CubeSat designers are increasingly looking to incorporate delicate structures and optics into their payloads. These delicate payloads, however, may not survive the required absolute-worst-case launch vibration testing needed for flight certification. To help address this problem, and to better match testing conditions to real-world launch environments, this thesis introduces Sync, a compact 1/4U CubeSat payload designed to collect data on the vibrations and thermal environments CubeSats experience inside a deployer on the way to orbit. This data can be used to better understand the launch environment for different vehicles, and help develop new, more realistic testing guidelines that could enable more delicate payloads to be launched.

To enhance mechanical and/or electrical properties of composite materials used in additive manufacturing, nanoparticles are often time deposited to form nanocomposite layers. To customize the mechanical and/or electrical properties, the thickness of such nanocomposite layers must be precisely controlled. A thickness model of filter cakes created through a spray assisted vacuum filtration is presented in this paper, to enable the development of advanced thickness controllers. The mass transfer dynamics in the spray atomization and vacuum filtration are studied for the mass of solid particles and mass of water in differential areas, and then the thickness of a filter cake is derived. A two-loop nonlinear constrained optimization approach is used to identify the unknown parameters in the model. Experiments involving depositing carbon nanofibers in a sheet of paper are used to measure the ability of the model to mimic the filtration process.

To date the commercial use of shape memory alloys (SMAs) has been mostly limited to binary NiTi alloys with transformation temperatures approximately in the -100 to 100 ºC range. In an ongoing effort to develop high-temperature shape memory alloys (HTSMAs), ternary and quaternary additions are being made to binary NiTi to form NiTi-X (e.g., X: Pd, Pt, Au and Hf) alloys. Stability and repeatability can be further increased at these higher temperatures by limiting the stress, but the tradeoff is reduced work output and stroke. However, HTSMAs operating at decreased stresses can still be used effectively in actuator applications that require large strokes when used in the form of springs. The overall objective of this work is to facilitate the development of HTSMAs for use as high-force actuators in active/adaptive aerospace structures. A modular test setup was assembled with the objective of acquiring stroke, stress, temperature and moment data in real time during joule heating and forced convective cooling of Ni19.5Ti50.5Pd25Pt5 HTSMA springs. The spring actuators were evaluated under both monotonic axial loading and thermomechanical cycling. The role of rotational constraints (i.e., by restricting rotation or allowing for free rotation at the ends of the springs) on stroke performance was also assessed. Recognizing that evolution in the material microstructure results in changes in geometry and vice versa in HTSMA springs, the objective of the present study also included assessing the contributions from the material microstructural evolution, by eliminating contributions from changes in geometry, to overall HTSMA spring performance. The finite element method (FEM) was used to support the analytical analyses and provided further insight into the behavior and heterogeneous stress states that exist in these spring actuators. iv Furthermore, with the goal of improving dimensional stability there is a need to better understand the microstructural evolution in HTSMAs that contributes to irrecoverable strains. Towards this goal, available Ni29.5Ti50.5Pd20 neutron diffraction data (from a comparable HTMSA alloy without the solid solution strengthening offered by the Pt addition) were analyzed. The data was obtained from in situ neutron diffraction experiments performed on Ni29.5Ti50.5Pd20 during compressive loading while heating/cooling, using the Spectrometer for Materials Research at Temperature and Stress (SMARTS) at Los Alamos National Laboratory. Specifically, in this work emphasis was placed on neutron diffraction data analysis via Rietveld refinement and capturing the texture evolution through inverse pole figures. Such analyses provided quantitative information on the evolution of lattice strain, phase volume fraction (including retained martensite that exists above the austenite finish temperature) and texture (martensite variant reorientation and detwinning) under temperature and stress. Financial support for this work from NASA’s Fundamental Aeronautics Program Supersonics Project (NNX08AB51A), Subsonic Fixed Wing Program (NNX11AI57A) and the Florida Center for Advanced Aero-Propulsion (FCAAP) is gratefully acknowledged. It benefited additionally from the use of the Lujan Neutron Scattering Center at Los Alamos National Laboratory, which is funded by the Office of Basic Energy Sciences (Department of Energy) and is operated by Los Alamos National Security LLC under DOE Contract DE-AC52-06NA25396.

