研究者業績
基本情報
- 所属
- 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 教授東京大学大学院 教授東京理科大学 理工学研究科機械工学専攻 客員教授
- 学位
- 博士（工学）(2000年3月 東北大学大学院)
- 研究者番号
- 10373440
- J-GLOBAL ID
- 200901044748363926
- researchmap会員ID
- 5000069161
- 外部リンク
宇宙科学航空研究開発機構宇宙科学研究所の大山です．
自分の研究分野にとらわれず，新しい研究分野にも挑戦していきたいと考えています．
自分の研究分野にとらわれず，新しい研究分野にも挑戦していきたいと考えています．
研究キーワード
17研究分野
6経歴
13-
2023年12月 - 現在
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2023年4月 - 現在
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2019年4月 - 現在
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2015年4月 - 2023年11月
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2010年4月 - 2023年3月
学歴
5-
1997年4月 - 2000年3月
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- 2000年
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1995年4月 - 1997年3月
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1991年4月 - 1995年3月
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- 1995年
委員歴
7-
2020年10月 - 現在
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2010年6月 - 現在
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2018年10月 - 2020年9月
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2017年4月 - 2019年3月
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2015年4月 - 2017年3月
受賞
13-
2022年4月
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2021年5月
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2020年11月
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2020年2月
論文
91-
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 67(1) 12-22 2024年
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60(5) 1721-1727 2023年9月
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Journal of Computational Physics 468 111505-111505 2022年11月 査読有り
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Physical Review Fluids 7(8) 2022年8月19日 査読有り
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Measurement Science and Technology 32(12) 125306-125306 2021年12月1日 査読有りThe methods that measure the turbulence statistics distribution of high-speed flow with high spatial resolution using particle-image-velocimetry images are proposed, and their performances are verified. Two methods are proposed, and the problems caused by blurring the particle shape due to the high-speed movements are resolved. While the conventional method approximates the unique particle shapes with a circular distribution for image pairs, the first proposed method approximates the different particle shapes with an ellipse for first and second images of pairs and reduces the effects of the blur in the flow direction of the particles. Meanwhile, the second proposed method treats the general particle shapes such as the blur of temporally changing laser intensity and adopts the deconvolution analysis using a Fourier transform. Synthetic particle images were created and a supersonic jet test was performed, and the proposed methods were evaluated to be superior to the previous method for the estimation of turbulent fluctuation using those data.
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Engineering Research Express 3(1) 015035-015035 2021年3月1日 査読有り<title>Abstract</title> A fixed-wing aircraft called Mars Airplane Balloon Experiment −2 (MABE2) developed by the authors was the subject of high-altitude flight demonstration test in this paper, which can simulate the near-actual environment of a Martian atmospheric flight. Although the flight condition is in the low-Reynolds-number region due to low density at high altitude, the wing suffers from aeroelastic deformation given the relatively high dynamic pressure load in the pull-up phase. Stereophotogrammetry was applied in the 6.5 m × 5.5 m low-speed wind tunnel at the Japan Aerospace Exploration Agency (JAXA), aiming to optically measure MABE2’s aeroelastic deformation under dynamic pressure loads equivalent to high-altitude flight test, with the MABE2’s reinforced structural strength. The results of the accuracy test indicated that stereophotogrammetry measures aeroelastic deformation at high accuracy of ±0.1 mm around the image center and ±0.3 mm around the edge. A slight deflection of up to 4 mm was observed on the main wing, whereas both the main and tail wings were hardly twisted. Compared with flight-simulation-assumed errors, these deformations are extremely small and have a negligible effect on the high-altitude flight test. The study results confirmed the practicality and efficiency of this optical measurement technique in aeroelastic deformation measurement for a real light aircraft.
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Accelerating Space Commerce, Exploration, and New Discovery conference, ASCEND 2021 2021年The Japan Aerospace Exploration Agency, in partnership with academia and industry, are developing the Air Turbo Rocket for Innovative Unmanned Mission (ATRIUM) engine: an air turboramjet + rocket combine cycle propulsion system intended to replace conventional liquid rocket engines in Vertical Takeoff Vertical Landing applications, such as reusable sounding rockets. A subscale Flight Test Bed (FTB) vehicle is also being developed to demonstrate the ATRIUM engine in a flight environment. In this paper, the ATRIUM engine and FTB vehicle are introduced, and current progress in their development is summarized. Future test plans and practical applications are also discussed.
