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.

Solar sailing is a promising propellantless propulsion method that employs large reflective surfaces to harness solar radiation pressure for spacecraft propulsion. Despite the fact that several solar-sail near-Earth missions will launch in the coming years, there is notable lack of published studies on the uncertainties associated with missions of this kind. This thesis addresses this gap in knowledge by quantifying uncertainties related to the solar sail's optical coefficients, structural deformations, and attitude profiles. Through two uncertainty propagation methods, namely Monte Carlo simulations and the Gauss von Mises method, the study reveals the significant impact of the optical coefficient uncertainties on mission performance. The results indicate a worst-case 3-sigma uncertainty of 8.1% in altitude gain and 16.5% uncertainty in inclination gain for the NEA Scout solar sail model. Specularity coefficient uncertainty emerges as the primary driver of performance uncertainty among the analyzed optical coefficients. Structural deformation, on the other hand, exerts minimal impact. Uncertainty in the attitude profile is modelled through Ornstein-Uhlenbeck processes and is found to impact mean mission performance as well as introduce performance uncertainty. Overall, this work demonstrates the critical importance of characterizing uncertainties and provides insights crucial for mission planning and decision-making.

The first mission proposals to visit the Alpha Centauri system use photon-sail acceleration as a mode of propulsion to reach this stellar system closest to our own Solar System. To prepare for a future mission, the photon-sail dynamics in the system is investigated. Planar Lyapunov orbits around the colinear classical Lagrange points are designed to explore the Alpha Centauri system. This has been done before in other systems like the Earth-moon and Sun-Earth systems, but not yet in an elliptical binary star system. Starting with an initial guess in the circular restricted three-body problem without photon-sail acceleration, a Multiple Shooting Differential Correction (MSDC) algorithm changes the trajectory to a periodic orbit. A continuation method increases the eccentricity to match e = 0.5208, which is the eccentricity of the inner binary system of Alpha Centauri. The lightness number of the photon sail is increased to add photon-sail acceleration to the model up to a defined maximum of ?ZNe = 2. A set of five constant steering laws is chosen to investigate its effect. Next to that, the moment at which the periodic orbit starts in terms of the true anomaly is varied as well. This results in a set of 40 families of periodic orbits with increasing lightness numbers. Depending on the orientation, the augmented Lyapunov orbit either shrinks into smaller orbits or expands into larger orbits when increasing the lightness number. If the orbit shrinks, it can either converge into an artificial equilibrium point or the photon-radiation pressure on the sail can become minimal. In that case, the Lyapunov orbit becomes (almost) independent on the lightness number and reaches ?ZNe = 2. If the orbit expands, the maximum velocity will eventually go to infinity. At this vertical asymptote, the maximum lightness number is found. The initial true anomaly of Alpha Centauri 0 has a great effect on the Lyapunov orbits around L2 and L3 in the classical ER3BP. For 0 = 0, the orbit either converges to an AEP or the maximum velocity goes to infinity. For 0 = , a few orientations can reach ?ZNe = 2. To further explore Alpha Centauri, an adaptive differential evolution algorithm is used to design trajectories between the Lyapunov orbits. The performance of the algorithm is expressed as the Euclidean difference between the states at the end of the departure leg and the beginning of the arrival leg. Three different lightness number of ? = 0.1, 0.5 and 2 are used for these trajectories. With a lightness number of 0.1, the dimensionless Euclidean error is in the range of 1E-1 to 1E-3, depending on the Lyapunov orbits. With this lightness number, the stars are also used as a gravity assist. For larger lightness numbers, the Euclidean error becomes negligible in the range 1E-7. With a lightness number of 2, the time of flight during the trajectory is significantly lower. In future research, this can be further decreased using an MSDC algorithm.

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.

Solar sailing is a flight-proven low-thrust propulsion technology with strong potential for innovative scientific missions. The heliogyro is promising sailcraft design that utilizes a set of long slender blades which are deployed and flattened by spin-induced tension and whose orientations can be individually controlled. The main advantages of such a design are the easier stowage and deployment, and potentially lower structural mass. The heliogyro’s translational and rotational motions are strongly coupled, with non-trivial relationships between the control inputs and the forces and moments produced by the sail. The purpose of this research is to investigate, for the first time, the coupled roto-translational motion of the heliogyro. Two dynamical models of the heliogyro motion are developed and applied to design Earth-to-Mars stopover cycler trajectories. The resulting heliogyro trajectories are then compared to those of a traditional fixed-area and flat sail-system design demonstrating potential advantages of the heliogyro.

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.

