arXiv subjects

João Luiz F. Azevedo

Publications and source records attributed to João Luiz F. Azevedo.

14 recordsLinked to original sources

arXiv2026

Influence of Park's Two-Temperature Model Control Temperature on the Flow Properties in Hypersonic Reentry Conditions

Numerical simulations of reactive hypersonic flows under thermochemical non-equilibrium conditions are presented for the FIRE II and Mars Pathfinder capsules. An 11-species chemical model is employed to simulate Earth's atmosphere, while an 8-species chemical model simulates Mars' atmosphere. The current formulation uses Park's two-temperature model to account for the non-equilibrium phenomena. The present work analyzes the impact of different sets of weight factors used in Park's model to calculate the control temperature. The code used to simulate the hypersonic flow addressed in this work solves the Navier-Stokes equations for reacting gas flows. The findings are depicted in terms of the Mach number, temperature modes, and mass fraction distributions along the stagnation streamline in a region closer to the shock wave. The study also includes results regarding the stagnation point convective heat flux. The results presented are encouraging and show that the weight factors significantly impact the FIRE II test cases while having little impact on the Mars Pathfinder flows. In all cases, it is possible to observe some effect of the weight factor selection on property distributions. In summary, the weight factors influence the flow behavior with varying intensities depending on the flow conditions.

physics.flu-dyn↗

arXiv2026

A Study of Improved Limiter Formulations for Second-Order Finite Volume Schemes Applied to Unstructured Grids

A general, compact way of achieving second-order in finite-volume numerical methods is to perform a MUSCL-like, piecewise linear reconstruction of flow properties at each cell interface. To avoid the surge of spurious oscillations in the discrete solution, a limiter function is commonly employed. This strategy, however, can add a series of drawbacks to the overall numerical scheme. The present paper investigates this behavior by considering three different limiter formulations in the context of a second-order, finite volume scheme for the simulation of steady, turbulent flows on unstructured meshes. Three limiter formulations are considered: the original Venkatakrishnan limiter, Wang's modification to the Venkatakrishnan limiter and Nishikawa's recently introduced R3 limiter. Three different configurations of the fully-developed, two-dimensional, transonic NACA 0012 airfoil are analyzed, configured with different angles of attack and similar freestream properties. The gas dynamics are modeled using the Reynolds-averaged Navier-Stokes (RANS) equations, where the negative Spalart-Allmaras turbulence model is used to solve the closure problem. All limiters are shown to yield similar results for all configurations of this case, although with different dissipative characteristics, provided their control constants are used within appropriate intervals. The presented numerical results are in good agreement with experimental data available in the literature.

physics.flu-dyn↗

arXiv2026

Study of the Effects of Artificial Dissipation and Other Numerical Parameters on Shock Wave Resolution

The effects induced by numerical schemes and mesh geometry on the solution of two-dimensional supersonic inviscid flows are investigated in the context of the compressible Euler equations. Five different finite-difference schemes are considered: the Beam and Warming implicit approximate factorization algorithm, the original Steger and Warming flux vector splitting algorithm, the van Leer approach on performing the flux vector splitting and two different novel finite-difference interpretations of the Liou AUSM+ scheme. Special focus is given to the shock wave resolution capabilities of each scheme for the solution of an external supersonic inviscid flows around a blunt body. Significant changes in the shock structure are observed, mainly due to special properties of the scheme in use and the influence of the domain transformation procedure. Perturbations in the supersonic flow upstream of the shock are also seen in the solution, which is a non-physical behavior. Freestream subtraction, flux limiting and the explicit addition of artificial dissipation are employed in order to circumvent these problems. One of the AUSM+ formulations presented here is seen to be particularly more robust in avoiding the appearance of some of these numerically-induced disturbances and non-physical characteristics in the solution. Good agreement is achieved with both numerical and experimental results available in the literature.}

physics.flu-dyn↗

arXiv2026

Numerical Aspects of Gradient Reconstruction Schemes Applied to Complex Geometries

