment which obscures any claw from view, incorporate Vickers-Armstrongs Modification G.1964 (remote position visual indicator) or an FAA-approved equivalent. (c) Within the next 1,000 landings after the effective date of this AD, and thereafter at intervals not to exceed 1,000 landings from the last inspection, inspect the door locking mechanism of each entrance door for condition and correct operation, in accordance with paragraph C of P.T.L. No. 112, Issue 2. Repair any found defective before further flight. (d) For the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing each aircraft's hours time in service by the operator's fleet average time from takeoff to landing for the aircraft type. (Vickers-Armstrongs Preliminary Technical Leaflet No. 112, Issue 2, dated August 6, 1964, (800-810 Series) and Modification G.1964 cover this subject.) This directive effective July 26, 1965.
92-17-06: 92-17-06 BRITISH AEROSPACE: Amendment 39-8335. Docket No. 92-NM-63-AD. Applicability: Model BAC 1-11-200 and -400 airplanes, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the fuselage pressure vessel, accomplish the following: (a) For airplanes operated up to a maximum cabin pressure differential of 7.5 pounds per square inch, accomplish the following in accordance with British Aerospace Alert Service Bulletin 53-A-PM5989, Issue No. 1, dated October 3, 1991: (1) For airplanes not having modification PM51 installed: Prior to the accumulation of 20,000 landings, or within 1,000 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual, dye penetrant, or eddy current inspection to detect cracks in the top and bottom corners of the passenger and service door apertures, in accordance withthe service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,600 landings, in accordance with the service bulletin. (ii) If the immediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 3,200 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 5,000 landings, in accordance with the service bulletin. (2) For airplanes having modification PM51 installed: Prior to the accumulation of 30,000 landings, or within 1,200 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual inspection, dye penetrant, or eddy current inspection to detect cracks in the top and bottom corners of the passenger and service door apertures, in accordance with the service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,600 landings, in accordance with the service bulletin. (ii) If the immediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 3,200 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 5,000 landings, in accordance with the service bulletin. (3) For airplanes repaired in accordance with Structural Repair Manual Chapter 53-02-0, Figure 74: Prior to the accumulation of 20,000 landings (for airplanes not having modification PM51 installed), or prior to the accumulation of 30,000 landings (for airplanes having modification PM51 installed), from the date of installation of the repair; or within 1,000 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual inspection, dye penetrant, or eddy current inspection to detect cracks of the fuselage skin repair plates at the passenger and service door apertures, in accordance with the service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,600 landings, in accordance with the service bulletin. (ii) If the immediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 3,200 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 5,000 landings, in accordance with the service bulletin. (b) For airplanes operated at a cabin pressure differential in excess of 7.5 pounds per square inch, but not exceeding 8.2 pounds per square inch, accomplish the following in accordance with British Aerospace Alert Service Bulletin 53-A-PM5989, Issue No. 1, dated October 3, 1991: (1) For airplanes not having modification PM51 installed: Prior to the accumulation of 14,000 landings, or within 1,000 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual inspection, dye penetrant, or eddy current inspection to detect cracks in the top and bottom corners of the passenger and service door apertures, in accordance with the service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,100 landings, in accordance with the service bulletin. (ii) If theimmediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 2,250 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 3,500 landings, in accordance with the service bulletin. (2) For airplanes having modification PM51 installed: Prior to the accumulation of 20,000 landings, or within 1,000 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual, dye penetrant, or eddy current inspection to detect cracks in the top and bottom corners of the passenger and service door apertures, in accordance with the service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,100 landings, in accordance with the service bulletin. (ii) If the immediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 2,250 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 3,500 landings, in accordance with the service bulletin. (3) For airplanes repaired in accordance with Structural Repair Manual Chapter 53-02-0, Figure 74: Prior to the accumulation of 10,000 landings (for airplanes not having modification PM51 installed), or prior to the accumulation of 15,000 landings (for airplanes having modification PM51 installed), from the date of installation of the repair; or within 500 landings after the effective date of this AD, whichever occurs later; and thereafter at intervals specified below; perform a close visual, dye penetrant, or eddy current inspection to detect cracks of the fuselage skin repair plates at the passenger and service door apertures, in accordance with the service bulletin. (i) If the immediately preceding inspection was performed using a close visual inspection technique, the next inspection must be performed within 1,100 landings, in accordance with the service bulletin. (ii) If the immediately preceding inspection was performed using a dye penetrant technique, the next inspection must be performed within 2,250 landings, in accordance with the service bulletin. (iii) If the immediately preceding inspection was performed using an eddy current technique, the next inspection must be performed within 3,500 landings, in accordance with the service bulletin. (c) If cracks are found as a result of any inspection required by paragraphs (a) or (b) of this AD, prior to further flight, repair any cracks found; and inspect the door surround structure for associated damage, and, prior to further flight, repair any damage found; in accordance with British Aerospace Alert Service Bulletin 53-A-PM5989, Issue No. 1, dated October 3, 1991. (d) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch. (e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (f) The inspections and repair shall be done in accordance with British Aerospace Alert Service Bulletin53-A-PM5989, Issue No. 1, dated October 3, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from British Aerospace, PLC, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, DC 20041-0414. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street NW., 7th Floor, Suite 700, Washington, DC. (g) This amendment becomes effective on September 29, 1992.
2016-12-03: We are superseding Airworthiness Directive (AD) 2011-17-10, for all Fokker Services B.V. Model F.28 Mark 1000, 2000, 3000, and 4000 airplanes. AD 2011-17-10 required inspecting for a by-pass wire between the housing of each in-tank fuel quantity indication (FQI) cable plug and the cable shield, and corrective actions if necessary. AD 2011-17- 10 also required revising the airplane maintenance program. This new AD removes certain airplanes from the applicability. This new AD applies only to Model F.28 Mark 1000 airplanes and also requires revising the airplane maintenance or inspection program by incorporating the instructions in revised service information. This AD was prompted by the issuance of revised service information to update the critical design configuration control limitations (CDCCLs) that address potential ignition sources inside fuel tanks. We are issuing this AD to prevent potential ignition sources inside the fuel tanks, which, in combination with flammable fuelvapors, could result in fuel tank explosions and consequent loss of the airplane.