<p dir="ltr">At the cusp of scientific discovery and innovation, mankind's next greatest challenge lies in developing capabilities to enable human presence in deep space. This entails setting up space infrastructure, travel pathways, managing spacecraft traffic, and building up deep space operation logistics. Spacecrafts that are a part of the infrastructure must be able to perform myriad of operations and transfers such as rendezvous and docking, station-keeping, loitering, collision avoidance etc. In support of this endeavour, an investigation is done to analyze and recreate the solution space for fuel-optimal trajectories and control histories required for onboard trajectory design of inexpensive spacecraft transfers and operations. This study investigates close range rendezvous (CRR), nearby orbital transfer, collision avoidance, and long range transfer maneuvers for spacecrafts whose highly complex and nonlinear behavior is modelled using the circular restricted three-body problem (CR3BP) dynamics and to which a finite-burn maneuver is augmented to model low-propulsion maneuvers. In order to study the nonlinear solution space for such maneuvers, this investigation contributes new formulations of nonlinear programming (NLP) optimal control problems solved to minimize fuel consumption, and validated by traditional methods already in use. This investigation proposes a Feature Learning based Optimal Control Method (L-OCM) to learn the solution space and recreate results in real-time. The NLP problem is solved off-line for a range of initial conditions. The set of solutions is used to generate datasets with initial conditions as inputs and the identified features of the optimal control solution as outputs. These features are inherent to reconstructing the optimal control histories of the solution and are selected keeping onboard computational capabilities in mind. Deep Neural Networks (DNNs) are trained to map the complex, nonlinear relationship between the inputs and outputs, and then implemented to find on-line solutions to any initial condition. The L-OCM method provides fuel-optimal, real-time solutions that can be implemented by a spacecraft performing operations in cislunar space.</p>

As resident space object populations grow, and satellite propulsion capabilities improve, it will become increasingly challenging for space-reliant nations to maintain space situational awareness using current human-in-the-loop methods. This dissertation develops several real-time adaptive approaches to autonomous sensor network management for tracking multiple maneuvering and non-maneuvering satellites with a diversely populated Space Object Surveillance and Identification network. The proposed methods integrate suboptimal Partially Observed Markov Decision Processes (POMDPs) with covariance inflation or multiple model adaptive estimation techniques to task sensors and maintain viable orbit estimates for all targets. The POMDPs developed in this dissertation use information-based and system-based metrics to determine the rewards and costs associated with tasking a specific sensor to track a particular satellite. Like in real-world situations, the population of target satellites vastly outnumbers the available set of sensors. Robust and adaptable tasking algorithms are needed in this scenario to determine how and when sensors should be tasked. The strategies developed in this dissertation successfully track 207 non-maneuvering and maneuvering spacecraft using only 24 ground and space-based sensors. The results show that multiple model adaptive estimation coupled with a multi-metric, suboptimal POMDP can effectively and efficiently task a diverse network of sensors to track multiple maneuvering spacecraft, while simultaneously monitoring a large number of non-maneuvering objects. Overall, this dissertation demonstrates the potential for autonomous and adaptable sensor network command and control for real-world space situational awareness. / Ph. D. / As the number of spacecraft in orbit increase, and satellite propulsion capabilities improve, it will become increasingly difficult for space-reliant nations to keep track of every object orbiting earth using human-in-the-loop methods. Already, the population of target satellites vastly outnumbers the available set of sensors. At any given time, a given network of sensors cannot observe every satellite in orbit, and must manage the available sensors effectively to keep track of every object of interest. The ability to maintain actionable knowledge of every orbiting object of interest is known as space situational awareness. Conventional tracking processes have generally not changed for decades, and were designed when there were far fewer satellites in orbit with little or no ability to maneuver. These methods involve large numbers of operators and engineers who schedule a network of sensors under the assumption that the satellites will not unexpectedly change their orbits for long periods of time. In the near future, traditional space surveillance approaches will become insufficient at maintaining space situational awareness, particularly if more satellites conduct unanticipated maneuvers. This dissertation develops several real-time approaches for controlling a diverse network of ground and space-based sensors that remove the need for human intervention. These fully computer-based command and control processes adapt to dynamic situations and automatically task sensors to rapidly track multiple maneuvering and non-maneuvering satellites. The decision processes used to determine which sensors should be tasked to observe a particular spacecraft compare the amount of information that can be collected in a single observation and the workload a sensor must execute to collect the observation. The command and control strategies developed in this dissertation successfully track 207 non-maneuvering and maneuvering spacecraft using only 24 ground and space-based sensors. The results show that adaptive, fully autonomous sensor network control processes can effectively and efficiently task a diverse set of sensors to track multiple maneuvering spacecraft, while simultaneously monitoring a large number of non-maneuvering objects. Overall, this dissertation demonstrates the potential for adaptive, computer-based sensor network command and control for real-world space situational awareness. This research was supported by the Virginia Tech New Horizons Graduate Scholar Program, the Ted and Karyn Hume Center for National Security and Technology, the DARPA Hallmark program, and the U.S. Joint Warfare Analysis Center.

