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1
O'Donnell, Kathryn. "Satellite orbits in resonance with tesseral harmonics : absolute and relative orbit analysis." Thesis, University of Surrey, 2006. http://epubs.surrey.ac.uk/842688/.
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A concise, novel description is presented of near-circular satellite motion at arbitrary inclination in resonance with a single dominant tesseral harmonic of a gravitational potential. A practical method is then given for determining the validity of the ideal resonance assumption in specific regions of phase space. The model has been designed to be potentially sufficiently accurate for use in orbit determination yet computationally concise enough for implementation on-board small satellites. Unlike more traditional, mathematically rigorous approaches the orbit description has a relatively simple geometric interpretation making it ideal for use in mission analysis and design. It also facilitates a summary of the factors determining and affecting the nature of resonant motion experienced by satellites. This resonance model is incorporated into a curvilinear relative orbit framework to characterize the effects of tesseral resonance on the relative motion of formations of satellites. The results show that these effects can be the same magnitude as that due to short periodic J2 motion, or secular motion due to small inclination differences for close LEO satellite formations. The analytical relative resonant orbit model also allows the key factors determining the relative resonant motion to be isolated. Finally, the intuitive nature of the resonant orbit model is exploited in two ways. Firstly, the analogy between the non-linear motion of a satellite in resonance to that of a simple pendulum is exploited to develop control strategies for maintaining both spatial and temporal separations between satellites in a resonant formation. Secondly, a simple mission analysis tool is developed to allow orbit analysts to determine whether a given satellite mission could encounter a resonance of significant strength. The output of this software tool is also used to identify the limitations of the resonance model for describing motion about other celestial bodies such as the Moon, Venus and Mars.
2
Rothacher, Markus. "Orbits of satellite systems in space geodesy /." Zürich : Institut für Geodäsie und Photogrammetrie, 1992. http://www.ub.unibe.ch/content/bibliotheken_sammlungen/sondersammlungen/dissen_bestellformular/index_ger.html.
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Sowter, Andrew. "Drag coefficients with applications to satellite orbits." Thesis, Aston University, 1989. http://publications.aston.ac.uk/12059/.
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In the last twenty or so years the results of theory and experiment have produced much information on the characteristics of gas-surface interactions relevant to a satellite in hyperthermal free-molecular flow. This thesis contains reviews of the rarefied gas dynamics applicable to satellites and has attempted to compare existing models of gas-surface interaction with contemporary knowledge of such systems. It is shown that a more natural approach would be to characterise the gas-surface interaction using the normal and tangential momentum accommodation coefficients, igma' and igma respectively, specifically in the form igma = constant , igma' = igma'0 -igma'1sec i where i is the angle subtended between the incident flow and the surface normal and igma,igma'0 and igma'1 are constants. Adopting these relationships, the effects of atmospheric lift on inclination, i, and atmospheric drag on the semi-major axis, a, and eccentricity, e, have been investigated. Applications to ANS-1 (1974-70A) show that the observed perturbation in i can be ascribed primarily to non-zero igma'1 whilst perturbations in a and e produce constraint equations between the three parameters. The numerical results seem to imply that a good theoretical orbit is achieved despite a much lower drag coefficient than anticipated by earlier theories.
4
Shen, Haijun. "Optimal scheduling for satellite refueling in circular orbits." Diss., Georgia Institute of Technology, 2003. http://hdl.handle.net/1853/12331.
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Конин, Валерий Konin Valeriy, Алексей Pogurelskiy Oleksey Погурельский, and Федор Shyshkov Fedir Шишков. "Model for navigation satellite availability on various orbits analysis." Thesis, Proceedings, THE SIXTH World Con-gress. Aviation in the XXI- st Century. Safety in Aviation and Space Technologies, September 23–25, 2014 .– Kyiv, 2014. – V.2, – P.3.2.22 – 3.2.25, 2014. http://er.nau.edu.ua/handle/NAU/25239.
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While facing different tasks in space, the necessity to create a reliable way of
navigation in space is required. The model is developed to analyze the availability of
the navigation satellites in space<br>National Aviation University
6
DAVIDZ, HEIDI L. "USE OF NEAR-FROZEN ORBITS FOR SATELLITE FORMATION FLYING." University of Cincinnati / OhioLINK, 2001. http://rave.ohiolink.edu/etdc/view?acc_num=ucin994696079.
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Gabrielsson, Jonas. "Estimation of satellite orbits using ground based radar concept." Thesis, Umeå universitet, Institutionen för fysik, 2021. http://urn.kb.se/resolve?urn=urn:nbn:se:umu:diva-185298.
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Today an abundance of objects are circulating in earth captured orbit. Monitoring these objects is of national security interest. One way to map any object in orbit is with their Keplerian elements. A method for estimating the Keplerian elements of a satellite orbit simulating a ground based radar station has been investigated. A frequency modulated continuous wave radar (FMCW) with a central transmitter antenna and a grid of receivers was modeled in MATLAB. The maximum likelihood estimator (MLE) was obtained to estimate the parameters from the received signal. The method takes advantage of the relations between the Cartesian position and velocity and the Keplerian elements to confine the search space. For a signal to noise ratio (SNR) of 10dB, the satellite was followed during a time period of 0.1s where the positions were found within average error of range: ±1.4m, azimuth: ±2.0·10−6 rad and elevation: ±8.4·10−7 rad. Using a linear approximation of the velocity the Keplerian elements were found within average error of i: ±0.0050 rad, Ω:±0.0050 rad, ω: ±0.0058 rad, a: ±2.60·105m, e:±0.0021 and ν: ±0.24 rad. A method to obtain more accurate estimates is proposed.
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Hugentobler, Urs. "Astrometry and satellite orbits: theoretical considerations and typical applications /." [S.l.] : [s.n.], 1997. http://www.ub.unibe.ch/content/bibliotheken_sammlungen/sondersammlungen/dissen_bestellformular/index_ger.html.
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Beutel, Kenneth L. "A standard library for modeling satellite orbits on a microcomputer/." Thesis, Monterey, California. Naval Postgraduate School, 1988. http://hdl.handle.net/10945/23404.
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Introductory students of astrodynamics and the space environment are required to have a fundamental understanding for the kinematic behavior of satellite orbits. This thesis develops a standard library that contains the basic formulas for modeling earth orbiting satellites. This library is used as a basis for implementing a satellite motion simulator that can be used to demonstrate orbital phenomena in the classroom. This thesis surveys the equations of orbital elements, coordinate systems and analytic formulas into a standard method for modeling earth orbiting satellites. The standard library is written in the C programming language and is designed to be highly portable between a variety of computer environments. The simulation draws heavily on the standards established by the library to produce a graphics-based orbit simulation program written for the Apple Macintosh computer. The simulation demonstrates the utility of the standard library functions but, because of its extensive use of the Macintosh user interface, is not portable to other operating systems. Keywords: Computer programming; Celestial mechanics; Theses. (KR)
10
Fedi, Casas Manrico. "Dynamics and control of tethered satellite formations in low-Earth orbits." Doctoral thesis, Universitat Politècnica de Catalunya, 2015. http://hdl.handle.net/10803/290859.
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This thesis is focused on the study of dynamics and control of a multi-tethered satellite formation, where a multi-tethered formation is made up with several satellites (agents) connected by means of cables (tethers).
The goal of the first part of the study is to evaluate the effect of tether mass on multi-tethered clusters. Due to the complexity of the formations analyzed, the stability of the formation is assessed through a numerical simulation. The behavior is evaluated in the ideal case of circular orbits, but also in non-ideal cases such as that of elliptical reference orbit or perturbed motion.
For circular reference orbits, the dynamics of tethered satellite formation is studied, showing that tether mass affects formation dynamics for closed configurations featuring external tethers. This leads to significant instability effects affecting the position of deputies with respect to the parent body neglected by a more elementary modeling approach.
