Asteroid missions stand at the forefront of space exploration endeavours. These face a number of challenges, with safe navigation being one of the main issues. The irregular nature of these bodies creates highly perturbed gravity fields that, if modelled incorrectly, can be a danger to the safe navigation of the spacecraft. Therefore, improving the gravity field modelling of asteroids is of key importance.

This research aims to improve the gravity field estimation of asteroids through the use of a satellite constellation consisting of a mothership and a set of CubeSats. In particular, it focuses on the modelling and estimation of the gravity field with spherical harmonics (SH), which makes the research only applicable to the navigation of the spacecraft outside of the Brillouin sphere.

The system design is based on a CubeSat constellation orbiting the asteroid that relays measurements back to a mothership that estimates its state and gravity field through an Unscented Kalman Filter (UKF). The research is centered around 433 Eros asteroid for the model implementation, being an irregular body with known characteristics from the Near-Shoemaker mission.

An end-to-end simulation environment is implemented for the system research. This allows for simulating the orbits of the satellites in a real-world environment considering SH gravity field up to degree and order 15, the perturbation effect of the Sun point mass, and the solar radiation pressure. Additionally, it includes a realistic polyhedral shape of the asteroid including its landmarks. The asteroid model is used together with the modelling of the satellite sensors to estimate the measurements that can be obtained in a realistic scenario including sensor errors, visibility, and communications constraints. Furthermore, the simulation environment contains a UKF filter capable of conducting the gravity-field estimates from the measured states of the constellation satellites.

The research conducts a thorough sensitivity analysis of the scenario evaluating how the constellation design impacts the estimates obtained. This analysis evaluates a number of filter designs, and determines that better performance is achieved when the design models the dynamics of several satellites at the same time. This is followed by an analysis of variance, conducted to obtain a better understanding of the constellation characteristics' effects on the filter estimates.


From the results obtained, a design synthesis is conducted to test the system implemented with an optimised constellation design. Furthermore, the filter design is re-evaluated and improved through the addition of better covariance matrix tuning and the addition of a degree-by-degree estimation procedure that allows to reduce computational load, a main constraint of the system. The results obtained show that the model is capable of accurately estimating the gravity field spherical harmonic coefficients with an error lower than 15% when the filter is tuned properly for SH up to degree and order nine. Additional verification of its applicability has been conducted by testing the system with extreme case asteroids, which show that the system requires a thruster and control system for the satellites to be capable of maintaining stable orbits around the asteroids when these have extremely irregular gravity fields.

Automated control of Earth observation missions does not only allow for a more efficient use of onboard resources but also improved quality on the scientific return and cost reductions in the operations. Current developments in space hardware computational capabilities allow the implementation of onboard optimisation algorithms that can use the state estimations available onboard to dynamically generate control sequences in adaptation to disturbances or unpredictable behaviour.

This study proposes an onboard target management software with the aim of observing a set of targets on Earth’s surface in an automated manner. The development must be carried out within the framework of AAC Hyperion’s integrated attitude determination and control system. The proposed solution is designed to autonomously manage observations and optimise reorientation manoeuvers to designated ground targets, reducing the need for continuous ground intervention and enhancing operational efficiency. The method proposed combines the solving of a Mixed-Integer Linear Programming (MILP) optimisation problem for the scheduling of the observations, while
using successive convex programming to perform optimal control for spacecraft constrained reorientation and target pointing in a real-time manner. These two primary components, together with another task for computing the target visibility windows, operate within a modular framework, facilitating communication with the existing ADCS architecture and making efficient use of onboard computational resources.

The MILP scheduling subproblem selects targets based on priority, visibility constraints, and the required slew times between observations. Initial results identified limitations in MILP performance, particularly as the number of targets or proximity constraints increased, where large computational loads exceeded real-time processing capabilities. Alternative scheduling strategies are suggested for highly complex or densely populated target scenarios. Convex optimisation is found to be a suitable method for the solving of highly constrained manoeuvres in real-time. The reformulation of non-linear systems and complex, combined with the use of successive convexification were used to efficiently solve the spacecraft attitude control problem. This method’s robustness and
capability to detect infeasibility were especially beneficial for the modular onboard software.

