Since Galileo Galilei’s first discovery of natural satellites orbiting around other planets, observing and reconstructing their dynamics has been at the core of our efforts to understand and characterise these distant worlds. Far from following perfect, frozen in time Keplerian orbits, the dynamics of these satellites keep evolving, with tides as driving mechanism. The dissipation of energy in natural bodies due to their visco-elastic response to tidal forcing both heats up the moons’ interiors and causes their orbits to expand or shrink, as well as to become more circular or elliptical. Refining the moons’ ephemerides (i.e., tabulated solutions of their motion as a function of time) is thus key to studying not only their present-day dynamics, but also the long-termthermal-orbital evolution of planetary systems....@en

The primary motivation for this work was the lack of open-source or clearly motivated studies describing the orbit estimation of natural satellites. The capability to accurately model the trajectory of the moons of Jupiter is crucial for properly understanding the evolution of the Jovian system and, by extension, the solar system. Considering the small Jovian moons remain relatively unknown, they will be the focus of this study. The goal of this project was to develop a framework for the full estimation process, based on the Tudat software library, starting from raw astrometric observations, to be made available publicly. The first step was designing a new data processing algorithm, capable of uniformising the observational data. These measurements are reported in a wide variety of formats, with differences in time format, time scale, orientation, and observation type, amongst others. The developed software was able to produce Tudat-compatible observations with limited user interaction. These processed measurements were then analyzed and validated, first by examining the mean residuals with respect to existing ephemerides to detect remaining biases in the data, before comparing the standard deviation of these residuals to those reported in their original publication, whenever available. Secondly, the estimation framework was set up. Based on the processed astrometric observations, the orbits of three minor Jovian satellites, Himalia, Elara, and Amalthea were estimated as case studies. Each of these posed its own set of challenges, but together they yield a good representation of the full range of outer moons. The SPICE kernel ephemerides were used to evaluate the quality of these orbital solutions. Estimating Himalia’s orbit proved the easiest, as it is the minor moon that is observed the most, in combination with its relatively slow dynamics. One problem was the fact that there was a remaining unmodeled bias in the oldest, i.e. pre-1960, observations due to errors in converting from old to new frames. Removing these observations and replacing them with simulated measurements with realistic residuals both proved to reduce the difference with respect to the SPICE benchmark to within the uncertainty level of this very benchmark. This proves that the framework allows to accurately estimate the ephemerides of Himalia, although no conclusions could be drawn on the exact contribution of the old observations. Elara’s orbit is in many ways similar to that of Himalia, as they are part of the same orbital group. However, a close approach with this more massive Himalia in 1949 had a significant impact on Elara’s trajectory. Therefore, estimating Elara’s orbit is not limited to its initial state. Instead, Himalia’s gravitational parameter is determined concurrently to optimize the solution. However, even when determining this gravitational parameter, it became apparent that the current dynamical model could not perfectly capture the dynamics at the close approach, due to the limited amount of data before the event, leading to an error of several hundred km. This is still well below the residual of the observations, however. Finally, Amalthea proved to be the most difficult body to get an accurate orbital solution for. Since Amalthea is an inner moon, as opposed to the outer moons Himalia and Elara, its angular velocity is over 500 times higher. This logically makes both timing and position errors significantly more impactful. Additionally, the very limited amount of data, which is constrained to just a few campaigns of several nights, further hamper accurate orbit determination. By limiting the step size of the update in the estimated initial state between iterations to a standard deviation of 100 km by defining the a-priori covariance, a stable solution, accurate to several 100 km could be found. A preliminary investigation of the contribution of simulated spacecraft data indicated that the few available spacecraft observations of Amalthea have a large impact, which warrants further studies. Thus, through determining and analyzing the orbit of these three case studies, the quality of the developed framework is ascertained. The solutions proved to lie within the uncertainty region of the ephemerides published in SPICE. Inaccuracies in the estimated orbit could always be linked to deficiencies in the input data, thus not taking away from the quality of the framework.

