With the increasing amount of plastic debris in Earth’s waters, concerns about its impact on aquatic ecosystems continue to grow. While satellites like Sentinel-2 have demonstrated the value of remote sensing, the lack of a dedicated mission limits detection to large-scale plastic accumulations and fails to capture their dynamic behavior. This thesis investigates the feasibility of using deployable optics in Very Low Earth Orbit (VLEO) to improve spatial and temporal data on marine plastic. Focusing on system volume—critical to mission viability—and building on the Deployable Space Telescope under development at TU Delft, satellite volume was modeled through Monte Carlo simulations. A volume envelope as a function of orbital altitude was generated, accounting for design and environmental uncertainties. Compared to a traditional Low Earth Orbit telescope, the VLEO concept achieved an 88% reduction in volume. This significant improvement highlights VLEO’s potential for advancing Earth observation missions across various domains.

The electric-pump cycle is a rocket engine configuration that uses an electric motor to power the pumps instead of a turbine. This offers several expected advantages such as simpler design, lower development costs, and easier restartability, but comes with reduced performance compared to conventional cycles. Previous research has primarily compared the electric-pump cycle to the gas generator cycle and has been limited to direct comparison. To extend the previous research a Rocket Cycle Analysis Tool was developed called RoCAT. It models the last-named cycles as well as the open expander cycle for a broad scope of thrusts, burn times, and chamber pressures, and several propellants. In addition, RoCAT optimizes several inputs for each cycles individually for a fairer comparison. Besides analyzing the electric-pump cycle's current performance, this research also estimates its performance in the future based on historic trends in its key technologies like the battery and electric motor.

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.

A miniaturized instrument was developed to enable in-situ measurements of the radiation environment at the Moon. The instrument is designated to function as the science payload for the first mission of the Lunar Zebro nano-rover. Characterization and testing of the Floating Gate Dosimeter (FGDOS) was advanced with an emphasis on its utilization as the core detector for this payload. A prototype of the Radiation Payload was designed, produced and tested. Radiation environment prediction and analysis was performed for various mission phases using SPENVIS and OLTARIS. A radiation transport model of the payload was prepared, as a foundation for more extensive simulations in the future.

This thesis highlights the need for further research and development of the FGDOS technology, including experimental mapping of the complete envelope of FGDOS sensitivity based on expected mission radiation environments. Additional noise reduction measures and thermal characterization at mission conditions are needed to iterate and improve the payload design such that it can be made flight-worthy.

The need for Earth observation telescopes with high spatial and temporal resolution is constantly increasing, however, current state-of-the-art telescopes are costly. The Deployable Space Telescope project at TU Delft aims to provide a solution with an Earth observation telescope that has the same optical performance as today's best ones, but at a reduced cost. A baffle is required to surround the telescope to provide stable thermal environment, limit stray-light, and protect the optical components from debris. A deployable baffle consisting of pantographic arms has been designed that has only one degree of freedom, therefore the whole structure follows the configuration change if one angle is changed in it, and the diameter and height change happen synchronously, successfully reducing the required number of actuators. The proposed thermal solution decreases the thermal gradients and temperature extremes within the baffle considerably, and successfully shifts the overall temperature of the telescope towards colder regions.

