The dynamics of rendezvous of spacecraft in circular orbits is a problem that is well-understood and that is regularly taught in orbital mechanics majors, as are the techniques of linearization and the applications of the state transition matrix. The design of strategies for rendezvous and formation flying often makes use of standard building blocks in the form of specific manoeuvres and trajectories. Typical rendezvous manoeuvres include the Hohmann transfer and the radial hop, a manoeuvre that can change the along-track separation but that does not change the semi-major axis. Typical rendezvous and formation flying trajectories include drift orbits, safe orbits and hold points on V-bar. In this study, the Hohmann transfer, the radial hop and the useful relative trajectories were generalized to elliptical orbits, and this has enabled the application of the insights gained from circular orbit rendezvous and formation flying to elliptical orbits. The insights gained from the theoretical developments have been applied to the mission analysis for Proba-3, a formation flying mission in a highly elliptical orbit.@en

This thesis aims to calculate optimal trajectories from a user-defined Earth-bounded orbit to a user-defined Moon-bounded orbit using a bi-impulse direct transfer ultimately under the influence of a full dynamical model with perturbations, hence reflecting the actual physical environment.

Two tools are developed to achieve this goal. The first tool employs a global optimization algorithm, in particular a Particle Swarm Optimizer (PSO), to find an initial guess within a simplified dynamics model, exploring the user-defined search space. The second tool employs a gradient-based Sequential Linear Least SQuares Programming (SLLSQP) optimizer to refine the initial guess and include the relevant perturbations that act in real life. Additionally, the tools are supported by methods for evaluating the results, providing plotting and analysis tools to make the most out of the obtained solutions.

For the initial guess calculation, the dynamics model includes the point-mass gravity field of Earth and the Moon. The output provides the required ΔV for the transfer and the epochs at which each maneuver should be performed. The SLLSQP optimizer subsequently corrects the initial guess considering the user-specified perturbations, optimizing the time in the first orbit, the different components of both maneuvers, and the time of flight to reach the required orbit in an optimal way.

The capabilities of the tools are demonstrated through several test cases. The first test involves transferring from a circular low Earth orbit (LEO) to a circular near-polar low lunar orbit (LLO), resulting in a total ΔV of 4716.62 m/s. A second and a third test case involving transfers from a LEO or a geostationary transfer orbit (GTO) to an eccentric lunar orbit are also conducted, obtaining a ΔV of 3859.81 m/s when transferring from the LEO and of 1512.95 m/s when doing so from a GTO, corresponding to a decrease of around 60%. The solution obtained from the transfer from the GTO leads to a 4.5% improvement compared to preliminary results found in literature. The forth test comprises transfers from another circular LEO orbit to a high-altitude lunar polar orbit, requiring a ΔV of 3996.44 m/s, being 4.6% higher than the solution found in literature.

These test cases validate the functionality of the code and showcase its versatility in handling various scenarios. In conclusion, the developed tools provide efficient and robust solutions for optimizing direct transfers from Earth to the Moon under the influence of real-life perturbations.

Insertion maneuvers are used to move a spacecraft from an open orbit (parabolic or hyperbolic) into a closed orbit around a target body. These maneuvers are key components in any space mission considering orbiting a body for a large amount of time, for exploration or landing; the hyperbolic orbit will be the one that will be used to transfer between Earth and the target, while the closed orbit will be the one on which the spacecraft will station. In preliminary mission studies, insertion maneuvers are often assumed as being performed at pericenter, and with the two velocity vectors (before and after the maneuver) having the same direction. However, this method does not account for the relative orientation of the two orbits, which are often constrained by separate optimization studies, which may not grant the necessary conditions for a tangential insertion. This study aims to provide a simple method to perform preliminary studies on insertion maneuvers, while ensuring the continuity between the two trajectories, even when those are subject to shape or orientation requirements. The objective is to optimize the insertion maneuver for a crewed mission to Mars, and via this case study gain insight in the best maneuver available (instead of assuming a pericenter, tangential insertion), as well as the best shape and orientation of the trajectories before and after the maneuver.

