An LED-based solar simulator has been designed, constructed, and qualified under ASTM standards for use in the Cal Poly Space Environments Laboratory. The availability of this simulator will enhance the capability of undergraduate students to evaluate solar cell and thermal coating performance, and offers further research opportunities. The requirements of ASTM E927-19 for solar simulators intended for photovoltaic cell testing were used primarily, supplemented by information from ASTM E491-73 for solar simulators intended for spacecraft thermal vacuum testing. Three main criteria were identified as design goals - spectral match ratio, spatial non-uniformity, and temporal instability. An electrical design for an LED-based simulator to satisfy these criteria was developed and implemented, making use of existing lab equipment where possible to minimize cost. The resulting simulator meets the desired spatial non-uniformity and temporal instability requirements of ASTM E927-19, but falls short of the spectral match ratio needed. This is shown to be due to a calibration issue that is easily amended via software. The simulator is overall Class UCB under ASTM E927, and Class CCC under ASTM E491. The simulator was used to conduct the same laboratory procedure for solar cell I-V curve testing as performed by undergraduate students, showing excellent promise as a course enhancement.

A light curve is a plot of brightness versus time of an object. Light curves are dependent on orbit, attitude, surface area, size, and shape of the observed object. Using light curve data, several analysis methods have been developed to derive these parameters. These parameters can be used for tracking orbital debris, monitoring satellite health, and determining the mission of an unknown spacecraft. This paper discusses the development, verification, and utilization of a tool that simulates light curve data. This tool models ground-based observations, space-based observations, self-shadowing geometry, tumbling debris, and controlled spacecraft. The main output from the tool is the pass prediction plot and the light curve plot. The author intends to publish the tool and supporting documents for future researchers to utilize. This will save researchers time developing their own models and the tool can act as a baseline for comparisons between analysis methods. For clarity, this paper does not develop nor implement a light curve analysis method, but rather creates a tool to simulate light curve observations and data. Each section of the tool was verified independently to ensure that the simulated light curves were correct. The tool was verified with STK, matlab, and simulink. It predicts the start and end times of passes, eclipses, and ground-site night cycles within 1% of the total event duration, when compared to STK. The attitude propagator predicts the attitude of the target with offsets less than 0.06 degrees on average and a maximum offset less than 0.6 degrees when compared to provided attitude code.

Attitude determination and estimation algorithms are developed and implemented in simulation for the Exocube satellite currently under development by PolySat at Cal Poly. A mission requirement of ±5˚ of attitude knowledge has been flowed down from the NASA Goddard developed payload, and this requirement is to be met with a basic sensor suite and the appropriate algorithms. The algorithms selected in this work are TRIAD and an Extended Kalman Filter, both of which are placed in a simulation structure along with models for orbit propagation, spacecraft kinematics and dynamics, and sensor and reference vector models. Errors inherent from sensors, orbit position knowledge, and reference vector generation are modeled as well. Simulations are then run for anticipated dynamic states of Exocube while varying parameters for the spacecraft, attitude algorithms, and level of error. The nominal case shows steady state convergence to within 1˚ of attitude knowledge, with sensor errors set to 3.5˚ and reference vector errors set to 2˚. The algorithms employed have their functionality confirmed with the use of STK, and the simulations have been structured to be used as tools to help evaluate attitude knowledge capabilities for the Exocube mission and future PolySat missions.

The Horizon Simulation Framework, or HSF, is a modeling and simulation framework compiled from C/C++ source code into a command line program. Picasso is an interface designed to control the input files to Horizon by providing visual tools to create and manipulate the XML files used to define an HSF system of assets, their environment, and other simulation parameters. Picasso also supports the visualization of Horizon output in several different forms, and import mechanics from online space object catalogues.

Because origami has proved to be a tremendously rich source of inspiration in engineering, interest in solving some of the challenges that affect origami-inspired design has been significant. One such challenge involves ensuring that origami-inspired mechanisms are rigid-foldable or capable of moving without requiring links to bend or distort. Because rigid-foldability is essential in mechanisms that are constructed using rigid materials, access to methods of engineering this characteristic are highly desirable. This research addresses this need by developing methods for the design of origami-inspired mechanisms that are rigid-foldable. Methods for modifying crease patterns to achieve this are described and compared. Methods for achieving rigid-foldability using thick materials are also developed. Proofs of a process for generating new variations of existing thick-origami models are developed and demonstrated on multiple models. The possibility of using compliant panels to create rigid-foldability is also studied.Because of the relationship between mechanism stiffness and rigid-foldability, means of managing the pattern stiffness are also examined. The design of compliant, deployable stiffeners is studied to permit a comparison of different stiffener types. This comparison is used to identify dominant configurations that are most advantageous for a deployable mechanism. The use of thick-origami models are also considered. The geometry of two varieties of a cantilever tube are optimized to support a cantilever beam.

