DOCSLIB.ORG

Moon & Mars Orbiting Spinning Tether Transport (MMOSTT) (PDF

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Moon & Mars Orbiting Spinning Tether Transport (MMOSTT) (PDF Tethers Unlimited, Inc. MMOSTT Final Report TABLE OF CONTENTS I. INTRODUCTION ..........................................................................................................................1 II. BACKGROUND .............................................................................................................................2 A. MOMENTUM-EXCHANGE TETHERS.................................................................................................2 B. ELECTRODYNAMIC REBOOST ...........................................................................................................2 C. KEY ADVANTAGES .............................................................................................................................3 D. SUMMARY OF PHASE I RESULTS .......................................................................................................4 III. PHASE II TECHNICAL OBJECTIVES......................................................................................6 IV. SUMMARY OF RESULTS OF THE PHASE II EFFORT........................................................6 A. ROADMAP FOR DEVELOPMENT OF A EARTH-MOON-MARS TETHER TRANSPORT SYSTEM ......6 B. DESIGN AND SIMULATION OF A TETHER BOOST FACILITY FOR LEO⇒GTO PAYLOAD TRANSPORT ........................................................................................................................................8 C. TETHER FACILITY REBOOST............................................................................................................ 11 D. DEVELOPMENT AND SIMULATION OF TETHER RENDEZVOUS METHODS .................................13 E. TETHER SYSTEMS FOR INTERPLANETARY TRANSPORT ................................................................15 F. µTORQUE:LOW-COST MOMENTUM-EXCHANGE/ELECTRODYNAMIC-PROPULSION DEMONSTRATION ............................................................................................................................17 V. PUBLICATIONS...........................................................................................................................18 VI. CONCLUSIONS ...........................................................................................................................19 Appendix A Tether Boost Facilities for In-Space Transportation, 2001 NIAC Fellows Conference Presentation. Appendix B. “Commercial Development of a Tether Transport System”, AIAA Paper 2000- 3842, Presented at the 36th Joint Propulsion Conference, Huntsville AL, July 18, 2000. Appendix C. “Design and Simulation of a Tether Boost Facility for LEO⇒GTO Payload Transport”, AIAA Paper 2000-3866, Presented at the 36th Joint Propulsion Conference, Huntsville AL, July 19, 2000. Appendix D. “The Cislunar Tether Transport System Architecture”, Paper presented at the 2nd Lunar Development Conference, Las Vegas, NV, July 20, 2000. Appendix E. Tether Boost Facility Design Study Interim Report, by Boeing/RSS. Appendix F. Tether Boost Facility Design Study Final Report, by Boeing/RSS. Appendix G. Tether Rendezvous Studies. Appendix H. 2000 NIAC Fellows Conference Presentation. Appendix I. “Rapid Interplanetary Tether Transport System”. Paper IAF-99-A.5.10, presented at the 50th International Astronautical Congress, 4-8 Oct 1999, Amsterdam, The Netherlands. This paper summarized the results of the Phase I effort. Appendix J. “Tether Systems for Satellite Deployment and Disposal”, Paper IAF-00-S.6.04, presented at the 51st International Astronautical Congress, 2-6 Oct 2000, Rio de Janiero, Brazil. Appendix K. Tether Facility Reboost Study. 1 Tethers Unlimited, Inc. MMOSTT Final Report Appendix L. The µTORQUE Momentum-Exchange Tether Experiment, Paper submitted to the 2002 STAIF Conference. Appendix M. TetherSim™: Tether Transport System Dynamics Verification Through Simulation. Appendix N. Momentum-Exchange/Electrodynamic Reboost Tether Facility for Deployment of Microsatellites to GEO and the Moon, paper presented at the 2001 STAIF Conference, Albuquerque, NM. Appendix O. µSatellite Tether Boost Facility, 2001 STAIF Conference presentation , Albuquerque, NM. Appendix P. Interplanetary Tether Transport Overview, Paper presented at the 2001 Space Mechanics Conference. Appendix Q. Application of Synergistic Multipayload Assistance with Rotating Tethers (SMART) Concept to Outer Planet Exploration, forum on innovative approaches to outer planetary exploration 2001-2020, 21-23 February 2001, Lunar and Planetary Institute, Houston, Texas. Appendix R. Mars-Earth Rapid Interplanetary Tether Transport (MERITT) Architecture, Paper presented at 2001 STAIF Conference. 