Integrated Research on Carbon Composite Materials. Part 4. Volume 3. Structural Component Development
Abstract
The final design of the representative fuselage component was determined, and the component was fabricated with treated Thornel 50 graphite- fiber, FRL2256 epoxy matrix composites. The structure consists of a tapered cylindrical skin, 48 inches long with end diameters of 24 inches and 20 inches; the fiber lay-up orientation of the skin is (90, + 15, 90). The skin is stiffened by 31 longitudinal stringers having (+ 10, -10, +10) fiber orientation and by three segmented ring stiffeners consisting of a balsa wood core reinforced with panels of (0, +45) orientation. The fuselage component structure was analyzed by the discrete element method; stresses, displacements, and margin of safety predictions were obtained for various loading conditions; failure levels were well above design requirements. The 90 layers of the skin were constructed by wet-winding and the inner 15 layers by hand lay-up of prepregged sheet. The stringers and the ring stiffener panels were molded from pre- stiffener panels were adhesively bonded to the component skin. Both ends of the component were reinforced with a lay-up of fiberglass tape and bonded to segmented aluminum rings for attachment of the component to the test stand. The NDT INSPECTION OF THE COMPONENT REVEALED DEFECTS CONSISTING PRIMIARILY OF PARTIAL DEBONDING OF THE STRINGERS AND RING-STIFFENERS FROM THE SKIN. Extensive analysis indicated that the debonding very likely resulted from thermal degradation of the adhesive and from cure of the end attachments. Novel repair techniques for composite structures were established, and the component was successfully repaired. Detailed plans for the various response tests and the final destruct test of the component are also presented.
Document Details
- Document Type
- Technical Report
- Publication Date
- Jul 01, 1970
- Accession Number
- AD0873386