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Physics of Fluids 32(4) 2020年4月1日 査読有り
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Journal of Verification, Validation and Uncertainty Quantification 5(4) 041004 2020年 査読有り
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Transactions of the Japan society for aeronautical and space sciences 63(4) 172-184 2020年 査読有り
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進化計算学会論文誌 11(2) 2020年 査読有り
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Aerospace technology Japan 18(2) 17-24 2020年 査読有り
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AIAA Journal Vol. 58(No. 10) 4260-4270 2020年 査読有りThe effectiveness of closed-loop flow control using a deep Q network (DQN), such as deep reinforcement learning algorithm, was experimentally researched for a NACA0015 airfoil. The system used a dielectric barrier discharge plasma actuator as the flow control device, and the experiment was conducted at the chord Reynolds number of 6.3 x 10(4). The closed-loop control system selected a nondimensional burst frequency of the actuator by analyzing the time series of the surface pressure data. The neural network of the DQN was sequentially trained at the angles of attack of 12 and 15 deg. As a result, the closed-loop control successfully generated a higher control gain compared with the open-loop control using a fixed burst frequency. Particularly, at 15 deg there was a significant control gain that could not be obtained by the open-loop control. The closed-loop control keeps the flow attached and preserves it for a longer time by periodically switching the actuator on and off.
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Transactions of the Japan Society for Aeronautical and Space Sciences 63(1) 8-17 2020年 査読有り© 2020 The Japan Society for Aeronautical and Space Sciences Aerodynamics of an owl-like wing model at low Reynolds numbers (Re = O(104-5)) are investigated using large-eddy simulations with high-resolution computational schemes. The airfoil shape of the owl-like wing model is constructed based on a cross-sectional geometry of the owl wing at 40% wingspan from the root. The chord-based Re ranges from 1.0 © 104 to 5.0 © 104 and the angle of attack (¡) varies from 0 to 14 deg. The time-averaged lift (Cl) and drag coefficients computed are in reasonable agreement with the results of force measurement. The results computed clarify a nonlinear change in the Cl curve slope, which is due to an increase in the suction peaks in conjunction with the change in type of separation, the formation of a laminar separation bubble (LSB), and pressure recovery on the pressure side. The generation of the LSB on the suction and/or pressure sides at the Re of 2.3 © 104 and 4.6 © 104 are seen, while reattachments are observed only on the pressure side at the Re of 1.0 © 104 due to the camber of the wing. Furthermore, the owl-like wing model demonstrates favorable aerodynamic performance in terms of a maximum lift-to-drag ratio in comparison with several airfoils at the Re range considered. This is due to the strong suction peaks and distribution of surface pressure on the pressure side. It is emphasized that the concave lower surface enhances the time-averaged aerodynamic performance at all of the ¡ even though the LSB is generated and fluctuation in lift history is induced at low ¡.
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Trans. JSASS Aerospace Tech. Japan 17(4) 512-518 2019年7月 査読有り
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Journal of Flow Control, Measurement & Visualization, Science Research Publishing 7(2) 2019年4月 査読有りThis paper reports the results of the aerodynamic deformation measurements of the meter-scale, entire shape, actual UAV in the wind tunnel using a video grammetry technique. The measured airplane was the airplane for Mars exploration being developed by Japan Aerospace Exploration Agency (JAXA) and Japanese universities. Its main wing span length was 2.4 m. The video grammetry measurement was performed using VICON’s system. Retroactive markers and stickers were put on the airplane. JAXA’s 6.5 m × 5.5 m Low-Speed Wind Tunnel was used. The airplane was mounted on the strut support with pitch-free or pitch-locked conditions. The deformations of the main wing bending, the main wing twisting, the tail boom bending, and the elevator deflection angle change were revealed quantitatively. The bending stiffness of a main wing spar that was designed as a safety factor of 2.8 at load factor of 5 was sufficient. The main wing spar was located around a center of pressure of an airfoil and it showed enough stiffness for twisting at nominal condition. The effects of the main wing bending and twisting, and the tail boom bending on the aerodynamic performance were estimated but they were in an acceptable range from the standpoint of the controllability of the aerodynamic performance using control surfaces. Even though the servo motor was located near the elevator and the linkage between the servo motor and the elevator was short, the measured elevator deflection angle was at most 4% smaller than the angle at no-wind condition. The obtained results and presented method are useful for control, flight data analysis, and design of lightweight airplanes.
