Showing papers by "Kozo Fujii published in 2008"

Compact Scheme with Filtering for Large-Eddy Simulation of Transitional Boundary Layer

Abstract: Bypass transition induced by freestream turbulence is numerically simulated by compressible large-eddy simulation and implicit large-eddy simulation using a compact differencing scheme with spatial filtering. Simulated transitional flowfields qualitatively reproduce the physics of the transition and are appropriate for the discussion of transitional flowfields. Key coherent structures, such as low-speed streaks, longitudinal vortex pairs, and hairpin vortices, which play important roles in determining the behavior of transition, need to be directly resolved to properly simulate the physics of transition. Only a slight numerical damping of these coherent structures by the spatial filtering procedure causes the delay of transition. On the other hand, once the flow develops into a fully turbulent state, the spatial filtering has little influence on the flowfield. At the late stage of transition, the coherent structures break down into finer scales which require finer grid resolution to accurately resolve. However, the underresolution of the finer scale structures has little influence on the overall results if the key coherent structures are adequately resolved. Under the condition examined, the tenth-order filtering with α f = 0.495 does not act as an implicit subgrid-scale model. The present results reasonably illustrate the capability of compact/filter-based compressible large-eddy simulation and the guidelines regarding how to properly simulate transitional boundary layers using the large-eddy simulation.

Electric polarization induced by a proper helical magnetic ordering in a delafossite multiferroic Cu Fe 1 − x Al x O 2

Abstract: We have established the relationship between the electric polarization vector and the local spin arrangement including vector spin chirality in delafossite multiferroic $\mathrm{Cu}{\mathrm{Fe}}_{1\ensuremath{-}x}{\mathrm{Al}}_{x}{\mathrm{O}}_{2}$. The results of polarized neutron diffraction and pyroelectric measurements demonstrate that proper helical magnetic ordering in $\mathrm{Cu}{\mathrm{Fe}}_{1\ensuremath{-}x}{\mathrm{Al}}_{x}{\mathrm{O}}_{2}$ induces a spontaneous electric polarization parallel to the vector spin chirality. This result cannot be explained by the Katsura-Nagaosa-Balatsky model, which successfully explains the ferroelectricity in recently explored ferroelectric helimagnets, such as $\mathrm{Tb}\mathrm{Mn}{\mathrm{O}}_{3}$. We thus conclude that $\mathrm{Cu}{\mathrm{Fe}}_{1\ensuremath{-}x}{\mathrm{Al}}_{x}{\mathrm{O}}_{2}$ is another type of multiferroic.

Generation and propagation of pressure waves from H-IIA launch vehicle at lift-off

Abstract: Numerical simulation was carried out to investigate the generation mechanism of pressure waves radiated from the H-IIA launch vehicle at lift-off. It was revealed that the Mach wave due to the large-scale structure of the unsteady supersonic exhaust plumes is the dominant noise source. The Mach wave propagating obliquely downstream is reflected from the constructions of the launch-pad, and then, turns to reach the vehicles, resulting in the acoustic loading. It was also found that the fluctuating supersonic plume entering into the flame duct is the dominant noise source that appears in the flame duct. Then, the pressure wave propagates through the flame duct and is ejected outside to the vehicle.

Development of multi-objective six-sigma approach for robust design optimization

Abstract: In this study, a new optimization approach for robust design, design for multi-objective six sigma, has been developed and applied to three robust optimization problems.The design for multi-objective six sigma builds on the ideas of design for six sigma, coupled with multiobjective evolutionary algorithm, for an enhanced capability to reveal tradeoff information considering both optimality and robustness of design. While design for six sigma requires careful input parameter setting, design for multi-objective six sigma needs no such prior tuning, plus it can reveal the tradeoff information in a single optimization run. Three robust optimization problems were taken as to demonstrate the capabilities of design for multiobjective six sigma. Results indicate that design for multi-objective six sigma has a more practical and more efficient capability than the design for six sigma to reveal tradeoff design information considering both optimality and robustness of design.

