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52-02-01: 52-02-01 Convair: Applies to All Model BT-13, -13A, -13B and BT-15 Aircraft With Bronze Rear Spar to Center Section Attachment Fittings. \n\n\tCompliance required as indicated. \n\n\tWhen it has been determined that the rear spar to center section attachment fittings are bronze castings, compliance with this Airworthiness Directive should be effected every 24 hours flying time or every 6 months nonflying time, whichever occurs first. \n\n\t(1)\tInspect visually the rear spar to center section attachment fittings on both sides of the rear spar for evidence of cracks. All cracked fittings (see Figure 1) should be replaced with new fittings of equivalent or greater strength. \n\n\n\n\nFIGURE 1 \n\n\n\t(2)\tA magnet may be used to determine if the fitting is a steel casting. \n\n\t(3)\tFittings that have been made from aluminum alloy forgings can be readily differentiated from those made from bronze castings by visual inspection.
2021-24-04: The FAA is adopting a new airworthiness directive (AD) for certain Bell Textron Canada Limited (type certificate previously held by Bell Helicopter Textron Canada Limited) Model 505 helicopters. This AD was prompted by the determination that reducing the pressure altitude limitations for certain fuel types is necessary. This AD requires revising the existing Rotorcraft Flight Manual (RFM) for your helicopter. The FAA is issuing this AD to address the unsafe condition on these products.
2010-04-10: The FAA is superseding an existing airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: During the flight test campaign of the A380-861 model (Engine Alliance powered), some cracks were found on the Movable Flap Track Fairing number 6 (MFTF6). These cracks were located at the pivot attachment support-ring and at the U-frame in the attachment area to aft-kinematic. In addition, delamination has been observed within the monolithic Carbon Fibre Reinforced Plastic (CFRP) structure around the pivot support-ring. This condition, if not corrected, could lead to in-flight loss of the MFTF6, potentially resulting in injuries to persons on the ground. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
86-19-10: 86-19-10\tMOONEY AIRCRAFT CORPORATION: Amendment 39-5408. Applies to Models M20 and M20A (all serial numbers) airplanes certificated in any category.\n\n\tCompliance: As indicated in the body of the AD.\n\n\tTo preclude structural failure due to deteriorated wooden structures, accomplish the following:\n\n\t(a)\tWithin the next 30 days after the effective date of this AD, if AD 76-15-01 has not been previously complied with, or within 12 calendar months after the last 12 month repetitive inspection required by AD 76-15-01, whichever is applicable, and each 12 months thereafter accomplish the wood structure proof load tests, modifications and visual inspections specified in paragraphs (d), (e), (f) and (g) of this AD and repair all discrepancies found prior to further flight.\n\n\t(b)\tThe empennage proof load tests and modification and inspection requirements of paragraphs (a), (d) and (e) are not required on airplanes modified with an all-metal empennage installed per Mooney Service Bulletin (S/B) Kit No. M20-170-1.\n\n\t(c)\tWithin the sixth month after the accomplishment of the inspections required by paragraph (a) of this AD, or within 12 calendar months after the last six month interval repetitive inspection required by AD 76-15-01 and each 12 months thereafter, whichever is applicable, accomplish the empennage and wing inspections specified in Parts IIA, III8, III9, and III10 of Mooney S/B No. M20-170A dated February 24, 1969, and repair all discrepancies found prior to further flight.\n\n\t(d)\tOn airplanes not equipped with an all-metal empennage, proof load test the empennage and supporting structure as follows:\n\n\t\t(1)\tApply proof loads to the vertical fin spar as shown in Figure 1 of this AD. Apply the load to the right side and then to the left side. Apply proof loads to the rudder hinges of the fin as shown in Figure 3 of this AD. Apply hinge proof loads to the right side and then to the left side.\n\n\t\t(2)\tAfter the initial proof load testing in paragraph(d)(1) of this AD, at intervals not to exceed one year, apply the proof loads to the vertical fin spar and rudder hinges as shown in Figures 1 and 3 of this AD. Apply the loads in one direction only during each 12 month inspection cycle by applying the loads to the right side at the end of the first interval and alternating load direction similarly thereafter.