When combined effect of orbit eccentricity and tether mass on tethered formations is analyzed, the most noticeable effect due to eccentricity is the increase in the variation of the local spin rate of the cluster between perigee and apogee passes of the reference elliptical orbit. This effect has consequences over the elongation of tethers, shape of tether oscillations and angular separation between adjacent tethers especially for open formations. When taking into account the J2 effect on massive tethered satellite formations, in the Earth¿facing scenario, the trajectory of the parent body presents oscillations of increasing amplitude in the direction perpendicular to the orbital plane.
The second part of the study is focused on deriving a control law for position and attitude control of an Earth-facing double pyramid multi-tethered cluster. The control problem is decomposed in two levels: A first level to perform position and attitude coarse control of the formation as a whole, and a second level to achieve accurate position and control of each agent of the cluster.
For the purpose of attitude control, and taking advantage again of the similarities between a tethered cluster and a rigid body, the virtual structure approach is chosen as the most suitable option. The formulation shown in this thesis augments the general virtual structure equations valid for a static formation by adding the kinematics of a spinning formation, since the original formulation is valid only to achieve a static final state. The controller is designed to modify the spin rate and the moment of inertia of the formation through a reeling mechanism, and therefore to be able to control the Likins-Pringle tilting angle of the cluster.
For the derivation of the accurate positioning control law, the study initially discusses different alternatives based on the state of the art of the robotics control literature. After evaluating other alternatives, two control laws are chosen for this application: One based on a PID controller and one based on the sliding mode control technique. For the sliding mode based control, a proof of semi-global exponential stability is provided. Results of a representative simulation assess the viability of the control approach proposed leading to a submillimetric positioning accuracy.<br>La tesi es centra en l'estudi de la dinàmica i control de formacions de satèl·lits connectats per tethers. Aquestes formacions estan compostes per diversos satèl·lits (agents) connectats per cables (tethers). L'objectiu de la primera part de l'estudi, és l'avaluació de l'efecte de la massa a clústers connectats per múltiples tethers. Degut a la complexitat de les formacions analitzades, l'estabilitat de la formació s'analitza a través de simulacions. S'estudia el comportament pel cas ideal d'orbites circulars, així com en casos no ideals tals com orbites de referència el·liptiques, o moviment sota l'efecte de pertorbacions. La tesi analitza la dinàmica de les formacions per òrbites circulars, mostrant que l'efecte de la massa dels tethers afecta la dinàmica de formacions de geometria tancada (on el perímetre extern esta definit per tethers) amb un satèl·lit central. Aquest efecte dóna lloc a una clara inestabilitat que afecta la posició dels agents respecte a l'objecte central. Aquest efecte no és apreciable en models simplificats on s'ignora l'efecte de la massa al model. Quan es combina una òrbita de referència el·liptica amb un model que incorpora la massa dels tethers, l'efecte més notori és la variació de la velocitat de rotació local del clúster entre el pas per l'apogeu i perigeu de l'òrbita de referència. Aquest efecte té conseqüències sobre l'elongació dels tethers, la forma de les oscil·lacions, i la separació entre tethers adjacents (especialment en el cas de formacions obertes). Quan es té en compte l'efecte de la pertorbació J2, en el cas de formacions orientades envers la Terra, la trajectòria de l'objecte central presenta oscil·lacions d'amplitud creixent en la direcció perpendicular al pla orbital. La segona part de l'estudi es centra en la definició d'una llei de control per regular la posició i orientació d'un clúster amb geometria de doble piràmide orientat envers la Terra. El problema de control es descompon en dos nivells. Un primer nivell per un control primari de posició i orientació del cluster, i un segon nivell per un control de posició precís per a cada agent del cluster. Per tal d'aconseguir el primer nivell de control, i aprofitant les similituds entre un cluster connectat per tethers i un sòlid rígid, s'utilitza la tècnica d'estructura virtual. La formulació utilitzada en aquest estudi amplia el model general d'estructura virtual utilitzat per formacions estàtiques, tot afegint les equacions necessàries per a una formació que gira sobre un eix propi. El controlador esta dissenyat per permetre el canvi de la velocitat de gir i el moment d'inèrcia de la formació a través d'un sistema que permet modificar la longitud dels tethers. D'aquesta forma es permet controlar l'angle d'inclinació de Likins-Pringle del clúster. Per a la definició del control de precisió, l'estudi avalua inicialment diferents alternatives basades en l'estat de l'art de sistemes de control aplicats a robòtica. Després de descartar altres alternatives, es proposen dues lleis de control : Una primera basada en un controlador PID, i una basada en control lliscant. Per l'opció de control lliscant es presenta una demostració d'estabilitat exponencial semiglobal. Els resultats de simulacions confirmen la viabilitat de la solució de control que permet posicionament amb precisió submil·limetrica
11
Amier, Zine-Eddine. "On some transportation problems involving tethered satellite systems." Thesis, McGill University, 1987. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=66256.
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Chadwell, C. David. "Investigation of stochastic models to improve the global positioning system satellite orbits /." The Ohio State University, 1995. http://rave.ohiolink.edu/etdc/view?acc_num=osu1487861796818699.
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胡玉蓉 and Yurong Hu. "Datagram routing for low earth orbit satellite networks." Thesis, The University of Hong Kong (Pokfulam, Hong Kong), 2001. http://hub.hku.hk/bib/B31224441.
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Salazar, Kardozo Alexandros. "A High-Level Framework for the Autonomous Refueling of Satellite Constellations." Thesis, Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/14534.
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Satellite constellations are an increasingly attractive option for many commercial and military applications. They provide a robust and distributed method of accomplishing the goals of expensive monolithic satellites. Among the many challenges that satellite constellations engender (challenges in control, coordination, disposal, and other areas), refueling is of particular interest because of the many methods one can use to refuel a constellation and the lifetime implications on the satellites.
The present work presents a methodology for carrying out peer-to-peer refueling maneuvers within a constellation. Peer-to-peer (P2P) refueling can be of great value both in cases where a satellite unexpectedly consumes more fuel than it was alloted, and as part of a mixed refueling strategy that will include an outside tanker bringing fuel to the constellation. Without considering mixed-refueling, we formulate the peer-to-peer refueling problem as an assignment problem that seeks to guarantee that all satellites will have the fuel they need to be functional until the next refueling, while concurrently minimizing the cost in fuel that the refueling maneuvers entail. The assignment problem is then solved via auctions, which, by virtue of their distributed nature, can easily and effectively be implemented on a constellation without jeopardizing any robustness properties.
Taking as a given that the P2P assignment problem has been solved, and that it has produced some matching among fuel deficient and fuel sufficient satellites, we then seek to sequence those prescribed maneuvers in the most effective manner. The idea is that while a constellation can be expected to have some redundancy, enough satellites leaving their assigned orbital slots will eventually make it impossible for the constellation to function. To tackle this problem, we define a wide class of operability conditions, and present three algorithms that intelligently schedule the maneuvers. We then briefly show how combining the matching and scheduling problems yields a complete methodology for organizing P2P satellite refueling operations.
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Habana, Nlingilili Oarabile Kgosietsile. "Gravity Recovery by Kinematic State Vector Perturbation from Satellite-to-Satellite Tracking for GRACE-like Orbits over Long Arcs." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1578042687104082.
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Belli, Jacob. "Mission Analysis for Pico-Scale Satellite Based Dust Detection in Low Earth Orbits." Master's thesis, University of Central Florida, 2013. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/5764.
