高木 亮治

This study presents a fully automated Cartesian-grid-based compressible flow solver, named FrontFlow/Violet Hierarchical Cartesian for Aeronautics based on Compressible-flow Equations (FFVHC-ACE), for large-eddy simulation (LES) and its aeronautical applications. FFVHC-ACE enables high-fidelity LES of high-Reynolds-number flows around complex geometries by adopting three key numerical methods: hierarchical Cartesian grids, wall modeling in LES, and the kinetic-energy and entropy preserving (KEEP) scheme. The hierarchical Cartesian grids allow fully automated grid generation for complex geometries in FFVHC-ACE, and high-fidelity LES of high-Reynolds-number flows around complex geometries is realized by the wall modeling and the KEEP scheme on the non-body-fitted hierarchical Cartesian grids. We apply FFVHC-ACE to wall-modeled LES around high-lift aircraft configurations at wind-tunnel-scale Reynolds number [Formula: see text] and real-flight Reynolds number [Formula: see text], demonstrating the capability of FFVHC-ACE for fully automated grid generation and high-fidelity simulations around complex aircraft configurations. The computed flowfield and aerodynamic forces at the wind-tunnel-scale Reynolds number agree well with the experimental data provided in the past AIAA High Lift Prediction Workshop (Rumsey et al., Journal of Aircraft, Vol. 56, No. 2, 2019, pp. 621–644). Furthermore, the wall-modeled LES at the real-flight Reynolds number shows good agreement of the lift coefficient with flight-test data.

Electrohydrodynamic (EHD) thrusters can silently propel small unmanned aerial vehicles without moving parts using corona discharges. Computational fluid dynamics would be a powerful tool to model the EHD thrusters and then optimize them. The drift-diffusion-Poisson equations govern corona discharges; hence, the equations can predict the current-voltage characteristics curves of EHD thrusters. However, the equations are too stiff to analyze EHD thrusters in the time domain. Here, we propose a perturbation technique to efficiently solve the stiff drift-diffusion-Poisson system in global (i.e., full two-dimensional or three-dimensional) and nonlinear (i.e., applied voltages higher than the corona inception voltage) regimes. Furthermore, we validated the method with the experimental results of a previous study. (C) 2022 Author(s).All article content, except where otherwise noted, is licensed under a Creative Commons Attribution (CC BY)license (http://creativecommons.org/licenses/by/4.0/).

To realize wall-modeled large-eddy simulations on Cartesian grids, a wall-boundary condition for non-body-conforming grids is developed. A slip-wall velocity boundary condition is introduced to reduce the error of the conservation law caused by the under-resolved velocity profile near the wall boundary. Corresponding to the slip velocity, the stress balance near the wall boundary is analyzed. To satisfy the stress balance, an anisotropic expression of the viscous-stress tensor is introduced that tends not to inhibit the wall-normal mixing by the turbulence. In the simulation of the flat-plate turbulent boundary layer, the result obtained by the developed wall-boundary condition shows a significant improvement from the results using the non-slip boundary condition in predicting both velocity and Reynolds shear stress profiles in the boundary layer.

This paper describes LS-FLOW-HO, a high-order compressible flow solver based on the Flux Reconstruction(FR) method, and its performance optimization. The Flux Reconstruction method achieves arbitrary high-order accuracy on unstructured grids and is suitable for many core architectures because of the local data sets (Stencil) involved in spatial discretization. This study focuses on the performance optimization of the PRIMEHPC FX100, a Fujitsu scalar supercomputer. First, the execution time of sample code that uses the BLAS library is compared with that of code that uses a sparse matrix multiplication which calculates only non-zero values. It is found that the sparse matrix multiplication takes less time than using DGEMM for hexahedral elements when the degree of interpolation polynomial is higher than 2. Using sparse matrix multiplication, hot spot tuning was done by extracting each subroutine code from LS-FLOW-HO. The speedup was confirmed by changing the array structure in the cell boundary, improving the memory/cache access latency by the sequential memory access, and increasing loop length by loop collapsing. Applying these tunings to LS-FLOW-HO, execution time was reduced by up to 40%, and reached 10.23% of the theoretical FLOPS peak using 16 threads of OpenMP on a single node. The performance on Intel Haswell was also shown as the execution time is reduced by about 49%. It was confirmed that the proposed techniques are effective on other processors. Finally, sustained strong scaling performance for real application to supersonic jets is demonstrated using 32 to 3200 nodes (1024 to 102400 cores).

