青山 剛史

We simulate unsteady flows over a 30P30N slat by taking account of velocity fluctuations inside the turbulent boundary layer (TBL) on the slat. Embedded Large Eddy Simulation method is employed to generate pseudo turbulences inside the TBL. Also, we conduct additional simulation by Delayed Detached Eddy Simulation (DDES), which has been made without the pseudo turbulence, for making comparisons with the ELES case. Comparisons between the ELES and DDES results suggests that the pseudo turbulence makes the vortex structures of a shear layer extending from the cusp more three-dimensionally. The difference of vortex structure leads to the shear layer in the ELES case curves with a smaller radius of curvature. The time-averaged flow field of the ELES case is closer to the experimental result than those in the DDES case. While the shear layer on the cusp side is affected by the pseudo turbulence, there is no significant difference between the ELES and DDES cases in the shear layer extending from the TE of the slat.

The Japan Aerospace Exploration Agency (JAXA) and The Korea Aerospace Research Institute (KARI) jointly started activities related to the research and development of "Active Tab," a helicopter noise reduction technique. KARI constructed the analytical methodology consisting of the aerodynamic, structural dynamic and acoustic codes for defining the requirements to be used in evaluating the performance of Active Tab when installed in a Mach scaled assumed blade. Based on the requirements defined, JAXA carried out a conceptual design study, developed the Active Tab drive mechanism and evaluated its performance. The analytical results show Active Tab satisfying the requirements has sufficient noise reduction capability. Evaluation for the Active Tab drive mechanism demonstrated the dynamic performance and durability required practical use installed in helicopter blades.

The influence of walls on the performance of multiple rotor type drones is numerically simulated. With the current wide-spread applications of autonomously flyable UAVs, there are potential needs to use the drones for inspections and observation near various structures, such as buildings and bridges. The flow fields around multiple rotors are influenced significantly by the existence of an upper or a side wall, so that the thrust, required torque, and other forces and moments are changed. It is found that when a rotor approaches an upper wall, as the distance between the rotor and the upper wall is less than a diameter of the rotor, the thrust suddenly increases, which causes the rotors to collide with the upper wall. When an isolated rotor is operated near a side wall, the thrust decreases and a rolling moment appears to tilt the rotor toward the wall as the distance becomes shorter. For a multiple rotor drone near a side wall, the rotors have different distances from the wall, which causes the whole aircraft tilts toward the wall. From the perspective of safety operations, the multi-rotor drone should be kept away from both the upper wall and the side wall at a distance of at least 1.5 times of the rotor diameter to prevent unexpected motions of the aircraft caused by the wall during hovering flight.

Copyright © 2018 by the American Institute of Aeronautics and Astronautics, Inc. Three mixed-element unstructured Reynolds-averaged Navier-Stokes solvers, the TAS code, FaSTAR, and Cflow, were applied to the Sixth AIAA Drag Prediction Workshop test cases to investigate the accuracy of the solvers with committee-provided grids and custom participant-generated grids by MEGG3D, BOXFUN, and the Cflow built-in grid generator. Solutions were obtained for the verification study using the two-dimensional NACA 0012 airfoil (test case 1), the NASA common research model nacelle-pylon drag incremental study (test case 2), and the common research model wing/body static aeroelastic effect study (test case 3). In test case 1, the three solvers used the same turbulence model, the Spalart-Allmaras one-equation turbulence model without the trip term for transition and without the ft2 function, and their grid convergence trends for lift, drag, and pitching moment coefficients were similar to that of FUN3D, CFL3D, and TAU. In test case 2, the TAS code with the NASA GeoLab grids and with MEGG3D grids, FaSTAR with the NASA GeoLab grids, and Cflow with its own grids predicted nacelle-pylon drag increment well. In test case 3, the three solvers predicted force and moment fairly comparable with experiment if support system interference effects were considered.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Predicting a transonic buffet over an aircraft is one of the challenging problems for the current CFD community due to the unsteady shock/boundary layer interaction involving flow separation. We have simulated the transonic buffet on NASA-CRM with a Zonal DES (ZDES) method implemented in an unstructured grid CFD code FaSTAR. The previous results showed that it is difficult to predict the shock location with the ZDES method. We have investigated the effects of turbulence modeling and grid resolution to improve the buffet prediction accuracy. First, we compared two ZDES with the Spalart-Allmaras (SA) and Menter’s Shear Stress Transport (SST) models. Although we found the buffet simulation depends on the turbulence model employed for ZDES, we could not predict the shock location accurately for both models. Next, we employed finer grids to investigate the effect of grid resolution. We found the small vortex structure is resolved with the fine grids. However, the shock wave location did not change largely. Then, we implemented the wall model with ZDES. The number of cells was reduced by one-third, and this is useful to shorten the computational time required for unsteady simulation. The shock wave location predicted with wall model is close to the experiment. We observed the buffet cell convection in the spanwise direction. The power spectral density of pressure is also close to the experiment. In this study, we found that the ZDES with wall model can predict the three-dimensional buffet reasonably well.

This paper presents a numerical analysis on the air intake of HIMICO (High-Mach Integrated Control experiment). Mass capture ratio (MCR) of the intake in the supersonic wind tunnel test is different from that we analytically expected quantitatively and qualitatively. CFD results cleared that the reason of the difference is the flow separation on the second ramp surface located at the upstream of the throat. The flow separation is caused by the spillage from clearances between the side walls and 2nd/3rd ramps. Some treatments on the intake configuration such as removing the clearances, adding the additional bleed holes with demerging bleed plenum chamber can improve the inlet performance and prevent the flow separations.

