Simulation Of Hypersonic Flow Research Articles

Aerodynamics simulations of three-dimensional inviscid flow using curvilinear discontinuous Galerkin method on unstructured meshes

Over the last decades, the discontinuous Galerkin (DG) method has demonstrated its excellence in accurate, higher-order numerical simulations for a wide range of applications in aerodynamics simulations. However, the development of practical, computationally accurate flow solvers for industrial applications is still in the focus of active research, and applicable boundary conditions and fluxes are also very important parts. Based on curvilinear DG method, we have developed a flow solver that can be used for solving the three-dimensional subsonic, transonic and hypersonic inviscid flows on unstructured meshes. The development covers the geometrical transformation from the real curved element to the rectilinear reference element with the hierarchical basis functions and their gradient operation in reference coordinates up to full third order. The implementation of solid wall boundary conditions is derived by the contravariant velocities, and an enhanced algorithms of Harten-Lax-van Leer with contact (HLLC) flux based on curved element is suggested. These new techniques do not require a complex geometric boundary information and are easy to implement. The simulation of subsonic, transonic and hypersonic flows shows that the linear treatment can limit the accuracy at high order and demonstrates how the boundary treatment involving curved element overcomes this restriction. In addition, such a flow solver is stable on a reasonably coarse meshes and finer ones, and has good robustness for three-dimensional flows with various geometries and velocities. For engineering practice, a reasonable accuracy can be obtained at reasonably coarse unstructured meshes.

Application of a Gas-Kinetic BGK Scheme in Thermal Protection System Analysis for Hypersonic Vehicles

One major problem in the development of hypersonic vehicles is severe aerodynamic heating; thus, the implementation of a thermal protection system is required. A numerical investigation on the reduction of aerodynamic heating using different thermal protection systems is conducted using a novel gas-kinetic BGK scheme. This method adopts a different solution strategy from the conventional computational fluid dynamics technique, and has shown a lot of benefits in the simulation of hypersonic flows. To be specific, it is established based on solving the Boltzmann equation, and the obtained gas distribution function is used to reconstruct the macroscopic solution of the flow field. Within the finite volume framework, the present BGK scheme is specially designed for the evaluation of numerical fluxes across the cell interface. Two typical thermal protection systems are investigated by using spikes and opposing jets, separately. Both their effectiveness and mechanisms to protect the body surface from heating are analyzed. The predicted distributions of pressure and heat flux, and the unique flow characteristics brought by spikes of different shapes or opposing jets of different total pressure ratios all verify the reliability and accuracy of the BGK scheme in the thermal protection system analysis.

Numerical simulation of hypersonic thermochemical nonequilibrium flows using nonlinear coupled constitutive relations

To predict aeroheating performance of hypersonic vehicles accurately in thermochemical nonequilibrium flows accompanied by rarefaction effect, a Nonlinear Coupled Constitutive Relations (NCCR) model coupled with Gupta’s chemical models and Park’s two-temperature model is firstly proposed in this paper. Three typical cases are intensively investigated for further validation, including hypersonic flows over a two-dimensional cylinder, a RAM-C II flight vehicle and a type HTV-2 flight vehicle. The results predicted by NCCR solution, such as heat flux coefficient and electron number densities, are in better agreement with those of direct simulation Monte Carlo or flight data than Navier-Stokes equations, especially in the extremely nonequilibrium regions, which indicates the potential of the newly-developed solution to capture both thermochemical and rarefied nonequilibrium effects. The comparisons between the present solver and NCCR model without a two-temperature model are also conducted to demonstrate the significance of vibrational energy source term in the accurate simulation of high-Mach flows.

