Showing papers by "Kozo Fujii published in 2006"

Experimental Study of Underexpanded Supersonic Jet Impingement on an Inclined Flat Plate

Abstract: Flowfields of the M2.2 jet impinging on an inclined flat plate at various plate angles, nozzle-plate distances, and pressure ratios are experimentally investigated using pressure-sensitive paints and Schlieren flow visualization. The effect of temperature variation on the flat plate is eliminated by the calibration using temperature-sensitive paints. A comparison with the former experiment shows that the current pressure-sensitive paint measurement produces reliable data sufficient for the flowfield discussion under consideration. The pressure-sensitive paints and Schlieren images obtained in the experiment suggest that the flowfields at various flat plate angles, nozzle-plate distances, and pressure ratios can all be classified into three types of flow structure. Within the range of geometrical and flow conditions considered in the present paper, the flowfield patterns can be predicted without experiments if the shock cell length in the freejet for various pressure ratios is known in advance. Extensive studies for a wide range of geometrical and physical parameters became feasible with a new efficient pressure/temperature-sensitive paint measurement technique showing surface pressure map with much less effort compared to conventional pressure tap measurement.

Time-series and time-averaged characteristics of subsonic to supersonic base flows

Abstract: Physics of cylindrical base flows ranging from subsonic to supersonic speeds at zero angle of attack are computationally investigated. Time-series and time-averaged investigations of base flows show distinctive characteristics at subsonic, transonic, and supersonic regimes. Normalized time-averaged base pressure decreases proportionally with respect to increasing freestream dynamic pressure in the subsonic regime of M ∞ 1.5. Normalized base pressure fluctuations sharply increase at transonic speeds, whereas they decrease with increasing freestream Mach number at subsonic and supersonic speeds. Appearance of unsteady local shock waves change the characteristics of base pressure distinctively at the transonic speeds. Spectra of the base pressure show one clear peak at subsonic speeds (related to the shear layer dynamics), two clear peaks at transonic speeds (related to the shear layer dynamics and its subharmonic), and three major peaks at supersonic speeds (related to the shear layer dynamics, its subharmonic, and an additional mechanism). Instability of the free shear layers has dominant influence on the overall base flowfield over a wide range of Mach numbers ranging from subsonic to supersonic speeds. However, at supersonic speeds, an additional mechanism of instability within the recirculating region is possibly at work and has dominant influence on the flowfield. The dominant mechanisms significantly cause the strong Mach number dependence of the high-pressure region which is strongly related to the base pressure. The spectrum analysis suggests that the substantial base pressure fluctuations are caused by the pulsing of the flow inside the recirculating region.

A Study on Airfoil Design for Future Mars Airplane

Abstract: An optimum airfoil design for future Mars airplane for Mars exploration is obtained by evolutionary computation coupled with a two-dimensional Reynolds-averaged Navier-Stokes solver. The optimized airfoil design is also compared with other airfoil designs optimized at different Reynolds number or at different Mach number to discuss Reynolds number and Mach number effects on airfoil design. The present results indicate 1) optimum airfoil shape in the sense of maximization of lift-to-drag has thin thickness distribution, 2) Optimum camber distribution depends on both Reynolds number and Mach number of the design condition, 3) Optimum leading edge shape depends on Reynolds number of the design condition.

Computational aerodynamic analysis of capsule configurations toward the development of reusable rockets

Abstract: Flowfields around Apollo-type conical configurations are numerically simulated by Reynolds-averaged Navier-Stokes computations at a wide range of angles of attack under subsonic through supersonic flows. Effects of the configuration parameters on the aerodynamic characteristics and their flow mechanisms were clarified based on the computed results. It turns out that the aerodynamic characteristics are mainly influenced by the separation position of the flow, the pressure level behind the body, and the wind-side high pressures. The shoulder radius has a great influence on the pressure level behind the body and the separation position, so that the aerodynamic characteristics, especially at base-first conditions, significantly change. The drag coefficient at base-first conditions becomes small for configuration with the large shoulder radius because of the increase of the lee-side pressure level. At both nose-first and base-first conditions, the fineness ratio has great influence on the location and the pressure level of the wind-side high pressures, which appear on the windward cone part, so that the lift, drag, and pitching-moment coefficients significantly change.

