Modeling of coupled aerogasdynamics and heat transfer processes on the thermal protection surface of a future hypersonic aircraft
To create future high-speed hypersonic aircrafts, it is necessary to predict with sufficient accuracy heat flows and temperature fields on the surface of the aircraft caused by aerodynamic heating. The surface temperature is currently calculated in two steps: the calculation of the aerogasdynamic flow parameters of a “cold” surface and the calculation of the surface temperature of the real structure. The method that is currently used for calculating the surface temperature of hypersonic vehicles can give significant errors because of the intensive aerodynamic heating of ablative beat shield materials throwing the multicomponent gas mixture into the incident flow. In this paper, a new numerical method for solving the coupled problem of aerogasdynamics and internal heat transfer in future hypersonic aircraft structures is proposed. The method implies an iterative solution of the following three independent problems: ideal gas dynamics, viscous gas dynamics on the basis of Navier-Stokes equations for a three-dimensional boundary layer, and heat conduction in a hypersonic aircraft shell. Algorithms for the numerical analysis of these problems in non-orthogonal curvilinear coordinates are developed. The TVD-type finite difference schemes were modified to be applied to geometrically adaptive grids. The results of simulation of the flow about an ellipsoidal nose segment of a model hypersonic aircraft are presented. The results of calculation of the adiabatic wall temperature in the case of heat transfer between the gas and the wall showed that the heat transfer must be taken into account when calculating the temperature distribution on the surface of a hypersonic aircraft. The presented numerical method can be applied to the analysis of aerodynamics of hypersonic aircrafts.