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Dissertations / Theses on the topic 'Astrodynamics'
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1
Pérez, Palau Daniel. "Dynamical transport mechanisms in celestial mechanics and astrodynamics problems." Doctoral thesis, Universitat de Barcelona, 2015. http://hdl.handle.net/10803/362369.
Full textAbstract:
L’objectiu d’aquesta tesi és afegir una petita fulla a l’arbre del coneixement. En particular a la branca del sistemes dinàmics. La teoria de sistemes dinàmics és la branca de les matemàtiques que estudia l’evolució del que ens envolta. Un dels objectius de la teoria dels sistemes dinàmics és estudiar com evoluciona amb el temps un cert procés evolutiu, és a dir, donades unes condicions inicials per a un cert estat, quin serà l’estat del sistema “t” unitats de temps. En alguns problemes és possible trobar estructures que ens separen diferents tipus de moviment. Per exemple, un moviment fitat d’un de no fitat. Aleshores, aquestes estructures determinen com evolucionà el sistema sota estudi. En aquest cas parlem de mecanismes dinàmics de transport. És a dir, quines són les possibles maneres que té un cert estat d’arribar a un altre. La teoria de sistemes dinàmics treu models i problemes gran varietat d’àmbits científics. En aquesta tesi ens centrarem en problemes de mecànica celeste i astrodinàmica.
L’estructura de la present tesi és com segueix:
− El Capítol 1 està dedicat a introduir alguns dels conceptes que es fan servir en els capítols posteriors, així com qüestions de notació i la definició dels sistemes dinàmics que s’empraran.
− En el Capítol 2 s’introdueix l’eina principal de la tesi, el Jet Transport. Per fer-la servir cal implementar una àlgebra de polinomis. El capítol explica com fer aquesta implementació. Les primeres seccions es dediquen a explicar com fer un ús eficient de la memòria i a introduir les operacions bàsiques amb polinomis (el producte per un escalar, la suma, el producte, la divisió de dos polinomis). També s’explica com realitzar altres operacions elementals com l’exponencial, el logaritme, el sinus i el cosinus així com la derivació i la integració de polinomis. A les darreres seccions s’explica com implementar operacions més complexes com la propagació de fluxos (incloent el càlcul d’aplicacions de Poincaré i altres tècniques per a millorar els resultats obtinguts), el càlcul de la inversa funcional d’un polinomi i la transformació de densitats mitjançant una aplicació.
− El Capítol 3 està dedicat a parlar sobre indicadors dinàmics. Primer es repassen els exponents de Lyapunov a temps finit i les estructures lagrangianes coherents. Fruit d’aquestes reflexions es desenvolupen algorismes per tal de disminuir el temps de còmput. Tot seguit, es donen quatre indicadors de la dinàmica alternatius basats en el Jet Transport: la màxima mida de la caixa inicial, la màxima relació d’expansió, la màxima relació de contracció i la màxima relació d’expansió a l’espai normal. El capítol segueix desenvolupant un algorisme d’extracció d’estructures per tal d’extreure i resumir la informació donada pels indicadors dinàmics. Finalment, es fan servir els indicadors dinàmics introduïts per tal de determinar zones d’estabilitat efectiva en el problema restringit de tres cossos.
− En el Capítol 4 s’estudia la col·lisió de satèl·lits artificials. Primerament s’estudien les diferents per torbacions que afecten al moviment de satèl·lits al voltant de la terra. Es considera un problema de dos cossos amb pertorbacions degudes al potencial terrestre, a la força de fregament atmosfèric i a la gravetat de la Lluna i el Sol. S’estudien els efectes d’aquestes pertorbacions i també com realitzar l’implementació mitjançant el Jet Transport. El capítol acaba amb algunes simulacions de Monte Carlo per extreure informació d’una col·lisió semblant a la produïda entre els satèl·lits Iridium-33 i el Kosmos-2251 l’any 2009.
− L’annex A explica breument les funcions desenvolupades per a aquesta tesi i s’introdueixen unes petites notes sobre paral·lelització de codis en C mitjançant open MP.
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Ya´rnoz, Daniel Garci´a. "Exploiting astrodynamics for the manipulation and exploration of asteroids." Thesis, University of Strathclyde, 2015. http://oleg.lib.strath.ac.uk:80/R/?func=dbin-jump-full&object_id=24963.
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The accessibility of minor bodies, the impact threat they pose, their scientific interest and the theorised potential for material extraction have pinpointed asteroids and comets as attractive targets for current and future space exploration. The manipulation of these minor bodies has been discussed for over a century, mainly with the aim of planetary protection, and in general based on the use of artificial external forces. Within this thesis, the manipulation of asteroids has been addressed across a range of length-scales, from orbit to dust particle manipulation, placing emphasis on exploiting natural astrodynamics. On the macro-scale regime, the capture of Near-Earth Objects into libration point orbits of the Sun-Earth system was investigated by exploiting manifold dynamics to obtain low-costs transfers. At middle scales or meso-scales, this thesis proposes the use of tidal torques acting on captured asteroids during swing-bys to manipulate the asteroid's rotational state. Possibilities included induced asteroid spin-up, de-spin, rotational fragmentation or binary break-up. In addition, the exploitation of solar radiation pressure was analysed with the purpose of generating new orbiting strategies around minor bodies. Finally, at the smallest scales or micro-scales, a novel asteroid regolith separation method based on the exploitation of differential solar radiation pressure has been proposed.
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Whiting, James K. (James Kalani) 1980. "Path optimization using sub-Riemannian manifolds with applications to astrodynamics." Thesis, Massachusetts Institute of Technology, 2011. http://hdl.handle.net/1721.1/63035.
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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2011.Cataloged from PDF version of thesis.
Includes bibliographical references (p. 131).
Differential geometry provides mechanisms for finding shortest paths in metric spaces. This work describes a procedure for creating a metric space from a path optimization problem description so that the formalism of differential geometry can be applied to find the optimal paths. Most path optimization problems will generate a sub-Riemannian manifold. This work describes an algorithm which approximates a sub-Riemannian manifold as a Riemannian manifold using a penalty metric so that Riemannian geodesic solvers can be used to find the solutions to the path optimization problem. This new method for solving path optimization problems shows promise to be faster than other methods, in part because it can easily run on parallel processing units. It also provides some geometrical insights into path optimization problems which could provide a new way to categorize path optimization problems. Some simple path optimization problems are described to provide an understandable example of how the method works and an application to astrodynamics is also given.
by James K. Whiting.
Ph.D.
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Farahmand, Mitra. "ORBITAL PROPAGATORS FOR HORIZON SIMULATION FRAMEWORK." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/167.
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This thesis describes the models of four common orbital propagators and outlines the process of integrating them into the Horizon Simulation Framework (HSF). The results of the Two-Body, J2, and J4 propagators from the HSF are then compared against the outcomes of these propagators in MATLAB and Satellite Toolkit (STK). The MATLAB algorithms verify the functionality of the propagators and determine the accuracy of the HSF implementation. The compassion against STK validates the formulation of the HSF propagators. In order to equip the HSF with a more precise means of orbit determination, adding the Simplified General Perturbations 4 (SGP4) propagator to the HSF has been the principal goal of this project. A brief description of the algorithm explains the process of configuring the original code into a format compatible with the HSF. Further, the orbital data from the SGP4 propagator across different implementations are examined. The outcomes demonstrate that the HSF algorithm generates reasonably accurate orbital data.
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Graef, Jared. "B-Plane Targeting with the Spacecraft Trajectory Optimization Suite." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2251.
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In interplanetary trajectory applications, it is common to design arrival trajectories based on B-plane target values. This targeting scheme, B-plane targeting, allows for specific target orbits to be obtained during mission design. A primary objective of this work was to implement B-plane targeting into the Spacecraft Trajectory Optimization Suite (STOpS). This work was based on the previous versions of STOpS done by Fitzgerald and Sheehan, however STOpS was redeveloped from MATLAB to python. This updated version of STOpS implements 3-dimensional computation, departure and arrival orbital phase modeling with patched conics, B-plane targeting, and a trajectory correction maneuver. The optimization process is done with three evolutionary algorithms implemented in an island model paradigm.
The algorithms and the island model were successfully verified with known optimization functions before being used in the orbital optimization cases. While the algorithms and island model are not new to this work, they were altered in this redevelopment of STOpS to closer relate to literature. This enhanced literature relation allows for easier comprehension of the both the formulation of the schemes and the code itself. With a validated optimization scheme, STOpS is able to compute near-optimal trajectories for numerous historical missions. New mission types were also easily implemented and modeled with STOpS. A trajectory correction maneuver was shown to further optimize the trajectories end conditions, when convergence was reached. The result is a versatile optimization scheme that is highly customization to the invested user, while remaining simple for novice users.
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Strange, Michael R. "Orbital Determination Feasibility of LEO Nanosatellites Using Small Aperture Telescopes." DigitalCommons@CalPoly, 2017. https://digitalcommons.calpoly.edu/theses/1714.
