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RUSSIAN ACADEMY OF COSMONAUTICS NAMED AFTER K.E. TSIOLKOVSKY

MANNED MISSION

TO MARS Dedicated to 50 anniversary of the irst satellite launch th

Moscow-Korolyov 2007

UDK 629.788:523.43 BBK 39.67 P32

Editorial Board: Editor-in-Chief: Academician of the RAS, A.S. Koroteyev Deputy Editors-in-Chief: N.N. Sevastianov, L.A. Gorshkov, V.F. Semyonov Academician of the RAS, A.I. Grigoriev, Corresponding Member of the RAS, L.M. Zelyoniy; N.M. Ivanov, A.N. Potapov, V.P. Smetannikov Authors: R.M. Abdulkhalikov, A.A. Adov, V.N. Akimov, P.O. Andreichuk, P.V. Andreyev, A.N. Astakhov, G.B. Astashev, R.I. Beglov, M.A. Bek, V.S. Belyakov, L.A. Besedina, L.V. Bobrysheva, A.N. Bogachov, I.B. Braverman, N.A. Brukhanov, V.S. Vasilkovskiy, I.N. Gansvindt, A.D. Egorov, O.I. Egorova, I.O. Eliseyev, U.A. Gashkov, A.N. Glukhov, I.A. Golov, L.M. Zelyoniy, A.S. Zernov, N.M. Ivanov, M.N. Kazakov, S.S. Klimov, I.B. Kozlovskaya, N.V Kolesnik, U.F Kolyuka, G.M Komarova, A.S. Koroteyev, V.I. Lukiashchenko, A.N. Krylov, D.V. Kutkin, V.I. Kucherenko, I.F. Lendrasova, V.M. Linkin, O.N. Logachov, M. A. Levinskikh, N.V. Maksimovskiy, M.I. Malenkov, N.G. Medvedev, A.I. Mezentsev, M.V. Mikhailov, V.A. Muravlyov, N.F. Moiseyev, A.A. Nesterenko, V.M. Nesterov, N.D. Novikova, S.N. Obukhov, O. I. Orlov, V.A. Pavshuk, V.M. Petrov, L.I. Podolskaya, N.B. Ponomaryov, A.N. Potapov, O.F. Prilutskiy, K.V. Psyanin, N.N. Ponomaryov-Stepnoy, V.G. Rodin, A.N. Rumynskiy, E.L. Romadova, S.U. Romanov, T.I. Rozhkova, V.P. Salnitskiy, N.N. Sevastyanov, V.F. Semyonov, U.P. Semyonov, A.V Semyonkin, U.E. Sinyak, L.D. Skotnikova, V.P Smetannikov, V.V. Suvorov, V.G Sobolevskiy, B.I. Sotnikov, S.I. Stepanova, S.F. Stoyko, O.G. Sytin, V.N. Sychov, S.O. Tverdokhlebov, E.V Timofeyeva, V.A. Usov, G.N. Ustinov, I.I. Fedik, A.I. Fedosova, I.I. Khamits, V.V. Tsvetkov, O.S. Tsigankov, A.G. Chernyavskiy, M.A. Shutikov, A.G. Yakushev, S.V. Yaroshenko

Manned Mission to Mars /Edited by A.S. Koroteyev. – M.: Russian Academy of Cosmonautics named after K.E. Tsiolkovsky, 2006, 320 pages, illustrated his work is the first systematic presentation of various concepts and projects of a manned mission to Mars, including the latest Russian project of a manned expedition to Mars – ‘MEC’. he book highlights the key challenges of the mission, including medical issues, and offers solutions based on the Soviet, Russian and international experience in the field of space exploration. It also shows how the development of Mars-targeted technologies, including those implementing nuclear energy, has influenced the development of lunar base and efficient space transport system projects. Some of the authors have been investigating the problem of a trip to Mars for about 50 years. hey regard this book as an opportunity to share their experience. he book may be of interest to rocket-and-space specialists, undergraduate and graduate students majoring in relevant academic subjects and to the broad reading public interested in the history and the future of space research and exploration. ISBN 5-9900783-1-5

© Russian Academy of Cosmonautics named after K.E. Tsiolkovsky

FOREWORD Mars has been the subject of human fascination since ancient times. he planetary processes, which we can observe on Mars, have a lot in common with those on our own planet. herefore, the study of Mars helps to reveal the laws, under which the observed processes progress, and enables the scientists to make a reliable forecast of their evolution on Earth. Exploration of Mars is a challenging task, which demands participation of all developed countries in possession of modern technologies. Russia has accumulated a huge intellectual and technological potential for the development and implementation of a man-controlled Mars-bound mission. Our leading research institutes and design bureaus are currently working at the project of a manned mission to Mars within the framework of the Federal Space Program. At this initial stage, their task consists in defining the general profile of the mission and some of its specific features considering the Russian scientific tradition and the existing national research and technology potential Recently, a manned Mars mission has started to turn into a reality, in a great measure due to the achievements connected with the operation of space stations, which can play an important part in the preparation of a long-term space mission. hus, they can be used to carry out research aimed at the prolongation of the equipment service life, solution of the problems connected with its maintainability and handling of open-space operations. Besides, they will be used to continue investigation of the possibility to extend the duration of living in the gravity-free environment, and to develop and optimize technologies, systems and assemblages to be used in future inter-planetary complexes. his book is a summary of space exploration activities, which have been carried out in Russia for over half a century. It was written by a team of leading specialists and is sure to arouse keen interest in Russia and worldwide. A.N. PERMINOV Head of the Russian Federal Space Agency

TABLE OF CONTENTS FOREWORD ……………………………………………………………………………………………………………3 LIST OF COMMON ABBREVIATIONS ………………………………………………………………………7 INTRODUCTION ……………………………………………………………………………………………………9 Chapter 1. Status and Mars exploration streamlines ……………………………………………………… 1.1. Introduction ………………………………………………………………………………………………… 1.2. Current state of exploration ……………………………………………………………………………… 1.3. Prospective Mars exploration program ……………………………………………………………… 1.4. Predecessors of the manned mission ………………………………………………………………… 1.5. Conclusions ………………………………………………………………………………………………… 1.6. List of References ………………………………………………………………………………………… 1.7. Annex 1. Mars as compared to Earth ………………………………………………………………… 1.8. List of References …………………………………………………………………………………………

10 10 10 22 27 29 30 35 41

Chapter 2. Historical Overview of the Manned Mars Mission: Concepts, Projects and Programs ……………………………………………………………………………… 2.1. Overview of Manned Mars Mission Key Concepts ………………………………………………… 2.2. Evolution of Russian Mars mission project …………………………………………………………… 2.3. Conclusions ………………………………………………………………………………………………… 2.4. List of References …………………………………………………………………………………………

42 42 45 48 48

Chapter 3. Interplanetary Mars Mission Complex ………………………………………………………… 3.1. Challenges of a Mars Mission …………………………………………………………………………… 3.2. Alternative Manned Martian Mission Concepts. Conceptual Decisions ……………………… 3.3. Ballistic Design of the Martian Mission ……………………………………………………………… 3.4. Conclusions ………………………………………………………………………………………………… 3.5. List of References …………………………………………………………………………………………

49 49 50 67 85 85

Chapter 4. Interplanetary Orbital Vehicle …………………………………………………………………… 86 4.1. General Design Requirements and Composition of an Interplanetary Orbital Vehicle ……… 86 4.2. Design and Configuration ……………………………………………………………………………… 87 4.3. Interplanetary Orbital Vehicle Onboard Systems …………………………………………………… 91 4.4. Conclusions ……………………………………………………………………………………………… 106 4.5. List of References ……………………………………………………………………………………… 106

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Chapter 5. Power and Propulsion System ………………………………………………………………… 5.1. Background ……………………………………………………………………………………………… 5.2. LPE Version ……………………………………………………………………………………………… 5.3. Nuclear-Powered Propulsion Version ……………………………………………………………… 5.4. Version of Propulsion System Based on Solar Cell Arrays and Electric Propulsion ……… 5.5. Version of Combined Propulsion System Based on Cell Solar Arrays, Electric Propulsion and Liquid-Propellant Rocket Engines ………………………………………… 5.6. Conclusions ……………………………………………………………………………………………… 5.7. List of References ………………………………………………………………………………………

107 107 110 113 137 166 166 167

Chapter 6. Martian Ascent-Descent Vehicle ……………………………………………………………… 6.1. Designation and Configuration of the Ascent-Descent Vehicle ……………………………… 6.2. Aerothermoballistic Design of the Ascent-Descent Vehicle …………………………………… 6.3. Descent Module (DM) ………………………………………………………………………………… 6.4. Ascent Module (AM) ………………………………………………………………………………… 6.5. Living Module (LM) …………………………………………………………………………………… 6.6. Main Characteristics of the Ascent-Descent Vehicle …………………………………………… 6.7. Conclusions ……………………………………………………………………………………………… 6.8. List of References ………………………………………………………………………………………

170 170 173 187 188 192 193 195 195

Chapter 7. Crew Return Vehicle……………………………………………………………………………… 7.1. Designation ……………………………………………………………………………………………… 7.3. Configuration and Design …………………………………………………………………………… 7.4. ‘Soyuz’-Based Version of the Crew Return Vehicle ……………………………………………… 7.5. Conclusions ……………………………………………………………………………………………… 7.6. List of References ………………………………………………………………………………………

196 196 199 201 202 202

Chapter 8. Deployment of an Interplanetary Mission Complex on the Earth Orbit. Martian Mission Program …………………………………………………………………………………… 8.1. Assembly of the Interplanetary Mission Complex ……………………………………………… 8.2. he Interplanetary Mission Complex Flight to Mars and its Return to Earth ……………… 8.3. Conclusions ……………………………………………………………………………………………… 8.4. List of References ………………………………………………………………………………………

203 203 204 207 207

Chapter 9. Martian Mission Equipment Flight-Development Tests ……………………………… 9.1. Flight Tests Conducted in Russia for the Purposes of the Martian Mission ………………… 9.2. Flight development tests of the Martian mission complex systems, units and modules … 9.3. Conclusions ……………………………………………………………………………………………… 9.4. List of References ………………………………………………………………………………………

208 208 211 215 215

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Chapter 10. Martian base. Planetary facilities ………………………………………………………… 10.1. Purpose and Structure ……………………………………………………………………………… 10.2. Living Complex ……………………………………………………………………………………… 10.3. Energy Complex ……………………………………………………………………………………… 10.4. Transport and Technological Complex …………………………………………………………… 10.5. Conclusions …………………………………………………………………………………………… 10.6. List of References ………………………………………………………………………………………

216 216 220 221 228 233 233

Chapter 11. Implementation of the Martian Mission Technologies for the Exploration of the Moon …………………………………………………………………………… 11.1. Possible Scenario of Moon Exploration for 3Не Production ………………………………… 11.2. Possible Scenario of Moon Exploration for Oxygen Production …………………………… 11.3. Possible Stages of the Moon Base Development ……………………………………………… 11.4. Conclusions …………………………………………………………………………………………… 11.5. List of References ………………………………………………………………………………………

236 236 240 242 244 244

Chapter 12.Biomedical maintenance of space missions. ……………………………………………… 12.1. Factors and conditions of manned missions. …………………………………………………… 12.2. Objectives and structure of biomedical maintenance of the mission ……………………… 12.3. Medical maintenance of the mission ……………………………………………………………… 12.4. Psychological Maintenance of the Mission ……………………………………………………… 12.5. Life support issues of the Martian mission crew ………………………………………………… 12.6. A green house on board of the first manned vehicle to Mars ………………………………… 12.7. Microbiological safety maintenance of the mission …………………………………………… 12.8. Radiation safety maintenance of the mission …………………………………………………… 12.9. Earth-based simulation tests ……………………………………………………………………… 12.10. Conclusions …………………………………………………………………………………………… 12.11. List of References ……………………………………………………………………………………

245 245 247 248 261 270 272 277 284 294 299 300

Chapter 13. Concept of a Space Transport System ……………………………………………………… 13.1. Designation …………………………………………………………………………………………… 13.2. Basic Requirements …………………………………………………………………………………… 13.3 Appearance and Specifications of an Expendable Launch-Vehicle of ‘Angara’ Type ……… 13.4. Appearance and Specifications of a Partly Reusable ‘RN-35’ Launch Vehicle ……………… 13.5. Appearance and Specifications of a Reusable Solar Tug MSB-1 (RST-1) …………………… 13.6. Foreign Space Transport Systems ………………………………………………………………… 13.7. Conclusions …………………………………………………………………………………………… 13.8. List of References ………………………………………………………………………………………

304 304 305 306 307 308 311 312 312

AFTERWORD …………………………………………………………………………………………………… 313

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LIST OF COMMON ABBREVIATIONS AB — AC — ADV — ACS — AES — AJE — AM — AMS — AP — APDA — a. u. — CIV — CLSS — Cospar — CRPS — CRV — CSR — cSv —

CUU — CWRS — DLR — ECD — ECG — EDB — EF — EP — EPAL — EPS — EPU — ERB — ERES — ESA — NSS — EVA — FA — FSUE — FTI — GCR — GDC — GIRD — GNCS — FA — HIF — HR — HSC — IAEA — IBMP — ICRP — ICV — IEP — IMC — IOV — IOVIE — IPS — ISS — ITMC — JLP — LPE — LSS — LV — MADV — mbar — MCC — MDM —

acceleration block aerodinamic container ascend-descend vehicle Automatic control system artiicial Earth satellite arc jet engine ascend module artiicial Mars satellite arterial blood pressure androgynous peripheral docking assembly astronomical unit of length equal to the average distance from the Earth to the Sun, 1 a.u. = 149.6 mln. km cargo interplanetary vehicle crew life support system Committee on Space Research of the International Council of Scientiic Unions crew radiation protection system Crew return vehicle, in which the crew returns to the earth orbit after on the completion of an interplanetary mission cosmic solar rays with a radiation efect (Е>10 million electron volts) Centisievert, 1 cSv = 1/100 Sievert – an in-system unit for measuring equivalent ionizing radiation dose received by a living organism, a Sievert unit is equal to the radiation absorbed dose unit measured in Gray (Gr) multiplied by the radiation danger coeicient K (К = 1-20); the former equivalent dose unit 1 rem = 1 cSv; I Gray = 1 Joule of energy absorbed by 1 kg of the body weight cryogenic upper-stage unit condensed water regeneration system Deutsche Forschungsanstalt fur Luft und Raumfahrt – German Aeronautics and Astronautics Research Society electrocardiogram electrochemical generator experimental design bureau eiciency factor electric propulsion electric propulsion with anodic layer electric propulsion system, which integrates the power plant with the propulsion unit electric propulsion unit earth radiation belt (Van Allen belt) earthbound renewable energy source European Space Agency Nozzle passage shut-of system extravehicular activity of the vehicle crew fuel assembly federal state unitary enterprise Flight-Test Institute galactic cosmic rays gas discharge chamber of an ion electric propulsion engine Group of Jet Propulsion Exploration – a public organization formed in Moscow in 1931 guidance, navigation and control system heat-generating reactor fuel assembly, which contains the issionable material harmful impurities air ilter heart rate heat-shedding coolers, which are used to shed heat into the open space International Atomic Energy Agency the RF Institute of Biomedical Problems International Commission on Radiological Protection interplanetary cargo vehicle ion electric propulsion interplanetary mission complex interplanetary orbital vehicle ion optical system of an ion engine integrated propulsion system international space station intelligent telemedical circuit system NASA Jet Propulsion Laboratory, manned up with California Technology Institute research staf . The main sphere of interest is connected with the outer space research carried out with the help of automated spacecraft liquid-propellant engine, which uses liquid working components life support system launch vehicle Martian ascent-descent vehicle Millibar – an of-system unit for measuring pressure, 1mbar=1/1000 bar, 1bar=0.1MN/m2=1.0197kg/cm2 mission control centre Martian descent module

7

MF — MFA — MIV — MPDE — MRS — MSO — MW — NASA — nl — nm — NPB — NPP — NPS — NPPU — NRE — OARC — oersted — OS — PHS — PLF — PP — PS — PSS — PSU — PU — RAC — RAS — RD — RDC — RF — RPE — RRE — RS — RS — RSC — RST — RTC — SA — SAC — SC — SDB — SPE — SPO — SPP — SPS — SPU — SSC — STS — TCG — TGI — TLC — TME — TPEU — UES — UHF — VHF — WM —

8

magnetic ield main fuel assembly manned interplanetary vehicle magnetoplasmodynamic rocket engine multipurpose rocket stage Mars satellite orbit Megawatt=106 watt National Aeronautics and Space Administration normal litre nanometer, 1nm=10-9 m nuclear power-block nuclear power plant nuclear power station nuclear power and propulsion unit nuclear rocket engine orbital assembly and refueling complex Oersted – an in-system unit to measure magnetic intensity 1 Oersted=79.58 A/m orbital station pneumatic and hydraulic system payload fairing power plant propulsion system, which includes the engine, the working-medium tanks and the automatic control and fault detection system power support system power supply unit propulsion unit Russian Academy of Cosmonautics named after K.E. Tsiolkovsky Russian Academy of Sciences return drum – reactor power control efector research and development centre Russian Federation research and production enterprise retrothrust rocket engine ionizing radiation shield Russian segment of the International Space Station Rocket-and-Space reusable solar tug research and technological centre solar array short arm centrifuge a spacecraft, unmanned vehicle special design bureau stable plasma jet engine solar proton occurrence accompanied by a chromospheric discharge (lare) of highest-energy particles called cosmic solar rays (CSL) solar power plant space power station solar propulsion unit state scientiic center space transport system turbo-compressor generator turbo-generator installation take-of and landing complex telemetric medical equipment harmful trace-pollutant extraction unit Unmanned explorer spacecraft ultra-high radio frequency ranging from 300MHz to 300 GHz very high frequency radio waves with the wavelength of 10–0.10 m working medium (propellant)

INTRODUCTION Exploration of Mars, which arouses keen interest of scientists, as well as of the broad public, has recently become one of the mainstream trends of space research. To some extent this interest has been supported by the still existing promise of discovering some life forms or at least some traces of life once inhabiting Mars. If it were to happen it would be a real breakthrough, which could throw light on the problem of the origin of life. Mars gives us a unique opportunity to study the problems of the Solar System planetary evolution at large and to make a forecast of the terrestrial biosphere evolvement in particular. But the main point is that Mars is the only planet suitable for colonization, the need of which might arise for the sake of preserving terrestrial civilization. his is probably the main objective of a prospective Martian mission. Lately we have become acutely aware of the possibility of global disasters. Should they happen, life on Earth would be at stake and the price of our survival would be too high. Under these circumstances, putting off the preparation of a Martian mission would be unwise, for it is a challenging task demanding considerable time. here have been numerous attempts to develop a project of a manned mission to Mars. Space agencies of the leading world powers, including the Russian Space Agency, regard this issue as one of the most promising tasks of space exploration. Manned missions have always been high on the list of Russia’s space research priorities, whose exploration of outer space is based on the most advance technologies. Leading Russian research institutes and design bureaus, which determine Russian industrial development, have been considering the issues of a manned mission to Mars from the very start of the Space Age. he development of an interplanetary manned vehicle for a round trip to Mars, which will undoubtedly be a most sophisticated manmade space object, will require synergy of the most advanced technologies. At the same time, implementation of this task will open a way to using these technologies for public ends and contribute to the World stability. his book presents the current Russian concept of a manned mission to Mars. It is intended for all those who are interested in the mainstream trends of scientific and technological development in the XXI century.