Simulations of Dusty Flows over Full-Scale Capsule During Martian Entry

Suspended dust particles can have adverse effects on heat shields during Mars atmospheric entry. For instance, they can increase both surface heating and surface recession. Current studies on this subject often make many simplifications, focus on a specific physical process, or include flow conditions different from those on Mars. To address this, we perform computational fluid dynamics simulations of hypersonic dusty flows over a full-scale entry capsule based on a realistic flight trajectory. The simulations are performed with an Euler–Lagrange methodology in which the carrier gas solution is obtained with a discontinuous Galerkin method. We employ a recently developed physics-based drag correlation that obtains better agreement with experimental data than popular existing correlations. To improve understanding of the relevant physical processes, which are still not very well understood, the effects of various factors, such as particle size distribution, angle of attack, and two-way coupling, on dust-induced heating augmentation and erosion are assessed.

Stagnation point heat flux characterization under numerical error and boundary conditions uncertainty

The numerical simulation of hypersonic atmospheric entry flows is a challenging problem. Prediction of quantities of interest, such as surface heat flux and pressure, is strongly influenced by the mesh quality using conventional second-order spatial accuracy schemes while depending on boundary conditions, which may generally suffer from uncertainty. This paper illustrates one of the first systematic quantification of the numerical error and the uncertainty-induced variability for the simulation of hypersonic flows. Specifically, a mesh-convergence study using grid adaptation tools is coupled with surrogate-based approaches to Uncertainty Quantification. The illustrative example is the simulation of the EXPERT vehicle of the European Space Agency employing the US3D solver. First, we show the benefits in using mesh adaptation to simulate hypersonic flows under uncertainty. On the one hand, this practice reduces the numerical uncertainty associated with each prediction and, on the other hand, allows us to obtain a more reliable surrogate model for Uncertainty Quantification by preventing non-physical heat flux values.Secondly, we perform a sensitivity analysis to compare the numerical uncertainty associated with a given mesh with the UQ-induced variability for a specific quantity of interest. In the case considered, the impact of the numerical uncertainty turned out to be at least one order of magnitude less than the quantity of interest variability. This result indicates the possibility of using coarse and adapted meshes for future UQ studies.

Unsteady behavior and thermochemical non equilibrium effects in hypersonic double-wedge flows

The paper presents highly detailed two-dimensional numerical simulations of hypersonic nitrogen and air flows around a double wedge, characterized by a complex shock wave/boundary layer interaction (SWBLI). The aim of the study is on one hand to investigate the unsteady behavior of such an interaction and on the other to evaluate thermo-chemical non-equilibrium effects. Two different thermo-physical models have been considered, the classical multi-temperature, or mode approximation, and the state-to-state vibrational kinetics. It is shown that for the geometric configuration under investigation, the occurrence of unsteady periodic behavior depends on the flow conditions and is inhibited in the high enthalpy case. Furthermore, non-equilibrium effects become relevant in the case of high enthalpy flows and discrepancies between the results obtained with the two models are observed.

Numerical simulation of hypersonic flow with non-equilibrium chemical reactions around sphere

High-temperature effects have a significant impact on the characteristics of aircraft moving at super- and hypersonic speeds. Due to the complexity of setting up a physical experiment, methods of mathematical modelling play an important role in finding the characteristics of hypersonic aircraft. Numerical modelling of super- and hypersonic air flow around the sphere is carried out taking into account high-temperature effects and non-equilibrium chemical reactions. The mathematical model includes equations of gas dynamics, equations of the turbulence model and equations of chemical kinetics. A critical review of the various models that are used to find the stand-off distance is provided. The results of numerical calculations on the distribution of flow quantities in the shock layer and the stand-off distance at different freestream Mach numbers are presented. The results of numerical calculations in a wide range of freestream Mach numbers are compared with the experimental data and computational results of other researchers. Standoff distance is computed at various Mach number and heights during a flight in the atmosphere at hypersonic speed.

