This thesis presents a method for managing periods of communication loss between multiple spacecraft in formation flying (MSFF), and analyzes the effects of this method on the stability of the formation keeping control algorithm. The controller of interest in this work in an adaptive nonlinear controller, where synchronization is also incorporated to force the position tracking errors to converge to zero at the same rate. The communication loss compensation technique proposed in this thesis is to use the previously communicated data in lieu of the lost data, which is an effective and computationally-efficient technique that is advantageous for small satellites. The performance parameter of interest in this research is the maximum rate of communication loss that an MSFF system can withstand before going unstable, and this is analyzed theoretically and through simulations. Finally, experiments involving multiple robots in formation with communication loss are conducted, and the results are presented.

This thesis is concerned with input-output stability theory. Within this framework, it is of interest how inputs map to outputs through an operator that represents a system to be controlled or the controller itself. The Small Gain, Passivity, and Conic Sector Stability Theorems can be used to assess the stability of a negative feedback interconnection involving two systems that each have specific input-output properties. Our first contribution concerns characterization of the input-output properties of linear time-varying (LTV) systems. We present various theorems that ensure that a LTV system has finite gain, is passive, or is conic. We also consider the stability of various negative feedback interconnections. Motivated by the robust nature of passivity-based control, we consider how to overcome passivity violations. This investigation leads to the hybrid conic systems framework whereby systems are described in terms of multiple conic bounds over different operating ranges. A special case relevant to systems that experience a passivity violation is the hybrid passive/finite gain framework. Sufficient conditions are derived that ensure the negative feedback interconnection of two hybrid conic systems is stable. The input-output properties of gain-scheduled systems are also investigated. We show that a gain-scheduled system composed of conic subsystems has conic bounds as well. Using the conic bounds of the subsystems along with the scheduling signal properties, the overall conic bounds of the gain-scheduled system can be calculated. We also show that when hybrid very strictly passive/finite gain (VSP/finite gain) subsystems are gain-scheduled, the overall map is also hybrid VSP/finite gain. Being concerned with the control of aerospace systems, we use the theory developed in this thesis to control two interesting plants. We consider passivity-based control of a spacecraft endowed with magnetic torque rods and reaction wheels. In particular, we synthesize a LTV input strictly passive controller. Using hybrid theory we control single- and two-link flexible manipulators. We present two controller synthesis schemes, each of which employs numerical optimization techniques and attempts to have the hybrid VSP/finite gain controllers mimic a H2 controller. One of our synthesis methods uses the Generalized Kalman-Yakubovich-Popov Lemma, thus realizing a convex optimization problem.

This thesis is concerned with input-output stability theory. Within this framework, it is of interest how inputs map to outputs through an operator that represents a system to be controlled or the controller itself. The Small Gain, Passivity, and Conic Sector Stability Theorems can be used to assess the stability of a negative feedback interconnection involving two systems that each have specific input-output properties. Our first contribution concerns characterization of the input-output properties of linear time-varying (LTV) systems. We present various theorems that ensure that a LTV system has finite gain, is passive, or is conic. We also consider the stability of various negative feedback interconnections. Motivated by the robust nature of passivity-based control, we consider how to overcome passivity violations. This investigation leads to the hybrid conic systems framework whereby systems are described in terms of multiple conic bounds over different operating ranges. A special case relevant to systems that experience a passivity violation is the hybrid passive/finite gain framework. Sufficient conditions are derived that ensure the negative feedback interconnection of two hybrid conic systems is stable. The input-output properties of gain-scheduled systems are also investigated. We show that a gain-scheduled system composed of conic subsystems has conic bounds as well. Using the conic bounds of the subsystems along with the scheduling signal properties, the overall conic bounds of the gain-scheduled system can be calculated. We also show that when hybrid very strictly passive/finite gain (VSP/finite gain) subsystems are gain-scheduled, the overall map is also hybrid VSP/finite gain. Being concerned with the control of aerospace systems, we use the theory developed in this thesis to control two interesting plants. We consider passivity-based control of a spacecraft endowed with magnetic torque rods and reaction wheels. In particular, we synthesize a LTV input strictly passive controller. Using hybrid theory we control single- and two-link flexible manipulators. We present two controller synthesis schemes, each of which employs numerical optimization techniques and attempts to have the hybrid VSP/finite gain controllers mimic a H2 controller. One of our synthesis methods uses the Generalized Kalman-Yakubovich-Popov Lemma, thus realizing a convex optimization problem.