Some propellants currently used in space propulsion have the disadvantage of being dangerous to humans and the environment. A wellknown example of this are hydrazine and its derivatives which are toxic, corrosive, and carcinogenic. These disadvantages have led to a search for less hazardous storable liquid propellants, often referred to as green propellants. A green propellant is safer to use and handle and will therefore bring down the costs related to production, storage and handling. The main reason for the use of hydrazine is its hypergolicity with common oxidizers like NTO and nitric acid. Hypergolicity is the property that a fuel and oxidizer ignite spontaneously when brought in contact without the need for an external source. This property is beneficial since it eliminates the need for an ignition system, thereby making the propulsion system more simple, reliable, and cheaper. It is therefore desired that a replacement propellant also shows this property with common oxidizers.
During this thesis two methods are explored with the goal of creating a green hypergolic propellant combination. One based on catalytically enhanced ethanol and high concentration hydrogen peroxide. The other based on a pyrophoric liquid that is added to ethanol. There have already been efforts made before by adding catalyst particles or a strong reducer to a hydrocarbon fuel like kerosene or ethanol. The problem with using catalyst or strong reducer particles is the difficulty of creating a homogeneous mixture. Due to the liquid nature of the parent fuel, the particles will start to separate from the fuel and sink to the bottom of the container. To overcome this problem an organic gelling agent is added to the fuel to increase its viscosity, thereby increasing the sedimentation time. This will not only increase the shelf life of the fuel but also decrease the vapor pressure, making it less flammable and safer to work with, and reduce storage problems like propellant sloshing and spilling. Increasing the viscosity also makes it more difficult to transfer the fuel through the feed system and achieve proper atomization. However, by applying shear force on the fuel it shows shear thinning behavior, decreasing the viscosity close to that of the parent properties making it easy to use in existing propulsion systems designed for liquids. A second effect of the organic gelling agent is that due to its energetic nature it participates in the combustion. Therefore, the amount of catalyst needed can be reduced to a negligible amount while still achieving good performance properties like ignition delay time. This is verified by means of a drop test.
As a second approach, instead of adding solid catalyst particles a pyrophoric liquid is added to ethanol. A liquid allows for easier mixing and creating of a homogeneous mixture resulting in a longer shelf life compared to using catalyst particles. The pyrophoric liquid also fully participates in combustion resulting in increased performance compared to catalyst particles that do not combust. Due to the reactive nature of the pyrophoric liquid this propellant formulation is expected to be able to achieve hypergolicity with multiple common oxidizers. This results in a versatile system which can be used in current propulsion systems without extensive modification required. By eliminating the need for an ignition system, the propulsion system is simplified and its reliability is increased.

An important topic in the field of rocket propulsion is the development of a new green advanced hypergolic rocket propulsion system. The current hypergolic propellants have one large drawback, namely the toxic properties. The injector has a crucial role in the performance of a liquid engine. Therefore, the next step in the development of this novel propulsion system is the development of the injector. The purpose of this master thesis is to design, manufacture, and test this novel injector. The injector is designed to operate in monopropellant and bi-propellant modes, which is supported by the new green hypergolic propellants. In total, four different configurations of the injector are developed. Different testing objectives are defined for the experiments to test all the injector configurations. These testing objectives are atomization behavior, atomization performance, the mixing process in bi-propellant mode operation, and atomized flow in combination with the heating element.

The supersonic split line (SSSL) nozzle has been considered for years in the use of solid rocket motors. The SSSL nozzle is an attractive alternative to the more conventional submerged movable (SM) nozzle. The SSSL nozzle can save up to 43% on nozzle mass and 25% on the thrust vector control (TVC) system mass. The SSSL nozzle also has the benefit of the amplification factor, which is a result of the SSSL nozzle turning the flow in the supersonic region. The shock waves formed deflect the flow more than the mechanical deflection. Only limited research is available on the SSSL nozzle. This thesis aims to expand on the knowledge of the SSSL nozzle in cooperation with Nammo Raufoss. The aspects that are researched in detail are as following: The effect on the amplification factor when moving the split and changing the expansion ratio of the nozzle. Furthermore, to determine how well the SSSL nozzle compares to a more traditional SM nozzle, a comparison between the SSSL and the SM nozzle is performed at equal performance. This comparison has a focus on the mass differences. The performances of the SM nozzle and the SSSL nozzle are determined by simulations with computational fluid dynamics (CFD). From the CFD simulations, the amplification factor of the SSSL nozzle was determined. In order to compare the SSSL nozzle and the SM nozzle a mass model is developed. The model uses the calculated performance from the CFD simulations to calculate the mass difference between the two nozzle and TVC systems. The results show that the amplification factor increases when moving the split further down stream. Furthermore, when increasing the expansion ratio the amplification factor decreases. For the SSSL nozzle with a total expansion ratio of 12 and the split location at an expansion ratio of 1.75, the maximum amplification factor is 1.57. Additionally, a linear relationship between the amplification factor and the ratio of the expansion ratio at the split to the total expansion ratio of the nozzle is observed. The SSSL nozzle experiences increasing thrust and Isp losses while vectoring when the split is located further downstream, while receiving only a slight increase in the amplification factor. The results from the mass model show that the SSSL nozzle is on average 37% lighter than the SM nozzle for a vectoring duty cycle of 0.25. To conclude, for an increasing ratio between the expansion ratio at the split to the total expansion ratio of the nozzle, the amplification factor increases. Additionally, the SSSL nozzle is the lighter option when only the aerodynamic effects are considered. It was also seen that the temperature and heat flux around the split area was high, to overcome this the SSSL nozzle might need additional thermal protection which could potentially lead to a higher mass. It was observed that the largest mass difference between the SSSL and SM nozzle system was found when the split of the SSSL nozzle is located at an expansion ratio that gives the highest possible Isp.

