Numerical Study of Plasma-Assisted Aerodynamic Control for Hypersonic Vehicles

Bisek, Nicholas J.

doi:10.2514/1.39032

Cited by 45 publications

(27 citation statements)

References 104 publications

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“…26 Note that the formulas for G, G v , H, and H v are similar to those specified in Eqn. (2). The source vector, S c , on the right side of Eqn.…”

Section: Methodsmentioning

confidence: 99%

High-Order Implicit Large-Eddy Simulations of Supersonic Corner Flo

Bisek

2014

52nd Aerospace Sciences Meeting

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Large-eddy simulations are used to investigate a supersonic wall-bounded turbulent corner flow. Solutions are obtained using a high-fidelity time-implicit numerical scheme and an implicit large-eddy simulation approach. The inclusion of the sidewall leads to the development of a corner core flow which grows in size as the flow progresses downstream. A grid resolution study, with over a billion cells for the fine grid, is performed and both mean and time-accurate statistics are analyzed. The solutions are compared to a spanwiseperiodic flat-plate turbulent boundary layer developed at the same conditions and yield similar results when measured sufficiently far from the corner. Two-point autocorrelations verify that the domain's cross-sectional area is sufficient to de-correlate the corner core flow front eh turbulent equilibrium boundary-layer developed on each of the adjoining flat-plate walls. In addition, triple products are collected and demonstrate that the corner-dominated flow is significantly different from the rest of the domain. Nomenclaturewhere s is the wall normal direction E = total specific energy F, G, H = inviscid vector fluxes F v , G v , H v = viscous vector fluxes J = transformation Jacobian = geometry length M = Mach number p = nondimensional static pressure Re = Reynolds number, ρ ∞ u ∞ /µ ∞ t = nondimensional time T = nondimensional static temperature U = conserved variable vector u, v, w = nondimensional Cartesian velocity components in the x, y, z directions x, y, z = streamwise, normal, and spanwise directions in nondimensional Cartesian coordinates y + = nondimensional wall distance normalized by local inner scales, u τ ρ w y/µ w δ = boundary-layer thickness, 0.99 u ∞ ξ, η, ζ = computational coordinates θ = compressible boundary-layer momentum thickness, ∞ 0 ρ u ρ∞ u∞ 1 − u u∞ dy µ = dynamic viscosity ρ = nondimensional density τ ij = components of the viscous stress tensor Subscript w = wall ∞ = freestream * Research Aerospace Engineer, AFRL/RQHF. Senior Member AIAA.

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“…26 Note that the formulas for G, G v , H, and H v are similar to those specified in Eqn. (2). The source vector, S c , on the right side of Eqn.…”

Section: Methodsmentioning

confidence: 99%

High-Order Implicit Large-Eddy Simulations of Supersonic Corner Flo

Bisek

2014

52nd Aerospace Sciences Meeting

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“…The model was implemented through source terms in the momentum equation, vibrational energy equation, and total energy equation, based on the formulations in previous publications. 19,21,24,35,36 The spatial distribution function was taken to have the following form:…”

Section: Actuator Modelmentioning

confidence: 99%

Control of separated flow in a reflected shock interaction using a magnetically-accelerated surface discharge

2012

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A numerical investigation was carried out to explore the effects of a magnetically-accelerated surface discharge on a separated, turbulent boundary layer in supersonic flow. The geometry and test conditions were chosen for comparison to experiments carried out at Princeton University. For those studies, a reflected shock interaction was created using a 14° shock generator acting on an incoming turbulent boundary layer with a Reynolds number based on momentum thickness of 1 × 104 and a freestream Mach number of 2.6. Three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) calculations were carried out to simulate the experiments, using the US3D code developed at the University of Minnesota. The baseline code was modified to include a semi-empirical model of the surface discharge actuator, implemented through source terms in the momentum equation, vibrational energy equation, and total energy equation. The computational results for the baseline flow and several control cases were compared to experimental measurements of mean surface pressure. The level of discrepancy was typical of well-resolved RANS computations of three-dimensional, separated flows: qualitative agreement was obtained, and the general experimental trends were captured by the numerical model. Substantial three-dimensionality was observed even in the baseline flow, and significant changes in the flow topology were observed with the application of the actuator. Because of the highly three-dimensional nature of this shock interaction, the initial interpretation of the experiments may need to be revisited.

