AD 45-04-01

45-04-01 BOEING: (Was Service Note 1 of AD-704-1.) Applies to 314 Aircraft. The 24SRT aluminum alloy tubular members must be inspected for stress corrosion and fatigue cracks by visual and x-ray methods in accordance with the instructions listed below: STRESS CORROSION CRACKS INSPECTION PERIODS AND LOCATIONS (a) Inspection required every 250 hours of operation or 60 days, whichever occurs first. Inspect the visible portion of all readily accessible aluminum alloy 24SRT members for cracks. (b) Inspection required every 750 hours of operation or 120 days, whichever occurs first. Inspect the visible faces of all aluminum alloy 24SRT tubing structure for cracks. In addition, inspect by x-ray the inaccessible face of the spar chord members from Station 6 to Station 13 which is hidden by the wing skin attach to the chord (i.e., chord face areas hidden by gusset plates used to attach web members are excluded). (c) Inspection required annually. Inspect by x-ray all inaccessible portions of 24SRT spar chord members for their entire length. This inspection period may coincide with the inspection periods in paragraph (b) above. INSPECTION PROCEDURES The required visual inspection for new or elongated cracks shall be done in a manner satisfactory to FAA. The following procedure is an acceptable method for making these visual inspections: (a) Clean the surfaces of the members with a rag as necessary and closely examine the members (especially around gussets) with the naked eye. Direct a light on each member at varying angles so that no defects will be overlooked. Make certain to inspect all sides of each member using a mirror where necessary. (b) Examine any suspicious indication with a magnifying glass (10 power or over preferred). A crack will appear to have jagged edges and considerable depth. A scratch will appear to have smooth edges and the bottom of the groove should be visible. (c) If a new crack is found, the finish (if otherRoxalin Clear Primer) should be removed around the crack to facilitate inspection. Extreme care should be exercised while stripping areas immediately adjacent to gusset plates in order to prevent the stripping solvent from entering the inaccessible regions between the gussets and members. The crack should be further inspected for corrosion and its length measured to the nearest 1/16 inch. The two ends of a stress corrosion crack should be marked with a sharp indelible pencil, and Roxalin Clear Primer No. 3200 brushed over the stripped area. (d) Inspect known cracks for elongation by noting the pencil lines placed at the previous ends of each crack the same as for new cracks. (See (c) above.) The required x-ray inspections should be done with suitable equipment and by a company or personnel that have demonstrated to the FAA that their procedure will adequately show the condition of the hidden faces of the chord members. IDENTIFICATION AND LIMITS Stress corrosion typesof failures are denoted by longitudinal fissures in the members. These cracks may have a small transverse component. They vary in length and, as time elapses, may run together or continue from one rivet hole to another. If stress-corrosion cracks are within certain limits the airplane may be operated without reinforcing the affected member; however, if the magnitude, direction, or location of the crack is such as to violate any of the following provisions, the affected member shall be reinforced or replaced in a manner satisfactory to the FAA. (1) No crack should be allowed to exceed 8 inches in length. Diagonal (or transverse) cracks should in no case extend transversely in the member for a distance greater than the largest rivet or bolt diameter in the nearest fitting. (2) Cracks should not be allowed in joints, fittings, rivet holes, reduced sections, etc., unless it can be determined that the affected area is not critical or that adequate margins of safety exist to compensate for such cracks. (3) If two or more parallel cracks exist in the same face, none should exceed 6 inches in length. (4) If numerous small longitudinal cracks exist in one face of a member but are not joined by diagonal or transverse cracks, the length of the member so affected should not exceed 12 inches. FATIGUE CRACKS INSPECTION PERIODS AND LOCATIONS. Inspection required every 35 hours of operation. Inspect the visible portions of all the wing spar 24SRT diagonal tube members, between Stations 1 and 30 on the front spar and between Stations 5 and 23 on the rear spar, for fatigue cracks at intervals not to exceed 35 hours flight time. INSPECTION PROCEDURES. Same as parts (a) and (b) of the inspection procedures for stress-corrosion cracks. IDENTIFICATION AND LIMITS. Fatigue types of failures are denoted by fine hairline transverse cracking in the members. These cracks generally emanate from rivet holes under gussets and progress transversely or diagonally around the periphery of the tube. They may also originate from a longitudinal crack, scratch or other stress-raising discontinuity. In case fatigue cracks of any length are found, the defective member should be reinforced before flight is continued and upon arrival at the home base, the member should replaced. A proposed repair to take care of this contingency should be submitted to the FAA for approval. It shall also be the operators' responsibility to keep a record of all the cracks on each airplane. This record shall be revised periodically to show the status of existing cracks and to record newly developed cracks. Copies of the original reports and all revised pages should be submitted to the FAA for examination.