The Role of Nuclear Powerand Nuclear Propulsion in thePeaceful Exploration of Space

THE ROLE OF NUCLEAR POWER AND NUCLEAR PROPULSION IN THE PEACEFUL EXPLORATION

OF SPACE

The following States are Members of the International Atomic Energy Agency:

The Agencthe IAEA held atThe Headquartersenlarge the contrib

AFGHANISTANALBANIAALGERIAANGOLAARGENTINAARMENIAAUSTRALIAAUSTRIAAZERBAIJANBANGLADESHBELARUSBELGIUMBENINBOLIVIABOSNIA AND HERBOTSWANABRAZILBULGARIABURKINA FASOCAMEROONCANADACENTRAL AFRICA REPUBLICCHILECHINACOLOMBIACOSTA RICACTE DIVOIRECROATIACUBACYPRUSCZECH REPUBLICDEMOCRATIC RE OF THE CONGODENMARKDOMINICAN REPECUADOREGYPTEL SALVADORERITREAESTONIAETHIOPIAFINLANDFRANCEGABONGEORGIAGERMANYGHANA

GREECEGUATEMALAHAITIHOLY SEEHONDURASHUNGARYICELANDINDIAINDONESIAIRAN, ISLAMIC REPUBLIC OF IRAQIRELANDISRAEL

PAKISTANPANAMAPARAGUAYPERUPHILIPPINESPOLANDPORTUGALQATARREPUBLIC OF MOLDOVAROMANIARUSSIAN FEDERATIONSAUDI ARABIASENEGALys Statute was approved on 23 October 1956 by the Conference on the Statute of United Nations Headquarters, New York; it entered into force on 29 July 1957. of the Agency are situated in Vienna. Its principal objective is to accelerate and ution of atomic energy to peace, health and prosperity throughout the world.

ZEGOVINA

N

PUBLIC

UBLIC

ITALYJAMAICAJAPANJORDANKAZAKHSTANKENYAKOREA, REPUBLIC OFKUWAITKYRGYZSTANLATVIALEBANONLIBERIALIBYAN ARAB JAMAHIRIYALIECHTENSTEINLITHUANIALUXEMBOURGMADAGASCARMALAYSIAMALIMALTAMARSHALL ISLANDSMAURITANIAMAURITIUSMEXICOMONACOMONGOLIAMOROCCOMYANMARNAMIBIANETHERLANDSNEW ZEALANDNICARAGUANIGERNIGERIANORWAY

SERBIA AND MONTENEGROSEYCHELLESSIERRA LEONESINGAPORESLOVAKIASLOVENIASOUTH AFRICASPAINSRI LANKASUDANSWEDENSWITZERLANDSYRIAN ARAB REPUBLICTAJIKISTANTHAILANDTHE FORMER YUGOSLAV REPUBLIC OF MACEDONIATUNISIATURKEYUGANDAUKRAINEUNITED ARAB EMIRATESUNITED KINGDOM OF GREAT BRITAIN AND NORTHERN IRELANDUNITED REPUBLIC OF TANZANIAUNITED STATES OF AMERICAURUGUAYUZBEKISTANVENEZUELAVIETNAMYEMENZAMBIAZIMBABWE

THE RAND NTHE P

INOLE OF NUCLEAR POWER UCLEAR PROPULSION IN EACEFUL EXPLORATION

OF SPACE

TERNATIONAL ATOMIC ENERGY AGENCY

VIENNA, 2005

IAEA

The reA

SII

1Ato

IAE

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ole of nuclear power and nuclear propulsion in the peaceful xploration of space. Vienna : International Atomic Energy gency, 2005.

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. Nuclear propulsion. 2. Outer space Exploration. I. International mic Energy Agency.

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Printed by the IAEA in AustriaSeptember 2005STI/PUB/1197

FOREWORD

This publication has been produced within the framework of the IAEAs innovative reactor and fuel cycle technology development activities. It elucidates the role that peaceful space related nuclear power research and development could play in terrestrial innovative reactor and fuel cycle technology development initiatives. This review is a contribution to the Inter-Agency Meetinof the ScientifiUses of Outerorganizations i

Apart froorganizations, innovative reastatus and revispace, it also arelated nucledevelopment o

The IAEpublication, inStates of AmeRussian FedeRussian Feder

The IAEthe Division ofg on Outer Space Activities, and reflects the stepped up efforts c and Technical Subcommittee of the Committee on the Peaceful Space to further strengthen cooperation between international n space related activities.m fostering information exchange within the United Nations this publication aims at finding new potential fields for

ctor and fuel cycle technology development. In assessing the ewing the role of nuclear power in the peaceful exploration of ims to initiate a discussion on the potential benefits of space

ar power technology research and development to the f innovative terrestrial nuclear systems.A expresses its appreciation to all those who contributed to this particular to J. Graham (ETCetera Assessments LLP, United rica), V. Ionkin (Institute for Physics and Power Engineering, ration) and N.N. Ponomarev-Stepnoi (Kurchatov Institute, ation). A officer responsible for this publication was A. Stanculescu of Nuclear Power.

Although contained in thresponsibility fo

The use ojudgement by theof their authoriti

The mentias registered) doconstrued as an

The authoIAEA to reprocopyrights.EDITORIAL NOTE

great care has been taken to maintain the accuracy of information is publication, neither the IAEA nor its Member States assume any r consequences which may arise from its use.f particular designations of countries or territories does not imply any publisher, the IAEA, as to the legal status of such countries or territories, es and institutions or of the delimitation of their boundaries.on of names of specific companies or products (whether or not indicated es not imply any intention to infringe proprietary rights, nor should it be endorsement or recommendation on the part of the IAEA.rs are responsible for having obtained the necessary permission for the duce, translate or use material from sources already protected by

CONTENTS

1. INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

2. REGIMES FOR THE USE OF NUCLEAR POWER IN SPACE EXPLORATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

3. RADIOI

3.1. TE3.2. The3.3. Sov3.4. Saf3.5. Fut

4. REACTO

4.1. US4.1.

4.2. Sov4.2.4.2.4.2.4.2.

5. NUCLEA

5.1. US5.1.5.1.

5.2. Sov5.2.5.2.5.2.

5.3. Fut

6. SAFETY

7. OTHERSOTOPE POWER DEVICES . . . . . . . . . . . . . . . . . . . . . . . 7

Gs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7rmionic converters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13iet/Russian TEG developments . . . . . . . . . . . . . . . . . . . . . . . 14ety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17ure applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

RS IN SPACE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

experience . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 181. Studies of on-board nuclear reactors . . . . . . . . . . . . . . . 20iet/Russian experience . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231. Romashka NPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 232. BUK NPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 263. TOPAZ NPS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 294. Yenisey (TOPAZ-2) NPS . . . . . . . . . . . . . . . . . . . . . . . . . 31

R PROPULSION SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . 34

directions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 361. Safe affordable fission engine (SAFE) . . . . . . . . . . . . . . 372. Heat pipe operated Mars exploration reactor

(HOMER) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38iet/Russian directions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 401. IGR reactor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 412. IVG-1 experimental bench reactor . . . . . . . . . . . . . . . . . 423. IRGIT reactor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45ure applications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

INTERNATIONAL SPACE PROGRAMMES . . . . . . . . 50

7.1. China . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 507.2. France . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 527.3. India . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 537.4. Italy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 547.5. Japan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

8. THE FUTURE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

8.1. Nuc8.2. Inte

8.2.8.2.

8.3. The8.4. US

8.4.

8.4.8.5. Rus

8.5.8.5.8.5.

8.5.

8.5.8.5.

