DOCSLIB.ORG

CHARACTERISTICS and DESIGN STRATEGIES for NEAR RECTILINEAR HALO ORBITS WITHIN the EARTH-MOON SYSTEM a Thesis Submitted to the Fa

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
CHARACTERISTICS and DESIGN STRATEGIES for NEAR RECTILINEAR HALO ORBITS WITHIN the EARTH-MOON SYSTEM a Thesis Submitted to the Fa CHARACTERISTICS AND DESIGN STRATEGIES FOR NEAR RECTILINEAR HALO ORBITS WITHIN THE EARTH-MOON SYSTEM A Thesis Submitted to the Faculty of Purdue University by Emily M. Zimovan In Partial Fulfillment of the Requirements for the Degree of Master of Science in Aeronautics and Astronautics August 2017 Purdue University West Lafayette, Indiana ii iii THE PURDUE UNIVERSITY GRADUATE SCHOOL STATEMENT OF THESIS APPROVAL Professor Kathleen C. Howell, Chair School of Aeronautics and Astronautics Professor James M. Longuski School of Aeronautics and Astronautics Professor Carolin E. Frueh School of Aeronautics and Astronautics Approved by: Professor Weinong Wayne Chen Aeronautics and Astronautics Associate Head for Graduate Education iv To the women in science, technology, engineering, and mathematics that have come before me, and to those that will come after me. v ACKNOWLEDGMENTS First, a huge thank you for my parents. You instilled in me from a young age an invaluable love of math and science. Thank you for not just pushing me to question the world around me, but to find the answers. Without your encouragement and support, I would not be where I am today. From the bottom of my heart, I want to thank my fianc´e,Chris. When it felt like no one else in the world could understand where I was coming from or what I was going through, I knew that you actually did. I truly couldn't have done this without you. Your constant love and support mean the world to me. I'm grateful for you in so many ways. I want to extend my immense gratitude to my advisor, Professor Kathleen C. Howell. Thank you for giving me so many opportunities throughout my time in the research group. Your thoughts, encouragement, advice, and support have been invaluable to my research and time at Purdue. The members of my research group deserve a thank you, as well. You've listened to me present countless times over the past two years and have given me advice and thoughts that have led my research to where it is now. Thanks for helping me to learn all that I have. I would also like to thank my committee members, Professors James M. Longuski and Carolin Frueh. I appreciate the time you took to review my thesis and provide me with advice and feedback. Additionally, for helping fund my studies at both Purdue University and NASA Johnson Space Center, I would like to thank the School of Aeronautics and Astronautics, the College of Engineering, and the Purdue University vi Graduate School for awarding me with the Charles C. Chappelle Fellowship and the Ross Fellowship. Also, support provided by the NASA JSC NNX13AK60A grant is greatly appreciated. I am also grateful for the use of the outstanding facilities provided by the Rune and Barbara Eliasen Visualization Laboratory. vii TABLE OF CONTENTS Page LIST OF TABLES :::::::::::::::::::::::::::::::::: x LIST OF FIGURES ::::::::::::::::::::::::::::::::: xi ABSTRACT ::::::::::::::::::::::::::::::::::::: xvi 1 INTRODUCTION :::::::::::::::::::::::::::::::: 1 1.1 Previous Contributions ::::::::::::::::::::::::::: 2 1.1.1 Multi-Body Dynamics ::::::::::::::::::::::: 2 1.1.2 Differential Corrections ::::::::::::::::::::::: 3 1.1.3 Periodic Orbits ::::::::::::::::::::::::::: 4 1.2 Current Work :::::::::::::::::::::::::::::::: 6 2 DYNAMICAL MODELS ::::::::::::::::::::::::::::: 9 2.1 The Circular Restricted Three-Body Model :::::::::::::::: 10 2.1.1 Assumptions for the Circular Restricted Three-Body Model :: 10 2.1.2 Equations of Motion in the Circular Restricted Three-Body Model11 2.1.3 Integrals of the Motion ::::::::::::::::::::::: 17 2.2 The N-Body Ephemeris Model :::::::::::::::::::::: 25 2.2.1 Assumptions for the N-Body Ephemeris Model ::::::::: 25 2.2.2 Equations of Motion in the N-Body Ephemeris Model :::::: 27 viii Page 3 DIFFERENTIAL CORRECTIONS ::::::::::::::::::::::: 31 3.1 State Transition Matrix :::::::::::::::::::::::::: 31 3.1.1 STM for the Circular Restricted Three-Body Model ::::::: 33 3.1.2 STM for the N-Body Ephemeris Model :::::::::::::: 35 3.1.3 Monodromy Matrix ::::::::::::::::::::::::: 37 3.2 Differential Corrections Strategies - Shooting Methods ::::::::: 38 3.2.1 Simple Targeting Scheme Formulation :::::::::::::: 40 3.2.2 Multi-Segment Corrections Algorithm Formulation ::::::: 43 3.2.3 Additional Constraints ::::::::::::::::::::::: 46 3.3 Coordinate Frames ::::::::::::::::::::::::::::: 52 3.3.1 Coordinate Frame Transformation: Rotating Frame to Arbi- trary Inertial Frame :::::::::::::::::::::::: 52 3.3.2 Coordinate Frame Transformation: Rotating Frame to J2000 Inertial Frame :::::::::::::::::::::::::::: 55 4 TRAJECTORY CONSTRUCTION: CR3BP AND THE EPHEMERIS MODEL61 4.1 Periodic Orbit Computation in the CR3BP :::::::::::::::: 61 4.1.1 Continuation of an Orbit Family ::::::::::::::::: 68 4.1.2 Examples of Periodic Orbit Families ::::::::::::::: 71 4.2 Poincar´eMapping and Surfaces of Section :::::::::::::::: 84 4.3 Computation of Periodic Orbit Invariant Manifolds ::::::::::: 85 4.4 Transitioning CR3BP Solutions into the Ephemeris Model ::::::: 87 4.4.1 Apse Angle ::::::::::::::::::::::::::::: 88 5 MISSION APPLICATIONS: ORBITAL CHARACTERISTICS AND TRANS- FER DESIGN ::::::::::::::::::::::::::::::::::: 93 ix Page 5.1 Stability ::::::::::::::::::::::::::::::::::: 94 5.1.1 CR3BP: Stability Index :::::::::::::::::::::: 96 5.2 Eclipsing :::::::::::::::::::::::::::::::::: 104 5.2.1 Resonance ::::::::::::::::::::::::::::: 105 5.2.2 Eclipse Duration ::::::::::::::::::::::::: 112 5.3 Transfers :::::::::::::::::::::::::::::::::: 121 5.3.1 Accessing NRHOs from Low Earth Orbit :::::::::::: 121 5.3.2 Accessing Distant Retrograde Orbits from NRHOs :::::: 126 6 CONCLUDING REMARKS :::::::::::::::::::::::::: 145 6.1 Summary of the Present Investigation :::::::::::::::::: 145 6.2 Concluding Remarks ::::::::::::::::::::::::::: 146 6.3 Recommendations for Future Work ::::::::::::::::::: 148 REFERENCES ::::::::::::::::::::::::::::::::::: 149 x LIST OF TABLES Table Page 2.1 Characteristic Quantities for Nondimensionalization ::::::::::::: 15 2.2 Rotating Coordinates of the Libration Points in the Earth-Moon and Sun- Earth Systems ::::::::::::::::::::::::::::::::: 22 2.3 Value of the Jacobi Constant for each Libration Point in the Earth-Moon and Sun-Earth Systems :::::::::::::::::::::::::::: 22 4.1 Initial Conditions for the Distant Retrograde Orbits ::::::::::::: 73 4.2 Initial Conditions for