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A kriging-based genetic algorithm called efficient global optimization (EGO) was employed to optimize the parameters for the operating conditions of plasma actuators. The aerodynamic performance was evaluated by wind tunnel testing to overcome the disadvantages of time-consuming numerical simulations. The proposed system was used on two design problems to design the power supply for a plasma actuator. The first case was the drag minimization problem around a semicircular cylinder. In this case, the inhibitory effect of flow separation was also observed. The second case was the lift maximization problem around a circular cylinder. This case was similar to the aerofoil design, because the circular cylinder has potential to work as an aerofoil owing to the control of the flow circulation by the plasma actuators with four design parameters. In this case, applicability to the multi-variant design problem was also investigated. Based on these results, optimum designs and global design information were obtained ...

The purpose of the research is to develop the method of designing and manufacturing a sensitivity-adjustable three component force balance utilized in a wind tunnel experiment. This paper describes the characteristics of the force balance and its application to wind tunnel testing. In this study, an oval-shape force balance is newly introduced and the characteristics are investigated by FEM analysis and force balance calibration. Also, the validity of the force balance is investigated in a wind tunnel experiment. As a result, the sensitivity of the oval-shape force balance can be adjusted by changing the aspect ratio and aerodynamic forces can be measured in a wind tunnel experiment using it. In conclusion, we can develop the sensitivity-adjustable force balance by applying oval shape and the force balance can be used in a wind tunnel experiment. It is recommended that the strain gauge position and its paste procedure be still improved.

The advantage in the thrust performance of the trielectode (TED) plasma actuator with serrated electrode edge was investigated. The trielectrode actuators have an additional exposed electrode, to which high DC voltage is applied, on the opposite edge of the buried electrode of the exposed AC electrode of SDBD plasma actuator. To prevent the uneven distributions of the discharge, the intentional dent that is evenly spaced on the exposed electrode was introduced. In this research, the induced body force of the TED plasma actuator was investigated on quiescent air by measuring thrust of the induced jet. A timeresolved PIV was used to acquire a velocity field in the quiescent air to measure the induced jet structure. To clarify the performance improvement mechanism, discharge plasma and flow field were numerically simulated. From the results of the thrust measurement, it is confirmed that the serrated TED plasma actuator enhances the thrust compared to the straight-edge TED plasma actuator especially at high DC voltage regimes. The absolute thrust attained 82.7mN/m, which was about 2.1 times larger than the straight-edge TED actuator and 17.5 times larger than the standard SDBD plasma actuator at the same conditions. The power efficiency became twice of the straight-edge TED plasma actuator and more than six times better than the SDBD plasma actuator at the same time. The PIV analysis clearly show that the serrated TED plasma actuator generate a uniform pattern of facing jet from both of the electrode for the positive DC voltage case, because of the uniform generation of the plasma from the sawtooth vertexes. For the negative DC voltage case, the uniformity of the induced flow was not improved.

To reduce aerodynamic drag, reduction of skin friction is required. Vibrant flexible wall has been proposed to reduce skin friction drag for a long time. However, there are many problems for the flexible wall to get practical durability, easy maintenance and reasonable cost. On the other hand, vibrant solid wall, which is much adequate for a car body, has been rarely proposed. In this research, mechanism of drag reduction by a vibrating solid wall is considered from the velocity distribution near the wall surface. Reduction of skin friction drag on a vibrant wall is mainly caused by decrease of Reynolds shear stress, and the effect of shear stress from velocity gradient near the wall is small. The optimum condition of vibration for drag reduction is also discussed in this paper.

This study focuses on the aerodynamic performance of the supersonic biplane at the low-speed region. The performance was evaluated and discussed through Computational Fluid Dynamics (CFD) and Experimental Fluid Dynamics (EFD). The result of the CFD simulation was compared with the experimental result to validate the simulation and confirmed to be reliable. Therefore, the CFD results were employed to derive the aerodynamic performance coupled with the theoretical equations. In the wind tunnel experiment, the three-component force measurement was conducted to obtain lift, drag and pitching moment coefficients. The wake survey was conducted to measure the drag in detail. The results proved the low-speed aerodynamic performance of the supersonic biplane can be described by the classical ``general biplane theory'' reasonably well.