The heat fluxes in (liquid) rocket engines can go up to high values and they need active cooling to prevent the chamber wall from failing. Transpiration cooling is identified as a way to achieve lower wall temperatures than commonly used regenerative cooling and regenerative cooling with additional film cooling (referred to as film cooling from now on). This can lead to higher performance of the engine. However, transpiration cooled engines are not in use today and this is mainly attributed to problems with the required porous wall materials. Additive manufacturing (AM) is a promising solution for these material problems. Better cooling is especially useful for Inconel additively manufactured rocket engines as such engines experience higher wall temperatures due to the low thermal conductivity of the material.
This thesis has two purposes. Firstly, an analysis was performed to see if transpiration cooling actually performs better than regenerative and film cooling in total engine performance. Secondly, it was investigated if AM can be used to create the porous walls required for transpiration cooling.
To compare the cooling techniques, a simplified model for each cooling technique was developed. These models were verified and validated using data from literature. A new (transcritical) film cooling model was created by combining three existing film cooling models. The wall temperatures obtained from the three cooling methods were compared by applying them to a reference liquid rocket engine with either Inconel or copper as wall material. Subsequently, the losses in specific impulse and dry mass were determined. Then, a delta-v calculation for each cooling method was made to objectively compare them.
The conclusions on the comparison of the cooling techniques are that the regenerative cooled engine reaches wall temperatures above the material limits. Therefore, it is not a feasible to use this cooling method for the reference engine. Film and transpiration cooling can both achieve temperatures below the limit. Transpiration cooling requires less coolant than film cooling to achieve the same temperature. This will result in lower losses in specific impulse compared to film cooling. However, to achieve these low coolant mass flows, thick chamber walls are required to achieve the specified pressure drop over the wall. When comparing transpiration cooling to film cooling on the total delta-v achieved, it is found that the Inconel chambers perform better than the copper ones. However, it depends on the pore size if transpiration cooled engines outperform film cooled ones. A smaller pore size is better.
Additionally, it was found that the pore sizes producible with AM are an order of magnitude larger than required. Therefore, experiments were performed on the pressure drop over AM porous walls with different geometries. With these experiments, it was found that a new porous wall geometry made using AM techniques has a lower pressure drop than the geometry used in the calculations, being 1.32 lower. A geometry designed to achieve an as large as possible pressure drop increased the pressure drop 24.9 times. However, this geometry does not achieve a uniform coolant injection required for transpiration cooling, so further research is required.

The need for higher spatio-temporal resolution Earth observation increases rapidly. To fill this need, the Deployable Space Telescope (DST) project aims to make a light-weight, low-volume deployable telescope. In doing so, the achievable ground resolution is high while the cost per telescope stays low. This allows for the DST to be used in constellations, thus effectively achieving a high spatio-temporal resolution. One of the issues of this design are the relative translation and rotation of the secondary mirror due to temperature fluctuations. This thesis work focused on first finding these movements by identifying the temperature variations of the secondary mirror support structure using ESATAN TMS simulations, and subsequently designing a system to keep these movements within the allowed budgets. The end-result is a novel design in which all displacements are measured by means of 4 Displacement Measuring Interferometers and corrected by means of 4 linear piezo actuators.

The German Aerospace Centre (DLR) has in the past years been performing test runs with their experimental sub-scale combustion chambers to measure and investigate the effects of unstable combustion. Some of these investigations are focused on the heat flux to the chamber walls, changes in which can cause a significant shift in chamber wall strength and even melting, and can thus lead to structural failures. As a continuation of research regarding this phenomenon, the DLR combustion chamber version D (BKD) is used to investigate changes in heat flux, and its test data have been made available for this research. Using the test data, finding a relation between measured steady state heat flux and controlled chamber conditions became feasible, and changes in heat flux related to combustion instabilities can be identified. However, although these changes in heat flux are often observed and are described in literature as a known effect of combustion instabilities, there are currently no validated explanations as to why these effects take place. In an effort to provide this explanation, the results of a previously completed investigation regarding the DLR combustion chamber version H (BKH) are evaluated, as they allowed determination of a relation between combustion length and magnitudes of acoustic pressure oscillations. In this research, the results obtained for combustion length are converted to a model regarding varying heat flux profiles and modified as to implement them for BKD. Finally, the model predicted increases in heat flux during time periods when increases are observed experimentally, showing agreement with the experimental data. Besides this, a conversion in the model allowed reconstruction of a localised heat flux profile, which showed an increase in local heat flux near the injector face plate. This matches effects described in literature and observed in the experimental set-up after a test run with combustion instabilities. With this validation of the model, the theory developed and tested in this thesis work could prove a significant advancement in the understanding of effects of combustion chamber instabilities and their impact on the engine integrity and performance.