the Period-3 Distant Retrograde Orbits :::::::: 76 4.3 Initial Conditions for the L1 Halo Orbits ::::::::::::::::::: 79 4.4 Initial Conditions for the L2 Halo Orbits ::::::::::::::::::: 82 5.1 Shadow events for a January 25, 2025 apoapsis insertion into a 4:1 synodic resonant NRHO. ::::::::::::::::::::::::::::::: 117 5.2 Shadow events for a January 25, 2025 apoapsis insertion into a 9:2 synodic resonant NRHO. ::::::::::::::::::::::::::::::: 118 5.3 Shadow events for a January 1, 2025 apoapsis insertion into a 9:2 synodic resonant NRHO. ::::::::::::::::::::::::::::::: 118 xi LIST OF FIGURES Figure Page 2.1 Geometry of a Three-Body System Defined in an Inertial Frame :::::: 12 2.2 Inertial and Rotating Coordinate Frames Definitions :::::::::::: 13 2.3 Location of the Libration Points as Viewed in the Rotating Frame ::::: 21 2.4 The ZVS and corresponding ZVC for JCL3 < JC < JCL2. The forbidden regions are shaded in grey. ::::::::::::::::::::::::::: 23 2.5 The ZVCs for decreasing values of Jacobi constant in the Earth-Moon three-body system. Earth and Moon are plotted twice their actual size. The forbidden regions are shaded in grey. ::::::::::::::::::: 26 2.6 Inertial Frame Geometry of an N-Body System ::::::::::::::: 28 3.1 Single Shooter Corrections Algorithm Schematic ::::::::::::::: 41 3.2 Multiple Shooter Corrections Algorithm Schematic ::::::::::::: 43 4.1 A distant retrograde orbit about the Moon in the Earth-Moon rotating frame, computed using a single shooting algorithm. ::::::::::::: 66 4.2 An L2 southern halo orbit computed using a single shooting method. ::: 67 4.3 Family of distant retrograde orbits about the Moon computed in the Earth- Moon system, colored by Jacobi constant value. ::::::::::::::: 74 4.4 Jacobi constant value and orbital period as a function of perilune radius over a range of periodic orbits in the distant retrograde orbit family in the Earth-Moon system. :::::::::::::::::::::::::::::: 74 xii Figure Page 4.5 Family of period-3 distant retrograde orbits computed in the Earth-Moon system, colored by Jacobi constant value. ::::::::::::::::::: 77 4.6 Jacobi constant value and orbital period as a function of maximum ex- cursion in thex ^ direction for representative orbits in the period-3 distant retrograde periodic orbit family in the Earth-Moon system. ::::::::: 77 4.7 Family of L1 halo orbits computed in the Earth-Moon system, colored by Jacobi constant value. ::::::::::::::::::::::::::::: 80 4.8 Jacobi constant value and orbital period as a function of perilune radius for the periodic L1 halo orbit family in the Earth-Moon system. :::::: 80 4.9 Family of L2 halo orbits computed in the Earth-Moon system, colored by Jacobi constant value. ::::::::::::::::::::::::::::: 83 4.10 Jacobi constant value and orbital period as a function of perilune radius for the periodic L2 halo orbit family in the Earth-Moon system. :::::: 83 4.11 Poincar´emap schematic, adapted from Perko. [46] :::::::::::::: 85 4.12 Transition of four revolutions of a CR3BP L2 halo orbit (red dashed) to the EMSJ ephemeris model (blue solid). ::::::::::::::::::: 89 4.13 An ephemeris trajectory with a wider spread relative to a corresponding CR3BP orbit. ::::::::::::::::::::::::::::::::::
Load more

Recommended publications
  • The Utilization of Halo Orbit§ in Advanced Lunar Operation§
