VASIMR VX-200 Performance and Near-Term SEP Capability for Unmanned Mars Flight

Total Page:16

File Type:pdf, Size:1020Kb

VASIMR VX-200 Performance and Near-Term SEP Capability for Unmanned Mars Flight VASIMR VX-200 Performance and Near-term SEP Capability for Unmanned Mars Flight Future In-Space Operations Seminar January 19, 2011 presented by Tim Glover Director of Development [email protected] Ad Astra Rocket Company www.adastrarocket.com 1 Ad Astra Rocket Company Notes and Acronyms Notes: - solar array power values are for 1 AU - a number of publications on VASIMR R&D are available on the company’s website: http://www.adastrarocket.com Acronyms: SEP solar electric propulsion NTR nuclear thermal rocket VASIMR Variable Specific Impulse Magnetoplasma Rocket TMI Trans-Mars Insertion MOI Mars Orbit Insertion SOI sphere of influence IMLEO initial mass in low Earth orbit 2 Ad Astra Rocket Company Outline 1. VASIMR Prototype Performance 2. Simplified Earth‐Mars Trajectories 3. Chemical and NTR Hohmann Transfers 4. SEP: Initial Mass‐to‐Power Ratio and Payload Fraction 5. Near‐term SEP Mars Capabilities 6. Backup slides: • Propellant and transit time variation for actual orbits of Earth and Mars •Atlas V 551 performance curve 3 Ad Astra Rocket Company My Background • B.S., Physics, New Mexico State U. • M.S., Aerospace Engineering (Orbital Mechanics), UT-Austin • M.S., Physics, University of Pittsburgh • five years teaching high school physics, Ethical Culture Schools, New York • Ph.D., Applied Physics, Rice University, 2002 − thesis research at Johnson Space Center: designed and built plasma diagnostics to measure exhaust velocity in early VASIMR prototypes • Research Scientist, MEI Technologies 2003-2005 − continued experimental work on VASIMR up to 50 kW • Director of Development, Ad Astra Rocket Company 2005 – present − business development, external relations 4 Ad Astra Rocket Company VASIMR Operating Principles Superconducting Magnets typical path of an ion Helicon coupler (30 kW) through the rocket Ion cyclotron (170 kW) coupler i gas cold accelerated plasma plasma POWER 1. Ionize 2. Energize 3. Accelerate 4. Detach first stage second stage Caption: Argon gas is fed into the first stage of the VASIMR® engine, which converts the argon from a neutral gas to a fully ionized plasma. At this stage, the plasma is described as “cold” . Although the electrons are hot (10,000 degrees celsius or more), they constitute less than 0.1% of the mass of the plasma; the ions are roughly at room temperature (relatively cold) and make up more than 99.9% of the mass. Once the plasma leaves the first stage, it flows along the magnetic field lines produced by the superconducting magnets and makes no further contact with solid parts of the engine. As the plasma flows through the second stage, the ions are accelerated to much higher energies than the electrons by waves created in the plasma by the ion cyclotron coupler. Due to the strong magnetic field in the second stage, the ions are confined to move in tight helical orbits whose diameter is much smaller than the diameter of the plasma column. As the plasma flows downstream, the nozzle action of the expanding magnetic field accelerates the plasma to a high exhaust velocity. Like the plasma leaving the Sun in a Coronal Mass Ejection, the VASIMR exhaust plasma separates from the magnetic field of the rocket as it accelerates through the magnetic nozzle to high speed. 5 Ad Astra Rocket Company VX-200 Testing 6 Ad Astra Rocket Company VX-200 in Ad Astra Test Facility top view VX-200 A motion table can move the diagnostics through the exhaust plume over an axial range of 500 cm and a diameter range of 200 cm. Exhaust plume diagnostics are located side view through port downstream from the VX-200 exit plane. 