High Mach Number Research Articles

Inefficient star formation in high Mach number environments

Context. Predicting the star formation rate (SFR) in galaxies is crucial to understand their evolution and morphology. To do so requires a fine understanding of how dense structures of gas are created and collapse. In that, turbulence and gravity play a major role. Aims. Within the gravo-turbulent framework, we assume that turbulence shapes the interstellar medium (ISM), creating density fluctuations that, if gravitationally unstable, will collapse and form stars. The goal of this work is to quantify how different regimes of turbulence, characterized by the strength and compressibility of the driving, shape the density field. We are interested in the outcome in terms of SFR and how it compares with existing analytical models for the SFR. Methods. We ran a series of hydrodynamical simulations of turbulent gas. The simulations were first conducted without gravity, so that the density and velocity were shaped by the turbulence driving. Gravity was then switched on, and the SFR was measured and compared with analytical models. The physics included in these simulations was very close to the one assumed in the classical gravo-turbulent SFR analytical models, which makes the comparison straightforward. Results. We found that the existing analytical models convincingly agree with simulations at low Mach number, but we measure a much lower SFR in the simulation with a high Mach number. We develop, in a companion paper, an updated physically motivated SFR model that reproduces well the inefficient high Mach regime of the simulations. Conclusions. Our work demonstrates that accurate estimations of the turbulent-driven replenishment time of dense structures and the dense gas spatial distribution are necessary to correctly predict the SFR in the high Mach regime. The inefficient high-Mach regime is a possible explanation for the low SFR found in dense and turbulent environments such as the centre of our Milky Way and other galaxies.

Supersonic turbulence simulations with GPU-based high-order Discontinuous Galerkin hydrodynamics

ABSTRACT We investigate the numerical performance of a Discontinuous Galerkin (DG) hydrodynamics implementation when applied to the problem of driven, isothermal supersonic turbulence. While the high-order element-based spectral approach of DG is known to efficiently produce accurate results for smooth problems (exponential convergence with expansion order), physical discontinuities in solutions, like shocks, prove challenging and may significantly diminish DG’s applicability to practical astrophysical applications. We consider whether DG is able to retain its accuracy and stability for highly supersonic turbulence, characterized by a network of shocks. We find that our new implementation, which regularizes shocks at subcell resolution with artificial viscosity, still performs well compared to standard second-order schemes for moderately high-Mach number turbulence, provided we also employ an additional projection of the primitive variables on to the polynomial basis to regularize the extrapolated values at cell interfaces. However, the accuracy advantage of DG diminishes significantly in the highly supersonic regime. Nevertheless, in turbulence simulations with a wide dynamic range that start with supersonic Mach numbers and can resolve the sonic point, the low-numerical dissipation of DG schemes still proves advantageous in the subsonic regime. Our results thus support the practical applicability of DG schemes for demanding astrophysical problems that involve strong shocks and turbulence, such as star formation in the interstellar medium. We also discuss the substantial computational cost of DG when going to high order, which needs to be weighted against the resulting accuracy gain. For problems containing shocks, this favours the use of comparatively low DG order.

Validation of a Methodology to Assess The Flutter Limit Cycle Oscillation Amplitude of Low-Pressure Turbine Bladed-Disks - Part I: Mach Number Effects

Abstract This paper presents a methodology to estimate the vibration amplitude of fluttering low-pressure turbine blades saturated due to friction effects. The study utilizes an analytical model that balances aerodynamic work and dry-friction work. The analytical predictions are compared against experimental results to validate the model. The first part of this paper focuses on the influence of the Mach number on the work balance between aerodynamic and mechanical components. It is observed that the vibration amplitude of low-pressure turbine rotor blades notably increases with higher Mach numbers. In addition, numerical simulations are employed to assess the influence of the Mach number on the critical damping ratio. The results demonstrate that an appropriate scaling of the critical damping ratio with the exit Mach number collapses all the damping versus inter-blade phase angle curves into a single curve. This finding validates the scaling of the aerodynamic damping for different pressure ratios. Unsteady pressure measurements were acquired, carefully post-processed to extract their flutter-induced peak components, and presented in a nodal diameter by nodal diameter basis. The post-processed data were then used to characterize the vibration amplitude observed in the experiments. The trends of the measured unsteady pressure on the casing of a rotating rig and the proposed model with the Mach number for different shaft speeds are in good agreement. The vibration amplitude and the mean unsteady pressure increase with the Mach number and exhibit a maximum with the shaft speed.