This thesis serves to provide a conceptual mission design for a Titan-orbiting stellar occultation mission. Titan has a significant atmosphere much like Earth's. An improved understanding of Titan's atmosphere could provide valuable information about the evolution of Earth's climate. Titan's atmosphere is known to be in a state of superrotation, wherein the atmosphere rotates significantly faster than the surface beneath. The details of the creation and sustainment of this extreme state on Titan in terms of angular momentum exchange remain unknown despite current theories and models. These unknowns, alongside inconsistencies between current models with observations from the Cassini mission, call for an urgent need for Titan atmospheric observations able to resolve atmospheric waves. The science objectives driving the mission design include maximizing the number of measurements, the latitude versus longitude coverage, the latitude versus local solar time coverage, and the mission duration. These measurement needs can be met by a Titan orbiter utilizing a refractive stellar occultation technique. Refractive stellar occultation observes starlight bending through an atmosphere as stars set behind a body. The observed bending profile can be inverted to infer density, temperature, and pressure profiles. This research uses Systems Tool Kit (STK) as a simulation tool to predict measurement coverage for various orbits. The orbital radius was determined to be the driving independent variable which set all other design variables, including the orbital plane which was uniquely selected for a given orbital radius to maximize the number of occultations. The results of this study show that a lower orbital radius is desired as this produces the best combination of measurement number and distribution. This orbital plane should be closely aligned with the Milky Way galactic plane to see the most stars occult. For the lowest sustainable orbital altitude, Low Titan Orbit (LTO) at 1200 km, the orbital plane should be nearly polar to maximize the number of occultations and latitude coverage. The optimal orbit selection (defined by orbital elements a = 3775 km, e = 0, i = 85 degrees, Ω = 87 degrees, ω = 0 degrees, and ν = 0 degrees) for a single satellite can produce nearly 400 stellar occultation opportunities per orbit and provide full latitude versus longitude coverage. A single satellite shows gaps in latitude versus local solar time coverage at mid-latitudes normal to the satellite ground track which may inhibit the diagnosis of the angular momentum flux associated with thermal tides. If necessary, a second satellite in an orbit orthogonal to the first is suggested to close coverage gaps to provide full local time coverage over a Titan day. The optimal orbit selection of this second satellite (defined by orbital elements a = 3775 km, e = 0, i = 5.3 degrees, Ω = 5.9 degrees, ω = 0 degrees, and ν = 0 degrees) provides an additional 343 occultation opportunities per orbit and increases latitude versus local solar time coverage by a factor of 1.5. The understanding of Titan's Earth-like atmosphere could provide insight into climate evolution here on Earth. This concept proposes a novel approach to improving this understanding. / Master of Science / This thesis serves to provide a conceptual mission design for a Titan-orbiting stellar occultation mission. Titan, one of Saturn's 82 moons, has a significant atmosphere much like Earth's. An improved understanding of Titan's atmosphere could provide valuable information about the evolution of Earth's climate. Titan's atmosphere is known to be in a state of superrotation, wherein the atmosphere rotates significantly faster than the surface beneath. The details of the creation and sustainment of this extreme state on Titan remain unknown despite current theories and models. These unknowns, alongside inconsistencies between current models with observations from the Cassini mission, call for an urgent need for Titan atmospheric observation. The science objectives driving the mission design include maximizing the number of measurements, the latitude versus longitude coverage, the latitude versus local solar time coverage (on a 24-hour scale), and the mission duration. These measurement needs can be met by a Titan orbiter utilizing a refractive stellar occultation technique. Refractive stellar occultation observes starlight bending through an atmosphere as stars set behind a body. The observed bending profile can be inverted to infer density, temperature, and pressure profiles. This research uses a simulation tool to predict measurement coverage for various orbits. The radius of the orbit was determined to be the driving independent variable which set all other design variables, including the orbital plane which was uniquely selected for a given orbital radius to maximize the number of occultations. The results of this study show that a lower orbital radius is desired as this produces the best combination of measurement number and distribution. This orbital plane should be closely aligned with the Milky Way galactic plane to see the most stars occult. For the lowest sustainable orbital altitude, Low Titan Orbit (LTO) at 1200 km, the orbital plane should be nearly polar to maximize the number of occultations and latitude coverage. The optimal orbit selection for a single satellite can produce nearly 400 stellar occultation opportunities per orbit and provide full latitude versus longitude coverage. A single satellite shows gaps in latitude versus local solar time coverage at mid-latitudes normal to the satellite ground track which may inhibit the diagnosis of atmospheric waves tied to Titan's night and day cycle. If necessary, a second satellite in an orbit orthogonal to the first is suggested to close coverage gaps to provide full local time coverage over a Titan day. The optimal orbit selection of this second satellite provides an additional 343 occultation opportunities per orbit and increases latitude versus local solar time coverage by a factor of 1.5. The understanding of Titan's Earth-like atmosphere could provide insight into climate evolution here on Earth. This concept proposes a novel approach to improving this understanding.