The research tested and validated the proposed algorithms using a software-based real-time simulation environment, allowing a preliminary assessment of the observation manager’s performance. On the one hand, the method used for approximating the slew rate was found innaccurate, worsening the performance of the convex optimisation due to its interdependency with the MILP optimisation. On the other hand, results showed that the convex optimization method was effective for achieving precise target orientations, but achieving target accuracy of 0.1° often required refinement through additional onboard controllers, such as a PID controllers, to maintain
target lock during observation. Additionally, the trajectory optimiser demonstrated robustness at an update rate of 0.2 seconds, allowing the spacecraft to realign to targets successfully despite potential deviations in spacecraft state. However, the limited memory and slower processing speed of the ADCS hardware present challenges in extending these methods to onboard use.

Based on the results, the study suggests the modification of the slew approximation, and the use of alternative methods to the MILP, such as dynamic programming, to improve scheduling efficiency in future work. Moreover, testing on the ADCS hardware processor through the use of HIL simulations is essential to faithfully reproduce the onboard performance for analysis. A shift to higher-performance onboard processors or offloading some computational tasks to external processors is recommended to fully support the demands of automated Earth observation operations. This thesis demonstrates the feasibility and potential of integrating advanced optimisation methods onboard Earth observation satellites, providing a foundation for future advancements in autonomous spacecraft operations and contributing to the field of real-time, dynamic scheduling, and control in space systems.

This thesis describes the implementation of an innovative experimental set-up to perform simulations of laser link acquisitions between two spacecraft, in the context of laser ranging interferometry missions, which are increasingly relevant in the fields of gravimetry and gravitational waves detection. The objective of the project is to reproduce realistic laser beam characteristics, demonstrate the ability of the Optical Formation Metrology Testbed (OFMT) to perform autonomous link acquisition procedures and define requirements for the OFMT architecture based on the experimental results. To achieve the goals set, the OFMT set-up has been assembled and aligned, the relevant beam properties have been measured, the supporting experimental infrastructure to conduct the acquisition tests has been integrated, the necessary calibration procedures have been carried out and finally the acquisition tests have been defined and performed, followed by a comprehensive analysis of the results.
From the evaluation of the test runs and the measurements of the beam characteristics, it can be concluded that the OFMT set-up represents a new way of successfully recreating laser beams between spacecraft that are placed far apart from each other, following a design path that had not previously been covered by other experimental architectures of this kind. Additionally, the work outlined in this thesis demonstrates the ability of the set-up to carry out autonomous link acquisition procedures, paving the way for further developments of the instruments.

Small satellite systems such as CubeSats and PocketQubes have strict requirements in terms of size, weight, and power available onboard. In light of these constraints, small satellite systems typically omit the inclusion of a ranging system due to its power and specialized hardware requirements. However, new research in satellite ranging invented the telemetry ranging class of techniques which do not place any such requirements on the satellite. The thesis explores the possibility of implementing a telemetry ranging technique on the Delfi-PQ satellite, which is already in orbit using only software modifications. The working and performance of the technique in determining the position of the satellite were explored using its engineering model and have positive implications for small satellite navigation and science.

Crater counting is the main method used to determine the age of a planetary surface. This method relies on knowing the cratering rate to estimate the absolute age of a surface. However, two theories currently exist for impact cratering of the moons in the Saturnian system, one suggesting that the majority of the impactors are of heliocentric origin (i.e. they come from orbits around the Sun), and the other suggesting that the majority of impactors are of planetocentric origin (i.e. they originate from orbits around Saturn). The different theories result in very different surface ages of the icy moons. According to the heliocentric model, the surface of Titan would date to ~3 Ga, while according to the planetocentric model the surface of Titan could be dated between ~15 Ma and ~4 Ga, allowing for much more erosion of the surface according to Bell (2020). Given that these two orbits result in different impact velocities, this work attempts to discover if the impact velocity can be determined for impactors on icy moons based on crater characteristics, such as crater depth and diameter, by recreating impacts in the lab.