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A conceptual dust detection mission, KnightSat III, using pico-scale satellites is analyzed. The purpose of the proposed KnightSat III mission is to aid in the determination of the size, mass, distribution, and number of dust particles in low earth orbits through a low cost and flexible satellite or a formation of satellites equipped with a new dust detector. The analysis of a single satellite mission with an on-board dust detector is described; though this analysis can easily be extended to a formation of pico-scale satellites. Many design aspects of the mission are discussed, including orbit analysis, power management, attitude determination and control, and mass and power budgets. Two of them are emphasized. The first is a new attitude guidance and control method, and the second is the online optimal power scheduling. It is expected that the measurements obtained from this possible future mission will provide insight into the dynamical processes of inner solar system dust, as well as aid in designing proper micro-meteoroid impact mitigation strategies for future man-made spacecraft.<br>M.S.A.E.<br>Masters<br>Mechanical and Aerospace Engineering<br>Engineering and Computer Science<br>Aerospace Engineering; Space System Design and Engineering
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Polaha, James Henry. "An analysis of low-earth-orbit-satellite communication systems." Thesis, Virginia Polytechnic Institute and State University, 1989. http://hdl.handle.net/10919/74533.
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There is an ever increasing need for low-cost communication systems in the world. One such system, low-earth-orbit satellites, can provide store-and-forward, as opposed to real time, communication for many earth stations. The advantages and disadvantages of such a system is presented. Material covering protocols and communications architectures is elaborated upon for the use of amateur radio communications. Doppler shift and its effect on satellites in low-earth-orbit is examined. Efficiency and throughput of the Amateur X.25 Protocol will be explored. The last chapter entails the analysis of the PACSAT experiment.<br>Master of Science
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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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Mannheimer, Elias. "Dangerous Orbits : Applying the Law of Self-defence to Hostile Acts Against Satellite Systems." Thesis, Uppsala universitet, Juridiska institutionen, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:uu:diva-321385.
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The world has found itself in the unsatisfactory position of depending greatly upon the services of satellites, all while the risk of satellites becoming targets during conflict looms ever greater. This paper assesses the lex lata of the law of self-defence as enshrined in the Charter of the United Nations, focusing on the rationae materiae aspect of the armed attack concept. It thereafter applies general conclusions in this regard to the specific context of hostile acts against satellite systems, with an aim to clarify under what conditions such hostile acts justify the exercise force in self-defence.
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Thackrey, Keith R. "A geometric approach to determination of satellite ephemeris over a limited area." Thesis, Virginia Tech, 1988. http://hdl.handle.net/10919/43058.
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Range and,interferometric observations have been
examined for their potential, application in a geometric
approach to determination of satellite ephemeri. The
approach differs from the normal (dynamic) approach in that
each satellite position is treated as an independent state
variable or benchmark.
Programs have been developed that simulate and format
the input, data for the least squares estimation routines,
and perform statistical analyses of those results. Random error, tropospheric refraction errors, and atomic clock errors have been considered, and the range observation adjustment program directed to solve for clock errors.<br>Master of Science
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Patel, Ekta, Gurtina Besla, and Sangmo Tony Sohn. "Orbits of massive satellite galaxies – I. A close look at the Large Magellanic Cloud and a new orbital history for M33." OXFORD UNIV PRESS, 2017. http://hdl.handle.net/10150/623269.
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The Milky Way (MW) and M31 both harbour massive satellite galaxies, the Large Magellanic Cloud (LMC) and M33, which may comprise up to 10 per cent of their host's total mass. Massive satellites can change the orbital barycentre of the host-satellite system by tens of kiloparsec and are cosmologically expected to harbour dwarf satellite galaxies of their own. Assessing the impact of these effects crucially depends on the orbital histories of the LMC and M33. Here, we revisit the dynamics of theMW-LMC system and present the first detailed analysis of the M31-M33 system utilizing high-precision proper motions and statistics from the dark-matter-only Illustris cosmological simulation. With the latest Hubble Space Telescope proper motion measurements of M31, we reliably constrain M33' s interaction history with its host. In particular, like the LMC, M33 is either on its first passage (t(inf) < 2 Gyr ago) or if M31 is massive (>= 2 x 10(12) M-circle dot), it is on a long-period orbit of about 6 Gyr. Cosmological analogues of the LMC and M33 identified in Illustris support this picture and provide further insight about their host masses. We conclude that, cosmologically, massive satellites such as the LMC and M33 are likely completing their first orbits about their hosts. We also find that the orbital energies of such analogues prefer an MW halo mass similar to 1.5 x 10(12) M-circle dot and an M31 halo mass >= 1.5 x 10(12)M(circle dot). Despite conventional wisdom, we conclude it is highly improbable that M33 made a close (< 100 kpc) approach to M31 recently (t(peri) < 3 Gyr ago). Such orbits are rare (< 1 per cent) within the 4s error space allowed by observations. This conclusion cannot be explained by perturbative effects through four-body encounters amongst the MW, M31, M33, and the LMC. This surprising result implies that we must search for a new explanation for M33' s strongly warped gas and stellar discs.
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Feiger, Martin Earl. "An evaluation of semianlaytical satellite theory against long arcs of real data for highly eccentric orbits." Thesis, Massachusetts Institute of Technology, 1987. http://hdl.handle.net/1721.1/14905.
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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1987.<br>MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERO.<br>Bibliography: p. 209-212.<br>by Martin Earl Fieger.<br>M.S.
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Mouret, Serge. "Investigations on the dynamics of minor planets with Gaia : Orbits, masses and fundamental physics." Observatoire de Paris (1667-....), 2007. https://hal.science/tel-02071419.
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La mission de l’agence spatiale européenne Gaia, dont le lancement est prévu en 2011, observera 350 000 astéroïdes avec une précision inégalée (au niveau du milli-arseconde). Le but de la thèse est l’investigation de la science que les données très précises de Gaia permettront de traiter et l’élaboration des méthodes de réduction notamment pour la détermination de masses des astéroïdes. Dans le premier chapitre, nous présentons la mission Gaia et une synthèse de la détermination de masses des astéroïdes jusqu’à nos jours. La seconde partie décrit la sélection des passages proches entre les astéroïdes massifs (perturbateurs) et les plus petits (cible) afin d’analyser les perturbations orbitales résultantes, qui permettront d’estimer la masse des perturbateurs à partir d’un traitement global, détaillé dans le chapitre 3. Le nombre de masses attendues et la précision de leur mesure sont donnés à partir de données simulées réalistes. De plus, la précision sur les éléments osculateurs des astéroïdes cibles est présentée. Les chapitres 4 et 5 abordent des questions de physique fondamentale, avec la détermination de l’aplatissement dynamique du soleil J2, des paramètres relativistes β, γ et celui de Nordtvedt η la variation de la constante de gravitation, et aussi de paramètres physiques de certains astéroïdes à partir de leur sensibilité à l’effet Yarkovsky. Les deux derniers chapitres illustrent la contribution de Gaia avec les masses d’astéroïdes dans différents domaines de l’astronomie : une étude a été menée sur les astéroïdes les plus perturbateurs de Mars et sur l’amélioration de la loi hypothétique entre densité et classe taxonomique<br>The ESA astrometric mission Gaia, due for a launch in 2011, will observe 350,000 asteroids brighter than V = 20 with an unprecedented positional precision (at the sub-milliarcsecond level for single observation). The aim of this thesis is the investigation of the science that Gaia will allow with such precise observations of asteroids and the elaboration of reduction methods for treating this future huge pool of data, in particular regarding mass determination. In the first part, we relate the characteristics of the Gaia mission and the state-of-the-art of the current asteroid mass determination. The second part describes the selection of the interesting close approaches between the few massive asteroids (perturbers) and numerous smaller asteroids (targets) in order to analyse the resulting orbital perturbations and thus, to derive the masses of the perturbers from global processing detailed in the third chapter. The expected number of asteroid masses and their associated precisions are given from realistic simulated data between 2011 and 2016. Besides, the expected precisions on the orbital elements of the targets are presented. The fourth and fifth chapters open tracks on the fundamental science from the asteroid positions with the possible derivation of the solar quadrupole moment J2, the PPN parameters β, γ, the Nordtvedt parameter η, the variation of the gravitational constant dG/dt and a study of the so-called Yarkovsky effect. The two last chapters illustrate the contribution of asteroid mass determination by Gaia to several fields of the astronomy through a study about the perturbers of Mars and the hypothetical law between density and taxonomic class
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Du, Toit Daniel N. J. "Low Earth orbit satellite constellation control using atmospheric drag." Thesis, Link to the online version, 1997. http://hdl.handle.net/10019/2999.