With the recent trend for systems to be more and more autonomous, there is a growing need for cooperative trajectory planning. Applications that can be considered as cooperative systems such as surveying, formation flight, and traffic control need a method that can rapidly produce trajectories while considering all of the constraints on the system. Currently most of the existing methods to handle cooperative control are based around either simple dynamics and/or on the assumption that all vehicles have homogeneous properties. In reality, typical autonomous systems will have heterogeneous, nonlinear dynamics while also being subject to extreme constraints on certain state and control variables. In this thesis, a new approach to the cooperative control problem is presented based on the bio-inspired motion strategy known as local pursuit. In this framework, decision making about the group trajectory and formation are handled at a cooperative level while individual trajectory planning is considered in a local sense. An example is presented for a case of an autonomous farming system (e.g. scouting) utilizing nonlinear vehicles to cooperatively accomplish various farming task with minimal energy consumption or minimum time. The decision making and trajectory generation is handled very quickly while being able to consider changing environments laden with obstacles

Current trends in spacecraft are leading to smaller, more inexpensive options whenever possible. This shift has been primarily pursued for the opportunity to open a new frontier for technologies with a small financial obligation. Limited power, processing, pointing, and communication capabilities are all common issues which must be considered when miniaturizing systems and implementing low-cost components. This thesis addresses some of these concerns by applying two methods, in attitude estimation and control. Additionally, these methods are not restricted to only small, inexpensive satellites, but offer a benefit to large-scale spacecraft as well. First, star cameras are examined for the tendency to generate streaked star images during maneuvers. This issue also comes into play when pointing capabilities and camera hardware quality are low, as is often the case in small, budget-constrained spacecraft. When pointing capabilities are low, small residual velocities can cause movement of the stars in the focal plane during an exposure, causing them to streak across the image. Additionally, if the camera quality is low, longer exposures may be required to gather sufficient light from a star, further contributing to streaking. Rather than improving the pointing or hardware directly, an algorithm is presented to retrieve and utilize the endpoints of streaked stars to provide feedback where traditional methods do not. This allows precise attitude and angular rate estimates to be derived from an image which, with traditional methods, would return large attitude and rate error. Simulation results are presented which demonstrate endpoint error of approximately half a pixel and rate estimates within 2% of the true angular velocity. Three methods are also considered to remove overlapping star streaks and resident space objects from images to improve performance of both attitude and rate estimates. Results from a large-scale Monte Carlo simulation are presented in order to characterize the performance of the method. iii Additionally, a rapid optimal attitude guidance method is experimentally validated in a groundbased, pico-scale satellite test bed. Fast slewing performance is demonstrated for an incremental step maneuver with low average power consumption. Though the focus of this thesis is primarily on increasing the capabilities of small, inexpensive spacecraft, the methods discussed have the potential to increase the capabilities of current and future large-scale missions as well.

A conceptual dust detection mission, KnightSat III, using pico-scale satellites is analyzed. The purpose of the proposed KnightSat III mission is to aid in the determination of the size, mass, distribution, and number of dust particles in low earth orbits through a low cost and flexible satellite or a formation of satellites equipped with a new dust detector. The analysis of a single satellite mission with an on-board dust detector is described; though this analysis can easily be extended to a formation of pico-scale satellites. Many design aspects of the mission are discussed, including orbit analysis, power management, attitude determination and control, and mass and power budgets. Two of them are emphasized. The first is a new attitude guidance and control method, and the second is the online optimal power scheduling. It is expected that the measurements obtained from this possible future mission will provide insight into the dynamical processes of inner solar system dust, as well as aid in designing proper micro-meteoroid impact mitigation strategies for future man-made spacecraft.

