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InternacionalTipo de proyectoEntidad financiadoraNacionalidad EntidadTamao de la entidadFecha concesin
ParticipantesOtros ParticipantescAPiS: Anlisis dinmico, propagacin orbital avanzada y simulacin de sistemas espaciales complejos0<Proyectos y convenios en convocatorias pblicas competitivasSin nacionalidad
17/11/2014gDirector: Jesus Pelaez Alvarez//Participante: Oscar Lopez Rebollal//Participante: Manuel Ruiz Delgado//8Stardust: StardustThe Asteroid and Space Debris NetworkjAsteroids and space debris represent a significant hazard for space and terrestrial assets; at the same time asteroids represent also an opportunity. In recent years it has become clear that the increasing population of space debris could lead to catastrophic consequences in the near term. The Kessler syndrome (where the density of objects in orbit is high enough that collisions could set off a cascade) is more realistic than when it was first proposed in 1978. Although statistically less likely to occur, an asteroid impact would have devastating consequences for our planet. Although an impact with a large (~10 km) to medium (~300 m) sized, or diameter, asteroid is unlikely, still it is not negligible as the recent case of the asteroid Apophis has demonstrated. Furthermore impacts with smaller size objects, between 10 m to 100 m diameter, are expected to occur more frequently and hence are, proportionally, equally dangerous for humans and assets on Earth and in space. Asteroids and space debris share a number of commonalities: both are uncontrolled objects whose orbit is deeply affected by a number of perturbations, both have an irregular shape and an uncertain attitude motion, both are made of inhomogeneous materials that can respond unexpectedly to a deflection action, for both, accurate orbit determination is required, both need to be removed before they impact with something valuable for us.
The observation, manipulation and disposal of space debris and asteroids represent one of the most challenging goals for modern space technology. It represents a key scientific and commercial venture for the future in order to protect the space and Earth environment. Such a significant multidisciplinary technical challenge, with real societal benefit for the future, represents a compelling topic for a training network.
1
14/09/2012Participante: Davide Amato //Participante: Jesus Pelaez Alvarez//Participante: Hodei Urrutxua Cereijo//Director: Claudio Bombardelli //ISBN
Tipo de TesisCalificacinFechaHAdvanced Orbit Propagation Methods Applied to Asteroids and Space Debris9
Asteroids and space debris pose relevant menaces to civilization, both on ground and in space. Simultaneously, they present a number of common engineering and scientific challenges that must be tackled in the realm of Space Situational Awareness (SSA). As to improve current and future SSA technologies, robust and efficient orbit propagation methods are required. The main goal of the present thesis is to demonstrate that regularized for mulations of dynamics entail significant advantages in the most demanding orbit propagation problems for asteroids and space debris.
Regularized formulations are obtained by eliminating the 1/r^2 singularity in Newtonian equations of motion through an analytical procedure. The resulting regularized equations exhibit an excellent numerical performance. In this thesis, we consider the KustaanheimoStiefel formulation and several methods of the Dromo family, which represent the trajectory through a set of nonclassical orbital elements. In the first part, we focus on the orbit propagation of planetary close encounters, taking into account several test cases. As scenarios of relevant practical importance, we propagate resonant returns of several fictitious asteroids and measure the error in the bplane coordinates. To generalize the results, we carry out largescale simulations in the Circular, Restricted ThreeBody Problem by means of a bidimensional parametrization. We analyse the case of the asteroid (99942) Apophis, devoting particular attention to the amplification of the numerical error consequent to its deep close encounter in 2029. The second part is dedicated to the longterm prediction of Earth satellite orbits. We compare regularized formulations to a semianalytical method based on equinoctial elements for several orbital regimes and perturbations. The parameters affecting the semianalytical propagation efficiency are finetuned by analysing the different contributions to the integration error, which also gives insight on the limits of applicability of semianalytical methods. Regularized formulations compare very favourably for highly elliptical and supersynchronous orbits, which encourages their application to lifetime analyses and numerical explorations of the cislunar space. Applications to asteroid impact avoidance are presented in the third part. We show the results of a geographical deflection of the fictitious asteroid 2015PDC obtained with an Ion Beam Shepherd spacecraft. Finally, we perform a systematic study of the potential resonant returns of the fictitious asteroid 2017PDC after its deflection by a nuclear device.Doctoral
Sobresaliente
10/07/20176Autor: Davide Amato //Director: Claudio Bombardelli //)Director: Giulio Ba Universit di Pisa//PPropagation and Optimal Control of Space Trajectories Using Perturbation Methods
The development of lowthrust, highspecificimpulse thrusters, such as ionic and Hall effect thrusters, has brought new possibilities and challenges into the field of space mission design. This is highlighted by missions such as Deep Space 1, SMART1 and BepiColombo. The high specific impulse of these thrusters allows for important savings in propellant mass, while their capability to operate continuously confers greater flexibility and robustness to the design process. However, this flexibility comes at the cost of a greater design complexity compared to impulsive thrusters, since the control laws must now be expressed as time functions (instead of as a set of discrete maneuvers). Furthermore, this problem is aggravated by the fact that the low magnitude of the thrust implies a smaller control authority and longer mission times. In order to successfully address this challenges and fully exploit the advantages of lowthrust, highspecificimpulse thrusters, new mathematical tools have to be developed, both numerical and analytical. The aim is not just to be able to computationally solve increasingly complex practical problems, but also to gain a better insight into the underlying physics and to develop approximate analytical solutions for preliminary design and estimation.