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Chapter 5. Power and Propulsion System 5.1. Background he first projects of a manned Martian mission suggested by F.A. Tsander in Russia in 1924, and by Werner von Braun in the USA in 1952, were based on the employment of liquid propellant engines, see Chapter 2. In 1959, S.P. Korolyov started work on a manned mission to Mars. In 1960, ‘OKB-1’ presented the first Soviet project of a manned mission to Mars, which contemplated the employment of a nuclear propulsion system. It must be noted that by 1960 both Russia and the USA had obtained the first results in the field of nuclear space engineering research. In Table 5.1 you can see the types of nuclear power systems, which have been developed or contemplated in Russia. Table 5.1. Progress of the Nuclear Space Power Systems Development in Russia [5.1] Power system type designation

Main characteristics

Development progress

3 ‘IR-100’ reactors and 17 hydrogen delivery systems for an experimental nuclear propulsion with the A nuclear rocket engine with a solid-core reactor, propulsion power of 3.5 tons and Isp = 910sec. have which is used to heat the hydrogen lowing out been already developed and tested. Test duration of the nozzle. Propulsion power – from 3.5 to А – 1 hour. Another step in this direction was the 35 tons. Speciic impulse Isp = 750–920 sec. development and testing of a closed gas-turbine Endurance – 5–6 hours. Generated electric output loop for the power level of Nel = 10 KW. Test duration Nel = 10–50 KW for a reactor working with a reduced output capacity amounts to 6,000 hours A nuclear rocket engine with a solid core reactor, which is used for hydrogen heating and further A developed engine concept. B burning together with oxygen. Propulsion power - 30 tons. Isp = 2,000 sec. A nuclear rocket engine with a gaseous core, A developed engine concept. Models of some of the which is used to heat the hydrogen lowing out of C engine and reactor components have been already the nozzle. Propulsion power – 60 tons. made and tested. Isp = 1,000 sec. A nuclear engine, in which hydrogen is heated Assessment computation of the design engine D with the help of the heat emitted during atomic parameters has been carried out with a view to mini-bombs explosions using this type of engine for asteroid transportation 32 unmanned spacecrafts with the output of Nel = A nuclear power plant with thermoelectric energy E-T 2.5 KW and the endurance of 0.5 year have already conversion Nel = 2–100 KW. Endurance – 10 years been in trial operation E 2 unmanned spacecrafts with the output of A nuclear power plant with thermionic energy N = 5.0 KW and the endurance of 1 year have E-I conversion Nel = 2–7,500 KW. Endurance – 3 years el already been in trial operation A radioisotope energy source with thermoelectric This power system is used on almost 100 R energy conversion. spacecrafts, robotic explorers and planet rovers. Nel = 3–12,000 W. Endurance – 15 years Nel = 150–300 W, Endurance – 15 years Assessment computation of the engine parameters 7 F A nuclear photon engine Isp = 3 •10 sec. shows that its Nel = 1, 000MW, Tem. = 2,430 K

Rank of availability

4

5

6

7

2

3

1 8

107

Fig. 5.1 shows a photo of an ‘IR-100’ reactor, while Table 5.2 presents the results of the tests carried out on reactor No1 for type-A nuclear propulsion. Table 5.2. Results of the ‘Energy Startup’ (ES) and ‘Hot Firing’ (HF) Tests carried out on a ‘IR-100’ Reactor [5.2.] Type of test, date Heat output, MW Test duration, sec. Mean temperature at the at the fuel assembly output, K Pressure at the reactor vessel input, MPa Hydrogen consumption through: – vessel – relector – delay element, kg/sec – fuel assemblies, kg/sec

ES 27.03.78 24 70 1670 6.04

HF-1 03.07.78 33 93 2630 9.46

HF-2 11.08.78 42 90 2600 10.65

1.72 1.18

3.23 1.46

3.51 2.01

During the tests carried out on ‘IR-100’ development reactor No 2 (25.12.1981), the researches maintained the heat output of 63 MW for 38 seconds. he duration of the tests performed on development reactors No- s 1 and 2 depended on the available amount of the working medium – hydrogen. he results of the tests and of the further research activities carried out after the reactor start-up proved the correctness of the selected design concepts and confirmed that the reactor could be used as a basis for the development of space power systems and nuclear propulsion. It should be noted that from the very beginning, the designers concentrated their efforts on the development of a two-mode nuclear engine. When working at full capacity for a short span of 5–6 hours it could provide propulsive power, and when Fig. 5.1. Miniature nuclear engine reactor working at a much lower capacity (about 1,000 times lower) it could supply electric power for onboard needs for about a year of continuous work. To prove the operability of a nuclear reactor in the low-energy mode, the third development reactor of that series – ‘IR-100’ No 3 (RA reactor) was subjected to long-term hot firing tests (which lasted for about 6,000 hours). A special tester unit was constructed to test the ‘IR-100’ reactor, as well as power systems and propulsion units developed on its basis, see Fig. 5.2. At first, the reactor exhaust was discharged directly into the atmosphere. However, after the Chernobyl disaster, the safety requirements were toughened. Now, nuclear plants can be tested only under the condition that the tester unit is fitted out with the exhaust gases purification equipment. Fig. 5.3 features a nuclear engines test setup, which includes purification of the exhaust. Naturally, the necessity to comply with this requirement will increase the nuclear engines development costs. he concept for the Fig. 5.2. Tester unit for nuclear power plants and engines. Semipalatinsk, the Republic of Kazakhstan

108

DEVELOPMENT REACTOR PROGRAM

Fig. 5.4. History of the major nuclear rocket engine reactors tests in the USA

DEVELOPMENT ENGINE ‘NERVA’ PROGRAM

Fig. 5.3. Diagram of a tester unit with an exhaust gas treatment installation, Russia

development of a nuclear propulsion formulated by 1992 [5.3.], included the following conceptual tasks: • development of the rationale for the achievement of the NRE propulsion power at the level of 68.6 kN (7 tons), including the assessment of the costs for the construction and operation of a tester unit and an exhaust gas treatment system for the engineering development of the NRE; depending on the task, NREs can be employed in clusters of 2–4 engines; • development of a two-mode NRE with an electric output within the range of 20–100 KW; • development of a NRE with a specific impulse of about 950 sec, at the desired endurance of 5–6 hours; • development of a NRE with the thrust to mass ratio = 2 with allowances for protection and further improvement of this ratio to 4 at Isp ~ 1,000 sec. Russian NRE studies were based on a heterogeneous arrangement of the reactor core, where uranium- Fig. 5.5. Kiwi-B4E reactor bearing material and the inhibitor are placed separately. his approach permitted to develop a miniature NRE with the propulsion power of 3.5 tons, which was very important at the first development stages. Meanwhile, Americans were working at the development of a reactor with a homogeneous arrangement of the reactor core, with the propulsion power of at least 30 tons. Fig. 5.4 Fig. 5.6. Power unit construction diagram shows the history of NRE tests in the USA, [5.4.]. Fig 5.5 shows a Kiwi-B4E reactor prepared for a test. One can clearly see that the exhaust gas went straight into the atmosphere without undergoing any treatment. In Russia, starting from the middle of the 1950 s studies of the gas-core reactor, based on the employment of uranium-235 plasma, were carried out at Keldysh Center, NIKIET [5.5]. Fig. 5.6 shows a construction diagram of a nuclear rocket engine with a gas-core reactor. 109

Starting from the middle of the 1960s, a concept of a nuclear propulsion unit with the electric output of 5–7.5 MW based on the thermionic energy conversion principle has been developed at RSC Fig.5.8. General view of ‘BUK’ NPP ‘Energiya’ (Russia) [5.6], see Fig. 5.7. he task was Fig. 5.7. General view of the development set in connection with the reactor FS-1-4.19 realization of future manned 1 – reactor core; 2 – a model of a Martian mission. power-generating assembly with 36 EGC models; 3 – side core relector; In Russia, further 4 – rotary cylinders of the reactor development of the nuclear protection and control system propulsion unit major design Fig.5.9 General view of ‘Topaz’ NPP and operation principles was carried out on the basis of miniature nuclear plants with thermoelectric energy conversion – ‘BUK’ (see Fig. 5.8) and with thermionic energy conversion – ‘Topaz’ (Fig. 5.9). hey were given a trial run in space in the period from 1970 to 1988. he studies carried out by NPO ‘Krasnaya Zvezda’ demonstrated that it was necessary to pay due attention to the problems of the nuclear safety and radiation protection connected with the operation of nuclear propulsion from the very start of the project development activities. Keldysh Research Center was appointed principal research establishment responsible for the safe use of nuclear power units for space exploration purposes [5.7]. Of all the nuclear propulsion units only radioisotope thermoelectric generators (RTG) are now currently in use. Table 5.3 presents the main characteristics of RTGs developed in Russia and in the USA, particularly for ‘Apollo’, ‘Viking’ and ‘Voyager’ programs. Table 5.3. Main characteristics of RTGs, running on 238Pu, t1/2 = 87.7 years Characteristic Eiciency factor, % Speciic output, Watt/kg Endurance , years * Note: the RTG uses the thermionic conversion

30 10 5.0 10

Russia, Nel, W 60 100 5.5 5.5 2.5 2.6 10 10

230 12* 6.6 10

28 6 2.1 15

USA, Nel, W 42 73 6.3 5.0 3 2.3 15 15

160 6.7 4.2 15

5.2. LPE Version By 1992 Russia had developed a project of a manned Martian mission on the basis of an interplanetary vehicle [5.2.], which employed an LPE with cryogenic components – O2 and H2. he useful load included: Interplanetary orbital vehicle (IOV), 6 crewmembers .......... 80 tons Ascend-descend vehicle, 3 crewmembers ................................ 60 tons Crew return vehicle....................................................................... 10 tons Total mass: ...................................................................................... 150 tons 110

PPS and IMC characteristics PU propellant – Н2+О2 Thrust – 4×2000 kN 4×75 kN 1×75 kN Speciic impulse – 479 sec. PP output – 150 KW IMC initial mass – 1350 tons Mission duration time – 660 days IMC maximum length – 140 m IMC maximum cross dimension – 60 m Fig. 5.10. Lay-out diagram of IMC with LPE and SPP energy sources

According to the project the mission was to start from a low-earth orbit in 2010. he total mission duration time amounted to 660 days, including 30 days on the Mars orbit. he initial mass was 1350 tons on the condition that all the LPEs ran on the O2 + H2 components and provided the specific impulse Isp = 479 sec. he total propellant margin was equal to 1040 tons, including 850 tons requested for a breakoff from Earth. In this variant, one of the major challenges was the provision of appropriate conditions for a long-term storage of liquid hydrogen: for 1 year during the mission complex assembly on the low-earth orbit and for at least 1.8 years during the interplanetary transfer phase. he problem of the cryogen components preservation can be solved by the installation of onboard refrigeration machines to the total capacity of 150 KW. Fig. 5.10 presents the layout of an LPE-based interplanetary vehicle. If we assume that the liquid O2 + H2 components were used only for gaining a get-way acceleration at the start of the mission, and that a stable-component LPE (Isp = 320 sec.) was used for deceleration when the mission approached Mars and for acceleration when the mission started on a return trip to Earth, we will see, that the interplanetary vehicle initial mass would increase to 1,700 tons. It is necessary to specify that after a respective acceleration or deceleration maneuver, the WM tanks were to be shot off. he project didn’t consider the possibility of the IMC return to a low-earth orbit, so the IOV under consideration was expendable. here was another problem connected with placing the cryogen propellant storage tanks into the low-earth assembly orbit during the preparatory phase. We have already said that the IMC would need 850 tons of fuel to get away from Earth. Now, if the propellant were delivered to the vehicle by heavy launch vehicles with the load-carrying capacity of 100 tons, each of which carried 85 tons of cryogen components (12 tons of liquid hydrogen and 73 tons of liquefied oxygen), about 10 launches would be necessary to accomplish the task. he requested storage tank volumes would amount to 170 m3 for hydrogen and 64.3 m3 for oxygen. Assuming that the tank diameter is 5.5 m, the total tank length would be 10 m. he employment of a smaller launch vehicle, e.g. with the carrying load capacity of 42.5 tons, will double the number of tanks and, consequently, of launches, making the task unrealistic. From the above it follows that the realization of this LPE-based EPU variant will unquestionably need a launch vehicles with the load-carrying capacity of 85–100 tons. In this case, the total number of launches requested for the 111

assembly of the whole interplanetary mission complex with the initial mass of 1350 tons will amount to 15. Remarkably, 12 of them, or the absolute majority, will be launches of the vehicles carrying the tanks with the cryogen propellant. Currently, a country, no matter how big or developed it is, can afford no more than 6 launches of the carrier vehicles with the carrying load capacity of about 100 tons [5.8]. It means that at least two countries should possess such vehicles. Alternatively, a country sending an interplanetary vehicle could extend the assembly period to at least 2.5 years. But this measure would again raise the problem of protecting the cryogen propellant from boiling through the employment of special refrigeration machines. In 2001–2004, the European Space Agency carried out intensive studies of a manned Martian mission, performed on a LPR propulsion vehicle. It was assumed that the LPE used O2+H2 cryogen components for gaining a get-away acceleration with the specific impulse of 450 sec. at the start of the mission, while a stable component (Isp = 325 sec.) was used for deceleration when the mission approached Mars and for acceleration when the mission started on a return trip to Earth [5.9]. he results of the studies showed that the useful load would amount to 124.4 tons and include: 66.7 tons; • interplanetary orbital vehicle for a crew of 6 46.7 tons; • Martian ascend-descend vehicle for a crew of 3 11 tons. • crew return vehicle he mission will start from a low-earth orbit in 2033. he total mission duration time amounts to 963 days, including 30 days on the Mars orbit. he initial mass of the interplanetary complex on the low-earth orbit is 1357 tons. he duration of the on-orbit assembly time amounts to 4.6 years. A large amount of the cryogen propellant is likely to evaporate over this period, therefore, the initial mass of the interplanetary vehicle has to be increased to 1541 tons. he interplanetary vehicle will be assembled from several modules. Each of the modules will be launched to the assembly orbit by a respective launch vehicle: − 80 tons, ‘Energiya’ launch vehicle; − 20 tons, ‘Proton’ launch vehicle; − 20 tons, ‘Arian-5’ launch vehicle; − 11.2 tons, ‘Soyuz’ launch vehicle; − 20 tons, ‘Space-Shuttle’ launch vehicle. he configuration of the interplanetary vehicle at the moment of its start from the low-earth orbit is shown in Fig. 5.11. It was suggested to reduce the IMC initial mass by the employment of double-dip aerodynamic deceleration in the Martian atmosphere for the transfer to the desired 112

Fig. 5.11. LPE-propulsion interplanetary vehicle (ESA project)

near-mars orbit [5.10]. Because of the low atmospheric density, the depth of the dip should be equal to 30 km from the Martian surface. he LPE will use O2+H2 cryogen components for gaining a get-away acceleration both from Earth and from Mars. his measure will provide Isp = 480 sec. he Martian ascend-descend vehicle uses a stable –component LPE with Isp = 330 Fig. 5.12. IMC, which employs aerodynamic braking to get into a sec. In this case, the initial mass of the near-mars orbit: 1 – irst acceleration block used to get away from the interplanetary spacecraft will be equal low-earth orbit and start on a trip to Mars; 2 – MDM; 3 – aerodynamic to 775 tons against 1350 tons requested shield; 4 – CRV; 5 – IOV; 6 – second acceleration block used to get away for the version, which contemplates jet from the near-mars orbit and start on a trip to Earth; 7 – attitudecontrol engines braking for the transfer to the assigned near-mars orbit. Fig. 5.12 shows the 100.00 −g configuration of the interplanetary ±g vehicle. It is remarkable for the size of +g 10.00 its aerodynamic shield (22.5×26.2 m) −g +g and for the integration of a acceleration block into an interplanetary complex, 1.00 4.00 6.00 8.00 10.00 12.00 0.00 2.00 which undergoes aerodynamic Acceptable interaction time (min.) deceleration in the Martian atmosphere. Fig. 5.13. Acceptable g-factor exposure of the crewmembers he acceleration block has a cryogen- contemplated for diferent axes component (O2+H2) LPE. he total propellant mass amounts to 95 tons, including 13.5 tons of liquid hydrogen. he aerodynamic deceleration maneuver in the Martian atmosphere is a challenging task in itself. It needs great accuracy with regard to the entry speed (0.05–0.1 m/ sec.) and constant aerodynamic shield centering control. he presence of liquid hydrogen puts the life of the crewmembers at risk. It also very important to comply with the acceptable g-factor in view of the crewmembers’ long exposure (about 5 months) to zero gravity. In Fig. 5.13. you can see acceptable g-factor values, which depend on: • direction of the g-force; • duration of the g-force exposure; • duration of the preceding exposure to zero gravity. 2

g-force

y z

x x

5.3. Nuclear-Powered Propulsion Version 5.3.1. Development of the baseline nuclear space engineering technologies he development of a new class of space reactor units based on the principles that make them quite different from their ground prototypes, demanded for theoretic and practical conceptualization of all the processes connected with design engineering, selection and development of new engineering procedures for the production of the reactor and nuclear power plant components, development and implementation of new testing methods for individual components and integrated functional system tests for trials in the maximum real environment. 113

Since the system will be inaccessible for direct interference of the crew during the mission, the key tasks solved by space-system engineering include assurance of the highest level of reliability with regard to all the system components, development of an automated control system and elimination of a potential need for repairs when in service, to name just a few. Besides, the nuclear power plant performance parameters fall within the maximum permissible limits. his fact requires for original solutions at the product design and engineering development stages. Special importance is attached to the selection of the structural materials for the whole power plant at large and for its reactor core in particular. It must be noted that a number of technical solutions requested for the development of a space NPP, particularly those, which refer to its control system, propellant composition, structural materials and some types of equipment are really unique both in nuclear reactor industry and in other industrial sectors. First and foremost, all space-bound equipment should comply with flight, operational, radiation and nuclear safety requirements and also with requirements for mass and size parameters. Neither the development of the first generation of nuclear power plants based on the principle of direct conversion of heat into energy, nor the subsequent development of the first Russian prototypes of nuclear rocket engines would have been possible without the development and realization of radically new technologies. hese technologies and production lines were built up at various research institutes alongside with the testing and evaluation facilities for the engineering development of the power plants and their components. By 1990, after a 20-year period of continuous development and growth, the new production and research facilities had integrated into an independent industrial segment, which carried out the whole complex of activities connected with design, production, engineering development and testing of space-oriented nuclear power plants [5.11–5.16]. A significant difference between a NRE reactor and other space-oriented reactors demanded for the solution of several scientific and technical problems. he most critical was the development of the structures, which would work in the hydrogen environment and withstand temperature and pressure differences ranging from cryogenic to 3,000 K and from vacuum condition to hundreds of atmospheres respectively. he development of nuclear propulsion units (NPU) generates additional problems, particularly in connection with the conversion of thermal energy into electric energy, including the development of a reliable heat-shedding cooling device with optimal mass and size parameters and a recuperative heat exchanger, which would ensure a sufficiently high thermodynamic efficiency of the nuclear propulsion unit. A number of new purpose-oriented technologies and complexes, which ensured the development and realization of baseline techniques for the design and production of space-related nuclear engines and propulsion units, were expected to facilitate the solution of the most complicated problems arising in connection with NRE and NPU projects. hey are listed below: • a complex of programs and guidelines and a design and procedural support technology for NRE and NPPU engineering development. his complex is intended for the development of design baselines and concept. he concept 114

















should proceed from the assumption that since the number of the developed objects is limited the major amount of testing in the course of the engineering development of the propulsion unit with regard to target reliability of the its components and assemblies will be carried out in simulated or full-scale conditions; production and technological complex, for the realization of a production technology, which will ensure reliability of the propulsion unit components and systems when exposed to hydrogen or other heat carriers used as working medium in the process of operation; production and testing complex for the realization of an assembly technology applicable to the assembly of the propulsion unit components and for the final end-product assembly, including installation of measuring devices and hydrodynamic tuning of the cooling-fluid circuits for the target working medium consumption distribution , followed by the integrated end-product tests and production of the model fuel assemblies; research and testing complex and technologies for engineering development of the NRE parts and components carried out through a series of tests with the employment of real and model working mediums (mock hydrodynamic tests and trials, high temperature tests with the employment of ohmic heaters and plasma generators) and for substantiation and engineering development of radiation safety measures applicable during NREs and NPPUs transfer to the low-earth orbit; production complex and technology for the production of NRE and NPPU heat exchange assemblies based on compact plate-type heat exchanges with a specific heat-exchange area of 1.000–1,500 m2/m3, advanced droplet radiator-coolers and radiator-coolers based on the employment of heat transfer pipes; production complex and technology for the production of solid-fuel reactor core components from carbide and carbonitrite compositions, which sustain operability for several hours at the hydrogen temperature of 3,000 K and higher; production complex and technology for advanced purification of the power circuit inert gases (He, Kr, Xe, Ar, etc.), and laser optical diagnostic aids for the monitoring and analysis of the structural materials and fuel under service conditions; a complex of technical resources and a technology for substantiation of structural (various types of steel and heat-resistant alloys) and special (beryllium, zirconium, etc.) materials applicability in the working environment and under the conditions of radiation exposure; experimental and production complex for the production and testing of high temperature measurement devices

he development of the space-related nuclear engineering baseline technologies permitted to build up the production of unique nuclear power plants and carry out ground tests of prototype nuclear rocket engines, whose technical characteristics are shown in Tables 5.4 and 5.5. It would be appropriate to say that these characteristics by far surpass those achieved at the US nuclear research centers. 115

he whole complex of work performed in the course of the development of baseline technologies and serial production of space-oriented nuclear power plants supported by successful ground tests of the prototype NRE paved the way for the development of nuclear propulsion units for a manned mission to Mars and to other planets, for the deployment of the lunar base, etc. Next, we would like to draw you attention to some peculiarities, which characterize different types of nuclear units. Table 5.4. Comparative Figures for the Achievements of the USSR and US Direct Conversion Nuclear Propulsion Units Development Programs Characteristics Total NPPU mass, kg Useful electric output, KW Thermal output, KW Endurance: – speciied, years – achieved, years Launch position size: – length, m – diameter, m Reactor type Energy conversion technique Speciied NPPU mass, kg/KW el. Development progress Number of light items Development costs, mln USD

BUK 1,450 (inc. de-orbiting system) 2.6 90

USSR TOPAZ-1

TOPAZ-2

SNAP-2

USA SNAP-10A

SNAP-8

1350

1000

545

295

4545

5 150

6 160

6 55

0,5 30

35 600

0.25 0.96

1.5 – 3 1.5

I 1.2

1 0.12

1 0.95

3.8 3.9 1.3 1.37 Intermediate spectrum

– –

0.5 0.75 4.79 1.3 Fast-neutron

– – – – Intermediate spectrum Rankine Rankine cycle Thermocycle Thermo-electric Thermionic mercury electric mercury vapour vapour 560 270 170 182 590 130 Ground Ground Ground Flight tests Flight tests tests were tests were Withdrawn from were carried tests were were carried carried out in carried out in carried out in use in 1993 out in out in 1965 1965–1968 1987–1988 1980–1983 1961–1965 32 2 – – 1 – 466 224 418 490 243 417

Table 5.5. Comparative Figures for the Achievements of NRE Development Programes Characteristics Period of increased activity in the question area Total expenditures, bln. USD Number of reactor facilities made Engineering development and design principles Fuel composition Reactor core thermal factor, mean/maximum, MW/l Maximum achieved working medium temperature, K Speciic impulse, sec. Endurance at the maximum working medium temperature, sec.

