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NASA NTRS1985

NASA's hypersonic fluid and thermal physics program (Aerothermodynamics)

This survey paper gives an overview of NASA's hypersonic fluid and thermal physics program (recently renamed aerothermodynamics). The purpose is to present the elements of, example results from, and rationale and projection for this program. The program is based on improving the fundamental understanding of aerodynamic and aerothermodynamic flow phenomena over hypersonic vehicles in the continuum, transitional, and rarefied flow regimes. Vehicle design capabilities, computational fluid dynamics, computational chemistry, turbulence modeling, aerothermal loads, orbiter flight data analysis, orbiter experiments, laser photodiagnostics, and facilities are discussed.

Graves, R. A.↗
NASA NTRS1992

Development of a 3-D upwind PNS code for chemically reacting hypersonic flowfields

Two new parabolized Navier-Stokes (PNS) codes were developed to compute the three-dimensional, viscous, chemically reacting flow of air around hypersonic vehicles such as the National Aero-Space Plane (NASP). The first code (TONIC) solves the gas dynamic and species conservation equations in a fully coupled manner using an implicit, approximately-factored, central-difference algorithm. This code was upgraded to include shock fitting and the capability of computing the flow around complex body shapes. The revised TONIC code was validated by computing the chemically-reacting (M(sub infinity) = 25.3) flow around a 10 deg half-angle cone at various angles of attack and the Ames All-Body model at 0 deg angle of attack. The results of these calculations were in good agreement with the results from the UPS code. One of the major drawbacks of the TONIC code is that the central-differencing of fluxes across interior flowfield discontinuities tends to introduce errors into the solution in the form of local flow property oscillations. The second code (UPS), originally developed for a perfect gas, has been extended to permit either perfect gas, equilibrium air, or nonequilibrium air computations. The code solves the PNS equations using a finite-volume, upwind TVD method based on Roe's approximate Riemann solver that was modified to account for real gas effects. The dissipation term associated with this algorithm is sufficiently adaptive to flow conditions that, even when attempting to capture very strong shock waves, no additional smoothing is required. For nonequilibrium calculations, the code solves the fluid dynamic and species continuity equations in a loosely-coupled manner. This code was used to calculate the hypersonic, laminar flow of chemically reacting air over cones at various angles of attack. In addition, the flow around the McDonnel Douglas generic option blended-wing-body was computed and comparisons were made between the perfect gas, equilibrium air, and the nonequilibrium air results.

Tannehill, J. C.↗
NASA NTRS1974

A review of turbulence measurements in compressible flow

A survey of turbulence measurements in compressible flows is presented. The majority of turbulence measurements at super- and hypersonic speeds were made for the zero pressure gradient, turbulent boundary layer. It was found that the nondimensional turbulent stress terms for the zero pressure gradient flow appear to agree closely with equivalent incompressible measurements in the outer part of the boundary layer. The stress terms were nondimensionalized by the wall value of shear stress and plotted versus the distance from the wall, nondimensionalized by the boundary-layer thickness. Indirect evaluation of the total shear stress distribution from mean velocity measurements for both super- and hypersonic flows (zero pressure gradient, two-dimensional flows) indicate a near universal distribution. These total shear stress curves also agree very closely with measured incompressible shear stress distributions. Recent laser anemometer measurements of the turbulent Reynolds shear stress (puv), reported by Johnson and Rose for a Mach number 2.9 flow, are in reasonable agreement with the expected total shear stress curve over the outer 60% of the boundary layer.

Sandborn, V. A.↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results for a Blunt Mars Entry Vehicle Concept during Supersonic Retropropulsion

Simulations of unsteady supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle were performed using the OVERFLOW Computational Fluid Dynamics (CFD) solver. High-fidelity flow solver techniques, including Detached Eddy Simulation (DES) turbulence modeling and Adaptive Mesh Refinement (AMR), were employed to obtain improved realism in CFD predictions. Simulation conditions and geometry configurations were designed to match specific runs in the Descent System Study (DSS) wind tunnel testing (WTT) campaign. The accuracy of each simulation is assessed by direct comparison to experimental data. Comparisons of computational predictions of the SRP flowfield and bow shock shape to experimental schlieren imaging show reasonable prediction of mean shock shape, with approximately 10% similarity in shock standoff distance for selected conditions, as well as similarity in local, time-varying fluctuations of the shock-plume interaction. Comparisons of discrete measurements of surface pressure coefficient (Cp) indicate CFD accuracy within approximately 10% of the experiment across the majority of the model heatshield, with larger variations at some of the heatshield edge locations with stronger flow unsteadiness. Simulated unsteadiness of these chaotic flows, which were highly dynamic and multi-modal, was shown to be within 20-40% of experimentally-measured pressure standard deviation (SD) for the majority of the sampled locations.