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Journal of Fluid Science and Technology 14(3) JFST0017-JFST0017 2019年 査読有り<p>We redesigned the Mars Airplane Balloon Experiment Two (MABE-2) based on MABE-1 to improve the vehicle's stability and controllability. Following the redesign, the MABE-2 vehicle had a larger horizontal tail volume than that of MABE-1 for improved stability performance. In addition, to further improve the stability and control characteristics, a rectangular planform was employed for the horizontal tail wing; in contrast, MABE-1 had a tapered planform. The vertical tail position of MABE-2 was moved to the end of the horizontal tail wing, because the vertical tail of MABE-1, which was positioned at the mid span of the horizontal tail wing, showed aerodynamic interaction with the horizontal tail wing. In this paper, we discussed the aerodynamic performance of a control surface based on computational fluid dynamics with variation in the deflection angle between the control surface and the horizontal tail (elevator), and we examined the effects of this redesign on longitudinal control characteristics. Numerical investigations confirmed the linear variation in the pitching moment and the aerodynamic force with the changing elevator deflection angle in MABE-2. Surface pressure observations indicated that MABE-2 shows a smooth variation in the pressure distribution with changing elevator deflection angle, while MABE-1 does not. These results demonstrate that the aerodynamic control characteristics of MABE-2 were improved in comparison to those of MABE-1.</p>
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Theoretical and Computational Fluid Dynamics 32 805-820 2018年12月 査読有り
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AIAA Aerospace Sciences Meeting, 2018 *(210059) *-* 2018年 査読有り© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. The numerical analysis of the supersonic jet is conducted using large eddy simulations (LES) with high order schemes and the grid of approximately 650 million points. A Mach number and a Reynolds number are set to be 2.0 and 9.0×105, respectively. At first, we confirm the azimuthal grid resolution. As a result, it seems that the flow field and the acoustic field near the jet flow are slightly affected by changing the grid resolution, while the sound pressure level at far-field converges sufficiently with the present grid number. The computational flow field shows good agreement as compared with the experimental data. Moreover, it is shown that the sound pressure level at far-field can be predicted within 4dB difference as compared to the experimental data. Next, the effect of the jet temperature of the supersonic jet on the acoustic waves is investigated. The temperature ratio of the chamber to ambient air is set to be 1.0, 2.7, and 4.0 for the cold, mid-hot and hot jets, respectively. Mach waves are radiated from the supersonic jet toward downstream. we confirmed that the shorter potential core length, the higher sound pressure level, the larger angle of Mach waves with increasing jet temperature.
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AIAA Aerospace Sciences Meeting, 2018 (210059) 2018年 査読有り© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The relation between the parameters of the body-force field generated by a plasma actuator and the maximum induced velocity in quiescent air is investigated by expressing the body-force distribution as the Gaussian function of the spatial coordinates. The aim of this study is to identify the dominant parameters for modeling of the body-force distribution. For that purpose, the parametric study using numerical simulations and dimensional analysis are conducted to derive the nondimensional key parameters. It is found that the nondimensional maximum induced velocity is determined by the Reynolds number calculated by three parameters: the total induced momentum per unit time, the height of the center of gravity of the body-force distribution, and the standard deviation from the center of gravity. In addition, the relation for the Gaussian body-force distribution turns out to be applicable to a conventional model, i.e, the Suzen model, even though the shapes of the distribution differ. Thus, we conclude that the three body-force parameters above are the key parameters for the maximum velocity induced by a plasma actuator.
