DOCSLIB.ORG

Development of an Earth Smallsat Flight Test to Demonstrate Viability of Mars Aerocapture

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text

Load more

SmallSat Aerocapture Demonstration Development of an Earth SmallSat Flight Test to Demonstrate Viability of Mars Aerocapture Michael Werner / Bryce Woollard / Anirudh Tadanki / Swapnil Pujari / Dr. Robert Braun Space Systems Design Laboratory, Georgia Institute of Technology Rob Lock / Adam Nelessen / Ryan Woolley Jet Propulsion Laboratory, California Institute of Technology Aerocapture Benefits & Control Flight System Overview TrajectoryTrajectory Analysis Modeling Benefits of Aerocapture Aeroshell Design L Trajectory Modeling • Minimal propellant requirements • Aeroshell and drag skirt made from machined R • End-to-end simulation models • Potential large savings in mass and cost Example Aerocapture (“A/C”) Mass Savings aluminum vehicle trajectory and Skirt Jettison • Studies have shown conceptual & • PICA selected as TPS material uncertainties Entry Exit Mission A/C Mass (kg) Non-A/C Mass (kg) % Increase r technical viability ψ • 3-DOF integration of Venus Low-Circular Orbit 5078 2834 79 β β2 atmospheric equations of 1 Lack of flight demonstration inhibits Venus Elliptical Orbit 5078 3542 43 Example Trajectory Outputs Aeroshell & Drag Skirt Properties motion Nominal Atmospheric Trajectory Plots use of aerocapture on actual missions Mars Low Circular Orbit 5232 4556 15 Location Dimension Value • Key outputs: Altitude Jupiter Low Circular Orbit 2262 N/A Enabling Drag Skirt Outer Radius (R) 25 cm Aerocapture Control: Drag Modulation • Deceleration (km) • Ballistic coefficient changes used to alter Saturn Circular Orbit 494 N/A Enabling Blunted Cone Angle 60° • Heat Rates Inner Radius (r) 10 cm vehicle’s trajectory Titan Circular Orbit 2630 691 280 Foreshell • Orbit uncertainties Nose Radius 5.625 cm • Simple to implement compared to Deceleration Uranus Elliptical Orbit 1966 618 218 Shoulder Radius 1 cm traditional lifting methods: (g’s) Neptune Elliptical Orbit 1680 180 832 Length (L) 22.7 cm • No CG offset required Backshell Backshell Taper (ψ) 5.3° • Simple avionics algorithms Adapted from: Hall et al, “Cost-Benefit Analysis of the Aerocapture Mission Set”, Journal of Spacecraft & Rockets, 2005 Heat Rate So easy, a SmallSatcan do it (W/cm2) Discrete Single-Event Drag Modulation Drag Skirt & Jettison Mechanism Planet-Relative Velocity (km/s) • Drag skirt broken into 4 quadrants • Jettison accomplished via 3 staggered M3 Control Scheme pyrotechnic bolts per quadrant • Jettison timing controls amount of End-to-End Monte Carlo Results: 10000 Samples, Target = 1760 km • Spring loaded at interface to ensure safe aerocapture ΔV detachment • Guidance algorithm: numerical predictor-corrector • Only atmospheric accelerometer measurements required • Monte Carlo results show robustness to uncertainties in entry FPA, atmosphere, and burns Number of Samples Mission Concept Static Stability Static Margin Timeline Pre-Jettison Configuration Time (s) Static Margin (cm) • CG Location: 7.65 cm aft of nose vertex Monte Carlo Summary Purpose: Develop SmallSat mission concept for flight test demonstrating aerocapture at Earth 0 N/A • Backshell is shadowed from hypersonic flow by drag Parameter Value Philosophy: Prioritize simplicity & affordability Final Apogee Altitude (km) skirt 50 63.42 Mean 1746.0 km Requirements: Pre-Jettison • Ample static margin and stability 1. Aerocapture ΔV: ~2 km/s 100 58.99 3σ Error 169.4 km Jettison guidance algorithm is able to 2. Successful raise maneuver 150 58.64 Minimum 1596.1 km effectively target desired orbit Post-Jettison 3. Telemetry data return Maximum 1813.7 km • 150 7.98 Deliverable: Fully-documented mission concept in anticipation of future SmallSat proposal opportunities CG Location: 10.6 cm aft of nose vertex • Exposure of backshell to flow leads to reduced static 200 8.52 Mission Timeline stability Post-Jettison 250 11.34 Conclusions # Event Details 5. Atmospheric Interface 300 -3.31 1 GTO Period: 10 hrs 4. Descent Orbit Takeaways 350 -243.88 • A SmallSat mission featuring single-event jettison drag modulation is a promising architecture to Separation from 2 1 hr before apogee demonstrate a ~2 km/s aerocapture maneuver at Earth Host 9. Final Orbit • Structural and mission design analyses have led to a <25 kg flight system featuring COTS hardware Perigee Lowering Nominal Flight System 3 ΔV: ~10 m/s • Trajectory analyses and Monte Carlo results show robustness of system design to uncertainties Maneuver (PLM) 8. Perigee Raise • Aerojet MPS-120XL Hydrazine Propulsion 3. Perigee Maneuver • Sensonor IMU 4 Descent Orbit Duration: 5 hrs Lowering • ISIS On-Board Computer Atmospheric Altitude: 125 km Maneuver 5 • GomSpace EPS & Batteries Interface Velocity: 10.3 km/s 6. Aerocapture 7. Post Aerocapture Orbit • Astrodev UHF Radio For more information: 6 Aerocapture ΔV: ~2000 m/s • GomSpace UHF Dipole Antenna Post-Aerocapture Michael Werner 7 60 km x 1750 km 2. Separation Orbit 1. Geosynchronous [email protected] Transfer Orbit (GTO) from Host Perigee Raise 8 ΔV: ~40 m/s Maneuver (PRM) GTO to LEO A <25 kg flight system based on COTS CubeSat 9 Final Orbit 200km x 1750 km Aerocapture ΔV: 2 km/s hardware meets all mission requirements.
Recommended publications
  • An Assessment of Aerocapture and Applications to Future Missions