\n\n\t\t(3)\tIf the empennage fails during the proof loading specified in paragraphs (d)(1) or (d)(2) of this AD, prior to further flight, replace the wood empennage with an all-metal empennage in accordance with the instructions contained in Mooney S/B Kit No. M20-170-1. The empennage is considered to have failed when complete separation of the vertical fin from the horizontal stabilizer occurs, fin spar cracks occur, the hinge separates or loosens from the fin, other wood failures occur, glue joint failures occur, or permanent deformation occurs as shown in Figure 2 of this AD.\n\n\t(e)\tModify and inspect the wood empennage as follows:\n\n\t\t(1)\tIf fin failure did not occur during the proof load application specified in paragraphs (d)(1) and (d)(2) of this AD, prior to further flight, modify the vertical fin and visually inspect the empennage as follows:\n\n\t\t\t(i)\tModify the vertical fin by adding inspection access holes and reinforcing straps in accordance with Parts IB and IC of Mooney S/B No. M20-170A dated February 24, 1969, unless previously accomplished.\n\n\t\t\t(ii)\tVisually, and if necessary using the methods in paragraph (h) of this AD, inspect the empennage in accordance with S/B No. M20-170A dated February 24, 1969, Parts IA and II and repair any discrepancies found prior to further flight.\n\n\t(f)\tPrepare the wing and wing carry-thru structure for inspection in accordance with instructions in Part III of Mooney S/B No. M20-170A dated February 24, 1969, and as follows:\n\n\t\t(1)\tRemove the wing to fuselage fairings and fillets.\n\n\t\t(2)\tRemove all the wing and center section access doors and panels. (Refer to Figure 5 of Mooney Service Bulletin (S/B) No. M20-170A).\n\n\t\t(3)\tRemove the sealing tape at the wing-fuselage joint.\n\n\t\t(4)\tRemove the rear seat and auxiliary fuel tank for access to the wing center section.\n\n\t\t(5)\tRemove flap gap metal seal strips from the top trailing edge of both wings.\n\n\t\t(6)\tDisconnect and remove the wing flaps and ailerons.\n\n\t(g)\tVisually, and if necessary using the methods in paragraph (h) of this AD, inspect the wing and wing carry-thru structure and repair any discrepancies found prior to further flight in accordance with instructions in Part III of Mooney S/B No. M20-170A dated February 24, 1969, and as follows:\n\n\t\t(1)\tInspect the areas around wing to fuselage attach fittings for evidence of deterioration or joint separation.\n\n\t\t(2)\tInspect the flap and aileron attach bolts, bearings, bushings, and hinge fitting attach bolts and bushings for evidence of rust, corrosion, and wear. See Figure 6 in Mooney S/B No. M20-170A dated February 24, 1969.\n\n\t\t(3)\tVisually inspect the wood end-grain surrounding bolt holes for evidence of rust, discoloration, deterioration, and evidence of moisture accumulation at the trailing edges of the wings.\n\n\t\t(4)\tVisually inspect the rear stub spar for glue bond separations, water stains, and wood rot. If these inspections identify any questionable areas in which possible deterioration may exist in the concealed spar caps, prior to further flight, determine the condition of the internal spruce core in accordance with paragraph (h) of this AD. See Note 1 of this AD.\n\n\t\t(5)\tVisually inspect all accessible areas of the main spar from the fuselage center line (BL 0.0) out to left and right wing Station 59.25 for glue bond separations, water stains, and wood rot. If these inspections identify any questionable areas in which possible deterioration may exist in the concealed spruce spar prior to further flight determine the condition of the internal spruce core in accordance with paragraph (h) of this AD. See Note 1 of this AD.\n\n\t\t(6)\tVisually inspect the accessible interior of the wing using a flashlight and mirror, for wood decay, water and/or wood stains, pooled dust/dirt which may indicate evidence of previous standing water, rust or corrosion on metallic surfaces, wood discoloration, and detectable moisture. See Note 1 of this AD.\n\n\t\t(7)\tInspect the upper and lower exterior surfaces of the wing, including the wheel well, for the following: See Note 1 of this AD.