116

USSR 1961–1989 ~ 0.3 5 Component-based approach Solid solution UC-ZrC, UC-ZrC-NbC 15/33 3100 ~ 940 4000

USA 1959–1972 ~ 2.0 20 Integrated approach UC2 in graphite matrix 2.3/5.1 2550 ~ 850 2400

NRE – is an assembly, in which the working medium (WM), being heated to the requested high temperature at the expense of the energy released in the course of the nuclear fuel fission reaction, flows out of the nozzle and provides jet propulsion during the flight. he NRE nuclear reactor fulfills only one function – it heats the working medium produce the propulsive burn, and in this way, realizes the propulsion mode. NPPU – is a nuclear power and propulsion unit provides the vehicle transfer and generates the electric power. here are two NPPU modifications: A bi-modal NRE – is an upgraded NRE version. In a bi-modal (dual-mode) NRE, the nuclear reactor heats the working medium (hydrogen) and provides the propulsive burn for the vehicle transfer in space, realizing its propulsion mode. At the same time, it heats the heat-carrying medium of the power conversion system, which generates the electric power for the interplanetary vehicle onboard needs (energy mode). NPP combined with NREs is a system, which combines a nuclear power plant and electric rocket propulsion. he nuclear reactor heats the working medium of the energy-conversion system. he latter generates the energy used for propulsion and for the interplanetary vehicle onboard needs. 5.3.2. Cruise NPPU designation and specifications A cruise NPPU solves the task of delivering useful cargo, whose mass is determined by the selected Martian mission scenario [5.17], from the nuclear-safe low-earth orbit to the target Martian orbit. In the case of a two-vehicle Martian mission complex (MMC) a scheme, each of the vehicles – the manned interplanetary vehicle (MIV) and the cargo interplanetary vehicle (CIV) – has a cruise NPPU. he employment of a nuclear reactor as a cruise propulsion unit energizer for the purposes of interplanetary missions to Mars and to the other planets of the Solar System has a number of advantages over the other existing types of propulsion units. First of all, it’s the compact size, which the gives the interplanetary vehicle mobility and maneuverability. Another advantage is the vehicle self-sufficiency irrespective of its spatial position (remoteness from the Sun, or ingress into the planetary shading area). And last but not least is its reusability. One and the same nuclear and energy propulsion unit can be used on several Martian missions. However, for this option the NPPU will need additional maintenance on the radiation-safe low-earth orbit carried out in strict compliancy with the implicit requirements general and radiation safety. In the first place, these requirements apply to working-medium (hydrogen, helium, xenon, neon) loading operations. he following are the prevailing requirements for a cruise nuclear energy and propulsion unit: a high specific impulse value, a low mass-to-power ratio, endurance at continuous-duty rating and compliancy with the requirements for the mass and size parameters conditioned by the cargo compartment and the load-lifting capacity 117

of the launch vehicle used for the delivery of the of the interplanetary vehicle component to the low-earth assembly orbit. hese basic requirements determine the parameters of the nuclear reactor and other components with regard to all aspects, including the provision of radiation and nuclear safety, and compliance with the requirements for admissible radiation load on the crew quarters and instrumentation compartments. Radiation (biological) protection should provide acceptable conditions for radiosensitive instruments, crewmembers and vehicle equipment accommodated in the shaded area, in compliance with the international standards for radiation load limitation. he mass and size parameters of the NPPU under consideration (bi-modal NRE and NPP combined with EP) and of the respective mated vehicle components should comply with the load capability of the prospective heavy (35 tons) and super-heavy (70 tons) launch vehicles with a payload fairing (PLF) diameter of 6.5 m and a useful load area of 6 m in diameter and 22 m long (17 m in the cylindrical part) in the first case, and 35 m long (30 m in the cylindrical part) in the second case. In the context of a potentially long NPPU service life, we suggest using a solid core gas-cooled fast-neutron reactor as its basic version nuclear-heat source. When developing cruise NPPUs, special attention is paid to the provision of nuclear and radiation safety at all the stages of the Martian mission life cycle [5.19, 5.20]. he underlying principle of the modern concept of a safe use of nuclear energy sources for the purposes of space exploration is minimization of the radiological impact on the population and the environment. he key points of this concept are based on: • the principles of using nuclear energy sources in space approved by the UNO General Assembly Resolution 47/68 of 14 December, 1992; • recommendations of the International Commission on Radiological Protection, which specify permissible irradiation levels; • International Atomic Energy Agency documents; • national documents – radiation safety standards, sanitary regulations, etc. Currently, two basic versions of cruise NPPUs are being considered for the employment on the MIV and CIV for the purposes of the manned Martian mission realization: • a NPPU based on a bi-modal nuclear rocket engine, which can work in the propulsion and in the energy operation modes; • a NPPU based on a nuclear power plant and a number of electric propulsion rocket engines he bi-modal NRE, which provides propulsion and energy generation, combines two sufficiently different functions: • it functions as an engine with its special working medium, with a top-of-therange temperature at the reactor output, but with a comparatively short total time of engine operation; • it also functions as a usual electric power station with a different working medium and a different thermodynamic cycle; with a moderate working temperature and a materially longer operation time. 118

Since the reactor output in these two operation modes differs by more than two orders and due to a considerable difference in the character of its performance in each of the respective modes, the requirements imposed on the reactor and on the NPPU equipment have no analogues either in rocket production or in the atomic power-plant engineering industry. Naturally, this peculiarity, which leads to a more sophisticated structural configuration of the energy and propulsion unit, must ensure the realization of the technical requirements imposed on the spacecraft at large. he suggested cruise energy and propulsion unit based on the bi-modal NRE with a turbo-machine energy conversion consists of a cluster of 3–5 individual modules (the basic version contemplates 4 modules) with the propulsive burn of 68 kN each, which work simultaneously (both near Earth and near Mars) by 30–60 minute pulses. In this way, the employment of the NRE technology ensures a quick passage through the Earth radiation belts (ERB) (in about 5 days) and builds up the initial velocity gain when the IMC reaches the Martian flight path [5.18]. In the case of a bi-modal NRE, special measures should be taken to prevent or minimize the mutual influence of individual reactor modules (in a cluster). An interplanetary mission complex with bi-modal NRE modules is assembled on the radiation-safe low-earth orbit. he number of the launch vehicle shots requested for the assembly of the IMC will depend on the launch vehicle (LV) load-lifting capability. he most efficient way of the assembly works organization is based on the employment of multipurpose rocket stage (MRS) in the function of an inter-orbital docking tug. he correction of the interplanetary trajectory section of the Earth-Mars transfer can be carried out with the help of a bi-modal NRE, while an independent LPE-based corrective propulsion unit can be used to correct the Mars-Earth interplanetary transfer trajectory. he total mission duration, including the projected stay of the crew on the Martian surface, will amount to 460 days. he other cruise NPPU modification presents a combination of nuclear power plant and electric propulsion rocket engines [5.21]. he thermal energy generated by the NPP gas-cooled nuclear reactor is converted into the electric power, which is used to energize the EPs and to meet the onboard demand, by the gas-turbine plant generator. his cruise NPPU version has good prospects due to the availability of state-of-the-art power and propulsion systems. he high specific impulse of the electric propulsion and the combination of EPs with the NPP permits, among other things, to minimize the Martian mission complex initial mass. he assembly of the Martian mission complex, which employs and integrated NPP-EP propulsion system, is carried out on the radiation-safe low-earth orbit with the help of LPEs. he energy and ballistic evaluation shows that the employment of an integrated NPP-EP propulsion system can reduce the mission duration by about 50% as compared to the version, which contemplates the employment of the bi-modal-based cruise NPPU. However, the realization of this version depends on the development of an NPP (or a an NPP cluster) with the total power output of 50MW and the total mass-to-power ration of 1.5–2 kg/kWel, which will be capable of providing energy for the following two types of EPs: 119

• arc-jet engines with Isp =1,500 sec. for acceleration within the Earth’s sphere of

activity; • ion-engines with Isp =10,000 sec.

Compliance with the above parameters will reduce the duration of the Earth – Mars – Earth mission to 328 days (including a 1-month stay on Mars) see Chapter 13. 5.3.3. Nuclear power and propulsion unit based on the NRE technology and turbo-machine energy conversion Bi-modal NRE design concept and technical features Basic technical solutions for the NPPU concept based on the employment of a bi-modal NRE were selected with a glance to the following factors: • power plants are believed to be the most costly of the vehicle systems; • when designing a power plant it is necessary to anticipate its reusability and a long service life; • when operating in the propulsion mode, the power plants must ensure speedy delivery of the maximum useful load to the target destination and at the same time generate electric power for the interplanetary vehicle onboard needs; • when operating in the energy mode, the power plants must generate rated power output; • when designing power plants, it is necessary to anticipate redundancy and ensure high reliability of the power and propulsion unit operation in both design modes With the above factors in view, the nuclear power and propulsion unit will be fulfilled on the basis of a cluster of 4 individual bi-modal NRE modules, each of which has a heterogeneous fast neutron reactor as the energy source. he main characteristics of the bi-modal NRE are shown in Table 5.5. he suggested bi-modal NRE module concept is based upon the following assumptions [5.18]: • he reactor active-zone structure will be arranged upon the heterogeneous principle. In accordance with this principle, the nuclear fuel is encased in a fuel assembly. he heterogeneous principle of the reactor active zone arrangement allows for a free choice of materials without their rigid dependence on temperature resistance and permits to achieve optimal mass, size and propulsion performance engine characteristics. • he materials for the shady shielding were selected proceeding from their

efficiency and from the availability of the process technology for the production of respective protection items at the Russian production facilities. Lithium hydride possesses a sufficiently high degree of protection efficiency with regard to reactor radiation alongside with high thermal and radiation resistance and compatibility with the structural materials. he use of depleted uranium as the radiation protection heavy component in combination with lithium hydride permits to minimize the radiation protection structure mass characteristics. A rational arrangement of the support equipment in the propulsion module 120

and the liquid hydrogen in the bi-modal NRE starter tank provide additional reduction in the neutron and γ-radiation density, in this way contributing to the reduction of the radiation protection layers mass proper. Table 5.5. Bimodal NRE module characteristics Parameter Vacuum propulsive power, kN Speciic impulse (propulsive burn), sec. Reactor thermal output, MW Propulsion module working medium Working medium temperature in front of the main reactor FA nozzle cluster. K Conversion of thermal energy to electrical Working medium of the energy conversion loop Temperature entry temperature, K Nominal output power, KW : – of the CIV propulsion unit – of the MIS propulsion unit Bi-modal NRE total burning duration : – in the propulsion mode, hours; – in the energy mode, years

Value 68 ~ 940 340 Hydrogen 3000–3100 Turbo- machine, based on Brayton cycle Helium-xenon composition 1500 15 up to 50 At least 5–6 Up to 10

• A combined cycle (which consists in a simultaneous employment of the

preliminary hydrogen heating recuperative heat exchanger and pre-heating fuel assemblies, located in the reactor core) is used in the bi-modal NRE for preliminary heating of the working medium to the acceptable gas temperature at the reactor core inlet (not less than 300 K to ensure reliable behaviour of ceramic materials) and to the working temperature (~550-650 K) before the turbo-pump assembly turbine. his method is characteristic of a fast neutron reactor. • In a bi-modal NRE module, the integration of the reactor with a dynamic

energy converter, which makes use of Brayton gas turbine cycle, is the most successful combination. he gas-turbine plant working medium, a helium-xenon composition, is heated directly in the reactor energy-conversion-circuit passages (in the reactor core inter-casing space). he unused excessive heat is removed by way of dissipation in the space environment with the help of a radiator-cooler with heat transfer tubes. • Propulsive power is provided through the employment of a one-nozzle

conversion system, which converts the working medium thermal energy into the jet stream momentum with the help of a supercritical propelling nozzle. Proceeding from the NPPU functionality, the bi-modal NRE module should consist of two major parts – the propulsive and the energy-generating energy conversion systems. Importantly, respective hydraulic ducts of the propulsive and energy-generating systems within the reactor and in the module at large should be hydraulically independent over the unit operational period. his was one of the conceptual aspects in the process of the reactor core design. 121

he development of a nuclear reactor is the most sophisticated, multi-aspect and important task in the process of work on the development of a bi-modal nuclear power and propulsion unit for an interplanetary vehicle. he reactor design should comply with the following requirements: • it should ensure operability of all the reactor structural elements in the energy mode for at least 10 years; • it should ensure the maximum possible propulsive power of the bi-modal NRE module with the specific impulse of ~ 940 sec.; • it should possess a certain degree of flexibility with regard to the employed energy conversion system; • its technical solutions should allow for building up output; • it should have minimal mass and size parameters. he following principles were selected for the development of the bi-modal NRE reactor technical features. Neutron-physical characteristics To provide a 10-year employment in the interplanetary mission complex, it is advisable to use a fast neutron reactor, which has proved to have the minimum fuel burnout effect. At the same time, the reactor should be fitted out with some means of nuclear safety activated in case of accidents, connected with the launch vehicle failure during its flight to the low-earth assembly orbit, which might result in the reactor getting into water (or water-containing environment) or getting buried in the soil [5.22]. Fuel and structural materials he suggested type of nuclear fuel is carbon-nitride uranium compositions, which comply with the requirements for compatibility with working media and high-temperature resistance, provide compactness of fuel loading and the minimum size of the core [5.23, 5.24]. he core structural materials are heat-resistant wolfram and molybdenum based alloys and nickel-rich alloys. he core reflector structural material is beryllium [5.25]. Core structure In the propulsion mode, the main mechanism of heat transfer from the fuel elements is convective heat exchange with the working medium, while in the energy mode, it’s the thermal conduction of the core structural elements, particularly of the fuel elements array, and thermal radiation of the free surfaces. For the above reasons, the increase of the power output can be achieved by the improvement of the core elements thermal conduction. Technical solutions used in the concept of the reactor core must comply with the requirement for the possibility of performing a pre-launch ground check and blow-down firing tests of the bi-modal NRE before its delivery to the assembly orbit. In this way, a nuclear power and propulsion unit should include: • a reactor with a propelling nozzle in the propulsion module and with a

radiation protection module; 122

a propulsion mode working medium supply system; energy conversion and heat release loops; a system of working medium tanks; a system of automatic control, diagnostics and nuclear protection; a system for long-term onboard storage of the cryogen components; a bearing structure he nuclear reactor is the vehicle NPPU key component. It should be noted that the reactor is a sophisticated thermodynamic structure and that its thermodynamic component together with the neutron physics determines its structural peculiarities and forms its technical features. he NPPU hydraulic system is an elaborate network of ducts, which simultaneously carry several working mediums, whose thermodynamic parameters vary significantly depending on the unit’s operating mode and are determined by the boundary conditions of the reactor loops to which the ducts belong. When selecting a reactor concept for a two-mode NPPU, it is necessary to proceed from the principle of achieving the maximum possible technical parameters for the operation in each of the target modes. It means that in practice all the reactor structure elements will work under the conditions of the maximum permissible temperatures in both modes, hence the need for a reliable cooling system. his factor must be taken into consideration when designing the NPPU pneumatichydraulic system (PHS). Each module of the bi-modal NRE can either generate relatively short propulsive burns and simultaneously supply electric power (a combined propulsion and energy mode) or continuously generate electric power (energy mode) requested for refrigerating hydrogen in the fuel storage tanks, supplying power to the IMC target equipment and its life support and service systems throughout the Earth – Mars – Earth flight. On account of its functional application area, a bi-modal NRE has two energy conversion systems – one of them converts the thermal energy generated by the nuclear reactor to the jet propulsion, and the other – to the electric power. When selecting the structure of a bi-modal NRE module pneumatic and hydraulic system, it was necessary to comply with the functional specifications. Besides, it was necessary to meet the requirements for the provision of appropriate cooling of the reactor’s structural components in the propulsion (combined) and energy operating modes and for the maintenance of a high thermodynamic efficiency of the energy conversion process. In this connection it was decided that the reactor’s low-temperature elements (side beryllium reflector, radiation protection) would be cooled with the help of an independent auxiliary circuit coolant, since the heat emission of the side beryllium reflector and the radiation protection (RP) sufficiently increases in the propulsion mode and approaches the heat emission values characteristic of the ground nuclear power stations (NPS) fuel elements. It has already been mentioned that a bi-modal NRE consists of 4 identical modules. Fig. 5.14 shows the structural diagram of an individual module pneumatic and hydraulic system arrangement. Hydrogen is the working medium of the reactor block propulsion passages, while the energy conversion loops and coolant circuits employ a gas mixture, which consists of 92.83% Xe +7.17% He (mass) and the neutral helium respectively. • • • • • •