7 Ad Astra Rocket Company Exhaust Plasma Diagnostics and Translation Stage Table RPA Force Impact Target #1 Force Impact Target #2 3-axis Magnetometer Langmuir Probe Ion Flux Probe Array #1 Ion Flux Probe Array #2 8 Ad Astra Rocket Company Thrust Target Measurements • We take diameter scans (31 shots, shot to shot in this case) looking at force density directly • Integrate the force density to get a total force number • This method has been found to yield identical results to a thrust stand, by taking simultaneous measurements on a Hall thruster at University of Michigan 9 cm Corresponding size of target 9 1/18/2011 Ad Astra Rocket Company VX-200 Performance in Nov. 2010 Experiment Campaign • In the Thrust graph above, first stage power • The Efficiency graph is derived from the previous Thrust is held fixed at 30 kW for optimal ionization of graph; it does NOT represent VASIMR efficiency as a argon propellant at a flow rate of 130 mg/s, function of exhaust velocity except at the nominal operating and second stage power is raised from zero to point for this propellant flow rate, which is 50 km/s. 170 kW. • a single 400 V power supply powers both stages of the rocket; thruster efficiency is calculated as the jet power in The VX-200 prototype data gives us the plume (determined by the thrust target), divided by the measured power output of the RF amplifiers. confidence that we can build a single 200 kW thruster that achieves 70% thruster • measured DC to RF power conversion efficiency is 95%, efficiency at an Isp of 5000 seconds. leading to a system efficiency of 60% (DC to jet power) for 50 km/s (5000 s Isp) operation. 10 Ad Astra Rocket Company Performance of Chemical, NTR and SEP Systems • to evaluate the performance of SEP relative to chemical and NTR propulsion for Mars cargo/unmanned flights, we consider a simple hypothetical space flight • circular, coplanar orbits for Earth and Mars are used to eliminate variations due to the eccentricity and relative inclination of the planets’ orbits • while not accurate for actual flight studies, this comparison illustrates the relative merits of the three types of propulsion, and estimates the technological requirements needed for SEP to play a useful role 11 Ad Astra Rocket Company A Simplified Problem Relevant to Mars Flights The blue and red circles in the figure at right represent the orbits of Earth and Mars, respectively. Although Mars’ actual orbit is an ellipse, it is represented here by a circle with the same average distance from the Sun as Mars, to remove the effect of launch date on the trajectories. This circular approximation allows us to compare propulsion systems and trajectories for a problem that is independent of launch date and phasing of the planets. This simple situation will provide insights that will be helpful in selecting appropriate SEP vehicle parameters for cargo flights using the actual orbits of the Earth and Mars. To estimate the performance of two propulsion systems considered for Mars flights – chemical and nuclear thermal rockets - we calculate the maximum mass that can be transferred from the Earth’s orbit to an orbit about Mars with two impulsive burns. At Mars, we assume that the cargo vehicle enters a 1-sol Mars orbit. To minimize the delta-v required, this Mars Orbit Insertion (MOI) burn is done at an altitude of only 250 km above the surface of Mars, and results in a highly elliptical orbit with a semi-major axis of 20,000 km. (This is the assumption made in NASA’s Design Reference Architecture for the crewed vehicles envisioned to support human exploration of Mars – see backup slides). A Hohmann maneuver. The transit between the circular approximations to the orbits of Earth and Mars takes 260 For comparison, we calculate the performance of solar days. In the “all-propulsive” case, two burns are necessary. electric propulsion for fixed specific impulse, over a range of If aerocapture at Mars is presumed feasible, only a single power levels. These trajectories have been calculated using burn in LEO is needed. the Copernicus software package, for artificial planets in circular, co-planar orbits having radii equal to the semi-major axes of the orbits of Earth and Mars. 12 Ad Astra Rocket