Shunsuke Koike

Experimental Investigation of Vortex Generator Effect on Two- and Three-Dimensional NASA Common Research Models

Aerodynamic characteristics of twoand three-dimensional NASA common research model (2D-CRM and 3D-CRM) with co-rotating vortex generators (VGs) were investigated to clarify the influence of the three-dimensionality of the wings on the VGs effect. The base height of the VGs was 1.5 times of the boundary layer thickness at the VGs location. The direction of the VGs on the 3D-CRM was toe-out which meant the leading edge of the VGs turned to the wing tip. The Mach numbers in the 2Dand 3D-CRM experiment were 0.74 and 0.85 considering the sweepback angle of the 3D-CRM. The lift coefficient and the oil flow visualization showed that the effect of the VGs on the 3D-CRM was much larger than that on the 2D-CRM. From the comparison between the experiments and the CFD results, we concluded that the difference between 2Dand 3D-CRM was mainly caused by the crossflow due to the swept wing. The cross-flow enhances the effect of the co-rotating toe-out VGs on the swept wings. The installation drag of VGs was also investigated for the 3D-CRM and validated an empirical method to estimate the installation drag. At CL conditions below the design CL = 0.5, the VGs increased the total drag as expected, while at CL conditions above the design CL, the VGs decreased the total drag because the VGs suppressed the separation and the effect exceeded the installation drag of the VGs.

Effect of Vortex Generators on Transonic Swept Wings

This paper examines the effects of corotating blade-type vortex generators on transonic sweptback wings using computational fluid dynamics studies. Infinite-span (two-dimensional) swept wings are first considered to understand the basic physics of vortex generators. Sweep angles are given virtually to the wings by changing the freestream direction. Vortex generators are placed on the wings, and the visualized interactions of their tip vortices with the boundary layer reveal the relationship between the effect of the vortex generators and the wing sweep angle. The physics of vortex-generator tip vortices are then described to explain how vortex generators on swept wings efficiently suppress shock-induced separation by mixing boundary layers. It is also shown that the vortex-generator angle of incidence to the local flow can slightly improve the effect of the vortex generators but that wing sweep angle has a greater influence on their effect. Based on the discussion of infinite-span wings, the computational...

Unsteady PSP Measurement of Transonic Buffet on a Wing

In this study, unsteady pressure field caused by transonic buffeting phenomena on a generic transport model was analyzed using unsteady PSP. The tests were conducted in the JAXA 2-m Transonic Wind Tunnel at M=0.85 and the model angle of attack was varied between 2.5 and 6.8 degrees. From the obtained unsteady PSP data, the RMS of pressure fluctuation was calculated to detect the areas with strong shock oscillation. Spectral analysis was also conducted in order to find frequencies of shock oscillation and their correlation with pressure fluctuation. As the angle of attack increases, the shock position moves upstream and shock oscillation becomes stronger. At 6.8 degrees, two distinct areas with strong shock oscillation were observed. The spectral analysis shows that, under buffeting conditions, shock oscillation frequencies are distributed around either of 490 Hz or 365 Hz. Noticeable effects from the PSP coating on the measured aerodynamic forces were noticed, indicating that even the small roughness of the unsteady PSP coating has an influence on the transonic buffeting phenomena.

Unsteady Pressure Measurement of Transonic Buffet on NASA Common Research Model

Experimental investigation of transonic buffet was conducted in JAXA 2m×2m transonic wind tunnel in order to obtain the validation data for unsteady computational fluid dynamics and to clarify the buffet phenomena of an 80% scaled NASA common research model. Unsteady pressure distributions on the two lines of the main wing were successfully measured on the transonic buffet condition. Mach number of the uniform flow was 0.85. Reynolds numbers based on the reference chord length were 1.515×10 and 0.947×10. The shockwave oscillation on the wing can be classified into three regions, a small oscillation region without separation, an oscillation region with bump in the power spectrum, and a large oscillation region with broadband power spectrum. The Strouhal number based on the bump peak frequency was about 0.3. The cross-correlation and the phase analysis revealed that the pressure fluctuation of the bump frequency propagated from the wing root side to the wing tip side.

