Preliminary Electrical Designs for CTEx and AFIT Satellite Ground Station

This thesis outlines the design of the electrical components for the space-based ChromoTomography Experiment (CTEx). CTEx is the next step in the development of high-speed chromotomography at the Air Force Institute of Technology. The electrical design of the system is challenging due to the large amount of data that is acquired by the imager and the limited resources that is inherent with space-based systems. Additional complication to the design is the need to know the angle of a spinning prism that is in the field of view very precisely for each image. Without this precise measurement any scene that is reconstructed from the data will be blurry and incomprehensible. This thesis also outlines how the control software for the CTEx space system should be created. The software ow is a balance of complex real time target pointing angles and simplicity to allow the system to function as quick as possible. This thesis also discusses the preliminary design for an AFIT satellite ground station based upon the design of the United States Air Force Academy\u27s ground station. The AFIT ground station will be capable of commanding and controlling satellites produced by USAFA and satellites produced by a burgeoning small satellite program at AFIT

Design of a micro-Pulsed Plasma Thruster for a 3U Cubesat

This project presents the conceptual design of a micro-Pulsed Plasma Thruster (microPPT) and a propulsion module consisting of eight micro-PPTs for applications on a three-unit Cube Satellite (3U CubeSat). The computer aided mechanical design of the micro-PPT, the propulsion module and its integration in the 1U are presented. Material selection identifies Torlon 4203 for housing and Tungsten-plated copper for electrodes. Finite element stress analysis shows that the micro-PPT will sustain expected launch loads. The electric circuit for the main and initial discharge are designed and simulated for power and voltage levels expected in a 3U Cubesat. Finite element thermal analysis provides estimates of the temperature distribution in the micro-PPT and module for a typical discharge

Air Force Institute of Technology Research Report 2014

This report summarizes the research activities of the Air Force Institute of Technology’s Graduate School of Engineering and Management. It describes research interests and faculty expertise; lists student theses/dissertations; identifies research sponsors and contributions; and outlines the procedures for contacting the school. Included in the report are: faculty publications, conference presentations, consultations, and funded research projects. Research was conducted in the areas of Aeronautical and Astronautical Engineering, Electrical Engineering and Electro-Optics, Computer Engineering and Computer Science, Systems Engineering and Management, Operational Sciences, Mathematics, Statistics and Engineering Physics

Dynamic neural network-based pulsed plasma thruster (PPT) fault detection and isolation for the attitude control subsystem of formation flying satellites

The main objective of this thesis is to develop a dynamic neural network-based fault detection and isolation (FDI) scheme for the Pulsed Plasma Thrusters (PPTs) that are used in the Attitude Control Subsystem (ACS) of satellites that are tasked to perform a formation flying mission. In order to accomplish these objectives three fault detection and isolation (FDI) approaches based on dynamic neural networks (DNN) are developed: (i) a "Low Level" FDI scheme, (ii) a "High Level" FDI scheme, and (iii) an "Integrated" FDI scheme. Based on data that is collected from the electrical circuit of the PPTs, our proposed "Low Level" FDI scheme can detect and isolate faults in the PPT actuators. Using a Confusion Matrix evaluation system we demonstrate that can achieve a high level of accuracy but the precision level is below expectations and the misclassification rate is expressed as False Healthy and False Faulty parameters is significant. Our proposed "High Level" FDI scheme utilizes data collected from the relative attitudes of the formation flying satellites. According to the simulation results, our proposed FDI scheme can detect the pair of thrusters which is faulty. It represents a promising detection capability, however its isolation capabilities are not adequate. Finally, our proposed "Integrated" FDI scheme takes advantage of the strengths of each of the previous schemes and at same time reduces their individual weaknesses. To demonstrate its capabilities, various fault scenarios were simulated. The results demonstrate a high level of accuracy (99.79%) and precision (99.94%) with a misclassification rate that is quite negligible. Furthermore, our proposed "Integrated" FDI scheme provides additional and interesting information related to the effects of faults in the thrust production levels that would not be available from simply the low and high levels separately

Beamed-Energy Propulsion (BEP) Study

The scope of this study was to (1) review and analyze the state-of-art in beamed-energy propulsion (BEP) by identifying potential game-changing applications, (2) formulate a roadmap of technology development, and (3) identify key near-term technology demonstrations to rapidly advance elements of BEP technology to Technology Readiness Level (TRL) 6. The two major areas of interest were launching payloads and space propulsion. More generally, the study was requested and structured to address basic mission feasibility. The attraction of beamed-energy propulsion (BEP) is the potential for high specific impulse while removing the power-generation mass. The rapid advancements in high-energy beamed-power systems and optics over the past 20 years warranted a fresh look at the technology. For launching payloads, the study concluded that using BEP to propel vehicles into space is technically feasible if a commitment to develop new technologies and large investments can be made over long periods of time. From a commercial competitive standpoint, if an advantage of beamed energy for Earth-to-orbit (ETO) is to be found, it will rest with smaller, frequently launched payloads. For space propulsion, the study concluded that using beamed energy to propel vehicles from low Earth orbit to geosynchronous Earth orbit (LEO-GEO) and into deep space is definitely feasible and showed distinct advantages and greater potential over current propulsion technologies. However, this conclusion also assumes that upfront infrastructure investments and commitments to critical technologies will be made over long periods of time. The chief issue, similar to that for payloads, is high infrastructure costs