Scaling of coherent structures in compressible wall-bounded turbulence

Semi-local scales have been widely used in compressible wall-bounded turbulence, but it is still unclear whether they are applicable to the scaling of coherent structures, especially under conditions of high Mach number and cold wall temperature. By scrutinizing the direct numerical simulation dataset at different Mach numbers and wall temperatures, this paper demonstrates that the coherent structures normalized by semi-local scales are universal in size. In addition to this, we find that the ratios of Kolmogorov scales to semi-local scales are independent of Mach number and wall temperature. Thus, Kolmogorov scales can achieve the same scaling effect as the semi-local scales. The velocity spectra are also compared to verify the current scaling method quantitatively. A method to determine the threshold for the vortex identification criterion is proposed, allowing the same threshold for different cases to obtain vortices of similar size. The scaling of other statistics including turbulent kinetic energy, streamwise Reynolds normal stress, and root mean square of fluctuating vorticity is also investigated. A new velocity scale is proposed based on the total-stress-based transformation for mean streamwise velocity, which can collapse the profiles of these statistics more accurately than the semi-local velocity scale. The present paper demonstrates that through appropriate normalization, the structures and statistics of compressible turbulence become universal, reaffirming the validity of Morkovin's hypothesis even for the present high Mach number and cold wall cases.

Study on the influence of high-temperature thermochemical nonequilibrium effects on the flow field characteristics around shaped wing

With advances in research on hypersonic vehicles, the precise simulation of the effects of thermochemical non-equilibrium has become increasingly important in their design. In light of this, this study explores the influence of high-temperature thermochemical non-equilibrium on the characteristics of the flow field around the hypersonic wings of an aircraft. We initially conducted a numerical simulation by using the model of flow through a cylinder to validate the accuracy and reliability of an 11-species gas model in representing high-enthalpy flow fields. Subsequently, a systematic analysis was conducted on the impact of thermochemical nonequilibrium effects on the temperature, pressure, and enthalpy distribution in the flow field around a symmetric diamond wing under different Mach numbers and angles of attack. The research results indicated that the deeper reason behind the differing thermochemical nonequilibrium effects in the flow field at various Mach numbers lies in the distinct distribution of enthalpy of the air components at different locations, which provided a new perspective for understanding flow field variations from the standpoint of enthalpy. It is disclosed that the thermochemical non-equilibrium significantly altered the characteristics of the flow field, particularly at high Mach numbers and angles of attack, with a significant impact on the aerodynamic parameters of both the windward and the leeward sides of the wing. Through explaining the mechanisms of thermochemical non-equilibrium under flow fields with different structures, this study provides a theoretical foundations for and a fresh perspective on the design of hypersonic vehicles.

Thermodynamic cycle performance modeling and numerical simulation of higher Mach scramjet with inlet pre-injection

This paper focuses on the thermodynamic cycle performance modeling and numerical investigation of a high Mach number scramjet with inlet pre-injection of hydrogen. The scramjet model used in this study was the Hyshot-Ⅱ engine model. Flight experiments using this model at high Mach number have been conducted previously, which provides a unique testbed for validating the computational prediction of supersonic combustion. A thermodynamic cycle performance model with inlet pre-injection was developed for operating characteristics studies of the scramjet engine. A set of three-dimensional Reynolds average Navier-Stockes (RANS) simulations of the reactive flow was performed with a 9-specie and 19-step kinetic mechanism for hydrogen combustion. The turbulence is modeled by k-ω shear stress transport (SST) turbulence model. The thermodynamic cycle performance analysis results indicate that inlet pre-injection can enhance the hydrogen mixing efficiency, and a small amount of advanced heat release can improve the overall thermal cycle efficiency. Detailed flow analysis from the three-dimensional simulation results indicates flow can be ignited in the first shock wave induced boundary layer separation (SIBLS) region. Inlet pre-injection enables the propagation of the flame toward the core flow compared to the throat injection. The inlet fuel pre-injection proved to be a promising injection method due to the higher combustion efficiency, elevated from 0.75 to 0.85, and higher specific impulse potential, increased from 1570 s to 1810 s. As the ER of inlet pre-injection increases within the examined range, the performance curve exhibits an optimal point. As the geometric compression ratio increases, the entire combustion process advances, leading to a significant increase in the performance potential of the engine. The comparative analysis of the thermodynamic cycle performance modeling and computational fluid dynamics (CFD) modeling indicates the model shows a good prediction on the performance of the scramjet engine with the inlet pre-injection, with the same trend of performance curves and an error of less than 10 % compared to the numerical simulation.