As development of CubeSat based architectures increase, methods of deploying constellations of CubeSats are required to increase functionality of future systems. Given their low cost and quickly increasing launch opportunities, large numbers of CubeSats can easily be developed and deployed in orbit. However, as secondary payloads, CubeSats are severely limited in their options for deployment into appropriate constellation geometries. This thesis examines the current methods for deploying cubes and proposes new and efficient geometries using secondary launch opportunities. Due to the current deployment hardware architecture, only the use of different launch opportunities, deployment direction, and deployment timing for individual cubes in a single launch are explored. The deployed constellations are examined for equal separation of Cubes in a single plane and effectiveness of ground coverage of two regions. The regions examined are a large near-equatorial zone and a medium sized high latitude, high population density zone. Results indicate that simple deployment strategies can be utilized to provide significant CubeSat dispersion to create efficient constellation geometries. The same deployment strategies can be used to develop a multitude of differently dispersed constellations. Different launch opportunities can be utilized to tailor a constellation for a specific region or mission objective. Constellations can also be augmented using multiple launch opportunities to optimize a constellation towards a specific mission or region. The tools developed to obtain these results can also be used to perform specific analysis on any region in order to optimize future constellations for other applications.

CubeSat satellites have redefined the standard solution for conducting missions in space due to their unique form factor and cost. The harsh environment of space necessitates examining features that improve satellite robustness and ultimately extend lifetime, which is typical and vital for mission success. The CubeSat development team at Cal Poly, PolySat, has recently redefined its standard avionics platform to support more complex mission capabilities with this robustness in mind. A significant addition was the integration of the Linux operating system, which provides the flexibility to develop much more elaborate protection mechanisms within software, such as support for remote on-orbit software updates. This thesis details the design and development of such a feature-set with critical software recovery and multiple-mission single-CubeSat functionality in mind. As a result, features that focus on software update usability, validation, system recovery, upset tolerance, and extensibility have been developed. These include backup Linux kernel and file system image availability, image validation prior to boot, and the use of multiple file system devices to protect against system upsets. Furthermore, each feature has been designed for usability on current and future missions.

This thesis examines the feasibility of substituting the system input script of Cal Poly’s Horizon Simulation Framework (HSF) with a Model Based Systems Engineering (MBSE) model designed with the Systems Modeling Language (SysML). A concurrent student project, SysML Output Interface Creation for the Horizon Simulation Framework, focused on design of the HSF Translator Plugin which converts SysML models to an HSF specific XML format. A SysML model of the HSF test case, Aeolus, was designed. The original Aeolus HSF input script and the translated SysML input script retained the format and dependency structure required by HSF. Both input scripts returned identical results and thus validated the feasibility of linking SysML with HSF through the HSF Translator Plugin. A second SysML model of the Cal Poly CubeSat mission, ExoCube, was also designed and converted into an HSF input script. The ExoCube input script also retained the format and dependency structure required by HSF. This demonstrated that future SysML models can be used in conjunction with the HSF Translator Plugin to create a functional HSF system input script.

There is an increased need for spacecraft to quickly and efficiently de-orbit themselves as the amount of debris in orbit around Earth grows. Defunct spacecraft pose a significant threat to the LEO environment due to their risk of fragmentation. If these spacecraft are de-orbited at the end of their useful life their risk to future spacecraft is greatly lessened. A proposed method of efficiently de-orbiting spacecraft is to use an inflatable thin-film envelope to increase the body's area to mass ratio and thusly shortening its orbital lifetime. The system and analysis presented in this project is sized for use on a CubeSat as they are an effective utility as a technology demonstration platform. Analysis has been performed to characterize the orbital dynamics of high area to mass ratio spacecraft as well as the leak rate of such an inflatable device in a vacuum environment. Results show that a 1U CubeSat can be de-orbited using a 1.7 meter diameter spherical device in just under one year while using 0.7 grams of inflating gas, this is compared to over 25 years without any method of post-mission disposal.