With the development of space research into novel areas, new complex problems arise. The interest in solving space routing problems considering large numbers of targets has recently grown. This paper proposes a novel method to solve the optimal trajectory in such combinatorial space routing problems. This paper focuses on a global optimisation algorithm implemented to solve the problem posed in the 4th Global Trajectory Optimisation Com- petition (GTOC4). The solution is a trajectory of multiple legs, where each leg links two targets and has a specific flight time. To enable the use of the powerful mixed-integer linear problem solver software, Solving Constraint Integer Programs (SCIP), the routing problem concerned with visiting as many target bodies with a predetermined fuel and time budget is split into linear sub-problems. The Fixed Budget sub-problem selects a subset of the given set of targets. The Full Tour sub-problem orders the targets in the subset, and the Fixed Tour sub-problem optimises the flight time for every leg of the given trajectory to find the solution with the lowest total fuel consumption. Each of these sub-problems is formulated in a linear form and is solved using SCIP. The global optimisation algorithm evolves a population where every individual exists of a set of initial guesses for the time of flight values. Analysis shows that initialising this population with a mix of randomly generated individuals and individuals containing a constant value for all entries leads to the fastest convergence towards the optimal solution. In a population of 20, seeding ten individuals is found to be optimal. It is also found that the algorithm performance can be further increased by evolving individuals with infeasible solutions instead of iterating them until a feasible solution is found and eliminating the Full Tour sub-problem. These simplifications allow for an increase in the cost budget multiplier, which leads to finding better objective values without further increasing computational time. The best-performing setup, which uses a cost budget multiplier of 10, can find the optimal solution to the test problem in 100% of the runs, on average in 9 iterations, with a computation time of 5.82 seconds per evaluation. The results show that the global optimisation algorithm produces results that closely match known results for GTOC4 consistently and accurately.

Small solar system bodies have received increasing scientific attention over the past decades. Studying their primitive origins can reveal important insights on the formation of planets, as well as the general evolution of the Solar System. Also their potential threat to life on Earth upon collision and their potential to act as stepping stones for deep space exploration make these bodies interesting to explore. Especially CubeSats, being small and lightweight, are promising candidates for such missions.
To facilitate the design of small body missions, this research investigates the non-linear effects of uncertainties in an asteroid’s environment on the orbital motion of a CubeSat. Uncertainties in the asteroid’s mass, irregular gravity field and solar radiation pressure have been studied prior. Rotational state uncertainties have not been researched, but do affect the CubeSat’s motion indirectly through the orientation of the irregular gravity field. These are therefore selected as the subject of this work. This aids in identifying orbits that are robust against these uncertainties, thereby minimizing the required fuel for trajectory corrections and maximizing the mass for science instruments or increasing the mission duration.
The study of the non-linear effects of rotational state uncertainties requires the application of non-linear uncertainty propagation methods. Non-Intrusive Polynomial Chaos was selected for its ease of implementation, promising computational efficiency and ability to provide statistical information directly. In this method, a set of samples from the uncertain domain are propagated to a desired time according to the black-box dynamics that govern the orbital motion. Subsequently, a polynomial approximation, a so-called Polynomial Chaos Expansion, is constructed for these final states as a function of the uncertain variables. This then allows for finding the final states for all possible characterisations of these uncertain variables, in a Monte Carlo like fashion, without further numerical integrations. Above that, the Polynomial Chaos Expansion terms can be used to compute statistics, such as the mean and covariance, analytically. In this work, different initial conditions are propagated and the states at various times are approximated by Polynomial
Chaos Expansions. All Polynomial Chaos Expansions were verified by comparison with Monte Carlo simulations.
A study on the settings of Non-Intrusive Polynomial Chaos was conducted. It showed that the required settings, such as polynomial order, number of samples and method of solving for the coefficients can vary significantly, depending on the studied case. In addition, limitations in the application of this method to
Kepler elements were encountered for orbits that approach the singularities in this element set. In general, the results show an increase in the trajectory dispersions and non-linearities encountered with an increase in propagation times and with a decrease in orbital altitude. However, exceptions to this
trend were encountered. In the case of a retrograde equatorial orbit, the inclination was found to reach its maximum dispersion already within 5 days and stagnates thereafter. This is a result of the accelerations exerted by the asteroid’s gravitational bulges in its equatorial plane, which varies in space along with changes in the rotation pole under these uncertainties.
A comparison to uncertainties in the asteroid’s mass revealed that the effects of rotational state uncertainties are relatively small, especially considering the small uncertainty in mass that was used. However, again an exception was encountered. The right ascension of the ascending node of a polar orbit at 5 km was found more sensitive to changes in the asteroid’s rotation pole than in its mass. Thus, depending on the objective of an analysis, mission designers could be required to include rotational state uncertainties in their analyses.
Finally, a broader study of different initial orbital geometries revealed that retrograde orbits are more stable against rotational state uncertainties. However, depending on the exact initial orbital geometry, in terms of inclination and right ascension of the ascending node, prograde orbits can be just as stable. Also
here it was found, though, that the inclination is more stable for polar orbits than for inclined and equatorial orbits.
In conclusion, the finding that retrograde orbits are more stable against rotational state uncertainties helps mission designers to select promising orbits for actual missions. As these orbits are also beneficial for geodetic parameter estimation, they both minimise the required fuel for trajectory corrections and maximise the scientific return. Nonetheless, a wide variety of non-linear effects due to rotational state uncertainties can be encountered in the asteroid’s environment. Therefore, their influence should always be checked in mission design studies, even though in general their effects are relatively small.

