James R. Hayes

Hypersonic Flow over a Blunt Body with Plasma Injection

The drag reduction of blunt body in hypersonic flow via plasma injection has been investigated by a combined experimental and computational effort. The counterflow plasma jet generated by a plasma torch has a vibrionic temperature of 4400 K, an electronic temperature around 20,000 K, and electron density greater than 3 × 1012/cm3. At a fixed injection stagnation pressure and in the absence of an applied magnetic field, the plasma injection actually increases drag above that of room-temperature air due to a decreased mass flow rate at the elevated temperature. However, at an identical mass flow rate, the plasma injection reveals a greater drag reduction than the room-temperature air counterpart through thermal energy deposition. From experimental measurements, an overwhelming major portion of the drag reduction is derived from the viscous-inviscid interaction of the counterflow jet and thermal energy deposition. The numerical results of Navier-Stokes equations with a local equilibrium plasma composition also confirm this observation.

HIFiRE-1 Boundary Layer Transition Experiment Design

*† ‡ § ** The HIFiRE (Hypersonic International Flight Research and Experimentation) project will develop and demonstrate fundamental hypersonic technologies deemed critical to the realization of next generation aerospace weapon systems. Flight one focuses primarily on integration of instrumentation on the test vehicle, with application to future flights. Boundary layer transition has been chosen as one aerodynamic parameter to measure in order to assess instrumentation performance. Analysis has been performed to place requirements on the payload. Results show that the temperature of the aeroshell is relatively benign, easing temperature requirements on transducers. Reynolds number on the vehicle is relatively low, making it marginal for smooth-body boundary layer transition. Roughness must be placed on the vehicle to ensure transition and to obtain rough-wall data. I. Introduction The HIFiRE (Hypersonic International Flight Research and Experimentation) project will develop and demonstrate fundamental hypersonic technologies deemed critical to the realization of next generation aerospace weapon systems. The research effort will consist of a series of focused tasks to resolve hypersonic phenomena through validation of computational analysis; comparison with performance predictions, enhancement of design data bases; and development of correlations with ground test. Each research effort will culminate with a flight experiment to be launched to representative flight conditions (Reynolds number and Mach) employing low cost sounding rockets. The primary objective of flight one is to determine the feasibility of applying high-bandwidth instrumentation to aerothermal flight measurements for short duration hypersonic flights at Mach numbers up to eight. The experiment will determine the suitability and survivability of multiple instrumentation types in this environment. The experiment will be incorporated into the nose cone of a Terrier-Orion launch vehicle. This paper will refer to the nose cone as the “payload.” The measurement of three aerothermal phenomena will serve to assess the performance of the instrumentation. These sub-experiments are, in order of priority, boundary layer transition (BLT), turbulent separated shock boundary layer interaction (SBLI), and optical measurement of mass capture (OMC) in a duct. Measurement of these phenomena will serve as secondary objectives of the experiment. This paper describes analysis related to the BLT experiment to determine requirements for flight one. Boundary layer transition has an important impact on hypersonic vehicle aerodynamics and aerothermodynamics. It is an inherently challenging problem due to its nonlinear nature and sensitivity to initial and boundary conditions. Ground test of boundary layer transition is generally unsatisfactory due to the high noise levels of ground facilities compared to flight. Computation is challenging since it requires resolution of instabilities of relatively high wavenumber and frequency over large spatial domains. Flight provides the best environment for measuring transition. Instrumentation for the flight environment however, is challenging. The highest level of validation for transition prediction requires high bandwidth instrumentation. The current tests will explore the use of high-bandwidth instrumentation for flight measurement of transition. Careful measurement and computation of boundary conditions will provide added value in interpreting the data. Data from numerous past flight tests of supersonic and hypersonic transition exist. Schneider 1 offers a detailed review of transition flight tests. Vehicles were frequently thin-skin models. Backface thermocouples monitored transition, and heat transfer rates were inferred by lumped-capacitance heating models. Transition criteria based on

Rotational and Vibrational Temperature Distributions for a Dielectric Barrier Discharge in Air

Spatially resolved rotational and vibrational temperatures for N 2 and rotational temperatures for N + 2 , as a function of voltage, have been obtained for an asymmetric surface mode dielectric barrier discharge using emission spectroscopy. The rotational temperatures were obtained from a nonlinear least-squares fit of a two-temperature theoretical spectrum with the measured spectra of the N 2 (C 3 Π u ― B 3 Π g ) and N + 2 (B 2 Σ + u ― X 2 Σ + g ) electronic band systems. The vibrational temperatures were obtained by applying the Boltzmann plot method to the Δv = ―2 sequence of the N 2 (C 3 Π u ― B 3 Π g ) electronic band system. It was observed that the rotational temperatures for N 2 ? and N + 2 decreased in the induced flow direction and increased with increasing voltage. Values started at 390 ± 10 K and decreased to 340 ± 10 K for N 2 and started at 500 ± 15 K and decreased to 450 ± 15 K for N + 2 . The vibrational temperatures also decreased in the induced flow direction from 3250 to 2850 ± 300 K. A difference in rotational temperatures between N 2 and N + 2 was observed for all voltages studied, and these differences increased with increasing voltage. The rotational temperatures of both species fluctuated in the spanwise direction. These fluctuations damped out in the streamwise direction and were weakly correlated with the attachment points of the microdischarges on the edge of the exposed electrode.