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“…14,15 LeMANS is a general 2D/axisymmetric/3D, parallel, unstructured finite-volume CFD code and has been used previously in numerous studies of hypersonic flows. [12][13][14][15] LeMANS may be employed with any of three thermodynamic models: perfect gas, equilibrium, and nonequilibrium thermochemistry. LeMANS employs a two-temperature model to account for thermal nonequilibrium and a standard finite-rate chemistry model for nonequilibrium chemistry.…”

Section: Methodsmentioning

confidence: 99%

“…This low power budget contrasts with previous bow shock control studies, typically on the order of 10 kW, using pulsed DC discharge, 8 pulsed microwave discharge, 9 and laser optical breakdown. 10,11 In this paper, the demonstrated effect will be reproduced using the LeMANS code, 12,13 developed at University of Michigan. 14,15 LeMANS was previously used for hypersonic flow-control by energy deposition, 12 therefore rapid energy coupling by the nanosecond pulse DBD is modeled in the same phenomenological approach.…”

Section: Introductionmentioning

confidence: 99%

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Computational and Experimental Analysis of Mach 5 Air Flow over a Cylinder with a Nanosecond Pulse Discharge

Bisek

Poggie

Nishihara

et al. 2012

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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A computational study is performed for Mach 5 air flow over a cylinder with a dielectric barrier discharge actuator set into the cylinder surface. The actuator is pulsed at nanosecond time scales, which rapidly adds energy to the flow, thereby creating a shock wave that travels away from the pulse source. As the shock wave travels upstream, it interacts with the standing bow-shock and momentarily increases the bow-shock standoff distance. This phenomenon is also observed in phase-locked schlieren photography captured during the experiment. The focus of this paper is to reproduce flow phenomena observed in the experiment using high-fidelity computations in order to provide additional insight into the shock-shock interaction, the effect the dielectric barrier discharge pulse has on the surface properties of the cylinder, and develop a reduced-order phenomenological model representative of the nanosecond pulse discharge system. Experimental and high-fidelity modeling studies of the nanosecond pulse dielectric barrier discharge plasma actuators are known to operate with relatively low temperatures. This work explores the possibility that the induced compression wave is generated by rapid thermalization of the discharge which results in a local temperature rise occurring on longer time scales. Two-dimensional simulations are performed and provide many useful details about the discharge event while comparing with many measurements captured by the experiment. However, the simulations indicate the experiment experiences significant three-dimensional effects, thus requiring a three-dimensional simulation of the entire experiment to accurately capture the complex cylinder/tunnel-sidewall interaction and replicate the resultant flow. Three-dimensional results of the discharge event reveal the discharge pulse produces a compression wave that interacts with the standing bow-shock and that the momentary increase in the bow-shock standoff distance is not due to the interaction the three-dimensional compression wave has with the cylinder/tunnel-sidewall boundary layer. AIAA 2012-0186This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Report Documentation Page Form Approved OMB No. 0704-0188Public reporting burden for the collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to a penalty for failing to comply with a collection of information if it does not display...

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Numerical Study of Plasma-Assisted Aerodynamic Control for Hypersonic Vehicles

Cited by 45 publications

References 104 publications

High-Order Implicit Large-Eddy Simulations of Supersonic Corner Flo

High-Order Implicit Large-Eddy Simulations of Supersonic Corner Flo

Control of separated flow in a reflected shock interaction using a magnetically-accelerated surface discharge

Computational and Experimental Analysis of Mach 5 Air Flow over a Cylinder with a Nanosecond Pulse Discharge

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