Description of the Model
Total Page:16
File Type:pdf, Size:1020Kb
Load more
Recommended publications
-
Reduction of Saturn Orbit Insertion Impulse Using Deep-Space Low Thrust
Reduction of Saturn Orbit Insertion Impulse using Deep-Space Low Thrust Elena Fantino ∗† Khalifa University of Science and Technology, Abu Dhabi, United Arab Emirates. Roberto Flores‡ International Center for Numerical Methods in Engineering, 08034 Barcelona, Spain. Jesús Peláez§ and Virginia Raposo-Pulido¶ Technical University of Madrid, 28040 Madrid, Spain. Orbit insertion at Saturn requires a large impulsive manoeuver due to the velocity difference between the spacecraft and the planet. This paper presents a strategy to reduce dramatically the hyperbolic excess speed at Saturn by means of deep-space electric propulsion. The inter- planetary trajectory includes a gravity assist at Jupiter, combined with low-thrust maneuvers. The thrust arc from Earth to Jupiter lowers the launch energy requirement, while an ad hoc steering law applied after the Jupiter flyby reduces the hyperbolic excess speed upon arrival at Saturn. This lowers the orbit insertion impulse to the point where capture is possible even with a gravity assist with Titan. The control-law algorithm, the benefits to the mass budget and the main technological aspects are presented and discussed. The simple steering law is compared with a trajectory optimizer to evaluate the quality of the results and possibilities for improvement. I. Introduction he giant planets have a special place in our quest for learning about the origins of our planetary system and Tour search for life, and robotic missions are essential tools for this scientific goal. Missions to the outer planets arXiv:2001.04357v1 [astro-ph.EP] 8 Jan 2020 have been prioritized by NASA and ESA, and this has resulted in important space projects for the exploration of the Jupiter system (NASA’s Europa Clipper [1] and ESA’s Jupiter Icy Moons Explorer [2]), and studies are underway to launch a follow-up of Cassini/Huygens called Titan Saturn System Mission (TSSM) [3], a joint ESA-NASA project. -
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 -
The J–2X Engine Powering NASA’S Ares I Upper Stage and Ares V Earth Departure Stage
The J–2X Engine Powering NASA’s Ares I Upper Stage and Ares V Earth Departure Stage The U.S. launch vehicles that will carry nozzle. It will weigh approximately 5,300 explorers back to the moon will be powered pounds. With 294,000 pounds of thrust, the in part by a J–2X engine that draws its engine will enable the Ares I upper stage to heritage from the Apollo-Saturn Program. place the Orion crew exploration vehicle in low-Earth orbit. The new engine, being designed and devel oped in support of NASA’s Constellation The J–2X is being designed by Pratt & Program, will power the upper stages of Whitney Rocketdyne of Canoga Park, Calif., both the Ares I crew launch vehicle and Ares for the Exploration Launch Projects Office at V cargo launch vehicle. NASA’s Marshall Space Flight Center in Huntsville, Ala. The J–2X builds on the The Constellation Program is responsible for legacy of the Apollo-Saturn Program and developing a new family of U.S. crew and relies on nearly a half-century of NASA launch vehicles and related systems and spaceflight experience, heritage hardware technologies for exploration of the moon, and technological advances. Mars and destinations beyond. Fueled with liquid oxygen and liquid hydro The J–2X will measure 185 inches long and gen, the J–2X is an evolved variation of two 120 inches in diameter at the end of its historic predecessors: the powerful J–2 upper stage engine that propelled the Apollo-era Shortly after J–2X engine cutoff, the Orion capsule will Saturn IB and Saturn V rockets to the moon in the separate from the upper stage. -
E-Sail for Fast Interplanetary Travel. M
Planetary Science Vision 2050 Workshop 2017 (LPI Contrib. No. 1989) 8056.pdf E-SAIL FOR FAST INTERPLANETARY TRAVEL. M. Aru1, P. Janhunen2, 1University of Tartu, Estonia, 2Finnish Meteorological Institute, Finland Introduction: Propulsion is a significant factor for areas of scientific research and new types of missions our access to the Solar System and the time con- could be imagined and created, improving our unders- sumption of the missions. We propose to use the elect- tanding of the Solar System. ric solar wind sail (E-sail), which can provide remar- Manned presence on Mars. A spacecraft equipped kable low thrust propulsion without needing propellant with a large E-sail, that provides 1 N of thrust at 1 au [1, 2]. from Sun, can travel from Earth to the asteroid belt in a The E-sail is a propellantless propulsion concept year. One such spacecraft can bring back three tonnes that uses centrifugally stretched, charged tethers to of water in three years, and repeat the journey multiple extract momentum from the solar wind to produce times within its estimated lifetime of at least ten years thrust. Over periods of months, this small but conti- [9, 12]. The water can be converted to synthetic cryo- nuous thrust can accelerate the spacecraft to great genic rocket fuel in orbital fuelling stations where speeds of approximately 20 to 30 au/year. For examp- manned vehicles travelling between Earth and Mars le, distances of 100 au could be reached in <10 years, can be fuelled. This dramatically reduces the overall which is groundbreaking [1]. mission fuel ratio at launch, and opens up possibilities The principles of operation: A full-scale E-sail for affordable continuous manned presence on Mars includes up to 100 thin, many kilometers long tethers, [13]. -