Previous impact experiments have been performed with different impactors and target materials. This study aims at expanding previous research by performing tests on icy particles, something not done before. The data obtained from experiments are then used to obtain the estimated impactor diameter that created craters on the surface of these icy moons based on the expected impact velocities. This data can then be used to predict the likelihood of an impactor originating from either a heliocentric or planetocentric orbit.

This work developed several methods to produce impact craters using different instruments and different surfaces. The best results were obtained either with the gas gun and ice blocks or with the drop tower and icy particles. Using the scaling relationships found in this work, it was shown that the required impactor diameter can be obtained from craters on icy moons and that a distinction can be made between planetocentric and heliocentric impactors. However, more work is needed to narrow down the scaling relationship to obtain a reasonable impactor diameter given a certain impact crater. More information about the crater, such as the transition of ice type during an impact, should be used in the future to further constrain the velocity ranges.

Satellite tracking is used to predict a satellite’s orbit to communicate and perform other key functions. It is commonly performed using large ground-based radars with an accuracy of 1 to 10 km or specialized onboard satellite systems. Ground optical systems are a proliferating technology for satellite tracking, identification, and communications. Such optical systems have a narrow field of view, and to ensure smooth and quick satellite acquisition, the satellite must be tracked with higher accuracy (in the order of 100 m). Hence, to allow for optical ground systems to function smoothly they require satellite tracking systems providing sufficient accuracy which is not common for small satellites.
This work analyses two radio-based methods to provide sufficiently accurate satellite tracking for the next TU Delft satellite to be acquired optically. These methods utilize the innate satellite radio communications subsystem of the satellite without modifications. To optimize the analysis work, the model complexity was gradually built up. The first method analysed is phase interferometry, which based on a first-order model, turned out to be non-viable. The second method investigated is time difference of arrival. Increasingly more sophisticated models were developed to understand its performance and limitations. For example, for Delfi-PQ, by utilizing 25 ground stations in a square configuration with 1000 km baseline and a noise level of 236 ns (71 m), the target satellite in a 500 km circular orbit can be acquired, with at least the required accuracy, for 18% of the area where the satellite can be received by a minimum of 4 ground stations. It is expected that this performance can be improved through future work. Specifically, by utilizing multiple measurements taken at different times, the satellite orbital parameters can be estimated directly.
We conclude time difference of arrival to be a promising method for further research, development, and eventual implementation by TU Delft. The approach and structure of an even more sophisticated model is detailed for future work.

Stereo matching, the process of inferring depth maps from stereo images, is one of the most heavily investigated topics in computer vision. It is part of the first module of navigation systems of planetary rovers, e.g., NASA’s Mars Exploration Rover (MER) missions, NASA’s Mars Science Laboratory (MSL) mission, and ESA’s ExoMars mission. Many stereo matching algorithms, traditional and deep learning based, are available and their performance is typically evaluated on indoor objects or outdoor city scenes. In this thesis, our goal is to improve insight into the performance of stereo matching algorithms for the Martian surface on a single-board computer. First, we search for and compare stereo matching algorithms ranked on stereo vision datasets to obtain a manageable set of algorithms and verify their implementations. Second, we derive performance metrics from the requirements and validate our set of verified algorithms on the Katwijk Beach Planetary Rover dataset. Through experimental evaluation, we gain insight into the performance of four stereo matching algorithms.

