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2022-03-19: The FAA is adopting a new airworthiness directive (AD) for all General Electric Company (GE) Passport 20-17BB1A, Passport 20-18BB1A, and Passport 20-19BB1A model turbofan engines. This AD was prompted by a report of a manufacturing quality escape that requires a reduction to the life limit of certain high-pressure turbine (HPT) rotor stage 1 disks. This AD requires revising the airworthiness limitations section (ALS) of the existing maintenance manual and the operator's existing approved continuous airworthiness maintenance program (CAMP) to incorporate a reduced life limit for certain HPT rotor stage 1 disks. The FAA is issuing this AD to address the unsafe condition on these products.
91-17-03: 91-17-03 FOKKER: Amendment 39-8001. Docket No. 91-NM-68-AD. Applicability: Model F-27 series airplanes; serial numbers 10102, 10105 through 10684, 10686, 10687, and 10689 through 10692; certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent collapse of the main landing gear (MLG), accomplish the following: A. Within 180 days after the effective date of this AD, or prior to the accumulation of 500 landings after the effective date of this AD, whichever occurs first, perform a high frequency eddy current inspection of both sides of the actuating ram attachment lug in accordance with Part 1 of the Accomplishment Instructions of Fokker Service Bulletin F27/54-47, dated November 30, 1990. B. If cracks are found, prior to further flight, replace the MLG drag strut attachment fitting in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/54-47, dated November 30, 1990.C. An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. E. The inspection and replacement requirement shall be done in accordance with Fokker Service Bulletin F27/54-47, dated November 30, 1990. 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 Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401, Washington, D.C. This amendment (39-8001, AD 91-17-03) becomes effective on September 20, 1991.
78-07-05: 78-07-05 MCDONNELL DOUGLAS: Amendment 39-3176 as amended by amendment 39-3467. Applies to Model DC-9 and C-9 series airplanes, certificated in all categories, which correspond to the factory serial numbers as listed below. \n\n\n45695 thru 45749\n47389 thru 47397 \n45770 thru 45799\n47399 thru 47482 \n45825 thru 45847\n47484 thru 47514 \n45863 thru 45876\n47516 thru 47557 \n47000 thru 47386\n47559 thru 47764 \n\n 47769 \n\n 47771 \n\t\t\t\n\tCompliance required as indicated, unless already accomplished: \n\n\tTo prevent possible failure of the over-wheelwell fuselage frame(s) lower fittings manufactured with 7075-T6 aluminum alloy material, accomplish the following: \n\n\ta.\tFor airplanes with 12,000 or more hours time in service on the effective date of this AD, which have not had new frames installed, within the next 3400 hours time in service or 12 calendar months, whichever occurs earlier, and thereafter at intervals not to exceed 8000 hours time in service or 37 calendar months, whichever occurs earlier, comply with the program of inspections established in Paragraph (c), below. \n\n\tb.\tFor airplanes with less than 12,000 hours time in service as of the effective date of this AD, comply with Paragraph (c), below, prior to the accumulation of 15,400 hours time in service or 50 calendar months, whichever occurs earlier, and thereafter at intervals not to exceed 8000 hours time in service or 37 calendar months, whichever occurs earlier. \n\n\tc.\tVisually inspect for evidence of cracking, using optical inspection aids with a minimum of 2X magnification, the fore and aft webs, flanges and radii, in the areas of the frame lower fitting, as shown in Sketch No. 2765 of McDonnell Douglas Service Bulletin 53-131 dated February 24, 1978 or later FAA approved revision, (hereinafter referred to as SB 53-131). \n\n\td.\tIf no cracks are found: \n\n\t\t(1)\tClean and spray the pocket and adjacent areas of the frame fitting with corrosion inhibiting compound, per DAC SRM, Chapter 51-10-3, and perform repetitive inspections and corrosion inhibiting treatment at intervals not to exceed 8,000 hours time in service or 37 calendar months, whichever occurs earlier; or \n\n\t\t(2)\tRepair per Option 1 of SB 53-131 and conduct repetitive inspection at intervals not to exceed 16,000 hours time in service or 54 calendar months whichever occurs earlier; or \n\n\t\t(3)\tReplace with new part(s) per Paragraphs f.2 or f.3 and comply with the applicable program of repetitive inspections and/or corrective actions per this AD. \n\n\te.