Several targeting algorithms are developed and analyzed for possible future use onboard a spacecraft. Each targeter is designed to determine the appropriate propulsive burn for translunar injection to obtain desired orbital parameters upon arrival at the moon. Primary design objectives are to minimize the computational requirements for each algorithm but also to ensure reasonable accuracy, so that the algorithm’s errors do not force the craft to conduct large mid-course corrections. Several levels of accuracy for dynamical models are explored, the convergence range and speed of each algorithm are compared, and the possible benefits of the Broyden and trust-region targeters are evaluated. These targeters provide a proof of concept for the feasibility of a translunar injection targeting algorithm. Anticipating some future improvements, these algorithms could serve as a viable alternative to uploading ground-based targeting solutions and bypass the problems of delays and disruptions in communication, enabling the craft to conduct a translunar injection burn autonomously.

Space debris growth and an influx in space traffic will create a need for increased space traffic management. Due to orbital population density and likely future growth, the implementation of a slot architecture to Sun-synchronous orbit is considered in order to mitigate conjunctions among active satellites. This paper furthers work done in Sun-synchronous orbit slot architecture design and focuses on two main aspects. First, an in-depth relative motion analysis of satellites with respect to their assigned slots is presented. Then, a method for developing a slot architecture from a specific set of user defined inputs is derived.

The goal of this thesis is to use the Cal Poly Spacecraft Attitude Dynamics Simulator to provide proof of concept of two adaptive control theories developed by former Cal Poly students: Nonlinear Direct Model Reference Adaptive Control and Adaptive Output Feedback Control. The Spacecraft Attitude Dynamics Simulator is a student-built air bearing spacecraft simulator controlled by four reaction wheels in a pyramidal arrangement. Tests were performed to determine the effectiveness of the two adaptive control theories under nominal operating conditions, a “plug-and-play” spacecraft scenario, and under simulated actuator damage. Proof of concept of the adaptive control theories applied to attitude control of a spacecraft is provided. The adaptive control theories are shown to attain similar or improved performance over a Full State Feedback controller. However, the measurement capabilities of the simulator need to be improved before strong comparisons between the adaptive controllers and Full State Feedback can be achieved.

Challenging science requirements and complex space missions are driving the need for fast and robust space trajectory design and simulation tools. The main aim of this thesis is to develop new and improved high performance algorithms and solution techniques for commonly encountered problems in astrodynamics. Five major problems are considered and their state-of-the art algorithms are systematically improved. Theoretical and methodological improvements are combined with modern computational techniques, resulting in increased algorithm robustness and faster runtime performance. The five selected problems are 1) Multiple revolution Lambert problem, 2) High-fidelity geopotential (gravity field) computation, 3) Ephemeris computation, 4) Fast and accurate sensitivity computation, and 5) High-fidelity multiple spacecraft simulation. The work being presented enjoys applications in a variety of fields like preliminary mission design, high-fidelity trajectory simulation, orbit estimation and numerical optimization. Other fields like space and environmental science to chemical and electrical engineering also stand to benefit.