    An Assessment of Aerocapture and Applications to Future Missions

    Post-Exit Atmospheric Flight Cruise Approach An Assessment of Aerocapture and Applications to Future Missions February 13, 2016 National Aeronautics and Space Administration An Assessment of Aerocapture Jet Propulsion Laboratory California Institute of Technology Pasadena, California and Applications to Future Missions Jet Propulsion Laboratory, California Institute of Technology for Planetary Science Division Science Mission Directorate NASA Work Performed under the Planetary Science Program Support Task ©2016. All rights reserved. D-97058 February 13, 2016 Authors Thomas R. Spilker, Independent Consultant Mark Hofstadter Chester S. Borden, JPL/Caltech Jessie M. Kawata Mark Adler, JPL/Caltech Damon Landau Michelle M. Munk, LaRC Daniel T. Lyons Richard W. Powell, LaRC Kim R. Reh Robert D. Braun, GIT Randii R. Wessen Patricia M. Beauchamp, JPL/Caltech NASA Ames Research Center James A. Cutts, JPL/Caltech Parul Agrawal Paul F. Wercinski, ARC Helen H. Hwang and the A-Team Paul F. Wercinski NASA Langley Research Center F. McNeil Cheatwood A-Team Study Participants Jeffrey A. Herath Jet Propulsion Laboratory, Caltech Michelle M. Munk Mark Adler Richard W. Powell Nitin Arora Johnson Space Center Patricia M. Beauchamp Ronald R. Sostaric Chester S. Borden Independent Consultant James A. Cutts Thomas R. Spilker Gregory L. Davis Georgia Institute of Technology John O. Elliott Prof. Robert D. Braun – External Reviewer Jefferey L. Hall Engineering and Science Directorate JPL D-97058 Foreword Aerocapture has been proposed for several missions over the last couple of decades, and the technologies have matured over time. This study was initiated because the NASA Planetary Science Division (PSD) had not revisited Aerocapture technologies for about a decade and with the upcoming study to send a mission to Uranus/Neptune initiated by the PSD we needed to determine the status of the technologies and assess their readiness for such a mission.
  • Deep Space 2: the Mars Microprobe Project and Beyond

