Accurately tracking the trajectory of launching vehicles during the boost phase is crucial for a variety of reasons. First, an accurate burnout point location and  velocity of the launch vehicle will facilitate the estimation of its trajectory during the following phases of ballistic flight and the impact location. This has important implications for safety purposes. The first stage of many launcher systems such as the Russian Soyuz or the Chinese Long March rockets commonly impact on land, sometimes on populated areas. Furthermore, the proliferation of suborbital rocket launches presents a hazard if regulations are not properly followed. Therefore, it becomes critical to estimate the impact point of any vehicle that reenters the lower layers of the atmosphere beforehand, during the boost phase, in order to prevent personal and material damage. Second, a critical aspect in any space mission is to insert the payload into the required orbit. An accurate tracking of the launcher system during the boost phase allows carrying out early orbit determination in order to successfully complete the payload insertion. In this Master Thesis, an analysis and performance comparison of different tracking algorithms for launching vehicles during boost (e.g. launcher systems or suborbital rockets) has been carried out. This type of tracking problem has several difficulties that must be overcome. First, the unknown thrust profile – thrust magnitude and orientation – of a launching vehicle makes difficult to develop motion models that can accurately describe its behavior. Second, the observations of the plume of the launching vehicle obtained from two line-of-sight passive sensors located in geostationary satellites are used. The way these observations are obtained does not allow to accurately measuring the initial state and trajectory of the vehicle, complicating the initialization and the trajectory estimation processes of the tracking filters. Finally, the nonlinearities present in the system and measurements models of the tracking filters compel us to devise linearization schemes or the implementation of alternative filtering methods.The work presented in this document has been developed at the Military and  Security department of Airbus Defense and Space, as a continuation of an ongoing project. To solve the aforementioned difficulties of the tracking problem, a tracking system – EKF-Tool – was developed at ADS previous to the start of this Thesis. Nevertheless, this tool shows several limitations that hinder its performance, even preventing from successful tracking. In this Thesis, the limitations and potential areas of improvement of the EKF-Tool have been analyzed. Based on the results of this analysis, an alternative tracking algorithm – PKF-Tool – is introduced with the aim of overcoming the limitations of the EKF-Tool. The tracking performance of both tools has been tested using several indicators: stability, consistency, credibility, accuracy, precision, and versatility. 

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.

Understanding the role of ocean swell, monitoring and modelling it is critical for a diverse range of oceanographic research, coastal management initiatives and for better weather forecasting and climate modelling. Currently, swell measurements are obtained through in-situ measurements, wave models, and satellite measurements. So far, swell wave parameters have been retrieved from orbit using optical instruments, imaging radars and wave spectrometers. The potential to retrieve swell wave spectra from radar altimeters was recently demonstrated using fully focused synthetic aperture radar (FFSAR) signal processing. Sentinel-6 (S6) would be a good candidate for investigating the potential of swell parameter retrieval, however, no study has been performed regarding the impact that the on- board data compression/ waveform truncation mode of S6 could have on the retrieval of swell parameters. Swell wave period estimates were thus derived from the raw and compressed S6 FFSAR waveforms collected over a study area in the Channel Islands of California during a 22 month- period. A performance analysis was carried out comparing the period estimate pairs from a given timestamp with the records obtained by a buoy in the study area at the same time. The study demonstrated the presence of differences in the performance of swell period retrieval for the different operating modes of S6. It was found that swell period estimates can be derived also from the compressed waveform signal, obtaining however a lower accuracy compared to raw waveforms.

The military use of space brings interesting challenges for both technology and operational aspects. The Netherlands armed forces are a small player within the military world. However, in 2013 the beginning of a space program was started by the Royal Netherlands Air Force (RNLAF). One of the key mission statements stated: “show the military relevance of Nano Satellites". A feasibility study was conducted and the realisation of the first satellite, the BRIK-II,was started. One of the payloads of this satellite is a store and forward system. In this thesis work a systems engineering approach towards, requirements definition, design, verification and testing is provided.

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.

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.

