Given the potential of solar sails for applications around Earth, this paper investigates their performance in executing collision avoidance maneuvers. An extensive set of conjunctions (e.g., varying orbital regimes, sail control authorities, and collision geometries) is built. An "ideal" and a "real" case are examined, with the latter accounting for greater uncertainties (modeled via a covariance-mapping algorithm) and stricter collision probability requirements. The optimal control problem is derived and solved through direct collocation, seeking time optimality while constraining the probability. For all scenarios, maneuvers are completed within the one-day threshold, proving solar sails fit the current operational framework. Key findings are: 1) Better debris state knowledge (and later warnings) reduce maneuver time; 2) Higher altitudes increase the maneuver time, while greater sail control authority decreases it; 3) Eclipses influence results; 4) The Earth-Sun configuration only weakly affects the maneuver time; and 5) In-plane control is always non-zero. The locally optimal steering laws to change the semi-major axis and eccentricity best approximate the optimized maneuvers: especially for longer maneuver times, these straightforward laws comply with collision probability requirements.

Solar-sailing is a promising propellant-free propulsion method leveraging the momentum of photons to generate a thrust force, making them attractive for long-term missions both in Earth-bound and interplanetary space. In Earth orbit, solar-sails have been envisioned for space debris removal missions aiming to de-orbit multiple defunct satellites. However, their large sail area make them vulnerable to hypervelocity impacts with debris, potentially causing loss of attitude control. Therefore, this thesis presents a study of the long-term effects of tumbling on a sail's orbit and of the capability of a modern vane attitude control system to time-optimally stabilise the attitude motion. The tumbling dynamics result in an orbital eccentricity growth which is independent of the tumbling rate, potentially leading to a re-entry of the sail. The vane system is capable of detumbling rotational velocities up to 26 deg/s. For rotational velocities up to 8 deg/s, this system is capable of detumbling the sailcraft at a linear rate of 2 deg/s per day. For larger rotational velocities, the duration of the detumbling manoeuvre grows non-linearly. These results are considered for the specific case study of hypervelocity impacts and the sensitivity of the results to the sail reflectance model, the orbital regime, and the number of degrees of freedom of the vanes is assessed.

This work investigates the possibility of setting up an Earth-asteroid cycler to asteroid 2001AE2 using solar sail technology. An ideal solar sail model with near-future technology levels (𝛽 = 0.05) is implemented alongside the circular restricted three-body problem, two-body problem and the elliptic Hill problem. The spacecraft cycles between the Sun-Earth 𝐿2 point (SE − L2) and the Sun-asteroid 𝐿1 point (Sa − L1). From these equilibrium points, the spacecraft state vector is propagated forwards and backwards with a constant sail attitude to generate solar-sail assisted invariant manifolds. Initial guess trajectories for the outbound and inbound sections of the cycle are generated with a genetic algorithm by searching for optimal sail attitudes for the different dynamical models and departure, connection and arrival times. These initial guess trajectoryies are used to seed the optimization software PSOPT. This pseudospectral collocation method using Legendre-Chebyshev polynomials transforms the infinite dimensional control problem into a finite dimensional non-linear programming problem. This way a continuous control profile can be found that justifies the dynamical models and results in a time-optimal trajectory. The resulting cycler is designed to have a cycle time of 11.04 years, with an outbound trajectory departing from SE − L2 on 01 − 03 − 2036 and arriving at Sa − L1 on 02 − 04 − 2039 and an inbound trajectory departing from Sa − L1 on 04 − 08 − 2043 and arriving back at SE − L2 on 31 − 03 − 2046. This results in dwelling times at SE − L2 of 351 days (0.95y) and at Sa − L1 of 1585 days (4.34y). The trajectories are assumed time-optimal given the assumptions made in the work, but require further refining for non-ideal properties of the solar sail and other higher fidelity models.