Supersonic Retropropulsion↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results for a Blunt Mars Entry Vehicle Concept during Supersonic Retropropulsion

Simulations of unsteady supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle were performed using the OVERFLOW Computational Fluid Dynamics (CFD) solver. High-fidelity flow solver techniques, including Detached Eddy Simulation (DES) turbulence modeling and Adaptive Mesh Refinement (AMR), were employed to obtain improved realism in CFD predictions. Simulation conditions and geometry configurations were designed to match specific runs in the Descent System Study (DSS) wind tunnel testing (WTT) campaign. The accuracy of each simulation is assessed by direct comparison to experimental data. Comparisons of computational predictions of the SRP flowfield and bow shock shape to experimental schlieren imaging show reasonable prediction of mean shock shape, with approximately 10% similarity in shock standoff distance for selected conditions, as well as similarity in local, time-varying fluctuations of the shock-plume interaction. Comparisons of discrete measurements of surface pressure coefficient (Cp) indicate CFD accuracy within approximately 10% of the experiment across the majority of the model heatshield, with larger variations at some of the heatshield edge locations with stronger flow unsteadiness. Simulated unsteadiness of these chaotic flows, which were highly dynamic and multi-modal, was shown to be within 20-40% of experimentally-measured pressure standard deviation (SD) for the majority of the sampled locations.

Supersonic Retropropulsion↗
NASA NTRS1993

Near wake structure for a generic ASTV configuration

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using the direct simulation Monte Carlo method. Particular emphasis is given to the near wake flow and its sensitivity to rarefaction and other parametric variations. The flow conditions simulated are attainable in existing low-density hypersonic wind tunnels; that is, Mach 20 nitrogen flow encompassing freestream Knudsen numbers of 0.03 to 0.001. A stable vortex forms in the near wake at and below a freestream Knudsen number of 0.01 and the size of the vortex increases with decreasing freestream Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases. There is no formation of a lip separation shock or a distinct wake shock at these rarefied conditions.

Dogra, Virendra K.↗
NASA NTRSDate not supplied

Comparison Between DSMC and CFD for Hypersonic Planetary Entry Simulations

Hypersonic planetary entry flows span a wide range of Knudsen numbers between rarefied and continuum flows. While computational fluid dynamics (CFD) techniques cannot provide an accurate solution for flows in the rarefied regime, the direct simulation Monte Carlo (DSMC) method is capable of providing accurate solutions for flows in both in the rarefied and continuum regimes but becomes prohibitively expensive as the Knudsen number decreases. For the purpose of thermal protection systems (TPS) design and post-flight reconstruction, various selected points along an entry trajectory are often solved using hypersonic solvers. The quantities of interest that are obtained from that exercise are generally surface quantities, such as pressure, heat flux and enthalpy. Then, material response solvers are used to either design the heat shield to an optimal thickness based on a choice of material, or to provide in-depth heating profiles through the material at various select locations, and compare with flight instrumentation such as the ones that flew on NASA’s two most recent Mars missions, MSL and Mars2020. While most of the heating is generally experienced during the continuum part of the entry, the heating within the rarefied regime is significant for some atmospheres, and hence the flow solutions need to be computed using the DSMC method. Ensuring consistency between hypersonic CFD and the DSMC is crucial so that reliable surface quantities can be passed to material response solvers. Studies were performed to compare the two methods at various select locations, for both non-reacting argon flows as well reacting CO2/N2 flows. Preliminary conclusions show that, for non-reacting flows, the agreement between the two methods for surface heating is excellent (within expected uncertainties) for a freestream Knudsen number of 0.0006, and gets progressively worse as the Knudsen number increases to 0.06. Continuum breakdown analyses were performed and showed that, in general, the Gradient Length Local Knudsen number (KnGLL) from Boyd [1] and associated criteria (KnGLL > 0.05) can be accurately used to predict zones of breakdown in the flow, and associated errors for surface quantities. Furthermore, when studying reacting flow, our studies showed that while flow quantities are highly dependent on relaxation parameters and chemistry rates, it is possible to obtain a good agreement for surface heating, as long as the continuum breakdown in minimal (see Fig. 1). Finally, we studied the impact of the grid refinement on the surface quantities in DSMC.