Assessment of a high-order shock-capturing central-difference scheme for hypersonic turbulent flow simulations

High-speed turbulent flows are encountered in most space-related applications (including exploration, tourism and defense fields) and represent a subject of growing interest in the last decades. A major challenge in performing high-fidelity simulations of such flows resides in the stringent requirements for the numerical schemes to be used. These must be robust enough to handle strong, unsteady discontinuities, while ensuring low amounts of intrinsic dissipation in smooth flow regions. Furthermore, the wide range of temporal and spatial active scales leads to concurrent needs for numerical stabilization and accurate representation of the smallest resolved flow scales in cases of under-resolved configurations. In this paper, we present a finite-difference high-order shock-capturing technique based on Jameson’s artificial diffusivity methodology. The resulting scheme is ninth-order-accurate far from discontinuities and relies on the addition of artificial dissipation close to large gradient flow regions. The shock detector is slightly revised to enhance its selectivity and avoid spurious activations of the shock-capturing term. A suite of test cases ranging from 1D to 3D configurations (namely, perfect-gas and chemically reacting shock tubes, Shu–Osher problem, isentropic vortex advection, under-expanded jet, compressible Taylor–Green Vortex, supersonic and hypersonic turbulent boundary layers) is analyzed in order to test the capability of the proposed numerical strategy to handle a large variety of problems, ranging from calorically-perfect air to multi-species reactive flows. Results obtained on under-resolved grids are also considered to test the applicability of the proposed strategy in the context of implicit Large-Eddy Simulations.

Gas-Kinetic Scheme Coupled with Turbulent Kinetic Energy Equation for Computing Hypersonic Turbulent and Transitional Flows

ABSTRACT For simulation of hypersonic turbulent and transitional flows at high Reynolds numbers, a gas-kinetic scheme (GKS) strongly coupled with the turbulent kinetic energy equation in shear stress transport (SST) k−ω turbulence model is developed. To extend the current method to transitional flows, Langtry-Menter transition model is added together with SST k−ω turbulence model. Comparisons among the numerical solutions from current GKS method coupled with turbulent kinetic energy equation, GKS method uncoupled with turbulent kinetic energy equation, Reynolds-Averaged Navier-Stokes (RANS) solutions with turbulence/transition models, and experimental data are conducted. It is concluded that the current method can significantly improve the accuracy of numerical results.

Ab initio simulation of hypersonic flows past a cylinder based on accurate potential energy surfaces

For the first time in the literature, we present 2D simulations of hypersonic flows around a cylinder obtained from accurate ab initio potential energy surfaces (PESs). We compare results obtained from a low fidelity (empirical) and a high fidelity (ab initio) PES, thus demonstrating the impact of PES accuracy on the entire aerothermodynamic field around the body. We observe that the empirical PES is not adequate to accurately reproduce rotational and vibrational relaxation in the hypersonic flow, both in the compression and expansion regions of the flow field. This approach, enabled by advancements in large-scale computing, paves the way to the direct simulation of hypersonic flows where the only modeling input is the PES that describes molecular interactions between the various air constituents. Such flow field simulations provide benchmark solutions for the validation of reduced-order models.

RESEARCH ON THE AUSM SPLITTING DSMC-IP METHOD FOR HYPERSONIC RAREFIED FLOW FIELD

The Information Preservation (DSMC-IP) method is improved by using Advection Upstream Splitting Method (AUSM) splitting scheme, to solve the problems that the Direct Simulation of Monte Carlo (DSMC) method is effected by a large statistical dissipation, and that he traditional IP method cannot simulate the strong shock wave accurately. The fluxes of correlation terms in governing equations are reconstructed by local Mach numbers as the standard, so that the calculation is more accurate in accordance with the flow characteristics on both sides of the shock wave, thus forming a new DSMC-IP method with high statistical accuracy and hypersonic flow simulation ability. This method is utilized to simulate supersonic flow around the cylinder and hypersonic flow around the nozzle with divergent angle, the results of the AUSM splitting DSMC-IP method is basically consistent with the results of DSMC method, however the former has lower statistical scatter, and the flow characteristics are clearer. The surface characteristic coefficients differences between the results and the experiment or reference values are less than 5.5%. It is proved that the AUSM splitting DSMC-IP method is accurate and effective in the hypersonic flow simulation.

