Estimation of the total mission life of a spacecraft is an important issue for the communication satellite industries. For accurate determination of the remaining mission life of the satellite it is essential to estimate the amount of propellant present in the propellant tank of the spacecraft at various stages of its mission life. Because the annual revenue incurred from a typical commu-nication satellite operating at its full capacity is on the order of millions of dollars, premature removal of spacecraft from their orbits results in heavy losses. Various techniques such as the bookkeeping method, the gas law method, numerical modeling techniques, and use of capacitive sensors have been employed in the past for accurate determination of the amount of propellant present in a spacecraft. First half of the thesis is concerned with sensitivity analysis of the various propellant gauging techniques, that is, estimating the effects of the uncertainty in the instruments employed in the propellant gauging system on the onboard propellant estimation. This sensitivity analysis is done for three existing propellant gauging techniques – gas injection method, book-keeping method and the propellant tank heating method. A comparative study of the precision with which the onboard propellant is estimated by the three techniques is done and the primary source of uncertainty for all the three techniques is identified. It is illustrated that all the three methods — the gas injection method, the book-keeping method and the propellant tank heating method — are inherently indirect methods of propellant gauging, as a consequence of which, the precision with which the three techniques estimate the residual propellant decreases towards the end of mission life of the spacecraft. The second half of the thesis explores the possibility of using a new propellant tank configuration, consisting of a truncated cone centrally mounted within a spherical propellant tank, to measure the amount of liquid propellant present within the tank. The liquid propellant present within the propellant tank orients itself in a geometry, by virtue of its dominant surface tension force in zero-g condition, which minimizes its total surface energy. Study reveals that the amount of liquid propellant present in the tank can thus be estimated by measuring the height of the propellant meniscus within the central cone. It is also observed that, unlike gas law method, bookkeeping method or the propellant tank heating method, where the precision of the estimated propellant fill-fraction decreases towards the end-of-life of the spacecraft, for the proposed new configuration the precision increases.

Coulomb forces acting between close flying charged spacecraft provide near zero propellant relative motion control, albeit with added nonlinear coupling and limited controllability. This novel concept has numerous potential applications, but also many technical challenges. In this dissertation, two- and three-craft Coulomb formations near GEO are investigated, using a rotating Hill frame dynamical model, that includes Debye shielding and differential gravity. Aspects of dynamical systems theory and optimization are applied, for insights regarding stability, and how inherent nonlinear complexities may be beneficially exploited to maintain and maneuver these electrostatic formations. Periodic relative orbits of two spacecraft, enabled by open-loop charge functions, are derived for the first time. These represent a desired extension to more substantially studied, constant charge, static Coulomb formations. An integral of motion is derived for the Hill frame model, and then applied in eliminating otherwise plausible periodic solutions. Stability of orbit families are evaluated using Floquet theory, and asymptotic stability is shown unattainable analytically. Weak stability boundary dynamics arise upon adding Coulomb forces to the relative motion problem, and therefore invariant manifolds are considered, in part, to more efficiently realize formation shape changes. A methodology to formulate and solve two-craft static Coulomb formation reconfigurations, as parameter optimization problems with minimum inertial thrust, is demonstrated. Manifolds are sought to achieve discontinuous transfers, which are then differentially corrected using charge variations and impulsive thrusting. Two nonlinear programming algorithms, gradient and stochastic, are employed as solvers and their performances are compared. Necessary and sufficient existence criteria are derived for three-craft collinear Coulomb formations, and a stability analysis is performed for the resulting discrete equilibrium cases. Each specified configuration is enabled by non-unique charge values, and so a method to compute minimum power solutions is outlined. Certain equilibrium cases are proven maintainable using only charge control, and feedback stabilized simulations demonstrate this. Practical scenarios for extending the optimal reconfiguration method are also discussed. Lastly, particular Hill frame model trajectories are integrated in an inertial frame with primary perturbations and interpolated Debye length variations. This validates qualitative stability properties, reveals particular periodic solutions to exhibit nonlinear boundedness, and illustrates higher-fidelity solution accuracies. / text

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.

