Combustion issues and approaches for chemical microthrusters

In the present paper, results from the development of a meso/micro scale liquid-propellant thruster are reported along with a discussion of the usage of microthrusters in small spacecraft and the effects of downsizing on combustion performance. In particular, combustion of liquid nitromethane in a thruster combustion chamber with a volume of 108 mm³ and diameter of 5 mm was experimentally investigated. The meso-scale combustor utilized a vortex combustion concept, in which monopropellants were injected tangentially from the front end of the cylindrical combustor and combustion products exited the chamber tangentially at the other end. Although combustion was achievable at pressures as low as 150 psig, pressures as high as 350 psig were required for complete combustion. Combustion at atmospheric pressure was achieved only with the addition of small quantities of oxygen. To complement the experiments, a comprehensive numerical analysis was developed to study the combustion of liquid monopropellant in a small-volume vortex chamber based on a two-phase flow analysis using the level-set approach. The model allows for a detailed investigation of the liquid-film motion and gas-phase flow development. Combustion chambers and nozzles were fabricated from ceramics (alumina) because of the high flame temperatures required to sustain gas-phase reactions. Operation of thrusters on hydrogen-air-oxygen mixtures at a combustion chamber pressure of 40 psig was demonstrated to produce characteristic velocity efficiencies of approximately 90%.