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.

This paper investigates the cost­-benefit, payload performance and technical feasibility of green storable upper stage design concepts on their potential implementation with existing launchers. Since the demand for reliable, safe, and cost­-effective transport to space is growing, it is necessary to find sustain­able rocket propellant solutions. Green fuels that react hypergolic with hydrogen peroxide as oxidisers can provide distinct advantages in the performance, complexity and reliability when integrated into sys­tems with severe geometrical constraints, such as upper­ stage propulsion modules. This research aims to quantify the cost­-effectiveness of green storable upper stage concepts on the commercial market prospect. This is done through detailed research into design concepts operating on various storable propellant combinations. These various design concepts for different propellant combinations are described by the hypothetical storable upper stage called the “Prototype X”. As an oxidizer, concentrated hydrogen peroxide (HTP), combined with non-­carcinogenic, non­-toxic, and non­ corrosive green storable fuels, can provide reliable and environmentally friendly propulsion solutions. Moreover, green hypergolic storable propellants will lower the operational cost, reduce system complexity and are safer to procure, store and handle. These ”Prototype X” upper stage design concepts will be analysed and optimised on their cost­-per­-kg payload characteristic. Detailed mass and cost optimization analysis was performed for a green storable launcher upper­ stage concept. It was found that HTP together with screened storable fuels shows promising and acceptable performance characteristics in terms of specific impulse and density specific impulse. Im­plementation of these green storable propellants into the upper stage design results in dry mass re­ductions of over 20% compared to the conventional cryogenic hydrolox upper stage design. Although current ’green’ storable propellants can reduce the cost­-per­-flight of the upper stage by 8.9%, resulting from constructive cost reductions in development, manufacturing and (pre­launch) operation phases, their wet mass burden reduces the payload performance by at least 308.09%. This is mainly due to the reduced performance of current ’green’ storable propellants, compared to cryogenic propellants. Therefore, a cross­-over analysis was conducted for the “Prototype X” storable design concept. Here the focus was on finding the required performance of a hypothetical storable “Fuel X” that would make the storable concept economically attractive. Calculations suggest that further development in storable propellant engineering can improve the payload performance of the storable upper stage concept by +37.79% for the same wet mass, compared to the cryogenic design, while reducing the geometrical dimensions of the upper stage by roughly 33%. Storable ’green’ propellants significantly reduce com­plexity and add reliability and safety to the system, potentially improving the payload performance even further. Detailed analysis points out that a hybrid launch vehicle, comprised of a cryogenic first stage and a ’green’ storable upper stage, would be the most cost-­effective expendable launch vehicle in the West having a payload performance of 3.80 𝑘€/𝑘𝑔. Taking into account the added safety, higher reli­ability and reduced complexity of the storable systems, it is expected that the proposed hybrid launch vehicle (either expendable or reusable) will outperform all medium to heavy launch vehicles on the market today. These findings indicate that the green (hypergolic) storable upper stage can be a cost­-effective and reliable solution to future space flight applications.

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.

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.

The human eye has turned itself back to the sky with the commercialisation of the space industry, and a new goal has been set. Setting foot on the Red Planet is the next stage of the human exploration of the universe. The travel to Mars is very lengthy and costly, nonetheless the planet still shows great potential for sustaining human life. To make this a possibility, there is a need for locally sourced energy. The presence of (re-)usable resources on Mars could pave the way to further expand the exploration to an interplanetary scale, and successfully maintain a human presence outside the Earth's atmosphere. The availability of energy will be a key indicator for the success of the human race in the colonisation of Mars. To answer this call for the need to generate locally sourced energy, the design of a renewable energy system was started by a team of students and staff from the faculty of Aerospace Engineering at Delft University of Technology: The Arcadian Renewable Energy System (ARES). The energy system will power the construction and operations of a Mars habitat, to support the livability of humans. The system will use complementary renewable energy sources integrated into a microgrid, to sustainably harvest energy from local Martian environment and resources. To ensure the design will be able to fulfil its purpose, a mission need statement and a project objective statement are generated: Mission Need Statement: To provide renewable energy supply of 10 kW to a Mars habitat. Project Objective Statement: Design a renewable energy supply system, primarily focusing on wind energy, which provides 10kW to a Mars habitat, by 10 students in 10 weeks. Synthesis Exercise (DSE) will last a total of 10 weeks, beginning on the 20th of April, ending on the 2nd of July, with a poster session and symposium. The DSE is in collaboration with the Architectural faculty, where a separate team of students is working on a rhizomatic Mars habitat project as part of an ESA competition, which has an ESA-ESTEC feasibility study proposal incorporated. Due to the multi-disciplinary nature of this project, it is important that the DSE team produces a complete and verified design as the outcome. The Design The design the DSE has come up with consists of two energy production systems, namely the primary and secondary energy system providing wind and solar energy, respectively. In addition the system also consists of a power management and energy storage system.