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AIAA Aerospace Sciences Meeting, 2018 (210059) 2018年 査読有りAerodynamics optimization of airfoil for flying-wing type Mars airplane is conducted to show tradeoff between maximization of aerodynamic performance and stability requirement. To see the tradeoff between the two objectives, a multiobjective evolutionary algorithm is used where aerodynamic performance of each airfoil is evaluated with a two-dimensional laminar Navier-Stokes flow solver. Some design constraints are applied to obtain feasible airfoil shapes. The obtained Pareto-optimal airfoils show tradeoff between the two design objectives. Analysis of the flow field of the obtained airfoils leads to some useful knowledge in designing airfoil for flying-wing type Mars airplane.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 61(6) 258-267 2018年 査読有り
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Transactions of the Japan Society for Aeronautical and Space Sciences 61(6) 281-284 2018年 査読有り
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Fluid Mechanics Research International Journal 1(3) 2017年11月 査読有り
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JOURNAL OF SPACECRAFT AND ROCKETS 54(4) 796-807 2017年7月 査読有りThe present study focuses on exploring optimal transfer solutions for the Demonstration and Experiment of Space Technology for INterplanetary voYage (DESTINY) mission, which was proposed as an Institute of Space and Astronautical Science (ISAS) Epsilon-class small program in 2013 based on the "Space Science & Exploration Roadmap" proposed by the ISAS and later approved by the government committee on space policy. During the DESTINY mission, the spacecraft will first be injected into a low elliptical orbit by an Epsilon rocket, and its altitude will then be increased to approach the Moon using an ion engine system. Next, the spacecraft will be injected via a transfer orbit to the L2 Halo orbit of the Sun-Earth system by way of gravity assist from the Moon. As the spacecraft revolves around the Earth several hundred times, it gradually increases its altitude. Thus, the launch time and thrust profile must be chosen properly. Note that there are several conflicting requirements, including the reduction of fuel consumption, total flight time, and maximum eclipse time, that must be taken into consideration. To satisfy these requirements, many-objective evolutionary computation is applied to find better orbital designs.
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A Dual-Polarization X-Band Traveling-Wave Antenna Panel for Small-Satellite Synthetic Aperture RadarIEEE TRANSACTIONS ON ANTENNAS AND PROPAGATION 65(5) 2144-2156 2017年5月 査読有りThis paper investigated the possibility of using an X-band traveling-wave antenna for application in synthetic aperture radar (SAR) on a small-satellite platform. A rectangular slotted parallel-plate antenna panel capable of radiating dual polarization over the same physical aperture was selected for this purpose. The design procedure, which provides for an optimal antenna pattern in a SAR application, is described. A metaheuristic multiobjective optimization algorithm was applied to synthesize an appropriate far-field pattern, which enhances the quality of the SAR image. Subsequently, the optimization results were used to build a more realistic model of the antenna panel in electromagnetic simulation software. An antenna panel was fabricated, and its electrical characteristics were measured. The fabricated prototype weighed 1310 g and measurements showed nearly 50% aperture efficiency for both circular polarizations [34.6-dBic right-hand circular polarized (RHCP), 34.5-dBic left-hand circular polarized (LHCP), and peak gain at 9.65 GHz], with low side-lobes. A beam shift of 2.1 degrees was observed between the RHCP and LHCP beams at the design center frequency and was later analyzed to be the result of an error in the modeling of the adhesive layer.
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INTERNATIONAL JOURNAL OF COMPUTATIONAL FLUID DYNAMICS 31(1) 57-67 2017年 査読有りThis study investigates the predictability of the aerodynamic performance of some numerical methods at low Reynolds numbers and their dependency on the geometric shape of airfoil. We conducted three-dimensional large-eddy simulations (3-D LES), two-dimensional laminar simulations (2-D Lam), and Reynolds-averaged Navier-Stokes simulations with Baldwin-Lomax (2-D RANS(BL)) and Spalart-Allmaras (2-D RANS(SA)) turbulence models. Although there is little discrepancy between the 3-D LES, 2-D Lam, and 2-D RANS(SA) results in terms of the lift and drag characteristics, significant differences are observed in the predictability of the separation and reattachment points. The predicted lift, separation, and reattachment points of the 2-D Lam are qualitatively similar to those of the 3-D LES, except for high angles of attack at which a massive separation occurs. The 2-D RANS(SA) shows good predictability of the lift and separation points, but it does not estimate reattachment points accurately. The 2-D RANS(BL) fails to predict the precise separation points, which results in a poor lift predictability. These characteristics appear regardless of the airfoil geometry shapes. The results suggest that a 2-D Lam without any turbulence models can be used to estimate qualitative airfoil aerodynamic characteristics at the low Reynolds numbers.
