Michael Barnhardt

Unstructured grid approaches for accurate aeroheating simulations

The use of tetrahedral, prismatic, and hybrid hexahedral-prismatic-tetrahedral grids for the accurate prediction of aerodynamic heating at hypersonic conditions is investigated. We find that tetrahedral grids introduce significant error in the vicinity of strong shock waves, which results in unacceptable aeroheating predictions. The source of this error is studied with an idealized model, and it is found that a large spurious component of post-shock velocity is generated by triangular and tetrahedral elements. This type error is much smaller and easier to control on quadrilateral or hexahedral grids. Thus, we are very skeptical about the utility of tetrahedral grids for accurate hypersonic aeroheating predictions. Several comparisons of heating predictions for a three-dimensional sphere are made, and it is found that the stagnation region results are very sensitive to the grid design. Based on this work and our experience, we advocate the use of unstructured hexahedral grids which increase the grid design space, reduce the element count for many geometries, and result in accurate aeroheating predictions.

Development of the US3D Code for Advanced Compressible and Reacting Flow Simulations

Aerothermodynamics and hypersonic flows involve complex multi-disciplinary physics, including finite-rate gas-phase kinetics, finite-rate internal energy relaxation, gas-surface interactions with finite-rate oxidation and sublimation, transition to turbulence, large-scale unsteadiness, shock-boundary layer interactions, fluid-structure interactions, and thermal protection system ablation and thermal response. Many of the flows have a large range of length and time scales, requiring large computational grids, implicit time integration, and large solution run times. The University of Minnesota / NASA US3D code was designed for the simulation of these complex, highly-coupled flows. It has many of the features of the well-established DPLR code, but uses unstructured grids and has many advanced numerical capabilities and physical models for multi-physics problems. The main capabilities of the code are described, the physical modeling approaches are discussed, the different types of numerical flux functions and time integration approaches are outlined, and the parallelization strategy is overviewed. Comparisons between US3D and the NASA DPLR code are presented, and several advanced simulations are presented to illustrate some of novel features of the code.

An Overview of Technology Investments in the NASA Entry Systems Modeling Project

The Entry Systems Modeling Project, within the NASA Game Changing Development Program, is in its third year conducting mid-TRL research in the disciplines of entry aerosciences and entry thermal protection materials. The Project team is working a variety of challenging problems ranging from the delivery of new aerothermal CFD codes, to the development of the first truly new ablation material response model in more than 40 years, to new conformal and truly flexible thermal protection materials, using novel polymer resins and advanced multi-layered concepts, that will revolutionize entry system designs for future NASA missions. This paper briefly summarizes the achievements to date of the ESM project and provides a full bibliography of papers published by the project over its first two years for the interested reader.

Understanding High Recession Rates of Carbon Ablators Seen in Shear Tests in an Arc Jet

High rates of recession in arc jet shear test s of Phenolic Impregnated Carbon Ablator (PICA) inspired a series of tests and analysis on FiberForm (a carbon preform used in the fabrication of PICA). Arc jet tests were p erformed on FiberForm in both air and pure nitrogen for stagnation and shear configurations . The nitrogen tests showed little or no recession, while the air tests of FiberForm show ed recession rates similar to that of PICA (when adjusted for the difference in density). While mechanical erosion can not be ruled out, this is the first step in doing so. Anal ysis using a carbon oxidation boundary condition within DPLR was used to predict the recession rate of FiberForm. The analysis indicates that much of the anomalous recession behavior seen in shear tests may simply be an artifact of the non-flight like test conf iguration (copper upstream of the test article) a result of dis- similar enthalpy and oxygen concentration profiles on the copper. Shape change effects were also investigated and shown to be relatively small.