One of the most critical technical issues with regard to supersonic commercial transportation is the sonic boom that occurs during supersonic cruising flight, which causes impulsive noise on the ground. The “supersonic biplane theory” has been proposed to reduce the sonic boom. Shock wave interaction and cancellation between the wings of a supersonic biplane can be realized at a specific design Mach number, but does not work at off-design values. Here, the low-speed aerodynamic performance, as off-design performance, of a baseline supersonic biplane was investigated and discussed using experimental and computational fluid dynamics approaches. The thin airfoil stall characteristics of a supersonic biplane were shown to be caused by the stall of both upper and lower wings at an angle of attack of 20°. Although there was leading flow separation of the upper wing at lower angles of attack, the stall of the lower wing was suppressed by interference with the upper wing. The lift of the lower wing was almost dominant to produce the lift of the supersonic biplane in the low-speed range. However, the lower wing caused greater drag than the upper wing at higher angles of attack.

The aerodynamic forces and moments of a flexible delta wing in pitching motion were experimentally studied in a low-speed wind tunnel. Three types of flexible delta wing were investigated, the flexible parts of which were 44, 70 and 99% of the delta wing. Aerodynamic characteristics were different among the three types of flexible and completely hard delta wing, and it was found that the winding-up of the leading edge of the delta wing is the key factor for determining the leading edge vortex on the upper side of the wing and the pressure distribution on the windward side. Lift, drag, and pitching moment formed a hysteresis loop with an angle of attack in pitching motion, particularly in a region with a large attack angle, accompanied by leading edge vortex breakdown. The flow visualization of leading edge vortices was also carried out to explain the dynamic characteristics of the delta wings.

Dynamic properties of aerodynamic forces and moments acting on a delta wing were experimentally studied in a low-speed wind tunnel. The wing could make a smooth and periodic arbitrary motion, such as pitching, yawing, rolling, translation motions, and their combinations by a manipulator equipped with stepping motors controlled by a microcomputer, which had six degrees of freedom. The objective of the research is to analyze unsteady aerodynamics of the rolling delta wing with regard to the altitude from the ground. It was found that the hysteresis loops of lift, drag and rolling moment were drawn with the roll angles and amplified owing to the ground effect. The factors to determine the wing properties in the large angle of attack are the pressure on the windward side of the delta wing and the leading edge vortices which were broken down and reconstructed according to the wing motion. The pressure on the lower wing surface was also influenced by the clearance between the delta wing and the ground plate, which deduces the ventulli or stagnant effects.

This study was conducted for an aerospike nozzle and to analyze the phenomena of the jet impingement, which was spout out from a circular nozzle with four or eight holes to a flat plate. The impingment was undertaken to be perpendicular and oblique to the plate. If the distance between the flat plate and the nozzle becomes close, an usual PSP analysis is difficult because the source of light, a camera and the nozzle can not be set up for the PSP measurements. Therefore, the pressure on the plate was observed from the back side of the a flat plate. Furthermore, the visualization of the impingement jet flow was studied by the oil film method.

The flow control performance of the pulse-actuated plasma actuator and their mechanism was investigated. To clarify the flow mechanism of the separation control on a circular cylinder when pulsed plasma actuator is applied, the characteristics of the induced jet and their influence on the separated flow were separately examined. The characteristics of the induced jet from the pulsed plasma actuator were examined from the results of jet thrust measurement and high-speed Schlieren flow visualization. The effect of the frequency of pulsed plasma actuator on the flow separation on a circular cylinder model is examined in the viewpoint if aerodynamic force and the flow field measured with time-resolved PIV at the low speed wind tunnel test. For the induced jet analysis, the strength of the induced jet is found to be widely varied with the modulation frequency. When the modulation frequency is high, over 250Hz, the thrust of the induced jet increases with the modulation frequency increased. For the separation control test, the drag is significantly dropped with increasing the modulation frequency. From these results, it turns out that the enhancement of the induced jet generated from plasma actuator in high modulation frequency regime is the main cause of the enhancement of drag reduction performance.

