Experimental and Computational Study of the Vortical Flow over a Delta Wing at High Angles of Attack
The high speed flow over a sharp-edged plane delta wing with 65° sweep has been investigated experimentally as well as computationally. The experimental program contains surface pressure measurements, oil flow and Schlieren visualization at Mach numbers from 0.7 to 0.9 and angles of attack up to 20°. Attention was paid to the presence of embedded shocks, vortex breakdown and the interaction between them. Special emphasis is given to the case with a free stream Mach number of 0.85. At this Mach number and angles of attack of 15° and higher, a shock is observed across the wing symmetry plane. Beyond 18° angle of attack the breakdown of both leading edge vortices is observed going along with a double shock system across the symmetry plane: one at about 50% chord position and one at 80%. Downstream of the first shock a strong expansion is present.
In the numerical program a 3D Euler code is used of a flux-difference-splitting type upwind scheme. The solution of the discretized equations is achieved by an efficient non linear multigrid procedure. The computational results reveal the existence of a conical shock between the vortex core and the wing surface. At an angle of attack of 18.0° numerical vortex breakdown is observed, because at that incidence the axial velocity along the vortex core is found to be zero at 121% chord position. Beyond 18.0° no steady solutions could be achieved. Since the computational angle of attack for onset of vortex breakdown is very close to the experimental value of 18.3°, at which vortex breakdown is observed at 70% chord position, a tentative conclusion may be that Euler methods are able to predict the onset of vortex breakdown.
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