DOCSLIB.ORG

Design and Thermal Analysis of a Solar Thermal Thruster

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Design and Thermal Analysis of a Solar Thermal Thruster Design and Thermal Analysis of a Solar Thermal Thruster Krishti Das Technische Universiteit Delft Design and Thermal Analysis of a Solar Thermal Thruster by Krishti Das in partial fulfillment of the requirements for the degree of Master of Science in Aerospace Engineering at the Delft University of Technology, to be defended publicly on Wednesday 28th of November, 2018 at 9:00 AM Student number: 4624408 Supervisor: Dr. Angelo Cervone Daily Supervisor : Fiona Leverone Thesis committee: Prof. dr. E.K.A Gill, TU Delft Dr. Angelo Cervone, TU Delft Fiona Leverone, TU Delft Dr. M.J Heiligers, TU Delft An electronic version of this thesis is available at http://repository.tudelft.nl/. Acknowledgement I would like to start with expressing my deepest gratitude and regard to Dr. Angelo Cervone for his constant guidance, support and answering my silly doubts throughout. At this point, I am genuinely grateful to have had Fiona Leverone as my daily supervisor. She not only supervised me during my thesis and encouraged me to work better but she has helped me become a better researcher. Though I have a long way to go, I have been really fortunate to study at TU Delft and working with professors and students of great caliber. Last but not least, I would like to thank my mom, dad, and sister who have been my backbone throughout from the other side of the world. I am extremely grateful for the friends I have made during my studies and I would thank Maria, Shreyas, and Amantrita. Their unconditional help and support in my life and studies, in general, has been remarkable. I am also deeply grateful to Stanislaus and Jayraj for their critical and supportive feedback on my work. Finally, I owe it to the boys in the MSc student room, especially Mathijs Van de Poel and Sean Pepper, and to our countless late night conversations about space news and sharing ideas with each other. I would also like to thank Samiksha for her constant encouragement and tea breaks. iii Contents Acknowledgement . vii Nomenclature x Glossary xi List of Figures xiii List of Tables xvii 1 Introduction 1 2 Research Framework 3 2.1 Literature Study . 3 2.2 Existing STP Engine Concepts and Designs . 4 2.3 Preliminary Mission Analysis . 8 2.4 System Requirements . 9 2.5 Research Question, Aims, and Objectives . 10 2.5.1 Research Objective . 10 2.5.2 Research Question . 11 2.6 Design Methodology . 12 3 Conceptual Designs 15 3.1 Design Option Tree . 15 3.1.1 Concentrator - Receiver Interface . 15 3.1.2 Heating Method . 17 3.1.3 Cavity Shape . 18 3.1.4 Nozzle Shape. 19 3.2 Conceptual Designs. 19 3.2.1 Heat Exchanger Concept 1. 20 3.2.2 Heat Exchanger Concept 2. 20 3.2.3 Heat Exchanger Concept 3. 21 3.3 Concept Trade-off . 23 3.3.1 Trade-off Criteria . 23 3.3.2 Preliminary Channel Length. 24 3.3.3 Graphical Trade-off . 26 4 Baseline Design 29 4.1 Micro-nozzle Design . 29 4.1.1 Rocket Propulsion Fundamentals . 30 4.1.2 Actual Performance . 33 4.1.3 Effect of Nozzle Geometry on Nozzle Efficiency . 36 4.1.4 Thruster Performance at High Chamber Temperatures . 38 v vi Contents 4.2 Materials. 39 4.3 Heat Exchanger . 41 4.3.1 Flow Regime in Rectangular Micro-channel . 42 4.3.2 Pressure Loss in Rectangular Micro-channels . 43 4.3.3 Channel Wall Thickness . 45 4.3.4 Cavity Dimensioning. 46 4.4 Optical Fiber Cable . 46 4.5 Transmission Efficiency and Attenuation. 47 4.5.1 Optical Fiber Efficiency . 47 4.5.2 Sources of Attenuation . 48 4.6 Optical Fiber Candidates . 52 4.7 Concentrator Design . 53 4.7.1 Mass Estimation and Material . 55 4.7.2 Concentrator - Optical Fiber Requirement. 57 4.7.3 Bundled and Single Optical Fibers . 58 4.8 Initial Design . 62 5 Thermal Modelling 65 5.1 Geometrical Mathematical Model . 66 5.2 Heat Loss Model . 67 5.2.1 Input Thermal Power . 67 5.2.2 Case I: Without Insulation . 67 5.2.3 Case II: With Insulation. 70 5.2.4 Power Lost to Propellant Flow . 