Recent launches of satellite constellations in the Low Earth Orbit (LEO) region have increased the collision probability of existing debris objects with active satellites. Monitoring the trajectories of these debris objects is crucial for Space Situational Awareness (SSA) to prevent the creation of more debris due to unwanted collisions. Much focus is on the LEO regime, with little awareness of the higher Geostationary orbit (GEO) debris population. To date, the explosion of the Russian Ekran 2 satellite in 1978 as well as the disintegration of the Titan IIIC Trans-stage in 1992, have been recorded. These incidents have increased the number of small-sized debris objects in GEO. More unnoticed fragmentation events have been speculated to have occurred, which pose a significant risk of collisions and damage to all weather and communication satellites in use today. The NASA Debris Office confirms that current ground-based radar or optical sensing methods can only be performed for objects of size 1 m and larger, leaving a gap in the precise orbit determination of sub-meter-sized objects in GEO. Moreover, limited observations and atmospheric losses hinder the quality of orbit determination, thus limiting present ground-based SSA techniques. Attempting to bridge this gap in current space surveillance and tracking methods is the objective of this thesis. It evaluates the feasibility of using space-based sensing methods to enhance SSA in the GEO regime. In this research, a satellite in a sub-GEO orbit is deployed to collect in situ radar measurements, which are processed to determine the orbit of a single object in GEO. Different satellite geometries (altitudes and inclinations) and measurement types such as range, range-rate, and direction (azimuth and elevation angles) and combinations thereof have been analysed. A simple grid search optimisation has been performed to assess the feasibility of such a technique and propose a possible favourable observation configuration, which improves the quality and accuracy of orbit determination. It also analyses the uncertainties in the debris state for future epochs to assess the errors in orbit prediction. The limitations of the geometry and measurement model are identified in this study and provided as recommendations and suggestions for further research. INDIGO is hence a feasibility study or a proof-of-concept of space-based debris state observations in GEO. It can be considered a stepping stone towards inventorying the small-sized GEO debris population catalogue and exploring enhanced SSA techniques in the future.

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.

Apophis, an Aten-type asteroid, was considered a significant threat to Earth upon its discovery due to its potential impact with Earth in 2029. Although the collision has now been ruled out, Apophis will still perform an exceptionally close flyby of Earth that same year at a distance of 32,000 km from the surface. This event presents a unique opportunity for the investigation of rotational changes due to tidal forces and post-encounter ephemerides for planetary defence purposes. In fact, the upcoming OSIRIS-APEX mission will enter orbit around Apophis shortly after the 2029 flyby. Even though orbits around near-Earth asteroids like Apophis face challenges due to the asteroid's low gravity and the strong perturbations induced by solar radiation pressure, frozen orbits - a specialised orbit with constant eccentricity and argument of periapsis on average - can achieve orbital stability in such a complex dynamical environment.

Frozen orbits were successfully employed for some of the mission phases of the OSIRIS-REx mission, and past research has greatly focused on the investigation of frozen orbits around Apophis and other small bodies through the use of analytical and numerical methods. However, no prior research has addressed the design of frozen orbits that can survive the close approach in 2029 without orbital correction manoeuvres. The aim of this research is thus to investigate the stability of control-free frozen orbits around Apophis during the 2029 Earth flyby.

To fulfil this goal, both analytical and numerical methods are employed. The analytical analysis involves averaging of Lagrange's Planetary Equations including perturbations from solar radiation pressure and Apophis' zonal gravity up to degree four in combination with a Lyapunov stability analysis and a comparison to numerical simulations. Assuming an argument of periapsis and longitude of the ascending node of +-90 degrees, the analytical method identifies two main solution families: near-equatorial heliotropic/anti-heliotropic orbits and near-polar Sun-terminator orbits. However, the stability analysis predicts only half of the sampled solutions to be stable. The comparison to numerical simulations shows that both analytical techniques fail to identify stable, frozen orbits. The stability index correctly identifies stability for 66.67% of the results that reach the end of a numerical propagation without surface impact or orbital escape. More significantly, 42% of the results are identified as false positives. The variations in eccentricity and argument of periapsis for the solutions that reach the end of a 28-day simulation are approximately 0.69 and 600 degrees respectively for the
near-equatorial solutions and, at best, 0.89 and 215 degrees for the near-polar solutions, which are too large to be considered frozen orbits.

