Mars entry guidance faces a critical challenge: navigating hypersonic velocities within the thin atmospheric layers (120 – 45 𝑘𝑚 altitude) while balancing conflicting objectives of precision targeting, thermal survival, and structural integrity. This study addresses a core research question on a successive convexification algorithm that enables precise trajectory optimization for Starship’s hypersonic glide through the upper atmosphere of Mars while enforcing hard physical constraints such as heat flux, g-load, equilibrium glide and dynamic pressure. By formulating such successive convexification - based framework, the inherently non-convex entry problem is decomposed into iteratively refined convex sub-problems, enabling computational tractability under Mars’ variable CO₂-rich atmosphere. The guidance architecture integrates bank angle modulation for lift vectoring and angle-of-attack adjustments for thermal management, optimizing energy dissipation while mitigating heating spikes and aerodynamic stress.

Simulations demonstrate that the collocation discretization strategy used ensures trajectory adherence within the entry corridor, achieving terminal positioning errors below 3 𝑘𝑚 at 45 𝑘𝑚 altitude. The algorithm’s robustness is validated under ±10% dispersions in initial velocity (4.3 𝑘𝑚/𝑠) and flight-path angle (−15°) from a parking orbit around the planet, with heat flux, dynamic pressure, and g-load profiles remaining within mission-critical limits. Sensitivity analyses reveal that atmospheric density uncertainties induce predictable deviations compensated by rapid convex optimizations. These results align and improve on previous NASA mission data.

The study bridges theoretical convex optimization with operational reality, demonstrating that modern computational guidance outperforms legacy predictor-corrector methods in handling nonlinear dynamics and path constraints. By extending the convex framework with adaptive trust regions and sequential convex programming, the proposed method reduces terminal errors by 40% compared to state-of-the-art approaches (Mars 2020). This advancement not only enhances Starship’s capability to deliver crewed and cargo payloads to predefined Martian coordinates but also establishes a foundation for integrating the hypersonic glide phase with the subsequent powered descent phases. As humanity strides toward sustained Mars exploration, this work underscores the viability of successive convexification as a paradigm for achieving precise atmospheric glide through the Martian atmosphere.

This study investigates the calibration of accelerometer data for the Next Generation Gravity Mission (NGGM), proposed by the European Space Agency. With improved precision, NGGM aims to continue gravity field observations beyond the Gravity Recovery and Climate Experiment Follow-On (GRACE-FO) mission. The mission consists of a satellite pair measuring Earth's gravity field using an onboard laser tracking instrument. To isolate the gravity field signature in these observations, each satellite carries accelerometers to estimate non-gravitational accelerations. This thesis supports the accelerometer calibration process by applying lessons from previous gravity field missions.

A historical review of gravity missions highlights the evolution of scientific and hardware requirements. The study examines accelerometer principles, sources of instrumental imperfections, and existing data calibration techniques. NGGM’s preliminary design includes multiple accelerometers placed away from the satellite’s center of mass, allowing the use of shaking manoeuvres—first introduced in the GOCE mission—for calibration.

A comprehensive model is developed that can generate shaking manoeuvres with varying thrust magnitudes, shaking durations, and shaking frequencies to excite the satellite. This model is used in conjunction with various accelerometer units (two, three, and four accelerometer layouts are considered) and their placement in the satellite's body frame to evaluate the calibration quality against the scientific requirements posed for the mission.

Results indicate that along-track accelerometer placement minimizes non-gravitational acceleration measurement errors due to enhanced centrifugal acceleration from the satellite’s pitch rate during calibration. Furthermore, the along-track placement performs better than radial placement, even though it has the same centrifugal acceleration boost. The suspected cause is the electrode layout of the accelerometer, which boosts the acceleration signal due to the projection of the angular acceleration about the y-axis onto the z component of the linear acceleration. The radial placement of the accelerometers provides no additional signal to the x component of the linear acceleration due to a lack of projection. Lower shaking frequencies improve calibration by accumulating higher angular rates over time. However, due to volume constraints imposed by the laser tracking instrument, cross-track placement may be more favourable. This configuration requires higher thrust levels, as the absence of a pitch rate signal on the cross-track axis worsens the signal-to-noise ratio of the observations, which warrants a revision of the thruster requirements and accelerometer performance. Moreover, more than two accelerometers reduce measurement errors by providing redundancy in the observations. Even with three accelerometers placed in the along-track direction, at least 24 hours of shaking at maximum thrust, as stated by the thruster requirement, is required for effective calibration. Lower thrust or shorter shaking durations would necessitate four accelerometers—two on the x-axis and two on the y-axis. Finally, the accelerometer pair’s arm length is treated as a free variable, as it has minimal impact on calibration performance.

