Akira Kakami
(Kyushu Institute of Technology, Japan)

The current paper describes experimental results for a hydroxyl ammonium nitrate (HAN)-based monopropellant thruster using arc-discharge assisted combustion and experimental results for a designed thruster. The HAN-based propellant is a liquid mixture, comprised of HAN, methanol, ammonium nitrate and water, and is neither toxic nor reactive unlike a conventional monopropellant hydrazine. This propellant suppresses a discontinuous increase in burning rate, which conventional HAN-based propellants show at approximately 8 atm back pressure. The steep augmentation in burning rate causes sudden expansion in thrust chamber pressures and resultantly damages thrusters. With the designed thruster, the combustion was initiated and stabilized by arc discharge to avoid the problems that conventional Iridium-based particulate catalyst is relatively fragile especially at elevated temperatures. Experiments with a 1-N class prototype thruster using nitrogen as both a pressurant and an arcjet working fluid yielded a thrust chamber pressure of 0.4 MPa and a C* efficiency is approximately equal to 86 % at a propellant mass flow rate of 0.70 g/s with a discharge power of 1700 W

Nobuyoshi Suzuki
(Tokyo Metropolitan University, Japan)

We have developed a propulsion system for 50kg-class microsatellites since 2004 in order to broaden and deepen the use of microsatellites. We considered chemical propulsion system is the most suitable for microsatellites because of its high thrust density and low power consumption. However, the conventional propulsion system for satellites is based on hydrazine and it is difficult to handle the propellant due to its toxicity, so that microsatellites developed in universities and companies cannot install such a propulsion system. Furthermore, the conventional propulsion system is developed as a unique or identifiable product by proper specialists with quite high reliability, and it is awfully expensive for a microsatellite project. Here, we are developing a propulsion system with intermediate concentration hydrogen peroxide based on the policies of SAFTY FIRST and EFFECTIVE COTS. We completed a suitable mono-propellant propulsion system for microsatellites, and conducted its captive test in vacuum to obtain its thrust, specific impulse, and impulse bit. In this paper, we present results of the captive test and expand the propulsion to be a bi-propellant system.

Domenico Simone
(University of Rome "La Sapienza", Italy)

Aim of this work is to develop a study of silanes combustion in carbon dioxide and to provide an analysis suited to investigate their potential use for propulsion applications in the Martian atmosphere, where 95% of the atmospheric gas is made up of CO2. The great amount of silica sand in Martian atmosphere leads to an important consequence: sand entering the engine could damage mechanical parts installed such as compressors or turbines; for this reason it is impossible to consider turbojets as a choice. On the other hand, it could be possible to use silica-derived silanes in ramjets or scramjets. Equilibrium composition of the combustion products of silanes, from monosilane up to pentasilane were calculated. Ideal ramjet and scramjet performance (specific thrust and specific impulse) were evaluated along constant dynamic pressure trajectories in the Mars atmosphere and compared to that with CH4/Air and H2/Air mixtures. High specific thrust is obtained as the equivalence ratio is increased; the Isp trend is the reverse, but still very appealing when weighted with the bulk density of silanes.

Daisuke Goto
(JAXA, Japan)

The Monopropellant Hydrazine Thruster is widely used in many satellites and explorers propulsion system and launch vehicle RCS for its simple and low cost features. The specific impulse (ISP) is approximately 220sec and the thrust range is mainly 1-100N. The 1N Thruster is the work horse in the monopropellant thruster family, because it is suitable for attitude control or small orbit control of the small-medium size satellite. The current Japanese 1N thruster becomes difficult to fulfill the mission request with these changes of satellite mass and lifetime. We thought that the current 1N thruster cannot produce as much as 70,000 Nsec cumulative impulse safely. In 2010, the 1N thruster EM was manufactured and tested. The purpose of the EM phase tests is preparing the PM (prototype model) design determination and the Qualification Test for the main stream thruster of Japanese future satellite programs. We prepared six thrusters with three types of injectors and two type of catalysts. Finally, three thrusters completed the whole scheduled test sequence and all of them achieved over 200,000 Nsec cumulative impulse and 1,100,000 cumulative pulses. It is the top level value in the world. We are going to choose one design and prepare Qualification Test.