Donald P. Rizzetta

Numerical Simulation of Supersonic Flow Over a Three-Dimensional Cavity

A numerical solution is presented for the unsteady flow over a three-dimensio nal cavity at a freestream Mach number of 1.5 and Reynolds number of 1.09 x 10 6. The self-sustained oscillatory motion within the cavity is generated numerically by integration of the time-dependent compressible three-dimensional Reynolds averaged Navier-Stokes equations. Effects of fine-scale turbulence are simulated via a simple algebraic closure model. Details of the flowfield structure are elucidated, and it is verified that the fundamental behavior of the unsteady phenomena is two dimensional. Comparison with experimental data is made in terms of the mean static pressure and overall acoustic sound pressure levels within the cavity, as well as with the acoustic frequency spectra of the oscillation along the cavity floor and rear bulkhead.

Numerical simulation of slot injection into a turbulent supersonic stream

Steady flowfields resulting from slot injection at the surface of a flat plate in a freestream with a Mach number of 3.7 and a unit Reynolds number of 5.83 x 10 6 /m were simulated numerically by integration of the time-dependent compressible mass-averaged Navier-Stokes equations. Effects of fine scale turbulence were represented by a two-equation (k-e) closure model that included a generalized formulation, low Reynolds number terms, and a compressibility correction

Numerical simulation of turbulent cylinder juncture flowfields

Steady high-Reynolds-number subsonic and supersonic flowfields about a circular cylinder mounted upright on a flat plate were simulated numerically by integration of the time-dependent three-dimensional compressible mass-averaged Navier-Stokes equations. Effects of turbulence were represented by a two-equation (k-e) closure model that included a generalized formulation and low-Reynolds-number terms. For the supersonic case, the turbulence equations incorporated a compressibility correction. Grid mesh step-size studies were performed to assess resolution requirements of the solutions. Comparison is made with experimental data in terms of static pressure and total pressure loss coefficients, velocity distributions, and surface limiting streamline patterns

Numerical simulation of leading-edge vortex flows

Steady flowfields describing respectively the distinguished structure for subsonic, sonic, and supersonic leading-edge flow about a thin delta wing at angle of attack in a supersonic freestream are calculated numerically. Solutions of the steady three-dimensional compressible laminar Navier-Stokes equations are obtained by time integration. Details of these solutions demonstrate that the essential physical behavior of such flows, including both primary and secondary vortex motions, has been simulated. For purposes of comparison, a corresponding inviscid numerical solution was generated for the case of a subsonic leading edge. It is shown that, although the secondary features are absent, the gross dominant characteristics of the flowfield have been reproduced by the Euler equations. Effects of turbulence are assessed by incorporating a simple closure model in the viscous computation. Comparison between the numerical solutions and experimental data is provided for all flow regimes.

APPLICATION' OF A HIGH-ORDER COMPACT DIFFERENCE SCHEME TO LARGE-EDDY AND DIRECT NUMERICAL SIMULATION

This work investigates the application of a high- order compact difference scheme to a number of representative large-eddy and direct numerical sim- ulations. The scheme employs an implicit approx- imately factored finite-difference algorithm, which is used in conjunction with a tenth-order spatial filter. Newton-like subiterations are applied to achieve second-order temporal and sixth-order spa- tial accuracy. For large-eddy simulations, both the Smagorinsky and dynamic subgrid-scale stress mod- els are incorporated in the computations, and used for comparison along with direct numerical simula- tions. Details of the method are summarized, and a series of classic validating computations are per- formed. These include the decay of compressible isotropic turbulence, turbulent channel flow, and the subsonic flow past a circular cylinder. For each of these cases, comparison is made between the respec- tive computations as well as with available exper- imental data and with previous existing numerical results. It was found for both the channel and cylin- der simulations that the dynamic model provided better overall descriptions of the flowfield than the Smagorinsky model. cylinder lift coefficient cylinder pressure coefficient Smagorinsky model constant total specific energy three-dimensional energy spec-

Evaluation of Explicit Algebraic Reynolds-Stress Models for Separated Supersonic Flows

High-Reynolds-number supersonic e owe elds were generated numerically to assess the performance of three explicit algebraic Reynolds-stress turbulence models. The cone gurations consist of a shock/boundary-layer interaction and a 24-deg compression ramp, both of which exhibit an appreciable region of separated e ow. Solutions were also obtained using standard zero-equation and k‐ models. Details of the computations are summarized, and the accuracy of numerical results is established via grid resolution studies. Comparisons are made with experimental data in terms of surface pressure and skin friction, as well as off-surface proe les of mean velocity and componentsoftheReynolds-stresstensor. Forthee owsconsidered here, it isfound that thealgebraic-stressmodels offer little improvement over existing closures.

Numerical Simulation of Vortex-Induced Oblique Shock-Wave Distortion

We simulate oblique shock-wave/wing-tip vortex interactions via solution to the compressible three-dimensional mass-averaged turbulent Navier-Stokes equations. A two-equation (k-e) turbulence model is employed that includes low-Reynolds terms and a compressibility correction.

Numerical simulation of oblique shock-wave/vortex interaction

High Reynolds number supersonic flowfields were generated numerically by integration of the time-dependent three-dimensional compressible Euler and mass-averaged Navier-Stokes equations in order to simulate the interaction of a streamwise vortex with an oblique shock wave. The vortex develops at the tip of a vortex-generating fin which is suspended from the ceiling of a wind tunnel. Downstream, an oblique shock wave is produced by a wing surface which spans the width of the wind tunnel and has a two-dimensional sharp leading-edge wedge as its airfoil section. Grid resolution studies are provided in order to assess accuracy of the solutions. Resultant features of the numerical flowfields are discussed ,and comparison is made with experimental data in terms of static pressure distributions on the wing surface.

Numerical simulation of the Navier-Stokes equations for an F-16A configuration

The results of a numerical simulation are presented for the steady flow over an F-16A aircraft configuration at a freestream Mach number of 1.2, a Reynolds number of 12.75 x 10, and an angle of attack of 6 deg. The three-dimensional Navier-Stokes equations in mass-averaged variables were integrated numerically using the MacCormack explicit algorithm with an algebraic turbulence model to provide closure of the system of equations. The grid structure, boundary conditions, and solution procedure are discussed in detail for this complex aircraft geometry. The results are then compared to experimental data in terms of surface pressure coefficients, lift coefficient, and drag coefficient with reasonable agreement. Finally, details of the flow are discussed, such as the strake vortex and the wing vortex structures.

Exploration of Plasma Control for Supersonic Turbulent Flow over a Compression Ramp

The Navier-Stokes equations are solved using a high-fidelity time-implicit numerical scheme and an implicit large-eddy simulation approach to investigate plasmabased flow control for supersonic flow over a compression ramp. The configuration includes a flat-plate region to develop an equilibrium turbulent boundary layer at Mach 2.25, which is validated against a set of experimental measurements. The fully turbulent boundary-layer flow interacts with a 24◦ ramp and produces an unsteady shock-induced separation. A control strategy to suppress the separation through a magnetically-driven gliding-arc actuator is explored. The size, strength, and placement of the actuator are developed based on recent experiments. Three control scenarios were examined: steady control, pulsing with a 50% duty cycle, and Joule heating. The results show the control mechanism is very effective at reducing the mean separation length for all three situations. The case without pulsing and Joule heating was the most effective, with a reduction in the separation length by more than 75%. Control was also found to significantly reduce the low-frequency content of the turbulent kinetic energy spectrum.