Development of High-Density Propulsion System Technologies for Interplanetary Small Satellites and CubeSats

The goal of this research was to support the development of a novel propulsion system for small satellites (\u3c180 kg) and CubeSats. This was pursued by conducting a collection of studies that were designed to provide engineering data that would be critical in designing a functional prototype. The novel propulsion system was conceived by the author to provide best-in-class performance for the small satellite and CubeSat families of spacecraft. This context presents specific design requirements that the presented technology attempts to satisfy. The most critical among these is high density; the propellant was designed to be stored with high density and the thruster was designed to be as compact as possible. The propulsion system is composed of two primary elements, a propellant generator and a thruster. The propellant generator works by sublimating a solid crystal into vapor and then using this vapor to etch a dense metal. The resulting gaseous byproducts of this reaction are the propellant. This dissertation used xenon difluoride (XeF2) vapor to etch tungsten (W) which react to form xenon gas (Xe) and tungsten hexafluoride (WF6). This approach gave a theoretical propellant storage density 5.40 g/cm3; and 5.17 g/cm3 was demonstrated. The sublimation dynamics of the XeF2 were studied as a function of surface area and temperature and it was found to be suitable for the intended application due to its high effluence rate; that is, it sublimates fast enough to be useful. The sublimation rates are on the order of 10’s of µg/s. The etch rate of XeF2 on W was also studied and found to be suitably fast to provide useful amounts of reactants for use as a propellant, again on the order of 1’s of µg/s. The thruster is an electrostatic radio frequency (RF) ion thruster design and is manufactured with Low Temperature Co-Fired Ceramic (LTCC) materials system and manufacturing technology. Manufacturing samples of the thruster were built at the University of Arkansas in July 2015 and tested at NASA’s Marshall Space Flight Center in May 2018. Testing validated the viability of the LTCC thruster and provided valuable information on how to improve the thruster’s design

Performance Characterization of a Low Power Magnetic Nozzle

The thrust and efficiency performance of a low-power magnetic nozzle test article is analytically and experimentally investigated. In the last two decades the demand for new forms of in-space propulsion for small spacecraft has increased interest in low-power (< 200 W) magnetic nozzle thrusters. The inherent advantages of these devices, including the electrodeless design and the potential to be propellant-agnostic, coupled with the potential to efficiently accelerate the propellant makes low-power magnetic nozzles attractive propulsion options for small satellites. However, the measured performance of both the low and moderate power versions of these thrusters has compared unfavorably to existing state-of-the-art propulsion technologies. A theoretical model was developed to predict low-power magnetic nozzle performance and identify fundamental differences in operation between these devices and their higher power counterparts. An experiment was designed to inform the theoretical model and to provide insight into the fundamental dynamics of plasma flowing through a low-power magnetic nozzle. This test article consisted of a reconfigurable inductively-coupled plasma source and an electromagnet. A suite of electrostatic probes and laser induced fluorescence is used to measure the plasma properties throughout the nozzle and map the plasma structures present in the plume. Using the experimental measurements, it is found that the plasma expansion follows a polytropic law, as predicted in the literature. The observed increase in the ion velocity confirms that the test article accelerates the propellant. By coupling the experimental results with the theoretical framework, two novel effects that reduce device performance are identified: 1) a low ion fraction, and corresponding neutral-collisional effects, impedes ion acceleration and shifts the nozzle throat downstream, and 2) non-uniform power deposition enhances the plasma density adjacent to the liner wall, resulting in degraded source (the ratio of power flowing into the diverging nozzle section to the total power deposited in the plasma) and divergence efficiency (the fraction of the kinetic energy in the thrust direction). These effects arise from the low input power and the thruster design parameters. Experimental characterization of a reconfigured test article demonstrates that performance can be recovered by accounting for these two effects when designing the thruster and selecting the operating parameters.PHDAerospace EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttps://deepblue.lib.umich.edu/bitstream/2027.42/149990/1/collardt_1.pd

