In the last few decades, developments within the space industry have led to faster iterations in the design process of space products, where the emphasis has been laid on testing rapidly rather than perfecting a design first. Especially with thrusters, where the complex combustio
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In the last few decades, developments within the space industry have led to faster iterations in the design process of space products, where the emphasis has been laid on testing rapidly rather than perfecting a design first. Especially with thrusters, where the complex combustion and gas dynamics cannot be perfectly simulated, rigorous testing is necessary in order to obtain a good design. Building on this method of engineering, certain components of a thruster need to be changed in order to accommodate this. One critical component is the propellant. Where previously the objective was selecting a cheap, high-performing propellant, a shift has occurred where now safety, sustainability, and reduced complexity are the most critical aspects.
In this thesis, together with the European Space Agency (ESA), a new thruster has been developed using nitrous oxide and ethane as its propellants. Since both propellants are non-toxic, self-pressurizing, and easily obtainable, they make for ideal candidates in the NewSpace economy.
Using rocket engine theory, a modular 20 N regeneratively cooled thruster has been developed. CFD and FEM simulations indicate that the thruster is capable of firing in steady-state mode without failure; however, physical verification still has to take place in order to fully verify the thruster.
In order to demonstrate the modularity of the thruster, three configurations have been developed. While the design of the thrust chamber has been kept constant, the injector and igniter are different across the configurations. At atmospheric conditions, based on numerical analysis, the thruster has a specific impulse of 160 s.
For the material of the thruster wall, a recently developed additive manufacturing nickel alloy called ABD900AM is used. Together with seventeen cooling channels of 0.5 mm width and 0.5 mm height, a maximum inner wall temperature of 936 K is reached, well below the temperature after which the properties of ABD900AM start to degrade. Using FEM analysis, a safety factor with respect to the yield strength of 1.5 was found as the lowest value in the thrust chamber. At the corners of the fixed flange to which the thrust chamber was attached, a value of 1.0 was found.
As solely verifying the thruster via computational analysis is not sufficient, a verification plan in line with ECSS guidelines was written. This verification plan lays out the tests that need to be performed in order to fully verify and validate the thruster. It includes tests such as cold flows, actuation tests, and hot-fire tests. These tests have not been performed in this thesis due to time constraints but will be performed in the future by the European Space Agency.
Furthermore, the initial steps of the manufacturing process have also been performed. The CAD models have been developed and are shown in this thesis, as well as the associated technical drawings. Next to that, iterations have been made during the thesis based on feedback given by machinists at ESA, as well as external suppliers and manufacturers.
