Dissertations / Theses on the topic 'Circular Restricted Three Body Problem'

Near-Earth Asteroids (NEAs) can provide useful resources in terms of feedstock for spacecraft propellant, crew logistic support and a range of useful metals. The possibility of capturing small NEAs using low energy transfers would therefore be of significant scientific and commercial interest. Although NEAs may make close approaches to the Earth, and so represent a potential impact threat, the exploitation of their resources has long been proposed as a necessary element for future space exploration. The objective of the research presented in this thesis is to develop methodologies for the trajectory design of capturing NEAs in the neighbourhood of the Earth. Firstly aimed at capturing NEAs around the Earth-Moon L2 point, a new type of lunar asteroid capture is defined, termed direct capture. In this capture strategy, the transfer trajectory for capturing an NEA into the Earth-Moon system is modelled in the Sun-Earth-Moon restricted four-body. A Lambert arc in the Sun-asteroid two-body problem is used as an initial guess and a differential corrector used to generate the transfer trajectory from the asteroid’s initial obit to the stable manifold associated with Earth-Moon L2 point. The direct asteroid capture strategy requires a shorter flight time compared to an indirect asteroid capture strategy, which couples capture in the Sun-Earth circular restricted three-body problem and subsequent transfer to the Earth-Moon circular restricted three-body problem. Finally, the direct and indirect asteroid capture strategies are also applied to consider capture of asteroids at the triangular libration points in the Earth-Moon system. As ideal locations for space science missions and candidate gateways for future crewed interplanetary missions, the Sun-Earth libration points L1 and L2 are also preferred locations for the captured asteroids. Therefore, the concept of coupling together a flyby of the Earth and then capturing small NEAs onto Sun–Earth L1 or L2 periodic orbits is proposed. A periapsis map is then employed to determine the required perigee of the Earth flyby. Moreover, depending on the perigee distance of the flyby, Earth flybys with and without aerobraking are investigated to design a transfer trajectory capturing a small NEA from its initial orbit to the stable manifolds associated with Sun-Earth L1 and L2 periodic orbits. NEA capture strategies using an Earth flyby with and without aerobraking both have the potential to be of lower cost in terms of energy requirements than a direct NEA capture strategy without the Earth flyby. Moreover, NEA capture with an Earth flyby also has the potential for a shorter flight time compared to the NEA capture strategy without the Earth flyby. Following by this work, a more general analysis of aerobraking is undertaken and the low energy capture of near-Earth asteroids into bound orbits around the Earth using aerobraking is then investigated. Two asteroid capture strategies utilizing aerobraking are defined, termed single-impulse capture and bi-impulse capture, corresponding to two approaches to raising the perigee height of the captured asteroid’s orbit after the aerobraking manoeuvre. A Lambert arc in the Sun-asteroid two-body problem is again used as an initial estimate for the transfer trajectory to the Earth and then a global optimization is undertaken, using the total transfer energy cost and the retrieved asteroid mass ratio (due to ablation) as objective functions. It is shown that aerobraking can in principle enable candidate asteroids to be captured around the Earth with, in some cases, extremely low energy requirements. The momentum exchange theory is also applied to the capture of small near-Earth asteroids into bound periodic orbits at the Sun-Earth L1 and L2 points. A small asteroid is first manoeuvred to engineer a flyby with a larger asteroid. Two strategies are then considered: when the small asteroid approaches the vicinity of the large asteroid, it will either impact the large asteroid or connect to it with a tether. In both strategies, momentum exchange can be used to effect the capture of one of the asteroids. Then, a two-impulse Lambert arc is utilized to design a post-encounter transfer trajectory to the stable manifolds of the Sun-Earth L1 or L2 points. By investigating the outcome of the impact on the small asteroid, or the tension of the tether, the maximum velocity increment available using these momentum exchange strategies is investigated. Again the capture strategies using momentum exchange in principle have the potential to deliver low-energy capture of asteroids. The methods presented in this thesis are intended to be used as a preliminary analysis for these asteroid capture strategies. Although some significant practical challenges remain, the transfer in the CRTBP models can serve as a good approximation for the trajectory in a more accurate dynamical model.