IAF Space Propulsion Symposium 2019
IAF Space Propulsion Symposium 2019 Held at the 70th International Astronautical Congress (IAC 2019) Washington, DC, USA 21 -25 October 2019 Volume 1 of 2 ISBN: 978-1-7138-1491-7 Printed from e-media with permission by: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 Some format issues inherent in the e-media version may also appear in this print version. Copyright© (2019) by International Astronautical Federation All rights reserved. Printed with permission by Curran Associates, Inc. (2020) For permission requests, please contact International Astronautical Federation at the address below. International Astronautical Federation 100 Avenue de Suffren 75015 Paris France Phone: +33 1 45 67 42 60 Fax: +33 1 42 73 21 20 www.iafastro.org Additional copies of this publication are available from: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 USA Phone: 845-758-0400 Fax: 845-758-2633 Email: [email protected] Web: www.proceedings.com TABLE OF CONTENTS VOLUME 1 PROPULSION SYSTEM (1) BLUE WHALE 1: A NEW DESIGN APPROACH FOR TURBOPUMPS AND FEED SYSTEM ELEMENTS ON SOUTH KOREAN MICRO LAUNCHERS ............................................................................ 1 Dongyoon Shin KEYNOTE: PROMETHEUS: PRECURSOR OF LOW-COST ROCKET ENGINE ......................................... 2 Jérôme Breteau ASSESSMENT OF MON-25/MMH PROPELLANT SYSTEM FOR DEEP-SPACE ENGINES ...................... 3 Huu Trinh 60 YEARS DLR LAMPOLDSHAUSEN – THE EUROPEAN RESEARCH AND TEST SITE FOR CHEMICAL SPACE PROPULSION SYSTEMS ....................................................................................... 9 Anja Frank, Marius Wilhelm, Stefan Schlechtriem FIRING TESTS OF LE-9 DEVELOPMENT ENGINE FOR H3 LAUNCH VEHICLE ................................... 24 Takenori Maeda, Takashi Tamura, Tadaoki Onga, Teiu Kobayashi, Koichi Okita DEVELOPMENT STATUS OF BOOSTER STAGE LIQUID ROCKET ENGINE OF KSLV-II PROGRAM ....................................................................................................................................................... -
The Advanced Cryogenic Evolved Stage (ACES)- a Low-Cost, Low-Risk Approach to Space Exploration Launch
The Advanced Cryogenic Evolved Stage (ACES)- A Low-Cost, Low-Risk Approach to Space Exploration Launch J. F. LeBar1 and E. C. Cady2 Boeing Phantom Works, Huntington Beach CA 92647 Space exploration top-level objectives have been defined with the United States first returning to the moon as a precursor to missions to Mars and beyond. System architecture studies are being conducted to develop the overall approach and define requirements for the various system elements, both Earth-to-orbit and in-space. One way of minimizing cost and risk is through the use of proven systems and/or multiple-use elements. Use of a Delta IV second stage derivative as a long duration in-space transportation stage offers cost, reliability, and performance advantages over earth-storable propellants and/or all new stages. The Delta IV second stage mission currently is measured in hours, and the various vehicle and propellant systems have been designed for these durations. In order for the ACES to have sufficient life to be useful as an Earth Departure Stage (EDS), many systems must be modified for long duration missions. One of the highest risk subsystems is the propellant storage Thermal Control System (TCS). The ACES effort concentrated on a lower risk passive TCS, the RL10 engine, and the other subsystems. An active TCS incorporating a cryocoolers was also studied. In addition, a number of computational models were developed to aid in the subsystem studies. The high performance TCS developed under ACES was simulated within the Delta IV thermal model and long-duration mission stage performance assessed. -
Determination of Optimal Earth-Mars Trajectories to Target the Moons Of