This thesis project investigates navigational aspects of the upcoming JUICE (JUpiter ICy moons Explorer) mission, which is expected to deliver valuable data products for the scientific study of the Jovian system, and the three Galilean moons Europa, Ganymede and Callisto in particular. One such data product comes from the 3GM radio science payload, which is expected to contribute towards more accurate ephemerides of the Galilean moons and a refined study of their gravity fields. Accurate and reliable acquisition of these data depends greatly on the successful Guidance, Navigation and Control (GNC) operations, where an efficient GNC performance can unlock excess ∆V capabilities. It is the goal of this work to identify statistical ∆V savings from integration of in-mission radio science ephemeris improvements into the GNC operations. The resulting excess ∆V could be used to effectuate one of the mission enhancement and extension options (e.g. coverage of the Jupiter’s mid-high latitudes or a lower-altitude Ganymede orbit), the enablement of which ultimately motivates this thesis project.
A navigational orbit determination (OD) framework was set up, modelled closely after the OD setup from the JUICE mission analysis team. An interface for simulating moon state knowledge updates from external ephemeris products was implemented. This setup was then extended to support a simplified statistical ∆V analysis. The external ephemeris products were modelled using an adopted high-fidelity OD setup with a coupled filter configuration, which simulates the uncertainty of in-mission ephemeris updates from the 3GM radio science data. These results were then used to simulate external ephemeris updates to the navigation OD and to evaluate their impact on the statistical ∆V expenditure.
It was found that external moon ephemerides in general, and the simulated science data based OD solu- tion in particular, are not suitable for reducing the statistical ΔV cost of post-flyby correction manoeuvres. Within one to two encounters of a given flyby body, the nominal (update-free) navigation OD has improved moon state knowledge to a point where it no longer contributes significantly to size of the correction ma- noeuvres. To generate statistical ΔV savings, the moon state knowledge must be improved for the navigation of the mission’s first encounters. It was thus concluded that moon ephemeris improvements should not be provided from JUICE science data, but from observations that have been collected prior to mission. It was furthermore found that due to its extensive coverage, the navigation tracking data captures system informa- tion that the radio science data is less sensitive to. It could therefore be considered in the post-mission efforts to improve the high-accuracy ephemerides of the Galilean moons.

Laser communication offers high data rates and more bandwidth capacity compared to radio but is also advantageous in terms of power and volume. However, the required pointing accuracy and atmospheric influence pose stringent requirements on the system which form the main challenges for this type of communication. This is where laser ranging could make a difference. Laser ranging is more mature than laser communication but both technologies share many commonalities. A synergy of the two could therefore be beneficial to small satellites that are constrained by volume, power or cost. Alternatively, it could expand the capabilities of existing optical communication systems without significant modifications. The latter is considered as the main driver for investigating combined laser ranging and communication capabilities between CubeCat and the TNO optical ground station. CubeCat, currently being developed by TNO, is an example of a cubesat-sized terminal that will demonstrate in-orbit high-speed satellite-to-Earth laser communication.