Company Hohmann Transfer Assumptions: • 400 km altitude initial low Earth orbit • 100 mT vehicle in LEO (wet mass) • chemical Isp = 450 s (LOX/LH2) • NTR Isp = 925 s (based on Peewee reactor, final NERVA variant) Transfer orbit: • semi-major axis = 1.89 x 1011 m • transfer time (1/2 period) = 260 days • perihelion velocity = 32,670 m/s • aphelion velocity = 21,490 m/s Two maneuvers: • Trans-Mars Insertion (TMI) burn modeled as a single impulse in the 400 km LEO parking orbit (v1) • Mars Orbit Insertion (MOI) burn modeled as a single impulse performed at the 250 km altitude periapsis of the 1-sol (0.82 eccentricity) Mars parking orbit (v2) 13 Ad Astra Rocket Company TMI Burn for Chemical and NTR • 400 km altitude circular LEO velocity (pre- TMI burn) Excess velocity needed for = 7,676 m/s Hohmann transfer (perihelion) • transfer orbit’s perihelion velocity = 32,670 m/s • Earth’s orbital velocity = 29,750 m/s Earth’s sphere of influence • excess velocity needed from TMI burn: 32,670 – 29,750 = 2,920 m/s at Earth SOI • this corresponds to a post- TMI burn velocity in the 400 km altitude parking orbit 400 km parking orbit of: 2 2 v vLEO GM E E 2 2 rLEO 2GM v v2 E 11,241 m/s v1 LEO rLEO • v1 = 11,240 – 7,676 = 3,565 m/s 14 Ad Astra Rocket Company MOI Burn for Chemical and NTR • transfer orbit’s aphelion velocity = 21,490 m/s • Mars orbital velocity = 24,130 m/s • Mars approach velocity: 24,130 – 21,490 = 2,640 m/s at Mars SOI v2 • this corresponds to a pre-MOI burn velocity at the 250 km periapsis of the 1-sol km parking orbit of: v2 v2 GM E 250 M 2 2 r250 2 2GM M v250 v = 5,513 m/s r250 • 250 km periapsis velocity (post- TMI burn): 1‐sol orbit 1sol GM M ra vperi = 4,619 m/s a rp Mars approach velocity • v2 = 5,513 – 4,619 = 894 m/s 15 Ad Astra Rocket Company Validity of Circular Approximation TMI 3,565 m/s MOI 894 m/s from NASA DRA 2007, B.
Recommended publications
  • 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]
  • 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]
  • Utilizing Deep Reinforcement Learning to Effect Autonomous Orbit Transfers and Intercepts Via Electromagnetic Propulsion
    Problem Utilizing Deep Reinforcement Learning to Effect Autonomous Results The growth in space-capable entities has caused a rapid rise in Orbit Transfers and Intercepts via Electromagnetic Propulsion ➢ Analysis the number of derelict satellites and space debris in orbit Gabriel Sutherland ([email protected]), Oregon State University, Corvallis, OR, USA ❖ The data suggests spacecraft highly capable of around Earth, which pose a significant navigation hazard. Frank Soboczenski ([email protected]), King’s College London, London, UK neutralizing/capturing small to intermediate debris Athanasios Vlontzos ([email protected]), Imperial College London, London, UK ❖ In simulations, spacecraft was able to capture and/or Objectives neutralize debris ranging in mass from 1g to 100g ❖ Satellites suffered damage in some Develop a system that is capable of autonomously neutralizing simulations multiple pieces of space debris in various orbits. Software & Simulations ❖ Simulations show that damaged segments of satellite were vaporized, which means Background that no additional debris was added to orbit ➢ Dangerous amounts of space debris in orbit, estimates vary ❖ Neural Network based on Deep Deterministic Policy ➢20,000+ tracked objects larger Gradient (DDPG) effective at low-thrust orbit transfers in than 10 cm diameter 2D Hohmann transfer problem ➢Est 500,000 objects larger ❖ Using ASTOS simulation software, mission-representative than 1 cm in diameter model of spacecraft and experimental propulsion system ➢Est 100 million objects ❖ Interfaced with DRL to train the Neural smaller than 1cm in diameter ➢ Orbital speeds of these objects Network with realistic data vary from 100+ kph to 28,100 kph ➢ Future Studies ➢ Spacecraft collisions due to space ❖ Train DDPG for 3 dimensional orbit transfer problems debris have been sparse so far Tracked spacecraft in Earth orbit.
