Symplectic Map Description of Halley's Comet Dynamics
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Low Thrust Manoeuvres to Perform Large Changes of RAAN Or Inclination in LEO
Facoltà di Ingegneria Corso di Laurea Magistrale in Ingegneria Aerospaziale Master Thesis Low Thrust Manoeuvres To Perform Large Changes of RAAN or Inclination in LEO Academic Tutor: Prof. Lorenzo CASALINO Candidate: Filippo GRISOT July 2018 “It is possible for ordinary people to choose to be extraordinary” E. Musk ii Filippo Grisot – Master Thesis iii Filippo Grisot – Master Thesis Acknowledgments I would like to address my sincere acknowledgments to my professor Lorenzo Casalino, for your huge help in these moths, for your willingness, for your professionalism and for your kindness. It was very stimulating, as well as fun, working with you. I would like to thank all my course-mates, for the time spent together inside and outside the “Poli”, for the help in passing the exams, for the fun and the desperation we shared throughout these years. I would like to especially express my gratitude to Emanuele, Gianluca, Giulia, Lorenzo and Fabio who, more than everyone, had to bear with me. I would like to also thank all my extra-Poli friends, especially Alberto, for your support and the long talks throughout these years, Zach, for being so close although the great distance between us, Bea’s family, for all the Sundays and summers spent together, and my soccer team Belfiga FC, for being the crazy lovable people you are. A huge acknowledgment needs to be address to my family: to my grandfather Luciano, for being a great friend; to my grandmother Bianca, for teaching me what “fighting” means; to my grandparents Beppe and Etta, for protecting me -
Osculating Orbit Elements and the Perturbed Kepler Problem
Osculating orbit elements and the perturbed Kepler problem Gmr a = 3 + f(r, v,t) Same − r x & v Define: r := rn ,r:= p/(1 + e cos f) ,p= a(1 e2) − osculating orbit he sin f h v := n + λ ,h:= Gmp p r n := cos ⌦cos(! + f) cos ◆ sinp ⌦sin(! + f) e − X +⇥ sin⌦cos(! + f) + cos ◆ cos ⌦sin(! + f⇤) eY actual orbit +sin⇥ ◆ sin(! + f) eZ ⇤ λ := cos ⌦sin(! + f) cos ◆ sin⌦cos(! + f) e − − X new osculating + sin ⌦sin(! + f) + cos ◆ cos ⌦cos(! + f) e orbit ⇥ − ⇤ Y +sin⇥ ◆ cos(! + f) eZ ⇤ hˆ := n λ =sin◆ sin ⌦ e sin ◆ cos ⌦ e + cos ◆ e ⇥ X − Y Z e, a, ω, Ω, i, T may be functions of time Perturbed Kepler problem Gmr a = + f(r, v,t) − r3 dh h = r v = = r f ⇥ ) dt ⇥ v h dA A = ⇥ n = Gm = f h + v (r f) Gm − ) dt ⇥ ⇥ ⇥ Decompose: f = n + λ + hˆ R S W dh = r λ + r hˆ dt − W S dA Gm =2h n h + rr˙ λ rr˙ hˆ. dt S − R S − W Example: h˙ = r S d (h cos ◆)=h˙ e dt · Z d◆ h˙ cos ◆ h sin ◆ = r cos(! + f)sin◆ + r cos ◆ − dt − W S Perturbed Kepler problem “Lagrange planetary equations” dp p3 1 =2 , dt rGm 1+e cos f S de p 2 cos f + e(1 + cos2 f) = sin f + , dt Gm R 1+e cos f S r d◆ p cos(! + f) = , dt Gm 1+e cos f W r d⌦ p sin(! + f) sin ◆ = , dt Gm 1+e cos f W r d! 1 p 2+e cos f sin(! + f) = cos f + sin f e cot ◆ dt e Gm − R 1+e cos f S− 1+e cos f W r An alternative pericenter angle: $ := ! +⌦cos ◆ d$ 1 p 2+e cos f = cos f + sin f dt e Gm − R 1+e cos f S r Perturbed Kepler problem Comments: § these six 1st-order ODEs are exactly equivalent to the original three 2nd-order ODEs § if f = 0, the orbit elements are constants § if f << Gm/r2, use perturbation theory -
Astrodynamics
Politecnico di Torino SEEDS SpacE Exploration and Development Systems Astrodynamics II Edition 2006 - 07 - Ver. 2.0.1 Author: Guido Colasurdo Dipartimento di Energetica Teacher: Giulio Avanzini Dipartimento di Ingegneria Aeronautica e Spaziale e-mail: [email protected] Contents 1 Two–Body Orbital Mechanics 1 1.1 BirthofAstrodynamics: Kepler’sLaws. ......... 1 1.2 Newton’sLawsofMotion ............................ ... 2 1.3 Newton’s Law of Universal Gravitation . ......... 3 1.4 The n–BodyProblem ................................. 4 1.5 Equation of Motion in the Two-Body Problem . ....... 5 1.6 PotentialEnergy ................................. ... 6 1.7 ConstantsoftheMotion . .. .. .. .. .. .. .. .. .... 