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This thesis is directed toward the feasibility of observing satellites on the nano scale and determining an accurate propagated orbit using a Meade LX600-ACF 14” diameter aperture telescope currently located on the California Polytechnic State University campus. The optical telescope is fitted with an f/6.3 focal reducer, SBIG ST-10XME CCD camera and Optec TCF-S Focuser. This instrumentation allowed for a 22’ X 15’ arcminute FOV in order to accurately image passing LEO satellites. Through the use of the Double-r and Gauss Initial Orbit Determination methods as well as Least Squared Differential Correction and Extended Kalman Filter Orbit Determination methods, an accurate predicted orbit can be determined.
These calculated values from observational data of satellites within the Globalstar system are compared against the most updated TLEs for each satellite at the time of observation. The determined differential errors from the well-defined TLEs acquired via online database were used to verify the feasibility of the accuracy which can be obtained from independent observations. Through minimization of error caused from imaging noise, pointing error, and timing error, the main determination of accurate orbital determination lies in the instrumentation mechanical capabilities itself. With the ability to acquire up to 7 individual satellite observations during a single transit, the use of both IOD and OD methods, and the recently acquired Cal Poly telescope with an increased 14” aperture, the feasibility of imaging and orbital determination of nanosatellites is greatly improved.
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Polzine, Benjamin. "The Collisional Evolution of Orbital Debris in Geopotential Wells and Disposal Orbits." DigitalCommons@CalPoly, 2017. https://digitalcommons.calpoly.edu/theses/1703.
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This thesis investigates the orbital debris evolution in the geosynchronous disposal orbit regime and within geosynchronous orbits effected by the geopotential wells. A propagator is developed for the accurate simulation of GEO specific orbits and the required perturbations are determined and described. Collisions are then simulated in the selected regimes using a low velocity breakup model derived from the NASA EVOLVE breakup model. The simulations described in this thesis consider a set of perturbations including the geopotential, solar and lunar gravity, and solar radiation pressure forces. This thesis is based on a prior paper and additionally seeks to address an issue in simulating East-West trapped objects. The results show that this propagator successfully simulates the presence of all wells and the East-West entrapment, and the required perturbations are outlined. Five collision test cases were simulated, one for each type of entrapment and an additional for the disposal orbit.
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Bae, Sungkoo. "GLAS spacecraft attitude determination using CCD star tracker and 3-axis gyros /." Digital version accessible at:, 1998. http://wwwlib.umi.com/cr/utexas/main.
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Iuliano, Jay R. "A Solution to the Circular Restricted N Body Problem in Planetary Systems." DigitalCommons@CalPoly, 2016. https://digitalcommons.calpoly.edu/theses/1612.
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This thesis is a brief look at a new solution to a problem that has been approached in many different ways in the past - the N body problem. By focusing on planetary systems, satellite dynamics can be modeled in a fashion similar to the Circular Restricted Three Body Problem (CR3BP) with the Circular Restricted N Body Problem (CRNBP). It was found that this new formulation of the dynamics can then utilize the tools created from all the research into the CR3BP to reassess the possibility of different complex trajectories in systems where there are more than just two large gravitational bodies affecting the dynamics, namely periodic and semi-periodic orbits, halo orbits, and low energy transfers It was also found that not only system dynamics, but models of the Jacobi constant could also be formulated similarly to the CR3BP. Validating the authenticity of these new sets of equations, the CRNBP dynamics are applied to a satellite in the Earth-Moon system and compared to a simulation of the CR3BP under identical circumstances. This test verified the dynamics of the CRNBP, showing that the two systems created almost identical results with relatively small deviations over time and with essentially identical path trends. In the Jovian system, it was found the mass ratio required to validated the assumptions required to integrate the equations of motion was around .1$\%$. Once the mass ratio grew past that limit, trajectories propagated with the CRNBP showed significant deviation from trajectories propagated with a higher fidelity model of Newtonian motion. The results from the derivation of the Jacobi constant are consistent with the 3 body system, but they are fairly standalone.
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Korn, Steven M. "An Alternative Dual-Launch Architecture for a Crewed Asteroid Mission." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/862.
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This thesis is a feasibility study for a crewed mission to a Near Earth Asteroid (NEA). An alternate dual-launch architecture is proposed and analyzed against a more established architecture. Instead of a rendezvous in a low-Earth parking orbit, the new architecture performs the rendezvous while the two spacecraft are on an Earth-escape trajectory to the destination NEA. After selecting a target asteroid, 2000 SG344, each architecture will have its best mission compared to the best mission of the other architecture.
Using the new architecture, a mission is created to the chosen NEA, 2000 SG344. A back-up Orion MPCV and a Habitation Module are launched first on a cargo configuration SLS. A crew of two astronauts is launched two hours later in the primary Orion MPCV by a crewed configuration SLS. Both of these launches are on an Earth-escape trajectory and begin rendezvous after two full days in outer space. The completed spacecraft journeys the rest of the trip to the NEA. For a period of eight days, the spacecraft remains in a tight control sphere near the asteroid by using a control algorithm and the rendezvous thrusters. The astronauts have this period to perform their EVAs and accomplish their mission objectives at the NEA. The spacecraft then departs the NEA and returns to Earth. The entire mission is 134 days and requires 2.054 km/s of Delta-v maneuvers to complete.
An analysis of multiple Lambert's methods is also done due to their extensive use in this thesis. Many of the most popular Lambert algorithms are compared by evaluating each on its accuracy, speed, and singularities. The best Lambert method to use for the orbital analysis in this paper is Battin's method because it is accurate, quick, and robust for all cases that will be observed.
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Bae, Gyoung Hyun. "Optimal control of aero-assisted orbit transfer vehicles." Diss., Georgia Institute of Technology, 1988. http://hdl.handle.net/1853/13064.
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Azari, Pouyan. "An Orbit Control System for UWE-4 Using the High Fidelity Simulation Tool Orekit." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-61409.
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Cubesats are picosatellites that have a mass of less than 1.3kg and have a shape of acube. As a result of their low cost of development and launch, cubesats are gainingpopularity in industry and academia. These satellites are also a cost-efective way forspace technology demonstrations. University of Würzburg has a longstanding cubesatprogram started with the launch of UWE-1 in 2005. This was followed by UWE-2 andUWE-3. Several technologies were tested and validated using the UWE platform. Thelast mission UWE-3 has successfully tested an attitude control system.In the next mission, UWE-4 will demonstrate an orbit control system. Being a picosatellite as small as this one (10 x 10 x 10cm 3 and 1kg) brings new challenges intodi↵erent aspects of satellite design, development, control and operation. The orbit con-trol of such a satellite is one of the problems that should be tackled. Being such a smallsatellite means having less propellant mass and much smaller thrusters than conventionalsatellites. These should be addressed in the orbit control. UWE-4 will take advantage of four NanoFEEP thrusters, on one side. Because of theiraccuracy and functionality, these thrusters can be used to implement a continuous thrustsystem. They are also a good choice because of their low energy usage. This work startswith the preparation that was needed to implement a control system. Then explains thestate of the art for continuous thrust control systems. Implements two di↵erent methods,based on perfect control and discusses the outcome. It discuses the limiting factors, likefuel mass, available electrical energy and their e↵ect on the controller performance andconcludes with recommendation for the future researches.UWE-4
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Halter, Ronald Vaughn. "A universal time of flight equation for space mechanics." Thesis, Virginia Tech, 1988. http://hdl.handle.net/10919/43406.
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A universal time of flight equation for any orbit is developed as a
function of the initial and final radius, the change in true anomaly
and the initial flight path angle. Lambert's theorem, a new corollary
to this theorem, a trigonometric variable substitution and a continuing
fraction expression are used in this development. The resulting
equation is not explicitly dependent upon eccentricity and is
determinate for -2n < (change in true anomaly) < 2n. A method to
make the continuing fraction converge rapidly is evaluated using a top
down algorithm. Finally, the accuracy of the universal time of flight
equation is examined for a representative set of orbits including near
parabolic and near rectilinear orbits.
Master of Science
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Williams, Adrian Michael. "Thermal Vacuum Chamber Refurbishment and Analysis." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1860.
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Spacecraft are subject to different environments while on orbit around the Earth and beyond. One of the most critical of these environments that must be counteracted is the thermal environment. Each spacecraft has an operating temperature that is specified in the mission requirements. The requirement stems from internal component operating temperatures that are critical to mission success. Prior to placing the spacecraft in orbit, engineers must be sure that the spacecraft will survive or risk losing the mission entirely.
The primary way to mitigate this risk is to use a thermal vacuum chamber (TVAC). The chamber is designed to resemble a space environment by reducing the pressure within the chamber to 1e-6 Torr. The differentiating factor between a vacuum chamber and a thermal vacuum chamber is the ability for the TVAC to complete a process known as thermal cycling using a temperature controller. Thermal cycling begins at a set temperature and increases within the chamber to a designated hot temperature expected to be seen on orbit. After the maximum temperature is reached, it remains there for a specified amount of time in what is called a soak. The controller then reduces the temperature to a specified cold temperature where a second soak takes place. Finally, the temperature is returned to the initial temperature and the process is repeated for a number of cycles until testing is complete.
