Journal articles: 'Onera M6'

1

Алексеенко, С. В. "ЧИСЛЕННОЕ МОДЕЛИРОВАНИЕ ОБТЕКАНИЯ КРЫЛА ONERA M6." Bulletin of Dnipro University. Series: Mechanics 26, no. 5 (April 26, 2018): 57. http://dx.doi.org/10.15421/371806.

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Mayeur, J., A. Dumont, D. Destarac, and V. Gleize. "Reynolds-Averaged Navier–Stokes Simulations on NACA0012 and ONERA-M6 Wing with the ONERA elsA Solver." AIAA Journal 54, no. 9 (September 2016): 2671–87. http://dx.doi.org/10.2514/1.j054512.

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Balan, Aravind, Michael A. Park, William K. Anderson, Dmitry S. Kamenetskiy, Joshua A. Krakos, Todd Michal, and Frédéric Alauzet. "Verification of Anisotropic Mesh Adaptation for Turbulent Simulations over ONERA M6 Wing." AIAA Journal 58, no. 4 (April 2020): 1550–65. http://dx.doi.org/10.2514/1.j059158.

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NAMBU, Taisuke, Atsushi HASHIMOTO, Takashi AOYAMA, and Tetsuya SATO. "Numerical Analysis of the ONERA-M6 Wing with Wind Tunnel Wall Interference." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 58, no. 1 (2015): 7–14. http://dx.doi.org/10.2322/tjsass.58.7.

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Batina, John T. "Accuracy of an unstructured-grid upwind-Euler algorithm for the ONERA M6 wing." Journal of Aircraft 28, no. 6 (June 1991): 397–402. http://dx.doi.org/10.2514/3.46040.

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Goodheart, Kevin A., and Gunter H. Schnerr. "Condensation on ONERA M6 and F-16 Wings in Atmospheric Flight: Numerical Modeling." Journal of Aircraft 42, no. 2 (March 2005): 402–12. http://dx.doi.org/10.2514/1.5137.

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Nejati, A., and K. Mazaheri. "Application of the adjoint optimisation of shock control bump for ONERA-M6 wing." European Journal of Computational Mechanics 26, no. 5-6 (October 12, 2017): 557–83. http://dx.doi.org/10.1080/17797179.2017.1386022.

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Johan, Zdeněk, Kapil K. Mathur, S. Lennart Johnsson, and Thomas J. R. Hughes. "A case study in parallel computation: Viscous flow around an ONERA M6 wing." International Journal for Numerical Methods in Fluids 21, no. 10 (November 30, 1995): 877–84. http://dx.doi.org/10.1002/fld.1650211008.

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Li, Zong Zhe, Zheng Hua Wang, Lu Yao, and Wei Cao. "A Combined Global Coarsening Method for 3D Multigrid Applications." Applied Mechanics and Materials 236-237 (November 2012): 1049–53. http://dx.doi.org/10.4028/www.scientific.net/amm.236-237.1049.

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An automatic agglomeration methodology to generate coarse grids for 3D flow solutions on anisotropic unstructured grids has been introduced in this paper. The algorithm combines isotropic octree based coarsening and anisotropic directional agglomeration to yield a desired coarsening ratio and high quality of coarse grids, which developed for cell-centered multigrid applications. This coarsening strategy developed is presented on an unstructured grid over 3D ONERA M6 wing. It is shown that the present method provides suitable coarsening ratio and well defined aspect ratio cells at all coarse grid levels.

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Hejranfar, Kazem, and Ramin Kamali Moghadam. "A Comparative Study of Two Preconditioners for Solving 3D Inviscid Low Speed Flows." Applied Mechanics and Materials 110-116 (October 2011): 423–30. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.423.

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In the present study, two preconditioners proposed by Eriksson, and Choi and Merkel are implemented on a 3D upwind Euler flow solver on unstructured meshes. The mathematical formulations of these preconditioning schemes for the set of primitive variables are drawn and their eigenvalues and eigenvectors are compared with each others. A cell-centered finite volume Roe's method is used for discretization of the 3D preconditioned Euler equations. The accuracy and performance of these preconditioning schemes are examined by computing low Mach number flows over the ONERA M6 wing for different conditions.

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Yao, Jianyao, Weimin Wu, Kun Zhang, Dongyang Sun, Yaolu Liu, Huiming Ning, Ning Hu, and G. R. Liu. "Development of Three-Dimensional GSM-CFD Solver for Compressible Flows." International Journal of Computational Methods 14, no. 04 (April 18, 2017): 1750037. http://dx.doi.org/10.1142/s0219876217500372.