    NASA TECHNICAL NOTE NASA TN 0-6365 VI *o M *o d a THE UTILIZATION OF HALO ORBIT§ IN ADVANCED LUNAR OPERATION§ by Robert W. Fmqcbar Godddrd Spctce Flight Center P Greenbelt, Md. 20771 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. JULY 1971 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA TN D-6365 5. Report Date Jul;v 1971. 6. Performing Organization Code 8. Performing Organization Report No, G-1025 10. Work Unit No. Goddard Space Flight Center 11. Contract or Grant No. Greenbelt, Maryland 20 771 13. Type of Report and Period Covered 2. Sponsoring Agency Name and Address Technical Note J National Aeronautics and Space Administration Washington, D. C. 20546 14. Sponsoring Agency Code 5. Supplementary Notes 6. Abstract Flight mechanics and control problems associated with the stationing of space- craft in halo orbits about the translunar libration point are discussed in some detail. Practical procedures for the implementation of the control techniques are described, and it is shown that these procedures can be carried out with very small AV costs. The possibility of using a relay satellite in.a halo orbit to obtain a continuous com- munications link between the earth and the far side of the moon is also discussed. Several advantages of this type of lunar far-side data link over more conventional relay-satellite systems are cited. It is shown that, with a halo relay satellite, it would be possible to continuously control an unmanned lunar roving vehicle on the moon's far side. Backside tracking of lunar orbiters could also be realized.
    [Show full text]
  • 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
    [Show full text]
  • Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit Emmanuel Blazquez, Laurent Beauregard, Stéphanie Lizy-Destrez, Finn Ankersen, Francesco Capolupo
    Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit Emmanuel Blazquez, Laurent Beauregard, Stéphanie Lizy-Destrez, Finn Ankersen, Francesco Capolupo To cite this version: Emmanuel Blazquez, Laurent Beauregard, Stéphanie Lizy-Destrez, Finn Ankersen, Francesco Capolupo. Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit. International Symposium on Space Flight Dynamics (ISSFD), Feb 2019, Melbourne, Australia. pp.1408-1413. hal-03200239 HAL Id: hal-03200239 https://hal.archives-ouvertes.fr/hal-03200239 Submitted on 16 Apr 2021 HAL is a multi-disciplinary open access L’archive ouverte pluridisciplinaire HAL, est archive for the deposit and dissemination of sci- destinée au dépôt et à la diffusion de documents entific research documents, whether they are pub- scientifiques de niveau recherche, publiés ou non, lished or not. The documents may come from émanant des établissements d’enseignement et de teaching and research institutions in France or recherche français ou étrangers, des laboratoires abroad, or from public or private research centers. publics ou privés. Open Archive Toulouse Archive Ouverte (OATAO) OATAO is an open access repository that collects the wor of some Toulouse researchers and ma es it freely available over the web where possible. This is an author's version published in: h ttps://oatao.univ-toulouse.fr/22865 To cite this version : Blazquez, Emmanuel and Beauregard, Laurent and Lizy-Destrez, Stéphanie and Ankersen, Finn and Capolupo, Francesco Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit.