For the purpose of this thesis, only the initial temperature increase and the first soak are being investigated. The chamber being used to run these experiments was graciously donated by MDA US Systems, however, no additional documentation was provided with the chamber. The Two identical black coated aluminum and brass cylinders have been chosen to be run with three different temperature profiles. The profiles are manually designed in the temperature controller on the chamber and vary by final soak temperature. To supplement the testing, simulations have been created for each test case in order to verify the computer model of the chamber. The simulations utilize AutoCad and Thermal Desktop to provide the results for comparison.
Each of the tests were completed successfully and produced good results that corresponded well to the simulation. The largest difference between the simulation cylinder temperature and the experimental cylinder temperature was 1.9 $^{\circ}$C. The effectiveness and efficiency of the blue chamber was compared to the other thermal vacuum chamber in the Space Environments lab. Overall, the Blue Chamber proved to be more robust and much easier to operate than the HVEC thermal vacuum chamber.
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Jackson, Daniel J. "Formulation of an Optimal Search Strategy for Space Debris At GEO." DigitalCommons@CalPoly, 2011. https://digitalcommons.calpoly.edu/theses/656.
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The purpose of this thesis is to create a search strategy to find orbital debris when the object fails to appear in the sky at its predicted location. This project is for NASA Johnson Space Center Orbital Debris Program Office through the MODEST (Michigan Orbital Debris Survey Telescope) program. This thesis will build upon the research already done by James Biehl in “Formulation of a Search Strategy for Space Debris at GEO.” MODEST tracks objects at a specific right ascension and declination. A circular orbit assumption is then used to predict the location of the object at a later time. Another telescope performs a follow-up to the original observation to provide a more accurate orbit predication. This thesis develops a search strategy when the follow-up is not successful. A general search strategy for finding space debris was developed based on previous observations. A GUI was also generated to find a search strategy in real-time for a specific object based upon previous observations of that object.
Search strategies were found by adding a 2% mean random error to the position and velocity vectors. Adding a random error allows for finding the most likely location of space debris when the orbital elements are slightly incorrect. A bivariate kernel density estimator was used to find the probability density function. The probability density function was used to find the most probable location of an object. A correlation between error in the orbital elements and error in right ascension and declination root mean square (RMS) error was investigated. It was found that the orbital elements affect the RMS error nonlinearly, but the relation between orbital element and error depended on the object and no general pattern was found. It was found that how long after the original object was found until the follow-up was attempted did not have a large impact on the probability density function or the search strategy.
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Rund, Megan S. "Interplanetary Transfer Trajectories Using the Invariant Manifolds of Halo Orbits." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1853.
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Throughout the history of interplanetary space travel, the Newtonian dynamics of the two-body problem have been used to design orbital trajectories to traverse the solar system. That is, that a spacecraft orbits only one large celestial body at a time. These dynamics have produced impressive interplanetary trajectories utilizing numerous gravity assists, such as those of Voyager, Cassini, Rosetta and countless others. But these missions required large amounts of delta-v for their maneuvers and therefore large amounts of fuel mass. As we desire to travel farther and more extensively in space, these two-body dynamics lead to impossibly high delta-v values, and missions become infeasible due to the massive amounts of fuel that they would need to carry. In the last few decades a new dynamical system has been researched in order to find new ways of designing mission trajectories: the N-body problem. This utilizes the gravitational acceleration from multiple celestial bodies on a spacecraft, and can lead to unconventional, but very useful trajectories.
The goal of this thesis is to use the dynamics of the Circular Restricted Three-Body Problem (CRTBP) to design interplanetary transfer trajectories. This method of modelling orbital dynamics takes into account the gravitational acceleration of two celestial bodies acting on a spacecraft, rather than just one. The invariant manifolds of halo orbits about Sun-planet Lagrange points are used to aid in the transfer from one planet to another, and can lead into orbital insertion about the destination planet or flyby trajectories to get to another planet. This work uses this method of dynamics to test transfers from Earth to both Jupiter and Saturn, and compares delta-v and time of flight values to traditional transfer methods. Using the CRTBP can lead to reduced delta-v amounts for completing the same missions as two-body dynamics would. The aim of this work is to research if using manifolds for interplanetary transfers could be superior for some high delta-v missions, as it could drastically reduce the required delta-v for maneuvers. With this method it could be possible to visit more distant destinations, or carry more mass of scientific payloads, due to the reduced fuel requirements.
Results of this research showed that using manifolds to aid in interplanetary transfers can reduce the delta-v of both departure from Earth and arrival at a destination planet. For transfers to Jupiter the delta-v for the interplanetary transfer was reduced by 4.12 km/s compared to starting and ending in orbits about the planets. For a transfer to Saturn the delta-v required for the interplanetary transfer was reduced by 6.77 km/s. These delta-v savings are significant and show that utilizing manifolds can lead to lower energy interplanetary transfer trajectories, and have the potential to be useful for high delta-v missions.
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Fitzgerald, Timothy J. "Spacecraft Trajectory Optimization Suite (STOpS): Optimization of Multiple Gravity Assist Spacecraft Trajectories Using Modern Optimization Techniques." DigitalCommons@CalPoly, 2015. https://digitalcommons.calpoly.edu/theses/1503.
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In trajectory optimization, a common objective is to minimize propellant mass via multiple gravity assist maneuvers (MGAs). Some computer programs have been developed to analyze MGA trajectories. One of these programs, Parallel Global Multiobjective Optimization (PaGMO), uses an interesting technique known as the Island Model Paradigm. This work provides the community with a MATLAB optimizer, STOpS, that utilizes this same Island Model Paradigm with five different optimization algorithms. STOpS allows optimization of a weighted combination of many parameters. This work contains a study on optimization algorithm performance and how each algorithm is affected by its available settings.
STOpS successfully found optimal trajectories for the Mariner 10 mission and the Voyager 2 mission that were similar to the actual missions flown. STOpS did not necessarily find better trajectories than those actually flown, but instead demonstrated the capability to quickly and successfully analyze/plan trajectories. The analysis for each of these missions took 2-3 days each. The final program is a robust tool that has taken existing techniques and applied them to the specific problem of trajectory optimization, so it can repeatedly and reliably solve these types of problems.
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Smiroldo, Jordan. "Investigation into the Mitigation of the Effects of Uncertain Optical Degradation on an Interplanetary Solar Sail Mission Using a Single Model Update." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1135.
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The renewed academic interest in using solar sails as a source of spacecraft propulsion has been accompanied by a recent fervor of investigations into non-ideal and off-nominal sail performance considerations. One of the most influential considerations, uncertain optical degradation, has been shown to present significant trajectory design difficulties. This paper investigates the potential of using a mid-course degradation model update to mitigate the risk of missing the target destination in a sample 300 day Earth-Venus trajectory. Using a range of potential degradation profiles, it is shown that correcting in the first half of the mission is highly likely to result in a trajectory that arrives sufficiently close to Venus at the end of the mission timeframe. Depending on the exact extent of the uncertainty, the data suggests that the latest a correction should take place ranges from 150 to 240 days into the mission. The influence of two different parameters, the extent and rate of degradation, are compared to show that the former of the two is more impactful on correcting timing than the latter.
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Josselyn, Scott B. "Optimization of low thrust trajectories with terminal aerocapture." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Jun%5FJosselyn.pdf.
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Thesis (Aeronautical and Astronautical Engineer)--Naval Postgraduate School, June 2003.Thesis advisor(s): I. Michael Ross, Steve Matousek. Includes bibliographical references (p. 149-150). Also available online.
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Malloy, Michael G. "Spacecraft Trajectory Optimization Suite (STOpS): Design and Optimization of Multiple Gravity-Assist Low-Thrust (MGALT) Trajectories Using Modern Optimization Techniques." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2247.
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The information presented in the thesis is a continuation of the Spacecraft Trajectory Optimization Suite (STOpS). This suite was originally designed and developed by Timothy Fitzgerald and further developed by Shane Sheehan, both graduate students at California Polytechnic State University, San Luis Obispo. Spacecraft utilizing low-thrust transfers are becoming more and more common due to their efficiency on interplanetary trajectories, and as such, finding the most optimal trajectory between two planets is something of interest. The version of STOpS presented in this thesis uses Multiple Gravity-Assist Low-Thrust (MGALT) trajectories paired with the island model paradigm to accomplish this goal. The island model utilizes four different global search algorithms: a Genetic Algorithm, Differential Evolution, Particle Swarm Optimization, and Monotonic Basin Hopping. The first three algorithms were featured in the initial version of STOpS written by Fitzgerald [1], and were subsequently modified by Sheehan [2] to work with a low-thrust adaptation of STOpS. For this work, Monotonic Basin Hopping was added to aid the suite with the MGALT trajectory search.