This report provides foundational insight for future gravity missions. Smart accelerometer placement and shaking manoeuvre parameters can improve the measurement quality of the non-gravitational forces and subsequently improve gravity field recovery, which is crucial for tackling the climate crisis.

This thesis investigates how variations in the Earth-Sun distance influence global temperatures, by comparing a simplified model of the solar system with an existing paper from V.V. Zharkova, claiming that increasing temperatures can be explained naturally. Over a 5000-year period, numerical simulations including planetary gravitational influences, solar inertial motion, and Milankovitch cycles, this study looks at distance variations and Earth hemispheric differences in solar intensity due to albedo differences, to asses this statement. The result shows that while orbital mechanics influence the global temperature, Their role is minimal. It should see a slight decrease in temperature, and thus V.V. Zharkova’s research does not represent the actual situation. This offers valuable insight into the relationship between the Earth's orbital mechanics and climate. However, further research into the accuracy of the model is required.

This thesis aims to calculate optimal trajectories from a user-defined Earth-bounded orbit to a user-defined Moon-bounded orbit using a bi-impulse direct transfer ultimately under the influence of a full dynamical model with perturbations, hence reflecting the actual physical environment.

Two tools are developed to achieve this goal. The first tool employs a global optimization algorithm, in particular a Particle Swarm Optimizer (PSO), to find an initial guess within a simplified dynamics model, exploring the user-defined search space. The second tool employs a gradient-based Sequential Linear Least SQuares Programming (SLLSQP) optimizer to refine the initial guess and include the relevant perturbations that act in real life. Additionally, the tools are supported by methods for evaluating the results, providing plotting and analysis tools to make the most out of the obtained solutions.

For the initial guess calculation, the dynamics model includes the point-mass gravity field of Earth and the Moon. The output provides the required ΔV for the transfer and the epochs at which each maneuver should be performed. The SLLSQP optimizer subsequently corrects the initial guess considering the user-specified perturbations, optimizing the time in the first orbit, the different components of both maneuvers, and the time of flight to reach the required orbit in an optimal way.

The capabilities of the tools are demonstrated through several test cases. The first test involves transferring from a circular low Earth orbit (LEO) to a circular near-polar low lunar orbit (LLO), resulting in a total ΔV of 4716.62 m/s. A second and a third test case involving transfers from a LEO or a geostationary transfer orbit (GTO) to an eccentric lunar orbit are also conducted, obtaining a ΔV of 3859.81 m/s when transferring from the LEO and of 1512.95 m/s when doing so from a GTO, corresponding to a decrease of around 60%. The solution obtained from the transfer from the GTO leads to a 4.5% improvement compared to preliminary results found in literature. The forth test comprises transfers from another circular LEO orbit to a high-altitude lunar polar orbit, requiring a ΔV of 3996.44 m/s, being 4.6% higher than the solution found in literature.

These test cases validate the functionality of the code and showcase its versatility in handling various scenarios. In conclusion, the developed tools provide efficient and robust solutions for optimizing direct transfers from Earth to the Moon under the influence of real-life perturbations.

