A comprehensive large-scale power system modeling is developed to facilitate the design and analysis of present and future spacecraft power systems. A two-port coupling method is utilized to provide a modularity in model building and analysis of the system. The modular approach allows the model to be flexible, verifiable and computationally efficient. A methodology for the system level analysis is presented with the ability to focus on the performance characteristics of an arbitrary component or subsystem. The system performance parameters are derived explicitly in terms of the two-port hybrid g-parameter representation of the component or subsystem, and impedances of its terminating subsystems. From this, the stability of the system is analytically determined and the subsystem interaction criteria is observed. Also presented is a model development from the empirical data employing the complex curve fitting technique. The technique is especially powerful for large scale system modeling and analysis where certain components and subsystems are viewed as black boxes with measurable terminal characteristics. The technique can also be used to realize a reduced order model of a complex subsystem. The Direct Energy Transfer (DET) spacecraft power system is modeled to demonstrate the versatility of the comprehensive system model by performing various DC, small-signal and large-signal analyses. Of particular interest is the analysis of the large-signal behavior of the nonlinear solar array system by employing the state-plane method. The analysis of the solar array system operation focused on the transition mode between the shunt mode and the battery discharging mode is presented. The subsystem interaction problems in the local component and global system are illustrated. A methodology for the design and trouble-shooting of a system dealing with the interaction problems using the g-parameters is described. Finally, a system level analysis of the DET system using an empirical data modeling technique is performed. / Ph. D.

High Mach number, low-Reynolds number (high-altitude) reentry flowfield predictions are an important problem area in computational aerothermodynamics. Available numerical tools for handling such flows are very few and significantly limited in their applicability. A new implicit fully-iterative Parabolized Navier-Stokes (PNS) scheme is developed to accurately predict such low-Reynolds number flows. In this new approach the differential equations governing the conservation of mass, momentum and energy, and the algebraic equation of state for a perfect gas are solved simultaneously in a coupled manner. The idea is presented that by treating the governing equations in this manner (rather than eliminating the pressure terms in the governing equations by using appropriate differentiated forms of the equation of state) it may be possible to have an unconditionally time-like numerical scheme. The stability of a simplified version of this new PNS scheme is also studied, and it is demonstrated that these simplified equations are unconditionally time-like in the subsonic as well as the supersonic flow regions. A pseudo-time integration approach is used in addition to a new second-order accurate fully-implicit smoothing, to improve the efficiency of the solution algorithm. The new PNS scheme is used to predict the flowfield around a seven-deg sphere-cone vehicle under high- and low-Reynolds number conditions. Two test case, Case A and Case B, are chosen such that Case A has a large freestream Reynolds number (2.92x10⁵), whereas Case B has a freestream Reynolds number of 1.72x10³, which is smaller than the usual limit of applicability of the non-iterative PNS schemes (Re~10⁴ or larger). Comparisons are made with other available numerical schemes, and the results substantiate the stability, accuracy and efficiency claims of the new Parabolized Navier-Stokes scheme. / Ph. D.

Attitude control remains one of the top engineering challenges faced by small satellite mission planning and design. Conventional methods for attitude control include propulsion, reaction wheels, magnetic torque coils, and passive stabilization mechanisms, such as permanent magnets that align with planetary magnetic fields. Drawbacks of these conventional attitude control methods for small satellites include size, power consumption, dependence on external magnetic fields, and lack of full control authority. This research investigates an alternative, novel approach to attitude-control method for small satellites, utilizing the noncommutative property of rigid body rotation sequences. Piezoelectric bimorph actuators are used to induce sinusoidal small-amplitude satellite oscillations on two of the satellites axes. While zero net change occurs on these signaled axes, the third axis can develop an average angular rate. This noncommutative attitude control methodology has several advantages over conventional methods, including scalability, power consumption, and operation outside of Earth's magnetic field. This research looks into the feasibility of such a system, and lays the foundation for a simple control system architecture.