流動層の底部に設けたガス流れを伴う鉛直円筒移動層 (スタンドパイプと呼ぶ) について, 粉体とガスの連続の式, 壁面の摩擦力を考慮した運動量平衡の式および相対速度によるガスの圧力損失の式と, さらに, 軸方向のガス圧力分布に関する実験データから, スタンドパイプ軸方向の粉体圧σzと空隙率eの関係式を求めた.この結果によれば (1) |dσz/de|はスタンドパイプ入口の空隙率が大きい部分で大きく, 中間で一度低下し, さらに下方では再び増大する.すなわち, 中間部で粉体圧の変曲点が存在する. (2) スタンドパイプ入口部近傍で|dσz/de|が大きくなるのは, 主としてスタンドパイプ壁面の摩擦力の影響による可能性が高い. (3) 粉体圧は空隙率のみならず粉体粒子径の関数でもある.また, この粉体圧, 空隙率, 粒子径の関係式と他の理論式から, スタンドパイプを流れるガス量, および軸方向のガス圧力, 粉体圧, 空隙率などの分布が求められることを示した.

A Kriging-based genetic algorithm (GA) called efficient global optimization (EGO) was employed to optimize the parameters for the operating conditions of a plasma actuator (PA). The aerodynamic performance was evaluated by wind tunnel testing to overcome the disadvantages of time-consuming numerical simulations. The developed optimization system explores the optimum waveform of parameters for AC voltage by changing the waveform automatically. The proposed system was used on the drag minimization problem around a semicircular cylinder to design the power supply for a PA. Based on the results, the optimum design and global design information were obtained while drastically reducing the number of experiments required compared to a full factorial experiment.

The objectives of the present study are to obtain unsteady pressure distributions on a freely oscillating body and to examine its relation with the body motion. In the present experiment, a low speed wind tunnel was used, where the exit was directed in a vertical direction. The shape of the model is similar to a Chinese pan with a diameter of 480mm. This experiment was performed at two flow velocities: 10m/s and 20m/s, and the model was suspended from a bar above the wind tunnel exit. The model was freely oscillated in the wind tunnel flow under the gravity and aerodynamic forces. The amplitude of this oscillation was damped due to the aerodynamic forces. The pressure distributions on the body were successfully measured, which clarified the damping mechanism of the body motion. To compare with the experimental result, the unsteady numerical calculations were also performed with the boundary element method.

The purpose of the research is to characterize the ionization process behind a shock wave with precursor photoionization in argon. In this study, the H-β line is observed by spectroscopic measurements using a hypersonic shock tube and the Stark broadening of the H- β line is evaluated to obtain the electron density behind a shock wave. As a result, the electron density is on the order of 1021 m-3 and tends to decrease with increasing the distance from the shock front. The decrease of the measured electron density is caused by the radiative energy loss from the test gas because the radiative energy loss corresponds to a decrease of the temperature or the density of the test gas, both of which can decrease the electron density. Previous study showed that photoionization occurred ahead of the shock wave, making use of the radiation energy emitted from the region behind the shock wave. The fact might be related to the radiative energy loss obtained in the present study. From the present study, it is found that the radiative energy loss is dominant behind the shock wave under the condition that precursor photoionization occurs ahead of the shock wave. In future, we should investigate the radiative transfer phenomena around the shock wave for further clarification.