2 Tethers Unlimited, Inc. MMOSTT Final Report I. INTRODUCTION Momentum-Exchange/Electrodynamic-Reboost (MXER) Tethers can provide a fully reusable, zero-propellant infrastructure for in-space transportation that will reduce by an order of magnitude or more the costs of delivering payloads to geostationary orbit, the Moon, Mars, and other destinations. This Phase II NIAC research program has continued the development of a tether-based architecture for in-space propulsion to service transportation needs in the Earth-Moon-Mars system. This tether architecture will utilize momentum-exchange techniques and electrodynamic tether propulsion to transport multiple payloads with little or no propellant consumption. The tether transport architecture is designed to be deployed incrementally, with each component able to perform a useful revenue-generating mission to help fund the deployment of the rest of the system. The Phase II effort has focused on the design of the first component of this architecture, a Tether Boost Facility optimized for transferring payloads from low Earth orbit (LEO) to geostationary transfer orbit (GTO). The resultant system concept uses a modular design that enables a single launch to deploy a fully-operational tether boost facility which can later be augmented to increase its payload capacity. The first component of the tether boost facility will be able to toss 2,500 kg payloads from a low-LEO initial orbit to GTO. This same facility will also be capable of boosting 1,000 kg payloads to lunar transfer orbit (LTO) or to escape via a lunar swingby. Additional launches of essentially identical modules can increase the payload capacity of the Tether Boost Facility to enable it to boost larger satellites and, eventually, manned spacecraft. This Tether Boost Facility can, in turn, be used to deploy components of additional tether facilities at the Moon and Mars, providing an infrastructure for frequent, low-cost transport between the Earth, the Moon, and Mars. 1 Tethers Unlimited, Inc. MMOSTT Final Report II. BACKGROUND Space tethers can accomplish propellantless propulsion through two mechanisms, through momentum-exchange between two space objects, and through electrodynamic interactions with a planetary magnetic field. A. Momentum-Exchange Tethers In a momentum-exchange tether system, a long, thin, high-strength cable is deployed in orbit and set into rotation around a central body. If the tether facility is placed in an elliptical orbit and its rotation is timed so that the tether is oriented vertically below the central body and swinging backwards when the facility reaches perigee, then a grapple assembly located at the tether tip can rendezvous with and capture a payload moving in a lower orbit, as illustrated in Figure 1. Half a rotation later, the tether can release the payload, tossing it into a higher energy orbit. This concept is termed a momentum-exchange tether because when the tether picks up and tosses the payload, it transfers some of its orbital energy and momentum to the payload, resulting in a drop in the tether facility’s apogee. Figure 1. Concept of operation of a momentum-exchange tether facility. Orbits are depicted conceptually from the perspective of an observer on the Earth. B. Electrodynamic Reboost In order for the tether facility to boost multiple payloads, it must have the capability to restore its orbital energy and momentum after each payload transfer operation. If the tether facility has a power supply, and a portion of the tether contains conducting wire, then the power supply can drive current along the tether so as to generate thrust through electro- dynamic interactions with the Earth's magnetic field, as illustrated in Figure 2. By properly controlling the tether current during an orbit, the tether facility can reboost itself to its original orbit. The tether facility essentially serves as a large "orbital energy battery," allowing solar energy to be converted to orbital energy gradually over a long period of time and then rapidly transferred to the payload. 2 Tethers Unlimited, Inc. MMOSTT Final Report orbital thrust velocity applied voltage current magnetic field line Figure 2. Electrodynamic tether thrust generation. C. Key Advantages A tether transportation system has several advantages compared to conventional and other advanced in-space propulsion systems: 1. (Near) Zero Propellant Usage Chief among these advantages is the ability to eliminate the need for propellant expenditure to perform payload transfers. Of course, some propellant expenditure will be needed for trajectory corrections and rendezvous maneuvering, but with proper system design these requirements will be very small, a few tens of meters per second. The ability to cut several thousands of meters per second from the ∆V needed to deliver a payload to its destination can enable customers to utilize much smaller launch vehicles than would be required with a rocket- only
Load more

Recommended publications
  • Machine Learning Regression for Estimating Characteristics of Low-Thrust Transfers
