During a solar eclipse, sunlight incident on the Earth is reduced due to the (partial) shadow of the Moon. Atmospheric trace gas concentrations which are influenced by the amount of available sunlight, such as nitrogen dioxide (NO₂), may be affected due to the disrupted photolysis processes. Large-scale observations of the increased NO₂ concentrations caused by the solar eclipse would improve our understanding of the sensitivity of NO₂ in the atmosphere to short-term variations in sunlight. Spaceborne
measurements can provide valuable information about the large-scale spatial distribution of NO₂, which is provided daily by the TROPOMI instrument aboard the Sentinel-5 Precursor satellite by measuring and retrieving locally reflected sunlight. However, the TROPOMI NO₂ retrieval is unable to derive reliable concentrations during a solar eclipse, as solar eclipses are not taken into account in its retrieval algorithm. In this research, we have adjusted the NO₂ retrieval of TROPOMI such that it can handle solar eclipses and study the large-scale response of NO₂ during two solar eclipses over Europe in 2021 and 2022. We found a large-scale increase of NO₂ in the adjusted measurements, which linearly correlated with the degree of obscuration. We compared the measured NO₂ increase with the values from the atmospheric chemistry model TM5 including an applied eclipse implementation and we found a close agreement in most areas that are not highly polluted. Our measurements and model predict a NO₂ increase of 60%±12% and 70%±7% for an obscuration fraction of 1, respectively. More advanced chemistry modelling work is needed to explain the measurements in highly populated areas. We conclude that our results demonstrate that the TROPOMI algorithm is capable of correctly measuring NO₂ after an adjustment of the NO₂ retrieval. We have shown that it is possible to adjust an atmospheric trace gas retrieval for the influence of a solar eclipse. Moreover, we are the first to provide evidence for an increase in NO₂ during a solar eclipse using space-based measurement techniques and to quantify this increase on a large scale with the same instrument. Our measurements can be used to test atmospheric chemistry models, possibly improving their sensitivity to solar eclipses but also artificial shadows on the Earth induced by sunlight-intercepting geoengineering approaches.

In this work the possibility for use of the satellite remote sensing instrument TROPOMI to detect NOx emissions from en-route aviation is assessed. Analysis of aviation intensity, quality of satellite retrievals, and emissions from non-aviation anthropogenic sources showed that the North Atlantic, including the North Atlantic Flight Corridor, would have high detection potential. Over this corridor, emissions from telemetric observed aircraft are computed under assumption of stable meteorology leading to an upper bound for the average accumulation of emissions for the months of April 2020 and April 2021. These emissions are compared with TROPOMI NO2 observations retrieved over all sky and cloudy sky scenes. In order to reduce advection influences, TROPOMI measurements at locations of aviation observations close to the overpass time are analysed separately.

General conclusions of this first assessment are not optimistic for detection of en-route aviation emissions by means of TROPOMI observations. The maximum NOx emissions attributable to en-route aviation over the North Atlantic Flight Corridor during April 2020 and April 2021 were found to be lower than detection limits associated with TROPOMI for NO2 retrieval. Neither did inspection of TROPOMI measurements at locations of recent aircraft observations retrieve any evidence of aviation-attributable emissions. Therefore detection, documentation and quantification of actual NOx emissions from en-route aviation remains a
challenge.

The Kuiper belt is one of the last mostly unexplored regions in the Solar System. Exploration of the Kuiper belt can greatly increase humanity's understanding of the Solar System's formation and evolution. The use of low-thrust propulsion for Kuiper belt missions has the potential to improve the payload mass of the mission due to the high efficiency of its propellant. This research looks at the required methodology to optimize low-thrust trajectories with Kuiper belt object flybys. The trajectory is modelled using Tudat with spherical shaping as the trajectory parameterization method. A methodology is constructed which switches between high-thrust and low-thrust legs to constrain the input space for the low-thrust optimization problem. By using close-approach graphs and optimizing multiple flybys at the same time a trajectory with two Kuiper belt object flybys is found. The result is a robust method to find Kuiper belt object flyby trajectories with low-thrust propulsion.