Study on the particle traceability in transonic and supersonic flows using molecular tagging velocimetry

AbstractnParticle image velocimetry (PIV) has become a powerful tool for flow velocity measurements in wind tunnel testing. However, it is generally difficult to apply the PIV technique to supersonic flows because of unreliable particle traceability. In the present study, the PIV and MTV (molecular tagging velocimetry) techniques are applied to transonic and supersonic flows, in which a normal shock wave appears, to evaluate particle traceability. Based on this work, it is found that the PIV data largely deviate from the MTV data behind a normal shock wave for both flows. The drag coefficient is also estimated for a particle from the velocity data measured by the two techniques. Its value is then compared to the drag coefficient value calculated from an empirical formula for particle Mach numbers ranging between 0.1 and 0.9. Based on the results, it is found that the experimental data can be reproduced reasonably well by the formula for particle Reynolds numbers higher than ~1. However, the data associated with particle Reynolds numbers lower than ~1 deviate largely from the formula.Graphical abstract

Wavelet analysis of unsteady shock-wave motion on two-dimensional airfoil with vortex generators

We visualized the shock wave-boundary layer interaction on a two-dimensional transonic airfoil with and without vortex generators (VGs) by using a fast-framing focusing schlieren imaging. Image processing extracted a time-space trajectory of the shock wave motion from a time-series of the schlieren images. The shock trajectories were analyzed by using the Morlet wavelet. The shock motions in the cases without VGs were quite periodic. The Fourier analysis well extracted the characteristics of the shock motion such as frequency etc. However, the shock motions in the cases with VGs were not periodic. Therefore, the Fourier analysis is not applicable for this case. The wavelet analysis with the statistical significance test gave quantitative measure of change in the shock oscillation around the buffet frequency due to installation of VGs, such as amplitude and intermittency. The wavelet spectrograms revealed that the installation of VGs did not prevent from generating the buffet but just reduced the amplitude of the buffet oscillation.

Polymer/Ceramic PSP with Reduced Surface Roughness for Unsteady Pressure Measurement in Transonic Flow

Polymer/ceramic pressure-sensitive paints (PC-PSPs) with reduced surface roughness were developed for measuring unsteady pressure fields in transonic flow. Four types of PC-PSPs, each having different particle size, mass content, solvent, and so on, were formulated and applied to transonic wind-tunnel tests of a Common Research Model airfoil. The effects of surface roughness on unsteady transonic flow on the airfoil were evaluated at Mach 0.74 and Reynolds number of 5.0 x 10 6 . It was found that all four PC-PSPs had capabilities to measure time-series pressure distributions, but the location of a shock wave and the root-mean-square pressure fluctuations differed depending on types of PC-PSP. Among all tested PC-PSPs, the PC-PSP having arithmetic surface roughness of 0.5 μm and cutoff frequency of 3 kHz yielded data practically the same as that of a clean airfoil. Using this PC-PSP, propagation of pressure waves and oscillation of shock waves on the airfoil were clearly captured. A spectral analysis showed that the fundamental frequency of shock-wave oscillation agreed very well with that calculated based on the mechanism proposed by Lee. These results show that the selected PC-PSP can offer a powerful means to study transonic buffeting on airfoils and 3D wings.

In-Flight Wing Deformation Measurement

This study has been conducting research on in-flight measurement for deformation of main wing of an aircraft. In this paper, a measurement method which optically measures the deformation of the main wing using high-resolution cameras is discussed. It also discusses a measurement system and a camera calibration. The measurement technique was applied to JAXA’s experimental aircraft, a twin-propeller plane with sweepback angle of zero degree, and it demonstrates to measure the quantitative changes of bending and twist angle of the main wing in flight.

Particle Image Velocimetry for Transonic Unsteady Flow Field around a Rocket Fairing Model

Velocity distributions around a rocket fairing model were measured with PIV to obtain a validation data for numerical simulation and to clarify the influence of the expansion waves on the turbulent boundary layer. Mach number of the uniform flow was 0.8. Reynolds number based on the diameter of the fairing cylinder was 2.6 x 10. Velocity profiles in the boundary layer around the shoulder of the cone-cylinder body were successfully measured using a 2D-PIV system and a specially designed seeding rake. The velocity data was evaluated considering the influence of the traceability of the particles. The profiles of the velocity fluctuation around the shoulder were complex due to the interaction between the boundary layer and waves. The Reynolds stress of the boundary layer was reduced by the expansion waves emanating from the model shoulder.

PIV around a NACA0012 Airfoil with a Plasma Actuator for Noise Reduction

A dielectric barrier discharge plasma actuator (PA) was applied to reduce trailing edge noise and vortices emitted from a NACA0012 airfoil model. The PA in length of 1 m was installed at 60 % of the chord length on the pressure side of the model. The velocity field around the trailing edge of the model was measured via 2D-PIV at the uniform flow velocity of 15, 30, and 40 m/s. The Reynolds numbers based on the code length of the model for the three cases were 2.0x10 5 , 4.0x10 5 , and 5.4x10 5 , respectively. When the PA was off, the measured velocity and the result of the proper orthogonal decomposition showed that the strong vortices were periodically emitted from the trailing edge of the model at the uniform flow velocity of 15 m/s. These trailing vortices were dramatically reduced by the PA. As the uniform flow velocity increased, the trailing edge vortices decreased. As the result, the effect of the PA was decreased as the uniform flow velocity increased.