Effect of nonlinear terms in piston theory on characteristics of panel flutter

Currently, there is no systematic summary of the influence of the nonlinear terms in piston theory on the panel flutter. The two-dimensional nonlinear panel flutter equations under supersonic airflow based on the first, second, and third-order piston theories is established. The stability of the heated panel is analyzed by using the Lyapunov's indirect method, and the panel flutter equations are numerically solved based on the numerical analysis method to investigate the influence of the nonlinear terms in piston theory on the panel flutter. The results show that: ①Under a small temperature rise ratio, the panel response under the first-order piston theory only exhibits the convergent motion and single-period limit cycle flutter. While under higher-order piston theories, the panel response is more complex, exhibiting more complex nonlinear dynamic phenomena such as multi-period limit cycle flutter and chaotic motion in addition to the aforementioned characteristics. ②As the Mach number increases, the required dynamic pressure and temperature rise ratio decrease gradually when the dynamic response of the panel under the first-order piston theory and higher-order piston theories exhibit significant differences. ③The significant differences in the computational results of the second-order and third-order piston theory appear under the high Mach numbers and relatively high dynamic pressures, and the same phenomenon occurs underthe high temperature rise ratios. ④When the dynamic response characteristics of the panel are basically consistent, the displacement response peak calculated by using the higher-order piston theory is usually smaller than the result calculated by using the first-order piston theory, and the maximum error can reach about 16.66%. The present results have certain reference value for selecting the appropriate analysis method for the panel flutter under different conditions in practical applications.

Linear stability analysis of compressible boundary layer over an insulated wall using compound matrix method: Existence of multiple unstable modes for Mach number beyond 3

The linear spatial stability of a parallel two-dimensional (2D) compressible boundary layer on an adiabatic plate is investigated by considering both 2D and three-dimensional (3D) disturbances. The compound matrix method is employed here, for the first time, for compressible flows, which, unlike other conventional techniques, can efficiently eliminate the stiffness of the equations governing the spectral amplitudes. The method is first validated with published results in the literature corresponding to spatial and temporal instability of flows ranging from low subsonic to high supersonic Mach numbers (M), which shows a good match depending upon the proper choice of free-stream temperature and the wall dispersion relation. Subsequently, flow compressibility effects and the spanwise variation of disturbances are also investigated for M ranging from low subsonic to high supersonic cases (from M = 0.1 to 6). Mack (AGARD Report No. 709, 1984) reported the existence of two unstable modes for M > 3 from viscous calculations (the so-called “second mode”) that subsequently fuse to create only one unstable zone when M increases. Our calculations show a series of higher-order unstable modes for M > 3 in addition to the findings of Ma and Zhong [“Receptivity of a supersonic boundary layer over a flat plate. Part 1. Wave structures and interactions,” J. Fluid Mech. 488, 31–78 (2003)], where such higher-order modes for supersonic boundary layers are all noted to be spatially stable. The number and the frequency extent of the corresponding unstable zones increase with an increase in M beyond 3 while propagating downstream at a higher speed than those corresponding to incompressible, subsonic, and low supersonic (M < 2) cases.