To answer the question of habitability of other planets, it is crucial to find liquid water. As a planet’s surface might be difficult to characterise through observations, the observation of cloud composition and coverage could possibly reveal the presence of large bodies of surface water. Climate code SPEEDY is used to investigate relations between cloud patterns on rocky exoplanets with oceans for various planet parameters, such as obliquity and incident stellar flux, and the observable signals of such exoplanets are computed. SPEEDY was chosen for the modelling of rocky exoplanets, because of its speed, since our aim is to run simulations for various planet parameters, and its flexibility, since it allows the adaptation of planet properties such as the presence and distribution of continents. The planet’s rotational period is found to have the most obvious influence on the cloud pattern: with increasing rotational speed, bands of clouds form, parallel to the equator, with the number of bands increasing with the rotational speed. The total flux and polarisation of starlight that is reflected by the planets with cloud bands as functions of the wavelength and the planetary phase angle are computed. Also the influence of the integration time of the observations on the reflected light signals is studied. Our main recommendation for further research is to broaden the applicability of SPEEDY for exoplanet research by first gaining more insight into the parametrisations and then by adapting them where necessary to allow wider parameter settings.

The exponentially increasing number of Earth-orbiting satellites, as well as the pronounced importance for space military operations, impose new challenges on a functional Space Domain Awareness (SDA). A key factor for an effective SDA is the successful identification of an Earthorbiting satellite, or object in general. Current methods are usually limited to either time-resolved photometry, spectrometry, or polarimetry. In this thesis, we present the Spectropolarimeter for Satellite Identification (SSI), whose design combines these detection methods with the aim to provide a unique ‘fingerprint’ of the satellite, thereby expanding the range of satellite identification methods. The SSI combines a dual-channel spectrometer with off-the-shelf polarization optics. Using a combination of an achromatic quarter-wave retarder, a highly chromatic multiple order retarder, a polarizing beam splitter, and two linear polarizers, all operating in the visible spectrum, a sinusoidal modulation of the linear polarization information is imprinted onto the measured spectrum. The amplitude of the signal scales with the degree of linear polarization and the phase with the angle of linear polarization. Furthermore, by using a dual-channel setup, the intensity spectrum can be obtained for the full spectrograph resolution of 9 nm. We introduce and verify the retrieval algorithm to attain the linear polarization information with an approximate resolution of 135 nm due to the modulation approach. Differential transmission between both channels is corrected by a proposed iterative transmission correction which is verified to have an effect of ≤ 1% on the accuracy of the retrieved linear polarization information. This thesis concentrates on geostationary satellites as observation targets, making use of the stable observation geometry to simplify calibration and analysis. We reason that the phase angle has a strong influence on the retrieved spectropolarimetric data and received photon flux. A thorough photon budget estimation is conducted with the conclusion that we are able to observe faint geostationary satellites with an apparent visual magnitude of 10, especially if the observed phase angle is kept minimal. An analysis of the most probable errors is conducted. We identify the convolution of the instrument response function with the modulated signal and the induced polarization by the telescope and instrument as the most severe ones. Laboratory wavelength and polarimetric calibration are conducted. Using a double glass plate setup to induce low levels of degree of linear polarization we performed laboratory validation measurements. We determined the overall laboratory polarimetric accuracy to be ≤ 1.2%. Additionally, we are able to distinguish signals ≥ 1.3 %. Preliminary observations of three bright stars are conducted. We are able to retrieve instrumentinduced polarization between 1% and 2.5% which is in the expected range. We conclude that there has been a misalignment between the telescope and instrument which causes a large discrepancy between expected and received photons and propose future measures to resolve the issue. The proof of concept to use linear channelled spectropolarimetry for satellite identification with the SSI will open up the way for further technical refinement of the instrument, an extension to low Earth orbit satellites, and the possible addition of phase angle resolved spectropolarimetric measurements to further improve the concept of a satellite‘s spectropolarimetric ‘fingerprint’. Our work is part of a collaboration between Leiden observatory and Delft University of Technology commissioned by the Royal Netherlands Air Force. Access to the telescope and support with the alignment activities were kindly provided by TNO.