This work primarily focuses on the study of three gradient reconstruction techniques applied to the calculation of viscous terms in a cell-centered, finite volume formulation for general unstructured grids. The work also addresses different ways of formulating the limiter functions necessary to maintain stability in the presence of flow discontinuities. The flows of interest are simulated using the compressible Reynolds-averaged Navier-Stokes equations, and the negative Spalart-Allmaras model is used for turbulence closure. Definition of interface inviscid terms uses the Roe approximate Riemann solver, whereas the interface viscous terms are calculated with a standard centered scheme together with appropriate definitions of the interface gradients. Steady state solutions are obtained using an implicit time-integration method, together with a novel convergence acceleration technique. This new approach defines a set of three simple rules for controlling the global CFL number based on the residue evolution. The work considers three test cases, namely, the subsonic bump-in-channel flow, the subsonic NASA high-lift Common Research Model multielement airfoil and the transonic ONERA M6 wing. Present results are compared to experimental and numerical data available in the literature. Severe numerical instabilities are observed when the simplest gradient reconstruction technique is used, while more sophisticated formulations are able to provide excellent agreement with the existing literature. Current results are demonstrated to be highly insensitive to modifications made to the numerical flux entropy fix terms. Integrated aerodynamic forces are shown to be mildly dependent on the limiter formulation used, even in the absence of shock waves. The proposed convergence acceleration procedure manages to quickly drive the residue terms to machine zero, provided no major instabilities are present.

physics.flu-dyn↗

arXiv2025

Influence of the Control Temperature of Park's Two-temperature Model on the Mars Pathfinder Reactive Hypersonic Flow

Numerical simulations of reactive hypersonic flow under thermodynamic and chemical non-equilibrium conditions are presented for the Mars Pathfinder capsule. An 8-species chemical model is employed to simulate Mars' atmosphere. Park's two-temperature model is used to account for the thermal non-equilibrium phenomena. The present work analyzes the impact of different values of the weight factors used in Park's model, aiming to broaden the understanding of the weight factors influence. The code used to simulate the flows solves the Navier-Stokes equations modified to account for reacting gas mixtures. The findings are depicted in terms of the Mach number and temperature modes along the stagnation streamline in a region close to the shock wave. The present analysis also includes results regarding the stagnation point convective heat flux. The results indicate that varying the weight factors yields negligible differences in the shock wave position and stagnation point convective heat flux. The changes in the weight factors cause variations in the maximum temperature mode values in the non-equilibrium region. The results presented are in good agreement with experimental data present in the literature. The present work indicates that Park's two-temperature model weight factors can substantially affect the temperature mode distributions in the flow non-equilibrium region.

physics.flu-dyn↗

arXiv2025

A Comparison of LES Inlet Boundary Conditions for Supersonic Jet Flows

The present work evaluates the effects of three inflow boundary conditions on large-eddy simulations of supersonic jet flows. The three inlet flow configurations considered are an inviscid profile, a stationary turbulent profile extracted from a RANS calculation, and a time-dependent tripped turbulent profile. The study applies the nodal discontinuous Galerkin spatial discretization to perform the supersonic jet flow simulations. Initially, a mesh and polynomial resolution study is performed to identify the requisites for simulating the supersonic jet flow. The velocity profiles resulting from the three numerical simulations with the different inflow conditions are compared with experimental data. The velocity profiles extracted from the calculation results indicate a major role of the steady turbulent inlet profile on the jet flow behavior, even for the tripped inlet condition. The imposition of a steady turbulent inlet profile could approximate the velocity distributions in the region close to the jet inlet section with the experimental reference. Negligible changes were observed to the velocity profiles far from the inlet section, where a good agreement with the experimental reference was already obtained from the simulations even with the inviscid profile.

physics.flu-dyn↗

arXiv2024

Accuracy Assessment of Discontinuous Galerkin Spectral Element Method in Simulating Supersonic Free Jets