71 5.3 Thermal Study of the Optical System . 71 5.3.1 Heat Loss in Optical Fiber . 72 5.3.2 Optical Fiber Power Transmission . 73 5.4 Time Dependent Model . 74 5.5 Flow and Channel Wall Temperature Model . 75 6 Final Design and Results 81 6.1 Heat Exchanger . 82 6.1.1 Ray Tracing Model . 82 6.1.2 Ray Optics Modelling in COMSOL . 85 6.1.3 Conjugate Heat Transfer Model. 89 6.2 Fibre Optic-Receiver Coupling . 94 6.2.1 Optical Fiber Temperature due to Hot Cavity . 94 6.2.2 Fibre Optic-Receiver Coupling Concepts . 96 6.3 Verification of the Analytical model . 103 6.4 Thrust Performance. 103 6.5 Final Design Render . 104 6.6 System Requirements Verification . 108 7 Conclusions and Recommendations 111 Appendices 115 A Optical Fiber Specifications Sheet 117 B Advanced Ceramics Property Sheet 119 C Nusselt Number of Three Side Heated Walls 121 D Verification of Analytical and Comsol Model 123 Contents vii Bibliography 125 Nomenclature 훼 Absorption coefficient of the cavity material Δ푉 Velocity change ̇푚 Mass flow rate 휖 Emissivity of a surface 휂 Receiver efficiency Γ Vandenkerckhove constant 훾 Specific heat ratio 휇 Dynamic viscosity at a propellant temperature 휙 Solar flux at LEO 휌 Reflectivity Factor 퐴/퐴 Cavity opening area to internal area Ratio 퐴 Nozzle exit area 퐴/퐴 Expansion ratio 퐴, 퐴 Throat area of the nozzle 퐵 Gebhart Factor 퐶 Concentration ratio 퐶 Conductive Coupling 퐷 Throat diameter of the nozzle 푒 Cavity internal surface emissivity 푒 Cavity external surface emissivity 퐹 Thrust 퐹 View Factor 푔 Gravitational Acceleration 퐼 Specific Impulse 푘 Thermal Conductivity 푀 Propellant Mass 푀 Satellite Wet Mass ix x Contents 푃 Power collected 푃 Chamber pressure 푃/푃 Pressure ratio 푃 Power collected by Concentrator 푄 Incoming solar energy 푄 Heat flow to the propellant 푄 Hear lost through re-radiation 푄 Heat lost from the cavity external surface 푄 Heat lost by reflection 푄 Solar radiation absorbed by the absorber cavity 푅 Specific gas constant 푅 Radiative Coupling 푅푒 Reynolds number at nozzle throat 푆 Cross-sectional area of the concentrator 푇 Temperature 푇 Internal surface temperature of cavity 푇 Chamber temperature 푇 External surface temperature of cavity Glossary AFRPL Air Force Rocket Propulsion Laboratory. 4 CSF Conduction Shape Factor. 69 LEO Low Earth Orbit. 1 LUMIO Lunar Meteoroid Impact Observer. 9 MEMS Micro-Electro-Mechanical-Systems. 2 OH Hydroxyl Ion. 52 OW Optical Wave. xiii, xvii, 15, 17 STP Solar Thermal Propulsion. 2 STS Space Transportation System. 5 xi List of Figures 2.1 Solar thermal propulsion mechanism ........................... 3 2.2 Performance of solar thermal propulsion systems vs chemical, electric, and nuclear thermal propulsion options ................................ 4 2.3 Conceptual design of a solar thermal propulsion system ................ 5 2.4 Single crystal molybdenum solar thermal thruster sketch ................ 6 2.5 Mk.I thruster design ................................... 7 2.6 Final RAC design with nozzle ............................... 7 2.7 STP engine concept for PocketQubes ........................... 8 2.8 Design methodology .................................... 13 3.2 Left: Schematic of conventional solar thermal propulsion [1]; Right: Optical wave (OW) guide solar thermal propulsion system with integrated optical wave guide ACS [2] 15 3.1 Design option tree ..................................... 16 3.3 Net irradiance distributions on corrugated and flat samples [3] ............. 19 3.4 Concept 1 - trapezoidal cavity and serpentine channels ................. 20 3.5 Concept 2 - rectangular corrugated cavity and rectangular channels .......... 21 3.6 Platelet heat exchanger .................................. 22 3.7 Concept 3 - platelet heat exchanger ........................... 22 3.8 Graphical trade-off between conceptual designs ..................... 27 4.1 Conical nozzle configuration ............................... 30 4.2 Effect of divergence angle on thrust efficiency ...................... 38 4.3 The insulation thickness vs temperature for Pyrogel ................... 