Le premier objectif de cette thèse est de bien comprendre les propriétés de la dynamique du problème circulaire restreint des trois corps et de les utiliser pour calculer des missions pour satellites pourvus de moteurs à faible poussée. Une propriété fondamentale est l'existence de variétés invariantes associées à des orbites périodiques autour des points de \bsc{Lagrange}. En suivant l'idée de l'\emph{Interplanetary Transport Network}, la connaissance et le calcul des variétés invariantes, comme courants gravitationnels, sont cruciaux pour le \emph{design} de missions spatiales. Une grande partie de ce travail de thèse est consacrée au développement de méthodes numériques pour calculer le transfert entre variétés invariantes de façon optimale. Le coût que l'on cherche alors à minimiser est la norme $L^{1}$ du contrôle car elle est équivalente à minimiser la consommation des moteurs. On considère aussi la norme $L^{2}$ du contrôle car elle est, numériquement, plus facile à minimiser. Les méthodes numériques que nous utilisons sont des méthodes indirectes rendues plus robustes par des méthodes de continuation sur le coût, sur la poussée, et sur l'état final. La mise en œuvre de ces méthodes repose sur l'application du Principe du Maximum de Pontryagin. Les algorithmes développés dans ce travail permettent de calculer des missions réelles telles que des missions entre des voisinages des points de \bsc{Lagrange}. L'idée principale est d'initialiser un tir multiple avec une trajectoire admissible composée de parties contrôlées (des transferts locaux) et de parties non-contrôlées suivant la dynamique libre (les variétés invariantes). Les méthodes mises au point ici, sont efficaces et rapides puisqu'il suffit de quelques minutes pour obtenir la trajectoire optimale complète. Enfin, on développe une méthode hybride, avec à la fois des méthodes directes et indirectes, qui permettent d'ajuster la positions des points de raccord sur les variétés invariantes pour les missions à grandes variations d'énergie. Le gradient de la fonction valeur est donné par les valeurs des états adjoints aux points de raccord et donc ne nécessite pas de calculs supplémentaire. Ainsi, l'implémentation de algorithme du gradient est aisée
The first objective of this work is to understand the dynamical properties of the circular restricted three body problem in order to use them to design low consumption missions for spacecrafts with a low thrust engine. A fundamental property is the existence of invariant manifolds associated with periodic orbits around Lagrange points. Following the Interplanetary Transport Network concept, invariant manifolds are very useful to design spacecraft missions because they are gravitational currents. A large part of this work is devoted to designing a numerical method that performs an optimal transfer between invariant manifolds. The cost we want to minimize is the $L^{1}$-norm of the control which is equivalent to minimizing the consumption of the engines. We also consider the $L^{2}$-norm of the control which is easier to minimize numerically. The numerical methods are indirect ones coupled with different continuations on the thrust, on the cost, and on the final state, to provide robustness. These methods are based on the application of the Pontryagin Maximum Principal. The algorithms developed in this work allow for the design of real life missions such as missions between the realms of libration points. The basic idea is to initialize a multiple shooting method with an admissible trajectory that contains controlled parts (local transfers) and uncontrolled parts following the natural dynamics (invariant manifolds). The methods developed here are efficient and fast (less than a few minutes to obtain the whole optimal trajectory). Finally, we develop a hybrid method, with both direct and indirect methods, to adjust the position of the matching points on the invariant manifolds for missions with large energy gaps. The gradient of the value function is given by the values of the costates at the matching points and does not require any additional computation. Hence, the implementation of the gradient descent is easy

The goal of this thesis is to show global instability or Arnold's diffusion in the elliptic restricted three body problem (ERTBP) by proving the existence of pseudo-trajectories diffusing along the phase space for certain ranges of the eccentricity of the primaries (e), the angular momentum of the comet (G) and the parameter of mass (µ). More precisely, the results presented in his thesis, are valid for G big enough, eG bounded and µ small enough. The thesis is divided in two chapters and two appendices. The chapter one, contains all the main results. After introducing the ERTBP, we use McGehee coordinates to define the infinity manifold, which turn to be a three dimensional invariant manifold in the extended phase space which behaves topologically as a Normally Hyperbolic Invariant Manifold (NHIM), although it is of parabolic type. This means that the rate of approach and departure from it along its invariant manifolds is polynomial in time, instead of exponential-like as happens in a standard NHIM. On the other hand, the inner dynamics is trivial, since it is formed by a two-parameter family of 2p-periodic orbits in the 5D extended phase space which correspond to constant solutions in the 4D phase space. As a consequence, the stable and unstable manifold of the infinity manifold are union of the stable and unstable manifolds of its periodic orbits, and as long as these manifolds intersect along transversal heteroclinic orbits, the scattering map can be defined, as De la Llave, Seara and Delshams did. Unfortunately, since the inner dynamics of the infinity manifold is so simple, the classical mechanisms of diffusion, consisting of combining the inner and outer dynamics, do not work here. Instead, as a novelty, we will be able to find two different scattering maps which will be combined in a suitable way to provide orbits whose angular momentum increases. The asymptotic formula of the scattering map relies entirely in the computation of the so called Menikov potential as defined in the works of Delshams, Gutiérrez and Seara. The first derivative of the Melnikov potential gives the first order approximation of the distance between the stable and unstable invariant manifolds of the infinity manifold whenever the parameter of mass is exponentially small. Given this setting, a series of lemmas and propositions will lead to a formula of the dominant terms of this Melnikov potential. The key idea is to compute its Fourier coefficients which will be exponentially small when the angular momentum is large and an explicit formula will be not possible, therefore and effective computation will be necessary. To do so the product eG will play a key role which lead to theorems 1.5 and 1.6, the former gives an asymptotic formula for the Melnikov potential whenever eG is samll, and the latter whenever eG is finite. Both of them requires µ to be exponentially small with respect to G, and G to be big enough. These theorems naturally produce asymptotic formulas for the scattering maps in both cases and are the base for theorems 1.15 and 1.16 which formulate the existence of pseudo-trajectories in the ERTBP. In chapter two, we provide the details and the proofs of the results concerning the asymptotic formulas, given in chapter one, for the Melnikov potential and the scattering maps, including effective bounds of every error function involved. The appendices have the more technical results needed to complete in a rigorous way every proof, but because of its nature, can be relegated to the end, to make easier to follow up the main proofs.
El objetivo de esta tesis es mostrar inestabilidad global o difusión de Arnold en el problema restringido de tres cuerpos elíptico (PTCRE) mostrando la existencia de pseudo-trayectorias difusivas en el espacio fase para ciertos rangos de la excentricidad (e), el momento angular del cometa (G) y el parámetro de masa (µ). Mas precisamente, los resultados presentados, son válidos para G suficientemente grande, eG acotado y µ suficientemente pequeño. La tesis está dividida en dos capítulos y dos apéndices. El capítulo 1, contiene todos los resultados principales. Después de introducir el PTCRE, usamos coordenadas de McGehee para definir la variedad de infinito, que será de dimensión tres en el espacio fase extendido y que topológicamente se comporta como una variedad invariante normalmente hiperbólica (NHIM), aunque es de tipo parabólico. Esto significa que la tasa de acercamiento y alejamiento de ella a lo largo de sus variedades invariantes es polinomial, en lugar de exponencial como sucede en una NHIM estándar. Por otra parte, la dinámica interior es trivial ya que está formada por una familia de orbitas con 2 parámetros y de período 2p en el espacio extendido 5D que corresponden a soluciones constantes en el espacio reducido 4D. Como consecuencia, las variedades estables e inestables de la variedad de infinito son la unión de las variedades estables e inestables de sus orbitas periódicas y siempre que estas variedades se intersequen sobre orbitas heteroclínicas transversales, el scattering map puede ser definido como hicieron De la Llave, Seara y Delshams . Desafortunadamente, ya que la dinámica interior de la variedad de infinito es muy simple, el mecanismo de difusión clásico, que consiste en combinar la dinámica interior con la exterior, no funciona aquí. En su lugar, como una novedad, seremos capaces de encontrar dos scattering maps diferentes que serán combinados de manera adecuada para producir orbitas cuyo momento angular crezca. La fórmula asintótica del scattering map recae enteramente en el cálculo del llamado potencial de Melnikov, como es definido en los trabajos de Delshams, Gutiérrez y Seara. La primer derivada del potencial de Melnikov da la aproximación a primer orden de la distancia entre las variedades estable e inestable de la variedad de infinito cuando el parámetro de masa es exponencialmente pequeño. Con este planteamiento, una serie de lemas y proposiciones conducirán a la fórmula de los términos dominantes del potencial de Melnikov. La idea clave es calcular sus coeficientes de Fourier, que serán exponencialmente pequeños cuando el momento angular es grande y una fórmula explícita no será posible, así que un cálculo efectivo será necesario. Para hacerlo, el producto eG jugará un papel clave que conducirá a los teoremas 1.5 y 1.6, el primero da una fórmula asintótica del potencial de Melnikov cuando eG es pequeño y el segundo cuando eG es finito. Ambos requieren que µ sea exponencialmente pequeño con respecto a G, y G suficientemente grande. Estos teoremas naturalmente producirán las fórmulas asintóticas de los scattering maps para ambos casos y son la base de los teoremas 1.15 y 1.16, que formulan la existencia de pseudo-trayectorias en el PTCRE. En el capítulo 2, damos los detalles y las pruebas de los resultados concernientes a las formulas asintóticas, dadas en el capítulo 1, para el potencial de Melnikov y los scattering maps, incluyendo las cotas efectivas de cada error involucrado. Los apéndices tienen los resultados mas técnicos que son necesarios para completar de forma rigurosa cada prueba, pero que por su naturaleza, pueden ser relegados al final para hacer seguir las pruebas con mas facilidad.