123

he NPPU propulsion module is formed by the reactor block, which consists of a reactor, a nozzle 1 cluster and a radiation protection module, a hydrogen recuperative heat-exchanger, a working medium storage and supply system and a manifold with stop and control valves and cutoff devices. he working medium storage and supply system includes a hydrogen storage tank, two types of pressurization units to pressure up the working 2 medium line pressure – a lowspeed low-pressure booster pump with a hydro-turbine drive (booster turbo-pump unit – BTPU) and the main turbo-pump unit (TPU). 3 he acceptable carbon nitride fuel temperature regimes for the NPPU operation in the continuous energy mode will be maintained by filling the passages with neon Fig. 5.14. Structural diagram of a bi-modal NRE individual module pneumatic and hydraulic system arrangement: from the respective neon storage 1 – electric power generating loop equipment; tank, which forms part of the this 2 – propulsion module equipment; unit. his measure is necessary, 3 – auxiliary circuit equipment because when the unit functions in the continuous energy-generating mode, its hydrogen passages are filled with a neutral gas. he NPPU energy-converting module is a Brayton-cycle turbine-generator installation (TGI) with heat recovery, which includes two single-line turbocompressor generator units (with respective turbines, electric power generators and compressors), two recuperative heat exchangers, the main radiator-cooler and a manifold with stop valves and cutoff devices. he TGI working medium is heated in the reactor core inter-casing space. he cooling system of the reactor block outer components (radiation protection, side beryllium reflector unit and casing and the fire floor) includes an auxiliary radiator-cooler, a turbo-compressor unit, a gas-circulator, a heat-exchanger and a system of and a manifolds with stop and control valves. he heat-exchanging unit removes the excessive heat from the auxiliary circuit coolant to the bi-modal NRE propellant module passages, when the latter operates in the propulsion mode. Let’s study the principal operating modes of a bimodal NRE. he operating mode involves the operation of the reactor block, the propulsion loop heat-exchanging system, the working medium storage and supply system and the gas-turbine power plant (in the partial energy generation mode). 124

he booster turbo-pump unit and the main TPU pump hydrogen from the storage tank into the propulsion passages [5.26]. After passing through the main pump, the cumulative hydrogen flow is divided into three streams: the first stream goes through the recuperative heat exchanger, where it is heated to the target temperature, part of the cooled hydrogen flow goes to the nozzle cluster coolant passage and the remaining part is discharged through a bypass manifold. From the pre-heating reactor-fuel assemblies, the heated hydrogen gets into the recuperative heat exchanger hot duct and then to the TPU turbine, where it interfuses with the cold hydrogen stream and goes first into the heat-exchanging unit integrated with the auxiliary circuit and then, through the coolant passages to the reactor for the cooling of the return drums and the vessel walls. After that, it interfuses with the hydrogen, which has been admitted from the nozzle cluster jacket, and enters the fire floor chamber and then goes through the vessel annulus, formed by the reactor vessel and the main fuel assembly (MFA). he hydrogen heated in the MFA flows out of the supercritical propelling nozzle into space and creates jet propulsion. he energy conversion system operation based on the employment of one of its turbine-generator installations is the same both in the propulsion and in the energy modes. he helium-xenon composition of the energy-conversion loop is heated in the reactor core with the heat emitted by the main fuel assembly jacket, the pre-heating assembly and the central pre-heating assembly and gets into the turbine, which actuates the electric power generator and the compressor attached to the same shaft. From the turbine, the gas mixture, which has yielded its share of heat in the recuperating unit, gets into the main radiator-cooler, where it is cooled through emitting the remaining heat into space. he compressor feeds the cooled TGI working medium to the input of the reactor core inter-casing space. he other TGI remains inactive and serves to backup the operating installation in case of the latter’s failure. When the NPPU reactor operates in the propulsion and energy modes, the heat-transfer medium (helium) from the auxiliary circuit of the outer components coolant system gets into the auxiliary radiator-cooler and then with the help of the gas-circulator it circulates through the outer reactor units, radiation protection shield, side reflector, reactor vessel, fire floor, and, finally, it leaves the reactor block through the central gas passage of the central fuel assembly and a standpipe in the radiation protection unit. When the NPPU works in the energy mode, the TGI operates at the electric power generation nominal level, while the equipment, which provides the storage, circulation and heating of hydrogen is switched over to the standby mode. In this case, a stop valve shuts off the hydrogen loop from the servicing tank at one end, while engine-exit shut-off system (ESS) cuts it off from the space environment at the other end. he engine working medium passages are filled with neon from the neon tank, which is closed when the NPPU operates in the propulsion mode. he bi-modal NRE thermodynamic balance must be calculated separately for each of the two main operating modes. In the combined energy and propulsion NPPU operating mode (cruise mode), the nuclear reactor thermal power is used for the following types of useful work:

125

• generation of propulsive power, which enables the vehicle to transfer to its

destination; • generation of the electric power to meet the needs of the vehicle support systems

and of the NPPU electrical equipment. When the NPPU operates in the energy mode, the heat generated by the reactor is mostly consumed by its outer thermal-electric heat conversion system. he electric power generated by this system is used for the vehicle useful load performance, as well as for the performance of the NPPU own electrical equipment, and for the performance of the vehicle support systems, which maintain the vehicle and the crew during the flight. he unused thermal energy amount should be estimated separately for each of the NPPU modules – the reactor block and the turbine-generator installation. It should be noted that the heat-release process in the TGI module does not depend on the NPPU operating mode. According to the selected pneumatic and hydraulic system arrangement [Fig. 5.14], the excessive TGI heat is shed into the space environment through the main radiator-cooler. As for the reactor block, the heat release in the process of the outer reactor units cooling in both operating modes is carried out by the auxiliary circulation loop. In this case, while in the operating mode excessive reactor heat is shed into the space environment largely through the auxiliary radiator-cooler and partially through the reactor vessel and radiation protection unit outer surface, in the combined energy and propulsion operating mode (cruise mode), in addition to the above said, part of the excessive heat gets into the working medium through the inner reactor surfaces and in this way participates in the creation of the propulsive power. he main performance parameters resulting from the thermal and hydraulic analysis of the reactor performance are shown in Table 5.7. Table 5.7. Reactor Characteristics Parameter Value Electric output, including self consumption, MW 0.060 Thermal output, MW: – energy mode 0.226 – propulsion mode 340 Total operating time: – energy mode, years 2* – propulsion mode, hours 6* Brayton cycle eiciency factor, % 26.5 Turbine entry temperature, K 1,500 Distance from the reactor to the instrument compartment and hydrogen storage tanks, m 31 *The bi-modal NRE operating time is given on the basis of one interplanetary Earth – Mars – Earth transfer

he reactor block of an individual bi-modal NRE includes a reactor, a shady shielding and a propelling nozzle. he general view of the reactor block fitted out with the radiation protection unit, reactivity operating actuators and the nozzle unit is shown in Fig. 5.15. 126

• • • •

A bi-modular NRE module reactor is a duct-vessel unit with a fast neutronfission spectrum. Its core is formed by two modifications of fuel assemblies accommodated in the regular triangular lattice points. he first modification – the main fuel assembly – serves to heat the engine working-medium – hydrogen, to the temperatures, which ensure the generation of the target propulsive burn. he second modification – the preheating fuel assembly – serves to heat the hydrogen, which then gets into the TNI turbine in accordance with the pneumatic and hydraulic system arrangement. he preheating FA annular cavity has a safety rod, which is installed prior to the reactor delivery to the assembly low-earth orbit, and is taken out before the reactor is activated. he central pre-heating FA is mounted in the reactor core center. he suggested reactor structure consists of four assembly components: ready-assembled core floor; central fuel assembly; combined radiation and nuclear shielding; pressure vessel. he assembly unit reactor design permits to carry out individual or simultaneous experimental engineering development of the reactor units (with regard to their durability, vacuum tightness and resistance) and the prelaunch setting of the hydraulic passage flow. 1- nozzle cluster; 2 – nozzle passage shutof system; 3 – rotary drum; 4 – main FA; 5 – core barrel; 6 – rotary drum liner; 7 – He-Xe composition reactor outlet pipe; 8 – upper delector; 9 – ire loor with insulation; 10 – He-Xe composition reactor inlet pipe; 11 – safety rod; 12 – side relector; 13 – pressure vessel; 14 – pre-heating FA; 15 – central pre-heating duct; 16 – H2 pre-heating FA outlet pipe; 17 – combined radiation protection; 18 – rotary drum drive; 19 – safety rod drive; 20 – shut-of system drive

Fig. 5.15.Bi-modal NRE reactor block (longitudinal and cross section views)

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he fire floor (FF), a two-chamber hollow plate, is the main support structure of the reactor core. he main FA casings and the central liner for the nozzle passage shut-off system (2) are hermetically fastened to the two FF plates with the help of special pressure tight joints [5.27]. he hot gas from the main fuel assemblies is fed into the propulsion nozzle to generate the propulsive burn when the unit operates in the propulsion mode. he arrangement, under which hydrogen flows (after the nozzle cluster cooling) through the top FF chamber, while the auxiliary loop heat-carrying medium (helium) circulates in the bottom FF chamber, provides effective heat withdrawal from the FF both in the propulsion and in the energy reactor operating modes. he nozzle cluster is hermetically fastened to the fire floor rear end through a special adaptor block with the help of pressure-tight bolts. he working medium flow in the main fuel assembly is organized in accordance with Field flow diagram. Hydrogen flows from the top FF chamber through the vessel annulus formed by the flanged vessel walls and the FA casing. Once heated in the pre-heating FAs, hydrogen is removed from the reactor core through outlet pipes (16). he central fuelling assembly includes the core barrel (5) and the side beryllium reflector barrel with their respective in-core elements. he core barrel is a one-layer hexagonal section structure coaxial with the side beryllium reflector barrel. his layout guarantees an even vacuum gap, which reduces heat leaks from the core to the side beryllium reflector cooling passages. Pre-heating fuel-assemblies bottom nozzles (14) with safety rods (11), as well as the central FA with the nozzle passage shutoff system shaft and the central safety rod are fastened to the barrel bottom. he inflow and the withdrawal of the power loop heat-carrying medium are carried out with the help of the inlet and outlet pipes. he side beryllium reflector barrel is a drum-like structure, which consists of the inner and outer manacle rings connected by twelve evenly positioned liners. hey accommodate rotary drums. In the side beryllium reflector barrel, between the rotary drums liners, are beryllium spacers. he rotary drums are cooled by the engine’s working medium. From the outside, the rotary drums liners and the side beryllium reflector spacers are cooled by the reactor auxiliary cooling circuit. In the reactor upper and lower sectors, the core and the side beryllium reflector barrels are connected to each other with weld joints through flexible contraction bellows installed to compensate the linear expansion of the manacle rings provoked by big temperature differences. he pressure vessel is a two-layer structural barrel cooled by hydrogen in the propulsion mode. From the outside, it has a welded shaped manacle ring, which serves to arrange the auxiliary circuit coolant flow passage. he combined radiation protection (RP) structure is based on the so-called ‘monoblock’, which consists of packaged lithium hydride flat layers run through with a supporting frame, which is rigidly fastened to the RS section floor. From the end surface, the monoblock top surface is enhanced with additional sections made from several layers of depleted uranium and coated lithium hydride. he nozzle cluster (1) is a two-casing metal structure cooled with hydrogen. 128

he nozzle cluster structure was modeled after the LPE nozzles, which had undergone experimental research. he performed hydrodynamic and thermo hydraulic calculations showed that the suggested construction of the nozzle cluster can provide the structural materials working temperature of up to 900 K. After the suspension of the NPPU propulsion mode, the fuel elements, in the absence of special measures, will continue operation in vacuum. In this case the heat withdrawal will be carried out through the vacuum gaps of the propulsion section cooling passages only by the heat emission; at the same time, the achieved temperature level of the fuel elements does is not sufficient for their continuous operation in the energy mode. his circumstance resulted in the necessity to take special measures to eliminate fuel elements operation in the NPPU energy mode. Experimental research was carried out to substantiate the need for a shutoff nozzle device, which would serve to prevent fuel elements operation in the vacuum mode during the continuous energy mode phase of the NPPU performance. his work resulted in the development of an experimental installation for the NSS model refinement. he researchers produced and tested sealing elements from the GRAFLEX material with the density ρ = (1.2–1.6) gr/cm3 and studied the tightness of the ‘NSS- nozzle’ pair within the temperature range from 360 to 773 K. he experiments confirmed the correct ness of the suggested technical solutions [5.28]. he main FA is the key functional unit of the reactor core. It consists of the heating sections. he heating sections form the heating block, which is placed into the FA thin-shelled cylindrical casing together with the support and exhaust block and the end reflector. he FA shroud together with the FA casing form the annular cooling passage, in which the heat is transferred to the helium-xenon coolant of the energy-converting loop. As it has been already said, the NPPU works in two modes: the combined propulsion and energy mode, in which it provides the propulsive burn and generates electric power to meet the interplanetary vehicle energy demand, and the energy mode, in which the unit generates electric power. It should be noted that the NPPU operation regimes are very specific. In the first place it refers to a quick reactor switchover from the energy to the propulsion (propulsion and energy) mode, i.e. the unit continues to generate electric power in the propulsion mode [5.29]. • • • •

he following are the bi-modal NRE transit operating modes: the NPPU switchover from reactor subcriticality to the propulsion mode; the NPPU switchover from the energy to the combined propulsion-energy mode; the NPPU switchover from the combined propulsion-energy to the energy mode; scheduled NPPU shutdown.

he transit operating mode, in which the bi-modal NRE is switched over from the propulsion-energy to energy mode, treated as reactor-cooling in the reactor technology, takes a special position with regard to the procedure for the calculation of the required energy consumption and the working medium stock. Hydrogen requested for reactor cooling after the NPPU operation in the propulsion mode forms a considerable part of the onboard hydrogen stock. As a matter of fact energy 129

output continues after the switchover due to the fission-product decay process. In the reactor unit under consideration, this cooling continues until the power output reaches the values characteristic of the energy mode. he results of the evaluation showed that the amount of hydrogen requested for cooling, which is necessary to ensure the total propulsive burn over the total propulsion operating mode period is equal to 3,000 kg. According to the energy and ballistic estimates, the initial mass of the manned IMC will amount to about 770 tons with the total hydrogen consumption ~550 tons. It is necessary to remember that when reactors are run in a cluster they are exposed to reciprocal interference caused by additional external neutron ‘charging’. his interference can be significantly reduced by mounting a borocarbon shield around each of the reactors. Development and operation of space power units with nuclear energy sources demand special attention to safety issues, in the first place, to the issues of radiation and nuclear safety. he modern concept of space NPP safety is based on strict compliance with the regulations for the organization of the NPPU radiation and nuclear safety throughout their life cycle, subject to the requirements of respective national and international regulatory documents. Structural measures, such as the employment of highly effective rotary drums and of an additional safety rods system, optimization of the reactor core configuration and material composition and the maximum improvement of its operating parameters permit to develop a bi-modal NRE modification, which will meet modern requirements for nuclear and radiation safety in all specified NPPU operating modes. Table 5.8 gives a summary of the suggested NPPU mass parameter, and Fig. 5.16 presents the general view of the Martian mission complex fitted out with an NPPU. In this way we have proved that the development of a nuclear engine with the propulsive power of 68 kN and the specific impulse of 940 seconds is a challenging but a viable task.

Fig. 5.16. Martian mission complex

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Table 5.8. Aggregate Mass Parameter of the NRE Nuclear Power Block (NPB) NPPU component Bi-modal NRE 4module cluster Radiation protection Cross-over truss Total mass

Mass, tons 14.2 21.6 0.5 36.1

5.3.4. Cruise nuclear electric propulsion unit with a gas-cooled reactor and a gas-turbine converter For the purposes of interplanetary transfer and maneuver, advanced nuclear power plants can be used as rocket engines to generate propulsive power directly (bi-modal NRE), or, alternatively, they can be employed as power plants with intermediate conversion of heat into electric power for electric propulsion rocket engines [5.21, 5.30, 5.31]. We have already said that a nuclear power plant integrated with electric propulsion rocket engines can be regarded as the second type of a cruise nuclear power propulsion unit. he main characteristics of a cruise nuclear electric propulsion unit were computed for the mission scenario, which contemplated the employment of a manned vehicle and of a cargo vehicle launched to Mars from the radiation-safe orbit on different dates. To comply with all biomedical requirements for the mission duration and for the time of passage through Van Allen belt, the manned vehicle nuclear propulsion unit (NPU) can be formed of three 15 MWel modules. At the same time, one 15 MWel module will be quite sufficient for the cargo vehicle, which will be launched about 720 days before the start of the manned spacecraft. • • • •

Each NPU module will include: power supply system; electric propulsion unit; onboard control system; NPU equipment temperature control system

he presence of a power supply system is the peculiar feature of the nuclear electric propulsion unit. he NPP, which forms an integrate part of the NPU module, includes a nuclear power generating block and the NPP automatic control system. In its turn, the nuclear power generating block consists of: • nuclear reactor, which serves as an electric power source; • power conversion system; • hear withdrawal system, which disposes of unused thermodynamic-cycle heat with the help of a droplet radiator-cooler; • primary structural components; • power supply and information network cable systems.

131

he main parameters of an individual NPU module 15 MW power plant with a closed Brayton cycle gas-turbine installation are given in Table 5.9. Table 5.9. Main Parameters of an Individual Nuclear Propulsion Unit NPP Parameter Electric power output, MW Reactor thermal output, MW Conversion eiciency factor Gas-turbine installation working medium Turbine entry temperature, K Fuel composition Uranium mass 235U, kg Uranium enrichment with 235U isotope, %

Value 15 29.6 0.507 Neon 1500 Uranium carbon nitride 250 90

Heat is generated by a fast-neutron nuclear reactor with spherical fuel elements (FE) in the reactor core. he reactor structure (Fig. 5.17) [5.21] includes: a reactor core, a side beryllium reflector, a pressure vessel, the top and the bottom radiation protection units, reactivity control and nuclear safety systems with their actuators. Technical solutions used in reactor design fully comply with the nuclear safety requirements for space-oriented nuclear installations. To improve the reactor’s thermo-hydraulic characteristics, the designers decided on the structural arrangement with a radial heat-carrying medium admission. he reactor core consists of a ring fuel assembly, in which 3-5mm globular fuel elements made from coated carbon-nitride compositions are either laid in a regular pattern or poured at random between two cylindrical porous molybdenum alloy walls. he ring cavity is divided into sectors each of which has a space for flat safety rods placed there before the reactor is launched to the assembly orbit and removed before its activation. he safety rods are kinematically interlocked into one group and are activated by a common step-motor drive, mounted and fixed at the top end of the top radiation protection unit (3). In the center of the reactor core there is an axial cavity, which serves as the heatcarrying medium return header before the latter is withdrawn from the reactor core. Should the need arise, the return header can be fitted out with an additional fixed safety rod. Further from the center of the reactor core there is a ring cavity formed by the outer cylindrical wall of the fuel assembly and the inner manacle ring of the side beryllium reflector barrel. It serves as a distribution header. In the reactor core, the heat-carrying medium flows in accordance with a pi-circuit arrangement from the periphery to the center. At the end walls of the reactor core there are end-wall beryllium oxide reflectors. he side beryllium reflector barrel is a drum-like structure, which consists of the inner and outer manacle rings connected by evenly positioned twelve liners. hey accommodate the reactor reactivity control devices - rotary drums (7) with sector linings made from a neutron-absorbing material, particularly from boron carbide of a natural isotope composition.

