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 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 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.

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.

This thesis reports on the reconstruction of key orbital elements of the Earth’s orbit around the Sun from a miniaturized temperature sensor onboard of theDelfi-C3 Cubesat, a novel approach never explored before to the best of our knowledge. Delfi-C3 is a triple-unit CubeSat with a mass of 2.4 kg, developed by Delft University of Technology, which was launched on April 28th in 2008. Despite of its required lifetime of less than four months, Delfi-C3 is still operational and has been beaconing data for 8.6 years, which makes this CubeSat mission unique in terms of a continuous record of telemetry data. Recent inspections of Delfi-C3 telemetry data over five years from different temperature sensors have revealed surprisingly systematic patterns of periodic nature with an amplitude of 3.1 K and a period of one synodic year. First speculations associated this behavior to be correlated to the orbit of the Earth around the Sun. An analytical model of the temperature fluctuations has been established which associates the amplitude, phase and period of the observed temperature to the eccentricity, the argument of periapsis, and the period of the Earth’s orbit around the Sun, respectively. This analysis represented the first attempt to reconstruct the Earth’s orbit from satellite temperature measurements. To quantitatively estimate the Earth orbit parameters from in-flight telemetry data, a numerical least-squares estimator has been developed and applied to Delfi-C3 temperature data over the period 2008-2015. Preliminary results by using temperature data from the On Board Computer (OBC) Printed Circuit Board already showed that the estimated semi-major axis, eccentricity, and argument of periapsis of the Earth orbit differ from the true values of less than 1%, 3%, and 5%, respectively. An additional filtering of raw data further demonstrated that the achievable accuracy of the estimation can be refined up to better than 0.05%, 2% and 2.5%, respectively. In this thesis, the sensitivity of those results to initial conditions, filtering schemes and estimator settings are addressed. Furthermore, a refined analytical thermal model is created for Delfi-C3 internal stack, in order to check if a comparable accuracy can be obtained also for other instruments that are less sensitive to external temperature fluctuations, in cases where the internal dissipation requires a better representation of the thermal paths in the internal stack. In addition to the above analysis, the attitude determination systemon board Delfi-C3 is also reviewed by analyzing the telemetry data fromthe AutonomousWireless Sun Sensor (AWSS) and from the four photodiodes located on the deployable solar panels. By adopting an analytical model for the Earth albedo, the possibility of retrieving a three-axis attitude determination from these simple sensor is assessed for the case on hand. This ismade bymonitoring the influence of the albedo termon the determination accuracy for several periods in Delfi-C3 telemetry. It is expected that this analysis will contribute to the understanding of the accuracies of a coarse three-axis attitude determination that makes use of only Sun sensors. Furthermore, the results can be used to determine key factors in the hysteretic process in-orbit, which occur when a PassiveMagnetic Attitude Stabilization (PMAS) is chosen, by comparing the analytical models for the satellite dynamics with the in-orbit attitude obtained from the deterministic method herewith reviewed. This is expected to contribute to a better understanding and improvement of models used in the design of future CubeSat missions.

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.

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.

This dissertation focuses on intra-spacecraft wireless communication as a solution for reducing the spacecraft onboard harness. Despite outstanding advances in aerospace industry, the cost of accessing space is still very high and the amount of engineering work required for spacecraft design and development is enormous. The key elements which increase the development and launch cost of a spacecraft are size, mass, and the necessity of a tailored design for each mission. Studies show that the contribution of onboard harness to spacecraft mass is about 6% to 10%. Any effort to reduce harness can directly result in reducing the launch cost and arriving to a more modular and flexible design. This thesis aims to answer the following questions: 1. What are the problems of onboard wired standards and what are the benefits and characteristics of wireless network onboard spacecraft? 2. Which spacecraft subsystems could benefit most from a wireless onboard communication paradigm? 3. What is the major challenge regarding employing a wireless standard onboard a spacecraft? 4. How can we solve the identified system level design challenge? To answer these questions, this dissertation reviews the existing wired spacecraft data bus standards and major commercial off the shelf (COTS) wireless communication solutions to identify and characterize their architectures. These wireless standards are Wi-Fi, Bluetooth and ZigBee. Categorizing different onboard data types aids to identify a suitable COTS wireless communication solution for each application category. Specifically, sensors of attitude determination and control system (ADCS) can greatly benefit from a low power and low data rate wireless communication solution such as ZigBee, however, the major challenge is conserving energy on the sensors to enable a wireless architecture and achieve an adequate battery life without compromising the system performance. This dissertation proposes two onboard energy managers based on sensor scheduling schemes to tackle the energy conservation challenge. These solutions are tailored to ADCS and aim to reduce the overall ADCS energy consumption without affecting the required accuracy of attitude determination. Both energy managers use similar design elements and decision making algorithms while one of them presents a centralized scheme and the other one employs a decentralized architecture. A unique characteristic of these designs is that the energy management solution is fully integrated with the onboard attitude determination system of the spacecraft. Simulation results show that enabling the energy managers result in total energy saving between 20.9% to 51% (depending on the scenario) without compromoising accuracy of attitude determination.@en

Spacecraft orbit and attitude dynamics have been classically seen as two separated subjects, since the effect of attitude in orbit dynamics was deemed too small to be considered, or simply, modeled as noise in the estimation process for most satellites. In the eighties, a study by [1] showed that for a very large spacecraft, approximately the size of the International Space Station, the effect of the attitude in the orbit dynamics should be considered.@en

Exploring and utilising space to benefit humankind is undeniably a costly venture where the higher the mass of the satellite, the higher the cost to launch. However, the rewards of space-related achievements are indisputable. These achievements range from increasing our understanding of the universe using satellites, such as the Hubble Space Telescope,to improving the predictive ability of weather, climate, and natural disaster phenomena with the help of Earth observation satellites.@en