The main objective of this dissertation is the study of the ejection-collision (EC) orbits in the circular and planar Restricted Three Body Problem (RTBP from now on). In particular, we will focus on the analytical and numerical study of a very specific type of EC orbits, that we denote as n-EC orbits. An n-EC orbit is an orbit such that the particle ejects from one primary and reaches n times a relative maximum in the distance with respect to the primary from which it ejected before colliding with it. In this way, we will study numerically in depth this kind of orbits and we will show analytically that for a sufficiently large value of the Jacobi constant (for which we will give an expression in terms of the mass parameter and the value of n) there exist exactly four n-EC orbits with well-defined characteristics. These results generalize and improve the previous results for the particular case of n=1, and we will see that they can be easily extrapolated to the Hill problem. Besides, we will observe numerically that the evolution of these four original families of n-EC orbits present a very rich dynamics.It is well-known that the system that defines the motion of the particle is not well defined at the points where the primaries are located. For this reason, we have used two different techniques to regularize the collision, the McGehee regularization and the Levi-Civita regularization. Thus, in this dissertation we have analyzed the advantages and disadvantages of each regularization and the different methods that can be used to detect collisions. Since this dissertation will be mainly focused on values of the Jacobi constant greater than those associated to the equilibrium point L1, these two local regularizations will be enough. For less restrictive values of the Jacobi constant we will see that there exist other global regularizations or alternatively, we can simply work with local regularizations in a neighbourhood of each primary.On the other hand, from the numerical point of view we have analyzed the global behaviour of the ejection orbits in the RTBP. We have studied the relation between the family of Lyapunov periodic orbits around the equilibrium point L1 and the ejection orbits for values of the Jacobi constant such that the associated Hill regions only allow a bounded motion for these orbits. In particular, we have seen that a chaotic infinity of heteroclinic connections between one primary and the Lyapunov periodic orbits around the equilibrium point L1 are obtained. As a consequence a chaotic infinity of ejection-collision orbits is also derived. Besides, we will see that we can construct colour diagrams that allow to describe the global dynamics of the ejection orbits given a range of time. These colour diagrams provide a very precise understanding of the dynamics of these orbits.Finally, we have made a first exploration of the spatial case of the circular restricted three body problem (RTBP 3D). In this first approach we have not used the classical Kustaanheimo–Stiefel regularization, instead we have decided to use a 3D version of the McGehee regularization. This presents some problems that we have analyzed and addressed,
L'objectiu principal d'aquesta dissertació és l'estudi de les òrbites d'ejecció-col·lisió (EC) al problema restringit de tres cossos circular i pla (RTBP a partir d'ara). En particular, ens centrarem en l'estudi analític i numèric d'unes òrbites d'EC molt particulars, a les quals hem anomenat òrbites de n-EC. Aquestes òrbites de n-EC, són òrbites tal que la partícula ejecta d'un primari, assoleix n màxims en la distància respecte al primari del qual han ejectat per a continuació tornar a col·lisionar amb ell. D'aquesta forma numèricament estudiarem en profunditat aquest tipus d'òrbites i analíticament demostrarem que per un valor prou gran de la constant de Jacobi (per la qual donarem una expressió en termes del paràmetre de masses i el valor de n) existeixen exactament quatre òrbites de n-EC amb unes característiques ben determinades. Aquests resultats generalitzen i milloren els resultats previs pel cas particular de n=1, i veurem que es poden extrapolar fàcilment al problema de Hill. A més, numèricament veurem que l'evolució d'aquestes quatre famílies d'òrbites de n-EC originals presenta una dinàmica molt rica.És ben sabut, que el sistema que defineix el moviment de la partícula no està ben definit als punts on es troben situats els primaris. Per aquest motiu hem utilitzat dues tècniques de regularització de la col·lisió, la regularització de McGehee i la regularització de Levi-Civita. D'aquesta forma, en aquesta memòria hem analitzat els avantatges i els inconvenients de cada regularització, i els diferents mètodes que es poden utilitzar per detectar col·lisions. Com que gran part d'aquesta memòria es focalitzarà en valors de la constant de Jacobi més grans que l'associat al punt d'equilibri L1 aquestes dues regularitzacions de caràcter local seran suficients. Per valors menys restrictius de la constant de Jacobi veurem que existeixen altres regularitzacions de caràcter global o que simplement podem treballar amb regularitzacions locals a l'entorn de cada primari.Per altra banda, numèricament hem analitzat el comportament global de les òrbites d'ejecció al RTBP. Hem estudiat la relació entre la família de les òrbites periòdiques de Lyapunov al voltant del punt d'equilibri lineal L1 i les òrbites d'ejecció que es duu a terme al rang de valors de la constant de Jacobi tals que les regions de Hill associades només permeten un moviment fitat per a aquestes òrbites. En particular, hem vist que s'obté una infinitat caòtica de connexions heteroclíniques entre un primari i l'òrbita periòdica de Lyapunov al voltant del punt d'equilibri lineal L1. Com a conseqüència, també es deriva una infinitat caòtica d'òrbites d'ejecció-col·lisió. A més, veurem que podem construir uns diagrames de color que ens permeten descriure la dinàmica global de les òrbites d'ejecció donat un interval de temps. Aquests diagrames proporcionen una comprensió molt precisa de la dinàmica d'aquestes òrbites.Finalment, hem fet una primera exploració del cas espacial del problema restringit de tres cossos circular (RTBP 3D). En aquesta primera aproximació no hem utilitzat la clàssica regularització de Kustaanheimo-Stiefel i hem decidit utilitzar una versió 3D de la regularització de McGehee. Això presenta alguns problemes, que hem analitzat i abordat, però aquesta aproximació és suficient per obtenir un primer resultat numèric sobre òrbites de 1-EC i per il·lustrar la complexitat del cas 3D.
Matemàtica aplicada

