Abstract

Re-entry of a spacecraft occurs at the hypersonic regime where the flow field is extremely complex: high temperature gradients occurring in the shock-layer region ionize and dissociate the air. Even if a large portion of heat generated during this process is convected away in the surrounding air, a fraction of it is still transferred to the vehicle. Therefore, it is important to protect the vehicle with a suitable kind of shielding. Of the many techniques available today, use of ablative material is gaining popularity. The basic idea behind an ablating heat shield is that the energy incident on the spacecraft is used to vaporized the material, thus preventing a significant part of the heat to be transferred into the structure. The available literature indicates that most of the past investigations either do not consider the actual physical processes taking place during ablation, or are limited to a one-dimensional model. The present investigation shows the development of a numerical model for simulating the multi-dimensional heat transfer phenomena that occurred in a typical ablative TPS. The newly developed model is verified using closed form analytical solutions and validated with available data. This effort consists of the first steps of an ongoing project to develop a comprehensive multi-scale, multi-physics and multi-dimensional material response code aimed at modeling charring and surface ablators.

Document Type

Conference Proceeding

Publication Date

6-2012

Notes/Citation Information

Published in the Proceedings of the 43rd AIAA Thermophysics Conference, Paper 2012-2748, p. 1-13.

Copyright © 2012 by Haoyue Weng, Huaibao Zhang, Ovais U. Khan, and Alexandre Martin.

The copyright holders have granted the permission for posting the article here.

Digital Object Identifier (DOI)

http://dx.doi.org/10.2514/6.2012-2748

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