\n\n\t\t\t(i)\tIndications that the wood immediately below the fabric is soft or contains excessive moisture (i.e., swollen). Soft wood may be located and/or confirmed by depressing the wing's surface in the vicinity of the area in question with a rounded, blunt instrument and comparing its hardness to that of good wood. Note that any areas being compared must have identical substructure.\n\n\t\t\t(ii)\tIndications that the fabric/paint is delaminating from the wood surface (bubbles, discoloration, boils, softspots and other surface flaws).\n\n\t\t\t(iii)\tCracks or breaks in the paint which could allow water to enter the wing.\n\n\t\t\t(iv)\tAny other exterior damage which would allow water to penetrate the fabric/paint barrier and enter the wood.\n\n\t\t(8)\tVisually inspect the rear spar in all areas it is accessible from the fuselage center line out to the left and right wing tips for wood rot, water stains in wood and glue joint separation. Pay special attention to the area around all flap and aileron hinge supports, including the support ribs, lower wing skins, spars and closeout strips at wing stations 22.0 and 147.0 (flap inboard and outboard hinge support ribs).\n\n\t\t(9)\tInspect all drain holes on the bottom of the wing to ensure they are completely open and free of burrs and/or pieces of fabric.\n\n\t\t(10)\tVisually inspect the fuel scupper areas of the main and auxiliary fuel tank fillers for sealant condition between scupper boxes and wing structure.\n\n\t\t(11)\tCheck main and auxiliary fueltank scupper drains to be sure they are not clogged.\n\n\t\t(12)\tVisually inspect aileron and flap fabric covering under metal gap strips in accordance with Mooney S/B No. M20-29 dated December 4, 1957.\n\n\t\t(13)\tVisually inspect the areas of the upper wing surface trailing edge under flap gap metal seals for fabric or wood deterioration. Be alert for deteriorated wood around screw holes used in holding the metal strip to the wing.\n\n\t\t(14)\tVisually inspect, if necessary repair, and refinish the main landing gear wheel well area in accordance with Mooney S/B No. M20-67 dated February 15, 1960.\n\n\t(h)\tIf during any inspections specified in paragraphs (e) and (g) of this AD there are visual indications of wood deterioration below the surface, prior to further flight, inspect and test these areas to assure their structural integrity by using one or more of the following:\n\n\t\t(1)\tTest for soft/decayed wood with a sharp probe such as an awl or sharp pocket knife.\n\n\t\t(2)\tDisassemble thestructure as necessary to gain access to the area and perform a detailed visual inspection.\n\n\t\t(3)\tTap the wood area in question with a small rounded blunt instrument approximately the size of a small pocket knife. Compare the sound to similar areas that are not suspect. Assuming similar understructure, an abrupt change in sound to a less or non-resilient sound may indicate decay below the surface.\n\n\t(i)\tIf significant structural repair of the wing main spar, rear spar empennage is found necessary as a result of the inspections and tests of the preceding paragraphs, prior to initiation of the repair, contact Mooney Aircraft Corporation, Post Office Box 72, Kerrville, Texas 78028; Telephone (512) 896-6000, or the local Mooney Aircraft Repair Center, or FAA Airplane Certification Office, ASW-150, FAA, Southwest Region, Post Office Box 1689, Fort Worth, Texas 76101; Telephone (817) 624-5164, to arrange for engineering review and approval of the repair design.\n\n\t(j)\tWithin 30 daysafter the accomplishment of the first inspection required by paragraph (a) of this AD, the appropriately rated airframe mechanic who performed the inspection shall fill out and sign the one time reporting form included as Attachment 1 to this AD and mail it to the following address: DOT/FAA, Airplane Certification Branch, ASW-150, Post Office Box 1689, Fort Worth, Texas 76101. (Reporting approved by the Office of Management and Budget (OMB) under OMB No. 2120-0056).\n\n\tNOTE: This is a one time only reporting requirement.\n\n\t(k)\tAn equivalent method of compliance with this AD may be used if approved by the Manager, Airplane Certification Branch, ASW-150, FAA, Southwest Region, Post Office Box 1689, Fort Worth, Texas 76101.\n\n\tAll persons affected by this directive may obtain copies of the documents referred to herein upon request to Mooney Aircraft Corporation, Post Office Box 72, Kerrville, Texas 78028, or FAA, Office of the Regional Counsel, Room 1558, 601 East 12th Street, KansasCity, Missouri 64106.