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Shabanloui, Akbar [Verfasser]. "A new approach for a kinematic-dynamic determination of low satellite orbits based on GNSS observations / Akbar Shabanloui." Bonn : Universitäts- und Landesbibliothek Bonn, 2019. http://d-nb.info/1199005290/34.
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Svehla, Drazen [Verfasser], Reiner [Akademischer Betreuer] Rummel, Reiner [Gutachter] Rummel, and Markus [Gutachter] Rothacher. "Geometrical Theory of Satellite Orbits and Gravity Field / Drazen Svehla ; Gutachter: Reiner Rummel, Markus Rothacher ; Betreuer: Reiner Rummel." München : Universitätsbibliothek der TU München, 2017. http://d-nb.info/1148650040/34.
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Svehla, Drazen Verfasser], Reinhard [Akademischer Betreuer] [Rummel, Reiner [Gutachter] Rummel, and Markus [Gutachter] Rothacher. "Geometrical Theory of Satellite Orbits and Gravity Field / Drazen Svehla ; Gutachter: Reiner Rummel, Markus Rothacher ; Betreuer: Reiner Rummel." München : Universitätsbibliothek der TU München, 2017. http://nbn-resolving.de/urn:nbn:de:bvb:91-diss-20170529-1355925-1-6.
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Patel, Ekta, Gurtina Besla, and Kaisey Mandel. "Orbits of massive satellite galaxies - II. Bayesian estimates of the Milky Way and Andromeda masses using high-precision astrometry and cosmological simulations." OXFORD UNIV PRESS, 2017. http://hdl.handle.net/10150/624428.
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In the era of high-precision astrometry, space observatories like the Hubble Space Telescope (HST) and Gaia are providing unprecedented 6D phase-space information of satellite galaxies. Such measurements can shed light on the structure and assembly history of the Local Group, but improved statistical methods are needed to use them efficiently. Here we illustrate such a method using analogues of the Local Group's two most massive satellite galaxies, the Large Magellanic Cloud (LMC) and Triangulum (M33), from the Illustris dark-matter-only cosmological simulation. We use a Bayesian inference scheme combining measurements of positions, velocities and specific orbital angular momenta (j) of the LMC/M33 with importance sampling of their simulated analogues to compute posterior estimates of the Milky Way (MW) and Andromeda's (M31) halo masses. We conclude that the resulting host halo mass is more susceptible to bias when using measurements of the current position and velocity of satellites, especially when satellites are at short-lived phases of their orbits (i.e. at pericentre). Instead, the j value of a satellite is well conserved over time and provides a more reliable constraint on host mass. The inferred virial mass of the MW(M31) using j of the LMC (M33) is M-vir,M- MW = 1.02(-0.55)(+0.77) x 10(12) M-circle dot (M-vir,M- M31 = 1.37(-0.75)(+1.39) x 10(12) M-circle dot). Choosing simulated analogues whose j values are consistent with the conventional picture of a previous (<3 Gyr ago), close encounter (<100 kpc) of M33 about M31 results in a very low virial mass for M31 (similar to 10(12) M-circle dot). This supports the new scenario put forth in Patel, Besla & Sohn, wherein M33 is on its first passage about M31 or on a long-period orbit. We conclude that this Bayesian inference scheme, utilizing satellite j, is a promising method to reduce the current factor of 2 spread in the mass range of the MW and M31. This method is easily adaptable to include additional satellites as new 6D phase-space information becomes available from HST, Gaia and the James Webb Space Telescope.
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Carvalho, Jean Paulo dos Santos [UNESP]. "Pertubação orbital devida a um terceiro corpo com distribuição não uniforme de massa e em órbita elíptica." Universidade Estadual Paulista (UNESP), 2011. http://hdl.handle.net/11449/102491.
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carvalho_jps_dr_guara.pdf: 22860427 bytes, checksum: f429c89456cd5f54445e1cf96bf99ff7 (MD5)<br>Fundação de Amparo à Pesquisa do Estado de São Paulo (FAPESP)<br>Neste trabalho, apresentamos uma teoria analítica com simulações numéricas para estudar o movimento orbital de satélites artificiais em torno de satélites planetários. Consideramos o problema de um satélite artificial perturbado pela distribuição não uniforme de massa do corpo principal e por um terceiro corpo (assume-se em uma órbita circular ou elíptica). Polinômios de Legendre são expandidos em potências da excentricidade até o quarto grau e são usados para o potencial perturbador devido ao terceiro corpo. As condições para obter órbitas congeladas são apresentadas. O modelo analítico de média, simples e dupla, é considerado para analisar o movimento orbital dos satélites artificiais. Uma comparação entre os modelos de média, simples e dupla, é apresentada. O método de perturbação de Lie-Hori, até a segunda ordem, é aplicado para eliminar os termos de curto período do potencial perturbador. Termos de acoplamento são analisados. É dada ênfase para o caso de órbitas congeladas, inclinação crítica e ressonâncias. Mostramos uma nova equação aproximada para calcular o semi-eixo maior crítico para a órbita do satélite. Uma abordagem para estudar o comportamento da longitude do nodo ascendente de uma órbita lunar quase polar, hélio-síncrona é apresentada. As simulações numéricas para satélites artificiais hipotéticos são feitas considerando as perturbações acopladas ou isoladas.<br>In this work, we present an analytical theory with numerical simulations to study the orbital motion of artificial satellites around planetary satellites. We consider the problem of an artificial satellite perturbed by the non-uniform distribution of mass of the main body and by a third-body (assumed to be in a circular or elliptical orbit). Legendre polynomials are expanded in powers of the eccentricity up to the degree four and are used for the disturbing potential due to the third-body. The conditions to get frozen orbits are presented. The average analytical model, simple and double, is considered to perform an analysis of the orbital motion of the artificial satellites. A comparison between the averaged models, simple and double, is presented. Lie- Hori perturbation method up to the second-order is applied to eliminate the terms of shortperiod of the disturbing potential. Coupling terms are analyzed. Emphasis is given to the case of frozen orbits, critical inclination and resonances. We show a new approximated equation to compute the critical semi-major axis for the orbit of the satellite. An approach to studying the behavior of the longitude of the ascending node for a near Sun-synchronous polar lunar orbit is presented. Numerical simulations for hypothetical artificial satellites are performed considering the perturbations combined or isolated.
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Rodríguez, Solano Carlos Javier [Verfasser], Urs [Akademischer Betreuer] Hugentobler, Adrian [Akademischer Betreuer] Jäggi, and Oliver [Akademischer Betreuer] Montenbruck. "Impact of non-conservative force modeling on GNSS satellite orbits and global solutions / Carlos Javier Rodríguez Solano. Gutachter: Urs Hugentobler ; Adrian Jäggi ; Oliver Montenbruck. Betreuer: Urs Hugentobler." München : Universitätsbibliothek der TU München, 2014. http://d-nb.info/105985693X/34.
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McNally, Ian J. "Orbital and rotational dynamics of solar power satellites in geosynchronous orbits." Thesis, University of Glasgow, 2018. http://theses.gla.ac.uk/30628/.