    [Show full text]
  • Magnetoshell Aerocapture: Advances Toward Concept Feasibility
    Magnetoshell Aerocapture: Advances Toward Concept Feasibility Charles L. Kelly A thesis submitted in partial fulfillment of the requirements for the degree of Master of Science in Aeronautics & Astronautics University of Washington 2018 Committee: Uri Shumlak, Chair Justin Little Program Authorized to Offer Degree: Aeronautics & Astronautics c Copyright 2018 Charles L. Kelly University of Washington Abstract Magnetoshell Aerocapture: Advances Toward Concept Feasibility Charles L. Kelly Chair of the Supervisory Committee: Professor Uri Shumlak Aeronautics & Astronautics Magnetoshell Aerocapture (MAC) is a novel technology that proposes to use drag on a dipole plasma in planetary atmospheres as an orbit insertion technique. It aims to augment the benefits of traditional aerocapture by trapping particles over a much larger area than physical structures can reach. This enables aerocapture at higher altitudes, greatly reducing the heat load and dynamic pressure on spacecraft surfaces. The technology is in its early stages of development, and has yet to demonstrate feasibility in an orbit-representative envi- ronment. The lack of a proof-of-concept stems mainly from the unavailability of large-scale, high-velocity test facilities that can accurately simulate the aerocapture environment. In this thesis, several avenues are identified that can bring MAC closer to a successful demonstration of concept feasibility. A custom orbit code that dynamically couples magnetoshell physics with trajectory prop- agation is developed and benchmarked. The code is used to simulate MAC maneuvers for a 60 ton payload at Mars and a 1 ton payload at Neptune, both proposed NASA mis- sions that are not possible with modern flight-ready technology. In both simulations, MAC successfully completes the maneuver and is shown to produce low dynamic pressures and continuously-variable drag characteristics.
    [Show full text]
  • Space Sector Brochure
    SPACE SPACE REVOLUTIONIZING THE WAY TO SPACE SPACECRAFT TECHNOLOGIES PROPULSION Moog provides components and subsystems for cold gas, chemical, and electric Moog is a proven leader in components, subsystems, and systems propulsion and designs, develops, and manufactures complete chemical propulsion for spacecraft of all sizes, from smallsats to GEO spacecraft. systems, including tanks, to accelerate the spacecraft for orbit-insertion, station Moog has been successfully providing spacecraft controls, in- keeping, or attitude control. Moog makes thrusters from <1N to 500N to support the space propulsion, and major subsystems for science, military, propulsion requirements for small to large spacecraft. and commercial operations for more than 60 years. AVIONICS Moog is a proven provider of high performance and reliable space-rated avionics hardware and software for command and data handling, power distribution, payload processing, memory, GPS receivers, motor controllers, and onboard computing. POWER SYSTEMS Moog leverages its proven spacecraft avionics and high-power control systems to supply hardware for telemetry, as well as solar array and battery power management and switching. Applications include bus line power to valves, motors, torque rods, and other end effectors. Moog has developed products for Power Management and Distribution (PMAD) Systems, such as high power DC converters, switching, and power stabilization. MECHANISMS Moog has produced spacecraft motion control products for more than 50 years, dating back to the historic Apollo and Pioneer programs. Today, we offer rotary, linear, and specialized mechanisms for spacecraft motion control needs. Moog is a world-class manufacturer of solar array drives, propulsion positioning gimbals, electric propulsion gimbals, antenna positioner mechanisms, docking and release mechanisms, and specialty payload positioners.