Monotonic Basin Hopping was successfully validated against four different test functions which had been used to validate the other three algorithms. The purpose of this validation was to ensure Monotonic Basin Hopping would work as intended, ensuring it would work in cooperation with the other three algorithms to produce a near optimal solution. After verifying the addition of Monotonic Basin Hopping, all four algorithms were used in the island model paradigm to verify MGALT STOpS’ ability to solve two known orbital transfer problem. The first verification case involved an Earth to Mars transfer with fixed thruster parameters and a predetermined time of flight. The second verification case involved a transfer from Earth to Jupiter via a Mars gravity assist; two different versions of the verification case were solved against trajectories produced by industry optimization software, the Satellite Tour Design Program Low-Thrust Gravity Assist and the Gravity Assisted Low-thrust Local Optimization Program. In the first verification case, MGALT STOpS successfully validated the Earth to Mars trajectory problem and found results agreeable to literature. In the second verification case, MGALT STOpS was partially successful in validating the Earth to Mars to Jupiter trajectory problems, and found results similar to literature. The final software produced for this work is a trajectory optimization suite implemented in MATLAB, which can solve interplanetary low-thrust trajectories with or without the inclusion of gravity assists.
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Rogers, Andrew Charles. "Optimization-Based Guidance for Satellite Relative Motion." Diss., Virginia Tech, 2016. http://hdl.handle.net/10919/79455.
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Spacecraft relative motion modeling and control promises to enable or augment a wide range of missions for scientific research, military applications, and space situational awareness. This dissertation focuses on the development of novel, optimization-based, control design for some representative relative-motion-enabled missions. Spacecraft relative motion refers to two (or more) satellites in nearly identical orbits. We examine control design for relative configurations on the scale of meters (for the purposes of proximity operations) as well as on the scale of tens of kilometers (representative of science gathering missions). Realistic control design for satellites is limited by accurate modeling of the relative orbital perturbations as well as the highly constrained nature of most space systems. We present solutions to several types of optimal orbital maneuvers using a variety of different, realistic assumptions based on the maneuver objectives.
Initially, we assume a perfectly circular orbit with a perfectly spherical Earth and analytically solve the under-actuated, minimum-energy, optimal transfer using techniques from optimal control and linear operator theory. The resulting open-loop control law is guaranteed to be a global optimum. Then, recognizing that very few, if any, orbits are truly circular, the optimal transfer problem is generalized to the elliptical linear and nonlinear systems which describe the relative motion. Solution of the minimum energy transfer for both the linear and nonlinear systems reveals that the resulting trajectories are nearly identical, implying that the nonlinearity has little effect on the relative motion. A continuous-time, nonlinear, sliding mode controller which tracks the linear trajectory in the presence of a higher fidelity orbit model shows that the closed-loop system is both asymptotically stable and robust to disturbances and un-modeled dynamics.
Next, a novel method of computing discrete-time, multi-revolution, finite-thrust, fuel-optimal, relative orbit transfers near an elliptical, perturbed orbit is presented. The optimal control problem is based on the classical, continuous-time, fuel-optimization problem from calculus of variations, and we present the discrete-time analogue of this problem using a transcription-based method. The resulting linear program guarantees a global optimum in terms of fuel consumption, and we validate the results using classical impulsive orbit transfer theory. The new method is shown to converge to classical impulsive orbit transfer theory in the limit that the duration of the zero-order hold discretization approaches zero and the time horizon extends to infinity. Then the fuel/time optimal control problem is solved using a hybrid approach which uses a linear program to solve the fuel optimization, and a genetic algorithm to find the minimizing time-of-flight. The method developed in this work allows mission planners to determine the feasibility for realistic spacecraft and motion models.
Proximity operations for robotic inspection have the potential to aid manned and unmanned systems in space situational awareness and contingency planning in the event of emergency. A potential limiting factor is the large number of constraints imposed on the inspector vehicle due to collision avoidance constraints and limited power and computational resources. We examine this problem and present a solution to the coupled orbit and attitude control problem using model predictive control. This control technique allows state and control constraints to be encoded as a mathematical program which is solved on-line. We present a new thruster constraint which models the minimum-impulse bit as a semi-continuous variable, resulting in a mixed-integer program. The new model, while computationally more expensive, is shown to be more fuel-efficient than a sub-optimal approximation. The result is a fuel efficient, trajectory tracking, model predictive controller with a linear-quadratic attitude regulator which tracks along a pre-computed ``safe'' trajectory in the presence of un-modeled dynamics on a higher fidelity orbital and attitude model.Ph. D.
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Noyes, Connor David. "Characterization of the Effects of a Sun-Synchronous Orbit Slot Architecture on the Earth's Orbital Debris Environment." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1026.
Full textAbstract:
Low Earth orbit represents a valuable limited natural resource. Of particular interest are sun-synchronous orbits; it is estimated that approximately 44% of low Earth satellites are sun-synchronous. A previously developed sun-synchronous orbit slot architecture is considered. An in-depth analysis of the relative motion between satellites and their corresponding slots is performed. The long-term evolution of Earth's orbital environment is modeled by a set of coupled ordinary differential equations. A metric for quantifying the benefit, if any, of implementing a sun-synchronous architecture is developed. The results indicate that the proposed slot architecture would reduce the frequency of collisions between satellites in sun-synchronous orbits.
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Peixoto, Leandro Nogueira. "Estudo numérico da captura gravitacional temporária utilizando o problema de quatro corpos /." Guaratinguetá : [s.n.], 2006. http://hdl.handle.net/11449/91837.
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Orientador: Ernesto Vieira NetoBanca: Othon Cabo Winter
Banca: Helio Koiti Kuga
Resumo: Com o lançamento do primeiro satélite artificial da Terra, Sputnik I, surgiu a necessidade do desenvolvimento de satélites mais eficientes e mais econômicos. Um dos mecanismos utilizados para economizar combustível numa transferência completa de um veículo espacial em órbita da Terra para uma órbita em torno da Lua, é o fenômeno de captura gravitacional temporária. Nesse trabalho é feita a análise numérica de diversas trajetórias em torno da Lua, considerando-se as dinâmicas de três e quatro corpos, com o objetivo de estudar o fenômeno da captura gravitacional temporária, através do monitoramento do sinal da energia relativa de dois corpos partícula-Lua e das componentes radiais das forças gravitacionais da Terra, da Lua e do Sol. Através desses estudos também foram obtidos diversos mapas de escape e colisão, considerando-se os movimentos prógrado e retrógrado.
Abstract: With the launch of the first artificial satellite of the Earth, Sputnik I, arose the necessity of the development of the satellites more efficient and more economic. One of the mechanisms used to save fuel in a complete transference of one spacecraft in orbit of the Earth to an orbit around the Moon, is the phenomena of the temporary gravitational capture. In this paper is made the numerical analysis of the several trajectories around the Moon, considering the dynamics of the three and four-bodies, with the objective of studying the phenomena of temporary gravitational capture, through monitoring the sign of the relative two-body energy particle-Moon and the radial component of the force of attraction, gravitational of the Earth, of the Moon and of the Sun. Though of these studies also were obtained several maps of the escape and collision, considering the prograde and retrograde movements.
Mestre
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Richardson, Matthew. "Mass Estimation through Fusion of Astrometric and Photometric Data Collection with Application to High Area-to-Mass Ratio Objects." DigitalCommons@CalPoly, 2017. https://digitalcommons.calpoly.edu/theses/1742.
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This thesis work presents the formulation for a tool developed in MATLAB to determine the mass of a space object from the fusion of astrometric and photometric data. The application for such a tool is to better model the mass estimation method used for high area-to-mass ratio objects found in high altitude orbit regimes. Typically, the effect of solar radiation pressure is examined with angles observations to deduce area-to-mass ratio calculations for space objects since the area-to-mass ratio can greatly affect its orbital dynamics. On the other hand, photometric data is not sensitive to mass but is a function of the albedo-area and the rotational dynamics of the space object. Thus from these two data types it is possible to disentangle intrinsic properties using albedo-area and area-to-mass and ultimately determine the mass of a space object. Three case studies were performed for the different orbit regimes: geosynchronous, highly elliptic, and medium earth orbit. The position states were either initialized with a two line element set or with initial orbit determination methods to simulate data which was run through an unscented Kalman filter to estimate the translational and rotational states of the space object as well as the mass an albedo area. In the geosynchronous and highly elliptic cases the tool was able to accurately predict the mass value to within 5kg of the true value based on a 95% confidence interval which will allow applications to understanding high area-to-mass objects with high certainty.
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Stoker, Kyle. "Initial Orbit Determination Error Analysis of Low-Earth Orbit Rocket Body Debris and Feasibility Study for Debris Cataloguing from One Optical Facility." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2153.
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This paper is predicated on determining the effectiveness of angles-only initial orbit determination (IOD) methods when limited observational data is available for low-Earth orbit (LEO) rocket body debris. The analysis will be conducted with data obtained from Lockheed Martin Space’s Space Object Tracking (SpOT) facility, focusing on their observational data from 2018 that contains tracking of rocket body debris for less than one minute per overhead pass. After the IOD accuracies are better understood, a feasibility study will follow that investigates the possibility of cataloguing LEO orbital debris from a single optical observation facility with similar observational capabilities as that of the SpOT facility.
