Direct numerical simulation is performed on a 38.1% scale Hypersonic International Flight Research Experimentation Program Flight 5 forebody to study stationary crossflow instability. Computations use the US3D Navier-Stokes solver to simulate Mach 6 flow at Reynolds numbers of 8.1 × 106/m and 11.8 × 106/m, which are conditions used by quiet-tunnel experiments at Purdue University. Distributed roughness with point-to-point height variation on the computational grid and maximum heights of 0.5-4.0 μm is used with the intent to emulate smoothbody transition and excite the naturally occurring most unstable disturbance wavenumber. Cases at the low-Reynolds number condition use three grid sizes, and hence three different roughness patterns, and demonstrate that the exact flow solution is dependent on the particular roughness pattern. The same roughness pattern is interpolated onto each grid, which yields similar solutions, indicating grid convergence. A steady physical mechanism is introduced for the sharp increase in wall heat flux seen in both computations and experiment at the high-Reynolds number condition. Evolution of disturbance spanwise wavelength is computed, and is found to be more sensitive to Reynolds number than roughness, indicating that the disturbance wavelength is primarily the naturally occurring, flow-selected wavelength for these cases.
Bibliographical noteFunding Information:
This work was sponsored by the Air Force Office of Scientific Research under grant FA9550-12-1-0064. The views and conclusions contained herein are those of the authors and should not be interpreted as necessarily representing the official policies or endorsements, either expressed or implied, of the AFOSR or the U.S. Government.