This study presents a simulation-based damage modeling and fatigue risk assessment of a reusable ceramic matrix composite thruster designed for short-duration, green bipropellant propulsion systems. The thruster is constructed from a fiber-reinforced ultra-high temperature ceramic matrix composite composed of zirconium diboride, silicon carbide, and carbon fibers. Time-resolved thermal and structural simulations are conducted on a validated thruster geometry to characterize the severity of early-stage thermal shock, stress buildup, and potential degradation pathways. Unlike traditional fatigue studies that rely on empirical fatigue constants or Paris-law-based crack-growth models, this work introduces a simulation-derived stress-margin envelope methodology that incorporates ±20% variability in temperature-dependent material strength, offering a physically grounded yet conservative risk estimate. From this, a normalized risk index is derived to evaluate the likelihood of damage initiation in critical regions over the 0–10 s firing window. The results indicate that the convergent throat region experiences a peak thermal gradient rate of approximately 380 K/s, with the normalized thermal shock index exceeding 43. Stress margins in this region collapse by 2.3 s, while margin loss in the flange curvature appears near 8 s. These findings are mapped into green, yellow, and red risk bands to classify operational safety zones. All the results assume no active cooling, representing conservative operating limits. If regenerative or ablative cooling is implemented, these margins would improve significantly. The framework established here enables a transparent, reproducible methodology for evaluating lifetime safety in ceramic propulsion nozzles and serves as a foundational tool for fatigue-resilient component design in green space engines.

This study presents a comprehensive finite element investigation into the design optimization of an ultra-high temperature ceramic matrix composite thruster for green bipropellant systems. Focusing on ZrB2–SiC–Cfiber composites, it explores their thermal and mechanical response under realistic transient combustion conditions. Two geometries, a simplified and a complex full-featured model, were evaluated to assess the impact of geometric fidelity on stress prediction. The complex thruster model (CTM) offered improved resolution of temperature gradients and stress concentrations, especially near flange and convergent regions, and was adopted for optimization. A parametric study with nine wall thickness profiles identified a 2 mm tapered configuration in both convergent and divergent sections that minimized mass while maintaining structural integrity. This optimized profile reduced peak thermal stress and overall mass without compromising safety margins. Transient thermal and strain analyses showed that thermal stress dominates initially (≤3 s), while thermal strain becomes critical later due to stiffness degradation. Damage risk was evaluated using temperature-dependent stress margins at four critical locations. Time-dependent failure maps revealed throat degradation for short burns and flange cracking for longer durations. All analyses were conducted under hot-fire conditions without cooling. The validated methodology supports durable, lightweight nozzle designs for future green propulsion missions.

This study investigates the impact of different powder milling methods on the densification and mechanical properties of ZrB₂-SiC ceramic composites processed via spark plasma sintering (SPS). Powders were prepared using two ball milling techniques: tungsten carbide (WC) and conventional ZrO₂. The densification behavior during SPS was monitored, and the sintered samples were evaluated for their relative density, hardness, fracture toughness, and flexural strength. Results show that WC milling significantly enhances densification, achieving 99.2 % relative density at 2100 °C/65 MPa/15 min, compared to 96.5 % for ZrO₂-milled samples. This improvement is due to WC's sintering aid effect, which promotes grain boundary diffusion and particle packing. However, ZrO₂-milled composites exhibit superior hardness (17.38 GPa) and fracture toughness (3.97 MPa m¹/²), attributed to their refined grain structure and the absence of softer ZrO₂ phases. Conversely, WC-milled samples show slightly higher flexural strength (384–516 MPa), likely due to the transformation toughening effect of the secondary ZrO₂ phase. Overall, WC milling improves densification and flexural strength, while ZrO₂ milling yields finer-grained composites with higher hardness and toughness, making it better suited for wear-resistant and mechanically demanding applications.

This study concerns the ABS-Nitrous hybrid engine development performed at Delft Aerospace Rocket Engineering (DARE). DARE is a student rocketry society associated with the Delft University of Technology. Its goal is to provide hands-on experience to complement the theoretical material taught at the university's faculties. This project started during the development of DARE's Stratos IV rocket, directly after the breakup of Stratos III. A small propulsive roll control system was suggested to remedy the problem. The high-power requirement of a monopropellant system encouraged the DARE team to explore the restartable ABS-Nitrous hybrid system option as a low-power alternative. Some key features are it is non-toxic, requires no pyrotechnics for ignition, utilizes a low-power ignition source, has a simple system architecture, is restartable through a hydrocarbon seeding effect, and has consistent fuel grain production through the FDM process. It comes at a performance marginally lower than HTPB. These unique properties of 3D-printed ABS make it a suitable candidate for applications where hybrids typically are not. For example, in an engine ignition system, a satellite attitude control, orbit maintenance, or orbit transfer system; a sounding rocket roll control system, or its (restartable) main engine. This study aims to make ABS-Nitrous hybrid engines more accessible for future engineering applications by developing a validated preliminary design tool to generate the required geometry for a particular application. Different existing engineering models in literature have been collected that include models of self-pressurized propellant tank dynamics, multi-phase injection models, and several regression rate models. The infrastructure and hardware required to fire variable motor sizes are present within DARE and have been expanded and tailored for the needs of this system. 27 tests have been performed of which 10 were successfully ignited hot fires. Analysis of this data yielded some boundaries of ignitability, given insights in promoting restartability and showed a low ignition delay of 100ms at a consistent stable ignition input power. Two tests featured helical ports showcasing control of the grain burn profile. The data of the hot fires was used to validate the preliminary design tool. Modelled values for the feed system pressures and thrust output have been proven to be within ±10% of the experimental data. The validated rapid development tool enables future projects to use this concept and lower the threshold required to get started on a design while getting new students acquainted with the topic and expanding the body of knowledge.