The Brik-II satellite is a Nano-Satellite which has the main objective to be a technology demonstrator and prove the use of self owned military satellites. One of the payloads onboard the Brik-II satellite is called the Store and Forward payload. It has the capability to Store and Forward messages from and to different military assets. To demonstrate this capability it has been purposed to design and install a communication system onboard the Royal Netherlands Navy submarines enabling communication with Brik-II. Subsequently, this would be a good way to demonstrate the performance and operational relevance of such system.
The first step was to identify the problem that should be solved. The main objective of this MSc thesis was to design, manufacture and test an operationally deployable communication system for the submarines operated by the Royal Netherlands Navy. Based on the main objective and consultation with the Navy the requirements of the system were compiled. From the requirements it was concluded that there are a number of similarities between the Remote Radio Station (RRS) and the submarine communication system. However, the antenna system installed on the RRS cannot meet the size requirements. Therefore a new antenna systems had to be design that would comply with the requirements. This has been the main focus of this MSc thesis.
To find the optimal design for the antenna, a simulation tool for the communication between the submarine and Brik-II was developed. In the model different antennas were tested to identify the optimal design. The type of antenna used in the optimisation is a helical antenna. Once the optimal design was identified it was manufactured used the 3D printing facilities at the 982 Squadron in Dongen. Hereafter, measurements were conducted to verify that the radiation pattern of the antenna corresponds with the predicted radiation pattern. The similarities in shape between the predicted and measured pattern were evident. However, the measured gain of the antenna was lower than the predicted gain. To meet the requirements the losses will have to be mitigated. Multiple solutions were purposed to reduce the losses. A communication test between the antenna and the satellite could not be conducted due to technical difficulties during testing.
Part of the main objective was to develop an operationally deployable system. This first prototype developed in this MSc thesis provides a basis for further iteration. The current design imposes two constraints on the deployability. First the antenna gain losses have to be mitigated since the lower gain results in less than required data throughput. The losses can be reduced by implementing the proposed solutions. Second, the choice of a helical antenna with an omnidirectional radiation pattern results in an increased chance of detection by possible adversaries. This functional constraint cannot be remedied due to the size requirements. In conclusion, the antenna design proposed in this thesis gives a basis for a system that can provide added value to the operations of the submarine.

CryoSat-2 is a European Space Agency (ESA) altimeter mission with an objective to study the connection between cryosphere melting and global sea level rise. The satellite carries a Doppler navigation system (DORIS) and a Satellite Laser Ranging system (SLR) to aid the precise computation of orbits down to centimetre level.
CNES/IDS releases DORIS tracking data in two formats for CryoSat-2. One is the raw format called RINEX and other is the pre-processed Doppler format called version 2.2. V2.2 data contains all information necessary for straightforward usage in orbit determination - measurements time-tagged in TAI, range-rate measurements, ionospheric correction, tropospheric correction, antenna corrections and flags that indicate unusable measurements. RINEX does not contain any corrections and has the phase and pseudorange measurements at short latency allowing users to have flexibility in processing. Additionally, data required for formulating the corrections are present in RINEX. For missions in and after 2016, CNES/IDS supplies tracking data only in RINEX and not in V2.2. Analysis centres using DORIS data now have to independently develop processing strategies to process RINEX data. This problem is the main objective of this research. In this thesis, a pre-processor called RX2RR (RINEX to Range-Rate) has been built in Fortran in an attempt to process the raw data and compute all the necessary corrections. RX2RR converts RINEX format to a format exactly similar to V2.2 such that RINEX can now be used in any orbit determination tool that has been previously using V2.2. In this processor, clock synchronisation of on-board clock to International Atomic Time is performed. A new approach of utilizing Meteorological data in RINEX for troposphere delay corrections is implemented. Use of real time data from numerical weather models is also presented for tropospheric correction. Ionospheric delay and antenna phase centre corrections are performed using iono-free phase centre. An editing strategy to remove outliers in Doppler data is implemented and tested. To demonstrate the performance of the tool, we perform orbit determination using NASA Goddard Space Flight Center’s orbit computation software GEODYN-II. We use RX2RR processed RINEX data and CNES processed V2.2 data of CryoSat-2 for year 2016. Tracking residuals from both POD runs are compared and average difference in R.M.S residual is found to be approximately 0.011 mm/s over a year. This validates that the corrections are formulated and implemented correctly. The result proves our capability to process the RINEX measurements independently and the tool developed has extended the capability of GEODYN-II to process RINEX observations from DORIS system.

This master thesis addresses the need for a telemetry processing for Delfi Space spacecraft operations, a nano-satellite program of the Delft University of Technology. It lays the groundwork for the telemetry processing system (TPS) implementation and explores the design guidelines for a highly scalable but robust and adaptive architecture.

