Hiroto HABU

Nitrous Oxide (N2O) / ethanol propulsion system is distinguished as the liquid propulsion with non-toxic, user-friendly, and storable bipropellant. The current target of the present study is to build a quick-response and maneuverable main engine of a sounding-rocket like flying test bed which will be applied to the hypersonic air-breathing propulsion researches in the near future. The application to the spacecraft propulsion is also considered due to its compatibility in low-temperature operation environment. Two series of static firing tests were performed with 700 N class gas generator models so far. Current test results showed that valuable design data were collected and operational procedure was verified. Potential of application of composite materials to the combustion chamber was also examined from the chamber wall heat flux data obtained and the result of firing test using a thick SFRP chamber.

In this study, reactive B/Ti multilayer igniters were investigated for the noncontact ignition of a micro solid rocket array thruster in vacuum. When current is supplied to the B/Ti multilayer igniter, the chemical reaction: 2B + Ti -> TiB2 + 1320 cal/g occurs, and high temperature plasma is discharged to a distance of several millimeters or more. The B/Ti multilayer igniters with 3 sizes were fabricated, and tested in 6 configurations of solid propellant. Although one rocket with ignition charge was ignited successfully, the noncontact ignition of the solid propellant was not achieved.

Simulated solid propellant slurry containing lead sphere tracers is experimentally cast into a double circular cylinder container. During the casting, the temperature and the pressure environment has been mimicked to an actual composite solid propellant casting of solid rocket motors. X-rays are projected on to the slurry flow from two directions and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the slurry flow, which is invisible usually, has been estimated.

The static environment of the solid rocket motor chamber was simulated by the improved QPCB. This paper discussed about the combustion products collection method and the particle size distribution of Al/Alumina. Based on the combustion mechanism of the Al agglomeration particles, the shift of the aluminum/alumina particle size distribution in the combustion gas of the solid propellants depending on the traverse distance is estimated experimentally. Further, the correlation between the AP size distribution in the propellant and the particle size distribution of CCP was investigated.

As one of the experimental techniques to reduce the HCl concentration in the combustion gas of AP-Based composite propellants, we tried to replace aluminum as metal fuel by magnalium (Mg/Al alloy). The concentration of HCl in the combustion gas and the combustion characteristics were examined for AP-based propellants containing Magnalium. It was evaluated using TG experiments that the Mg/Al-based composite propellants appeared equivalent to the current aluminized propellants for the ignitability as a safety test. It was found that the HCl concentration in the combustion gas of 25%Mg/Al-AP propellants was over 90% lower than that of AP-based propellants without metals and over 40% low compared to Al-AP ones. Moreover, decreasing the particle size of Mg/Al alloy or increasing the apparent concentration of Mg in Mg/Al alloy decreased the HCl content in the combustion gas. Both burning rates and flame temperatures for Mg/Al-based propellants were found to be high compared to aluminized propellants. The pressure exponents of Mg/Al-based propellants and aluminized ones represented the nearly same values, n=0. 48, 0. 47, respectively, but the burning rates of Mg/Al-based propellants were about 50% higher than those of aluminized ones.