Leading Edge Bluntness Effects on Shock Wave/Boundary Layer Interactions Near a Convex Corner

Abstract

The cowl shock/ingested forebody boundary layer interaction is usually encountered near the hypersonic inlet shoulder, a convex corner, and sometimes causes the inlet to unstart. The bluntness of the forebody leading edge also affects the downstream convex corner flow, and that should be carefully considered to improve the inlet performance. A twodimensional inlet with a sharp leading edge and a series of blunt leading edges at Mach 5 were numerically studied. The relative positions of the cowl shock and the convex corner were changed to reveal the characteristics of the shock wave/boundary layer interactions with the corner expansion wave interference. Flow patterns and the distributions of surface pressure, heat flux, and friction coefficient involved were thoroughly studied. The results indicate that the boundary layer is even thicker and the flow is easier to separate when the bluntness of the leading edge increases. The variation trends of the separation extent with a sharp leading edge agree reasonably well with the estimations on the basis of inviscid analysis. However, the inviscid formulas established in the sharp leading edge case are not suitable for the blunt leading edge cases. With the aid of viscous simulations, it is found that the upstream effect of the corner extend a little farther and the dimensionless relative position range that can largely reduce the separation narrows with the increasing of the leading edge bluntness. To take into account the nonuniform flow in the boundary layer, a rapid calculation method of the incident shock (B. L.) was established. It is shown that the shock (B. L.) impinges on the corner can minimize the separation in the sharp leading edge case. However, the shock (B. L.) has to move downstream to minimize the separation when the leading edge is blunt. NomenclatureC * = Chapman-Rubensin constant, (µ/µ∞)(T∞/T) CD = drag coefficient of a hemicylinder cp = specific heat capacity of air at constant pressure d1 = distance between the convex corner and the first demarcation point of the inviscid shock impingement d2, d2', d2" = distance between the convex corner and the second demarcation point of the inviscid shock impingement d3 = distance between the convex corner and the third demarcation point of the inviscid shock impingement dn = diameter of the blunt leading edge Ls, Lsx = separation bubble length along the wall and along the x-axis, respectively Ma = Mach number M1, M2 = Mach number upstream and downstream of the incident shock, respectively M3 = Mach number downstream of the reflected shock at the flat plate Mc = Mach number downstream of the corner expansion M∞ = freestream Mach number p∞ = freestream static pressure q = heat flux rate Rn = radius of the blunt leading edge Re = Reynolds number 2 St = Stanton number, St = q/[ρ∞U∞cp (T0-Tw)] T0, T∞, Tw = Total temperature, freestream temperature, and wall temperature, respectively U∞ = freestream flow velocity x = x-axis coordinate xc = distance from the convex corner to the leading edge of the flat plate xi, xr,...

Modelling and shock control for a V-shaped blunt leading edge

“…The IS may cause a strong SWBLI on the impingement. Many studies (White & Ault 1996;Li & Yang 2015;Sathia Narayanan & Verma 2015;Zhang et al 2021b) on the SWBLI near an expansion corner show that if the IS impinges on the EC, the SWBLI-induced separation bubble will not be eliminated. The subsonic layer can propagate the adverse pressure gradient forwards.…”