Spacecraft on interplanetary missions or advanced satellites orbiting the Earth all require propulsion systems to complete their missions. Introducing microelectromechanical systems technology to the space industry will not only reduce size and weight of the propulsion system, but can also increase the performance of the mission. Fluid handling systems are used in chemical and electric propulsion. Some components incorporated in a fluidic handling system are presented and evaluated in this work. Microsystems are very sensitive to contamination. Reliable, robust, and easily integrated filters were modeled, manufactured, and experimentally verified. A fluid connector, designed to withstand large temperature variations and aggressive propellants was manufactured and characterized. Similar designs was also be used as a thermally activated minute valve. The feasibility of a cold gas system for precise attitude control has been demonstrated. Steps towards improving the performance (from specific im-pulse 45 s) have been taken, by the integration of suspended heater elements. For electric propulsion, two thermally regulated flow restrictors have been characterized. These devices can fine-tune the propellant flow to e.g. an ion engine. A single-use valve using a soldered seal has also been successfully dem-onstrated within a pressure range of 5 to 100 bar. The microsystem-based propulsion systems of tomorrow’s spacecraft need to be demonstrated in space, in order to gain necessary credibility.

Observing the relative state of two space vehicles has been an active field of research since the earliest attempts at space rendezvous and docking during the 1960's. Several techniques have successfully been employed by several space agencies and the importance of these systems has been repeatedly demonstrated during the on-orbit assembly and continuous re-supply of the International Space Station. More recent efforts are focused on technologies that can enable fully automated navigation and control of space vehicles. Technologies which have previously been investigated or are actively researched include Video Guidance Systems (VGS), Light Detection and Ranging (LIDAR), RADAR, Differential GPS (DGPS) and Visual Navigation Systems. The proposed system leverages the theoretical foundation which has been advanced in the development of VisNav, invented at Texas A & M University, and the miniaturized commercially available Northstar sensor from Evolution Robotics. The dissertation first surveys contemporary technology, followed by an analytical investigation of the coarse sun sensor and errors associated with utilizing it in the near-field. Next, the commercial Northstar sensor is investigated, utilizing fundamentals to generate a theoretical model of its behavior, followed by the development of an experiment for the purpose of investigating and characterizing the sensor's performance. Experimental results are then presented and compared with a numerical simulation of a single-sensor system performance. A case study evaluating a two sensor implementation is presented evaluating the proposed system's performance in a multisensor configuration. The initial theoretical analysis relied on use of the cosine model, which proved inadequate in fully capturing the response of the coarse sun sensor. Fresenel effects were identified as a significant source of unmodeled sensor behavior and subsequently incorporated into the model. Additionally, near-field effects were studied and modeled. The near-field effects of significance include: unequal incidence angle, unequal incidence power, and non-uniform radiated power. It was found that the sensor displayed inherent instabilities in the 0.3 degree range. However, it was also shown that the sensor could be calibrated to this level. Methods for accomplishing calibration of the sensor in the near-field were introduced and feasibility of achieving better than 1 cm and 1 degree relative position and attitude accuracy in close proximity, even on a small satellite platform, was determined.

This thesis considers the Perspective-N-Point (PNP) problem with orthogonal target geometry, as seen in the problem of cubesat relative navigation. Cubesats are small spacecraft often developed for research purposes and to perform missions in space at low cost. Sensor systems for cubesats have been designed that, by providing vector (equivalently line-of-sight, angle, and image plane) measurements, equate relative navigation to a PNP problem. Much study has been done on this problem, but little of it has considered the case where target geometry is known in advance, as is the case with cooperating cubesats. A typical constraint for cubesats, as well as other PNP applications, is processing resources. Therefore, we considered the ability to reduce processing burden of the PNP solution by taking advantage of the known target geometry. We did this by considering a specific P3P solver and a specific point-cloud correspondence (PCC) solver for disambiguating/improving the estimate, and modifying them both to take into account a known orthogonal geometry. The P3P solver was the Kneip solver, and the point-cloud-correspondence solver was the Optimal Linear Attitude Estimator (OLAE). We were able to achieve over 40% reduction in the computational time of the P3P solver, and around 10% for the PCC solver, vs. the unmodified solvers acting on the same problems. It is possible that the Kneip P3P solver was particularly well suited to this approach. Nevertheless, these findings suggest similar investigation may be worthwhile for other PNP solvers, if (1) processing resources are scarce, and (2) target geometry can be known in advance.

Control algorithm is developed for a satellite with flexible appendages to achieve a good pointing performance. Detailed modeling activity was carried out that consists of sensor and actuator models, disturbances and system dynamics. Common hardware found in the spacecraft such as reaction wheels, gyroscopes, star trackers etc. were included in the model. Furthermore, the Newton-Euler method is employed for the derivation of multi-body equations of motion. Evaluation of the pointing accuracy with proper pointing performance metrics such as accuracy, jitter and stability during slew maneuvers are obtained through simulations. Control strategies are proposed to improve pointing performance.