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.

The scientific interest in space exploration is driven by the desire to answer fundamental questions relating to the formation of our solar system and life on Earth. Space agencies are currently pushing the boundaries of space mission design to meet scientific goals. Thus, space missions require novel trajectories to further human space exploration. A modern approach that has arisen in space mission design is to use dynamical system tools that exploit the natural dynamics of the solar system. A spacecraft's natural dynamics are affected by environmental perturbations such as Solar Radiation Pressure(SRP). Traditionally, the design of space missions requires any perturbations to be counteracted through corrective manoeuvres. However, these corrective manoeuvres require propellant and therefore the pre-storing of fuel. This thesis investigates fuel-free propulsion for harnessing SRP in the design of space missions of the Sun-Earth restricted three-body problem. SRP propulsion is applied to the spacecraft's orbit control and furthermore to create the propulsion required for the design of transfers between quasi-periodic orbits and end-of-life disposal trajectories. The advantage of SRP manoeuvres is that the spacecraft can have access to an unlimited source of propellant (the Sun's radiation) consequently extending its life and reducing the overall mission costs; where the advancement in space technology makes harnessing SRP devices possible for future missions design. SRP manoeuvres are triggered by light and extended reflective deployable structures (i.e., mirror-like surfaces). The magnitude of the SRP acceleration is a function of the spacecraft's area-to-mass ratio, its reflectivity properties, mass and orientation of the reflective surface to the Sun-line direction. This thesis demonstrates that SRP manoeuvres are an effective and an effcient approach to stabilise the natural dynamics of the spacecraft in the Sun-Earth system. The size of the required reflective deployable area and spacecraft pointing accuracy are the ultimate outcomes of this research. Along with the design of the reflective area, the definition of a new control law, a method to perform transfers between quasi-periodic orbits and a strategy for the end-of-life disposal are the major important research findings.

