The water spray nozzle is one of the major part of the flash evaporating system. In the present work, experiments and analysis of the flow rate distribution onto the heating surface from the Swirl type nozzle were carried out under low pressure circumstances. The results show  the spray flow rate distribution under the low pressure circumstances has much narrower band of the distribution and much higher peak value compared with that at under the atmospheric pressure circumstances,  the inviscid analysis can describe the peak value of the spray flow rate distribution under low pressure circumstances, but cannot describe the profile of the flow rate distribution, whereas the viscous analysis can describe the profile of the flow rate distribution, too.
This paper describes the water flash evaporation under low pressure circumstances, which considers the applicability to the practical use for spacecrafts of the data obtained in ref. 1). In the present work, experiments and analysis were carried out to entimate the performance of the water flash evaporating in the unsteady state with the fin type heating surface to be used for spacecrafts, which leads to the following results, (1) the water evaporating condition on the heating surface obtained under the steady state can be regarded to be realized instantaniously, (2) the fin-type heating surface with large angle of water impinging has heat rejection capability corresponding to that obtained with horizontal smooth fiat plate, (3) the plump portion in flash evaporator should be avoided to inhibit icing in a pressure control type flash evaporator.
To clarify the action of the rudder of HSTs, a hypersonic concave flow separation was studied. In hypersonic flows the rudder can not act efficiently because of the boundary layer separation, so that studies of that are needed. In previous studies, the hypersonic concave flow separation was studied to clarify the mechanisms of the separation region. But, the relation between the separation point location and the rudder angle was not cleared. Therefore, experimental and analytical studies were carried out to clarify this relation. The first, the flow visualization experiment by schlieren method was carried out. Then the transfer of the separation point due to the variation of the rudder angle was observed. Next, the separation point transfer in the boundary layer was calculated analytically by potential flow theory. The analytical result agreed well with the experimental observation. Consequently, the correlation between the separation point transfer and the rudder angle variation was confirmed.
To clarify the action of the rudder of HSTs, a hypersonic convex flow separation was studied. In a hypersonic flow the rudder can not act efficiently because of the boundary layer separation, so that studies of that are needed. In previous studies, the hypersonic convex flow separation was not studied because the surface pressure gradient of the convex corner has a minus value. But, it is not realistic that the hypersonic stream flows around the convex corner with large rudder angles without separation. Then, experimental and analytical studies were carried out to clarify this separation phenomenon. The first, the flow visualization by schlieren method was carried out to clarify the separation point and to observe this transfer due to the variation of the rudder angle. Next, pitot pressures in this region were measured to confirm the separation on the convex corner. Finally, the separation point transfer was calculated analytically by characteristics method and potential flow theory. In these results, the hypersonic convex flow separation was clarified.
Spatially developing compressible square jet was simulated by numerically solving the three-dimensional Navier-Stokes equations. A fourth-order non-oscillatory scheme and Roe's approximate Riemann solver are used in space discretization, and the third-order Runge-Kutta scheme is utilized as time integration. This code has an advantage that it can be easily extended to any order of accuracy in space and time. In the case of subsonic jet, two types of vortex ring behavior were observed. One vortex expands outward from the centerline, while the other slightly contracts inward. Vortex pairing was not observed unlike the round jet, and a symmetric vortex ring was broken into several longitudinal vortices in the downstream. In the case of supersonic jet, vortex ring was asymmetric even in the roll-up, and no more vortex pairing was observed than in the subsonic case. In both subsonic and supersonic jets, longitudinal vortices in the downstream slow down the growth rates of shear layer and three-dimensionalization. The growth of shear layer in square jet is greater than that of round jet, although the core of square jet is maintained up to further downstream.
The reusable regeneratively cooled thrust engine is a candidate for the orbital maneuvering system (OMS) engine of H-II orbiting plane. The high performance thrust chamber, however, coupled with extended reuse requirements impose difficult cooling requirements particularly at throat region. To meet the cooling and life enhance requirements, the inner wall of the thrust chamber are fabricated with thermal barrier, such as ceramic coating. However, the main drawback to the use of these ceramic coatings are the delamination between the ceramic side and the metal one during repeated thermal heat load. One promising method of improving adhesion of the ceramic coating to the metal wall in the thrust chamber is to apply the functionally graded materials (FGM) composed of ZrO2 and nickel. The FGM chamber consists of different materials components of ZrO2 and Ni with sequential variation of characteristics from one side of ZrO2 to the other side of nickel. In this study, the long firing tests and the durability tests of the regeneratively cooled 1, 200N thrust engine composed of ZrO2/Ni FGM were conducted with nitrogen tetroxide (NTO)/monomethyl hydrazine (MMH) propellant. The film cooling fraction were reduced to 0% to obtain high performance. An optical multichannel analyzer was used to acquire the exhaust plume spectral data in the firing test. A total of 260 firing tests were performed to evaluate the performance of the engine and to obtain thermal data, and also to inspect the damage of the thermal barrier coatings by using the Reprica method.