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At least 775 records · Page 43

Injection slot location for boundary-layer control in shock-induced separation

An experimental investigation of the effect of tangential air injection, when the injection slot is located inside of what would otherwise have been the dead air zone in a separated flow, in controlling shock-induced turbulent boundary layer separation is presented. The experiments were carried out at a free-stream Mach number of 2.5 in the separated flow induced by a compression corner with a 20 deg angle. The observations made were wall static pressures, pitot profiles, and schlieren visualizations of the flow. The results show that the present location for injection is more effective in suppressing boundary-layer separation than the more conventional one, where the slot is located upstream of where separation would occur in the absence of injection.

Viswanath, P. R.↗

A time-accurate high-resolution TVD scheme for solving the Navier-Stokes equations

A total variation diminishing (TVD) scheme has been developed and incorporated into an existing time-accurate high-resolution Navier-Stokes code. The accuracy and the robustness of the resulting solution procedure have been assessed by performing many calculations in four different areas: shock tube flows, regular shock reflection, supersonic boundary layer, and shock boundary layer interactions. These numerical results compare well with corresponding exact solutions or experimental data.

Kim, Hyun Dae↗

A time-accurate high-resolution TVD scheme for solving the Navier-Stokes equations

A total variation diminishing (TVD) scheme has been developed and incorporated into an existing time-accurate high-resolution Navier-Stokes code. The accuracy and the robustness of the resulting solution procedure have been assessed by performing many calculations in four different areas: shock tube flows, regular shock reflection, supersonic boundary layer, and shock boundary layer interactions. These numerical results compare well with corresponding exact solutions or experimental data.

Kim, Hyun D.↗

Electron Scattering by Low-frequency Whistler Waves at Earth’s Bow Shock

Electrons are accelerated to nonthermal energies at shocks in space and astrophysical environments. While shock drift acceleration (SDA) has been considered a key process of electron acceleration at Earth’s bow shock, it has also been recognized that SDA needs to be combined with an additional stochastic process to explain the observed power-law energy spectra. Here, we show mildly energetic (∼0.5 keV) electrons are locally scattered (and accelerated while being confined) by magnetosonic-whistler waves within the shock transition layer, especially when the shock angle is large (θ(sub Bn) approximately equal or greater than 70°). When measured by the Magnetospheric Multiscale mission at a high cadence, ∼0.5 keV electron flux increased exponentially in the shock transition layer. However, the flux profile was not entirely smooth and the fluctuation showed temporal/spectral association with large-amplitude (δB/B ~ 0.3), low-frequency (approximately equal or less than 0.1 Ω(sub ce) where Ω(sub ce) is the cyclotron frequency), obliquely propagating (θ(sub kB) ~ 30°–60°, where θ(sub kB) is the angle between the wave vector and background magnetic field) whistler waves, indicating that the particles were interacting with the waves. Particle simulations demonstrate that, although linear cyclotron resonances with ∼0.5 keV electrons are unlikely due to the obliquity and low frequencies of the waves, the electrons are still scattered beyond 90° pitch angle by (1) resonant mirroring (transit-time damping), (2) non-resonant mirroring, and (3) subharmonic cyclotron resonances. Such coupled nonlinear scattering processes are likely to provide the stochasticity needed to explain the power-law formation.

Oka, M.↗

Control of shock wave-boundary layer interactions by bleed in supersonic mixed compression inlets

An experimental investigation was conducted to determine the effect of bleed on a shock wave-boundary layer interaction in an axisymmetric mixed-compression supersonic inlet. The inlet was designed for a free-stream Mach number of 2.50 with 60-percent supersonic internal area contraction. The experiment was conducted in the NASA Lewis Research Center 10-Foot Supersonic Wind Tunnel. The effects of bleed amount and bleed geometry on the boundary layer after a shock wave-boundary layer interaction were studied. The effect of bleed on the transformed form factor is such that the full realizable reduction is obtained by bleeding of a mass flow equal to about one-half of the incident boundary layer mass flow. More bleeding does not yield further reduction. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise.

Fukuda, M. K.↗

Laminar Boundary Layer Behind a Strong Shock Moving into Air

The laminar wall boundary layer behind a strong shock advancing into stationary air has been determined. Numerical results have been obtained for shock Mach numbers up to 14 using real gas values for density and viscosity and assuming Prandtl and Lewis numbers of 0.72 and 1, respectively. The numerical results for shear and heat transfer agree, within 4 percent, with a previously presented approximate analytical expression for these quantities. A slight modification of this expression results in agreement with the numerical data to within 2.5 percent. Analytical expressions for boundary-layer thickness and displacement thickness, correct to within 4 percent for the present data, have also been obtained.

Mirels, Harold↗

Turbulence modeling of shock separated boundary-layer flows

Computations of transonic and hypersonic shock-separated boundary-layer flows using zero-equation (algebraic), one-equation (kinetic energy), and two-equation (kinetic energy plus length scale) turbulence eddy viscosity models are described and compared with measurements. The computations make use of a new Navier-Stokes computer algorithm that has reduced computing times by one to two orders of magnitude. The algorithm, and how the turbulence models are incorporated into it, are described. Results for the transonic flow show that the unmodified one-equation model is superior to the zero-equation model in skin-friction predictions. For the hypersonic flow, a highly modified one-equation model that accurately predicts surface pressure and heat transfer is described. Preliminary two-equation model results are also presented.