Comparative studies of shock-wave boundary-layer interactions in Earth and Mars atmospheres

Investigations of ramp-induced shock-wave boundary-layer interaction have been carried out for real gas flows of air and carbon dioxide through hypersonic laminar flow simulations corresponding to Earth and Mars atmospheres. An in-house-developed solver, which accounts for the real gas effects, has been employed for these studies. Effects of various parameters like wall temperature, freestream stagnation enthalpy, freestream Mach number, and blunt leading edge are explored on the intensity of shock-wave boundary-layer interaction (SWBLI). In either case, an increase in separation length is observed with an increase in wall temperature and a decrease in Mach number as well as freestream stagnation enthalpy. Here, the intensity of alteration is always noted to have a higher percentage for the Mars gas model. Further, separation length is found to be almost equal for the same wall to total temperature ratio in both of the flow mediums. The present study also affirms the fact that the leading edge bluntness can be used as a tool to reduce the size of the separation region in these planetary atmospheres. Revised correlations have been proposed for hypersonic Earth atmospheric flow with real gas effects to predict the extent of upstream influence and separation bubble size. The outcomes of simulations have also helped to device new correlations for these flow features of SWBLI for Mars atmospheric conditions. In all, the need for consideration of real gas effects and an exclusive real gas flow solver for the Mars atmosphere are the prominent recommendations of current studies.

Numerical simulation of hypersonic reaction flows with nonlinear coupled constitutive relations

As all known, there exists extremely thermal and chemical non-equilibrium phenomenon in hypersonic flows. Much effort has been put into the development of computational models for the prediction of such non-equilibrium flow. In order to obtain the real physical solution of the challenging non-equilibrium problems, Eu proposed a set of generalized hydrodynamic equations (GHE) from the viewpoint of generalized hydrodynamics. Based on Eu's GHE equations, Myong developed an efficient computational model, named as the nonlinear coupled constitutive relations (NCCR). The present study is based on the 3-D FVM NCCR model in conjunction with 7-species chemical reaction model to consider the reacting gas effect for hypersonic reentry vehicles. For further examining the performance of the proposed computational model, the typical engineering cases of Radio Attenuation Measurements (RAM)-C II vehicle are adopted. The flow contours, distribution profiles along stagnation line and surface properties are compared between nonlinear coupled constitutive model and those of the Navier-Stokes equation. The accuracy and physical consistency of the NCCR model are critically examined and validated for hypersonic reaction flow in the continuum region. For the near continuum region, the NCCR model can predict non-equilibrium phenomenon more accurately than NS equation through the comparison of shock stand-off distance and electron number density. The comparisons between NCCR model with one component perfect gas and multi-species reacting gas effect are also made to show the importance of chemical reaction source term.

Immersed boundary conditions for hypersonic flows using ENO-like least-square reconstruction

In this study, we present an immersed boundary reconstruction method for the simulation of viscous compressible flows presenting strong discontinuities in the vicinity of the immersed body. The technique we propose is based on a weighted least-square reconstruction. Its major feature is that the weights are automatically adapted to the presence of a discontinuity within the least-square stencil by using a shock-sensor approach. In so doing, the reconstruction becomes ENO-like and mathematically consistent even in the presence of strong discontinuities. The objective of this paper is to demonstrate that it then becomes possible to use an immersed boundary method for the simulation of hypersonic flows and that it is robust, even for flows around complex geometries. To do so, two-dimensional compressible Navier-Stokes equations are solved to study canonical hypersonic flows and it is showed that the results are in good agreement with either experimental data and/or body-fitted approaches and that the impact of the immersed boundary technique on the overall spatial order of convergence is reasonable.