Nonlinear propagation through a relaxing atmosphere of pressure disturbances extracted from a computational fluid dynamics (CFD) solution of the flow around a supersonic aircraft is simulated using an augmented Burgers equation. The effects of nonlinearity, geometrical spreading, atmospheric inhomogeneity, thermoviscous attenuation, and molecular vibration relaxation are taken into account. The augmented Burgers equation used for sonic boom propagation calculations is often solved by the operator splitting method, but numerical difficulties arise with this approach when dissipation is not effective. By re-examining the solution algorithms for the augmented Burgers equation, a stable method for handling the relaxation effect has been developed. This approach can handle the Burgers equation in a unified manner without operator splitting and, therefore, the resulting scheme is twice as fast as the original one. The approach is validated by comparing it with an analytical solution and a detailed CFD of dispersed plane wave propagation. In addition, a rise time prediction of low-boom supersonic aircraft is demonstrated. (c) 2015 Acoustical Society of America.

The flow around an ONERA-M6 wing, including the effect of wind tunnel wall interference, is computed using CFD analysis with a porous wall model. The computational domain sets porous walls at the top and bottom of the wing section similar to actual wind tunnel experiments. The computational result captures almost the same shock wave shape as the wind tunnel. This could not be computed exactly in previous works that did not include wall effects. The interference of the porous walls reduces the Mach number and attack angle of the flow, and these effects alter the swept angle of front shock wave and the location of rear shock wave. The aerodynamics coefficients are also affected by the wall interference. The lift coefficient becomes smaller due to the reduction in attack angle. The lower Mach number decreases the drag coefficient, while the reduction in attack angle causes additional drag by a mechanism similar to induced drag.

In the present study, the interaction between a wind turbine tower and blades and its aerodynamic effect on the wake structure are investigated using the rFlow3D CFD code, which was developed by JAXA. NREL Phase VI experimental wind turbine is selected as the computational test case. The result shows the shed vortex from the tower does not affect the total wind turbine performance significantly. However, cyclic fluctuation of the aerodynamic load and the change of vortex structure behind the tower are captured clearly. The rotational flow and shed vortex from the tower cause the increase of turbulence intensity and irregular velocity distributions. This is considered as one of the key features for capturing the wake vortex breakdown accurately. It is shown that the tower and blade interaction is important for detailed wake investigation even in this upwind wind turbine case.

To model porous walls used in transonic wind tunnels, flow through a hole is investigated using Computational Fluid Dynamics (CFD). First, we analyze the relation between flow rate and differential pressure across the hole. At low differential pressures, such as for wind tunnel porous walls, the flow rate is found to increase linearly with differential pressure. We therefore propose a new model based on a linear relationship between flow rate and differential pressure. The effects of hole shape and boundary layer conditions near the hole are then investigated. In the outflow case (i.e., wind tunnel to plenum chamber), the flow rate increases as the ratio of hole depth to diameter becomes large due to variation of the flow separation area at the hole exit. Boundary layer thickness also affects the flow field: when the ratio of boundary layer thickness to hole diameter becomes small, the flow rate decreases, because the flow along wind tunnel side wall interacts more strongly with the flow through the hole.

Insects generate sound by their flapping wings as a consequence of spatial and temporal changes of pressures on the wing surface and vortices generated by the wing motion. To clarify the mechanism of sound generation, hybrid method combining CFD techniques and acoustic analysis is incorporated here and detailed characteristics of flapping sound, e.g. directivity of transmission or spectrum distributions, are clarified.

JAXAではヘリコプタ飛行時の主要な騒音であるBVI(Blade Vortex Interaction)ノイズの予測システムであるMENTORを開発している。その中で先行するロータブレードが放出した翼端渦をBeddoseの渦モデルを用いてモデル化しているが、このモデルはブレードが作る誘導流中の渦の幾何学的な配置を実験結果に基づいてモデル化しており、計算コストを抑えながら実用的な精度で渦の影響を考慮できるメリットを持つ。しかしながら、渦の減衰の効果を含まないため、予測されたBVIノイズは実際の計測値に比べて過大評価される傾向があった。そこで、本研究ではBeddoseの渦モデルに渦の減衰の効果を取り入れて計算を行い、渦の減衰がBVI騒音に与える影響を評価した。

Two types of prediction tools for helicopter noise have been developed under the cooperative research between Advanced Technology Institute of Commuter Helicopter Ltd. (ATIC) and National Aerospace Laboratory (NAL) . One of them is a combined method of CAMRAD II, interpolation code for blade motion and wake geometry, aerodynamic code of 3D unsteady Euler solver, and aeroacoustic code based on Ffowcs Williams and Hawkings (FW-H) formulation. The other consists of CAMRAD II, 3D unsteady Euler solver using moving overlapped grid method, and FW-H code. The acoustic waveform of Blade-Vortex Interaction (BVI) noise predicted by the former tool is in good agreement with the experimental data of 1/7-scale model AH-1 Operational Loads Survey (OLS) rotor. This method is applied to investigate the effect of blade-tip shape on the intensity of BVI noise. As a result, it is shown that anhedral and swept-forward tip shapes effectively reduce the BVI noise of OLS rotor in a descent flight condition. The predicted Effective Perceived Noise Level (EPNL) of a helicopter is also compared with the experimental data obtained by ATIC and reasonable correlation is obtained. The latter tool successfully predicts the distinct spikes in the BVI wave-form of ATIC model rotor tested in German-Dutch Wind Tunnel (DNW) . In the comparison of measured and calculated carpet noise contours, reasonable agreement is obtained. The present tools are expected to be useful for the design of low-noise helicopters in the future.