In the numerical analysis, the frozen orbit problem is defined as a multi-objective optimisation problem with two objectives: minimisation of the maximum variation in eccentricity and argument of periapsis. Trajectories with different orbital injection parameters are simulated to find the optimal initial state leading to a frozen orbit. First, the results are focused exclusively on the pre-flyby period with no constraint on surviving the flyby. The best solutions lead to a maximum variation in eccentricity and argument of periapsis of approximately 0.047 and 66 degrees respectively over a 28-day period. However, these orbits all eventually collide with the asteroid at the time of the flyby. Imposing a constraint on survival increases the maximum variation in eccentricity and argument of periapsis to ranges of 0.08-0.21 and 103-110.5 degrees in the pre-flyby period. The behaviour post-flyby is stable for some of these solutions but no longer corresponds to the frozen configuration. In both cases, the solutions are categorised under the near-circular, near-polar, Sun-terminator frozen orbit family, the same type of orbit employed for the frozen orbit phases of the OSIRIS-REx mission. Despite the limitations of this work, the numerical pre-flyby results exhibit robustness against uncertainties in modelling parameters and orbital injection inaccuracies.

Temporarily Captured Orbiters (TCOs) - also known as Earth’s mini-moons – are meter-size asteroid fragments temporarily trapped in the Earth-Moon system. TCOs are challenging to identify due to their small size and high speed. While only two TCOs have been confirmed so far, studies suggest a constant presence of at least one TCO at any given moment. This research aims to analyze transfer possibilities to these objects using a spacecraft powered by low-thrust electric propulsion departing from hibernating orbits near Sun-Earth L1. The study focuses on the TCO 2006 RH120, but it intends to develop tools that can be used to analyze other TCOs as they are discovered by advanced survey systems under development. A fast and robust optimization algorithm is developed, which successfully analyzes various departure orbits and identifies low delta-v transfers in the order of 200 m/s.

Space debris has become a rising problem in the aerospace community, leading to the need for effective spacecraft collision avoidance processes. Currently, these processes can be called unilateral as only one object in conjunction is considered maneuverable. This thesis focuses on the implementation of a combined action approach to collision avoidance and proposes a cooperation process between operators. The research objective is to optimize the dual maneuver solutions and explore negotiation proposals within a Space Traffic Management system.

The study utilizes an optimization algorithm based on multiple objectives, including propellant mass consumption, collision probability, and mission disturbance. The decision variables used in the optimization are related to the three-direction maneuvering within both objects in conjunction. The optimization is first carried out to minimize the three objectives listed above. These optimization results are considered preliminary as they do not allow for a proper trade-off for the operators. Hence, a review of the objectives used in the optimization algorithm yields the two new criteria used for the final results: the Collision parameter and the Cost parameter. The latter combines propellant mass consumption and mission disturbance.

The results, displayed as Pareto fronts, demonstrate that these objectives allow for the identification of optimal maneuver solutions. Adding on, a sensitivity analysis highlights the importance of precise maneuver timing and lower ΔV contributions within the solutions. The operator is recommended to re-analyze the CAM if the maneuver timing varies by more than 5 minutes. Through the exploration of various study cases and scenarios, insights are provided into the interaction between different systems in space. In general, the chaser showed higher values of ΔV magnitude than the target but the optimization results showed that both interacted together to reach the collision avoidance solution. The I_sp factor proved to not affect the optimization results significantly, and the single maneuvering spacecraft scenarios were successfully solved with the optimization method. This scenario led to higher Cost parameters and higher Collision parameter, the P_c could only be lowered slightly further than 10^-10. As this is below the defined threshold, the results were accepted.