    MACHINE LEARNING REGRESSION FOR ESTIMATING CHARACTERISTICS OF LOW-THRUST TRANSFERS A Thesis Presented to The Academic Faculty By Gene L. Chen In Partial Fulfillment of the Requirements for the Degree Master of Science in the School of Aerospace Engineering Georgia Institute of Technology May 2019 Copyright c Gene L. Chen 2019 MACHINE LEARNING REGRESSION FOR ESTIMATING CHARACTERISTICS OF LOW-THRUST TRANSFERS Approved by: Dr. Dimitri Mavris, Advisor Guggenheim School of Aerospace Engineering Georgia Institute of Technology Dr. Alicia Sudol Guggenheim School of Aerospace Engineering Georgia Institute of Technology Dr. Michael Steffens Guggenheim School of Aerospace Engineering Georgia Institute of Technology Date Approved: April 15, 2019 To my parents, thanks for the support. ACKNOWLEDGEMENTS Firstly, I would like to thank Dr. Dimitri Mavris for giving me the opportunity to pursue a Master’s degree at the Aerospace Systems Design Laboratory. It has been quite the experience. Also, many thanks to my committee members — Dr. Alicia Sudol and Dr. Michael Steffens — for taking the time to review my work and give suggestions on how to improve - it means a lot to me. Additionally, thanks to Dr. Patel and Dr. Antony for helping me understand their work in trajectory optimization. iv TABLE OF CONTENTS Acknowledgments . iv List of Tables . vii List of Figures . viii Chapter 1: Introduction and Motivation . 1 1.1 Thesis Overview . 4 Chapter 2: Background . 5 Chapter 3: Approach . 17 3.1 Scenario Definition . 17 3.2 Spacecraft Dynamics . 18 3.3 Choice Between Direct and Indirect Method . 19 3.3.1 Chebyshev polynomial method . 20 3.3.2 Sims-Flanagan method .
    [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]
  • Call for M5 Missions
    ESA UNCLASSIFIED - For Official Use M5 Call - Technical Annex Prepared by SCI-F Reference ESA-SCI-F-ESTEC-TN-2016-002 Issue 1 Revision 0 Date of Issue 25/04/2016 Status Issued Document Type Distribution ESA UNCLASSIFIED - For Official Use Table of contents: 1 Introduction .......................................................................................................................... 3 1.1 Scope of document ................................................................................................................................................................ 3 1.2 Reference documents .......................................................................................................................................................... 3 1.3 List of acronyms ..................................................................................................................................................................... 3 2 General Guidelines ................................................................................................................ 6 3 Analysis of some potential mission profiles ........................................................................... 7 3.1 Introduction ............................................................................................................................................................................. 7 3.2 Current European launchers ...........................................................................................................................................
    [Show full text]
  • Launch and Deployment Analysis for a Small, MEO, Technology Demonstration Satellite
    46th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2008-1131 7 – 10 January 20006, Reno, Nevada Launch and Deployment Analysis for a Small, MEO, Technology Demonstration Satellite Stephen A. Whitmore* and Tyson K. Smith† Utah State University, Logan, UT, 84322-4130 A trade study investigating the economics, mass budget, and concept of operations for delivery of a small technology-demonstration satellite to a medium-altitude earth orbit is presented. The mission requires payload deployment at a 19,000 km orbit altitude and an inclination of 55o. Because the payload is a technology demonstrator and not part of an operational mission, launch and deployment costs are a paramount consideration. The payload includes classified technologies; consequently a USA licensed launch system is mandated. A preliminary trade analysis is performed where all available options for FAA-licensed US launch systems are considered. The preliminary trade study selects the Orbital Sciences Minotaur V launch vehicle, derived from the decommissioned Peacekeeper missile system, as the most favorable option for payload delivery. To meet mission objectives the Minotaur V configuration is modified, replacing the baseline 5th stage ATK-37FM motor with the significantly smaller ATK Star 27. The proposed design change enables payload delivery to the required orbit without using a 6th stage kick motor. End-to-end mass budgets are calculated, and a concept of operations is presented. Monte-Carlo simulations are used to characterize the expected accuracy of the final orbit.