Recent launches of satellite constellations in the Low Earth Orbit (LEO) region have increased the collision probability of existing debris objects with active satellites. Monitoring the trajectories of these debris objects is crucial for Space Situational Awareness (SSA) to prevent the creation of more debris due to unwanted collisions. Much focus is on the LEO regime, with little awareness of the higher Geostationary orbit (GEO) debris population. To date, the explosion of the Russian Ekran 2 satellite in 1978 as well as the disintegration of the Titan IIIC Trans-stage in 1992, have been recorded. These incidents have increased the number of small-sized debris objects in GEO. More unnoticed fragmentation events have been speculated to have occurred, which pose a significant risk of collisions and damage to all weather and communication satellites in use today. The NASA Debris Office confirms that current ground-based radar or optical sensing methods can only be performed for objects of size 1 m and larger, leaving a gap in the precise orbit determination of sub-meter-sized objects in GEO. Moreover, limited observations and atmospheric losses hinder the quality of orbit determination, thus limiting present ground-based SSA techniques. Attempting to bridge this gap in current space surveillance and tracking methods is the objective of this thesis. It evaluates the feasibility of using space-based sensing methods to enhance SSA in the GEO regime. In this research, a satellite in a sub-GEO orbit is deployed to collect in situ radar measurements, which are processed to determine the orbit of a single object in GEO. Different satellite geometries (altitudes and inclinations) and measurement types such as range, range-rate, and direction (azimuth and elevation angles) and combinations thereof have been analysed. A simple grid search optimisation has been performed to assess the feasibility of such a technique and propose a possible favourable observation configuration, which improves the quality and accuracy of orbit determination. It also analyses the uncertainties in the debris state for future epochs to assess the errors in orbit prediction. The limitations of the geometry and measurement model are identified in this study and provided as recommendations and suggestions for further research. INDIGO is hence a feasibility study or a proof-of-concept of space-based debris state observations in GEO. It can be considered a stepping stone towards inventorying the small-sized GEO debris population catalogue and exploring enhanced SSA techniques in the future.

RLV development can be considered as the modern step towards mission design due to financial and strategic decisions. In the past, reusability has been addressed however the level of maturity of the technology, both in terms of hardware and software was not yet reached. There are several aspects to developing a RLV, and these can be categorized into optimization of the LV, optimization of the trajectory, and cost analysis. TO be able to determine the feasibility of the mission, it is not just necessary to develop a suitable configuration, but to also determine the physical feasibility of the trajectory. Several methods exist of which convex optimization is selected. This class of algorithms have risen in popularity in the regime of powered descent guidance, and present a desirable trade-off between performance and computational cost. An already existing algorithm, DESCENDO, for a two-staged vehicle CALLISTO purposed for a mission to a geo-synchronous orbit, is taken as reference. The algorithm is rewritten in YALMIP allowing it to perform more efficiently by saving computation time through creation of multiple controllers based on a discretized burn schedule. A potential candidate for reusability in the future is selected, which is a VEGA variant, considered as a two-staged vehicle with set requirements on the mission and configuration. Through closed-loop simulations, the feasibility of RTLS for a particular mission of this VEGA variant can be studied. The disciplines involved in the study include the launch vehicle optimization, engine sizing, preliminary ascent & descent, and 3-DoF simulations. Previous research at TU Delft on RLV has included the work Rozenmeijer, Vandamme, Van Kesteren, Miranda, and Contant, graduate students of the TU Delft Aerospace Engineering faculty. The work relied on the usage of the TUDAT C++ software environment and based its feasibility or reusability of operations through a cost-analysis. A shift in direction is taken away from cost-analysis to examine at a greater detail the physical feasibility of the trajectory for a nominal candidate RLV. This is done by examining the influence of simulator to guidance algorithm dynamics and guidance algorithm parameters. Moreover, a nominal payload class between 100 and 500 kg is selected to determine the configuration of the vehicle ideal for this mission. To be able to determine feasibility of RTLS, three metrics are considered, which are the final landing velocity, final landing position, and maximumdynamic pressure. The study performs higher fidelity analysis only on the return phase, and as such the starting conditions for descent are determined through a preliminary design process by considering a drag-less ascent. This returns a starting altitude of around 26-30 km, with similar values for starting downrange position, and varying conditions of initial mass and velocity. For the preliminary descent, it is found that the metric of dynamic pressure does not reach more than around 60% of the limit imposed by the VEGA-C, which is similar for other VEGA variants. This coincides with research done with the CALLISTO vehicle. All in all, the best cases for these metrics and one included as an overall best case where candidates for the 400 kg payload class. The selection criteria for best cases of the preliminary descent involved the dynamic pressure, final velocity, and required propellant mass for descent. Moreover, the vehicle optimization results showed that this contained the most variation of vehicle characteristics, and as such it was deemed as a desirable class to work with for its flexibility in design. The best case burnt mass result was selected as the best case velocity required extensive propellant mass to burn for only a less than 7 m/s difference in result, which would not be indicative of what the convex algorithm could achieve due to dynamics involved in the preliminary experiment. This nominal candidate is then tested for various variations of controller tuning parameters combinations, burn schedules, and simulator/guidance frequencies. The results showed a clear desirable region for the final time of just between 300 and 310 seconds for return, favouring shorter burns. The solution envelope for the burn schedule showed gaps in zones of feasibility as well as optimality, suggesting some performance issues with the algorithm due to it failing to find solutions. Nevertheless this envelope is well defined and several feasible solutions existed. The influence of parameter tuning and simulator frequency was studied. It was determined that no set of guidance parameters could give an advantage over the other, but that some values did favour feasibility more. This is somewhat in contrast to the selection of frequencies, as despite the fact there was also a large difference for higher frequency ratios between the Q3 to Q4 and Q0 to Q2, there is a noticeably trend that higher ratios are favoured. Moreover, a local optimal ratio of frequency of 10-1 was also found, and being the same ratio used for the other experiments as well as the CALLISTO study, provides more evidence that this effect is intended. The uncertainties studied are for the initial state variations, errors in reading of position and velocities, and process time delays. It was noted that almost all the errors in the initial state variations shared similar distributions for ranges of values of the metrics. The overall majority returned feasible as well as the large minority of this returned optimal. Of little to no significance was the processing time delay, of which the overwhelming majority returned optimal results, and the rest where outliers whose process time factor where beyond the 3σ limit imposed in the creation of normal random variables. The largest errors arose from the real-time uncertainty in the velocities and position, modelled after pseudo-range errors. Although results showed a high density in the feasible and optimal regions for metrics of final time and position, there was also a high density past the feasible regions. It can be considered that the feasibility of RTLS operations for such a VEGA vehicle is restricted by such errors as expected, but nevertheless results are promising in what can be the main steps to lead to an error analysis study by the use of state estimation techniques and testing various modifications made to the SOCP problem to improve performance.