The main objective of this thesis is to develop mathematical methods and tools for the propagation and optimization of lowthrust trajectories, both Earthbound and interplanetary, and apply them to different practical test cases. These families of optimal control problems (OCPs) are studied both numerically, by implementing algorithms for their accurate and efficient resolution, and analytically, by searching for approximate solutions.
This works covers a wide range of techniques for solving OCPs, including both direct and indirect methods. With direct methods, the OCP is transcribed as a discrete nonlinear programming (NLP) problem, which is then solved numerically using iterative algorithms starting from an initial guess of the solution. Conversely, in indirect methods the solution is sought for by imposing the first order optimality conditions, derived from Pontryagin's Maximum Principle or using the calculus of variations, which yields a twopoint boundary value problem (TPBVP). Said TPBVP is studied in this thesis in two different, yet complementary ways. On the on< e hand, approximate analytical solutions are sought for using perturbation methods. On the other hand, numerical solutions are obtained using iterative algorithms.
The dissertation is structured around three different formulations for orbital dynamics: a novel relative motion formulation in curvilinear coordinates; the modified equinoctial elements, originally introduced by P. Cefola; and Dromo, an elementbased orbital propagator developed by the Space Dynamics Group (Technical University of MadridUPM). The section devoted to the relative motion formulation focuses on obtaining approximate analytical solutions for the phase change, radius change and inclination change problems, identifying different operation regimes. In the next section a general framework for indirect optimization with elementbased formulations is developed, which is then applied, together with the modified equinoctial elements, to the design of endoflife dispossal maneuvers for satellites in the Galileo constellation. The final sections focuses on Dromo, studying its applicability to OCPs and presenting a multiplescales solution to the radial thrust problem.Sobresaliente cum laude
18/09/2017`Autor: Juan Luis Gonzalo Gomez//Director: Claudio Bombardelli //Director: Jesus Pelaez Alvarez//JCR del ISITtulo de la revistaISSNFactor de impacto JCRInformacin de impactoVolumenDOINmero de revistaDesde la pginaHasta la pginaMesRanking6A Hopf variables view on the libration points dynamicsThe dynamics about the libration points of the Hill problem is investigated analytically. In particular, the use of perturbation theory allows to reduce the problem to a one degree of freedom Hamiltonian depending on two dynamical parameters. The invariant manifolds structure of the Hill problem is then disclosed, yet accurate computations are limited to energy values close to that of the libration points.)Celestial Mechanics & Dynamical Astronomy 092329581,58412910.1007/s10569017977843285306SIN MES Autor: Martin Fidel Lara Coira//HAccurate orbit propagation in the presence of planetary close encounters1MONTHLY NOTICES OF THE ROYAL ASTRONOMICAL SOCIETY 003587114,95247010.1093/mnras/stx12542207920993Autor: Davide Amato //Autor: Claudio Bombardelli //Autor: giulio bau //<Alternative Set of Nonsingular Quaternionic Orbital ElementsnQuaternionic elements in orbital mechanics are usually related to the Kustaanheimo?Stiefel transformation or to the definition of the orbital plane. The new set of regular elements presented in this paper stems from the form of the equations of motion of a rotating solid, which model the evolution of a quaternion defining the orientation of a bodyfixed frame and the change in the angular velocity of such frame. By replacing the bodyfixed frame with a special orbital frame and accounting for the radial motion separately, an equivalent solution to orbital motion can be constructed. The variation of parameters technique furnishes a new set of elements that is independent from the orbital plane. A secondorder Sundman transformation introduces a fictitious time that replaces the physical time as the independent variable. This technique improves the numerical performance of the method and simplifies the derivation. The use of a time element yields an even smoother evolution of the orbital elements under perturbations. Once the Lagrange and Poisson brackets are obtained, the most general nonosculating version of the set of elements is presented. Regarding the performance, numerical experiments show that the method is comparable to other formulations involving similar stabilization and regularization techniques.
