James L. Brown

Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

Modeling of Shock Tunnel Aeroheating Data on the Mars Science Laboratory Aeroshell

A series of shots are run in the T5 shock tunnel at California Institute of Technology to measure heating levels on a 70 blunt cone at angle of attack in an environment representative of the Mars Science Laboratory entry. Twenty shots are obtained in CO 2 over a range of enthalpies and pressures chosen to span the laminar and turbulent flow regimes. The data indicate that the lee side turbulent heating augmentation predicted by flight simulations is valid and must be accounted for during the design of the thermal protection system. Computational fluid dynamic simulations are generally in good agreement with the laminar data when employing a supercatalytic wall model, whereas turbulent simulations are in reasonable agreement when a noncatalytic wall model is used. The reasons for this discrepancy are unknown at this time. The turbulent heating augmentation is shown to be inversely related to freestream enthalpy. Changes in angle of attack between 11 and 16 are shown to have minimal impact on measured and computed heating. A transition criterion based on momentum thickness Reynolds number, analogous to that used in flight predictions, predicts onset with reasonable accuracy, although transition is observed to occur later than the current design criterion indicates.

Hypersonic Shock Wave Impingement on Turbulent Boundary Layers: Computational Analysis and Uncertainty

Current status of computational uncertainty for impinging hypersonic shock wave turbulent boundary-layer interactions (SWTBLIs) is evaluated by comparison of computational results with vetted experiments. Employed is one of NASA’s production real gas Reynolds-averagedNavier–Stokes finite volume codes, DPLR, alongwith several commonly used turbulence models. Uncertainty and residual errors, inherent to the analysis and turbulence model implementation, are numerically evaluated for physics quantities of interest. These uncertainty results should prove of value to computational practitioners and developers and to designers making use of modern computational methods to innovate and develop hypersonic hardware such as prototype scramjet engines. Reported are statistical means, variances, and confidence limits of uncertainty measures for the physics quantities of interest to reveal the certitude with which computations of impinging hypersonic SWTBLIs can be relied. A hybrid computational fluid dynamics correlation approach yields improved peak heating estimates in the vicinity of SWTBLIs.

Computational Fluid Dynamics for Winged Re-entry Vehicles at Hypersonic Conditions

The present work reviews the status of computational fluid dynamics methods for predicting the hypersonic flow environments around winged re-entry vehicles. Special attention is paid to the applicability of current flow solvers for use within the context of future aerothermodynamic shape optimization methods. Recent experiences using hypersonic computational fluid dynamic methods for the Columbia accident investigation and the Shuttle Return to Flight Program form the basis of the current assessment. Attention is focused on real-gas Navier-Stokes solvers appropriate for flow regimes typically experienced by winged earth-entry vehicles. Each of the major physical and numerical modeling issues for these solvers is examined individually: numerical methods, thermodynamic properties, laminar transport modeling of mass, momentum and energy, chemistry kinetics, and gas-surface interaction modeling. The variability and uncertainty levels of real-gas Navier- Stokes solvers, in terms of mesh refinement and physical modeling choices, are explored for different flow regions including, wind-side acreage, wing leading edges, lee sides, coves and aft bodies. The feasibility of employing computational fluid dynamics methods for the design optimization of hypersonic entry vehicle shapes is also addressed.

An Asymmetric Capsule Vehicle Geometry Study for CEV

An asymmetric heatshield configuration named Asymmetric Capsule Vehicle, intended for use as the basis for an atmospheric entry, is introduced. The aerodynamic and aerothermodynamic behavior of this new class of heatshield geometries is examined in the supersonic and hypersonic flow regimes using both high-fidelity and engineering analysis techniques, and compared to the performance of heatshields derived from the asymmetric Aeroassist Flight Experiment (AFE) and the symmetric Apollo configurations. The geometry of this new class of heatshield is described analytically and involves a small set of parameters suitable for inclusion in an optimization process. Two particular optimized configurations, the ACVOpt and ACVeOpt2 heatshield shapes, prove to have ‐ (1) higher L/D aerodynamic performance, and (2) lower convective and radiative heating compared to the Apollo-derived symmetric shape. These configurations also do not have the supersonic trim stability issue found to occur with the asymmetric AFE-derived shape.

Flow Topology About An Orbiter Leading Edge Cavity At STS-107 Reentry Conditions

The topology of the three-dimensional flow generated by the Shuttle Orbiter undergoing hypersonic reentry is examined. Simulated damage, modeled as a cavity in the leading edge, leads to an external flow that differs considerably on the leeward side as compared to the external flow over the smooth body Orbiter. Flow topology analysis provides an organizing principle to clarify these complex 3 D flowfields. Surface shear stress maps of the solid surfaces facilitate identification of topological singularities, such as saddle points, nodes and foci. Associated with these surface singularities are flow structures in the 3D flow regions above the solid surface. The relationship of identified vortical structures, shocks, and streamlines to surface property changes are c larified. Real-Gas Navier-Stokes s imulations, with and without the damage cavity, at a selected Orbiter reentry trajectory point during the period o f peak heating provide the flow field descriptions upon which flow topological analyses are performed. The topological analyses of these CFD flow simulations lead to an improved understanding of the nature of the flow differences arising from the inclusion of the leading edge c avity and their relation to changes in the Orbiter surface pressure and temperature fields.