Computational analysis of viscous hypersonic flow over delta wings
Numerical simulations for the laminar hypersonic flow past blunt edged delta wings have been performed by solving the Navier-Stokes equations for four different angles of attack (0°, 15°, 25°, and 30°) and flow conditions M ∞ = 7.15, Re c = 5.85 × 106, and for three different chord Reynolds numbers (5.625 × 105, 5.625 × 106, and 10.0 × 106), and flow conditions M ∞ = 8.7, α = 30°. It is known that for blunt edged delta wings at high angles of attack, the leeside hypersonic flow is dominated by a shear layer that separates just past the blunt leading edge and forms a more distributed vortical region over the wing, rather than a concentrated vortex structure as observed at lower speeds. Our calculations here show that the size of the vortical regions increases with increasing angle of attack, and as the Reynolds number increases, the character of the separation on the leeside of the wing changes from a primary separation inboard to one closer to the leading edge but with the addition of a secondary separation near midspan. The shape of the wing apex determines the location where the separation begins. The solutions are analyzed and compared with available experimental data.
KeywordsShear Layer Hypersonic Flow Reynolds Number Increase Delta Wing Stanton Number
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