In this thesis, we provide a method for reconstructing a planet's surface map from its reflected light curve. We are going to derive an equation for the reflective light-curve under the assumption that the surface map is characterized by four different surface types (ocean, vegetation, sand and snow), is stationary (no clouds), and that all reflection is diffuse (Lambertian). We will show that the transformation is a linear function of the surface map and we will work out the transformation for arbitrary observer inclination and axial tilt. Using this knowledge, we create mock light curve data of self-generated planets. Afterwards, the transformation is inverted using the Moore-Penrose pseudo-inverse and the mock data will be used to demonstrate the surface map recovery for edge-on and face-on observations of planets with different axial tilts. Furthermore, we also provide a method for recovering the planet's axial tilt from its reflected light curve.

Even when a realistic amount of photon shot noise is added to the light curve, we are able to retrieve the planet's surface map and axial tilt fairly well, especially when the planet's tilt has larger axial tilt.

For a human Mars exploration mission, it is required to minimize the time-of-flight of the interplanetary trajectories to mitigate the adverse effects of the radiation environment and prolonged weightlessness conditions on the health of the crew. Moreover, having a fail-safe provision to return the crew back to Earth in case of a mission abort situation is highly desirable. This thesis research investigated several optimal/sub-optimal and feasible solutions for the high-thrust interplanetary transfer and abort trajectories of a reconnaissance human Mars mission. A conjunction-class mission architecture has been analyzed to transport the crew to and from Mars during the transfer opportunities in 2028 and 2030, respectively. Baseline requirements of such a mission were defined by finding the optimal ballistic Earth-to-Mars and Mars-to-Earth transfer trajectory solutions. A number of propulsive abort trajectory solutions were then computed that can return the crew back to Earth when the nominal mission is aborted either during the Earth-to-Mars transfer or during the Mars surface stay period. A semi-analytic trajectory model was used for the design of such abort trajectories that can include one or two deep space maneuvers and a powered Mars swing-by. With the use of a meta-heuristic global optimization algorithm, multi-objective optimization problems were solved to minimize the ΔV cost and the interplanetary transit duration of the mission. By comparing the transfer and abort trajectory solutions, it was concluded that various abort trajectory options can be provided for a reconnaissance human Mars mission, that also satisfy the imposed constraint on the arrival excess velocity. The effects of such abort trajectory options on the ΔV cost and other baseline mission requirements (such as the total mission duration) were analyzed. Critical challenges of such a safe human Mars mission architecture have been identified and discussed.

This thesis studies the radiation dose deposited on a 6U CubeSat in interplanetary orbit. The European Space Agency’s radiation data tool SPENVIS has been used to determine the proton flux in interplanetary space during a solar particle event which occurs only once every one hundred years. SPENVIS has further been applied to obtain a dose depth curve for this specific 1-in-100 years proton flux. Simultaneously, 6U simplified models of several materials are created as well as a realistic 6U CubeSat structure which represents the state-of-the-art in CubeSat industry. Using radiation simulation tool Fastrad, two simulations have been performed: ray-tracing simulation and forward Monte Carlo simulation. Ray-tracing simulation provides a sector shielding analysis of the models, giving an overview of the received radiation dose for each specific point in the models. Forward Monte Carlo gives an overview of the deposited radiation for each part of the model, which has shown especially beneficial for the state-of-the-art model, in which deposited dose has been obtained for each part in the model. This has led to a dataset that shows the radiation dose deposited at each location and in each part of the 6U CubeSat structure during a 1-in-100 years solar particle event that strikes in interplanetary space.

