The flow field in a micro arcjet thruster using argon as propellant was analyzed numerically using the Direct Simulation Monte Carlo (DSMC) method. Subsonic inlet conditions were imposed to correctly simulate the actual flow inside the thruster. It was assumed that the input power was partitioned to propellant molecules to change their thermal velocities. Wall temperature was determined by the balance of conductive heat flux and radiative heat release. The primary purpose of the present work is to investigate the effects of the nozzle half-cone angle and input power on thruster performance characteristics. The results provide engineering direction to the micro arcjet thruster design.