In the present work, two numerical tools are under scrutiny. Both were made to study the atmospheric re-entry of a spacecraft: DEBRISK computes the trajectory and the survivability of the spacecraft as well as its fragments, and ELECTRA calculates the trajectory of the spacecraft and its fragments, as well as the associated on-ground risk of human casualty. However, they differ in some of their functionalities and their physical models, leading to a difference in the trajectories, thus in the impact points locations for the same spacecraft. This work has multiple purposes. First, the influence of several simulation parameters are studied in both tools in order to determine a correction law for the trajectory of the spacecraft in ELECTRA, making it imitate the DEBRISK trajectory. To do so, a large dataset is built then manipulated, and a verification process is realised to quantify the accuracy of the correction law. Successive iterations of the method show a decent improvement in the ELECTRA trajectory, yet uncertainties around the correction and the low applicability of the law lead to try a new promising method based on a live data reading of the flight parameters from DEBRISK to ELECTRA. Finally, the influence of the shielding of the buildings on the human casualty risk computation, symbolised by a protection coefficient in ELECTRA is studied. Results show that considering this, protection coefficients can multiply up by five the risk of casualty. A technical documentation was written for potential future works on the same subject. / I detta arbete studeras två numeriska verktyg som utformats för att studera det atmosfäriska återinträdet av en rymdfarkost: DEBRISK beräknar rymdfarkostens bana och överlevnadsförmåga såväl som dess fragment, och ELECTRA beräknar rymdfarkostens bana och dess fragment, samt tillhörande risk för olycksfall på marken. De skiljer sig åt i vissa av sina funktioner och sina fysiska modeller, vilket leder till skillnader i banorna, alltså i nedslagspunkterna för samma rymdfarkost. Detta arbete har flera syften. Först studeras påverkan av flera simuleringsparametrar i båda verktygen för att bestämma en korrigeringslag för rymdfarkostens bana i ELECTRA, vilket gör att den imiterar DEBRISK-banan. För att göra detta byggs en stor datamängd som sedan manipuleras, och en verifieringsprocess realiseras för att kvantifiera korrigeringslagens korrekthet. Successiva iterationer av metoden visar en viss förbättring av ELECTRA-banan, men osäkerhet kring korrigeringen och den låga tillämpligheten av lagen leder till att en ny lovande metod, baserad på en direkt dataavläsning av flygparametrarna från DEBRISK till ELECTRA, provats. Slutligen studeras inverkan av avskärmningen av byggnaderna på riskberäkningen av mänskliga olyckor, symboliserad med en skyddskoefficient i ELECTRA. Resultaten visar att med tanke på detta kan skyddskoefficienter multiplicera upp med en faktor fem risken för olyckor. En teknisk dokumentation skrevs för potentiella framtida arbeten om samma ämne.

A spacecraft formation possesses several benefits over a single-satellite mission. However, launching a fleet of satellites is a high-cost, high-risk venture. One way to mitigate much of this risk is to demonstrate hardware and algorithm performance in groundbased testbeds. It is typically difficult to experimentally replicate satellite dynamics in an Earth-bound laboratory because of the influences of gravity and friction. An air bearing provides a very low-torque environment for experimentation, thereby recapturing the freedom of the space environment as effectively as possible. Depending upon con- figuration, air-bearing systems provide some combination of translational and rotational freedom; the three degrees of rotational freedom provided by a spherical air bearing are ideal for investigation of spacecraft attitude dynamics and control problems. An interest in experimental demonstration of formation flying led directly to the development of the Distributed Spacecraft Attitude Control System Simulator (DSACSS). The DSACSS is a unique facility, as it uses two air-bearing platforms working in concert. Thus DSACSS provides a pair of "spacecraft" three degrees of attitude freedom each. Through use of the DSACSS we are able to replicate the relative attitude dynamics between nodes of a formation such as might be required for co-observation of a terrestrial target. Many dissertations present a new mathematical technique or prove a new theory. This dissertation presents the design and development of a new experimental system. Although the DSACSS is not yet fully operational, a great deal of work has gone into its development thus far. This work has ranged from configuration design to nonlinear analysis to structural and electrical manufacturing. In this dissertation we focus on the development of the attitude determination subsystem. This work includes development of the equations of motion and analysis of the sensor suite dynamics. We develop nonlinear filtering techniques for data fusion and attitude estimation, and extend this problem to include estimation of the mass properties of the system. We include recommendations for system modifications and improvements. / Ph. D.