DSMC↗
NASA NTRSDate not supplied

Comparison Between DSMC and CFD for Hypersonic Planetary Entry Simulations

Hypersonic planetary entry flows span a wide range of Knudsen numbers between rarefied and continuum flows. While computational fluid dynamics (CFD) techniques cannot provide an accurate solution for flows in the rarefied regime, the direct simulation Monte Carlo (DSMC) method is capable of providing accurate solutions for flows in both in the rarefied and continuum regimes but becomes prohibitively expensive as the Knudsen number decreases. For the purpose of thermal protection systems (TPS) design and post-flight reconstruction, various selected points along an entry trajectory are often solved using hypersonic solvers. The quantities of interest that are obtained from that exercise are generally surface quantities, such as pressure, heat flux and enthalpy. Then, material response solvers are used to either design the heat shield to an optimal thickness based on a choice of material, or to provide in-depth heating profiles through the material at various select locations, and compare with flight instrumentation such as the ones that flew on NASA’s two most recent Mars missions, MSL and Mars2020. While most of the heating is generally experienced during the continuum part of the entry, the heating within the rarefied regime is significant for some atmospheres, and hence the flow solutions need to be computed using the DSMC method. Ensuring consistency between hypersonic CFD and the DSMC is crucial so that reliable surface quantities can be passed to material response solvers. Studies were performed to compare the two methods at various select locations, for both non-reacting argon flows as well reacting CO2/N2 flows. Preliminary conclusions show that, for non-reacting flows, the agreement between the two methods for surface heating is excellent (within expected uncertainties) for a freestream Knudsen number of 0.0006, and gets progressively worse as the Knudsen number increases to 0.06. Continuum breakdown analyses were performed and showed that, in general, the Gradient Length Local Knudsen number (KnGLL) from Boyd and associated criterion (KnGLL > 0.05) seems conservative in predicting zones of breakdown in the flow, and associated errors for surface quantities. Updated criteria of KnGLL = 2.0 and 0.5 appear to be more appropriate for surface and flow quantities, respectively. Furthermore, when studying reacting flows, our studies showed that while flow quantities are highly dependent on relaxation parameters and chemistry rates, it is possible to obtain a good agreement for surface heating, as long as the continuum breakdown is minimal.

DSMC↗
NASA NTRS1971

An Analytical Investigation of an Oscillating Wedge in a Supersonic Perfect Gas Flow

Several aspects of the oscillating wedge are investigated to evaluate both the resulting trends for the wedge and methods of analyzing unsteady flows. An existing hypersonic small disturbance theory for an oscillating thin wedge is extended and applied. A perturbation method involving linearization about the known flow is then derived and discussed. Subsequently, a finite difference technique for calculating the complete unsteady flow field of the wedge in motion is presented and discussed in conjunction with some calculated quasi-static nonlinear trends.

Bennett, R. M.↗
NASA NTRS2012

Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

Georgiadis, Nicholas J.↗
NASA NTRS1986

Numerical analysis of some supersonic viscous flows related to inlet and nozzle systems

A numerical method originally developed for three-dimensional supersonic inlet flow calculations is extended and applied to the study of two-dimensional and three-dimensional flows associated with arbitrary propulsion systems. The method is based on the forward spatial marching solution of a reduced form of the three-dimensional steady Navier-Stokes equations in which streamwise pressure gradients are retained in both the subsonic and supersonic regions. The present paper briefly describes the analysis and then shows three applications. In the first application, a wall transpiration study has been performed for the two-dimensional shock wave/turbulent boundary layer interaction flow field with application to an inlet configuration. The second application treats a rectangular high speed inlet with a swept sideplate including the effects of sideplate spillage. Finally, the method is utilized to analyze the interaction of an under-expanded supersonic jet with an ambient flow. Computed results are examined and compared with available experimental measurements. It is demonstrated that the present numerical method is capable of numerically simulating complex two- and three-dimensional flows relevant to hypersonic propulsion systems in a manner which both shows good agreement with data when such data is available, and which shows the complex flow features in the absence of data.

Kim, Y.-N.↗
NASA NTRS1991

Simultaneous measurements of temperature and density in air flows using UV laser spectroscopy

The simultaneous measurement of temperature and density using laser-induced fluorescence of oxygen in combination with Q-branch Raman scattering of nitrogen and oxygen is demonstrated in a low-speed air flow. The lowest density and temperature measured in the experiment correspond to the freestream values at Mach 5 in the Ames 3.5-Foot Hypersonic Wind Tunnel for stagnation conditions of 100 atm and 1000 K. The experimental results demonstrate the viability of the optical technique for measurements that support the study of compressible turbulence and the validation of numerical codes in supersonic and hypersonic wind tunnel flows.

