Simulation and Study of Gravity Assist Maneuvers
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Electric Propulsion System Scaling for Asteroid Capture-And-Return Missions
Electric propulsion system scaling for asteroid capture-and-return missions Justin M. Little⇤ and Edgar Y. Choueiri† Electric Propulsion and Plasma Dynamics Laboratory, Princeton University, Princeton, NJ, 08544 The requirements for an electric propulsion system needed to maximize the return mass of asteroid capture-and-return (ACR) missions are investigated in detail. An analytical model is presented for the mission time and mass balance of an ACR mission based on the propellant requirements of each mission phase. Edelbaum’s approximation is used for the Earth-escape phase. The asteroid rendezvous and return phases of the mission are modeled as a low-thrust optimal control problem with a lunar assist. The numerical solution to this problem is used to derive scaling laws for the propellant requirements based on the maneuver time, asteroid orbit, and propulsion system parameters. Constraining the rendezvous and return phases by the synodic period of the target asteroid, a semi- empirical equation is obtained for the optimum specific impulse and power supply. It was found analytically that the optimum power supply is one such that the mass of the propulsion system and power supply are approximately equal to the total mass of propellant used during the entire mission. Finally, it is shown that ACR missions, in general, are optimized using propulsion systems capable of processing 100 kW – 1 MW of power with specific impulses in the range 5,000 – 10,000 s, and have the potential to return asteroids on the order of 103 104 tons. − Nomenclature -
Optimisation of Propellant Consumption for Power Limited Rockets
Delft University of Technology Faculty Electrical Engineering, Mathematics and Computer Science Delft Institute of Applied Mathematics Optimisation of Propellant Consumption for Power Limited Rockets. What Role do Power Limited Rockets have in Future Spaceflight Missions? (Dutch title: Optimaliseren van brandstofverbruik voor vermogen gelimiteerde raketten. De rol van deze raketten in toekomstige ruimtevlucht missies. ) A thesis submitted to the Delft Institute of Applied Mathematics as part to obtain the degree of BACHELOR OF SCIENCE in APPLIED MATHEMATICS by NATHALIE OUDHOF Delft, the Netherlands December 2017 Copyright c 2017 by Nathalie Oudhof. All rights reserved. BSc thesis APPLIED MATHEMATICS \ Optimisation of Propellant Consumption for Power Limited Rockets What Role do Power Limite Rockets have in Future Spaceflight Missions?" (Dutch title: \Optimaliseren van brandstofverbruik voor vermogen gelimiteerde raketten De rol van deze raketten in toekomstige ruimtevlucht missies.)" NATHALIE OUDHOF Delft University of Technology Supervisor Dr. P.M. Visser Other members of the committee Dr.ir. W.G.M. Groenevelt Drs. E.M. van Elderen 21 December, 2017 Delft Abstract In this thesis we look at the most cost-effective trajectory for power limited rockets, i.e. the trajectory which costs the least amount of propellant. First some background information as well as the differences between thrust limited and power limited rockets will be discussed. Then the optimal trajectory for thrust limited rockets, the Hohmann Transfer Orbit, will be explained. Using Optimal Control Theory, the optimal trajectory for power limited rockets can be found. Three trajectories will be discussed: Low Earth Orbit to Geostationary Earth Orbit, Earth to Mars and Earth to Saturn. After this we compare the propellant use of the thrust limited rockets for these trajectories with the power limited rockets. -
Asteroid Retrieval Mission
Where you can put your asteroid Nathan Strange, Damon Landau, and ARRM team NASA/JPL-CalTech © 2014 California Institute of Technology. Government sponsorship acknowledged. Distant Retrograde Orbits Works for Earth, Moon, Mars, Phobos, Deimos etc… very stable orbits Other Lunar Storage Orbit Options • Lagrange Points – Earth-Moon L1/L2 • Unstable; this instability enables many interesting low-energy transfers but vehicles require active station keeping to stay in vicinity of L1/L2 – Earth-Moon L4/L5 • Some orbits in this region is may be stable, but are difficult for MPCV to reach • Lunar Weakly Captured Orbits – These are the transition from high lunar orbits to Lagrange point orbits – They are a new and less well understood class of orbits that could be long term stable and could be easier for the MPCV to reach than DROs – More study is needed to determine if these are good options • Intermittent Capture – Weakly captured Earth orbit, escapes and is then recaptured a year later • Earth Orbit with Lunar Gravity Assists – Many options with Earth-Moon gravity assist tours Backflip Orbits • A backflip orbit is two flybys half a rev apart • Could be done with the Moon, Earth or Mars. Backflip orbit • Lunar backflips are nice plane because they could be used to “catch and release” asteroids • Earth backflips are nice orbits in which to construct things out of asteroids before sending them on to places like Earth- Earth or Moon orbit plane Mars cyclers 4 Example Mars Cyclers Two-Synodic-Period Cycler Three-Synodic-Period Cycler Possibly Ballistic Chen, et al., “Powered Earth-Mars Cycler with Three Synodic-Period Repeat Time,” Journal of Spacecraft and Rockets, Sept.-Oct. -