The IOD accuracy analysis will investigate nine different rocket bodies, with a total of 50 orbital passes of data included in the research. Three main IOD approaches will be tested for each data set to determine the best method in achieving high levels of IOD accuracy: a traditional three-point method, an iterative method, and an assumed-circular orbit method. Application of the iterative approach results in increased accuracy for the resultant initial orbit determination as compared to the three-point IOD method, and an assumed-circular orbit assumption allows for a further increase in accuracy, especially for observed objects in near-circular orbits. The feasibility of cataloguing debris from a singular optical facility shows promise, as subsequent target acquisition after an object’s initial observation is determined to be achievable under the correct circumstances. By choosing a correct telescope pointing angle based on the IOD results from one pass of data, an observed rocket body debris object would pass through the field of view of SpOT’s spotter scope (0.7-degrees) during its next overhead pass for two different test cases. An increase field of view would increase both the likelihood of acquiring the target object and the amount of time the object is visible by the telescope.
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Mages, Declan Moore. "NAVIGATIONAL FEASIBILITY OF FLYBY / IMPACT MISSIONS TO INTERSTELLAR OBJECTS." DigitalCommons@CalPoly, 2019. https://digitalcommons.calpoly.edu/theses/2221.
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In October 2017, the first interstellar object, designated 1I/2017 U1 and more commonly referred to as Oumuamua, was detected passing through our solar system by the Pan-STARRS telescope, followed recently by the detection of 2I/Borisov in August 2019. These detections came much sooner than thought possible, and have redefined our understanding of the population of interstellar objects. With the construction of the next generation of powerful observatories, future detections are estimated to occur as frequently as two per year, and while there is significant scientific understanding to be gained from observing these objects remotely, a spacecraft sent to intercept one might be the only way to collect up-close, detailed information on the composition of extra solar object. The ideal mission scenario would be a combination flyby and impact as performed and proven feasible by the Deep Impact encounter with the comet Temple 1. A study has already been done showing that trajectories to interstellar objects are feasible with current chemical propulsion and a “launch on detection” paradigm, with an estimated 10 year wait time between favorable mission opportunities, assuming future detection capabilities.
However, while a trajectory to one of these objects might be feasible, accurately performing a flyby and impacting an object with a hyperbolic orbit presents unprecedented navigational challenges. Spacecraft-target relative velocities can range between 10 km/s to 110 km/s with high phase angles between 90° and 180°. The goal of this thesis is to determine the required navigation hardware – an optical navigation camera and attitude determination system – which could provide high mission success probability for many potential encounter scenarios. This work is performed via a simulation program developed at the Jet Propulsion Laboratory that generates simulated images of a target during the terminal guidance phase of a mission, and feeds them into the algorithms behind autonomous navigation software (AutoNav) used for the Deep Impact mission. Observations are derived from the images and used to perform target-relative orbit determination and calculate correction maneuvers.
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Bilal, Mohd. "A Heuristic Search Algorithm for Asteroid Tour Missions." Thesis, Luleå tekniska universitet, Rymdteknik, 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-71361.
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Since the discovery of Ceres, asteroids have been of immense scientific interest and intrigue. They hold answers to many of the fundamental questionsabout the formation and evolution of the Solar System. Therefore, a missionsurveying the asteroid belt with close encounter of carefully chosen asteroidswould be of immense scientific benefit. The trajectory of such an asteroidtour mission needs to be designed such that asteroids of a wide range ofcompositions and sizes are encountered; all with an extremely limited ∆Vbudget.This thesis presents a novel heuristic algorithm to optimize trajectoriesfor an asteroid tour mission with close range flybys (≤ 1000 km). The coresearch algorithm efficiently decouples combinatorial (i.e. choosing the asteroids to flyby)and continuous optimization (i.e. optimizing critical maneuversand events) of what is essentially a mixed integer programming problem.Additionally, different methods to generate a healthy initial population forthe combinatorial optimization are presented.The algorithm is used to generate a set of 1800 feasible trajectories withina 2029+ launch frame. A statistical analysis of these set of trajectories isperformed and important metrics for the search are set based on the statistics.Trajectories allowing flybys to prominent families of asteroids like Flora andNysa with ∆V as low as 4.99 km/s are obtained.Two modified implementations of the algorithm are presented. In a firstiteration, a large sample of trajectories is generated with a limited numberof encounters to the most scientifically interesting targets. While, a posteriori, trajectories are filled in with as many small targets as possible. Thisis achieved in two different ways, namely single step extension and multiplestep extension. The former fills in the trajectories with small targets in onestep, while the latter optimizes the trajectory by filling in with one asteroid per step. The thesis also presents detection of asteroids for successfullyperforming flybys. A photometric filter is developed which prunes out badlyilluminated asteroids. The best trajectory is found to perform well againstthis filter such that nine out of the ten planned flybys are feasible.
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Morrison, Oliver K. "Use of Manifolds in the Insertion of Ballistic Cycler Trajectories." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1937.
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Today, Mars is one of the most interesting and important destinations for humankind and copious methods have been proposed to accomplish these future missions. One of the more fascinating methods is the Earth-Mars cycler trajectory which is a trajectory that accomplishes repeat access to Earth and Mars with little to no fuel-burning maneuvers. This would allow fast travel to and from Mars, as well as grant the possibility of multiple missions using the same main vehicle.
Insertion from Earth-orbit onto the cycler trajectory has not been thoroughly ex- plored and the only existing method so far is a Hohmann-esque transfer via direct burn. The use of manifolds from gravitational equilibrium points has not been con- sidered for low energy transfer to the cycler trajectory. This work is primarily focused on closing this gap and analyzing the feasibility of this maneuver.
To accomplish this, a study of the cycler trajectory – and the S1L1-B class specif- ically – was completed. The required gravity assist maneuvers at each planet was analyzed through V∞ matching and the entire trajectory was generated over the re- quired inertial period. This method allowed for the generation of 2 cycler trajectories of the inbound and outbound classes, which combine to allow for a reduction in the amount of time the astronauts spend in space.
The Earth-Sun L2 point is analyzed as a potential hub for the maneuver and a halo orbit about this libration point is optimized for low energy transfer from and Earth parking orbit. The associated invariant manifold is then optimized for launch date and distance to the first trajectory on the cycler in order to burn from a trajectory on the manifold to the cycler trajectory.
iv
The comparisons of this work lie in the required ∆V to perform each maneuver compared to a direct burn onto the cycler trajectory. These values are compared and the practicality of this maneuver is drawn from these comparisons. It was found that the total required ∆V for the manifold method is larger than a direct burn from Earth orbit. However, this considers the trajectory from Earth to the halo orbit and if this is removed from consideration the ∆V is significantly reduced.
It was shown that the feasibility of this method relies heavily on the starting position of the cycler vehicle. If the vehicle begins in Earth-orbit, a direct burn is preferred, however, if the vehicle began in a halo orbit (say it was assembled there) the manifold maneuver is largely preferable.
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Nastasi, Kevin Michael. "Autonomous and Responsive Surveillance Network Management for Adaptive Space Situational Awareness." Diss., Virginia Tech, 2018. http://hdl.handle.net/10919/84931.
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As resident space object populations grow, and satellite propulsion capabilities improve, it will become increasingly challenging for space-reliant nations to maintain space situational awareness using current human-in-the-loop methods. This dissertation develops several real-time adaptive approaches to autonomous sensor network management for tracking multiple maneuvering and non-maneuvering satellites with a diversely populated Space Object Surveillance and Identification network. The proposed methods integrate suboptimal Partially Observed Markov Decision Processes (POMDPs) with covariance inflation or multiple model adaptive estimation techniques to task sensors and maintain viable orbit estimates for all targets. The POMDPs developed in this dissertation use information-based and system-based metrics to determine the rewards and costs associated with tasking a specific sensor to track a particular satellite. Like in real-world situations, the population of target satellites vastly outnumbers the available set of sensors. Robust and adaptable tasking algorithms are needed in this scenario to determine how and when sensors should be tasked. The strategies developed in this dissertation successfully track 207 non-maneuvering and maneuvering spacecraft using only 24 ground and space-based sensors. The results show that multiple model adaptive estimation coupled with a multi-metric, suboptimal POMDP can effectively and efficiently task a diverse network of sensors to track multiple maneuvering spacecraft, while simultaneously monitoring a large number of non-maneuvering objects. Overall, this dissertation demonstrates the potential for autonomous and adaptable sensor network command and control for real-world space situational awareness.Ph. D.
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Mehrparvar, Arash. "ATTITUDE ESTIMATION FOR A GRAVITY GRADIENT MOMENTUM BIASED NANOSATELLITE." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1097.
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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.
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Li, Zhenhua. "Modeling and Simulation of Autonomous Thermal Soaring with Horizon Simulation Framework." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/442.
