The COLISEUM framework is a tool designed for electric propulsion plume interactions. Virginia Tech has been developing a module for COLISEUM called DRACO, a particle-in-cell based code capable of plume modeling for geometrically complex spacecraft. This work integrates a charging module into DRACO. Charge is collected via particle impingement on the spacecraft surface and converted to potential. Charge can be stored in the surface, or added to a local ground potential. Current can flow through the surface and is governed by the internal electric field in the spacecraft. Several test cases were run to demonstrate the code's capabilities. The first was a plume impingement of a composite spherical probe by a xenon thruster. It was shown that the majority of current conducted will reach the interior of the spacecraft, not other surface elements. A conductive interior will therefore result in a uniform surface potential, even for low surface conductivities. A second test case showed a composite spacecraft exposed to a solar wind. This test showed that when a potential gradient is applied to a conductive body, the ground potential of the spacecraft will lower significantly to compensate and maintain a zero net current. The case that used the semiconductive material showed that the effect of the potential gradient can be lowered in cases such as this, where some charge will always be stuck in the surface. If a dielectric material is used, the gradient will disappear altogether. The final test case showed the effect of charge exchange ion backflow on the potential of a spacecraft similar to the DAWN spacecraft. This case showed that CEX ion distribution is not even along the spacecraft and will result in a transverse potential gradient along the panel. / Master of Science

Electrodynamic tethers have a wide range of proposed applications in the fields of satellite propulsion and space plasma research. The fundamental purpose of this dissertation is to improve the understanding of the behavior of an electrodynamic tether (EDT) system in Earth's ionosphere. An electrodynamic tether system consists of two satellites connected by a long tether that generates current to produce either power or thrust via the system's electromagnetic interaction with the space environment. Previous electrodynamic tether investigations decouple the interaction between the tether and the constantly changing plasma environment. The limiting factor inhibiting the development of a full system model that has an accurate characterization of the tether/plasma interaction is that the understanding of that interaction is not well developed over a wide range of system parameters. The EDT system model developed in this study uses a high fidelity dynamics model that includes a tether current described by an analytical current collection model whose plasma parameters are determine by the International Reference Ionosphere. It is first shown that new instabilities are induced in the system dynamics under a basic analytical current model versus a constant current model. A 2-D3$v$ Particle-in-Cell (PIC) code has been developed to study the plasma dynamics near a positively charged EDT system end-body and their impact on the current collected. Simulations are run over a range of system parameters that occur throughout a LEO orbit. The azimuthal current structures observed during the TSS-1R mission are found to enhance the current collected by the satellite when the magnetic field is slightly off of perpendicular to the orbital velocity. When the in-plane component of the magnetic field becomes large, the electrons are not able to easily cross the field lines causing plasma lobes form above and below the satellite. The lobes limit the current arriving to the satellite and also cause an enhanced wake to develop. A high satellite bias causes a stable bow-shock structure to form in the ram region of the satellite, which limits the number of electrons entering the sheath region and thus limiting the current collected. Electron-neutral collisions are found to destabilize the bow-shock structure and remove its current limiting effects. Additionally, as the magnetization of the plasma is increased, the current becomes limited by the charged particle's inability to cross magnetic field lines. Analytical curve fits based on the simulation results are presented that characterize the dependence of the average current collected on the local magnetic field orientations, space plasma magnetization and satellite potential. The results from the PIC simulations characterizing the magnetic field's influence on the tether's current are incorporated into the system dynamics model to study the behavior of the EDT system over a range of inclinations. The magnetic field is found to limit the diurnal variations in the current collected by the system throughout its orbit. As the inclination of the system's orbit is increased, the impact of the magnetic field becomes more pronounced as its orientation sweeps through a larger range of angles. The impact of the magnetic field on the collected current is, therefore, found to limit the ability of an EDT system to boost the system's orbit as the orbit's inclination is increased. In summary, new system dynamics have been observed due to the previously unobserved behavior of the current over a range of end-body configurations. / Ph. D.

Secondary electron emission (SEE) and incident-particle backscattering are important processes accompanying the impact of energetic electrons and ions on surfaces. The phenomena play a key role in the buildup of electrical charge on spacecraft surfaces, and are therefore of particular interest to scientists attempting to model spacecraft charging. In response to a demonstrated need for data, techniques for determining total secondary electron (SE) and backscatter (BS) yields (del) and (neu), and associated scattering-angle-resolved,scattering-energy-resolved, and simultaneous angle-energy-resolved yields have been developed. Further, an apparatus capable of making the necessary measurements for experimental determination of these quantities---for conducting materials in an ultra-high vacuum environment-has been designed, constructed, and partially tested. The apparatus is found to be in working order, though in need of fine-tuning, and the measurement technique successful. Investigations using a 1-3 Kev beam of monoenergetic electrons normally incident on bulk AI have been undertaken with the new apparatus. Electron-stimulated desorption of surface contaminants has been observed, as has been beam-induced carbon deposition, and an empirical model describing the resulting dynamic evolution of (del)is presented. Totalb and 11 values obtained in the present investigation are found to be in qualitative agreement with the results of previously reported investigations, though quantitative disagreement of b-values is substantial. Specifically, evidence is presented suggesting that previously reported SE yields for clean AI under electron bombardment (in the 1-3 Kev energy range) are in error by as much as 30 %.