Coakley, T. J.↗

Prediction of secondary separation in shock wave boundary-layer interactions

A complex three-dimensional shock-wave/turbulent boundary layer interaction at Mach 4 has been investigated experimentally and computationally, using two turbulence models with substantial refinement. With the use of a fine grid model, secondary flow separation was successfully computed, and grid changes did not improve agreement with experiment for the extent of upstream influence. The use of a non-isotropic turbulence model gave a slight improvement in upstream influence, but the size of the interaction was still significantly less than in the experiment.

Horstman, C. C.↗

An exploratory study of a three-dimensional shock wave turbulent boundary layer interaction in a corner

Three-dimensional shock wave turbulent boundary layer interactions in a corner have been studied at M about 3. The first phase, detailed mean wall static pressure distributions, has shown the interaction of a two-dimensional flow on one wall with the three-dimensional flow generated by a fin. Major three-dimensional effects on both flows were observed, extending from the interaction of the initial region of the three-dimensional flow and the corner. A large part of the resulting corner flow has little resemblance to either the two-dimensional or three-dimensional interacting flows.

Bogdonoff, S. M.↗

Development of Langmuir and Spatially Resolved OES Diagnostics for Arc Jet Shock and Boundary Layer Measurements Relevant to Atmospheric Re-Entry Environments

The second-generation miniature Arc Jet Research Chamber (mARC) ground testing facility at NASA Ames requires physical characterization of its supersonic plasma flow to verify numerical models and material sample testing. Two types of plasma diagnostics will be utilized: Langmuir probes and optical emission spectroscopy (OES). Several single and triple Langmuir probes have been constructed and will be inserted parallel to the flow via a sweep arm to measure post-shock electron temperature and number density along radial profiles at different axial distances. Due to the high heat flux (200-3000 W/cm2) sweep durations are necessarily short (< 0.1 s). Measuring these properties near the probe surface provide characterization of the post-shock environment. This is an important metric for validating coupled CFD-material codes, and will provide insight into longstanding thermocouple anomalies attributed to charging of heatshield surfaces. An OES system has also been assembled to obtain spatially resolved density and temperature for various flow species. The system consists of a convex lens that forms an image of the plasma onto a 16-channel linear fiber array, which will simultaneously capture the spectra from sixteen different locations. The light emission from the plasma is then focused by a spherical mirror into a spectrometer and recorded with a camera. These measurements non-intrusively provide a new spatially resolved flow characterization capability. This will initially be used to resolve issues with shock radiation interference in pyrometer measurements of material surface temperature. The design of each diagnostic and any initial results will be discussed.

Ethan Main-Loam Leong↗

Experimental study of three shock wave/turbulent boundary layer interactions

The paper presents a systematic study of the supersonic flow of a turbulent boundary layer over several compression-corner models. The wind tunnel and the compression-corner models (ramps fitted with aerodynamic fences to minimize three-dimensional effects) were identical with those used by Settles et al. (1979); constant-temperature hot-wire anemometry was used for the mass-flow measurements. The turning angles used for the compression corners were 8, 16, and 20 deg. In all three flow cases, the shock wave/turbulent flow interaction did amplify the turbulent stresses dramatically, with amplification increasing with increasing turning angle. However, different stress components were amplified by different amounts.

Smits, Alexander J.↗

Inception length to a fully-developed fin-generated shock wave boundary-layer interaction

An experimental study of fin-generated shock wave turbulent boundary-layer interactions confirmed previous observations that, sufficiently far from the fin apex, such interactions become conical. The inception length to conical symmetry was found to increase weakly with Mach number for Mach numbers from 2.5 to 4 and fin angles from 4 to 22 deg. For the range of interactions examined, the inception length was found to depend primarily upon the inviscid shock angle, this angle ranging from 21 to 40 deg. The behavior of the inception length with shock angle can be broadly divided into two categories. For 'weak' interactions with shock angles less than about 35 deg, the inception length decreased as the shock angle increased. For 'strong' interactions with shock angles greater than about 35 deg, the inception region was small and was approximately constant at three boundary-layer thicknesses in length. In the latter, strong interaction case, the inception length was an order of magnitude smaller than that found in the weakest interactions examined, to the extent that strong interactions were practically fully-developed from the apex.

Lu, Frank K.↗

Numerical investigation of turbulence models for shock separated boundary-layer flows

Numerical solutions of the Navier-Stokes equations for shock separated turbulent boundary-layer flows are presented. Several turbulence models are investigated and assessed by their ability to predict the physical phenomena associated with two extensively documented experiments. The experimental flows consist of shock-wave boundary-layer interactions in axisymmetric internal and external geometries at Mach numbers of 1.5 and 7, respectively. Algebraic and one-equation eddy viscosity models are used to describe the Reynolds shear stress. Calculated values of skin friction, wall pressure distribution, kinetic energy of turbulence, and heat transfer are compared with measurements.

Viegas, J. R.↗

Three-dimensional separation for interaction of shock waves with turbulent boundary layers

For the interaction of shock waves with turbulent boundary layers, obtained experimental three-dimensional separation results and correlations with earlier two-dimensional and three-dimensional data are presented. It is shown that separation occurs much earlier for turbulent three-dimensional than for two-dimensional flow at hypersonic speeds.

Goldberg, T. J.↗