Now that a rocky planet is confirmed to orbit in the habitable zone of our closest stellar neighbor Proxima Centauri, the interest in visiting that system is growing; especially since Breakthrough Starshot proposed a fly-through mission of the Alpha Centauri system by sending a swarm of laser-driven photon sails. While many engineering problems still need to be solved for such a mission to succeed, research has shown that futuristic, theoretical photon-sail configurations can reach the Alpha Centauri system within 75-80 years while also getting captured in a bound orbit about one of the binary stars. This paper investigates trajectories from the binary star system towards planet Proxima b. A mission to Proxima b is scientifically grounded since measurements or pictures could help us better comprehend the evolution of rocky planets and potential life-formation in our Universe. The classical Lagrange points in the binary system (AC-A/AC-B) and the system Proxima Centauri-Proxima b (AC-C/Proxima b) are used to find possible trajectories towards Proxima b. The transfer is divided into a departure phase from AC-A/AC-B and an arrival phase to AC-C/Proxima b. Heteroclinic connections are then exploited using a patched restricted three-body problem method to connect the two phases. A grid search is applied on the optimization parameters to explore the design space, after which a genetic algorithm is applied to further optimize the link, focusing on minimization of the position, velocity, and time error at linkage. Futuristic sail configurations are used, including double-sided reflective sails and lightness numbers up to ß = 1779. The design space exploration shows that a double-sided sail provides little improvement over a one-sided sail, mainly due to the constant sail attitude along the trajectories. Results from the genetic algorithm show that a transfer from the L2-point in the AC-A/AC-B system to the L1-point in the AC-C/Proxima b can be accomplished with a transfer time of 235 years for the one-sided graphene-based sail with a surface of 315x315 m^2 carrying a payload of 10 grams. A transfer from the L2-point in the AC-A/AC-B system to the L3-point in the AC-C/Proxima b, with a smaller one-sided graphene-based sail (75x75 m^2, carrying a payload of 10 grams), results in a transfer time of 1025 years. For both sail configurations, the position error at linkage is kept below 1% of the total travel distance, the velocity error below 1% of the velocity at linkage, and the time error below 1% of the total transfer time.

Recent studies have shown the feasibility of differential dynamic programming (DDP) in optimizing Earth-centered solar-sail trajectories. In order to further demonstrate the ability of DDP in the optimization of solar-sail trajectories, this work investigates the performance of DDP for optimizing interplanetary solar-sail trajectories. The selected dynamical framework is based on the two-body problem, augmented with an ideal solar-sail force model. A superior numerical performance is obtained for the optimization algorithm by propagating the state in modified equinoctial elements and applying a Sundman transformation to change the independent variable from time to the true anomaly. The developed algorithm finds similar or more optimal solutions than locally optimal steering laws for the maximization of different orbital elements. In addition, constrained time-optimal Earth-Mars orbital transfers are investigated for different sail performance levels. The DDP algorithm is proven to be efficient and robust for different optimization settings and initial guesses for solar-sail trajectory optimization in the interplanetary regime.

This paper investigates the use of solar sailing propulsion to visit as many co-orbital near-Earth asteroids (NEAs) as possible, within a fixed time-frame. This research builds on previous publications, which have shown solar sailing to be a suitable propulsion method to visit NEAs. The dynamics of this problem are modelled within the Solar Sail Augmented Circular Restricted Three-Body Problem (CR3BPS), and assume a near-term solar sailing technology level. A sequence generation algorithm is developed which generates trajectories to visit multiple co-orbital NEAs beginning at either the artificial co-linear equilibrium point SL1 or SL2. This algorithm develops trajectories with fixed controls to transfer between target asteroids, using Monte Carlo simulations to propagate a wide range of random combinations of settings before selecting those that perform the best. It is shown that the tuning performed within this research can generate a trajectory that enables 18 asteroid fly-bys within the selected nominal mission lifetime of ten years. Following this sequence generation, the first fly-by of the trajectory is optimised as proof of concept that each leg of the trajectory can be optimised for fly-by distance and velocity. An optimal control problem is developed, which is then implemented and solved using direct pseudospectral methods. The solution to this optimal control problem reduces the fly-by distance by 99.95 %, down to 158.73 km, while reducing the fly-by velocity by 9.68 % to 4.33 km/s.

Operating spacecraft in a perturbed environment of a binary asteroid system is a challenging task. In light of the near-future exploration of the 65803 Didymos system by the Hera probe and the lack of study of orbital evolution of naturally-levitated regolith particles in this system, a method is here proposed to identify regions of high risk of collision with the levitated regolith grains. Regions of regolith levitation are identified, periodic orbits and regions of stable motion are computed through a grid search method, and the distance between trajectories leading from the off-surface levitation of the grains from the primary body and the trajectories of bounded motion is then assessed to determine the occurrence of temporary capture. A qualitative evaluation of the expected patterns of motion of regolith particles is presented together with a discussion of the key conclusions in the context of in situ operations planning for the Hera probe.

