| タイトル | An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design |
| 本文（外部サイト） | http://hdl.handle.net/2060/20080013535 |
| 著者（英） | Dec, John A.; Braun, Robert D. |
| 著者所属（英） | NASA Langley Research Center |
| 発行日 | 2006-01-09 |
| 言語 | eng |
| 内容記述 | A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds. |
| NASA分類 | Composite Materials |
| レポートNO | AIAA Paper 2006-0780 |
| 権利 | Copyright, Distribution as joint owner in the copyright |
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