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The detailed mechanism of supersonic mixing enhancement by a novel device is investigated computationally based on the Navier-Stokes equations. The device is a wall-mounted three-dimensional cavity, several parts of the rear face of which are cut out, and the lower wall of the cut-out part has a slope directed toward the injector from which the injectant is discharged. The flow structures and behaviors observed in the computational results are confirmed in the experimental results. Experimental information about the high-frequency (several tens of kHz) pressure oscillation in the cut-out part is acquired in this study using a pressure-sensitive paint (PSP) responding to rapid change in pressure. The computational results reveal that a hairpin-like vortex is formed inside the cavity. The vortex works on the shear layer across the cavity to deflect it upward. The deflected shear layer shields the injectant from the primary flow, which causes the injectant to penetrate deep into the primary flow. The computational results also clarify that the expansion and compression waves produced in the rectangular part of the cavity are delivered effectively to the injector along the slope of the cut-out part, which is confirmed in the PSP measurement results. The delivered waves actuate the jet issuing from the injector to fluctuate up and down periodically with a large amplitude.

In this study, skin friction around a ½-scale Ahmed body was measured experimentally at a Reynolds number of Re = 2 × 105. The slant angle of the Ahmed body was 25 deg, and the yaw angles ranged from 0 deg to 8 deg. This study focused on the flow structure on the slant surface under different cross-wind conditions. A force balance system was applied to measure the aerodynamic drag of the model. The global skin-friction topology was measured by applying a luminescent oil layer with a subgrid data processing algorithm. The method used to measure the skin friction was conducted for the first time on the Ahmed body. The results indicated that the technique is highly capable of extracting the skin-friction topology. For a yaw angle below 3 deg, the flow on the slant surface was not significantly affected by the cross-wind condition, and the drag of the model was nearly constant. However, at yaw angles above 3 deg, the flow on the slant surface was highly affected by the roof longitudinal vortexes on the windward side, leading to a dramatic increase in the drag of the model. High consistency in the drag and skin-friction fields was observed. The detailed skin-friction structure at different yaw angles will be discussed in this study.

In this study, we apply the optical flow method to the time-series shadowgraph images of impinging jets using a high-speed video camera with high spatial and temporal resolution. This image analysis provides quantitative velocity vector fields in the object space without tracer particles. The analysis results clearly capture the details of the coherent vortex structure and its advection from the shear layer of the free jet. Although the results still leave challenges for the quantitative validation, the results show that this analysis method is effective for understanding the details of the physical phenomenon based on the quantitative values extracted from the shadowgraph images.

In this study, we experimentally investigate the effects of mainstream turbulence intensity (Ti) on a leading-edge separation bubble under low-Reynolds number (Rec) conditions range of 2.0 × 104 to 6.0 × 104. We used a flat plate to fix a separation point at the leading edge. Also, we visualized the behavior of the leading-edge separation bubble using the smoke wire technique and Particle Image Velocimetry (PIV) measurement. Furthermore, we measured the effect of Ti on the turbulent transition process in the separated shear layer using a hot-wire anemometer. The results indicate that the bypass transition for large Ti causes the turbulent transition, and so accelerates the reattachment of the separated shear layer. The results show that the bypass transition promotes the reattachment of the separated shear layer to maintain the leading-edge separation bubble on the upper surface even at high angles of attack, increasing the stall angle.

In the low Reynolds number range below Re SD and Ishii airfoils are known as high performance low Reynolds number airfoils with a relatively high lift to drag ratios Although the aerodynamic characteristics at particular Reynolds numbers have been thoroughly studied research is lacking on how Reynolds numbers affect the aerodynamic characteristics that become pronounced for general low Reynolds number airfoils This study investigates the Reynolds number dependence on the aerodynamic performance of these two airfoils The results demonstrate that the Ishii airfoil is particularly less dependent on the Reynolds number of the lift curve compared to the SD airfoil Although there is a slight difference in the high angles of attack above the maximum lift coefficient the lift slope hardly changes even when the Reynolds number varies