7 1.8 TrajectoryEquation .............................. .... 8 1.9 ConicSections ................................... 8 1.10 Relating Energy and Semi-major Axis . ........ 9 2 Two-Dimensional Analysis of Motion 11 2.1 ReferenceFrames................................. 11 2.2 Velocity and acceleration components . ......... 12 2.3 First-Order Scalar Equations of Motion . ......... 12 2.4 PerifocalReferenceFrame . ...... 13 2.5 FlightPathAngle ................................. 14 2.6 EllipticalOrbits................................ ..... 15 2.6.1 Geometry of an Elliptical Orbit . ..... 15 2.6.2 Period of an Elliptical Orbit . ..... 16 2.7 Time–of–Flight on the Elliptical Orbit . .......... 16 2.8 Extensiontohyperbolaandparabola. ........ 18 2.9 Circular and Escape Velocity, Hyperbolic Excess Speed . .............. 18 2.10 CosmicVelocities -
Orbit Options for an Orion-Class Spacecraft Mission to a Near-Earth Object
Orbit Options for an Orion-Class Spacecraft Mission to a Near-Earth Object by Nathan C. Shupe B.A., Swarthmore College, 2005 A thesis submitted to the Faculty of the Graduate School of the University of Colorado in partial fulfillment of the requirements for the degree of Master of Science Department of Aerospace Engineering Sciences 2010 This thesis entitled: Orbit Options for an Orion-Class Spacecraft Mission to a Near-Earth Object written by Nathan C. Shupe has been approved for the Department of Aerospace Engineering Sciences Daniel Scheeres Prof. George Born Assoc. Prof. Hanspeter Schaub Date The final copy of this thesis has been examined by the signatories, and we find that both the content and the form meet acceptable presentation standards of scholarly work in the above mentioned discipline. iii Shupe, Nathan C. (M.S., Aerospace Engineering Sciences) Orbit Options for an Orion-Class Spacecraft Mission to a Near-Earth Object Thesis directed by Prof. Daniel Scheeres Based on the recommendations of the Augustine Commission, President Obama has pro- posed a vision for U.S. human spaceflight in the post-Shuttle era which includes a manned mission to a Near-Earth Object (NEO). A 2006-2007 study commissioned by the Constellation Program Advanced Projects Office investigated the feasibility of sending a crewed Orion spacecraft to a NEO using different combinations of elements from the latest launch system architecture at that time. The study found a number of suitable mission targets in the database of known NEOs, and pre- dicted that the number of candidate NEOs will continue to increase as more advanced observatories come online and execute more detailed surveys of the NEO population. -
Orbit Determination Using Modern Filters/Smoothers and Continuous Thrust Modeling
Orbit Determination Using Modern Filters/Smoothers and Continuous Thrust Modeling by Zachary James Folcik B.S. Computer Science Michigan Technological University, 2000 SUBMITTED TO THE DEPARTMENT OF AERONAUTICS AND ASTRONAUTICS IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF SCIENCE IN AERONAUTICS AND ASTRONAUTICS AT THE MASSACHUSETTS INSTITUTE OF TECHNOLOGY JUNE 2008 © 2008 Massachusetts Institute of Technology. All rights reserved. Signature of Author:_______________________________________________________ Department of Aeronautics and Astronautics May 23, 2008 Certified by:_____________________________________________________________ Dr. Paul J. Cefola Lecturer, Department of Aeronautics and Astronautics Thesis Supervisor Certified by:_____________________________________________________________ Professor Jonathan P. How Professor, Department of Aeronautics and Astronautics Thesis Advisor Accepted by:_____________________________________________________________ Professor David L. Darmofal Associate Department Head Chair, Committee on Graduate Students 1 [This page intentionally left blank.] 2 Orbit Determination Using Modern Filters/Smoothers and Continuous Thrust Modeling by Zachary James Folcik Submitted to the Department of Aeronautics and Astronautics on May 23, 2008 in Partial Fulfillment of the Requirements for the Degree of Master of Science in Aeronautics and Astronautics ABSTRACT The development of electric propulsion technology for spacecraft has led to reduced costs and longer lifespans for certain -
