<div> <div> <div> <p>As cislunar and outer space exploration regains worldwide popularity, the low-thrust spacecraft technology, whether in the form of solar sails, electric propulsion or nuclear propulsion, has seen a major increase in the last two decades, as new technologies arise that not only seek for a reduction of the size of the spacecraft —and/or the payloads— but also to minimize the cost of spaceflights, while trying to approach further destinations in our solar system. Mission designers are being challenged with the need to develop new strategies to generate rapid and informed initial guesses for low-thrust spacecraft trajectory design, that are easily converged into fully continuous solutions in position, velocity and mass states, in a high-fidelity dynamical model that incorporates the true ephemerides and perturbations of the gravitational attracting bodies acting on the spacecraft as it navigates through space. </p> <p>In an effort to explore further mission options for spacecraft traveling in the lunar vicinity, new interest arises into the problem of constructing a general framework for the initial guess generation of low-thrust trajectories in cislunar space, that is independent of the force models in which the orbits of interest are de ned. Given the efficiency of the low-thrust engines, most vehicles are equipped to perform further exploration of the cislunar space after completion of their primary science and technology demonstrations in orbits around the Moon. In this investigation, a generalized strategy for constructing initial guesses for low-thrust spacecraft traveling between lunar orbits that exist within the context of multiple dynamical models is presented. These trajectories are converged as mass-optimal solutions in lower fidelity model, that are easily transitioned and validated in the higher-fidelity ephemeris model, and, achieve large orbital plane changes while evolving entirely within the cislunar region. </p> <p>The robustness of the initial guess generation of the spacecraft’s path, depends highly on the fidelity of the dynamical model utilized to construct such trajectories, as well as on the numerical techniques employed to converge and propagate them into continuous solutions. Other researchers have extensively investigated novel techniques for the generation of initial guesses for the low-thrust spacecraft trajectory design problem including, but not limited to, patched conics strategies, methodologies for the transformation of impulsive burns into nite burns, the orbit chaining framework and, more recently, artificial intelligence schemes. This investigation develops an adaptive orbit chaining type approach that relies on the energy parametrization of periodic orbits that exist within the context of the circular restricted three-body problem, to construct informed initial guess for the low-thrust spacecraft trajectory.</p> <div> <div> <div> <p>A variety of multiple transfer applications for vehicles traveling between orbits in the cislunar region is explored for a wide range of low-thrust spacecraft with varying thrust acceleration magnitude. The examples presented in this investigation are consistent with the low-thrust parameters of previously own missions that utilized the same propulsion capabilities, such as, the DAWN mission and the Japanese Hayabusa missions 1 and 2. The trajectories presented in this work are optimized for either propellant consumption or time- of-flight in the lower-fidelity model, and later transitioned into a higher-fidelity ephemeris model that includes the gravitational attraction of the Sun, the Earth and the Moon. </p> <p>Two strategies are explored for the transition of trajectories from a lower-fidelity model to the higher-fidelity ephemeris model, both of which are successful in retaining the transfer geometry. The framework presented in this investigation is further applied to the upcoming NASA Lunar IceCube (LIC) mission to explore possible extended mission options once its primary science and technology demonstration objectives are achieved. It is demonstrated in this investigation that the strategies developed and presented in this work are not only applicable to the specific low-thrust vehicles explored, but it is applicable to any spacecraft with any type of propulsion technology. Furthermore, the energy-informed adaptive algorithm is easily transition to generate trajectories in a range of varying dynamical models. </p> </div> </div> </div> </div> </div> </div>

While the interest in the Martian moons increases, the low-thrust propulsion technology is expected to enable novel mission scenarios but is associated with unique trajectory design challenges. Accordingly, the current investigation introduces a multi-phase low-thrust design framework. The trajectory of a potential spacecraft that departs from the Earth vicinity to reach both of the Martian moons, is divided into four phases. To describe the motion of the spacecraft under the influence of gravitational bodies, the two-body problem (2BP) and the Circular-Restricted Three Body Problem (CR3BP) are employed as lower-fidelity models, from which the results are validated in a higher-fidelity ephemeris model. For the computation and optimization of low-thrust trajectories, direct collocation algorithm is introduced. Utilizing the dynamical models and the numerical scheme, the low-thrust trajectory design challenge associated each phase is located and tackled separately. For the heliocentric leg, multiple optimal control problems are formulated between the planets in heliocentric space over different departure and arrival epochs. A contour plot is then generated to illustrate the trade-off between the propellant consumption and the time of flight. For the tour of the Martian moons, the science orbits for both moons are defined. Then, a new algorithm that interfaces the Q-law guidance scheme and direct collocation algorithm is introduced to generate low-thrust transfer trajectories between the science orbits. Finally, an end-to-end trajectory is produced by merging the piece-wise solutions from each phase. The validity of the introduced multi-phase formulation is confirmed by converging the trajectories in a higher-fidelity ephemeris model.<br>

<p dir="ltr"> With NASA's Artemis program and international collaborations focused on building a sustainable infrastructure for human exploration of the Moon, there is a growing demand for lunar exploration and complex spaceflight operations in cislunar space. However, designing efficient transfer trajectories between the Earth and the Moon remains complex and challenging. This investigation focuses on developing a dynamically informed framework for constructing low-energy transfers in the Earth-Moon-Sun Bicircular Restricted Four-body Problem (BCR4BP). Techniques within dynamical systems theory and numerical methods are exploited to construct transfers to various cislunar orbits. The analysis aims to contribute to a deeper understanding of the dynamical structures governing spacecraft motion. It addresses the characteristics of dynamical structures that facilitate the construction of propellant-efficient pathways between the Earth and the Moon, exploring periodic structures and energy properties from the Circular Restricted Three-body Problem (CR3BP) and BCR4BP. The investigation also focuses on constructing families of low-energy transfers by incorporating electric propulsion, i.e., low thrust, in an effort to reduce the time of flight and offer alternative transfer geometries. Additionally, the investigation introduces a process to transition solutions to the higher fidelity ephemeris force model to accurately model spacecraft motion through the Earth-Moon-Sun system. This research provides insights into constructing families of ballistic lunar transfers (BLTs) and cislunar low-energy flight paths (CLEFs), offering a foundation for future mission design and exploration of the Earth-Moon system.</p>