72-07-09: Cracks And Loose Bolts In Fin & Rudder

72-07-09 CESSNA: Amendment 39-1423 as amended by Amendment 39-1988. Applies to all 205, 206, P206, TP206 series; 182 series (Serial Numbers 18253599 through 18261528); U206 and TU206 series (Serial Numbers U206-0276 through U20601905); 207 and T207 series (Serial Numbers 20700001 through 20700216); 210 series (Serial Numbers 21057841 through 21059739); and T210 series (Serial Numbers T210-0001 through T210-0454 and 21059200 through 21059739) airplanes. Compliance: Required as indicated, except that the inspections accomplished per AD 71-21-1 are acceptable in lieu of the initial inspections required herein. To detect cracks and bolt looseness which could lead to inflight separation of the fin and rudder, within the next 50 hours' time in service after the effective date of this AD on all airplanes with 1000 or more hours' time in service, and thereafter at intervals not to exceed 1000 hours', except that 206 series airplanes with the floatplane vertical stabilizer are to be inspected initially at 500 hours and at 500 hour intervals thereafter, accomplish the following: A) Remove stinger and fairing between the vertical and horizontal stabilizer. B) Visually inspect the aft fin spar attachment bolts for indications of looseness (i.e. cracked inspection putty or signs of working). C) If any indications of looseness of the bolt attachments are found during the inspection required by Paragraph B, prior to further flight, accomplish the following: 1) Remove and reinstall bolts one at a time while dye penetrant inspecting attachment holes for cracks. 2) Check attachment holes for excessive elongation. AN5 bolt holes can be redrilled for AN6 bolts and AN6 bolt holes can be redrilled for AN7 bolts. Acceptable limits for bolt holes are as follows: a) 0.313 to 0.319 inches for AN5 bolts b) 0.375 to 0.381 inches for AN6 bolts c) 0.437 to 0.444 inches for AN7 bolts If reaming does not remove all signsof bolt hole elongation, repair or replace damaged parts, as appropriate. 3) When reinstalling bolts use the following torque values: a) AN5 140-225 inch pounds b) AN6 190-390 inch pounds c) AN7 500-840 inch pounds Add washers as required to provide sufficient threads when torquing the bolts. D) Visually inspect the forward and aft vertical fin attachment structure for cracks and dye penetrant or equivalent inspect the following for cracks: 1) the lower inboard corner of each rudder cable cutout in the aft bulkhead. 2) the area in the aft bulkhead between the rudder cable cutout and the bulkhead flange. E) If cracks are found during the inspection required by Paragraph D, prior to further flight: 1) Replace complete aft bulkhead assembly. 2) Dye penetrant or equivalent inspect forward face of the front fin spar attach fitting for cracks from the top of the mount holes to one inch above same. If cracks are found, thefront fin spar attach fitting must be replaced. 3) Dye penetrant or equivalent inspect the front fin spar attach bulkhead attachment holes for cracks. If cracks are found, accomplish one of the following: a) Replace the bulkhead. b) Ream the bolt holes to the next larger size (AN6) in accordance with the tolerances set forth in Paragraph C(2) provided all evidence of cracks is removed. c) An FAA-approved equivalent repair. F) Install inspection putty on all nuts and bolts. G) Replace stinger and fairings. H) Equivalent methods of compliance with this AD must be approved by Chief, Engineering and Manufacturing Branch, FAA, Central Region. I) Report any discrepancies on a Malfunction or Defect Report, FAA Form 8330.2 (Form 1226). (Reporting approved by the Bureau of the Budget under BOB No. 04-R0174). Cessna Service Letter SE72-3, dated February 11, 1972, pertains to this subject. NOTE: It is recommended that the inspection required by this AD, be accomplished any time the airplane has been exposed to severe wind or gust conditions. This AD supersedes AD 71-21-1. Amendment 39-1423 became effective April 1, 1972. This Amendment 39-1988 becomes effective October 17, 1974.