The Royal Netherlands Airforce (RNLAF) is currently supporting the Feasibility study for Optical Tracking of Orbital Satellites (FOTOS), which aims to create a satellite observation instrument, FOTOS1, for the optical tracking of satellites, which would be a key tool towards maintaining a sustainable use of space. This thesis was tasked with producing an improved image processing pipeline with the aim of detecting more objects and accurate endpoints for orbit determination. Additionally, it was important for the pipeline to maintain the cost and time efficiency of FOTOS1. For time and cost efficiency, data was firstly reduced; the data used originated from the MASCARA instrument in Chile, which takes subsequent exposures at a fixed exposure length from dusk until dawn. The instrument consists of five cameras that virtually covers the whole local horizon. The number of frames to be processed varies depending on the time of year, but is in the range of 23, 500 to 34, 000 images. Since optical satellite tracking requires that the satellites are sunlit, we could reduce the number of images for processing by specifying an observable altitude limit of ℎ = 2, 000 kilometers. This reduced ranges from 4.35% up to 48.38% depending on the day of the year. For higher altitude limits the reduction becomes significantly smaller as the earthshadow becomes less of an effect. To increase the detection performance of the pipeline we combined 50 images to create a single detection image. Several operations were performed to highlight dynamic features among the starry background. The astrometric solution was used to align the images together and subtract subsequent images from one another. These difference images were stacked together by their maximum value to highlight longer streaking features. For the detection, the Probabilistic Hough Transform worked efficiently and returned the correct positions on the image frame. The created detection images did not lead to an increase in the number of uniquely detected objects, but it did extend the altitude range of detected objects and the number of endpoints per unique object. Since the features in the stacks are longer the detection method was able to detect objects up to GEO altitude. Also since the stacks consisted of 50 images, a single detection of a feature could contain data from multiple images and thus contain multiple endpoints. Because of this approach the number of endpoints per unique object was increased almost fourfold whilst only detecting 20% less unique objects in total. Our last task was to determine the endpoints within one track; two novel methods were produced and tested with the aim to use position and discrete time data to create an overall better representation. However, it turned out that fitting noisy data was difficult for the tested regressors (least squares, TheilSen and RANSAC). It oftentimes caused the determinations to be off by 15 or more pixelsruling those results useless for orbit determination. Both the index prediction and index tracing methods had trouble defining the shape due to noise and thus can benefit from a possible iterative approach for data selection. The double index prediction method combined with a TheilSen regressor was selected for the pipeline as it showed to be the best performing method being the most robust combination and had relatively little error compared to the other combinations. When comparing the best performing new method to the existing method, it turned out that the existing method performed better. The endpoint accuracy of the new method was centered around 2 pixels but was more distributed than the existing methods. However the quantity of endpoints was almost doubled. How these results translate into quality of orbit determination is a recommended topic for future work. During this thesis project we therefore produced a new pipeline which can be implemented for instruments with the same observation strategy as MASCARA and is compatible with different lenses and exposure times. It showed proofofconcept that stacking images and through the data reduction simulation allowed for the processing of all the images within 24 hours such that a backlog of data will be avoided.