8.6. Neweng

8.7. The8.7.8.7.

8.8. Pro

9. CONCLU

10. LOOKIN

APPENDIX I

APPENDIX Ilear energy for interplanetary missions . . . . . . . . . . . . . . . . 55rnational projects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

1. Mars Together . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 572. TOPAZ-2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57 road ahead . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

A: Future directions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 661. Variable specific impulse magneto-plasma rocket

(VASIMR) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 682. Ion engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68sian Federation: Future directions . . . . . . . . . . . . . . . . . . . . 691. Transport and energy module TEM . . . . . . . . . . . . . . . . 702. Advanced thermionic NPS . . . . . . . . . . . . . . . . . . . . . . . . 713. Advanced NPSs using the external energy

conversion systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 744. Advanced nuclear systems using lithiumniobium

technology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 765. Gas core NTP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 776. Nuclear photon engines for deep space exploration . . . 79

technologies through nuclear space systems ineering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80 value of space technology in terrestrial applications . . . . . 911. Research and development . . . . . . . . . . . . . . . . . . . . . . . 912. Products, equipment and materials . . . . . . . . . . . . . . . . . 92blems to be solved in space by the use of nuclear power . . 98

SIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101

G AHEAD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104

: EVALUATION OF MARS MISSIONS . . . . . . . . . . 107

I: SPACECRAFT LAUNCHES INVOLVING RADIOISOTOPE SYSTEMS . . . . . . . . . . . . . . . . . . . 108

APPENDIX III: SPACE ACHIEVEMENTS . . . . . . . . . . . . . . . . . . . . . 111

APPENDIX IV: CASSINI SPACECRAFT . . . . . . . . . . . . . . . . . . . . . . . 114

APPENDIX V: AMTEC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115

APPENDIX VI: PROPULSION PERFORMANCE . . . . . . . . . . . . . . . 116

APPENDIX V

APPENDIX V

APPENDIX I

APPENDIX X

APPENDIX X

ACKNOWLEBIBLIOGRACONTRIBUTII: THE ROVER PROGRAMME . . . . . . . . . . . . . . . . . . 117

III: COMPARISON OF REACTOR SIZES. . . . . . . . . . . 118

X: SOVIET NUCLEAR POWER SYSTEMS IN SPACE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119

: SPACE EXPLORATION AGENCIES AND CONTRACTORS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121

I: INTERNET REFERENCES . . . . . . . . . . . . . . . . . . . . 125

DGEMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129PHY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131ORS TO DRAFTING AND REVIEW . . . . . . . . . . . . . . . . 133

1. INTRODUCTION

It is more than 100 years since the Russian theoretician Konstantin Eduardovitch Ziolkovsky advocated the use of liquid fuel rockets for space exploration and almost 80 years since Robert Hutchings Goddard launched the first liquid fuel rocket at Auburn, Massachussetts, in the United States of America. Sincethe endeavourOberth, Wernhfor Space Trav

Rocket rSecond Worldmotive force oresult was the

The trajethe border ofDevelopment War when bothprogrammes, c1957 and the fi

Artificialsatellites this generated by tsilicon and gergenerating powcube of the dineighbourhoodwould receivenecessity be so

Such conpower and healternative invuse of radioiselectric generisotope heaterSoviet/Russianpower sourcesVoyager, Galiillustrate the uprobe to use p1

then, rocket development has continued apace, largely through s of experimentalists such as Goddard, Willy Ley, Hermann er von Braun and other pioneers in both the German Society

el and the American Rocket Society. esearch and development was given a major boost during the War when the potential of the rocket engine to provide the f a long range weapon delivery system was recognized. The

German V-2.ctory of the V-2 took it to the edge of the upper atmosphere and space; it can be regarded as being the first space rocket. of rocket technology gained momentum after the Second World the USA and the former Soviet Union embarked on extensive ulminating in the first satellite launch (Sputnik 1) in October rst moon landing in July 1969. satellites of necessity require their own power source. For many has taken the form of solar panels, whereby electricity is he photovoltaic effect of sunlight on certain substrates, notably manium. For satellites in earth orbit this a common method of er. However, the intensity of sunlight varies inversely with the

stance from the sun, which means that a probe sent out to the of Jupiter would only receive a few per cent of the sunlight it

were it in earth orbit. In this case the solar panels would of large as to be entirely impractical.siderations lead to the development of alternative sources of ating which are completely independent of solar energy. One olves the use of nuclear power systems (NPSs). These rely on the otopes and are generally referred to as radioisotope thermo-ators (RTGs), thermoelectric generators (TEGs) and radio- units (RHUs). These units have been employed on both US and spacecraft for more than 40 years. Examples of the use of these on US probes over this period include Apollo, Viking, Pioneer, leo, Ulysses and Cassini missions. None of these missions tility of RTGs better than Pioneer 10, which was the first such

ower supplied solely from a radioisotope (238Pu).

2Pioneer 10 was launched from Cape Kennedy on 2 March 1972. It was the first interplanetary probe, successfully navigating the asteroid belt before making rendezvous with Jupiter and Saturn. The probe was equipped with an array of instruments for measuring such phenomena as the solar wind and the magnetic and radiation fields surrounding Jupiter. In fact, its discovery of the intense radiation fields surrounding Jupiter influenced the design of the Voyager and Galileo probes. Regarding Saturn, its instruments detected another ring planets magne

What waJanuary 2003 attempt to conthis time the stfurther commuhad been in coinformation onis now some 13

None ofprovide electritheir operation

The use electrical powenuclear thermtransferring sipossible using

This pupropulsion sysunits salient chistory of the advantages ove

2. REGIM

A space elaunch of the station and comlighting; for exand discovered two new satellites, as well as measuring the tic field. s apparently the spacecrafts last signal was received on 22

by the Jet Propulsion Laboratorys Deep Space Network. An tact Pioneer 10 was made on 3 February 2003, but this failed. By rength of the probes signal had degraded to such an extent that nication was impossible. For much of its 30-year life Pioneer 10 ntact with earth and during this time had transmitted valuable Jupiter and Saturn and the outer reaches of the solar system. It billion kilometres from earth.

this would have been possible without the use of RTGs to cal power and to maintain the components temperatures within al ranges.

of space NPSs is not restricted to the provision of thermal and r. Considerable research has been devoted to the application of

al propulsion (NTP). Such propulsion units will be capable of gnificantly heavier payloads into earth orbit than is currently conventional chemical propellants.blication reviews the development of NPSs and nuclear tems used in several national space programmes and details the haracteristics and other data (Appendices IXI). It provides a missions on which they were deployed and summarizes their r other systems.

ES FOR THE USE OF NUCLEAR POWER IN SPACE EXPLORATION

xploration mission requires power at many stages: for the initial pace vehicle and for subsequent manoeuvering; for instrumen-munication systems; for warming or cooling vital systems; for

periments and many more uses, especially in manned missions.

To date, chemical rocket thrusters have been used for launching. It would be tempting to believe that all power could be supplied by solar means since the sun is available and free. However, in many cases the mission may take place in the dark and large solar panels are not always suitable for a mission. Figure 1 shows the regimes of possible space power applicability.

For short durations of up to a few hours, chemical fuels can provide energy of up to 60 000 kW, but for durations of a month use is limited to a kilowatt or lessto provide rappower is most needed.

Nuclear duration. HowRadioisotopes

105

104

103

102

101

100

10-1

1 HOUR

ELE

CTR

ICA

L P

OW

ER

LE

VE

L (k

W(e

))

FIG. 1. RegimLaboratory.3

. Owing to the diffuse nature of solar power, it is not practicable id surges of large amounts of energy. On the other hand, solar efficient for power levels of some 1050 kW for as long as it is

reactors can provide almost limitless power for almost any ever, they are not practicable for applications below 10 kW.

are best used for continuous supply of low levels (up to 5 kW) of

NUCLEAR REACTORS

SOLAR RADIOISOTOPES

CHEMICAL

1 DAY 1 MONTH 1 YEAR 10 YEARS

DURATION OF USE

es of possible space power applicability. Source: Los Alamos National

4power or in combinations up to many times this value. For this reason, especially for long interplanetary missions, the use of radioisotopes for commu-nications and the powering of experiments is preferred.

Figure 2 shows that from any nuclear process, heat is emitted. This heat can then either be converted into electricity or it can be used directly to supply heating or cooling. The initial decay produces some decay products and the use of the thermal energy will provide some additional excess thermal energy to be rejected.