Integrated Waverider Forebody/Inlet Fusion Method Based on Discrete Point Cloud Reconstruction

The integrated design of waverider forebodies and inlets is considered a critical challenge in high Mach number vehicle development. To facilitate the rapid construction of integrated geometrical models for waverider forebodies and inlets during the conceptual design phase, a method based on discrete point cloud reconstruction has been proposed. In this method, the geometries of the waverider body and inlet are used as inputs and decomposed into the point cloud under discrete rules. This point cloud is refitted to generate new section lines, which are then lofted into an integrated shape under the constraints of guide curves. By modifying the coordinates of the point cloud positions, the geometric configuration of the integrated shape can be rapidly adjusted, providing initial support for subsequent aerodynamic optimization and thermal protection. Using this method, an integrated approach was applied to a waverider forebody and inward-turning inlet in a tandem configuration. This achieved body-inlet matching and integration, resulting in a 15.6% improvement in the inlet’s total pressure recovery coefficient. The integration time was reduced to just 3.18% of the time required for traditional manual adjustments. Additionally, optimization based on the discrete point cloud enhanced the lift-to-drag ratio by 7.83%, demonstrating the feasibility of the proposed method.

A Structure-Preserving Semi-implicit IMEX Finite Volume Scheme for Ideal Magnetohydrodynamics at all Mach and Alfvén Numbers

We present a divergence-free semi-implicit finite volume scheme for the simulation of the ideal magnetohydrodynamics (MHD) equations which is stable for large time steps controlled by the local transport speed at all Mach and Alfvén numbers. An operator splitting technique allows to treat the convective terms explicitly while the hydrodynamic pressure and the magnetic field contributions are integrated implicitly, yielding two decoupled linear implicit systems. The linearity of the implicit part is achieved by means of a semi-implicit time linearization. This structure is favorable as second-order accuracy in time can be achieved relying on the class of semi-implicit IMplicit–EXplicit Runge–Kutta (IMEX-RK) methods. In space, implicit cell-centered finite difference operators are designed to discretely preserve the divergence-free property of the magnetic field on three-dimensional Cartesian meshes. The new scheme is also particularly well suited for low Mach number flows and for the incompressible limit of the MHD equations, since no explicit numerical dissipation is added to the implicit contribution and the time step is scale independent. Likewise, highly magnetized flows can benefit from the implicit treatment of the magnetic fluxes, hence improving the computational efficiency of the novel method. The convective terms undergo a shock-capturing second order finite volume discretization to guarantee the effectiveness of the proposed method even for high Mach number flows. The new scheme is benchmarked against a series of test cases for the ideal MHD equations addressing different acoustic and Alfvén Mach number regimes where the performance and the stability of the new scheme is assessed.

Considerations for Compressible Film Cooling: A Computational Study of the Effects of Transonic Flows and Varying Mainstream Mach Number

Abstract Recent evidence suggests that film cooling flows with engine-realistic mainstream Mach number have declined performance in comparison to those with conventional low-speed laboratory conditions. Consideration and understanding of these effects are fundamental to improving film cooling research. This computational study investigates the film cooling performance of a 7-7-7 shaped film cooling hole with respect to varying mainstream Mach number, with constant Reynolds number. The cases studied include mainstream Mach numbers from 0.15 to 0.75, with a fixed, engine realistic, hole Reynolds number of Red=10,100. Significant results are then evaluated against varying stagnation temperature ratio and blowing ratio. The results showed that at a blowing ratio of 1.75, the adiabatic effectiveness declines significantly with increasing mainstream Mach number. The decreased performance is due to supersonic flows and shocks in the film cooling hole that disrupt flow in the diffuser section of the hole. These characteristics are observed across all stagnation temperature ratios considered. In addition to these insights, the study discusses the importance of proper thermal scaling and definition of adiabatic effectiveness when operating at high mainstream Mach number. It is demonstrated that the effects of high-speed flow challenge the efficacy of the conventional parameters used to characterize film cooling performance.

A novel simplified simulation method for the precooler based on the virtual source term: modeling, validation and application