This thesis studies single events (SE) in an SRAM from the identical satellites Sentinel 2A and 2B, in order to find links between SE, space weather, satellite operation and device properties from time and position of event occurrence. A study of theory on single event rate prediction and space radiation environments lies the foundation for the subsequent data analysis. A focus is placed on aspects relevant with the Sentinel 2 satellites. Using Python libraries, raw satellite housekeeping data is pre-processed to make available time and place of event occurrence. Then, a moving-average filter is applied to reveal the long-term behavior of SE position and rate. The filtered SE rate data shows a linear trend, attributed to solar cycle and device degradation. Further, there are features attributed to memory hot spots, South Atlantic Anomaly (SAA) semi-annual behavior and geomagnetic storms. Future event rate data and data on affected memory addresses will help to further distinguish these influences. Filtered event positions track the drift of the SAA, but contain suspicious differences between the satellites. These are attributed to memory hot spots as well. Spectral analysis reveals that the only short-term influences in the unfiltered data arise from the orbits of the satellites. Within the SAA, event rate is similar for eclipsed and sunlit, but outside of the SAA, event rate is approximately twice as high in sunlit. This difference is currently unexplained.

Geolocation of Radio Frequency (RF) emitters, is finding the geographical location of said emitter. This thesis explores the concepts of geolocation, specifically, Space-Based geolocation of earth-surface emitters by means of measuring the Angle-of-Arrival using a Uniform-Rectangular Antenna Array. Initially the literature is explored, the significance of Geolocation is made clear and the research outline and scope is defined. Afterwards, Angles-of-Arrival finding algorithms are investigated and implemented, specifically the subspace algorithms: "MUltiple SIgnal Classification" (MUSIC) and "Estimation of Signal Parameters via Rotational Invariance"(ESPRIT). These algorithms are thoroughly explained, a model is obtained and implemented into a Simulation framework. The Simulation Framework is used to create a Graphical Tool used to calculate performance figures of Space-Based Geolocation. By using the obtained Simulation, a sensitivity analysis is produced. This sensitivity analysis observes the effect of design parameters of a satellite platform and geolocation payload on the geolocation performance. This sensitivity analysis is used to find the optimal approach when designing a geolocation payload. This approach can be used to synthesize a geolocation payload design that maximizes performance.

As the new space movement develops, so comes the request for cheaper space solar panels with a shorter delivery time. This master thesis investigates the possibilities of using a Semi-flexible terrestrial solar panel in a low Earth orbit space environment. A Semi-flexible solar panel is a sandwich of polymer films with the solar cells in between, these layers are laminated to create the end product. This, for space, new process is applicable to all major solar cell types. It enables a wide range of possible designs and could result in a drastically lower price, shorter production time and a lighter and smaller end product. By using system engineering methodology, the risks of Semi-flexible solar panels in space are identified after which the major risks have been tested. The risks investigated in this report are: outgassing, temperature cycling, vacuum UV radiation, charged particle radiation and the stiffness in deployed and stowed position. The results show no red flags for a seven-year 600 km Earth orbit. A transmission degradation of 10% is observed, resulting in lower power output for the selected material. When entering higher orbits, the amount of radiation leads to delamination and potentially no power output. When comparing the Semi-flexible solar panel to a conventional solar panel, a price difference of factor 4 and a power output difference of 33% is expected. The vast price difference shows the potential for the concept, but further investigation is needed to find out if the transmission degradation could be mitigated and whether designs comply with stiffness and vibration requirements.

This thesis shows the mass reduction of a mechanical bracket, without compromising on thermal, EMC and structural aspects. The focus was on analysing thermal aspects related to bracket design. Models were made to show the deflection and solder layer stresses due to CTE mismatches between the bracket and PCB. A thermal heat transfer model was built to analyse the temperature distribution over the bracket and PCB. Using these models, the performance of different materials can be analysed. The most promising candidate for bracket design was aluminum silicon. This material was used in the design of a new bracket. The resulting bracket has a mass of 74 g, which is only 8 % of the mass of the original copper bracket.