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A three-dimensional (3D) Computational Fluid Dynamics (CFD) solver based on the gradient smoothing method (GSM) is developed for compressible flows based on previous research. The piecewise constant smoothing function with one-point integration scheme is implemented for gradient approximation of field variables and convective fluxes. The matrix-based method for gradient approximations is also developed to improve the numerical efficiency. Numerical examples of gradient approximations of several given functions have shown that the proposed GSM is more accurate and robust to mesh distortion. A transonic ONERA M6 wing is used to demonstrate the effectiveness of the proposed GSM-CFD solver.

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Ma, Wenpeng, Wu Yuan, and Xiaodong Hu. "Implementation and Optimization of a CFD Solver Using Overlapped Meshes on Multiple MIC Coprocessors." Scientific Programming 2019 (May 27, 2019): 1–12. http://dx.doi.org/10.1155/2019/4254676.

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In this paper, we develop and parallelize a CFD solver that supports overlapped meshes on multiple MIC architectures by using multithreaded technique. We optimize the solver through several considerations including vectorization, memory arrangement, and an asynchronous strategy for data exchange on multiple devices. Comparisons of different vectorization strategies are made, and the performances of core functions of the solver are reported. Experiments show that about 3.16x speedup can be achieved for the six core functions on a single Intel Xeon Phi 5110P MIC card, and 5.9x speedup can be achieved using two cards compared to an Intel E5-2680 processor for two ONERA M6 wings case.

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Aftab, Syed Mohammed Aminuddin, and P. Srinivasa Murthy. "Comparative Study of Vortex Generator Orientation on Wing Surface Considering Delta Vortex Generators." Applied Mechanics and Materials 225 (November 2012): 79–84. http://dx.doi.org/10.4028/www.scientific.net/amm.225.79.

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Flow over the ONERA M6 wing with vortex generators using more accurate higher order numerical schemes is studied using computational methods. In this paper, the effect of delta vortex generator orientation on the wing and its implication on wing performance is computed more accurately using second order upwinding scheme. Turbulence modeling used is k-omega sst. It has been found that numerical results are comparable and close to the experiment. The analysis results show that the co-rotating clockwise position of vortex generators is more effective than co-rotating (anticlockwise) or counter rotating position. The vortex generators have been found to control the boundary layer separation and give improvement in lift at high angle of attack.

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Ma, Xinrong, Sanyang Liu, and Gongnan Xie. "Predictor-Corrector LU-SGS Discontinuous Galerkin Finite Element Method for Conservation Laws." Mathematical Problems in Engineering 2015 (2015): 1–11. http://dx.doi.org/10.1155/2015/940257.

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Efficient implicit predictor-corrector LU-SGS discontinuous Galerkin (DG) approach for compressible Euler equations on unstructured grids is investigated by adding the error compensation of high-order term. The original LU-SGS and GMRES schemes for DG method are discussed. Van Albada limiter is employed to make the scheme monotone. The numerical experiments performed for the transonic inviscid flows around NACA0012 airfoil, RAE2822 airfoil, and ONERA M6 wing indicate that the present algorithm has the advantages of low storage requirements and high convergence acceleration. The computational efficiency is close to that of GMRES scheme, nearly 2.1 times greater than that of LU-SGS scheme on unstructured grids for 2D cases, and almost 5.5 times greater than that of RK4 on unstructured grids for 3D cases.

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Xiuling, Sun, Li Liang, and Li Guojun. "Transonic Flow of Moist Air around the ONERA M6 Wing with Non-equilibrium and Homogeneous Condensation." Research Journal of Applied Sciences, Engineering and Technology 6, no. 10 (July 20, 2013): 1825–33. http://dx.doi.org/10.19026/rjaset.6.3910.

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16

Borisov, V. E. "Simulation of flow over the ONERA M6 wing using a parallel implementation of an implicit scheme." Moscow University Mechanics Bulletin 70, no. 4 (July 2015): 97–100. http://dx.doi.org/10.3103/s0027133015040044.

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17

Kazim, Ali Hussain, Abdullah Hamid Malik, Hammad Ali, Muhammad Usman Raza, Awais Ahmad Khan, Tauseef Aized, and Aqsa Shabbir. "CFD analysis of variable geometric angle winglets." Aircraft Engineering and Aerospace Technology 94, no. 2 (October 15, 2021): 289–301. http://dx.doi.org/10.1108/aeat-10-2020-0241.

