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Mubody is an astrodynamics open-source Python library focused on the libration points. Such points result from the equilibrium of the gravitational forces between two massive bodies as the Sun and Earth, for example. The library is mainly intended for the generation of orbits in these regions, which is not a straightforward process, specially if perturbations are considered. Currently, the library allows to generate Lissajous orbits in the second Lagrange point of the Sun-Earth system under the influence of perturbations such as the Earth orbit eccentricity. The next milestone, as a result of a master student work, is the incorporation of Halo orbits and the expansion to all three collinear libration points from any two massive bodies of the Solar System. This tool has been developed as part of a PhD, motivated by the need of performing mission analysis in libration point regions. Nevertheless, since its creation it has also proven to be an excellent academic tool for both enhancing the library itself and using its results for further studies (collision risk, thermal analysis, formation flight control, etc). As a result, the tool has rapidly evolved, building onto the knowledge and experience that the students gather while working on their academic projects (bachelor’s degree dissertations, master theses, subjects, internships). The participation on the library development provides students with experience in orbital mechanics, software design, version control and it compels them to ensure that their work can be readily used by others as it is properly documented. The project is hosted in GitLab under a MIT licence

The Satellite for Orbital Aerodynamics Research (SOAR) is a CubeSat mission, due to be launched in 2021, to investigate the interaction between different materials and the atmospheric flow regime in very low Earth orbits (VLEO). Improving knowledge of the gas-surface interactions at these altitudes and identification of novel materials that can minimise drag or improve aerodynamic control are important for the design of future spacecraft that can operate in lower altitude orbits. Such satellites may be smaller and cheaper to develop or can provide improved Earth observation data or communications link-budgets and latency. Using precise orbit and attitude determination information and the measured atmospheric flow characteristics the forces and torques experienced by the satellite in orbit can be studied and estimates of the aerodynamic coefficients calculated. This paper presents the scientific concept and design of the SOAR mission. The methodology for recovery of the aerodynamic coefficients from the measured orbit, attitude, and in-situ atmospheric data using a least-squares orbit determination and free-parameter fitting process is described and the experimental uncertainty of the resolved aerodynamic coefficients is estimated. The presented results indicate that the combination of the satellite design and experimental methodology are capable of clearly illustrating the variation of drag and lift coefficient for differing surface incidence angle. The lowest uncertainties for the drag coefficient measurement are found at approximately 300 km, whilst the measurement of lift coefficient improves for reducing orbital altitude to 200 km., Comment: 19 pages, 11 figures. Accepted for publication in Acta Astronautica (2020-12-14)

Sending spacecrafts to any other planet or satellite within the Solar System is a difficult task. To be able to reach it at a precise time, it is crucial to know at which position the target planet is going to be, as well as a good study of the trajectory that the spacecraft is going to follow, which shall not be too close to any other object in the Solar System. It must be also be taken into account that the spacecrafts have a limited fuel capacity. In order to avoid using a high amount of fuel, multi-gravity assist trajectories are useful to provide additional manoeuvering to the spacecrafts without consuming fuel. Given an arbitrary sequence of planets for the swing-by manoeuvres, the next step is to determine when the spacecraft must be launched and how long will it take to it to perform the transfer trajectory. In this preliminary stage of the mission analy-sis, optimization algorithms can be used to obtain the most efficient trajectory, for instance, the trajectory at which the spacecraft requires the least amount of fuel to reach its destination. From the wide variety of heuristic optimization algorithms, in this Master’s thesis four of them have been chosen (Genetic Algorithm, Differential Evolution Algorithm, Particle Swarm Algorithm and the simulated annealing algorithm). The aim of this thesis is to perform a hyperparameter optimization, that is, finding whether a certain combination of parameters for each optimizer works properly for different trajecto-ries’ scenarios: a direct launch to an outer planet and a multi-gravity assist scenario to outer planets as well as to inner planets. The results show that it exists a region of parameters for the Genetic Algorithm and the Differential Evolution Algorithm where the best solutions can be found. Simulated Annealing Algorithm shows a dependance on the scenario, as well as some parameters of the Particle Swarm Algorithm.

With the miniaturization of the electrical components and the advance in aerospace technology, new methods of spatial propulsion have been developed. This thesis project is focused on solar radiation pressure modelling, solar sail kinematics and dynamics and attitude determination and control. Solar sailing is based on the same principle as any conventional propulsion system, Newton's third law. However, solar sails gain momentum from Sunlight photons and thus can provide continuous acceleration. The momentum carried by photons is small and for that reason solar sails must have a large area and be extremely light. Since conventional actuators are not suitable for attitude control in solar sailing, specific methods based on Reflectivity Control Devices (RCDs) have been discussed.

© 2017. This version is made available under the CC-BY-NC-ND 4.0 license http://creativecommons.org/licenses/by-nc-nd/4.0/ We present and discuss a solution to the growing demand for satellite telecommunication coverage in the high-latitude geographical regions (beyond 55°N), where the signal from geostationary satellites is limited or unavailable. We focus on the dynamical issues associated to the design, the coverage, the maintenance and the disposal of a set of orbits selected for the purpose. Specifically, we identify a group of highly inclined, moderately eccentric geosynchronous orbits derived from the Tundra orbit (geosynchronous, eccentric and critically inclined). Continuous coverage can be guaranteed by a constellation of three satellites in equally spaced planes and suitably phased. By means of a high-precision model of the terrestrial gravity field and the relevant environmental perturbations, we study the evolution of these orbits. The effects of the different perturbations on the ground track (which is more important for coverage than the orbital elements themselves) are isolated and analyzed. The physical model and the numerical setup are optimized with respect to computing time and accuracy. We show that, in order to maintain the ground track unchanged, the key parameters are the orbital period and the argument of perigee. Furthermore, corrections to the right ascension of the ascending node are needed in order to preserve the relative orientation of the orbital planes. A station-keeping strategy that minimizes propellant consumption is then devised, and comparisons are made between the cost of a solution based on impulsive maneuvers and one with continuous thrust. Finally, the issue of end-of-life disposal is discussed.

In the circular restricted three-body problem (CR3BP) the weak stability boundary (WSB) is defined as a boundary set in the phase space between stable and unstable motion relative to the second primary. At a given energy level, the boundaries of such region are provided by the stable manifolds of the central objects of the L1 and L2 libration points, i.e., the two planar Lyapunov orbits. Besides, the unstable manifolds of libration point orbits (LPOs) around L1 and L2 have been identified as responsible for the weak or temporary capture around the second primary of the system. These two issues suggest the existence of natural dynamical channels between the Earth's vicinity and the Sun–Earth libration points L1 and L2. Furthermore, it has been shown that the Sun–Earth L2 central unstable manifolds can be linked, through an heteroclinic connection, to the central stable manifolds of the L2 point in the Earth–Moon three-body problem. This concept has been applied to the design of low energy transfers (LETs) from the Earth to the Moon. In this contribution we consider all the above three issues, i.e., weak stability boundaries, temporary capture and low energy transfers, and we discuss the role played by the invariant manifolds of LPOs in each of them. The study is made in the planar approximation.

The objective of the present study is to analyze the evolution of low lunar orbits under the gravitational potential of the Moon and additional perturbations of gravitational (Earth and Sun gravity) and non-gravitational (radiation pressure) nature. Such evolution can be used to estimate the amount of propellant required to maintain a certain orbit during a given time or, alternatively as in the present case, to identify orbits which may offer a wide surface coverage if no orbit correction maneuvers are applied.