James N. Moss

Direct simulation of transitional flow for hypersonic reentry conditions

This paper presents results of flowfield calculations for typical hypersonic reentry conditions encountered by the nose region of the Space Shuttle Orbiter. Most of the transitional flow regime is covered by the altitude range of 150 to 92 km. Calculations were made with the Direct Simulation Monte Carlo (DSMC) method that accounts for translational, rotational, vibrational, and chemical nonequilibrium effects. Comparison of the DSMC heating results with both Shuttle flight data and continuum predictions showed good agreement at the lowest altitude considered. However, as the altitude increased, the continuum predictions, which did not include slip effects, departed rapidly from the DSMC results by overpredicting both heating and drag. The results demonstrate the effects of rarefaction on the shock and the shock layer, along with the extent of the slip and temperature jump at the surface. Also, the sensitivity of the flow structure to the gas-surface interaction model, thermal accommodation, and surface catalysis are studied.

Hypersonic viscous shock-layer solutions over long slender bodies. I - High Reynolds number flows

Numerical solutions from the time-steady viscous shock-layer (VSL) equations are presented for the hypersonic laminar and turbulent flow of a perfect gas over long slender bodies. These results are obtained from a spatial-marching, implicit, finite-difference technique, which includes coupling of the normal momentum and continuity equations and use of the Vigneron pressure condition in the subsonic nose region. Detailed comparisons have been made with other predictions and experimental data to assess the accuracy of the present numerical technique, especially for slender-body flows. The comparisons have been shown to yield accurate results. Two widely used algebraic turbulence models, namely the Cebeci-Smith (CS) and the Baldwin-Lomax (BL) models, have been analyzed with the present technique for application to long, slender bodies. The BL model has been modified in the present work for pressure-gradient effects and applied successfully at hypersonic flow conditions. Both of these models have been shown to result in similar heating predictions for the attached flows analyzed here.

Direct Simulation Monte Carlo Simulations of Hypersonic Flows With Shock Interactions

The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary-layer interactions, where boundary-layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made using the DS2V code of Bird, a general two-dimensional/axisymmetric time-accurate code. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. The freestream Knudsen numbers, with the characteristic length equal to the test model diameter, range from 0.0008 to 0.0004, consequently demanding computations for DSMC simulations. Results presented highlight the sensitivity of the calculations to grid resolution, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Experimental and Numerical Study of the Laminar Separation in Hypersonic Flow

Abstract This article is devoted to an experimental and numerical study of shock wave/boundary layer interaction in hypersonic laminar flow (M = 10). The experimental was performed in the ONERA R5Ch wind tunnel on a hollow cylinder flare. The flow stream delivered by the R5Ch wind tunnel produces physical conditions which justify both the theoretical approach using classical Navier-Stokes equations and the approach by Direct Simulation Monte-Carlo. So the aim of this study is to improve the capacity of Navier-Stokes and DSMC codes to predict high Mach number interactions. Pressure and heat flux have been measured at the wall and compared with the results obtained with Navier-Stokes and DSMC codes. Two different meshes have been considered with the three Navier-Stokes codes used. The conclusion is that the codes can give a good evaluation of the physical quantities at the wall.

Nonequilibrium thermal radiation for an aeroassist flight experiment vehicle

The direct-simulation Monte Carlo method incorporating a dissociating and ionizing gas model for air with thermal radiation is used to characterize the hypersonic flow about an axisymmetric representation of an aeroassist flight experiment (AFE) vehicle, whose freestream conditions correspond to selected points along the entry, aerobraking, and exit phases of the trajectory. Calculations for two trajectory conditions indicate that the radiative heating of the AFE forebody is lower than the convective heating, but becomes significant as the maximum convective heating rate condition is approached.