Fragments of asteroids and comets constantly encounter the Earth and Moon in their orbits, impacting them as meteoroids. Observations of meteor showers on Earth have been studied for at least 50 years, in order to construct accurate Solar System meteoroid models. More recently, Earth-based telescopic observations of the light flashes produced by lunar meteoroid impacts have revealed useful in the validation and improvement of such meteoroid models. However, Earth-based lunar observations are restricted by weather, geometric and illumination conditions. As such, it has been proposed that a lunar orbiter could improve the detection rate of lunar meteoroid impact flashes. Assessing which orbit a spacecraft should fly in order to detect these flashes and improve current Earth-based observation methods is the aim of this thesis. The study is restricted to spacecraft with the CubeSat format and its inherent limitations, since it is also inserted in the context of the feasibility study of the Lunar Meteoroid Impacts Observer (LUMIO) and ESA’s Lunar CubeSats for Exploration challenge.

A methodology of sequential orbital trade-offs was followed, taking into account acceptance criteria based on the mission requirements and selection criteria based on the research objective. The goal was maximise the number of meteoroid detections, during the mission lifetime, while minimising the mission ∆V budget. Circular Frozen Orbits, Earth–Moon L2 Lyapunov, Halo, Near-Rectilinear, Vertical, Distant-Retrograde and Low-Prograde orbits were selected as candidates, based on a preliminary orbital trade-off. In order to determine the total number of meteoroid detections possible from a certain orbit, the lunar meteoroid environment was modelled and a coverage analysis tool was developed to determine the payload FOV-area in the lunar nightside.

Frozen Orbits were found to allow the detection of meteoroid impacts with kinetic energies too small with respect to the desired, so the design space was restricted to CRTBP orbits. From Lyapunov, Halo, Near-Rectilinear and Vertical Orbits it would be possible to detect between 1000 and 10000 impacts during the mission lifetime, but detections from some DROs could be one order of magnitude larger. Nonetheless, since transfer costs to DROs are known to be high, a Near-Rectilinear Orbit, with a minimal ∆V budget, was chosen as the operational orbit.

Objects ingeostationary transfer orbit (GTO) can collide with operative satellites in lowEarth orbit (LEO) and geostationary orbit (GEO). Various organisationshave laid down debris-mitigation guidelines that will be enforced by law forfuture launchers. One of the guidelines entails proving that the generateddebris will re-enter in less than 25 years with a 90% probability. Naturalperturbations can be exploited to meet this requirement without the use ofextra propellant or complex de-orbiting systems, which is especially attractivefrom an economic point of view. Objects in GTO can undergo a resonancetriggered by an interplay between perturbations caused by the Sun’s gravity andthe irregularities in Earth’s gravity field, leading to a sudden re-entryor making the object stay in orbit for decades. This effect is very sensitiveto initial conditions because it depends on the relative positions of theperigee and the Sun when the semi-major axis is close to 15000 km.By simulating the orbital evolution of a representative GTO object—ballistic coefficient of 0.011 m²/kg, initial orbital inclination of 10degrees and initial perigee altitude of 200 km— for several initialepochs, it was found that favourable launch conditions take place twice perday during most part of the year, while for epochs close to the equinoxesof March and September they only happen once per day or not at all. Giventhe high sensitivity to initial conditions, the problem was studied from astatistical perspective, taking into account the uncertainties in thevalues of the relevant parameters. Semi-analytical techniques were used topropagate the mean equinoctial elements instead of the osculatingCartesian elements, which reduced computation times by a factor of 45 whilestill keeping proper levels of accuracy. Current practice for GTO launches fromKourou is to launch at around 6-7 PM. It was found that the launchtime leading to the highest probability of compliance with debris-mitigationguidelines for GEO launches from the European spaceport in Kourou isslightly later, regardless of the day of the year, although the value of theoptimal lifetime does vary slightly throughout the year. Thus, achange in procedures would be required in order to reach a higher degreeof compliance with debris-mitigation guidelines, which was below 10% for GTOlaunches carried out with Ariane 5 from in the period 2004-2012.