The flow around the wing of exploration aircraft flying in the very dilute Martian atmosphere experiences very specific low-Reynolds-number compressible flows. A fundamental understanding of compressibility effects, especially on a separation bubble, is an important issue in the aerodynamic wing design. We investigated the effect of compressibility on the laminar separation bubbles and aerodynamic properties of a blunt flat plate with a thickness of 5% of the chord length in unsteady two-dimensional laminar and large-eddy simulations at chord-based Reynolds numbers ( Re) of 6100 and 11 000 in the Mach number ( M) range from M = 0.1 to 0.8. In the incompressible flow below M = 0.2 at Re = 11 000, the separated shear layer rolled up due to the Kelvin–Helmholtz (KH) instability reattaches to the surface with turbulent transition and then forms a laminar separation bubble. In contrast, with increasing Mach number, the transition and reattachment are delayed, and the separation bubble is stretched. In particular, at M = 0.8, the rolled-up vortex is not distorted in the span direction and advects downstream with a strong spanwise two-dimensionality. This is due to the compressibility effect that suppresses the KH instability. On the other hand, at Re = 6100, the separated shear layer is stable without spanwise distortion even in the incompressible flow due to the stronger influence of viscous forces. With increasing Mach number, such stable separated shear layer reattaches without rolling up further downstream, and the Strouhal number of the vortex also gradually decreases. This leads that the large lift oscillation owing to periodic and two-dimensional spanwise vortex shedding is significantly suppressed with an increasing Mach number. This is also due to the weakening of the KH instability due to the compressibility effects. However, at Re = 11 000, the reattachment points move downstream more significantly with the increasing Mach number, indicating that the suppression of the KH instability due to the compressibility effect is more pronounced at Re = 11 000 than at Re = 6100. Consequently, the negative skin friction shear stress owing to the reverse flow inside the expanded separation bubble becomes relatively large, decreasing the total drag with the increasing Mach number.

This study experimentally investigates aerodynamic characteristics and flow fields of a smooth owl-like airfoil without serrations and velvet structures. This biologically inspired airfoil design is intended to serve as the main-wing for low-Reynolds-number aircrafts such as micro air vehicles. Reynolds number dependency on aerodynamics is also evaluated at low Reynolds numbers. The results of the study show that the owl-like airfoil has high lift performance with a nonlinear lift increase due to the presence of a separation bubble on the suction side. A distinctive flow feature of the owl airfoil is a separation bubble on the pressure side at low angles of attack. The separation bubble switches location from the pressure side to the suction side as the angle of attack increases and is continuously present on the surface within a wide range of angles of attack. The Reynolds number dependency on the lift curves is insignificant, although differences in the drag curves are especially pronounced at high angles of attack. Eventually, we obtain the geometric feature of the owl-like airfoil to increase aerodynamic performance at low Reynolds numbers.

Accepted: 2018-04-11, 資料番号: SA1180225000

A flow resonance accompanied by the emission of acoustic tones occurs in an over-expanded convergent-divergent (C-D) nozzle when operated at comparatively low pressure ratios. This phenomenon is distinguished from conventional screech tones and is referred to as “transonic tones”. In contrast to screech tones, the peak resonant frequency for transonic tones increases with pressure ratio; the peak sound pressure level exceeds 110 dB. In this study, we investigated the basic characteristics of transonic resonance and tones using a circular C-D nozzle in an anechoic room. The effects of the boundary layer trip were also evaluated using a tripping wire for the suppressing transonic resonance and tones. The results of acoustic measurements show that several predominant peaks correspond to transonic tones. However, the boundary layer trip inside the nozzle effectively eliminated these tones and suppressed the unsteadiness of the flow inside the nozzle.

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.

Flow over a flat plate with a 5% thickness ratio is investigated by schlieren visualization in compressible low-Reynolds-number conditions. The results show that flow separates at the leading-edge and laminar separation-bubble forms. The position of the maximum root mean square of the schlieren image which is related to the position of the vortex shedding moves downstream as a Mach number increases. Furthermore, the two-dimensional structure of generated vortices is maintained up to the trailing edge at the Mach number of 0.66. The frequency analysis of the time-series intensity value of the schlieren images also shows that the flow is stabilized with increasing the Mach number. The position of the end of the pressure plateau region matches the position where the root-mean-square value of the intensity image becomes a maximum due to vortex shedding.

Accepted: 2014-09-21, 資料番号: SA1150051000

This study proposes separation control investigation using a Dielectric Barrier Discharge (DBD) plasma actuator on a NACA0015 airfoil over a wide range of Reynolds numbers. The airfoil was a two dimensional NACA0015 wing model with chord length of 200mm. Reynolds numbers based on the chord length were ranged from 252,000 to 1,008,000. A plasma actuator was installed at the leading edge and driven with AC voltage. Burst mode (duty cycle) actuations, in which nondimensional burst frequency F+ was ranged in 0.1–30, were applied. Time-averaged pressure measurements were conducted with angles of attack from 14deg to 22deg. The results show that initial flow fields without an actuation can be classified into three types; 1) leading edge separation, 2) trailing edge separation, and 3) hysteresis condition between 1) and 2), and the effect of burst actuation is different for each above initial condition.

