C. C. Horstman

Computation of crossing shock/turbulent boundary layer interaction at Mach 8.3

A three-dimensional (3D) hypersonic crossing shock wave/turbulent boundary-layer interaction is examined numerically at Mach 8.3. The test geometry consists of a pair of opposing sharp fins of angle alpha = 15 deg, mounted on a flat plate. Two theoretical models are evaluated. The full (3D) Reynolds-averaged Navier-Stokes equations are solved using the Baldwin-Lomax and the Rodi (modified k-epsilon) turbulence models. Computed results for both cases show good agreement with experiment for flat plate surface pressure and for flowfield profiles of pitot pressure and yaw angle, indicating that the flowfield is primarily rotational and inviscid. Fair to poor agreement is obtained for surface heat transfer, indicating a need for more accurate turbulence models. The overall flowfield structure is similar to that observed in previous crossing shock interaction studies.

Three-dimensional shock wave-turbulent boundary layer interactions generated by a sharp fin at Mach 4

This paper describes a combined experimental and theoretical study of three-dimensional swept shock wave-turbulent boundary layer interactions at Mach 4 generated by a sharp fin of angles alpha equals 16 and 20 degrees. The theoretical model is the three-dimensional compressible Reynolds-averaged Navier-Stokes equations with turbulence incorporated through the algebraic eddy viscosity model of Baldwin and Lomax. Previous computations have been performed by Horstman using the Baldwin-Lomax, Cebeci-Smith and Jones Launder models. Computed results for the surface pressure, skin friction and streamline angles are compared with experiment and previous numerical results. The present results display good agreement with experimental data for surface pressure and surface flow direction. All turbulence models fail to accurately predict the peak skin friction. The computed flowfields are in agreement with many of the features of the quasi-conical flowfield model of Settles.

Measurements of the Triple Shock Wave/Turbulent Boundary-Layer Interaction

A joint experimental and computational study has investigated the flowfield structure created by three crossing oblique shock waves interacting with a turbulent boundary layer. Such an interaction is of practical importance in the design of high-speed sidewall-compression inlets. The interaction is created by a test model consisting of two vertical sharp fins mounted at 15-deg angle of attack to a horizontal flat plate. A third compression surface of 10-deg angle is mounted to the plate between the two vertical fins. The Mach 3.85 flowfield is examined experimentally through kerosene lampblack and planar laser scattering visualizations. The results are used to develop a flowfield model of the interaction structure. This structure is found to consist of a complex wave pattern overlying a viscous separated region. The presence of the 10-deg compression ramp reduces the severity of the boundary-layer separation compared to interactions created by the two fins alone. Additionally, the experimental flowfield data are compared with a computational solution using a modified κ-e (Rodi) turbulence model.

Flowfield surveys and computations of a crossing-shock wave/boundary-layer interaction

A joint experimental and computational study has been performed to investigate the flowfield structure created by two crossing oblique shock waves interacting with a turbulent boundary layer. Such an interaction is of practical importance in the design of high-speed sidewall-compression inlets. The interaction is created by a test model, consisting of two sharp fins mounted at 15-deg angle of attack to a flat plate, placed in a Mach 3.85 freestream flow with a unit Reynolds number of 76 X 106/m. Two computational solutions, one using a Baldwin-Lomax algebraic turbulent eddy viscosity model and one using a modified K-£ (Rodi) turbulence model, are compared with experimental flowfield data obtained from a fast-response five-hole probe. Both the experiment and the computations show that the flowfield is dominated by a large, low-Mach-number, low-total-pressure separated region located on the interaction centerline. A comparison of the results shows significant differences between experiment and computations within this separated region. Outside the separated region, the experiment and computations are in good agreement. Additionally, the comparison shows that both turbulence models provide similar results, with neither model being clearly superior in predicting the flowfield.

