The H2GO project presents the design of H2ERMES, a reusable launch vehicle developed to autonomously deliver liquid hydrogen (LH2) to orbital fuel depots, supporting the advancement of nuclear thermal rocket (NTR) propulsion for deep space missions. The vehicle, optimized for environmental and economic sustainability, features a reusable second stage with novel technologies, including a regeneratively cooled heat shield, aerospike-effect propulsion, and cryogenic LH2 storage integrated with structural elements. H2ERMES is designed for 25 or more missions with minimal refurbishment, meeting stringent performance, safety, and sustainability requirements. A comprehensive systems engineering approach defined user and stakeholder requirements, conducted risk analyses, and performed extensive subsystem trade-offs. The final design includes high-efficiency LH2 tanks, a 24-chamber aerospike engine using an expander bleed cycle, actively cooled heat shield, autonomous docking systems, and reliable recovery mechanisms. The project addresses growing market needs for sustainable space infrastructure, with first operational flights planned by 2032 to enable economical, reusable, and low-emission LH2 transport to orbit.

This thesis presents the development of an open-loop optimization framework for computing the optimal ascent trajectory of a multi-stage booster, with specific application to the insertion of a Hypersonic Glide Vehicle (HGV) into its glide phase. The study addresses the challenge of guiding a launch vehicle through atmospheric and exo-atmospheric regimes while satisfying physical and terminal constraints. The HGV insertion case provides a relevant example, given the growing interest in boost-glide systems for both defense and research applications.

The optimization problem is formulated using the Pontryagin Maximum Principle (PMP), enabling an indirect method that solves the coupled state--costate dynamics with strict boundary conditions. The ascent is modeled as a 3-DoF point-mass in Cartesian coordinates, taking into account aerodynamic forces, Earth's rotation, and path-and-control constraints. This study presents a suitable constraint set and problem definition for the optimization problem that improves numerical convergence. A multiple-shooting approach is used to ensure convergence for the nonlinear dynamics, while a homotopy continuation strategy gradually incorporates aerodynamic and path constraints to improve numerical robustness. Furthermore, the developed trajectory optimization model allows for the implementation of arbitrary boosters that include staging and throttle management.

Results demonstrate smooth, dynamically consistent ascent trajectories that meet target insertion conditions required for an efficient HGV glide. In addition, this report introduces various evaluation techniques for the computed trajectories, together with verification methods that ensure their physical accuracy. Overall, the work delivers a comprehensive and extensible framework for optimal ascent guidance under realistic aerodynamic and operational constraints, with direct applicability to modern boost-glide mission design.

Reliable propellant tank pressurization in a launcher is a critical requirement for mission success, as deviations in propellant or tank pressure can compromise feed system operation or even lead to structural failure of the tanks.
This thesis addresses the development of a pressure regulation strategy for the pressurization systems of the fuel and oxidizer tanks in the first stage of Maiaspace’s mini-launcher. The work focuses particularly on the creation of pressurization system models capable of replicating system behaviour with sufficient accuracy for design studies, while prioritizing fast development and computational efficiency. Simulations of the developed models served as the basis for defining regulation strategies for the oxygen and methane tank pressurization systems.

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.

Due to numerous disadvantages exhibited by hybrid rocket motors (HRMs) as a propulsive technology, adoption has not been wide-spread outside of amateurs and student rocketry teams. However, in recent years, additive manufacturing has revitalised this propulsive technology. This thesis aims to use this technology to overcome some of the disadvantages exhibited with HRMs.

By choosing regressive burn profile port geometries, shifting of the mixture ratio can be minimised, meaning parameters such as specific impulse to be more constant and – in the future – be engineered to operate at its peak efficiency. A multi-port geometry was also chosen to overcome HRMs' slower regression rate.

A practical investigation was carried out to determine the extent these changes affect the performance of HRMs compared to simulations, especially in the domain of the mixture ratio. The results show this technology to have potential, however certain testing limitations need to be overcome before a definitive conclusion can be drawn.

Sounding rockets are used in various research applications such as micro-gravity research, re-entry experiments and atmospheric measurements. They provide a low-cost alternative to orbital test platforms. A recent development in the field of sounding rocketry has been the development of steerable upper stages to enhance these capabilities.

To bring this capability to T-Minus Engineering and its miniaturised sounding rocket, the Barracuda, the development of an upper stage for this rocket was commenced. One of the key subsystems of this upper stage is propulsion system. This thesis addressed the design, modelling and mass-optimisation of this upper-stage propulsion system.