HTR-1.2 solver: Hypersonic Task-based Research solver version 1.2

We present an updated version of the open-source Hypersonics Task-based Research (HTR) solver for hypersonic aerothermodynamics. The solver, whose first version was presented in Di Renzo et al. (2020), is designed for direct numerical simulation (DNS) of canonical hypersonic flows at high Reynolds numbers in which thermo-chemical effects induced by high temperatures are relevant. The solver relies on high-order spatial discretization on structured meshes and efficient time integrators for stiff systems within the Regent/Legion software stack, which makes the code highly portable and scalable in CPU and GPU-based supercomputers. The new version herein presented includes several optimizations and new tools for data analysis, along with novel user option for hybrid skew-symmetric/targeted essentially non-oscillatory numerics, to offer higher computational efficiency and lower numerical dissipation at moderate Mach numbers, inclusion of a new combustion mechanism for methane and oxygen, and new recycling–rescaling inlet boundary conditions targeted to the simulation of fully developed turbulent boundary layers. New version program summaryProgram Title: Hypersonics Task-based Research solverCPC Library link to program files:https://doi.org/10.17632/9zsxjtzfr7.2Developer’s repository link:https://github.com/stanfordhpccenter/HTR-solver.gitLicensing provisions: BSD 2-clauseProgramming language: Regent, C++, and CUDAJournal Reference of previous version: Di Renzo, M., Fu, L., & Urzay, J. (2020). HTR solver: An open-source exascale-oriented task-based multi-GPU high-order code for hypersonic aerothermodynamics. Computer Physics Communications, 107262.Does the new version supersede the previous version?: YesReasons for the new version: Release of new featuresSummary of revisions:•New optional sixth-order hybrid scheme has been implemented (activated by the flag “hybridScheme” in the input file). The scheme combines the energy-preserving properties of a sixth-order skew-symmetric central difference scheme [1] in smooth flow regions with the shock-capturing properties of a sixth-order targeted essentially non-oscillatory (TENO) scheme at points where shocks are involved. The switch between the two schemes is controlled by a TENO sensor whose cutoff value is adapted based on the maximum value of a modified Ducros sensor [2] across the reconstruction stencil. If the flag “hybridScheme” is set to false, the numerical scheme will revert to the TENO6-A scheme released in the previous version of the solver [3];•New recycling–rescaling inflow boundary conditions [4,5] for the simulation of turbulent compressible boundary layers are now available to the user;•Support for Legion tracing, which significantly improves the strong scalability of the solver, has been implemented;•A diagnostic tool to monitor the time evolution of the flow variables in a subvolume of the computational domain is now available;•A single-step chemistry mechanism for methane/oxygen combustion has been added to the mixtures handled by the HTR solver;•Sample scripts for strong scaling have been added to the “testcases” directory;•Unit test and regression test suites have been added to the repository;•The input file scheme has been modified in order to reduce verbosity and increase flexibility in specifying the boundary conditions and type of gas mixture;•Hyperbolic sine stretching functions have been made available to users during the grid generation process;•Computationally intensive tasks have been ported to C++ and CUDA in order to achieve higher efficiency on all hardware;•Several optimizations of the tasks body and mapper have been implemented in order to increase the computational efficiency and reduce the memory footprint.Nature of problem: This code solves the Navier–Stokes equations at hypersonic Mach numbers including finite-rate chemistry for dissociating air and multicomponent transport. The solver is designed for direct numerical simulations (DNS) of transitional and turbulent hypersonic turbulent flows under high-enthalpy conditions, and it accounts for thermochemical effects (vibrational excitation and chemical dissociation).Solution method: This code uses a low-dissipation sixth-order schemes for the spatial discretization of the conservation equations on Cartesian stretched meshes. Time advancement is carried out by either an explicit method if chemistry is slow, hence not introducing additional stiffness, or by an operator-splitting algorithm whereby chemical production rates are handled implicitly.Additional comments including restrictions and unusual features: The HTR solver builds on the runtime Legion [6,7] and is written in the programming language Regent [8,9] developed at Stanford University. Instructions for installation of the components are provided in the README file enclosed with the HTR solver and in the Legion repository [6].