A planetary sunshade is a large, reflecting disk built to shield the Earth from a small fraction of solar irradiance, partly compensating global warming caused by greenhouse gas emissions. As a specific form of solar geoengineering, the sunshade is an emergency solution that would be implemented to prevent catastrophic climate change, while working towards the net-zero emission goal. In this paper, a dynamic sunshade is proposed. Such a system is capable of not only reducing the global mean surface temperature anomaly, but also minimizing regional climate changes by tailoring the sunshade's motion according to climate requirements. A sunshade orbiting in the vicinity of the Sun-Earth L1 point is able to reduce the global mean surface temperature from 16.39°C (scenario with 680 ppm of atmospheric CO2) to 14.13°C until equilibrium is reached. It also reduces the polar mean surface temperature by more than 2°C with respect to a scenario without sunshade.

This paper investigates the use of solar-sail technology to increase the warning time for Coronal Mass Ejections (CMEs) heading towards Earth. In addition, this research will build upon the current understanding of using solar-sail dynamics with regards to CME detection by providing insights into the problem characteristics. The warning time is proportional to the distance from the Earth to the spacecraft detecting the CME: a current warning time of 30 to 60 minutes is achieved by satellites at or near the Sun-Earth L1 point. By considering the actual shape of a CME, the continuous solar-sail acceleration from the solar sail can be used to find a periodic trajectory that travels further upstream of the CME-axis, thereby increasing the warning time with respect to current missions. Finding a periodic solar-sail trajectory can be regarded as an optimal control problem, which requires a near-feasible initial guess trajectory. the latter is found by generating heteroclinic connections between artificial equilibrium points in the vicinity of the sub-L1 and sub-L5 point through the use of a grid search and a genetic algorithm. The optimal control problem is solved with a direct pseudospectral method, resulting in four representative trajectories, each having specific (dis)advantages. The performance impact due to (the uncertainty of) non-ideal sail properties, change in lightness number, and variation in CME size are investigated. Ultimately, the most optimal trajectory increases the average and maximum warning time by a factor 20 and 30 with respect to current missions at L1, respectively, with a 90\% probability that the spacecraft detects the CME.

This thesis demonstrates the usability of differential dynamic programming (DDP) to obtain, for the first time, globally optimal Earth-centred solar-sail trajectories. To this end, DDP is combined with a global optimisation heuristic, monotonic basin hopping. The dynamical model is implemented as a two-body problem, augmented with an ideal solar-sail reflectance model and accounts for eclipses. The numerical performance of the optimisation algorithm is enhanced by integrating the sailcraft state in modified equinoctial elements and performing a Sundman transformation to change the independent variable from time to the true anomaly. The DDP algorithm is proven to be robust for trajectories extending up to 500 revolutions and, compared to known locally optimal steering laws, allows to obtain more or equally optimal solutions. The latter is demonstrated in this paper through a set of test cases that range from theoretical scenarios to realistic mission applications, including increasing the specific orbital energy of NASA’s upcoming ACS3 mission. Additionally, the algorithm's ability to cope with different optimisation settings, perturbing accelerations and constraints is demonstrated.

Space-weather events have a large impact on Earth. In particular, Coronal Mass Ejections (CME) pose huge potential dangers to human technology both in orbit and on the surface of the planet, such as disruptions to power grids, increased radiation doses to astronauts and damage to sensitive components of satellites. Warnings for space-weather events are currently given by satellites in the vicinity of the L1 point. When a CME passes by that point, the satellites emit a warning that reaches the Earth, on average, one hour before the CME. This thesis aims at using solar-sail technology to move a satellite closer to the Sun, detect the CME sooner, and thus increase the warning time.

Solar sails continuously generate thrust by reflecting solar photons off a large and highly reflective sail membrane. This continuous acceleration can be used to generate Artificial Equilibrium Points (AEPs) in the Circular Restricted Three-Body Problem (CRTBP) that are displaced away from the five classical Lagrange points. Like for the classical case, periodic orbits exist around these AEPs, enabling, for example, CME monitoring in a periodic orbit around an AEP that is located closer to the Sun than the Sun-Earth L1 point from where current satellites detect CMEs. There have been some theoretical mission designs taking advantage of this possibility, but the increase in warning time is modest for any near-term sail performance.

This thesis investigates the use of solar-sail technology to travel upstream of the CME and significantly increase the warning time. The study considers the actual shape of CMEs as a constraint for the solar-sail trajectory that surfs along invariant manifold-like structures emanating from AEPs and Lyapunov orbits around sub-L1 points, i.e., AEPs sunward of the classical L1 point, to travel upstream of the CME.

