This thesis aims to investigate the potential of a transfer orbit using successive aerocaptures at Saturn and Titan to establish a science orbit around Titan. Titan is an Earth-like moon with a dense atmosphere and organic compounds present. It has many similarities with Earth that are useful to study such as superrotation. Superrotation is when the atmosphere rotates faster than the body it surrounds. In order to study Titan, we need to establish an orbit around it. The Saturn system is distant from Earth, 8.5 Astronomical Units (AU) which makes it difficult to reach from a time and velocity point of view. We propose to use an aerocapture at Saturn to intercept Titan with lower relative velocity in order to perform an aerocapture at Titan. The analysis was performed in primarily MATLAB to simulate the orbits. The results of this showed that we can aerocapture a spacecraft at Saturn and arrive at Titan within roughly 4 to 8 km/s relative velocity regardless of the incoming hyperbolic excess velocity at the Saturn system. This can be improve upon by using intermediate transfer orbits, such as bi-elliptics, to arrive with even lower relative velocities to Titan of as low as 1 km/s. The drag acceleration experienced during the Saturn aerocapture had peak values of between 0.2 and 1.4 g's and acceleration over 50% of the peak is experienced between 6.8 and 8 minutes. This capture method has the potential to make Titan more easily accessible and allow for scientific study of a clear target for improving our understanding of Earth-like processes on other bodies in our solar system. / Master of Science / This thesis aims to investigate the potential of a transfer orbit using successive aerocaptures at Saturn and Titan to establish a science orbit around Titan. Aerocapturing is utilizing the atmosphere of a body to slow down a spacecraft. Titan is an Earth-like moon with a dense atmosphere and organic compounds present. It has many similarities with Earth that are useful to study such as superrotation. Superrotation is when the atmosphere of a body rotates faster than the body it surrounds. In order to study Titan, we need to establish an orbit around it. The Saturn system is distant from Earth, 8.5 Astronomical Units (AU) which makes it difficult to reach from a time and velocity point of view. It takes a large amount of time to get there so we attempt to get there faster by increasing velocity. This means we arrive at the Saturn system with a large amount of velocity that we need to counter-act in order to orbit. We propose to use an aerocapture at Saturn to intercept Titan with lower velocity in order to perform another aerocapture at Titan to slow into an orbit. The analysis was performed in primarily MATLAB to simulate the orbits. The results of this showed that we can aerocapture a spacecraft at Saturn and arrive at Titan within roughly 4 to 8 km/s regardless of the incoming velocity to the Saturn system. This can be improve upon by using intermediate transfer orbits, after capturing at Saturn, to arrive with even lower velocities at Titan of as low as 1 km/s. The drag acceleration experienced during the Saturn aerocapture had peak values of between 0.2 and 1.4 g's and acceleration over 50% of the peak is experienced between 6.8 and 8 minutes. This is relatively gentle for an aerocapture and means the spacecraft likely will not require significant structural support. This capture method has the potential to make Titan more easily accessible and allow for scientific study of a clear target for improving our understanding of Earth-like processes on other bodies in our solar system.

Low-thrust spacecraft trajectory optimization is often a difficult and time-consuming process. One alternative is to instead use a closed-loop, feedback-driven control law, which calculates the control using knowledge of only the current state and target state, and does not require the solution of a nonlinear optimization problem or system of nonlinear equations. Though generally suboptimal, such control laws are attractive because of the ease and speed with which they may be implemented and used to calculate feasible low-thrust maneuvers. This thesis presents the theoretical foundations for seven modern low-thrust control laws based on control law "blending" and Lyapunov control theory for a particle spacecraft operating in an inverse-square gravitational field. The control laws are evaluated critically to determine those that present the best combinations of thoroughness of method and minimization of user input required. The three control laws judged to exhibit the most favorable characteristics are then compared quantitatively through three numerical simulations. The simulations demonstrate the effectiveness of feedback-driven control laws, but also reveal several situations in which the control laws may perform poorly or break down altogether due to either theoretical shortcomings or numerical difficulties. The causes and effects of these issues are explained, and methods of handling them are proposed, implemented, and evaluated. Various opportunities for further work in the area are also described. / text