\n\n\tNote 1: The surface features described in the paragraphs of this AD may be accentuated by illuminating the surface with a light source at a shallow angle. The following technique may be used by an experienced inspector to detect soft and/or decayed wood in the wing spars. Tap the wing directly above and below the spars with a small rounded blunt instrument approximately the size of a small pocket knife. Start at the outboard end and move inboard, listening to the sound generated by the wing.\n\n\tThe sound quality will change slowly. If the change is abrupt or if the sound is not resilient, the wood directly below the surface may be deteriorated due to decay.\n\n\tNote 2: Shelter - Owners and operators are encouraged to shelter the airplanes, to keep the airplane out of rain storms, and to protect the fabric surface from unnecessary exposure to the deteriorating effects of the sun.\n\n\tNote 3: Maintenance - Owners and operators are encouraged to be selective in who performs maintenance on their airplane. Only personnel extremely experienced in wood airplane inspection and repair should be contacted.\n\n\tNote 4: The inspection intervals required by this AD differ from the inspection intervals shown in Mooney Service Bulletin No. M20-170A. The intervals in this AD are the same as AD 76-15-01 which this AD supersedes.\n\n\tNote 5: Repairs to primary and secondary structure may be accomplished with reference to:\n\n\t\t(a)\tFAA Advisory Circular No. 43-13-1A: Acceptable Methods, Techniques and Practices Aircraft Inspection and Repair, Department of Transportation, Federal Aviation Agency 1972; available through the Government Printing Office.\n\n\t\t(b)\tANC-18: Design of Wood Aircraft Structures, Chapter 4, Munitions Board Aircraft Committee June 1951.\n\n\t\t(c)\tMooney Aircraft Corporation Engineering Drawings; the specific drawings required will depend on the affected structural components.\n\n\t\t(d)\tMooney Service Bulletin No. M20-170A dated February 24, 1969.\n\n\tNote 6:\tDesign of major repairs to primary wood structure (main and stub spars and ribs receiving loads such as landing gear loads or loads related to attachment of moveable control surfaces to fixed surfaces or attachment of fixed surfaces to the fuselage) should be reviewed and approved by Mooney Aircraft Corporation or an FAA Designated Engineering Representative having appropriate ratings or by FAA Aircraft Certification Division engineers. This is not intended to apply to those situations wherein a deteriorated part is replaced with an entire new part of like design.\n\n\tThis AD supersedes AD 76-15-01, Amendment 39-2673.\n\n\tThis amendment becomes effective October 6, 1986.\n\n\n\n\n\n\n\nFIGURE NO. 1 - APPLICATION OF FIN SIDE\nPROOF LOAD ON MAIN SPAR\n\n\n\n\n\n\n\n\n\nFIGURE 1A\nRECOMMENDED FIN PROOF LOAD METHOD\n(INCLUDING LOWER RESTRAINT)\n\n\n\n\n\n\n\n\n\n\n\n\n\nVIEW LOOKING FORWARD\n\nFIGURE NO. 2 - FIN PERMANENT DEFORMATION\nLIMITSFIN SIDE VIEW\n\nFIGURE NO. 3 - FIN - RUDDER HINGE SIDE PROOF\nLOAD APPLICATION\n\n\n86-19-10 ATTACHMENT 1\n\nInspection Results Reporting Form, Mooney M20 and M20A Wood Structure Inspection.\n\nI.\tAirplane Model No. , Serial No. .\n\nII.\tAirplane N-Number .\n\nIII.\tDoes the airplane have a metal empennage installed:\n\tYes NO \t\n\tIf no, go to IV, if yes, skip to V.\n\nIV.\tFor airplanes that have the wooden empennage:\n\n\t- Did the visual inspection of the empennage result in discovery of any failed glue joints, rotted wood or delaminated wood? \n\tYes No \t\n\tIf yes, describe location and extent of problem.\n\t- Did the visual inspection of the empennage result in discovery of any deteriorated paint or fabric covering?\n\tYes No \t\n\tIf yes, describe location and extent.\n\t- Did the visual inspection of the empennage result in discovery of any indications of water inside the empennage?\n\tYesNo \t\n\tIf yes, describe where it got in and where it pooled.\n\t- Did the empennage pass or fail the proof load tests of paragraph (d)?\n\tYes No \t\n\t- This inspection and testing of the wood empennage on this airplane results in the conclusion that its overall general structural condition is:\n\tPoor \t\n\tFair \t\n\tGood \t\n\tExcellent \t\n\t- Make any comments you wish to make about adding or deleting inspections and testing of the wood empennage on M20 and M20A airplanes.\n\nV.\tWing inspection results.\n\t- Did the visual inspection of the wing and wing carry-thru result in discovery of any failed glue joints, rotted wood or delaminated wood?