Read More: https://arc.aiaa.org/doi/abs/10.2514/1.G002753*Journal of Guidance, Control, And Dynamics 07315090401127372751 NOVIEMBREAutor: Javier Roa Vicens//=An efficient code to solve the Kepler equation. Elliptic caseOA new approach for solving Kepler equation for elliptical orbits is developed in this paper. This new approach takes advantage of the very good behaviour of the modified Newton?Raphson method when the initial seed is close to the looked for solution. To determine a good initial seed the eccentric anomaly domain [0, ?] is discretized in several intervals and for each one of these intervals a fifth degree interpolating polynomial is introduced. The six coefficients of the polynomial are obtained by requiring six conditions at both ends of the corresponding interval. Thus the real function and the polynomial have equal values at both ends of the interval. Similarly relations are imposed for the two first derivatives. In the singular corner of the Kepler equation, M smaller than 1 and 1 ? e close to zero an asymptotic expansion is developed. In
most of the cases, the seed generated leads to reach machine error accuracy with the modified Newton?Raphson method with no iterations or just one iteration. This approach improves the computational time compared with other methods currently in use./Montly notice of the royal astronomical society4,96146710.1093/mnras/stx13817021713<Autor: Virginia Raposo Pulido//Autor: Jesus Pelaez Alvarez//]Approximate solutions of nonlinear circular orbit relative motion in curvilinear coordinates)CELESTIAL MECHANICS & DYNAMICAL ASTRONOMY1,612710.1007/s105690169716x49667Autor: Claudio Bombardelli //Autor: Javier Roa Vicens//Autor: juan luis gonzalo //IDeflection of fictitious asteroid 2017 PDC: Ion beam vs. kinetic impactorMission scenarios for the deflection of fictitious asteroid 2017 PDC are investigated. Two deflection options, kinetic impactor (KI) and ion beam shepherd (IBS), are studied and compared on the basis of deflection performance, safety, as well as mission schedule and political aspects. Firstly, we propose the launch of a mediumsize rendevous spacecraft equipped with at least two ionic thrusters that can serve as propulsion means for the interplanetary trajectory up to rendezvous with the asteroid and as contactless actuators for a possible followup deflection mission. The asteroid, whose uncertainty ellipsoid is initially too large to establish whether (and how) it should be deflected, is reached by the rendezvous spacecraft after a lowthrust interplanetary trajectory of reasonable duration. Following rendezvous the spacecraft is placed in the vicinity of the asteroid to estimate its mass, study its structure and composition and, crucially, reduce its uncertainty ellipsoid by ground tracking to confirm or rule out an impact. Assuming that an impact is confirmed two main deflection scenarios are considered based on the actual asteroid size. Ion beam deflection is considered with the possibility of full deflection (the asteroid misses the Earth by a safe margin) or impact location adjustment (the impact footprint is diplaced to the nearest unpopulated region) depending on the asteroid size and the predicted impact location. The launch of a kinetic impactor mission is also considered with the employment of the rendezvous spacecraft to measure the deflection outcome and possibly to refine the deflection in case it is needed. The deflection performance of the two methods is compared.Acta Astronautica 009457651,53610.10167=Autor: Juan Luis Gonzalo Gomez//Autor: Claudio Bombardelli //Autor: Emilio Jos Calero //2Environmental effect of space debris repositioningADVANCES IN SPACE RESEARCH 027311771,4096010.1016/j.asr.2017.03.0442837Autor: Claudio Bombardelli //BAutor: e. m. alessi //Autor: a. rossi //Autor: g. b. valsecchic //CIntermediary LEO propagation including higher order