Precooling technology is one of the promising approaches for expanding the flight envelope of conventional turbine engines and enhancing flight performance at high Mach numbers. Within this domain, a crucial challenge is to rapidly and accurately obtain the flow and heat transfer characteristics of the precooler. Currently, full-size simulation of the precooler is impractical, while simplified and fast simulation methods employed in previous research have exhibited certain deficiencies, including excessive simplification and unconvincing validation. In this study, a novel simplified simulation method for the annular precooler based on the virtual source term model was proposed. This method comprehensively considered the axial and radial non-uniformity of airflow caused by the coupling effect between the precooler and the inlet channel in the actual flow field. Subsequently, systematic validations of the reliability and superiority of this method were conducted, along with its application in the overall performance calculation of the precooled engine. The results demonstrate that the simplified method proposed in this paper showcases higher accuracy than the methods employed in previous research. Compared with the real model, the relative errors of the overall parameters are within ± 1 %, and the relative errors of the distortion indices for total temperature and total pressure are 5.1 % and −1.5 %, respectively. Meanwhile, the number of computational mesh cells is reduced by more than 95 %. Moreover, the first application of the simplified method in the overall performance calculation of the precooled engine reveals that the traditional low-dimensional model underestimates both the total pressure loss of the precooler and the fuel consumption for precooling, resulting in larger calculation deviations in the thrust and equivalence ratio, respectively. The values of thrust calculated with the high-dimensional virtual model show an average decrease of 2.01 % compared to those derived from the low-dimensional model. In addition, a higher Mach number results in lower effectiveness, higher air inlet temperature, and reduced resistance of the precooler to heat exchange imbalance caused by flow non-uniformity. These factors lead to an escalating relative deviation of the equivalence ratio between the two models, reaching 26.45 % at Mach 5.

Effect of non-equilibrium thermochemistry on Pitot pressure measurements in shock tunnels (or: Is 0.92 really the magic number?)

Pitot pressure is the most common measurement in high total enthalpy shock tunnels for test condition verification. Nozzle calculations using multi-temperature non-equilibrium thermochemistry are needed in conjunction with Pitot measurements to quantify freestream properties. Pitot pressure is typically matched by tuning the boundary layer transition location in these simulations. However, non-equilibrium thermochemistry effects on the Pitot probe are commonly ignored. A computational study was undertaken to estimate the effect of non-equilibrium thermochemistry on Pitot pressure and freestream conditions. The test flow was produced by a Mach 7 nozzle in a reflected shock tunnel for air at a relatively low total enthalpy of 2.67MJ/kg. Three different thermochemical models (equilibrium, finite-rate chemistry and two-temperature thermochemistry) were employed to compute flow variables at the nozzle exit and Pitot probe. Pitot pressures from these simulations were compared against those obtained via experiments. The results show a departure from the commonly utilized C of 0.92 in the reduced Rayleigh-Pitot equation form for high Mach numbers. Additionally, calculations were done with a sweep of free-stream conditions and resulting in values for one- and two-temperature models to use in future shock tunnel studies. Overall, our results show that the influence of finite-rate thermochemistry should be taken into account, even at relatively low flow enthalpies.

Optimization of Thrust of a Generic X-51 Hypersonic Vehicle

A prune-and-search optimization algorithm was combined with a reduced-order model of the scramjet engine within a generic X-51 vehicle. The goal was to maximize the thrust in order to achieve thrust-to-drag ratios that equal or exceed 1, even for the difficult cases of flight at high altitudes and high Mach numbers up to 10. For the inlet to the engine, the lengths and inclination angles of three wall panels were varied. For the combustor, three parameters were varied: the diameter and number of fuel injector ports and the wall divergence angle. These variables affect both the combustion efficiency and the finite-rate chemistry of the surrogate JP-7 fuel. The optimization algorithm showed that at higher flight Mach numbers an inlet that maximizes static pressure ratio is preferred over one that maximizes the stagnation pressure ratio, because good combustion efficiency requires a sufficient static pressure in the combustor. As the flight Mach number is increased, thrust decreases while the drag increases. Thus, it becomes difficult to achieve thrust that exceeds drag. A solution is shown to be increasing the engine inlet area above a critical value that was computed. Advantages of a reduced-order model over a high-fidelity CFD approach are discussed when thousands of computations of all components within an entire engine are required for optimization.

