Jonathan Poggie

Magnetic control of flow past a blunt body: Numerical validation and exploration

Computational and theoretical studies of a Mach 5 flow over a hemisphere were carried out to validate a new computer code for magnetogasdynamic simulation, and to examine the possibility of heat transfer mitigation through magnetic control. Sample calculations were made using a local solution developed by W. B. Bush [J. Aerosp. Sci. 25, 685 (1958); 28, 610 (1961)] for the stagnation point flow over an axisymmetric blunt body with an imposed dipole magnetic field. Numerical computations, which obviated many of the simplifications inherent in the Bush theory, were carried out employing the low magnetic Reynolds number approximation. Both models indicate that an imposed dipole field can slow the flow in the conductive shock layer and consequently reduce the wall heat flux in the vicinity of the stagnation point. The theoretical model predicts a slightly higher level of heat transfer than that obtained computationally, but there is good agreement between the two models in the fractional change in heat transfer with increasing strength of the applied magnetic field. For both models, nonuniform electrical conductivity was found to reduce the effectiveness of a given applied field. Magnetic flow control is seen to have a sound physical basis, and may prove to be a useful technology for heat transfer mitigation.

Numerical simulation of nanosecond-pulse electrical discharges

Recent experiments with a nanosecond-pulse, dielectric barrier discharge at the stagnation point of a Mach 5 cylinder flow have demonstrated the formation of weak shock waves near the electrode edge, which propagate upstream and perturb the bow shock. This is a promising means of flow control, and understanding the detailed physics of the conversion of electrical energy into gas motion will aid in the design of efficient actuators based on the concept. In this work, a simplified configuration with planar symmetry was chosen as a vehicle to develop a physics-based model of nanosecond-pulse discharges, including realistic air kinetics, electron energy transport, and compressible bulk gas flow. A reduced plasma kinetic model (23 species and 50 processes) was developed to capture the dominant species and reactions for energy storage and thermalization in the discharge. The kinetic model included electronically and vibrationally excited species, and several species of ions and ground state neutrals. The governing equations included the Poisson equation for the electric potential, diffusion equations for each neutral species, conservation equations for each charged species, and mass-averaged conservation equations for the bulk gas flow. The results of calculations with this model highlighted the path of energy transfer in the discharge. At breakdown, the input electrical energy was transformed over a time scale on the order of 1?ns into chemical energy of ions, dissociation products, and vibrationally and electronically excited particles. About 30% of this energy was subsequently thermalized over a time scale of 10??s. Since the thermalization time scale was faster than the acoustic time scale, the heat release led to the formation of weak shock waves originating near the sheath edge, consistent with experimental observations. The computed translational temperature rise (40?K) and nitrogen vibrational temperature rise (370?K) were of the same order of magnitude as experimental measurements (50?K and 500?K, respectively), and the approach appears promising for future multi-dimensional calculations. The effectiveness of flow control actuators based on nanosecond-pulse, dielectric barrier discharges is seen to depend crucially on the rapid thermalization of input energy, in particular the rate of quenching of excited electronic states and the rate of electron?ion recombination.

Traveling Instability Waves in a Mach 8 Flow over an Elliptic Cone

Simultaneous measurements were carried out with three hot-film probes in the Mach 8 flow over an elliptic cone of 2:1 aspect ratio, and the data obtained were compared to the results of computations using the parabolized Navier-Stokes equations and linear stability theory. The elliptic-cone flow was found to be significantly different from the flows studied in previous hypersonic-flow stability experiments, which have focused exclusively on wind-tunnel models with two-dimensional, planar or axial symmetry. At least two instability mechanisms appear to be active in the present flow: one associated with the region of maximum crossflow in the vicinity of the shoulder of the cone and the other associated with the inflectional velocity profiles on the top centerline. Between the shoulder and leading edge of the cone, the dominant flow instability occurred at relatively low frequency, and the direction of the phase velocity was significantly skewed from that of the boundary-layer-edge streamlines. The results were found to be in rough agreement with linear stability calculations and are suggestive of a traveling crossflow instability mode, which apparently has not heen observed before in hypersonic flow

Plasma-sheath transition in the magnetized plasma-wall problem for collisionless ions

A model for the collisionless plasma-wall problem under the action of an applied magnetic field is developed. The behavior of its solution is examined and found to be qualitatively consistent with experiment. The plasma and the sheath are then modeled separately to obtain the position of the quasi-neutral plasma boundary and the position of the edge of the electron-free sheath. It is shown that the plasma boundary can be specified as the point where the component of the ion velocity normal to the wall reaches the ion sound speed (Bohm criterion), and the sheath edge is specified as the point corresponding to Godyaks condition for the electric field. Studying the behavior near the plasma boundary and the sheath edge, the plasma solution and the solution of the space charge region are patched together to approximate the solution of the plasma-wall problem.