Thousands of space debris objects remain in space after decades of spaceflight without sustainability regulations. Some of these objects de-orbit naturally, but not objects in the geostationary region for which active debris removal is needed. This thesis presents the case of geostationary debris removal using solar sailing, a propulsion method where no propellant is consumed, which allows for repeatable debris removal. A focus is given to find minimum-time trajectories between the geostationary region and a so-called ‘graveyard orbit’ and back. Optimisation using numerical methods is time-consuming, as these trajectories include hundreds of orbit revolutions. This work utilises analytical control methods instead to determine solar-sail control, specifically the ’Accessibility-and-deficit blending method’ based on locally optimal steering laws. Resulting mission times are compared with literature for validation, and a catalogue of mission time is found for a wide range of mission parameters, including additional thrust from a solar-electric propulsion engine.

The development of solar-sail technology in combination with the rising interest in a mission to the Sun-Earth L5 region for heliophysics and the search for Trojan asteroids, raises the question of using solar sailing as the primary propulsion method to enable such a mission. This study therefore investigates different invariant objects and their properties in the neighbourhood of Earth and in the L5 region that could be used as departure and arrival conditions: equilibrium points, families of periodic orbits and families of invariant tori as well as the stable manifold of periodic orbits. Then, transfers between these invariant objects are studied using a hybridisation of different trajectory design techniques. A multi-objective genetic algorithm is applied to obtain near-feasible initial guesses, which are transformed into feasible transfers using a differential correction method. Through a continuation on the fixed time of flight, the differential corrector is subsequently used to reduce the transfer time. A pseudospectral optimisation method is finally taken at hand to obtain a smooth, continuous control profile, to, if possible, further reduce the transfer time. This approach results in fast solar-sail transfers between 391 and 1194 days, depending on the departure and arrival configuration and the assumed solar-sail technology. These results can serve as preliminary design solutions for a mission to the Sun-Earth L5 region.

This thesis presents the design of solar-sail transfer trajectories to a constellation of two spacecraft in displaced vertical Lyapunov orbits at the L2 point of the Earth-Moon system. The constellation provides continuous coverage of the Aitken basin and the lunar South Pole. Initial guesses for the transfers are generated using reverse time propagations of the dynamics, where the control is provided by a locally optimal steering law. These initial guesses are subsequently used to initialize a 12th-order Gauss-Lobatto collocation method. The minimum altitude with respect to the Earth and the Moon are constrained, as well as the maximum rotation rate of the solar sail. Sets of feasible trajectories for both spacecraft with identical launch conditions are sought, such that the constellation can be initiated using a single Soyuz launch. Such a Soyuz launch can deliver two 1160-kg spacecraft into the found transfer trajectories. The first spacecraft subsequently requires a transfer time of 53.06 days to enter its constellation orbit, while the transfer of the second spacecraft takes 67.89 days. This research demonstrates that solar-sail transfer trajectories are a feasible option for future missions in the Earth-Moon system.

Recent studies have shown the feasibility of (quasi-)pole-sitter orbits at Mars and Venus, which involves a satellite positioned along or near the polar axis of a planet in order to have a continuous, hemispherical view of the planet's polar regions. In order to further demonstrate the feasibility of this mission concept, this thesis investigates time-optimal solar sail transfers to these (quasi-)pole-sitters. In particular, (quasi-)pole-sitters which are achievable when assuming solar sail technology expected in a near- to mid-term time-frame. To reduce mission operational cost, the objective of this research is to minimize the time required for the transfer, which requires the solution to an optimal control problem. Initial guess solutions for this optimal control problem are provided through two completely different techniques, in order to compare and validate the individual performances: first, a technique derived from dynamical systems theory (a type of grid search) and second, a genetic algorithm. Subsequent optimization using a direct pseudospectral algorithm results in time-optimal transfers to the considered Mars (quasi-)pole-sitters that span 2.61 and 2.72 years, and 1.07 and 1.19 years to the considered Venus (quasi-)pole-sitters. Effects due to variations in performance of the ideal sail, non-ideal sail properties, and Earth departure orbit are investigated. In addition, this paper demonstrates that a genetic algorithm is well suited to generate initial guesses for similar interplanetary transfers in the inner solar system. It provides initial guesses that outperform the more conventional grid search technique, in terms of feasibility of the initial guess transfers, as well as in computation time and ease of implementation.