Over-expansion Effects on Mach 3.0 Supersonic Jet Acoustics

Abstract: For the prediction of acoustic waves from rocket plume, over-expansion effects on Mach 3.0 supersonic jet acoustics are investigated with monotonically integrated large eddy simulation. In this study, ideally-expanded Mach number and designed Mach number are used to express the jet conditions. Three designed Mach number 3.0, 3.5 and 4.0 are chosen, while ideally-expanded Mach number is constant 3.0. Reynolds number is 100000 and coldjet condition is adopted to reduce computational costs. With regard to computation, the seventh order weighted compact non-linear scheme and the tenth order compact scheme are used for solving the jet-flow and near fields acoustics propagation, respectively. Computational results of Mach wave emissions of these three conditions are almost same. This result corresponds to that of relatively low Mach number supersonic which is reported by Tam(AIAA paper 2005-2938). Present results show that it can be applied to high Mach number supersonic jet. For the the prediction of acoustic waves from rocket plume, this results imply that rocket parameters could be reduced.

CFD contributions to high-speed shock-related problems

Abstract: Computational Fluid Dynamics (CFD) has contributed extensively to high speed shock-wave research. With study examples by the author’s group in the past, effectiveness of CFD both for design of transportation vehicles and for understanding of fluid physics is discussed. Trends of CFD for further use are then discussed based on recent applications and three key features: computer progress, spectral-like high-resolution schemes and LES and LES/RANS hybrid methods are focused. Recent CFD research reveals that high-speed flows, even the ones considered to be steady state have inherently unsteady nature that requires LES-like computations for successful simulations. Such simulations require remarkably higher grid resolution, but emerging numerical techniques having spectral-like high resolution would help reducing the number of grid points required for such simulations and make them feasible. The paper is summarized by addressing issues of future CFD.

Numerical Study of Pressure Waves Generated by H-IIA Launch Vehicle at Lift-off

Abstract: Generation mechanisms of pressure waves from the H-IIA launch vehicle are analyzed numerically. The Mach wave radiated downstream from wavy shear-layer of supersonic exhaust plume is revealed to be the dominant noise source. Reflecting from the constructions of the launch-pad, the Mach wave turns to propagate to the vehicles. It was also found that the fluctuating supersonic plume entering into the flame duct is the dominant noise source that appears in the flame duct. Then, the pressure wave propagates through the flame duct and is ejected outside to the vehicle.

Aerodynamic Characteristics of a Delta Wing at High Angles of Attack

Abstract: any supersonic aircrafts use delta wing and they often fly at high angles of attack. For example, in landing or taking off phase, they need to fly at very high angle of attack due to their poor aerodynamic performance at low speeds. Future space plane may fly at high angle of attack even at transonic and supersonic speeds in the reentry phase. When an aircraft with delta wing flying at high angle of attack in low speeds, there appear two large counter-rotating leading edge vortices. However, when an aircraft with delta wing flies at much higher speeds, the

Comparative Study of Data Mining Methods for Aerodynamic Multiobjective Optimizations

Abstract: Practical aerodynamic design problems are typically multiobjective design optimization problems that have multiple contradicting objectives and many design parameters. Goal of multiobjective design optimization is to find Pareto-optimal solutions to reveal tradeoff information between the objectives and effect of each design parameters. Recently, idea of “multi-objective design exploration (MODE)” [1] was proposed by Obayashi et al. as an approach to find such design information. They proposed to use multiobjective evolutionary algorithm to find Pareto-optimal solutions and to use data mining methods such as self-organizing map (SOM) to extract design information from the Paretooptimal solutions. However, it has not been discussed yet which data mining method is suitable for analysis of Pareto-optimal solutions among many data mining methods. Therefore, the objective of the present study is to apply data mining methods such as SOM, clustering, and decision tree to an aerodynamic design optimization problem and to find the best data mining approach for aerodynamic multiobjective optimizations. Here, Pareto-optimal solutions of the multiobjective aerodynamic design optimization problem of flapping airfoil motion [3] are considered. Objectives are maximization of the time averaged lift (CL,ave), maximization of the time averaged thrust (CT,ave), and minimization of the time-averaged required power (CPR,ave) at the given cruising condition. The flapping motion of the airfoil is parameterized by frequency (k), plunge amplitude (h), pitch amplitude (α1) and offset (α0), and phase shift between plunging and pitching (φ). The objective values are evaluated using a two-dimensional imcompressible Navier-Stokes solver and a multiobjective evolutionary algorithm code is used to obtain Pareto-optimal solutions. As the result, 561 Pareto-optimal solutions is obtained. In this abstract, SOM, scatter plot matrix, and clustering are compared.