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A thermal is a column of warm rising air triggered by differential heating on the ground. In recent studies UAVs were programmed to exploit this free atmospheric energy from thermals to improve their range and endurance. Researchers had successfully flown UAVs autonomously with thermal soaring method. Most research involved some form of flight simulation. Improvements to the aircraft and thermal models for simulation purpose would enable researchers to better design their UAVs and explore any potential flaws in their designs. An aircraft simulation with a thermal environment was created in Horizon Simulation Framework, a modeling and verification framework that was developed by Cal Poly Space Technologies and Applied Research laboratory. The objective of this study is to enhance the fidelity of existing modeling and simulation methods on autonomous thermal soaring, and to advance and demonstrate the capabilities of Horizon Simulation Framework through such implementation. The geometry of a small remote controlled glider was used in this simulation. Aerodynamic prediction programs DATCOM+ and AVL were used to obtained stability and control derivatives for this glider. The induced roll effect caused by the asymmetric vertical velocity distribution of a thermal was included in the aerodynamic roll moment calculation. The autonomous guidance algorithm for the glider included a turn logic which would determine the correct turn direction for the glider when a thermal is detected. The thermal model developed in this thesis included the capabilities to vary the time dependent location, height, radius, and vertical velocity characteristics of naturally occurring thermals.
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Diaz, Christina R. "A STUDY OF THE COLLISIONAL EVOLUTION OF ORBITAL DEBRIS IN GEOPOTENTIAL WELLS AND GEO DISPOSAL ORBITS." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1063.
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This thesis will present the effects of the orbital debris evolution in two key areas: the geosynchronous disposal orbit regime known as “graveyard” and the two geopotential wells found in 105◦ W and 75◦ E longitude bins. After developing a GEO specific orbit propagator for NASA Johnson Space Center’s Orbital Debris Of- fice, collisions were simulated throughout these regimes using a low velocity breakup model. This model considered the effects of perturbations particularly non-spherical Earth effects (specifically sectorial and zonal harmonics), lunar effects, third body effects and solar radiation pressure effects. The results show that CDPROP does well in simulating the presence of the Eastern and Western geopotential wells, as well as catching drifting GEO objects. It does not do as well in catching East-West trapped objects. Three collision test cases were then simulated in graveyard and the East and West geopotential wells.
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Peixoto, Leandro Nogueira [UNESP]. "Estudo numérico da captura gravitacional temporária utilizando o problema de quatro corpos." Universidade Estadual Paulista (UNESP), 2006. http://hdl.handle.net/11449/91837.
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Made available in DSpace on 2014-06-11T19:25:30Z (GMT). No. of bitstreams: 0
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peixoto_ln_me_guara.pdf: 13781791 bytes, checksum: 7047ea962d175dfc7039c195697ef84d (MD5)Com o lançamento do primeiro satélite artificial da Terra, Sputnik I, surgiu a necessidade do desenvolvimento de satélites mais eficientes e mais econômicos. Um dos mecanismos utilizados para economizar combustível numa transferência completa de um veículo espacial em órbita da Terra para uma órbita em torno da Lua, é o fenômeno de captura gravitacional temporária. Nesse trabalho é feita a análise numérica de diversas trajetórias em torno da Lua, considerando-se as dinâmicas de três e quatro corpos, com o objetivo de estudar o fenômeno da captura gravitacional temporária, através do monitoramento do sinal da energia relativa de dois corpos partícula-Lua e das componentes radiais das forças gravitacionais da Terra, da Lua e do Sol. Através desses estudos também foram obtidos diversos mapas de escape e colisão, considerando-se os movimentos prógrado e retrógrado.
With the launch of the first artificial satellite of the Earth, Sputnik I, arose the necessity of the development of the satellites more efficient and more economic. One of the mechanisms used to save fuel in a complete transference of one spacecraft in orbit of the Earth to an orbit around the Moon, is the phenomena of the temporary gravitational capture. In this paper is made the numerical analysis of the several trajectories around the Moon, considering the dynamics of the three and four-bodies, with the objective of studying the phenomena of temporary gravitational capture, through monitoring the sign of the relative two-body energy particle-Moon and the radial component of the force of attraction, gravitational of the Earth, of the Moon and of the Sun. Though of these studies also were obtained several maps of the escape and collision, considering the prograde and retrograde movements.
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Watson, Eric. "Sun-Synchronous Orbit Slot Architecture Analysis and Development." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/760.
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Space debris growth and an influx in space traffic will create a need for increased space traffic management. Due to orbital population density and likely future growth, the implementation of a slot architecture to Sun-synchronous orbit is considered in order to mitigate conjunctions among active satellites. This paper furthers work done in Sun-synchronous orbit slot architecture design and focuses on two main aspects. First, an in-depth relative motion analysis of satellites with respect to their assigned slots is presented. Then, a method for developing a slot architecture from a specific set of user defined inputs is derived.
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Bryan, Jason M. "GLOBAL OPTIMIZATION OF MGA-DSM PROBLEMS USING THE INTERPLANETARY GRAVITY ASSIST TRAJECTORY OPTIMIZER (IGATO)." DigitalCommons@CalPoly, 2011. https://digitalcommons.calpoly.edu/theses/663.
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Interplanetary multiple gravity assist (MGA) trajectory optimization has long been a field of interest to space scientists and engineers. Gravity assist maneuvers alter a spacecraft's velocity vector and potentially allow spacecraft to achieve changes in velocity which would otherwise be unfeasible given our current technological limitations. Unfortunately, designing MGA trajectories is difficult and in order to find good solutions, deep space maneuvers (DSM) are often required which further increase the complexity of the problem. In addition, despite the active research in the field over the last 50 years, software for MGA trajectory optimization is scarce. A few good commercial, and even fewer open-source, options exist, but a majority of quality software remains proprietary.
The intent of this thesis is twofold. The first part of this work explores the realm of global optimization applied to multiple gravity assist trajectories with deep space maneuvers (MGA-DSM). With the constant influx of new global optimization algorithms and heuristics being developed in the global optimization community, this work aims to be a high level optimization approach which makes use of those algorithms instead of trying to be one itself. Central to this approach is PaGMO, which is the open-source Parallel Multiobjective Global Optimizer created by ESA's Advanced Concepts Team (ACT). PaGMO is an implementation of the Island Model Paradigm which allows the parallelization of different global optimizers. The second part of this work introduces the IGATO software which improves PaGMO by complementing it with dynamic restart capabilities, a pruning algorithm which learns over time, subdomain decomposition, and other techniques to create a powerful optimization tool. IGATO aims to be an open-source platform independent C++ application with a robust graphical user interface (GUI). The application is equipped with 2D plotting and simulations, real time Porkchop Plot generation, and other useful features for analyzing various problems. The optimizer is tested on several challenging MGA-DSM problems and performs well: consistently performing as well or better than PaGMO on its own.
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James, Karsten J. "Feasibility of Microsatellite Active Debris Removal Systems." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1047.
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Space debris has become an increasingly hazardous obstacle to continued spaceflight operations. In an effort to mitigate this problem an investigation of the feasibility of a microsatellite active debris removal system was conducted. Through proposing a novel concept of operation, utilizing a grapple-and-tug system architecture, and by analyzing each resultant mission phase in the frame of a representative example, it was found that microsatellite scale systems are capable of fulfilling the active debris removal mission. Analysis of rendezvous, docking, control and deorbit mission requirements determined that the design of a grapple-and-tug system will be driven by sizing of the propellant required to deorbit the target vehicle. Further sensitivity analysis determined that target altitude and mass are critical factors in determining the capabilities of a microsatellite mission. Preliminary sizing demonstrated that hardware considerations for both satellite core and mission related activities do not impede microsatellite feasibility. Further investigation of microsatellite debris removal missions including detailed design analysis and engineering is suggested.
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Hawkins, Robert A. Jr. "Analysis of an Inflatable Gossamer Device to Efficiently De-Orbit CubeSats." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/1139.
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There is an increased need for spacecraft to quickly and efficiently de-orbit themselves as the amount of debris in orbit around Earth grows. Defunct spacecraft pose a significant threat to the LEO environment due to their risk of fragmentation. If these spacecraft are de-orbited at the end of their useful life their risk to future spacecraft is greatly lessened. A proposed method of efficiently de-orbiting spacecraft is to use an inflatable thin-film envelope to increase the body's area to mass ratio and thusly shortening its orbital lifetime. The system and analysis presented in this project is sized for use on a CubeSat as they are an effective utility as a technology demonstration platform. Analysis has been performed to characterize the orbital dynamics of high area to mass ratio spacecraft as well as the leak rate of such an inflatable device in a vacuum environment. Results show that a 1U CubeSat can be de-orbited using a 1.7 meter diameter spherical device in just under one year while using 0.7 grams of inflating gas, this is compared to over 25 years without any method of post-mission disposal.
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Curiel, Luis R. III. "Investigation on the Use of Small Aperture Telescopes for LEO Satellite Orbit Determination." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2253.
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The following thesis regards the use of small aperture telescopes for space domain awareness efforts. The rapidly populating space domain was motivation for the development of a new operation scheme to conduct space domain awareness feasibility studies using small telescopes. Two 14-inch Schmidt-Cassegrain Telescopes at the California Polytechnic State University and the Air Force Research Lab in Kirtland AFB, NM, in conjunction with a dedicated CCD camera and a commercial DSLR camera, were utilized to conduct optical observations on satellites in Earth orbit.