Based on previously developed technologies within T-Minus Engineering, a high-test peroxide and gasoline thruster was designed, utilising catalytic decomposition of the peroxide to auto-ignite the propellants while also using the peroxide regeneratively cool the combustion chamber. The outcome of this project will further guide the development of bi-propellant thrusters at T-Minus Engineering.

With the urge of not only the aerospace propulsion system industry, but all industries of engineering to improve each system into a more sustainable solution, a lot of different green propellant combinations have been investigated that are less toxic, and more eco-friendly, but however efficient at the same time. This is true for liquid bi-propellant systems as well. Due to the emergence of new propellant combinations, or with the demand for more energy-efficient propulsion systems, new types of ignitor systems, which are a significant element in a bi-propellant space propulsion system, or the improvement of the conventional systems to a more efficient one are being investigated. Following the demand for research in these sectors of propulsion systems, this MSc thesis focuses on the development of a liquid bi-propellant space propulsion ignition concept, which is optimised from the perspective of power, cost, and performance.

During the literature study phase of the thesis, several ignitor systems were investigated and the thermal wire ignitor concept was chosen to be the most suitable system for the study due to its reusability function, cost-effectiveness, mass/volume, and handling features. On top of that, after understanding the features of several fuel and oxidiser combinations, ethanol and HTP were chosen as the most optimal propellant combinations. These were shown to provide performance similar to conventional propellant combinations, such as in density-specific impulse, and most importantly, were considered an eco-friendly combination.

During the experimental phase of the MSc thesis research, the first experiment aimed to measure the resulting HTP temperature when it passes through a NiCr thermal wire mesh in a glass chamber, where the chamber itself is heated up with a wire wound around the external wall. This experiment showed that a resulting temperature of almost 400 deg C was achievable with 20 W of power applied to the wire mesh down to a value of 294.37 deg C for 5 W of power applied. The most interesting point is at a power value of 12.5 W, in which the HTP reaches a temperature value right above the auto-ignition temperature of ethanol, 375.65 degrees C.

The second experiment aimed in investigating the ignition behaviour of the fuel and HTP in an open environment. The results showed that at 12 W of power applied to the NiCr thermal wire mesh, which was in contact with a premixed fuel and oxidiser pool this time, combustion and self-sustained ignition were achieved with a sufficiently short amount of IDT (around 200 ms). Also, at 10 W of power applied, combustion occurred with HTP and Jet A, which was a reference fuel used with the purpose of showing that fuel types that have an auto-ignition temperature lower than ethanol are able to ignite at lower power consumption values. The self-sustained ignition was obtained at power values slightly higher than this, 12.5 W.

The last experiment aimed in investigating the ignition behaviour of the fuel and HTP in a closed environment, which also serves as a pressurised system. The results showed that at 10 W, the ethanol and HTP combination was able to combust and self-sustain. Also, at a value of 7.5 W, the Jet A fuel and HTP combination were able to combust and provide a self-sustainable ignition. They both showed that in a pressurised system, lower levels of power values are required in order for the same fuel and oxidiser combination to achieve combustion.

Simultaneously with the experiment conducted, a simulation of the HTP decomposition temperature in a glass chamber was done using the cross-platform finite element analysis, solver and multi-physics model software, COMSOL. The simulation was able to show that at power values of 20 W, as the HTP is injected in the glass chamber, the liquid was able to achieve a temperature of 589 degrees C but decreased drastically down to the initial temperature of the volume that was inputted in the software. This simulation was done using a mesh configuration named Normal, which is an automatically available mesh quality in the COMSOL software. Together with the same mesh quality, if 15 W of power is applied, the initial temperature that the HTP achieves is 492 degrees C and for 10 W, it is 394 degrees C. All the results presented matched the results of the experiment conducted. Lastly, 2 sensitivity analysis were performed in order to prove that the simulation was done properly.

The Mars Sample Return (MSR) mission is a collaboration between NASA and ESA with the aim of retrieving the Martian rock samples gathered by the Perseverance Rover and sending them back to Earth for further study. This report outlines the design of a Mars Ascent Vehicle (MAV) and a Sample Return Lander (SRL) for this mission, with the additional consideration that in-situ resource utilization will be performed. This means that while on Mars, carbon dioxide will be captured, and converted into liquid oxygen and liquid carbon monoxide to be used as propellants. Here one trades off a lower mass of propellant to be brought, with a larger mass of additional systems for propellant generation. This has the potential to bring net mass benefits to the mission, hence justifying a study of its feasibility.