Full textAbstract:
Designs for geostationary (GEO) solar power satellites (SPS) are extremely large in scale, more than one order of magnitude larger than the International Space Station. In this thesis a detailed study of the orbit dynamics of SPS is performed. Analytical equations, derived by the process of averaging of the SPS equations of motion, are used to determine the long-term orbital evolution. Previous SPS studies have simply assumed a GEO as the operational orbit, and then designed control systems for maintaining the orbit within acceptable nominal values. It is found that an alternative SPS orbital location known as the geosynchronous Laplace plane orbit (GLPO) is superior to GEO in many aspects. An SPS in GLPO requires virtually no fuel to maintain its orbit, minimises the risk of debris creation at geosynchronous altitude, and is extremely robust operationally, i.e. loss of control is inconsequential. The GLPO SPS requires approximately 10^5 kg less fuel per year compared to a GEO SPS while providing near equivalent power delivery. Although savings in orbit control are achieved, depending on the mass distribution of the SPS, attitude control costs may be incurred by placing an SPS in GLPO. Consideration of the attitude dynamics of SPS has motivated the development of a model for the rotational dynamics of a body which includes energy dissipation and the effects of external torques. Multiple spring-damper masses are used to provide a mechanism for energy dissipation. This rotational dynamics model is used to assess the naturally stable attitude configurations of a SPS design in geosynchronous orbit subject to gravity gradient torque. It is found that for a large planar array, a dynamically stable configuration requiring nominal orbit-attitude control is possible. This involves rotating around the maximum axis of inertia at the orbit rate, with the minimal axis aligned in the radial direction. It will be shown that a SPS in this configuration while in GLPO requires virtually no orbit or attitude control. The most significant result of the research in this thesis is proving that a SPS can operate in GLPO with nominal orbit control and yet still deliver almost equivalent power to the Earth’s surface as the same SPS would in a controlled GEO.
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NASCIMENTO, MARCELLE SANTIAGO DO. "EFFICIENT USE OF THE GEOESTATIONARY SATELLITE ORBIT: ORBITAL POSITION OPTIMIZATION." PONTIFÍCIA UNIVERSIDADE CATÓLICA DO RIO DE JANEIRO, 2005. http://www.maxwell.vrac.puc-rio.br/Busca_etds.php?strSecao=resultado&nrSeq=6709@1.
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COORDENAÇÃO DE APERFEIÇOAMENTO DO PESSOAL DE ENSINO SUPERIOR<br>Este trabalho está relacionado ao problema do uso
eficiente da órbita de
satélite geoestacionário. A utilização eficiente da órbita
é obtida através de
um algoritmo de otimização que permite escolher as
posições orbitais para os
diversos sistemas de modo a reduzir ao máximo o percentual
do arco orbital
utilizado. Sendo assim, desenvolvido um modelo matemático
que considerou
além de aspectos de interferência, detalhes da geometria
envolvida no
problema (posições orbitais dos satélites, posições das
estações terrenas,
apontamento de antenas, etc.). Este modelo foi utilizado
na definição de
um problema de otimização com restrição cuja função
objetivo se baseia
na parcela do arco orbital utilizado. Neste problema de
otimização com
restrição foram consideradas restrições de níveis máximos
de interferência
(de entrada única e agregada) além de restrições de arcos
orbitais, impostas
por aspecto de propagação. O algoritmo de otimização
utilizado requer
o cálculo do Vetor Gradiente e da Matriz Hessiana. Para
evitar erros de
origem numéricos essas quantidades foram calculadas
utilizando expressões
analíticas desenvolvidas neste trabalho. O método
matemático foi aplicado
a situações específicas conduzindo a resultados que
mostraram um uso
eficiente da órbita de satélites geoestacionários através
de soluções onde
a parcela utilizada do arco é minimizada.<br>This work is related to the efficient use of the
geostationary satellite orbit. It
presents and describes an optimization model which chooses
the best orbital
position for each satellite so that the length used
orbital arc is minimized.
A mathematical model considering aspects such as
interference, geometry
details (orbital position of the systems, earth station
position, boresight of
the antenna, etc) is proposed. This model was used in the
definition of a
constrained optimization problem in which the cost
function is the length
of the used orbital arc. Constrained imposed by
propagation aspects (min-
imum elevation angle) and by the maximum allowable
interference levels
(aggregate and single-entry) are considered. The
optimization algorithm re-
quires the evaluation of the Gradient vector and the
Hessian matrix. To
avoid numeric problems, analytic expressions of these
quantities were de-
rived. Results of the application of this model to
specific situations involving
real data were also described and conducted to solutions
where the length
of the orbit used was minimized
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Dutta, Atri. "Optimal cooperative and non-cooperative peer-to-peer maneuvers for refueling satellites in circular constellations." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/28082.
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Thesis (M. S.)--Aerospace Engineering, Georgia Institute of Technology, 2009.<br>Committee Chair: Panagiotis Tsiotras; Committee Member: Eric Feron; Committee Member: Joseph Saleh; Committee Member: Ryan Russell; Committee Member: William Cook
34
Goodson, Troy D. "Numerical solutions to optimal low- and medium-thrust orbit transfers." Thesis, Georgia Institute of Technology, 1993. http://hdl.handle.net/1853/13393.
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Paffett, John. "VHF band interference measurement, analysis and avoidance." Thesis, University of Surrey, 2000. http://epubs.surrey.ac.uk/2105/.
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Dainty, Benjamin G. "Use of two-way time transfer measurements to improve geostationary satellite navigation :." Ft. Belvoir Defense Technical Information Center, 2007. http://handle.dtic.mil/100.2/ADA472457.
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Cutri, Anthony D. "Satellite servicing using the Orbital Maneuvering Vehicle in low earth orbit." Thesis, Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA236941.
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Thesis (M.S. in Systems Technology (Space Systems Operations))--Naval Postgraduate School, June 1990.<br>Thesis Advisor(s): Boger, Dan C. ; Loomis, Herschel H. "June 1990." Description based on signature page as viewed on October 21, 2009. DTIC Identifier(s): Maneuvering satellites, logistics support, space maintenance, OMV (Orbital Maneuvering Vehicles), space vehicles, modular space construction, cost estimates, polar orbit trajectories, space logistics support, theses. Author(s) subject terms: Orbital Maneuvering Vehicle, satellite servicing, polar satellites, expendible launch vehicles, low earth orbit. Includes bibliographical references (p. 106-109). Also available online.
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Zaheer, Muhammad. "Kinematic orbit determination of low Earth orbiting satellites, using satellite-to-satellite tracking data and comparison of results with different propagators." Thesis, KTH, Geodesi och geoinformatik, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-142627.
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The GPS data from Challenging Mini-satellite Payload (CHAMP) is used for its orbit determination for the epoch day of January 1st 2002. The orbit of CHAMP is computed from the GPS data and ionospheric effects are removed by frequency combination. Further, the orbits of CHAMP for the same epoch day are computed using the satellite tool kit (STK) employing simplified general perturbations (SGP4) and a high precision orbit propagator (HPOP). Results from both techniques (GPS computed orbit and STK computed orbit) are compared. Furthermore, orbits computed using GPS data are also compared with jet propulsion laboratory’s published CHAMP spacecraft orbit and we have found that root mean square difference in ECEF position X component is below 0.01km other than some spikes at poles. The standard deviation of the difference in ECEF position X coordinate is 11.7m. The accuracy of our computed satellite positions (using GPS data) is about 12 metres for other than polar areas. However there are some occasional spikes, especially at poles, having maximum errors (about 0.055 km).
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RIBEIRETE, ERICSON CARVALHO. "EFFICIENT USE OF THE GEOESTATIONARY SATELLITE ORBIT THROUGH THE OPTIMIZATION OF ORBITAL LOCATIONS." PONTIFÍCIA UNIVERSIDADE CATÓLICA DO RIO DE JANEIRO, 1987. http://www.maxwell.vrac.puc-rio.br/Busca_etds.php?strSecao=resultado&nrSeq=9477@1.