The study performs large-eddy simulations of supersonic free jet flows using the Discontinuous Galerkin Spectral Element Method (DGSEM). The main objective of the present work is to assess the resolution requirements for adequate simulation of such flows with the DGSEM approach. The study looked at the influence of the mesh and the spatial discretization accuracy on the simulation results. The present analysis involves four simulations, incorporating three different numerical meshes and two different orders of spatial discretization accuracy. The numerical meshes are generated with distinct mesh topologies and refinement levels. Detailed descriptions of the grid generation and refinement procedures are presented. The study compares flow property profiles and power spectral densities of velocity components with experimental data. The results show a consistent improvement in the computed data as the simulation resolution increases. This investigation revealed a trade-off between mesh and polynomial refinement, striking a balance between computational cost and the accuracy of large-eddy simulation results for turbulent flow analyses.

physics.flu-dyn↗

arXiv2023

Stability and Order of Accuracy Analysis of High-Order Schemes Formulated Using the Flux Reconstruction Approach

A stability analysis is performed on high-order schemes formulated using the Flux Reconstruction (FR) approach. The one-dimensional advection model equation is used for the assessment of the stability region of these schemes when coupled with Runge-Kutta-type time-march procedures. Schemes are created using different numbers of internal points for each cell of the discrete domain, in such a way that a broad spectrum of spatial discretization orders can be analyzed. Multiple correction functions are employed so that well-known schemes, for instance the Discontinuous Galerkin (DG) and the Staggered-Grid schemes, may also be included in the study. The stability analysis is performed by identifying the behavior, in the complex plane, of the eigenvalues associated with each one of the considered cases. It is observed that, as the order of the scheme increases, a significant decrease in its domain of stability occurs, followed by a significant reduction in the amount of artificial dissipation that is introduced to the solution.

physics.flu-dyn↗

arXiv2023

Aeroelastic Analysis of Transonic Flutter with CFD-Based Reduced-Order Model

The current work is concerned with studying processes for constructing reduced-order models capable of performing transonic aeroelastic stability analyses in the frequency domain based on computational fluid dynamics (CFD) techniques. The CFD calculations are based on the Euler equations, and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. The dynamic system considered in the present work is a NACA 0012 airfoil-based typical section in the transonic regime. Unsteady calculations are performed for mode-by-mode and simultaneous excitation approaches, the latter defined by orthogonal Walsh functions. System identification techniques are employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer functions. Generalized unsteady aerodynamic forces are approximated by a rational transfer function in the Laplace domain, in which nonlinear parameters are selected through a non-gradient optimization process. Different types of interpolating polynomials are tested and the results are compared. Results show that the proposed procedure can reproduce literature aeroelastic analysis data with a single unsteady CFD run.

physics.flu-dyn↗

arXiv2023

Development of Transonic Unsteady Aerodynamic Reduced-Order Models Using System Identification Techniques

The present paper develops a reduced-order model capable of modeling unsteady aerodynamic loads in the transonic regime using system identification techniques. The computational fluid dynamics (CFD) calculations are based on the Euler equations and the code uses a finite volume formulation for general unstructured grids. A centered spatial discretization with added artificial dissipation is used, and an explicit Runge-Kutta time marching method is employed. For comparison reasons, unsteady calculations are performed using mode-by-mode and simultaneous excitation approaches. System identification techniques are employed to allow the splitting of the aerodynamic coefficient time histories into the contribution of each individual mode to the corresponding aerodynamic transfer function. Such methodology is applied to model the aerodynamic terms of the aeroelastic state-space system particularly of a NACA 0012 airfoil-based typical section. Results demonstrate the importance of signal processing techniques to compute the aerodynamic transfer functions and also the advantageous applicability of transonic unsteady aerodynamic reduced-order models to perform aeroelastic analyses in the frequency domain.

physics.flu-dyn↗

arXiv2023

On the Influence of Gradient Reconstruction Procedures Over the Accuracy of Finite Volume Based Schemes