Multi-Body Trajectory Design Strategies Based on Periapsis Poincaré Maps
MULTI-BODY TRAJECTORY DESIGN STRATEGIES BASED ON PERIAPSIS POINCARÉ MAPS A Dissertation Submitted to the Faculty of Purdue University by Diane Elizabeth Craig Davis In Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy August 2011 Purdue University West Lafayette, Indiana ii To my husband and children iii ACKNOWLEDGMENTS I would like to thank my advisor, Professor Kathleen Howell, for her support and guidance. She has been an invaluable source of knowledge and ideas throughout my studies at Purdue, and I have truly enjoyed our collaborations. She is an inspiration to me. I appreciate the insight and support from my committee members, Professor James Longuski, Professor Martin Corless, and Professor Daniel DeLaurentis. I would like to thank the members of my research group, past and present, for their friendship and collaboration, including Geoff Wawrzyniak, Chris Patterson, Lindsay Millard, Dan Grebow, Marty Ozimek, Lucia Irrgang, Masaki Kakoi, Raoul Rausch, Matt Vavrina, Todd Brown, Amanda Haapala, Cody Short, Mar Vaquero, Tom Pavlak, Wayne Schlei, Aurelie Heritier, Amanda Knutson, and Jeff Stuart. I thank my parents, David and Jeanne Craig, for their encouragement and love throughout my academic career. They have cheered me on through many years of studies. I am grateful for the love and encouragement of my husband, Jonathan. His never-ending patience and friendship have been a constant source of support. Finally, I owe thanks to the organizations that have provided the funding opportunities that have supported me through my studies, including the Clare Booth Luce Foundation, Zonta International, and Purdue University and the School of Aeronautics and Astronautics through the Graduate Assistance in Areas of National Need and the Purdue Forever Fellowships. -
Orbit Conjunction Filters Final
AAS 09-372 A DESCRIPTION OF FILTERS FOR MINIMIZING THE TIME REQUIRED FOR ORBITAL CONJUNCTION COMPUTATIONS James Woodburn*, Vincent Coppola† and Frank Stoner‡ Classical filters used in the identification of orbital conjunctions are described and examined for potential failure cases. Alternative implementations of the filters are described which maintain the spirit of the original concepts but improve robustness. The computational advantage provided by each filter when applied to the one versus all and the all versus all orbital conjunction problems are presented. All of the classic filters are shown to be applicable to conjunction detection based on tabulated ephemerides in addition to two line element sets. INTRODUCTION The problem of on-orbit collisions or near collisions is receiving increased attention in light of the recent collision between an Iridium satellite and COSMOS 2251. More recently, the crew of the International Space Station was evacuated to the Soyuz module until a chunk of debris had safely passed. Dealing with the reality of an ever more crowded space environment requires the identification of potentially dangerous orbital conjunctions followed by the selection of an appropriate course of action. This text serves to describe the process of identifying all potential orbital conjunctions, or more specifically, the techniques used to allow the computations to be performed reliably and within a reasonable period of time. The identification of potentially dangerous conjunctions is most commonly done by determining periods of time when two objects have an unacceptable risk of collision. For this analysis, we will use the distance between the objects as our proxy for risk of collision. -
Perturbation Theory in Celestial Mechanics