As electronics become smaller and more capable, it has become possible to conduct meaningful and sophisticated satellite missions in a small form factor. However, the capability of small satellites and the range of possible applications are limited by the capabilities of several technologies, including attitude determination and control systems. This dissertation evaluates the use of image-based visual attitude propagation as a compliment or alternative to other attitude determination technologies that are suitable for miniature satellites. The concept lies in using miniature cameras to track image features across frames and extracting the underlying rotation. The problem of visual attitude propagation as a small satellite attitude determination system is addressed from several aspects: related work, algorithm design, hardware and performance evaluation, possible applications, and on-orbit experimentation. These areas of consideration reflect the organization of this dissertation. A “stellar gyroscope” is developed, which is a visual star-based attitude propagator that uses relative motion of stars in an imager’s field of view to infer the attitude changes. The device generates spacecraft relative attitude estimates in three degrees of freedom. Algorithms to perform the star detection, correspondence, and attitude propagation are presented. The Random Sample Consensus (RANSAC) approach is applied to the correspondence problem to successfully pair stars across frames while mitigating false-positive and false-negative star detections. This approach provides tolerance to the noise levels expected in using miniature optics and no baffling, and the noise caused by radiation dose on orbit. The hardware design and algorithms are validated using test images of the night sky. The application of the stellar gyroscope as part of a CubeSat attitude determination and control system is described. The stellar gyroscope is used to augment a MEMS gyroscope attitude propagation algorithm to minimize drift in the absence of an absolute attitude sensor. The stellar gyroscope is a technology demonstration experiment on KySat-2, a 1-Unit CubeSat being developed in Kentucky that is in line to launch with the NASA ELaNa CubeSat Launch Initiative. It has also been adopted by industry as a sensor for CubeSat Attitude Determination and Control Systems (ADCS).

"The ablative Teflon pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion enabling technology for small spacecraft missions. The integration of PPTs onboard spacecraft requires the understanding and evaluation of possible thruster/spacecraft interactions. To aid in this effort the work presented in this thesis is directed towards the development and application of Langmuir probe techniques for use in the plume of PPTs. Double and triple Langmuir probes were developed and used to measure electron temperature and density of the PPT plume. The PPT used in this thesis was a laboratory model parallel plate ablative Teflon® PPT similar in size to the Earth Observing (EO-1) PPT operating in discharge energies between 5 and 40 Joules. The triple Langmuir probe was operated in the current-mode technique that requires biasing all three electrodes and measuring the resulting probe currents. This new implementation differs from the traditional voltage-mode technique that keeps one probe floating and requires a voltage measurement that is often susceptible to noise in the fluctuating PPT plume environment. The triple Langmuir probe theory developed in this work incorporates Laframboise’s current collection model for Debye length to probe radius ratios less than 100 in order to account for sheath expansion effects on ion collection, and incorporates the thin-sheath current collection model for Debye length to probe radius ratios greater than 100. Error analysis of the non-linear system of current collection equations that describe the operation of the current-mode triple Langmuir probe is performed as well. Measurements were taken at three radial locations, 5, 10, and 15 cm from the Teflon® surface of the PPT and at angles of 20 and 40 degrees to either side of the thruster centerline as well as at the centerline. These measurements were taken on two orthogonal planes, parallel and perpendicular to the PPT electrodes. A data-processing software was developed and implements the current-mode triple Langmuir probe theory and associated error analysis. Results show the time evolution of the electron temperature and density. Characteristic to all the data is the presence of hot electrons of approximately 5 to 10 eV at the beginning of the pulse, occurring near the peak of the discharge current. The electron temperature quickly drops off from its peak values to 1-2 eV for the remainder of the pulse. Peak electron densities occur after the peak temperatures. The maximum electron density values on the centerline of the plume of a laboratory PPT 10 cm from the Teflon® surface are 6.6x10^19 +/- 1.3x10^19 m^-3 for the 5 J PPT, 7.2x10^20 +/- 1.4x10^20 m^-3 for the 20 J PPT, and 1.2x10^21 +/- 2.7x10^20 m^-3 for the 40 J PPT. Results from the double Langmuir probe taken at r=10 cm, theta perpendicular=70 degrees and 90 degrees of a laboratory PPT showed good agreement with the triple probe method."

"The National Aeronautics and Space Administration (NASA) plans to send a human exploration team to Mars within the next 25 years. In support of this effort Hamilton Standard Space Systems International (HSSSI), current manufacturers of the Space Shuttle spacesuit, began exploring alternative solutions for supporting an astronaut during a Martian surface exploration. A design concept was developed by HSSSI to integrate a minimally equipped Martian spacesuit with a robotic support cart capable of providing life support assistance, communications, and independent navigational functions. To promote NASA's visionary efforts and increase university relations, HSSSI partnered with Worcester Polytechnic Institute (WPI) to develop a proof-of-concept robotic support cart system, FIDOE - Fully Independent Delivery of Expendables. As a proof-of-concept system, the primary goal of this project was to demonstrate the feasibility of current technologies utilized by FIDOE's communication and controls system for future Martian surface explorations. The primary objective of this project was to procure selected commercial-off-the-shelf components and configure these components into a functional robotic support cart. The design constraints for this project, in addition to the constraints imposed by the Martian environment and HSSSI's Martian spacesuit, were a one-year time frame and a $20,000 budget for component procurement. This project was also constrained by the protocols defined by the NASA demonstration test environment. The final design configuration comprised of 37 major commercial off-the-shelf components and three individual software packages that integrated together to provide FIDOE's communications and control capabilities. Power distribution was internally handled through a combination of a main power source and dedicated power supplies. FIDOE also provided a stowage area for handling assisted life support systems and geological equipment. The proof-of-concept FIDOE system proved that the current technologies represented by the selected components are feasible applications for a Mars effort. Specifically, the FIDOE system demonstrated that the chosen technologies can be integrated to perform assisted life support and independent functions. While some technologies represented by the proof-of-concept system may not adequately address the robustness issues pertaining to the Mars effort, e.g., voice recognition and power management, technology trends indicate that these forms of technology will soon become viable solutions to assisting an astronaut on a Martian surface exploration."