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Purpose Winglets play a major role in saving fuel costs because they reduce the lift-induced drag formed at the wingtips. The purpose of this paper is to obtain the best orientation of the winglet for the Office National d’Etudes et de Recherches Aérospatiales (ONERA) M6 wing at Mach number 0.84 in terms of lift to drag ratio. Design/methodology/approach A computational fluid dynamics analysis of the wing-winglet configuration based on the ONERA M6 airfoil on drag reduction for different attack angles at Mach 0.84 was performed using analysis of systems Fluent. First, the best values of cant and sweep angles in terms of aerodynamic performance were selected by performing simulations. The analysis included cant angle values of 30°, 40°, 45°, 55°, 60°, 70° and 75°, while for the sweep angles 35°, 45°, 55°, 65° and 75° angles were used. The aerodynamic performance was measured in terms of the obtained lift to drag ratios. Findings The results showed that slight alternations in the winglet configuration can improve aerodynamic performance for various attack angles. The best lift to drag ratio for the winglet was achieved at a cant angle of 30° and a sweep angle of 65°, which caused a 5.33% increase in the lift to drag ratio. The toe-out angle winglets as compared to the toe-in angles caused the lift to drag ratio to increase because of more attached flow at its surface. The maximum value of the lift to drag ratio was obtained with a toe-out angle (−5°) at an angle of attack 3° which was 2.53% greater than the zero-toed angle winglet. Originality/value This work is relatively unique because the cant, sweep and toe angles were analyzed altogether and led to a significant reduction in drag as compared to wing without winglet. The wing model was compared with the results provided by National Aeronautics and Space Administration so this validated the simulation for different wing-winglet configurations.

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Zhang, G. Q., S. C. M. Yu, and A. Chien. "Investigation of the Three-Dimensional Hinge Moment Characteristics Generated by the ONERA-M6 Wing with an Aileron." Advances in Mechanical Engineering 5 (January 2013): 714168. http://dx.doi.org/10.1155/2013/714168.

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Zhao, Ran, Chao Li, Xiaowei Guo, Sijiang Fan, Yi Wang, and Canqun Yang. "A Block Iteration with Parallelization Method for the Greedy Selection in Radial Basis Functions Based Mesh Deformation." Applied Sciences 9, no. 6 (March 18, 2019): 1141. http://dx.doi.org/10.3390/app9061141.

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Greedy algorithm is one of the important point selection methods in the radial basis function based mesh deformation. However, in large-scale mesh, the conventional greedy selection will generate expensive time consumption and result in performance penalties. To accelerate the computational procedure of the point selection, a block iteration with parallelization method is proposed in this paper. By the block iteration method, the computational complexities of three steps in the greedy selection are all reduced from O ( n 3 ) to O ( n 2 ) . In addition, the parallelization of two steps in the greedy selection separates boundary points into sub-cores, efficiently accelerating the procedure. Specifically, three typical models of three-dimensional undulating fish, ONERA M6 wing and three-dimensional Super-cavitating Hydrofoil are taken as the test cases to validate the proposed method and the results show that it improves 17.41 times performance compared to the conventional method.

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Liu, Xiazhen, Zhonghua Lu, Wu Yuan, Wenpeng Ma, and Jian Zhang. "Massively Parallel CFD Simulation Software: CCFD Development and Optimization Based on Sunway TaihuLight." Scientific Programming 2020 (July 22, 2020): 1–17. http://dx.doi.org/10.1155/2020/8847481.

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A parallel framework software, CCFD, based on the structure grid, and suitable for parallel computing of super-large-scale structure blocks, is designed and implemented. An overdecomposition method, in which the load balancing strategy is based on the domain decomposition method, is designed for the graph subdivision algorithm. This method takes computation and communication as the limiting condition and realizes the load balance between blocks by dividing the weighted graph. The fast convergence technique of a high-efficiency parallel geometric multigrid greatly improves the parallel efficiency and convergence speed of CCFD software. This paper introduces the software structure, process invocations, and calculation method of CCFD and introduces a hybrid parallel acceleration technology based on the Sunway TaihuLight heterogeneous architecture. The results calculated by Onera-M6 and DLR-F6 standard model show that the software structure and method in this paper are feasible and can meet the requirements of a large-scale parallel solution.

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Khalid, Mahmood. "Crosswise Wind Shear Represented as a Ramped Velocity Profile Impacting a Forward-Moving Aircraft." International Journal of Aerospace Engineering 2019 (August 18, 2019): 1–18. http://dx.doi.org/10.1155/2019/7594737.

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Abrupt changes in wind velocities over small distances in a lateral or vertical direction can produce wind shear which is known to have serious effects upon the performance of an aircraft. Brought about by large-scale changes in the atmospheric conditions, it is a three-dimensional flow phenomenon imposing severe velocity gradients on an aircraft from all possible directions. While it would be difficult to model an instantaneous velocity gradient in a lateral plane, a vortical flow impinging from the sides which represents a wind shear in a vertical direction is imposed on a forward-moving aircraft to investigate the effect on the aerodynamic performance. The maximum shear wind speed from the side was fixed at 0.3 times the forward velocity. After due validations under no-wind shear conditions on simpler half-reflection plane models, a BGK airfoil-based full 3D wing and the ONERA M6 3D wing model were selected for preliminary studies. The investigation was concluded using the ARA M100 wing-fuselage model.