Future Mars exploration missions require safe and controlled landing on the planet’s surface. Conventional entry, descent and landing (EDL) technologies, such as parachutes and rigid aeroshells, face limitations in meeting increasing demands for heavier payloads, harsher entry conditions, and desired landing locations due to their limited deployment windows or geometry constraints set by current launchers. The stacked-toroid inflatable aerodynamic decelerator (IAD) has emerged as a promising EDL technology which has the potential to enable new and ambitious applications. Unlike conventional aeroshells, it utilizes flexible, high-temperature resistant materials that can be folded during orbital injection and transportation, and subsequently deployed before entering the Martian atmosphere. To address the complex interdependence of design variables and the multidisciplinary nature of stacked-toroid analysis, this research proposes a novel Multidisciplinary Design Analysis and Optimization (MDAO) framework. The framework integrates aerodynamics, aerothermodynamics, structural analysis, mass estimation, and Flexible-Thermal Protection System (F-TPS) sizing with trajectory simulations for a parametrized stacked-toroid. The major design variables are parameterized to trace model responses back to the design space. An additional novel contribution is the inclusion of a smaller torus on the IAD’s shoulder. For aerodynamics and aerothermodynamics modelling, the local-inclination panel method is implemented, separately addressing the continuum, rarefied, and transitional flow regimes using well-established analytical methods and bridging functions. The scalloping phenomenon of the deflected surface is accounted for through semi-empirical expressions and experimentally-fitted correlations, capturing the additional aerodynamic and aerothermal contribution. F-TPS sizing employs a 1D Finite Difference Method (FDM) with harmonic weighted averaging of material characteristics to accommodate abrupt thickness variations. Results from each discipline are compared to experimental, flight, and high-fidelity numerical data, showing consistent agreement under various conditions. All disciplines present mean percentage errors in the order of 10-20% which are deemed acceptable for early design stages. The framework is applied to the ESA MiniPINS study, to demonstrate its applicability to a novel EDL architecture for a penetrating probe. The proposed environment efficiently evaluates the stacked-toroid optimum design for minimum mass, weighing only 3.72 kg whilst complying with mission requirements and optimisation constraints. The design remains robust against variations in entry parameters and atmospheric density, requiring minor adjustments for the penetrator impact speed. The framework enables design space exploration, revealing trends favoring stacked-toroids with low inner torus radii and large numbers of tori to minimize aerothermal loads while ensuring sufficient aerobraking. The inflated radius remarkably results in a global variable that can further simplify the design space to a single input for near-optimum rapid evaluations. The feasible parameter ranges identified through the design space search expedite the evaluation and optimization of stacked-toroids’ multidisciplinary performance, aiding decision-making in early design stages of future Mars missions.

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.