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IEEE ACCESS 5 19597-19619 2017年 査読有りRecently, a large number of multi-objective evolutionary algorithms (MOEAs) for many-objective optimization problems have been proposed in the evolutionary computation community. However, an exhaustive benchmarking study has never been performed. As a result, the performance of the MOEAs has not been well understood yet. Moreover, in almost all previous studies, the performance of the MOEAs was evaluated based on nondominated solutions in the final population at the end of the search. Such traditional benchmarking methodology has several critical issues. In this paper, we exhaustively investigate the anytime performance of 21 MOEAs using an unbounded external archive (UEA), which stores all nondominated solutions found during the search process. Each MOEA is evaluated under two optimization scenarios called UEA and reduced UEA in addition to the standard final population scenario. These two scenarios are more practical in real-world applications than the final population scenario. Experimental results obtained under the two scenarios are significantly different from the previously reported results under the final population scenario. For example, results on the Walking Fish Group test problems with up to six objectives indicate that some recently proposed MOEAs are outperformed by some classical MOEAs. We also analyze the reason why some classical MOEAs work well under the UEA and the reduced UEA scenarios.
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INTERNATIONAL JOURNAL OF HEAT AND FLUID FLOW 62 10-23 2016年12月 査読有りHigh-fidelity computations of the flow control around a pitching NACA63(3) - 618 airfoil by a plasma actuator are conducted. The effectiveness of the plasma actuator and the effects of its installation position are investigated. The plasma actuator is installed at x/c = 0%, 10%, and 60% from the leading edge of the airfoil. The installation position of 60% is chosen based on the investigation of the uncontrolled flow field; the case with this position successfully enhanced the aerodynamic performances of the airfoil. The results show the importance of a priori investigation of the separation and the reattachment points for an uncontrolled flow-field. In addition, the results illustrate that a properly installed and actuated plasma actuator is capable of controlling the dynamic flowfields and improving the aerodynamic performances of an airfoil. (C) 2016 Elsevier Inc. All rights reserved.
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JOURNAL OF SPACECRAFT AND ROCKETS 53(4) 751-758 2016年7月 査読有りThis study explores possible design of an aeroacoustic flame deflector for a rocket launch pad. The design objectives are 1)minimization of the overall sound pressure level near the payload fairing, 2)minimization of the time-averaged maximum pressure on the flame-deflector surface, and 3)minimization of the shape difference from a flat plate inclined at 45deg. The acoustic wave characteristics associated with deflector shapes are identified by large-eddy simulations. To overcome difficulties of required computational time, the following are adopted: 1)a high-order scheme that reduces the computational cost of large-eddy simulations, 2)a multi-objective evolutionary algorithm for efficient parallelization, and 3)large-scale parallelization on the Japanese supercomputer K. Total computational time for optimization is approximately 350h with 6500 processors of the K computer. The analysis of nondominated (Pareto-optimal) solutions reveals a tradeoff relation and correlation among the objective functions. In the result, there appears a well-balanced solution that significantly reduces the overall sound pressure level. The shape difference is relatively minor, with a small bump located somewhat upstream of the impinging region. The result suggests that the local angle of the inclined deflector near the impinging region plays an important role for the reduction of overall sound pressure level near the rocket fairing.
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2016 IEEE CONGRESS ON EVOLUTIONARY COMPUTATION (CEC) 1139-1148 2016年 査読有りThe advantages of evolutionary computation with very large populations for many-objective optimization problems are investigated. The effects of a population size of up to 1,000,000 are studied, with the number of generations fixed at 100. To overcome difficulty in computational time, we use a many-objective evolutionary algorithm designed for massive parallelization (CHEETAH) on the K supercomputer. For unimodal test problems DTLZ2 and DTLZ4, the inverted generational distance (IGD) decreases as the population increases while the generational distance (GD) is saturated with a population size of 10,000. This means an evolutionary computation with massive population size mainly contributes to improvement of diversity of obtained non-dominated solutions. Even when the total number of evaluations is fixed, this conclusion is unchanged. For the multimodal test problems DTLZ1 and DTLZ3, GD and IGD are reduced with increasing population size of up to 10,000 but are not significantly improved with population sizes larger than this. This is probably due to the difficulty in obtaining good non-dominated solutions for DTLZ1 and DTLZ3 with current CHEETAH. Because CHEETAH is bases on NSGA-II (only the non-dominated sort portion is modified for more effective many-objective optimization and parallelization), we expect that the current conclusion qualitatively stays the same for other NSGA-II-based algorithms. To take advantage of the larger population size, development of operators such as selection and crossover designed for very large population size may be required.