\tIf cracks are found which are within the limits of Figure 1 of SB 53-131, aircraft may be continued in service provided that \n\n\t\t(1)\tRepetitive inspections are conducted at intervals not to exceed 3,400 hours time in service if an adjacent frame does not have cracks in pocket area or has been repaired per Option 1 of SB 53-131; or \t\t(2)\n\n\tRepetitive inspections are conducted at intervals not to exceed 1,700 hours time in service if adjacent frame(s) hasunrepaired flange cracks within limits as outlined on Figure 1 of SB 53-131. \n\n\tf.\tIf cracks are found, which are beyond the limits of Figure 1, but within the limits of Figure 2 of SB 53-131, before further flight. \n\n\t\t(1)\tRepair per Option 1 of SB 53-131 and conduct repetitive inspections at intervals not to exceed 8,000 hours time in service or 27 calendar months, whichever occurs earlier; or \n\n\t\t(2)\tReplace with a new part(s) of the same design made from 7075-T6 aluminum material; or \n\n\t\t(3)\tReplace with a new part(s) of the same design made from 7075-T73 aluminum alloy material. \n\n\tg.\tIf cracks are found which are beyond the limits of Figure 2 of SB 53-131, before further flight: \n\n\t\t(1)\tReplace with a new part(s) of the same design made from 7075-T6 aluminum alloy material; or \n\n\t\t(2)\tReplace with a new part(s) of the same design made from 7075-T73 aluminum alloy material. \n\n\th.\tThe requirements per this AD may be terminated for that frame(s) only, when both theright and left hand fittings, made from 7075-T73 aluminum alloy material, have been installed. \n\n\ti.\tIf new parts have been installed per f(2) or g(1) for the stations specified, the requirements of this AD may be discontinued for that part(s) only, until the new part(s) has accumulated 15,400 hours time in service or within 50 calendar months after the part(s) has been replaced, whichever occurs earlier, at which time reinstate the program of repetitive inspections and/or corrective action per this AD. \n\n\tj.\tEquivalent inspection procedures and repairs may be used when approved by the Chief, Aircraft Engineering Division, FAA, Western Region. \n\n\tk.\tSpecial flight permits may be issued in accordance with FAR's 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or maintenance required by this AD. \n\n\tl.\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 initial and 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 that operator. \n\n\tThis supersedes Amendment 39-3149 (43 FR 9587), AD 78-05-03. \n\n\tAmendment 39-3176 became effective April 6, 1978. \n\n\tThis amendment 39-3467 becomes effective May 18, 1979.
86-22-05: 86-22-05 GATES LEARJET: Amendment 39-5448. Applies to the following Gates Learjet series airplanes, models/serial number listed below, certificated in any category. MODEL SERIAL NUMBER 23 003 thru 089 (if equipped with dual flap actuators) 23 090 thru 099 24 100 thru 357 25 003 thru 373 28 001 thru 005 29 001 thru 004 35 002 thru 545, 589 thru 598 36 001 thru 053, and 055 55 001 thru 121 Compliance required as indicated, unless previously accomplished. To prevent impairment of flap operation, an asymmetric flap condition, false gear warning horn signals, or incorrect biasing of the stall warning system, due to flap sector upper mount bracket failures, accomplish the following: A. Within the next 50 hours time-in-service after the effective date of this AD, inspect the flap sector upper mounting brackets for cracks, in accordance with instructions in Gates Learjet Corporation Airplane Modification Kit (AMK) 86-4 or 55-86-2, as applicable. 1. If cracks are found in either the left-hand or right-hand flap sector upper mounting brackets (Figure I of AMK), prior to further flight, replace both brackets and install stiffeners in accordance with the applicable AMK. 2. If cracks are not found in the flap sector upper mounting brackets, replace the brackets and install stiffeners in accordance with the applicable AMK within 100 hours time-in-service after the effective date of this AD. B. If both flap sector upper mounting brackets have been previously replaced with steel brackets, compliance with this AD is not required. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. D. Alternate means of compliance with this AD, which provides an acceptable level of safety, may be used when approved by the Manager, Wichita Aircraft Certification Office, FAA, Central Region. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Gates Learjet Corporation, P.O. Box 7707, Wichita, Kansas 67277. These documents may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, Central Region, Wichita Aircraft Certification Office, 1801 Airport Road, Room 100, Wichita, Kansas. This supersedes AD 84-20-06, Amendment 39-4908. This amendment becomes effective November 10, 1986.