This thesis provides an innovative architecture for CubeSats and PocketQubes to improve their performance and reliability. CubeSats and PocketQubes are standard satellite form factors composed of one or more cubic units of 10 cm and 5 cm respectively. It is found that the current modular subsystem approach and the electrical interfaces are not optimal in terms of performance and use of technical resources. Reliability is also a concern for CubeSats. Only 35% are able to achieve full mission success. Available literature provides no comprehensive studies on these matters and there is thus a gap in knowledge to be filled. The objective of this study is to identify and quantify the performance and reliability issues related to the physical arrangement of subsystems and the electrical interfaces and to develop an innovative bus architecture which is reliable, flexible and allows for increased performance. The overarching research question is: ”Which satellite bus architecture provides a reliable solution to the needs and constraints of a CubeSat and a PocketQube mission?” @en

In recent years, there has been an increase in the number of small multi-mission platforms such as CubeSats, in an attempt to reduce costs of space missions. CubeSats have been used for different purposes including Earth observation, research and technology demonstration. However, a key technology that is still under development is the micropropulsion system that has the potential to significantly increase the capabilities of CubeSat missions. Micropropulsion has been recognized as one of the key development areas for the next generation of highly miniaturized spacecraft such as CubeSats and PocketQubes. It will extend the range of applications of this class of satellites to include missions that require, for example, orbital maneuvering or drag compensation. An interesting option for CubeSats and PocketQubes is the Vaporizing Liquid Microthruster (VLM) which has received increasing attention due to its ability to provide high thrust levels with relatively low power consumption. The thruster uses the vapor generated in the vaporization of the propellant to produce thrust using a nozzle. The vaporization is usually done by applying power to resistive heaters that could be integrated into the device or externally attached to it. The nozzle is usually a convergent-divergent nozzle that can accelerate the propellant to supersonic velocities. This thesis aims to develop modeling and control concepts for micropropulsion systems to allow the spacecraft to perform maneuvers of position and attitude control. The Vaporizing Liquid Microthruster has been selected due to its characteristics that suit the needs of very small spacecraft. The first part of the research is dedicated to an in-depth literature study of the currently available micropropulsion systems. Those that are manufactured with silicon and MEMS (Micro Electro-Mechanical Systems) technologies have been analyzed and compared in terms of their thrust, specific impulse, and power. A classification in terms of complexity is introduced in an attempt to identify the suitability of the devices for the current trend towards simplifying architectures. The analysis of development levels of different types of micropropulsion systems revealed that although the actual thrusters are significantly developed, the interfacing and integration to other components of the system are still to be further developed. The second part of the research focuses on the characterization and modeling of VLM systems. This is an extremely important step in the development of such systems since a proper model, i.e., one that sufficiently represents the dynamics of the system, is required during the design phase to help, for example, in designing controllers, and also during the operational phase to help reproducing the events happening when the satellite is in orbit. A comprehensive model has been developed using theoretical and empirical relations. The third part of the research addresses the problem of controlling multiple redundant devices allowing failures to occur. This is very important to guarantee the successful operation of VLM systems with many thrusters while performing combined attitude-position maneuvers. A fuzzy control system was developed introducing an automatic rule generation algorithm that allows the fuzzy controller to solve control allocation problems. Finally, the last part of the research investigates the possible applications of VLM systems. An example scenario is considered to analyze the performance required to execute different maneuvers and missions. The key contributions of the work presented in this thesis are related to the modeling and control of Vaporizing Liquid Microthrusters. A comprehensive model of the complete system has been proposed and used to develop control algorithms for individual thrusters and for a set of thrusters. A fuzzy control system has been developed to solve the problem of controlling multiple devices with redundant outputs. Finally, an in-depth literature study and an analysis on the possible applications allowed to put VLM systems into perspective offering a glimpse into the future development of such systems.@en

The repertoire of words in the English language to refer to groups of animals is quite fascinating to say the least – a congregation of alligators, an army of ants, a troop of baboons, a pride of lions, a train of camels, a destruction of cats, an intrusion of cockroaches, a mob of emus, a plague of insects, a drift of pigs, and so on. The intention of using such a wide range of terms is to associate an underlying emotion or meaning to the different kinds of groups. Therefore, without knowing much about choughs or goldfinches, one is more likely to appreciate a charm of goldfinches rather than a clattering of choughs. This thesis is about groups of small spacecraft – characterizing them and enhancing them. The aim of this thesis is to enable charms of CubeSats and prides of PocketQubes.@en

Satellite swarms are a novelty, yet promise to deliver unprecedented robustness and data-collection efficiency. They are so new in fact that even the definition of what a satellite swarm is is disputable, and consequently, the term "swarm" is used for practically any type of distributed space architecture. This thesis poses the proposed definition of a satellite swarm as "a space system consisting of many egalitarian,. spacecraft, cooperating to achieve a common global ,. goal". Methods for designing such swarms are proposed and analysed, as well as the purported robustness and reliability commonly associated with swarms. The investigations show that, like with many systems, it is possible to create a swarm that is less reliable than even a single satellite, yet it is also possible to create one that is more reliable. However, this requires a paradigm shift, as in order to achieve this goal, a satellite swarm's satellites should be built as simple as possible, and this implies without internally redundant systems. The OLFA'R (Orbiting Low Frequency Antennas for Radio astronomy) mission, studying astronomical phenomena at low frequencies* has been used as a test case throughout the thesis, and various technological hurdles required for achieving the OLFAR mission are investigated and solved. This shows that while the OLFAR swarm itself is still slightly beyond current-day technologies, it is not as far out as originally thought, and it could well serve as a prime example of a mission for which a satellite swarm not only would be beneficial, but almost imperative. @en