Design, manufacturing and characterisation of a water fed CubeSat micro-resistojet (Dondersteen)

To further the applicability of nano-satellites in large scale relevant missions, it is necessary that they have their own propulsion system for manoeuvring. Multiple nanosatellites flying in formation or even a constellation, with a distributed network of instruments can obtain simultaneous, distributed and guided measurements, which are not possible with conventional single-spacecraft architectures. A cubesat propulsion system technology demonstration has been shown on the Delffi-N3Xt satellite. Demonstrating cold gas generators, which generate self-pressurising nitrogen propellant from solid material. Now comes the time for an operational demonstration of powerful, innovative and effective propulsion for adequate manoeuvrability potential. This operational demonstration will be undertaken by the DelFFi mission, by having two nano-satellites demonstrating formation flying capabilities, such as inter-satellite distance maintenance. The DelFFi satellites will operate in a network of 50 nano-satellites, called the QB50 mission, with the goal to take multi-point in-situ measurements of the lower thermosphere and research re-entry. The gas generator based cold gas system have however been shown to be unable to produce the necessary total impulse within the stringent mass and volume constraints. The formation flying objective requires a new high performance propulsion system, which is designed during this master thesis. In order to design this propulsion system a number of tests were conducted on the candidate components and assemblies. After a survey of all commonly mentioned reaction engines, a resistojet was found to best meet the needs of this mission. Mainly because resistojets are relatively uncomplicated, do not require combustion and have a relatively low power consumption per amount of impulse generated. The best propellant for resistojets, selected from a large number of possibilities, was found to be water. Water was chosen because it is liquid around room temperature, It has a high potential impulse bit per unit of mass, it is benign, cheap and readily available. Based on the requirements on the propulsion system the following necessary design parameters were found: a water propellant mass of 50 [g], a nitrogen pressurant mass of 0.2 [g], a storage pressure of 4.5 [bar] to 2.5 [bar], a nozzle throat diameter of 25 [?m], a nozzle area ratio of 20 [-] and a final propellant heating temperature of 500 [°C]. The resulting propulsion system performance is: a total velocity increase ?V of 21.01 [m/s] (with a satellite mass of 3 [kg]), a total thrusting time of 17 [h] and 56 [min], a power consumption of 6.8 [W] to 3.7 [W] and a thrust force of 1.4 [mN] to 0.8 [mN]. A resistojet propulsion system requires pressure to transport the propellant from the storage tank to and through the thruster. Because pre-pressurisation is generally not allowed in CubeSats, gas generators were found to be the best option. These gas generators can be electrically initialised in orbit, where they produce gas from a solid. Gas generators used for airbag initiation in the automotive industry were found to be unsuitable for this mission because of the high shock pressure of possibly up to 95 [bar]. Additionally, three out of four gas generators that were tested malfunctioned. The reason for this malfunctioning is still unknown. For the DelFFi mission a waiver was granted, allowing pressurisation before launch. Therefore pre-pressurisation is considered to be the best option for propellant pressurisation. The most innovative part of this thesis is the development of the MEMS micro-thrusters called Dondersteen. These consist of an integrated fluidic inlet channel, heating chamber and rocket nozzle, which are all etched in silicon. The propellant is heated using heating elements made out of the very strong material silicon carbide. These elements are suspended in the middle of the fluid flow in order to maximise the heat transfer to the propellant, and minimise the heat loss to the surroundings. The novelty not only lies in the geometry, but also in the manufacturing process. Hereby the fluidic channels and suspended heating element are both etched out of the silicon carbide covered silicon wafer in the same two etching steps. The developed thruster consists of the following three distinct sections: inlet, heating chamber and nozzle. Multiple designs were made for each of these sections. The production process was chosen such that any combination of these sections can be etched behind one another. This creates the flexibility of being able to produce and test multiple thruster layouts, in order to find the one which best fits the requirements. Different heating chamber geometries were designed and tested. These enable the assessment of the geometrical influence on heat loss, pressure drop and propellant heating capabilities. Both the manufacturing process and the resulting heating chambers show great potential. The heating elements have been shown to be robust during the rough testing process. The propellant channels have very favourable rounded edges which reduce pressure and heat losses, while maximising the heating contact and propellant mixing. Due to the experimental nature of the manufacturing process the fluidic channel dept was found to be too large and the nozzle throat diameters were too large. The performed analysis of these discrepancies will be used to improve the manufacturing process. The tested resistance values of the resistive heater modules were found to be 200 to 600 times larger than designed. This increases the required input voltage from 5 [V] to the range from 70 [V] to 120 [V]. For the current testing phase this is not a problem since these supplies are available. The flight models will however have to be redesigned in order to comply with the 5 [V] requirement of the satellite power supply. The main reason of this discrepancy was found to be a calculation error. Some adjustments have already been listed with which this resistance can be decreased by a factor of 144. Further design efforts will have to increase this factor to obtain the real design value of the resistance. There are some tasks that still have to be performed by a succeeding master student to consolidate the work done in this master thesis to a propulsion system flight model. The developed thrusters have to be performance tested with propellant in multiple operating conditions. The results of these tests lead the final redesign of the developed thrusters. The propellant storage system needs to be built and tested. Finally the complete in flight propulsion systems has to be defined, built and integrated into the satellites.Space Systems EngineeringSpace EngineeringAerospace Engineerin