92-15-13: 92-15-13 BEECH: Amendment 39-8307. Docket No. 92-CE-02-AD. Supersedes AD 90-04-04; Amendment 39-6487. Applicability: 99 series airplanes (serial numbers U-1 through U-49, and serial numbers U-51 through U-164) that have 3,000 hours or more time-in-service (TIS), except those airplanes that have Beech Wing Modification Kit No. 99-4023-1S installed, certificated in any category. Compliance: Required as indicated, unless already accomplished. To prevent fatigue failure of the wing front spar lower cap and associated structure, accomplish the following: (a) For airplanes that do not have a spar reinforcing strap installed in accordance with the instructions in STC SA1178CE, accomplish the actions specified in paragraphs (a)(1) through (a)(4) of this AD using the criteria in the Beech Structural Inspection and Repair Manual (SIRM). (1) Upon the accumulation of 3,000 hours TIS on the front spar lower cap or within the next 100 hours TIS after the effective date of this AD, whichever occurs later, unless already accomplished within the last 500 hours TIS (the inspection interval established by either superseded AD 77-05-01 R3 or superseded AD 90-04-04), and thereafter at intervals not to exceed 500 hours TIS, inspect the areas of structure defined by Index Number 1 (lower forward fitting only) and Index Numbers 2 through 7 on Page 202, Section 57-15-00, of the Beech SIRM, using the visual, fluorescent penetrant, and eddy current methods as specified in the Beech SIRM. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (2) Upon the accumulation of 10,000 hours TIS on the nacelle splice plates, or within the next 100 hours TIS after the effective date of this AD, whichever occurs later, unless already accomplished (superseded AD 90-04-04), and thereafter at intervals not to exceed 1,000 hours TIS, inspect the nacelle splice plates as defined by Index Number 9 on Page 202, Section 57-15-00, of the Beech SIRM, using visual methods as specified in the Beech SIRM. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (3) Upon the accumulation of 10,000 hours TIS on the wing structure or within the next 100 hours TIS after the effective date of this AD, whichever occurs later, unless already accomplished within the last 500 hours TIS (the inspection interval established by either superseded AD 77-05-01 R3 or superseded AD 90-04-04), and thereafter at intervals not to exceed 500 hours TIS, inspect the wing structure components defined in paragraph (d) of this AD using visual and dye penetrant methods as indicated. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (4) Upon the accumulation of 10,000 hours TIS on the front spar lower cap or within the next 100 hours TIS after the effectivedate of this AD, whichever occurs later, unless already accomplished (superseded AD 90-04-04), and thereafter at intervals not to exceed 10,000 hours TIS, replace the structural components set forth on Page 203, Section 57-15-00, of the Beech SIRM, and summarized below: (i) Lower cap of the front spar, with attachment fitting, in each outer wing panel. (ii) Lower cap of the front spar, with left and right attachment fittings, in the center section. (b) For airplanes that have a spar reinforcing strap installed in accordance with Supplemental Type Certificate (STC) SA1178CE, accomplish the actions specified in paragraphs (b)(1) through (b)(5) using the Beech SIRM and Aerocon California, Inc., Engineering Order No. E.O. B-9975-2, dated November 14, 1975. Strap tension is to be adjusted in accordance with the instructions in Aerocon California Service Letter, dated May 25, 1976. (1) If the strap was installed before 1,000 hours TIS on the front spar lower cap, within the next 2,000 hours TIS after the effective date of this AD, unless previously accomplished within the last 2,000 hours TIS (the inspection interval established by either superseded AD 77-05-01 R3 or superseded AD 90-04-04), and thereafter at intervals not to exceed 2,000 hours TIS: (i) Remove and inspect the STC SA1178CE strap in accordance with the instructions in Aerocon California, Inc. Engineering Order No. E.O. B-9975-2, dated November 14, 1975. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace in accordance with the instructions in Aerocon California, Inc. Engineering Order No. E.O. B-9975-2. (ii) Inspect the following areas of structure using the visual, fluorescent penetrant, and eddy current methods as specified in the Beech SIRM. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (A) Areas defined by Index Number 1 (lower forward fitting only) and Index Numbers 2 through 7 on Page 202, Section 57-15-00, of the Beech SIRM. (B) Areas defined by paragraphs (d)(5) and (d)(8) of this AD. (iii) Reinstall the STC SA1178CE strap and adjust its tension in accordance with the instructions in Aerocon California Service Letter, dated May 25, 1976. (2) If the strap was installed at or after 1,000 hours TIS on the front spar lower cap, within the next 1,000 hours TIS after the effective date of this AD, unless previously accomplished within the last 1,000 hours TIS (the inspection interval established by either superseded AD 77-05-01 R3 or superseded AD 90-04-04), and thereafter at intervals not to exceed 1,000 hours TIS, accomplish the following: (i) Remove and inspect the STC SA1178CE strap in accordance with the instructions in Aerocon California, Inc. Engineering Order No. E.O. B-9975-2, dated November 14, 1975. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace in accordance with the instructions in Aerocon California, Inc. Engineering Order No. E.O. B-9975-2. (ii) Inspect the following areas of structure (specified in paragraphs (b)(2)(ii)(A) and (b)(2)(ii)(B) of this AD) using the visual, fluorescent penetrant, and eddy current methods as specified in the Beech SIRM. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (A) Areas defined by Index Number 1 (lower forward fitting only) and Index Numbers 2 through 7 on Page 202, Section 57-15-00, of the Beech SIRM. (B) Areas defined by paragraphs (d)(5) and (d)(8) of this AD. (iii) Reinstall the STC SA1178CE strap and adjust its tension in accordance with the instructions in Aerocon California Service Letter, dated May 25, 1976. (3) Upon the accumulation of 10,000 hours TIS on the nacelle splice plates or within the next 100 hours TIS after the effective date of this AD,whichever occurs later, unless already accomplished (superseded AD 90-04-04), and thereafter at intervals not to exceed 1,000 hours TIS, inspect the nacelle splice plates as defined by Index Number 9 on Page 202, Section 57-15-00, of the Beech SIRM, using the visual methods as specified in the Beech SIRM. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (4) Upon the accumulation of 10,000 hours TIS on the wing structure or within the next 100 hours TIS after the effective date of this AD, whichever occurs later, unless already accomplished within the last 500 hours TIS (the inspection interval established by either superseded AD 77-05-01 R3 or superseded AD 90-04-04), and thereafter at intervals not to exceed 500 hours TIS, inspect the wing structure components defined in paragraph (d) of this AD using visual and dye penetrant methods as indicated; compliance is not required with paragraphs (d)(5), (d)(8), and that portion of paragraph (d)(12) of this AD that refers to the lower spar cap and hinge. If a crack, loose fastener, or corrosion is found, prior to further flight, repair or replace as specified in the Beech SIRM. (5) Replace the structural components that are set forth on Page 203, Section 57-15-00, of the Beech SIRM (summarized in paragraphs (b)(5)(i) and (b)(5)(ii) of this AD) upon the accumulation of the front spar's allowable service life. Determine the allowable service life by subtracting the front spar lower cap hours TIS at which the strap was installed from 48,000 hours TIS. NOTE 1: For example, if the spar cap had been in service 5,000 hours TIS when the strap was installed, then the spar cap's allowable service life becomes 43,000 hours TIS (48,000 minus 5,000). (i) Lower cap of the front spar, with attachment fitting, in each outer wing panel. (ii) Lower cap of the front spar, with left and right attachment fittings, in the center section. (c) The inspection intervals established by superseded AD 77-05-01 R3 and superseded AD 90-04-04 may be substituted for the "unless already accomplished" statement in paragraphs (a)(1), (a)(3), (b)(1), (b)(2), and (b)(4) of this AD. (d) The items specified in paragraphs (d)(1) through (d)(13) of this AD define the additional structural items to be inspected as referenced by paragraphs (a)(3) and (b)(4) of this AD. (1) Inspect the lower fuselage skin at the attachment to the main spar for possible cracks or loose rivets. (2) Inspect the lower left hand (LH) and right hand (RH) nacelle skins for cracks or loose rivets. (3) Remove the aft fabric covers in the wheel wells and inspect for possible cracks in the center section skin under the top nacelle fairing. Check around the nacelle attach flange on the top side for possible loose rivets or cracks in the top skin. (4) Inspect the structure and attaching fasteners of both keel beam assemblies at ButtLine (BL) 68 inboard, BL 88 outboard, at the center section rear spar, Nacelle Station 160.50. (5) Inspect for possible cracks or loose rivets in the LH and RH dimpled skin attachment holes on the forward side of the main spar at the four countersunk screws and at all rivets between the fuselage and the nacelles. (6) Inspect for possible cracks or loose rivets along the top skin attachment to the aft spar. (7) Inspect for possible loose fasteners in the lower aft spar cap and skin. (8) Inspect for possible cracks or loose fasteners in the lower strap on the main spar at Wing Station 68.5. (9) Inspect the lower stringers running forward and aft between the main spar and the aft spar for possible cracks or loose fasteners to the lower fuselage skin. This area is to be checked from the center aisle and through access panels inside of the airplane. (10) Inspect for possible cracks or loose fasteners in frames and angle clips of the center wing/fuselageat Fuselage Stations 188, 197, and 207. (11) Using dye penetrant procedures outlined in AC 43.13-1A, inspect the four upper forward wing to center section fittings and the eight aft wing to center section fittings for possible cracks. Do not remove the wing attachment bolts unless cracks are indicated. (12) Inspect the outer wing upper and lower spar cap and hinge for possible cracks, loose rivets, or wear of hinge. (13) Lower the flaps and remove the lower aft access covers of the outer and center wing to inspect the aft spar and ribs for possible cracks near the inboard flaps. (e) Airplane maintenance record entries must be made and notification in writing sent to the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Room 100, Wichita, Kansas 67209, stating the location and length of any cracks found during inspections required by this AD and also the total hours TIS of the component at the time the crack was discovered. Reports may be submitted by letter or through M or D (Malfunction or Defect) or MRR (Maintenance Reliability Reports) procedures. (Reporting approved by the Office of Management and Budget under OMB No. 2120-0056). (f) The eddy current inspections required by this AD must be performed by personnel who have received training and are qualified in the operation of eddy current equipment that has been calibrated using a specimen obtained from the airplane manufacturer and simulates cracking of the spar cap. (g) 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. (h) An alternative method of compliance or adjustment of the compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Mid-Continent Airport, Wichita, Kansas 67209. The request shall be forwarded through an appropriateFAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office. NOTE 2: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Wichita Aircraft Certification Office. (i) The strap inspection or modification required by this AD shall be done in accordance with Aerocon California, Inc., Engineering Order No. E.O. B-9975-2, dated November 14, 1975; and Aerocon California Service Letter, dated May 25, 1976. 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 Western Aircraft Maintenance, 4444 Aeronca Street, Boise, Idaho 83705. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 1100 L Street, NW; Room8401, Washington, DC. (j) This amendment supersedes AD 90-04-04, Amendment 39-6487, which superseded AD 77-05-01 R3 and AD 75-27-10. (k) This amendment becomes effective on August 24, 1992.