Mach number effect on flowfield over a delta wing in supersonic region

Abstract: To understand Mach number effect on flow field over a delta wing with blunt leading edge in supersonic and high angle of attack region, wind tunnel experiments of a 65° delta wing are performed in supersonic and high angle of attack flow conditions at the JAXA’s transonic / supersonic wind tunnel. Oil flow for surface flow visualization, pressure sensitive paint for surface pressure distribution measurement, and Schlieren images for shock wave visualization are used. The present results indicate that a delta wing with blunt leading edge can be mixed flow of two different types of flow structure in supersonic and high angle of attack flow region and the location of the boundary of the two types of flow moves toward the apex of the wing as the free-stream Mach number increases.

Visualization and verification method for failure network analysis of space launch vehicles

Abstract: Comprehensive failure network analysis method was studied for liquid rocket engine development which includes failure propagation through various types of component interfaces in order to achieve exhaustive enumeration of possible failures and to identify actions to eliminate or reduce the potential failure. New failure network visualization method was developed in order to make it easier to understand complicated failure propagation mechanism among multiple system levels. Verification analysis method is developed in which it is verified all of user-specified component interfaces are contained in the failure network analysis result. The perceived component interface is specified by the analyzer and the failure propagation in the result of failure analysis is summarized in two separate N2 charts. By comparing with these two N2 charts, unperceived component interface and the unconsidered failure propagation can be found. It is found to be promising approach to achieve exhaustive enumeration especially for forgettable component interface.Copyright © 2008 by ASME

Unsteady Computational Analysis of Supersonic Underexpanded Jet Impinging on an Inclined Flat Plate

Abstract: Payload oscillation due to strong acoustics emitted from rocket plume is one of the most significant problems in rocket launch. It is known that the rocket plume produces strong acoustics when it impinges the rocket plume deflector. Therefore, it is important to understand this phenomenon, which can be modelled with a supersonic jet impinging on an included flat plate. Supersonic jet impinging on an inclined flat plate was experimentally studied by P. J. Lamont and B. L. Hunt 1 . They measured the pressure on the plate surface with pressure taps and visualized the flow fields with shadow graph method. However, the surface pressure data were limited because they are measured on discrete points. Nakai et al. 2 conducted experimental study of the jet impingement on an included flat plate using pressure sensitive paint and Schlieren method. Based on continuous surface pressure data on the plate and the Schlieren images, they classified the flow fields into three types, corresponding to the different shock wave structures (Figure. 1). This classification enables flow type prediction if pressure ratio, the inclined angle of the plate, and the nozzle-plate distance are given. These past researches, however, only steady phenomena are analyzed. Therefore, unsteady phenomena of the supersonic jet impingement, which are key mechanism of acoustics emission, are not well-known. An objective of the present research is to understand unsteady flow phenomena of the supersonic jet impingement using computational fluid dynamics. So far, steady flow structures have been clarified using RANS simulation, where unsteady flow structure will be analyzed using high resolution scheme. In this abstract, steady analysis of the jet impingement is shown. In Figure 2, pressure distributions at the pressure ratio PR=7.4 for four different plate angles and four different plate distances are shown as an example. Some cases have a single peak and some cases have a few different types of peaks. In these cases, there observed are four types of pressure peaks; (1) stagnation point of the main stream, (2) strong shock waves in the upstream area, (3) reattachment of the detached flow and (4) interaction between the intermediate tail shock and the boundary layer. Figure 3 shows