Planetary observational accuracy increases and the dynamical modeling of tidal behaviour has to follow. In this thesis, a newly developed method (Correia et al. 2014; Boue et al. 2016) to deal with tidal interactions in a 2D two body problem is validated and for the first time applied to Solar system dynamics on a short time-scale. This method uses a Maxwell rheology for the tidally perturbed body, to calculate the instantaneous deformation of the body with a differential equation at the same time as its position, spin vector and orientation. The coupling between the translational, rotational, and tidal dynamics is incorporated in a consistent way, taking into account the frequency dependency of classical tidal parameters. The coupled model is more general and is in contrast to currently used models valid for every eccentricity, spin rate and orientation. The model is applied to the Mars-Phobos and Earth-Moon systems and the tides are determined separately on both the central as the satellite body in various propagations. Several parameters of tidal effects of these systems are obtained and compared with currently available literature approximations and currently used tidal models. The instantaneous deformation of the tidal body and the evolution of the system's orbit for the tides on the primary compare to the literature approximations and currently used direct tidal force model (Lainey et al. 2007). However, the coupled model displays behaviour of the tidal time lag and angle, which are accompanied with the deformation, that cannot be captured by the classical method. Small differences between final states of the coupled model and the direct tidal force model are obtained that could potentially be important in future space missions and ephemeris determination. The tides on the locked secondary show a larger difference between the coupled model and the current tidal direct force models (Lainey et al. 2007; Lari et al. 2018) in the case of an existing libration. Phobos has such a libration and huge differences in evolution of the states and orbit are found. The coupled model's evolution values can be partly justified by the theoretical work of Efroimsky et al. (2018) that states the additional dissipation in these bodies with large librations. Altogether, currently used methods or literature approximations can still be used for inaccurate propagations. However, for more accurate propagations, and for bodies with large physical librations, it rewards to switch to a coupled method. Especially when regarding the tides on the satellite body a coupled model is beneficial.

Interplanetary Laser Ranging (ILR) is an experimental satellite navigation technique with the potential to produce cm- to mm-level accurate and precise range observations over interplanetary distances. This is an improvement over current radio ranging techniques, which can achieve m-level performance. One common application of navigation data is the estimation of physical and astronomical parameters. To utilize the improved observation quality of ILR, it is necessary to use sufficiently detailed observation- and dynamical models during parameter estimation. We investigate the influence of relativistic modelling on parameter estimation quality using ILR data. Simulated parameter estimation experiments involving a Juno-like Jovian orbiter were used to gauge the impact of 22 different relativistic effects. The spacecraft's initial state, the mass of Jupiter and two post-Newtonian parameters were estimated in each experiment. Parameter estimation error, observation errors and estimation residuals were used as metrics to quantify the influence of each effect. Rough preliminary estimates of an effect's magnitude were found to be a good predictor of the actual impact on estimation error, with some exceptions. Excluding the gravitational influence of Jupiter's velocity and the Jupiter-Sun coupled gravity resulted in almost no change in estimation error, likely due to compensation by the Newtonian gravity term. Similar compensatory relationships can be identified by comparing mean observation errors with mean estimation residuals. Including empirical accelerations in the estimation process reduces the impact of these outliers.

