A thermal computational analysis for the composite structure of a CubeSat is presented. The main purpose of this investigation is to study the thermal performance of carbon fibre/epoxy resin composite materials with Zinc Oxide nanoparticles in order to be used in the panels of the primary structure of a CubeSat. The radiative heat fluxes over each composite panel are computed according to the orbit trajectory and they are utilized as boundary conditions for the analysis. The direct solar, albedo and Earth infrared radiation fluxes are considered in this study. The model implementation, including the computation of the orthotropic thermal conductivity of the composite material is presented. The thermal simulations were performed for three different orbit inclination angles: the selected mission (
), the worst hot (
) and the worst cold (
). The temperature ranges in the electronic boards are analyzed in order to show that are into the operating limits of each electronic component.
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