40 4.4 Total internal reflection phenomena in an optical fiber ................. 47 4.5 Thermal power flow through the Heat Exchanger Receiver Cavity system without insulation .......................................... 49 4.6 Projection of rays into a rectangular closed cavity through bending in optical fibers . 50 4.7 Attenuation due to microbends [4] ............................ 50 4.8 Power output in optical fibers of different numerical apertures with varying core diameter 55 4.9 Inflatable Concentrator using torus and tripod tube as support structure [5] ..... 56 4.10 Possible configuration of single optical fiber system ................... 59 4.11 Bundled optical fibers ................................... 60 4.12 A possible configuration of bundled optical fiber system ................ 60 4.13 Optical system mass for single and bundled fiber system ................ 61 4.14 Trade-off matrix for optical system concept ....................... 62 4.15 Baseline design ...................................... 63 5.1 Structure of the thermal model for flow and channel wall temperature ........ 65 5.2 Thermal power flow through the Heat Exchanger Receiver Cavity system without insulation .......................................... 66 5.3 Schematic diagram of enclosed cavity and outer geometry with effective gap (thick- ness) for conduction .................................... 69 5.4 Total power output vs. No. of optical fibers plot ..................... 73 xiii xiv List of Figures 5.5 Cavity wall temperature with increasing no. of optical fibers .............. 75 5.6 Effect of channel height on the final flow temperature ................. 76 5.7 Effect of channel aspect ratio on the final flow temperature .............. 77 5.8 Determining the required (minimum) channel length to reach a flow temperature of 1000 K ........................................... 78 5.9 Pressure drop in the channels with respect to aspect ratio for a constant channel length of 4.023 cm .................................... 79 5.10 Baseline design dimensions: front view ........................
Load more

Recommended publications
  • 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.
    [Show full text]
  • MULTIPHYSICS DESIGN and SIMULATION of a TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET a Thesis by BRAD APPEL Submitted to the Office O
    MULTIPHYSICS DESIGN AND SIMULATION OF A TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET A Thesis by BRAD APPEL Submitted to the Office of Graduate Studies of Texas A&M University in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE August 2012 Major Subject: Nuclear Engineering Multiphysics Design and Simulation of a Tungsten-Cermet Nuclear Thermal Rocket Copyright 2012 Brad Appel ii MULTIPHYSICS DESIGN AND SIMULATION OF A TUNGSTEN-CERMET NUCLEAR THERMAL ROCKET A Thesis by BRAD APPEL Submitted to the Office of Graduate Studies of Texas A&M University in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE Approved by: Chair of Committee, Karen Vierow Committee Members, Shannon Bragg-Sitton Paul Cizmas Head of Department, Yassin Hassan August 2012 Major Subject: Nuclear Engineering iii iii ABSTRACT Multiphysics Design and Simulation of a Tungsten-Cermet Nuclear Thermal Rocket. (August 2012) Brad Appel, B.S., Purdue University Chair of Advisory Committee: Dr. Karen Vierow The goal of this research is to apply modern methods of analysis to the design of a tungsten-cermet Nuclear Thermal Rocket (NTR) core. An NTR is one of the most viable propulsion options for enabling piloted deep-space exploration. Concerns over fuel safety have sparked interest in an NTR core based on tungsten-cermet fuel. This work investigates the capability of modern CFD and neutronics codes to design a cermet NTR, and makes specific recommendations for the configuration of channels in the core. First, the best CFD practices available from the commercial package Star-CCM+ are determined by comparing different modeling options with a hot-hydrogen flow experiment.
