It is well known that the flow around a sphere becomes steady and non-axisymmetric when the freestream Reynolds number (based on the diameter) becomes greater than 210. However such steady asymmetric patterns cannot be found in the case of the two-dimensional flow around a circular cylinder. In this paper, the flow around a three-dimensional low aspect ratio cylinder is calculated, and it is found that the flow around a low aspect ratio cylinder also becomes steady and asymmetric under a certain range of freestream Reynolds numbers. The two-dimensional and three-dimensional incompressible Navier-Stokes equations are numerically solved for flows around these bodies, and the transition mode from steady symmetric flow to asymmetric flow is investigated by global stability analysis. These results show that in the two-dimensional case, the asymmetric non-oscillatory mode decays faster than the oscillatory modes. In the case of the asymmetric non-oscillatory mode of the two-dimensional flow, the region of its influence extends far downstream and reaches the outer boundary of the computational domain, while in the three-dimensional non-oscillatory mode the region of influence is observed only in the vicinity of the body. The similarities of the flow around a sphere and low aspect ratio cylinder are discussed from a viewpoint of the configuration of these transition modes.
The effects of various radical species, supplied from a plasma torch ignitor upon initiating hydrogen-air chemical reaction in a two-dimensional supersonic shear flow have been investigated by a new numerical approach based on the molecular gas dynamics. Direct Simulation Monte Carlo (DSMC) method has been employed to derive the microscopic reaction phenomena from the Boltzmann equation with a reactive inelastic molecular collision model (IE Model). The plasma species and injection mode into air stream have been varied. The result shows that fuel hydrogen initial combustion can be significantly enhanced by plasma injection of atomic oxygen supplying in diffusive mode.
The 1st report clarified the superiority of the atomic oxygen injected from a plasma torch with the diffusive mode operation for initiating supersonic combustion. In this paper the features of combustion enhancement by various combinations of plasma mole flux and heat flux are investigated. The effectiveness of each combination is evaluated by H2O production efficiency. A molecular dynamical numerical simulation technique with the reactional molecular collision model is applied for the analysis of multi-species reaction system. The results of parametric study show that the effect of atomic oxygen plasma is attributed to the mole flux, whilst the strong dependence on the heat flux is observed for the atomic hydrogen plasma.
Wind tunnel tests were conducted to investigate aerodynamic characteristics of a Supersonic Transport (SST) model with leading-edge and trailing-edge flaps. Force and surface pressure measurements were performed for the SST model either with leading-edge flaps or with trailing-edge flaps deflected and for the model with all the flaps deflected. The lift-to-drag ratio (L/D) can be improved by the leading-edge flap deflection. When the trailing-edge flap is deflected modestly, the L/D is also improved. According to the measured results, the best improvement of the L/D is attained when the leading-edge and trailing-edge flaps are deflected at the same time. This paper discusses how the combination of leading-edge and trailing-edge flaps improves the wing performance.
There are many design methods for the multiloop control system, but a method to exactly evaluate the stability margin for multiloop control systems has not been established yet. It is very important for the designers to exactly know the stability margin of the flight control system, because it directly comes to bear on flight safety. Previous research has proposed the “-1/ξ locus” analysis method for the exact evaluation of the stability margin of multiloop flight control systems. -1/ξ is an open loop transfer function of multiloop control systems that is considered as an extension of that for single-loop control systems. In this paper, the relation between the response of the closed loop and the stability margin is considered by using the -1/ξ locus. Lateral-directional flight control system was taken for instance, important points to pay attention to use the -1/ξ locus method for designing the flight control system are presented.
In this paper we propose a new method known as “the line detection method” in order to detect small pieces of LEO debris. Directions of the line created by small LEO debris on CCD images are supposed and values of pixels along the direction are accumulated to improve the signal-to-noise ratio. This method is able to detect 30 to 40 times darker LEO debris than usual methods. We tested this method by using the 35cm telescope and the back-illuminated CCD camera at the Mt. Nyukasa Astronomical Observatory. One small LEO debris whose radar cross section was 0.0047 square meter was detected. The observation facilities at the Bisei Spaceguard Center, the 1m telescope and the back-illuminated wide field camera, are expected to detect LEO debris of a few centimeter with this method. The line detection method will be used to probe the environment of small LEO debris and contribute to solve the space debris problem.
A method to exactly evaluate the stability margin for multiloop control systems has not been established yet. In this paper, it is shown that the results of the classical method of stability margin which is used by designers of aircraft manufacturers is conservative. Previous research has proposed the “-1/ξ locus” analysis method for the evaluation of the stability margin. Using this -1/ξ locus method, more exact evaluation of the stability margin of multiloop flight control systems is presented.
Recently, an airplane cruising at near-sonic regime is watched with keen interest. The Sonic-Cruiser, of which the Boeing Company has examined and challenged the development, is the most remarkable case. In this paper, motivated by this trend, aerodynamic performance optimization for an airplane cruising at near-sonic regime is discussed based on CFD simulations. NAL’s experimental supersonic airplane, called NEXST-1, was employed as the baseline model for optimization. Aerodynamic performance was evaluated by solving the Euler equations with the unstructured grid method. It was confirmed that the performance Euler simulation predicted was qualitatively correct. By the evaluation to select a baseline model for optimization, NEXST-1 was accepted as a candidate of sonic plane because of the existence of drag bucket at near-sonic regime. In the optimization, Genetic Algorithm was used with Euler simulations. The objective was to reduce drag keeping lift constant, at the flying speed of Mach 0.98. The optimized result showed L/D improvement not only for near-sonic regime but also for transonic regime. The mechanism of design to reduce drag force was found through the analysis and comparison of the geometries and aerodynamic phenomena about the baseline model and the optimized one.
Wind tunnel experiments have been made to investigate behaviors of laminar separation bubble formed on a NACA 631-012 airfoil at chord Reynolds numbers of Re=1.9×105–3.8×105. Results indicated that a short bubble is formed when the angle of attack α is lower than 10° at Re=1.9×105. The short bubble burst occurs at α=10°. It has been reported that the reattachment point of the short bubble moves upstream as α is increased at higher Reynolds numbers such as at Re=6×106. In this study, however, it is shown that the reattachment point of the short bubble moves downstream as α is increased just before the bubble burst.
This paper discusses an aircraft conceptual design sizing method that is lectured during university undergraduate education. A new two step sizing method is proposed here both by improving sizing methods currently used and by incorporating a spreadsheet method. This new method enables students to design an aircraft by understanding how the main design parameters such as a maximum take-off weight affect the design results.