An array of BlackGEM telescopes is currently being constructed in Chile. The Royal Netherlands Air Force wants to assess the suitability of the BlackGEM telescopes for initial orbit determination of unknown objects in high orbits around Earth. A set of short arc observations of geostationary satellites of no more than 5 minutes is available from a prototype of the BlackGEM telescope called MeerLICHT, which is located in South Africa. Using the classical initial orbit determination methods of Gauss and Gooding and a newly developed circular orbit method, orbit estimates are calculated based on the short arcs. In addition, two methods to reduce measurement error that are specific to the observation system are developed and tested. As a final outcome, recommendations for observation strategies and orbit determination methods are formulated to optimise BlackGEM telescopes for satellite orbit determination.

Mars is expected to become a focal point of exploration (human and robotic) in the near future. Extensive operations and continued human presence on the planet would require a robust space infrastructure. Be it navigation satellite constellations or scientific missions in low Mars orbits (LMO) and Areosynchronous orbits (ASO), every satellite would have a definitive period of operation after which it becomes derelict. At the end-of-life (EOL) the satellite could either be left unattended or dealt with in a sustainable manner. The first option is what has led to the problem of space debris in terrestrial orbits. To protect our access to Mars, proactive sustainability needs to be practised already in the design stages of such missions. This project aimed at providing graveyard orbit solutions in circummartian space for future Mars debris. 200-year period stability was studied for orbits using the symplectic integration technique. Extensive validations were performed and propagation settings were tuned to suit a variety of configurations. A plethora of graveyard orbit solutions were found and presented for orbits in ASO and LMO regimes. For example, it was found that transferring an ASO satellite to 400 km below the nominal orbit altitude would ensure a stability margin of +/-25 km for at least 200 years. The protected zones were found to be safe from debris even for an uncertainty in initial eccentricity of 0.01 and a tumbling satellite. Multiple orbital geometry orientations (combinations of semi-major axis, inclination, right ascension of the ascending node), satellite geometries (various values of area-to-mass ratios) and uncertainties were studied to produce a comprehensive analysis of long-term stability of potential graveyard orbits around Mars, making them attractive for such purposes.

The next decade will see an unprecedented growth in space activity and usher in a new era in space. Together commercial parties, the so-called NewSpace operators, have announced large constellations that will increase the number of satellites in Low Earth Orbit from almost 2,000 satellites to 20,000. Simultaneously, enhanced space object tracking capabilities are expected to increase the number of tracked objects from 23,000 to 100,000. These developments will lead to a strong increase in dangerous conjunction events and the required number of collision avoidance maneuvers as well as more complex conjunction geometries for satellites in Low Earth Orbit. This research proposes a new approach to collision avoidance maneuver planning that produces a set of maneuvers to reduce the collision probability to an acceptable level for complex conjunction geometries without requiring human interpretation beforehand. The approach combines an accurate numerical propagation procedure to take advantage of the small position uncertainties of the Conjunction Data Messages published by the Combined Space Operations Center with a multi-objective optimisation procedure to avoid having to assign relative importance to optimisation objectives beforehand. The approach is tested against four different conjunction scenarios ranging from two objects to involving four objects proving that it efficiently finds solutions for future complex conjunction geometries. It is shown that the new approach finds the global optima in the objective space spawned by collision probability. Applying an intrack maneuver is the most propellant-efficient method for reducing collision probability. In contrast to most contemporary research, this research finds that additional radial and crosstrack maneuver components can offer a significant decrease in collision probability. The quality of the product of many NewSpace operators depends on the service level that is offered: the maximum interval between two satellite visits. Intrack maneuvers significantly change the orbital period and negatively impact the service level. This research investigates extending the problem definition by introducing constellation performance loss as a third optimisation objective. Extending the objective space by an additional dimension increases the required computational effort, but provides an interesting new perspective on the true optimality of maneuvers. The approach uncovers the three-dimensional Pareto front at the expense of 7,875 function evaluations showing both a set of solutions that are propellant-efficient, which were mentioned earlier, as well as a set of solutions that inflict little constellation performance loss. The latter set of solutions has a large radial or crosstrack component and almost no intrack component. It affirms the finding that large radial and crosstrack maneuver components can offer promising solutions for collision avoidance maneuvers. As an alternative for reducing constellation performance loss, this work investigated the application of a second maneuver after the dangerous conjunction situation had passed to restore the original orbit. This concept can offer benefits for simple conjunction geometries, but is infeasible for complex geometries where a large amount of propellant is required to reduce the collision probability. It was also proven that errors in the actuators can cause significant deterioration in the resulting collision probability and constellation performance loss.