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Gao, Huanqin, Jiale Zhang, Hongquan Chen, Shengguan Xu, and Xuesong Jia. "A High-Order Discontinuous Galerkin Method for Solving Preconditioned Euler Equations." Applied Sciences 12, no. 14 (July 12, 2022): 7040. http://dx.doi.org/10.3390/app12147040.

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A high-order discontinuous Galerkin (DG) method is presented for solving the preconditioned Euler equations with an explicit or implicit time marching scheme. A detailed description is given of a practical implementation of a precondition matrix of the type of Weiss and Smith and of the DG spatial discretization scheme employed, with particular emphasis on the artificial viscosity-based shock capturing techniques. The curved boundary treatment is proposed through adopting a NURBS surface equipped with a radial basis function interpolation to propagate the boundary displacement to the interior of the mesh. The resulting methods are verified by simulating flows over two-dimensional airfoils, such as symmetric NACA0012 or asymmetric RAE2822, and over three-dimensional bodies, such as an academic hemispherical headform or aerodynamic ONERA M6 wing. Numerical results show that the present method functions for both transonic and nearly incompressible flow simulations, and the proposed treatment of curved boundaries, play an important role in improving the accuracy of the obtained solutions, which are in good agreement with available experimental data or other numerical solutions reported in literature.

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Li, Bowen, Qiangqiang Sun, Dandan Xiao, and Wenqiang Zhang. "Numerical Investigation of the Aerofoil Aerodynamics with Surface Heating for Anti-Icing." Aerospace 9, no. 7 (June 24, 2022): 338. http://dx.doi.org/10.3390/aerospace9070338.

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The aerodynamics of an aerofoil with surface heating was numerically studied with the objective to build an effective anti-icing strategy and balance the aerodynamics performance and energy consumption. NACA0012, RAE2822 and ONERA M6 aerofoils were adopted as the test cases and the simulations were performed in the subsonic flight condition of commercial passenger aircraft. In the first session, the numerical scheme was firstly validated with the experimental data. A parametric study with different heating temperatures and heating areas was carried out. The lift and drag coefficients both drop with surface heating, especially at a larger angle of attack. It was found that the separation point on the upper surface of the aerofoil is sensitive to heating. Higher heating temperature or larger heating area pushes the shock wave and hence flow separation point moving towards the leading edge, which reduces the low-pressure region of the upper surface and decreases the lift. In the second session, the conclusions obtained are applied to inform the design of the heating scheme for NACA0012. Further guidelines for different flight conditions were proposed to shed light on the optimisation of the heating strategy.

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Xu, Li, Fengfeng Zhao, and Jingjing Cai. "A Modified Fifth-Order WENO-HLLC Riemann Solver for Transonic Viscous Flows around Helicopter Rotor in Hover." Shock and Vibration 2022 (January 19, 2022): 1–23. http://dx.doi.org/10.1155/2022/2180845.

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This paper presents a modified fifth-order WENO-HLLC Riemann solver for the transonic compressible viscous flows around the helicopter rotor in hover. The HLLC approximate Riemann solver is proposed to discrete the convection term involving grid velocity of the Navier-Stokes equations. In order to solve the interface flow accurately, a modified fifth-order WENO scheme is presented by designing the smoothness indicators based on L 1 norm measurement. The improved WENO scheme can provide the optimal approximation order even at critical points. Numerical accuracy and robustness are validated by several benchmark inviscid flow problems. Then the numerical properties of the WENO-HLLC solver in conjunction with the implicit LU-SGS time integration method with high efficiency are further validated by simulating transonic viscous flows over RAE2822 airfoil and ONERA-M6 wing. The results show that the accuracy of calculating shock, discontinuity, and the vortex is significantly improved. Finally, the method is developed to compute the transonic vortex flow around the helicopter rotor with a domain discretized by overset grids. The results indicate that the proposed method is very robust and effective in acquiring high resolution for vortex wake.

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Veilleux, C., C. Masson, and I. Paraschivoiu. "A new induced-drag prediction method using Oswatitsch’s expression." Aeronautical Journal 103, no. 1024 (June 1999): 299–307. http://dx.doi.org/10.1017/s0001924000064861.