    [Show full text]
  • Astrox Studies and Experience with the Reusable Booster Systems and Two-Stage-To-Orbit Concepts
    Astrox Studies and Experience with the Reusable Booster Systems and Two-Stage-to-Orbit Concepts Presented to: National Research Council Aeronautics and Space Engineering Board February 2012 By: Astrox Corporation Dr. Christopher Tarpley Colorado Springs, CO 80917 Dr. Ajay P. Kothari College Park, MD 20740 Distribution A: Approved for public release; distribution unlimited Acknowledgements • Dr. Mark Lewis, Ex-Chief Scientist, Air Force • Dr. Werner Dahm, Ex-Chief Scientist, Air Force • Dr. Donald Paul, Chief Scientist – Rtrd, AFRL/RB • Mr. Bruce Thieman, AFRL/RB • Mr. Barry Hellman, AFRL/RB • Mr. John Livingston, ASC/XR • Mr. Glenn Liston, AFRL/RZ • Mr. Dan Risha, AFRL/RZ • Dr. Kevin Bowcutt, Boeing Huntington Beach • Dr. Ray Moszee, SAF/AQR, Pentagon 2 Distribution A: Approved for public release; distribution unlimited Background • Astrox experience with Access-to-Space (ATS) and high Mach cruise configurations covers almost two decades of work primarily with Air Force and NASA • Astrox has been developing tools for vehicle design and quantitative analysis since 1990 • Studies have covered: – SSTO and TSTO Systems – RP, JP, Methane and LH2 Systems – Payloads from 2,000 to 60,000 lbs – Rocket, Turbine, Ram/Scramjet Engines – Air Launch, Horizontal and Vertical Takeoff Configurations 3 Distribution A: Approved for public release; distribution unlimited Relevant Studies Performed Inward Turning Inlet 1990-1992 ASC/XR Inward Turning Flowpath and Vehicles 1993-2000 NASA/LaRC HADO and HySIDE Code 1995-1997 ASC/XR Inward Turning SSTO Designs 1997-1999 NASA/MSFC Access-to-Space / FAST* 1 2004 - 2006 AFRL/VA TSTO Architectures 2005 AFRL/VA Aerial Refueling 2006 AFRL/PRS Prompt Global Strike 2006 AFRL/PRS Hybrid Launch Study 2007 AFRL/PRS TSTO Study 2007-2008 AFRL/PRS FAST* 2 2008 AFRL/RB Joint System Study 2009 AFRL/RB *FAST – Fully Reusable Access-to-Space Technology 4 Distribution A: Approved for public release; distribution unlimited Recent Relevant Publications 1.
    [Show full text]
  • Aerocapture As an Enhancing Option for Ice Giants Missions
    WHITE PAPER FOR THE PLANETARY SCIENCE DECADAL SURVEY, 2023 - 2032 Aerocapture as an Enhancing Option for Ice Giants Missions Primary Author: Soumyo Dutta NASA Langley Research Center Phone: 757-864-3894 E-Mail: [email protected] Co-Authors: Gonçalo Afonso1 Jay Feldman5 Zachary R. Putnam11 Samuel W. Albert2 Roberto Gardi12 Jeremy R. Rea15 Hisham K. Ali3 Athul P. Girija13 Sachin Alexander Reddy21 Gary A. Allen4 Tiago Hormigo1 Thomas Reimer22 Antonella I. Alunni5 Jeffrey P. Hill5 Sarag J. Saikia23 James O. Arnold4 Shayna Hume2 Isil Sakraker Özmen22 Alexander Austin6 Christopher Jelloian14 Kunio Sayanagi24 Gilles Bailet7 Vandana Jha5 Stephan Schuster25 Shyam Bhaskaran6 Breanna J. Johnson15 Jennifer Scully6 Alan M. Cassell5 Craig A. Kluever16 Ronald R. Sostaric15 George T. Chen6 Jean-Pierre Lebreton17 Christophe Sotin6 Ian J. Cohen8 Marcus A. Lobbia6 David A. Spencer6 James A. Cutts6 Ping Lu18 Benjamin M. Tackett4 Rohan G. Deshmukh4 Ye Lu19 Nikolas Trawny6 Robert A. Dillman9 Rafael A. Lugo9 Ethiraj Venkatapathy5 Guillermo Dominguez Daniel A. Matz15 Paul F. Wercinski5 Calabuig10 Robert W. Moses9 Michael C. Wilder5 Sarah N. D’Souza5 Michelle M. Munk9 Michael J. Wright5 Donald T. Ellerby5 Adam P. Nelessen6 Cindy L. Young9 Giusy Falcone11 Miguel Pérez-Ayúcar20 Alberto Fedele12 Richard W. Powell4 1 Spin.Works S.A. 4 Analytical Mechanics Associates 2 University of Colorado, Boulder 5 NASA Ames Research Center 3 Georgia Institute of Technology Aerocapture for Ice Giants Missions 6 Jet Propulsion Laboratory/California 16 University of Missouri Institute of Technology 17 French National Centre for Scientific 7 University of Glasgow Research 8 John Hopkins University/Applied Physics 18 San Diego State University Laboratory 19 Kent State University 9 NASA Langley Research Center 20 Aurora Technology B.V.
