The Effect of Center Delamination on the Instability of Composite Cylindrical Panels.
Abstract
The buckling loads of 8-ply graphite-epoxy delaminated cylindrical panels were determined experimentally. The analysis included two different ply orientations, two different aspect ratios (length/chord), two different delamination sizes, and one set of boundary conditions; clamped along the top and bottom edges and simply supported along the vertical sides. The experimental test results were compared to the linear bifurcation loads obtained from the STAGSC-1 finite element computer code and the buckling loads of panels with square cutouts which were also obtained from STAGSC-1. The experimental testing was accomplished by the Air Force Flight Dynamics Laboratory. Experimental difficulties were encountered due to the inability of the test apparatus to evenly distribute the load through the simply supported boundary conditions. The uneven load distribution resulted in nonsymmetric buckling patterns, however, the buckling loads showed good correlation between analytical and experimental values. All experimental buckling loads were less than the analytical value predicted by STAGSC-1. The amount of decrease in load depended on ply orientation, aspect ratio, and delamination size. Additional keywords: Experimental shell analysis; and Manufacturing. (Author)
Document Details
- Document Type
- Technical Report
- Publication Date
- Dec 01, 1984
- Accession Number
- ADA153494
Entities
People
- G. R. Seifert
Organizations
- Air Force Institute of Technology