Full textAbstract:
COORDENAÇÃO DE APERFEIÇOAMENTO DO PESSOAL DE ENSINO SUPERIOR<br>Este trabalho está relacionado ao problema da utilização
eficiente da órbita de satélite geoestacionários. Neste
problema, os satélites devem ser alocados adequadamente na
órbita, obedecendo a limites de interferência máxima, de
modo a satisfazer os requisitos de desempenho dos sistemas.
Algumas técnicas para a determinação das posições orbitais
de satélites existentes e planejados são mencionadas. Um
novo modelo matemático para a otimização das posições
orbitais dos satélites, baseado na minimização de uma
função objetivo associada à interferência total que afeta
cada um dos sistemas, é proposto. Finalmente, alguns
resultados obtidos através da aplicação deste modelo são
analisados e comparados a outros resultados disponíveis.<br>This work is related to the problem of efficiently using
the geoestationary satellite orbit in an interference
limited environment. Some techniques to determine the
orbital positions of existing and planned satellites are
mentioned. A new mathematical model for optimizing
satellite orbital positions based on the minimization of an
objective function associated to the total interference
affecting each system is proposed. Finally, results
obtained through the application of this model are analysed
and compared to other available results.
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Bakari, Salim Rashid. "Solar panel development for high altitude and low earth orbit application." Thesis, Cape Peninsula University of Technology, 2010. http://hdl.handle.net/20.500.11838/2208.
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Thesis (MTech (Electrical Engineering))--Cape Peninsula University of Technology, 2010.<br>Stable and reliable source of electrical energy is a requirement for efficient operation of
satellites. Several sources of electrical power for satellites exist such as fuel cells,
nuclear or battery stored Direct Current energy but of late concentration has been on
solar cells as the advantages compared to the other sources are many. Solar cells are p-n
semiconductor devices which convert light energy into electrical energy by photovoltaic
effect. The biggest drawback of solar cell energy system is the low light to electricity
conversion efficiency. Apart from powering satellites, solar cells and panels have found
other numerous applications such as in water pumping systems, rural electrification,
street lightning. Photovoltaic principle of solar cells started way back in 1839 when
Alexandre Edmund Becquerel observed that electrical currents arose from certain light
induced chemical reactions. A comprehensive understanding of this phenomenon
became clear when the science of quantum theory was unveiled in the early parts of the
20th century. Most solar cells and panels available today in the market are silicon based
made of single junction technology. The disadvantage with single junction technology is
that the p-n junction is made of a single type of solar cell material which absorbs a
fraction of light wavelengths from the spectrum of light. The disability of the single p-n
junction to convert all the light energy to electricity accounts for the low efficiency for the
solar cells. One way to go around the problem of efficiency is to use multi-junction solar cells. Multijunction
solar cells are designed to absorb a large fraction of the light spectrum and
convert them to electrical energy. They are made of multiple p-n junctions made of
different solar cell materials which absorb different parts of light spectrum and convert
them to electrical energy. In this thesis, a design of a multi-junction solar cell for
developing space solar panel is presented. The multi-junction cell has been designed
from simulation results of different solar cell materials simulated with space conditions.
Ideas and recommendations for future work are also presented.
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SYED, ANEES. "COLLISON PREDICTION AND AVOIDANCE OF SATELLITES IN FORMATION." University of Cincinnati / OhioLINK, 2004. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1100034591.
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Afful, Michael Andoh. "Orbital lifetime predictions of Low Earth Orbit satellites and the effect of a DeOrbitSail." Thesis, Stellenbosch : Stellenbosch University, 2013. http://hdl.handle.net/10019.1/85862.
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Thesis (MEng)--Stellenbosch University, 2013.<br>ENGLISH ABSTRACT: Throughout its lifetime in space, a spacecraft is exposed to risk of collision with orbital
debris or operational satellites. This risk is especially high within the Low Earth Orbit
(LEO) region where the highest density of space debris is accumulated.
This study investigates orbital decay of some LEO micro-satellites and accelerating orbit
decay by using a deorbitsail. The Semi-Analytical Liu Theory (SALT) and the Satellite
Toolkit was employed to determine the mean elements and expressions for the time rates
of change. Test cases of observed decayed satellites (Iridium-85 and Starshine-1) are used
to evaluate the predicted theory. Results for the test cases indicated that the theory tted
observational data well within acceptable limits.
Orbit decay progress of the SUNSAT micro-satellite was analysed using relevant orbital
parameters derived from historic Two Line Element (TLE) sets and comparing with decay
and lifetime prediction models. The study also explored the deorbit date and time for a
1U CubeSat (ZACUBE-01).
A proposed orbital debris solution or technology known as deorbitsail was also investigated
to gain insight in sail technology to reduce the orbit life of spacecraft with regards to de-
orbiting using aerodynamic drag. The deorbitsail technique signi cantly increases the
e ective cross-sectional area of a satellite, subsequently increasing atmospheric drag and
accelerating orbit decay. The concept proposed in this work introduces a very useful
technique of orbit decay as well as deorbiting of spacecraft.<br>AFRIKAANSE OPSOMMING: Gedurende sy leeftyd in die ruimte word 'n ruimtetuig blootgestel aan die risiko van 'n
botsing met ruimterommel of met funksionele satelliete. Hierdie risiko is veral hoog in die
lae-aardbaan gebied waar die hoogste digtheid ruimterommel voorkom.
Hierdie studie ondersoek die wentelbaanverval van sommige Lae-aardbaan mikrosatelliete
asook die versnelde baanverval wanneer van 'n deorbitaal meganisme gebruik gemaak word.
Die Semi-Analitiese Liu Teorie en die Satellite Toolkit sagtewarepakket is gebruik om die
gemiddelde baan-elemente en uitdrukkings vir hul tyd-afhanlike tempo van verandering
te bepaal. Toetsgevalle van waargenome vervalde satelliete (Iridium-85 en Starshine-1) is
gebruik om die verloop van die voorspelde teoretiese verval te evalueer. Resultate vir die
toetsgevalle toon dat die teorie binne aanvaarbare perke met die waarnemings ooreenstem.
Die verloop van die SUNSAT mikrosatelliet se wentelbaanverval is ook ontleed deur gebruik
te maak van historiese Tweelyn Elemente datastelle en dit te vergelyk met voorspelde baan-
elemente. Die studie het ook ondersoek ingestel na die voorspelde baan-verbyval van 'n
1-eenheid cubesat (ZACUBE-01).
Die impak op wentelbaanverval deur 'n voorgestelde oplossing vir die beperking van
ruimterommel, 'n deorbitaalseil, is ook ondersoek. So seil verkort 'n satelliet se ruimte-
leeftyd deur sy e ektiewe deursnee-area te vergroot en dan van verhoogde atmosferiese
sleur en sonstralingsdruk gebruik te maak om die vervalproses te versnel. Hierdie voorgestelde
konsep is 'n moontlike nuttige tegniek vir versnelde baanverval en beheerde deorbitalering
van ruimtetuie om ruimterommel te verminder.
43
McCaine, Gina. "Halo orbit design and optimization." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2004. http://library.nps.navy.mil/uhtbin/hyperion/04Mar%5FMcCaine.pdf.
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Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, March 2004.<br>Thesis advisor(s): I. Michael Ross, Don Danielson. Includes bibliographical references (p. 39-40). Also available online.
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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.<br>UWE-4
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Nagarajan, N. "Autonomous Orbit Estimation For Near Earth Satellites Using Horizon Scanners." Thesis, Indian Institute of Science, 1994. http://hdl.handle.net/2005/155.
