Stephen P. Wilkinson

Electrohydrodynamic Flow Control with a Glow-Discharge Surface Plasma

Multiplee ow diagnostics have been applied to planarpanels covered by strips of glow-discharge surface plasma in atmospheric pressure air generated by the one atmosphere uniform glow discharge plasma. Direct drag measurements, smoke wire and titanium tetrachloride e ow visualization, and boundary-layer velocity proe les were obtained. The plasma generated along streamwise-oriented, symmetric strip electrodes is shown to cause a large increase in drag, whereas the plasma along spanwise-oriented, asymmetric strip electrodes can generate a signie cant thrust. Flow visualization and mean velocity measurements show the primary cause of the phenomena to be a combination of mass transport and vortical structures induced by strong electrohydrodynamic body forces on the e ow, known as paraelectric forcing.

Boundary-layer stability measurements in a hypersonic quiet tunnel

Hypersonic boundary-layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82 x 10 6 /ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary-layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode is the dominant mode of instability. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. Furthermore, the existence of higher harmonics of the fundamental suggests that nonlinear disturbances are not associated with high freestream disturbance levels.

NASA Langley Mach 6 Quiet Wind-Tunnel Performance

The flow in the NASA Langley Mach 6 quiet wind tunnel has been investigated to quantify the effectiveness of laminar-flow control techniques used to delay transition of the nozzle-wall boundary layer. The results of this investigation include an assessment of the mean and unsteady nozzle flow to define the quiet core length, and hence performance, over the operating range of the facility. A large, uniform region of Mach 5.91 flow was documented for a variety of unit Reynolds numbers. By using a prototype constant-voltage anemometer to measure the unsteady flowfield, acoustic radiation patterns from the transitional nozzle-wall boundary layers were mapped. These disturbances originating at the Irregular edge of the transitional nozzle-wall boundary layer were shown to follow Mach lines into the test section of the nozzle, thereby limiting the length of the quiet core. With a virtual origin downstream of the nozzle throat, a Reynolds number dependency was found for the amplitudes of the acoustic radiation.

Hypersonic Boundary Layer Stability over a Flared Cone in a Quiet Tunnel

Hypersonic boundary layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82x106/ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode disturbances were the most unstable and scaled with the boundary layer thickness. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode sub-harmonic. The sub-harmonic wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that non-linear disturbances are not associated with high free stream disturbance levels.

Effect of angle of attack on hypersonic boundary-layer stability

An experimental investigation of the effects of angle of attack on hypersonic boundary-layer stability on a ¯ ared-cone model was conducted in the low-disturbance Mach-6 Nozzle-Test-Chamber Facility at NASA Langley Research Center. Hot-wire anemometry diagnostics were applied to identify the boundary-layer instability mechanisms that lead to transition. The present results show that the boundary layer becomes more stable on the windward ray and less stableon the leeward ray relative to the zero-degree angle-of-attack case. The second-mode instability dominates the transition process at a 0-deg angle of attack; however, on the windward meridian at an angle of attack this mode was stabilized. On the leeward meridian the frequency of the dominant instability was higher than the estimated frequency of the second-mode disturbance; thus the dominant transition mechanism may beother than a second-modedisturbance. Nonlineareffects, such as growth saturation, harmonic generation,

Novel Technique to Determine SparkJet Efficiency

E NERGY deposition has recently received significant interest as a powerful technique in a variety of high-speed flow-control applications [1–5]. To achieve energy deposition, an electromagnetic local flow/flight control (ELFC) device generates pulsed electromagnetic fields. A wide variety of ELFC devices have been developed, including, for example, laser and/or microwave discharge, electron beam, and surface dc/ac discharge [2,3,6] and the nanosecond pulse mode of dielectric barrier discharge operation [7]. The Johns Hopkins University Applied Physics Laboratory has developed a unique ELFC device denoted the SparkJet for flow and flight control [8–14]. Figure 1 illustrates schematically the three stages of energy deposition by a SparkJet device. A spark is discharged within a typical volume of several cubic centimeters (stage 1). The high-pressure gas discharges through a converging nozzle, thereby generating an impulse (stage 2). Provided there is a mechanism for recharging the gas in the cavity (stage 3), the sequence can be repeated. Research has been conducted on the SparkJet by different groups. Narayanaswamy et al. [15] have performed experiments to investigate the effect of a SparkJet discharged at different frequencies and locations upstream of a 30 deg wedge on the separation induced by a shock/boundary-layer interaction. Caruana et al. [16] have carried out numerical and experimental research on a plasma synthetic jet generated by a similar device. They have shown that this method can be applicable for separation control and noise reduction. Anderson and Knight [17] have performed analytical and numerical studies to predict the impulse from the jet and the temporal pressure and temperature inside the cavity upon discharge. They have shown that an array of SparkJets can effectively replace a flap and hence be practical in flight control. The objective of this note is the determination of the electromechanical efficiency of a particular design of a SparkJet. The electromechanical efficiency is the fraction of the electrical energy that results in the generation of the SparkJet mechanical impulse based upon a perfect gas model. The spark discharge generates a plasma with excited electronic, rotational, and vibrational states. A portion of the energy dissipated across the spark gap is channeled into heating of the gas (i.e., increasing the translational–rotational temperature), which leads to a high pressure and subsequent jet exiting through the converging nozzle, thus creating the mechanical impulse. Various definitions of the SparkJet efficiency have been published. Haack et al. [13] have experimentally found an efficiency of 35% based on the measured peak pressure inside the SparkJet. In a later study, Haack et al. [14] have experimentally shown an average efficiency of 20–30% in different operating conditions based on the measured pressure and voltage. In this paper, a novel method for determining the electromechanical efficiency is proposed based upon comparison of predictions of a theoretical model and experimental measurements.