As a preliminary solution, two strategies are evaluated. The first strategy considers a series of heteroclinic connections between different AEPs in the sub-L1 region. The second strategy uses a homoclinic connection of a Lyapunov orbit around a sub-L1 point. The trajectories aim to travel upstream along the path of the CME and back to the initial AEP or Lyapunov orbit to guarantee periodicity and, therefore, CME monitoring for as long as the sail remains operational. The homo- and heteroclinic connections are sought for by looking for connections between the unstable and stable manifolds emanating from the AEPs and Lyapunov orbits. To minimize the discontinuity in states at the linkage of the unstable and stable manifolds, a genetic algorithm approach is used to optimize the piece-wise constant attitude of the sail along the manifold trajectories and the location, i.e., which AEP or where along the Lyapunov orbit, from where the manifolds emanate.

Though the homo- and heteroclinic connections exhibit a discontinuity in the attitude of the sail at the connection of the unstable and stable manifolds, they provide a good initial guess for further optimization with a direct pseudospectral method, implemented in the software tool PSOPT. In the optimal control approach, the attitude of the sail is allowed to vary along the trajectory (instead of the piece-wise constant sail attitude in the genetic algorithm approach) such that the sail travels as far as possible upstream of the CME while staying as close as possible to the central axis of the CME.

The genetic algorithm results from both strategies show an improvement in warning time with respect to the warning time achieved by current satellites in the environment of the L1 point. Under the assumptions taken in this research, the trajectories using the homoclinic connections from a Lyapunov orbit out-perform those that employ heteroclinic connections between AEPs. The best genetic algorithm solution offers an up to 10 times longer warning time than current satellites at L1.
This solution shows a small discontinuity in the states at the linkage of the unstable and stable manifold trajectories in the order of thousands of kilometers for the position and centimeters per second for the velocity. However, the
discontinuity in the attitude of the sail of approximately 70 degrees renders the trajectory unfeasible before further optimization.
Furthermore, some parts of the trajectory are too far from the axis of the CME to intercept CMEs approaching the Earth.

Finally, the trajectories obtained with PSOPT show that the sail remains within a defined distance to the axis of the CME while traveling upstream of the CME due to a control law that modifies the attitude of the sail at a rate achievable with state of the art technology. This strategy allows a 15 times longer average warning time compared to the warning time provided by current satellites at the L$_1$ point.

This paper presents time-optimal transfer trajectories between planar solar-sail displaced libration point orbits around the $L_1$ and $L_2$ points of the Sun-Earth and Earth-Moon systems. Initial guesses for these transfers are generated as two-segment trajectories, with a fixed sail attitude along each segment. The position and velocity discontinuity between the segments is minimized by means of a genetic-algorithm approach. The time of flight of these trajectories is then optimized by means of a direct pseudo-spectral collocation method as well as a multiple shooting differential correction (MSDC) method. The results of these methods are subsequently compared. Pseudo-spectral collocation was found to outperform MSDC in the solar-sail augmented Sun-Earth system, obtaining trajectories with transfer times as short as 101 days. Differential correction generally yields trajectories with slightly longer flight times than pseudo-spectral collocation, but converges more reliably. In the time-dependent solar-sail augmented Earth-Moon system, pseudo-spectral collocation fails to provide reproducible results. Transfer trajectories with flight times as short as 15.6 days are obtained in the Earth-Moon system using MSDC, but these trajectories feature control discontinuities that are likely to be problematic in practice.

Recognising the threat of near-Earth objects (NEOs), many ground-based surveys have been deployed worldwide. However, ~20% of these potential impactors are estimated to be approaching us from the day-side, and are thus very difficult to detect using ground surveys. Over the last decade, several space-based capabilities have emerged in an effort to discover and catalogue NEOs, yet little research has gone into dealing with imminent-impacting NEOs. The aim of this thesis is to perform a mission analysis of a space-based telescope that provides warning for Earth-impacting NEOs down to 20 m in size, by determining the performance of both a visible and an infrared (IR) space-based telescope used in two different mission candidates. The best space-based NEO survey system is concluded to be an IR telescope placed at the Sun-Earth solar-sail displaced L1 point due to the long warning times obtained and the beneficial contribution to existing ground-based surveys.

