A current trend, in recent decades, in spacecraft engineering is the miniaturization of satellites and their subsystems. One of these subsystems is for propulsion, which is important for orienting satellites and performing orbital manoeuvres. Currently one such micropropulsion system, being developed at the Space Engineering department at the Delft University of Technology, is called the ``Vaporizing Liquid Microthruster". These thrusters are manufactured using Deep Reactive Ion Etching on silicon wafers. They operate by resistively heating the propellant, vaporizing and expanding it throughout the nozzle to create thrust. Currently, several experimental investigations have been conducted to formulate models that can predict the performance of these microthrusters, on a wide range of operating conditions, without the use of expensive testing. These models can then be validated comparing them to experimental data. Thrust tests were performed with nitrogen, for three different thrusters in ambient, cold conditions, for operating pressures of 2-5bars. From these tests, the discharge coefficient was predicted within 10\% of the experimental values. So the model's prediction of the discharge coefficient was valid. It was also immediately clear that the thruster was not designed for ambient conditions due its low specific impulse values of 7.6-22.7s, resulting in incomplete flow expansion in the nozzle. The current 1st generation thruster interface used at the department is limited to 150ºC due to its teflon material. Therefore, an alternate design using a printable circuit board was investigated to reach at least 600K, where the thruster would be glued to it using a thermally resistant, ceramic based glue (Thermeez 7020). Unfortunately, the design was limited to 150ºC due to the low adhesive quality of the glue, which meant that the thruster inlet could not be sealed off and a less suitable replacement glue had to be used. Heated thrust tests were performed for two thrusters up to around 150ºC in ambient conditions. For this increase in temperature , throat Reynolds numbers decreased from 2915 to 779, while the discharge coefficient increased from 0.71 to near 1. However, a less suitable mass flow sensor had to be used, so the validation of the model based on the hot thrust tests was inconclusive for discharge coefficient and the Isp efficiency. Overall, the goals for this experimental research study were partially reached. Thrust tests were conducted, using nitrogen as propellant, for cold conditions and up to 150ºC. The experiments could not substantially validate the prediction of the Isp efficiency, however the model is accurate to within 10% of the experimental values for the discharge coefficient. Experimental data from literature also could not validate the Isp efficiencies of the model, unless the fitting constant of the thrust coefficient relation was altered. It is recommended to focus on experimenting with the use of water as a propellant and make minor improvements to the thruster design, to obtain accurate performance data at higher temperatures.