\n\tYes No \t\n\tIf yes, describe location and extent of problem.\n\t- Did the visual inspection of the wing and wing carry-thru result in discovery of any deteriorated paint, fabric covering plugged drain holes or deteriorated scupper box seal?\n\tYes No \t\n\tIf yes, describe problem, location and extent.\n\t- Did the visual inspection of the wing and wing carry-thru result in discovery of any indications of water inside the wing?\n\tYes No \t\n\tIf yes, describe where it got in and where it pooled.\n\t- This inspection of the wing on this airplane results in the conclusion that its overall general structural condition is:\n\tPoor \t\n\tFair \t\n\tGood \t\n\tExcellent \t\n\t- Make any comments you wish to make about adding or deleting inspections of the wing on M20 and M20A airplanes:\nMechanics Name: \t\nFAA Certificate Number: \t\nDate:
48-14-04: 48-14-04 DOUGLAS: Applies to All C-54 Series Aircraft Prior to Model C-54G. \n\n\tCompliance required by November 1, 1948. \n\n\tBecause of fire hazard install a tailpipe shroud deflector on each tailpipe shroud assembly to prevent flame from a zone 1 fire entering the engine accessory section through the space existing between the shroud and cowling. AN 3-3A bolts may be substituted for the AN 3C-3A bolts called out in the Service Bulletin. \n\n\t(Douglas Service Bulletin C-54-289 addendum covers this same subject.)
59-13-06: 59-13-06 DOUGLAS: Applies to the Following Aircraft: DC-6A Serial Numbers 43296, 43297, 43817-43819, 43839-43841, 44063, 44064, 44069-44073, 44076, 44257. DC-6B Serial Number 43257-43259, 43261-43276, 43291, 43292, 43298-43300, 43518-43537, 43539-43547, 43549-43555, 43557-43564, 43738-43741, 43743-43746, 43748-43750, 43820-43822, 43824-43826, 43828-43834, 43836, 43837, 43842, 43844-43847, 44056-44062, 44080-44083, 44087-44089, 44102-44113, 44165-44168 and 44251. \n\n\tCompliance required as indicated. \n\n\tSeveral instances of lower front spar cap cracks have been reported. The cracks involved are located at wing Station 30 at the intersection of lower front spar cap aft tang and wing-to-fuselage attach angle. To date no cracks have been found in the lower center spar cap; however, due to the similarity of the structure, it is logical to assume that cracks can occur in this cap as well as in the lower front spar cap. To detect cracking of the lower front and center spar cap tangs at intersection with lower fuselage attach angle, the following must be accomplished on affected aircraft having in excess of 12,000 flying hours. \n\n\t(a)\tInspect lower front spar caps at nearest maintenance inspection period to 200 flight-hours unless similar inspection has been conducted within last 1,250 flying hours. \n\n\t(b)\tInspect lower front and center spar caps at maintenance inspection period nearest to each succeeding 1,250 flying hours. \n\n\t\t(1)\tAt first 1,250-hour inspection period, holes in aft tank of front spar lower cap and fuselage attach angle should be enlarged and new attachments installed (KIT "A" of Douglas SB A-821 or equivalent). \n\n\t\t(2)\tAt next regularly scheduled overhaul period, holes located in forward tank of front spar lower cap should be enlarged and new attachments installed (KIT "A" of Douglas SB A-821 or equivalent). \n\n\t(c)\tIn event spar cap cracking is found at the 200-hour initial or 1,250-hour repetitive inspection periods, temporary rework per drawing No. 3645935 (KIT "B"), or equivalent, may be accomplished. \n\n\tWith temporary rework installed, inspection must be repeated at the 1,250-hour intervals for a maximum of 3,200 flight-hours at which time permanent rework per drawing No. 5761922 (KIT "C"), or equivalent, must be accomplished. \n\n\t(d)\tAll aircraft not already reworked per (c) must have permanent rework per drawing No. 5761922 (KIT "C"), or equivalent, accomplished within the next 6,400 flying hours. \n\n\t(e)\tAfter installation of KIT "C", operators may revert to normal repetitive inspection periods not to exceed 3,200 flying hours. \n\n\t(Douglas Service Bulletin DC-6 No. A-821 dated March 19, 1960, covers this same subject.) \n\n\tNOTE: This AD is not presently applicable to aircraft Serial Numbers 43548 and 43152 since they are not currently under U. S. registry. However, compliance with this AD will be required at the time application is made for recertification of such aircraft in the U.S.