2022-03-18: The FAA is adopting a new airworthiness directive (AD) for certain British Aerospace (Operations) Limited and British Aerospace Regional Aircraft Model Jetstream Series 200, Jetstream Model 3101, and Jetstream Model 3201 airplanes. 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 identifies the unsafe condition as a bent control rod within the gust lock system, which may enable both power levers to be pushed into the flight range with the gust lock lever fully engaged. This AD requires replacing the push rod assembly with a modified push rod assembly. The FAA is issuing this AD to address the unsafe condition on these products.
97-12-03: This amendment supersedes two existing airworthiness directives (AD) that are applicable to certain Boeing Model 747 series airplanes. One of those AD's currently requires inspections for cracking, corrosion, and fracturing of the lower horizontal clevis of the strut midspar fittings, and replacement of discrepant parts with new or serviceable parts, or repair, if necessary. That AD also requires inspection for removal of broken sealant of the clevis and the fasteners, and various follow-on actions. It also provides for optional terminating actions for the inspections. The other AD currently requires inspection for cracking of certain fastener holes of the upper and lower horizontal clevis legs. This amendment continues to require inspections to detect cracking, corrosion, and fracturing of the lower horizontal clevis; and adds corresponding inspections of the upper horizontal clevis, and replacement of discrepant parts with new parts, or rework, if necessary. This amendment also removes certain optional terminating actions. This amendment is prompted by reports of cracking of the lower and upper leg of the horizontal clevis of the midspar fitting. The actions specified in this AD are intended to detect and correct cracking and fracturing of the clevis, which could result in drooping of the strut at the strut-to-wing interface, and consequent separation of the engine and strut from the airplane.
94-22-04: This amendment supersedes an existing airworthiness directive (AD), applicable to Costruzioni Aeronautiche Giovanni Agusta S.p.A. Model A109A and A109AII series helicopters, that currently imposes a calendar life limit of 10 years and 6 months on the main rotor retention strap assemblies (strap assemblies). This amendment requires reducing the calendar life limit to 8 years. This amendment is prompted by additional service experience and analyses, that show the current life limit needs to be reduced from 10 years and 6 months to 8 years to prevent deterioration and subsequent failure of the strap assemblies. The actions specified by this AD are intended to prevent failure of the strap assemblies, loss of a main rotor blade, and subsequent loss of control of the helicopter.
2011-27-05: We are superseding an existing airworthiness directive (AD) for all Saab AB, Saab Aerosystems Model 340A (SAAB/SF340A) and SAAB 340B airplanes. That AD currently requires an inspection of the main landing gear (MLG) separation bolt harness for broken wires and corroded connectors, and corrective actions if necessary; and for certain airplanes, a modification of the MLG separation bolt's electrical harness. This new AD requires replacement of the separation bolt harness. This AD was prompted by reports of broken wires and corroded connectors in the SAAB 340 MLG emergency release system. We are issuing this AD to prevent improper release of the MLG during an emergency situation, possibly resulting in damage to the airplane during landing and injury to the occupants.
97-11-12: This amendment adopts a new airworthiness directive (AD) that applies to Aerospace Technologies of Australia Pty Ltd. (ASTA) Models N22B, N22S, and N24A airplanes. This action requires repetitively inspecting the stub wing upper front spar cap flanges for cracks, and repairing any cracked part. This AD results from fatigue tests that show that the stub wing upper front spar cap flanges could fail over time because of fatigue. The actions specified by this AD are intended to prevent structural failure of the front spar caused by cracks in the stub wing upper front spar cap flanges, which could result in loss of control of the airplane.