Perturbation Theory in Celestial Mechanics Alessandra Celletti Dipartimento di Matematica Universit`adi Roma Tor Vergata Via della Ricerca Scientifica 1, I-00133 Roma (Italy) ([email protected]) December 8, 2007 Contents 1 Glossary 2 2 Definition 2 3 Introduction 2 4 Classical perturbation theory 4 4.1 The classical theory . 4 4.2 The precession of the perihelion of Mercury . 6 4.2.1 Delaunay action–angle variables . 6 4.2.2 The restricted, planar, circular, three–body problem . 7 4.2.3 Expansion of the perturbing function . 7 4.2.4 Computation of the precession of the perihelion . 8 5 Resonant perturbation theory 9 5.1 The resonant theory . 9 5.2 Three–body resonance . 10 5.3 Degenerate perturbation theory . 11 5.4 The precession of the equinoxes . 12 6 Invariant tori 14 6.1 Invariant KAM surfaces . 14 6.2 Rotational tori for the spin–orbit problem . 15 6.3 Librational tori for the spin–orbit problem . 16 6.4 Rotational tori for the restricted three–body problem . 17 6.5 Planetary problem . 18 7 Periodic orbits 18 7.1 Construction of periodic orbits . 18 7.2 The libration in longitude of the Moon . 20 1 8 Future directions 20 9 Bibliography 21 9.1 Books and Reviews . 21 9.2 Primary Literature . 22 1 Glossary KAM theory: it provides the persistence of quasi–periodic motions under a small perturbation of an integrable system. KAM theory can be applied under quite general assumptions, i.e. a non– degeneracy of the integrable system and a diophantine condition of the frequency of motion. -
Satellite Orbits
Course Notes for Ocean Colour Remote Sensing Course Erdemli, Turkey September 11 - 22, 2000 Module 1: Satellite Orbits prepared by Assoc Professor Mervyn J Lynch Remote Sensing and Satellite Research Group School of Applied Science Curtin University of Technology PO Box U1987 Perth Western Australia 6845 AUSTRALIA tel +618-9266-7540 fax +618-9266-2377 email <[email protected]> Module 1: Satellite Orbits 1.0 Artificial Earth Orbiting Satellites The early research on orbital mechanics arose through the efforts of people such as Tyco Brahe, Copernicus, Kepler and Galileo who were clearly concerned with some of the fundamental questions about the motions of celestial objects. Their efforts led to the establishment by Keppler of the three laws of planetary motion and these, in turn, prepared the foundation for the work of Isaac Newton who formulated the Universal Law of Gravitation in 1666: namely, that F = GmM/r2 , (1) Where F = attractive force (N), r = distance separating the two masses (m), M = a mass (kg), m = a second mass (kg), G = gravitational constant. It was in the very next year, namely 1667, that Newton raised the possibility of artificial Earth orbiting satellites. A further 300 years lapsed until 1957 when the USSR achieved the first launch into earth orbit of an artificial satellite - Sputnik - occurred. Returning to Newton's equation (1), it would predict correctly (relativity aside) the motion of an artificial Earth satellite if the Earth was a perfect sphere of uniform density, there was no atmosphere or ocean or other external perturbing forces. However, in practice the situation is more complicated and prediction is a less precise science because not all the effects of relevance are accurately known or predictable. -
Statistical Orbit Determination
Preface The modem field of orbit determination (OD) originated with Kepler's inter pretations of the observations made by Tycho Brahe of the planetary motions. Based on the work of Kepler, Newton was able to establish the mathematical foundation of celestial mechanics. During the ensuing centuries, the efforts to im prove the understanding of the motion of celestial bodies and artificial satellites in the modem era have been a major stimulus in areas of mathematics, astronomy, computational methodology and physics. Based on Newton's foundations, early efforts to determine the orbit were focused on a deterministic approach in which a few observations, distributed across the sky during a single arc, were used to find the position and velocity vector of a celestial body at some epoch. This uniquely categorized the orbit. Such problems are deterministic in the sense that they use the same number of independent observations as there are unknowns. With the advent of daily observing programs and the realization that the