2021-23-12: The FAA is adopting a new airworthiness directive (AD) for all transport and commuter category airplanes equipped with a radio (also known as radar) altimeter. This AD was prompted by a determination that radio altimeters cannot be relied upon to perform their intended function if they experience interference from wireless broadband operations in the 3.7-3.98 GHz frequency band (5G C-Band). This AD requires revising the limitations section of the existing airplane/ aircraft flight manual (AFM) to incorporate limitations prohibiting certain operations requiring radio altimeter data when in the presence of 5G C-Band interference as identified by Notices to Air Missions (NOTAMs). The FAA is issuing this AD to address the unsafe condition on these products.
2021-24-03: The FAA is adopting a new airworthiness directive (AD) for all Airbus Helicopters Model AS355NP helicopters. This AD was prompted by a report of mechanical deformation found on the protective cover (also referred to as switch guard) of the ''SHEAR'' control pushbutton installed on a co-pilot collective stick of a Model EC225LP helicopter, caused by incorrect handling; due to having an identical design switch guard installed on the pilot collective stick, Model AS355NP helicopters are also affected. This AD requires replacement of the protective cover of the ''SHEAR'' control pushbutton, and re- identification of the pilot collective stick, as specified in a European Union Aviation Safety Agency (EASA) AD, which is incorporated by reference. The FAA is issuing this AD to address the unsafe condition on these products.
2010-04-16: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: The Civil Aviation Authority of the United Kingdom (UK) has informed EASA [European Aviation Safety Agency] that significant quantities of Halon 1211 gas, determined to be outside the required specification, have been supplied to the aviation industry for use in fire extinguishing equipment. * * * * * * * * * * * This Halon 1211 has subsequently been used to fill P/N [part number] 1708337B4 portable fire extinguishers that are now likely to be installed in or carried on board aircraft. The contaminated nature of this gas, when used against a fire, may provide reduced fire suppression, endangering the safety of the aircraft and its occupants. Inaddition, extinguisher activation may lead to release of toxic fumes, possibly causing injury to aircraft occupants. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
76-14-07 R2: 76-14-07 R2 CESSNA: Amendment 39-2760 as amended by Amendment 39-2778 is further amended by Amendment 39-5124. Applies to Models 210 thru 210J (Serial Numbers 57001 thru 57575, 21057576 thru 21059199) and Models T210F thru T210J (Serial Numbers T210-0001 thru T210-0454) airplanes. \n\n\tCompliance: Required as indicated, unless already accomplished. \n\n\tTo decrease the possibility of main landing gear extension failures, accomplish the following: \n\n\tA.\tOn Model 210 and 210A (Serial Numbers 21057001 thru 21057840) airplanes with 1,000 hours or more time in service or upon accumulation of 1,025 hours time in service on those aircraft with less than 1,000 hours time in service: \n\n\t\t1.\tWithin 25 hours time in service after the effective date of this AD and within each 25 hours time in service thereafter, inspect Part Numbers 1241004-1 and 1241004-2 landing gear saddles for cracks using dye penetrant procedures in accordance with the instructions outlined in Paragraph E. of this AD. Particular attention should be given to the critical areas shown in Figure 1 of this AD. When all modifications specified in Paragraph A.3. have been accomplished, the requirements of this Paragraph A.1. are no longer applicable. \n\n\t\t2.\tPrior to further flight, replace any cracked saddles found during any inspection required by Paragraph A.1. \n\n\t\t3.\tWithin 100 hours time in service after August 16, 1976, or prior to April 1, 1977, whichever occurs later, and thereafter at each 1,000 hours time in service replace P/N's 1241004-1 and 1241004-2 main landing gear saddles with new components having the same P/N's in accordance with Cessna Service Letter SE 75-26 dated December 5, 1975, or later approved revisions. \n\n\tB.