Aquesta tesi doctoral està emmarcada en el camp de l'astrodinàmica. Presenta solucions a problemes identificats en el disseny de missions que utilitzen òrbites entorn dels punts de libració, fent servir la teoria de sistemes dinàmics.
El problema restringit de tres cossos és un model per estudiar el moviment d'un cos de massa infinitessimal sota l'atracció gravitatòria de dos cossos molt massius. Els cinc punts d'equilibri d'aquest model, en especial L1 i L2, han estat motiu de nombrosos estudis per aplicacions pràctiques en les últimes dècades (SOHO, Genesis...).
Genèricament, qualsevol missió en òrbita al voltant del punt L2 del sistema Terra-Sol es veu afectat per ocultacions degudes a l'ombra de la Terra. Si l'òrbita és al voltant de L1, els eclipsis són deguts a la forta influència electromagnètica del Sol. D'entre els diferents tipus d'òrbites de libració, les òrbites de Lissajous resulten de la combinació de dues oscil.lacions perpendiculars. El seu principal avantatge és que les amplituds de les oscil.lacions poden ser escollides independentment i això les fa adapatables als requeriments de cada missió. La necessitat d'estratègies per evitar eclipsis en òrbites de Lissajous entorn dels punts L1 i L2 motivaren la primera part de la tesi. En aquesta part es presenta una eina per la planificació de maniobres en òrbites de Lissajous que no només serveix per solucionar el problema d'evitar els eclipsis, sinó també per trobar trajectòries de transferència entre òrbites d'amplituds diferents i planificar rendez-vous.
Per altra banda, existeixen canals de baix cost que uneixen els punts L1 i L2 d'un sistema donat i representen una manera natural de transferir d'una regió de libració a l'altra. Gràcies al seu caràcter hiperbòlic, una òrbita de libració té uns objectes invariants associats: les varietats estable i inestable. Si tenim present que la varietat estable està formada per trajectòries que tendeixen cap a l'òrbita a la qual estan associades quan el temps avança, i que la varietat inestable fa el mateix però enrera en el temps, una intersecció entre una varietat estable i una d'inestable proporciona un camí asimptòtic entre les òrbites corresponents. Un mètode per trobar connexions d'aquest tipus entre òrbites planes entorn de L1 i L2 es presenta a la segona part de la tesi, i s'hi inclouen els resultats d'aplicar aquest mètode als casos dels problemes restringits Sol Terra i Terra-Lluna.
La idea d'intersecar varietats hiperbòliques es pot aplicar també en la cerca de camins de baix cost entre les regions de libració del sistema Sol-Terra i Terra-Lluna. Si existissin camins naturals de les òrbites de libració solars cap a les lunars, s'obtindria una manera barata d'anar a la Lluna fent servir varietats invariants, cosa que no es pot fer de manera directa. I a l'inversa, un camí de les regions de libració lunars cap a les solars permetria, per exemple, que una estació fos col.locada en òrbita entorn del punt L2 lunar i servís com a base per donar servei a les missions que operen en òrbites de libració del sistema Sol-Terra. A la tercera part de la tesi es presenten mètodes per trobar trajectòries de baix cost que uneixen la regió L2 del sistema Terra-Lluna amb la regió L2 del sistema Sol-Terra, primer per òrbites planes i més endavant per òrbites de Lissajous, fent servir dos problemes de tres cossos acoblats. Un cop trobades les trajectòries en aquest model simplificat, convé refinar-les per fer-les més realistes. Una metodologia per obtenir trajectòries en efemèrides reals JPL a partir de les trobades entre òrbites de Lissajous en el model acoblat es presenta a la part final de la tesi. Aquestes trajectòries necessiten una maniobra en el punt d'acoblament, que és reduïda en el procés de refinat, arribant a obtenir trajectòries de cost zero quan això és possible.
This PhD. thesis lies within the field of astrodynamics. It provides solutions to problems which have been identified in mission design near libration points, by using dynamical systems theory.
The restricted three body problem is a well known model to study the motion of an infinitesimal mass under the gravitational attraction of two massive bodies. Its five equilibrium points, specially L1 and L2, have been the object of several studies aimed at practical applications in the last decades (SOHO, Genesis...).
In general, any mission in orbit around L2 of the Sun-Earth system is affected by occultations due to the shadow of the Earth. When the orbit is around L1, the eclipses are caused by the strong electromagnetic influence of the Sun. Among all different types of libration orbits, Lissajous type ones are the combination of two perpendicular oscillations. Its main advantage is that the amplitudes of the oscillations can be chosen independently and this fact makes Lissajous orbits more adaptable to the requirements of each particular mission than other kinds of libration motions. The need for eclipse avoidance strategies in Lissajous orbits around L1 and L2 motivated the first part of the thesis. It is in this part where a tool for planning maneuvers in Lissajous orbits is presented, which not only solves the eclipse avoidance problem, but can also be used for transferring between orbits having different amplitudes and for planning rendez-vous strategies.
On the other hand, there exist low cost channels joining the L1 and L2 points of a given sistem, which represent a natural way of transferring from one libration region to the other one. Furthermore, there exist hyperbolic invariant objects, called stable and unstable manifolds, which are associated with libration orbits due to their hyperbolic character. If we bear in mind that the stable manifold of a libration orbit consists of trajectories which tend to the orbit as time goes by, and that the unstable manifold does so but backwards in time, any intersection between a stable and an unstable manifold will provide an asymptotic path between the corresponding libration orbits. A methodology for finding such asymptotic connecting paths between planar orbits around L1 and L2 is presented in the second part of the dissertation, including results for the particular cases of the Sun-Earth and Earth-Moon problems.
Moreover, the idea of intersecting hyperbolic manifolds can be applied in the search for low cost paths joining the libration regions of different problems, such as the Sun-Earth and the Earth-Moon ones. If natural paths from the solar libration regions to the lunar ones was found, it would provide a cheap way of transferring to the Moon from the vicinity of the Earth, which is not possible in a direct way using invariant manifolds. And the other way round, paths from the lunar libration regions to the solar ones would allow for the placement of a station in orbit around the lunar L2, providing services to solar libration missions, for instance. In the third part of the thesis, a methodology for finding low cost trajectories joining the lunar L2 region and the solar L2 region is presented. This methodology was developed in a first step for planar orbits and in a further step for Lissajous type orbits, using in both cases two coupled restricted three body problems to model the Sun-Earth-Moon spacecraft four body problem. Once trajectories have been found in this simplified model, it is convenient to refine them to more realistic models. A methodology for obtaining JPL real ephemeris trajectories from the initial ones found in the coupled models is presented in the last part of the dissertation. These trajectories need a maneuver at the coupling point, which can be reduced in the refinement process until low cost connecting trajectories in real ephemeris are obtained (even zero cost, when possible).