    Deep Space 2: the Mars Microprobe Project and Beyond

    First International Conference on Mars Polar Science 3039.pdf DEEP SPACE 2: THE MARS MICROPROBE PROJECT AND BEYOND. S. E. Smrekar and S. A. Gavit, Mail Stop 183-501, Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasa- dena CA 91109, USA ([email protected]). Mission Overview: The Mars Microprobe Proj- System Design, Technologies, and Instruments: ect, or Deep Space 2 (DS2), is the second of the New Telecommunications. The DS2 telecom system, Millennium Program planetary missions and is de- which is mounted on the aftbody electronics plate, signed to enable future space science network mis- relays data back to earth via the Mars Global Surveyor sions through flight validation of new technologies. A spacecraft which passes overhead approximately once secondary goal is the collection of meaningful science every 2 hours. The receiver and transmitter operate in data. Two micropenetrators will be deployed to carry the Ultraviolet Frequency Range (UHF) and data is out surface and subsurface science. returned at a rate of 7 Kbits/second. The penetrators are being carried as a piggyback Ultra-low-temperature lithium primary battery. payload on the Mars Polar Lander cruise ring and will One challenging aspect of the microprobe design is be launched in January of 1999. The Microprobe has the thermal environment. The batteries are likely to no active control, attitude determination, or propulsive stay no warmer than -78° C. A lithium-thionyl pri- systems. It is a single stage from separation until mary battery was developed to survive the extreme landing and will passively orient itself due to its aero- temperature, with a 6 to 14 V range and a 3-year shelf dynamic design (Fig.
  • Aerothermodynamic Design of the Mars Science Laboratory Heatshield

    Aerothermodynamic Design of the Mars Science Laboratory Heatshield

    Aerothermodynamic Design of the Mars Science Laboratory Heatshield Karl T. Edquist∗ and Artem A. Dyakonovy NASA Langley Research Center, Hampton, Virginia, 23681 Michael J. Wrightz and Chun Y. Tangx NASA Ames Research Center, Moffett Field, California, 94035 Aerothermodynamic design environments are presented for the Mars Science Labora- tory entry capsule heatshield. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat flux, a first for Mars entry, and the heatshield environments were defined for a fully-turbulent heat pulse. The effects of distributed surface roughness on turbulent heat flux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W=cm2 for heat flux, 471 P a for shear stress, 0.371 Earth atm for pressure, and 5477 J=cm2 for total heat load. Time-varying conditions at fixed heatshield locations were generated for thermal protection system analysis and flight instrumentation development. Finally, the aerother- modynamic effects of delaying launch until 2011 are previewed. Nomenclature 1 2 2 A reference area, 4 πD (m ) CD drag coefficient, D=q1A D aeroshell diameter (m) 2 Dim multi-component diffusion coefficient (m =s) ci species mass fraction H total enthalpy
  • Aerocapture FS-Pdf.Indd

    Aerocapture FS-Pdf.Indd

    NASA Facts National Aeronautics and Space Administration Marshall Space Flight Center Huntsville, Alabama 35812 FS-2004-09-127-MSFC July 2004 Aerocapture Technology NASA technologists are working to develop ways to place robotic space vehicles into long-duration, scientific orbits around distant Solar System destinations without the need for the heavy, on-board fuel loads that have historically inhibited vehicle performance, mission dura- tion and available mass for science payloads. Aerocapture -- a flight maneuver that inserts a space- craft into its proper orbit once it arrives at a planet -- is part of a unique family of “aeroassist” technologies under consideration to achieve these goals and enable robust science missions to any planetary body with an appreciable atmosphere. Aerocapture technology is just one of many propulsion technologies being developed by NASA technologists and their partners in industry and academia, led by NASA’s In-Space Propulsion Technology Office at the Marshall Space Flight Center in Huntsville, Ala. The Center implements the In-Space Propulsion Technolo- gies Program on behalf of NASA’s Science Mission Directorate in Washington. Aerocapture uses a planet’s or moon’s atmosphere to accomplish a quick, near-propellantless orbit capture to place a space vehicle in its proper orbit. The atmo- sphere is used as a brake to slow down a spacecraft, transferring the energy associated with the vehicle’s high speed into thermal energy. Aerocapture entering Mars Orbit. The aerocapture maneuver starts with an approach trajectory into the atmosphere of the target body. The dense atmosphere creates friction, slowing the craft and placing it into an elliptical orbit -- an oval shaped orbit.
  • Up, Up, and Away by James J