    [Show full text]
  • Rideshare and the Orbital Maneuvering Vehicle: the Key to Low-Cost Lagrange-Point Missions
    SSC15-II-5 Rideshare and the Orbital Maneuvering Vehicle: the Key to Low-cost Lagrange-point Missions Chris Pearson, Marissa Stender, Christopher Loghry, Joe Maly, Valentin Ivanitski Moog Integrated Systems 1113 Washington Avenue, Suite 300, Golden, CO, 80401; 303 216 9777, extension 204 [email protected] Mina Cappuccio, Darin Foreman, Ken Galal, David Mauro NASA Ames Research Center PO Box 1000, M/S 213-4, Moffett Field, CA 94035-1000; 650 604 1313 [email protected] Keats Wilkie, Paul Speth, Trevor Jackson, Will Scott NASA Langley Research Center 4 West Taylor Street, Mail Stop 230, Hampton, VA, 23681; 757 864 420 [email protected] ABSTRACT Rideshare is a well proven approach, in both LEO and GEO, enabling low-cost space access through splitting of launch charges between multiple passengers. Demand exists from users to operate payloads at Lagrange points, but a lack of regular rides results in a deficiency in rideshare opportunities. As a result, such mission architectures currently rely on a costly dedicated launch. NASA and Moog have jointly studied the technical feasibility, risk and cost of using an Orbital Maneuvering Vehicle (OMV) to offer Lagrange point rideshare opportunities. This OMV would be launched as a secondary passenger on a commercial rocket into Geostationary Transfer Orbit (GTO) and utilize the Moog ESPA secondary launch adapter. The OMV is effectively a free flying spacecraft comprising a full suite of avionics and a propulsion system capable of performing GTO to Lagrange point transfer via a weak stability boundary orbit. In addition to traditional OMV ’tug’ functionality, scenarios using the OMV to host payloads for operation at the Lagrange points have also been analyzed.
    [Show full text]
  • Analysis of Interplanetary Transfers Between the Earth and Mars Halo Orbits
    Analysis of Interplanetary Transfers between the Earth and Mars Halo orbits M. Nakamiya (1), H. Yamakawa (2), D. J. Scheeres (3) and M. Yoshikawa (4) (1) Japan Aerospace Exploration Agency, 3-1-1 Sagamihara, Kanagawa 229-8510, Japan. nakamiya,[email protected] (2) University of Colorado, Boulder, Colorado 80309, USA. [email protected] (3) KyotoUniversity, Gokasho, Uji, Kyoto 611-0011, Japan. [email protected] (4) Japan Aerospace Exploration Agency, 3-1-1 Sagamihara, Kanagawa 229-8510, Japan. [email protected] Spacecraft interplanetary transportation systems using Earth and Mars Halo orbits are analyzed. This application is motivated by future proposals to place “deep space ports” at the Earth and Mars Halo orbits as space hub for planetary Mission. First, the feasibility of interplanetary trajectories between Earth Halo orbit and Mars Halo orbit was investigated, and such trajectories were designed with reasonable DV and flight time. Here, we utilize unstable and stable manifolds associated with the Halo orbit to approach the vicinity of the planet’s surface, and assume impulsive maneuver for escape and capture trajectories from/to Halo Orbits. Next, applying to Earth-Mars transportation system using spaceports on Halo orbits, we evaluated the system with respect to the mass of round-trip transfer. As a result, the transfer between the low Earth orbit and the low Mars orbit via the Earth Halo orbit and the Mars Halo orbit could be saved the fuel cost compared with a direct transfer between the low Earth orbit and the low Mars orbit 1. INTRODUCTION These days there has been great interest in the libration points of the circular restricted 3-body problem (CR3BP), where the gravity of the primary and secondary bodies and centrifugal force are balanced.