The Pennsylvania State University The Graduate School Department of Aerospace Engineering DETERMINATION OF OPTIMAL EARTH-MARS TRAJECTORIES TO TARGET THE MOONS OF MARS A Thesis in Aerospace Engineering by Davide Conte 2014 Davide Conte Submitted in Partial Fulfillment of the Requirements for the Degree of Master of Science May 2014 ii The thesis of Davide Conte was reviewed and approved* by the following: David B. Spencer Professor of Aerospace Engineering Thesis Advisor Robert G. Melton Professor of Aerospace Engineering Director of Undergraduate Studies George A. Lesieutre Professor of Aerospace Engineering Head of the Department of Aerospace Engineering *Signatures are on file in the Graduate School iii ABSTRACT The focus of this thesis is to analyze interplanetary transfer maneuvers from Earth to Mars in order to target the Martian moons, Phobos and Deimos. Such analysis is done by solving Lambert’s Problem and investigating the necessary targeting upon Mars arrival. Additionally, the orbital parameters of the arrival trajectory as well as the relative required ΔVs and times of flights were determined in order to define the optimal departure and arrival windows for a given range of date. The first step in solving Lambert’s Problem consists in finding the positions and velocities of the departure (Earth) and arrival (Mars) planets for a given range of dates. Then, by solving Lambert’s problem for various combinations of departure and arrival dates, porkchop plots can be created and examined. Some of the key parameters that are plotted on porkchop plots and used to investigate possible transfer orbits are the departure characteristic energy, C3, and the arrival v∞. -
The New Vision for Space Exploration
Constellation The New Vision for Space Exploration Dale Thomas NASA Constellation Program October 2008 The Constellation Program was born from the Constellation’sNASA Authorization Beginnings Act of 2005 which stated…. The Administrator shall establish a program to develop a sustained human presence on the moon, including a robust precursor program to promote exploration, science, commerce and U.S. preeminence in space, and as a stepping stone to future exploration of Mars and other destinations. CONSTELLATION PROJECTS Initial Capability Lunar Capability Orion Altair Ares I Ares V Mission Operations EVA Ground Operations Lunar Surface EVA EXPLORATION ROADMAP 0506 07 08 09 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 LunarLunar OutpostOutpost BuildupBuildup ExplorationExploration andand ScienceScience LunarLunar RoboticsRobotics MissionsMissions CommercialCommercial OrbitalOrbital Transportation ServicesServices forfor ISSISS AresAres II andand OrionOrion DevelopmentDevelopment AltairAltair Lunar LanderLander Development AresAres VV and EarthEarth DepartureDeparture Stage SurfaceSurface SystemsSystems DevelopmentDevelopment ORION: NEXT GENERATION PILOTED SPACECRAFT Human access to Low Earth Orbit … … to the Moon and Mars ORION PROJECT: CREW EXPLORATION VEHICLE Orion will support both space station and moon missions Launch Abort System Orion will support both space stationDesigned and moonto operate missions for up to 210 days in Earth or lunar Designedorbit to operate for up to 210 days in Earth or lunar orbit Designed for lunar -
Astrodynamics (AERO0024)
Astrodynamics (AERO0024) The twobody problem Lamberto Dell'Elce Space Structures & Systems Lab (S3L)0 Outline The twobody problem µ rÜ = – r Equations of motion r 3 Resulting orbits 1 Outline The twobody problem µ rÜ = – r Equations of motion r 3 Resulting orbits 2 What is the twobody problem (or Kepler problem)? F 21 F m1 12 m2 Motion of two point masses due to their gravitational interaction 3 Twobody problem vs real world 4 What is the interest in the twobody problem? 5 Gravitational force of a point mass F 21 F m1 12 m2 r Norm: m1 m2 kF 12k = kF 21k = G r 2 Direction: • Along the line joining m1 and m2 • Directed toward the attractor 6 Gravitational constant 7 Gravitational parameter of a Celestial body 8 Satellite laser ranging 9 Satellites as bodies in free fall 10 Is pointmass a good approximation for Earth gravity? 11 Gravitational potential of a uniform sphere 12 Gravitational potential of a sphericallysymmetric body 13 Outline The twobody problem µ rÜ = – r Equations of motion r 3 Resulting orbits 14 Dynamics of the two bodies 15 Motion of the center of mass 16 Equations of relative motion Assume m2 m1 17 Equations of relative motion 18 Integrals of motion: The angular momentum 19 Implication: Motion lies in a plane 20 Azimuth component of the velocity 21 Integrals of motion: The eccentricity vector 22 Relative trajectory 23 In summary 24 Outline The twobody problem µ rÜ = – r Equations of motion r 3 Resulting orbits 25 Conic sections in polar coordinates 26 Conic sections 27 Possible trajectories of the twobody -
Solar Electric Propulsion Sail