The paper reports a result of low temperature hypervelocity impact (HVI) tests of aluminum sphere against a 16-ply quasi-isotropic Carbon Fiber Reinforced Plastic (CFRP) laminate plate at speed ranging from 1.4 to 5.4 km/s in air at 10 Pa. The result was compared with room temperature impacts. At low speed impact on CFRP plates, fracture patterns of specimens varied depending on their temperatures, whereas at high-speed impact, any significant differences in the fracture patterns around penetration holes and independent of the temperatures.

This paper reports the results of hypervelocity impact (HVI) experiments on cryogenic aluminum alloys at 122 K. Target plates were impinged by room temperature aluminum spheres at speeds ranging from 1 to 3.7 km/s in air at 10 Pa. The results were compared with HVI results at room temperature. We visualized the HVI processes by shadowgraph and recorded them with a digital high-speed video camera. We showed temperature effects of target plate physical properties on the formation of debris clouds and impact craters and hence the shielding efficiency of aluminum alloys at low temperature.

Low-speed wind tunnel tests are carried out to investigate the control surface effectiveness at low Reynolds numbers (Re = 20,000-80,000). A thick airfoil, NACA0012, has a nonlinearity of the control surface effectiveness, which are associated with flow separation both on the upper and lower surfaces and a formation of a laminar separation bubble. Thin airfoils such as NACA0006 and a 3%c flat plate have a much smaller nonlinearity of the control surface effectiveness. These differences in the control surface effectiveness are due to the nonlinearity of the lift curve for each airfoil, and the flow separation behavior has a profound effect on the control surface effectiveness.

The aerodynamic characteristics of thin airfoils at low Reynolds number were investigated in a Mars Wind Tunnel at Tohoku University. The lift and the drag of five airfoils (NACA 0012, 5% flat plate, 1% flat plate, 3% circular arc, 6% circular arc) were investigated by using a two-component balance system at Re=1.1×10 and Mach number about 0.2. Also, Pressure-Sensitive Paint technique was used to measure pressure distributions around the airfoils. The results obtained by the balance system show that The maximum lift to drag ratio of the 3% circular arc is the largest among the airfoils tested in this study. And the results of PSP measurement show that the lift slope becomes large where a bubble with relatively short length appears on the airfoil surface, and the thicker airfoil has longer separation bubble than the thin airfoil at the same angle of attack. Furthermore, the airfoil surface having large curvature, the range of angles of attack where the separation bubble exists becomes narrow because the appearance of a separation bubble by an increase of angle of attack is delayed.

Strong acoustic waves emitted from rocket plume might damage to rocket payloads because rocket payloads consist of fragile structure. Therefore, it is important to predict acoustic directivity and reduce its intensity level. In this study, we conduct experiments of supersonic jet flows and investigate an influence of the nozzle geometry on acoustic waves by means of Schlieren method and microphone measurement. Three different nozzles are examined: a conical nozzle, a convergent-divergent nozzle (referred as C-D nozzle), and a tab-C-D nozzle. Tabs are equipped in the nozzle inside and turbulence is generated in the tab-C-D nozzle case. The Schlieren visualization shows that the strong shock trains are observed in the potential core of the jet for the conical nozzle case, while the shock waves are relatively weak since the nozzles are in the nearly ideal expanded condition in the C-D nozzle and tab-C-D cases. The distribution of near field OASPL (over all sound pressure level) obtained by microphone measure...

Pressure-Sensitive Paint (PSP) technique has been applied to measure pressure distribution on airfoil models at low Reynolds number (Re = 4.1x10 4.3x10) in the Mars Wind Tunnel (MWT) at Tohoku University. In this study, the capability of PSP for pressure measurement was evaluated with a particular emphasis on its applications to thin airfoil models. The obtained results show that PSP can clearly visualize pressure distributions on the airfoils, allowing us to analyze a behavior of the separated shear layer on the airfoil including laminar-to-turbulent transition and reattachment. A comparison between the PSP and balance measurements indicates that the evolution of a laminar separation bubble over the airfoil surface has a strong effect on the lift and drag characteristics. It has been also verified by this experiment that, using the proposed method for temperature correction, pressure distribution can be obtained quantitatively even on thin airfoils.