In addition, a proposal is drafted for a communication flow and cooperation framework. The Middle Man, acting as a central authority between the two parties, facilitates the cooperation process, ensuring fair and efficient collaboration between operators. The proposed framework for decision-making is called "rule and resource following shared approach". While specific rules and procedures are not defined in this thesis, the framework allows for them to be included once agreed upon by operators. This thesis concludes that the proposed combined action and cooperation process offers potential solutions to the challenges posed by space debris and contributes to the safety and sustainability of space activities.

While the space industry expands rapidly and space exploration becomes ever more relevant, this thesis aims to automate the design of Multiple Gravity-Assist (MGA) transfers using low-thrust propulsion. In particular, during the preliminary design phase of space missions, the combinatorial complexity of MGA sequencing is large and current optimisation approaches require extensive experience and can take days to simulate. Therefore, a novel optimisation approach is developed -- called the Recursive Target Body Approach -- that uses the hodographic-shaping low-thrust trajectory representation together with a combination of tree-search methods to automate the optimisation of MGA sequences. The approach gradually constructs the expected optimal MGA sequence by recursively evaluating the optimality of subsequent gravity-assist targets. Moreover, the approach includes novel figures of merit as well as parallelisation concepts to increase the robustness and accelerate the convergence. An Earth-Jupiter transfer with a maximum of three gravity assists is considered as a reference problem. Extensive tuning improved the quality of the MGA trajectory representations substantially and as a result, a robust low-thrust trajectory optimisation could be ensured. A distinct group of highly fit MGA sequences is consistently found that can be passed to a higher-fidelity method. In conclusion, the Recursive Target Body Approach can automatically and reliably be used for the preliminary optimisation of low-thrust MGA trajectories.

The current project investigated the feasibility of aerocapture at Jupiter, and the benefits in terms of payload mass fraction that could be achieved compared to traditional orbital insertion burns. A numerical simulation model has been set up, as well as an analytical formulation of the problem. The numerical verification of the analytical model showed that the analytical model still needs to be refined to produce accurate and useful results.
Thermal fluxes, a driving aspect of aerocapture, have been implemented by using correlation laws, as well as corrective terms, all retrieved from literature.
The aerocapture problem was numerically modeled and has been then optimized. However, the best trajectories provided a negative mass fraction benefit of −0.37 when compared to a traditional insertion burn. The best available mass fraction for the spacecraft's entry-unrelated subsystems was 0.44.
Therefore, apart from some niche applications, aerocapture at Jupiter can be considered unappealing at best in the near future.

Space debris fragments smaller than 10 cm cannot be tracked from Earth and are generally neglected in conjunction and risk analyses because of this. However, these fragments pose a great threat, as they can lead to collisions. Currently, the threat that space debris poses on the space environment is getting larger, so methods to mitigate said debris need to be explored. The technique that was studied here, consisted of a passive spacecraft with a circular cross-sectional surface of a material, such as aerogel or foam, with the ability to decelerate and catch the debris fragments encountered. Given this choice, a capture analysis was carried out for such a potential technique. It was concluded, that the most favorable orbital settings for such a method, would be in the case of an explosion of an active or defunct spacecraft, as a direct reaction device. This scenario was then simulated using the orbital parameters of the Kosmos 1408 anti-satellite missile test. The results showed that for a spacecraft with a collector radius of 20 meters, 2 to 3% of the newly created fragments were caught, whereas, for a larger spacecraft with a radius of 100 meters, that number increased to up to 12%. The ideal deployment time for a spacecraft of 20 meters radius was found to be 12 hours after the fragmentation, whereas for 100 meters it was 6 hours. It was found that the capability of such a method is highly dependent on the catcher size, whereas deployment time has a smaller impact. Moreover, it was concluded that the performance of this technique is very sensitive to injection inaccuracies, as the number of fragments caught would be close to zero.

