In this paper, the expansion of xenon, argon, krypton, and neon gases through a Laval nozzle is studied experimentally and numerically. The pressurized gases are accelerated through the nozzle into a vacuum chamber in an attempt to simulate the operating conditions of a cold-gas thruster for attitude control of a micro-satellite. The gases are evaluated at several mass flow rates ranging between 0.178 mg/s and 3.568 mg/s. The Re numbers are low (8–256) and the estimated values of Kn number lie between 0.33 and 0.02 (transition and slip-flow regime). Direct Simulation Monte Carlo (DSMC) and continuum-based simulations with a no-slip boundary condition are performed. The DSMC and the experimental results show good agreement in the range Kn > 0.1, while the Navier-Stokes results describe the experimental data more accurately for Kn < 0.05. Comparison between the experimental and Navier-Stokes results shows high deviations at the lower mass flow rates and higher Kn numbers. A relation describing the deviation of the pressure drop through the nozzle as a function of Kn is obtained. For gases with small collision cross sections, the experimental pressure results deviate more strongly from the no-slip assumption. From the analysis of the developed function, it is possible to correct the pressure results for the studied gases, both in the slip-flow and transition regimes, with four gas-independent accommodation coefficients. The thrust delivered by the cold-gas thruster and the specific impulse is determined based on the numerical results. Furthermore, an increase of the thickness of the viscous boundary layer through the diffuser of the micronozzle is observed. This results in a shock-less decrease of the Mach number and the flow velocity, which penalizes thrust efficiency. The negative effect of the viscous boundary layer on thrust efficiency can be lowered through higher values of Re and a reduction of the diffuser length.

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Fuel mass is one of the main economical and technical restrictions while designing space propulsion systems. Given the high costs related to the transport of mass into space, the necessary fuel mass for accomplishment of the mission should be minimised. For an optimum "thrust/fuel consumption ratio" the gas exit velocity must be maximised. In this research this is achieved through the heating of the micro gas flow by an electrical arc inside the sub-sonic region of the propulsion system. The electrical arc induces a partial ionisation of the propellant gas. Because of the very low mass flow, the gap of the plasma channel has a width of just a few hundred microns. The electrical arc consists of electrons being accelerated through this small gap by the charged walls of the microchannel. The electrons move in a cross flow compared to the propellant gas. The computational results are validated with the experimental data. Through this investigation a very efficient form of electrodynamic heating-modelling is developed. The very good results show the quality of the present method and encourage further utilisation and development. For this reason this model will be used for the optimisation and the computational engineering pre-development of future thermo-electric propulsion systems.

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