This thesis work has been motivated by the growing interest in small-spacecraft small-body missions and more specifically inspired by a proposed mission to the largest Martian Trojan (5261) Eureka. This binary asteroid represents a promising target for a planetary mission, as a better understanding of its formation and evolution would provide valuable insights into the history of both the Martian system and binary asteroids. This work has focused on identifying orbital designs which would optimise the geodetic parameters estimate, and thus best help characterise the asteroid's internal structure. The orbit determination solutions have been computed from simulated tracking data in various orbital configurations. To compensate the highly perturbed nature of the binary asteroid's environment, pseudo-periodic solutions identified around Eureka have been used as stable initial conditions for the spacecraft orbits. As Eureka's shape, gravity field and rotational state are not well characterised yet, large resulting uncertainties have been accounted for. They have been propagated to assess the robustness of the spacecraft orbits and subsequent orbit determination solutions to dynamical mismodelling, in order to eventually limit the science risk of the mission. For single CubeSat configurations, an optimal semi-major axis has been identified, being the one closest to the asteroid before orbital instability would start deteriorating the orbit determination solution. Non-equatorial retrograde orbits have been found to provide the lowest formal errors for geodetic parameters. Moreover, the benefit of simultaneously deploying two or even three CubeSats around the asteroid has been demonstrated. In both the two and three CubeSats configurations, promising orbital designs have been identified. They provide an estimation of the geodetic parameters which is accurate enough to bring insights into Eureka’s internal structure, and has been proven to be robust to dynamical mismodelling.
The methodology developed in this work is of interest to design other small-spacecraft small-body missions, and optimise their achievable geodetic parameters estimate. Additionally, Eureka's dynamical model and the large uncertainties assigned to it have been mostly based on models available for other binary asteroids. This brings confidence in the possible applicability of the identified optimal orbital designs to CubeSats missions targeting other binary asteroids.

Tests of General Relativity are often done in the Solar System by using tracking data from interplanetary space missions, to measure the perturbation on the orbits of the planets that is predicted by General Relativity. The next best opportunity for such a test is the BepiColombo spacecraft by the European Space Agency, which will arrive at Mercury in late 2025. However, a very similar perturbation is also produced by the gravitational oblateness of the Sun through the zonal coefficient J2. The exact value of this coefficient has been hard to determine despite centuries of observations of the solar shape, and besides the accuracy of the tracking data it is the main source of uncertainty for the tests of gravitational theory. Recent publications in heliophysics suggest that higher order effects might also be of relevant influence too, such as coefficient J2 being dynamic along the solar cycle, or the fourth zonal coefficient J4. This thesis project attempts to bring together two fields of research: the field that tries to test General Relativity (GR) in the Solar System, and the field of heliophysics that tries to unravel the structure of the Sun. The orbit of Mercury, as well as the observations of BepiColombo and it predecessor MESSENGER, are simulated in a virtual reality where settings of the solar shape are varied. Relevant parameters to tests of gravity are then estimated using a least-squares algorithm and their error is analysed. It is found that the amplitude of a periodic component of the solar oblateness can be found with an uncertainty of 0.02% of the value of J2. It is also found that if a periodic component exists with an amplitude higher than 1%, it can lead to errors in the experiments of GR to the point that results oppose the theory of General Relativity. Expected values for J4 from heliophysics currently do not influence the orbit of Mercury by a measurable amount. Based on this work, it is recommended to those in the field who test General Relativity using data from BepiColombo, to take this effect into consideration to prevent getting results in their experiments that can lead physicists in the wrong direction concerning the development of gravitational theory.

Currently, an increasing amount of debris is floating in Low Earth Orbit (LEO) endangering space operations. To limit the contribution of non-operational satellites to the space debris problem a semi-controlled end-of-life de-orbit strategy can be employed. In this strategy the satellite is controlled from the end-of-life orbit down to low altitudes (120 to 200 km). From this altitude the satellite follows a ballistic trajectory towards the surface of the Earth. The target of such maneuver is to phase the impact probability track, with the Earth, such that the resulting impact track covers mostly uninhabited areas.

The objective of this thesis, is to provide insights into the effects of assumptions made in recent research of the semi-controlled de-orbit strategy. Furthermore, the extent to which the casualty risk can be reduced is investigated. Two optimization strategies are carried out. First, the control profile for the complete semi-controlled de-orbit strategy is optimized for minimum casualty risk using a differential evolution algorithm. Second, the initial state of the ballistic trajectory, at the point where the controls are turned off, is also optimized for minimum casualty risk using a differential evolution algorithm. In case similar impact tracks are achieved for both optimizations, future analysis of the semi-controlled de-orbit strategy can be performed which less computational effort.