The architecture of a system which enables the cost-effective exploration of the solar system is proposed. Such a system will make use of the benefits of the natural dynamics represented in the Circular Restricted Three-Body Problem (CRTBP). Additionally, a case study of the first missions which apply the lessons from the CRTBP is examined. The guiding principle of the proposed system is to apply lessons learned from both the Apollo project for deep space exploration and the International Space Station for long term habitation in space as well as modular space vehicle design. From this preliminary system design, a number of missions are outlined. These missions form the basis of an evolvable roadmap to fully develop the infrastructure required for long-term sustained manned exploration of the solar system. This roadmap provides a clear and concise pathway from current exploration capabilities to the current long-term goal of sustained manned exploration of Mars. The primary method employed in designing the staging orbits is the "Single Lunar Swingby", each of the component segment trajectory design processes is explored in detail. Additionally, the method of combining each of these segments together in a larger End-to-End optimizer environment within the General Mission Analysis Tool (GMAT) is introduced, called the Multiple Shooting Method. In particular, a specific Baseline Parking Orbit, or BPO, is chosen and analyzed. This BPO serves as the parking home orbit of any assets not currently in use. A BPO of amplitude (14000, 28000, 6000) kilometers. The BPO has full coverage to both the Earth and the Moon and orbit station-keeping may be conducted at a cost of less than 1 m/s over a 14 year period. This provides a cost-effective platform from which more advanced exploration activities can be based, both robotic and manned. One of the key advanced exploration activities considered is manned exploration of Mars, one of the current long-term goals of NASA. Trajectories from the BPO to Mars and back to Earth are explored and show approximately 50% decrease in required ΔV provided by the spacecraft.

Thesis (Ph. D.)--University of Texas at Austin, 2008. / Vita. Includes bibliographical references.

In this work, capturing Near-Earth Asteroids (NEAs) into Near-Earth orbits is investigated. A general optimization strategy is employed whereby a genetic algorithm is used to seed a sequential quadratic programming (SQP) method for the first step, and then nearby solutions seed further SQP runs. A large number of solutions are produced for several asteroids with varying levels of thrust, and under the effects of various perturbations. Solutions are found over a range of epochs and times of flight as opposed to many traditional methods of optimizing point solutions. This methodology proved effective, finding low-thrust capture solutions within 10% of the required Delta V for analytically estimated transfers, and matching results from other optimization programs such as MALTO. It is found that the effects of solar radiation pressure (SRP) and n-body effects do not have a significant impact on the optimized transfer costs, nor do the perturbations significantly affect the shapes and trends of the optimized solution space. <p><br></p> <p>These optimized results are then used to develop analytic models for both optimized transfer costs and flight times. These models are then used to estimate the transfer costs and flight times for all listed Near Earth Asteroids from the JPL Small Body Database. This analysis is then used to determine the nominal properties of potentially capturable asteroids. The characteristics are then related to a series of different asteroid transfer technologies, elucidating each technology's capabilities and potential capture targets. Finally, this analysis concludes with a brief roadmap of the major decisions mission designers should consider for future asteroid capture missions.</p>

