Abstract
With advanced composite materials expected to appear to a greater and greater extent in aircraft primary structure, their inherent weight-savings attractiveness is enhanced by permitting relatively lightly loaded plate elements to operate in the postbuckled state. A theoretical development and analysis procedure is presented for prediction of buckling, postbuckling and crippling loads in laminated composite plates. Specific application is made to a number of simply supported, graphite epoxy plates with geometric and material properties corresponding to those included in several experimental programs, the results of which, in the form of load-shortening curves, have been reported in the literature. With the effects of transverse shear and material nonlinearity combined with the maximum-strain failure criterion included in the theoretical analysis, good agreement is obtained with the experimental results for initial buckling, postbuckling stiffness and failure (crippling). The theory and analysis described herein can be used as an aid in the design process, particularly in the isolation of candidate laminates without resorting to extensive and costly test programs.
Talk to us
Join us for a 30 min session where you can share your feedback and ask us any queries you have
Schedule a callDisclaimer: All third-party content on this website/platform is and will remain the property of their respective owners and is provided on "as is" basis without any warranties, express or implied. Use of third-party content does not indicate any affiliation, sponsorship with or endorsement by them. Any references to third-party content is to identify the corresponding services and shall be considered fair use under The CopyrightLaw.