An Assessment of Aerocapture and Applications to Future Missions
Post-Exit Atmospheric Flight Cruise Approach An Assessment of Aerocapture and Applications to Future Missions February 13, 2016 National Aeronautics and Space Administration An Assessment of Aerocapture Jet Propulsion Laboratory California Institute of Technology Pasadena, California and Applications to Future Missions Jet Propulsion Laboratory, California Institute of Technology for Planetary Science Division Science Mission Directorate NASA Work Performed under the Planetary Science Program Support Task ©2016. All rights reserved. D-97058 February 13, 2016 Authors Thomas R. Spilker, Independent Consultant Mark Hofstadter Chester S. Borden, JPL/Caltech Jessie M. Kawata Mark Adler, JPL/Caltech Damon Landau Michelle M. Munk, LaRC Daniel T. Lyons Richard W. Powell, LaRC Kim R. Reh Robert D. Braun, GIT Randii R. Wessen Patricia M. Beauchamp, JPL/Caltech NASA Ames Research Center James A. Cutts, JPL/Caltech Parul Agrawal Paul F. Wercinski, ARC Helen H. Hwang and the A-Team Paul F. Wercinski NASA Langley Research Center F. McNeil Cheatwood A-Team Study Participants Jeffrey A. Herath Jet Propulsion Laboratory, Caltech Michelle M. Munk Mark Adler Richard W. Powell Nitin Arora Johnson Space Center Patricia M. Beauchamp Ronald R. Sostaric Chester S. Borden Independent Consultant James A. Cutts Thomas R. Spilker Gregory L. Davis Georgia Institute of Technology John O. Elliott Prof. Robert D. Braun – External Reviewer Jefferey L. Hall Engineering and Science Directorate JPL D-97058 Foreword Aerocapture has been proposed for several missions over the last couple of decades, and the technologies have matured over time. This study was initiated because the NASA Planetary Science Division (PSD) had not revisited Aerocapture technologies for about a decade and with the upcoming study to send a mission to Uranus/Neptune initiated by the PSD we needed to determine the status of the technologies and assess their readiness for such a mission. -
Up, Up, and Away by James J
www.astrosociety.org/uitc No. 34 - Spring 1996 © 1996, Astronomical Society of the Pacific, 390 Ashton Avenue, San Francisco, CA 94112. Up, Up, and Away by James J. Secosky, Bloomfield Central School and George Musser, Astronomical Society of the Pacific Want to take a tour of space? Then just flip around the channels on cable TV. Weather Channel forecasts, CNN newscasts, ESPN sportscasts: They all depend on satellites in Earth orbit. Or call your friends on Mauritius, Madagascar, or Maui: A satellite will relay your voice. Worried about the ozone hole over Antarctica or mass graves in Bosnia? Orbital outposts are keeping watch. The challenge these days is finding something that doesn't involve satellites in one way or other. And satellites are just one perk of the Space Age. Farther afield, robotic space probes have examined all the planets except Pluto, leading to a revolution in the Earth sciences -- from studies of plate tectonics to models of global warming -- now that scientists can compare our world to its planetary siblings. Over 300 people from 26 countries have gone into space, including the 24 astronauts who went on or near the Moon. Who knows how many will go in the next hundred years? In short, space travel has become a part of our lives. But what goes on behind the scenes? It turns out that satellites and spaceships depend on some of the most basic concepts of physics. So space travel isn't just fun to think about; it is a firm grounding in many of the principles that govern our world and our universe. -
Future Space Launch Vehicles
Future Space Launch Vehicles S. Krishnan Professor of Aerospace Engineering Indian Institute of Technology Madras Chennnai - 600 036, India (Written in 2001) Introduction Space technology plays a very substantial role in the economical growth and the national security needs of any country. Communications, remote sensing, weather forecasting, navigation, tracking and data relay, surveillance, reconnaissance, and early warning and missile defence are its dependent user-technologies. Undoubtedly, space technology has become the backbone of global information highway. In this technology, the two most important sub-technologies correspond to spacecraft and space launch vehicles. Spacecraft The term spacecraft is a general one. While the spacecraft that undertakes a deep space mission bears this general terminology, the one that orbits around a planet is also a spacecraft but called specifically a satellite more strictly an artificial satellite, moons around their planets being natural satellites. Cassini is an example for a spacecraft. This was developed under a cooperative project of NASA, the European Space Agency, and the Italian Space Agency. Cassini spacecraft, launched in 1997, is continuing its journey to Saturn (about 1274 million km away from Earth), where it is scheduled to begin in July 2004, a four- year exploration of Saturn, its rings, atmosphere, and moons (18 in number). Cassini executed two gravity-assist flybys (or swingbys) of Venus one in April 1998 and one in June 1999 then a flyby of Earth in August 1999, and a flyby of Jupiter (about 629 million km away) in December 2000, see Fig. 1. We may note here with interest that ISRO (Indian Space Research Organisation) is thinking of a flyby mission of a spacecraft around Moon (about 0.38 million km away) by using its launch vehicle PSLV. -
Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion
AAS 05-398 Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion ∗ ϒ Daniel W. Parcher ∗∗ and Jon A. Sims ϒϒ This paper examines optimal low-thrust gravity-assist trajectories to Jupiter using nuclear electric propulsion. Three different Venus-Earth Gravity Assist (VEGA) types are presented and compared to other gravity-assist trajectories using combinations of Earth, Venus, and Mars. Families of solutions for a given gravity-assist combination are differentiated by the approximate transfer resonance or number of heliocentric revolutions between flybys and by the flyby types. Trajectories that minimize initial injection energy by using low resonance transfers or additional heliocentric revolutions on the first leg of the trajectory offer the most delivered mass given sufficient flight time. Trajectory families that use only Earth gravity assists offer the most delivered mass at most flight times examined, and are available frequently with little variation in performance. However, at least one of the VEGA trajectory types is among the top performers at all of the flight times considered. INTRODUCTION The use of planetary gravity assists is a proven technique to improve the performance of interplanetary trajectories as exemplified by the Voyager, Galileo, and Cassini missions. Another proven technique for enhancing the performance of space missions is the use of highly efficient electric propulsion systems. Electric propulsion can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems.1,2 This technology has been demonstrated on the Deep Space 1 mission 3 − part of NASA’s New Millennium Program to validate technologies which can lower the cost and risk and enhance the performance of future missions. -
Mscthesis Joseangelgutier ... Humada.Pdf
Targeting a Mars science orbit from Earth using Dual Chemical-Electric Propulsion and Ballistic Capture Jose Angel Gutierrez Ahumada Delft University of Technology TARGETING A MARS SCIENCE ORBIT FROM EARTH USING DUAL CHEMICAL-ELECTRIC PROPULSION AND BALLISTIC CAPTURE by Jose Angel Gutierrez Ahumada MSc Thesis in partial fulfillment of the requirements for the degree of Master of Science in Aerospace Engineering at the Delft University of Technology, to be defended publicly on Wednesday May 8, 2019. Supervisor: Dr. Francesco Topputo, TU Delft/Politecnico di Milano Dr. Ryan Russell, The University of Texas at Austin Thesis committee: Dr. Francesco Topputo, TU Delft/Politecnico di Milano Prof. dr. ir. Pieter N.A.M. Visser, TU Delft ir. Ron Noomen, TU Delft Dr. Angelo Cervone, TU Delft This thesis is confidential and cannot be made public until December 31, 2020. An electronic version of this thesis is available at http://repository.tudelft.nl/. Cover picture adapted from https://steemitimages.com/p/2gs...QbQvi EXECUTIVE SUMMARY Ballistic capture is a relatively novel concept in interplanetary mission design with the potential to make Mars and other targets in the Solar System more accessible. A complete end-to-end interplanetary mission from an Earth-bound orbit to a stable science orbit around Mars (in this case, an areostationary orbit) has been conducted using this concept. Sets of initial conditions leading to ballistic capture are generated for different epochs. The influence of the dynamical model on the capture is also explored briefly. Specific capture trajectories are then selected based on a study of their stabilization into an areostationary orbit. -
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. -
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. -
Interplanetary Cruising
Interplanetary Cruising 1. Foreword The following material is intended to supplement NASA-JSC's High School Aerospace Scholars (HAS) instructional reading. The author, a retired veteran of 60 Space Shuttle flights at Mission Control's Flight Dynamics Officer (FDO) Console, first served as HAS mentor in association with a June 2008 workshop. On that occasion, he noticed students were generally unfamiliar with the nature and design of spacecraft trajectories facilitating travel between Earth and Mars. Much of the specific trajectory design data appearing herein was developed during that and subsequent workshops as reference material for students integrating it into Mars mission timeline, spacecraft mass, landing site selection, and cost estimates. Interplanetary Cruising is therefore submitted to HAS for student use both before and during Mars or other interplanetary mission planning workshops. Notation in this document distinguishes vectors from scalars using bold characters in a vector's variable name. The "•" operator denotes a scalar product between two vectors. If vectors a and b are represented with scalar coordinates such that a = (a1, a2, a3) and b = (b1, b2, b3) in some 3- dimensional Cartesian system, a • b = a1 b1 + a2 b2 + a3 b3 = a b cos γ, where γ is the angle between a and b. 2. Introduction Material in this reference will help us understand the pedigree and limitations applying to paths (trajectories) followed by spacecraft about the Sun as they "cruise" through interplanetary space. Although cruise between Earth and Mars is used in specific examples, this lesson's physics would apply to travel between any two reasonably isolated objects in our solar system. -
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.