Numerical Investigation of Flow Pattern Variations Over a 60-deg delta Wing from Subsonic to Hypersonic

Steady flows over a 60º delta wing with sharp leading edge are numerically studied at different sound regimes. In the numerical analysis using the Fluent software, an unstructured grid, Roe’s upwind scheme, k-ω SST turbulence model, implicit solution and parallel processing are utilized. Visualization results obtained by numerical simulations show that a leading edge vortex is formed on the upper side of the wing with or without the presence of the secondary separation which gradually expands and becomes closer to the wing surface with increasing Mach number. At Mach numbers higher than 2, the leading edge becomes supersonic and the expansion wave emanating from the leading edge accelerates the flow. At this condition, shock waves are formed on the wing which interact with the vortices. With increasing Mach number up to the hypersonic region, the shock waves above the primary vortex disappear and the vortex size and strength are decreased due to the severe expansion of the flow over the lee-surface of the wing. The effect of Mach number on the vortex core position, lift coefficient, surface pressure distribution, and vortex breakdown location is investigated. Increasing Mach number causes that aerodynamic coefficients increase in subsonic range and decrease in supersonic range. Also, vortex breakdown location moves downstream, and vortex core location gets closer to the wing surface until M = 3 and gets far from it at higher Mach numbers.