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Abstract This paper presents a critical study related to the evaluation and breakdown of total wing drag based on solutions of the Euler equations for subsonic and transonic three-dimensional flows. This study clearly identifies the false entropy production present in all discretised Euler solutions as the main source of the discrepancies between the drag predictions produced by body-surface pressure integrations and far-field methods. A good understanding of the entropy production mechanisms present in discretised threedimensional Euler solutions has lead to an original procedure for the total drag breakdown and a new method for the evaluation of the induced drag. This new technique is based on body-surface pressure integrations and Oswatitsch’s expression integrated over suitable integration surfaces. In contrast with the various methods for the drag breakdown and the evaluation of the induced drag available in the literature, the proposed technique uses an exact formulation and is simple to implement. Using detailed drag calculations on an elliptic wing (the Xt = 1·0 wing) and the ONERA M6 wing, it has been shown that this new technique is not strongly sensitive to grid refinement and to the level of false entropy production in the calculation domain.

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Araya, Guillermo. "Turbulence Model Assessment in Compressible Flows around Complex Geometries with Unstructured Grids." Fluids 4, no. 2 (April 28, 2019): 81. http://dx.doi.org/10.3390/fluids4020081.

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One of the key factors in simulating realistic wall-bounded flows at high Reynolds numbers is the selection of an appropriate turbulence model for the steady Reynolds Averaged Navier–Stokes equations (RANS) equations. In this investigation, the performance of several turbulence models was explored for the simulation of steady, compressible, turbulent flow on complex geometries (concave and convex surface curvatures) and unstructured grids. The turbulence models considered were the Spalart–Allmaras model, the Wilcox k- ω model and the Menter shear stress transport (SST) model. The FLITE3D flow solver was employed, which utilizes a stabilized finite volume method with discontinuity capturing. A numerical benchmarking of the different models was performed for classical Computational Fluid Dynamic (CFD) cases, such as supersonic flow over an isothermal flat plate, transonic flow over the RAE2822 airfoil, the ONERA M6 wing and a generic F15 aircraft configuration. Validation was performed by means of available experimental data from the literature as well as high spatial/temporal resolution Direct Numerical Simulation (DNS). For attached or mildly separated flows, the performance of all turbulence models was consistent. However, the contrary was observed in separated flows with recirculation zones. Particularly, the Menter SST model showed the best compromise between accurately describing the physics of the flow and numerical stability.

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Crovato, Adrien, Hugo S. Almeida, Gareth Vio, Gustavo H. Silva, Alex P. Prado, Carlos Breviglieri, Huseyin Guner, et al. "Effect of Levels of Fidelity on Steady Aerodynamic and Static Aeroelastic Computations." Aerospace 7, no. 4 (April 11, 2020): 42. http://dx.doi.org/10.3390/aerospace7040042.

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Static aeroelastic deformations are nowadays considered as early as in the preliminary aircraft design stage, where low-fidelity linear aerodynamic modeling is favored because of its low computational cost. However, transonic flows are essentially nonlinear. The present work aims at assessing the impact of the aerodynamic level of fidelity used in preliminary aircraft design. Several fluid models ranging from the linear potential to the Navier–Stokes formulations were used to solve transonic flows for steady rigid aerodynamic and static aeroelastic computations on two benchmark wings: the Onera M6 and a generic airliner wing. The lift and moment loading distributions, as well as the bending and twisting deformations predicted by the different models, were examined, along with the computational costs of the various solutions. The results illustrate that a nonlinear method is required to reliably perform steady aerodynamic computations on rigid wings. For such computations, the best tradeoff between accuracy and computational cost is achieved by the full potential formulation. On the other hand, static aeroelastic computations are usually performed on optimized wings for which transonic effects are weak. In such cases, linear potential methods were found to yield sufficiently reliable results. If the linear method of choice is the doublet lattice approach, it must be corrected using a nonlinear solution.

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Li, Sibo, and Roberto Paoli. "Modeling of Ice Accretion over Aircraft Wings Using a Compressible OpenFOAM Solver." International Journal of Aerospace Engineering 2019 (June 3, 2019): 1–11. http://dx.doi.org/10.1155/2019/4864927.

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A method to simulate ice accretion on an aircraft wing using a three-dimensional compressible Navier-Stokes solver, a Eulerian droplet flow field model, a mesh morphing model, and a thermodynamic model, is presented in this paper. The above models are combined together into one solver and implemented in OpenFOAM. Two-way coupling is achieved between airflow field calculation and ice simulation. The density-based solver rhoEnergyFoam is used to calculate the airflow field. The roughness wall function is proposed to simulate the roughness effect caused by ice accretion. For droplet flow field calculation, the Eulerian model is applied and the permeable wall boundary condition is used on the wing to simulate the droplet impingement. The icing thermodynamic model is built based on the Messinger model. The mesh morphing model adjusts the wing’s shape every time step based on the amount of accreted ice so that the airflow field is updated during the simulation. The effect of the ice accretion on the airflow is studied by comparing the aerodynamic performance—with and without ice. The ice accretion on the ONERA M6 wing model under a specific condition has been simulated to validate the solver’s performance and investigate the effect of the accreted ice on the aerodynamic performance.