In the framework of JAXA’s Hayabusa2 mission, we study the dynamical environment around asteroid Ryugu to investigate whether ejecta particles can be temporarily trapped in periodic orbits following the Small Carry-on Impactor (SCI) operation. If these particles remain about the asteroid, they could potentially jeopardize the mission as, in the event of a collision with the Hayabusa2 spacecraft, the spacecraft’s functionality could be reduced. In
this paper, we make use of invariant manifold theory to assess the conditions - impact location, particle radius, ejection velocity - that cause ejecta particles to get captured in periodic orbits. The analysis is carried out within the dynamical framework of the perturbed Augmented Hill Problem, which takes into account the solar radiation pressure, the effect of eclipses, and the J2 and J4 terms of the asteroid’s gravity potential in its spherical harmonics expansion. We analyze millimeter to centimeter sized particles and captures into three families of periodic orbits that are robust to large values of the solar radiation pressure acceleration – the traditional a and g’ families of the Hill Problem and the southern halo orbits. We go on to find the impact locations for the SCI from where ejecta particles are most likely to be captured into periodic orbits via their stable manifolds. As such, we recover the sets of initial states that lead ejecta to temporary orbital capture and show that solar radiation pressure cannot be neglected in these analyses, identifying locations on the Sun side of the asteroid at medium latitudes as the best impact locations.

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.

The propellantless nature of solar-sail propulsion allows researchers to design completely new sets of orbits unreachable or not-maintainable by means of conventional propulsion. The advantage of this somewhat new type of low-thrust propulsion system becomes more evident for long-duration missions where chemical or electrical rocket engines would run out of onboard reaction mass. Solar sails happen to be particularly well-suited for polar observation missions. One can find several works concerning solar-sail orbits to establish a permanent communication link between a settlement on the lunar South Pole and the Earth. In addition, the literature also encourages the usage of solar sails to monitor the melting of the polar ice caps. However, these solutions employ either complicated steering laws or more than one sailcraft. In addition, little is known about how to perform necessary station-keeping maneuvers to keep the sailcraft bounded to the reference orbits under the effect of dynamic perturbations.

This thesis focuses on the orbit control of three types of solar-sail periodic orbits within the Earth-Moon system that have been shown potential for polar observation of either body. One of these orbit types, coined the distant-circular family, is newly introduced in this thesis together with transfers
between orbits within the family; its remote orbits can achieve continuous coverage of both the Earth's North Pole and the lunar South Pole with just a single sailcraft. Developed mainly under the simplified, but non-autonomous, dynamics of the solar-sail circular restricted three-body problem, the studied orbits are all unstable and require an active control strategy for station-keeping. By slowly migrating the dynamics to a higher fidelity model, we introduce several disturbances, namely those associated with the eccentricity of the Moon's orbit, the plane offset between the ecliptic and the Earth-Moon orbital plane, the Sun's gravity, and the heliocentric eccentricity of the system's barycenter. Two solar-sail architectures, a fixed-shape and heliogyro sails, are compared with respect to the objective of tracking the ideal reference orbits that all use a simple Sun-sail steering law as reference control. To do so, we develop a methodology based on multiple shooting differential correction, weighted least squares, particle swarm optimization (quantitative analysis) and acceleration bubble visualisation (qualitative analysis).

We find that the effect of the eccentricity of the Moon's orbit is too large to be considered a perturbation and should be included in the design of the reference orbits. Then, we show that other perturbations can be completely counteracted by the heliogyro, which suggests that pure solar-sail quasi-periodic orbits are maintainable.

With the increased interest for interstellar exploration after the discovery of exoplanets and the proposal of the Breakthrough Starshot project to perform a fly-through mission of Alpha Centauri to capture the first images of an Earth-like exoplanet, this paper investigates the optimization of photon-sail trajectories in Alpha Centauri for various mission applications. The prime objective is to find the optimal steering strategy for the photonic sail to get captured in Alpha Centauri after a minimum-time transfer from Earth. By extending the idea of the Breakthrough Starshot project with a deceleration phase at arrival, the mission's scientific yield is increased. In addition to capture into an orbit at the center of the habitable zone of one of the stars, a transfer trajectory to an orbit around the other star and an orbit-raising maneuver to the outer edge of the habitable zone are also investigated and added to the proposed mission scenario. The trajectories are optimized for minimum time of flight with the trajectory optimization tool InTrance, which makes use of evolutionary neurocontrol. The results show that an increase in technological development is required to obtain a feasible time of flight of approximately one century, using a two-sided reflective sail. The optimized time of flight from our Solar System until capture in Alpha Centauri ranges from 20,000 years for current sail technology to less than 80 years for a futuristic graphene-based ultralight sail. The latter allows a maximum injection speed into Alpha Centauri of 5.5\% of the speed of light. The results in this work show an average improvement of 30\% in terms of time of flight for lightness numbers smaller than a graphene-based sail. A sail as proposed by the Breakthrough Starshot project, however, would need over 2000 years to travel from our Solar System until capture about Alpha Centauri B, such that a fly-through mission that arrives within a lifetime is arguably the best option for a near-term mission to Alpha Centauri.