Fletcher, D. G.↗
NASA NTRS1991

A new technique for low-to-high altitude predictions of ablative hypersonic flowfields

A parabolized Navier-Stokes (PNS) technique for predicting chemically reacting reentry flows over a wide range of altitudes extending from zero to 259 kft is presented. This coupled-chemistry PNS scheme can use either a strongly implicit solution approach or a two-step solution approach and is used to study a hypersonic flow over various configurations over a Mach-number range of 6.89 to 20. It is demonstrated that the two-step solution provides a combination of accuracy, stability, and computational efficiency. Nonablating low-altitude calculations show that the flowfield is in chemical equilibrium only in the near-wall region, while finite-rate chemistry effects are significant over a major portion of the shock layer, and the Blottner-type linearization of the production terms is unstable for low-altitude high-Reynolds-number conditions.

Bhutta, Bilal A.↗
NASA NTRS1991

Constant-temperature anemometry measurements in hypersonic boundary layers

One of the major unresolved issues in fluid dynamics is the nature of apparent stresses, called Reynolds stresses, which occur in turbulent boundary layers. In hypersonic boundary layers, the flow physics is further complicated by the large temperature and density fluctuations and the concomitant contamination of the Reynold stresses by fictitious terms. Because of the severe flow environment and the extraordinary demands on sensors and instruments, the turbulence characteristics of hypersonic boundary layers were studied in only a cursory fashion. The plans for supersonic (HSCT) and hypersonic (NASP) vehicles made supersonic flow physics one of the critical pacing technologies in aerospace science. In particular, experimental data are needed to verify candidate computer models and to reach an important understanding of the turbulence physics. The presented research is the start of a substantial effort to refine existing instrumentation and develop experimental techniques to measure the various component of the Reynolds stress in hypersonic boundary layers.

Spina, Eric F.↗
NASA NTRS1993

Hypersonic engine component experiments in high heat flux, supersonic flow environment

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Even though progress has been made in the computational understanding of fluid dynamics and the physics/chemistry of high speed flight, there is also a need for experimental facilities capable of providing a high heat flux environment for testing component concepts and verifying/calibrating these analyses. A hydrogen/oxygen rocket engine heat source was developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to fulfill this need. This 'Hot Gas Facility' is capable of providing heat fluxes up to 450 w/sq cm on flat surfaces and up to 5,000 w/sq cm at the leading edge stagnation point of a strut in a supersonic flow stream. Gas temperatures up to 3050 K can also be attained. Two recent experimental programs conducted in this facility are discussed. The objective of the first experiment is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Macrophotographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight. The objective of the second experiment is to assess the capability of cooling a porous surface exposed to a high temperature, high velocity flow environment and to provide a heat transfer data base for a design procedure. Experimental results from transpiration cooled surfaces in a supersonic flow environment are presented.

Gladden, Herbert J.↗
NASA NTRS1975

Computer user's guide for a chemically reacting viscous shock-layer program

A description is given of the computer code for predicting viscous shock-layer flows over nonanalytic blunt bodies (Program VISLNABB) for hypersonic, low Reynolds number flows. Four specific and one general body geometries are considered. In addition to sphere-cones, cylinder wedges and geometries defined in tabular form, options for hyperboloids and paraboloids are included. Details of the theory and results are included in a separate engineering report. The program, subroutines, variables in common, and input and output data are described. Listings of the program code, output data for a sample case, and the input data for this sample case are included.

Miner, E. W.↗
NASA NTRS1980

Three-dimensional interactions and vortical flows with emphasis on high speeds

Diverse kinds of three-dimensional regions of separation in laminar and turbulent boundary layers are discussed that exist on lifting aerodynamic configurations immersed in flows from subsonic to hypersonic speeds. In all cases of three dimensional flow separation, the assumption of continuous vector fields of skin-friction lines and external-flow streamlines, coupled with simple topology laws, provides a flow grammar whose elemental constituents are the singular points: nodes, foci, and saddles. Adopting these notions enables one to create sequences of plausible flow structures, to deduce mean flow characteristics, expose flow mechanisms, and to aid theory and experiment where lack of resolution in numerical calculations or wind tunnel observation causes imprecision in diagnosing the three dimensional flow features.

Peake, D. J.↗
NASA NTRS1991

Hypersonic shock-wave turbulent-boundary-layer interaction flows - Experiment and computation

Solutions of the Reynolds-averaged Navier-Stokes equations are presented and compared with experimental surface data for a series of hypersonic shock-wave/turbulent-boundary-layer interaction flows. The turbulence models used include the standard k-epsilon two-equation eddy viscosity model, a two-layer modification to this model, and a third model with several modifications to account for compressibility effects. Both modified models gave significant improvements for all the test flows.

Horstman, C. C.↗