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1 – safety rods drive; 2 – safety rod; 3 – top radiation protection unit; 4 – reactor vessel; 5 – reactor core block; 6 - side relector block; 7 – rotary drum; 8 – bottom radiation protection unit; 9 – rotary drum drive

Fig. 5.17. General view of the NPP reactor block:

Each rotary drum is controlled by its own independent actuator. he RD drivers (9) are mounted in the propulsion module and are connected to the bottom radiation protection unit. Between the side beryllium reflector barrel liners there are beryllium spacers with through paths for the heat carrying medium flow. In the central part of the through paths there are beryllium plugs. A vacuum gap between the reactor core and the reflector barrels helps to reduce the outflow of heat from the core to the cooling ducts. From the outside, the side reflector barrel in enveloped by the two-layer pressure vessel (4). he top and the bottom parts of the reactor pressure vessel are hermetically connected to respective radiation protection units. he reactor core and the reactor nuclear safety physical parameters are optimized so as to minimize the NPPU module mass and size parameters, provide the necessary heat-carrying medium hydrodynamic parameters in the globular fuel elements sector and ensure the maximum possible level of nuclear safety in standard operating modes and in emergency conditions. he bottom radiation protection unit (8), the so called shady shielding, serves to reduce the reactor ionizing radiation, which may effect the IMC equipment, and to ensure the prescribed radiation environment of the IOV living module in full compliance with the established radiation protection requirements: 133

− the integral absorbed gamma radiation doze must not exceed 1.0·106 rads; − the fast neutron fluence from E > 0.1 MeV must not exceed 1.0·1012 cm-2; − exposure dose in the living module protected with the local radiation protection

against cosmic rays, must not exceed 0.1 Sv (10 rem). he key component of the shady shielding is a packaged lithium hydride monoblock run through with a supporting frame, which is rigidly fastened to the radiation protection unit floor (at the bottom end) and the reactor pressure vessel (at the top end). From top to bottom, the monoblock is perforated with special standpipes, which accommodate liners with minimum positive allowance. hese liners, in their turn, serve to accommodate the drum-driver bridges, and peripheral liners for the admission of the heat-carrying medium, which cools the pressurized vessel, the side reflector barrel and the top radiation protection unit. he top end surface of the monoblock is enhanced with central and side sections, which consist of packaged lithium hydride layers alternated with depleted uranium layers. he layers are run through with standpipes at the same points as the monoblock. he vacant space inside the liners is filled up with the same anti-radiation materials. he number of layers in the sections directly depends on whether the reactor will serve for a single Martian mission or will be used repeatedly (on a shuttle-type spacecraft). If the reactor serves for a single mission, two layers will suffice to comply with the radiation safety requirements, whereas a shuttle application will need four layers. It should also be noted that in case of a repeated use of an IMC with the NPPU of the above described configuration, it is necessary to anticipate for the installation of some additional means of reactive power compensation in the reactor core, for instance, shim rods, to compensate a sufficient fuel burn-up caused by an inevitable increase in the power output. he top radiation protection unit (3) serves to protect the safety rods drive and the core fuel elements compressing units from ionizing radiation exposure. he top radiation protection unit is made with the same composition of anti-radiation materials as the shady shielding. his multilayer structure consists of individual blocks with special cavities for safety rods, fuel elements compressing units and coolant flow ducts. Proceeding from the NPPU functional area, which covers electric power supply of the electric propulsion (EP) and of the IMC at large, the nuclear reactor heat is converted into the electric power with the help of a gas-turbine installation (GTI), which is designed in accordance with a two shaft recuperative structural scheme and operates in the closed thermodynamic cycle (Brayton cycle). In Fig. 5.18 you can see the NPPU basic pneumatic and hydraulic diagram, which illustrates the unit’s operating principle. he GTI working medium is neon. From the turbo-compressor compressor unit (1) pressure line, the working medium with the temperature of about 370 K flows to cool the low-temperature components of the reactor structure. After passing by and cooling the reactor vessel, the top reactor RP unit with safety rods, the side beryllium reflector block with rotary drums and the shady shielding RP, the neon gets into the recuperative heat-exchanger (2). here it is heated up to about 1315 K and flows to the reactor core distribution header, from which, through the perforated cylindrical manacle rings it radially flows by the reactor core fuel elements, gets to the return 134

header and flows out of the reactor. From the reactor, the neon, heated to 1500 K is pumped to the turbine-generator turbine (3) and then to the turbine-compressor turbine (4). From the turbines, the gas gets into the recuperative heat-exchanger (2) from the hot side and passing through the heatexchanger ducts, transfers the heat to the gas contained in the heat-exchanger cold ducts. After that, the neon cooled to about 430 K, passes to the heat interchanger (5), in which it transfers the remaining excessive heat to the droplet radiator-cooler circuit and reaching the temperature of 320 K goes on to the turbine-compressor compressor unit (1), thus forming Fig. 5.18. NPPU basic pneumatic and hydraulic diagram: a closed Brayton gas-turbine 1 – compressor unit; cycle. 2 – recuperative heat exchanger; he droplet radiator-cooler 3,4 – turbine; circuit includes a drop generator, 5 – heat interchanger; 6 – pump; a drop receiver, the heat RC – reactor core; SR – side relector; interchanger hydraulic ducts, a V – reactor vessel; SS – shady shielding; pump for pumping the working RP – top radiation protection unit; CR – safety rod medium and manifolds. he droplet radiator-cooler working medium is vacuum oil, because it meets the low temperature heat disposal conditions and low evaporation rate requirements. Circulation of the vacuum oil in the circuit is shown in Fig. 5.18. With the help of the pump (7), the pressurized working medium with the temperature of 310 K is pumped into the heat interchanger (6) , where it is heated to 369 K and ousted to the drop generator. From the drop generator, which is a perforated chamber, the working medium is injected under pressure in the direction of the drop receiver in the form of trickles, which burst into drops under the influence of the surface tension force. While the drops fly into the receiver they emit heat into the space environment. he collected working medium is transported to the pump, from where it again gets into the heat interchanger. his closed-cycle system provides high heat removal parameters, has a large radiating area and a low mass parameter. It is necessary to mention that currently, technologies requested for the realization of the suggested IMC version components are at different degrees of development. hus, production technologies for such critical items as fuel elements and fuel assemblies, reactor control system regulatory devices and radiation protection blocks were developed in the USSR within the framework of the nuclear rocket engine development program. A specific technological issue, which needs 135

experimental verification, is the development of a droplet radiator-cooler, whose bright prospects, especially with regard to megawatt power consumption, have been described at length [5.32-5.41]. Work on this issue is now being carried out within the framework of the research directed at the selection of the basic design engineering solutions. Obviously, the mass and size parameters of the Martian mission complex at large and of its nuclear power and propulsion unit in particular will be determined by the rational arrangement of the equipment and units within each mission module. he launch configuration of a manned interplanetary vehicle is shown in Fig.5.19. he reactor, the radiation protection unit and the propulsion module are structurally united into the so-called nuclear power block (NPB). he propulsion module, placed directly behind the shady shielding, accommodates actuators of the reactor control system regulatory devices, all installations and heat-exchange equipment. he estimated operating overall dimensions of the droplet radiatorcooler prove the necessity to have a deployment system as part of the vehicle NPPU. It should provide on-orbit NPB shifting and radiator-cooler deployment, since there are certain limitations for the overall dimensions of the launch vehicles, which deliver the IMC components to the assembly orbit. he NPPE subsystem modules, blocks and components are delivered to the assembly orbit under launch vehicle faring. he launch vehicle is 6.5 m in diameters (the useful load diameters must not exceed 6,000 mm), 22 m long (the cylindrical part is 20 m long). It can take on not more than 31,000 kg of useful load under the nose fairing and 35,000 without the nose fairing. Once on the orbit, the NPB is shifted away from the NPPE with the help of the system, which forms part of the NPPE and which actuates the NPB deployment. his operation is necessary because of the overall dimensions of the droplet radiatorcooler receiver unit, which should operate within the radiation protection shadow cone. In Table 5.10 you can find the main characteristics of a manned IMC power and propulsion system (PPS). NPP

Radiation protection Radiation protection shadow cone

Electric propulsion

EP propellant storage tanks

Trickle-drop receiver

Bearing structure ADV-35 IMC ADV-35

Trickle-drop generator Power block Support systems module

Fig. 5.19. Manned interplanetary complex with a turbo-machine cycle and a drop Reactor-cooler, EPU

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Table 5.10. Main Characteristics of a Manned IMC Power and Propulsion System PPS key assembly units Reactor plant Turbine generator Turbo-compressor Recuperative heat-exchanger Heat interchanger Droplet radiator-cooler Compensation volume tank for the droplet radiator-cooler circuit Multi-section bearing structures Control, protection and diagnostics system

Number of units 3 3 3 3 3 3 3 3 3

Assembly units total weight, tons 13 12.7 4.2 26.5 18.4 19.9 5.4 9.5 1.5

he mass characteristics given in Table 5.10 prove that one launch vehicle can deliver one NPPE module to the NPB assembly orbit. It means that an integrated ground functional check of NPB systems can be carried out before its launch to the low-earth orbit. he energy and ballistic analysis of the Martian complex within the framework of the suggested cruise electric propulsion unit shows that the mission duration of a two vehicles mission scheme will be equal to ~ 328 days. In this case, at the launch position it will be necessary to have onboard about 580 tons of the PPS working medium (xenon), and the total PPS weight will amount to ~ 700 tons, including the mass parameters shown in Table 5.10. Fig. 5.20 shows an interplanetary vehicle, which uses a power and propulsion system based on a Fig. 5.20. Interplanetary vehicle with a 50 MW NPPU closed gas-turbine cycle and a panel-type radiator-cooler with the temperature of about 650 K [5.42]. 5.4. Version of Propulsion System Based on Solar Cell Arrays and Electric Propulsion 5.4.1. Substantiation of the solution on the selection of the solar array type he results of the first studies of a manned mission to Mars based on the employment of high-power thin-film solar arrays and electric propulsion undertaken by S.P. Korolyov NPO ‘Energiya’ date back to 1988 [5.43]. Two configurations were presented in this work: one of them contemplated a rigid honeycomb structure with a thin-film face-sheet; the other was based on the use of the centrifugal force, i.e. it considered the rotation of the whole array as a means of maintaining its flatness (flexible blanket). At the same time, the researches continued work on other types and configurations of solar arrays. Today they can be summarized as follows: • a panel solar battery based on the employment of crystalline silicon, gallium arsenide (GaAs), or triple photocells of the GaInP2 / GaAs / Ge type; 137

• a panel solar array based on the above mentioned crystalline photocells, enhanced

with a solar concentrator; • a honeycomb-structure solar array with a thin-film face-sheet from amorphous

silicon alloys; • a thin-film flexible-blanket-type solar array with the centrifugal formation of

the flat surface; • a solar array with a solar concentrator and machine solar energy conversion.

In Table 5.11, you can find comparison characteristics of the photocells based on the technologies utilized by NPP ‘Kvant’ (Russia) by 2006. Table 5.11. Comparison Characteristics of Photocells at Nel = 32 kW and the output voltage of 120V (as received on 01.01.2006) Type of photocell Single-crystalline silicon

Total surface Total mass of Total cost of Eiciency, of photocells, photocells, photocells, % m2 kg mln USD 15 200 400 8

Gallium arsenide

26

100

200

12

Amorphous silicon alloys

8,6

300

110

3

Development prospects Tests of experimental models showed the eiciency factor of 17% Tests showed the prospects of achieving a 35 % eiciency through employment of new materials (USA) The latest developments include photocells with the EF of 13% (USA) and a substrate layer with δn =7 μ (USA)

he tests of amorphous silicone photocells in the space environment carried out on board ‘Mir’ orbital station for two years and on board ‘Tatiana’ unmanned spacecraft (as of today the tests have been carried out for over a year), Fig. 5.21, have shown that: − amorphous silicone photocells temperature dependence is half as much as that of single-crystalline silicon elements; − radiation degradation is lower that that of single-crystalline silicon elements. Table 5.12. World Market Price Movement with Regard to 1W of Rated Capacity Received at a Ground-Based Solar Power Plant, USD per 1 W Ground-Based Solar Power Plant using: Single-crystalline silicon Ribbon silicon Thin-ilm amorphous silicon

Years 1990 5.40 6.0 5.0

1995 5.0 5.0 4.5

2000 4.5 4.0 3.5

2010 3.0 2.50 1.5

Currently, the progress in the ground-based solar power generation industry is connected with the employment of third generation silicon photocells [5.45, 5.46]. In the last seven years , the global solar array market has been steadily growing by 30–40% annually. he leading international producers of the thin-film photocells are: Kaneka, United Solar, Mitsubishi, First Solar and Antec. he first three companies 138

Efficiency (%)

use amorphous silicon, while the remaining two producers have given preference to cadmium telluride. A number of companies are engaged in the research aimed at the achievement of a 15% or more efficiency of amorphous silicon photocells. he research concept is based on the idea of creating a quantum points (‘holes’) matrix in the amorphous silicon structure. his matrix is expected to provide resonant tunneling and increase the current carriers’ (electrons and ‘holes’) mobility. he work [5.45] forecasts the market appearance of such photocells in 2008–2009, and contains estimates of the ‘cost-efficiency’ limits on the photocells market, Fig.22. In Chapter 3, we gave the principal results of the ballistic calculations for a manned mission to Mars. hey showed that the requested electric power of 15 MW and the power and propulsion system mass to power ratio of not more than 5 kg/ kWel will be the same for both propulsion variants: ‘nuclear power plant + electric Fig. 5.21. Photocells unit on board ‘Tatiana’ SC propulsion’ and ‘solar power plant + electric propulsion. As it is, the main shortcoming of solar arrays is their big surface. US $ 0.10/W US $ 0.20/W US $ 0.50/W he following 100 problems arise in this Thermodynamic connection: limit 80 • achievement of the requested structural 60 US $ 1.00/W rigidity of the solar Expected limit III arrays and of the 40 whole power plant and propulsion 20 US $ 3.50/W I system at a given II mass perfection; 0 100 200 300 400 500 • delivery of the solar array modules ‘Cost-eiciency’ limits for the three generations to the low-earth, of solar arrays production technologies (as of the USD assembly orbit and rate in 2004) their assembly at the I generation – single-crystalline silicon slats interplanetary vehicle level; II generation – thin-ilm amorphous silicon • achievement of III generation – perspective thin-ilm silicon the solar-pointed Fig. 5.22. ‘Cost – efectiveness’ photocells market forecast attitude; 139

• risk of the solar arrays damage with meteorite chunks and space debris in the

space environment; • solar array cost.

he solution of the above problems needed a thorough analysis of solar arrays types and configurations. he use of gallium arsenide solar arrays will lead to a dramatic increase in the cost of the program, since the cost of the arrays requested for the rated power of 15 MW will be by 4 times higher than the cost of amorphous silicon solar arrays with the same power output. In absolute figures the difference may amount to 4.2 bln USD, whereas the total acceptable cost of the first mission is estimated at 16 bln USD (as of the USD rate in 2005). In this way, the share of the solar arrays will amount to 26% of the total cost, although their share in the total vehicle weight does not exceed 12%. On the other hand, although the power and propulsion system modification, which contemplates the use of amorphous silicon solar arrays, will be the cheapest, it will demand the biggest solar array surfaces, about three times as big as those of the gallium arsenide version. he total surface of amorphous silicon solar arrays will amount to 115,200 m2, against 38,400 m2 of their gallium arsenide counterparts. However, due to their big surfaces, both modifications present difficulties with regard to the achievement of the target rigidity. To reduce the cost of gallium arsenide solar arrays it was decided to modify the whole structure by introducing solar energy concentrators with the concentration factor K = 50–150. In this case the amounted of the requested expensive gallium arsenide reduces by the factor of K. In Fig. 5.23 you can see various configurations of solar concentrators. Fig. 5.24 shows an experimental panel with a gallium arsenide photocell and with the concentration factor K = 50. his panel was successfully tested on board the ‘AUOS-SM’ unmanned spacecraft. he test results showed that by developing a photocell for a given K value it is possible to increase the array efficiency by 4–5 % absolute. At the same time, the use of concentrators, especially with a high K value, may lead to additional problems: • it will increase the requested solar-oriented attitude accuracy to α = ±1°, while standard solar arrays need the accuracy of α = ±30°; the high solar-pointed attitude accuracy of the solar arrays will demand for a more rigid structure of the whole power and propulsion system and will lead to its heaviness; • it will demand thermal control of the photocell, since the increase in the K value leads to the increase in the photocell temperature and to the decrease in its efficiency; this factor is especially important in the context of a Martian mission, during which, on the return transfer to Earth, the vehicle passes at a distance of 0.57–0.58 a.u. from the Sun, when the solar radiant density triples as compared to its value near Earth – qsol = 1350 W/m3. hat is why this solar array configuration is not considered acceptable for a manned mission. he studies showed that at the requested power and propulsion system mass perfection of 5kg/kWel., almost 2 kg/kWel. fall to the share of the bearing structure, which bears the solar arrays, voltage converter modules, cables and actuators of 140

the orientation and stabilization system (OSS). To reduce this share (almost 40%), the designers studied the modification, which contemplated the use of the centrifugal force as the means of forming and maintaining a flat solar array [5.47], i.e. a frameless solar-array structure. Estimates were made for the configuration shown in Fig.5.25. Rk is the inner radius of a circular solar array, ω1 is the solar array angular rotational speed. At Nel. = 15 MW and the steel substrate thickness δn =12 μ, we get that Rk =114 m; R0 = 11.4 m;

ω

radian sec

Relection Selfindex shadow (light factor trans. factor)

LENS LFTI

FOCON KVANT, FOTON

Concentr. Radiation PV SolarSpeciic ratio Attenuation element pointed power factor Temper attitude acc., W/m3 °С degrees

0.85

1

0.9-0.95

1

2÷10

0.9

0.8-0.9

Speciic weight, W/m2

80-100

± 1.5





10

40-60

±7

150

10-30

50÷1000 50÷100

PARABOLOID KVANT 50÷6000

2-50

80-100

± 1.5





CASSEGRAIN SYSTEM VIKI

0.73-0.77 0.85-0.92 50-8000

20

90-125

± 0.5

82

16

OFF-AXIS WITH FACETS NIITP

0.88-0.9

0.98

50÷200 100÷200

90-150

± 1.5

200

30

0.88-0.93

0.99

90÷300 100÷300

100

± 1.0





3

OFF-AXIS WITH DUAL FACETS NIITP PLANAR SILICON PANEL KVANT

SiO2 Si

0.1

1

1

5

25-50

± 30

140÷160

30-40

PLANAR GALLIUM ARSENIDE PANEL KVANT

SiO2 GaAs

0.1

1

1

5

25-20

± 30

140÷220

30-50

Fig. 5.23. Solar system energy concentration geometry and parameters

rps

2 rpm

he angular speed ω1 was found from the condition that the centrifugal force overrides Coriolis force, which is brought about by the solar array orientation and stabilization maneuver carried out to achieve the right attitude with regard to the Sun. It was assumed that the solar array axis variation, brought about by the influence of Coriolis force must not exceed 3°. On February 4, 1993, ‘Energiya’ RSC carried out ‘Znamya’ (banner) experiment on the low-earth orbit near ‘Mir’ orbital station. he experiment consisted in Fig. 5.24. Experimental photovoltaic panel unfolding a thin-film structure with the radius Rk = 10m, using the centrifugal force. In Fig. 5.26 you can see the photo of the unfolded structure made from ‘Mir’ orbital station. It is clearly seen that the split petals failed to unfold in the tangential direction. he designers came to the conclusion that a thin-film structure should be made one-piece to employ the tangential force. However, in this case the unfolding process becomes more complicated and requires additional study. It must be noted that when the spacecraft rotates with the angular speed ω1 = 2·10-1 rad./ sec., the centrifugal acceleration of g0 = 9.81 m/sec2 can be achieved at the radius rg0 = 9.81/w2 ~200m. his fact can be regarded as a positive factor if the living module with the crew quarters is placed on the said radius. 141

Another important factor, which should be considered in connection with this configuration is the spatial arrangement of voltage converter modules. he maximum voltage generated by the solar arrays must not exceed 120V, since in the space environment, a higher voltage may provoke electrical discharges. At the same time, electric propulsion engines, Fig. 5.25. A frameless solar array coniguration especially ion engines, operate at 3,000–6,000V. Another advantage of high voltage is that it helps to reduce the weight of the cable assembly in large-size solar array structures. Besides, voltage converter modules must be placed as close to the generating photocells as possible. All factors considered, the voltage converter modules arrangement is of a paramount importance to this configuration of the solar array, since their specific weight is comparable to that of the photocells. One of the possible solutions for the centrifugal solar array arrangement configuration is the integration of electric power generation with its conversion into UHF radiation followed by its transmission to the Fig. 5.26. Unfolded thin-ilm structure (picture made from ‘Mir’ orbital station) propulsion units and reverse conversion. he practical aspects of this issue are currently being investigated. An attempt to reduce the solar array surface, lead to the idea of solar power plants, which use solar energy concentrators and a solar energy receiver. he latter is a heat exchanger with a neutral gas as a working medium, which is heated by the concentrated solar energy, and performs useful work in a closed Brayton or Stirling cycle. Fig. 5.27 shows the Brayton cycle diagram, developed for a power and propulsion system module of a manned Martian mission with the aggregate capacity of 6MW. Fig. 5.28 shows a schematic view of an interplanetary vehicle with 8 parabolic-type concentrators, 38m in diameter each. Although a gas-turbine installation can actually Fig. 5.27. SGTU-750 basic diagram. Parameters for the Mars/Earth orbit conditions N =345/750 kW el., ηΣ =0.60/0.57 ensure a rather high efficiency 1 142

57–60%, the request for a high solaroriented accuracy (the deviation permitted for this configuration is ±3°) makes it unacceptable. he measures taken to achieve the specified degree of accuracy will considerably increase the weight of the bearing structure. Proceeding from the results of the studies the thin-film amorphous silicon alloys configuration of the solar array was suggested as the basis for further development.