The nuclradioisotope fuconverted to thermionic coseveral heat cspace applicaworkhorses forradioisotopes i(and thereforemore by comb

Nuclear (SNAP-10A) sflew spacecrafAppendix IX)space is unthielectrical energ

FIG. 2.ear process shown in Fig. 2 can either be a critical reactor or el source such as plutonium oxide. In either case the heat can be electricity either statically through thermoelectrics or a

nverter, or dynamically using a turbine generator in one of ycles (Rankine, Stirling, Brayton). A classification of potential tions of nuclear power is shown in Table 1. The nuclear current space missions are the RTGs and the TEGs powered by n the Russian Federation that provide electricity through static reliable) conversion at power levels of up to half a kilowatt, or ining modules.reactors have also been used in space, one by the USA in 1965 uccessfully achieved orbit. The former Soviet Union routinely t powered by reactors: 34 had been launched prior to 1989 (see . A Soviet position paper stated that the investigation of outer nkable without the use of nuclear power sources for thermal and y. The USA agreed.

Generic space NPS. Source: Los Alamos National Laboratory.

The use ooptions. The chmuch more effpropulsion misratio efficiencietook a plannedcomparison, a n550 d on the p

TABLE 1. CLASSIFICATION OF NUCLEAR POWER TYPES BEING CONSIDERED FOR SPACE APPLICATION (Source: Los Alamos National Laboratory)

NPS type Electrical power range (module size)

Power conversion

RTG Up to 500 W(e) Static: thermoelectric

Radioisotope dyconversion gene

Reactor systemsHeat pipeSolid coreThermionics

Reactor system:Heat pipeSolid core

Reactor:Solid corePellet bedFluidized bedGaseous core

1 The specthe equivalent exhThe thrust (measuimpulse but the po5

f nuclear power in space is more than simply one of several power oice of nuclear power can make deep space missions possible and icient. For example, in a comparison between a typical chemical sion to Mars and one using nuclear propulsion, owing to the mass s and the larger specific impulse1, the chemically powered mission total of 919 d and provided a stay of 454 d on the planet. By uclear powered mission was completed in 870 d while it provided

lanet (see Appendix I). The outward bound and return journeys

namic rator

0.510 kW(e) Dynamic: Brayton Organic Rankine

: 101000 kW(e) Static: Thermoelectric ThermionicsDynamic: Brayton Rankine Stirling

110 MW(e) Brayton Rankine Stirling

10100 MW(e) Brayton (open loop)StirlingMagnetic hydrodynamic

ific impulse (a measure of rocket performance and measured in seconds) is aust velocity divided by the acceleration due to gravity at sea level (9.8 m/s2). red in Newtons or kilograms of force) is directly proportional to the specific wer needed to produce it is proportional to the square of the specific impulse.

6took 30% less time. In the map of possibilities involving time and a variety of payloads, nuclear power wins most of the time.

The prospects for using NPSs in space are determined by their advantages over conventional solar photovoltaic and other power sources, including:

(a) Independence of the distance to the sun and orientation with respect to the sun.

(b) Compactan area otheir size

(c) Better mbeginnin

(d) The capathe NPS

(e) Resistanc(f) The poss

give the power/prtwo to chemicalsupplyinginstrume

The expthrusters and new exploratiolance and glocommunicationuclear powerseveral hundreenvironmentallunar and Mar

As a mea15 kW in orbEstimates for aA baseline Mafaster transpor39 d. Such powness (a 10 MW solar array would require solar panels that cover f 68 000 m2 at the distance of Mars and 760 000 m2 at Jupiter and would render them impracticable).ass and size parameters when used on unmanned spacecraft,

g with a power level of several tens of kilowatts.bility of providing a power level two to three times greater with mass depending relatively weakly on the power improvement.e to the earths radiation belts.ibility of combining nuclear power with electrical thrusters to highest efficiency of specific impulse for thrust and of building opulsion systems on this basis to allow launch of payload masses three times greater than those possible with conventional propellant orbital boosters. This can be achieved while 50100 kW of electrical power and more for onboard

ntation over periods of 10 years or more.

erience accumulated in developing space NPSs, electrical NTPSs could, in the future, enable a number of quantitatively n missions, such as round the clock all-weather radar surveil-

bal telecommunication systems, including global systems for n with moving objects. In the future, space NPSs and combined /propulsion systems (NPPSs) with an electrical power level of d kilowatts will enable such long term space missions as global monitoring, production at facilities in space, supply of power for tian missions, and others.sure of power needs in space, a space shuttle consumes about

it while the International Space Station (ISS) uses 75 kW. Mars habitat range from 2060 kW not including propulsion. rs mission would require about 10 MW, but higher power means tation. Thus, a 200 MW engine could theoretically reach Mars in er is only available through advanced NPSs.

3. RADIOISOTOPE POWER DEVICES

3.1. TEGS

The basicthe German scdissimilar wirewhile the othejunctions is cal

The heatthe heat is statand continues energy. The hesotope is custoplanned missio

Figure 3 sconnecting thebottom and an

A compa12 years and it

RTGs haforty years, asinstalled in lonRTGs have sinRTG is a very cations.

An examFig. 6 shows it Gordon Bean units were usewere an ideal power during bof a year of ser

The US fuel pellets en572 multiply reEach thermocmissions, diffeMoreover, RT7

TEG is a simple device. It is based on an effect discovered by ientist Thomas Johann Seebeck in 1821. He found that when two s are connected at two junctions, and if one junction is kept hot r is cold, an electric current will flow in the circuit. Such a pair of led a thermocouple or thermoelectric couple. can be supplied from an isotope as in an RTG. The conversion of ic. The device has no moving parts and is, therefore, very reliable for as long as the radioisotope source produces a useful level of at production is, of course, continually decaying but the radioi-m selected to fit the intended use of the electricity and for its n duration.hows a hot shoe, through which radioisotopic heat is introduced, positive and negative legs. Some excess heat is rejected at the electric current is generated.rison between the predicted performance of a 150 W RTG over s actual performance during that time is shown in Fig. 4.ve been used in 26 US and many Russian missions over the past well as in the later French missions. They were originally g term remote navigational and meteorological satellites, but ce been used in a variety of lunar and planetary missions. An

versatile unit that can be custom designed for very specific appli-

ple of an RTG SNAP is shown in Fig. 5. This is the SNAP-27 and being removed from the Lunar Excursion Module by astronaut during the Apollo 12 mission to the moon in 1969. Five of these d to power experimental packages on the lunar surface. They choice for long missions that required the supply of continuous oth the lunar day and night. Each unit produced 63 W at the end vice.

designed general purpose heat source (GPHS) comprises 238Pu cased in iridium shells (4 pellets each weighing 151 g) and dundant thermocouples made of silicongermanium (see Fig. 7). ouple can produce more than half a watt. However, for other rent fuel and different thermocouple materials can be used. Gs can be used as modules of a total space auxiliary power

8system for botUlysses space lunar experimelectrical powemodules.

Another shown in Fig. 8components. Eclad in platinuthe event of aunits in addit

FIG. 3. Operatinh redundancy and for total power output. For the Galileo and missions, which had much higher power requirements than the ents, the GPHSRTG was designed to provide 300 W of r with a nominal fuel loading of 4.4 kW. It used 18 heat source

design, the lightweight radioisotope heater unit (RHU), is . These units provide temperature control for sensitive electrical ach includes a 2.68 g 238Pu dioxide fuel pellet producing 1 W,

mrhodium and encased in a graphite capsule for protection in n accident. The Galileo spacecraft had 120 of these lightweight ion to its GPHS. The Galileo spacecraft was launched on

g principle of the thermoelectric converter. Source: Rockwell International.

PO

WE

R O

UTP

UT

(W)

FIG. 4. Compara 12 year period.

HEA

HERMETIC

THERMOPILE

HOT FRAME

FUEL CAPLATCH PL9

TIME (Years)

ison between the predicted and actual performance of a 150 W RTG over

T REJECTION FINS OUTER CASE

SEAL

COLD FRAME

MOUNTING LEGS

HERMETIC SEAL

SULEATE

RADIOISOTOPEFUEL CAPSULE

FIG. 5. The SNAP-27 system. Source: NASA.