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14(30) Pk_111-Pk_116 2016年<p>DESTINY: the Demonstration and Experiment of Space Technology for Interplanetary Voyage, which is a candidate mission of Epsilon Launch Vehicle, plans to execute scientific observations using instruments with the mass of up to about 10 kg on the transfer and Halo orbit of the sun to earth Lagrangian point L1/L2 or on the fly-by orbit of near earth objects (NEO). Potential scientific objects include in-situ observation and remote sensing from these space are solar system explorations, such as, the observations of plasma and energetic particles around the terrestrial magnetosphere, inter-planetary and inter-stellar dust, and NEO. It is also considered to be useful for the pilot observations for future infrared, gamma-ray, and cosmic-ray space astronomical telescope in the deep space. Applied missions of DESTINY will be able to go to deep space with higher mass of payloads. Using the Epsilon Launch Vehicle, it will convey instruments of up to 50 kg to the space between Venus and Mars. DESTINY launched by the improved launch vehicle with the power of M-V rocket will carry payloads of up to 200 kg into the orbit of Venus and Mars. In these phases, Explorations for Venus, Mars, and multiple NEO, and astronomical observations from the deep space observatory will be realized by low cost deep space missions.</p>
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 59(6) 323-331 2016年 査読有り
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Transaction of the Japan Society for Aeronautical Sciences, Aerospace Technology Japan, Aerospace Technology 14(ists30) Pk_47-Pk_53 2016年 査読有り
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International Journal of Machine Learning and Computing 6(1) 2016年1月 査読有り
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14(30) Pk_83-Pk_88 2016年 査読有り<p>Airplanes have been expected as a new platform for Mars exploration. This paper presents conceptual design results of a Mars airplane for various entry capsule diameters, payload mass, and flight range. The conceptual design method includes considerations of the deployment mechanism using hinges and aerodynamic characteristics at low Reynolds number. The results reveal a relation among total mass of the airplane, entry capsule size, payload mass, and flight range. When the total mass and entry capsule diameter are fixed, the flight range is expressed as a linear function of the payload mass with negative slope. When the entry capsule diameter and the payload mass are fixed, optimal total mass from the viewpoint of the flight range is found for each entry capsule diameter. Finally a specification of the airplane at the optimal total mass is shown. An airplane weighing 34.0 kg shows a maximum range of 370 km under the conditions of an aeroshell diameter of 3 m and a payload mass of 3 kg.</p>
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PHYSICS OF FLUIDS 27(2) 2015年2月 査読有りMechanisms behind the pressure distribution and skin friction within a laminar separation bubble (LSB) are investigated by large-eddy simulations around a 5% thickness blunt flat plate at the chord length based Reynolds number 5.0 x 10(3), 6.1 x 10(3), 1.1 x 10(4), and 2.0 x 10(4). The characteristics inside the LSB change with the Reynolds number; a steady laminar separation bubble (LSB_S) at the Reynolds number 5.0 x 10(3) and 6.1 x 10(3), and a steady-fluctuating laminar separation bubble (LSB_SF) at the Reynolds number 1.1 x 10(4), and 2.0 x 10(4). Different characteristics of pressure and skin friction distributions are observed by increasing the Reynolds number, such that a gradual monotonous pressure recovery in the LSB_S and a plateau pressure distribution followed by a rapid pressure recovery region in the LSB_SF. The reasons behind the different characteristics of pressure distributions at different Reynolds numbers are discussed by deriving the Reynolds averaged pressure gradient equation. It is confirmed that the viscous stress distributions near the surface play an important role in determining the formation of different pressure distributions. Depending on the Reynolds numbers, the viscous stress distributions near the surface are affected by the development of a separated laminar shear layer or the Reynolds shear stress. In addition, we show that the same analyses can be applied to the flows around a NACA0012 airfoil. (C) 2015 AIP Publishing LLC.
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2015 IEEE CONGRESS ON EVOLUTIONARY COMPUTATION (CEC) 3256-3262 2015年 査読有りOptimization of interplanetary space mission trajectories have been a long standing challenge. Here a novel approach is presented that considers several aspects of the space mission simultaneously as many-objective problem. Such problem is then solved by a decomposition approach in combination with a (massive) parallelization framework employing instances of Ant Colony Optimization algorithms. Numerical results show that the here presented approach has advantages over a classical weighted sum approach and is very suitable to efficiently exploit massive parallelization.