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Zhang, Bin, Zhiwei Feng, Tao Yang, Boting Xu, and Xiaojian Sun. "Integrated improvement of the elasticity-based mesh deformation method based on robust parameters and mesh quality." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 8 (July 25, 2018): 3078–95. http://dx.doi.org/10.1177/0954410018790887.

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Highly robust mesh deformation methods are key techniques for solving unsteady flow field problems with moving or deforming boundaries. Because it is imperative to reduce the remeshing times, these methods are important in engineering applications, especially for complex geometric boundaries and large displacements. We introduce three classical elasticity-based mesh deformation methods and determine the limitations of the two nonlinear classical methods. Two steps were taken to achieve an integrated improvement: first, the robust power parameters a and b and the weighted parameter x are introduced to enhance the robustness of the basic elasticity equation. Second, a mesh quality parameter is implemented to prevent the large distortion of the poor elements and this parameter is added to the elasticity equation as a constraint. To validate the validity of the integrated improvement approach, several test cases of a moving or deforming two-dimensional flat plate are used. Additionally, two simulated engineering examples are used to demonstrate the application of the integrated improvement for practical problems, including the pitching oscillation of a National Advisory Committee for Aeronautics (NACA) 0012 airfoil and the ONERA M6 wing. The results show that the integrated improvement approach does not only allow for the selection of suitable robust parameters to achieve more robust deformed meshes but also reduces the distortion of the poor elements near the moving boundary, even when the deformation is severe.

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Kaya, Mustafa, and Munir Ali Elfarra. "Taper stacking for the aerodynamic performance of wings." Aircraft Engineering and Aerospace Technology 92, no. 7 (June 5, 2020): 1101–10. http://dx.doi.org/10.1108/aeat-12-2019-0257.

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Purpose The critical Mach number, lift-to-drag ratio and drag force play important role in the performance of the wings. This paper aims to investigate the effect of taper stacking, which has been used to generalize wing sweeping, on those parameters. Design/methodology/approach The results obtained are based on steady-state turbulent flowfields computations. The baseline wing is ONERA M6. Various wing planforms are generated by linearly or parabolically varying the spanwise stacking location. The critical Mach number is determined by changing the freestream Mach number for a fixed angle of attack. On the other hand, the analysis of the drag force is carried out by changing the angle of attack to keep the lift force constant. Findings By changing the stacking location, the critical Mach number and the corresponding lift-to-drag ratio have increased by around 7 and 3%, respectively. A reduction of 12.8% in total drag force has been observed in one of the analyzed cases. Moreover, there exist some cases in which the values of drag reduce significantly while the lift is the same. Practical implications The results of this new stacking approach have implied that the drag force can be decreased without decreasing the lift. This outcome is valuable for increasing the range and endurance of an aircraft. Originality/value This work generalizes wing sweeping by modifying the taper stacking along the span. In literature, wing sweep is enhanced using segmented stacking of taper distribution. The present study is further enhancing this concept by introducing continuous stacking (infinite number of stacking segments) for the first time.

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Li, Sibo, Roberto Paoli, and Michael D’Mello. "Scalability of OpenFOAM Density-Based Solver with Runge–Kutta Temporal Discretization Scheme." Scientific Programming 2020 (March 11, 2020): 1–11. http://dx.doi.org/10.1155/2020/9083620.

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Compressible density-based solvers are widely used in OpenFOAM, and the parallel scalability of these solvers is crucial for large-scale simulations. In this paper, we report our experiences with the scalability of OpenFOAM’s native rhoCentralFoam solver, and by making a small number of modifications to it, we show the degree to which the scalability of the solver can be improved. The main modification made is to replace the first-order accurate Euler scheme in rhoCentralFoam with a third-order accurate, four-stage Runge-Kutta or RK4 scheme for the time integration. The scaling test we used is the transonic flow over the ONERA M6 wing. This is a common validation test for compressible flows solvers in aerospace and other engineering applications. Numerical experiments show that our modified solver, referred to as rhoCentralRK4Foam, for the same spatial discretization, achieves as much as a 123.2% improvement in scalability over the rhoCentralFoam solver. As expected, the better time resolution of the Runge–Kutta scheme makes it more suitable for unsteady problems such as the Taylor–Green vortex decay where the new solver showed a 50% decrease in the overall time-to-solution compared to rhoCentralFoam to get to the final solution with the same numerical accuracy. Finally, the improved scalability can be traced to the improvement of the computation to communication ratio obtained by substituting the RK4 scheme in place of the Euler scheme. All numerical tests were conducted on a Cray XC40 parallel system, Theta, at Argonne National Laboratory.