    Up, Up, and Away by James J

    www.astrosociety.org/uitc No. 34 - Spring 1996 © 1996, Astronomical Society of the Pacific, 390 Ashton Avenue, San Francisco, CA 94112. Up, Up, and Away by James J. Secosky, Bloomfield Central School and George Musser, Astronomical Society of the Pacific Want to take a tour of space? Then just flip around the channels on cable TV. Weather Channel forecasts, CNN newscasts, ESPN sportscasts: They all depend on satellites in Earth orbit. Or call your friends on Mauritius, Madagascar, or Maui: A satellite will relay your voice. Worried about the ozone hole over Antarctica or mass graves in Bosnia? Orbital outposts are keeping watch. The challenge these days is finding something that doesn't involve satellites in one way or other. And satellites are just one perk of the Space Age. Farther afield, robotic space probes have examined all the planets except Pluto, leading to a revolution in the Earth sciences -- from studies of plate tectonics to models of global warming -- now that scientists can compare our world to its planetary siblings. Over 300 people from 26 countries have gone into space, including the 24 astronauts who went on or near the Moon. Who knows how many will go in the next hundred years? In short, space travel has become a part of our lives. But what goes on behind the scenes? It turns out that satellites and spaceships depend on some of the most basic concepts of physics. So space travel isn't just fun to think about; it is a firm grounding in many of the principles that govern our world and our universe.
  • Magnetoshell Aerocapture: Advances Toward Concept Feasibility

    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.
  • Orbital Fueling Architectures Leveraging Commercial Launch Vehicles for More Affordable Human Exploration

    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 .........................................................................
  • Reusable Rocket Upper Stage Development of a Multidisciplinary Design Optimisation Tool to Determine the Feasibility of Upper Stage Reusability L

    Reusable Rocket Upper Stage Development of a Multidisciplinary Design Optimisation Tool to Determine the Feasibility of Upper Stage Reusability L

    Reusable Rocket Upper Stage Development of a Multidisciplinary Design Optimisation Tool to Determine the Feasibility of Upper Stage Reusability L. Pepermans Technische Universiteit Delft Reusable Rocket Upper Stage Development of a Multidisciplinary Design Optimisation Tool to Determine the Feasibility of Upper Stage Reusability by L. Pepermans to obtain the degree of Master of Science at the Delft University of Technology, to be defended publicly on Wednesday October 30, 2019 at 14:30 AM. Student number: 4144538 Project duration: September 1, 2018 – October 30, 2019 Thesis committee: Ir. B.T.C Zandbergen , TU Delft, supervisor Prof. E.K.A Gill, TU Delft Dr.ir. D. Dirkx, TU Delft This thesis is confidential and cannot be made public until October 30, 2019. An electronic version of this thesis is available at http://repository.tudelft.nl/. Cover image: S-IVB upper stage of Skylab 3 mission in orbit [23] Preface Before you lies my thesis to graduate from Delft University of Technology on the feasibility and cost-effectiveness of reusable upper stages. During the accompanying literature study, it was determined that the technology readiness level is sufficiently high for upper stage reusability. However, it was unsure whether a cost-effective system could be build. I have been interested in the field of Entry, Descent, and Landing ever since I joined the Capsule Team of Delft Aerospace Rocket Engineering (DARE). During my time within the team, it split up in the Structures Team and Recovery Team. In September 2016, I became Chief Recovery for the Stratos III student-built sounding rocket. During this time, I realised that there was a lack of fundamental knowledge in aerodynamic decelerators within DARE.
  • Smallsat Aerocapture to Enable a New Paradigm of Planetary Missions

    Smallsat Aerocapture to Enable a New Paradigm of Planetary Missions

    SmallSat Aerocapture to Enable a New Paradigm of Planetary Missions Alex Austin, Adam Nelessen, Bill Ethiraj Venkatapathy, Robin Beck, Robert Braun, Michael Werner, Evan Strauss, Joshua Ravich, Mark Jesick Paul Wercinski, Michael Aftosmis, Roelke Jet Propulsion Laboratory, California Michael Wilder, Gary Allen CU Boulder Institute of Technology NASA Ames Research Center Boulder, CO 80309 Pasadena, CA 91109 Moffett Field, CA 94035 [email protected] 818-393-7521 [email protected] [email protected] Abstract— This paper presents a technology development system requirements, which concluded that mature TPS initiative focused on delivering SmallSats to orbit a variety of materials are adequate for this mission. CFD simulations were bodies using aerocapture. Aerocapture uses the drag of a single used to assess the risk of recontact by the drag skirt during the pass through the atmosphere to capture into orbit instead of jettison event. relying on large quantities of rocket fuel. Using drag modulation flight control, an aerocapture vehicle adjusts its drag area This study has concluded that aerocapture for SmallSats could during atmospheric flight through a single-stage jettison of a be a viable way to increase the delivered mass to Venus and can drag skirt, allowing it to target a particular science orbit in the also be used at other destinations. With increasing interest in presence of atmospheric uncertainties. A team from JPL, NASA SmallSats and the challenges associated with performing orbit Ames, and CU Boulder has worked to address the key insertion burns on small platforms, this technology could enable challenges and determine the feasibility of an aerocapture a new paradigm of planetary science missions.
  • Aas 19-725 Improved Atmospheric Estimation for Aerocapture Guidance