The U.S Department of Defense and the National Aeronautics and Space Administration continue to face common challenges in the development and acquisition of their space systems. In particular, space programs repeatedly experience significant schedule slippages, and spacecraft are often delivered on-orbit several months, sometimes years, after the initially planned delivery date. The repeated pattern of these schedule slippages suggests deep-seated flaws in managing spacecraft delivery and schedule risk, and an inadequate understanding of the drivers of schedule slippages. Furthermore, due to their long development time and physical inaccessibility after launch, space systems are exposed to a particular and acute risk of obsolescence, resulting in loss of value or competitive advantage over time. The perception of this particular risk has driven some government agencies to promote design choices that may ultimately be contributing to these schedule slippages, and jeopardizing what is increasingly recognized as critical, namely space responsiveness. The overall research objective of this work is twofold: (1) to identify and develop a thorough understanding of the fundamental causes of the risk of schedule slippage and obsolescence of space systems; and in so doing, (2) to guide spacecraft design choices that would result in better control of spacecraft delivery schedule and mitigate the impact of these "temporal risks" (schedule and obsolescence risks). To lay the groundwork for this thesis, first, the levers of responsiveness, or means to influence schedule slippage and impact space responsiveness are identified and analyzed, including design, organizational, and launch levers. Second, a multidisciplinary review of obsolescence is conducted, and main drivers of system obsolescence are identified. This thesis then adapts the concept of a technology portfolio from the macro- or company level to the micro-level of a single complex engineering system, and it analyzes a space system as a portfolio of technologies and instruments, each technology with its distinct stochastic maturation path and exposure to obsolescence. The selection of the spacecraft portfolio is captured by parameters such as the number of instruments, the initial technology maturity of each technology/instrument, the resulting heterogeneity of the technology maturity of the whole system, and the spacecraft design lifetime. Building on the abstraction of a spacecraft as a portfolio of technologies, this thesis then develops a stochastic framework that provides a powerful capability to simultaneously explore the impact of design decisions on spacecraft schedule, on-orbit obsolescence, and cumulative utility delivered by the spacecraft. Specifically, this thesis shows how the choice of the portfolio size and the instruments Technology Readiness Levels (TRLs) impact the Mean-Time-To-Delivery (MTTD) of the spacecraft and mitigate (or exacerbate) schedule risk. This work also demonstrates that specific combinations/choices of the spacecraft design lifetime and the TRLs can reduce the risk of on-orbit obsolescence. This thesis then advocates for a paradigm shift towards a calendar-based design mindset, in which the delivery time of the spacecraft is accounted for, as opposed to the traditional clock-based design mindset. The calendar-based paradigm is shown to lead to different design choices, which are more likely to prevent schedule slippage and/or enhance responsiveness and ultimately result in a larger cumulative utility delivered. Finally, missions scenarios are presented to illustrate how the framework and analyses here proposed can help identify system design choices that satisfy various mission objectives and constraints (temporal as well as utility-based).

Spacecraft trajectory optimization is a topic of crucial importance to space missions design. The less fuel required to accomplish the mission, the more payload that can be transported, and the higher the opportunity to lower the cost of the space mission. The objective is to find the optimal trajectory through space that will minimize the fuel used, and still achieve all mission constraints. Most space trajectories are designed using the simplified relative two-body problem as the base model. Using this patched conics approximation, however, constrains the solution space and fails to produce accurate initial guesses for trajectories in sensitive dynamics. This dissertation uses the Circular Restricted Three-Body Problem (CR3BP) as the base model for designing transfer trajectories in the Circular Restricted Four-Body Problem (CR4BP). The dynamical behavior of the CR3BP guides the search for useful low-energy trajectory arcs. Two distinct models of the CR4BP are considered in this research: the Concentric model, and the Bi-Circular model. Transfers are broken down into trajectory arcs in two separate CR3BPs and the stable and unstable manifold structures of both systems are utilized to produce low-energy transfer arcs that are later patched together to form the orbit-to-orbit transfer. The patched solution is then used as an initial guess in the CR4BP model. A vital contribution of this dissertation is the sequential process for initial guess generation for transfers in the CR4BP. The techniques discussed in this dissertation overcome many of the difficulties in the trajectory design process presented by the complicated dynamics of the CR4BP. Indirect optimization techniques are also used to derive the first order necessary conditions for optimality to assure the optimality of the transfers and determine whether additional impulses might further lower the total cost of the mission. / text

One of the most important aspects of preliminary interplanetary mission planning entails designing a trajectory that delivers a spacecraft to the required destinations and accomplishes all the objectives. The design tool described in this thesis allows an investigator to explore various interplanetary trajectories quickly and easily. The design tool employs the patched conic method to determine heliocentric and planetocentric trajectory information. An existing Lambert Targeting routine and other common algorithms are utilized in conjunction with the design tool’s specialized code to formulate an entire trajectory from Earth departure to arrival at the destination. The tool includes many options for the investigator to accurately configure the desired trajectory, including planetary gravity assists, deep space maneuvers, and various departure and arrival conditions. The trajectory design tool is coded in MATLAB, which provides access to three dimensional plotting options and user adaptability. The design tool also incorporates powerful MATLAB optimization functions that adjust trajectory characteristics to find a configuration that yields the minimum spacecraft propellant in the form of change in velocity. / text