The CubeSat community has begun to develop and implement propulsion systems. This movement represents a new capability which may satisfy mission needs such as orbital and constellation maintenance, formation flight, de-orbit, and even interplanetary travel. With the freedom and capability granted by propulsion systems, CubeSat providers must accept new responsibilities in proportion to the potential hazards that propulsion systems may present. The Cal Poly CubeSat program publishes and maintains the CubeSat Design Specification (CDS). They wish to help the CubeSat community to safety and responsibly expand its capabilities to include propulsive designs. For this reason, the author embarked on the task of developing a draft of safety standards CubeSat propulsion systems. Wherever possible, the standards are based on existing documents. The author provides an overview of certain concepts in systems safety with respect to the classification of hazards, determination of required fault tolerances, and the use of inhibits to satisfy fault tolerance requirements. The author discusses hazards that could exist during ground operations and through launch with respect to hazardous materials and pressure systems. Most of the standards related to Range Safety are drawn from AFSPCMAN 91-710. Having reviewed a range of hypothetical propulsion system architectures with an engineer from Range Safety at Vandenberg Air Force Base, the author compiled a case study. The author discusses many aspects of orbital safety. The author discusses the risk of collision with the host vehicle and with third party satellites along with the trackability of CubeSats using propulsion systems. Some recommendations are given for working with the Joint Functional Component Command for Space (JFCC SPACE), thanks to the input of two engineers who work with the Joint Space Operations Center (JSpOC). Command Security is discussed as an important aspect of a mission which implements a propulsion system. The author also discusses End-of-Life procedures such as safing and de-orbit operations. The orbital safety standards are intended to promote “good citizenship.” The author steps through each proposed standard and offers justification. The author is confident that these standards will set the stage for a dialogue in the CubeSat community which will lead to the formulation of a reasonable and comprehensive set of standards. The author hopes that the discussions given throughout this document will help CubeSat developers to visualize the path to flight readiness so that they can get started on the right foot.

Most present day spacecrafts have large interconnected solar panels. The dynamic behavior of the spacecraft in orbit can be modeled as a free rigid mass with flexible elements attached to it. The natural frequencies of such spacecrafts with deployed solar panels are very low. The low values of the natural frequencies pose difficulties for maneuvering the spacecraft. The control torque required to maneuver the spacecraft is influenced by the flexibility of the solar arrays. The control torque sets up transient oscillations in the flexible solar panels which in turn induces disturbances in the rigid satellite body and the payload within. Therefore the payload operations can be carried out only after the disturbances die out. For any reduction of the above disturbances it is necessary to understand the dynamic behavior of such systems to an applied torque. The present work first studies the nature of the disturbances. The influence of structural parameters on these disturbances is then investigated. Finally, the use of passive damping treatment using viscoelastic material is investigated for the reduction of the disturbances. In order to understand the nature of vibrations induced in the flexible appendages of a satellite during maneuvers, we model the maneuver loads in terms of applied angular acceleration as well as varying torque. The transient decay of the disturbance of the rigid element is characterized by the dynamic characteristics of the flexible panels or appendages. It is shown that by changing the stiffness of the panel the response of the rigid element can be modified. A simple model consisting of an Euler-Bernoulli beam attached to a free mass is next considered. The influence of various parameters of the EulerBernoulli beam in mitigating vibration and thereby the disturbance in the rigid mass is investigated. As the response of the rigid system mounted with the large flexible panels are influenced by the dynamics of the flexible panels, reduction of these disturbances can be achieved by reducing the vibration in the flexible panels. Therefore application of viscoelastic materials for passive damping treatment is investigated. The loss factor of a structure is significantly improved by using constrained viscoelastic layer damping treatment. However providing a constrained layer damping treatment on the entire structure is very inefficient in terms of the additional mass involved. Therefore damping material is applied at suitable optimal locations. In previous studies reported in literature, modal strain energy distribution in the viscoelastic material as well as the base structure is used as a tool to arrive at the optimum location for the damping treatment. It is shown in this study that such locations selected are not the optimum. A new approach is proposed in this study by which both the above shortcomings are overcome. It is shown that use of a parameter that is the ratio of the strain in the viscoelastic material to the angle of flexure is a more reliable measure in arriving at optimal locations for the application of constrained viscoelastic layers. The method considers the deformations in the viscoelastic material and it is shown that significant values of loss factors are achieved by providing material in a small region alone. We also show that loss factor can be improved by providing damping material near the interface region. The loss factor can be further improved by incorporating spacers by using spacer material having higher extensional modulus. Also shown is the fact that loss factor is unaffected by the shear modulus of the spacer material. Experiments have been conducted to validate these results. In a related study we consider honeycomb type flexible structures since in most of the spacecraft applications honeycomb sandwich constructions are employed. But loss factors of sandwich panels with constrained layer damping treatment are seldom discussed in the literature. Use of viscoelastic layers to improve the loss factors of the honeycomb sandwich beams is explored. The results show that the loss factors are enhanced by increasing the inplane stiffness of the constraining layer. These conclusions too are validated by experimental results. Finally a typical satellite with flexible solar panels is considered, and the use of the viscoelastic material for improving the damping is demonstrated.