In this dissertation, we develop two novel methods for studying the nature of the Milky Way's dark matter halo. In both cases, we rely on the relationship between the dark matter halo's gravitational potential and the orbital structure it supports. The first method explores the morphology of stellar streams orbiting in non-spherical gravitational potentials. When globular clusters or dwarf galaxies fall into the Milky Way, tidal forces shred them into long filaments of stars called stellar streams. We show that in non-spherical potentials, stream morphologies are heavily dependent on the characteristics of the progenitor's orbit. Flattened axisymmetric galactic potentials, for example, are known to host minor orbit families surrounding special orbits with commensurable frequencies. The behavior of orbits that belong to these orbit families is fundamentally different from that of typical orbits with non-commensurable frequencies. We show that streams evolving near the boundaries, or separatrices, between orbit families, may become fanned out, develop a bifurcation, or both. We utilize perturbation theory to estimate the timescale of this effect and the likelihood of a stream evolving close enough to a separatrix to be affected by it. Next, we study the dynamical reasons for stream fanning and bifurcations near resonances, and find that each morphological outcome has a slightly different dynamical cause. Using a novel numerical approach for measuring the libration frequencies of resonant and near-resonant orbits, we reveal that fans come about due to a large spread in the libration frequencies near a separatrix, whereas bifurcations arise when a separatrix splits the orbital distribution of the stellar stream between two (or more) distinct orbit families. We then demonstrate how these features can arise in streams on realistic galactic orbits, in realistic galactic potentials, over timescales as short as 2-3 Gyr, and discuss how this might be used to constrain the global shape of the Milky Way's gravitational potential. The second method studied in this dissertation enables dynamical tests of a dark matter candidate known as Fuzzy (or Ultra-Light) Dark Matter. Our method relies on a wave generalization of the classic Schwarzschild approach for constructing self-consistent halos -- such a halo consists of a suitable superposition of waves instead of particle orbits, chosen to yield a desired mean density profile. As an illustration, we apply the method to spherically symmetric halos. We derive an analytic relation between the particle distribution function and the wave superposition amplitudes, and show how it simplifies in the high energy (WKB) limit. We verify the stability of such constructed halos by numerically evolving the Schrodinger-Poisson system. The proposed algorithm provides an efficient and accurate way to simulate the time-dependent halo substructures from wave interference, and to test how they will affect dynamical tracers or other observables in a galaxy. The dissertation concludes with a brief discussion of the future prospects of these two methods, especially in the context of upcoming ground- and space-based missions like Rubin LSST and the Roman Space Telescope.

<p>Many opportunities for frequent transit between the lunar vicinity and the heliocentric region will arise in the near future, including servicing missions to space telescopes and proposed missions to various asteroids and other destinations in the solar system. The overarching goal of this investigation is the development a framework for periodic and transit options in the Earth-Moon-Sun system. Rather than overlapping different dynamical models to capture the dynamics of the cislunar and heliocentric region, this analysis leverages a four-body dynamical model, the Bicircular Restricted Four-Body Problem (BCR4BP), that includes the dynamical structures that exist due to the combined influences of the Earth, the Moon, and the Sun. The BCR4BP is an intermediate step in fidelity between the CR3BP and the higher-fidelity ephemeris model. The results demonstrate that dynamical structures from the Earth-Moon-Sun BCR4BP provide valuable information on the flow between cislunar and heliocentric spaces. </p> <p><br></p> <p>Dynamical structures associated with periodic and bounded motion within the BCR4BP are successfully employed to construct transfers between the 9:2 NRHO and locations of interest in heliocentric space. The framework developed in this analysis is effective for transit between any cislunar orbit and the Sun-Earth libration point regions; a current important use case for this capability involves departures from the NRHOs, orbits that possess complex dynamics and near-stable properties. Leveraging this methodology, one-way trajectories from the lunar vicinity to a destination orbit in heliocentric space are constructed, as well as round-trip trajectories that returns to the NRHO after completion of the objectives in heliocentric space. The challenges of such trajectory design include the phasing of the trajectory with respect to the Earth, the Moon, the Sun, on both the outbound and inbound legs of the trajectory. Applications for this trajectory include servicing missions to a space telescope in heliocentric space, where the initial and final locations of the mission is the Gateway near the Moon. Lastly, the results of this analysis demonstrate that the properties and geometry of the periodic orbits, bounded motion, and transfers that are delivered from the BCR4BP are maintained when the trajectories are transitioned to the higher-fidelity ephemeris model. </p>

No description available.