In the New Space era, private companies are driving rapid advancements in rocket propulsion. Among them, OPUS Aerospace is developing a reusable micro launcher that requires the design of a new liquid rocket engine, Elanion, to deliver 45 kN of thrust using liquid oxygen and propane as propellants. This thesis presents the preliminary design of the Elanion engine’s thrust chamber, spanning from contour generation to the optimization of its regenerative cooling system. Two in-house tools were developed to rapidly iterate and refine the chamber geometry and cooling system design. Additionally, a study of propane’s cooling properties led to the development of two Nusselt correlations for modeling its convective heat transfer. The thrust chamber designed in this research meets all specifications, leaves margin on the budgeted pressure loss, and was successfully converted into a CAD model, laying a strong foundation for further development of the Elanion engine.

Ball valves are a critical component in a plethora of liquid propellant propulsion systems. The Spanish company PLD Space is considering the usage of these valves in their new series of KER-LOX liquid propellant rocket engines, the TEPREL-C & TEPREL-C Vacuum engines. These will be operating on-board the Miura 5 launch vehicle. The purpose of this research is to re-design and analyze an unoptimized ball valve design, as to make it suitable for cryogenic service while minimizing its mass. The secondary objective of the research is to inquire into the aspects of ball valve operation. This includes investigating the required operating torque of a ball valve and the achievable flow factor as a function of a ball valve's opening angle. The results of this thesis are scripts which can estimate the torque and flow factor values, a methodology to mass-optimize a ball valve, and a re-designed ball valve suitable for cryogenic service.

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.

Cryogenic and semi-cryogenic propellants are the most commonly used liquid propellants for applications in medium-lift launch vehicles. Despite their high performance, the storage requirements for these propellants often lead to complex, heavy, and voluminous structures. The only storable propellant used in medium-lift launch vehicles, UDMH/NTO, comes with its own problems of high toxicity and reduced performance. A promising alternative to this could be storable fuels with highly concentrated hydrogen peroxide (HTP) as an oxidiser. Despite a shorter history of dedicated development, HTP has proved itself an effective oxidiser for in-space applications and small-lift launch vehicles. Therefore, the question could be raised towards the potential of this oxidiser for applications in medium-lift launch vehicles. In this study, the application potential of an HTP-based storable bi-propellant for medium-lift expendable launch vehicles was investigated. To this extent, a large selection of green storable fuels was considered to find the most suitable propellant for this application.

Both the integration and compatibility potential of the propellants and the propulsive and mass performance potential were investigated. The integration and compatibility potential were evaluated through a qualitative assessment based on non-performance-related propellant characteristics. Furthermore, eight fuels were subjected to a more detailed assessment covering the criteria of handling toxicity, environmental toxicity, material compatibility, handling and storage, development level, and coolant qualities. RP-1 was found to be the most suitable fuel with respect to the specific criteria, while ethanol, methanol, isooctane, and isopropanol were also found to be promising alternatives. A launch vehicle model was created to evaluate the propulsive and mass potential of twelve fuels proposed based on earlier findings. This model included a propulsion model, a mass and sizing model, and an aerodynamics and trajectory model, which were all connected through a global optimisation model. In terms of propulsive potential, the cryogenic propellant hydrolox was predicted to have a 25% higher vacuum specific impulse than the best-performing HTP-based propellant DMAZ/HTP. In terms of the specific impulse density, kerosene-derivative fuels in combination with HTP were predicted to have a better performance than hydrolox and than that other conventional storable propellant UDMH/NTO. The optimised gross lift-off mass for the launch vehicle concepts employing HTP was found to be 42-61% higher than the gross lift-off mass of Ariane 6 predicted through the model. Separately, the payload capability of the HTP-based launch vehicle concepts was predicted to be at least 38% lower. In both cases, RP-1/HTP was reported to be the HTP-based propellant with the best performance, while DMAZ, isooctane, and isopropanol could be regarded as suitable alternatives. All of these propellants also outperformed UDMH/NTO. Through a sensitivity analysis, it was discovered that up to 270kg additional payload could be taken to GTO upon considering elevated chamber pressures in the HTP-based engine design. In the end, the high potential and promise of HTP were confirmed as it was concluded that increased development efforts towards HTP-based storable bi-propellant rocket engines could not only lead to a promising alternative to cryogenic propellants but could also allow for the complete replacement of toxic hydrazine-derivative fuels.

