DOCSLIB.ORG

Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies- a Preliminary Assessment

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies- a Preliminary Assessment //1/--.:20 NASA Technical Memorandum 107095 5:3d-/ AIAA-95-2631 f lF' Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies­ A Preliminary Assessment Stanley K. Borowski Lewis Research Center Cleveland, Ohio Prepared for the 31st Joint Propulsion Conference and Exhibit cosponsored by AIAA, ASME, SAE, and ASEE San Diego, California, July 10-12, 1995 (NASA-TM-I07095) HU~AN LUNAR . --.................. N96-12353 ~... MISSION CAPA ILITIES USING SSTO ..... ~. ~~. ISRU AND lOX-AUGMENTED NTR • TECHNClOGIES: A PRELr~I ARY .' Unc)as National• Aeronautics and ASSESSMENT (NASA. Lewis Research Space Administration Center-) 15 p G3/20 0011561 HUMAN LUNAR MISSION CAPABILITIES USING SSTO, ISRU AND LOX-AUGMENTED NTR TECHNOLOGIES --A PRELIMINARY ASSESSMENT Stanley K. Borowski * NASA Lewis Research Center Cleveland, OH 44135 ABSTRACT The feasibility of conducting human missions rich" mare soils and impact / volcanic glass-- to to the Moon is examined assuming the use of three produce LUNOX will also be key to minimizing "high leverage" technologies: (1) a single-stage-to­ IMLEO, the size and number of LTS elements, and orbit (SSTO) launch vehicle, (2) .. in-situ" resource the cost for each lunar mission. At present, NASA is utilization (ISRU)--specifically "lunar-derived" liquid focusing considerable resources on the construction oxygen (LUNOX), and (3) LOX-augmented nuclear of the international space station (ISS) and on thermal rocket (LANTR) propulsion. Lunar development and demonstration of a reusable, transportation system elements consisting of a single-stage-to-orbit (SSTO) vehicle1 to replace the LANTR-powered lunar transfer vehicle (LTV) and a current Space Shuttle in the post-2005 timeframe. It chemical propulsion lunar landing / Earth return is envisioned that an industry-developed and­ vehicle (LERV) are configured to fit within the operated SSTO would have an average flight rate of "compact" dimensions of the SSTO cargo bay -40 missions per year with -50% of these flights (diameter: 4.6 m/length: 9.0 m) while satisfying an procured by NASA to provide logistics resupply to initial mass in low Earth orbit (IMLEO) limit of -60 t the ISS. (3 SSTO launches). Using -8 t of LUNOX to "reoxidize" the LERV for a "direct return" flight to Planning future human missions to the Moon Earth reduces its size and mass allowing delivery to and Mars is also beginning anew--albeit in a less LEO on a single 20 t SSTO launch. Similarly, the dramatic fashion than the previous Space LANTR engine's ability to operate at any Exploration Initiative2--in response to the "Human oxygen/hydrogen mixture ratio from 0 to 7 with high Exploration and Development of Space" (HEDS) specific impulse (-940 to 515 s) is exploited to component of NASA's just released strategic plan.3 reduce hydrogen tank volume, thereby improving Activities are focusing on the development of key packaging of the LANTR LTV's "propulsion" and "high leverage" technologies and systems, their "propellant modules". Expendable and reusable, demonstration on precursor missions, and on the piloted and cargo missions and vehicle designs are identification of viable mission architectures that presented along with estimates of LUNOX production could be supported by NASA in the future given the required to support the different mission modes. expected tight budgets and the trends toward Concluding remarks address the issue of lunar commercialization. transportation system costs from the launch vehicle pe rspective. This paper addresses the feasibility of conducting human lunar missions using SSTO, INTRODUCTION ISRU and LOX-augmented nuclear thermal rocket (lANTR) technologies/systems. For " access to Future human exploration missions to the space," the lure of the SSTO is its full reusability, Moon will require an "economical" lunar improved operability, high flight rate and reduced transportation system (L TS) providing efficient launch costs estimated at -40 M$ per flight. 