zonal harmonicsTwo new intermediary orbits of the artificial satellite problem are proposed. The analytical solutions include higher order effects of the geopotential, and are obtained by means of a torsion transformation applied to the quasiKeplerian system resulting after the elimination of the parallax simplification, for the first intermediary, and after the elimination of the parallax and perigee simplification< s, for the second one. The new intermediaries perform notably well for low Earth orbits propagation, are free from special functions, and result advantageous, both in accuracy and efficiency, when compared to the standard Cowell integration of the J2 problem, thus providing appealing alternatives for onboard, shortterm, orbit propagation under limited computational resources.10.1007/s10569016973664505526ABRILAutor: Denis Hautesserres //#Note on the ideal frame formulationAn implementation of the ideal frame formulation of perturbed Keplerian motion is presented which only requires the integration of a differential system of dimension 7, contrary to the 8 variables traditionally integrated with this approach. The new formulation is based on the integration of a scaled version of the Eulerian set of redundant parameters and slightly improves runtime performance with respect to the 8dimensional case while retaining comparable accuracy.10.1007/s105690179778z12137151
SEPTIEMBRE>Optimal ContinuousThrust Rephasing Maneuver in Circular Orbit(JOURNAL OF GUIDANCE CONTROL AND DYNAMICS1,65110.2514/1.G002305511551165Autor: juan l. gonzalo //cOrbit covariance propagation via quadraticorder state transition matrix in curvilinear coordinates1,59410.1007/s1056901797739215234Autor: javier hernandoayuso //\The theory of asynchronous relative motion I: time transformations and nonlinear corrections10.1007/s10569016972863013307Autor: Javier Roa Vicens//Autor: Jesus Pelaez Alvarez//NThe theory of asynchronous relative motion II: universal and regular solutions10.1007/s105690169730z343368Nombre congresoTipo de participacinLugar del congreso RevisoresISBN o ISSNFecha inicio congresoFecha fin congresoTtulo de las actasIDeflection of Fictitious Asteroid 2017 PDC: Ion Beam vs. Kinetic ImpactorMission scenarios for the deflection of fictitious asteroid 2017 PDC are investigated. Two deflection options, kinetic impactor (KI) and ion beam shepherd (IBS), are studied and compared on the basis of deflection performance, safety, as well as mission schedule and political aspects. Firstly, we propose the launch of a mediumsize rendevous spacecraft equipped with at least two ionic thrusters that can serve as propulsion means for the interplanetary trajectory up to rendezvous with the asteroid and as contactless actuators for a possible followup deflection mission. The asteroid, whose uncertainty ellipsoid is initially too large to establish whether (and how) it should be deflected, is reached by the rendezvous spacecraft after a lowthrust interplanetary trajectory of reasonable duration. Following rendezvous the spacecraft is placed in an orbit around the asteroid to estimate its mass, study its structure and composition and, crucially, reduce its uncertainty ellipsoid by ground tracking to confirm or rule out an impact. Assuming that an impact is confirmed two main deflection scenarios are considered based on the actual asteroid size. Ion beam deflection is considered with the possibility of full deflection (the asteroid misses the Earth by a safe margin) or impact location adjustment (the impact footprint is diplaced to the nearest unpopulated region) depending on the asteroid size. The launch of a kinetic impactor mission is also considered with the employment of the rendezvous spacecraft to measure the deflection outcome and possibly to refine the deflection in case it is needed. The deflection performance of the two methods is compared.
