The transition towards non-toxic, high-performance spacecraft propulsion has positioned highly concentrated hydrogen peroxide (HTP) and kerosene as a leading green propellant combination. However, achieving reliable hypergolic ignition in non-polar hydrocarbons remains a critical challenge due to significant physical mixing limitations and high chemical activation barriers. This study investigates the catalytic efficacy of Manganese(III) acetylacetonate (Mn(III)AA) dissolved in aviation-grade kerosene to enable rapid hypergolicity with 98% HTP. High-speed imaging and thermal diagnostics were employed to map the ignition delay time (IDT) across a range of catalyst loadings (0.5–10 wt%) and oxidizer-to-fuel ratios (4.5–7.5). The results demonstrate that Mn(III)AA is highly effective, achieving a minimum IDT of 25 ms at 50°C. Kinetic analysis revealed a significant reduction in apparent activation energy (9 to 14 kJ/mol), accelerating the chemical reaction rate until the system becomes limited by physical mixing processes. Notably, a non-linear performance trend was observed, where catalyst additions beyond 5 wt% yielded diminishing returns, suggesting a saturation threshold for practical engine design. These findings establish Mn(III)AA as a viable, high-efficiency additive for green bipropellant systems.

The urgent need for sustainable propulsion solutions has accelerated the exploration of green bipropellant thrusters using high-test hydrogen peroxide (HTP) with kerosene. In this study, a transient, high-fidelity CFD model coupling droplet-phase dynamics, real-gas behavior, and finite-rate chemical kinetics was developed to simulate the ignition and combustion processes in a coaxial-injected HTP-kerosene thruster. Simulations investigated the impact of oxidizer purity (95 % and 98 %) and mixture ratio variations, targeting a vacuum thrust of 100 N. Results revealed that stoichiometric mixtures with 98 % HTP delivered the most favorable balance of thrust (63.22 N at sea level) and thermal loads, with combustion temperatures aligning within 1 % of CEA predictions. Fuel-rich mixtures exhibited significant inefficiencies, with up to 18 % unburnt kerosene detected at the nozzle exit. Wall temperatures peaked at 3271 K under adiabatic assumptions, exceeding material safety thresholds, highlighting the necessity of advanced thermal management strategies. Observations of flow separation, shock structures, and model-predicted oxygen backflow further reinforced the realism of the simulations. This study advances green propulsion by linking combustion dynamics with structural viability. It provides new insights into propellant formulation, thermal management, and injector optimization for future environmentally compliant engines.

To address concerns regarding toxicity inherent in conventional storable brpropellants, the propulsion community is actively exploring "green" alternatives. Hydrogen peroxide (HTP), paired with eco-fnendly fuels such as kerosene or ethyl alcohol, is emerging as a promising option due to its potential for cost reduction in space launch, enhanced safety, ease of handling, and favorable density-impulse characteristics. This study investigates the hypergohcity and combustion dynamics of HTP with kerosene doped with organic manganese-based additives, targeting application in a 100 N class upper-stage thruster. The experimental campaign employs 95% and 98% HTP combined with variable catalyst loadings in kerosene to optimize ignition behavior and combustion performance. Drop tests were performed under controlled conditions to characterize ignition delay times (IDT) and post-ignition flame temperatures, supported by high-speed imaging and infrared diagnostics. Two O/F ratios (6.5 and 7.5) were explored to balance stoichiometric efficiency and ignition responsiveness. Comparative evaluation of Mn(II)AA and Mn(III)AA catalysts was conducted across a unified experimental matrix, with Mn(III)AA consistently outperforming Mn(II)AA by enabling faster ignition and higher combustion efficiency. Furthermore, the demonstration of catalytic hypergohcity eliminates the need for a conventional HTP decomposition catalyst bed, simplifying propulsion system architecture and reducing engine mass and cost. These findings provide critical data for the development of next-generation green propulsion systems, contributing to lighter, simpler, and safer thrusters for space applications. This research is an integral part of the EU Horizon Mane Sklodowska-Cune Actions (MSCA) funded initiative GREENLAM project, which aligns with the overarching goals of the EU Horizon initiative, facilitating technological innovation and advancements in the aerospace industry for the benefit of space exploration and satellite deployment.

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.