Employing a semi-controlled low-thrust de-orbit strategy can reduce the casualty risk with 2 to 3 orders of magnitude with respect to an uncontrolled de-orbit strategy. Large ballistic coefficients are beneficial for the reduction of the casualty risk. Furthermore, orbits with high inclinations results in larger reductions of the casualty risk with respect to low inclination orbits. Also the results for low inclination orbits are less sensitive to errors in the re-entry state caused by imperfections in the GNC system of the satellite during the de-orbit maneuver. The lower sensitivity is the result of a better configuration of the impact track with respect to the large land masses on Earth. To find optimal impact tracks on the Earth, it is sufficient to optimize only the last 150 km of the trajectory. In case an accurate estimation for the value for the casualty risk is needed, higher fidelity models must be applied and the uncertainties in initial state at 150 km need to be investigated in more detail. The assumptions made in previous research influence the results with less than 33% and can easily be applied when an optimal impact track is sought for with low fidelity models. For investigating the range of re-entry conditions for which the resulting casualty risk is below the set requirement, some of the assumptions cannot be applied since it limits the solution space for the GNC system to target. Furthermore, in case an accurate value for the risk needs to be found, high-fidelity models are required and in that case, the effect if the assumptions is too large.

Currently one of the major drawbacks of interplanetary travel is the cost of these types of missions. Mission to extra-terrestrial bodies require a large amount of propellant to complete, which often takes up a large part of the mass budget of the spacecraft. This requires larger launchers to be used to get the spacecraft into the required orbit, which increases the cost of these missions. An often used method to decrease the propellant used during an interplanetary mission is the gravity assist. The gravity assist uses the gravitational influence of a planetary body to increase its heliocentric velocity. This is done by exchanging some of the planets momentum as it orbits the Sun to the spacecraft. The increase in velocity is limited by several factors: the gravitational influence of the body, the closest distance possible between the body and the spacecraft, and the orientation of the planetary bodies relative to each other. A new method has been proposed by McRonald (1990) called the aerogravity assist, which improves upon the gravity assist by increasing the angle over which the spacecraft can turn using the aerodynamic forces it experiences while travelling through the atmosphere of the planetary body. This increase in the velocity bending angle increases the heliocentric velocity compared to the gravity assist, and also allows for changes in inclination of the interplanetary orbit using less propellant. This report investigates both the atmospheric and interplanetary trajectory of the aerogravity assist, and combines them to understand the benefit of the aerogravity assist for interplanetary trajectories.

A modular simulation environment is designed using the TU Delft Astrodynamics Toolbox to simulate the atmospheric trajectory of the aerogravity assist. This simulator allows the selection of different types of acceleration and environment models. The waverider is selected to be the vehicle used during the aerogravity assist, as it is able to obtain high lift-over-drag ratios which reduce the drag losses during the atmospheric section of the maneuver. The specific aerodynamic characteristics and heating models of the selected vehicle are evaluated and incorporated into the simulator. From this simulation environment, the optimal atmospheric trajectories were investigated using an optimization algorithm. It was found that this specific problem is extremely sensitive to the selected optimization algorithm, the tuning parameters of the optimization algorithm, and the objecive and decision variables that are used. The MOEA/D algorithm was selected to find the optimal atmospheric trajectories for Mars, Earth, and Venus, where the objective variables were chosen to be the atmospheric bending angle and the incoming and out-going inclination difference. It was found that using the angle-of-attack and bank angle as control variables, an atmospheric bending angle of 120.5 degrees and an inclination difference of 0.2 degrees could be obtained at Mars. Using the control action instead of the inclination difference as objective variable at Mars showed that the atmospheric bending could be increased to 151.2 degrees while maintaining an inclination difference of below 3 degrees, and smoothed the trajectory in the process. For Earth and Venus the same optimization scheme was employed. They were able to obtain atmospheric bending angles between 70 and 50 degrees, while obtaining inclination differences below 1.5 degrees. For Venus, the heat load was also optimized instead of the inclination difference and it was found that the peak heat flux decreased by around 1000 W/cm$^2$, while decreasing the atmospheric bending angle by around 10 degrees.

