Thesis (M.S. in Space Systems Operations)--Naval Postgraduate School, September 2002. / Thesis advisor(s): Stephen A. Whitmore, Michael G. Spencer. Includes bibliographical references (p. 73-74). Also available online.

Mon travail de recherche a porté surl’étude de la réponse magnétique orbitale d’un gazd’électrons dans le potentiel cristallin d’un solide.Cette étude est dans la continuité du travail deLandau (diamagnétisme de Landau) et de Peierls,ce dernier ayant développé une formule de susceptibilitéorbitale valable pour les modèles cristallinsà une bande. L’objet de ma thèse a été degénéraliser cette formule à un nombre quelconquede bandes, travail réalisé à l’aide d’une théorie deperturbation invariante de jauge. Je me concentreparticulièrement sur l’étude des modèles à deuxbandes, afin de mettre en évidence l’importancedes couplages interbandes dans la susceptibilitéorbitale. Cela fait intervenir la courbure de Berry,grandeur classiquement associée à ces couplages,mais également le tenseur métrique qui joue unrôle crucial. En particulier, je montre qu’un isolantde bandes peut avoir une réponse magnétiquemême si sa bande de valence est remplie, et je metsen évidence un modèle où les propriétés géométriquespeuvent être variées tout en maintenantfixes les propriétés spectrales, ce qui induit desmodifications importantes de la réponse magnétique. / My research project has been to studythe orbital magnetic response of a electron gas inthe periodic potential of a crystal. Its purpose isto generalize Landau’s diamagnetism and Peierls’formula for one-band crystals. The main goal wasto generalize Peierls’ work to any number of bands.Then, I applied the obtained formula to 2-bandsystems in order to highlight the role of interbandeffects in the orbital susceptibility. The susceptibilitycan be written using Berry curvature, quantityassociated to interband effects, as well as the metrictensor. In particular, I show that a band isulatorcan have a non-vanishing magnetic responseeven if the chemical potential lies in the gap. Moreover,I study a model where the geometric propertiescan be tuned without changing the dispersionrelation. This tuning can drastically modifythe orbital magnetic response.

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A simple theoretical method for calculating electronic band structures of wurtzite materials based on the bond-orbital models is presented. This method can be used to study many problems such as band mixing and effects of external fields (electric field, magnetic field, and unaxial stress, etc.), since it can reproduce fairly accurate lowest conduction-band and top three valence-band structures. This method is very similar to LCAO method; however, it is much simpler and requires less computational effort than LCAO method.

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In this thesis, we investigate the orbit control strategies of small satellites in Low Earth Orbits (LEO) where the disturbance effects are significant, in particular the nonspherical Earth and atmospheric drag effects. These orbits are not suitable to be controlled by using traditional ground-based control strategies which generally require high-thrust propulsion systems and other expensive resources, both onboard and in the ground segment. In order to react to those disturbances spontaneously and keep a small satellite at a pre-defined station using its limited resources, autonomous orbit control technology needs to be enabled. With the current advances in navigation and propulsion technology, as well as onboard computation systems, the only key issue that needs further investigations for practical implementation of an autonomous orbit operation system is the control algorithm. The orbit control strategies we investigate here are treated separately for each of the orbital control phases, i.e. orbit deployment and acquisition, orbit transfer and orbit maintenance. We present various forms of the solutions of the epicycle motion which allow us to treat each control problem according to the control requirements, nature of perturbations, control time scales and available resources. Although applied in different manners, the optimal low-thrust control scheme is a common aim for all control problems investigated here, as we mainly focus upon applications for low cost small satellites in LEO. The verifications of the strategies proposed in this thesis have been demonstrated not only via computer simulations, but also successfully demonstrated on in-orbit small satellite platforms thanks to an active small satellite programme at Surrey Space Centre. The success of this study is hoped to provide a valuable basis for satellite orbit operations which will involve larger number of satellites with more complex configurations in the future.

In this research study, several new in-orbit algorithms are proposed to improve the performance of Attitude Determination and Control System (ADCS) by estimating the inertia matrix and calibrating the cold gas thruster system of the UoSAT-12 spacecraft. Computer-based simulation models will be constructed using MATLAB and SIMULINK in order to evaluate the expected performance. The first focus is on the identification of the satellite inertia matrix. A new algorithm based on a Recursive Least Square (RLS) estimation technique is proposed for in-orbit use to estimate the inertia matrix (moments and products of inertia parameters) of a satellite. To facilitate this, one attitude axis is disturbed using a reaction wheel whilst the other two axes are controlled to keep their respective angular" rates small. Within a fraction of an orbit three components of the inertia matrix can be accurately determined. This procedure is then repeated for the other two axes to obtain all nine elements of the inertia matrix. The procedure is designed to prevent the build up of momentum in the reaction wheels, whilst keeping the attitude disturbance to the satellite within acceptable limits. It can also overcome potential errors introduced by unmodeled external disturbance torques and attitude sensor noise. The second focus is on a new algorithm for in-orbit use to calibrate thruster coefficients for thrust level and alignment, using three reaction wheel actuators. These algorithms will ensure robustness against modeling errors. The algorithms assume no prior knowledge of the thruster parameters and only an initial guess of the inertia matrix. It is proposed that this calibration can be used during normal mission conditions when the satellite is stabilised. The final goal of this research study was to apply the proposed algorithms in real-time. Firstly, the thruster calibration algorithm was tested on an air-bearing table. Finally, both thruster calibration and moment of inertia algorithms were tested using data generated by UoSAT-12 while in orbit. The practical estimation results proved the feasibility of proposed algorithms.