Spacecraft design is inherently difficult due to the nonlinearity of the systems involved as well as the expense of testing hardware in a realistic environment. The number and cost of flight tests can be reduced by performing extensive simulation and analysis work to understand vehicle operating limits and identify circumstances that lead to mission failure. A Monte Carlo simulation approach that varies a wide range of physical parameters is typically used to generate thousands of test cases. Currently, the data analysis process for a fully integrated spacecraft is mostly performed manually on a case-by-case basis, often requiring several analysts to write additional scripts in order to sort through the large data sets. There is no single method that can be used to identify these complex variable interactions in a reliable and timely manner as well as be applied to a wide range of flight dynamics problems. This dissertation investigates the feasibility of a unified, general approach to the process of analyzing flight dynamics Monte Carlo data. The main contribution of this work is the development of a systematic approach to finding and ranking the most influential variables and combinations of variables for a given system failure. Specifically, a practical and interactive analysis tool that uses tractable pattern recognition methods to automate the analysis process has been developed. The analysis tool has two main parts: the analysis of individual influential variables and the analysis of influential combinations of variables. This dissertation describes in detail the two main algorithms used: kernel density estimation and nearest neighbors. Both are non-parametric density estimation methods that are used to analyze hundreds of variables and combinations thereof to provide an analyst with insightful information about the potential cause for a specific system failure. Examples of dynamical systems analysis tasks using the tool are provided.

In this thesis, recently developed Blended Inverse (B-inverse) steering law is applied to two different redundant actuator systems. First, repeatability of Binverse is demonstrated on a redundant robotic manipulator. Its singularity avoidance and singularity transition performance is also demonstrated on the same actuator system. It is shown that B-inverse steering law provides singularity avoidance, singularity transition and repeatability. Second, its effectiveness is demonstrated for an Integrated Power and Attitude Control - Control Moment Gyroscope (IPAC-CMG) cluster, which can perform energy management and attitude control functions simultaneously. For this purpose, an IPAC-CMG flywheel is conceptually designed. A control policy is developed for the energy management.

The Square Kilometre Array (SKA) Project is an international effort to build the world’s largest radio telescope. It will be 100 times more sensitive than any other radio telescope currently in existence and will consist of thousands of dishes placed at baselines up to 3000 km. In addition to its increased sensitivity it will operate over a very wide frequency range (current specification is 100 MHz - 22 GHz) and will use frequency bands not primarily allocated to radio astronomy. Because of this the telescope needs to be located at a site with low levels of radio frequency interference (RFI). This implies a site that is remote and away from human activity. In bidding to host the SKA, South Africa was required to conduct an RFI survey at its proposed site for a period of 12 months. Apart from this core site, where more than half the SKA dishes may potentially be deployed, the measurement of remote sites in Southern Africa was also required. To conduct measurements at these sites, three mobile measurement systems were designed and built by the South African SKA Project. The design considerations, implementation and RFI measurements recorded during this campaign will be the focus for this dissertation.

The successful prediction of spacecraft motion is often heavily based upon assumptions used to simplify the problem without compromising solution accuracy. For many analyses, a primary assumption used is the decoupling of trajectory and attitude dynamics when calculating trajectories. In cases where spacecraft or objects have high area to mass ratios, non-conservative effects such as atmospheric drag and solar radiation pressure can greatly perturb spacecraft translational motion based on rotational state. A modular, six degree of freedom (6DOF) simulation with coupled orbit and attitude dynamics has been developed to model spacecraft and orbits of arbitrary geometries. First, the basis for the modular rotational and translational equations of motion are introduced. Next, formulations are provided for the gravity gradient torque, solar radiation pressure, aerodynamic, and non-spherical gravity potential sources of perturbations, and the Marshall Engineering Thermosphere atmospheric model used is described. A first test case is performed using the 6DOF simulation to simulate the deorbit of the spacecraft Lightsail 1, which flew in 2015. Next, predictive cases are demonstrated using the simulation for a theoretical sail-boom-rocket combination representative of a debris removal scenario, and for the Aerodynamic Deorbit Experiement, which will demonstrate a passively stable drag sail technology and characterize its effectiveness on orbit. All simulation cases have had aerodynamic perturbation formulations compared against high fidelity Direct Simulation Monte Carlo runs, and suggestions have been made for the future development of the simulation tool.