Full textAbstract:
Autonomous navigation is the determination of satellites position and velocity vectors onboard the satellite, using the measurements available onboard. The orbital information of a satellite needs to be obtained to support different house keeping operations such as routine tracking for health monitoring, payload data processing and annotation, orbit manoeuver planning, and prediction of intrusion in various sensors' field of view by celestial bodies like Sun, Moon etc. Determination of the satellites orbital parameters is done in a number of ways using a variety of measurements. These measurements may originate from ground based systems as range and range rate measurements, or from another satellite as in the case of GPS (Global Positioning System) and TDUSS (Tracking Data Relay Satellite Systems), or from the same satellite by using sensors like horizon sensor^ sun sensor, star tracker, landmark tracker etc. Depending upon the measurement errors, sampling rates, and adequacy of the estimation scheme, the navigation accuracy can be anywhere in the range of 10m - 10 kms in absolute location.
A wide variety of tracking sensors have been proposed in the literature for autonomous navigation. They are broadly classified as (1) Satellite-satellite tracking, (2) Ground- satellite tracking, (3) fully autonomous tracking. Of the various navigation sensors, it may be cost effective to use existing onboard sensors which are well proven in space. Hence, in the current thesis, the Horizon scanner is employed as the primary navigation sensor-. It has been shown in the literature that by using horizon sensors and gyros, a high accuracy pointing of the order of .01 - .03 deg can be achieved in the case of low earth orbits. Motivated by such a fact, the current thesis deals with autonomous orbit determination using measurements from the horizon sensors with the assumption that the attitude is known to the above quoted accuracies.
The horizon scanners are mounted on either side of the yaw axis in the pitch yaw plane at an angle of 70 deg with respect to the yaw axis. The Field Of View (FOV) moves about the scanner axis on a cone of 45 deg half cone angle. During each scan, the FOV generates two horizon points, one at the space-Earth entry and the other at the Earth-space exit. The horizon points, therefore, lie• on the edge of the Earth disc seen by the satellite. For a spherical earth, a minimum of three such horizon points are needed to estimate the angular radius and the center of the circular horizon disc. Since a total of four horizon points are available from a pair of scanners, they can be used to extract the satellite-earth distance and direction.These horizon points are corrupted by noise due to uncertainties in the Earth's radiation pattern, detector mechanism, the truncation and roundoff errors due to digitisation of the measurements. Owing to the finite spin rate of the scanning mechanism, the measurements are available at discrete time intervals. Thus a filtering algorithm with appropriate state dynamics becomes essential to handle the •noise in the measurements, to obtain the best estimate and to propagate the state between the measurements. The orbit of a low earth satellite can be represented by either a state vector (position and velocity vectors in inertial frame) or Keplerian elements. The choice depends upon the available processors, functions and the end use of the estimated orbit information. It is shown in the thesis that position and velocity vectors in inertial frame or the position vector in local reference frame, do result in a simplified, state representation. By using the f and g series method for inertial position and velocity, the state propagation is achieved in linear form.
i.e. Xk+1 = AXK
where X is the state (position, velocity) and A the state transition matrix derived from 'f' and 'g' series. The configuration of a 3 axis stabilised spacecraft with two horizon scanners is used to simulate the measurements.
As a step towards establishing the feasibility of extracting the orbital parameters, the governing equations are formulated to compute the satellite-earth vector from the four horizon points generated by a pair of Horizon Scanners in the presence of measurement noise. Using these derived satellite-earth vectors as measurements, Kalman filter equations are developed, where both the state and measurements equations are linear. Based on simulations, it is shown that a position accuracy of about 2 kms can be achieved. Additionally, the effect of sudden disturbances like substantial slewing of the solar panels prior and after the payload operations are also analysed. It is shown that a relatively simple Low Pass Filter (LPF) in the measurements loop with a cut-off frequency of 10 Wo (Wo = orbital frequency) effectively suppresses the high frequency effects from sudden disturbances which otherwise camouflage the navigational information content of the signal. Then Kalman filter can continue to estimate the orbit with the same kind of accuracy as before without recourse to re-tuning of covariance matrices.
Having established the feasibility of extracting the orbit information, the next step is to treat the measurements in its original form, namely, the non-linear form. The entry or exit timing pulses generated by the scanner when multiplied by the scan rate yield entry or exit azimuth angles in the scanner frame of reference, which in turn represents an effective measurement variable. These azimuth angles are obtained as inverse trigonometric functions of the satellite-earth vector. Thus the horizon scanner measurements are non-linear functions of the orbital state. The analytical equations for the horizon points as seen in the body frame are derived, first for a spherical earth case. To account for the oblate shape of the earth, a simple one step correction algorithm is developed to calculate the horizon points. The horizon points calculated from this simple algorithm matches well with the ones from accurate model within a bound of 5%. Since the horizon points (measurements) are non-linear functions of the state, an Extended Kalman Filter (EKF) is employed for state estimation. Through various simulation runs, it is observed that the along track state has got poor observability when the four horizon points are treated as measurements in their original form, as against the derived satellite-earth vector in the earlier strategy. This is also substantiated by means of condition number of the observability matrix. In order to examine this problem in detail, the observability of the three modes such as along-track, radial, and cross-track components (i.e. the local orbit frame of reference) are analysed. This difficulty in observability is obviated when an additional sensor is used in the roll-yaw plane. Subsequently the simulation studies are carried out with two scanners in pitch-yaw plane and one scanner in the roll-yaw plane (ie. a total of 6 horizon points at each time). Based on the simulations, it is shown that the achievable accuracy in absolute position is about 2 kms.- Since the scanner in the roll-yaw plane is susceptible to dazzling by Sun, the effect of data breaks due to sensor inhibition is also analysed. It is further established that such data breaks do not improve the accuracy of the estimates of the along-track component during the transient phase. However, filter does not diverge during this period.
Following the analysis of the' filter performance, influence of Earth's oblateness on the measurement model studied. It is observed that the error in horizon points, due to spherical Earth approximation behave like a sinusoid of twice the orbital frequency alongwith a bias of about 0.21° in the case of a 900 kms sun synchronous orbit. The error in the 6 horizon points is shown to give rise to 6 sinusoids. Since the measurement model for a spherical earth is the simplest one, the feasibility of estimating these sinusoids along with the orbital state forms the next part of the thesis. Each sinusoid along with the bias is represented as a 3 state recursive equation in the following form
where i refers to the ith sinusoid and T the sampling interval. The augmented or composite state variable X consists of bias, Sine and Cosine components of the sinusoids. The 6 sinusoids together with the three dimensional orbital position vector in local coordinate frame then lead to a 21 state augmented Kalman Filter. With the 21 state filter, observability problems are experienced. Hence the magnetic field strength, which is a function of radial distance as measured by an onboard magnetometer is proposed as additional measurement. Subsequently, on using 6 horizon point measurements and the radial distance measurements obtained from a magnetometer and taking advantage of relationships between sinusoids, it is shown that a ten state filter (ie. 3 local orbital states, one bias and 3 zero mean sinusoids) can effectively function as an onboard orbit filter. The filter performance is investigated for circular as well as low eccentricity orbits. The 10-state filter is shown to exhibit a lag while following the radial component in case of low eccentricity orbits. This deficiency is overcome by introducing two more states, namely the radial velocity and acceleration thus resulting in a 12-state filter. Simulation studies reveal that the 12-state filter performance is very good for low eccentricity orbits. The lag observed in 10-state filter is totally removed. Besides, the 12-state filter is able to follow the changes in orbit due to orbital manoeuvers which are part of orbit acquisition plans for any mission.
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Su, Hua. "Precise orbit determination of global navigation satellite system of second generation (GNSS-2) orbit determination of IGSO, GEO and MEO satellites /." [S.l.] : [s.n.], 2000. http://deposit.ddb.de/cgi-bin/dokserv?idn=962099635.