Developing a predictive model for low-temperature Laval nozzles with applications in chemical kinetics

Laval nozzles are used in the CRESU (“Cinétique de Réaction en Écoulement Supersonique Uniforme”) method to generate a collimated low temperature (5–200 K), low pressure (30–500 Pa), high Mach number (1 < M < 20) supersonic jet. Laval nozzles have been designed using the Method of Characteristics (MOC) since the development of CRESU, which is an analytical method that assumes inviscid, isentropic flow, and is routinely used to design nozzle profiles for a particular gas and temperature with a uniform shock free exit. This study aims to provide a robust computational framework to overcome the limitations of the MOC while also providing recommendations on the numerical model setup required to model a low-temperature supersonic jet. It also discusses the blockage effects when using the Pitot tube method for flow characterization, the influence of inlet turbulence and reservoir size. Numerical results are validated using two different experimental apparatuses from research groups at the University of Leeds and the University of Birmingham. Finally, a MATLAB framework was developed and has been provided as an open source toolbox to allow any user to perform computational fluid dynamics on any Laval nozzle, with the ability to change nozzle geometry, operating conditions and bath gas. The toolbox has been rigorously tested against many benchmark cases, which shows that steady-state Reynolds-averaged Navier–Stokes with the k-omega-shear stress transport turbulence model can be used to accurately predict global quantities, such as average temperature in the stable region of the supersonic jet.

An all Mach number scheme for visco-resistive magnetically-dominated MHD flows

In this work we present a novel conservative, semi-implicit, finite-volume scheme, designed for the numerical simulation of magnetically-dominated plasma rapidly evolving from a near-equilibrium state into a time-dependent compressible regime. The proposed scheme utilises a well-studied flux vector splitting approach, where the ideal magnetohydrodynamic (MHD) system is partitioned into two sub-systems, one containing all the advective terms and the other containing the rest of the terms, including not only the gas pressure but also magnetic pressure terms. The former is discretised explicitly as part of a standard finite-volume update, while the latter is treated implicitly by means of a nested Picard algorithm. As a result, the algorithm is constrained only by a mild velocity CFL stability condition, rendering it highly suitable for low Mach number flows in both gas and magnetic pressure dominated conditions, and for the incompressible limit of the MHD equations. The visco-resistive dissipative terms are also treated implicitly so that the numerical stability of the algorithm remains dependent only on the fluid velocity, maintaining its efficiency even for visco-resistive dominated problems. The divergence constraint of the magnetic field is handled through either staggered or unstaggered adaptations of the constrained transport methodology, which are both shown to preserve the divergence error up to machine precision levels. The algorithm is validated thoroughly by means of several benchmark problems for high and low Mach number flows, as well as multiple visco-resistive test cases. The scheme is found to perform well for both regimes, providing accurate and computationally efficient results in low Mach number flows in both gas and magnetic pressure dominated scenarios, while also demonstrating excellent shock-capturing capacity. A novel test case is finally introduced to truly demonstrate the all Mach number capabilities of the proposed scheme.

Investigation of hydrogen/air co-flow jet flame propagation mechanism in supersonic crossflow

The design of fuel injection schemes is crucial for improving the combustion performance of high-Mach number scramjet. To clarify the feasibility of the coaxial jet injection scheme, high fidelity Large Eddy Simulation of the supersonic coaxial jet flame is conducted. The simulations are in good agreement with the experimental results in terms of time-averaged velocity, temperature, and species distribution. Auto-ignition phenomenon and the characteristics of partially premixed flame are well captured. The introduction of co-flow air increases the vorticity magnitude close to the injection port and downstream near-wall region, which results in a 400 K rise in the time-averaged temperature on the downstream near-wall region and a 4% increase in the proportion of premixed combustion near the injection port. Moreover, the instantaneous distribution of hydroxy radical indicates that the spanwise width of the windward reaction shear layer is reduced utilizing the coaxial jet scheme. Chemical kinetic analysis is applied to reveal the propagation mechanism of partially premixed flames. Thermal explosion is the chemical explosion mode for all coaxial jet flame front, which are dominated by a high-temperature reaction path. The low-temperature reaction path mainly exists in the transverse jet injection port, downstream near-wall region of the single transverse jet and co-flow lifted flame base. These significant findings provide valuable insights for the potential engineering application of the coaxial jet injection scheme to a high Mach number scramjet.

Electron Heating in High Mach Number Collisionless Shocks.