    [Show full text]
  • Station-Keeping and Momentum-Management on Halo Orbits Around L2: Linear-Quadratic Feedback and Model Predictive Control Approaches
    MITSUBISHI ELECTRIC RESEARCH LABORATORIES http://www.merl.com Station-keeping and momentum-management on halo orbits around L2: Linear-quadratic feedback and model predictive control approaches Kalabi´c,U.; Weiss, A.; Di Cairano, S.; Kolmanovsky, I.V. TR2015-002 January 11, 2015 Abstract The control of station-keeping and momentum-management is considered while tracking a halo orbit centered at the second Earth-Moon Lagrangian point. Multiple schemes based on linear-quadratic feedback control and model predictive control (MPC) are considered and it is shown that the method based on periodic MPC performs best for position tracking. The scheme is then extended to incorporate attitude control requirements and numerical simulations are presented demonstrating that the scheme is able to achieve simultaneous tracking of a halo orbit and dumping of momentum while enforcing tight constraints on pointing error. AAS/AIAA Space Flight Mechanics Meeting This work may not be copied or reproduced in whole or in part for any commercial purpose. Permission to copy in whole or in part without payment of fee is granted for nonprofit educational and research purposes provided that all such whole or partial copies include the following: a notice that such copying is by permission of Mitsubishi Electric Research Laboratories, Inc.; an acknowledgment of the authors and individual contributions to the work; and all applicable portions of the copyright notice. Copying, reproduction, or republishing for any other purpose shall require a license with payment of
    [Show full text]
  • Utilizing the Deep Space Gateway As a Platform for Deploying Cubesats Into Lunar Orbit
    Deep Space Gateway Science Workshop 2018 (LPI Contrib. No. 2063) 3138.pdf UTILIZING THE DEEP SPACE GATEWAY AS A PLATFORM FOR DEPLOYING CUBESATS INTO LUNAR ORBIT. Fisher, Kenton R.1; NASA Johnson Space Center 2101 E NASA Pkwy, Houston TX 77508. [email protected] Introduction: The capability to deploy CubeSats from the Deep Space Gateway (DSG) may significantly increase access to the Moon for lower- cost science missions. CubeSats are beginning to demonstrate high science return for reduced cost. The DSG could utilize an existing ISS deployment method, assuming certain capabilities that are present at the ISS are also available at the DSG. Science and Technology Applications: There are many interesting potential applications for lunar CubeSat missions. A constellation of CubeSats Figure 1: Nanoracks CubeSat Deployer (NRCSD) being could provide a lunar GPS system for support of positioned by the Japanese Remote Manipulator System surface exploration missions. CubeSats could be used (JRMS) for deployment on ISS [4]. to provide communications relays for missions to the Present Lunar CubeSat Architecture: lunar far side. Potential science applications include While there are many opportunities for deploying using a CubeSat to produce mapping of the magnetic CubeSats into Earth orbits, the current options for fields of lunar swirls or to further map lunar surface deployment to lunar orbits are limited. Propulsive volatiles. requirements for trans-lunar injection (TLI) from International Space Station CubeSats: The Earth-orbit and lunar orbital insertion (LOI) International Space Station (ISS) has been a major significantly increase the cost, mass, and complexity platform for deploying CubeSats in recent years.
    [Show full text]
  • Aas 19-266 Designing Low-Thrust Enabled Trajectories for A
    AAS 19-266 DESIGNING LOW-THRUST ENABLED TRAJECTORIES FOR A HELIOPHYSICS SMALLSAT MISSION TO SUN-EARTH L5 Ian Elliott,∗ Christoper Sullivan Jr.,∗ Natasha Bosanac,y Jeffrey R. Stuart,z and Farah Alibayx A small satellite deployed to Sun-Earth L5 could serve as a low-cost platform to observe solar phenomena such as coronal mass ejections. However, the small satellite platform introduces significant challenges into the trajectory design pro- cess via limited thrusting capabilities, power and operational constraints, and fixed deployment conditions. To address these challenges, a strategy employing dynam- ical systems theory is used to design a low-thrust-enabled trajectory for a small satellite to reach the Sun-Earth L5 region. This procedure is demonstrated for a small satellite that launches as a secondary payload with a larger spacecraft des- tined for a Sun-Earth L2 halo orbit. INTRODUCTION Low-thrust spacecraft trajectory design has the potential to enable targeted science missions for small satellites to explore solar processes. In fact, the Sun-Earth (SE) L5 equilibrium point, located at the vertex of an equilateral triangle formed with the Sun and the Earth, is a candidate for locating a SmallSat for a heliophysics mission. At SE L5, a spacecraft would have a currently unavailable perspective of the three-dimensional structure of the coronal mass ejections off the Sun-Earth line.1 Additionally, due to the rotation of the Sun, early warning of solar storms is possible from SE L5.2 This equilibrium point has previously been visited by the Solar TErrestrial RElations Observatory B (STEREO-B) spacecraft which drifted through the SE L5 region, but did not maintain long- term bounded motion.3;4 With the increasing availability of rideshare opportunities and the recent demonstration of CubeSats operations in deep space during the Mars Cube One (MarCO) mission,5 small satellites may soon have the potential to be deployed to SE L5 to perform heliophysics-based science objectives.