This dissertation develops a general framework for designing stabilizing feedback controllers and observers for dynamics with state/time dependent gains on the control signals and measured outputs. These gains have potential singularity periods but satisfy a technically non-trivial condition referred to as persistence of excitation. A persistence filter design constitutes the primary theoretical innovation of this work around which the controller and observer development is centered. Application areas of singular gain systems considered in this study include robotics, biomechanics, intelligent structures and spacecrafts. Several representative problems involving singular, time-dependent gains are addressed. The specific contributions of this dissertation are outlined as follows: (i) a stabilizing feedback for linear, single-input systems with time-varying, singular control scaling is designed that allows arbitrary exponential convergence rate for the closed-loop dynamics. An adaptive control generalization of this result allows asymptotic convergence in presence of unknown plant parameters. An extension to a special, single-input nonlinear system in the controller canonical form is also proposed. It is proven that this control design results in bounded tracking error signals for a trajectory tracking objective; (ii) observer design for linear, single-output systems with time-varying, singular measurement gains is considered. A persistence filter similar in structure to the control counterpart aids an observer design that guarantees exponential state reconstruction with arbitrary convergence rates; (iii) the observer and controller designs are combined to obtain an exponentially stabilizing output feedback controller for linear, single-input, single-output dynamics with singular gains on both the control and measurements. A novel separation property is established as a consequence. The construction motivates applications to stabilization with reversible transducers which can switch between sensor and actuator modes. The results are verified on two illustrative applications, vibration control using piezoelectric devices and inverted pendulum stabilization with a DC motor. The linear result is further generalized to include state dependent gains; (iv) application of the persistence filter theory to spacecraft attitude stabilization using intermittent actuation is explored. The intermittence is characterized by a time-varying, periodically singular control gain. A nonlinear persistence filter allows construction of an exponentially stabilizing controller and simulations verify convergence with intermittent actuation where conventional proportional-derivative control fails; (v) a stabilization result for a special multi-input, linear system with time-varying matrix control gains is presented. The matrix gain is assumed to be diagonal but allows fewer controls than states subject to a controllability assumption in absence of the singular gain matrix. The single-input adaptive control results are shown to extend to the multi-input case. An application to angular velocity stabilization of an underactuated rigid spacecraft is considered. / text

The incorporation of inflatable structures into aerospace systems can produce significant advantages in stowed volume to mechanical effectiveness and overall weight. Many applications of these ultra-lightweight systems are designed to precisely control internal or external surfaces, or both, to achieve desired performance. The modeling of these structures becomes complex due to the material nonlinearities inherent to the majority of construction materials used in inflatable structures. Furthermore, accurately modeling the response and behavior of the interfacing boundaries that are common to many inflatable systems will lead to better understanding of the entire class of structures. The research presented involved using nonlinear finite element simulations correlated with photogrammetry testing to develop a procedure for defining material properties for commercially available polyurethane-coated woven nylon fabric, which is representative of coated materials that have been proven materials for use in many inflatable systems. Further, the new material model was used to design and develop an inflatable pathfinder system which employs only internal pressure to control an assembly of internal membranes. This canonical inflatable system will be used for exploration and development of general understanding of efficient design methodology and analysis of future systems. Canonical structures are incorporated into the design of the phased pathfinder system to allow for more universal insight. Nonlinear finite element simulations were performed to evaluate the effect of various boundary conditions, loading configurations, and material orientations on the geometric precision of geometries representing typical internal/external surfaces commonly incorporated into inflatable pathfinder system. The response of the inflatable system to possible damage was also studied using nonlinear finite element simulations. Development of a correlated material model for analysis of the inflatable pathfinder system has improved the efficiency of design and analysis techniques of future inflatable structures.