Pulsed plasma thrusters for small satellites

Since the Russian launch of the Zond-2 satellite in 1964 there have been over fifty years of research dedicated to the understanding of the first electric propulsion device to be flown in space, the Pulsed Plasma Thruster. The Pulsed Plasma Thruster originates from the evolution of the vacuum arc switch, but due to its microsecond operation time, the internal dynamics and nature of operation have remained unclear. The Pulsed Plasma Thruster is generally cheap to manufacture and to operate, which keeps it a popular device to research within institutes worldwide and has contributed to its longevity. As a satellite propulsion device it has unique capabilities that other propulsion systems cannot provide. The thruster operates by accelerating plasma formed in the accelerating electrodes (or nozzle) in short discrete packets of thrust or impulse. The pulsed nature of the thruster means that between pulses energy can be stored in capacitors, ready for the next discharge. The storage of energy over time means the power draw is variable and is only dependant on the frequency that the system is pulsed at. This property of the thruster makes the Pulsed Plasma Thruster extremely versatile, allowing the thruster to perform both velocity correction and control manoeuvres and attitude control manoeuvres. The Pulsed Plasma Thruster is mechanically scalable but the performance of the thruster has been shown to depend linearly on the energy storage ability of the thruster’s capacitor. The work presented here covers two areas. Firstly is the critical analysis of the physical mechanisms that occur within a Pulsed Plasma Thruster through a review of literature, experimentation and the development of a high current plasma flow model. The second area is the design, development, manufacture and evaluation of the Pulsed Plasma Thruster for use on a nanosatellite platform known as the CubeSat. Several novel observations and contributions were made during the critical analysis of the physical mechanisms of the Pulsed Plasma Thruster. The most significant was realising how the erosion of the metal electrodes affected the overall discharge process. It is postulated that the expulsion of material from emission sites (or cathode spots), the ionisation of that material and the resulting freed electrons, create a pinched plasma column between the electrodes. It is postulated that the interaction of the electrode sheath region and the intersecting plasma column cause the current flow to become limited. This was then shown to affect the efficiency with which the stored energy of the capacitor was converted to energy to accelerate the plasma. Understanding this issue is key in improving future designs of the Pulsed Plasma Thruster. The observations and conclusions made during this work were put into practice to create an eight μPPT propulsion module for a 3U CubeSat. Initial results show that a μPPT with a specific impulse of 321s, an impulse bit of 0.56μNs and a mass bit of 0.17μg has been developed. The thruster was developed for two technology demonstration CubeSats. STRaND-1 is a joint collaboration between Surrey Space Centre and Surrey Satellite Technology Limited and UKUBE-1 is a joint collaboration between Surrey Space Centre and the UK Space Agency. Both CubeSats are scheduled for launch late 2011, early 2012. The propulsion module for the STRaND-1 CubeSat will be the first to provide full axis control and the first to provide electric propulsion on this class of satellite, showing the advantages of the Pulsed Plasma Thruster for Small Satellites.EThOS - Electronic Theses Online ServiceGBUnited Kingdo

GSFC Heliophysics Science Division FY2010 Annual Report

This report is intended to record and communicate to our colleagues, stakeholders, and the public at large about heliophysics scientific and flight program achievements and milestones for 2010, for which NASA Goddard Space Flight Center's Heliophysics Science Division (HSD) made important contributions. HSD comprises approximately 323 scientists, technologists, and administrative personnel dedicated to the goal of advancing our knowledge and understanding of the Sun and the wide variety of domains that its variability influences. Our activities include: Leading science investigations involving flight hardware, theory, and data analysis and modeling that will answer the strategic questions posed in the Heliophysics Roadmap; Leading the development of new solar and space physics mission concepts and support their implementation as Project Scientists; Providing access to measurements from the Heliophysics Great Observatory through our Science Information Systems; and Communicating science results to the public and inspiring the next generation of scientists and explorers

Design and implementation of small satellite inspection

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2012.Cataloged from PDF version of thesis.Includes bibliographical references (p. 179-181).For a variety of missions, vision-based navigation and similar architectures provide the advantage of detailed measurements for a fraction of the size and complexity of ground-based imagers. This thesis provides a simple navigation algorithm using no more than a visual centroid measurement to enable in-situ inspection of space objects. This work evaluates those inspection maneuvers using the Synchronize Position Hold Engage Reorient Experimental Satellites, known as SPHERES. Evaluation of hardware performance was done using data from the International Space Station, in concert with ground-based simulations. Ultimately, this work is in preparation for future experimentation using the VERTIGO vision-navigation payload for SPHERES. The first step presented is an analysis of the measurement capabilities of the SPHERES system and the predicted performance of the VERTIGO system. Using this analysis it is shown that tests run using the former system are applicable to the latter in terms of accuracy, precision, and observability. The second step is an analysis of the tests run on the Space Station, a comparison to those predicted by simulation, and an extension of those results to simulations of more complex maneuvers. Further, a determination of the robustness of the control to disturbances is also performed. Finally, this thesis reflects on the technical and programmatic challenges of developing the VERTIGO payload. From these challenges, lessons are drawn which may guide future developers and program managers, particularly in the university engineering environment.by Michael Christopher O'Connor.S.M