Fig. 5.28 Schematic view of an interplanetary vehicle with 8 parabolictype concentrators

5.4.2.Configuration and main characteristics of a power and propulsion system he configuration of the power and propulsion system (PPS) based on the employment of thin-film solar arrays with amorphous silicon photocells was defined proceeding from the results of the functional and morphological analysis of its components. he final version included: • power source; • energy conversion and voltage control system; • commutation and power distribution system (cable system); • bearing structure; • structural stiffness setting system; • power users, including electric propulsion rocket engines (EP); • EP conversion and control system; • EP propellant (working medium) storage and feed system. Fig. 5.29 shows the division of an integrated solar power plant (SPP) into structural components. When developing the system composition, special attention was paid to the modularity principle, according to which the vehicle PPS is built up from system units of optimal dimensions. SPP Conversion and Control System

Solar Array (SA) SA Wing

SA Wing

Module

Module

Module

Module

Section

Section

Photocell

Photocell

Cable System Bearing Structure Truss Setting system

Beams

Fig. 5.29 The sceme of solar powwer plant division

143

Crew return vehicle (CRV) he optimal dimensions of the system units were determined proceeding from the following criteria: • convenience of ground development considering the state of the existing tester units; • convenience of Solar tug delivery to a lowInterplanetary orbital vehicle (IOV) earth orbit • possibility of effective Fig. 5.30. General coniguration of an interplanetary mission complex (IMC) robotic assembly on the low-earth orbit; • assurance of the target PPS reliability; • financial viability in the context of maximum cost-effectiveness.

In Table 5.13 you will find the main characteristics of a one-vehicle interplanetary mission complex scheme, which does not contemplate a landing on the Martian surface. he general configuration of the interplanetary vehicle is shown in Fig. 5.30. In Table 5.13, the marked out values still need special mathematical and experimental justification, including flight tests. Fig. 5.31 shows an assembly of 22 series-connected photocells with a 25 μ steel substrate. In Fig. 5.32 you can see a 5.6m-long model solar array. Fig. 5.33 features thin-film diodes used in solar arrays. he photocells, represented on this page, have the efficiency of 8.6%. he required 10% efficiency level can be achieved by : • selecting the photocells with the required efficiency level; • introduction of new technologies mentioned in subsection 5.4.1.

Fig. 5.31. Assembly of 22 series-connected photocells

144

Fig. 5.32. Model of a solar array section (Σli=5.6 m)

Fig. 5.33. Polyimide ilm diode

Table 5.13. Main Characteristics of a One-Vehicle Interplanetary Mission Complex Scheme without a Martian Landing (as of 01.01.2006) Unit No

Name of the unit or system

Symbol

1

Interplanetary orbital vehicle

IOV

2

Power and propulsion system 2.1. Solar array

PPS SA

2.2. Solar array voltage converter modules

2.3. Bearing structure 2.4. Propulsion modules with EP-50 (500 EP-50 engines)

2.5. EP-50 working medium storage modules (14 modules) 2.6. Orientation and stabilization systems 2.7. PPS automated control system

Characteristic 1. Initial mass on the low-earth orbit (inc. propellant) 2. Crew size (in prospect)

1. Total power output on the near-earth orbit 2. Total SA surface 3. Eiciency factor 4. Total mass 5. SA mass-to-power ratio VCM 1. Voltage converter eiciency 2.Total mass with argon (xenon) as the EP working medium (propellant) 3. Mass-to-power ratio BS 1. Total mass 2. Mass-to-power ratio PM 1. Speciic impulse control band (for Ar) 2. EP-50 eiciency 3. Total mass 4. Mass-to-power ratio WMM 1. Working medium 2. Working medium total mass 3. ‘Dry’ storage tanks total mass OSS Initial mass on the low-earth orbit (inc. propellant) PPS ACS

Mass

3

Martian descend module

MDM

Initial mass on the low-mars orbit

4

Crew return vehicle

CRV

1. Initial mass on the low-earth orbit (inc. propellant) 2. Crew size (in prospect)

The total initial mass of the Interplanetary Mission Complex on the low-earth orbit

IMC

IMC total mass

Value, dimensions 60 tons 4 persons (6) 15 MW 115, 200 m2 10% 15 tons ~1kg/kW, el. 97% 15 (20) tons ~1kg/kW, el 30 tons ~2 kg/kW, el. 6800-9000s > 65% 15 tons ~1kg/kW, el. Argon 200 tons 6 tons 6 tons 4 tons 0 tons for the irst mission 15 tons 4 persons (6) 366 tons for a scenario without a landing on Mars

Modern technologies permit to produce photocells, which have a δn = 20– 25 μ steel substrate with a 1μ-thick amorphous silicon film upper layer. his type of substrate results in the photocell mass-to-power ratio of 1.57–1.96 kg/kW. Currently there are two possible ways of reducing the photocell mass-to-power ratio: • changeover to nickel band with the thickness δn= 10-12 μ; • changeover to titan band with the thickness δn= 20-25 μ; the current technology

permits to produce 30 μ thick titan band.

145

Fold 1 without a solar cell

Unfolded position of an SS module (back view)

Total fold thickness Transport position of an SS module Crosscut view of the qualiied semi-unfolded SS position

F0ld 103 with a solar cell Fold 104 without a solar cell

Fig. 5.34. General view of a standard 15 kW SA section

he general view of a standard 15 kW section is shown in Fig. 5.34. A solar array generates the voltage of 120V, while the argon-based EP requires 2,600V. he onboard commutation and power distribution system also needs high voltage to ensure that its share in the total PPS mass does not exceed 1%. hat is why the system includes voltage converter modules, which are installed next to the solar battery section and can boost Fig. 5.35. Coniguration of a BM-12 voltage converter module (VCM) Mass – 1480g; the DC voltage from 110V to 2,600V. Overall dimensions – 240×140×35mm; Fig. 5.35 shows a voltage converter Volume – 1.2 dm3; module unit with the efficiency factor Nominal power – 1.2 kW; of 94% (the requested efficiency factor Peak power output – 1.5 kW; Eiciency – 94% must be at least 97%). he efficiency factor can be improved with the help of the following measures: • changeover to a synchronous transistor AC/DC converter; • transition to a new voltage transformers printedcircuit technique; • changeover to a new-brand stranded conductor he following should be said with regard to the bearing farm. According to the estimates, its massto-power ration should be less than 2 kg/kW.el. he truss must be capable of bearing an overload of up to 7·10-2. In the normal EP operating mode, the overload will vary from 7·10-5 at the beginning of the mission to 18·10-5 at the end of the mission. hat is why the suggested parameters of the bearing structure have a certain mass redundancy, which can be used to advantage after making allowances for off-nominal situation. 146

Fig. 5.36. A model bearing structure component (Σli = 12 m)

he bearing structure consists of: • trusses, each of which has a 4×4 m square cross section and is assembled from

longitudinal pipes with the outer diameter of 206mm and the wall thickness of 2mm made from ‘UKN-500’ carbon fiber-reinforced plastic. A 12 m-long longitudinal pipe is shown in Fig. 5.36. • 0.5×0.5×0.5 m delta-cross-section beams, refer to subsection 3.2.6. for details he PPS assembly on the low-earth orbit starts with the assembly of the bearing construction carried out with the help of assembling robots and special jigs. 5.4.3. Electric propulsion unit 5.4.3.1. Selection of the EP type One of the works [5.48] reviews the progress made with regard to the development of different types of cruise electric propulsion engines, including: • arc-jet engine (AJE); • magnetoplasmodynamic engine (MPD); • Hall EP, including SPE and EPAL engines; • ion EP engine Designers at Keldysh Center have developed a project design of a plasmatron with the power of 1 MW, the chamber temperature of 6,000 K and the endurance of 500 hours, which uses air as the working medium, and which can serve a prototype of an arc-jet (plasma-jet) engine. Currently they are working at a project design for the development of a 6 MW AJE. Starting from the late 1950s and up to 1975, Keldysh Center carried out research on MPD rocket engines with the power ranging from hundreds of kilowatts to 1MW and with the specific impulse of 5,000–7,000 seconds, which used lithium or potassium as their working medium. Flight tests of MPD with Nel. = 5 kW were carried out on board spacecraft ‘Cosmos-728’ and ‘Cosmos-760’ to study the problems of MPD-spacecraft integration. Hall EPs have become especially popular in Russia, which is particularly true with regard to SPE-100 EP (stable plasma jet engine), developed by ‘Fakel OKB’. his engine has Nel.= 1.35 kW and Isp. =1,500 sec. More powerful engines are in the process of development. One of them is a xenon-based SPE-140 EP, which will have Nel.= 4.5kW and Isp. =2,000 sec. Xenon is believed to be capable of providing a specific impulse of up to 4,000 sec. his project is currently being developed at CNIIMASH under the name of ‘D-200’ – electric propulsion with anodic layer (EPAL). According to another work [5.49], designers are planning to achieve Isp. = 7,000 sec., using a two-stage EPAL configuration. However, the problem of the engine operational endurance needs special technical solutions [5.50]. In Russia (then, the USSR) successful short-run flight tests of an ion EP were carried in 1968 on board the ‘Yantar’ ionospheric laboratory. he EP working medium was argon. And the engine achieved Isp. = 4,000 sec. During further tests the nitrogen-based EP demonstrated Isp. =12,000 sec., and the air-based EP even demonstrated Isp. = 14,000 sec. Currently, American designers have been the most successful in the development of ion EPs. hus, an ion EP developed within the framework of the ‘NSTAR’ program 147

worked for 16, 265 hours as the ‘Deep Space-1’ spacecraft cruise engine. It had Nel. = 2.3kW, Isp. = 3.170 sec and used xenon as a working medium. he European Space Agency (ESA) is working at the project of an ion ‘ESA-XX’ electric propulsion unit with radio frequency ionization, Nel. = 6kW, Isp. = 5,000 sec, which will use xenon as a working medium. In Japan, an ion electric propulsion unit with UHF ionization is currently undergoing flight tests. Previously, in bench-run tests it has already demonstrated an18,000- hour endurance. By June 2005, the engine had worked 10,000 hours. he results of the ballistic estimates (see Chapter 3) showed that an electric propulsion unit for a Martian mission should meet the following requirements: • working medium: xenon or argon; • EP-50 individual module power 25–50 kW; • running time in flight ~ 10,000 hours; • design endurance redundancy margin 1.5; • in flight controllable EP specific impulse: for xenon EPs 5,000–9,000 sec.; for argon EPs 6,800–9,000 sec. ; • EP control algorithm: − for xenon EPs: propulsive power stability (P = const); − for argon EPs: medium consumption stability ( = const). Arguments in favour of different EP types are given in Table 5.14. he measures required for the achievement of the target parameters are given in Table 5.15. Table 5.14. EPs Comparison Characteristics No Type

Argument for its employment in the program High eiciency, which increases with the increase of accelerating potential. Good endurance. Good predictability of speciic parameter. Unlimited peak voltage.

1

Ion

2

EPAL Multifunctional propulsion unit with high eiciency at the extreme points of the speciic impulse band. Unlimited peak voltage. Easily scaled. Low efect on the onboard systems. Simple design. Low-power engines produced in Russia for 30 years. High endurance was demonstrated at the set point of 300V.

3

SPE

4

MPD High propulsive power density. One module has the power of 1MW.

148

Arguments against its employment in the program Propulsive power augmentation becomes problematic at reduced speciic impulse values. Structural complexity of the power supply source. Low propulsive power density. A drop in eiciency when using working mediums with the atomic mass value below 80. Insuicient volume of ground endurance tests. The development of a 50Kw EP is connected with a number of technological problems. The problems connected with the achievement of a high speciic impulse seem unsolvable. Russian projects are based on the use of lithium and potassium. Low eiciency.

Remarks In ground light tests ion EPs have demonstrated the endurance of over 10.000 hours The 25 kW EPAL prototype operation has been demonstrated at low and high speciic impulse values. The operating voltage of the most high-performance engine – SPE-290 – is limited by 700V. The laboratory model has been tested at 400 kW

Table 5.15. Measures Required for the Achievement of the Target Parameters EP type, Endurance No w. medium reached, hours 1

Ion, Xe

30000

2

Ion, Ag EPAL, Xe

15 000 *

3

4 5

6 7

Speciic impulse reached, seconds 3200

Measures required for the achievement of the target vehicle engine parameters

Remarks

It is necessary to develop a xenon engine with the Propulsive power power of no less than 50kW and demonstrate its augmentation capability is endurance and propulsive power variation capabilities. conceptually limited. Current projects are limited by the engine power of 30 kW 7000–9000 It is necessary to develop a 30-50 kW argon engine

≥ 5000*

3500

EPAL, Xe SPE, Xe

≥ 5000*

7000

9000

1600

SPE, Xe MPD, Li

data is not available 1000

> 3100 4000

It is necessary to demonstrate the endurance of >5,000 hours under the conditions of the full cyclic graph reproduction at the extreme points of the speciic impulse band ** It is necessary to conirm the two-stage EPAL coniguration It is necessary to develop a physical rationale of the engine operating capability at high (>3,000 sec.) speciic impulse values It is necessary to develop an engineering model to justify the parameters It is necessary to achieve the engine endurance exceeding 1,000 hours and conirm the possibility of increasing the eiciency to >50% by changing over to argon

The maximum reached value does not exceed 3,100 sec.

The central cathode erosion reaches critical values in the maximum propulsive power eiciency mode.

* design estimate **All EP types require modernization of testing facilities

5.4.3.2. Exploratory tests of a 30-cm argon ID-300 ion engine laboratory model To evaluate the basic operational parameters of an individual EP-50 module, it is necessary to set the cost of ion and gas efficiency. Unfortunately, few argon ion engines have been developed internationally. his statement is especially true with regard to big-diameter ion engines. Some experiments with a 5cm ion engine have been carried out in Japan [5.51]. American engineers have tested a 12 cm ion engine [5.53]. In Table 5.16 you can find the data on the efficiency of different gas discharge chambers (GDC). he table includes the data on xenon, krypton and argon engines. It has Fig. 5.37 Laboratory model sceme for ID-300 ion engine already been mentioned that only data on low-power engines are available with regard to argon. he GDC efficiency grows proportionally with the increase in the engine power. 149

Table 5.16. Ion Engines Eiciency Depending on the Working Medium Beam diameter, mm 50 120 280 350 400 500* * - design parameters

Power, kW ~0.1 ~1 2.3 3.3 7.3 3-11*

Ar Ci, V/A 520 280–350

Кr ηg 0.53 0.7–0.78

Xe

Ci, V/A 500

ηg 0.66

325*

0.89*

Ci, V/A 465 200–300 174 140 150–200 ~250

ηg 0,75 0,85–0,95 0,94 0,9 0,85–0,92 ~0,9

Reference 5.51 5.53 5.54 5.55 5.56 5.52

Keldysh Center in cooperation with MAI (Moscow Aeronautical Institute) developed a laboratory model of a 30cm xenon ion engine with the nominal power of 2 kW and the specific impulse of 3,000 sec. To obtain the missing data required for the development of the EP-50 , this model was tested employing argon within the specific impulse band of 5,000–7,000 sec. he ID-300 gas discharge chamber (1. Fig.5.37) has the form of a cylinder with a conical back wall. Anodes (2) are fixedly attached to the side and the conical GDC walls via insulators. Magnetic field is generated with the help of electromagnets (3) placed outside the gas discharge chamber. he magnetic field pattern is determined by three pole pieces (4). he cathode pole piece contains a cathode pack (5), formed by a diaphragm hollow cathode. Lanthanum hexaboride is used as an emitter. he working medium is let into the gas discharge chamber through a header (6) located next to the ion-optic system (IOS) (7). About 10% of the total medium consumption gets into the chamber through the cathode. he ion-optic system extracts the ions from the gas discharge chamber plasma and then forms and accelerates an ion beam. he IOS consists of an emitting (8), accelerating (9) and decelerating (10) electrodes. he engine uses a three-electrode IOS, in which the decelerating electrode has the form of a ring, embracing the whole beam. he emitting and the accelerating electrodes are made from titanium and are 0.5 mm and 1.0 mm thick respectively. In the emitting electrode, 3.0mm apertures alternate with 0.6 mm links, and provide transparency at the level of 0.6. Apertures in the accelerating electrode have a diameter of 2.0 ( transparency – 0.28). he electrodes are shaped as big-radius sphere segment and have an outwardly directed initial deflection. he initial deflection of the accelerating electrode is bigger than that of the emitting electrode. hat is why, when the electrodes are cold, the gap between the electrodes in the central IOS section is by 0.5–0.8 mm bigger than on the periphery. However, as in operation the temperature of the emitting electrode is higher, the gap between the electrodes becomes almost homogeneous Fig. 5.38. Laboratory model he pressure inside the vacuum chamber of an ID-300 ion engine -4 of a running engine was equal to ~10 Torr. 150

Performance prediction calculations were adjusted for the backward current of argon atoms from the vacuum chamber to the gas discharge chamber. For this particular IOS geometry this value amounted to 725·p·104 mA equiv., where p is the pressure inside the vacuum chamber, measured in Torrs. During the experiments, the positive terminal of the main high voltage power supply source was hooked up to the IOS emitting electrode, which, in its turn, was electrically connected to the gas discharge chamber cathode. his connection diagram is more convenient at the stage of the engine engineering development. However, when analyzing the output characteristics, it is necessary to remember that ions are generated within the GDC volume, where the potential value is close to the anode potential, and that they pass though the voltage difference set up by the ‘Discharge’ and ‘Emitting Electrode’ voltage power supply sources. hat is why the calculation of the engine’s basic integral parameters was carried out with the help of the following formulas:

P = Ib Ci =

2 m(U b + U d ) e

I dU d Id

−Ud ,

,

I sp =

ηg

2 e (U b + U d )

g

m

ηT = η g

I b (U b + U d ) I bU b + I d U d

,

,

where Id and Ud are the discharge current and voltage; Ib and Ub are the current and voltage in the ‘Emitting Electrode’ voltage power supply source; P is propulsive power, Isp is the specific impulse; ηg is gas efficiency; Ci is the ion cost; ηm is the propulsive efficiency; m is a xenon ion mass; e is an electron charge and g=9.81 m/ sec2. Calculations of the laboratory model integral parameters were made without regard to power expenditures required for the magnetic system operation, which amounted to 50–80 W. At this stage, magnetic system optimization aimed at the reduction of its power consumption was beyond the scope of the project group research activities. At a more advanced stage of the project development, power consumption can be reduced through the reduction in the number of solenoids and their diameter. Alternatively, it is possible to change over to permanent magnets. Experiments were conducted without the employment of the cathode neutralizer. he results of the experiments carried out with an ID-300 engine, which ran on the argon working-medium, are given in Table 5.17. he following comments should be made with regard to the said results. he experiments showed the gas efficiency of 0.56–0.78 at the ion cost of 215– 295 W/A. he gas efficiency of the values higher than 0.70 can be achieved at high (over 50V) discharge voltage parameters. However, the discharge voltage boost will lead to the increase in the atomization rate of the GDC surfaces under the cathode potential and reduce the engine service life. Unfortunately there are no reliable data on the cathode-atomization yield for the values below 100electronvolt. herefore, it’s impossible to predict the GDC endurance. It should be noted that high pressure in the vacuum chamber during the test performance results in the increase of the 151

gas efficiency error. Consequently, the obtained results should be confirmed by the results of the tests carried out with the employment of more efficient exhaust facilities. In the table below Ub is the emitting electrode potential; Ib is the ion beam current; m is the working medium consumption; Ud1 and Ud2 are discharge 1 and 2 voltage respectively; Ci is the ion cost; ηg is gas efficiency; P is propulsion; Isp is specific impulse; ηm is the propulsive efficiency and N is the engine power. In this way, proceeding from experimental results and the data given in Table 5.16, for the purpose of further argon EP-50 properties estimates , we will assume that gas efficiency ηg ~ 0.80 and the ion cost Ci ~ 350 W/A Table 5.17. Experimental Results Mode No Ub, V Ib, mА m, eq.А Ud1, V Ud2, V Сi, W/А ηg Р, mN Isp, sec ηт N, W