Nonequilibrium effects for hypersonic transitional flows

Presented are the results of numerical simulations of hypersonic flow about blunt cones and hemispherical nose configurations for reentry velocities of 7.5 and 10 km/s. Cone half angles 0, 5, and 10 deg are considered at zero angle of incidence; however, the focus is for the 5 deg cone. The body size and altitude ranges considered (70 to 110 km) are such that the flow is in the transitional regime. Translational, thermodynamic, and chemical nonequilibrium effects are considered in the numerical simulation by utilizing the direct simulation Monte Carlo (DSMC) method of Bird. The DSMC results are compared with those obtained with viscous shock-layer and Navier-Stokes methods. Comparisons between the DSMC and continuum calculations show the altitude range where differences in flowfield structure and surface quantities become significant. The current calculations show that the binary scaling similitude provides a means of correlating the blunt body surface quantities in the hypersonic, transitional regime. Furthermore, for the higher velocity entry conditions, the results highlight some of the concerns in the application of multitemperature continuum formulations, particularly the use of some proposed functional relations for the chemical rate constants under thermodynamic nonequilibrium conditions.

Assessment of thermochemical nonequilibrium and slip effects for Orbital Reentry Experiment (OREX)

Results are provided from a viscous shock layer (VSL) analysis of the reentry flowfield around the forebody of the Japanese Orbital Reentry Experiment (OREX) vehicle. This vehicle is a 50 deg. spherically blunted cone with a nose radius of 1.35 m and a base diameter of 3.4 m. Calculations are done for the OREX trajectory from 105 to 48.4 km altitude range. A 7-species chemical model is found adequate for the flowfield analysis. However, for altitudes greater than 84 km, the low density effects (such as thermal nonequilibrium and slip) are to be implemented for good agreement between the predictions and flight inferred heat-transfer rate data. Further, at altitudes lower than 84 km, a finite surface recombination probability is to be employed in place of a non-catalytic surface for better comparison between the calculations and data. VSL results are also compared with the direct simulation Monte Carlo (DSMC) predictions at high altitudes (greater than 80 km) and the electron number density data for three altitudes in the OREX trajectory. Overall, there is a good comparison between the flight data and calculated results. With the ongoing refinements in data extraction procedures, the OREX data should prove valuable for validating theoretical models employed in flowfield codes for calculation of reacting-gas flowfields.

Effects of Chemistry on Blunt-Body Wake Structure

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Viscous Shock Layer Analysis of the Martian Aerothermal Environment

Detailed surface heating and flowfield results have been obtained for the stagnation region of a planetary exploration vehicle entering the Martian atmosphere. A viscous shock layer analysis (which includes an absorbing boundary layer) is used to obtain solutions with and without coupled ablation injection. Recently developed curve fits for the transport and thermodynamic properties of Martian atmospheric and ablation species as well as for the absorption coefficient for CO(4 + ) are employed. Extensive results are provided at altitudes of 30, 36, and 50 km for bodies with nose radii of 1, 2.3, and 23 m at freestream velocities of 6, 8,10, and 12 km/s. Sublimation temperature is employed with coupled ablation injection cases, whereas radiative equilibrium wall temperature is used without injection. Only for bodies with large nose radii (23 m or larger) and for velocities of approximately 6 km/s can a reusable heat shield (with the currently available materials) be used. For higher velocities or vehicles with smaller nose radii, an ablative thermal protection system will be required. A comparison with thermochemical nonequilibrium calculations suggests that much of the flow in the shock layer is in thermochemical equilibrium for the cases analyzed. This is one of the first studies for the Martian entry conditions of large size bodies with coupled radiation and ablation injection.

Direct simulation of diatomic gases using the generalized hard sphere model

The generalized hard sphere model that incorporates the effects of attraction and repulsion is examined. First, the model is used to study simple adiabatic heat baths. Then, it is used to predict flow measurements in tests involving extremely low freestream temperatures. For the two cases considered, a Mach 26 nitrogen shock and a Mach 20 nitrogen flow over a flat plate, only rotational excitation is deemed important, and appropriate modifications for the Borgnakke-Larsen procedure are developed. In general, for the cases considered, the present model showed a modest improvement over the variable hard sphere model.