Laser communication provides numerous benefits over typical Radio Frequency communication, such as lower power, not occupying regulated frequency bands, possibilities for much higher data rates and resistance to jamming. Combined with satellite constellations, Laser Inter-satellite Links (LISL) can enable global connectivity. However, satellites move at fast relative velocities, while optical beam divergence angles are in micro-radian levels. Thus, low-latency and precise position data between linking satellites is crucial. This thesis investigates the LISL conditions in a combined LEO/MEO constellation and the applicability of on-board GNSS-based Orbit Determination (OD) and Orbit Prediction (OP). Novel methods, such as Preprocessing Extended and Single-propagation Unscented Kalman Filters are tested and compared to typical GNSS-OD methods. Analyzing Pointing Uncertainty contributions in various link cases, results indicated that fully-kinematic methods could support LISL for 100-s periods, whereas reduced-dynamic OD-OP methods performed more consistently.

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.

CubeSat missions have been deployed to cislunar space and beyond, paving the way for the next significant advancement: a dedicated CubeSat mission to explore a planet near Earth. To contribute to this goal, the German Aerospace Center (DLR) has developed radiation-hardened small satellite technologies, including communications, power and onboard computer subsystems. This study presents a stand-alone Mars exploration mission using the CubeSat standard that will demonstrate DLR’s in-house technologies. This mission will perform an independent transfer to the Red Planet, achieve orbital insertion, and conduct measurements on its lower atmosphere and gravity field. A system concept has been created by integrating the in-house technologies, investigating the necessary Commercial-Off-The-Shelf (COTS) components, and performing the mission analysis to assess the feasibility of the mission. The resulting 12U CubeSat has a 20.8 kg wet mass, 6.3 km/s low-thrust maneuvering capability and can generate up to 90 W of power at Mars. The proof-of-concept mission is planned for a 4-year duration.

Distant Retrograde Orbits (DROs) are special orbits for third bodies in two-body systems. The third body revolves around the secondary – the smaller of the two primaries – in a retrograde way, meaning the direction is opposite to the direction that the primaries revolve around each other. DROs are not close to either of the primaries, making it difficult to model them as perturbed two-body orbits.

There is no analytical solution for the initial conditions of DROs. This thesis presents a novel method of calculating an initial velocity guess which is then fed into a differential corrector that is able to calculate the initial conditions. In contrast to the state-of-the-art, this happens without the method of incremental steps in the initial position, which requires to go through all possible DROs for a specific two-body system first.

For the calculation of DROs, numerical integration is done. Optimal integrator settings are determined, which is in this case an eighth-order Runge-Kutta method (RK8). By setting the tolerance to the lowest possible value, the accuracy requirements are satisfied.

Furthermore, this thesis explores a different method of modeling DROs that makes use of Fourier series and polynomials, which had already been proposed by Hirani in 2006 for a different set of parameters. By exploiting explicit knowledge about the shape of DROs, this approach is made more efficient in terms of accuracy per Fourier/polynomial parameters needed and thus the computation time is enhanced.

The second part of this study addresses the stability of DROs. This is analyzed in order to get an idea of what DROs would be suitable for future missions. For mass ratios of primary and secondary that realistically occur in the Solar System, all DROs that are closer to the secondary than the primary turn out to be stable when disregarding perturbations. Perturbations are modeled as a constant external acceleration with a constant direction, which is only a first step towards modeling the Sun's and other planet's point mass gravity (p.m.g.), the solar radiation pressure (s.r.p.), and other perturbations, as they are usually depending on time and position. With this rough estimate, only the Sun's p.m.g. is identified as a possible source of instability for DROs in the Earth-Moon system, as all other perturbations are too small.

