大山 聖

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

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved. A Monte-Carlo simulation is performed to evaluate the success probability of the flight test of the Mars airplane at high altitude atmosphere on Earth. The number of the uncertainty parameters is 136. The simulation result shows that the success probability is 95%. It is clarified that the most severe criterion is a maximum Mach number.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. We propose to use a helicopter for exploration of pit craters and caves on Mars. Helicopter does not contaminate inside of pit craters and caves because it does not emit gas including carbon compounds. In addition, a helicopter can explore caves even if the bottom of the caves is covered by rocks. We propose to explore the pit craters found near Elysium Mons because they are suitable for the aerial exploration because (1) there is higher possibility of existence of life and (2) air density is relatively high. We consider two specific missions using a helicopter; one is a mission which explores only a pit crater and the other is a mission which explores both a pit crater and a cave. The present analysis shows both missions are feasible.

This paper investigates the robustness of the aerial deployment behavior of the foldable-wing airplane for Mars exploration especially focused on the effect of the hinge axis tilting in a yaw direction. This study deals with four dispersive parameters for the robustness evaluation: drop velocity, surrounding gust velocity, initial pitch angle, and height. The robustness of several tilted and non-tilted hinge axis designs are calculated and then compared. The result clearly shows that the tilted hinge axis design can deploy with lower torque than the torque of the non-tilted hinge axis design. The increase of sideslip angle due to the hinge axis tilting suppressed an aerodynamic force on the deploying wing.

Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. Mars is the next milestone in our exploration of the solar system. The presence of an atmosphere on Mars signifies that an airplane could travel in its atmosphere using the aerodynamic forces of flight. The airplane allows for a platform that can cover a larger area of exploration than is currently available. A reconnaissance airplane offers the possibility to obtain high-resolution data on a regional scale of several hundreds to thousands of kilometers, which cannot be achieved with rovers or satellites. There is an extremely high demand for the exploration of Mars using an airplane that can fly in its atmosphere. One of the big problems for a Mars Airplane is the very low atmospheric density on Mars. So, it is difficult to obtain the required lift, as the wing area required to generate enough lift is inversely proportional to the density. So in order to reduce the required lift, thorough weight reduction is needed. Even so, a Mars Airplane needs a large wing area, which leads to another problem. To transport to Mars, a Mars Airplane must be small and compact. As a way to solve this conflicting problem, the Mars Airplane needs some deployment mechanisms. Various hurdles, including those described above, must be overcome in order to realize the flight exploration of Mars and all of them require innovative technological solutions. Hence, the Mars Airplane Working Group was established in 2010 with the aim of conducting flight technology validations for the MELOS1 mission using a compact airplane in JAXA/ISAS. The working group aims to realize Mars exploration using an airplane for the first time ever. At present, the mission being considered for the Mars exploration plane is "flying over a range of about 100 km to capture ground surface images and observing high-resolution images of residual magnetic fields." Cameras and magnetic field observation equipment are mounted as payloads on the airplane, which is expected to fly over a range of 100 km at a speed of 60 m/s. In the conceptual design, the weight of the airplane is about 4.0 kg, the span length is about 2.5 m, and the total length is about 2.0 m. This paper provides a summary of the Mars airplane development being considered by the working group in Japan and discusses the technical issues addressed in order to realize a Mars airplane.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. A folding wing is an effective deployment mechanism for the airplane that is used for Mars exploration. A spring loaded hinge is considered as a deployment actuator for Mars airplane in Japan. A hinge torque is one of the primitive design variables to control the aerial deployment behavior. The required hinge torque for deployment is directly concerned with the deployment mechanism mass. Since the Mars airplane requires thorough mass reduction, it is necessary to reduce the required hinge torque while keeping high robustness for the aerial deployment. This paper investigates the robustness of the aerial deployment behavior especially focused on the effect of the hinge axis tilting. The hinge axis of the non- tilted hinge axis design is defined to be parallel to the X-axis of the center body coordinates. The conditions to judge whether the deployment succeeded or failed are defined for the state of the airplane. The margins of the airplane state for the conditions are set to the evaluation functions of the safety. The robustness of the safety is evaluated using the sigma level where the sigma level is a function of the average and standard deviation of the evaluation functions. For the robustness evaluation, this study deals with four dispersive parameters: drop velocity, surrounding gust velocity, initial pitch angle, and height. The robustness of several tilted and non-tilted hinge axis designs are calculated and then compared. The result clearly shows that the tilted hinge axis design can deploy with lower torque than the torque of the non-tilted hinge axis design. The motions of the individual cases are then studied to reveal the effect of the hinge axis tilting. It is clarified that the tilted hinge axis design is able to set the angle of attack of the outer wing positive under the wide range of conditions. Therefore, the aerodynamic force assists the deployment. In the appropriate condition, the wings deploy without torque of the deployment actuator.