In this study, we have investigated the sensitivity and the time response characteristics of thermal anemometers such as a hot wire in low-pressure environment. To achieve this purpose, we developed the calibration device to calibrate a thermal anemometer under lowpressure condition. By using this device, we investigated the relationship between Reynolds number and Nusselt number at 1 kPa to 80 kPa and Mach number 0.1 to 0.95, for a standard hot-wire having a diameter of 5 µm. We found that the sensitivity of the hot-wire anemometer reduces with decreasing pressure and the time response characteristics get drastically worsen. However, the results of calibration with higher overheat ratios show that the sensitivity and the time response can be improved by increasing wire temperature, particularly in low-pressure condition. These results would provide guidelines for measuring turbulent intensity in a low-density wind tunnel such as a Mars wind tunnel to simulate the flight environment of a Mars airplane.

The aerodynamic characteristics of a 5% flat plate and NACA0012-34 airfoil in low Reynolds number (Re=0.43x10 4 ~4.1x10 4 ) and high subsonic flow (M=0.1~0.6) were investigated in the Mars Wind Tunnel (MWT) at Tohoku University. A two-component balance system and Pressure-Sensitive Paint (PSP) technique have been developed to measure the lift and drag forces and pressure profiles on the model. For the flat plate, Mach number effect does not have much effect on its aerodynamic performance while Reynolds number affects the lift slope and the drag characteristics. On the contrary, for NACA001234 airfoil, both Reynolds and Mach number effects become more prominent. The lift curves are highly nonlinear and the drag polars are affected by behaviors of a laminar separation bubble in trans-critical condition. A comparison of the results obtained at different Mach numbers has suggested that the compressibility has an effect to stabilize separated shear layer. It has been verified by this experiment that the MWT can offer a unique capability to investigate airfoil performance in low Reynolds number and high Mach number flow.

The Mars Wind Tunnel at Tohoku University is a low-density tunnel that can simulate Martian atmospheric flight condition. Up to the present, this tunnel has been operated using air as the working gas. To expand the capability of this tunnel and to study the effects of compressibility on airfoil performance at low Reynolds number, a modification has been made to allow the tunnel to be operated using CO 2, that is the main constituent of the Martian atmosphere. In this study, the operational characteristics as well as the flow quality in the test section are investigated in CO2 mode and compared with the results obtained in air mode. It is found that a higher Mach number can be achieved in CO 2 mode than in air mode as the speed of sound decreases with the molecular weight. No negative effect of CO 2 such as freezing inside the ejector nozzles has been observed over the entire operational conditions of the tunnel. Mach number profile is found to be uniform over a wide range of Reynolds number except inside the boundary layers on the walls. The static pressure gradient along the test section is almost zero at the total pressure of 1kPa. These results prove that the Mars Wind Tunnel has a capability to simulate the real Martian atmospheric condition.

A new type of wind tunnel that can simulate airfoil flow in the Mars atmosphere has been designed and constructed at Tohoku University. This wind tunnel is driven by ejector and designed to simulate low-density Martian atmosphere consisting of CO2. The preliminary performance tests were conducted using air to investigate the operating envelop of the tunnel. It was demonstrated that this wind tunnel could cover the Reynolds number range from 10 4 to 10 5 and the Mach number range up to 0.71, allowing us to simulate low Reynolds number and high subsonic flow on Mars airplanes. Calibration tests were also conducted to investigate the flow characteristics in the test section. It was confirmed that the flow was kept uniform in the test section even when the total pressure was 1 kPa. There were significant effects of the boundary layer development on static pressure gradient along the test section. It was found that these effects of boundary layers could be corrected by inclining both the upper and lower walls of the test section.

The fragmentation of solid plates impacted by a high-speed projectile has been studied intensively in the past. However, the ejection and shattering process of fragmented particles have not been well resolved yet. To simulate the earlier stage of particle ejection, we performed an analogue experiment of the ejection of aluminum spheres attached on a thin aluminum plate in a row and impinged at tis reverse side by a high-spee plastic cylinder of 51 mm in diameter. The ejection of spheres was observed by using a high-speed digital camera and also holographic interefrometry. These image analysis revealed the motion of ejected spheres.

We have developed a computer-aided analysis tool for shadowgraph and schlieren optical setups with a graphical user interface by handling both the optical setups and the computational fluid dynamics. It has been demonstrated that our simulation can display simple shadowgraph and schlieren images for two-dimensional flows. In this paper, the technique is further used to analyze the detailed effects of slits and cutoffs used in a typical schlieren optical setup for axisymmetric flows. The present study successfully reveals the influences of the two factors on the schlieren image of the supersonic flow over a sphere.