Kenichi Takahashi

Combustion Characteristics of Boron Particles in The Secondary Combustor of Ducted Rockets

Ducted rockets are one kind of ramjet engines which are air-breathing propulsion system suited for supersonic flight. Fuel components used for ducted rockets are fuel-rich propellant. To obtain high specific impulse, it is desired that the fuel materials evolve high combustion heat per unit mass in the air. Boron is the most suitable material of the fuel of ducted rockets in terms of its high combustion heat. Obtaining the combustion characteristics of boron particles is the key to improve boron particles combustion. In this study, combustion times of single particles and flame areas around the particles were measured by direct photography and spectroscopic analysis, and flame structure of burning particles was observed by schlieren photography. It was obtained that boron particles burn in two stages separately and the combustion time decrease with increasing temperature. The flame areas are composed with boron dioxide.

Behavior of Charcoal in Black Powder on Burning Surface

Nomenclature Cp = specific heat of propellant Pg = pressure Qs = heat of reaction at the surface R = gas constant Tf = flame temperature Ts = surface temperature T0 = initial temperature ug = gas velocity δ = reaction length λg = heat conductivity in the gas phase ρg = gas density ρp = density of propellant τc = chemical ignition delay time τig = ignition delay time τp = physical ignition delay time Φs+ = temperature gradient near the burning surface in the gas phase

Combustion characteristic of solid propellants used H2O

It has been known by past explorations that metal and water (H2O) exists on the moon and Mars. And H2O is expected as an oxidizer of the in-situ solid propellant. To use H2O as oxidizer, the high temperature metal which reacts readily with H2O is needed. Terefore, we selected Magnesium (Mg)/Teflon/PEO/H2O as sample propellant. In this study, we obtain the burning rate of the sample propellant, the combustion efficiency of C*, and the hydorogen gas generation ratio . The maximum combustion efficiency of C* is 94% obtained at the concentration of H2O equal 15 parts. The concentration of generated H2 gas increases with increasing the concentration of H2O , and the H2 gas generation ratio is over 82% all of the concentration of H2O.

Burning rate characteristics of AN composite propellant with addition of Al particles

In this study, we evaluated effect of Aluminum (Al) particles for combustion mechanism of Ammonium Nitrate (AN) composite propellant. In the result of the combustion experiments, the burning rate of AN composite propellants increased with increasing concentration of Al particles. Aluminum particles do not ignite and combust at solid phase including burning surface of AN composite propellants from the result of the experiment of using Differential Scanning Calorimeter (DSC) and temperature history. Aluminum particles ignite and combust in the gas phase of AN composite propellants from the result of temperature history. Temperature gradients of AN composite propellants increase with increasing concentration of Al particles. Reaction heat in gas phase increase with increasing concentration of Al particles from analysis result of temperature history. Therefore, temperature gradients increase with increasing heat of reaction in gas phase.

Reaction of Aluminum in Gas Phase Near Burning Surface of AP Composite Propellants

To design a rocket motor, thrust and burning rate of propellant are important factors. The combustion efficiency of a rocket motor is determined by combustion of aluminum (Al). It is needed to ignite aluminum particles near the burning surface of propellant. The temperature in the reaction zone increases with increasing the distance from burning surface and reaches high temperature about 2000K, therefore it is possible to ignite aluminum particles near the burning surface in the reaction zone. The reaction zone length at burning surface of propellant is less than 1mm at 0.1MPa. It is difficult to observe the reaction zone with a video camera and to evaluate the gas reaction. We measured temperature histories in this reaction zone with a small thermo-couple. The temperature fluctuation occurred near the burning surface. The temperature fluctuation appeared by combustion of solid particles like Al. Heat of reaction in the gas phase is proportional to the partial derivative of the second order with respect to the distance (heat flow) at steady state. Aluminum ignited in the reaction zone, and burned in the luminous flame zone. Hydrogen and carbon reacted easily in the reaction zone, and generated hydrogen oxide and carbon dioxide. In the luminous flame zone there are no oxygen, but only H2O and CO2, aluminum burned in high temperature gases of carbon dioxide and hydrogen oxide. Aluminum for propellant is used small particles, because they first melt and vaporize and mix with the surrounding gas. They need a time to combust completely. The combustion efficiency of Al increases with increasing pressure in a rocket motor, and also increases with increasing stay time in the motor.