10

18 October 198extended throdates and the iboard sources

AppendiRTGs (or radnumbers of RTthe successes successes, with

FIG. 6. RemovaBean during the9 and arrived at Jupiter on 7 December 1995. The mission was ugh 1999 to allow it to fly past Europa, Callisto and Io. These nvaluable information fed back indicate the reliability of its on-of thermal control and electricity generation.x II shows a listing of US and Russian spacecraft that have used ioisotope powered TEGs in the Russian Federation), the G systems and the reasons for those missions. Appendix III lists of programmes supported by those power systems. These

requirements for the supply of steady and reliable power for up

l of SNAP-27 from the Lunar Excursion Module by astronaut Gordon Apollo 12 mission to the moon in 1969. Source: NASA.

to 14 years in power, would n

The intewith 3 RTGs (smission and 6335 more on thelectronics war

2 Partners

Space Agency (Ewere a total of 17

FUEL CLAD

GRAPHITE AEROSHELL CAP

LO

FLOATINGMEMBRANE

FUELPELLET

GRAPHITE IMPACTSHELL

FIG. 7. GPHS m11

locations well beyond those which would allow the use of solar ot have been possible without RTGs.rnational2 Cassini mission to Jupiter and Saturn was equipped ee Appendix IV) which produced 885 W at the beginning of the 3 W at the end. Cassini also had 82 small RHUs and there were e Huygens probe, each producing 1 W of heat to keep nearby m. These contained a total of about 0.32 kg of 238Pu.

: The National Aeronautics and Space Administration (NASA), the European SA), the Italian Space Agency (Agenzia Spaziale Italiana ASI) and there countries involved.

GRAPHITEIMPACT SHELL

CBCFDISC

LOCKMEMBER

AEROSHELL

CK SCREW

CBCFDISC

CBCFSLEEVE

odule assembly. Source: US Department of Energy/General Electric Co.

12

For the designed. It utechnology to AMTEC cell (tubes connectehot end of thestate and the stubes and in charged sodiumcold end of theartery towardthermal shieldthe hot side co

HEAT SHIELDEND CAP

CAPSULE

H

INSULATOR

INSULATORPLUGfuture, a new advanced radioisotopic power system has been ses alkali metal thermal to electric conversion (AMTEC) convert the heat produced by its plutonium heat source. The Appendix V) is made up of eight beta alumina solid electrolyte d in series. The end of the cell with the tubes is adjacent to the heat source. At this end, liquid sodium is heated to a vapour odium atoms in the vapour are driven through the walls of the so doing are stripped of an electron, thus creating positively

ions. The vapour is cooled and collected in a condenser at the cell and the cycle is repeated as the sodium flows through the s the hot surface at the other end of the cell. The cell uses

s in its upper section to reduce radiative bypass heat losses from mponents to the cold side condenser.

INSULATORMIDDLE TUBE

INSULATORPLUG

EAT SHIELD

INSULATOROUTER TUBE

INNER TUBE

FIG. 8. Lightweight RHU.

Leads are taken from the first and eighth tubes in series as the positive and negative leads for the cell. An explanatory cutaway diagram of the AMTEC system is shown in Appendix V.

This is an area of space research and development in which the latest ideas can be beneficial to various ongoing international innovative reactor technology research and development initiatives for terrestrial applications, particularly because older versions of these devices have already been used to provide power

3.2. THERM

Thermioninto electricity(typically at 18to a cooler collan electrical epreventing thicycle. Among t

(a) Radiant (b) Space ch(c) Energy lo

Much of The USA hadwith lifetimes materials). Ththermionic reundertaken sinthe more effici

Thermionsurface and sconnections at

A thermcoolant but thewith a typical drums. The temreactor powerlike torch battabout 2.5 cm in13

in remote situations, e.g. lighthouses and in the Arctic.

IONIC CONVERTERS

ic energy conversion is another method of transforming heat . It comprises a static device with a very hot emitter surface 00 K) that boils electrons across a small space (about 0.5 mm) ector surface (typically at 1000 K). This action essentially creates ngine with the electrons as a working fluid. There are factors s engine from achieving its ultimate efficiency of the Carnot hem are:

heat transfer between the hot emitter and the cool collector;arge effects between the plates; sses to the environment.

the development programmes aim to overcome these difficulties. a development programme targeting a 120 kW(e) power level of 10 00020 000 h (limited by heat induced effects on the e programme first tested converters in the reactor core (a actor) but this programme was terminated in 1970. Work ce has addressed usage separate from the reactor, resulting in ent use of both the reactor and the thermionic converter.ic diodes include fuel that is firstly surrounded by the emitter

econdly surrounded by the collector surface with electrical the bottom to connect to the next diode in series.ionic reactor does not contain fuel rods releasing heat to a rmionic fuel elements (TFEs) directly generating electricity. As

reactor, the fuel is critical and is controlled by rotating control perature of the hot emitter plates is in turn dependent upon the

level. These thermionic fuel rods are packaged in series, much eries. Designs have been produced showing these TFEs to be diameter and up to 40 cm long.

14

Gulf General Atomics tested the Thermionic Test Reactor employing carbide and oxide fuels between 1962 and 1973. Combinations of thermionic diodes and heat pipes (similar to the SP-100, see Section 4.1.1) provide challenging development problems but offer great potential.

This is another area of space research and development that can be beneficial to various ongoing international innovative reactor and fuel cycle research and development initiatives with terrestrial applications (see Section 8.7).

3.3. SOVIET

In Septelaunched into a90 satellites. Thhalf-life of 138semiconductorwas determine

In the m(radioisotope)research of Maelectrical powtransferred todevelopment wfabrication of sthis era, key pTEGs, or RTG

(a) The prod(b) The prod(c) The struc

radionucover a wi

(d) The RHU(e) Bench te

In 1992, aradionuclide) wand tested (seegermanium ba/RUSSIAN TEG DEVELOPMENTS

mber 1965, TEGs (Orion-1 and Orion-2) based on 210Po were near earth orbit as components of the Cosmos-84 and Cosmos-e choice of 210Po (with a specific thermal power of 141 W/g and d) allowed for a compact design which incorporated silicon

converters with an electrical output of ~20 W. The service life d mainly by the half-life of 210Po and this could reach ~3000 h.id-1970s, research and development on a complex radionuclide power system using 238Pu was initiated to support long term rs. This power system, named VISIT, included an RTG with an er output of about 40 W, the excess heat from which was a heat exchanger by pipes. However, VISIT power system

as limited to conducting terrestrial tests of its design and the cale models, as well as thermal and electrical prototypes. During roblems connected with the creation of radioisotope powered s, for space were solved, namely:

uction and processing of the 238Pu;uction of the cermet tablet fuel based on plutonium dioxide;tural materials for the manufacture of the RHU (capsules with

lide), as well as their compatibility with the fuel composition de temperature range; single elements design and the production process;

sting of the RHU.

thermoelectric mock-up (using an electric heater instead of the ith an electrical output of 3.75 W at end-of-life was fabricated Fig. 9). Its thermal power was 100 W. The tellurium, lead and sed alloy semiconductors in the thermoelectric battery were

medium tempribbed casing. ESAs Leda lu

In the latresearch probemission. The m(see Fig. 10). Rtures, to powethe orbiting spAngel were deand RHU are for equipmentflow to a thermRHUs and twoan electrical poRHUs with thewas developed

FIG. 9. Thermoinsulator, (4) cas15

erature ones with heat removed by thermal radiation from a This RTG was proposed for use as a lander power source under nar programme.e 1990s, RTGs were used as the electrical power supply for the s to be landed on Mars as part of the Mars-96 international ission included long life small autonomous stations and probes TGs were needed to maintain equipment at design tempera-

r equipment and to recharge a battery for communication with acecraft. Thus, 8.5 W 238Pu RHUs and a 200 mW(e) RTG named veloped for the small autonomous station spacecraft. The Angel unified products destined to be used both as self-contained units heating and as the initial heat source to provide the steady heat oelectric converter. The small autonomous station included two RTGs. The complex monoblock radionuclide power system had wer output of about 400 mW and includes a storage battery, two thermal power of 8.5 W each and a thermoelectric converter. It for the probes.

electric mock-up: (1) electric heater, (2) thermoelectric battery, (3) heat ing, (5) rib. Source: Kurchatov Institute.