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Sulistiya, Sulistiya, and Alief Sadlie Kasman. "Validasi Model Turbulensi pada Simulasi Numerik Menggunakan Software Fluent dengan Sayap Onera M6." Journal of Aero Technology 2, no. 1 (November 22, 2019). http://dx.doi.org/10.29122/joat.v2i1.3817.

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AbstractNumerical simulation using Computational Fluid Dynamics (CFD) method is one way of predicting airflow characteristics on the model. This method is widely used because it is relatively inexpensive and faster in getting desired results compared with performing direct testing. The correctness of a computational simulation output is highly dependent on the input and how it was processed. In this paper, simulation is done on Onera M6 Wing, to investigate the effect of a turbulence model’s application on the accuracy of the computational result. The choice of Onera M6 Wing as a simulation’s model is due to its extensive database of testing results from various wind tunnels in the world. Among Turbulence models used are Spalart-Allmaras, K-Epsilon, K-Omega, and SST.Keywords: CFD, fluent, Model, Turbulence, Onera M6, Spalart-Allmaras, K-Epsilon, K-Omega, SST.AbstraksSimulasi numerik dengan menggunakan metode Computational Fluid Dynamics (CFD) merupakan salah satu cara untuk memprediksi karakteristik suatu aliran udara yang terjadi pada model. Metode ini banyak digunakan karena sifatnya yang relatif murah dan cepat untuk mendapatkan hasil dibandingkan dengan melakukan pengujian langsung. Benar tidak hasil sebuah simulasi komputasi sangat tergantung pada inputan yang diberikan serta cara memproses data inputan tersebut. Pada tulisan ini dilakukan simulasi dengan menggunakan sayap onera M6 dengan tujuan untuk mengetahui pengaruh penggunaan model turbulensi terhadap keakuratan hasil komputasi. Pilihan sayap onera M6 sebagai model simulasi dikarenakan model tersebut sudah memiliki database hasil pengujian yang cukup lengkap dan sudah divalidasi dari berbagai terowongan angin di dunia. Model turbulensi yang digunakan diantaranya Spalart-Allmaras, K-Epsilon, K-Omega dan SST.Kata Kunci : CFD, fluent, Model, Turbulensi, Onera M6, Spalart-Allmaras, K-Epsilon, K-Omega, SST.

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Hart, Pierce L., and Sven Schmitz. "Drag Decomposition Using Partial-Pressure Fields: ONERA M6 Wing." AIAA Journal, December 13, 2021, 1–12. http://dx.doi.org/10.2514/1.j061152.

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Lakhshmanan, D., P. Vadivelu, G. Sivaraj, and M. S. Prasath. "Computational fluid dynamics simulation on aerodynamic characteristics of ONERA M6-wing." Materials Today: Proceedings, July 2021. http://dx.doi.org/10.1016/j.matpr.2021.06.042.

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Zhang, Yang, Jun-Qiang Bai, and Jing-Lei Xu. "A Scale-Adaptive Turbulence Model Based on the k-Equation and Recalibrated Reynolds Stress Constitutive Relation." Journal of Fluids Engineering 138, no. 6 (March 24, 2016). http://dx.doi.org/10.1115/1.4032535.

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An algebraic relationship between turbulent dissipation rate and von Karman length are used to dismiss the transport equation of turbulent dissipation rate in standard k−ε (SKE) turbulence model. Meanwhile, a recalibrated Bradshaw's assumption is built based on the data from a boundary layer flow of turbulent flat plate simulated by direct numerical simulation (DNS). The JL model is reformed to a one-equation model which only depends on the turbulent energy, so the new model can also be called kinetic-energy dependent only (KDO) turbulence model. As the KDO model is using the von Karman length scale, it can automatically adjust to fit the resolved structures of the local flow. Results will be shown for the boundary layer flow on a turbulent flat plate, and the external flows of an NACA4412 airfoil, an ONERA-M6 wing, a three dimension delta wing, and an NACA0012 airfoil at deep stall.

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Dam, Burak, Tolga Pirasaci, and Mustafa Kaya. "Artificial neural network based wing planform aerodynamic optimization." Aircraft Engineering and Aerospace Technology, May 12, 2022. http://dx.doi.org/10.1108/aeat-10-2021-0311.