    [Show full text]
  • Study of a Crew Transfer Vehicle Using Aerocapture for Cycler Based Exploration of Mars by Larissa Balestrero Machado a Thesis S
    Study of a Crew Transfer Vehicle Using Aerocapture for Cycler Based Exploration of Mars by Larissa Balestrero Machado A thesis submitted to the College of Engineering and Science of Florida Institute of Technology in partial fulfillment of the requirements for the degree of Master of Science in Aerospace Engineering Melbourne, Florida May, 2019 © Copyright 2019 Larissa Balestrero Machado. All Rights Reserved The author grants permission to make single copies ____________________ We the undersigned committee hereby approve the attached thesis, “Study of a Crew Transfer Vehicle Using Aerocapture for Cycler Based Exploration of Mars,” by Larissa Balestrero Machado. _________________________________________________ Markus Wilde, PhD Assistant Professor Department of Aerospace, Physics and Space Sciences _________________________________________________ Andrew Aldrin, PhD Associate Professor School of Arts and Communication _________________________________________________ Brian Kaplinger, PhD Assistant Professor Department of Aerospace, Physics and Space Sciences _________________________________________________ Daniel Batcheldor Professor and Head Department of Aerospace, Physics and Space Sciences Abstract Title: Study of a Crew Transfer Vehicle Using Aerocapture for Cycler Based Exploration of Mars Author: Larissa Balestrero Machado Advisor: Markus Wilde, PhD This thesis presents the results of a conceptual design and aerocapture analysis for a Crew Transfer Vehicle (CTV) designed to carry humans between Earth or Mars and a spacecraft on an Earth-Mars cycler trajectory. The thesis outlines a parametric design model for the Crew Transfer Vehicle and presents concepts for the integration of aerocapture maneuvers within a sustainable cycler architecture. The parametric design study is focused on reducing propellant demand and thus the overall mass of the system and cost of the mission. This is accomplished by using a combination of propulsive and aerodynamic braking for insertion into a low Mars orbit and into a low Earth orbit.
    [Show full text]
  • + Part 17: Acronyms and Abbreviations (265 Kb PDF)
    17. Acronyms and Abbreviations °C . Degrees.Celsius °F. Degrees.Fahrenheit °R . Degrees.Rankine 24/7. 24.Hours/day,.7.days/week 2–D. Two-Dimensional 3C. Command,.Control,.and.Checkout 3–D. Three-Dimensional 3–DOF . Three-Degrees.of.Freedom 6-DOF. Six-Degrees.of.Freedom A&E. Architectural.and.Engineering ACEIT. Automated.Cost-Estimating.Integrated.Tools ACES . Acceptance.and.Checkout.Evaluation.System ACP. Analytical.Consistency.Plan ACRN. Assured.Crew.Return.Vehicle ACRV. Assured.Crew.Return.Vehicle AD. Analog.to.Digital ADBS. Advanced.Docking.Berthing.System ADRA. Atlantic.Downrange.Recovery.Area AEDC. Arnold.Engineering.Development.Center AEG . Apollo.Entry.Guidance AETB. Alumina.Enhanced.Thermal.Barrier AFB .. .. .. .. .. .. .. Air.Force.Base AFE. Aero-assist.Flight.Experiment AFPG. Apollo.Final.Phase.Guidance AFRSI. Advanced.Flexible.Reusable.Surface.Insulation AFV . Anti-Flood.Valve AIAA . American.Institute.of.Aeronautics.and.Astronautics AL. Aluminum ALARA . As.Low.As.Reasonably.Achievable 17. Acronyms and Abbreviations 731 AL-Li . Aluminum-Lithium ALS. Advanced.Launch.System ALTV. Approach.and.Landing.Test.Vehicle AMS. Alpha.Magnetic.Spectrometer AMSAA. Army.Material.System.Analysis.Activity AOA . Analysis.of.Alternatives AOD. Aircraft.Operations.Division APAS . Androgynous.Peripheral.Attachment.System APS. Auxiliary.Propulsion.System APU . Auxiliary.Power.Unit APU . Auxiliary.Propulsion.Unit AR&D. Automated.Rendezvous.and.Docking. ARC . Ames.Research.Center ARF . Assembly/Remanufacturing.Facility ASE. Airborne.Support.Equipment ASI . Augmented.Space.Igniter ASTWG . Advanced.Spaceport.Technology.Working.Group ASTP. Advanced.Space.Transportation.Program AT. Alternate.Turbopump ATCO. Ambient.Temperature.Catalytic.Oxidation ATCS . Active.Thermal.Control.System ATO . Abort-To-Orbit ATP. Authority.to.Proceed ATS. Access.to.Space ATV . Automated.Transfer.Vehicles ATV .