Teleoperation allows the use of human intelligence and decision making in remote tasks which are too dangerous for humans to perform. Technologies such as force feedback and haptic guidance have shown to increase task efficiency during teleoperation. In an unmodeled environment, sensors provide input for haptic guidance or present extra information about the environment to the user in order to make decisions and to perform the tasks. These sensors come with inherent errors and uncertainties which propagate through the teleoperation system. The absence of knowledge of these errors has been shown to cause deterioration in the task performance. These errors can further cause the application of forces on the environment by the robot without the knowledge of the user while using haptic guidance. Thereby, the strategies being used to increase task performance can have some adverse hidden effects. Thus, it crucial to have an understanding of the behaviour of the errors and uncertainties in the system. It is considered critical for making decisions about how the robot system can be controlled and used to manipulate objects remotely.

In this thesis, a novel framework for estimating the uncertainties in a vision-aided teleoperation system in real-time is introduced. The uncertainty estimate can then be used by the control system or communicated to the user. Methods to use the uncertainty estimate for haptic guidance and for user display are proposed. Furthermore, the thesis analyzes the behaviour of the uncertainties in the system and the sensitivity of the system to individual component errors. It evaluates the uncertainties in individual components of the system and implements an uncertainty model for each of them. It then provides a method to propagate these uncertainty models through the system. This results in a final uncertainty estimate in the frame of reference of interest for the task. Experiments were performed to validate the component uncertainty models, the propagation method, and the system as a whole. Additionally, an inverse of the propagation method is also conceived so as to obtain the component accuracy specification from system uncertainty requirements. This can be used in the design of future teleoperation systems.

Unlike a liquid rocket-engine, it is not possible to actively throttle the overall combustion of a solid rocket-motor. It is however possible to alter the thrust of an individual thruster by varying its throat area. The throat area can be altered by inserting a pintle of a conical or other shape into the nozzle throat. A setup which consist of multiple thrusters each with a pintle to alter its throat area, and which are connected to a single solid-propelled combustion chamber, can be used for attitude control of a rocket or spacecraft.

The goal of this thesis was to investigate the flow in a single thruster of such a system. For this purpose, a simple 1D simulation was developed. Additionally, multiple CFD simulations both static and dynamic were setup. The results of the simulations were then analysed in detail and compared amongst each other and with externally obtained test-data.

The influence of using different chemistry model and different chemical kinetics schemes on modeling combustion inside a rocket engine igniter are investigated. An igniter for a LOX/CH4 rocket engine is designed and a CFD model is developed which is used to model the designed igniter. Verification and validation methods are discussed.

There are several benefits of using autonomous sensors in spacecraft. Avoidance of wired connections reduces cost, mass, and increases the flexibility and reliability of the system. The impact of wire reduction can be significant, especially for small satellites with many sensors, like temperature and sun sensors. Previous research has already focused on wireless intra- spacecraft communications. This research tests the self-powering capabilities of a system based on a COTS thermoelectric generator connected to a Bluetooth Low energy communication system, with a built-in controller and temperature sensor, and a power management interface. The system will be considered as a candidate for an autonomous temperature sensor in a future PocketQube mission of the university.

Controlled temperature differences can be achieved in a test environment, allowing the measurement of the generator power capabilities. It is tested that the system requires, for operation, a minimum temperature difference of 2.31 degrees between the extremes of the thermoelectric generator. It generates a peak power of 234 μW for that difference. In addition, the voltage difference obtained of 35.5 mV exceeds the minimum voltage required by the power management subsystem to be used. The power management sub-system consists of an ultra-low power converter that provides an output voltage of 4.1 V and a measured power efficiency of 32 % Moreover, thanks to the management of the Bluetooth sleeping modes, with the built-in controller and several operational amplifier comparators, an average power consumption of 5 μW is required during operation. The case studied would allow measuring temperature and sending the data over a Bluetooth link to the on-board computer every 16.2 seconds

It is concluded that the technology, based on COTS components, can be implemented and considered as the first step for a fully autonomous sensor with thermoelectric power generation in small satellites. Its implementation may provide substantial advantages for remote or/and locations where wiring is difficult to integrate. The tested performance values provide the foundation to develop the technology further.