1 1,700 2,140 3.82 42 58 220 0.56 81.5 5,230 0.50 4,215

2 1,600 2,400 3.82 46 52 220 0.63 88.7 5,700 0.56 4,485

3 1,700 2,400 3.82 46 52 215 0.63 91.3 5,870 0.56 4,710

4 1,600 2,100 2.68 75 67 295 0.78 78.1 7,150 0.67 4,125

Efficiency and stability of the ion engine operations largely depend on the geometric parameters and manufacturing quality of the ion-optic system. hus unit possesses the highest degree of technological intensity. To ease the production of the ID-300 laboratory model they were made from titanium and their geometrical parameters were not optimal. he production of new IOS electrodes, with smaller apertures, increased transparency of the emitting electrode and lower accelerating voltage will permit to bring down the cost of ion and increase gas efficiency. 5.4.3.3. Evaluation of EP working parameters Apart from the requirements imposed on the engine by the mission task, it is necessary to take into consideration EP-50 components and units fabricability, as well as availability of the testing facilities for the further experimental engineering development of the engine. he engine’s ion-optic system (IOS) is its most sophisticated component in terms of fabricability. Its production and alignment quality determine the engine’s stability of operation and endurance. Today, in different countries there are well developed ion engines with the overall working size of up to 35 cm. Bigger engines are represented only by individual unique laboratory models. he IOS endurance of 10,000–15,000 hours can be achieved at the current density of 3–4 mA/cm2 (with molybdenum or titanium electrodes). If the significant accelerating potential is equal to 1.5kW, the power of an individual module with a 30cm IOV diameter will not exceed 4–5 kW and the EP will have to be formed of at 152

least 3,000 modules. his number of engine modules can hardly be called acceptable. One of the possible solutions is the employment of an engine with a segmented IOS. In this case, several ion-optic systems are mounted at the engine’s gas discharge chamber (GDC) exit. his configuration will simplify the IOS alignment procedure due to the reduction of the electrode gap to electrode diameter ratio. his approach was applied at Glenn Center (USA) to a 76cm ion engine at the stage of the laboratory model engineering development [5.52]. hree 30cm IOS units were mounted on its GDC. However, the employment of a segmented IOS will lead to the reduction of the engine’s efficiency, since in this case the IOS becomes less transparent for ions. he transparency of this IOS configuration will be equal to less than 50%, whereas the transparency of modern IOS amounts to 65–70%. Besides, IOS segmentation will require accommodation of high-voltage insulating units inside the gas-discharge chamber and, consequently, will increase the probability of a discharge breakdown. he ion engine specific impulse Isp is found with the help of the following formula (unadjusted for the ion beam divergence loss and the presence of divalent ions): I sp =

ηg

2 eU b

g

m

,

(1)

where Ub is the ion energy (emitting electrode potential); m, e are the ion’s mass and charge; g=9.81 m/sec2. By inserting Isp = 7,000 sec. and ηg ~0.8 into (1) we will get Ub~1540V for argon. he total power of the engines N is equal to N ≈ I b ⋅ (U b+C i ) .

(2)

By inserting N=15MW and C1=350 W/A into (2), we will find that the total ion beam voltage from all the engine modules will amount to Ib=7,938A. In this case, according to (3), the EP propulsive power near Earth (unadjusted for the ion beam divergence loss and the presence of divalent ions) will be equal to P~285 N and its efficiency ηm will be ~0.65.

P = Ib ηт =

2mUb

e g ⋅ P ⋅ I sp 2N

, .

(3) (4)

he evaluation of an individual EP-50 module starts with the characterization of the ion-optic system parameters and operating conditions.

153

A traditional configuration of an ion engine IOS system presents a set of two or three thin plates perforated with coaxial circular apertures positioned at hexagonal mesh points. he first plate, which limits the gas discharge chamber volume is placed under the cathode potential and is known as the emitting electrode. Its apertures serve to extract ions. Negative voltage is applied to the second (accelerating) electrode. his acceleration speeds up the ions and, at the same time, creates a potential barrier, which prevents the beam plasma electrons from getting into the GDC. he third (decelerating) electrode is under the potential, similar to the potential of the plasma around the engine. In some cases, this electrode can be shaped as a ring, which goes around all IOS components. Sometimes, to improve the cut-off of outer electrons, this electrode is made in the form of a perforated plate. he current density, which the IOS can extract from the GDC is determined by the type of the engine’s working medium, the distance between the emitting and the plasma electrodes and the accelerating voltage applied between the said electrodes. In one-dimensional approximation, this dependency goes under the name of ChildLangmuir law and can be presented by the following formula [5.57]:

j=

4 9

32

ε0

2e U a M

2

,

(5)

la

where e is the ion charge; M is the ion mass; Ua is the voltage between the emitting and the accelerating electrodes; la is the distance between the said electrodes and ε0 is the electric constant. he circular apertures and the final thickness of the electrodes result in the reduction of the electric field as compared to the ideal case of a plane onedimensional model. When doing the IOS EP calculations this factor is included by introducing the so-called effective acceleration length. he following formula was suggested for this length [5.58]: 2

l eff =

(l a + ts )

2

⎛d ⎞ +⎜ s ⎟ , ⎝ 2 ⎠

(6)

where ts is the emitting electrode thickness and ds is the diameter of the electrode apertures. One of the most labour intensive technological operations performed in the process of the ion engine development is the production of the IOS plates perforated with numerous circular apertures. he attempt to reduce the total number of the holes by the enlargement of their diameter faces limitations imposed by the above-described dependency (6). he enlargement of the apertures diameter up to the values sufficiently exceeding the length of the electrode gap will increase acceleration effective length and result in the reduction of the current density extracted by IOS from GDC. Experiments carried out with IOS showed that for the 154

purposes of effective ion extraction the diameter of electrode apertures must not exceed the length of the electrode gap by more than 3–3.5 times. he proportion between the thickness of the emitting electrode and the diameter of its apertures is determined proceeding from its effect on the extraction of ions from the GDC plasma. he results of the numerical simulation of the elementary ion beams formation process in individual IOS apertures show that the ratio between the ion flow in the acceleration area and the flow that comes to the emitting electrode from the GDC plasma reduces exponentially with the growth of the emitting electrode thickness to aperture diameter ratio [5.59]. For this reason, today, emitting electrodes are made as thin as the requirements for their mechanical strength and service life would permit. Currently, the typical electrode thickness to aperture diameter ratio, which ensures sufficiently effective ion beam extraction from the gas discharge camera, is equal to 0.15–0.25. Considering all the above proportions, the effective acceleration length exceeds the actual distance between the emitting and the accelerating electrodes by 2.5– 3 times. After inserting the physical constants, the ion beam density, found from Child-Langmuir law for the specified effective acceleration length, will take on the following form:

j≈

10 M

32



Ua

2

.

(7)

la

For practical considerations, in this formula ion mass M is expressed in a.m.u. (atomic mass units), accelerating voltage Ua – in kV, the length of the accelerating gap la – in mm and the current density j – in mA/cm2. As it has been shown before, to achieve the target specific impulse, argon ions must be accelerated to 1,540 electron-volt. he total accelerating voltage must exceed this value, since it is necessary to apply a negative voltage to the accelerating electrode in order to cut-off the beam plasma electrons. To increase the IOS service life, the negative potential bias value should be kept as low as possible. Experiments show that reliable cutoff of secondary electrons is possible if the ratio between the accelerating voltage and the voltage of the ion beam is not less than 1.25. It means that in the IOS under development the total accelerating voltage between the emitting and the accelerating electrodes will amount to Ua=1,925V. he maximum electric field intensity of the accelerating gap, which provides its resistance to vacuum voltage breakdowns is equal to about 2,500 V/mm. hat is why, the minimum permissible accelerating gap length is equal to 0.77 mm. his size of the accelerating gap permits to achieve the current density of 7 mA/cm2 for argon ions in an individual emitting electron aperture. However, the maximum current density for real ion engines (IE), intended for long service is often established proceeding from the requirements for the IOS service life. he fact is that during the ion engine operation, the ion emission from the gas discharge chamber is accompanied by the outflow of the working medium neutral atoms. he volume of this outflow is determined by the engine’s gas efficiency. Collisions between the accelerated ions and the neutral atoms provoke a charge exchange reaction, which results in the 155

appearance of fast neutral atoms and slow positively charged ions, which move at a heat flow rate. hese ions head for the negatively charged accelerating electrode and atomize its surface. he accelerating electrode erosion rate is the key factor, which determines the IOS service life. An increase in the ion current density is followed by an increase in the erosion rate. hat is why, the IE ion current density is selected proceeding from the requirements for the IOS service life Exact computation of the secondary recharged ions flow parameters, adjusted for their spatial and energy distribution, is a difficult task, because it requires computation of the beam plasma parameters. Today, even numerical simulation techniques give only approximate estimates of the IOS service life. Besides, experimental evidence of the relationship between the electrode surface atomization ratio and the ions drop energy and tilt angle is scarce and often contradictory. Up to the present moment, the endurance of the ion engines intended for long-term operation in space is determined by way of conducting full-scale ground tests. One of the most thoroughly studied ion engines is a 30 cm NSTAR (NASA’s Solar Electric Propulsion Technology Application Readiness) [5.54], which first underwent 8,200hour ground endurance tests, and then worked in space for 16, 265 hours as the main propulsion unit of the ‘Deep Space 1’ orbiter [5.60]. To a certain extent, the results of the test confirmed the correctness of the engine’s estimated design characteristics, according to which the NSTAR XIPS-30 prototype IOS service life should amount to 10,500 hours in laboratory conditions and to 25,000 hours in the conditions of hard space vacuum [5.61]. On the strength of these data, the authors of the article [5.51] dedicated to the prospects of high specific impulse 10–30 kW ion engines application for the purposes of deep space exploration suggested the following formula for the computation of the current density, which will ensure the target IOS endurance parameters:

j max = j NSTAR ⋅

M Mo ⋅ YXe− Mo ⋅ ρ a ⋅ ta ⋅ TNSTAR M a ⋅ Ya ⋅ ρ Mo ⋅ t NSTAR ⋅ Ta

,

(8)

where the ‘a’ index refers to the developed IOS parameters and the ‘NSTAR’ index refers to the operating parameters of the NTSR engine, working at the nominal input power of 2.32 kW. he above life service estimate was made for these two engines: jNSTAR = 2.55 mA/ cm is the mean density of the ion current over the NSTAR engine cross-section; Ma and ρa are the atomic mass and the accelerating electrode density respectively (currently titanium and carbon-carbon composites are regarded as the most promising electrode materials); MMo and ρMo are the atomic mass and the density of molybdenum, which served as a material for the NSTAR engine accelerating electrode; ta is the emitting electrode thickness; tNSTAR = 0.38 mm; Ya is the factor, which shows the rate of the accelerating electrode atomization (erosion) under the influence of the working medium ions (this value was calculated for the mean tilt angle and for the energy corresponding to the acceleration electrode negative potential); YXe-Mo ≈ 0.2 atoms/ion is the factor, which shows atomization of molybdenum under the influence of xenon ions with the power of about 180 electron-volt; Ta is the service life of the developed IOS; TNSTAR = 20,000 hours. 156

When calculating the estimated current density (8), it was assumed that the flow density of the slow secondary ions heading for the accelerating electrode is proportional to the density of the primary ion beam current. However, it seems more logical to suppose that the flow density of the slow secondary ions is proportional to their volume generating velocity, i.e. to the product of three values: 1) primary ions flow density, 2) beam plasma neutral atoms density in the near –IOS area; 3) charge-exchange cross-section αex. If we now assume that the neutral atoms density Nn is proportional to the density of their flow jn, which can be found proceeding from the known source gas efficiency value ηg with the help of the following formula

Nn ∝

jn =

1 − ηg ηg

j,

(9)

proportion (8) will take the following form:

j max

= j NSTAR

σ NSTAR (1 − η NSTAR ) ⋅ η g ⋅ M Mo ⋅ Y Xe− Mo ⋅ ρ a ⋅ ta ⋅ TNSTAR

(

σ ex 1 − η g

)⋅ η

NSTAR

⋅ M a ⋅ Ya ⋅ ρ Mo ⋅ t NSTAR ⋅ Ta



(10)

he latter proportion was used to evaluate the parameters of an IOS with the service life of 15,000 hours. he evaluation was carried out for two types of the accelerating electrode material – titanium and carbon-carbon composite, which are currently regarded as the most promising substitutes of the traditionally used molybdenum. heir main advantages are a low mass and a low atomization factor. Evaluation results are shown in Table 5.18, where jb is the mean density of the ion current over the NSTAR engine cross-section; jh is the current density in an individual IOS aperture and Ea is the intensity of the electric field in the accelerating gap. he evaluation results show that titanium electrodes will provide the target service life at the electric field intensity value close to the permissible maximum. On the other hand, the use of the carbon-carbon composite appears unpractical, since it is more difficult to make. Besides, the increase in the current density and the matching IOS diameter reduction, which could be achieved due to its higher atomization resistance properties, are impossible for the reasons connected with the maintenance of the accelerating gap electric strength.

157

Table 5.18. Evaluated IOS Parameters Parameters Ма, a.m.u. ρа, kg/m3 ta, mm Ub, V αex, m2 ηg Ua, V Ya, , atoms/ion Тa, hours jb, mА/cm2 jh, mА/cm2 la, mm Еа, V/mm

NSTAR 95.9 10220 0.51 1 100 3.9·10-19 0.90 -180 0.18 20000 2.49 3.56 0.58 2249

Ti 47.9 4505 1 1 540 1.9·10-19 0.80 -385 0.19 15000 3.55 5.07 0.87 2210

С-С 12.0 1 890 1 1 540 1.9·10-19 0.80 -385 0.05 15000 8.95 12.79 0.55 3510

he evaluation results permit to find the required accelerating gap length and other geometrical characteristics of an individual aperture and of the IOS as a whole. he summary of the IOS geometrical parameters is shown in Table 5.19. Table 5.19. Geometrical Parameters of an Individual IOS Aperture Parameters Emitting electrode thickness, mm Emitting electrode aperture diameter, mm Accelerating gap length, mm Accelerating electrode thickness, mm Accelerating electrode aperture diameter, mm

Value 0.5 3.0 0.9 1.0 1.8

he distance between the aperture centers is determined proceeding from technological practicability and requirements for the emitting electrode strength. If this distance is equal to 3.5 mm (the thickness of the space between the apertures being equal to 0.5 mm), the geometrical transparency of the emitting electrode will amount to 0.67, which is very close to the value applicable in the existing IOS. Focus quality control of the ion beam for the selected IOS cell geometrical parameters was carried out with the help of the ‘GASEL’-program-based calculation of the ion fields and trajectories [5.56]. the calculation was carried out proceeding from the following set-up parameters: • • • • • •

emitting electrode potential GDC plasma potential accelerating electrode potential GDC electrons temperature density of the ion flow from the GDC volume beam plasma potential

158

1,540 V; 1,570 V; -385 V; 10 electron-volt; 5.07 mA/cm2; 20 V.

In Fig. 5.39 you can see lines of equal potentials and ion trajectories received in the course of calculations. he calculated elementary beam current value amounted to 436 μA. It corresponds to the effective IOS ion transparency Fig. 5.39. Equal potential lines and ion trajectories in an individual IOV aperture of 0.815. he following conclusions can be drawn from the above calculations: 1. he mean ion beam density over the IOS cross-section jb = 3.55mA/cm2 is the maximum permissible value proceeding from the requirements for the accelerating gap electric strength. 2. Titanium is the most suitable material for the IOS electrodes. According to preliminary estimates, at the selected current density titanium electrodes will ensure the target service life (engine endurance) T = 15,000 hours. 3. An individual IOS cell size is determined practically singularly proceeding from its compliance with the requirement for the elementary beam target focusing. In doing so the hole diameter in the emissive electrode is equal to 3 mm, the area of an elementary cell hexagonal segment is equal to 0.1055 cm2 is equal to 436 μA. 5.4.3.4. Evaluation of argon EP-50 parameters at Isp =7,000 sec. To provide a 25–50 kW engine capacity of an individual module it is necessary to have a 70 cm IOS. Its electrodes will have about 36,500 apertures and the ion beam total current will amount to 15.9A. At the ion energy parameter of 1,540 electronvolt, the ion beam power will ignitor electrode be equal to 24.5 kW. Design power supply unit (PSU) parameters of an individual EP-50 module are given in cathode heater PSU Table 5.20. Fig. 5.40 features a connection diagram for the emitting connection of an individual electrode PSU module to the power supply system. In Table 5.21 you discharge PSU can find requirements for individual power supply acceleration electrode PSU units. he requirements for the low-power supply units neutralizer heater are provisional and will be PSU specified after the development of these EP components. he neutralizer ignitor electrode PSU total integrated propulsion unit will consist of ~500 individual EP- 50 l modules Fig. 5.40 Connection Diagram for the Connection of an Individual EP-50 Module to the Power Supply Units

159

Table 5.20. Design Parameters of an Individual EP-50 in the Nominal Duty Cycle Working medium Power, kW Propulsive thrust, mN Speciic impulse, sec. Eiciency factor

Аг 30.1 570 7000 0.65

Emitting electrode potential, V Beam current, A Accelerating electrode potential, V Accelerating electrode current, mA Discharge voltage, V Discharge current, A

1540 15.9 -385 150 50 111

Ar consumption, equiv. A Neutralizer consumption, equiv. mA Gas eiciency Ion cost, W/A Electrical eiciency Propulsive power cost, W/mN

19.4 400 0.80 350 0.81 52.8

Table 5.21. Power Supply Units Parameters Power supply unit Ignitor electrode Cathode heater Emitting electrode Discharge Accelerating electrode Neutralizer heater Neutralizer ignitor electrode

Current, A 10 15 16 90–140 0.5 15 3

Voltage, V 300 20 1540 40–60 500 20 300

Power, kW 0.2 0.3 25 6 0.2 0.3 0.1

Stabilization, operating mode Current (only at launch0 Current (only before launch0 Gas Voltage Current Voltage Current (only before launch) Current

Ion cost, W/A

5.4.3.5. Possible alternative modifications In the above-described modification, the EP-50 design efficiency amounted to 65% (unadjusted for the losses connected with the ion beam divergence) against the target 70%. he EP efficiency was evaluated proceeding from the estimated ion cost of 350W/A and from the gas efficiency of 0.8. It has been already mentioned that today there are no experimental data on the big argon GDC efficiency. hat is why, experimental research, will be the next step of this project development. he efficiency forecast will be adjusted for the future experimental findings. In Fig. 5.41 (the bottom curve) you can see the correlation between the maximum ion cost and gas efficiency, which ensures the EP-50 efficiency of not less than 70% (provided that the EP Gas eiciency runs on argon and that its specific Fig. 5.41 Correlation between the maximum ion cost and gas eiciency for impulse amounts to 7,000 sec.). At the EP eiciency of not less than 70% the gas efficience equal to 0.8, the ion cost schould hot exceed 200 W/A. he possible alternative steps may include: • increasing the specific impulse to 9,000 sec. To carry out this measure it will be necessary to increase the emitting electrode potential to 2,500–2,600 V. At the same time, its realization will permit to lower the ion cost requirements (see the middle curve in Fig. 5.41); • using krypton as the working medium. In this case, to ensure the specific impulse of 7,000 sec; it will be necessary to increase the emitting electrode potential to 2,800–3,200V. his measure will undoubtedly help to meet the ion cost requirements (see the top curve in Fig. 5.41) 160

For the analysis of the suggested alternatives it will be necessary to check their feasibility in terms of ballistics, carry out new computation of the ion-optic system geometry and make a new evaluation of the individual EP-50 module parameters.