The formulation of the RANS equations for supersonic and hypersonic turbulent flows

AbstractAccurate prediction of supersonic and hypersonic turbulent flows is essential to the design of high-speed aerospace vehicles. Such flows are mainly predicted using the Reynolds-Averaged Navier–Stokes (RANS) approach in general, and in particular turbulence models using the effective viscosity approximation. Several terms involving the turbulent kinetic energy (k) appear explicitly in the RANS equations through the modelling of the Reynolds stresses in such approach, and similar terms appear in the mean total energy equation. Some of these terms are often ignored in low, or even supersonic, speed simulations with zero-equation models, as well as some one- or two-equation models. The omission of these terms may not be appropriate under hypersonic conditions. Nevertheless, this is a widespread practice, even for very high-speed turbulent flow simulations, because many software packages still make such approximations. To quantify the impact of ignoring these terms in the RANS equations, two linear two-equation models and one non-linear two-equation model are applied to the computation of five supersonic and hypersonic benchmark cases, one 2D zero-pressure gradient hypersonic flat plate case and four shock wave boundary layer interaction (SWBLI) cases. The surface friction coefficients and velocity profiles predicted with different combinations of the turbulent kinetic energy terms present in the transport equations show little sensitivity to the presence of these terms in the zero-pressure gradient case. In the SWBLI cases, however, comparisons show that inclusion of k in the mean flow equations can have a strong effect on the prediction of flow separation. Therefore, it is highly recommended to include all the turbulent kinetic energy terms in the mean flow equations when dealing with simulations of supersonic and hypersonic turbulent flows, especially for flows with SWBLIs. As a further consequence, since k may not be obtained explicitly in zero-equation, or certain one-equation, models, it is debatable whether these models are suitable for simulations of supersonic and hypersonic turbulent flows with SWBLIs.

A robust and contact preserving flux splitting scheme for compressible flows

Low-dissipation shock-capturing methods usually suffer from various forms of shock instabilities, such as the notorious carbuncle phenomenon, in simulations of hypersonic flows. Many strategies for tackling the shock instability of flux difference splitting schemes, such as HLLC and Roe schemes, have been proposed while little work has been undertaken on the shock instability of flux splitting schemes. In this paper, we are concerned with the shock instability of a convection-pressure flux splitting scheme. By means of a linearized perturbation analysis, it shows that all of the perturbations in the streamwise direction are damped, while the perturbations of density and shear velocity in the transverse direction are undamped. Due to the special symmetry, the two-dimensional Sedov blast wave problem is studied to demonstrate the relation between the transverse flux and shock instability. Based on the theoretical analyses, a shock stable flux splitting scheme is presented through introducing additional viscosities of entropy wave and shear wave to damp out all perturbations in the transverse direction. For restoring the contact surface and shear layer, a pressure-based function depending on the shock-wave strength is defined to control the amount and location of additional viscosities. Numerical results of several challenging test problems demonstrate that the new scheme has greatly improved the robustness without undermining the accuracy.

Modified Nonlinear Coupled Constitutive Relations Model for Hypersonic Nonequilibrium Flows

To obtain accurate physical description of nonequilibrium flows, Eu proposed a set of generalized hydrodynamic equations (GHEs) using a nonequilibrium ensemble method. However, Eu’s equations are very difficult to solve by employing the well-established numerical schemes for hyperbolic partial differential equations. Recently, a simplified nonlinear model based on Eu’s equations has been proposed by Myong for GHEs in conjunction with hyperbolic conservation laws, which is called the nonlinear coupled constitutive relations (NCCR) model. However, in the NCCR model, two specific terms in the evolution equation of heat flux are omitted for providing simplification, which reduces the accuracy of GHEs to some degree especially for hypersonic flows. To address this deficiency in NCCR, a modified NCCR+ model is proposed in this paper, which includes the two omitted terms in the NCCR model. NCCR+ model is applied to compute the hypersonic flow past an Apollo capsule at different altitudes and in expansion corners in a tube with various expansion angles. It is found that the computed NCCR+ solutions are almost identical to the Navier–Stokes (NS) solutions at low Knudsen numbers, but as the Knudsen number increases, the difference between the NCCR+ model solutions and the NS results increases rapidly and the NCCR+ solutions become closer in agreement with the direct simulation Monte Carlo data. By adding the two omitted terms in the NCCR model, the NCCR+ model also improves the overall accuracy of the original NCCR model. The results presented in this paper demonstrate the potential of the NCCR+ model in simulation of nonequilibrium hypersonic flows in both compression and expansion regions at moderate Knudsen numbers.

