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The mechanism causing high-frequency combustion instability was presumed to be the progression of combustion at an off-design location; that is, in the vicinity of the faceplate. Experiments were carried out to prove this. In the experiments, water was used to simulate the propellant and the increase in pressure caused by off-design combustion was simulated by injecting nitrogen gas. Pressure increase and decay were measured to within approximately 1 ms, and the maximum local pressure was approximately twice the pressure prior to increasing the pressure. To estimate the pressure change in actual engines, a simplified simulation calculation model was constructed. After calibration using the experimental results, the time and amplitude of pressure change were on the same order of magnitude of those of the engines. The present study results show that the off-design combustion model can be a cause of the high-frequency combustion instability experienced when using liquid rocket engines. [ABSTRACT FROM AUTHOR]

A newly constructed model calculates the starting position of the pseudo-shock, or shock train, in the divergent ducts. The model is based on the conservation laws of mass, energy and momentum of the inflow and outflow into/from the pseudo-shock region, and further uses a sonic condition in the core flow of the pseudo-shock. The model adopts an empirical growth rate of the low-velocity region, and the divergent angle of the nozzles becomes a design parameter. The lengths of the pseudo-shocks calculated show reasonable agreement with the experimental results. Pressure distribution with quick increase near the starting position is simulated. The supersonic core flow cross-section becomes smaller than that at the starting position of the pseudo-shock. The pressure at the sonic point rises to 0.7 of that at the duct exit. The length of the quick pressure-increase region becomes longer as the Mach number increases at the starting position. [ABSTRACT FROM AUTHOR]

Numerical simulation of aerodynamics around the spaceplane was conducted in the equivalent condition to the transonic wind tunnel tests at ISAS/JAXA. Numerical results reproduced experimental results, and they are useful to discuss the experimental results. Two configurations, namely, the experimental model and the original model were investigated. In all Mach numbers, drag coefficient of the experimental model was larger than that of the original model. Especially the wing and the tail made large drag. The different cross section of the wing and the tail caused the larger drag in the experimental model. [ABSTRACT FROM AUTHOR]

This paper describes the characteristics of methane-fueled rocket engines and compares these characteristics with those of hydrogen-fueled engines in terms of both expander and coolant-bleed cycles. Methane vaporizes in a cooling jacket under low-pressure operating conditions, whereas it liquefies in the turbine in the coolant-bleed cycle. The thermodynamic property of methane limits the operating range of methane-fueled engines. When the coolant-bleed cycle is used, the specific impulse degradation of methane-fueled engines becomes larger compared to that of hydrogen-fueled engines. This is due to methane having a lower specific heat and temperature after regenerative cooling. Even though the heat absorbing ability of methane is much lower than that of hydrogen, methane-fueled engines can operate with higher chamber pressures using the expander cycle. This is due to the larger density and the higher temperature after the regenerative cooling of liquid methane. Throttling of the methane-fueled engines does not have a great impact on the pump exit pressure in the expander cycle, whereas it increases the bleed ratio and degrades the specific impulse in the coolant bleed cycle of methane-fueled engines. [ABSTRACT FROM AUTHOR]

The laminar to turbulent transition in a pipe flow is studied from the viewpoint of momentum conservation. The transition is presumed to happen downstream of the laminar flow development. The inflow conditions to the transition region are presumed not to change at the moment of transition. The exit pressure is presumed not to change either. In the present model, the turbulent transition is a function of the ratio of the pipe length to its diameter, and becomes larger as the pipe length ratio increases. The natural transition Reynolds number calculated shows reasonable agreement with previous experimental results. The critical transition Reynolds number calculated is 1,752, and is close to the critical number of 1,760 measured previously. The distribution of difference in pressure drop between the laminar and turbulent flows corresponds to the reported location of the puffs and slugs in the Reynolds number under forced transition. [ABSTRACT FROM AUTHOR]

The effects of radiation heat transfer in rocket engine combustion chambers are studied analytically. The fuel is hydrogen, methane or ethanol, and the oxidizer is oxygen. Radiative heat fluxes are estimated using empirical equations, and convective heat flux is estimated using flux on a flat plate, with modification of circumferential length in convergence to, or divergence from, the throat. The calculated total heat flux including radiation and convection showed reasonable agreement with the flux measured experimentally. The ratio of radiative heat flux to total flux is increased up to 30% in the cylindrical section, whereas it is less than 10% at the throat. The effect of radiation on the total amount of heat transferred to the chamber is remarkable when increasing the length of the cylindrical section and diameter, respectively. It is also made clear that the conventional estimation method of heat flux based on the pipe flow model can estimate larger heat flux in small chambers, in spite of ignoring the effects of radiation. [ABSTRACT FROM AUTHOR]

The prediction method for heat flux in the pseudo-shock region is improved based on data measurements taken during ramjet-mode experiments. Combustion experiments using a ramjet-mode combustor are conducted to investigate heat flux in the pseudo-shock region. The heat flux is measured in the region with and without combustion. The heat transfer coefficient is modified using the Stanton number and specific heat under combustion. Using modified coefficients, the modified heat transfer coefficient ratios show the same relation to the pressure ratios as those without combustion. The ratio of heat transfer coefficients is calculated using a single equation. The calculation method can predict the experimental results of heat flux in the pseudo-shock region with combustion in the divergent duct. [ABSTRACT FROM AUTHOR]