With the ever-growing number of spacecraft launching into orbit, alongside the growing desire for propulsion systems on many of these craft, an emerging opportunity is present in the potential servicing and refuelling of these spacecraft. Proposals for propellant resupply services are growing in abundancy, primarily with architectures involving craft launching from the Earth's surface to transfer their cargo of propellant to in-orbit customers. Current state of the art solutions utilise hydrazine, however its popularity is dwindling with the European Chemicals Agency considering outlawing it due to its toxic and carcinogenic nature. For this reason, SolvGE, a start-up based at TU Delft in the Netherlands, has proposed a sustainable architecture involving the production of high concentration hydrogen peroxide (H2O2) from water-ice present at off-Earth locations. While H2O2 does possess a lower specific impulse than hydrazine, it has several other attractive characteristics such as its high density, storability, and nontoxicity. To investigate the viability of such an architecture, a Single Stage to Orbit refuelling craft is sized, showing that with a structural coefficient of 0.3 and an I_sp of 330 s - 340 s, a refuelling craft could launch from the Moon to low lunar orbit, and from Deimos and Phobos to low Martian orbit, to refuel spacecraft with 200 L of H2O2. A trade-off of potential refuelling mechanisms shows that a piston-based system, used in conjunction with pressurant gas, a gas generator, or a pump, is a good candidate for high cycle usage. Comparison of permutations of these systems shows low variation in total system mass (< 2%) over different transfer masses and transfer rates. A prototype test set up of the transfer mechanism using pressurant gas is created to investigate the functioning of the piston system and the relevant pressurant parameters. Transfer tests with and without the piston head show a 17%-25% increase in pressurant mass required when a piston head is added. Testing with the system inverted shows 9% more pressurant is required due to the adverse gravity gradient. Thus, a spacecraft capable of refuelling 200 L of H2O2 weighs approximately the same as the wet mass of the Apollo lunar lander (~16000 kg) and is able to serve customers in low lunar and Martian orbits in a reusable manner. Further research on high cycles of the piston as well as possible mass savings, and actuation using H2O2 gas generators will further assess the applicability of piston transfer systems for in-orbit refuelling.

This research investigated how the variation of temperature and shear rate affects the viscosity of ethanol gel propellants that use methyl cellulose as gellant and, in parts, use boron as energetic additive. Using a rotational viscometer in a cone-and-plate configuration, propellant viscosity data was recorded across a range of temperatures and applied shear rates. The temperaturedependence of the viscosity was modelled using an Arrhenius-type equation. For the high shear rates, the data was modelled using the Power Law, Herrschel–Bulkley model, Carreau model, and Cross model. For low shear rates the used model was the rearranged Herrschel–Bulkley model. The temperature investigation suggested that the trend of decreasing viscosity with increasing temperature, predicted by the Arrhenius-type equation, is only applicable until approximately 320 K, after which the gel viscosity increased strongly. At high shear rates, the gel behaved in a shear thinning manner and was modelled most accurately by the Cross model. At low shear rates, the gel was shear thickening up to its elastic limit, which was found to lie at 0.41 s^–1.

Since there is a high interest in the use of green propellants, hydrogen peroxide is coming back after once making place for the rise of Hydrazine in monopropellant propulsion systems. Typically, these thrusters are outfitted with catalyst beds. A fully modular 1N thruster is designed to provide the capability of testing and comparing the performance of different concentrations of hydrogen peroxide, different catalysts as well as new technologies in an attempt to resolve the disadvantages associated with the use of catalyst beds. A preliminary baseline design of a catalytic thruster has been created. This will be followed by the design of a secondary decomposition chamber for new technologies, a propellant feed system, a test setup and a test plan.

In regard to propulsion system applications, the stability of liquid propellants in long-term storage is of increasing importance, and this had led to a greater interest in gelation technology. As part of a preliminary test to determine the feasibility of using a gel propellant in a rocket with a catalyst bed, a hybrid rocket with a catalyst reactor using a gel propellant as an oxidizer was tested for the first time in this study. Experiments were conducted with two different oxidizers: one with liquid phase hydrogen peroxide and the other with gel phase hydrogen peroxide, as well as high-density polyethylene as fuel for a 250N class hybrid thruster performance test. The thruster was designed with the catalyst ignition system, and a catalyst was manufactured to be inserted into the catalyst reactor to facilitate oxidizer decomposition. While the test result with neat hydrogen peroxide indicated sufficient decomposition efficiency using a manganese dioxide/alumina catalyst and successful autoignition of the fuel via the decomposed product, gel hydrogen peroxide exhibited insufficient decomposition and there were difficulties in operating the thruster as a part of the catalyst was covered in the gelling agent. This preliminary study identifies the potential challenges of using a gel phase oxidizer in a catalyst ignited hybrid thruster and discusses the technical issues that should be addressed in regard to a gel propellant hybrid thruster design with a catalyst reactor.