    [Show full text]
  • 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
    [Show full text]
  • AFSPC-CO TERMINOLOGY Revised: 12 Jan 2019
    AFSPC-CO TERMINOLOGY Revised: 12 Jan 2019 Term Description AEHF Advanced Extremely High Frequency AFB / AFS Air Force Base / Air Force Station AOC Air Operations Center AOI Area of Interest The point in the orbit of a heavenly body, specifically the moon, or of a man-made satellite Apogee at which it is farthest from the earth. Even CAP rockets experience apogee. Either of two points in an eccentric orbit, one (higher apsis) farthest from the center of Apsis attraction, the other (lower apsis) nearest to the center of attraction Argument of Perigee the angle in a satellites' orbit plane that is measured from the Ascending Node to the (ω) perigee along the satellite direction of travel CGO Company Grade Officer CLV Calculated Load Value, Crew Launch Vehicle COP Common Operating Picture DCO Defensive Cyber Operations DHS Department of Homeland Security DoD Department of Defense DOP Dilution of Precision Defense Satellite Communications Systems - wideband communications spacecraft for DSCS the USAF DSP Defense Satellite Program or Defense Support Program - "Eyes in the Sky" EHF Extremely High Frequency (30-300 GHz; 1mm-1cm) ELF Extremely Low Frequency (3-30 Hz; 100,000km-10,000km) EMS Electromagnetic Spectrum Equitorial Plane the plane passing through the equator EWR Early Warning Radar and Electromagnetic Wave Resistivity GBR Ground-Based Radar and Global Broadband Roaming GBS Global Broadcast Service GEO Geosynchronous Earth Orbit or Geostationary Orbit ( ~22,300 miles above Earth) GEODSS Ground-Based Electro-Optical Deep Space Surveillance
    [Show full text]
  • Aas 11-509 Artemis Lunar Orbit Insertion and Science Orbit Design Through 2013
    AAS 11-509 ARTEMIS LUNAR ORBIT INSERTION AND SCIENCE ORBIT DESIGN THROUGH 2013 Stephen B. Broschart,∗ Theodore H. Sweetser,y Vassilis Angelopoulos,z David C. Folta,x and Mark A. Woodard{ As of late-July 2011, the ARTEMIS mission is transferring two spacecraft from Lissajous orbits around Earth-Moon Lagrange Point #1 into highly-eccentric lunar science orbits. This paper presents the trajectory design for the transfer from Lissajous orbit to lunar orbit in- sertion, the period reduction maneuvers, and the science orbits through 2013. The design accommodates large perturbations from Earth’s gravity and restrictive spacecraft capabili- ties to enable opportunities for a range of heliophysics and planetary science measurements. The process used to design the highly-eccentric ARTEMIS science orbits is outlined. The approach may inform the design of future eccentric orbiter missions at planetary moons. INTRODUCTION The Acceleration, Reconnection, Turbulence and Electrodynamics of the Moons Interaction with the Sun (ARTEMIS) mission is currently operating two spacecraft in lunar orbit under funding from the Heliophysics and Planetary Science Divisions with NASA’s Science Missions Directorate. ARTEMIS is an extension to the successful Time History of Events and Macroscale Interactions during Substorms (THEMIS) mission [1] that has relocated two of the THEMIS spacecraft from Earth orbit to the Moon [2, 3, 4]. ARTEMIS plans to conduct a variety of scientific studies at the Moon using the on-board particle and fields instrument package [5, 6]. The final portion of the ARTEMIS transfers involves moving the two ARTEMIS spacecraft, known as P1 and P2, from Lissajous orbits around the Earth-Moon Lagrange Point #1 (EML1) to long-lived, eccentric lunar science orbits with periods of roughly 28 hr.