The energy partition in high Mach number collisionless shock waves is central to a wide range of high-energy astrophysical environments. We present a new theoretical model for electron heating that accounts for the energy exchange between electrons and ions at the shock. The fundamental mechanism relies on the difference in inertia between electrons and ions, resulting in differential scattering of the particles off a decelerating magnetically dominated microturbulence across the shock transition. We show that the self-consistent interplay between the resulting ambipolar-type electric field and diffusive transport of electrons leads to efficient heating in the magnetic field produced by the Weibel instability in the high Mach number regime and is consistent with fully kinetic simulations.

URANOS-2.0: Improved performance, enhanced portability, and model extension towards exascale computing of high-speed engineering flows

We present URANOS-2.0, the second major release of our massively parallel, GPU-accelerated solver for compressible wall flow applications. This latest version represents a significant leap forward in our initial tool, which was launched in 2023 (De Vanna et al. [1]), and has been specifically optimized to take full advantage of the opportunities offered by the cutting-edge pre-exascale architectures available within the EuroHPC JU. In particular, URANOS-2.0 emphasizes portability and compatibility improvements with the two top-ranked supercomputing architectures in Europe: LUMI and Leonardo. These systems utilize different GPU architectures, AMD and NVIDIA, respectively, which necessitates extensive efforts to ensure seamless usability across their distinct structures. In pursuit of this objective, the current release adheres to the OpenACC standard. This choice not only facilitates efficient utilization of the full potential inherent in these extensive GPU-based architectures but also upholds the principles of vendor neutrality, a distinctive characteristic of URANOS solvers in the CFD solvers' panorama. However, the URANOS-2.0 version goes beyond the goals of improving usability and portability; it introduces performance enhancements and restructures the most demanding computational kernels. This translates into a 2× speedup over the same architecture. In addition to its enhanced single-GPU performance, the present solver release demonstrates very good scalability in multi-GPU environments. URANOS-2.0, in fact, achieves strong scaling efficiencies of over 80% across 64 compute nodes (256 GPUs) for both LUMI and Leonardo. Furthermore, its weak scaling efficiencies reach approximately 95% and 90% on LUMI and Leonardo, respectively, when up to 256 nodes (1024 GPUs) are considered. These significant performance advancements position URANOS-2.0 as a state-of-the-art supercomputing platform tailored for compressible wall turbulence applications, establishing the solver as an integrated tool for various aerospace and energy engineering applications, which can span from direct numerical simulations, wall-resolved large eddy simulations, up to most recent wall-modeled large eddy simulations. Program summaryProgram title: Unsteady Robust All-around Navier-StOkes Solver (URANOS)CPC Library link to program files:https://doi.org/10.17632/pw5hshn9k6.2Developer's repository link:https://github.com/uranos-gpu/uranos-gpu, https://github.com/uranos-gpu/uranos-gpu/tree/v2.0Licensing provisions: BSD License 2.0Programming language: Modern Fortran, OpenACC, MPINature of problem: Solving the compressible Navier-Stokes equations in a three-dimensional Cartesian framework.Solution method: Convective terms are treated with high-resolution shock-capturing schemes. The system dynamics is advanced in time with a three-stage Runge-Kutta method. Parallelization adopts MPI+OpenACC.

Characteristics and transformations of dust acoustic double-layers in plasmas with adiabatically trapped-nonextensive ions using Tsallis-Gurevich distribution

The investigation focuses on examining the characteristics of dust acoustic double-layers (DA-DLs) in the presence of adiabatically trapped-nonextensive ions featuring Tsallis-Gurevich distribution. Indeed, the ion-density and pseudo-potential expressions are formulated in terms of transcendental functions, allowing for a detailed analysis of Dust Acoustic Double Layers associated with complex plasmas in the presence of non-extensive adiabatically trapped ions. The study then delves into the impact of background ion nonextensivity on key DA-DLs properties. The findings reveal that our plasma model can exhibits both compressive and rarefactive DA-DLs, depending on the nonextensive parameter q. Also, the results indicate that in both sub-sonic and supra-sonic scenarios, the amplitude of the DA-DLs structure rises as the nonextensive parameter decreases. Additionally, it is observed that, for a constant value of the nonextensive parameter q, the amplitude of the DA-DLs structure increases with higher Mach numbers. This investigation, prompted by observations in space and laboratory plasmas containing non-Maxwellian particles alongside trapped particles, has the potential to complement and offer fresh insights into previous works on solitary waves in plasma.