Multiple temperature model of nonlinear coupled constitutive relations for hypersonic diatomic gas flows

The rotational energy of diatomic gases would be activated by the process of intermolecular collisions in high-temperature hypersonic flows. In this paper, a multi-temperature nonlinear coupled constitutive model has been proposed for simulating the transfer of energy between translational and rotational motions in hypersonic non-equilibrium flows. In this model, the nonlinear coupled constitutive equations are modified by introducing a rotational energy relaxation model and a changeable viscosity ratio related to local temperature. To confirm its accuracy, the new model is applied to investigate steady shock wave structures and high-speed gas flows around a cylinder and across a flat plate. The computational results are compared with the multi-temperature Navier–Stokes (NS) equations, the direct simulation Monte Carlo (DSMC) solutions, and the experiment data. The final results show the new model would reproduce the NS results at low Knudsen numbers but behave quite differently from the NS results as the non-equilibrium degree is enhanced. The new model is in better agreement with the DSMC solutions and the experimental data than the NS solutions in the far-from-equilibrium regions, which demonstrates the potential of the new relaxation model in the simulation of hypersonic non-equilibrium flows.

HTR solver: An open-source exascale-oriented task-based multi-GPU high-order code for hypersonic aerothermodynamics

In this study, the open-source Hypersonics Task-based Research (HTR) solver for hypersonic aerothermodynamics is described. The physical formulation of the code includes thermochemical effects induced by high temperatures (vibrational excitation and chemical dissociation). The HTR solver uses high-order TENO-based spatial discretization on structured grids and efficient time integrators for stiff systems, is highly scalable in GPU-based supercomputers as a result of its implementation in the Regent/Legion stack, and is designed for direct numerical simulations of canonical hypersonic flows at high Reynolds numbers. The performance of the HTR solver is tested with benchmark cases including inviscid vortex advection, low- and high-speed laminar boundary layers, inviscid one-dimensional compressible flows in shock tubes, supersonic turbulent channel flows, and hypersonic transitional boundary layers of both calorically perfect gases and dissociating air. Program summaryProgram Title: Hypersonics Task-based Research solverProgram Files doi:http://dx.doi.org/10.17632/9zsxjtzfr7.1Licensing provisions: BSD 2-clauseProgramming language: RegentNature of problem: This code solves the Navier–Stokes equations at hypersonic Mach numbers including finite-rate chemistry for air dissociation along with multicomponent transport. The solver is designed for direct numerical simulations (DNS) of transitional and turbulent hypersonic turbulent flows at high enthalpies, and accounts for thermochemical effects such as vibrational excitation and chemical dissociation.Solution method: This code uses a low-dissipation sixth-order targeted essentially non-oscillatory (TENO) scheme for the spatial discretization of the conservation equations on Cartesian stretched grids. The time advancement is performed either with an explicit method, when the chemistry is slow and therefore does not introduce additional stiffness in the integration, or with an operator-splitting method that integrates the chemical production rates with an implicit discretization.Additional comments: The HTR solver builds on the runtime Legion [1] and is written in the programming language Regent [2] developed at Stanford University. Instructions for the installation of the components are provided in the README file enclosed with the HTR solver and in the Legion repository [1].