For space plane with air-breathing engines, equipping Single Expansion Ramp Nozzle on its aft-body is beneficial for thrust augmentation. The cowl of this nozzle is truncated (thus, termed as an external nozzle), in order to reduce its weight, friction drag, and cooling requirement, and also to attain pressure recovery (so-called altitude compensation) on the ramp wall by the impingement of pressure waves generated by the interference between external nozzle flow and ambient flow on ramp wall surface, in case of over-expanded condition. The present study is to evaluate the effects of ambient flow on the external nozzle flow field, as little research has been conducted to elucidate the effects. Room temperature nitrogen gas was injected to simulate air-breathing engine exhaust in cases without and with ambient flow at a supersonic semi-free jet wind tunnel. A straight expansion ramp was employed as the external nozzle. Pressure distributions on the ramp wall were measured and results without and with ambient flow were compared. Comparison showed that presence of ambient flow reduced the magnitude of ramp wall pressure induced by the impingement of pressure waves, and thus, the strength of pressure waves generated by the interference. Flow characterization by pressure wave tracking showed that spatial variation in ambient flow pressure caused this weakening of the incident pressure wave upon the ramp surface. Thus, altitude compensation effect in the over-expanded condition became smaller than that without ambient flow.
This paper reports on the conceptual design of a three-stage launch vehicle (LV) with a clustered hybrid rocket engine (HRE) through multi-disciplinary design optimization. This LV is a space transportation concept that can deliver micro-satellites to sun-synchronous orbits (SSOs). To design a high-performance LV with HRE, the optimum size of each component, such as an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank, and a nozzle, should be determined. In this study, paraffin (FT-0070) is used as a propellant for the HRE, and three cases are compared: In the first case, HREs are optimized for each stage. In the second case, HREs are optimized together for the first and second stages but separately for the third stage. In the third case, HREs are optimized together for each stage. The optimization results show that the performance of the design case that uses the same HREs in all stages is 40% reduced compared with the design case that uses optimized HREs for each stage.
Uniform suction or blowing from the wall is one of the methods to reduce the friction drag. The uniform suction improves the stability of a laminar boundary layer: the transition will be delayed and the overall friction drag will be reduced due to the extended laminar region. In contrast, the uniform blowing is known to reduce the drag in the fully-turbulent regime. Therefore, a combination of uniform suction and blowing is expected to be effective for flows involving transition, such as the flow around an airfoil, by delaying the transition near the trailing edge and by reducing the turbulent drag in the post-transition (i.e., turbulent) region. The objective of this study is to investigate the friction drag reduction effect of such a combined uniform suction and blowing. The Reynolds-Averaged Navier-Stokes simulation is used to deal with a spatially developing boundary layer on a flat plate at a practically high Reynolds number. As a result, the combined control is found to reduce the global skin friction coefficient by 44.1%, whereof the contribution of transition delay by the uniform suction is about 90%, and that of turbulent drag reduction by the uniform blowing is about 10%. It is also found that the position of the blowing region should better be located in the upstream side of the turbulent region because the drag reduction effect is sustained for a while even after the blowing is terminated.
The purpose of the research is to characterize the ionization process behind a shock wave with precursor photoionization in argon. In this study, the H-β line is observed by spectroscopic measurements using a hypersonic shock tube and the Stark broadening of the H- β line is evaluated to obtain the electron density behind a shock wave. As a result, the electron density is on the order of 1021 m-3 and tends to decrease with increasing the distance from the shock front. The decrease of the measured electron density is caused by the radiative energy loss from the test gas because the radiative energy loss corresponds to a decrease of the temperature or the density of the test gas, both of which can decrease the electron density. Previous study showed that photoionization occurred ahead of the shock wave, making use of the radiation energy emitted from the region behind the shock wave. The fact might be related to the radiative energy loss obtained in the present study. From the present study, it is found that the radiative energy loss is dominant behind the shock wave under the condition that precursor photoionization occurs ahead of the shock wave. In future, we should investigate the radiative transfer phenomena around the shock wave for further clarification.
For the next generation space launch vehicle, fully reusable spaceplane with rocket and ramjet combined propulsion system is proposed. That type of spaceplane can drastically reduce the cost per unit mass of the payload. One of the main technical challenges of the spaceplane is the propulsion system. Thus, many component tests have been carried out using hydrogen as fuel. On the other hands, in these days, some of the projects uses hydrocarbon(HC) fuel, because of the ease of treatment, and the high density Isp performance. Because of these backgrounds, pressurized high temperature HC fuel feeding device for the component tests of space propulsion system is required to simulate re-generative cooling. We started to develop the fuel heating device aiming to feed ethanol at supercritical condition for the component tests of the RBCC engine. The device is needed to heat up the fuel higher than 520 K and pressurize the fuel higher than 7 MPa. In the development process, we experienced some troubles, but we improved the composition of the fuel heating device, and most of the troubles were settled. We achieved the main goal of the development, to feed supercritical ethanol (7.5 MPa, 520 K at the exit of fuel heating device) to the test equipment of the dual-mode ramjet combustor. In this paper, we report the construction of the developed fuel heating device, the detail of the troubles and countermeasures, the result of the unit tests and the result of an actual use test.
Relationships between local flow topology (geometry) of vortical flow and pressure minimum feature derived from its flow kinematics are investigated in isotropic homogeneous turbulence, in a new topological point of view with respect to vortical flow symmetry in a swirl plane. Consistency of the pressure minimum plane and the swirl plane, and relationships between the pressure minimum and vortical flow symmetry in their development are also analyzed. The pressure minimum feature is specified by the λ2 definition and also the λˇ2 definition that is the integrated definition of the Δ, Q and λ2 definitions and specifies the pressure minimum in the swirl plane. The swirlity φ that represents the geometrical average of the intensity of swirling flow is applied to the statistical analysis of the flow topology and pressure minimum. It shows that the pressure minimum requires a certain flow symmetry, and that the development or decay of the pressure minimum is associated with that of the vortical flow symmetry especially in the state (process) of a vortex attaining or losing the pressure minimum feature. The vortices with high intensity of φ have the feature that the effect of vorticity components parallel to the swirl plane decreases in specifying the pressure minimum plane by the λ2 definition, and then this pressure minimum plane tends to approach the swirl plane.