Off-Body Boundary-Layer Measurement Techniques Development for Supersonic Low-Disturbance Flows

*† ‡ Investigations were performed to develop accurate boundary-layer measurement techniques in a Mach 3.5 laminar boundary layer on a 7ϒ half-angle cone at 0ϒ angle of attack. A discussion of the measurement challenges is presented as well as how each was addressed. A computational study was performed to minimize the probe aerodynamic interference effects resulting in improved pitot and hot-wire probe designs. Probe calibration and positioning processes were also developed with the goal of reducing the measurement uncertainties from 10% levels to less than 5% levels. Efforts were made to define the experimental boundary conditions for the cone flow so comparisons could be made with a set of companion computational simulations. The development status of the mean and dynamic boundary-layer flow measurements for a nominally sharp cone in a lowdisturbance supersonic flow is presented. Nomenclature

Control of Stationary Cross-flow Modes Using Patterned Roughness at Mach 3.5

Spanwise-periodic roughness designed to excite selected wave lengths of stationary cross-flow modes was investigated in a 3-D boundary layer at Mach 3.5. The test model was a sharp-tipped 14° right-circular cone. The model and integrated sensor traversing system were placed in the Mach 3.5 Supersonic Low Disturbance Tunnel (SLDT) equipped with a “quiet design” nozzle at NASA Langley RC. The model was oriented at a 4.2° angle of attack to produce a mean cross-flow velocity component in the boundary layer over the cone. Three removable cone tips have been investigated. One has a smooth surface that is used to document the baseline (“natural”) conditions. The other two have minute “dimples” that are equally spaced around the circumference, at a streamwise location that is just upstream of the linear stability neutral growth branch for cross flow modes. The azimuthal mode numbers of the dimpled tips were selected to either enhance the most amplified wave numbers or to suppress the growth of the most amplified wave numbers. The results indicate that the stationary cross-flow modes were highly receptive to the patterned roughness.

Crossflow Instability on a Wedge-Cone at Mach 3.5

As a follow-on activity to the HyBoLT flight experiment, a six degree half angle wedge-cone model at zero angle of attack has been employed to experimentally and computationally study the boundary layer crossflow instability at Mach 3.5 under low disturbance freestream conditions. Computed meanflow and linear stability analysis results are presented along with corresponding experimental Pitot probe data. Using a model-mounted probe survey apparatus, data acquired to date show a well defined stationary crossflow vortex pattern on the flat wedge surface. This effort paves the way for additional detailed, calibrated flow field measurements of the crossflow instability, both stationary and traveling modes, and transition-to-turbulence under quiet flow conditions as a means of validating existing stability theory and providing a foundation for dynamic flight instrumentation development. Nomenclature

Effect of Off-Body Laser Discharge on Drag Reduction of Hemisphere Cylinder in Supersonic Flow

The interaction of an off-body laser discharge with a hemisphere cylinder in supersonic flow is investigated. The objectives are 1) experimental determination of the drag reduction and energetic efficiency of the laser discharge, and 2) assessment of the capability for accurate simulation of the interaction. The combined computational and experimental study comprises two phases. In the first phase, laser discharge in quiescent air was examined. The temporal behavior of the shock wave formed by the laser discharge was compared between experiment and simulation and good agreement is observed. In the second phase, the interaction of the laser discharge with a hemisphere cylinder was investigated numerically. Details of the pressure drag reduction and the physics of the interaction of the heated region with the bow shock are included. The drag reduction due to this interaction persisted for about five characteristic times where one characteristic time represents the time for the flow to move a distance equal to the hemisphere radius. The energetic efficiency of laser discharge for the case with 50 mJ energy absorbed by the gas is calculated as 3.22.