For over 50 years, hydrazine has been the industry standard for monopropellant propulsion systems, widely used in satellite attitude and orbit control systems. However, hydrazine’s toxicity necessitates expensive handling procedures and may lead to a future ban of the propellant in Europe. This has motivated the development of novel monopropellants, featuring reduced toxicity compared to hydrazine. Separately, life cycle assessments (LCAs) are becoming increasingly prevalent in the space industry. As very few assessments have been made so far for monopropellant systems, this thesis performs a comparative life cycle sustainability assessment (LCSA) of a hydrazine and three novel monopropellant systems for a single use case, evaluating the environmental, economic and social sustainability of each. This research provides new insights into the life cycle impact of the differences between the various propulsion systems and identifies hotspots in each sustainability dimension, informing a more sustainable development of novel monopropellant systems in the future.

Highly concentrated hydrogen peroxide, known as high-test peroxide (HTP), is widely regarded as a greener and safer alternative to traditional hypergolic fuels. A recent development demonstrated the
potential to manufacture HTP in-orbit, bypassing prolonged storage challenges associated with the substance. This sparked interest in development of a space resident HTP manufacturing depot which could enable a next generation of satellites powered by greener propellants capable to routinely refuel.
One of the missing pieces is a refuelling interface that could be utilized in such a system. In this work, an initial set of conceptual requirements for the device are proposed together with a potential design based on existing solutions. The design is partially implemented and studied in simulations to verify the
major components are feasible to develop. Out of the components, the development of an all-aluminium construction quick-insert fluid coupler is found to be an important immediate target for future research.
While assumptions are made, this work eliminates an area of otherwise pure speculation within the depot concept by establishing a feasible set of capabilities of the interface as well as identifying specific targets for future developments.

An ambient drop test set-up was developed and used to obtain data about the hypergolic performance of TNO's novel hypergolic propellant combination. Said propellant combination was found to have an ignition delay time of 50 ms before further optimization, making it a promising candidate to replace the current toxic state-of-the-art hypergolic propellant combinations. Furthermore, based on the knowledge gained during the two ambient test campaigns a hermetic drop test set-up has been proposed and is under development.

The use of additive manufacturing (AM) can potentially offer significant benefits to the design and performance of resistojet components but has so far only seen limited research in this application. The goal of this study was to investigate the possibility of 3D printing a functional resistojet heating chamber and the potential benefits that this brings, by using an existing water resistojet as a reference design. To achieve these objectives a prototype chamber was designed, manufactured and tested, while in parallel a numerical model was developed to predict the performance of future designs. The tests were successful in demonstrating the functionality of the AM heating chamber, which allowed the thruster to produce up to 17.3 mN of thrust. However, the results did not allow for the validation of the numerical model, thus preventing its use as a tool to investigate the benefits of using AM for resistojet heating chambers.

The current state of in-orbit refuelling involves launching propellant from the Earth’s surface in a single use refuelling craft, often to transfer hydrazine to the customer. This method is a logical first step however this architecture is not reusable, and it centres around a toxic and carcinogenic propellant. An architecture is proposed where hydrogen peroxide, a green oxidiser useful in both propulsion and power systems, is created from water ice in the solar system and refuelled with a reusable refuelling craft. First order sizing of the craft is conducted showing the viability of refuelling routes from Deimos, Phobos and the Moon. Prototype testing of a propellant transfer mechanism has shown the promise of using a piston-based transfer system, and the results of testing are used to better estimate the mass of a potential reusable refuelling craft.

With hydrazine on the list of "substances of very high concern" by the European Chemicals Agency; hence alternative (storable) liquid rocket propellants are being developed which are less hazardous. During this feasibility study, experiments have been performed with a thermal ignition system in combination with hydrogen peroxide and ethanol as green propellants to achieve ignition followed by self-sustained combustion at the lowest possible power consumption. The experiments have been followed up by a computer model that has been simulated in COMSOL.

After the breakup of the Stratos III vehicle, inertial roll coupling was pinpointed to be the root cause. In order to battle the problem a propulsive roll control system was suggested to keep the roll rate under acceptable bounds. A restartable hybrid rocket engine using nitrous oxide and ABS plastic was proposed. 27 tests have been performed of which 10 were hot fires that had successful ignition. 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. In addition, a preliminary design tool was made and validated using the data of the hot fires. Modelled values for the feed system pressures and thrust output have been proven to be within ±10% of the experimental data.