A new dynamic mode decomposition (DMD) method is introduced for simultaneous system identification and denoising in conjunction with the adoption of an extended Kalman filter algorithm. The present paper explains the extended-Kalman-filter-based DMD (EKFDMD) algorithm which is an online algorithm for dataset for a small number of degree of freedom (DoF). It also illustrates that EKFDMD requires significant numerical resources for many-degree-of-freedom (many-DoF) problems and that the combination with truncated proper orthogonal decomposition (trPOD) helps us to apply the EKFDMD algorithm to many-DoF problems, though it prevents the algorithm from being fully online. The numerical experiments of a noisy dataset with a small number of DoFs illustrate that EKFDMD can estimate eigenvalues better than or as well as the existing algorithms, whereas EKFDMD can also denoise the original dataset online. In particular, EKFDMD performs better than existing algorithms for the case in which system noise is present. The EKFDMD with trPOD, which unfortunately is not fully online, can be successfully applied to many-DoF problems, including a fluid-problem example, and the results reveal the superior performance of system identification and denoising.

A novel dynamic mode decomposition (DMD) method based on a Kalman filter is proposed. This paper explains the fast algorithm of the proposed Kalman filter DMD (KFDMD) in combination with truncated proper orthogonal decomposition for many-degree-of-freedom problems. Numerical experiments reveal that KFDMD can estimate eigenmodes more precisely compared with standard DMD or total least-squares DMD (tlsDMD) methods for the severe noise condition if the nature of the observation noise is known, though tlsDMD works better than KFDMD in the low and medium noise level. Moreover, KFDMD can track the eigenmodes precisely even when the system matrix varies with time similar to online DMD, and this extension is naturally conducted owing to the characteristics of the Kalman filter. In summary, the KFDMD is a promising tool with strong antinoise characteristics for analyzing sequential datasets. (c) 2018 Author(s).

<p>An adaptive estimation method for nonlinear structural deformations is presented. The method is based on the ensemble Kalman filter (EnKF), which can effectively handle nonlinearities in structural models by using the Monte-Carlo simulation. In this study, a self-tuning algorithm for the system noise in the conventional Kalman filter is extended and applied for tuning in the EnKF. To verify the effectiveness of the presented method, a numerical experiment was performed for a deployable frame structure system that contains the typical nonlinearities of space deployable structures, that is, the geometrical nonlinearity in flexible members and the cable nonlinearity resulting from its slackened state.</p>

Transonic flowfield of a generic payload fairing with cone-cylinder configuration at Mach 0.8 is computed with the hybrid of Large-Eddy/Reynolds-Averaged Navier-Stokes simulations based on the Improved Delayed Detached Eddy Simulation. Two types of the grid having 38M and 166M points are employed. It is found that turbulent boundary layer developed at the nose cone is relaminarized through expansion fan generated at shoulder of the fairing. Then, shock wave/boundary layer interaction followed by separation bubble is formed. At the downstream, isotropic turbulent flow is obtained. However, it is unclear if laminar to turbulent transition occurs before or after the shock wave/boundary layer interaction within the present numerical result. For the simulation of transonic flowfield around payload fairing, further study is required to investigate characteristics of the Improved Delayed Detached Eddy Simulation on the relaminarization, laminar to turbulent transition of the boundary layer, and shock wave/boundary layer interaction.

The thermal condition of ASTRO-H under air-cooled environment before launch is investigated with a thermal testing and a computational analysis. The thermal testing shows that the temperatures of devices are confirmed to be within the operating range if the additional fans are used. Moreover, the results of the thermal testing are compared with those of computational results. The computational results of temperature of the devices around the dewar with the additional fans are in good agreement with those of the thermal testing. The good agreement in the condition with the additional fans is because the forced convection, which is a dominant effect, is well captured in the computational analysis. Meanwhile, the computational results of temperature on the side panels are in very good agreement with thermal testing despite the difference in the flow outside satellite by air conditioner: computational analysis models the air flow from the air-conditioner while thermal testing does not. This is because the air-flow is very slow (0.1[m/s] at the side panel locations) and forced convection effects are very small.