16

The Angshield case andThe ampoule campoule has alloys are usedof 238Pu dioxidhas a release malpha decay ofstrength tantacoated with muconstruction haadditionally presistant carbo

The Angsame basis. Stelluride alloysof about 200 mpower for the through a nick

SMALL STATION

RTG(2) + RHU(2)

FIG. 10. Mars-els cylindrical heat unit (d = 4 cm, h = 6 cm) includes a heat carbon heat insulation surrounding the radioisotopes ampoule. ontains about 17 g of 238Pu dioxide with an activity of 260 Ci. The a dual structure. High corrosion resistant platinumrhodium for the inner ampoule casing that contains two ceramic tablets e clad in iridium. The inner ampoule is hermetically welded and echanism for dealing with radiogenic helium resulting from the

238Pu. The load bearing outer cladding is fabricated from high lumtungsten alloys. After hermetically welding, its surface is ltilayer refractory materials. Thus, the radionuclide heat source s a double containment in each capsule and the ampoule itself is

rotected against outer thermal and impact attacks by heat n materials.el radioisotope powered TEG or RTG was developed on the emiconductor thermoelectric materials based on bismuth are used as a converter. The RTG generates an electrical power W at an operating voltage of 15 V at room temperature. The

small autonomous station equipment is supplied from the RTG elcadmium buffer battery.

96 project: small autonomous station layout. Source: Kurchatov Institute.

Radioisotope powered TEGs (or RTGs) of milliwatt electrical power for space application, such as the Angel RTG and its modifications, are compact, reliable in operation and have low mass and size, which makes them convenient for probes. The RTG waste heat is enough to maintain design temperatures for equipment working in deep space environments.

Further, the RHU is also used for heating gas to warm the instrument module of the Lunokhod-1 and Lunokhod-2 stations. The heat sources thermal power

Researchthe 238Pu baseplaced on impSuch generatorwere proposedLeda program

3.4. SAFETY

The safenormal operattrolled descenunder all possAngel RHUs wPertaining to tby the United

3.5. FUTURE

Althoughpower devicespowerful and mcations and thewith innovativ17

is 900 W. and development work on thermionic converters, together with

d radionuclide heat sources, is ongoing with an emphasis being roving energy conversion efficiencies from 810% to 1014%. s using a thermionic converter with an electrical power of 75150 W as the electrical power sources for a rover vehicle under the

me.

ty of all RHU applications using 238Pu must take into account ion, emergency conditions during the launch and an uncon-

t from the circular orbit. Maintaining the 238Pu capsule airtight ible conditions is the basis for ensuring radiation safety. The ere developed and designed to be consistent with the Principles

he Use of Nuclear Power Sources in Space that were approved Nations General Assembly in 1992.

APPLICATIONS

the principles of thermoelectric and thermionic heating and are very simple, continued development is producing more ore compact designs. These designs have many terrestrial appli-refore the research and development in this area has synergies

e reactor technology development activities.

18

4. REACTORS IN SPACE

As discussed in Section 2, while radioisotope powered systems are ideal for long term low power functions, nuclear reactors have the capability of producing almost unlimited power above a kilowatt for any length of mission. The USA used one in 1965 in its SNAP-10A probe and the Soviet/Russian programme hspacecraft wer

4.1. US EXPE

Early inlaunches and tsystems led to salient details

EXPANSIO

SUPPORT LEG

FIas routinely used them. Thirty-four nuclear powered Soviet e launched between 1970 and 1989.

RIENCE

tentions were to use nuclear reactors both to power space o supply onboard power needs. Considerable research on SNAP the launch of SNAP-10A (see Fig. 11) on an Atlas launch vehicle, of which are as follows:

T/E CONVERTER RADIATORS THERMOELECTRIC PUMP

N COMPENSATOR

REACTOR

SHIELD

STRUCTURE 8-RING STIFFENERS

LOWER SODIUMPOTASSIUM MANIFOLD

INSTRUMENTATION COMPARTMENT

G. 11. SNAP-10A system. Source: Atomics International.

Launch date: 3 April 1965 21:24 GMT. Launch site: Vandenberg. Launch complex: PALC2-4. Launch vehicle: SLV-3 Atlas/Agena D.

Payload: SNAP-10A/Agena D. Mass: 440 kg. Class: Technology. Type: Ion engine. Agency: USAF/AEC. Perigee: 1270 km. Apogee: 1314 km. Incli-nation: 90.3 deg. Period: 111.4 min. COSPAR: 1965-027A. The spacecraft carried a SNAP-10A nuclear power source. The onboard nuclear reactor provided electrical power for a 1 kg force ion engine. The crafts telemetry

Howeverwere launchedinto space, theother reactors,all RTG auxilinumerals werebut had differe

TABLE 2. COREACTORS

Characteristic

Power (kW)

Design lifetime

Reactor power (

Reactor outlet (

Fuel and spectru

Coolant

Power conversio

Hot junction (K

Cold junction (K

Turbine inlet tem (K)

Condenser temp (K)

Unshielded weig19

failed but the reactor itself operated well.

, US Government policy changed and no more nuclear reactors . Although the US sent only one nuclear reactor power source SNAP-10A, considerable work had already been done on two SNAP-2 and SNAP-8. In the convention adopted by the USA, ary power systems were denoted by odd numerals, while even reserved for nuclear reactors. All three reactors were similar nt power levels (see Table 2).

MPARISON OF SNAP-2, SNAP-10A AND SNAP-8

SNAP-2 SNAP-10A SNAP-8

3 0.58 35

(a) 1 1 1

kW) 55 43 600

K) 920 833 975

m UZrH thermal UZrH thermal UZrH thermal

NaK-78 NaK-78 NaK-78

n Rankine (Hg) Thermoelectric (SiGe)

Rankine (Hg)

) 777

) 610

perature 895 950

erature 590 645

ht (kg) 545 295 4545

20

Unsolved design issues such as mercury corrosion and crud, and protection of stator windings, bearings and the pump made the reliability of SNAP-2 and SNAP-8 potentially poor and it was for this reason that the SNAP-10A with its thermoelectric power conversion system was used in flight. The SNAP-10A power conversion system is shown in Fig. 12.

4.1.1. Studies of on-board nuclear reactors

There arAdvanced spacoupled with cspectrum with reactor, which

This desliquid metal comission is pow

The reacfor its controlrotating beryll

833 K

FIG. 12. SNe many possible designs of nuclear reactors for use in space. ce mission requirements for high power levels (25 500 kW(e)) ompact size and long lifetimes favour the use of the fast reactor highly enriched fuel. One design for a liquid metal cooled space is still a major contender for the future, is shown in Fig. 13.ign is heavily dependent upon the designs of terrestrial oled fast reactors but is adapted for spacecraft in which the er production rather than breeding or waste reduction. tor needs to be small, restrained and not dependent upon gravity , which would be normal on earth. Therefore, the design uses ium control drums that have boron carbide absorber segments.

FLOW 14.5 gpmP 9 kPaTHERMAL POWER 1000 WAVE. RADIATOR TEMP. 583 K

ELECTRIC POWER 580 WAVE. HOT JUNCTION TEMP. 777 KAVE. RADIATOR TEMP. 610 KEFFICIENCY, CONVERSION 1.43%VOLTAGE 30 V

POWER CONVERSION SYSTEM

761 K

829 K831 K PUMP

POWER 43 kWP 72 K

REACTOR

AP-10A power conversion system. Source: Atomics International.

However, the completed in t

This exammany candidatTable 3).

For a mamust be of thethe form of th0.55.0 MW(ehave been stud

Apart fropower converssystem.

In 1983, Nagreed to fundreactor systemincluded a lithconverters. In spacecraft.

