The Dutch/New Zealand based aerospace company Dawn aerospace has developed a 0.5N green bi-propellant propulsion system for use in CubeSat applications called the PM200. This thruster uses gaseous nitrous oxide and propylene as propellants. Currently, the burn time of this thruster is limited to 10 seconds after which cool down of the thruster is required. This burn time limit was set quite arbitrarily as no thermal analysis nor any thermal experimentation was performed on the thruster. In this study the thermal limits of the PM200 and similar thrusters was explored. In doing so designs with two particular cooling methods were analysed in detail: designs with Regenerative cooling and designs with Radiation cooling. The overall goal of the project was to assess the thermal performance of the thruster and to see which of the two cooling methods would be most beneficial to implement. In order to perform the thermal analysis a transient numerical model was created based on the finite volume method in combination with ideal rocket theory and several semi-empirical relationships. The model is able to calculate the temperature distribution and heating rate for several different thruster- and cooling system designs. The model results were verified by comparing the results of the model with the results from commercially available software. A good agreement was found with all results matching within 5-8% or less. Because the model relied on several semi-empirical relationships, test firings were performed using the PM200 and a heat-sink/radiation cooled version of the PM200 specifically designed for this study to determine two empirical constants required to calibrate the model. In total, 325 tests were performed. During these tests, temperature measurements were taken on various locations on the thruster. After the model was calibrated, the model was validated by checking the model results for different starting conditions with temperature measurements from test firings with corresponding starting conditions. It was found that the model was able to reproduce the temperature distribution measured in the tests within 15% for all cases (with the exception of one particular 2s burn case).
With the validated model, a number of regenerative cooling channel designs and radiation cooled designs was simulated for a reference thruster similar to the PM200. It was determined that with regenerative cooling the maximum wall temperature could be lowered by up to 23% for the reference thruster. This reduction in temperature was sufficient to lower the maximum wall temperature (to ∼1090 K) below the maximum operating temperature (1150K) of the stainless steel alloy used. A simulation of the PM200 design was also performed and a similar result was found; with regenerative cooling a maximum wall temperature of ∼1020 K is predicted for the PM200. For the radiation cooled designs, the wall temperatures exceeded the allowable temperature of 1150 K for both cases. The reference thruster design reached a maximum temperature of ∼1420 K and the PM200 reached a maximum temperature of ∼1320 K. For radiation cooling to be feasible a different more temperature resistant wall material is thus required. As this would come at an increase in cost, it was concluded that it is more beneficial to use regenerative cooling for the PM200.
While regenerative cooling was found to be effective, it was found that for a given type of cooling channel the design parameters have relatively little effect on the cooling performance due to the small size of the thruster. For all designs (with ribs) the variation in temperature was less than 82 K. The main drivers for determining the feasibility of a cooling method are thus not the cooling channel design parameters, but rather the heat input from the combustion. The heat input is mainly determined by the chamber pressure. As a result, the chamber pressure thus determines whether or not regenerative cooling is feasible. It was found that for the PM200 regenerative cooling is feasible as long as the chamber pressure stays below 7.1 bar.

This report presents the Final design of the Design Synthesis Exercise (DSE) to 'Capture a Small Asteroid and Change its Orbit' at the Faculty of Aerospace Engineering at Delft University of Technology. The bachelor programme 'Aerospace Engineering' comprises several projects enabling students to explore aeronautics and space from different kinds of perspectives. The Design Synthesis Exercise serves as the conclusion to this programme. During this final project students integrate their previously obtained knowledge and skill to examine a specific design problem in groups of ten students for the duration of eleven weeks. This final report is the last in a series of four and documents the detailed design of the concept that was chosen in the mid-term report.