Central configurations provide the only closed-form analytical solutions of the n-body problem. All possible central configurations of three bodies have been extensively studied along with the stability of the associated periodic orbits. Stable cases have been found for the Lagrangian triangle configuration, which we see occurring with the Trojan asteroids. However, the knowledge about four-body central configurations remains limited. An explicit parameterization of a family of kite shaped four-body central configurations has recently been published. The present research investigates the stability of periodic solutions provided by these central configurations. An analytical treatment of linear stability is carried out and the eigenvalues for circular periodic orbits are calculated. This is complemented with a numerical estimation of Floquet multipliers to determine the linear stability of eccentric periodic orbits. While most of the kite configurations are found to be unstable, regions of linearly stable cases are discovered for both circular and eccentric orbits. Further, numerical simulations of the non-linear system are performed as an independent approach to validate the linear stability results. Perfect agreement with the linear analysis is found, suggesting that stable kites may be observed in the universe.

The number of objects in space is increasing over time, and therefore it is desired to find more efficient propagation models in terms of accuracy and computational speed. This thesis project focuses on the propagation of the state and uncertainties of Potential Hazardous Asteroids, to predict potential Earth impacts. A machine learning technique is proposed to reduce the computational expensiveness of current propagation techniques. The PHA position and the error made by the uncertainties are predicted using a neural network, which uses data from currently known PHAs. Several options are considered regarding the input and output variables and the networks are also tuned. The aim for the best model is to find a reasonable accuracy for the prediction of the position and the uncertainty of newly discovered PHAs.

Quasi-satellite orbits (QSOs) are orbits around the smaller primary in three-body systems. The MMX mission (developed by JAXA, to be launched in 2024) will use QSOs to orbit Phobos, before landing on its surface. This work studies the feasibility of using the invariant manifolds of three-dimensional QSOs for designing landing trajectories, using the Mars-Phobos system as a case study.

Applying continuation and bifurcation-analysis techniques, different families of QSOs were computed and their invariant manifolds propagated. Using a set of favorable manifolds as a reference, multiple maneuvers were designed via multiple-shooting and optimization techniques, producing feasible landing trajectories. The robustness of these trajectories was analyzed using a stability index and Monte Carlo simulations.

It was found that using the invariant manifolds of QSOs to land is only possible for some families of QSOs. However, for these, the manifolds allow generating robust and propellant-efficient landing trajectories that are able to reach most of Phobos' surface.

This thesis focused on the relevance and feasibility of an uncontrolled cubesat mission in a binary asteroid environment, highlighting the example of the representative Near-Earth Asteroid 1999 KW4, modelled using the TU Delft Astrodynamics Tool (TUDAT). A clear distinction was made between orbits of the cubesat starting in the interior and the exterior ring of the binary for numerical computations. Lifetimes over 100 days were found in the interior ring through a family of so-called pseudo-periodic orbits around the primary asteroid, with some that withstand the introduction of various uncertainties in the model. Reasonable lifetimes in the exterior ring were found to be around 60 days in the exterior ring, but could no longer be guaranteed after the sensitivity analysis. This suggests that targeting pseudo-periodic orbits is the best strategy to adopt. Velocity uncertainties in the initial state of the cubesat were observed to be the driving aspect in both the interior and exterior ring, as managing to lower them could shift a seemingly too risky run into an acceptable one w.r.t. our criteria.

Near-Earth Asteroids (NEAs) have attracted the attention of the scientific community in the last few years. Not only because of their importance to life on Earth but also their scientific potential and possible economic returns. This work explores the use of quasi-periodic orbits to bound the motion of NEAs close to the Earth’s vicinity for their exploitation. The invariant manifolds emanating from these quasi-periodic tori are used to design NEA high-thrust, low-energy retrieval trajectories. A thorough characterization of the two-dimensional space in which the invariant tori can exist is conducted. Three promising NEAs are selected, from which only two of them (2006 RH120 and 2020 CD3) permit transfers at extremely low ΔVs. For 2006 RH120, transfers that require between 20 and 2 times less ΔV than the existing results from literature were found. We prove that quasi-periodic orbits allow for better transfers than just considering manifolds from periodic families or not using manifolds at all. The use of quasi-periodic tori also permits extended transfer windows and more flexibility in the mission design.