/5th IAA Planetary Defense Conference  PDC 2017960Tokio, Japn
15/05/2017
19/05/201714=Autor: Claudio Bombardelli //Autor: Juan Luis Gonzalo Gomez//*Autor: Emilio Jos Calero estudiante UPM//cIndirect Optimization of EndofLife Disposal for Galileo Constellation Using Low Thrust PropulsionIn this work, the endoflife disposal of satellites in the Galileo constellation using low thrust propulsion is studied. Indirect optimization methods are employed to design transfer maneuvers to remove the satellite from its original operational orbit into previously computed orbits leading to its natural reentry within 100 years due to lunisolar perturbation effects. The dynamics are formulated using the modified equinoctial elements, which allow expressing the boundary conditions in a simple way at the cost of more complex equations compared to the use of Cartesian coordinates. A special focus is placed in defining an efficient and robust algorithm for solving the two point boundary value problem arising from the first order optimality conditions, including the integration of the analyticallyderived variational equations to obtain the State Transition Matrix, and the accurate detection of thrustswitching events. The numerical results obtained for several test cases show the practical feasibility of this endoflife disposal approach at thrust levels compatible with electric thrusters likely to be used by the next generation of Galileo satellites.
=26th International Symposium on Space Flight Dynamics (ISSFD)Matsuyama, Japn
03/06/2017
09/06/20176 Autor: Juan Luis Gonzalo Gomez//]Autor: Francesco Topputo Politecnico di Milano//Autor: Roberto Armellin Surrey Space Center//DLastminute semianalytical asteroid deflection by nuclear explosion_ In this paper, a semianalytical method to calculate impulsive asteroid deflection maneuvers is pre
sented. It can be applied to the design of any impulsive asteroid deflection mission, and is particularly
useful when optimizing lastminute deflection by nuclear explosion as the optimum impulse direction
can be far from tangential. The method hinges on minimizing the impulse size under a constraint on the
bplane coordinates. The optimization is based on a fast, semianalytical algorithm developed for Low
Earth Orbit optimal collision maneuvers design. Several deflection strategies can be selected following
this algorithm. Additionally, an analysis of resonant returns and keyholes is performed, in order to detect
and prevent deflecting the asteroid into an Earthimpacting orbit in the near future..
To method is applied to the fictitious asteroid 2017PDC. First we constrain the bplane intersection
point to be at a distance from the Earth center greater than a threshold. The minimumimpulse maneu
vers with varying anticipation show nontrivial properties. For integer multiples of the asteroid period,
the optimal maneuver is purely prograde or retrograde. An increasing tendency, superimposed to peri
odic variations, can be observed on the required impulse size as the deflection is performed closer to
the conjunction. The impulse size abruptly grows shortly before the conjunction. The optimal maneu
ver orientation may present discontinuities as the maneuver anticipation changes, and this will happen
when different local optima exchange global optimality. We present a transfer maneuver that arrives to
the asteroid about a whole orbit before the conjunction. For this case, we present a sensitivity analysis
of the optimal deflection maneuver. We also compute the resonant circles corresponding to returns
of the asteroid in up to 20 years after the 2027 encounter, and calculate to a first approximation the
associated keyholes to evaluate and mitigate the risk of an impact in a subsequent encounter.
Finally, we present a lastminute deflection strategy for a transfer maneuver that arrives months
before the collision with Earth. We perform a geographical deflection, shifting the impact point on Earth
surface from a densely populated area to a water body. We show evidence that the optimal impulse is
far from tangential, and has an important outofplane component.!Planetary Defense Conference 2017Tokyo, Japn6Autor: Javier HernandoAyuso The University of Tokyo//<On the resolution of the Lambert's problem with the SDGcodehBased on the Lambert's problem, once identified the oneparameter family of orbits that verify the geometric constraints of the problem, we must express the orbits based on a single parameter that allows to select thos<e that satisfy the kinematic condition. The aim of this paper is to reformulate the problem choosing as parameter the true anomaly of the bisector defined by the directions of the two position vectors. The algorithm applied is the SDGcode, developed by the Space Dynamics Group at UPM, which has already been assessed on the resolution of the Kepler equation proving its stability and reliability.,27th AAS/AIAA Space Flight Mechanics Meeting730San Antonio (Texas) 10816003
05/02/2017
09/02/2017412541420Advances in the Astronautical Sciences, Vol. 160Solving the Kepler equation with the SDGcodeA new code to solve the Kepler equation for elliptic and hyperbolic orbits has been developed. The motivation of the study is the determination of an appropriate seed to initialize the numerical method, considering the optimization already tested of the well known NewtonRaphson method. To do that, we take advantage of the full potential of the symbolic manipulators. The final algorithm is stable, reliable and solves successfully the solution of the Kepler equation in the singular corner (M << 1 and e ~ 1).41434160FUnified Formulation for ElementBased Indirect Trajectory OptimizationA general mathematical framework is presented to treat low thrust trajectory optimization problems using the indirect method and employing a generic set of orbital elements (e.g. classical elements, equinoctial, etc.). An algebraic manipulation of the optimality conditions stemming from Pontryagin Maximum Principle reveals the existence of a new quadratic form of the costate, which governs the costate contribution in all the equations of the first order necessary optimality conditions. The quadratic form provides a simple tool for the mathematical development of the optimality conditions for any chosen set of orbital elements and greatly simplifies the computation of a state transition matrix needed in order to improve the convergence of the associated twopoint boundary value problem. Objective functions corresponding to minimumtime, minimumenergy and minimumfuel problems are considered.