Ballistic capture is a transfer method which was first applied in 1990. It allows a spacecraft to approach a target celestial body and enter a (temporary) orbit around it without requiring manoeuvres in between. Ballistic capture is a promising concept, as it is expected to be safer, cheaper, and more flexible in terms of launch windows than a traditional Hohmann transfer. Currently, a computationally efficient method which simulateneously allows for an inisghtful description of the dynamics of the ballistic capture problem remains to be found. A potential solution lies within the field of Lagrangian Coherent Structures (LCS). LCS is defined as a separatrix of regions in a flow with distinct dynamics. It may be possible that LCS around a planet have some correspondence to results found using stable set manipulation, a classic technique for obtaining capture trajectories. In this research three new areas within the field relating LCS to ballistic capture are explored. Firstly, it has not yet been shown what LCS can be found in an area around a planet, without making use of a priori stable set information. Furthermore, it is unclear what the effect is of changing the integration time in the procedure of extracting LCS. Finally, there has not yet been an analysis to show how the LCS relate to stable sets with different number of revolutions n. In this work two algorithms for extracting LCS have been developed. One is based on the simple but efficient computation of the Finite Time Lyapunov Exponent (FTLE). Another is based on the more involved Variational Theory. Both algorithms are validated on a toy problem used frequently in LCS extraction studies, and are then applied to the Elliptic Restricted Three Body Problem (ERTBP). It is shown that LCS around a planet yield resemblance with stable set results. The FTLE-based algorithm is able to quickly and efficiently identify the shape of the stable set. The Weak Stability Boundary, however, can not be extracted distinctly. The Variational Theory-based algorithm yields more distinguishable results for the Weak Stability Boundary. It is shown that large and constant integration times are beneficial. It is shown that extracted LCS form an approximation of the average resulting WSB for all stable sets.

The past decade has seen a continuous increase of Earth observation missions, since they are regarded as an important tool to address global problems such as climate change or disaster mitigation. A commercial trend exists now towards higher resolution imagery, which drives the use of agile satellites. Nevertheless, a disadvantage of agile satellites is the increased complexity to compute the optimal imaging schedule. In fact, the problem can be interpreted as a time-dependent selective travelling salesman problem for which the travel between the cities is also a constrained optimal control problem. Considering the simplifying assumptions in literature, this MSc thesis presents a novel method which approximates the optimal control problem between the targets using an artificial neural network. This approach provides a significant improvement over any previous research resulting in an increase in scheduling performance of around 10%. Considering the high cost of agile Earth observation satellites, this advancement can offer important profit increases for satellite operators. Additionally, a new exact scheduling algorithm was developed based on dynamic programming logic. The algorithm is shown to be able to plan up to one hundred targets for a given restricted number of visible targets at each epoch, while previous research was only able to solve problem sizes of up to twelve targets. Furthermore, the new exact scheduling algorithm is also shown to have unmatched performance for cases of up to twelve targets, such that this should not be considered a difficult problem anymore.

Meteors have been observed by accident from geostationary orbit (GEO). It is of interest for surveillance, meteor science and the updating of empirical models to know what instruments are needed on a GEO satellite to observe meteors and re-entering capsules in Earth ‘s atmosphere. A re-entry simulation code is written to fill the gap of missing observation data of meteors and re-entry capsules. The simulated re-entry data is used to determine what optical instrument is needed to track or detect these objects. The calculations take into account the reflected sunlight, Earth’s thermal irradiance and atmospheric transparency, and the requirement of a sufficient signal-to-noise ratio and signal-to-background ratio. The investigated meteors, with a minimal absolute magnitude of -2.9, and the two investigated Hayabusa and Stardust re-entry capsules can be tracked from GEO for at least 3 consecutive observations with a sub 1 meter aperture diameter optical instrument. A 2 to 5 second tracking time of the investigated meteors requires aperture diameters between 0.1 m and 0.87 m. For 20 second tracking of the re-entry capsules a 0.1 m aperture diameter is needed, which increases to 0.27 m when tracking for 50 seconds. To observe the whole Earth multiple optical instruments would be needed on one GEO satellite. This requires in most cases less than 1 cubic meter of space on the satellite.