INSULATION(ZrO2)

TIE-DOWN BARS(MoRe)

HEAT PIPE(MoRe/Li)

FUEL LAYERS(UO2/MoRe)

DRUM BEARING

REFLECTOR (Be)

SAFETY PLUG C

RETAINER BANDS (Mo)

DRIVE SHAFT

REFLECTOR (BeO)

FIG. 13. DistribLaboratory.21

materials technology and proof of that technology have been he non-space liquid metal fast breeder reactor programme.

ple of a distributed liquid metal cooled reactor is merely one of e systems that include several variants of solid core reactors (see

ss density of 30 kg/kW(e) in a small reactor, outlet temperatures order of 12001500 K. This temperature objective defines both e fuel and the coolant. For higher power requirements, in the) range, fluidized bed and pellet bed reactors with gas cooling ied.m the nuclear reactor, a power plant includes shielding and a ion system, including converters and an excess heat rejection

ASA, the US Department of Energy and several other agencies a joint programme, named SP-100 (see Fig. 14), to develop technology. This programme developed a power system that ium cooled reactor coupled by heat pipes to thermoelectric this way the reactor could be used remotely from a manned

AP (Be) THERMAL INSULATION (MULTIFOIL)

CONTROL DRUMS (Be)

ABSORBER SEGMENT (B4C)

FUEL MODULES

CORE CONTAINMENT(MoRe)

uted cooled (liquid metal) space reactor. Source: Los Alamos National

22

The therradiated to themolybdenum converters arethe heat rejectuser spacecraftbasic technoloUS programm

TABLE 3. EXAMPLES OF SOLID CORE NUCLEAR REACTOR SYSTEMS

Solid core type Variant 1 Variant 2 Variant 3

Integral heat transport reactor

Matrix fuel, gaseous (He) coolant

Pin fuel, NaLi coolant

In-core cylindrical thermionics,NaK coolant

Distributed heattransport reacto

FIG. 14. SP-10Los Alamos Namoelectric panels contain converters, which use heat directly hot shoes from the heat pipes. The hot shoes are made from

or lightweight carboncarbon composites. The thermoelectric distributed throughout the panels with the cold shoes serving as ion surfaces. The whole power system would be secured to the by a boom running through the centre of the heat pipes. These gies and their evaluation are already contributing to the future e.

r

Heat pipewafer or coated particle fuel with heat pipes

Liquid metalwafer or coated particle fuel with electromagnetic pumps

In-core thermionicswafer or coated particle fuel with either electromagnetic pumps or heat pipes

0 nuclear power system (radioactively coupled system design). Source: tional Laboratory.

4.2. SOVIET/RUSSIAN EXPERIENCE

The development of space NPSs with direct conversion of nuclear fission thermal power into electrical power started in the early to mid-1950s. The former Soviet Unions first NPS with the direct (thermoelectric) conversion of nuclear fissionNPS first operaenergy over 15

The BUKelectrical powe1970s this NPSthere were 32 lCosmos series

The deveearly 1960s. Tthermionic reaFurther testinprototypes andTOPAZ in flcomponents ofplace in Februaand testing of t1960s. Since ththose of the TO

On the b(BUK, TOPAZsystems (NPS-1980s. The parof the electricaas well as the l

4.2.1. Roma

The mainNPS is a convereactor core isreflector outerthe segmented90% enriched 23

heat into electricity was the terrestrial Romashka NPS. This ted in August 1964 and generated about 6100 kWh of electrical

000 h. space thermoelectric NPS was created in the 1960s and has an r output of about 3 kW. After the conclusion of tests in the early was put into operation in near earth orbits. From 1970 to 1988 aunches of these power systems (reactors) as a component of the of spacecraft (see Appendix IX).lopment of a space thermionic NPS was also undertaken in the he first successful power test of a terrestrial prototype of a ctor converter for the TOPAZ NPS was completed in 1970.

g over the next two decades of thermionic reactor converter TOPAZ NPS terrestrial prototypes made it possible to test the

ight. The first test flights of two TOPAZ NPS models as the Plasma-A spacecraft (Cosmos-1818 and Cosmos-1867) took ry and July 1987. Along with the TOPAZ NPS, the development he Yenisey thermionic NPS was started in the second half of the e power, mass and size parameters of this NPS were similar to PAZ NPS, in the West it was referred to as TOPAZ-2.asis of the experience gained with first generation of NPSs , Yenisey), the development of the next generation thermionic

25, NPS-50 (Space Star) and NPS-100) was started in the mid-ameters of these systems meet the higher requirements in terms l power and lifetimes imposed by new space exploration targets

atest safety requirements.

shka NPS

unit of the Romashka NPS is shown in Fig. 15. The Romashka rter based on a fast reactor, in which the heat generated in the conducted to a coaxially arranged TEG located on the radial surface. The reactor core comprises a stack of 11 fuel elements; fuel elements consisting of discs of uranium bicarbide with 235U. This is located within a graphite package, so built that a

24

significant parreducing the te

A radial located betweethe high operaand beryllium The reactor enheat insulation

FIG. 15. The Relectric elements(7) radial reflectt of the heat from the core goes through the package body, thus mperature drop in the uranium bicarbide.beryllium reflector encloses the reactor. Graphite bushings are n the core and the reflector to prevent reflector deformation at ting temperatures. The bushings are coated with silicon carbide oxide to protect them from chemical interaction with beryllium. d reflectors are also made of beryllium. The high temperature made of foam graphite and multilayer graphitized fabric is

omashka NPS reactor converter layout: (1) radiator ribs, (2) thermo-, (3) control rod, (4) reactor vessel, (5) upper reflector, (6) reactor core, or. Source: Kurchatov Institute.

mounted on the reactor end walls to reduce the heat transfer. The combination allows the reactor to operate with a temperature of up to 2173 K in the centre of the core and between 1273 K and 1373 K on the reflector outer surfaces.

The reactor control system consists of four rods located in the radial reflector and in the lower end reflector. Two rods are used for automatic and manual control, whereas the other two, together with the movable end reflector, are used for reactor protection in case of emergency.

High te(Si: 85% by minside the hermindependent pp-conductivityOn the cold sidbridge onto a c

To preveberyllium oxidclearances betare filled withenamel coateddetails of the RTable 4.

TABLE 4. THCHARACTER

Characteristic

Reactor core dia

Radial reflector

Reactor load ma

Total mass of thand reactor (

Reactor converttaking into ac

Reactor convert

Electrical power

Reactor convertof thermoelec

Number of therm25

mperature semiconductor grade silicongermanium alloy ass, Ge: 15% by mass) is used in the TEGs. These are mounted

etically sealed steel vessel in four groups, each group having an ower outlet. The cell comprises two thermopiles with the n- and joined together on the hot side by the molybdenum keyboard. e, separate pairs are joined with each other in series by a copper ommon arm running the height of the generator. nt the thermoelectric converters shorting, insulating plates of e are used on the hot and cold sides. To reduce heat loss, all ween the thermoelectric cell and hollows in the TEG structure a cotton-like quartz and a helium atmosphere. A total of 192 fins, with an emissivity of at least 0.9, reject excess heat. Basic omashka NPS reactor converter characteristics are presented in

E ROMASHKA NPS REACTOR CONVERTER ISTICS

Value

meter/height (by package) (mm) 241/351

outer diameter/height (mm) 483/553

ss by uranium-235 (kg) 49

e TEG (with the casing and radiator) without drives and control rods) (kg) 635

er effective thermal power (without count the end wall spread of heat) (kW) 28.2

er electrical output (at start-of-life) (W) 460475

reduction over a lifetime of 15 000 h 80%

er terminal operating voltage (four groups tric converters connected in series) (V) 21

oelectric converters in a TEG 3072

26

The reactor converter generated an electric power of 460475 W at a constant optimum outer load at start-of-life. By the end of testing (after approximately 15 000 h) the reactor converter electrical power had decreased to 80% of its initial power. This electrical power loss was mainly accounted for by an increase in the thermoelectric converter inner resistance owing to the diffusion processes operating at the graphite disc/silicongermanium alloy interface resuland, partly, to

4.2.2. BUK

The BUcylindrical radsystem of ribbeis located on th

The BUKfuel is a highly30 kg. Longitureflector. A twsodium and p973 K, is supplof which are hradiator, behinexcess heat toradiator inlet bfor the spacecrelectromagnetshown in Fig. 1

There armade of a highalloy. The nucwhich the maxefficiency of 3%by which timconversion effi

Radiatio

(1) The basicspacecrafmore thasafely toting in the formation of a high resistance silicon carbide layer the failure of contacts on the hot side.