In this study, the effects of the computational spanwise domain length on the flowfield with massive separation and on the flowfield with dynamic stall are investigated by large-eddy simulation. The objective airfoil is NACA0012 and the chord-based Reynolds number is of 2.56 × 105. The objective flowfields are that around a fixed angle of attack of 10 and 25 degrees, and that around a pitching airfoil between AoA of 5 degrees and 25 degrees. The spanwise length effect become significant after the stall, as observed as the attenuation of the large vortices. Observations of the flowfield clarified that the undulation of two large vortices from the leading edge and the trailing edge is one of the mechanisms for the spanwise length effects. The qualitative analysis for this mechanism is performed to address the sufficient spanwise length, and the spanwise length have to be at least 1.0c for the flowfield with large vortex structures so as to resolve its spanwise distribution.

In this study, AC DBD plasma actuator is applied to control the flow around NACA0015 and Ishii airfoils in a low Reynolds number condition (Re = 63,000). Here, the Ishii airfoil is a high performance airfoil at the low Reynolds number condition. The DBD plasma actuator is located at x/c = 5% and is actuated in burst mode with the nondimensional burst frequency F+ from 0.1 to 20. Maximum control authority is achieved with Vpp = 6kV and F+higher than 6 for both airfoils. Results show that different effect of separation control between NACA0015 airfoil and Ishii airfoil.

The convective Mach number and density ratio dependences of sound sources and flow structures in a compressible mixing layer are investigated by direct numerical simulations. Characteristics of sound sources are analyzed using the source terms of Lighthill equation. As the Mach number increases sound source strength decreases, because vortex motion is weakened by compressibility. For density ratio dependence, the emission angle of Mach waves becomes shallower and vortices show sparse structures as density ratio increases. In addition, larger vortex structures appear at lower density side for higher density case.

Large-eddy simulations of the separated flow over an NACA0015 airfoil controlled by the DBD plasma actuator are conducted and the flow fields and the aerodynamic performances are compared with the Ishii airfoil, one of the high performance airfoil at the low Reynolds number. The DBD plasma actuator is set at the 5% chord length from the leading edge of NACA0015 airfoil and operated in burst mode at the Reynolds number Re=63,000. In both cruise and post stall angle of attack, Ishii airfoil show higher aerodynamic performance than NACA0015 airfoil when DBD plasma actuator is OFF. However, when the DBD plasma actuator is activated, NACA0015 show higher aerodynamic performance.

The main aim of this paper is to elucidate the mechanism of massive separation control by using a dielectric barrier discharge plasma actuator (DBDPA). A technique of design exploration is applied to find good operating-parameter combinations for the DBDPA. We consider a NACA 0015 airfoil with 16° angle of attack and Reynolds number Re = 63000. The flow without the control is massively separated, however we can suppress the separation using the DBDPA with the relevant operating parameters. Using good parameter combinations obtained by design exploration technique, the nature of the flow around the airfoil with and without control is explored in detail.

DESTINY is injected to long elliptical orbit by Epsilon rocket launcher. If the apogee altitude of the injected orbit is high enough, it is achieved to ease the requirements for design and operation of the spacecraft. This paper investigates the ability of trajectory injection by means of 4-stage Epsilon rocket using the method of multi objective optimization under several flight constraints.

"DESTINY" is an acronym of "Demonstration and Experiment of Space Technology for INterplanetary voYage", which is proposed by JAXA/ISAS as "ISAS Small Scientific Satellite" mission. In this mission, trajectory design is one an important technical element because of its many revolution low-thrust orbits with many mission objectives and constraints. Evolutionary computation is utilized to find candidates for the orbit.

本記事に「抄録」はありません。

A numerical study of the effects of airfoil shape on low Reynolds number aerodynamics is presented. The large-eddy simulations are performed with 6 ^th-order compact finite difference scheme and 10^th-order low pass filter, and 2^nd-order backward implicit time integration with inner iterations. Systematic numerical excesses show the feasibility of the current simulations to predict flow fields around fixed-wing configurations involving a laminar separation and laminar-to-turbulence transition at low Reynolds number. At the Reynolds number of 2.3×10⁴, two types of thin and asymmetric airfoils as a target airfoil shape of micro-size air vehicle are considered. The results show that the airfoil cross section affects the formation of a laminar separation bubble and the transition to turbulence in the three-dimensional flow around the wings at low angle of attack and hence significant influence on the aerodynamic performance. © Civil-Comp Press, 2012.