The flow control performance and jet vectoring characteristics of trielectrode (TED) plasma actuators on quiescent air were analyzed. To examine the body force generated from the actuator, the thrust of the induced jet was measured, as a reaction of the aerodynamic force from the actuator. The corresponding induced jet was visualized by means of highspeed Schrielen Visualization with the discharge conditions to understand physical phenomena on the actuator. The experiments show that the thrust generated by the TED actuator depended on the applied DC voltage. By applying high negative DC voltage, with the existence of sliding discharge channels between the two exposed electrodes, the induced flow becomes in the pronounced vertical jet. The overall thrust is greatly improved from SDBD actuator at these regimes. The efficiency of SDBD actuator also shows substantial improvement. These confirm that the TED actuator could improve both body force and their power efficiency at the same time.

Characteristics of plasma flow in a 10.5kW arc-heated wind tunnel are investigated by spectroscopic measurement. The test gases are 100 % CO2 and 90 % CO2 - 10% N2 mixture which simulates the Martian atmosphere. In CO2 plasma, radiation of C2 Swan bands system is predominant and strong radiation of atomic oxygen is seen in the near-infrared light region. In CO2 - N2 plasma, radiation of CN violet band system is much stronger than those of C2 Swan band system and atomic oxygen. By using the multipoint spectroscopic system, spatial profiles of spectra are obtained with high accuracy. With approaching the stagnation point, radiation intensity increases caused by the formation of shock wave around the hemisphere model. Thickness of shock layer is estimated to be 3mm from the stagnation point. The rotational and vibrational temperatures of C2 and CN are estimated by the spectrum fitting method using the SPRADIAN 2 and spatial distribution of temperatures are obtained. For C2, the rotational temperature is higher than vibrational temperature ahead of shock wave. These temperatures are almost same in the shock layer. For CN, the rotational temperature is lower than vibrational temperature, showing the opposite tendency observed in C2. However, in the shock layer, rotational and vibrational temperatures are almost same.

The flow control performances of the trielectrode plasma actuators are examined. By the balance measurement the TED-DBD (positive DC voltage was applied to the additional electrode) actuator enhances the thrust as the DC voltage increase, and the maximum thrust for the TED-DBD reaches 270% of the one for the SDBD actuator in the same conditions. The TED-SD (negative DC voltage to the additional electrode) actuator generates strong negative thrust up to 400% of SDBD case. The TED-SD actuator generated recirculation bubble on the electrode. The recirculation area becomes expanded and diverts their direction upward and reversed as the DC voltage increases. For the delta wing vortex control, the TED-SD plasma actuator was found to be feasible to control the flows that include massive three dimensional separations. The change in the direction of the induced jet generates peculiar control performance of the flow. Nomenclature fac = frequency of the AC voltage T = Thrust generated by plasma actuator Vax = AC voltage applied to the DBD electrode of plasma actuator, peak-to-peak Vdc = DC voltage applied to the Third (DC) electrode of plasma actuator

The aerodynamic characteristics and the effects of tip plates of a tapered supersonic biplane wing during the starting process have been investigated through Experimental and Computational Fluid Dynamics (EFD/CFD). Three types of the wing model were used: without tip plate (type-N); with the tip plate which covers only the aft-half of the wing tip (type-A); with the tip plate which covers the entire wing tip (type-B). Experiment was conducted in the supersonic blowdown wind tunnel with 600 mm × 600 mm cross section located at the High-speed Wind Tunnel Facility of Institute of Space and Astronautical Science (ISAS/JAXA). The flow conditions covered from M∞ = 1.5 to 1.9 with increments of 0.1. Pressure-Sensitive Paint was applied to measure pressure distributions on the surface of the wing. CFD simulations were conducted to compare with experiments and to investigate effects of the Mach numbers in detail. The tapered biplane wing without the tip plate was found to start between M∞ = 1.8 and 1.9. The difference of the starting Mach numbers between type-N and type-A was small. On the other hand, the starting Mach number of type-B was about 0.05 higher than that of type-N.Copyright © 2011 by JSME