\tOn Models 210B thru 210G (Serial Numbers 21057841 thru 21058936) and T210F and T210G (Serial Numbers T210-0001 thru T210-0307) airplanes with 1,000 or more hours time in service or upon accumulation of 1,025 hours time in service on those aircraft with less than 1,000 hours time in service: \n\n\t\t1.\tWithin 25 hours time in service after the effective date of this AD and within each 25 hours time in service thereafter, inspect P/N's 1241423-1 and 1241423-2 main landing gear saddles for cracks using dye penetrant procedures in accordance with the instructions outlined in Paragraph E. of this AD. Particular attention should be given to the critical area shown in Figure 2 of this AD. When all modifications specified in Paragraph B.3. have been accomplished, the requirements of this Paragraph B.1. are no longer applicable. \n\n\t\t2.\tPrior to further flight, replace any cracked saddles found during any inspection required by Paragraph B.1. \n\n\t\t3.\tWithin 100 hours time in service after August 16, 1976, or prior to April 1, 1977, whichever occurs later, replace P/N's 1241423-1 and 1241423-2 main landing gear saddles with improved saddles of the same part number in accordance with Cessna Service Letter SE 75-26 dated December 5, 1975, or later approved revisions. \n\n\tNOTE (1): The improved main landing gear saddle for Models 210B thru 210G, T210F and T210G aircraft is identified in Figure 3 accompanying this AD. \n\n\tC.\tOn those airplanes having improved main landing gears saddles installed per Paragraph B. and on Models 210H and 210J (Serial Numbers 21058937 thru 21059199) and Models T210H and T210J (Serial Numbers T210-0308 thru T210-0454) airplanes, within the next 100 hours time in service after the effective date of this AD, for airplanes with over 1,200 hours time in service or upon accumulation of 1,300 hours time in service for those airplanes with less than 1,200 hours time in service, and at each annual inspection thereafter, inspect the P/N's 1241423-1 and 1241423-2 main landing gear saddles for cracks using dye penetrant procedures in accordance with the instructions outlined in Paragraph E. of this AD. Particular attention should be given to the critical area shown in Figure 2 accompanying this AD. Replace any saddles showing evidence of cracks. \n\n\tD.\tOn those airplanes on which main landing gear saddles have been replaced, base the compliance time for Paragraphs A., B., and C. on the new saddles time in service rather than the airplane time in service. \n\n\tE.\tPerform the dye penetrant inspections required by Paragraphs A.1., B.1., and C. of this AD as outlined in either procedure 1 or 2 below: \n\n\t\t1.\tProcedure 1: Place the airplane on jacks, disconnect the main landing gear doors, retract \nthe landing gear and perform dye penetrant inspection of the saddle fittings from underneath the airplane, or: \n\n\t\t2.\tProcedure 2: With the airplane in normal ground attitude, remove the inspection cover in the floorboard area of the airplane and perform dye penetrant inspection of the saddle fittings from inside the airplane. \n\n\t\t\tRefer to applicable Cessna Maintenance Manual instructions for disconnecting main landing gear doors and removal of inspection cover in floorboard area. \n\n\tF.\tAn equivalent means of compliance with this AD may be used if approved by the Manager, Aircraft Certification Office, Federal Aviation Administration, 1801 Airport Road, Room 100, Mid-Continent Airport, Wichita, Kansas 67209; telephone (316) 946-4400. \n\n\tG.\tInstallation of main landing gear saddles part numbers 1294151-1 and 1294151-2 in lieu of part numbers 1241423-1 and 1241423-2 constitutes an equivalent means of compliance for this AD. \n\n\tNOTE (2): A significant savings of manhours will result if initial compliance with this AD and modifications required by AD 76-04-01 are accomplished at the same time. \n\n\tNOTE (3): It is imperative that new saddles required to comply with this AD be ordered immediately to assure that a sufficient supply of saddles will be available for modification of all aircraft on or before April 1, 1977. The purpose of this admonishment is to forewarn owners/operators to avoid grounding of their aircraft for failure to comply with this AD by April 1, 1977.Amendment 39-2760 became effective August 16, 1976. \n\n\tAmendment 39-2778 became effective December 2, 1976. \n\n\tThis Amendment 39-5124 becomes effective August 28, 1985.