The Dutch/New Zealand based aerospace company Dawn aerospace has developed a 0.5N green bi-propellant propulsion system for use in CubeSat applications called the PM200. This thruster uses gaseous nitrous oxide and propylene as propellants. Currently, the burn time of this thruster is limited to 10 seconds after which cool down of the thruster is required. This burn time limit was set quite arbitrarily as no thermal analysis nor any thermal experimentation was performed on the thruster. In this study the thermal limits of the PM200 and similar thrusters was explored. In doing so designs with two particular cooling methods were analysed in detail: designs with Regenerative cooling and designs with Radiation cooling. The overall goal of the project was to assess the thermal performance of the thruster and to see which of the two cooling methods would be most beneficial to implement. In order to perform the thermal analysis a transient numerical model was created based on the finite volume method in combination with ideal rocket theory and several semi-empirical relationships. The model is able to calculate the temperature distribution and heating rate for several different thruster- and cooling system designs. The model results were verified by comparing the results of the model with the results from commercially available software. A good agreement was found with all results matching within 5-8% or less. Because the model relied on several semi-empirical relationships, test firings were performed using the PM200 and a heat-sink/radiation cooled version of the PM200 specifically designed for this study to determine two empirical constants required to calibrate the model. In total, 325 tests were performed. During these tests, temperature measurements were taken on various locations on the thruster. After the model was calibrated, the model was validated by checking the model results for different starting conditions with temperature measurements from test firings with corresponding starting conditions. It was found that the model was able to reproduce the temperature distribution measured in the tests within 15% for all cases (with the exception of one particular 2s burn case). With the validated model, a number of regenerative cooling channel designs and radiation cooled designs was simulated for a reference thruster similar to the PM200. It was determined that with regenerative cooling the maximum wall temperature could be lowered by up to 23% for the reference thruster. This reduction in temperature was sufficient to lower the maximum wall temperature (to ∼1090 K) below the maximum operating temperature (1150K) of the stainless steel alloy used. A simulation of the PM200 design was also performed and a similar result was found; with regenerative cooling a maximum wall temperature of ∼1020 K is predicted for the PM200. For the radiation cooled designs, the wall temperatures exceeded the allowable temperature of 1150 K for both cases. The reference thruster design reached a maximum temperature of ∼1420 K and the PM200 reached a maximum temperature of ∼1320 K. For radiation cooling to be feasible a different more temperature resistant wall material is thus required. As this would come at an increase in cost, it was concluded that it is more beneficial to use regenerative cooling for the PM200. While regenerative cooling was found to be effective, it was found that for a given type of cooling channel the design parameters have relatively little effect on the cooling performance due to the small size of the thruster. For all designs (with ribs) the variation in temperature was less than 82 K. The main drivers for determining the feasibility of a cooling method are thus not the cooling channel design parameters, but rather the heat input from the combustion. The heat input is mainly determined by the chamber pressure. As a result, the chamber pressure thus determines whether or not regenerative cooling is feasible. It was found that for the PM200 regenerative cooling is feasible as long as the chamber pressure stays below 7.1 bar.

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.

Various different challenges are being faced in the space industry, concerning the increased difficulty level of missions that are planned, down-scaling of the satellites and especially considering the toxicity of fuels used in the spacecraft. To systematically tackle these problems, requirements are set for a ‘CubeSat’ sized satellite for which a sustainable and environmental friendly power and propulsion system has to be developed. Many ‘green’ propellant options are there to deliver the power, of which mainly Hydrogen Peroxide (H2O2) is considered a suitable option, because of its high energy density and reaction products being only water and oxygen. Besides being used in the propulsion system as propellant, the H2O2 could then also be used as fuel and oxidizer to generate electric power in an additional system. The exothermic decomposing liquid is currently under investigation to provide thrust and electricity using catalyst materials, which experiences degradation of performance over time. Therefore, different long-lasting solutions have been experimented with, based on thermal decomposition of the H2O2 and improved electro-reduction methods. This way, a green solution would be present to undertake various missions using small spacecraft that need around 1N of thrust, while simultaneously delivering up to 60W of extra power during its operation. Multiple Proof of Concept (PoC)’s were therefore developed to understand the processes and acquire the voltage, current and temperature data, generated and required by the systems. The energy required for ignition of the H2O2 using various resistance wires was recorded visually and with thermocouples. A drop test study was done for better understanding first, after which a safe injection setup was constructed where different conditions could be looked into. Various fuel cell structures were tested in a different experimental setup before this, to try and reach the extra power the thermal ignition requires. A demonstration of the various ways this system could be optimized was then done, mainly by changing input parameters. Cooperation of the systems was looked into by combining them and checking feasibility of incorporation in a CubeSat. Several electrochemical related properties prove to have large effects on the power density and potential and a maximum of 4.5mW was reached. Similarly, the ignition conditions of the propellant will have to be looked into more, since the 24.8W currently needed for small mass flow rate decomposition, is excessive already. This makes the incorporation in a combined system design difficult, but worth investigating, because of the many possibilities it brings forth. Based on the study performed so far it is concluded that providing enough sustainable electrochemical power for an efficient thermal decomposition is challenging, but feasible. The required conditions and achieved results with H2O2 will be presented in this research as well as recommendations to future work on this upcoming field of study. Utilizing a second propellant such as ethanol would be the next step in upgrading the system performances and lower the ignition energy required, whilst increasing the electrical energy obtained from the fuel cell.