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Carnochan, Stuart. "Orbit and altimetric corrections for the ERS satellites through analysis of single and dual satellite crossovers." Thesis, Aston University, 1996. http://publications.aston.ac.uk/14262/.
Full textAbstract:
Due to the failure of PRARE the orbital accuracy of ERS-1 is typically 10-15 cm radially as compared to 3-4cm for TOPEX/Poseidon. To gain the most from these simultaneous datasets it is necessary to improve the orbital accuracy of ERS-1 so that it is commensurate with that of TOPEX/Poseidon. For the integration of these two datasets it is also necessary to determine the altimeter and sea state biases for each of the satellites. Several models for the sea state bias of ERS-1 are considered by analysis of the ERS-1 single satellite crossovers. The model adopted consists of the sea state bias as a percentage of the significant wave height, namely 5.95%. The removal of ERS-1 orbit error and recovery of an ERS-1 - TOPEX/Poseidon relative bias are both achieved by analysis of dual crossover residuals. The gravitational field based radial orbit error is modelled by a finite Fourier expansion series with the dominant frequencies determined by analysis of the JGM-2 co-variance matrix. Periodic and secular terms to model the errors due to atmospheric density, solar radiation pressure and initial state vector mis-modelling are also solved for. Validation of the dataset unification consists of comparing the mean sea surface topographies and annual variabilities derived from both the corrected and uncorrected ERS-1 orbits with those derived from TOPEX/Poseidon. The global and regional geographically fixed/variable orbit errors are also analysed pre and post correction, and a significant reduction is noted. Finally the use of dual/single satellite crossovers and repeat pass data, for the calibration of ERS-2 with respect to ERS-1 and TOPEX/Poseidon is shown by calculating the ERS-1/2 sea state and relative biases.
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Wolf, Robert. "Satellite orbit and ephemeris determination using inter satellite links." [S.l.] : [s.n.], 2001. http://deposit.ddb.de/cgi-bin/dokserv?idn=961611820.
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Zhou, Ning. "Onboard orbit determination using GPS measurements for low Earth orbit satellites." Thesis, Queensland University of Technology, 2005. https://eprints.qut.edu.au/16076/1/Ning_Zhou_Thesis.pdf.
Full textAbstract:
Recent advances in spaceborne GPS technology have shown significant advantages in many aspects over conventional technologies. For instance, spaceborne GPS can realize autonomous orbit determination with significant savings in spacecraft life cycle, in power, and in mass. At present, the onboard orbit determination in real time or near-real time can typically achieve 3D orbital accuracy of metres to tens metres with Kalman filtering process, but 21st century space engineering requires onboard orbit accuracy of better than 5 metres, and even sub-metre for some space applications. The research focuses on the development of GPS-based autonomous orbit determination techniques for spacecraft. Contributions are made to the field of GPS-based orbit determination in the following five areas:
Techniques to simplify the orbital dynamical models for onboard processing have been developed in order to reduce the computional burden while retaining full model accuracy. The Earth gravity acceleration approximation method was established to replace the traditional recursive acceleration computations. Results have demonstrated that with the computation burden for a 55× spherical harmonic gravity model, we achieve the accuracy of a 7070× model. Efforts were made for the simplification of solar & lunar ephemerides, atmosphere density model and orbit integration. All these techniques together enable a more accurate orbit integrator to operate onboard.
Efficient algorithms for onboard GPS measurement outlier detection and measurement improvement have been developed. In addition, a closed-form single point position method was implemented to provide an initial orbit solution without any a priori information.
The third important contribution was made to the development of sliding-window short-arc orbit filtering techniques for onboard processing. With respect to the existing Kalman recursive filtering, the short-arc method is more stable because more measurements are used. On the other hand, the short-arc method requires less accurate orbit dynamical model information compared to the long-arc method, thus it is suitable for onboard processing. Our results have demonstrated that by using the 1 ~ 2 revolutions of LEO code GPS data we can achieve an orbit accuracy of 1 ~ 2 metres. Sliding-window techniques provide sub-metre level orbit determination solutions with 5~20 minutes delay.
A software platform for the GPS orbit determination studies has been established. Methods of orbit determination in near-real time have been developed and tested. The software system includes orbit dynamical modelling, GPS data processing, orbit filtering and result analysis modules, providing an effective technical basis for further studies.
Furthermore a ground-based near-real time orbit determination system has been established for FedSat, Australia's first satellite in 30 years. The system generates 10-metre level orbit solution with half-day latency on an operational basis. This system has supported the scientific missions of FedSat such as Ka-band tracking and GPS atmosphere studies within the Cooperative Research Centre for Satellite System (CRCSS) community. Though it is different from the onboard orbit determination, it provides important test-bed for the techniques described in previous section.
This thesis focuses on the onboard orbit determination techniques that were discussed in Chapter 2 through Chapter 6. The proposed onboard orbit determination algorithms were successfully validated using real onboard GPS data collected from Topex/Poseidon, CHAMP and SAC-C satellites.
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Zhou, Ning. "Onboard Orbit Determination Using GPS Measurements for Low Earth Orbit Satellites." Queensland University of Technology, 2005. http://eprints.qut.edu.au/16076/.
Full textAbstract:
Recent advances in spaceborne GPS technology have shown significant advantages in many aspects over conventional technologies. For instance, spaceborne GPS can realize autonomous orbit determination with significant savings in spacecraft life cycle, in power, and in mass. At present, the onboard orbit determination in real time or near-real time can typically achieve 3D orbital accuracy of metres to tens metres with Kalman filtering process, but 21st century space engineering requires onboard orbit accuracy of better than 5 metres, and even sub-metre for some space applications. The research focuses on the development of GPS-based autonomous orbit determination techniques for spacecraft. Contributions are made to the field of GPS-based orbit determination in the following five areas: Techniques to simplify the orbital dynamical models for onboard processing have been developed in order to reduce the computional burden while retaining full model accuracy. The Earth gravity acceleration approximation method was established to replace the traditional recursive acceleration computations. Results have demonstrated that with the computation burden for a 55× spherical harmonic gravity model, we achieve the accuracy of a 7070× model. Efforts were made for the simplification of solar & lunar ephemerides, atmosphere density model and orbit integration. All these techniques together enable a more accurate orbit integrator to operate onboard. Efficient algorithms for onboard GPS measurement outlier detection and measurement improvement have been developed. In addition, a closed-form single point position method was implemented to provide an initial orbit solution without any a priori information. The third important contribution was made to the development of sliding-window short-arc orbit filtering techniques for onboard processing. With respect to the existing Kalman recursive filtering, the short-arc method is more stable because more measurements are used. On the other hand, the short-arc method requires less accurate orbit dynamical model information compared to the long-arc method, thus it is suitable for onboard processing. Our results have demonstrated that by using the 1 ~ 2 revolutions of LEO code GPS data we can achieve an orbit accuracy of 1 ~ 2 metres. Sliding-window techniques provide sub-metre level orbit determination solutions with 5~20 minutes delay. A software platform for the GPS orbit determination studies has been established. Methods of orbit determination in near-real time have been developed and tested. The software system includes orbit dynamical modelling, GPS data processing, orbit filtering and result analysis modules, providing an effective technical basis for further studies. Furthermore a ground-based near-real time orbit determination system has been established for FedSat, Australia's first satellite in 30 years. The system generates 10-metre level orbit solution with half-day latency on an operational basis. This system has supported the scientific missions of FedSat such as Ka-band tracking and GPS atmosphere studies within the Cooperative Research Centre for Satellite System (CRCSS) community. Though it is different from the onboard orbit determination, it provides important test-bed for the techniques described in previous section. This thesis focuses on the onboard orbit determination techniques that were discussed in Chapter 2 through Chapter 6. The proposed onboard orbit determination algorithms were successfully validated using real onboard GPS data collected from Topex/Poseidon, CHAMP and SAC-C satellites.