    [Show full text]
  • Design of the Trajectory of a Tether Mission to Saturn
    DOI: 10.13009/EUCASS2019-510 8TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: Design of the trajectory of a tether mission to Saturn ⋆ Riccardo Calaon , Elena Fantino§, Roberto Flores ‡ and Jesús Peláez† ⋆Dept. of Industrial Engineering, University of Padova Via Gradenigio G/a -35131- Padova, Italy §Dept. of Aerospace Engineering, Khalifa University Abu Dhabi, Unites Arab Emirates ‡Inter. Center for Numerical Methods in Engineering Campus Norte UPC, Gran Capitán s/n, Barcelona, Spain †Space Dynamics Group, UPM School of Aeronautical and Space Engineering, Pz. Cardenal Cisneros 3, Madrid, Spain [email protected], [email protected], rfl[email protected], [email protected] †Corresponding author Abstract This paper faces the design of the trajectory of a tether mission to Saturn. To reduce the hyperbolic excess velocity at arrival to the planet, a gravity assist at Jupiter is included. The Earth-to-Jupiter portion of the transfer is unpropelled. The Jupiter-to-Saturn trajectory has two parts: a first coasting arc followed by a second arc where a low thrust engine is switched on. An optimization process provides the law of thrust control that minimizes the velocity relative to Saturn at arrival, thus facilitating the tethered orbit insertion. 1. Introduction The outer planets are of particular interest in terms of what they can reveal about the origin and evolution of our solar system. They are also local analogues for the many extra-solar planets that have been detected over the past twenty years. The study of these planets furthers our comprehension of our neighbourhood and provides the foundations to understand distant planetary systems.
    [Show full text]
  • The Minor Planet Bulletin
    THE MINOR PLANET BULLETIN OF THE MINOR PLANETS SECTION OF THE BULLETIN ASSOCIATION OF LUNAR AND PLANETARY OBSERVERS VOLUME 36, NUMBER 3, A.D. 2009 JULY-SEPTEMBER 77. PHOTOMETRIC MEASUREMENTS OF 343 OSTARA Our data can be obtained from http://www.uwec.edu/physics/ AND OTHER ASTEROIDS AT HOBBS OBSERVATORY asteroid/. Lyle Ford, George Stecher, Kayla Lorenzen, and Cole Cook Acknowledgements Department of Physics and Astronomy University of Wisconsin-Eau Claire We thank the Theodore Dunham Fund for Astrophysics, the Eau Claire, WI 54702-4004 National Science Foundation (award number 0519006), the [email protected] University of Wisconsin-Eau Claire Office of Research and Sponsored Programs, and the University of Wisconsin-Eau Claire (Received: 2009 Feb 11) Blugold Fellow and McNair programs for financial support. References We observed 343 Ostara on 2008 October 4 and obtained R and V standard magnitudes. The period was Binzel, R.P. (1987). “A Photoelectric Survey of 130 Asteroids”, found to be significantly greater than the previously Icarus 72, 135-208. reported value of 6.42 hours. Measurements of 2660 Wasserman and (17010) 1999 CQ72 made on 2008 Stecher, G.J., Ford, L.A., and Elbert, J.D. (1999). “Equipping a March 25 are also reported. 0.6 Meter Alt-Azimuth Telescope for Photometry”, IAPPP Comm, 76, 68-74. We made R band and V band photometric measurements of 343 Warner, B.D. (2006). A Practical Guide to Lightcurve Photometry Ostara on 2008 October 4 using the 0.6 m “Air Force” Telescope and Analysis. Springer, New York, NY. located at Hobbs Observatory (MPC code 750) near Fall Creek, Wisconsin.
    [Show full text]
  • Orbital Fueling Architectures Leveraging Commercial Launch Vehicles for More Affordable Human Exploration
    ORBITAL FUELING ARCHITECTURES LEVERAGING COMMERCIAL LAUNCH VEHICLES FOR MORE AFFORDABLE HUMAN EXPLORATION by DANIEL J TIFFIN Submitted in partial fulfillment of the requirements for the degree of: Master of Science Department of Mechanical and Aerospace Engineering CASE WESTERN RESERVE UNIVERSITY January, 2020 CASE WESTERN RESERVE UNIVERSITY SCHOOL OF GRADUATE STUDIES We hereby approve the thesis of DANIEL JOSEPH TIFFIN Candidate for the degree of Master of Science*. Committee Chair Paul Barnhart, PhD Committee Member Sunniva Collins, PhD Committee Member Yasuhiro Kamotani, PhD Date of Defense 21 November, 2019 *We also certify that written approval has been obtained for any proprietary material contained therein. 2 Table of Contents List of Tables................................................................................................................... 5 List of Figures ................................................................................................................. 6 List of Abbreviations ....................................................................................................... 8 1. Introduction and Background.................................................................................. 14 1.1 Human Exploration Campaigns ....................................................................... 21 1.1.1. Previous Mars Architectures ..................................................................... 21 1.1.2. Latest Mars Architecture .........................................................................