In this thesis, the continuation and bifurcation software AUTO is used to compute periodic solutions of the circular restricted Three-Body problem (CR3BP). Periodic solution families for the Sun-Earth, the Earth-Moon, and the Sun-Jupiter system are studied in detail. Bifurcation diagrams for these systems are presented. Corresponding periodic orbits are also shown. To understand the solution structure better, a new data visualization package, PLAUT04, has been developed for AUTO. It reads AUTO data files and creates solution diagrams and bifurcation diagrams. This new package can also be used to animate solutions. A special version of PLAUT04, called PLAUT04/r3b, has been developed for the CR3BP. Using PLAUT04/r3b, we can animate solutions both in a rotating frame and in an inertial frame. These new graphics packages for AUTO have good rendering speed, flexibility, and display quality. A user-friendly interface makes both easy to learn and use.

With the growth of human interest in the Lunar region, methods of enabling Lunar access including surface and Low Lunar Orbit (LLO) from periodic orbit in the Lunar region is becoming more important. The current investigation explores the Lunar access capabilities of these periodic orbits. Impact trajectories originating from the 9:2 Lunar Synodic Resonant (LSR) Near Rectilinear Halo Orbit (NRHO) are determined through explicit propagation and mapping of initial conditions formed by applying small maneuvers at locations across the orbit. These trajectories yielding desirable Lunar impact final conditions are then used to converge impacting transfers from the NRHO to Shackleton crater near the Lunar south pole. The stability of periodic orbits in the Lunar region is analyzed through application of a stability index and time constant. The Lunar access capabilities of the Lunar region periodic orbits found to be sufficiently unstable are then analyzed through impact and periapse maps. Using the impact data, candidate periodic orbits are incorporated in the the NRHO to Shackleton crater mission design to control mission geometry. Finally, the periapse map data is used to determine periodic orbits with desirable apse conditions that are then used to design NRHO to LLO transfer trajectories.

An increasing interest in lunar exploration calls for low-cost techniques of reaching the Moon. Ballistic lunar transfers are long duration trajectories that leverage solar perturbations to reduce the multi-body energy of a spacecraft upon arrival into cislunar space. An investigation is conducted to explore methods of constructing ballistic lunar transfers. The techniques employ dynamical systems theory to leverage the underlying dynamical flow of the multi-body regime. Ballistic lunar transfers are governed by the gravitational influence of the Earth-Moon-Sun system; thus, multi-body gravity models are employed, i.e., the circular restricted three-body problem (CR3BP) and the bicircular restricted four-body problem (BCR4BP). The Sun-Earth CR3BP provides insight into the Sun’s effect on transfers near the Earth. The BCR4BP offers a coherent model for constructing end-to-end ballistic lunar transfers. Multiple techniques are employed to uncover ballistic transfers to conic and multi-body orbits in cislunar space. Initial conditions to deliver the spacecraft into various orbits emerge from Periapse Poincaré maps. From a chosen geometry, families of transfers from the Earth to conic orbits about the Moon are developed. Instantaneous equilibrium solutions in the BCR4BP provide an approximate for the theoretical minimum lunar orbit insertion costs, and are leveraged to create low-cost solutions. Trajectories to the L2 2:1 synodic resonant Lyapunov orbit, L2 2:1 synodic resonant Halo orbit, and the 3:1 synodic resonant Distant Retrograde Orbit (DRO) are investigated.

This thesis evaluates decommissioning strategies for the International Space Station ISS. A permanent solution is attempted by employing energy efficient invariant manifolds that arise in the circular restricted three body problem CRTBP to transport the ISS from its low Earth orbit LEO to a lunar orbit. Although the invariant manifolds provide efficient transport, getting the the ISS onto the manifolds proves quite expensive, and the trajectories take too long to complete. Therefore a more practical, although temporary, solution consisting of an optimal re-boost maneuver with the European Space Agency's automated transfer vehicle ATV is proposed. The optimal re-boost trajectory is found using control parameterization and the sequential quadratic programming SQP algorithm. The model used for optimization takes into account the affects of atmospheric drag and gravity perturbations. The optimal re-boost maneuver produces a satellite lifetime of approximately ninety-five years using a two ATV strategy.