KNO3-Sugar propellants, also known as rocket candy, form a group of simple, cheap and safe solid propellants that are used extensively in student and amateur rocketry communities. One of the most frequently used compositions is KNO3-sorbitol (KNSB) with a typical 65/35 ratio by mass. This composition is used extensively by Delft Aerospace Rocket Engineering (DARE) for small experimental launches including the record-breaking flight of Stratos I in 2009 to 12.3 km altitude. However, KNSB propellant quality has been inconsistent: propellant density is occasionally below 85% in combination with large surface defects. Besides a very high grain rejection rate this has resulted in several explosive failures of new experimental motors in recent years.

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 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.

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.

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.

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.

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.

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 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 the boom of space activities in recent years and the increasing concern for clean technologies, green space propulsion is becoming a topic of interest within the space sector. Unfortunately, most of the propellants used for space activities tend to be highly toxic, corrosives, and even carcinogenic. Toxicity is especially critical for the hypergolics, which in most cases are the preferred propellant due to their simple application as they ignite upon contact. Throughout the history of space propulsion, some chemicals have been left aside due to their low performance when benchmarked against these more powerful hypergolic propellants. However, the trend of "green'' technologies is pushing research towards green lower performance propellants. This is the case of hydrogen peroxide, H2O2, which was left aside due to its reduced effectiveness and the complexity to exploit its energy. Nonetheless, the work with hydrogen peroxide has recently acquired a new focus. Its green perspective has instigated large projects by the European Commission, such as GRASP, and approaches to exploit its energetic content are a topic of research. Currently, the use of catalysts to decompose the chemical is the usual procedure. However, their complexity and effect on performance raise some problems. In this thesis, a new approach through thermal energy supply aims at avoiding the catalyst approach to decompose hydrogen peroxide and overcome their related problems. The feasibility of this approach is tested and characterized for different combinations of temperatures and concentrations of hydrogen peroxide. In order to characterize this, a first look into the procurement of the chemical is done. The different approaches presented lead to a successful refining technique, which is now patented. This innovative approach allows acquiring hydrogen peroxide in concentrations in excess of 99% in the span of 1-2 days for a reduced price. The simplicity of this technique is unrivaled to the stringent regulations surrounding its commercial availability. The resulting solution is investigated in terms of stability over time and concentration. The final selected concentrations for this study are 75, 80, 85, 90, and 95% H2O2. The thermal activation of hydrogen peroxide is first investigated. Through an open-air drop-test study, the droplets of H2O2 are generated and precipitated over a heating plate at various temperatures. The combinations of temperature and concentration, which lead to decomposition, are recorded and analyzed through a set of thermocouples and a high-speed camera at 6400 fps. For the valid combinations, the analysis returns their maximum temperatures and decomposition delay times. Values of up to 800 celsius for 90 and 95% H2O2 are achieved, and low decomposition delay times below 100 ms were measured for heating plate temperatures of 250 and 270 celsius. This information is put against the decomposition with catalyst MnO2 for comparison, showing a more significant temperature released for the thermal approach, but a slower decomposition. With the valid decomposition combinations, an ignition study is conducted with ethanol as the fuel and hydrogen peroxide as the oxidizer. Through a drop test on a heating plate, the mechanisms leading to the ignition are researched. The tests showed that ignition is pseudo-hypergolic and can achieve increases in temperature of up to 700 celsius for 95% H2O2 and ethanol, reaching maximum temperatures over 1000 celsius at points. The tests showed that combinations with 90 and 95% H2O2 result on average ignition delay times below 100 ms. The conducted test cases showed that the ignition temperature profiles can be correlated with the recorded filming and that the lessons learned can be applied for a future combustion chamber. In order to test performance-boosting techniques, the gelling of the fuel was done and tested. The results are promising, the ignition of ethanol with 1.8% wt. of hydroxypropyl cellulose reached such temperatures that melted a thermocouple, proving the ability to raise the performance. Lastly, the gel sample was characterized in terms of a rheology study by performing a strain sweep to find its yield point. Moreover, a shear rate analysis was performed, and its thixotropic nature was tested through a hysteresis loop. The effect of temperature was also analyzed and fitted to an exponential Arrhenius-like form.