Aerothermal analysis of hypersonic flows using generalized thermo-fluid dynamic equations

Abstract

Accurate prediction of the flowfields and aerothermodynamic performance of air-breathing hypersonic vehicles, space launch vehicles, and airbreathing propulsion systems are critical to their efficient design. Such flowfields are characterized by strong shocks and nonequilibrium processes, and difficult to accurately model and predict. Recently, the authors have derived two thermo-fluid dynamic equations: a generalized equation of state and a finite amplitude wave propagation speed that manifests the nonequilibrium effects (entropy production) and provides the correct eigenvalues or Reimann invariants. These equations will allow more accurate modeling and prediction of the effects of real (thermally imperfect) gas, high temperature (calorically imperfect) and nonequilibrium processes, and the aero thermal characteristics and performance of hypersonic vehicle and propulsion systems. Implementation of this approach in a CFD code was made by calculating a Mach 19.79 sphere-cylinder chemically reacting flowfield test problem. Two full Navier-Stokes solutions were obtained: one with an isothermal wall boundary condition, and the other with an adiabatic wall boundary condition. For the solutions with isothermal wall boundary condition, comparisons of the solutions to those obtained without the generalized equations (equation of state) show dramatic differences in convergence rates of better than a factor of four. For the case with the adiabatic wall boundary, the solution with the generalized equations gave wall temperatures in the stagnation region of the order of 1000AT higher, which is quite significant in terms of aerothermal loads. Thus, for the Mach 19.79 sphere-cylinder test problem, the new generalized thermo-fluid dynamic equations are seen to have significant impact on the obtained solutions.