Investigation of a hypersonic crossing shock wave/turbulent boundary layer interaction

A combined theoretical and experimental study is presented for the interaction between crossing shock waves generated by (10°, 10°) sharp fins and a flat plate turbulent boundary layer at Mach 8.3. The theoretical model is the full 3-D mean compressible Reynolds-averaged Navier-Stokes RANS) equations incorporating the algebraic turbulent eddy viscosity model of Baldwin and Lomax. A grid refinement study indicated that adequate resolution of the flowfield has been achieved. Computed results agree well with experiment for surface pressure and surface flow patterns and for pitot pressure and yaw angle profiles in the flowfield. The computations, however, significantly overpredict surface heat transfer. Analysis of the computed flowfield results indicates the formation of complex streamline and wave structures within the interaction region.

Supersonic turbulent flow past a 3-D swept compression corner at Mach 3

Experimental and theoretical studies are presented on the three-dimensional shock wave-turbulent boundary layer interaction generated by a swept compression corner at Mach 3 for compression angle of 24 deg, sweep angle of 60 deg, and Reynolds numbers from 140,000 to 900,000. Two theoretical approaches were used, both of which utilize the full mass-averaged compressible three-dimensional Navier-Stokes equations but differ in the choice of turbulence model (the Baldwin-Lomax, 1978, and the Jones-Launder, 1972, model, respectively). The features of the computed mean flow structure were found to be qualitatively the same for both the Baldwin-Lomax and Jones-Launder models.

The hypersonic shock wave-turbulent boundary layer interaction generated by a sharp fin at Mach 8.2

A combined experimental and numerical study has been conducted on the hypersonic shock-wave turbulent-boundary layer interaction at Mach 8.2 generated by a single fin of angles alpha = 10 and 15 deg. Three models are considered: (1) the 3D compressible Reynolds-averaged Navier-Stokes (RANS) equations using the k-epsilon turbulence model, (2) the 3D RANS using the Rodi turbulence model, and (3) the conical RANS using the Baldwin-Lomax algebraic turbulence model. The computations are compared with various experimental data. The computations using models (1) and (2) show quantitatively very similar results and very good agreement with experimental data for surface pressure and skin friction. Comparison with boundary layer profiles of pitot pressure and yaw angle are also generally good, but the peak surface heat transfer is overestimated by up to 48 percent. The effect of the laminar boundary layer on the fin is restricted to the immediate vicinity of the fin surface. Conical calculations using model (3) show substantially poorer agreement with experiment.

Crossing shock wave-turbulent boundary layer interactions

Three-dimensional interactions between crossing shock waves generated by symmetric sharp fins and a turbulent boundary layer on a flat plate are investigated experimentally and theoretically at Mach number 2.95 and freestream unit Reynolds number 1.96 x 10 to the 7th/ft. The incoming boundary layer has a thickness of 4 mm at the location of the fin leading edges. A comparison of experimental and computational results for two sets of fin angles (11 x 11 and 9 x 9 deg) shows general agreement with regard to surface pressure measurements and surface streamline patterns. The principal feature of the streamline structure is a collision of counterrotating vortical structures emanating from near the fin leading edges and meeting at the geometric centerline of the interaction.

Structure of supersonic turbulent flow past a swept compression corner

The structure of the shock wave/turbulent boundary-layer interaction generated by a 3D swept compression corner has been investigated through a combined experimental and theoretical research program. The flowfield geometry is defined by the streamwise compression angle alpha and the sweep angle lambda of the corner. The present study examines two different configurations, namely (alpha, lambda) = (24 deg, 40 deg) and (24 deg, 60 deg) at Mach 3 and Re sigma infinity about 9 x 10 exp 5. The theoretical model is the 3D Reynolds-averaged compressible Navier-Stokes equations with turbulence incorporated using a turbulent eddy viscosity. The calculated flowfields display general agreement with experimental data for surface pressure and good agreement with experimental flowfield profiles of pitot pressure and yaw angle. The principal feature of the flowfield is a large vortical structure approximately aligned with the corner. The entrainment of incoming fluid into the vortical structure is strongly affected by the sweep angle lambda. Viscous (turbulent and molecular) effects appear to be important only in the immediate vicinity of the surface and in an isolated region within the interaction and near the corner.