Shock-Shock Interaction over a Hemisphere in Hypersonic Flow (original) (raw)

Simulation of Hypersonic Shock-Shock Interaction over a Hemisphere

AIAA 2018-3708, 2018

The interaction of an oblique shock with the bow shock in front of a hemisphere in Mach 14.6 flow is examined. The hypersonic laminar simulation is performed using a C ++ code with the implementation of MPI written by the authors. The simulation is based on a set of experiments performed at the 48-inch Shock Tunnel at Calspan University at Buffalo Research Center (CUBRC). The Edney III interaction exhibits a statistically stationary behavior with dominant Strouhal number of 0.346. The surface pressure coefficient and di-mensionless heat transfer are in general agreement with experiment; however, the location of the interaction of the incident shock with the bow shock is evidently slightly different than in the experiment. More calculations are in progress.

Numerical Prediction of Laminar Shock/Shock Interactions in Hypersonic Flow

Journal of Spacecraft and Rockets, 2003

Results are presented of a series of numerical simulations of experiments conducted at the ONERA Chalais-Meudon Research Center and at the Calspan—University at Buffalo Research Center on shock/shock interactions. The flowfield characteristics are described with the aid of the numerical predictions, and the computed values of surface pressure and heat transfer are compared with the experimental measurements for purposes of verification and validation.Issues related to boundaryconditions, grid convergence and time unsteadiness of the computational fluid dynamics simulations are addressed, and the difficulties that characterize the validation of the computational results are put in evidence.

Numerical study of unsteady viscous hypersonic blunt body flows with an impinging shock

Lecture Notes in Physics, 2000

A complex two-dimensional, unsteady, viscous hypersonic shock wave interaction is numerically simulated by a high-resolution, second-order fully implicit shock-capturing scheme. The physical model consists of a non-stationary oblique shock impinging on the bow shock of a blunt body. Studies indicate that the unsteady flow patterns are slightly different from their steady counterparts. However, for the sample cases investigated the peak surface pressures for the unsteady flows seem to occur at very different impingement locations than for the steady flow cases.

Transitional shock-wave/boundary-layer interactions in hypersonic flow

Journal of Fluid Mechanics, 2014

Strong interactions of shock waves with boundary layers lead to flow separations and enhanced heat transfer rates. When the approaching boundary layer is hypersonic and transitional the problem is particularly challenging and more reliable data is required in order to assess changes in the flow and the surface heat transfer, and to develop simplified models. The present contribution compares results for transitional interactions on a flat plate at Mach 6 from three different experimental facilities using the same instrumented plate insert. The facilities consist of a Ludwieg tube (RWG), an open-jet wind tunnel (H2K) and a high-enthalpy free-piston-driven reflected shock tunnel (HEG). The experimental measurements include shadowgraph and infrared thermography as well as heat transfer and pressure sensors. Direct numerical simulations (DNS) are carried out to compare with selected experimental flow conditions. The combined approach allows an assessment of the effects of unit Reynolds ...

Numerical Study of Shock-Boundary Layer Interaction in Hypersonic Flow

The aim of the present work is to access the ability of the three-dimensional RANS code MB-EURANIUM is predicting shock wave boundary-layer interaction. Results are presented for two cases in this paper. The first is the turbulent boundary-layer shock interaction on a twodimensional ramp. The second case is that of laminar boundary-layer shock interaction over a blunted cone-flare.

Shock Wave/Transitional Boundary-Layer Interactions in Hypersonic Flow

AIAA Journal, 2006

The experimental and numerical transitional interactions in hypersonic flow are studied. The experiments were performed on a hollow cylinder-flare model in the ONERA R2Ch wind tunnel at a Mach number of 5 and for varying stagnation pressure. Wall pressure and heat-flux measurements, laser Doppler velocimetry, pitot boundary-layer surveys, surface flow visualizations, and schlieren photographs provide a precise and complete description of the flowfield. In all of the cases examined here, grid-converged axisymmetric mathematical solutions of the problem were obtained by use of the two-dimensional numerical simulation, but it was found that these solutions do not fit experiments when the Reynolds number is increased. A purely three-dimensional organization of the flow then appears, characterized by the Görtler vortices. Two families of solutions were thus evidenced, and the precise calculation of the physical one remains a numerical challenge. The prediction of transition by use of stability calculations is only partly possible because the waves used do not have a sufficiently general form to model such a complex physical problem. New information on the true nature of what is commonly called a transition mechanism in this kind of flow is deduced from these results.

THREE-DIMENSIONAL STEADY SHOCK WAVE INTERACTIONS. NUMERICAL SIMULATIONS AND EXPERIMENTAL VALIDATION

Numerical simulations and wind tunnel experiments have been performed to investigate 3D shock configurations in the supersonic flow around two symmetrical finitewidth wedges. Regular and irregular (Mach) interactions of shock waves have been studied and a number of unexpected details have been revealed. The existence of a new type of 3D shock interaction, combined reflection, has been numerically predicted. Experiments performed with the laser sheet visualization technique have confirmed all these numerical findings. Close qualitative and quantitative agreement between the results of computations and experiments has been established.

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

AIAA Journal, 1993

A three-dimensional 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 a = 15 deg mounted on a flat plate. Two theoretical models are evaluated. The full three-dimensional Reynolds-averaged Navier-Stokes equations are solved using the Baldwin-Lomax and the Rodi (modified A>e) 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.

Simulation of Hypersonic Shock/Turbulent Boundary-Layer Interactions Using Shock-Unsteadiness Model

In hypersonic flows, the interaction of a shock wave with a turbulent boundary layer can result in flow separation and high aerothermal loads. In this paper, cone–flare configurations with different flare angles and freestream Mach numbers are simulated using Reynolds-averaged Navier–Stokes method, and results are compared with experimental data. The standard Spalart–Allmaras and k-!turbulence models do not predict flow separation at the cone–flare junction, and therefore yield a large deviation from the surface pressure measurements. Sinha et al. (“Modeling Shock-Unsteadiness in Shock/Turbulence Interaction,” Physics of Fluids, Vol. 15, No. 8, 2003, pp. 2290– 2297) proposed a shock-unsteadiness model to account for the effect of unsteady shock motion in a steady mean flow. The shock-unsteadiness correction damps turbulence amplification at the shock and results in significant improvement in predicting flow separation and reattachment. The flow topology in the interaction region, in terms of the pattern of shocks and expansion waves, is predicted correctly by the modified turbulence models. The resulting surface pressure distribution matches experimental data well.

Experimental investigation of shock–shock interactions with variable inflow Mach number

Shock Waves, 2021

Experiments on shock-shock interactions were conducted in a transonic-supersonic wind tunnel with variable free-stream Mach number functionality. Transition between the regular interaction (RI) and the Mach interaction (MI) was induced by variation of the free-steam Mach number for a fixed interaction geometry, as opposed to most previous studies where the shock generator angles are varied at constant Mach number. In this paper, we present a systematic flow-based postprocessing methodology of schlieren data that enables an accurate tracking of the evolving shock system including the precise and reproducible detection of RI MI transition. In line with previous experimental studies dealing with noisy freestream environments, transition hysteresis was not observed. However, we show that establishing accurate values of the flow deflections besides the Mach number is crucial to achieve experimental agreement with the von Neumann criterion, since measured flow deflections deviated significantly, up to 1.2 • , from nominal wedge angles. We also report a study conducted with a focusing schlieren system with variable focal plane that supported the image processing by providing insights into the three-dimensional side-wall effects integrated in the schlieren images.