This thesis aims to design a reliable CubeSat platform, including the avionic subsystems that can sustain a high radiation environment for a mission having a lifetime of at least six months. The science instruments put stringent requirements on the platform to achieve and maintain the desired spin rate. The simulation background is set up in Systems Tool Kit (STK). A trade-off analysis for the Attitude and Orbit Control System (AOCS) of FORESAIL 2 was done, focusing on the actuators and their ability to offer the right amount of torque to fulfill the tether deployment. Mission design analyses were performed to conclude on the form factor of the CubeSat, its ability to generate power, its compliance with the Space Debris Mitigation (SDM) technical requirements, and the total radiation dose accumulated. It was found that a 6U form factor is preferred to allocate more space for each subsystem, alongside with generating enough power for the satellite to work in all modes wanted. The mission is compliant with European Cooperation for Space Standardization (ECSS) and International Organization for Standardization (ISO) standards if the CubeSat is to be launched in September 2022. To allow a threshold limit of 10 krads on the components of the satellite, a shielding wall of 7 mm should be implemented on the CubeSat’s structure. Major requirements for the designed mission were written to initialize the investigation on the sensors and actuators. The results showed that only a propulsion system provided the necessary angular momentum to deploy the tether. The lack of magnetic field makes magnetorquers almost unusable in the desired orbit, leaving reaction wheels as the only option remaining to assist the propulsion units. The different analyses and simulations led to a final AOCS configuration composed of five various sensors (Sun sensors, magnetometers, a GPS, an IMU, and housekeeping sensors) for the attitude determination. A propulsion system and reaction wheels will provide the necessary control over the satellite.

abstract: It remains unquestionable that space-based technology is an indispensable component of modern daily lives. Success or failure of space missions is largely contingent upon the complex system analysis and design methodologies exerted in converting the initial idea into an elaborate functioning enterprise. It is for this reason that this dissertation seeks to contribute towards the search for simpler, efficacious and more reliable methodologies and tools that accurately model and analyze space systems dynamics. Inopportunely, despite the inimical physical hazards, space systems must endure a perturbing dynamical environment that persistently disorients spacecraft attitude, dislodges spacecraft from their designated orbital locations and compels spacecraft to follow undesired orbital trajectories. The ensuing dynamics’ analytical models are complexly structured, consisting of parametrically excited nonlinear systems with external periodic excitations–whose analysis and control is not a trivial task. Therefore, this dissertation’s objective is to overcome the limitations of traditional approaches (averaging and perturbation, linearization) commonly used to analyze and control such dynamics; and, further obtain more accurate closed-form analytical solutions in a lucid and broadly applicable manner. This dissertation hence implements a multi-faceted methodology that relies on Floquet theory, invariant center manifold reduction and normal forms simplification. At the heart of this approach is an intuitive system state augmentation technique that transforms non-autonomous nonlinear systems into autonomous ones. Two fitting representative types of space systems dynamics are investigated; i) attitude motion of a gravity gradient stabilized spacecraft in an eccentric orbit, ii) spacecraft motion in the vicinity of irregularly shaped small bodies. This investigation demonstrates how to analyze the motion stability, chaos, periodicity and resonance. Further, versal deformation of the normal forms scrutinizes the bifurcation behavior of the gravity gradient stabilized attitude motion. Control laws developed on transformed, more tractable analytical models show that; unlike linear control laws, nonlinear control strategies such as sliding mode control and bifurcation control stabilize the intricate, unwieldy astrodynamics. The pitch attitude dynamics are stabilized; and, a regular periodic orbit realized in the vicinity of small irregularly shaped bodies. Importantly, the outcomes obtained are unconventionally realized as closed-form analytical solutions obtained via the comprehensive approach introduced by this dissertation. / Dissertation/Thesis / Doctoral Dissertation Systems Engineering 2019

