Renewed interest in returning to the Moon, emboldened by recent directives and missions by NASA, has necessitated the establishment of lunar infrastructure to support continuous human presence. With that, the objective of making this return more cost effective has gained significant importance. Low energy lunar transfers are more efficient ways to reach the Moon than the traditional Hohmann-type transfer. These trajectories leverage the multi-body gravitational effects to reduce overall delta-v requirements, in some cases removing the capture delta-v completely. While the time of flight for these transfers can be much longer than a Hohmann-type transfer, the chaotic design space of these transfers can enable large changes in arrival conditions at the Moon for small changes in initial conditions. Many investigations of these transfers take place in simplified dynamical models, such as the Planar Circular Restricted Three Body Problem, with very few higher-fidelity models being implemented. This approach is good to understand the dynamics of these trajectories as well as provide initial guesses for higher-fidelity models; but approximating the dynamics heavily make these models less applicable to mission design. This thesis aims to investigate the application of a higher-order model to simulate these trajectories. STK Astrogator was used to recreate the NASA GRAIL trajectory; and from the recreated trajectory, a nominal trajectory absent of mid-course corrections was established. This nominal trajectory was used to perform parametric and variational studies of departure and arrival conditions as well as compare to a nominal trajectory in a reduced-fidelity model. An investigation into the post launch correction burn requirements following launch vehicle under-performance was completed. Utilizing low energy transfers proved beneficial to adjusting arrival conditions for low delta-v requirements. All arrival inclinations are reasonably achievable for around 255 m/s. Using 255 m/s as a baseline, right ascension of the ascending node could be reached in a 40 degree range and argument of periapsis in a 50 degree range. Lunar insertion arrival can be varied by 7 hours on either side for less than 80 m/s. Trans-lunar injection epoch can be varied by 7 hours on either side of nominal departure for less than 4 m/s. Orbit radius and initial velocity are the most expensive errors to correct. These trajectories can be tuned to reduce the overall mid-course correction delta-v requirement for differing arrival inclinations if other orbital elements are relaxed. A relationship between placement of post-launch correction maneuver for velocity or radius errors was found. Comparing the trajectory in STK to the Inclined Bi-Elliptic Restricted Four Body problem, revealed that timing of the trajectory is variable while keeping the same arrival and departure conditions. However, solar radiation pressure cannot be ignored for more accurate simulation of these trajectories. This investigation has shown that low energy lunar transfers are a viable method to reach the Moon and their chaotic nature can be leveraged to relax restrictions in the design space. / Master of Science / Returning to the Moon has become a more important goal within the space industry. This has required more cost-efficient ways to reach the Moon; an important cost savings being fuel. Traditional ways to reach the Moon required large amounts of fuel to be expended to remain around the Moon after launch. Low energy lunar transfers aim to reduce fuel usage while still reaching the Moon, although they take much longer to reach their destination. Fuel and energy have direct comparisons and are used to evaluate these transfers. These transfers are highly susceptible to changes in their trajectory making them ideal for transferring to the Moon in different orientations. These changes can be made using very little fuel, allowing for more resources to be brought to the Moon. Navigating these transfers to the Moon require an accurate model of space for mission design.

<p>Secondary payloads have a rich and successful history of utilizing cheap rides to orbit to perform outstanding missions in Earth orbit, and more recently, in cislunar space and beyond. New launch vehicles, namely the Space Launch System (SLS), are increasing the science opportunity for rideshare class missions by providing regular service to the lunar vicinity. However, trajectory design in a multi-body regime brings a host of novel challenges, further exacerbated by constraints generated from the primary payload’s mission. Often, secondary payloads do not possess the fuel required to directly insert into lunar orbit and must instead perform a lunar flyby, traverse the Earth-Moon-Sun system, and later return to the lunar vicinity. This investigation develops a novel framework to construct low-cost, end-to-end lunar transfer trajectories for secondary payload missions. The proposed threephase approach provides unique insights into potential lunar transfer geometries. The phases consist of an arc from launch to initial perilune, an exterior transfer arc, and a lunar approach arc. The space of feasible transfers within each phase is determined through low-dimension grid searches and informed filtering techniques, while the problem of recombining the phases through differential corrections is kept tractable by reducing the dimensionality at each phase transition boundary. A sample mission demonstrates the trajectory design approach and example solutions are generated and discussed. Finally, alternate strategies are developed to both augment the analysis and for scenarios where the proposed three-phase technique does not deliver adequate solutions. The trajectory design methods described in this document are applicable to many upcoming secondary payload missions headed to lunar orbit, including spacecraft with only low-thrust, only high-thrust, or a combination of both. </p>