    [Show full text]
  • Mercury Orbit Insertion March 18, 2011 UTC (March 17, 2011 EDT)
    Mercury Orbit Insertion March 18, 2011 UTC (March 17, 2011 EDT) A NASA Discovery Mission Media Contacts NASA Headquarters Policy/Program Management Dwayne C. Brown (202) 358-1726 [email protected] The Johns Hopkins University Applied Physics Laboratory Mission Management, Spacecraft Operations Paulette W. Campbell (240) 228-6792 or (443) 778-6792 [email protected] Carnegie Institution of Washington Principal Investigator Institution Tina McDowell (202) 939-1120 [email protected] Mission Overview Key Spacecraft Characteristics MESSENGER is a scientific investigation . Redundant major systems provide critical backup. of the planet Mercury. Understanding . Passive thermal design utilizing ceramic-cloth Mercury, and the forces that have shaped sunshade requires no high-temperature electronics. it, is fundamental to understanding the . Fixed phased-array antennas replace a deployable terrestrial planets and their evolution. high-gain antenna. The MESSENGER (MErcury Surface, Space . Custom solar arrays produce power at safe operating ENvironment, GEochemistry, and Ranging) temperatures near Mercury. spacecraft will orbit Mercury following three flybys of that planet. The orbital phase will MESSENGER is designed to answer six use the flyby data as an initial guide to broad scientific questions: perform a focused scientific investigation of . Why is Mercury so dense? this enigmatic world. What is the geologic history of Mercury? MESSENGER will investigate key . What is the nature of Mercury’s magnetic field? scientific questions regarding Mercury’s . What is the structure of Mercury’s core? characteristics and environment during . What are the unusual materials at Mercury’s poles? these two complementary mission phases. What volatiles are important at Mercury? Data are provided by an optimized set of miniaturized space instruments and the MESSENGER provides: spacecraft tele commun ications system.
    [Show full text]
  • Launch and Deployment Analysis for a Small, MEO, Technology
    Launch and Deployment Analysis for a Isp = specific impulse, sec Small, MEO, Technology Demonstration k = spring constant, Nt/mm KLC = Kodiak Island Launch Complex Satellite MTO = medium transfer orbit transfer orbit 1 LEO = low earth orbit Tyson Karl Smith 2 LMA = Lockheed Martin Aerospace Stephen Anthony Whitmore M = final mass after insertion burn, kg Utah State University, Logan, UT, 84322-4130 final Mpayload = mass of payload after kick motor jettison, kg A trade study investigation the possibilities of delivering M = propellant mass consumed during a small technology-demonstration satellite to a medium prop earth orbit are presented. The satellite is to be deployed insertion burn, kg in a 19,000 km orbit with an inclination of 55°. This Mstage = mass of expended stage after jettison, payload is a technology demonstrator and thus launch kg and deployment costs are a paramount consideration. NRE = non-recurrent engineering, kg Also the payload includes classified technology, thus a Nspring = number of springs in Lightband® USA licensed launch system is mandated. All FAA- separation system licensed US launch systems are considered during a MTO = MEO transfer orbit preliminary trade analysis. This preliminary trade OSC = Orbital Sciences Corporation analysis selects Orbital Sciences Minotaur V launch R = apogee radius, km vehicle. To meet mission objective the Minotaur V 5th a stage ATK-Star 37FM motor is replaced with the Rp = perigee radius, km smaller ATK- Star 27. This new configuration allows R⊕ = local earth radius, km for payload delivery without adding an additional 6th SDL = Space Dynamics Laboratory stage kick motor. End-to-end mass budgets are SLV = space launch vehicle calculated, and a concept of operations is presented.
    [Show full text]