NPS

K NPS includes the reactor, the shielding and the conic/iator located in series along the axis. The radiator comprises a d pipes for coolant flow united by input and output collectors. It e load bearing frame structure that is joined to the spacecraft. NPS uses a small fast reactor which contains 37 fuel rods. The

enriched uraniummolybdenum alloy; the 235U load being about dinally movable control rods are placed in the beryllium side o-loop liquid metal heat removal system uses a eutectic alloy of otassium as coolant. The first loops coolant, heated to about ied to the outer casing of the TEG. The TEG, the inner cavities ermetically sealed and filled with inert gas, is located under the d the reactor shielding. The second circuit coolant removes the the radiator with the coolant maximum temperature at the eing about 623 K. The TEG has two independent sections: one aft users and an auxiliary one for the power-to-conduction type ic pumps used for both coolant loops. The BUK NPS layout is 6.e two cascade thermoelectric converters in the TEG; the first temperature alloy, the second made of a medium temperature lear reactor thermal power is limited to about 100 kW, from imum electrical power generated is about 3 kW, representing an

. The BUK NPS lifetime was extended in operation to 4400 h e the electrical parameters of the TEG had degraded. The ciency at 4400 h was about 90% of its start-of-life value. n safety is provided by two diverse systems:

safety system, the spacecraft component, relied on moving the t into a long term burial orbit, close to circular, at a height of n 850 km. There, nuclear reactor fission products can decay

the level of natural radioactivity. The orbit change system is

located inpower uoperationsystem wpower.

(2) The backproducts the earthin the opreactor, eheating, expectedpose no eThis backthat defofrom cylifrom the

The backfailure of Cosdescent resulteline across noshown in Table

FIG. 16. The Breactor shieldingbearing frame st27

the spacecraft module and is mechanically joined to the nuclear nit and separated from the spacecraft service module in low al orbit. The orbit change system includes an off-line propulsion ith its own control systems and an off-line source of electrical

-up emergency system provides for the dispersion of fuel, fission and other materials with induced activity into the upper layers of s atmosphere. This system ejects the fuel element assembly either erational orbit or when the object, which includes the nuclear nters denser atmospheric layers. During the descent, aerodynamic thermal destruction, melting, evaporation, oxidation, etc., are to disperse the fuel into particles that are sufficiently small as to xcess radiological hazard to the population or to the environment. up system consists of control devices and actuating mechanisms

rm and destroy special flexible elements by the pressure of gases nders. A diagram of the fuel element assembly ejection system reactor core is shown in Fig. 17.

up safety system was introduced into the BUK NPS after the mos-954 spacecrafts change of orbit system. The spacecrafts d in large radioactive fragments of wreckage being strewn in a rthern Canada in 1978. Characteristics of the BUK NPS are 5.

UK NPS layout: (1) nuclear reactor, (2) liquid metal circuit pipeline, (3) , (4) liquid metal circuit expansion tanks, (5) radiator, (6) TEG, (7) load ructure. Source: Kurchatov Institute.

28

TABLE 5. CH

Characteristic

Power (kW(e))

Design lifetime

Reactor power (

Reactor outlet t

Fuel and spectru

Coolant

Power conversio

Hot junction tem

Unshielded weig

FIG. 17. Diagraplate, (2) fuel el(6) side reflectorARACTERISTICS OF THE BUK NPS

Value

4.2.3. TOPAZ NPS

The TOPAZ NPS includes a thermionic reactor converter with a caesium vapour supply system and control drum drive unit, the reactor shielding, the radiator and the frame by which the system is joined to the spacecraft service module (Fig. 18). The automatic control system is placed in the hermetically sealed service module and connected to the related nuclear power unit systems by electrical se

The coreThe TFEs and form a systemcollector stackinterelectrode working sectiosection, whereconduction tyremoval systemcaesium vapouat a voltage of1.1 V. Before level, the electstorage battery

Twelve rthermal powe

FIG. 18. The TOdrive unit, (2) thshielding, (5) liqKurchatov Instit29

rvice lines. consists of 79 TFEs and four zirconium hydride moderator discs. cooling channels are located in the moderator disc openings and of five concentric rows. Five-cell TFEs with a three layer

are used, with fission gas vented from emitter assemblies to the gap. The TFEs are electrically connected so that they form the n of 62 TFEs and the pump section of 17 TFEs. The pump the TFEs are connected in parallel, is intended to energize the pe electromagnetic pump of the nuclear power units heat . The TFEs within this section are connected at both ends in the

r. The operating section terminal electrical output is about 6 kW ~32 V. The pumping section current is ~1200 A at a voltage of the reactor converter is brought to the rated electrical power romagnetic pump is fed from the startup unit by a high current located behind the radiation shielding.

otating cylinders (drums) located in the side reflector provide r control, reactivity compensation and emergency shutdown.

PAZ NPS layout: (1) caesium vapour supply system and control drum ermionic reactor converter, (3) liquid metal circuit pipeline, (4) reactor

uid metal circuit expansion tank, (6) radiator, (7) frame structure. Source: ute.

30

These beryllium cylinders have sector cover plates of boron carbide and are divided into four groups of three cylinders. Each group is controlled by its own drive.

The caesium vapour supply system pumps vapour through the TFE interelectrode space at a flow rate of about 10 g/d. A pyrolized graphite trap absorbs used caesium and non-condensing impurities are ejected into space. The NPS uses a lithium hydride reactor shield located in a hermetically sealed steel container

The singradiator that hThe radiator isparallel. The supported by ta steel radiatoabout 7 m2, wtemperature o

The autopower levels, mthe rated levesupply lines an

A high thermionic reacontrols the vooperating sectcorrecting therrises to 880 Kcontrol systempractically conwhich the onbshutdown. NPSwell as by radi

The TOefficiency of aautomatic conthas a design limaximum diam

In the pcombination wautomatic modwith TFEs usincrystal tungst with the inner load bearing elements.le circuit sodiumpotassium heat removal system includes a as load carrying capacity and also serves as a structural member. designed as a system of D-shaped tubes placed hydraulically in tubes are welded into the radiator O-ring collectors and

he load bearing elements. The tubes plane surface is soldered to r which has a high emissivity coating. The area of the radiator is

hich ensures rejection of at least 170 kW(th) at a coolant f 880 K. matic system controls the NPS to rated thermal and electrical aintains the working section current or coolant temperature at

l, maintains the voltage of ~28 V for the on-board equipment d can provide shutdown of the thermionic reactor converter. speed controller that redistributes the direct current of the ctor converter section between the spacecraft and ballast loads ltage. In the nominal operating mode, the rated current of the ion and, consequently, its electrical power are sustained by mal power. As the efficiency degrades the coolant temperature . After that, instead of maintaining the current, the automatic limits the coolant temperature. The thermal power then remains stant, whereas the operating section current will fall to values at oard network voltage exceeds allowable limits, requiring NPS shutdown is also provided for specific emergency situations, as

o command from earth.PAZ NPS generates approximately 6 kW at a start-of-life bout 5.5%. Its mass, including the nuclear power unit, the rol system and the coupling service lines, is about 1200 kg and it fetime of 4400 h. The nuclear power unit is 4.7 m long with a eter of 1.3 m.eriod 19821984, two power tests of the TOPAZ NPS in ith the automatic control system were performed in the e to prepare for flight tests. The first of these was performed g emitter assemblies of single crystal molybdenum with a single en coating; the second with TFEs using single crystal