Body-fitted cartesian grid method for complex high reynolds number flows

Abstract: In order to improve turnaround time and usability of computational fluid dynamics (CFD) for conceptual aerodynamic design of space launch vehicles, CFD analysis method based on the body-fitted Cartesian grid is developed. As comparing with conventional unstructured hybrid grid method, this approach has strong advantages such as efficiency of filling space, fas ter convergence rate and robustness to handle complicated geometry[1, 2, 3, 4, 5, 6, 7]. In this approach computational grid is generated in the following process, 1) generation of volume Cartesian grids with keeping near surface space, 2) generation of the grid front which covers body, 3) smoothing of grid front, 4) grid front projection onto the surface, 5) geometric feature preservation and 6) clustering of laye r grid. In order to receive the full benefit of this approach robustness of the feature preservation is key issue, and its improvement was conducted in past studies by present authors[8]. In this study, further improvement to handle small gap problem and to realize explicit grid resolution control is carried out. Body-fitted Cartesian grid generated over the roc ket geometry is shown in Fig. 1. In order to capture small gap area such as inside of the nozzle and chamber, such area is detected and size of the Cartesian cells near this area is controlled to be sufficiently small. By using this treatment, this grid generation method can also be used for internal flows which include s small gaps such as rocket engine inducer. Flow computations based on this grid system is applied to high Reynolds number fl ows such as airfoil, massive separated re-entry capsule configuration and aerodynamic an alysis of an aircraft geometry. Computed pressure distributions for RAE2822 airfoil is compared with the experimental data in Fig 2. In the computation, solution adaptive grid refinement method is used to capture the local shock wave accurately, which results in the good agreement with experimental data. Most of the cells are Cartesian cell, and thus, grid quality does not degenerate by cell refinement. There fore, it is easy to conduct grid refinement depending on the flow-field, which is strong advantages in the s ituation that flow structure is unpredictable.

Computational analysis of noise sources inside the high speed flow over a generalized bump

Abstract: The aerodynamic noise sources around a three dimensional bump are studied. Firstly in this paper, validation of the numerical method with ILES using symmetric bump is discussed. Cp distribution on the bump is discussed and ILES simulation result shows good agreement with Visbal’s numerical result 1 . The gradient of spanwise velocity fluctuation spectrum of present result agrees with gradient of spanwise velocity fluctuation spectrum of Gwibo Byun’s experimental result 2-4 . At the inertial sub-range, the gradient of sound pressure fluctuation spectrum of present result agrees with the theoretically-determined gradient. Second, the noise level around the bump which is generalized by geometrical parameters is investigated with LES. On the head of bump, wall pressure fluctuation spectrum has a peak besides at other measuring points’ don’t have. At the front of bump and the back of bump, SPL becomes large below St = 0.79. Moreover longitudinal vortex at side of head of bump generates the loudest noise for all St number.

Multiobjective design exploration and its application to an aerodynamic flapping airfoil design

Abstract: Aerodynamic knowledge for flapping airfoil is obtained by application of the multi-objective design exploration framework to a multiobjective aerodynamic flapping airfoil design optimization problem, where the airfoil oscillates in plunging and pitching modes. Pareto-optimal solutions are obtained by a multiobjective evolutionary optimization and analyzed with the self-organizing map. Aerodynamic performance of each flapping airfoil is evaluated by a two-dimensional Navier-Stokes solver. Analysis of the flow over the extreme Pareto-optimal flapping airfoils provides insights into flow mechanism for thrust maximization, lift maximization, and required power minimization. Analysis of the design objectives and design parameters with the self-organizing map leads to useful guidelines for practical flapping-wing micro air vehicles. Nomenclatures c = airfoil chord CL(t) = lift coefficient CPR(t) = required power coefficient CT(t) = thrust coefficient h = plunge amplitude nondimensionalized with c