4 Its lift "access to and through space: Utilizing locally capability is limited, however, and varies from -25 available resources--specifically "iron oxide (FeO)- klbm (-11 .3 t) to 45 klbm (-20.4 t) depending on the *Ph .D.lNuclear Eng ineering, Senior Member AIAA low Earth orbit (LEO) altitude and its inclination. The SSTO, LUNOX, AND LANTR TECHNOLOGY SSTO's cargo bay volume is also small (-50 to 75% DESCRIPTION that of the current Space Shuttle) which makes packaging of payload and LTS elements a particular The SSTO Reusable Launch Vehicle challenge. In January 1993, NASA initiated a The development of LUNOX production, comprehensive "in-house" study called "Access to storage and tanker transport / transfer technologies Space," The goals of the study were to identify and systems for "reoxidizing" expendable lunar attractive options for the next generation of U.S. landing/Earth return vehicles (LERVs) initially, and launch vehicles which would : (1) allow major then reusable, lunar landing and transfer vehicles reductions in the cost of space transportation (by at (LLVs and LTVs) is expected to dramatically reduce least 50%); (2) provide an order of magnitude IMLEO requirements, the number of SSTO launches increase in crew flight safety; and (3) substantially and the cost of delivering payload to the lunar improve overall system operability. The study surface. For example, with -7 kilograms of mass examined three major architectural options which required in LEO for each kilogram of mass landed on included: retaining and upgrading the current Space the lunar surface, the ability to reoxidize an LERV Shuttle and expendable launch vehicle fleet with -8 tons of LUNOX for a "direct Earth return" (Option 1); developing new expendable vehicles mission would save three 20 t-class SSTO using conventional technologies (Option 2); and launches compared to an "all Earth-supplied" developing new reusable vehicles using advanced mission requiring twice that number. technology (Option 3). A conservative "launch needs" mission model was defined and used in the To improve "access through space," a LANTR­ study.4 It included all U.S. civilian (NASA), defense based LTV mission scenario has been assumed. and commercial missions anticipated during the The ability to alter the engine's thrust and specific period 1995 through 2030, but did not consider impulse (/sp) by varying the liquid oxygen-to-liquid future human missions to the Moon and Mars--a hydrogen (LOXlLH2) mixture ratio (MR) in LANTR's major omjssion. "LOX-afterburner" nozzle leads to small engines (both in terms of physical size and reactor power In its final summary report4 issued in January output) capable of providing "big engine" 1994, NASA concluded that a fully reusable, pure­ performance. Supersonic combustion of LOX with rocket SSTO (SSTO-R) launch vehicle was the most reactor preheated hydrogen emerging from the attractive/beneficial option for development. The engine's sonic throat also results in higher Isp values preferred Option 3 reference vehicle described in the than LOX/LH2 chemical engines operating at the "Access to Space Study" Summary Report4 is shown same mixture ratio (-100 s at MR = 6) . Lastly, the in Figure 1. It is a vertical takeoff, horizontal-landing increased use of high-density LOX in place of low­ winged concept with a circular cross-section density LH2 reduces hydrogen tank volume and fuselage for structural efficiency (other configurations improves LTV component packaging within the are presently being examined under NASA­ "volume-constrained" SSTO cargo bay. supported industry contracts). It uses aluminum­ lithium (Al/Li) cryotanks, graphite composite material for all non-pressurized primary structure and a This paper describes results of preliminary tripropellant (LOX/LH2/RP) propulsion system system and mission analysis conducted by Lewis employing seven RD-704 engines. The payload bay Research Center for NASA Headquarters' Advanced is located horizontally between the forward LOX tank Concepts Office over the past two months. The and the LH2 tank positioned in the vehicle's paper first discusses the relevant technologies and midsection. The crew cabin is situated on top of the describes their characteristics. Mission and vehicle next to an airlock/work-station area which transportation system ground rules and assumptions provides crew access to both the payload bay and are then presented along with a description of the the ISS through an overhead hatch. A "LOXlLH2" reference lunar mission scenario. Next, a version of the SSTO-R would have an overall length comparison of different LTS element options--both and diameter -15% larger than its tripropellant for the LERV and L TV--is provided and the counterpart and a gross liftoff weight (GLOW) of requirements/performance gains from transitioning -2.48 million pounds (-1125 t) . from an expendable to reusable LTS architecture are discussed. The paper concludes with a discussion of The reference SSTO-R cargo bay is 15 feet lunar transportation system and launch vehicle costs. (-4.6 m) in diameter and 30 feet (-9.15 m) long. A 2 c ___________________________________________________________ r --- -- ------_._--- l'lu:lluuJ.' fllib.tJ:.w!: fini.l..Qr!2i1 2Sklbm None Station: 220'nmi cire@ 51.6" IS.Sklbm 2 Person Station: 220 Ilmi eire @ 51.6" 45klbm . No"" 100 nmi eire @ 28.5· 3H lbm None 220 nmi eire @ 28.5· 24 klbm No .. 