Study of Plasma Electrode Arrangements for Optimum Lift in a Mach 5 Flow

Abstract : This work is an experimental effort to study the power efficiency of using a plasma discharge to alter the lift on a body or surface. In this paper several electrode geometries are considered in an effort to reduce the plasma power required for a given change in lift. The cathode electrode position and electrode size are studied. For all cases studied the anode electrode is kept the same. Results are presented for four different size cathodes and four different cathode positions. The primary result presented is the lift change produced by the discharge per unit power input. The lift is determined by measuring the deflection of the model under the applied plasma. This type of a measurement system has some advantages and disadvantages compared to a load cell lift measurement system used by the authors in past work. Results from each of these lift measurement tools compare well. Results for 9 and 24 mA DC discharges are shown in this paper. For the conditions utilized in this work the results indicate that both cathode position and cathode size affect the lift change caused by a plasma discharge per unit of power input.

Application of Plasma Discharge Arrays to High-Speed Flow Control

Surface DC plasma discharges were created in the boundary layer of a plate in a Mach 5 flow. The electrodes consisted of three circular cathodes and three pin anodes arranged to create transverse discharges. The cathodes were arranged linearly along the centerline of the plate, with the anodes displaced laterally. The cathodes and anodes were each held at common voltages. The discharge was run in two configurations, one with a single cathodeanode pair lit, and one in which all three cathode-anode pairs were lit. The discharge from these electrode configurations and its effect on the flow were strikingly different from the streamwise discharge created by linear electrodes. Flow modification from the circular electrode discharges was confined primarily to the boundary layer. The circular electrode discharges act similarly to crossflow jets or bumps , creating a weak wave structure with a pronounced vortex in the boundary layer downstream of the cathode. In contrast to the linear electrode discharges, surface static pressures and the inviscid flow above the boundary layer were largely unaffected by the circular electrode discharge. Substantial boundary layer distortion from the circular electrodes occurred at powers as low as 15 Watts. The circular cathode arrays have potential applications as high -bandwidth vortex generators.

Development of a Langmuir Probe for Plasma Diagnostic Work in High Speed Flow

At the present time there is an interest in controlling high-speed flow over a body utilizing a plasma and an applied magnetic field. In order to critically understand what is happening in this process it is necessary to characterize the plasma. This paper outlines the development of a Langmuir probe for this task and presents some data taken in a quiescent plasma. A Langmuir probe can be used to determine the electron number density, the electron temperature, and the electric field as a function of position in the plasma. From this information the electrical conductivity can be determined utilizing drift velocity data available in the literature. Both a single and a double Langmuir probe were developed and tested.

Rotational and Vibrational Termperatures for a Dielectric Barrier Discharge in Air Using Emission Epectroscopy

Spatially resolved rotational and vibrational temperatures have been obtained for a dielectric barrier discharge (DBD) using emission spectroscopy. The temperatures were obtained by ma tching the measured nitrogen second positive 0 -2, 1 -3 and 2 -4 transitions with a calculated one. The temperature had a periodic profile in the spanwise direction where the peaks correlated with the strong attachment points within the discharge. These resul ts, as well as visual observations, indicate that there are variations in the structure of the DBD and it is not a fully uniform glow discharge. It may be that these temperature high points indicate localized flow structure in the form of small plasma jets . The maximum rotational temperature in the DBD was 410 K located at the interface between the exposed and buried electrode. The minimum value of temperature measured was 310 K located at the edge of the discharge region. The vibrational tem peratures had a maximum of 6200 K at the i nterface, and a minimum of 5700 K at the edge of the discharge region.

Application of a Spectroscopic Measuring Technique to Plasma Discharge in Hypersonic Flow

Abstract : Spatially resolved rotational temperatures have been obtained within the boundary layer of a flat plate model in a Mach 5.1 flow using emission spectroscopy. The temperatures were obtained by matching the measured nitrogen second positive 0-2 rovibrational band with a calculated one. Temperature profiles are given above the cathode and the anode. The maximum temperature obtained above the cathode did not correspond to the surface of the model, but rather at an elevation 0.55 mm above the surface. This indicates that heat is traveling from the discharge into the plate for a period of time after the discharge is ignited. This characteristic in the temperature profile sheds light on some of the effects a plasma has on the flow field as determined by Pitot probe measurements and total lift measurements.