Carbon-fiber-reinforced silicon-carbide (C/SiC) composite technologies are being developed and advanced with the notion that they will find widespread use in the aerospace industry in the near future. Potential applications include turbine blades, combustion chambers, control surfaces, and thermal protection systems. Indeed, C/SiC composites are an attractive option for designers of advanced spacecraft and advanced space propulsion systems, since C/SiC composites are lightweight and maintain their strength and stiffness at high temperatures.
However, many challenges must be overcome to advance C/SiC composite technologies. These composites are vulnerable to carbon fiber oxidation from environmental oxygen attack. Matrix microcracks formed during the fabrication process and a significant pore volume due to insufficient infiltration provide a free path for oxygen to attack the carbon fibers, resulting in the loss of carbon fibers and thus a reduction in composite strength and stiffness. Other obstacles are that the material’s stress-strain response is nonlinear even at low stress levels and that the material responds to stress differently in tension than compression. These problems are not unrelated to the carbon oxidation problem. They all arise from the fact that the SiC matrix in C/SiC composites is usually severely cracked in the as-processed state. The nonlinear stress-strain response, the temperature-dependent properties, and the dissimilar response to tension and compression loading all complicate the structural analysis of components made of C/SiC composites.
The objective of this work was to develop analysis methods that incorporate analytical material models into a structural solution routine. In this way, the complex material behavior that is inherent to C/SiC composite materials is accounted for in the structural analysis solution. Micromechanics-based material models were developed at the NASA Glenn Research Center for predicting the response properties of two C/SiC composites: a two-dimensional plain weave and a three-dimensional woven angle interlock architecture. The micromechanics-based material models were calibrated by correlating the predicted material response properties with the measured properties.
Four-point flexure subelement showing loading and strain gauge locations. All dimensions given in inches unless indicated otherwise.
Four-point beam-bending subelement specimens were fabricated with these two fiber architectures, and four-point bending tests were performed at room temperature and at 2000 °F. Displacements and strains were measured at various locations along the beam and recorded as a function of load magnitude. The calibrated material models were used in concert with a nonlinear finite-element solution using the ABAQUS (ABAQUS, Inc., Providence, RI) finite-element code to simulate the structural response of these two materials in the four-point beam-bending tests. The structural response predicted by the nonlinear analysis method, limited to an inert environment in this work, compared favorably with the measured response for both materials and for both test temperatures. Results showed that the material models scale up fairly well from the coupon to the subcomponent level. This approach presents an integrated analysis, material model development, and test approach that is providing the basic understanding necessary to characterize and utilize these materials in various aerospace applications.
Load versus midpoint displacement of a two-dimensional, quasi-isotropic, four-point flexure specimen at room temperature.
Load versus strain at gauges 2, 3, and 5 for a two-dimensional, quasi-isotropic, four-point flexure specimen at room temperature.
Last updated: October 11, 2006
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