The leading-edge receptivity to acoustic disturbances of supersonic/hypersonic boundary layers on a cylinder-wedge of 20° half-wedge angle and 0.1 mm nose radius is numerically investigated for a set of six different cases with Mach number ranging from 3.0 to 7.3, through direct numerical simulation (DNS) of the two-dimensional (2D) Navier-Stokes equations. Two angles of attack (0°, 10°), and two inclination angles of the acoustic waves (0°, 10°) are considered among the different numerical cases. For the Mach 3.0 case both fast and slow planar acoustic waves with multiple frequencies are inserted into the flow-field of the steady state solution in order to carry out unsteady computations, while for the remaining cases the unsteady computations are performed only for fast waves in the freestream. The results show that the response along the wall is stable in the nose region, up to 400 nose radii downstream, and that at Mach 3.0 there is a higher amplitude for the fast mode than for the slow mode. The wall pressure and heat flux perturbation spectra show that the receptivity is higher at the higher Mach numbers and at the higher frequencies, while for the lower Mach numbers (Mach 6.0 and 3.0) a frequency-dependent oscillatory behaviour is shown by the pressure perturbation distribution along the wall. The angle of incidence of the acoustic waves seems to slightly increase the amplitude of the response along the wall on the top (lee) side of the body at the higher frequencies, and to produce a flatter response on the same side for the lower frequencies. Including an angle of attack decreases the receptivity along the top side, as the shock is weaker here, and amplifies the response on the windward side. These results serve to quantify the relationship between the disturbances measured by sensors near the leading-edge of a measurement probe and the freestream disturbances in hypersonic wind tunnels.
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