The Laser Interferometer Space Antenna (LISA) is a European mission for the detection of gravitational waves in space set to be launched in 2034. The mission will see the deployment of 3 spacecraft in heliocentric orbit keeping a triangular formation with side length of 2.5 million km. Laser beams are exchanged between the spacecraft by means of suitably mounted telescopes (2 per spacecraft), with the objective of synthesizing a very-large baseline interferometer. The interferometric measurements are taken between free-floating test-masses placed inside the spacecraft. Due to the nature of the scientific objectives, the mission requirements on spacecraft-spacecraft pointing precision are exceptionally strict. Moreover, the formation needs to operate in almost perfect free-fall, therefore the solar radiation pressure needs to be continuously compensated for by the on-board thrusters. Gravitational wave signals are measured in the frequency bandwidth of 20 μHz to 1 Hz, requiring the vibrations in that domain to also be eliminated both for the attitude and the displacement. The task is made possible by the gravitational reference system, a complex device that keeps the test-masses from touching the walls of the spacecraft by applying on the latter an external force through μNewton thrusters. This mode of operation is called Drag-Free and Attitude Control System (DFACS). In this thesis we attempt to study and design a DFACS for LISA using a technique called Quantitative Feedback Theory (QFT). The design process starts from the definition of the orbits, the goal orientation of the spacecraft, the sizing of the solar radiation pressure induced disturbances and the derivation of the dynamics of the 19 degrees of freedom to be controlled. Using QFT, the design process is carried out on the DFACS using separation of the dynamics. As a result, analytical equations for the calculation of the LISA commands are derived and the methods to design a control system compliant to the scientific requirements imposed on the sensitivity are shown.

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.

Comets, the sporadic visitors from the outer edges of the Solar System, are considered to hold the key for understanding the formation of planets and the origin of life on Earth. Having spent the majority of time away from the radiative environment of the inner Solar System, the chemistry of the comets has remained unaltered, making them the pristine samples of the matter from the ancient Solar nebula. A mission to bring cometary particles back to Earth enables the examination of the materials in well equipped laboratories and saves the mass of the instruments to be carried on board. As conventional propulsion methods require a large quantity of propellant for this type of mission, the feasibility of using the novel propulsion technique of solar sailing is explored in this thesis. In order to return the comet samples to Earth within a reasonable time period, the orbit transfer is considered as an optimal control problem with constraints placed on the sailcraft’s position and velocity. The Differential Evolution (DE) algorithm was used to search for time-optimal trajectories that minimize the approach distance and the relative velocity with respect to the comet during sample collection. The optimal trajectory obtained predicts the solar sailcraft to reach the comet, collect the samples and return back to Earth in 6.8 years. The time of arrival at the comet was found to match with the comet's perihelion passage, enabling effective sample collection. The outcome of the trajectory analysis, thus successfully demonstrates the applicability of solar sailing to comet sample return missions in the near future.

DIMITRI is a tool that has been developed for monitoring the radiometric performance of optical instruments onboard Earth observation satellites. It includes a reflectance model through which a calibration coefficient is computed which is used for performance monitoring. In this thesis, the uncertainty that is associated with this coefficient is researched and a classification in random and systematic uncertainty components is attempted. Throughout this research it became clear that such a classification is somewhat ambiguous and can be misleading because systematic uncertainties are not always constant. In order to identify such uncertainties, the entire modelling / measurement process with respect to the input variables of the reflectance model, and the algorithm of the reflectance model itself, have been researched for potential uncertainty sources. What became clear is that input variable uncertainties are of both nature; random and systematic. In addition, a sensitivity analysis has been conducted to assess how the calibration coefficient depends on changes in the input variables. Out of these, the Chlorophyll variable causes the highest deviations due to the large uncertainties associated with this parameter. This research has been able to identify uncertainty sources but has not been able to assign any value to those sources. Any subsequent research could use the identified uncertainty sources described here and attempt to quantify their value to yield a total uncertainty that can be associated to the calibration coefficient. This will increase the usability of this performance monitoring coefficient which means that the user will obtain more information regarding the accuracy of the performance degradation of the optical instrument onboard Earth observation satellites.