This dissertation investigates attitude transition maneuvers of a gyrosat-satellite between relative equilibria. The primary challenge in transitioning between relative equilibria is the proper adjustment of the system angular momentum so that upon completing the transition maneuver the gyrostat-satellite will satisfy all the requirements for a relative equilibrium. The system angular momentum is a function of the attitude trajectory taken during the transition maneuver. A new concept, the constant orbital momentum equilibrium trajectory or COMET, is introduced as a means to a straight-forward solution to a subset of the possible transitions between relative equilbria. COMETs are a class of paths in SO(3) that a gyrostat-satellite may travel along that maintain a constant system angular momentum. The primary contributions of this dissertation are the introduction and analysis of COMETs and their application to the problem of transitioning a gyrostat-satellite between two relative equilibria. The current work introduces, defines, and analyzes COMETs in detail. The requirements for a path in SO(3) to be a COMET are defined. It is shown via example that COMETs are closed-curves in SO(3). Visualizations of families of COMETs are presented and discussed in detail. A subset of COMETs are shown to contain critical points that represent isolated relative equilibrium attitudes or furcations of the COMET. The problem of transitioning between two relative equilibria is split into the sub-problems of transitioning between relative equilibria on the same COMET and transitioning between relative equilibria on different COMETs. For transitions between relative equilibria on the same COMET, an open-loop control law is developed that drives a gyrostat-satellite along the COMET until the target relative equilibrium is reached. For transitions between relative equilibria on different COMETs, an open-loop control law is developed that transfers a gyrostat-satellite from the initial relative equilibrium to a relative equilibrium that resides on the same COMET as the target relative equilbrium. Acquisition of the target relative equilibrium is then accomplished via the application of the open-loop control law for transitions between relative equilibria on the same COMET. The results of numeric simulations of gyrostat-satellites executing these transitions are presented. / Ph. D.

Spacecraft formation flying (SFF) offers robust observations of multi-scale ionospheric space weather. A number of hardware-in-the-loop (HIL) SFF simulation testbeds based on Global-Navigation-Satellite-Systems (GNSS) have been developed to support GNSS-based SFF mission design, however, none of these testbeds has been directly applied to ionospheric space weather studies. The Virginia Tech Formation Flying Testbed (VTFFTB), a GNSS-based HIL simulation testbed, has been developed in this work to simulate closed-loop real-time low Earth orbit (LEO) SFF scenarios. The final VTFFTB infrastructure consists of three GNSS hardware signal simulators, three multi-constellation multi-band GNSS receivers, three navigation and control systems, an STK visualization system, and an ionospheric remote sensing system. A fleet of LEO satellites, each carrying a spaceborne GNSS receiver for navigation and ionospheric measurements, is simulated in scenarios with ionospheric impacts on the GPS and Galileo constellations. Space-based total electron density (TEC) and GNSS scintillation index S4 are measured by the LEO GNSS receivers in simulated scenarios. Four stages of work were accomplished to (i) build the VTFFTB with a global ionospheric modeling capability, and (ii) apply the VTFFTB to incubate future ionospheric measurement techniques. In stage 1, a differential-TEC method was developed to use space-based TEC measurements from a pair of LEO satellites to determine localized electron density (Ne). In stage 2, the GPS-based VTFFTB was extended to a multi-constellation version by adding the Galileo. Compared to using the GPS constellation only, using both GPS and Galileo constellations can improve ionospheric measurement quality (accuracy, precision, and availability) and relative navigation performance. Sensitivity studies found that Ne retrieval characteristics are correlated with LEO formation orbit, the particular GNSS receivers and constellation being used, as well as GNSS carrier-to-noise density C/N0. In stage 3, the VTFFTB for dual-satellite scenarios was further extended into a 3-satellite version, and then implemented to develop a polar orbit scenario with more fuel-efficient natural motion. In stage 4, a global 4-dimensioanl ionospheric model (TIE-CGM) was incorporated into the VTFFTB to significantly improve the modelling fidelity of multi-scale ionospheric space weather. Equatorial and polar space weather structures (e.g. plasma bubbles, tongues-of-ionization) were successfully simulated in 4-dimensional ionospheric scenarios on the enhanced VTFFTB. The dissertation has demonstrated the VTFFTB is a versatile GNSS-based SFF mission incubator to study ionospheric space weather impacts and develop next-generation multi-scale ionospheric observation missions. / Doctor of Philosophy / Spacecraft formation flying (SFF) is a space mission architecture with a group of spacecraft flying together and working as a team. SFF provides new opportunities for robust, flexible and low-cost observations of various phenomena in the ionized layer of Earth's atmosphere (called the ionosphere). Several hardware SFF simulation platforms based on Global Navigation Satellite Systems (GNSS) have been established to develop GNSS-based SFF missions, however, none of these platforms has ever directly used on-board GNSS receivers to study the impact of space weather on ionospheric density structures. The Virginia Tech Formation Flying Testbed (VTFFTB), a hardware simulation infrastructure using multiple GNSS signals, has been built in this work to emulate realistic SFF scenarios in low altitude orbits. The overall VTFFTB facility comprises three GNSS hardware signal emulators, three GNSS signal receivers, three navigation and control components, a software visualization component, and an ionospheric measurement component. Both Global-Positioning-System (GPS) and Galileo (the European version GNSS) are implemented in the VTFFTB. The objectives of this work are to (i) develop the VTFFTB with a high-fidelity ionospheric modeling capability, and (ii) apply the VTFFTB to incubate future ionospheric measurement techniques with GNSS receivers in space. A fleet of two or three spacecraft, each having a GNSS receiver to navigate and sense the ionosphere is emulated in several space environments. The electron concentration of the ionosphere and the GNSS signal fluctuation are measured by the GNSS receivers from space in simulated scenarios. These measurements are advantageous to study the location, size and structure of irregular ionospheric phenomena nearby the trajectory of spacecraft fleet. The culmination of this study is incorporation of an external global ionospheric model with temporal variations into the VTFFTB infrastructure to model a variety of realistic ionospheric structures and space weather impacts. Equatorial and polar space weather phenomenon were successfully simulated on the VTFFTB to verify a newly developed space-borne electron density measurement technique in the 3-dimensional ionosphere. Overall, it was successfully demonstrated that the VTFFTB is a versatile GNSS-based SFF mission incubator to study multiple kinds of ionospheric space weather impacts and develop next-generation space missions for ionospheric measurements.