The general station-keeping problem is a focal topic when considering any spacecraft mission application. Recent missions are increasingly requiring complex trajectories to satisfy mission requirements, necessitating the need for accurate station-keeping controllers. An ideal controller reliably corrects for spacecraft state error, minimizes the required propellant, and is computationally efficient. To that end, this investigation assesses the effectiveness of several controller formulations in the circular restricted three-body model. Particularly, a spacecraft is positioned in a L₁ southern halo orbit within the Sun-Earth Moon Barycenter system. To prevent the spacecraft from departing the vicinity of this reference halo orbit, the Floquet mode station-keeping approach is introduced and evaluated. While this control strategy generally succeeds in the station-keeping objective, a breakdown in performance is observed proportional to increases in state error. Therefore, a new hybrid controller is developed which leverages Floquet mode and reinforcement learning. The hybrid controller is observed to efficiently determine corrective maneuvers that consistently recover the reference orbit for all evaluated scenarios. A comparative analysis of the performance metrics of both control strategies is conducted, highlighting differences in the rates of success and the expected propellant costs. The performance comparison demonstrates a relative improvement in the ability of the hybrid controller to meet the mission objectives, and suggests the applicability of reinforcement learning to the station-keeping problem.

A dynamical understanding of orbits in the Earth-Moon neighborhood that can sustain long-term activities and trajectories that link locations of interest forms a critical foundation for the creation of infrastructure to support a lasting presence in this region of space. In response, this investigation aims to identify and exploit fundamental dynamical motion in the vicinity of a candidate ‘hub’ orbit, the L2 southern 9:2 lunar synodic resonant near rectilinear halo orbit (NRHO), while incorporating realistic mission constraints. The strategies developed in this investigation are, however, not restricted to this particular orbit but are, in fact, applicable to a wide variety of stable and nearly-stable cislunar orbits. Since stable and nearly-stable orbits that may lack useful manifold structures are of interest for long-term activities in cislunar space due to low orbit maintenance costs, strategies to alternatively initiate transfer design into and out of these orbits are necessary. Additionally, it is crucial to understand the complex behaviors in the neighborhood of any candidate hub orbit. In this investigation, a bifurcation analysis is used to identify periodic orbit families in close proximity to the hub orbit that may possess members with favorable stability properties, i.e., unstable orbits. Stability properties are quantified using a metric defined as the stability index. Broucke stability diagrams, a tool in which the eigenvalues of the monodromy matrix are recast into two simple parameters, are used to identify bifurcations along orbit families. Continuation algorithms, in combination with a differential corrections scheme, are used to compute new families of periodic orbits originating at bifurcations. These families possess unstable members with associated invariant manifolds that are indeed useful for trajectory design. Members of the families nearby the L2 NRHOs are demonstrated to persist in a higher-fidelity ephemeris model.

Transfers based on the identified nearby dynamical structures and their associated manifolds are designed. To formulate initial guesses for transfer trajectories, a Poincaré mapping technique is used. Various sample trajectory designs are produced in this investigation to demonstrate the wide applicability of the design methodology. Initially, designs are based in the circular restricted three-body problem, however, geometries are demonstrated to persist in a higher-fidelity ephemeris model, as well. A strategy to avoid Earth and Moon eclipse conditions along many-revolution quasi-periodic ephemeris orbits and transfer trajectories is proposed in response to upcoming mission needs. Lunar synodic resonance, in combination with careful epoch selection, produces a simple eclipse-avoidance technique. Additionally, an integral-type eclipse avoidance path constraint is derived and incorporated into a differential corrections scheme as well. Finally, transfer trajectories in the circular restricted three-body problem and higher-fidelity ephemeris model are optimized and the geometry is shown to persist.

An integral part of NASA's vision for the coming years is a sustained infrastructure in cislunar space. The current baseline trajectory for this facility is a Near Rectilinear Halo Orbit (NRHO), a periodic orbit in the Circular Restricted Three-Body Problem. One of the goals of the facility is to serve as a proving ground for human spaceflight operations in deep space. Thus, this investigation focuses on transfers between the baseline NRHO and a family of periodic orbits that originate from a period-doubling bifurcation along the halo family. This new family of orbits has been termed the ``butterfly" family. This investigation also provides an overview of the evolution for a large subset of the butterfly family. Transfers to multiple subsets of the family are found by leveraging different design strategies and techniques from dynamical systems theory. The different design strategies are discussed in detail, and the transfers to each of these regions are compared in terms of propellant costs and times of flight.

In this thesis, the continuation and bifurcation software AUTO is used to investigate elemental periodic orbits, certain secondary bifurcating orbit families, and homoclinic orbits, associated with the libration points of the Circular Restricted 3-Body Problem, for values of the mass ratio between 0 and 0.5. Periodic solution families for representative values of the mass ratio, and corresponding bifurcation diagrams, are studied in detail. To understand the solution structure of the Circular Restricted 3-Body Problem better, a new data visualization package, DR Orbits, has been developed. It reads AUTO data files and creates solution diagrams. The package may be useful for space mission design. This new graphics package has good rendering speed, flexibility, and display quality. A user-friendly interface makes DR Orbits easy to use

Thesis (Ph. D.)--University of Wisconsin--Madison, 1988.
Typescript. Vita. eContent provider-neutral record in process. Description based on print version record. Includes bibliographical references (leaves 161-185).