In the context of the cell centered finite volume approach, care must be taken when performing the reconstruction of property gradients at cell interfaces. The present work analyzes three different gradient reconstruction procedures, using three different turbulent simulation test cases, namely the zero-gradient flat plate, the subsonic NACA 0012 airfoil and the transonic OAT15A airfoil. The analysis is concerned mainly with the usage of quadrilateral meshes. The gas dynamics equations are solved using an implicit implementation of Roe's second-order upwind scheme. The RANS closure problem is solved by using the negative Spalart-Allmaras turbulence model. The solution quality of each gradient discretization procedure is analyzed and compared to experimental data and other numerical solutions available in the literature. For the cases considered here, excellent agreement is obtained between the computed solutions and the expected results, regardless of which gradient reconstruction scheme is used.

physics.flu-dyn↗

arXiv2022

Influence of different subgrid-scale models in low-order LES of supersonic jet flows

The present work is concerned with a study of large eddy simulations (LES) of unsteady turbulent jet flows. In particular, the present analysis is focused on the effects of the subgrid-scale modeling used when a second-order spatial discretization methodology is employed for the numerical simulations. The present effort addresses perfectly expanded supersonic jets, because the authors want to emphasize the effects of the jet mixing phenomena. The LES formulation is discretized using the finite difference approach, after the equations are rewritten in a generalized coordinate system. Both space and time discretizations are second-order accurate and an explicit time march is adopted. Special care is dedicated to the discretization of the energy equation to appropriately model the set of filtered equations appearing in the LES formulation. The classical Smagorinsky, the dynamic Smagorinsky and the Vreman models are the subgrid-scale closures selected for the present work. The computational results are compared to data in the literature to validate the present simulation tool. Results indicate that the characteristics of numerical discretization can be as important as the effects of the subgrid-scale models for such low-order spatial discretization schemes. A detailed analysis is presented for the performance of each subgrid closure in the numerical context here considered.

physics.flu-dyn↗

arXiv2020

Strong Scaling of Numerical Solver for Supersonic Jet Flow Configuration

Acoustics loads are rocket design constraints which push researches and engineers to invest efforts in the aeroacoustics phenomena which is present on launch vehicles. Therefore, an in-house computational fluid dynamics tool is developed in order to reproduce high-fidelity results of supersonic jet flows for aeroacoustic analogy applications. The solver is written using the large eddy simulation formulation that is discretized using a finite-difference approach and an explicit time integration. Numerical simulations of supersonic jet flows are very expensive and demand efficient high-performance computing. Therefore, non-blocking message passage interface protocols and parallel input/output features are implemented into the code in order to perform simulations which demand up to one billion degrees of freedom. The present work evaluates the parallel efficiency of the solver when running on a supercomputer with a maximum theoretical peak of 127.4 TFLOPS. Speedup curves are generated using nine different workloads. Moreover, the validation results of a realistic flow condition are also presented in the current work.

cs.CE↗

arXiv2020

On the scalability of CFD tool for supersonic jet flow configurations

New regulations are imposing noise emissions limitations for the aviation industry which are pushing researchers and engineers to invest efforts in studying the aeroacoustics phenomena. Following this trend, an in-house computational fluid dynamics tool is build to reproduce high fidelity results of supersonic jet flows for aeroacoustic analogy applications. The solver is written using the large eddy simulation formulation that is discretized using a finite difference approach and an explicit time integration. Numerical simulations of supersonic jet flows are very expensive and demand efficient high-performance computing. Therefore, non-blocking message passage interface protocols and parallel Input/Output features are implemented into the code in order to perform simulations which demand up to one billion grid points. The present work addresses the evaluation of code improvements along with the computational performance of the solver running on a computer with maximum theoretical peak of 2.727 PFlops. Different mesh configurations, whose size varies from a few hundred thousand to approximately one billion grid points, are evaluated in the present paper. Calculations are performed using different workloads in order to assess the strong and weak scalability of the parallel computational tool. Moreover, validation results of a realistic flow condition are also presented in the current work.

cs.DC↗