100 nmi chi: @900 · 'ndudes Payload and ASE Vehicle: GLOW: 1.96 Mib Dry Mass: 159.51db Propellant Mass: 1.74 Mlb CltXXi Propellant Type: LoxllHiRP Main Engine Type/No.: RD-704!7 Vac Thrust (ea.) Mode 1/2: 441.4/175.5 klb Var. ISP Mode 1/2: 4071452 10( 150ft-----J~~1 Area Ratio: 74:1 OMS Engine Type/No ~ New LoxllH2 Notes: • Vertical takeofflhorizorrtaJ landing Vat; Thrust (ea.): 6 klbf • 15 percent dry weight margin Vat; Isp: 462s • Option for carrying two on-orbit opel1i!ions Area Ratio: 100:1 crew for 7 days CryoTanks A1-U • Option for carrying four 5.5. Freedom rotation Primary Structure: Gr-COmposite crew in payload bay Control Surfar:.es: ACC • Option for carrying high energy transfer stage TPS: ACC/TABIIAFRSI for Atlas class (5k Ib) GEO missions Payload Bay-Usable Volume 15ftDx30ftL Fig .
Load more

Recommended publications
  • 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]
  • 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]
  • 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]
  • 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]
  • 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]
  • (NTR) Artificial Gravity Mars Transfer Vehicle Concepts
    AIAA-2014-3623 Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts Stanley K. Borowski1, David R. McCurdy2 and Thomas W. Packard2 NASA Glenn Research Center, Cleveland, OH, 44135 A variety of countermeasures have been developed to address the debilitating physiological effects of “zero-gravity” (0-g) experienced by cosmonauts and astronauts during their ~0.5 – 1.2 year long stays in LEO. Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to ~2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity (AG) Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced “bimodal” nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (Isp) capability of ~900 s – twice that of today’s best chemical rockets. The AG/MTV concepts using conventional NTP carry twin cylindrical “ISS-type” habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules.
    [Show full text]
  • Two-Stage-To-Orbit Transporting System Combining Microwave Rocket and Microwave Thermal Rocket for Small Satellite Launch
    Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30, pp. Pb_99-Pb_103, 2016 Two-Stage-to-Orbit Transporting System Combining Microwave Rocket and Microwave Thermal Rocket for Small Satellite Launch By Kaoru KAKINUMA,1) Masafumi FUKUNARI,1) Toshikazu YAMAGUCHI,1) Yusuke NAKAMURA,1) Hiroyuki KOIZUMI,1) Kimiya KOMURASAKI1) and Kevin PARKIN2) 1) The University of Tokyo, Kashiwa, Chiba, Japan 2) Carnegie Mellon University, Silicon Valley Campus, California, USA (Received August 1st, 2015) This paper proposes a two-stage-to-orbit launch system comprised of a Microwave Rocket 1st stage and a Microwave Thermal Rocket 2nd stage. The air-breathing 1st stage improves payload fraction relative to a single-stage-to-orbit system and carries the 2nd stage above the atmosphere and into range of its beam director. For the 1st stage task, a Microwave Rocket is superior to an unmanned aerial vehicle because it is simpler, faster, and reaches higher altitude at higher speed. In addition, we present a new trajectory that eliminates power beaming at low elevation angles and improves system performance. This combination of factors reduces the propellant needed in the 2nd stage, which in turn increases payload fraction by a remarkable factor of 3 times. Key Words: Beamed Energy Propulsion, Gyrotron, Microwave, Wireless Power Transfer Nomenclature problems of this approach have been pointed out, the most critical one is the cost. The launch cost remains around $ 10,000 A : cross sectional area per kilogram of payload, and a breakthrough of further cost CD : drag coefficient reduction cannot be detected with the chemical rocket at present. D : drag The specific impulse of conventional chemical rockets is about m : rocket mass 450 s at maximum, resulting in low payload fraction of about p : pressure 3 %.