    [Show full text]
  • Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion
    AAS 05-398 Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion ∗ ϒ Daniel W. Parcher ∗∗ and Jon A. Sims ϒϒ This paper examines optimal low-thrust gravity-assist trajectories to Jupiter using nuclear electric propulsion. Three different Venus-Earth Gravity Assist (VEGA) types are presented and compared to other gravity-assist trajectories using combinations of Earth, Venus, and Mars. Families of solutions for a given gravity-assist combination are differentiated by the approximate transfer resonance or number of heliocentric revolutions between flybys and by the flyby types. Trajectories that minimize initial injection energy by using low resonance transfers or additional heliocentric revolutions on the first leg of the trajectory offer the most delivered mass given sufficient flight time. Trajectory families that use only Earth gravity assists offer the most delivered mass at most flight times examined, and are available frequently with little variation in performance. However, at least one of the VEGA trajectory types is among the top performers at all of the flight times considered. INTRODUCTION The use of planetary gravity assists is a proven technique to improve the performance of interplanetary trajectories as exemplified by the Voyager, Galileo, and Cassini missions. Another proven technique for enhancing the performance of space missions is the use of highly efficient electric propulsion systems. Electric propulsion can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems.1,2 This technology has been demonstrated on the Deep Space 1 mission 3 − part of NASA’s New Millennium Program to validate technologies which can lower the cost and risk and enhance the performance of future missions.
    [Show full text]
  • 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 .......................................................................................................................................................
    [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]
  • 1 Iac-06-C4.4.7 the Innovative Dual-Stage 4-Grid Ion
    IAC-06-C4.4.7 THE INNOVATIVE DUAL-STAGE 4-GRID ION THRUSTER CONCEPT – THEORY AND EXPERIMENTAL RESULTS Cristina Bramanti, Roger Walker, ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], Roger. Walker @esa.int Orson Sutherland, Rod Boswell, Christine Charles Plasma Research Laboratory, Research School of Physical Sciences and Engineering, The Australian National University, Canberra, ACT 0200, Australia [email protected], [email protected], [email protected]. David Fearn EP Solutions, 23 Bowenhurst Road, Church Crookham, Fleet, Hants, GU52 6HS, United Kingdom [email protected] Jose Gonzalez Del Amo, Marika Orlandi ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, The Netherlands [email protected], [email protected] ABSTRACT A new concept for an advanced “Dual-Stage 4-Grid” (DS4G) ion thruster has been proposed which draws inspiration from Controlled Thermonuclear Reactor (CTR) experiments. The DS4G concept is able to operate at very high specific impulse, power and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. A small low-power experimental laboratory model was designed and built under a preliminary research, development and test programme, and its performance was measured during an extensive test campaign, which proved the practical feasibility of the overall concept and demonstrated the performance predicted by analytical and simulation models. In the present paper, the basic concept of the DS4G ion thruster is presented, along with the design, operating parameters and measured performance obtained from the first and second phases of the experimental campaign.
    [Show full text]
  • Single-Stage Drag Modulation GNC Performance for Venus Aerocapture Demonstration
    See discussions, stats, and author profiles for this publication at: https://www.researchgate.net/publication/330197775 Single-Stage Drag Modulation GNC Performance for Venus Aerocapture Demonstration Conference Paper · January 2019 DOI: 10.2514/6.2019-0016 CITATIONS READS 5 32 3 authors, including: Evan Roelke University of Colorado Boulder 9 PUBLICATIONS 18 CITATIONS SEE PROFILE Some of the authors of this publication are also working on these related projects: Multi-Event Drag Modulated Aerocapture Guidance View project All content following this page was uploaded by Evan Roelke on 14 September 2020. The user has requested enhancement of the downloaded file. Single-Stage Drag Modulation GNC Performance for Venus Aerocapture Demonstration Roelke, E.∗, Braun, R.D.†, and Werner, M.S.‡ University of Colorado Boulder The future of space exploration is heavily reliant on innovative Entry, Descent, and Landing (EDL) technologies. One such innovation involves aeroassist technology, mainly aerocapture, which involves a single pass through the upper atmosphere of a planetary body from an incom- ing hyperbolic trajectory to capture into an elliptical orbit. Previous aerocapture guidance studies focused on lift modulation using reaction-control system (RCS) thrusters. This pa- per investigates the guidance law performance of drag-modulated aerocapture for a single, discrete, jettison event. Venus is selected as the target planet to observe performance in the most challenging environment. Three different guidance algorithms are studied, including a deceleration curve-fit, a pure state-predictor, and a numerical predictor-corrector utilizing both bisection and Newton-Raphson root-finding methods. The performance is assessed us- ing Monte Carlo simulation methods. The deceleration curve-fit achieves a capture rate of over 55% for a 2000km apoapsis target and 1000km tolerance.