0Autor: Francesco Topputo Politecnico di Milano//EstadoReferencia Patente PrioritariaEn explotacinLicenciatariosFecha solicitudReferencia PCTReferencia EPOReferencia EEUUReferencia JaponTitulares aparte de la UPM_Sistema de generacin de potencia elctrica en rbita por medio de cables conductores flotantes Concedida
P201531648
13/11/2015!Fundacin Universidad Carlos III;)Inventor Contacto: Claudio Bombardelli //EInventor: Gonzalo Snchez Arriaga Fundacin Universidad Carlos III//ISistema de propulsin en rbita por medio de cables conductores flotantesEl sistema de propulsin en rbita por medio de cables conductores flotantes embarcado en un vehculo espacial comprende dos conjuntos de cables conductores electrodinmicos
conectados respectivamente a cada uno de los dos polos de una fuente generadora de potencia elctrica, y en donde cada conjunto est formado por al menos un cable conductor.
En presencia de un plasma y un campo magntico, como es el caso de un satlite orbitando en la ionosfera terrestre, una corriente elctrica fluye de forma natural a lo largo de los cables conductores. Como resultado de la interaccin del campo
magntico con dicha corriente, se genera una fuerza de Lorentz sobre los cables. Dicha fuerza se puede utilizar para controlar la rbita del vehculo espacial y puede ser variada, en sentido y magnitud, por medio de la fuente generadora de potencia elctrica que permite modificar la intensidad y sentido de la corriente a lo largo de los cables. La eficiencia del sistema depender del ambiente espacial, de la velocidad del satlite relativa al plasma, y de las propiedades y diseo de los cables (longitud, seccin y material). El sistema se podr optimizar aislando parcialmente los cables a lo largo de su longitud y empleando materiales que favorezcan la emisin de electrones por efecto terminico o fotoelctrico. Dichos materiales se utilizarn en la composicin de los cables o para revestir su superficie a lo largo de su extensin completa o de partes de ella. A diferencia de los sistemas de propulsin convencionales, como
cohetes qumicos o elctricos, o los cables electrodinmicos propuestos hasta la fecha, el sistema propuesto no requiere ni propulsante ni expelante.
P201531649DInventor: Gonzalo Snchez Arriaga Fundacin Universidad Carlos III//Entidad premiadaEntidad concedenteLugar donde se premiJBest Paper Award del 26th International Symposium in Space Flight Dynamics Premio otorgado a la comunicacin "Graphical Methodology for the Preliminary Design of an IonBeam Shepherd Mission" en el 26th International Symposium in Space Flight Dynamics celebrado en Matsuyama, JapnGrupo de Dinmica EspacialUPMMatsuyama, Japon<Autor: Hodei Urrutxua Cereijo//Autor: Claudio Bombardelli //LugarTipoParticipacin en el comit cientifico del congreso Key Topics in Orbit Propagation Applied to Space Situational Awareness (KEPASSSA 2017)$Organizacin cientifica del Congreso
24/07/2017ESTEC (ESA), Nordwijk, HOLANDA,,Responsabilidades en cmites internacionales=Autor: Martin Fidel Lara Coira//Autor: Jesus Pelaez Alvarez//Nombre Apellidos categoraentidadTutorDavidMorante GonzlezBECARIO (OTROS)Claudio Bombardelli VIRGINIA
RAPOSO PULIDOCONTRATADOS PREDOCTORALESJesus Pelaez Alvarez JUAN LUIS
GONZALO GOMEZ
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