    [Show full text]
  • Movement and Maneuver in Deep Space: a Framework to Leverage Advanced Propulsion
    Movement and Maneuver in Deep Space A Framework to Leverage Advanced Propulsion Brian E. Hans, Major, USAF Christopher D. Jefferson, Major, USAF Joshua M. Wehrle, Major, USAF Air University Steven L. Kwast, Lieutenant General, Commander and President Air Command and Staff College Thomas H. Deale, Brigadier General, Commandant Bart R. Kessler, PhD, Dean of Distance Learning Robert J. Smith, Jr., Colonel, PhD, Dean of Resident Programs Michelle E. Ewy, Lieutenant Colonel, PhD, Director of Research Liza D. Dillard, Major, Series Editor Peter Garretson, Lieutenant Colonel, Essay Advisor Selection Committee Kristopher J. Kripchak, Major Michael K. Hills, Lieutenant Colonel, PhD Barbara Salera, PhD Jonathan K. Zartman, PhD Please send inquiries or comments to Editor The Wright Flyer Papers Department of Research and Publications (ACSC/DER) Air Command and Staff College 225 Chennault Circle, Bldg. 1402 Maxwell AFB AL 36112-6426 Tel: (334) 953-3558 Fax: (334) 953-2269 E-mail: [email protected] AIR UNIVERSITY AIR COMMAND AND STAFF COLLEGE MOVEMENT AND MANEUVER IN DEEP SPACE A FRAMEWORK TO LEVERAGE ADVANCED PROPULSION Brian E. Hans, Major, USAF Christopher D. Jefferson, Major, USAF Joshua M. Wehrle, Major, USAF Wright Flyer Paper No. 67 Air University Press Curtis E. LeMay Center for Doctrine Development and Education Maxwell Air Force Base, Alabama Accepted by Air University Press April 2017 and published May 2019. Project Editor Dr. Stephanie Havron Rollins Copy Editor Carolyn B. Underwood Cover Art, Book Design, and Illustrations Leslie Fair Composition and Prepress Production Megan N. Hoehn AIR UNIVERSITY PRESS Director, Air University Press Lt Col Darin Gregg Air University Press Disclaimer 600 Chennault Circle, Building 1405 Maxwell AFB, AL 36112-6010 The views expressed in this academic research paper are those of https://www.airuniversity.af.edu/AUPress/ the author and do not reflect the official policy or position of the US government or the Department of Defense.
    [Show full text]
  • Nuclear Thermal Rocket Simulation in NPSS
    Nuclear Thermal Rocket Simulation in NPSS Michael L. Belair,∗ Charles J. Sarmiento,y and Thomas M. Lavellez NASA Glenn Research Center, Cleveland, OH, 44135 Four nuclear thermal rocket (NTR) models have been created in the Numerical Propul- sion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design. Nomenclature AR Area Ratio CERMET Ceramic-Metallic fuel element DRA Design Reference Architecture g Acceleration of gravity at sea level γ Ratio of specific heats GRC Glenn Research Center Isp Specific Impulse kN kilownewtons force lbf pounds force MCNP Monte Carlo N-Particle transport code MW Megawatts MW Molecular
    [Show full text]
  • Nuclear Thermal Propulsion Systems (Last Updated in January 2021) Eric PROUST Eric Proust
    Lecture Series on NUCLEAR SPACE POWER & PROPULSION SYSTEMS -2- Nuclear Thermal Propulsion Systems (Last updated in January 2021) Eric PROUST Eric Proust To cite this version: Eric Proust. Lecture Series on NUCLEAR SPACE POWER & PROPULSION SYSTEMS -2- Nuclear Thermal Propulsion Systems (Last updated in January 2021) Eric PROUST. Engineering school. France. 2021. hal-03147500 HAL Id: hal-03147500 https://hal.archives-ouvertes.fr/hal-03147500 Submitted on 19 Feb 2021 HAL is a multi-disciplinary open access L’archive ouverte pluridisciplinaire HAL, est archive for the deposit and dissemination of sci- destinée au dépôt et à la diffusion de documents entific research documents, whether they are pub- scientifiques de niveau recherche, publiés ou non, lished or not. The documents may come from émanant des établissements d’enseignement et de teaching and research institutions in France or recherche français ou étrangers, des laboratoires abroad, or from public or private research centers. publics ou privés. FROM RESEARCH TO INDUSTRY LECTURE SERIES ON NUCLEAR SPACE POWER & PROPULSION SYSTEMS -2- Nuclear Thermal Propulsion Systems Eric PROUST Commissariat à l’énergie atomique et aux énergies alternatives - www.cea.fr Last update: January 2021 Lecture Series on SPACE NUCLEAR POWER & PROPULSION SYSTEMS -2- Nuclear Thermal Propulsion Systems (last updated in January 2021) Eric PROUST Nuclear Space Power & Propulsion in the last 2 month news Nuclear Thermal Propulsion Nuclear Thermal Propulsion Nuclear Electric Propulsion Space Nuclear Power Reactor