    [Show full text]
  • Energy Sources for Propulsion To, and for Distributed Use On, Mars
    id10495406 pdfMachine by Broadgun Software - a great PDF writer! - a great PDF creator! - http://www.pdfmachine.com http://www.broadgun.com Mehtapress 2015 Print - ISSN : 2319–9814 Journal of Online - ISSN : 2319–9822 SpFauclle pEapxerploration Full Paper WWW.MEHTAPRESS.COM Mitchell Swartz Energy sources for propulsion to, and for SPSR distributed use on, Mars Received : July 06, 2015 Accepted : August 12, 2015 Published : October 14, 2015 *Corresponding author’s Name & Add. Mitchell Swartz SPSR MORE POWER IS NEEDED Curiosity, in space, measured 1.8 millisieverts per day, with 97% from galactic cosmic rays and the rest from Mars is the future. How do we know that? Because solar particles. This is a very high level which is six times as many people signed up (202,586) for a associated with carcinogenesis and other unwanted one-way ride than did for the Affordable Care Act effects. In Mars, this exposure decreased to circa 0.7 which means that people would rather buy a ticket to millisieverts per day, similar to the exposure received the vast unknown than get healthcare as they age. on the International Space Station. What is holding back the transport of humans to Mars, Other dangers beyond the unwanted irradiation and its colonization, is energy. We need more powerful include biological changes from bone loss and muscle rockets to take people there quicker, and we need more atrophy to electrolyte changes. The goal is to decrease accessible, more portable distributed power sources the transit to Mars in less time than 3 months each for exploring, making outposts on, and terraforming way, compared to the 12 to 16 months, as it is now.
    [Show full text]
  • Space Propulsion Technology for Small Spacecraft
    Space Propulsion Technology for Small Spacecraft The MIT Faculty has made this article openly available. Please share how this access benefits you. Your story matters. Citation Krejci, David, and Paulo Lozano. “Space Propulsion Technology for Small Spacecraft.” Proceedings of the IEEE, vol. 106, no. 3, Mar. 2018, pp. 362–78. As Published http://dx.doi.org/10.1109/JPROC.2017.2778747 Publisher Institute of Electrical and Electronics Engineers (IEEE) Version Author's final manuscript Citable link http://hdl.handle.net/1721.1/114401 Terms of Use Creative Commons Attribution-Noncommercial-Share Alike Detailed Terms http://creativecommons.org/licenses/by-nc-sa/4.0/ PROCC. OF THE IEEE, VOL. 106, NO. 3, MARCH 2018 362 Space Propulsion Technology for Small Spacecraft David Krejci and Paulo Lozano Abstract—As small satellites become more popular and capa- While designations for different satellite classes have been ble, strategies to provide in-space propulsion increase in impor- somehow ambiguous, a system mass based characterization tance. Applications range from orbital changes and maintenance, approach will be used in this work, in which the term ’Small attitude control and desaturation of reaction wheels to drag com- satellites’ will refer to satellites with total masses below pensation and de-orbit at spacecraft end-of-life. Space propulsion 500kg, with ’Nanosatellites’ for systems ranging from 1- can be enabled by chemical or electric means, each having 10kg, ’Picosatellites’ with masses between 0.1-1kg and ’Fem- different performance and scalability properties. The purpose tosatellites’ for spacecrafts below 0.1kg. In this category, the of this review is to describe the working principles of space popular Cubesat standard [13] will therefore be characterized propulsion technologies proposed so far for small spacecraft.
    [Show full text]
  • Feasibility and Performance of the Microwave Thermal Rocket Launcher
    Feasibility and Performance of the Microwave Thermal Rocket Launcher Kevin L.G. Parkin, Fred E.C. Culick Division of Engineering and Applied Science, California Institute of Technology, 1200 East California Boulevard, Pasadena, CA 91125, USA Abstract. Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit (SSTO) microwave thermal rocket. We present an SSTO concept employing a scaled X-33 aeroshell. The flat aeroshell underside is covered by a thin-layer microwave absorbent heat-exchanger that forms part of the thruster. During ascent, the heat-exchanger faces the microwave beam. A simple ascent trajectory analysis incorporating X-33 aerodynamic data predicts a 10% payload fraction for a 1 ton craft of this type. In contrast, the Saturn V had 3 non-reusable stages and achieved a payload fraction of 4%. INTRODUCTION In the year 2003, payloads are launched into orbit the same way they were in 1963: by chemical rockets. Traditional expendable multistage rockets usually achieve payload fractions of less than 5%. As described by the rocket equation, this is due partly to the structural limits of existing materials, and partly to the limited specific impulse (Isp) of chemical propellants, which have reached a practical limit of 480 seconds. The structural economies made to preserve these minute payload fractions result in fragile rockets that are expensive to build.