The investment of space commerce is skyrocketing and it is predicted to be a nascent business in the future. The spacecraft demand has been growing not only for NASA and other space agency’s mission but also collaboration business between small space industries, academia, and scientific community. This glimpse brought an interest to a new investor, government, military, and manufacturing company to deliver their objectives efficiently. Nowadays, many startups compete embracing innovation and pioneering the novelty of space project beyond prodigious vision in an unprecedented way. Many players foresee that decreasing size of the rocket is an important key to survive and succeed in the space business. One of the efficient acts is lowering the launch cost. This can be achieved by designing a small size, lightweight and affordable spacecraft. Within this context, a Beyond Atlas Spacecraft which will be sent to Asteroid 2016-HO3, has achieved a wet mass of 20.85 kg with the size of 24.7 x 42.2 x 40.8 cm in stowed mode and 84 x 399 x 40.8 cm in unstowed mode. However, the drawback being light and small may lead to catastrophic failure due to resonance frequency events. According to past experience, the gyro of the Swedish national satellite was damaged during ground testing and it was suspected due to high amplification when the natural frequency coincides to the main structure resonance. Therefore, this work is focusing on a spacecraft development and a non-destructive structural analysis. The coupled-load analysis of a preliminary spacecraft design including sinusoidal, random vibration and shock analysis are calculated using FEM. This effort can reduce the risk of component destruction before laboratory testing as well as understand better the dynamic behavior of the spacecraft. The critical frequency in each orthogonal axis with base input from launch environment of the LM-3A Launch Vehicle was devised. The maximum stress, amplitude, and acceleration in accordance of qualification test criteria were evaluated and discussed. / Investeringen av rymdhandeln är skyrocketing och det förväntas bli en växande verksamhet i framtiden. Efterfrågan på rymdfarkoster har ökat inte bara för NASA och andra rymdorganisationens uppdrag utan även samarbete mellan små rymdindustrier, akademin och det vetenskapliga samfundet. Denna glimt väckte intresse för en ny investerare, regering, militär och tillverkningsföretag för att effektivt kunna leverera sina mål. Idag konkurrerar många startups om att omfatta innovation och banbrytande rymdprojektets nyhet bortom en fördärvad vision på ett aldrig tidigare skådat sätt. Många spelare förutser att minskad storlek på raketen är en viktig nyckel för att överleva och lyckas i rymdverksamheten. En av de effektiva handlingarna sänker lanseringskostnaden. Detta kan uppnås genom att utforma en liten storlek, lätt och prisvärd rymdfarkost. Inom detta sammanhang har en Beyond Atlas Spacecraft som skickas till Asteroid 2016-HO3, uppnått en våt massa på 20,85 kg med storleken 24,7 x 42,2 x 40,8 cm i stuvningsläge och 84 x 399 x 40,8 cm i ostoppat läge. Nackdelen som är ljus och liten kan emellertid leda till katastrofalt fel på grund av resonansfrekvenshändelser. Enligt tidigare erfarenhet skadades gyroen i den svenska nationella satelliten under marktestning och det misstänktes på grund av hög förstärkning när den naturliga frekvensen sammanföll med huvudstrukturen resonans. Därför fokuserar detta arbete på rymdskeppsutveckling och en icke-destruktiv strukturanalys. Den kombinerade belastningsanalysen av en preliminär rymdfarkostkonstruktion inklusive sinusformad, slumpvibration och chockanalys beräknas med användning av FEM. Denna insats kan minska risken för komponent förstörelse före laboratorietestning samt förstå bättre rymdskeppets dynamiska beteende. Den kritiska frekvensen i varje ortogonal axel med basinmatning från startmiljön för LM-3A-startkärlet utformades. Den maximala spänningen, amplituden och accelerationen i enlighet med kvalifikationstestkriterierna utvärderades och diskuterades.