    [Show full text]
  • Rendezvous in Cis-Lunar Space Near Rectilinear Halo Orbit: Dynamics and Control Issues
    aerospace Article Rendezvous in Cis-Lunar Space near Rectilinear Halo Orbit: Dynamics and Control Issues Giordana Bucchioni † and Mario Innocenti *,† Department of Information Engineering, University of Pisa, 56126 Pisa, Italy; [email protected] * Correspondence: [email protected] † These authors contributed equally to this work. Abstract: The paper presents the development of a fully-safe, automatic rendezvous strategy between a passive vehicle and an active one orbiting around the Earth–Moon L2 Lagrangian point. This is one of the critical phases of future missions to permanently return to the Moon, which are of interest to the majority of space organizations. The first step in the study is the derivation of a suitable full 6-DOF relative motion model in the Local Vertical Local Horizontal reference frame, most suitable for the design of the guidance. The main dynamic model is approximated using both the elliptic and circular three-body motion, due to the contribution of Earth and Moon gravity. A rather detailed set of sensors and actuator dynamics was also implemented in order to ensure the reliability of the guidance algorithms. The selection of guidance and control is presented, and evaluated using a sample scenario as described by ESA’s HERACLES program. The safety, in particular the passive safety, concept is introduced and different techniques to guarantee it are discussed that exploit the ideas of stable and unstable manifolds to intrinsically guarantee some properties at each hold-point, in which the rendezvous trajectory is divided. Finally, the rendezvous dynamics are validated using available Ephemeris models in order to verify the validity of the results and their limitations for Citation: Bucchioni, G.; Innocenti, M.
    [Show full text]
  • Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit
    Student Paper NON-PEER REVIEW Rendezvous Design in a Cislunar Near Rectilinear Halo Orbit Emmanuel Blazquez 1, Laurent Beauregard 1, Stéphanie Lizy-Destrez 1, Finn Ankersen 2 and Francesco Capolupo 3 1 Institut Supérieur de l’Aéronautique et de l’Espace (ISAE-SUPAERO), 10 Avenue Edouard Belin, 31400 Toulouse, France 2 European Space Agency, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands 3 Airbus Defence and Space, 31 rue des Cosmonautes, 31400 Toulouse, France Abstract In the context of future Human Spaceflight exploration missions, Rendezvous and Docking (RVD) activities are critical for the assembly and maintenance of cislunar structures. The scope of this research is to investigate the specifics of orbits of interest for RVD in the cislunar realm and to propose novel strategies to safely perform these kinds of operations. This paper focuses on far rendezvous approaches and passively safe drift trajectories in the ephemeris model. The goal is to exhibit phasing orbit requirements to ensure safe far approach. Ephemeris representations of Near Rectilinear Halo Orbits (NRHOs) were derived using multiple shooting and adaptive receding-horizon targeting algorithms. Simulations showed significant drift and overlapping properties for phasing and target orbits of interest, motivating the search for safe natural drift trajectories, using impact prediction strategies. Keywords: Rendezvous, Near Rectilinear Halo Orbits, Ephemeris, Trajectory design, Safety. Introduction The future Lunar Orbital Platform Gateway will require cargo delivery and crew exchange operational activities, both which rely critically on Rendezvous and Docking operations (RVD) with modules such as Orion [1]. The RVD problem in Keplerian dynamics has been extensively discussed. However, to this date, no operational RVD has yet been performed in the vicinity of a Lagrangian point.
