This paper presents results from a finite element micromechanics analysis of thermally induced stresses in composites at cryogenic temperatures typical of spacecraft operating environments. The influence of microstructural geometry, constituent and interphase properties, and laminate orientation were investigated. Stress field information was used to predict damage initiation temperatures and locations. The results indicated that significant matrix stresses occur in composites exposed to typical spacecraft thermal excursions. These matrix stresses varied with laminate orientation and circumferential position around the fiber. The fiber thermoelastic properties had a minimal effect on the overall magnitudes of these stresses. The major difference in the predicted response of unidirectional and multidirectional laminates was the presence of tensile radial stresses, at the fiber/matrix interface, in multidirectional laminates with off-axis ply angles greater than 15 deg. Damage initiation predictions were made by comparing this radial interfacial stress component with a radial fiber/matrix interface strength. The predicted damage initiation temperatures and modes were in good agreement with experimental data for both low (207 GPa) and high (517 GPa) modulus carbon fiber/epoxy composites. Including an interphase region with tailored properties in the analysis reduced the magnitude of some of the matrix stress components, but had only a small effect on the interfacial radial stress.
Micromechanics thermal stress analysis of composite for space structure applications
Mikromechanische Analyse von Wärmespannungen an Verbundwerkstoffen für Anwendungen in Raumfahrtstrukturen
1991
12 Seiten, 12 Bilder, 3 Tabellen, 14 Quellen
Conference paper
English
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