<p>Increasingly, space flight missions are planned to traverse regions of space with complex dynamical environments influenced by multiple gravitational bodies. The nature of these systems produces motion and regions of sensitivity that are, at times, unintuitive,</p> <p>and the accumulation of trajectory dispersions from a variety of sources guarantees that spacecraft will deviate from their pre-planned trajectories in this complex environment, necessitating the use of a targeting process to generate a new feasible reference path. To ensure mission success and a robust path planning process, trajectory designers require insight into the interaction between the targeting process, the baseline trajectory, and the dynamical environment. In this investigation, the convergence behavior of these targeting processes is examined. This work summarizes a framework for characterizing and predicting the convergence behavior of perturbed targeting problems, consisting of a set of constraints, design variables, perturbation variables, and a reference solution within a dynamical system. First, this work identifies the typical features of a convergence basin and identifies a measure of worst-case performance. In the absence of an analytical method, efficient numerical discretization procedures are proposed based on the evaluation of partial derivatives at the reference solution to the perturbed targeting problem. A method is also proposed for approximating the tradespace of position and velocity perturbations that achieve reliable</p> <p>convergence toward the baseline solution. Additionally, evaluated scalar quantities are introduced to serve as predictors of the simulation-measured worst-case convergence behavior based on the local rate of growth in the constraints as well as the local relative change in the targeting-employed partial derivatives with respect to perturbations.</p> <p><br></p> <p>A variety of applications in different dynamical regions and force models are introduced to evaluate the improved discretization techniques and their correlation to the predictive metrics of convergence behavior. Segments of periodic orbits and transfer trajectories from past and planned missions are employed to evaluate the relative convergence performance across sets of candidate solutions. In the circular restricted three-body problem (CRTBP), perturbed targeting problems are formulated along a distant retrograde orbit and a near-rectilinear halo orbit (NRHO) in the Earth-Moon system. To investigate the persistence of results from the CRTBP in an ephemeris force model, a targeting problem applied to an NRHO is analyzed in both force models. Next, an L1 -to-L2 transit trajectory in the Sun-Earth system is studied to explore the effect of moving a maneuver downstream along</p> <p>a trajectory and altering the orientations of the gravitational bodies. Finally, a trans-lunar return trajectory is explored, and the convergence behavior is analyzed as the final maneuver time is varied.</p>

As an alternative solution to global warming, this thesis explores the possibility of aspace-based geoengineering scheme that may prove worthwhile to implement in parallel toother environmental efforts that help mitigate impact of climate change. One suggestionof a geoengineering solution is deploying a large number of sunshades in the vicinity ofthe first Lagrange point of the Sun-Earth system, and this prospective sunshade projectwould serve to shield Earth from incident solar radiation. This thesis is an extension ofa feasibility study for the implementation of this large-scale mission, and has a focus oncomparing electric thrusters to solar sailing as a means of propulsion. Background onelectric propulsion systems and spaceflight mechanics is provided. The investigation wasperformed by defining the spacecraft configurations, and then computing trajectories toa point of escape from Earth and from there to the final equilibrium point.Our results show that in order to meet the propellant demands of the electric thrusters,the launch mass would need to increase by around 15-25 % compared to the solar sailingimplementation, equating to around 1010 kg. Nevertheless, electric propulsion could stillbe a beneficial choice since it would allow shorter transfer times for each shade whichreduces the radiation exposure and subsequent degradation of the spacecraft’s systems.It was found that the transfer time with electric propulsion would be about one-half orone-fifth that of solar sailing, depending on spacecraft parameters. Additionally, electricpropulsion allows a much lower initial parking orbit, and while this would increase the ra-diation exposure it would also reduce the launch costs due to the higher payload capacityto lower altitudes. However, electric propulsion of this scale require prior advancementsin xenon or other inert propellant extraction methods and possibly a wide-scale construc-tion of air separation plants.