    [Show full text]
  • Liquid Oxygen Augmented Gas Core Nuclear Thermal Rocket
    Advances in Aerospace Science and Applications. ISSN 2277-3223 Volume 3, Number 3 (2013), pp. 239-244 © Research India Publications http://www.ripublication.com/aasa.htm Liquid Oxygen Augmented Gas Core Nuclear Thermal Rocket V. Krishnamurthy1 and A. Vignesh2 1,2Department of Aeronautical Engineering Rajalakshmi Engineering College Chennai-602105, Tamil Nadu, India. Abstract Conventional propulsion technology (chemical and electric) currently limits the possibilities for human space exploration to the neighborhood of the Earth. If farther destinations (such as Mars) are to be reached with humans on board, a more capable interplanetary transfer engine featuring high thrust and high specific impulse is required. The source of energy which could in principle best meet these engine requirements is nuclear thermal energy. So an innovative gas core nuclear thermal rocket concept which combines conventional liquid hydrogen cooled nuclear thermal rocket and supersonic combustion ramjet (scramjet) technologies is described. Known as the Liquid oxygen augmented Gas core Nuclear Thermal Rocket (LAG- NTR), this concept utilizes the large divergent section of the gas core nuclear thermal rocket’s nozzle as an afterburner into which liquid oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the LAG-NTR’s choked sonic throat- scramjet propulsion in reverse. By varying the oxygen-to- hydrogen mixture ratio, the LAG-NTR can operate over a wide range of thrust and specific impulse values while the gaseous core reactor’s power level remains relatively constant. This thrust augmentation feature means that a larger engine performance can be obtained with a comparatively smaller LAG-NTR engine. Keywords: Nuclear thermal rocket, gas core, afterburner, scramjet, thrust, specific impulse, supersonic combustion.
    [Show full text]
  • Effects and Solutions on the Human Body After Long-Duration Space Flights
    McNair Research Journal SJSU Volume 15 Spring 2019 Article 9 2019 Effects and Solutions on the Human Body After Long-Duration Space Flights Jose Jaime Esquivel Patricio San Jose State University Follow this and additional works at: https://scholarworks.sjsu.edu/mcnair Part of the Other Aerospace Engineering Commons Recommended Citation Esquivel Patricio, Jose Jaime (2019) "Effects and Solutions on the Human Body After Long-Duration Space Flights," McNair Research Journal SJSU: Vol. 15 , Article 9. Available at: https://scholarworks.sjsu.edu/mcnair/vol15/iss1/9 This Article is brought to you for free and open access by SJSU ScholarWorks. It has been accepted for inclusion in McNair Research Journal SJSU by an authorized editor of SJSU ScholarWorks. For more information, please contact [email protected]. Esquivel Patricio: Effects and Solutions on the Human Body After Long-Duration Space Biography Jaime is a first-generation student pursuing a B.S. in Aerospace Engineering with a Minor in Astronomy. His research focuses on the effects of microgravity on astronauts in order to find solutions for long-duration space flights to Mars. After graduating, he plans to attend graduate school to continue his dream of making humans an interplanetary species. He believes the purpose of engineering goes beyond making people’s lives easier. When used sensibly, human history demonstrates engineering can propel humanity into unimaginable new boundaries. He Jose Jaime Esquivel developed his mission to promote higher education among Patricio underrepresented groups in his community and advocate for more Major: students to pursue a career in Aerospace Engineering STEM. His belief is that if we only have one life; to make our lives Mentor: more meaningful, we should strive Taylor-Dawn Francis to make this world better and to serve the community.
    [Show full text]