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This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators that are designed to control shock-wave induced boundary-layer separation in a Mach 2.8 supersonic flow. By using two dielectric barrier surface discharge actuator configurations, parallel and canted with respect to the flow velocity vector, a previous experiment suggested that the actuator worked in two ways to influence the interaction: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the boundary-layer separation stronger, while the canted electrode produced vorticity and suppressed the boundary-layer separation due to the momentum transfer from the core flow. Because the detailed physical processes are still unclear, in this paper a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the plasma actuation, in which the dielectric barrier discharge produced plasma is approximated as a high temperature region. The flow characteristics without the plasma actuation correspond to the experimental observation, indicating that the numerical method successfully resolves the shock-wave/boundary-layer interaction. With the plasma actuation, complete agreement between the experiment and calculation has not been obtained in the size of shock-wave/boundary-layer interaction region. Nevertheless, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the strength of the interaction, while the canted electrode leads to a reduction of the interaction strength, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the shock induced boundary layer separation.

During rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency (RF) transmission under certain conditions. To clarify the physical process involved and to establish the estimation methodology, a plume-RF interference experiment during a sea-level static firing test of a full-scale solid rocket motor was conducted. The result of the ground experiment was adequately matched by a computational fluid dynamics (CFD) model of the plume flow field coupled to a finite-difference timedomain (FDTD) model of RF transmission. The CFD/FDTD coupling method was then refined for predicting interference and RF attenuation levels during an actual rocket flight. The calculated far-field received levels were compared with the in-flight attenuation data at different look angles (angles between the vehicle axis and the line-of-sight of the antennas). The calculated results showed good agreement with the flight data over a wide range of look angles. An adaptation of the model, based on the diffraction theory, proved appropriate both for rough estimation of attenuation and for conducting a preliminary analysis of signal/rocket plume interactions. Nomenclature A = pre-exponential factor for Arrhenius expression, consistent unit c = speed of light, m/s d = width of plasma slab, m e = elementary charge, C E = electric field vector, V/m E a = activation energy for Arrhenius expression, J/mol E y = y-component of electric field, V/m H = magnetic field vector, V/m k = rate coefficient, (m 3 /mol) N−1 /s k B = Boltzmann constant, J/K m e = electron mass, kg n = time step for FDTD method N = reaction order N h = heavy-particle number density, m −3 2 N e = electron number density, m −3 P c = combustion chamber pressure, Pa Q e = electron collision cross section, m 2 R = universal gas constant, J/K/mol T = flow temperature, K T e = electron temperature, K V = received radio wave voltage with motor plume, V V 0 = received radio wave voltage without motor plume, V α = look angle of launch vehicle, deg (shown in Fig. 1) β = roll angle of launch vehicle, deg (shown in Fig. B1) χ = index of attenuation χ 0 = relative electric susceptibility ∆t = time step used in FDTD method, s ∆x, ∆y, ∆z = grid sizes in x, y and z directions, respectively, used in FDTD method, m Φ Φ Φ Φ = recursive accumulator, C/m 2 ε = relative permittivity of free electron ε 0 = vacuum permittivity, F/m η = temperature exponent for Arrhenius expression ϕ c = critical angle, rad λ = radio wave wavelength, m µ = index of refraction ν e = electron collision frequency, s −1 ω = radio wave angular frequency, s −1 ω p = electron plasma frequency, s −1

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