Fig. 5.42. Model EP-50 (cross inter-sectional drawing)

To facilitate the assembly of a big number of EP-50 vehicle engine modules, the EP-50 gas discharge chamber can have a rectangular crosssection. he slot-type IOS can reduce the labor input required for the ion-optic system manufacturing. Today there are no experimental data on the tests carried out with regard to the employment of slot IOS in standardsize engines. hat is why, it will be logical to make a slot IOS trial model and test it in the existing ID-300 gas discharge chamber. In conclusion it is necessary to mention that: 1. A 30 cm argon ID-300 laboratory model was tested in the

Fig. 5.43. EP-50 cathode assembly, model

course of the ‘Manned Martian Mission’ project development. he engine was tested at the power of about 4kW. he tests proved attainability of duty cycles with the ion cost of 220–300 W/A at the gas efficiency of 0.55–0.78. he project team has outlined the ways of the further engine development and has planned the measures for modernization of the high-power engines testing facilities. 2. he project team has developed the EP-50 bench-engine drawings, Fig. 5.42. 3. he project group has fabricated and tested the EP-50 cathode assembly, Fig. 5.43 5.4.4. EP working medium unit he EP-50 working medium (argon) feed circuit is shown in Fig. 5.44. he main issue of the Working Medium Unit (WMU) development is the problem of the working medium (propellant) storage conditions. Currently, EP units used for the vehicle orientation, stabilization and orbit correction maneuvers run on xenon, which is stored onboard in a supercritical state and the tank mass ratio at (tank ‘dry mass’/working medium ‘dry mass’) is equal to at = 0.10–0.13. In the case of a manned Martian mission, in which EP will be used as a cruise propulsion unit, especially in the case of its one-vehicle configuration, in which the EP propulsion unit will be used at all the flight stages, the total onboard working medium reserves may amount to 200–300 tons. For the purposes of 161

ballistic computations (see Chapter 3) it was assumed that at = 0.03. Preliminary estimates have proved that in a Martian mission, both xenon and argon can be kept in the storage tanks in a supercritical state. However, the argon feed circuit will be more complicated because of the following optimum argon storage conditions: − maximum propellant-tank pressure: 5 MPa; − minimum propellant tank temperature: 150 K. he project team has studied the following techniques, which can provide the required argon storage minimum temperatures: • employment of the thermal blanket, proceeding from the fact that during the flight there is a constant argon flow, which can block heat penetration, Fig. 5.45; • employment of an onboard refrigerating unit, which would run on the solar energy; • boosting of the inside tank pressure and, consequently, of the minimum temperature value through the introduction of an advanced tank structure on the basis of new composite materials.

Ar

Ar

Ar Ar

Ar

КЗ

ПД1 ПД

КЗ1

Ф1

Ф2 ЗК ЗК4

ЗК ЗК3

ТД1 ТД

КС1

С

КС2

С

Д1

С

КС3

С

КС4

ТДПД1 ТД

ТД ТДПД4

ТДПД2 ТД

ПД

ТДПД3 ТД

ПД

ПД ПД

А1

А2

Г-Ar

Г-Ar

Ф3

Ф4 ЗК ЗК1

ТД

ТД ТДПД5 ТД ТДПД6

ПД

ЗК

ПД

ТД-ПУР1

ТД ТД2

Ф5

ПУР1

ПД

Pressure transformer

КЗ

Fill and service valve Argon ilter

ЗК ЗК2

ТД-ПУР2

ТД

Ф6 ПУР2

Ф7 ТД ТД-ПУР3 ПУР3

КЗ

КЗ

КЗ3

КЗ4

С

Solenoid valve

ЗК

Cutout valve Hose line

Flow control device Г Тр1

Г Тр2

Voltage insulator

Г Тр3 ТД

Главный расход

Temperature transducer

Расход на Расход на катод нейтрализатор

Figure 5.44. EP argon feed circuit

Fig. 5.45. Combination Brayton cycle for long-term space storage of: liquid O2, liquid CH4 and liquid argon

162

5.4.5. 10 years of solar arrays operating experience on board ‘Mir’ orbital station In subsection 5.4.1 we have already dwelt on the necessity of achieving the EP mass-to-power ratio of 5 kg/kW el. Research carried out in this direction in the past few years has proved that the said mass perfection can be achieved by using new materials, such as: − polymer films; − fiber glass composites; − carbon-carbon composites; − organic fiber composites; − aluminum and titanium alloys Many of the above materials have already been used on board the ‘Mir’ orbital station and the analysis of the ‘Mir’ solar arrays (SA) operating experience presents indisputable value. All the above materials are made in ground conditions. Naturally, for this reason they are saturated with atmospheric gases and water to say nothing of the gaseous products emitted due to the polymerization of different glues, coatings and adhesives. In the space environment these gaseous products emitted by the said materials will form a local atmosphere around the vehicle. Its other components will be: • cosmic plasma on low-earth orbits; • products of the used materials radiolytic decomposition resulting from ionizing

radiation exposure (solar ultraviolet radiation, Van Allen belt, solar bursts, galactic rays); • fine particles, generated by the micrometeorite and space debris impact on the EP structural elements. One should also remember that on the return trip to Earth the mission interplanetary vehicle will fly at the distance of only 0.57 a.u. from the Sun, and the solar energy flow will triple as compared to its intensity near Earth (which is located at the distance of 1a.u. from the Sun). A fifty-year experience of space exploration has demonstrated that unless special measures are taken, the spacecraft structure will be exposed to static electricity with the voltage of up to 20kV. It often leads to electric discharges on the spacecraft and solar array surfaces and also in the instrument compartments interior. According to the statistics [5.62], 50% of the spacecraft failures are provoked by electric discharges, ~30% fall to the share of ionizing radiation, the remaining failures are caused by design and fabrication technique errors. Onboard electric charge and discharge depend on several factors and need special attention. In January 1998, STS-89 Space-Shuttle orbiter acting within the framework of the Russian ‘Fragment’ space experiment brought back to Earth 8 MSB17KS5810- 0 solar array sections from the ‘Mir’ orbital station, where they had worked for 10 years. he retrieved SA fragments permitted to carry out a comprehensive analysis of the following factors with regard to their share in the SA deterioration: 163

• • • • • • • •

thermo cycling; contaminating impurity; radiation; micrometeorites and space debris; UV radiation; atomic oxygen; shadowing of individual solar cells during the SA operation; static electricity.

hermo cycling, which was connected with the alternation of day and night on the low-earth orbit, provoked cyclic compressive stress at the points where the conductor contacted with the solar photocell. Contaminating impurity mostly affected the solar cells front surface. It consisted of amorphous silicone dioxide (25% SiO2 + 15% Si) and was 3 μ thick. Contaminating impurity resulted in the reduction of the SA cover glass transparency by 5–10%. Radiation on the low-earth orbit (Hmean ≈ 400 km) was caused by the Earth’s proton radiation belt bottom boundary and resulted in a 2–3% power loss over the period of 10 years. Micrometeorites and space debris resulted in a 5mm reach-through breakdown of one of the SA sections, which had a total surface area of 0.812 m2, see Fig. 5.46. 20 more micro craters ranging from 16 to 200 μ in diameter were found on the surface of the same SA section. he total area under through holes and craters amounted to 1.1% of the section surface. Traces of erosion, namely micro craters with the diameters of 4–8 μ, were also found on photocells front surfaces. he eroded area amounted to 34% of their total surface. he photocells surface was also polluted with 2–200 μ debris particles, which included:

Fig. 5.46. SA section breakdown

− glass-like micro chips; − dark dust particles; − threadlike micro fragments.

Debris residue covered from 20 to 40% of the photocells surface. UV radiation darkened the SA cover glass and provoked gas release from polymeric materials.

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Fig. 5.47. SA section with darkened solar photocells

Atomic oxygen. Initially, the photocells front surface was polluted with silicone (Si) released from the SKTIF glue, which eventually oxidized to SiO2. Different degradation factor on the SA windward and leeward sides can be probably also explained by the atomic oxygen exposure. Shadowing. he results of the ground tests prove that partial strip shadowing of some photocells on the low-earth orbit, which induces the photocells to consume power and leads to their heating, was the main reason of the SA degradation. Partial strip shadowing of the photocells within the orbit daylight section was caused by the pantograph link bars, which unfolded the solar arrays and maintained their configuration in-flight. In Fig. 5.47 you can see a SA section with darkened photocells. Discoloration was caused by the excessive heating of the photocells followed by the distortion of glue optical properties. Static electricity. Although our American colleagues named electrical discharges as one of the reasons of the photocells darkening and degradation [5.63], no evidence was found in favour of this hypothesis with regard to SA degradation on low-earth orbits. Dangers connected with static radiation appear on intermediate heights and persist up to stationary near-earth orbits. his problem needs special consideration and special solutions aimed at the prevention of impurities on the SA surfaces, since impurities provoke electrostatic discharges and breakdowns. Table 5.23 shows MSB17KS5810-0 SA degradation factor after 10 years of operation on board the ‘Mir’ space station Table 5.23 SA section No 1 2

Degradation factor 39% 22.7%

3

24.4%

Remarks The section was the closest to the ‘Mir’ station main module Destruction of commutation node points and damage of the photocell structure at the points of contact with the conductor

4 54.2% 5 70% Destruction of commutation node points 6 70.6% Destruction of commutation node points 7 24.7% 8 48% The section was handed over to American scientists Excluding sections 3,5,6, which had an inherent structural defect, the mean degradation Mean degradation 50.4% factor value will be 37.7% factor Note: the examined samples demonstrated linear service-life dependence

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5.5. Version of Combined Propulsion System Based on Cell Solar Arrays, Electric Propulsion and Liquid-Propellant Rocket Engines his modification is a backup version of the EP model presented in Section 5.4 applicable in case this baseline model fails to meet the requirements for a number of key specific characteristics highlighted in Table 5.13. Naturally, the SPP+EP+LPE modification contemplated in a two-vehicle Martian mission configuration, which includes: • an interplanetary cargo vehicle (ICV); • a manned interplanetary vehicle (MIV) will considerably complicate a manned mission to Mars. On the other hand, it has a number of advantages. First of all, the required SPP output can be lowered to 5 MW both for the ICV and for the MIV. his factor together with their separate sequential assembly on the low-earth orbit will facilitate the assembly process. Secondly, the SPP+EP+LPE combination will be applied only to the MIV enabling it to accelerate from the low-earth orbit with the help of LPE and pass Van Allen belts at a maximum possible speed, softening the radiation effect on the crew and on the instruments and reducing the total time of the mission duration from 2.5 to 2.0 years. hirdly, the ICV arrives at the near-mars orbit a month before the MIV, using an energy-optimal Earth-Mars trajectory with the minimum reference speed builtup, and in this way creates reserves on the near-mars orbit: EP working-medium reserves, other expendable supplies and materials, maintenance equipment reserves, backup PPS and life support system units. In Fig. 5.48 you can see a MIV with the SPP+EP+LPE combination for the twovehicle mission scheme. 5.6. Conclusions In this chapter we have analyzed power and propulsion systems based on: • LPE; • solar power plants of different configurations for the EP power supply; • nuclear power plants for the EP power supply; • nuclear rocket engines; • combination of the above engine types. he complex analysis of power and propulsion systems for a Martian Mission Complex and its ballistic efficiency evaluation have elicited two preferred PPS versions based on the employment of: • a solar power plant with thin-film solar photocells from amorphous silicon alloys with the total output of 15 MW; Fig. 5.48. Manned Martian vehicle for the two-vehicle mission scheme

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• A modular-type nuclear power plant with gas-turbine conversion of heat into

electric power for the EP energy supply with the total output of 50 MW. 5.7. List of References 5.1. A. Koroteev, V. Semenov, ‘Nuclear Propulsion Perspectives in Russia’, 52rd IAC, Toulouse, France, 02.Oct., 2001. 5.2. Yu. G. Demyanko, G.V. Konukhov, A.S. Koroteev, E.P. Kuzmin, A.A. Paveliev, ‘Nuclear Rocket Propulsion’, Norma-Inform LLC, 2001. 5.3. A.I. Gorin, G.V. Konukhov, A.S. Koroteev, S.A. Popov, V.F. Semenov, ‘NRE Development Concept 1992’, 3rd Industry Conference ‘Space Nuclear Engineering’, Semipalatinsk, he Republic of Kazakhstan, 22-26 September, 1992. 5.4. E. Wahlquist, ‘US Space Reactor Programs’, May. DOE. 1990 5.5. R.A. Glinik, Yu. G. Demyanko, B.I. Katorgin, N.G. Pulkhrova and others, ‘Nuclear Power and Propulsion Unit Based on a High-Temperature Gas-Cycle Reactor for a Manned Martian Mission’, RKT Scientific and Technical Works, Issue 1 (134), 1992. 5.6. V.P Ageev, P.I. Bystrov, A.V. Vizgalov, L.A. Gorshkov, V.Ya. Pupko, Yu.P. Semenov, V.V. Sinyavsky, Yu.A. Sobolev, Yu.I. Sukhov, ‘Electric Propulsion Unit Based on a hermionic Nuclear Electric Propulsion Installation for a Martian Mission Complex’ RKT Scientific and Technical Works, Issue 1 (134), NIITP, 1992. 5.7. ‘M.V. Keldysh Research Center. 70 Years on the Forward Line of Rocket and Space Technology’, under the editorship of A.S. Koroteev, Mashinostroyenie, 2003. 5.8. S.P. Umansky, ‘Launch Rockets. Space Launch Complexes’, ‘Restart+’ Publishers, 2001. 5.9. ESA ‘Human Mission to Mars’, Report: CDF-20(A), Feb., 2004. 5.10. A.A. Nesterenko. ‘A Martian Mission Complex Version with a Mars Liquid Propellant Rocket Units and Aerodynamic Shield’. RKT Scientific and Technical Works, Issue 1 (134), NIITP, 1992. 5.11. V.N. Akimov, A.A. Gafarov, A.S. Koroteev, A.B. Prishletsov, Nuclear Engineering in Cosmonautics. M. ‘Polyot’ No10. pp. 3-11, 2000. 5.12. Machine Building. Encyclopedia. Volume IV-25, Book2 ‘Nuclear Engineering’, p. 496. M.: Mashinostroyenie, 2005. 5.13. S.P. Zatserkovny, A.I. Kuzin, K.A. Pavlov, G.A. Shevtsov, ‘Employment of TEM for the Solution of Long-Run Space Exploration Tasks’ – Article in the ‘Aerospace Engineering and Technology’ science and technology journal, Russian Engineering Academy, Moscow, No2, 2000. 5.14. Nuclear Rocket Engines, under the editorship of Academician A.S. Koroteev, M., ‘Norma-Info’, 2001. 5.15. N.N. Ponomarev, V.M. Talyzin, V.A. Pavshuk, V.K. Ulasevich, V.P. Smetannikov, Yu.S. Cherepenin, I.I. Fedik, V.P.Deniskin, E.K. Diakov, Sh.T. Tukhvatulin, Research High-Temperature Reactor (dedicated to the 30th anniversary of IVG1 power operation start-up) - Article in the ‘Atomic Power Engineering’ Journal’, vol. 98, Issue 2, March, 2005. 5.16. V.P.Deniskin, E.K. Diakov, Yu. S. Vasiliev, A.N. Kolbaenkov, A.A. Kolodeshnikov, V.A. Pavshuk, O.S. Pivovarov, N.N. Ponomarev-Stepnoi, V.P. Smetannikov, A.N. Tikhomirov, Sh.T. Tukhvatulin, V.K. Ulasevich, I.I. Fedik, Yu. S. Cherepenin, IVG-1 Reactor. Experience and Results of 30 Years of Operation: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 11-20, 2005. 5.17. A.A. Medvedev,A.I, Kuzin, A.A. Nesterenko, S.N, Lozin., Employment of Nuclear Power and Propulsion Units in Manned Mission Complexes for Martian and Lunar Exploration: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 21-24, 2005. 5.18. S.V. Barinov, M.S. Belyakov. O.N. Logachov, T.I. Rozhkova, et al., Concept of a Cruise Nuclear Propulsion Unit for the realization of a manned Martian mission: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 683-696, 2005. 5.19. S.V. Barinov, M.S. Belyakov, A.S. Kaminsky, V.S. Kuznetsov et al., Critical Solutions with Regard to the Martian Mission NPPU Nuclear and Radiation Safety.: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 679-682, 2005. 5.20. A.S. Kaminsky, V.S. Kuznetsov, V.A. Pavshuk, L.P. Bass, T.A. Germogenova, O.V. Nikolaeva, Calculation of the Contribution Made by the Fission Fragments Escaping with the Operating Engine Propellant into the Spacecraft Outward Radiation Environment: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 623-626, 2005. 5.21. S.V. Barinov, M.S. Belyakov, I.D. Daragan, A.S. Kaminsky, V.D. Kolganov, V.S. Kuznetsov, O.N. Logachev, V.A. Pavshuk, T.I. Rozhkova, V.P. Smetannikov, Yu.E. Khandamirov, Concept of a Nuclear Power Plant with Turbine-Machine Energy Conversion for the Purposes of an Electric-Propulsion-Unit-based Spacecraft Intended for the Exploration of the Distant Solar System Planets: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 659668, 2005.

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5.22. S.V. Barinov, O.N. Logachev, Summary of the Calculations Carried out with Regard to NPPU Reactor Core Areas Neutron and Physical Properties: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 627-634, 2005. 5.23. I.I. Fedik, Prospective Fuel and structural Materials for a Nuclear Power and Propulsion Unit: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 25-30, 2005. 5.24. V.Yu. Vishnevsky, I.D. Daragan, E.K. Diakov, V.N. Zagryazkin, V.A. Zaitsev, hermo-Dissociating Fuel for a TurbineMachine Energy Conversion NPPU: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 375-380, 2005. 5.25. A.D. Ivanov, O.N. Logachev, E.L. Romadova, Analysis and Classification of the Data Collected with Regard to NPPU Equipment Candidate Structural Materials: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 591-598, 2005. 5.26. Yu.V. Demianenko, A.I. Dmitrienko, Hydrogen Turbo-Pump Group for Pumping Liquid Hydrogen in Spacecraft Propulsion Units (NPE, LPE): report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 605-614, 2005. 5.27. V.T. Fedotov, O.N. Sevrukov, O.N. Logachev, Analysis of hread-Soldered Reactor Components Manufacturing Methods: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 615-622, 2005. 5.28. A.I. Belogurov, M.S. Belyakov, A.A. Kuleshov, P.C. Ozerov, O.N. Logachev, G.A. Ulanov, Design Engineering and Experimental Justification of the NPPU Nozzle Passage Shut-off System Development Feasibility: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 527-538, 2005. 5.29. M.S. Belyakov, L.A. Kanunnikov, Characteristics of NPE-based Space Nuclear Power and Propulsion Units Operation in Transient Modes: report at the international conference ‘Atomic Engineering in Space – 2005’, Collected reports, pp. 635-644, 2005. 5.30. Space Engines: Current Status and Prospects, under the editorship of L. Cavny, translated from English, Moscow. ‘Mir’, 1988. 5.31. S.D. Grishin, Yu.A. Zakharov, V.K. Odelevsky, Design of Low hrust Propulsion Spacecraft, M., ‘Mashinostroyenie’ , 1990. 5.32. Mattick A.T., Hertzberg A., Liquid Droplet Radiator Performance Studies. Acta Astronautica, 1985. 5.33. Mattick A.T., Hertzberg A., Liquid Droplet Radiator for Heat Rejection in Space, Journal of Energy, 1981. 5.34. 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