Toward Supersonic Retropropulsion CFD Validation

This paper begins the process of verifying and validating computational fluid dynamics (CFD) codes for supersonic retropropulsive flows. Four CFD codes (DPLR, FUN3D, OVERFLOW, and US3D) are used to perform various numerical and physical modeling studies toward the goal of comparing predictions with a wind tunnel experiment specifically designed to support CFD validation. Numerical studies run the gamut in rigor from code-to-code comparisons to observed order-of-accuracy tests. Results indicate that this complex flowfield, involving time-dependent shocks and vortex shedding, design order of accuracy is not clearly evident. Also explored is the extent of physical modeling necessary to predict the salient flowfield features found in high-speed Schlieren images and surface pressure measurements taken during the validation experiment. Physical modeling studies include geometric items such as wind tunnel wall and sting mount interference, as well as turbulence modeling that ranges from a RANS (Reynolds-Averaged Navier-Stokes) 2-equation model to DES (Detached Eddy Simulation) models. These studies indicate that tunnel wall interference is minimal for the cases investigated; model mounting hardware effects are confined to the aft end of the model; and sparse grid resolution and turbulence modeling can damp or entirely dissipate the unsteadiness of this self-excited flow.

Simulations of High-Speed Flow over an Isolated Roughness

A study has been performed to determine the feasibility of using computational fluid dynamics as a tool for predicting hypersonic boundary layer transition to turbulence and the resulting increase in heat transfer. Of particular interest is whether Detached Eddy Simulation can be used to overcome the scaling problems associated with Direct Numerical Simulation and Large Eddy Simulation of boundary layers. Fine-grid results for a boundary layer trip oriented 45 degrees to the flow inside a Mach 10 hypersonic wind tunnel show that Detached Eddy Simulation can predict transition in this perfect-gas flow. I. Introduction Prediction and control of boundary layer transition in hypersonic flows are of crucial importance for the design of planetary atmospheric entry vehicles as well as two-stage-to-orbit reusable launch systems. Since turbulent heat transfer rates can be up to five times higher than laminar heating rates, reductions in the weight of thermal protection systems can be realized with an improved understanding of the physics of transition from laminar to turbulent flow. The gap-filler incident during Space Shuttle mission STS-114 in 2005 was a potent reminder of the importance of accurate prediction of roughness-induced boundary layer transition and the subsequent increase in surface heating 1

Detached-eddy simulation of the reentry-F flight experiment

Results of detached-eddy simulations of the Reentry-F vehicle are presented and compared to the experimental data. Comparisons between the detached-eddy turbulence model and a traditional Reynolds-averaged Navier– Stokes model show a substantial improvement in the predicted base heating rates. A grid-resolution study is performed to examine the sensitivity of the detached-eddy formulation to the grid spacing. Initial results at a zero angle of attack indicate a discrepancy with the experiment that is accounted for by the inclusion of a small experimentally measured angle of attack. Moreover, it is observed that the simulations are very sensitive to the vehicle orientation. Upon accounting for this sensitivity, the resulting predicted base pressure and heating are shown to be in very good agreement with the flight data, typically within the stated experimental uncertainty.

Results from the Mars Science Laboratory Parachute Decelerator System Supersonic Qualification Program

In 2010 the Mars Science Laboratory (MSL) Mission will deliver the most massive and scientifically capable rover to the surface of Mars. To deliver this payload, an aerodynamic decelerator is required to decelerate the entry vehicle from supersonic to subsonic speeds, in advance of propulsive descent and touchdown on Mars. The aerodynamic deceleration will be accomplished by a mortar-deployed 21.5-m Viking-type disk-gap-band parachute (DGB), and will be the largest extra-terrestrial decelerator in the history of space exploration [1]. The parachute will deploy at up to Mach 2.2 and 750 Pa, resulting in the highest load and speed experienced by a parachute on Mars. The MSL parachute extends the envelope of the existing heritage deployment space in terms of load, size and Mach number. This has created the challenge of leveraging the existing heritage supersonic- high-altitude database, implementing a ground-based qualification program, and quantifying known aerodynamic instabilities associated with supersonic operation in the Mach regime of the MSL deployment. To address these challenges MSL has embarked upon a physics-based modeling and validation program to explore the fundamental physics associated with DGB-parachute operation in supersonic flow. The functional dependence of parachute performance and stability on Mach number, Reynolds number, parachute size, entry-vehicle size and parachute to entry vehicle proximity, is under investigation. The quantitative understanding garnered from this analytical effort will be used to leverage the existing heritage database of the Viking Lander, Viking Balloon Launched Decelerator Test (BLDT), Mars Pathfinder (MPF) and Mars Exploration Rover (MER) programs for the larger scale, deployment conditions, and modern construction techniques of the MSL parachute system. The physics-based modeling and validation effort includes the development of a coupled fluid and structural solver, i.e. fluid-structure-interaction code, and supersonic wind-tunnel experiments with subscale representations of the flight configuration.