Experiment using the 2m× 2m transonic wind tunnel in JAXA and numerical simulation based on the RANS are carried out to investigate transonic flowfield influenced by a protuberance attached to shoulder of a generic cone-cylinder type rocket faring. Thickness of the protuberance considered here is about 0.1% of the fairing diameter, which is 1.8 times thicker than the displacement thickness of the incoming boundary layer. Validation and verification of the numerical method are firstly performed based on the experimental result of the clean fairing without any protuberance. It is found that effect of the protuberance changes depending on the flow speed. If the free-stream Mach number is less than 0.8, separation shock wave is generated at the protuberance on the shoulder, although the separation shock wave appears at the cylinder section of the clean fairing. On the other hand, little influence on the flow structure is observed in the flow condition with higher free-stream Mach number. The effect revealed in this study indicates the importance of the protuberance to the design of rocket fairing.

The scalar tuning of a compressible fluid solver for a supercomputer with a distributed memory architecture is conducted. We use the K computer which is one of the peta-scale supercomputers recently developed in Japan. A computational code "LANS3D" and its high-order compact differencing option are tuned. The original version of the code achieves approximately 4.5% of full performance of CPU for the simple test case. Scalar tuning based on combining do-loops works well, and the tuned code attains about 10% of full performance for the same case. The reasons are the improvement in the use of the cache, the suppression of the data transfer, and the efficient use of the data that once transferred to the cache from the memory that results in hiding the low speed of data transfer. The tuned code becomes twice faster than the original one in the wall-clock time and enables us to perform over-160-case parametric study about airfoil flow computation by large-eddy simulations with high-order accurate and high resolution numerical scheme. © 2013 All rights reserved. ISSR Journals.

© 2012 7th International Conference on Computational Fluid Dynamics, ICCFD 2012. All rights reserved. Impacts on the spatial and temporal resolutions are discussed through the twodimensional model problem of jet impinging which is proposed by the present authors and Housman et al.[AIAA paper 2011-3650,2011]. The result shows that the high-resolution schemes improve the resolution of fine structures of vortices, though even a conventional scheme can predict the blast waves well. For solving the fine structure of vortices, high-order scheme is more than 10 times as efficient as conventional scheme.

The accurate evaluation of the ground reflection effect is of importance for acoustic measurement conducted during static-firing tests of rocket motors. The effectiveness of existing acoustic impedance models was examined by comparing experimental data collected from acoustically hard surfaces. Through these comparisons and the subsequent evaluation of the effect of meteorological conditions, it was confirmed that the existing impedance models are not suitable for the evaluation of long-distance propagation over a hard surface, which corresponds to the far-field conditions in the present static-firing tests of solid rocket motors. A new acoustic impedance model is proposed in this paper. Comparisons of the acoustic measurement data indicated that the new model is effective for both near- and far-field propagation of acoustic waves. The proposed model was applied to acoustic data collected during static-firing tests of solid rocket motors assuming distributed acoustic sources along the exhaust jet axis in order to remove the ground reflection effect.

The high-specification computational power of Japan Aerospace Exploration Agency's new supercomputer system, called Fujitsu FX1 cluster, enables us to perform really macroscale 3-D situations with full particle plasma simulation [particle-in-cell (PIC) method]. A fully 3-D kinetic approach to collisionless shock problems, which is one of the most important problems in the space plasma science, is possible, and a challenging run is being executed for a pioneering study of the topic. About 0.4 billion grids are allocated for the electromagnetic fields, and about 0.1 trillion particles are loaded into the simulation run. The computational efficiency of the PIC code is about 8% of the peak performance (4.6 Tflops) using 5776 CPU cores (57 Tflops). The simulation parameters were selected to simulate ESA's Cluster-II spacecraft observational result reported by Seki et al. (in 2009). The full mass ratio m(i)/m(epsilon) = 1840 was taken for this simulation, and almost one ion inertia length square could be allocated for the simulation. In this simulation, a quite complicated wave activity is found in the shock foot region. In this paper, comparing 3-D results with 2-D simulation results, a 3-D nature of shock transition region of quasi-perpendicular shock is reported.

