The hydrodynamic instability of a deflagration wave propagating with finite Mach number has been investigated. We dealt with a wave front of deflagration as a surface of hydrodynamic discontinuity, and considered the pressure change through the wave front. We analyzed the instability of the deflagration wave with respect to infinitesimal fluctuations, and obtained the relation between growth rates of fluctuations and their wave numbers (dispersion relation). The dispersion relation is consistent with the Darrieus-Landau solution when the Mach number is sufficiently small. Increased value of the Mach number causes larger pressure differences, and causes high growth rates compared with those obtained by Darrieus and Landau. Therefore, we should consider the pressure change in the stability analysis of the deflagration wave propagating with considerably fast velocity.
A solution-adaptive grid-generation method based on an elliptic equation is developed in this paper. Application of a centered-noncentered (upwind) hybrid finite-difference method to first derivative terms with combination of the under-relaxation technique makes the method robust. Test problems using asymptotically steep functions of radically changing first and second derivatives show that the method can generate adaptive grids with excellent quality.
A numerical analysis for a modeled plasma flow in an MPD thruster has been done, including nonequilibrium ionization/recombination processes of the propellant gas. According to the existing studies on MPD thrusters, current density distribution in a discharge chamber is strongly influenced by the nonequilibrium ionization processes. In order to determine the ionization fraction at the inlet of chamber, backstream of plasma from the arc region is taken into account. In this study, the flowfield based on this nonequilibrium and two-fluid model is theoretically analyzed.
Flutter calculations were carried out on the results of fundamental wind-tunnel experiments on tip-fin configuration wings. In the experiments, mild flutter has been observed within a certain range of flow speed which was well below the violent flutter point. This mild flutter seems to be peculiar to the tip-fin configuration because its flutter mode couples with the wing torsion and the motion of the rudder on the tip-fin. Since the wing has the non-coplanar configuration with a partial control surface, the non-planar doublet-point method was utilized to calculate the subsonic unsteady aerodynamic forces of the control surface deflection as well as of the wing deformation. The results of the analysis show excellent agreement with those of the wind-tunnel experiments. It clearly explains the mild flutter due to the rudder motion on a tip-fin.
A panel code using subpanel technique, based on the internal Dirichlet formulation has been developed. In this study, only the case where the internal region has no flow is considered. Every panel is devided into 9 flat subpanels. Each subpanel has 4 subgrid points which are located on the biquadratic curve defined between two neibouring grid points, which effects more accurate expression of the geometry. Each subpanel is assumed to have uniform density of singularity, the value of which is expressed at the center of the subpanel using two-dimensional quadratic interpolation functions defined with the geometry of the surrounding panels. The calculated results are compared with those obtained by the panel method based on the external Neumann formulation using the same subpanel technique. Both results show good agreements in external flow examples, but in the calculation of the flow suffering strong interaction such as a flow through a duct, the panel code of the internal Dirichlet formulation gives much better results than that of the external Neumann formulation. This suggests that the internal Dirichlet panel method could be applied more effectively to the problems, such as wings with wing-tip clearance or wing-tip vortex sheet, which have severe interactions with the nearby singularity surfaces.
We have been working on the boundary layer separation, in particular its control by vortex excitation. In this paper, acoustic forcing is applied for the separated flow over a NACA0015 airfoil to control the boundary layer separation at high angle of attack. The results show that the separation bubble can be successfully suppressed even at angle of attack 19° through exciting strong vortices acoustically. The importance of the matching between the scale of the excited vortices and the separation bubble height, which has been suggested in our previous experiment on the acoustic control of the leading-edge separation of a fiat-plate airfoil, is also supported by the present experiment.