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Purpose Environmental and operational restrictions increasingly drive modern aircraft design due to the growing impact of global warming on the ecology. Regulations and industrial measures are being introduced to make air traffic greener, including restrictions and environmental targets for aircraft design that increase aerodynamic efficiency. This study aims to maximize aerodynamic efficiency by identifying optimal values for sweep angle, taper ratio, twist angle and wing incidence angle parameters in wing design while keeping wing area and span constant. Design/methodology/approach Finding optimal wing values by using gradient-based and evolutionary algorithm methods is very time-consuming. Therefore, an artificial neural network-based surrogate model was developed. Computational fluid dynamics (CFD) analyses were carried out by using Reynolds-averaged Navier–Stokes equations to create a properly trained data set using a feedforward neural network. Findings The results showed how a wing could be optimized by using a CFD-based surrogate model. The two optimum results obtained resulted in increases of 10.7397% and 10.65% in the aerodynamic efficiency of the baseline design ONERA M6 wing. Originality/value The originality of this study lies in the combination of sweep angle, taper ratio, twist angle and wing incidence angle within the scope of wing optimization calculations.

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Xu, Lei, Wu Zhang, Yuhui Chen, and Rongliang Chen. "A Parallel Discrete Unified Gas Kinetic Scheme on Unstructured Grid for Inviscid High-speed Compressible Flow Simulation." Physics of Fluids, September 21, 2022. http://dx.doi.org/10.1063/5.0118179.

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The discrete unified gas kinetic scheme (DUGKS) is a recently devised approach to simulate multiscale flows based on the kinetic models, which also shows distinct features for continuum flows. Most of the existing DUGKS are sequential or based on structured grids, thus limiting their scope of application in engineering. In this paper, a parallel DUGKS for inviscid high-speed compressible flows on unstructured grids is proposed. In the framework of DUGKS, the gradients of the distribution functions are calculated by a least-square method. To parallelize the method, a graph-based partitioning method is employed to guarantee the load balancing and minimize the communication among processors. The method is validated by several benchmark problems, i.e., a two-dimensional (2D) Riemann problem, 2D subsonic flows passing two benchmark airfoils, a 2D regular shock reflection problem, 2D supersonic flows (Mach numbers are 3 and 5) around a cylinder, an explosion in a three-dimensional (3D) box, a 3D subsonic flow around the Office National d'Etudes et de Recherches Aérospatiales (ONERA) M6 wing, and a 3D hypersonic flow (Mach number is 10) around a hemisphere. The numerical results show good agreements with the published results and the present method is robust for a wide range of Mach number, from subsonic to hypersonic. The parallel performance results show that the proposed method is highly parallel scalable, where an almost linear scalability with 93% parallel efficiency is achieved for a 3D problem with over 55 million tetrahedrons on a supercomputer with up to 4800 processors.

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Cao, Cheng, Hongquan Chen, and Jiale Zhang. "Preconditioned gridless methods for solving three-dimensional Euler equations at low Mach numbers." International Journal of Modeling, Simulation, and Scientific Computing, October 19, 2020, 2050055. http://dx.doi.org/10.1142/s1793962320500555.

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In this paper, preconditioned gridless methods are developed for solving the three-dimensional (3D) Euler equations at low Mach numbers. The preconditioned system is obtained by multiplying a preconditioning matrix of the type of Weiss and Smith to the time derivative of the 3D Euler equations, which are discretized under the clouds of points distributed in the computational domain by using a gridless technique. The implementations of the preconditioned gridless methods are mainly based on the frame of the traditional gridless method without preconditioning, which may fail to have convergence for flow simulations at low Mach numbers, therefore the modifications corresponding to the affected terms of preconditioning are mainly addressed in the paper. An explicit four-stage Runge–Kutta scheme is first applied for time integration, and the lower-upper symmetric Gauss-Seidel (LU-SGS) algorithm is then introduced to form the implicit counterpart to have the further speed up of the convergence. Both the resulting explicit and implicit preconditioned gridless methods are validated by simulating flows over two academic bodies like sphere or hemispherical headform, and transonic and nearly incompressible flows over one aerodynamic ONERA M6 wing. The gridless clouds of both regular and irregular points are used in the simulations, which demonstrates the ability of the method presented for coping with flows over complicated aerodynamic geometries. Numerical results of surface pressure distributions agree well with available experimental data or simulated solutions in the literature. The numerical results also show that the preconditioned gridless methods presented still functions for compressible transonic flow simulations and additionally, for nearly incompressible flow simulations at low Mach numbers as well. The convergence of the implicit preconditioned gridless method, as expected, is much faster than its explicit counterpart.

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Journal articles on the topic 'Onera M6'

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