<pre>Because of the increasing population of space objects, there is an increasing necessity to monitor and predict the status of the near-Earth space environment, especially of critical regions like geosynchronous Earth orbit (GEO) and low Earth orbit (LEO) regions, for a sustainable future. Space Situational Awareness (SSA), however, is a challenging task because of the requirement for dynamically insightful fast orbit propagation models, presence of dynamical uncertainties, and limitations in sensor resources. Since initial parameters are often not known exactly and since many SSA applications require long-term orbit propagation, long-term effects of the initial uncertainties on orbital evolution are examined in this work. To get a long-term perspective in a fast and efficient manner, this work uses analytical propagation techniques. Existing analytical theories for orbital perturbations are investigated, and modifications are made to them to improve accuracy. While conservative perturbation forces are often studied, of particular interest here is the orbital perturbation due to non-conservative forces. Using the previous findings and the developments in this thesis, two SSA applications are investigated in this work. In the first SSA application, a sensor tasking algorithm is designed for the detection of new classes of GEO space objects. In the second application, the categorization of near-GEO objects is carried out by combining knowledge of orbit dynamics with machine learning techniques.</pre>

<p>A dynamical understanding of orbits in the Earth-Moon neighborhood that can sustain long-term activities and trajectories that link locations of interest forms a critical foundation for the creation of infrastructure to support a lasting presence in this region of space. In response, this investigation aims to identify and exploit fundamental dynamical motion in the vicinity of a candidate ‘hub’ orbit, the L2 southern 9:2 lunar synodic resonant near rectilinear halo orbit (NRHO), while incorporating realistic mission constraints. The strategies developed in this investigation are, however, not restricted to this particular orbit but are, in fact, applicable to a wide variety of stable and nearly-stable cislunar orbits. Since stable and nearly-stable orbits that may lack useful manifold structures are of interest for long-term activities in cislunar space due to low orbit maintenance costs, strategies to alternatively initiate transfer design into and out of these orbits are necessary. Additionally, it is crucial to understand the complex behaviors in the neighborhood of any candidate hub orbit. In this investigation, a bifurcation analysis is used to identify periodic orbit families in close proximity to the hub orbit that may possess members with favorable stability properties, i.e., unstable orbits. Stability properties are quantified using a metric defined as the stability index. Broucke stability diagrams, a tool in which the eigenvalues of the monodromy matrix are recast into two simple parameters, are used to identify bifurcations along orbit families. Continuation algorithms, in combination with a differential corrections scheme, are used to compute new families of periodic orbits originating at bifurcations. These families possess unstable members with associated invariant manifolds that are indeed useful for trajectory design. Members of the families nearby the L2 NRHOs are demonstrated to persist in a higher-fidelity ephemeris model. </p><p><br></p> <p>Transfers based on the identified nearby dynamical structures and their associated manifolds are designed. To formulate initial guesses for transfer trajectories, a Poincaré mapping technique is used. Various sample trajectory designs are produced in this investigation to demonstrate the wide applicability of the design methodology. Initially, designs are based in the circular restricted three-body problem, however, geometries are demonstrated to persist in a higher-fidelity ephemeris model, as well. A strategy to avoid Earth and Moon eclipse conditions along many-revolution quasi-periodic ephemeris orbits and transfer trajectories is proposed in response to upcoming mission needs. Lunar synodic resonance, in combination with careful epoch selection, produces a simple eclipse-avoidance technique. Additionally, an integral-type eclipse avoidance path constraint is derived and incorporated into a differential corrections scheme as well. Finally, transfer trajectories in the circular restricted three-body problem and higher-fidelity ephemeris model are optimized and the geometry is shown to persist.</p>

New Horizons, Voyager 1 & 2, and Pioneer 10 & 11 are the only spacecraft to ever venture past Pluto and provide information about space at those large distances. These spacecraft were very expensive and primarily designed to study planets during gravitational assist maneuvers. They were not designed to explore space past Pluto and their study of this environment is at best a secondary mission. These spacecraft rely on radioisotope thermoelectric generators (RTGs) to provide power, an expensive yet necessary approach to generating sufficient power. With Cubesats graduating to interplanetary capabilities, such as the Mars-bound MarCO spacecraft, matching the modest payload requirements to study the outer Solar System (OSS) with the capabilities of low-power nano-satellites may enable much more affordable access to deep space. This paper explores a design concept for a low-cost, small spacecraft, designed to study the OSS and satisfy mission requirements with solar power. The general spacecraft design incorporates a parabolic reflector that acts as both a solar concentrator and a high gain antenna. This paper explores a working design concept for a small spacecraft to operate up to 100 astronomical units (AU) from the sun. Deployable reflector designs, thermal and radiation environments, communications and power requirements, solar system escape trajectory options, and scientific payload requirements are detailed, and a working system is proposed that can fulfill mission requirements with expected near-future innovations in a few key technologies.