With a growing trend in the miniaturisation of satellites, there is an increasing need to develop micro-propulsion systems for these satellites. Since scaling down conventional propulsion systems is challenging and not always possible, new concepts need to be developed. These concepts, although often based on already known systems and principles, require significant modifications to make them meet the requirements of miniaturized propulsion. One such concept is the micro-resistojet, using an electrical resistance to increase the propellant temperature. At Delft University of Technology, two thrusters based on this concept were developed: the Vaporized Liquid Micro-resistojet (VLM) and the Low-Pressure Micro-resistojet (LPM), which were specifically designed for being demonstrated on-board a PocketQube satellite. To determine the operating regime of the thrusters for this demonstration, there is a need to develop simplified analytical models that accurately predict their performance without significant computational expenses. Although there were previous attempts to model these thrusters, they did not provide a complete representation of their performance. For the VLM thruster, the focus of the model presented in this paper was on coupling the heating chamber and the nozzle, to obtain a more accurate value for the mass flow rate through the thruster. The heating chamber section was discretized into finite one-dimensional cells and convective heat transfer equations were used to model parameters such as density, pressure, wall temperature and heat transfer coefficient. The nozzle was modelled based on ideal rocket theory corrected with adequate loss factors. The mass flow rate was calculated iteratively by coupling the two sections until it reached convergence. For the LPM thruster, the focus was on including an accommodation coefficient to account for heat transfer efficiency between thruster walls and propellant. Rarefied gas dynamics equations were used to calculate performance parameters due to the low-pressure conditions within the thruster. The models proved to produce realistic results when compared to available numerical and experimental values, although still with some limitations in modelling heat transfer, which could not be fully overcome yet due to the lack of available data for validation. Optimal operating points were determined for both thrusters by maximizing an objective function based on performance parameters such as thrust-to-power ratio, specific impulse, and mass flow rate. Constraints included thrust, power, and temperature requirements, which led to different optimal points for the thrusters under varying operational conditions.