Satellites were imaged during August 2019, and from January 2020 to March 2020, resulting in the collection of 77 valid images of 16 unique satellites. These images were used to obtain celestial spherical coordinates, which were used in Gauss and Double-R angles-only initial orbit determination methods. Initial orbit determination methods successfully produced valid results, reaffirming the feasibility of using small aperture telescopes for such methods. These orbit determinations were used to propagate orbit states forward in time to determine the feasibility of future imaging of the targets with the same apparatus. Propagation results demonstrated that initial orbit determinations rapidly decayed in accuracy over distant times and are most accurate for immediate satellite passes. In addition, an attempt to combine multiple initial orbit determinations using Lambert’s problem solutions was made. Combination of these multiple initial orbit determinations resulted in either no orbit state accuracy improvement compared to individual initial orbit determinations, or a decrease in accuracy compared to these methods. Ultimately, efforts demonstrated that small telescope usage is feasible for orbit determination operations, however there may be a need for hardware and operational revisions to improve the ability of the apparatus.
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Galles, Marc Alexander. "Passive Disposal of Launch Vehicle Stages in Geostationary Transfer Orbits Leveraging Small Satellite Technologies." DigitalCommons@CalPoly, 2021. https://digitalcommons.calpoly.edu/theses/2337.
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Once a satellite has completed its operational period, it must be removed responsibly in order to reduce the risk of impacting other missions. Geostationary Transfer Orbits (GTOs) offer unique challenges when considering disposal of spacecraft, as high eccentricity and orbital energy give rise to unique challenges for spacecraft designers. By leveraging small satellite research and integration techniques, a deployable drag sail module was analyzed that can shorten the expected orbit time of launch vehicle stages in GTO. A tool was developed to efficiently model spacecraft trajectories over long periods of time, which allowed for analysis of an object’s expected lifetime after its operational period had concluded. Material limitations on drag sail sizing and performance were also analyzed in order to conclude whether or not a system with the required orbital performance is feasible. It was determined that the sail materials and configuration is capable of surviving the expected GTO environment, and that a 49 m2 drag sail is capable of sufficiently shortening the amount of time that the space vehicles will remain in space.
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Miura, Nicholas Z. "COMPARISON AND DESIGN OF SIMPLIFIED GENERAL PERTURBATION MODELS (SGP4) AND CODE FOR NASA JOHNSON SPACE CENTER, ORBITAL DEBRIS PROGRAM OFFICE." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/86.
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This graduate project compares legacy simplified general perturbation model (SGP4) code developed by NASA Johnson Space Center, Orbital Debris Program Office, to a recent public release of SGP4 code by David Vallado. The legacy code is a subroutine in a larger program named PREDICT, which is used to predict the location of orbital debris in GEO. Direct comparison of the codes showed that the new code yields better results for GEO objects, which are more accurate by orders of magnitude (error in meters rather than kilometers). The public release of SGP4 also provides effective results for LEO and MEO objects on a short time scale. The public release code was debugged and modified to provide instant functionality to the Orbital Debris Program Office. Code is provided in an appendix to this paper along with an accompanying CD. A User’s Guide is presented in Chapter 7.
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Golden, Rory Martin Mr. "Design and Performance of Circulation Control Geometries." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/964.
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With the pursuit of more advanced and environmentally-friendly technologies of today’s society, the airline industry has been pushed further to investigate solutions that will reduce airport noise and congestion, cut down on emissions, and improve the overall performance of aircraft. These items directly influence airport size (runway length), flight patterns in the community surrounding the airport, cruise speed, and many other aircraft design considerations which are setting the requirements for next generation aircraft. Leading the research in this movement is NASA, which has set specific goals for the next generation regional airliners and has categorized the designs that meet the criteria as Cruise Efficient Short Takeoff and Land (CESTOL) aircraft.
With circulation control (CC) technology addressing most of the next generation requirements listed above, it has recently been gaining more interest, thus the basis of this research. CC is an active flow control method that uses a thin sheet of high momentum jet flow ejected over a curved trailing edge surface and in turn utilizes Coanda effect to increase the airfoil’s circulation, augmenting lift, drag, and pitching moment. The technology has been around for more than 75 years, but is now gaining more momentum for further development due to its significant payoffs in both performance and system complexity.
The goal of this research was to explore the design of the CC flap shape and how it influences the local flow field of the system, in attempt to improve the performance of existing CC flap configurations and provide insight into the aerodynamic characteristics of the geometric parameters that make up the CC flap. Multiple dual radius flaps and alternative flap geometry, prescribed radius, flaps were developed by varying specific flap parameters from a baseline dual radius flap configuration that had been previously developed and researched. The aerodynamics of the various flap geometries were analyzed at three different flight conditions using two-dimensional CFD. The flight conditions examined include two low airspeed cases with blown flaps at 60° and 90° of deflection, and a transonic cruise case with no blowing and 0° of flap deflection.
Results showed that the shorter flaps of both flap configurations augmented greater lift for the low airspeed cases, with the dual radius flaps producing more lift than the corresponding length prescribed radius. The large lift generation of these flaps was accompanied by significant drag and negative pitching moments. The incremental lift per drag and moment produced was best achieved by the longer flap lengths, with the prescribed radius flaps out-performing each corresponding dual radius. Longer flap configurations also upheld the better cruise performance with the least amount of low airspeed flow, drag, and required angle of attack for a given cruise lift coefficient. The prescribed radius flaps also presented a favorable trait of keeping a more continuous skin friction distribution over the flap when the flaps were deflected, where all dual radius configurations experienced a distinct fluctuation at the location where the surface curvature changes between its two radii. The prescribed radius flaps displayed a similar behavior when the flaps were not deflected, during the cruise conditions analyzed.
Performance trends for the different flap configurations, at all three flight conditions, are presented at the end of each respective section to provide guidance into the design of CC geometry. The results of the presented research show promise in modifying geometric surface parameters to yield improved aerodynamics and performance.
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Zohar, Guy G. "AD-HOC REGIONAL COVERAGE CONSTELLATIONS OF CUBESATS USING SECONDARY LAUNCHES." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/927.
Full textAbstract:
As development of CubeSat based architectures increase, methods of deploying constellations of CubeSats are required to increase functionality of future systems. Given their low cost and quickly increasing launch opportunities, large numbers of CubeSats can easily be developed and deployed in orbit. However, as secondary payloads, CubeSats are severely limited in their options for deployment into appropriate constellation geometries.
This thesis examines the current methods for deploying cubes and proposes new and efficient geometries using secondary launch opportunities. Due to the current deployment hardware architecture, only the use of different launch opportunities, deployment direction, and deployment timing for individual cubes in a single launch are explored. The deployed constellations are examined for equal separation of Cubes in a single plane and effectiveness of ground coverage of two regions. The regions examined are a large near-equatorial zone and a medium sized high latitude, high population density zone.
Results indicate that simple deployment strategies can be utilized to provide significant CubeSat dispersion to create efficient constellation geometries. The same deployment strategies can be used to develop a multitude of differently dispersed constellations. Different launch opportunities can be utilized to tailor a constellation for a specific region or mission objective. Constellations can also be augmented using multiple launch opportunities to optimize a constellation towards a specific mission or region. The tools developed to obtain these results can also be used to perform specific analysis on any region in order to optimize future constellations for other applications.
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Guzman, Esteban. "Generating Exploration Mission-3 Trajectories to a 9:2 NRHO using Machine Learning." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1953.
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The purpose of this thesis is to design a machine learning algorithm platform that provides expanded knowledge of mission availability through a launch season by improving trajectory resolution and introducing launch mission forecasting. The specific scenario addressed in this paper is one in which data is provided for four deterministic translational maneuvers through a mission to a Near Rectilinear Halo Orbit (NRHO) with a 9:2 synodic frequency. Current launch availability knowledge under NASA’s Orion Orbit Performance Team is established by altering optimization variables associated to given reference launch epochs. This current method can be an abstract task and relies on an orbit analyst to structure a mission based off an established mission design methodology associated to the performance of Orion and NASA's Space Launch System. Introducing a machine learning algorithm trained to construct mission scenarios within the feasible range of known trajectories reduces the required interaction of the orbit analyst by removing the needed step of optimizing the orbit to fit an expected translational response required of the spacecraft. In this study, k-Nearest Neighbor and Bayesian Linear Regression successfully predicted classical orbital elements for the launch windows observed. However both algorithms had limitations due to their approaches to model fitting. Training machine learning algorithms off of classical orbital elements introduced a repetitive approach to reconstructing mission segments for different arrival opportunities through the launch window and can prove to be a viable method of launch window scan generation for future missions.
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Robinson, John. "Analysis of the orbit lowering and attitude control performance of a magnetic coil-augmented gossamer sail." Honors in the Major Thesis, University of Central Florida, 2009. http://digital.library.ucf.edu/cdm/ref/collection/ETH/id/1314.
Full textAbstract:
This item is only available in print in the UCF Libraries. If this is your Honors Thesis, you can help us make it available online for use by researchers around the world by following the instructions on the distribution consent form at http://library.ucf.edu/Systems/DigitalInitiatives/DigitalCollections/InternetDistributionConsentAgreementForm.pdf You may also contact the project coordinator, Kerri Bottorff, at kerri.bottorff@ucf.edu for more information.Bachelors
Engineering and Computer Science
Aerospace Engineering
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Gagliano, Joseph R. "Orbital Constellation Design and Analysis Using Spherical Trigonometry and Genetic Algorithms: A Mission Level Design Tool for Single Point Coverage on Any Planet." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1877.