Computational Analysis of Mach Number Effects on Edgetone

Abstract: In this study, Mach number effect on edgetone is investigated to verify the feedback-loop of edgetone using the high-order computation. The computational results show three clear points. When the Mach number increases independently, 1) the edgetone phenomenon tends to cease 2) the frequency of edgetone becomes lower 3) the oscillation mode (which is named stage) of edgetone becomes lower. The second point shows that the edgetone mechanism is explained by the fluid-acoustic feedback-loop. As for the Powell's feedback-loop equation our computational results show that phase-lag p is constant. Therefore the feedback-loop equation is verified to be physically correct. However the computed value of p is -0.2 which does not correspond to that of Powell's suggestion.

Improvement of the Optimization Method of the TSTO Configuration — Application of Accurate Aerodynamics

Abstract: A new aerodynamics model was constructed using the data obtained by the contemporary CFD simulations of the TSTO spaceplane, and the effect of the aerodynamics model on the comprehensive optimization of TSTO was investigated The comparison of the aerodynamic coefficients by the CFD simulations and those by the conventional method shows that both the lift and drag of the TSTO configuration increases at transonic to supersonic flow region by the CFD simulations There are two keys fot these increases: the interaction of the flow fields over the Booster and the Orbiter, and low pressure region in the Orbiter’s base Therefore, these keys should be considered for an accurate optimization The CFD-based aerodynamics model evaluated aerodynamic performance of the TSTO more severely than the conventional model and the optimization using the CFD-based aerodynamics model results in the heavier gross take-off weight compared to the result using the conventional model

Large-eddy simulation of compressible transitional boundary layer

Abstract: Compressible large-eddy simulation (LES) is applied to the bypass transition of a compressible boundary layer over a flat plate. Simulated time-averaged and statistics flow properties and analyses of flowfields and flow spectra reasonably demonstrate that compressible LES with the adequate grid resolution and numerical scheme ensures qualitative discussions for the compressible transitional turbulent flowfields. The interactions of relatively large-scale coherent structures, such as backward flow streaks, pair of clockwise and counter-clockwise rotating longitudinal vortices and hairpin vortices play key roles in determining the behavior of the transition. Grid resolution and numerical scheme which can directly resolve the large-scale coherent structures are required for LES to properly simulate the physics of the bypass transition. Much finer grid resolution than that required to resolve the large-scale coherent structures is required to resolve the small structures at the latter part of the transitional zone. However, when the large-scale coherent structures are adequately resolved, the underresolution of the small structures has little influences to the results.

Robust Aerodynamic Airfoil Design Optimization against Wind Variations for Mars Exploratory Airplane

Abstract: Robust aerodynamic airfoil design optimizations of Mars exploratory airplane against wind variations have been carried out by using DFMOSS coupled with the CFD simulation. The present robust optimizations successfully found the airfoil designs with robust aerodynamic performances against wind variations. Obtained airfoil design information about the optimality and the robustness of aerodynamic performances indicated that an airfoil with smaller camber can improve the robustness in lift to drag ratio against the variation of flight Mach number, and an airfoil with larger curvature near the shock wave location can improve the robustness in pitching moment against the variation of flight Mach number.

Flow Field Analysis of Jet Impinging on an Inclined Flat Plate at High Plate Angles

Abstract: Flow fields of the supersonic jets impinging on an inclined flat plate at high plate-angles are experimentally investigated using surface pressure measurement with pressure sensitive paint and Schlieren flow visualization. While Type I flow type is dominant at high plate angles, the present research found a new flow type “TYPE I with bubble” at plate angle between 60 and 80 degrees. The flow classification according to L/Ls’ and plate angle indicated that the constant x/L’s curve doesn’t represent the boundary of Type I and Type II anymore at high plate angles between 60 and 90 probably because Type II flows at low plate angles and high plate angles is different phenomena. This study also indicates that the curve dividing Type I and Type I with bubble regions is same as the curve dividing Type II and Type II with bubble regions.

Numerical Approach to the Analysis of Internal Pressure of M–V Rocket Fairing

Abstract: Navier-Stokes simulations are performed for time variation of the pressure inside the nose-fairing of M–V rocket. As the altitude increases during the launch of the rocket, flight Mach number and ambient conditions change with time. The numerical method to treat such an unsteady phenomenon is described in detail. The external pressure around the venting hole rapidly drops due to the passage of the local shock-wave over the rocket surface at transonic speed. The pressure inside the nose-fairing changes, because the external pressure is weakened through the venting hole. The effect of the time-accuracy is investigated.