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2015 IEEE CONGRESS ON EVOLUTIONARY COMPUTATION (CEC) 965-969 2015年 査読有りWe propose a new algorithm FHV (Fast Hyper Volume) for exact hypervolume (HV) calculation using divide and conquer algorithm. FHV divides the original set of non-dominated solutions into several fractions first, calculate the value of HV of each fraction separately, sum up the each value, and finally obtain the value of HV of the original set. Therefore there are three very strong points: 1. Calculation cost is reduced significantly because each fraction contains less number of non-dominated solutions; 2. Complete parallelism is possible because each divided fraction is independent each other; 3. An arbitrary HV calculation method such as HOY (Hypervolume by Overmars and Yap) and WFG (Walking Fish Group) can be used together with FHV. Using these features computation time of HV becomes considerably reduced. For example, in a case number of objective functions is 5 and number of solutions is 1 x 10(6), it takes about 21 hours using WFG, however it takes only less than 30 minutes using FHV.
MISC
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宇宙航空研究開発機構特別資料: 第53回流体力学講演会/第39回航空宇宙数値シミュレーション技術シンポジウム論文集 = JAXA Special Publication: Proceedings of the 53rd Fluid Dynamics Conference / the 39th Aerospace Numerical Simulation Symposium JAXA-SP-21-008 149-155 2022年2月14日第53回流体力学講演会/第39回航空宇宙数値シミュレーション技術シンポジウム (2021年6月30日-7月2日. 日本航空宇宙学会 : 宇宙航空研究開発機構(JAXA)オンライン会議) The 53rd Fluid Dynamics Conference / the 39th Aerospace Numerical Simulation Symposium (June 30 - July 2, 2021. The Japan Society for Aeronautical and Space Sciences : Japan Aerospace Exploration Agency (JAXA), Online meeting) In this paper, aerodynamic characteristics around the blade of the hexacopter ''HAMILTON (HexAcopter for Martian pIt crater exploraTiON)'' for Mars exploration are investigated to obtain design knowledge regarding multicopter drone flying in Martian atmosphere. Reynolds-averaged Navier-Stokes simulation with the moving overlapped grid was employed for aerodynamic evaluation of two cases; one is hexa-rotor case and the other is single rotor case in order to compare single rotor case and hexa-rotor case and reveal unique characteristics of multirotor case. According to computational results, in both cases, hexa-rotor and single rotor, the maximum figure of merit could be observed at higher hovering thrust conditions. It is suggested that the baseline blade geometry could generate thrust efficiently at higher thrust conditions. The flow structure around the hexa-rotor can be classified into three groups; turn-in side where the flow was drawn the inside by blades rotation, turn-out side where the flow was put out to the outside by blades rotation, and the center side which was located between the turn-in and turn-out sides. The rotors of the center side took the low figure of merit compared with the other rotors because of aerodynamic interference from the turn-in side and the turn-out side rotors. Therefore, the total figure of merit of all rotors increased when the distance among rotors is increased. 形態: カラー図版あり Physical characteristics: Original contains color illustrations 資料番号: AA2130027012 レポート番号: JAXA-SP-21-008
書籍等出版物
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Springer Verlag 2009年 RefereedConstraint-handling techniques for evolutionary multiobjective aerodynamic and multidisciplinary designs are focused. Because number of evaluations is strictly limited in aerodynamic or multidisciplinary design optimization due to expensive computational fluid dynamics (CFD) simulations for aerodynamic evaluations, very efficient and robust constraint-handling technique is required for aerodynamic and multidisciplinary design optimizations. First, in Section 2, features of aerodynamic design optimization problems are discussed. Then, in Section 3 constraint-handling techniques used for aerodynamic and multidisciplinary designs are overviewed. Then, an efficient constraint-handling technique suitable to aerodynamic and multidisciplinary designs is introduced with real-world aerodynamic and multidisciplinary applications. Finally, an efficient geometry-constraint-handling technique commonly used for aerodynamic design optimizations is presented. © 2009 Springer-Verlag Berlin Heidelberg.
講演・口頭発表等
352共同研究・競争的資金等の研究課題
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日本学術振興会 科学研究費補助金 2014年4月 - 2017年3月
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日本学術会議 科学研究費補助金 2012年4月 - 2015年3月
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日本学術振興会 科学研究費補助金 2011年4月 - 2013年3月
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日本学術会議 科学研究費補助金 2008年4月 - 2011年3月
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日本学術振興会 科学研究費補助金 2008年4月 - 2010年3月