Full textAbstract:
Recent interest surrounding large scale satellite constellations has increased analysis efforts to create the most efficient designs. Multiple studies have successfully optimized constellation patterns using equations of motion propagation methods and genetic algorithms to arrive at optimal solutions. However, these approaches are computationally expensive for large scale constellations, making them impractical for quick iterative design analysis. Therefore, a minimalist algorithm and efficient computational method could be used to improve solution times. This thesis will provide a tool for single target constellation optimization using spherical trigonometry propagation, and an evolutionary genetic algorithm based on a multi-objective optimization function. Each constellation will be evaluated on a normalized fitness scale to determine optimization. The performance objective functions are based on average coverage time, average revisits, and a minimized number of satellites. To adhere to a wider audience, this design tool was written using traditional Matlab, and does not require any additional toolboxes.
To create an efficient design tool, spherical trigonometry propagation will be utilized to evaluate constellations for both coverage time and revisits over a single target. This approach was chosen to avoid solving complex ordinary differential equations for each satellite over a long period of time. By converting the satellite and planetary target into vectors of latitude and longitude in a common celestial sphere (i.e. ECI), the angle can be calculated between each set of vectors in three-dimensional space. A comparison of angle against a maximum view angle, , controlled by the elevation angle of the target and the satellite’s altitude, will determine coverage time and number of revisits during a single orbital period.
Traditional constellations are defined by an altitude (a), inclination (I), and Walker Delta Pattern notation: T/P/F. Where T represents the number of satellites, P is the number of orbital planes, and F indirectly defines the number of adjacent planes with satellite offsets. Assuming circular orbits, these five parameters outline any possible constellation design. The optimization algorithm will use these parameters as evolutionary traits to iterate through the solutions space. This process will pass down the best traits from one generation to the next, slowly evolving and converging the population towards an optimal solution. Utilizing tournament style selection, multi-parent recombination, and mutation techniques, each generation of children will improve on the last by evaluating the three performance objectives listed. The evolutionary algorithm will iterate through 100 generations (G) with a population (n) of 100.
The results of this study explore optimal constellation designs for seven targets evenly spaced from 0° to 90° latitude on Earth, Mars and Jupiter. Each test case reports the top ten constellations found based on optimal fitness. Scatterplots of the constellation design solution space and the multi-objective fitness function breakdown are provided to showcase convergence of the evolutionary genetic algorithm. The results highlight the ratio between constellation altitude and planetary radius as the most influential aspects for achieving optimal constellations due to the increased field of view ratio achievable on smaller planetary bodies. The multi-objective fitness function however, influences constellation design the most because it is the main optimization driver. All future constellation optimization problems should critically determine the best multi-objective fitness function needed for a specific study or mission.
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Villa, Jacopo. "Optical Navigation for Autonomous Approach of Unexplored Small Bodies." Thesis, KTH, Rymdteknik, 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-285863.
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This thesis presents an autonomous vision-based navigation strategy applicable to the approach phase of a small body mission, developed within the Robotics Section at NASA Jet Propulsion Laboratory. Today, the operations performed to approach small planetary bodies are largely dependent on ground support and human decision-making, which demand operational complexity and restrict the spectrum of achievable activities throughout the mission. In contrast, the autonomous pipeline presented here could be run onboard, without ground intervention. Using optical data only, the pipeline estimates the target body's rotation, pole, shape, and performs identification and tracking of surface landmarks, for terrain relative navigation. An end-to-end simulation is performed to validate the pipeline, starting from input synthetic images and ending with an orbit determination solution. As a case study, the approach phase of the Rosetta mission is reproduced, and it is concluded that navigation performance is in line with the ground-based state-of-the-art. Such results are presented in detail in the paper attached in the appendix, which presents the pipeline architecture and navigation analysis. This thesis manuscript aims to provide additional context to the appended paper, further describing some implementation details used for the approach simulations.Detta examensarbete presenterar en strategi för ett autonomt visionsbaserat navigationssystem för att närma sig en liten himlakropp. Strategin har utvecklats av robotikavdelningen vid NASA Jet Propulsion Laboratory i USA. Nuvarande system som används för att närma sig en liten himlakropp bygger till största delen på markstationer och mänskligt beslutsfattande, vilka utgör komplexa rutiner och begränsar spektrumet av möjliga aktiviteter under rymduppdraget. I jämförelse, det autonoma system presenterat i denna rapport är utformat för att köras helt från rymdfarkosten och utan krav på kontakt med markstationer. Genom att använda enbart optisk information uppskattar systemet himlakroppens rotation, poler och form samt genomför en identifiering och spårning av landmärken på himlakroppens yta för relativ terrängnavigering. En simulering har genomförts för att validera det autonoma navigationssystemet. Simuleringen utgick ifrån bilder av himlakroppen och avslutades med en lösning på banbestämningsproblemet. Fasen då rymdfarkosten i ESA:s Rosetta-rymduppdrag närmar sig kometen valdes som fallstudie för simuleringen och slutsatsen från denna fallstudie var att systemets autonoma navigationsprestanda var i linje med toppmoderna system. Den detaljerade beskrivningen av det autonoma systemet och resultaten från studien har presenterats i ett konferensbidrag, som ingår som bilaga till rapporten. Inledningen av rapporten syftar till att förtydliga bakgrunden och implementering som komplement till innehållet i bilagan.
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Kinnett, Ryan L. "System Integration and Attitude Control of a Low-Cost Spacecraft Attitude Dynamics Simulator." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/271.
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The CalPoly Spacecraft Attitude Dynamics Simulator mimics the rotational dynamics of a spacecraft in orbit and acts as a testbed for spacecraft attitude control system development and demonstration. Prior to this thesis, the simulator platform and several subsystems had been designed and manufactured, but the total simulator system was not yet capable of closed-loop attitude control. Previous attempts to make the system controllable were primarily mired by data transport performance. Rather than exporting data to an external command computer, the strategy implemented in this thesis relies on a compact computer onboard the simulator platform to handle both attitude control processing and data acquisition responsibilities. Software drivers were created to interface the computer’s data acquisition boards with Matlab, and a Simulink library was developed to handle hardware interface functions and simplify the composition of attitude control schemes. To improve the usability of the system, a variety of actuator control, hardware testing, and data visualization utilities were also created. A closedloop attitude control strategy was adapted to facilitate future sensor installations, and was tested in numerical simulation. The control model was then updated to interface with the simulator hardware, and for the first time in the project history, attitude control was performed onboard the CalPoly spacecraft attitude dynamics simulator. The demonstration served to validate the numerical model and to verify the functionality of the entire simulator system.
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Arora, Nitin. "High performance algorithms to improve the runtime computation of spacecraft trajectories." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/49076.
Full textAbstract:
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.
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Crowley, Kian Guillaume. "Exploring the Concept of a Deep Space Solar-Powered Small Spacecraft." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1936.
Full textAbstract:
New Horizons, Voyager 1 & 2, and Pioneer 10 & 11 are the only spacecraft to ever venture past Pluto and provide information about space at those large distances. These spacecraft were very expensive and primarily designed to study planets during gravitational assist maneuvers. They were not designed to explore space past Pluto and their study of this environment is at best a secondary mission. These spacecraft rely on radioisotope thermoelectric generators (RTGs) to provide power, an expensive yet necessary approach to generating sufficient power. With Cubesats graduating to interplanetary capabilities, such as the Mars-bound MarCO spacecraft, matching the modest payload requirements to study the outer Solar System (OSS) with the capabilities of low-power nano-satellites may enable much more affordable access to deep space. This paper explores a design concept for a low-cost, small spacecraft, designed to study the OSS and satisfy mission requirements with solar power. The general spacecraft design incorporates a parabolic reflector that acts as both a solar concentrator and a high gain antenna. This paper explores a working design concept for a small spacecraft to operate up to 100 astronomical units (AU) from the sun. Deployable reflector designs, thermal and radiation environments, communications and power requirements, solar system escape trajectory options, and scientific payload requirements are detailed, and a working system is proposed that can fulfill mission requirements with expected near-future innovations in a few key technologies.
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Conrad, Michael Curt. "COMET: CONSTRAINED OPTIMIZATION OF MULTIPLE-DIMENSIONS FOR EFFICIENT TRAJECTORIES." DigitalCommons@CalPoly, 2011. https://digitalcommons.calpoly.edu/theses/666.
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The paper describes the background and concepts behind a master’s thesis platform known as COMET (Constrained Optimization of Multiple-dimensions for Efficient Trajectories) created for mission designers to determine and evaluate suitable interplanetary trajectories. This includes an examination of the improvements to the global optimization algorithm, Differential Evolution, through a cascading search space pruning method and decomposition of optimization parameters. Results are compared to those produced by the European Space Agency’s Advanced Concept Team’s Multiple Gravity Assist Program. It was found that while discrepancies in the calculation of ΔV’s for flyby maneuvers exist between the two programs, COMET showed a noticeable improvement in its ability to avoid premature convergence and find highly isolated solutions.