A mission planner was designed to be able to investigate the effect of the aerogravity assist on interplanetary trajectories. To increase the accuracy of this mission planner, the numerically found atmospheric trajectories are used to determine the possible interplanetary trajectories. Pareto fronts were created for Mars, Earth, and Venus for a range of incoming velocities that showed the possible velocity bending angles and out-going velocities that can be achieved using an aerogravity assist. These Pareto fronts were then implemented in the mission planner to determine the influence of the AGA on an interplanetary trajectory. Several different interplanetary trajectories were investigated using both gravity and aerogravity assists. It was found that due to the fact that for Earth and Venus higher velocities are needed to travel through the atmosphere, the aerogravity assist would lose too much velocity and thus not improve upon the gravity assist. However, for Mars an increase was found of 4 km/s compared to the gravity assist for a trajectory to Saturn using fly-by's at Mars and Jupiter. Taking into account the extra mass of the thermal protection system, the decrease of mass was found to be around 60 percent, which could be used to lower the cost of the launcher or increase the scientific output of the mission by adding instruments to the payload.

The Jupiter Icy moons Explorer (JUICE) will generate ephemerides of the Galilean moons to assess their ability to sustain live. However, determining the ephemerides only using radiometric tracking (acquired using PRIDE and 3GM) results in an unstable solution which is related to the absence of flybys of especially Io, such that its dynamics can only be estimated indirectly (through the Laplace resonance). This problem is partially mitigated in this thesis by optimizing the observation schedule for optical astrometry of Io (provided by JANUS) using the NSGA-II algorithm. The optimal epochs for space-based astrometry are found to be preferentially distributed around the closest approaches of JUICE with respect to Io. During these events, the observations show a high sensitivity with respect to the dynamics of Io and have a low positional uncertainty. Using the optimized astrometry, both the condition number and formal errors can be substantially reduced.

As part of the ESA mission to Jupiter and Ganymede carried out by the Jupiter Icy Moons Explorer (JUICE) spacecraft, the Ganymede Laser Altimeter (GALA) will determine the topography and detect Ganymede’s tidal degree-2 signal in order to investigate the existence of a subsurface ocean and constrain its thickness. To deliver successful measurements GALA is to be calibrated for its misalignment with an accuracy of at least 14 arcsec. This misalignment cannot be measured on ground since it will be subject to vibrations, microgravity and settling after launch. Thus this calibration is to be performed while JUICE is on its interplanetary cruise to Jupiter.

This thesis aims to answer the question how this calibration can be performed using laser ranging to an Earth-based ground station. This calibration procedure consists of a scan performed by JUICE with a size as large as the maximum expected misalignment found at 333 arcsec. A model for the attitude during laser ranging was developed to be used to simulate the distribution and intensities of the laser pulses transmitted by GALA. During this scan of several hours and with a shot frequency of 30 Hz, the ground station will be able to detect 500 to 1500 pulses. This result was validated with the real life MLA and Hayabusa experiments already performed. The spatial distribution of these detected pulses is used to reconstruct the Gaussian pulse shape of the laser pulses. By determining its peak intensity, the ground station location can be estimated which is used to measure the misalignment. This estimation procedure was found to be robust and reliable enough to be incorporated in the simulation tool that can be run many times to achieve a statistically significant result on the estimation error. Since the estimation is based on the spatial distribution, the main drivers for the accuracy are the slew rate, scan pattern, distance, detection threshold and attitude model. These were investigated for their influence on the calibration and used to find their optimal values.

The trajectory of JUICE was analysed for the most optimal opportunities given various constrains to perform a laser ranging campaign leading to a selection of options to be analysed. Using the simulation it was found that interplanetary laser ranging can be used to calibrate GALA by employing 8-10 hours of scanning at a distance no further than around 0.5 AU by using ground station characteristics such as found at Wettzell with a telescope diameter of 0.75 metres and optical and quantum efficiencies of 0.5 and 0.2 respectively. For the most optimal opportunity, laser ranging will be performed at a distance of 0.39 AU for 10 hours with a slew rate of 0.0055 deg/s. This leads to a 2 sigma calibration residual of 3.62 arcsec resulting in a total calibration error of 11 arcsec. This confirms the assumption that a calibration accuracy of 14 arcsec is possible and thus a successful laser altimetry mission can be performed by GALA.