    Aas 19-725 Improved Atmospheric Estimation for Aerocapture Guidance

    AAS 19-725 IMPROVED ATMOSPHERIC ESTIMATION FOR AEROCAPTURE GUIDANCE Evan Roelke,∗ Phil D. Hattis,y and R.D. Braunz Increased interest in Lunar or Mars-sample return missions encourages considera- tion of innovative orbital operations such as aerocapture, which generally provides significant mass-savings for orbital insertion at Earth or Mars. Drag modulation architectures offer a straightforward approach to orbital apoapsis targeting by en- abling ballistic entry, among other benefits. A shortcoming of these architectures is the poor estimation of atmospheric density resulting in target apoapsis altitude errors. This research seeks to assess and improve upon current atmospheric den- sity estimation techniques in order to support the flight viability of discrete event drag modulated aerocapture. Three different estimation techniques are assessed in terms of estimation error and apoapsis altitude error: a static density factor, a density array interpolator, and an ensemble correlation filter. The density inter- polator achieves a 5% improvement in median apoapsis altitude over the density factor when entering at −5:9◦ and targeting a 2000km apoapsis altitude, while the ensemble correlation filter achieves a 7% improvement under identical simulation conditions. The ensemble correlation filter was found to improve with decreasing density search tolerance, achieving a 4:6% improvement in median apoapsis alti- tude for a tolerance of 1% over 5%. These improvements are dependent on entry and vehicle parameters and improve as the entry angle becomes more shallow or the target apoapsis is reduced. Errors in the density factor measurements are main contributors to the error in estimated versus true density profiles. INTRODUCTION The revitalized interest in Lunar missions as well as the drive for Mars sample-return missions in recent years encourages innovative solutions to orbital operations.
  • 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 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.
  • Author's Instructions For

    Author's Instructions For

    Feasibility Analysis for a Manned Mars Free-Return Mission in 2018 Dennis A. Tito Grant Anderson John P. Carrico, Jr. Wilshire Associates Incorporated Paragon Space Development Applied Defense Solutions, Inc. 1800 Alta Mura Road Corporation 10440 Little Patuxent Pkwy Pacific Palisades, CA 90272 3481 East Michigan Street Ste 600 310-260-6600 Tucson, AZ 85714 Columbia, MD 21044 [email protected] 520-382-4812 410-715-0005 [email protected] [email protected] Jonathan Clark, MD Barry Finger Gary A Lantz Center for Space Medicine Paragon Space Development Paragon Space Development Baylor College Of Medicine Corporation Corporation 6500 Main Street, Suite 910 1120 NASA Parkway, Ste 505 1120 NASA Parkway, Ste 505 Houston, TX 77030-1402 Houston, TX 77058 Houston, TX 77058 [email protected] 281-702-6768 281-957-9173 ext #4618 [email protected] [email protected] Michel E. Loucks Taber MacCallum Jane Poynter Space Exploration Engineering Co. Paragon Space Development Paragon Space Development 687 Chinook Way Corporation Corporation Friday Harbor, WA 98250 3481 East Michigan Street 3481 East Michigan Street 360-378-7168 Tucson, AZ 85714 Tucson, AZ 85714 [email protected] 520-382-4815 520-382-4811 [email protected] [email protected] Thomas H. Squire S. Pete Worden Thermal Protection Materials Brig. Gen., USAF, Ret. NASA Ames Research Center NASA AMES Research Center Mail Stop 234-1 MS 200-1A Moffett Field, CA 94035-0001 Moffett Field, CA 94035 (650) 604-1113 650-604-5111 [email protected] [email protected] Abstract—In 1998 Patel et al searched for Earth-Mars free- To size the Environmental Control and Life Support System return trajectories that leave Earth, fly by Mars, and return to (ECLSS) we set the initial mission assumption to two crew Earth without any deterministic maneuvers after Trans-Mars members for 500 days in a modified SpaceX Dragon class of Injection.