This thesis presents an analysis of CubeSat orbits for both Low Earth Orbit (LEO) and Sun-Synchronous Orbit (SSO) missions using Systems Tool Kit (STK). The study focuses on analyzing communication, power generation, and radiation exposure while considering various factors. The analysis is based on the 3U CubeSat called UT-ProSat-1, developed by students at Virginia Polytechnic Institute and State University (VT) for an upcoming mission. The orbit size and mass adjustments were made for the LEO mission to enhance communication performance. The influence of solar activity on CubeSat lifetime and access time was examined, highlighting the significance of mass and solar activity. The impact of increasing orbit size on communication time was analyzed, emphasizing the trade-offs between mass, orbit size, and communication performance. The SSO mission prioritized power generation optimization resulted in generating sufficient power for the nominal phase of the mission. It also considered the effects of the South Atlantic Anomaly (SAA) on radiation exposure. Effective risk management of increasing the shielding for the avionics were emphasized which consequently will stabilize the orbit and prolong its lifetime. Additionally, temperature dynamics were investigated, indicating the need for further analysis considering heat dissipation and utilizing a more accurate CubeSat model. The insights gained from this study contribute to the improved the performance of CubeSats and validate the mission results, providing valuable information for successful missions in the future. / Master of Science / This project explored the trajectories that small satellites, known as CubeSats, follow around Earth. Two main paths were investigated: the Low Earth Orbit (LEO), which is close to the Earth's surface, and the Sun Synchronized Orbit (SSO), which aligns with the Sun's movement. The software called Systems Tool Kit (STK) served as the simulation tool, helping to analyze the satellites' abilities to communicate, generate power, total space radiation, and satellite's temperature throughout the missions. The study was conducted on the satellite called UT-ProSat-1, a design by students from Virginia Polytechnic Institute and State University (VT). For the LEO path, changes to the satellite's size and weight were applied to analyze its effect on the communication capabilities. Also the Sun's effect on the satellite's operational life and communication windows was assessed. Changes in the satellite's orbit can influence its communication duration, and this necessitates a balance between its weight, trajectory, and communication capacity. Regarding the Sun-aligned path, SSO, the power generated from the Sun was sufficient for the satellite's power needs throughout its mission. A particular space zone with high radiation, the South Atlantic Anomaly (SAA), was evaluated. The majority of total radiation build up on the satellite was determined to came from this area. However, risks associated with this radiation can be minimized by enhancing protection for the satellite's electronics. Such measures not only safeguard the satellite but also increase its stability and longevity in space. The temperature behavior of the satellite was analyzed, underscoring the need for a deeper examination of its thermal patterns. Insights from this study will bolster CubeSat performance and provide valuable information for future successful space missions.

Made available in DSpace on 2014-06-11T19:25:29Z (GMT). No. of bitstreams: 0 Previous issue date: 2009-02-19Bitstream added on 2014-06-13T19:12:14Z : No. of bitstreams: 1 jacob_rr_me_guara.pdf: 637187 bytes, checksum: 24c38b78cd04acc4d1694e9912f95502 (MD5) / Coordenação de Aperfeiçoamento de Pessoal de Nível Superior (CAPES) / No presente trabalho é abordada a manobra orbital Terra-Lua - Terra com o objetivo de um menor consumo de combustível. Inicialmente o satélite executa uma órbita em torno da Terra. Em um certo instante um impulso é efetuado para efetuar uma manobra não con-focal em que o satélite é transferido para uma órbita de transferência geocêntrica até um ponto da esfera de influência da Lua. A partir deste ponto o satélite é transferido para uma órbita hiperbólica em torno da Lua, e, no perilúnio desta órbita um novo impulso é dado transferindo o satélite para uma órbita lunar. A seguir é efetuada a manobra de volta em que o satélite é transferido para a órbitra geocêntrica de transferência, e no pericentro desta o satélite é transferido para a sua órbita inicial em torno da Terra. O caso não coplanar e a influência das perturbações devidas ao achatamento da Terra e a atração gravitacional da Lua também são analisados. / In the present work an Earth-Moon-Earth orbital maneuver is studied with the purpose of minimum fuel consumption. Initially it is considered an artificial satellite orbiting around the Earth. In a certain instant an impulse is effected to effect a not cofocal maneuver where the satellite is transferred to an orbit of geocentric transference until a point of the sphere of influence of the Moon. From this point the satellite is transferred to a hyperbolic orbit around the Moon, and, in the perilúnio of this orbit a new impulse is given transferring the satellite to a lunar orbit. To follow the maneuver is effected in return where the satellite is transferred to geocentric orbit of transference, and in pericentro of this the satellite is transferred to its initial orbit around the Earth. The non-coplanar case and the influence of the disturbances due to the flattening of the Earth and the gravitational attraction of the Moon also are analyzed.

No description available.