Cassese, Sergio (2024) Experimental and Numerical Research on Small-Scale Monopropellant and Hybrid Space Thrusters. [Tesi di dottorato]

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Abstract

In a worldwide scenario which sees a growing interest of the space market in the use of small satellites (e.g. CubeSats), the scientific community is pushing towards the research of proper miniaturized propulsive systems for attitude, trajectory and orbit control. In particular, novel mission concepts may be unlocked providing the spacecrafts with very precise and rapid manoeuvres capability, that electric thrusters cannot guarantee because of their relatively low thrusts. In this context, chemical propulsion appears to be suitable for this kind of applications. Among the various options, monopropellants and hybrids are garnering significant interest, as they can ensure good specific impulse performance, re-ignition and throttling capability with relative system simplicity, due to the single flow feed line. Another fundamental aspect is the pursuit of sustainable solutions using green propellants. Among these, hydrogen peroxide is a viable option, as it can be used both as a monopropellant, overcoming the toxicity issues associated with most used hydrazine-based systems, and as an oxidizer in hybrid propellant thrusters. For this reason, in recent years, this propellant has been experiencing renewed interest from the research community, particularly for potential applications in small satellites. The performance of hydrogen peroxide-based thrusters is strictly related to the catalytic chamber: the catalyst’s task is to decrease the activation energy leading to the decomposition reaction acceleration. The heat released by decomposition can be then converted directly into propulsive energy for monopropellants, which can also provide the necessary power for mixture ignition in the case of hybrids. In the literature, there are numerous studies characterizing the decomposition reaction and analysing the performance of the monopropellant using various catalysts. However, it should be noted that hydrogen peroxide-based monopropellants were just utilized several decades ago before being replaced by hydrazine. Meanwhile, hybrid rocket engines, despite their great potential, remain a highly challenging technology that is currently not used in space missions. In this scenario, the University of Naples “Federico II” (UNINA) is involved in several projects in which hydrogen peroxide finds its application as a monopropellant or oxidizer in the case of the hybrid rocket, in combination with polymeric fuel grains. In particular, this doctoral research, at the outset, aimed to go beyond the current state of the art of hydrogen peroxide monopropellants through both experimental and numerical approaches. It proposed a new type of catalyst and sought to optimize the decomposition process to ensure the fastest possible reactivity suitable for space applications. Subsequently, the goal was to elevate the technological level of hybrid rockets for small satellites, as this technology is virtually nonexistent in the literature for a thrust class of 10N. This process was to involve identifying the major challenges of this system, then attempting to overcome them and characterizing the internal ballistics of different types of fuels, thereby providing the scientific community with useful information for potential future developments. Thus, this work presents, step by step, the results obtained, starting from the monopropellant and culminating in the hybrid rocket configuration. Initially, a monopropellant thruster intended for use on a 12-Unit CubeSat in Low Earth Orbit was designed, developed, and characterized. This prototype was assembled and tested on ground and the effects of several parameters including chamber pressure, mass flow rate and catalyst material on decomposition efficiency and performance were studied and discussed. This experience led to the identification of an optimal ignition procedure that ensures maximum reactivity during the transient phase and maximum efficiency in the steady state. As a second step, inspired by the emergent high-entropy (HE) concept, a new family of catalytically active noble-metal-free oxides has been developed. The primary aim was to go beyond commercially available inert pellets doped with rare precious metals and tailor novel candidates for the H2O2 decomposition in monopropellant thrusters. Attention was dedicated to the fastest decomposition response that represents a fundamental requirement for short-time manoeuvres. Different compositionally complex new compositions were prepared and screened: the fastest responsive (i.e., the most reactive) candidate was thus selected, and a batch of pellets manufactured. Then, its performance was compared with a commercial Pd/Al2O3 catalyst used as benchmark: comparatively excellent results in terms of both reactivity and decomposition efficiency were achieved. Subsequently, some steps were made towards testing in hybrid rocket configuration. Thus, the problem of ignition for a 10N hybrid rocket based on hydrogen peroxide was treated, using the monopropellant engine in a pulsed mode, useful for preheating and capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also shows a good potential for space applications. In the second part of this phase, the objective was to find an ignition procedure that reduces propellant consumption and eliminates the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicate that the concept of pulsed pre-heating is feasible with only a small propellant consumption, while substantially decreasing the ignition duration from approximately 45 minutes to just maximum 3 minutes. Finally, similar ignition procedures were adopted using different fuels. The results show that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition are also included. However, the catalysts demonstrated a rather short lifespan when used in hybrid propulsion. For this reason, commercial palladium catalysts were compared with our compositionally complex (CC) catalysts. The former showed excellent performance in two tests but failed in the third and exhibited some mechanical weaknesses. In contrast, the CC pellets, under the same operating conditions, exhibited gradually decreasing reactivity but successfully ignited the flame three times without showing any mechanical damage. Furthermore, it was demonstrated that the loss of reactivity does not have a significant impact on combustion efficiency, nor does the use of different catalysts. Finally, characterization of the internal ballistics of various fuels was performed, supported by numerical techniques and Computational Fluid Dynamics (CFD) analysis. The first goal was to find a regression rate law for (Polyvinyl Chloride) PVC coupled to hydrogen peroxide on a 10-N-scale hybrid thruster. Different techniques were initially analysed in order to reconstruct the ballistics inside the combustion chamber, starting from the experimental data. Subsequently, three different regression rate curves were obtained according to various criteria. The resulting regression rate laws coming from the ballistic reconstruction appear to be in agreement with the experimental results and constitute a powerful means for designing engines with this fuel. Then, High-Density Polyethylene (HDPE) was characterized. Various tests were successfully conducted, and the propulsive performance was evaluated. Afterwards, using ballistic reconstruction techniques, the regression rate law was obtained, and considerations were provided on the possible effect of pressure on the regression rate. Furthermore, using a Computational Fluid Dynamics (CFD) model, the test conditions were simulated. It was found that for this scale, apart from the very extensive recirculation bubble, thermal aspects that had been neglected for large-scale engines may now be taken into consideration. Therefore, a modification was made to the setup of the numerical model, which resulted in a negligible simulation error compared to the experiments. The same was done for the 3D-printed Acrylonitrile Butadiene Styrene (ABS), a fuel that, especially in terms of the physico-chemical parameters affecting regression rate and ballistics, is poorly characterized in the literature. Therefore, in addition to post-processing the data and ballistic reconstruction using experimental results and CFD analysis, we sought to identify the parameters that best describe the internal ballistics of this fuel, achieving a good correlation between experimental and numerical results. The work concludes with a numerical and experimental analysis of a hybrid engine with a geometry different from the previous one, which reveals new challenges related to the combustion chamber geometry. However, these tests demonstrated that the numerical model effectively reproduces the axial consumption of the grain near the post-chamber. Notably, by isolating the grain from the heat coming from the post-chamber, a lower regression rate was obtained compared to the earlier tests. This indicates that grain heating is a significant issue to address for engines of this scale and has a non-negligible impact on the regression rate.

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