or bits evolve continuously, the foundation of modem precision orbit determination evolved from the attempts to use a large number of observations to determine the orbit. Such problems are over-determined in that they utilize far more observa tions than the number required by the deterministic approach. The development of the digital computer in the decade of the 1960s allowed numerical approaches to supplement the essentially analytical basis for describing the satellite motion and allowed a far more rigorous representation of the force models that affect the motion. This book is based on four decades of classroom instmction and graduate- level research. -
On Optimal Two-Impulse Earth–Moon Transfers in a Four-Body Model
Noname manuscript No. (will be inserted by the editor) On Optimal Two-Impulse Earth–Moon Transfers in a Four-Body Model F. Topputo Received: date / Accepted: date Abstract In this paper two-impulse Earth–Moon transfers are treated in the restricted four-body problem with the Sun, the Earth, and the Moon as primaries. The problem is formulated with mathematical means and solved through direct transcription and multiple shooting strategy. Thousands of solutions are found, which make it possible to frame known cases as special points of a more general picture. Families of solutions are defined and characterized, and their features are discussed. The methodology described in this paper is useful to perform trade-off analyses, where many solutions have to be produced and assessed. Keywords Earth–Moon transfer low-energy transfer ballistic capture trajectory optimization · · · · restricted three-body problem restricted four-body problem · 1 Introduction The search for trajectories to transfer a spacecraft from the Earth to the Moon has been the subject of countless works. The Hohmann transfer represents the easiest way to perform an Earth–Moon transfer. This requires placing the spacecraft on an ellipse having the perigee on the Earth parking orbit and the apogee on the Moon orbit. By properly switching the gravitational attractions along the orbit, the spacecraft’s motion is governed by only the Earth for most of the transfer, and by only the Moon in the final part. More generally, the patched-conics approximation relies on a Keplerian decomposition of the solar system dynamics (Battin 1987). Although from a practical point of view it is desirable to deal with analytical solutions, the two-body problem being integrable, the patched-conics approximation inherently involves hyperbolic approaches upon arrival. -
Wobbling Stars and Planets in This Activity, You
Activity: Wobbling Stars and Planets In this activity, you will investigate the effect of the relative masses and distances of objects that are in orbit around a star or planet. The most common understanding of how the planets orbit around the Sun is that the Sun is stationary while the planets revolve around the Sun. That is not quite true and is a little more complicated. PBS An important parameter when considering orbits is the center of mass. The concept of the center of mass is that of an average of the masses factored by their distances from one another. This is also commonly referred to as the center of gravity. For example, the balancing of a ETIC seesaw about a pivot point demonstrates the center of mass of two objects of different masses. Another important parameter to consider is the barycenter. The barycenter is the center of mass of two or more bodies that are orbiting each other, and is the point around which both of them orbit. When the two bodies have similar masses, the barycenter may be located between NASA the two bodies (outside of either of bodies) and both objects will follow an orbit around the center of mass. This is the case for a system such as the Sun and Jupiter. In the case where the two objects have very different masses, the center of mass and barycenter may be located within the more massive body. This is the case for the Earth-Moon system. scienceprojectideasforkids.com Classroom Activity Part 1 Materials • Wooden skewers 1.