Particle and continuum method comparison of a high-altitude, extreme-mach-number reentry flow

Stardust reentry flows have been simulated at an altitude of 80 km for a freestream velocity of 12.8 km/s using direct simulation Monte Carlo (DSMC) and computational fluid dynamics (CFD). Five ions and electrons were considered in the flowfield, and ionization processes were modeled in DSMC. The ion-averaged velocity method in DSMC was validated to maintain charge neutrality in the shock. Collision and energy-exchange models for DSMC were reviewed to ensure adequacy for the high-energy flow regime. Accurate electron-heavy particle collision cross sections and an electron-vibration relaxation model using Lees relaxation time were implemented in DSMC. Although the DSMC results agreed well with CFD for the collision-only case, discrepancies between DSMC and CFD were observed in the shock with the relaxation model activated. Furthermore, with full chemical reactions and ionization processes, DSMC results were compared with CFD. It was found that the assumption of electron temperature is crucial for the prediction of degree of ionization. At 80 km, the degree of ionization predicted by DSMC was found to be approximately 5 %, but in CFD, the degree of ionization is greater than 25 % for the case of T e = T tr and 9% for the case of T e = T vib· In DSMC, the electron-vibration relaxation model was found to be important to predict electron and vibrational temperatures at this altitude, and the electron temperature is the same order as the vibrational temperature. Therefore, compared to the DSMC solution, the assumption of T e = T vib is preferable in CFD. In addition, using the Mott-Smith model, good agreement was obtained between the analytical bimodal distribution functions and DSMC velocity distributions. An effective temperature correction in the relaxation and chemical reaction models using the Mott-Smith model may reduce the continuum breakdown discrepancy between DSMC and CFD inside the shock in terms of degree of ionization and temperatures, but a general implementation is not clear.

CFD Analysis of CUBRC Base Flow Experiments

This paper presents results from a computational analysis of a series of experiments conducted in the CUBRC 48” reflected shock tunnel for the purpose of studying the transition of afterbody wake flows. The experiments examined the flow over a spherically-blunted capsule, roughly chosen to be a scale representation of the new Orion crew module currently being designed by NASA. In this study, we have focused on three test runs for analysis, corresponding to low, medium, and high Reynolds number conditions. Each case has been examined with the intent of understanding the influence of turbulence modeling, time accuracy, and flux discretization on solution accuracy. Numerically, we find a standard first order in time, second order in space (with modified Steger-Warming fluxes) to be adequate for the external, reentry type flows of interest. For the low Reynolds number case (ReD ≤ 10 6 ), a laminar Navier-Stokes simulation is sufficient to accurately capture the statistical character of the flowfield. At the medium Reynolds number (ReD =6 .3×10 6 ), the analysis indicates a possibly transitional wake where heat transfer predictions are bounded by laminar and turbulent simulations. At the high Reynolds number (ReD =1 0.8 × 10 6 ), the experimental data are best matched by a fully turbulent calculation using the Detached Eddy Simulation (DES) form of the Spalart-Allmaras turbulence model. Additionally, comparisons between Reynolds-averaged Navier-Stokes (RANS) and DES calculations reveal the inability of RANS models to correctly capture either the rapidly fluctuating flow transients or mean field. Repeating the DES calculation with an alternative second order in time, low-dissipation flux scheme shows a substantial improvement in the simulation’s ability to capture flow transients and may prove crucial for applications in which this is a priority. It is observed that the improvements seen in moving from RANS to DES and from Steger-Warming to the low-dissipation scheme are related in that both ultimately reduce the level of unnecessary dissipation present in the simulation.