    [Show full text]
  • Long Term Missions at the Sun-Earth Libration Point L1: Ace, Soho, and Wind Aas 11
    AAS 11- 485 LONG TERM MISSIONS AT THE SUN-EARTH LIBRATION POINT L1: ACE, SOHO, AND WIND Craig E. Roberts* Three heliophysics missionsthe Solar Heliospheric Observatory (SOHO), the Advanced Composition Explorer (ACE), and the Global Geoscience WINDhave been orbiting the Sun-Earth interior libration point L1 conti- nuously since 1996, 1997, and 2004, respectively. ACE and WIND (both NASA missions) and SOHO (an ESA-NASA joint mission) are all operated from the NASA Goddard Space Flight Center Flight Dynamics Facility. While ACE and SOHO have been dedicated libration point orbiters since their launches, WIND prior to 2004 flew a remarkable 10-year deep-space trajectory that featured 38 targeted lunar flybys. The L1 orbits and the mission histories of the three spacecraft are briefly reviewed, and the station-keeping techniques and orbit maneuver experience are discussed. INTRODUCTION Three heliophysics missionsthe Solar Heliospheric Observatory (SOHO), the Advanced Composition Explorer (ACE), and the Global Geoscience International Physics Laboratory (un- iversally referred to by its nickname, WIND)have been orbiting the Sun-Earth interior libration point L1 continuously since 1997, 1996, and 2004, respectively.† ACE and WIND (both NASA missions) and SOHO (an ESA-NASA joint mission) all have their maneuver operations con- ducted from the NASA Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF). While ACE and SOHO have been dedicated libration point orbiters since their launches, WIND has had also a remarkable 10-year career flying a deep-space, multiple lunar-flyby trajectory prior to 2004. That era featured 38 targeted lunar flybys during an initial Double-Lunar-Swingby (DLS) phase, a Distant Prograde Orbit (DPO) phase, and looping excursions to both the Sun- Earth L1 and L2 collinear points before its final insertion into an L1 orbit.
    [Show full text]
  • Generating Exploration Mission-3 Trajectories to a 9:2 Nrho
    GENERATING EXPLORATION MISSION-3 TRAJECTORIES TO A 9:2 NRHO USING MACHINE LEARNING A Thesis presented to the Faculty of California Polytechnic State University, San Luis Obispo In Partial Fulfillment of the Requirements for the Degree Master of Science in Aerospace Engineering by Esteban Guzman December 2018 c 2018 Esteban Guzman ALL RIGHTS RESERVED ii COMMITTEE MEMBERSHIP TITLE: Generating Exploration Mission-3 Trajectories to a 9:2 NRHO using Machine Learning AUTHOR: Esteban Guzman DATE SUBMITTED: December 2018 COMMITTEE CHAIR: Eric Mehiel, Ph.D. Professor of Aerospace Engineering COMMITTEE MEMBER: Kira Jorgensen Abercromby, Ph.D. Professor of Aerospace Engineering COMMITTEE MEMBER: Amelia Greig, Ph.D. Professor of Aerospace Engineering COMMITTEE MEMBER: Jeffrey P. Gutkowski, EG5 Deputy Branch Chief NASA Johnson Space Center iii Abstract Generating Exploration Mission-3 Trajectories to a 9:2 NRHO using Machine Learning Esteban Guzman The purpose of this thesis is to design a machine learning algorithm platform that provides expanded knowledge of mission availability through a launch season by im- proving trajectory resolution and introducing launch mission forecasting. The specific scenario addressed in this paper is one in which data is provided for four determin- istic translational maneuvers through a mission to a Near Rectilinear Halo Orbit (NRHO) with a 9:2 synodic frequency. Current launch availability knowledge un- der NASA’s Orion Orbit Performance Team is established by altering optimization variables associated to given reference launch epochs. This current method can be an abstract task and relies on an orbit analyst to structure a mission based off an established mission design methodology associated to the performance of Orion and NASA’s Space Launch System.
    [Show full text]