    [Show full text]
  • The Spacedrive Project – First Results on Emdrive and Mach-Effect Thrusters BARCELO RENACIMIENTO HOTEL, SEVILLE, SPAIN / 14 – 18 MAY 2018
    SP2018_016 The SpaceDrive Project – First Results on EMDrive and Mach-Effect Thrusters BARCELO RENACIMIENTO HOTEL, SEVILLE, SPAIN / 14 – 18 MAY 2018 Martin Tajmar(1), Matthias Kößling(2), Marcel Weikert(3) and Maxime Monette(4) (1-4) Institute of Aerospace Engineering, Technische Universität Dresden, Marschnerstrasse 32, 01324 Dresden, Germany, Email: [email protected] KEYWORDS: Breakthrough Propulsion, Propellant- Recent efforts therefore concentrate on using less Propulsion, EMDrive, Mach-Effect Thruster propellantless laser propulsion. For example, the proposed Breakthrough Starshot project plans to use ABSTRACT: a 100 GW laser beam to accelerate a nano- Propellantless propulsion is believed to be the spacecraft with the mass of a few grams to reach our best option for interstellar travel. However, photon closest neighbouring star Proxima Centauri in around rockets or solar sails have thrusts so low that maybe 20 years [2]. The technical challenges (laser power, only nano-scaled spacecraft may reach the next star steering, communication, etc.) are enormous but within our lifetime using very high-power laser maybe not impossible [3]. Such ideas stretch the beams. Following into the footsteps of earlier edge of our current technology. However, it is breakthrough propulsion programs, we are obvious that we need a radically new approach if we investigating different concepts based on non- ever want to achieve interstellar flight with spacecraft classical/revolutionary propulsion ideas that claim to in size similar to the ones that we use today. In the be at least an order of magnitude more efficient in 1990s, NASA started its Breakthrough Propulsion producing thrust compared to photon rockets. Our Physics Program, which organized workshops, intention is to develop an excellent research conferences and funded multiple projects to look for infrastructure to test new ideas and measure thrusts high-risk/high-payoff ideas [4].
    [Show full text]
  • Space Exploration and Development Systems (SEEDS) 2014
    SpacE Exploration and Development Systems (SEEDS) 2014 “Missions and Expeditions to Mars” Orpheus Executive Summary SEEDS Executive Summary 09/2014 Page 1 SEEDS Executive Summary 09/2014 Page 2 List of authors: Crescenzio Ruben Xavier AMENDOLA Portia BOWMAN Samuel BROCKSOPP Andrea D’OTTAVIO Alex GEE Samuel R. F. KENNEDY Antonio MAGARIELLO Adrian MORA BOLUDA Ignacio REY Alex ROSENBAUM Joachim STRENGE Aurthur Vimalachandran THOMAS JAYACHANDRAN SEEDS Executive Summary 09/2014 Page 3 SEEDS Executive Summary 09/2014 Page 4 ABSTRACT This six crew mission called Orpheus has been designed in order to fulfil some main objectives of space exploration: scientific advancements, technological progress, public outreach and international cooperation. This paper investigates the possibility of exploring the Martian system by using a manned spacecraft, the Crew Interplanetary Vehicle, (CIV) and a cargo vehicle, the Mars Automated Transfer Vehicle (MATV). The main payloads, carried by the high efficiency solar electric MATV, are a two-passenger spacecraft landing on Phobos; an orbital laboratory and a rover network for deployment to the surface of Mars. In order to cope with the constraints imposed for a human mission to deep space, the feasibility study has been performed using a human-centred design approach. The main output of the mission is the preliminary design of the CIV. Furthermore, the main parameters of the MATV as well as the orbital laboratory and the Phobos Lander were estimated. The manned spacecraft is designed to depart from LEO in 2036 using chemical propulsion. Once in Mars proximity, the main manoeuvres will be performed using nuclear thermal propulsion and a bi-propellant chemical system will perform the minor manoeuvres.
    [Show full text]