Modeling the heat and mass transfer in the pores of the thermal protection carbon-carbon frame during the gas-phase deposition of silicon carbide

Sergey V Reznik,Konstantin V Mikhailovskiy,Pavel V Prosuntsov,G.V Kuznetsov,P.A Strizhak,A.O Zhdanova

doi:10.1051/matecconf/20179201075

Sergey V Reznik, Konstantin V Mikhailovskiy + Show 4 more

Open Access

https://doi.org/10.1051/matecconf/20179201075

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Abstract

This paper is dedicated to the numerical simulation of the gas-phase deposition into the porous carbon-carbon frame for the gradient thermal protection of reentry vehicles. The paper presents the specifics of creating the representative volume elements of the porous carbon-carbon frames based on the microstructure data from the computed tomography scanner, electronic scanning microscope and porosimetry results. The finite-element models of the representative volume elements are created in order to obtain the net independent solution, including the special finite elements at the phase interfaces. The finite-element and finite-volume method was used for calculation, implemented in MSC.Digimat and ANSYS software packages. The effect of the reagent parameters on the gas-phase deposition process, uniformity and rate of silicon carbide deposition was simulated parametrically, and the results are presented.

Highlights

The rise in the international space programs and the development of new generation reentry space vehicles are calling for the design of heat-resistant thermal protection
The drawbacks of the ablative thermal protection coating (TPC) include the physical-chemical transformation in composite materials (CM) resulting in irreversible changes in the aerodynamic and thermal protection properties affecting their reusability
The following assumptions are introduced for the mathematical simulation of the deposition process: the reaction medium is optically transparent; the radiative heat transfer surfaces are grey; the surfaces are diffusely radiating and diffusely absorbing; the changes in the thermal physical properties in the representative volume elements of the porous ceramic composite materials (CCCM) are not taken into account when the SiC matrix is formed

Summary

Introduction

The rise in the international space programs and the development of new generation reentry space vehicles are calling for the design of heat-resistant thermal protection. The heat transfer through the coating is reduced by the formation of the porous frame and the increase in the radiative ability of the surface. The drawbacks of the ablative TPC include the physical-chemical transformation in CM resulting in irreversible changes in the aerodynamic and thermal protection properties affecting their reusability. Technological and economic reasons, reusable large spacecraft do not employ ablative TPC, since they alter their shape, dimensions in the operation process, and lose their heat-insulating ability as a result of physical chemical transformations. The main operation in the TPC parts fabrication process is the gas-phase deposition of silicon carbide into the porous carbon-carbon frame lasting hundreds of hours and determining the final physical mechanical, thermal physical and optical properties consistent with the operation requirements. The problems were solved in a one-dimensional or two-dimensional setting, the porous medium was viewed as a set of cylindrical pores, and the radiative heat transfer in the pores was not taken into account

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