The TU Delft Deployable Space Telescope proposes to meet increasing demand for high spatio-temporal resolution Earth observation imagery by reducing the mass, volume and launch cost of space telescopes by using segmented deployable optics. A 3DOF piston/ tip/ tilt fine positioning mechanism is proposed to actively align the primary mirror segments in orbit. A preliminary thermo-mechanical design of the mechanism was synthesised and its mechanical performance verified with an efficient, low order finite element model. The results showed that the mechanism can support the mirror through launch without a hold down mechanism; has a piston step size of 10 nm; a range of ±4 µm in piston and ±9 µrad in tip/ tilt with accuracies of 0.005±0.780 µm (2σ) in position and 0.000±0.041 µrad (2σ) in rotation. A Monte Carlo ray tracing analysis showed that the active optics could successfully align the mirror segments in all considered deployment scenarios.

As technology improves, increasingly higher resolution payload can be achieved using Cubesats for Earth observation. The diffraction limit prevents the resolution to up to a few meters for these missions and are confined to Very Low Earth Orbits (VLEO). At these altitudes, strong disturbances act on the system, limiting its lifetime and the pointing capabilities of CubeSats. As a solution, the Department of Space Engineering at the Delft University of Technology
has proposed a 6 unit CubeSat named the Stable and Highly Accurate Pointing Earth-imager (SHAPE) orbiting at a Sun synchronous VLEO which uses a momentum wheel to passively stabilize the system against the external environment within a competitive cost of less than 500 000e. By utilizing the dual-spin stabilization concept, composed of a stable platform and a spinning rotor, it is expected to perform pointing missions of less than 1 degree.

In this thesis, the SHAPE concept has been revisited and further developed based on the work of Kuiper and Dolkens to conduct whether these types of missions are feasible within the aspect of the Attitude Determination and Control Subsystem (ADCS). This thesis covers the base of this subsystem approached from a top-down methodology; designed from the final nominal mission mode to the detumbling mode on a system level.

The ADCS design will consist of a momentum wheel which has been determined to have an angular momentum of 1 Nms. This value is based on a prediction of the worst-case atmospheric density of the next solar cycle. The design
point, at which the momentum wheel has been sized, has been taken at 90% of SHAPE’s lifetime after several design iterations. Hereby, the last 10% of the mission has been partially forfeited with degraded performances due to the
exponential increase in disturbances acting on the spacecraft at lower altitudes. As therefore, the mass and size of the momentum wheel has been reduced with 41% and 20%, respectively. To re-align the angular momentum vector within
the 1 degree pointing requirement, a set of magnetorquers with a dipole moment of 0.5 Am2 has been chosen due to their low power consumption, mass, cost, and high reliability while capable of producing sufficient torque. Also, a
damper is to be integrated as it provides the system asymptotic reduction of the transverse momenta, thus increasing the image quality without expenditure of additional power.

To reach the nominal mission observation state, several momentum wheel spinup strategies have been investigated. Based on a trade-off between three spinup concepts, it was concluded that the major axis spinup is most suited. This
type is initiated after the spacecraft as a whole has attained an angular momentum equivalent to that of the desired end value of the momentum wheel. Then, a constant rotor torque is applied, providing a momentum transfer from the platform to the rotor. The disadvantage of this spinup procedure is that the system’s solar panels are aligned parallel to the orbital plane, meaning that power cannot be generated and batteries are required during the spinup. Despite this, it was found that after completing the spinup, the transverse angular momenta was minimized to marginal values in contrast to the other spins. The inclusion of the passive damper during the major axis spin further improves the ability of reducing the transverse momentum as the damper’s dissipative energy property adds an asymptotic attraction at the point of lowest energy, located at the spin axis near the end of the spinup.

The all-spun state is achieved using a set of thrusters. This choice was taken as the magnetorquers was found not to deliver sufficient torque. From the detumbling analysis, it was concluded that the magnetorquers are able to reduce
the tumbling rates with magnitudes of up to 35 deg/s to mean motion values in less than an orbit using a static gain B-dot controller.

Imperfections in the momentum wheel can cause static and dynamic imbalance, imparting internal disturbances to the system which affects the image quality detrimentally. Therefore, isolators are to be integrated within the momentum
wheel suspension subsystem. If these disturbances can be negated using isolators, it can be expected that the pointing error will stay within one degree and attitude stability can be achieved at least until the design altitude of 280 km.
However, the analysis and design of the isolators have yet to be done and thus the attainability of the pointing and image quality requirements are still inconclusive.