IOSR Journal of Electronics and Communication Engineering (IOSR-JECE) e-ISSN: 2278-2834,p- ISSN: 2278-8735.Volume 13, Issue 5, Ver. I (Sep.-Oct. 2018), PP 18-22 www.iosrjournals.org Solar Electric Propulsion Sail Dr. S.S. Subashka Ramesh [1] Sri Haripriya Nutulapti [2], Yashraj Sharma [3], Pratyush Kumar [4] [1], [2], [3], [4] Department of Computer Science and Engineering, SRM institute of science & technology Ramapuram Campus, Chennai-89, India. Corresponding Author: Dr. S.S. Subashka Ramesh Abstract: The solar expedition missions have been minimized mainly due to the performance of the space capsules, also because of the planned amount of fuel a space shuttle must carry without discharging in order to advance to an unfamiliar region. Traditionally, a solar electric propulsion sail is extremely deformable to drive a space rocket through outer space. Photon or electric sails are propounding medium of propulsion by utilizing the cosmic radiation endeavored because of sun’s rays on massive mirrors. It is a fusion of photo-voltaic cells and ions for the propelling and is also capable of enabling very fine maneuvering of the spacecraft by means of large sail-surface deformations. Solar electric propulsion sailutilizes the natural beams of sunlight to advancethe vehicles into and out of space, just the way wind helps to propel the sailboats over the water. NASA team claimed working on the start ofgrowth of technology on the assignment recognized as the solar sail demonstrator which proved thatmaking use of giant, weightless and unfurling objects float in universe would enhance the abilities of travelling deeper in space. -
Electric Sail Technology Demonstration Mission Spacecraft
https://ntrs.nasa.gov/search.jsp?R=20170001821 2019-06-22T08:27:43+00:00Z The Conceptual Design of an Electric Sail Technology Demonstration Mission Spacecraft Presentation at: 40th Annual AAS Guidance and Control Conference Breckenridge, CO, USA February 3-8, 2017 Bruce M. Wiegmann NASA-MSFC-ED04 [email protected] Presentation Agenda • HERTS/Electric Sail background information • Findings from the Phase I NIAC • This propulsion technology enables trip times to the Heliopause in 10 – 12 years • Fastest transportation method to reach Heliopause of near term propulsion technologies • Current Phase II NIAC tasks • Plasma chamber testing • Particle-in-cell (PIC) space plasma to spacecraft modeling • Tether material investigation • Conceptual design of a TDM spacecraft • Mission capture Image shown is copyright by: Alexandre Szames, Antigravite, Paris, and is used with permission National Aeronautics and Space Administration 2 Solar Wind Basics-> Solar Sail • The relative velocity of the Solar Wind through the decades The solar wind ions traveling at 400-500 km/sec are the naturally occurring (free) energy source that propels an E-Sail National Aeronautics and Space Administration 3 Electric Sail Origins The electric solar wind sail, or electric sail for short, is a propulsion invention made in 2006 at the Kumpula Space Centre by Dr. Pekka Janhunen. Image courtesy of: Dr. Pekka Janhunen Phase I Findings • Electric-Sail propulsion systems are the fastest method to get spacecraft to deep space destinations as compared to: • Solar sails, • All chemical propulsions, • Electric (ion) propulsion systems • Technology appears to be viable . • Technology Assessment – Most subsystems at high state of readiness except: • Wire-plasma interaction modeling, • Wire deployment, and • Dynamic control of E–Sail spacecraft… • These are the three areas of focus for the current Phase II NIAC National Aeronautics and Space Administration 5 Electric Sail – Concept of Operations • The E-sail consists of 10 to 20 conducting, positively charged, bare wires, each 1–20 km in length. -
Progress on the J-2X Upper Stage Engine for the Ares Launch Vehicles
https://ntrs.nasa.gov/search.jsp?R=20080036837mSFC-9;LO 2019-08-30T05:16:33+00:00Z From Concept to Design: Progress on the J-2X Upper Stage Engine for the Ares Launch Vehicles Thomas Byrd, Deputy Manager, J-2X Upper Stage Engine Element Ares Projects Office Marshall Space Flight Center Huntsville, AL 35812 Abstract In accordance with national policy and NASA's Global Exploration Strategy, the Ares Projects Office is embarking on development ofa new launch vehicle fleet to fulfill the national goals ofreplacing the space shuttle fleet, returning to the moon, and exploring farther destinations like Mars. These goals are shaped by the decision to retire the shuttle fleet by 2010, budgetary constraints, and the requirement to create a new fleet that is safer, more reliable, operationally more efficient than the shuttle fleet, and capable ofsupporting long-range exploration goals. The present architecture for the Constellation Program is the result ofextensive trades during the Exploration Systems Architecture Study and subsequent refinement by the Ares Projects Office at Marshall Space Flight Center. The vehicles selected to support those goals are the Ares I crew launch vehicle and the Ares V cargo launch vehicle, the first new human-rated launch vehicles developed by NASA in more than three decades (Figure 1). Ares I will begin its initial operating capability offlying up to six astronauts to the International Space Station (ISS) no later than 2015. Ares V is scheduled to be operational in the 2020 timeframe to support lunar missions. Figure 1. The Ares V Cargo Launch Vehicle (left) and Ares I Crew Launch Vehicle (right) will form the backbone of America's new space fleet.