Activity on numerical plasma simulations by JAXA&rsquo;s Engineering Digital Innovation (JEDI) Center is overviewed. Currently, R&D of two major numerical tools is conducted. First one is spacecraft charging analysis tool that can compute charging status of a spacecraft solving charged particle motions precisely. Using this information, we can evaluate onboard measurement of electrostatic probes or consider better configuration of onboard equipment of a spacecraft. Computation example of the interactions between solar wind plasma and a solar sail is shown in this paper. Second one is a numerical tool called JIEDI (JAXA&rsquo;s Ion Engine Development Initiative), which aims to reduce the cost and the time required for an ion thruster life test. The JIEDI tool can numerically estimate ion bombardment to an ion thruster&rsquo;s grid material to predict the erosion rate of the grid material, and preliminary analysis by the JIEDI tool showed good agreement with the real-time life test of a microwave ion thruster.

Evaluation of ground effect is important for acoustic measurement in static-firing tests on rocket motors. The effectiveness of existing acoustic impedance models is examined by comparing with some experimental results. Through the comparison and evaluation of effect of meteorological condition, it is confirmed that the existing impedance models are unsatisfactory for the evaluation of long distance propagation over a hard surface, which corresponds to the far field condition in the present static-firing tests of rocket motors. In this study, a new acoustic impedance model is proposed. From the comparison with the existing acoustic measurement data, it is shown that the new model is effective for both of near and far field propagation. The proposed model is applied to acoustic data measured in the static-firing tests of solid rocket motors, assuming distributed acoustic sources along the exhaust jet axis.

Aerodynamic characteristics of a vertical landing rocket vehicle are computationally investigated under subsonic and supersonic flight conditions as a preliminary study for the concept design using a light optimization method and a light CFD tool. The results show that the simulations with a coarse grid can accurately estimate the aerodynamic characteristics like axial force coefficient and the lift-to-drag ratio. The results of the light aerodynamic shape optimization indicate tradeoff information among objective functions, and the correlation between design variables and objective functions. The preliminary knowledge for the aerodynamic shape design is obtained. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

Aerodynamic characteristics of a reusable observation vehicle are computationally investigated under subsonic and supersonic flight conditions as a preliminary study for the concept design using a design exploration method and a light CFD tool. The results show that the simulations with a coarse grid can accurately estimate the aerodynamic characteristics like axial force coefficient and the lift to drag ratio. The results of the aerodynamic shape exploration indicate tradeoff information among objective functions, and the correlation between design variables and objective functions. The preliminary knowledge for the aerodynamic shape design is obtaind.

Attenuation of sound by water droplets in air is analytically investigated. This is one of the mechanisms of the noise reduction in actual rocket launch where water is injected to jet plume. The scattering and absorption by the droplets are considered as the attenuation mechanisms. It is found that the absorption is superior to the scattering for fine droplets, such as fog or mist. Then we estimate the attenuation of sound for the experiment using subscale solid rocket motors. The result agrees well with the experimental data within the frequency range where the assumptions of this study are valid.

Since acoustic wave radiated from the rocket plume causes severe acoustic loading on the payload at lift-off, prediction and reduction of acoustic level is an important issue in the design phase. Acoustic environment is empirically predicted as of now. However, the acoustic generation mechanism is not clearly understood, which results in unsatisfactory prediction accuracy. In order to overcome the limitations of the empirical method, CFD study is carried out in JAXA. Through the comparison with the measured result in the static-firing tests, it is found that prediction accuracy of the CFD superiors to the empirical method. In addition, the acoustic generation becomes clear, which gives us the idea of developing the launch-pad which reduces acoustic level of the vehicle.