Aerodynamic characteristics of reusable observation vehicle are computationally investigated under subsonic and supersonic flows using the RANS (Reynolds-averaged Navier-Stokes) simulations. The initial investigation for the concept design is done with the light optimization using the light CFD. The results show that the simulations using coarse grid estimate the axial force coefficient and the lift to drag ratio accurately except some cases. The results indicate the correlation between the supersonic lift to drag ratio and the axial force coefficient. The results show the correlation between the y-coordinate of the design variable and the volume. The required knowledge for the concept design in the near future is obtained.

PARSEC airfoil parameters often used for transonic airfoil design are re-examined by data-mining Pareto-optimal airfoil designs. The Pareto-optimal airfoils are obtained by using a multiobjective evolutionary algorithm. For data mining, scatter plot matrix coupled with correlation coefficient is used. The present result shows that the PARSEC airfoil parameters may not be the best choice for transonic airfoil design. The result also indicates that data mining from Pareto-optimal airfoils may give more information than data mining from all feasible airfoils.

The performance of the surface air-flow induced by non-thermal plasma is studied experimentally. The non-thermal plasma is generated by atmospheric dielectric-barrier discharge. The input discharge power was 1.8W. At first, flow induced by the discharge on a flat plate is investigated. Velocity profile is measured by a hot-wire anemometer. The maximum value 1.1 m/sec was observed on the plate surface. Secondary, separation control for wing surface flow is investigated using a 9cm chord NACA0015 in a wind tunnel at 20m/s of air stream velocity (Re〜1.2x10^5). Barrier discharge electrode is set on the leading edge of the wing. Separation angle is increased by 3.5 degrees and the maximum of the lift coefficient is improved by 12%.

This paper describes recent activity in JAXA aiming reformation of design and development (D & D) process by introduction of information technology, simulation technology, reliability engineering, etc for rocket valve reliability improvement. In this activity, JAXA's information system for rocket valve D & D will be developed by the end of FY2007. This information system consists of detailed FMEA/FTA utilization support tool, QFD utilization support tool, and material database system and material database utilization support tool. This information system will help to improve efficiency and reliability of D & D process of JAXA's rocket valves and other JAXA's products.

In order to improve an efficiency of failure analysis process by using failure mode and effect analysis (FMEA) and failure tree analysis (FTA), detailed FMEA/FTA support tool is developed. By using this tool, failure mechanisms can be analyzed with FMEA approach as well as FTA approach. In order to improve coverage of the extracted failure modes, new failure analysis approach is proposed, which is based on the interface N2 chart. Interface N2 chart is N2 chart which is extended to express the interface of mechanical system. Interface N2 chart can also be used to visualize the failure analysis result to describe system structure based on the knowledge obtained in failure analysis such as causal relationship of fault propagation. New schematic visualization method is discussed to realize useful visualization tool which promotes comprehensive understanding of failure mechanism.

Flow fields of the supersonic jets impinging on an inclined flat plate at high plate-angles are experimentally investigated using surface pressure measurement with pressure sensitive paint and Schlieren flow visualization. While Type I flow type is dominant at high plate angles, the present research found a new flow type "TYPE I with bubble" at plate angle between 60 and 80 degrees. The flow classification according to L/L_s' and plate angle indicated that the constant x/L'_s curve doesn't represent the boundary of Type I and Type II anymore at high plate angles between 60 and 90 probably because Type II flows at low plate angles and high plate angles is different phenomena. This study also indicates that the curve dividing Type I and Type I with bubble regions is same as the curve dividing Type II and Type II with bubble regions.

An efficient and useful robust optimization approach, design for multi-objective six sigma (DFMOSS), has been developed. The DFMOSS couples the ideas of design for six sigma (DFSS) and multi-objective genetic algorithm (MOGA) to solve drawbacks of DFSS. DFMOSS obtains trade-off solutions between optimality and robustness in one optimization. In addition, it does not need careful parameter tuning. Robust optimizations of a test function and welded beam design problem demonstrated that DFMOSS is more effective and more useful than DFSS.