京都大学 / 0048 / 新制・課程博士 / 博士(工学) / 甲第20375号 / 工博第4312号 / 新制||工||1668(附属図書館) / 京都大学大学院工学研究科電気工学専攻 / (主査)教授 山川 宏, 教授 引原 隆士, 准教授 海老原 祐輔 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DFAM

京都大学 / 0048 / 新制・課程博士 / 博士(工学) / 甲第21069号 / 工博第4433号 / 新制||工||1689(附属図書館) / 京都大学大学院工学研究科電気工学専攻 / (主査)教授 山川 宏, 教授 松尾 哲司, 准教授 海老原 祐輔 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DFAM

Les satellites en orbite terrestre se déplacent dans le plasma ionosphérique, un mélange de particules chargées, et éventuellement de particules neutres. Des électrons et des ions issus de ce plasma, ainsi que les émissions Ultra-Violets(UV) en provenance du soleil, interagissent avec les surfaces du satellite et modifient sa charge électrostatique. Cette chargement peut induire elle-même des décharges électrostatiques aux conséquences allant de perturbations électromagnétiques (fausses commandes par exemple) à la perte du satellite. En orbites de basse altitude (LEO) l'énergie cinétique et thermique du plasma est généralement faible et donc, les satellites vont rarement présenter des décharges importantes. Néanmoins, les missions scientifiques qui embarquent des instruments très performants et précis peuvent être affectées par cette interaction satellite-plasma-émissions UV. Cette thèse s'intéresse particulièrement à ces phénomènes de charge des structures externes du satellite et à l'impact de ce chargement sur les mesures scientifiques effectuées à bord, i.e. mesures du champ électrique et de la densité du plasma thermique. / Earth-orbiting satellites travel in ionospheric plasma, a mixture of charged particles, and possibly neutral particles. Electrons and ions from this plasma, as well as Ultra-Violet (UV) emissions from the sun, interact with the surfaces of the satellite and modify its electrostatic charge. This loading can itself induce electrostatic discharges to the consequences ranging from electromagnetic disturbances (false commands for example) to the loss of the satellite. In low-Earth orbits (LEO), the kinetic and thermal energy of the plasma is generally low and therefore satellites rarely exhibit large discharges. Nevertheless, scientific missions that carry high-performance and accurate instruments can be affected by this satellite-plasma-UV-emissions interaction. This thesis is particularly interested in these phenomena of charge of the external structures of the satellite and the impact of this load on the scientific measurements carried out on board, i.e. measures of the electric field and the density of the thermal plasma.

As an orbiter cruising to Jupiter will encounter different plasma environments, variety of spacecraft surface charging is expected. This surface potential can lead to inaccurate and wrong in-situ plasma measurements of on-board sensors, which explain the interest in simulating the charging.In this thesis the spacecraft-plasma interactions for a future mission to Jupiter are modelled with the help of the Spacecraft Plasma Interaction System, taking the case of a Jupiter Ganymede Orbiter (JGO) and a Jupiter Europa Orbiter (JEO) as an archetype for a future mission.It is shown that in solar wind at Earth and Jupiter, spacecraft potentials of about 8 V for the JEO, and 10 V to 11 V for the JGO are expected. Furthermore, at a distance of 15 Jupiter radii from Jupiter, the JGO is expected to charge to an electric potential of 2 V, except in the planetary shadow, where it will charge to a high negative potential of -40 V. Moreover, close to the orbit of Callisto, JGO will charge to 12 V in the sun and to 4.6 V in eclipse, due to a high secondary electron emission yield. / <p>Validerat; 20120115 (anonymous)</p>

In July 2031, a new chapter in the exploration of the Jovian system will begin with the arrival of the Jupiter Icy Moons Explorer (Juice) at Jupiter. Launched on April 14 2024 as part of ESA’s Cosmic Vision programme, the mission aims to study Jupiter and its icy Galilean moons Callisto, Europa, and Ganymede. Juice carries a whole suite of instruments for in-situ and remote ground observations, one of which is the Jovian plasma Dynamics and Composition analyser (JDC). As a part of the Particle Environment Package (PEP), the particle detector will measure the energy, mass, charge and arrival direction of ions and electrons in the Jovian magnetosphere. Spacecraft charging caused by interactions between the spacecraft and its surrounding plasma environment poses a significant problem for JDC because the electrostatic potential of the spacecraft accelerates/decelerates charged particles, resulting in distorted measurements, particularly for the lowest energy particles.  In this report, we show the results of spacecraft charging and instrument simulations performed in the Spacecraft Plasma Interaction System (SPIS) for the solar wind and Earth’s magnetosheath—two environments that Juice will encounter at the start of the cruise phase. We found that the conductive surfaces that cover the majority of the spacecraft become positively charged as a result of a large photoelectron current in both the solar wind and magnetosheath environments. We show that these surfaces are expected to reach potentials of 9 V in the solar wind and 4 V in the magnetosheath. The four radiators on Juice that are covered with dielectric paint and shaded by the sun shield become negatively charged in both simulated environments. The radiator potentials can be as low as -40 V in the solar wind and -100 V in the magnetosheath. We also conclude that due to blocking by the spacecraft main body, the ion population cannot be sampled in the solar wind unless a spacecraft roll is performed. Furthermore, due to the high ion f low energy, spacecraft charging will not influence JDC measurements in this environment.  In the magnetosheath, the ion population can be sampled by JDC, and we identified three distortion mechanisms: (1) repulsion by the main body, (2) attraction by two of the radiators, and (3) repulsion by the MAG boom. Of all the distortion modes, the one originating from a negatively charged (-67.8 V) radiator close to JDC is the strongest, affecting ions with energies above 80 eV. The least powerful but most prevalent mode is the repulsion of ions by the main body. Our results can be compared with future in-situ measurements to identify distortion mechanisms well ahead of the science phase in which the scientifically important measurements will be carried out.