Spacecraft trajectory optimization is a topic of crucial importance to space missions design. The less fuel required to accomplish the mission, the more payload that can be transported, and the higher the opportunity to lower the cost of the space mission. The objective is to find the optimal trajectory through space that will minimize the fuel used, and still achieve all mission constraints. Most space trajectories are designed using the simplified relative two-body problem as the base model. Using this patched conics approximation, however, constrains the solution space and fails to produce accurate initial guesses for trajectories in sensitive dynamics. This dissertation uses the Circular Restricted Three-Body Problem (CR3BP) as the base model for designing transfer trajectories in the Circular Restricted Four-Body Problem (CR4BP). The dynamical behavior of the CR3BP guides the search for useful low-energy trajectory arcs. Two distinct models of the CR4BP are considered in this research: the Concentric model, and the Bi-Circular model. Transfers are broken down into trajectory arcs in two separate CR3BPs and the stable and unstable manifold structures of both systems are utilized to produce low-energy transfer arcs that are later patched together to form the orbit-to-orbit transfer. The patched solution is then used as an initial guess in the CR4BP model. A vital contribution of this dissertation is the sequential process for initial guess generation for transfers in the CR4BP. The techniques discussed in this dissertation overcome many of the difficulties in the trajectory design process presented by the complicated dynamics of the CR4BP. Indirect optimization techniques are also used to derive the first order necessary conditions for optimality to assure the optimality of the transfers and determine whether additional impulses might further lower the total cost of the mission.
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This thesis describes the design, implementation and analysis for a preliminary study for DLR's MASCOT lander's next mission to Small Solar System Bodies (SSSB). MASCOT (Mobile Asteroid Surface Scout) is a nano-lander that flew aboard Hayabusa2 (JAXA) to an asteroid, Ryugu. It is a passive nano-spacecraft that can only be deployed ballistically from a hovering spacecraft. Current research focusses on optimizing similar close-approach missions for deploying landers or small cubesats into periodic orbits but does not provide solutions with semi-autonomous small landers deployed from farther distances. This study aims to overcome this short-coming by proposing novel yet simple Guidance, Navigation and Control (GNC) and Propulsion systems for MASCOT. Due to its independent functioning and customisable anatomy, MASCOT can be adapted for several mission scenarios. In this thesis, a particular case-study is modelled for the HERA (ESA) mission. The first phase of the study involves the design of a landing trajectory to the moon of the Didymos binary asteroid system. For a preliminary analysis, the system - Didymain (primary body), Didymoon (secondary body) and MASCOT (third body) - are modelled as a Planar Circular Restricted Three Body Problem (PCR3BP). The numerical integration methodology used for the trajectory is the variable-step Dormand–Prince (Runge Kutta) ODE-4,5 (Ordinary Differential Equation) solver. The model is built in MATLAB-Simulink (2019a) and refined iteratively by conducting a Monte Carlo analysis using the Sensitivity Analysis Tool. Two models - a thruster-controlled system and an alternative hybrid propulsion system of solar sails and thrusters - are simulated and proven to be feasible. The results show that the stable manifold near Lagrange 2 points proposed by Tardivel et. al. for ballistic landings can still be exploited for distant deployments if a single impulse retro-burn is done at an altitude of 65 m to 210 m above ground with error margins of 50 m in position, 5 cm/s in velocity and 0.1 rad in attitude. The next phase is the conceptual design of a MASCOT-variant with GNC abilities. Based on the constraints and requirements of the flown spacecraft, novel GNC and Propulsion systems are chosen. To identify the overriding factors in using commercial-off-the-shelf (COTS) for MASCOT, a market survey is conducted and the manufacturers of short-listed products are consulted. The final phase of the study is to analyse the proposed equipment in terms of parameter scope and capability-oriented trade-offs. Two traceability matrices, one for devised solutions and system and another for solutions versus capabilities, are constructed. The final proposed system is coherent with the given mass, volume and power constraints. A distant deployment of MASCOT-like landers for in-situ observation is suggested as an advantageous and risk-reducing addition to large spacecraft missions to unknown micro-gravity target bodies. Lastly, the implications of this study and the unique advantages of an enhanced MASCOT lander are explored for currently planned SSSB missions ranging from multiple rendezvous, fly-by or sample-return missions. Concluding, this study lays the foundation for future work on advanced GNC concepts for unconventional spacecraft topology for the highly integrated small landers.

This thesis is submitted as per the requirements for the Spacemaster (Round 13) dual master's degree under the Erasmus Mundus Joint Master's Degree Programme. 

MASCOT team, DLR

Within the phase space of the planar circular restricted three-body problem, stable and unstable manifolds of periodic orbits with a S x R (cylindrical) geometry are shown to exist. The periodic orbits considered reside in bottleneck regions of the energy manifold, separating large zones associated with motion about one mass, the other mass, or both masses.

The cylinders have the physical property that all motion through the bottleneck in which the periodic orbit resides must occur through the interior of these surfaces. The cylinders thus mediate the global transport of test particles between large zones of the energy surface which are separated by the bottlenecks.

By elucidating the structuring role of the cylinders, we provide a new language for discussing some important problems in celestial mechanics. Furthermore, we propose that these cylindrical structures are the natural objects of study for the design of space mission trajectories which take advantage of three-body effects.