Electric low-thrust propulsion has nowadays found wide application in space dynamics as it entails considerable savings in spacecraft propellant mass, thanks to the very high specific impulse that this kind of engine is able to generate. However, continuous thrust opens new extensive sets of feasible trajectories, and optimization algorithms are needed to mine the admissible search space and find optimal transfers. Practical methods to solve complex optimal control problems as low-thrust trajectory optimization typically involve high computational times, the major bottleneck of these techniques. The objective of this work is the development of a novel multiple-shooting optimization tool, employing a variational approach for quick derivative computation, and the assessment of its performance against a variety of test cases. Indeed, after a first analysis of practical optimization methods, the derivative estimation by finite-difference approximations has been found as the major contributor to the computational burden, and the propagation of the variational dynamics has been selected as an accurate approach to speed-up their computation. The theory of variational dynamics for multiple-shooting application has been analyzed in detail, and further developed for what concerns the second-order equations. After practical considerations on the method implementation (scaling procedures, sparsity patterns, et cetera) and its interface with WORHP, the selected non-linear programming solver, the tool has been applied to a broad range of test cases, spanning from elementary problems to practical applications. For what concerns the latter cases, two complex problems were analyzed and optimized: a CubeSat rendezvous departing from Earth-Moon L2 and arriving at asteroid 2000SG334, resulting in a propellant mass convenient trajectory suitable for asteroid reconnaissance well before a proposed NASA manned mission in 2069; The Kessler Run, i.e. the 9th Global Trajectory Optimization Competition (GTOC9), in which the developed tool, employed as last step of the optimization cascade of Strathclyde++ team, managed to make the solution constraint-feasible, valid and further mass-optimal.

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.

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.

The future space environment is predicted to grow in number of both operational and inactive man-made objects and the era of constellations is expected to arrive during the following years, with many telecom companies launching constellations of up to 12000 satellites. This situation will inevitably lead to over-population of the most demanded orbits making its exploitation a challenge to the scientific community as well as spacecraft operators. Regular products within the field of Space Surveillance and Tracking (SST) and Space Traffic Management (STM), such as high-risk collisions, upcoming re-entries or fragmentations, rely both on the estimated state and associated uncertainty of detectable Resident Space Objects (RSOs). Orbit Determination (OD) algorithms provide the required estimations, assuming that the uncertainty in the state of the object is properly characterized by its state vector covariance and assuming Gaussian processes. However, a common problem of OD processes is the misrepresentation of the RSOs uncertainty through the estimated and predicted covariance. Ultimately, this causes a great impact in the quality and accuracy of SST products as the covariance is overly optimistic (too small) and the true uncertainty of the object is not properly captured. The aim of this work is to devise a novel methodology to improve the covariance realism of OD and orbit propagation processes through the classical theory of consider parameters of batch least-squares estimators. The outcome of this project is a software application integrated as part of the GMV’s SST software suite that can deliver efficient and effective covariance realism improvement for a more accurate provision of SST products.

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.

Cryosat-2 altimetry helps to assess the state of the Cryosphere by measuring elevation and elevation change. The SARIn mode of the Cryosat-2 SIRAL instrument is specifically designed for rough topography that we encounter near ice sheet margins. Various SARIn data products (Full Bit Rate (FBR), Level 1b and Level 2) are provided by ESA, where the level denotes the amount of processing applied relative to the raw instrument data. Often research activities are based on L1b data; in this thesis, a new Cryosat-2 SARIn FBR-to-L1b processor is developed to evaluate whether elevation measurements improve when we reconstruct a L1b product from the FBR data. The L1b-to-L2 processing is performed by means of Swath Processing, by which an increase in amount of measurement points is achieved with respect to the amount in conventional L2 products. The analysis is limited to the Petermann glacier, located in the North-West of Greenland, which is known for its large calving events. The results are verified by comparing them to the ESA products. It is shown that the resulting elevations following from the FBR-to-L1b processor agree; we detect significant variability in topography as present in the Petermann Glacier region that can induce data gaps in the raw FBR data; with consequences for the L1b- and L2 products. Overall, it is shown that small improvements in terms of accuracy can be achieved with respect to the ESA L1b data; this conclusion follows from the consideration of independent IceBridge laser altimetry data in our study.

This report presents the Final design of the Design Synthesis Exercise (DSE) to 'Capture a Small Asteroid and Change its Orbit' at the Faculty of Aerospace Engineering at Delft University of Technology. The bachelor programme 'Aerospace Engineering' comprises several projects enabling students to explore aeronautics and space from different kinds of perspectives. The Design Synthesis Exercise serves as the conclusion to this programme. During this final project students integrate their previously obtained knowledge and skill to examine a specific design problem in groups of ten students for the duration of eleven weeks. This final report is the last in a series of four and documents the detailed design of the concept that was chosen in the mid-term report.

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.

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.

The Equivalence Principle is today challenged by some theories attempting to unify General Relativity and Quantum Mechanics. The MICROSCOPE mission thus intents to confirm or overturn this principle by testing the universality of free fall in space with an unrivaled precision objective of 10^-15. The principle of the test relies on the precise measurement of a gravitational signal by a differential electrostatic accelerometer (referred to as SAGE instrument) on board a drag-free microsatellite. The precision of the measurements exploited for the test is limited by several perturbations due to the high sensitivity of the instrument. The instrumental model implemented to process the collected data used overestimated models for some systematic errors such as the thermal sensitivity and coupling defects. The additional scientific measurements available by the end of the mission, will enable to perform a more in-depth exploration of these errors and thus to improve the current instrumental model. This is the purpose of this research work.