Solar radiation pressure (SRP) is one of the major non-gravitational forces acting on spacecraft. Acceleration by radiation pressure depends on the radiation flux; on spacecraft shape, attitude, and mass; and on the optical properties of the spacecraft surfaces. Precise modeling of SRP is needed for dynamic satellite orbit determination, space mission design and control, and processing of data from space-based science instruments. During Earth penumbra transitions, sunlight is passing through Earth's lower atmosphere and, in the process, its path, intensity, spectral composition, and shape are significantly affected. This dissertation presents a new method for highly physical SRP modeling in Earth's penumbra called Solar radiation pressure with Oblateness and Lower Atmospheric Absorption, Refraction, and Scattering (SOLAARS). The fundamental geometry and approach mirrors past work, where the solar radiation field is modeled using a number of light rays, rather than treating the Sun as a single point source. This dissertation aims to clarify this approach, simplify its implementation, and model previously overlooked factors. The complex geometries involved in modeling penumbra solar radiation fields are described in a more intuitive and complete way to simplify implementation. Atmospheric effects due to solar radiation passing through the troposphere and stratosphere are modeled, and the results are tabulated to significantly reduce computational cost. SOLAARS includes new, more efficient and accurate approaches to modeling atmospheric effects which allow us to consider the spatial and temporal variability in lower atmospheric conditions. A new approach to modeling the influence of Earth's polar flattening draws on past work to provide a relatively simple but accurate method for this important effect. Previous penumbra SRP models tend to lie at two extremes of complexity and computational cost, and so the significant improvement in accuracy provided by the complex models has often been lost in the interest of convenience and efficiency. This dissertation presents a simple model which provides an accurate alternative to the full, high precision SOLAARS model with reduced complexity and computational cost. This simpler method is based on curve fitting to results of the full SOLAARS model and is called SOLAARS Curve Fit (SOLAARS-CF). Both the high precision SOLAARS model and the simpler SOLAARS-CF model are applied to the Gravity Recovery and Climate Experiment (GRACE) satellites. Modeling results are compared to the sub-nm/s^2 precision GRACE accelerometer data and the results of a traditional penumbra SRP model. These comparisons illustrate the improved accuracy of the SOLAARS and SOLAARS-CF models. A sensitivity analyses for the GRACE orbit illustrates the significance of various input parameters and features of the SOLAARS model on results. The SOLAARS-CF model is applied to a study of penumbra SRP and the Earth flyby anomaly. Beyond the value of its results to the scientific community, this study provides an application example where the computational efficiency of the simplified SOLAARS-CF model is necessary. The Earth flyby anomaly is an open question in orbit determination which has gone unsolved for over 20 years. This study quantifies the influence of penumbra SRP modeling errors on the observed anomalies from the Galileo, Cassini, and Rosetta Earth flybys. The results of this study prove that penumbra SRP is not an explanation for or significant contributor to the Earth flyby anomaly. / Ph. D.