    [Show full text]
  • Cislunar Tether Transport System
    FINAL REPORT on NIAC Phase I Contract 07600-011 with NASA Institute for Advanced Concepts, Universities Space Research Association CISLUNAR TETHER TRANSPORT SYSTEM Report submitted by: TETHERS UNLIMITED, INC. 8114 Pebble Ct., Clinton WA 98236-9240 Phone: (206) 306-0400 Fax: -0537 email: [email protected] www.tethers.com Report dated: May 30, 1999 Period of Performance: November 1, 1998 to April 30, 1999 PROJECT SUMMARY PHASE I CONTRACT NUMBER NIAC-07600-011 TITLE OF PROJECT CISLUNAR TETHER TRANSPORT SYSTEM NAME AND ADDRESS OF PERFORMING ORGANIZATION (Firm Name, Mail Address, City/State/Zip Tethers Unlimited, Inc. 8114 Pebble Ct., Clinton WA 98236-9240 [email protected] PRINCIPAL INVESTIGATOR Robert P. Hoyt, Ph.D. ABSTRACT The Phase I effort developed a design for a space systems architecture for repeatedly transporting payloads between low Earth orbit and the surface of the moon without significant use of propellant. This architecture consists of one rotating tether in elliptical, equatorial Earth orbit and a second rotating tether in a circular low lunar orbit. The Earth-orbit tether picks up a payload from a circular low Earth orbit and tosses it into a minimal-energy lunar transfer orbit. When the payload arrives at the Moon, the lunar tether catches it and deposits it on the surface of the Moon. Simultaneously, the lunar tether picks up a lunar payload to be sent down to the Earth orbit tether. By transporting equal masses to and from the Moon, the orbital energy and momentum of the system can be conserved, eliminating the need for transfer propellant. Using currently available high-strength tether materials, this system could be built with a total mass of less than 28 times the mass of the payloads it can transport.
    [Show full text]
  • Method for Rapid Interplanetary Trajectory Analysis Using Δv Maps with Flyby Options
    View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by DSpace@MIT METHOD FOR RAPID INTERPLANETARY TRAJECTORY ANALYSIS USING ΔV MAPS WITH FLYBY OPTIONS TAKUTO ISHIMATSU*, JEFFREY HOFFMAN†, AND OLIVIER DE WECK‡ Department of Aeronautics and Astronautics, Massachusetts Institute of Technology, 77 Massachusetts Avenue, Cambridge, MA 02139, USA. Email: [email protected]*, [email protected]†, [email protected]‡ This paper develops a convenient tool which is capable of calculating ballistic interplanetary trajectories with planetary flyby options to create exhaustive ΔV contour plots for both direct trajectories without flybys and flyby trajectories in a single chart. The contours of ΔV for a range of departure dates (x-axis) and times of flight (y-axis) serve as a “visual calendar” of launch windows, which are useful for the creation of a long-term transportation schedule for mission planning purposes. For planetary flybys, a simple powered flyby maneuver with a reasonably small velocity impulse at periapsis is allowed to expand the flyby mission windows. The procedure of creating a ΔV contour plot for direct trajectories is a straightforward full-factorial computation with two input variables of departure and arrival dates solving Lambert’s problem for each combination. For flyby trajectories, a “pseudo full-factorial” computation is conducted by decomposing the problem into two separate full-factorial computations. Mars missions including Venus flyby opportunities are used to illustrate the application of this model for the 2020-2040 time frame. The “competitiveness” of launch windows is defined and determined for each launch opportunity. Keywords: Interplanetary trajectory, C3, delta-V, pork-chop plot, launch window, Mars, Venus flyby 1.
    [Show full text]