Numerical simulation was carried out to investigate the generation mechanism of pressure waves radiated from the H-IIA launch vehicle at lift-off. It was revealed that the Mach wave due to the large-scale structure of the unsteady supersonic exhaust plumes is the dominant noise source. The Mach wave propagating obliquely downstream is reflected from the constructions of the launch-pad, and then, turns to reach the vehicles, resulting in the acoustic loading. It was also found that the fluctuating supersonic plume entering into the flame duct is the dominant noise source that appears in the flame duct. Then, the pressure wave propagates through the flame duct and is ejected outside to the vehicle. Copyright © 2008 by American Institute of Aeronautics and Astronautics, Inc.

The generation mechanism of the pressure waves radiated from Japanese former solid rocket, M-V, at lift-off was studied numerically. Bowl-shaped flame deflector placed underneath the 1st stage motor was revealed to induce the unsteadiness of the plume, which generates strong pressure wave propagating directly to the vehicle. Promising launch-pad for the advanced solid rocket which utilizes the slope terrain of the launch-site was also analyzed, and compared with the result of the M-V. It turns out that the intense pressure wave is radiated from the flapping motion of the supersonic plume on the sloping deflector. Due to the inclination of the sloping deflector, the pressure wave propagates away from the vehicle, resulting in the better acoustic environment. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

Aeroacoustic mechanisms of an axisymmetric over-expanded supersonic jet impinging on a flat plate with and without hole are numerically investigated. High-order weighted compact nonlinear scheme is used to simulate the unsteady flow including shock waves and sound radiation in the near field of the jet. Analyses Of Unsteady flowfield and related near-sound field reasonably identify three major noise generation mechanisms, that is, noises from Mach wave, shock cell-shear layer interaction and small fluctuations of jet shear layer. Especially, intense noise radiation in the form of Mach waves and its reflection at the plate predominates the noises from the other two finer sources. The simulated distributions of sound source power and its frequency along the jet axis qualitatively well coincide with typical experimental data used in NASA SP-8072. Similar sound pressure spectrum shape is obtained both the cases of flat plate with and without hole, but the case of without hole shows higher SPL by several dB than that of with hole due to the stronger Mach wave radiation. Aeroacoustic flowfield is drastically affected by the Reynolds number because the jet shear layer instability directly causes the strength of acoustic waves.

Gas flow over a two-dimensional airfoil at very low Reynolds number is investigated in order to understand basic aerodynamic characteristics related to design of Micro Air Vehicle (MAV) for planetary exploration. Before the investigations, verification was conducted for the current numerical approach, which are commonly used and validated for high Reynolds number flow analysis, showing good applicability for low Reynolds number flow analysis. Flow around NACA4402 has been investigated for the condition of Mach number of 0.1 and Reynolds number of 1,000. Investigation shows that Reynolds number has a substantial influence on aerodynamic characteristics of the airfoil in low Reynolds number flow. In contrast, Mach number has a slight influence in comparison with Reynolds number.

UPACS, Unified Platform for Aerospace Computational Simulation, is a project to develop a common CFD platform since 1998 at National Aerospace Laboratory of Japan. The project aims not only to overcome the increasing difficulty in recent CFD code programming on parallel computers for aerospace applications, but also to accelerate the development of CFD technology by sharing a common base code among research scientists and engineers. The latest UPACS can solve not only compressible flows around fuselages but also flows of rotating cascade, very low speed flows, and flows coupled with heat conduction in objects.

A computational parametric study is presented illustrating the effects of upstream heat release on drag and heating rate in hypersonic flow over an axisymmetric blunt body. A chemical nonequilibrium viscous flow is considered with seven species; Park's two-temperature model is also used to take account of thermally nonequilibrium phenomena. Three parameters that control the heat release are introduced, and their effects on the flow structure, specifically with respect to drag and heating rate, are quantified. Results show that heat release upstream of the body can reduce not only the aerodynamic drag but also the aerodynamic heating rate at the body surface. Computed results achieved a reduction of 23% in drag and 74% in heating rate, relative to baseline values in the absence of heat release. The reduction mechanism is discussed, and dynamic pressure is shown to play an important role in the drag reduction mechanism.