Modeling low pressure collisional plasma sheath with space-charge effect

The present work develops a computationally efficient one-dimensional subgrid embedded finite element formulation for plasma-sheath dynamics. The model incorporates space-charge effect throughout the whole plasma and the sheath region using multifluid equations. Secondary electron emission is not considered. A third-order temperature dependent polynomial is used to self-consistently calculate the rate of ionization in the plasma dynamic equations. The applications include dc and rf sheath inside a glow discharge tube where the noble gas is immobile, and a partially ionized plasma sheath inside an electric propulsion thruster channel in which the gas flows. The electron and ion number densities of the numerical solution decrease in the sheath region as expected. The ion velocity and electron temperature profiles also exhibit the expected behavior. The computed sheath potential compares well with the available experimental data.

Spectral Characteristics of Separation Shock Unsteadiness

Spectra of wall-pressure fluctuations caused by separation shock unsteadiness were compared for data obtained from wind-tunnel experiments, the Hypersonic International Flight Research Experimentation flight test 1, and large-eddy simulations. The results were found to be in generally good agreement, despite differences in Mach number and two orders of magnitude difference in Reynolds number. Relatively good agreement was obtained between these spectra and the predictions of a theory developed by Plotkin. The predictions of this theory are also qualitatively consistent with the results of experiments in which the shock motion was synchronized to controlled perturbations. The results presented here support the idea that separation unsteadiness has common features across a broad range of compressible flows and that it behaves as a selective amplifier of large-scale disturbances in the incoming flow.

Shock unsteadiness in a reattaching shear layer

The origin of shock unsteadiness in a Mach 2.9 turbulent reattaching shear layer was investigated experimentally using temporally resolved flow visualization and measurements of wall pressure fluctuations. In this flow, the separation point of a turbulent boundary layer is essentially fixed at a backward-facing step, and the reattachment point is free to move along a ramp. In order to examine the influence of disturbances originating in the incoming shear layer, artificial disturbances were introduced into the flow through steady air injection in the vicinity of separation. The effect on the reattachment shock system was dramatic: the intensity of the pressure fluctuations and the amplitude of the shock motion increased substantially, and power spectra of the pressure fluctuations showed a distinct shift to lower frequency. The spectra collapsed onto a common curve in non-dimensional coordinates based on a length scale derived from two-point cross-correlations of the flow visualization data and a convection velocity derived from cross-correlations of the pressure measurements. The data were compared to a theory developed by Plotkin (1975), which is based on perturbation of a shock by random fluctuations in the incoming turbulent flow. Plotkins model mimics the manner in which relatively broad-band perturbations in the incoming turbulent flow lead to relatively low-frequency motion of the separation bubble and its associated shock system. It is an excellent fit to separation shock motion, such as that generated in a blunt fin flow (briefly illustrated here). In the present shear layer flow, this low-frequency motion was detectable in the spectra near reattachment, but contained considerably less energy relative to the shock motions caused by direct perturbations by the incoming turbulent structures. These results indicate that the shock motion in the reattaching shear layer is primarily caused by organized structures in the incoming turbulent flow.

Closed-Loop Stall Control System

A closed-loop, stall sense and control system was demonstrated on a morphing airfoil. The FlexSys, Inc. Mission Adaptive Compliant Wing was modified to accept a Boeing Co. dielectric barrier discharge actuator panel in a location immediately upstream of the trailing-edge morphing flap, and hot-film sensors were installed on the model surface. A signal analysis algorithm, developed by Tao Systems, Inc., was applied to the hot-film signals to detect separation and trigger activation of the dielectric barrier discharge actuators. The system was successfully demonstrated in the U.S. Air Force Research Laboratory Phillip P. Antonatos Subsonic Aerodynamics Research Laboratory wind-tunnel facility, and an improvement in lift of about 10% was observed at Mach 0.05 (chord Reynolds number 9 x 10 5 ) under closed-loop control and a turbulent boundary-layer state. Actuator effectiveness was demonstrated up to Mach 0.1, but must be extended to Mach 0.2-0.3 to enable a practical stall control system for takeoff and approach of large aircraft. It may be possible to obtain that level of performance by optimizing the actuator locations and input waveforms.

Experimental evidence for Plotkin model of shock unsteadiness in separated flow

Experimental evidence is presented in support of a model of separation shock unsteadiness developed by Plotkin [AIAA J. 13, 1036 (1975)]. Under this model, the position of the separation shock follows linearly damped Brownian motion. The model describes the manner in which relatively broad-band perturbations in the incoming flow lead to relatively low-frequency motion of the separation shock. Close agreement was found between the predictions of the model and the autospectra and autocorrelations of wall pressure fluctuations and shock position fluctuations for several blunt fin flows at Mach 3 and Mach 5. Given the similarity of the power spectra of wall-pressure fluctuations for a variety of separated, supersonic flows, this description may have broad applicability.

Large-scale coherent turbulence structures in a compressible mixing layer flow

Detailed flow visualization experiments were carried out in a self-similar turbulent mixing layer at a nominal convective Mach number of 1.1. The flow visualization technique was base on Rayleigh scattering from nanometer-scale contaminant particles present in the freestream flow. The interface marked by the vaporization of the particles revealed the large-scale organized turbulence structures in the mixing layer. Quantitative measures of the length scale, orientation, and speed of organized structures were derived from the flow visualization data, and were found to agree well with conventional point-probe measurements.