Ultra-high-temperature ceramic matrix composites (UHTCMCs) are promising materials for thruster design due to their durability and exceptional high-temperature properties. This study assesses their viability by exploring the loads on the thruster, the results of 3D finite element analysis (FEA), and the impact of prolonged operation. Using ANSYS 2023 R2 and expert forums, a detailed model was created to capture the necessary data. The optimal thickness configuration resulted in a mass of 0.31 kg. The analysis predicts a maximum nozzle temperature of 1643K and a stress level of 302 MPa at the flange after 10 seconds of operation. Peak nozzle stress occurs at 2.16 seconds, reaching 133 MPa. The results suggest the thruster is well-suited for short-burst attitude control manoeuvres. However, extended operations such as orbit control, reentry, and sudden manoeuvres increase the risk of sudden flange fractures, making these applications less viable in the long run.

Space exploration depends on materials that can withstand extreme conditions, particularly for rocket thrusters. Ceramic matrix composites (CMCs) like carbon-carbon (C/C) and carbon-silicon carbide (C/SiC) have been widely used for rocket nozzle applications ,due to their thermal and mechanical properties. However, the demand for materials capable of higher temperature tolerance and reusability has shifted focus to ultra-high temperature ceramics (UHTCs) and UHTC matrix composites (UHTCMCs). Among UHTCs, ZrB₂ stands out for its excellent mechanical properties, oxidation resistance, and lower cost compared to HfB₂. To overcome ZrB₂'s inherent brittleness and enhance properties like densification and oxidation resistance, additives such as SiC and carbon fibers are introduced. SiC enhances densification, controls grain growth, and improves oxidation resistance via the formation of a SiO₂ protective layer, while carbon fibers improve mechanical strength, oxidation resistance, and reduce density. Spark Plasma Sintering (SPS) is a preferred fabrication method for its ability to rapidly densify materials while maintaining fine microstructures.
Although ZrB₂-SiC composites are extensively studied for high-temperature applications, the relationships between sintering parameters (e.g., temperature, pressure, and dwell time) and densification, microstructure, and mechanical properties remain underexplored. This study investigates these correlations using SPS and examines the impact of milling methods—high-energy milling with WC balls versus regular milling with ZrO₂ balls—on final material properties. The feasibility of incorporating short carbon fibers into the ZrB₂-SiC matrix is also assessed, focusing on the effects of preparation techniques and fiber length.
Fabrication insights revealed that increased sintering temperature generally improved densification due to enhanced atomic diffusion, grain boundary migration, and mass transport. High-energy WC milling achieved superior densification compared to ZrO₂ milling, with ZSW samples reaching a maximum relative density of 99.2%, versus 96.5% for ZSZ samples under similar conditions. ZSW samples, however, developed a secondary ZrO₂ phase due to more intense abrasion and oxygen diffusion during milling, while ZSZ samples maintained a finer microstructure, with an average grain size of 2.65 μm compared to 2.91 μm for ZSW. Attempts to incorporate 35 vol% short carbon fibers were unsuccessful under current sintering conditions, but improvements in fiber distribution were achieved with a rotary evaporator. Shorter fibers showed better structural integrity by reducing stress concentrations.
Mechanical properties were strongly influenced by sintering temperature. Higher temperatures caused grain coarsening, leading to reductions in hardness, flexural strength, and fracture toughness. For instance, ZSW hardness decreased from 14.33 GPa at 1950°C to 13.92 GPa at 2050°C, flexural strength declined from 407 MPa to 384 MPa, and fracture toughness dropped from 3.71 MPa·m¹/² to 3.58 MPa·m¹/². Milling methods also played a critical role; ZSW samples showed lower hardness and toughness due to the softer ZrO₂ phase and coarser grain sizes, with a maximum fracture toughness of 3.76 MPa·m¹/² compared to 3.97 MPa·m¹/² for ZSZ samples. However, ZSW samples exhibited comparable or higher flexural strength (384–516 MPa) due to ZrO₂’s transformation toughening effect, while ZSZ samples ranged from 317 to 476 MPa.

With the space industry growing, environmental considerations become increasingly important, especially with respect to the propulsion systems used to launch satellites into space and control their position in orbit. Since traditional satellite propellants are highly toxic, there is an increased demand for green, i.e., environmentally friendly, substitutes like hydrogen peroxide.
This work explored the modelling of hypergolic bi-liquid thrusters in the framework of the Greenlam project, which aims to develop a 100N hydrogen peroxide kerosene thruster. While previous works were either experimental or focused on staged H2O2–RP-1 engines with a catalyst bed, this thesis investigated a numerical approach and focused on unstaged engines, aiming to identify and validate models viable to simulate the decomposition of hydrogen peroxide and subsequent combustion with kerosene with the aid of a catalyst.
Transient three-dimensional simulations were performed. k-ω SST, the Peng Robinson real-gas equation of state and Species Transport with Finite Rate chemistry were employed to model turbulence, gas properties and reactions, respectively. The effect of the catalyst was represented by adapting the Arrhenius rate parameters. Propellants were injected using the Discrete Phase Model. The Eulerian model was shown not to be suitable to simulate the propellant injection and atomisation.
A coaxial, an impinging-jet and a pintle injector were considered. Simulations with the coaxial injector showed good agreement with data obtained from CEA and with other rocket engines. Simulations with the impinging-jet and pintle injector failed to capture droplet impingement and consequent atomisation and thus could not be validated.
Both stoichiometric and fuel-rich propellant mixtures and H2O2 concentrations of 95% and 98% were simulated. Thrust was between 62 and 63N under sea-level conditions, equivalent to 103 to 105N in vacuum and hence approximately 3 − 5% higher than anticipated. Chamber temperature reached up to 2763K. Chamber pressure was 7.6bar. The stoichiometric mixtures showed higher thrust output, higher chamber temperature and higher wall temperature than the fuel-rich mixtures. The higher concentrations led to higher chamber and wall temperatures. Analysing the kerosene mass fraction in the exhaust showed that in any case at least 9% of the injected kerosene was ejected unburnt due to a lack of mixing, and most of the additional kerosene in the fuel-rich mixtures was also simply ejected. The chamber walls reached temperatures of up to 3271K, about 500K higher than bearable by the material. While the coaxial injector was shown to be a cause for the high wall temperatures due to unfavourable propellant distribution, an adiabatic wall boundary condition was assumed which likely also led to an overestimation of the temperature.
A set of models applicable for simulating hypergolic bi-liquid rocket engines was found and validated. More work is required in terms of injector design and modelling, confirmation of reaction rate parameters and wall modelling.