Simulations of a shock emanating from a compression corner and interacting with a fully developed turbulent boundary layer are evaluated herein. Mission-relevant conditions at Mach 7 and Mach 14 are defined for a precompression ramp of a scramjet-powered vehicle. Two compression angles are defined: the smallest to avoid separation losses and the largest to force higher temperature flow physics. The Baldwin-Lomax and the CebeciSmith algebraic models, the one-equation Spalart-Allmaras model with the Catrix-Aupoix compressibility modification, and two-equation models, including the Menter shear stress transport model and the Wilcox k-! 98 and k-! 06 turbulence models, are evaluated. Comparisons are made to existing experimental data and Van Driest theory to provide preliminary assessment of model-form uncertainty. A set of coarse-grained uncertainty metrics are defined to capture essential differences among turbulence models. There is no clearly superior model as judged by these metrics. A preliminary metric for the numerical component of uncertainty in shockturbulent-boundary-layer interactions at compression corners sufficiently steep to cause separation is defined as 55%. This value is a median of differences with experimental data averaged for peak pressure and heating and for extent of separation captured in new grid-converged solutions presented here. Nomenclature c = speed of sound, m=s E = metric of difference between computation and experiment e = static energy, J=kg f = dummy variable for p, q, or H = total enthalpy, J=kg k = turbulent kinetic energy, J=kg L = separation length, m M = Mach number M t = turbulence Mach number, 2k p =c M t0 = critical value of turbulence Mach number used in compressibility correction M = Mach number based on friction velocity, u =c w P = production term in turbulent kinetic energy equation Pr t = turbulent Prandtl number p = pressure, N=m 2 q = heat transfer rate, W=m 2 Re = momentum thickness Reynolds number Re = incompressible momentum thickness Reynolds number T = temperature, K U = u=u , dimensionless velocity u = velocity, m=s u i , u j = velocity component in i and j directions, respectively, m=s u = friction velocity, w = w p V = velocity in freestream, m=s x = distance along wall (flat plate), coordinate in streamwise direction, m x i , x j = coordinates in i and j directions, respectively, m y = distance normal to wall (flat plate), coordinate orthogonal to x, m y = u y=, normalized distance (flat plate) = angle of attack = viscosity, kg=m s = density, kg=m 3 = shear, N=m 2 ij = Reynolds stress tensor k = @u k 1 =@x k1 @u k1 =@x k 1 Subscripts conv = convective e = at edge of boundary layer i = component in i direction j = component in j direction t = turbulent value w = conditions at wall 1 = reference condition in freestream
Interaction between the external flowfield and the reaction control system (RCS) thruster plumes of the Phoenix capsule during entry has been investigated. The analysis covered rarefied, transitional, hypersonic and supersonic flight regimes. Performance of pitch, yaw and roll control authority channels was evaluated, with specific emphasis on the yaw channel due to its low nominal yaw control authority. Because Phoenix had already been constructed and its RCS could not be modified before flight, an assessment of RCS efficacy along the trajectory was needed to determine possible issues and to make necessary software changes. Effectiveness of the system at various regimes was evaluated using a hybrid DSMC-CFD technique, based on DSMC Analysis Code (DAC) code and General Aerodynamic Simulation Program (GASP), the LAURA (Langley Aerothermal Upwind Relaxation Algorithm) code, and the FUN3D (Fully Unstructured 3D) code. Results of the analysis at hypersonic and supersonic conditions suggest a significant aero-RCS interference which reduced the efficacy of the thrusters and could likely produce control reversal. Very little aero-RCS interference was predicted in rarefied and transitional regimes. A recommendation was made to the project to widen controller system deadbands to minimize (if not eliminate) the use of RCS thrusters through hypersonic and supersonic flight regimes, where their performance would be uncertain.
The supersonic transitional flow aerodynamics of the inflatable reentry vehicle experiment are simulated with the direct simulation Monte Carlo method. Also, results from Navier-Stokes calculations are presented that provide both a check on the direct simulation Monte Carlo results near continuum conditions and the general trend of the aerodynamic data at lower altitude conditions. Calculations of axial, normal, and static pitching coefficients are obtained for an angle-of-attack range of 0 to 180 deg. These results clearly demonstrate the strong sensitivity of the aerodynamic coefficients to the relatively low speeds encountered as the inflatable reentry vehicle experiment reenters the atmosphere, and that existing hypersonic aerodynamic data bases for similar geometric configurations are not appropriate for the inflatable reentry vehicle experiment environment. The current numerical simulations focus on the rigid body aerodynamics from 150 to 91 km altitude for the 0 to 180 deg angle of incidence sweep and to lower altitudes (46 km) while at zero incidence. Nomenclaturemaximum diameter of spacecraft, m Kn 1;D;HS = freestream hard sphere Knudsen number, 1 =D mcs = mean collision separation distance, m mfp = mean free path, m n = number density, m 3 p = pressure, N=m 2 q = wall heat transfer rate, W=m 2 T = temperature, K V 1 = freestream velocity, m=s x; y; z = model coordinates, m X = mole fractions = angle of incidence, deg 1 = mean free path in freestream, m 1 = freestream density, kg=m 3 Subscripts D = maximum spacecraft diameter HS = hard sphere W = wall 1 = freestream
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