molybdenum eand the seconalgorithms in parameters of no less than 5.6

In 1987, twith new gene

Safety waa circular operprovide sufficproducts. The attained. Deta

2 FebruaSatellite Mass: 380

10 July 19Mass: 380

The NPS142 d and theoperation was

The test presults confirmterrestrial conoperation of ththe presence o

4.2.4. Yenise

Figure 20packaged withthe shape of a

FIG. 19. Plasma-A satellite. Source: Kurchatov Institute.31

mitter assemblies. The first of the units was tested to ~4500 h d to ~7000 h. Results of testing fully corroborated the control startup and operating modes, as well as agreeing with output the NPS, its subsystems and components and electrical power of kW for the prescribed lifetime of 4400 h. wo experimental Plasma-A satellites (see Fig. 19) were launched ration TOPAZ reactors.s provided during the TOPAZ tests by placing the spacecraft in ational orbit at a height in excess of 800 km. This orbit would

ient decay time (350 a) for radioactive materials and fission reactor was only made critical once the safe orbit had been

ils are summarized as follows:

ry 1987 Cosmos-1818 Programme: Radar Ocean Reconnaissance RORSAT. Launch Site: Baikonur. Launch Vehicle: Tsyklon 2. 0 kg. Perigee: 790 km. Apogee: 810 km. Inclination: 65.0o.87 Cosmos-1867 Programme: RORSAT. Launch Site: Baikonur. 0 kg. Perigee: 797 km. Apogee: 813 km. Inclination: 65.0o.

used with the first of the spacecraft (Cosmos-1818) operated for second one (Cosmos-1867) for 342 d. In both cases the NPS terminated as planned when the caesium stock was exhausted. rogramme objectives were fulfilled for both units. The flight test ed the TOPAZ NPS output parameters and that operation in

ditions agreed with those in space. They attested to the stable e reactor converter and its support systems in space flight and in f operating plasma thrusters.

y (TOPAZ-2) NPS

shows a general view of the Yenisey NPS. All the equipment is in a single unit referred to as the reactor or head unit, which has truncated cone. The reactor is at the top, with the radiation

32

shielding locatthe shielding s

The TOarrangements. thermionic reahaving an outeof 23.7 mm (vecharacteristics

A singlecoating is usedalloy is used athrough whichTFEs are locat

FIG. 20ed immediately underneath and all other equipment arranged in hadow.

PAZ and Yenisey NPSs have similar structures and design The principal difference between them is that the Yenisey ctor converter employs a single unit TFE; the emitter unit r diameter of 19.6 mm and the collector pack an outer diameter rsus 10.0 mm and 14.6 mm, respectively, for TOPAZ). The main

of the Yenisey NPS are presented in Table 6. crystal molybdenum alloy with a single crystal tungsten 184 as the emitter material and the polycrystalline molybdenum

s the collector material. The emitter units have a central orifice the gaseous fission products are to be ejected into space. The ed in the thermionic reactor converter core tubes.

. General view of the Yenisey NPS. Source: Kurchatov Institute.

The smalIn the reactor clocated in orificonsists of 34

TABLE 6. MAIN CHARACTERISTICS OF THE YENISEY NPS

Description Value

Maximum electrical power at the reactor unit terminals supplied to consumer (kW) 5.5

Current type Direct

Voltage (V) 27

Reactor therma

Maximum coola

Maximum emitt

Lifetime corrob

Reactor unit ma

Dimensions of t Length (mm) Maximum dia

Radiation situat the core centr Fluence of n Gamma rad

Core diameter (

Core height (mm

Number of TFE

Number of rotat

Loading of uran

Effective neutro cold state) (keTotal reactivity t

Worth of 12 con

Peak to average Along to th Along to th

Lifetime ensure33

l clearance between the TFE and the tube is filled with helium. ore there are 37 TFEs with the O-ring channels for their cooling ces in zirconium hydride moderator discs. The operating section TFEs, the pumping section consisting of three. The electrical

l power (kW(th)) 135

nt temperature at the reactor outlet (C) 550

er temperature (C) 1650

orated by nuclear tests (a) 1.5

ss (kg) 1000

he reactor unit: 3900

meter (mm) 1400

ion over a plane of diameter 1.5 m at 6.5 m from e:eutrons with energy >0.1 MeV (n/cm2 ) 5 1012 iation exposure dose (R) 5 105

mm) 260

) 375

s in the core 37

ional control elements in the side reflector 12

ium-235 in the core (kg) 25

n multiplication factor (control elements out,

ff) 1.005

emperature effect (Dk/k) 0.012trol elements (Dk/k) 0.055 power density:e core radius 1.1e core height 1.26

d by the reactivity margin 3

34

power on the operating section terminals can vary through a range from 4.5 kW to 5.5 kW at about 30 V.

The Yenisey NPS, as does the TOPAZ, employs a one circuit heat removal system with a conduction type electromagnetic pump powered in startup by a heavy current storage battery.

The NPS has a design mass of 1000 kg and a lifetime of 1.5 a. It is 3.9 m long and has a maximum diameter of 1.4 m.

The Yenithe off-line tephysical tests preparatory pelectrical heate

Six terre1986, the threpreserved theistable within

The thermefficiency of abthe liquid meta

5

Nuclear power is used provide motiorocket propuls

Two parapropulsion enethe take-off mareaction using4300 s with a mfission reactorimpulse with adifferent coreswith a mass rat

The fundengines lies insey NPS went through a whole cycle of terrestrial tests, including st of the various units and systems, mechanical and thermal of full-scale mock-ups, and tests of a number of pre-launch rocesses and NPS prototypes. Testing was performed using rs in place of nuclear fuel.

strial nuclear power tests were performed between 1975 and e final tests being for 12 500, 8000 and 4700 h. All TFEs r full working capacity and their electrical characteristics were 3%.

ionic reactor converter operating section provides 4.5 kW at an out 4.5%. The tests were stopped owing to a loss of integrity of l circuits. No Yenisey NPS flight tests were made.

. NUCLEAR PROPULSION SYSTEMS

power can be used for a rocket propulsion system. The reactor to heat a propellant that is forced through the rocket nozzle to n in the opposite direction. Figure 21 shows a typical nuclear ion module.meters that provide a measure of the efficiency of a rocket rgy source are the theoretical specific impulse and the ratio of ss to the final mass in orbit. As shown in Appendix VI, chemical

hydrogen, oxygen or fluorine can achieve a specific impulse of ass ratio for earth escape of 15. However, hydrogen heated by a instead of a chemical reaction achieves twice the specific solid core while at the same time having a mass ratio of 3.2. With , the specific impulse can be as much as seven times greater again io of only 1.2.amental advantage of NTP units over liquid propellant rocket being able to use a single component working fluid with the

minimum molmaximum speproduced by N

Owing tomission with atheoretical basnuclear fusionthat is barely la

PROPELLANTTANK

FIG. 21. TypicaEnergy.35

ecular weight, which, all other factors being equal, provides a cific impulse. Thus, when using hydrogen the specific impulse TP can be more than twice as high as that of chemical engines. its higher specific impulse, NTP can perform the same space

smaller mass of propellant than a chemical engine. On a purely is it is possible to do even better with direct fission, or thermo-, with theoretical impulses rising to 36 106 s and a mass ratio rger than unity.

PROPELLANTFEED PUMP

NUCLEAR REACTORHEAT EXCHANGER

BLEED TURBINEDRIVE

HEATED PROPELLANT

l nuclear rocket propulsion module. Source: NASA/US Department of

36

5.1. US DIRECTIONS

The basic improvement in mass ratio to be obtained from a nuclear fission propulsion system was recognized very early on. In the USA, a test programme named Rover was conducted from 1955 to 1973 using various designs. Appendix VII lists the achievements that culminated in 1969 in the XE-prime engine (see Fig. 22), a first down firing prototype operating at 1100 MW. In 1972, a 44 MW nuclear furnactemperatures othe sizes of rea

The nuclcurrent dollarsthe US progra

FIG. 22. The gNo. 1 at the Nuce demonstrated peak fuel power densities of 4500 MW/m3 with f up to 2500 K for 109 min. Appendix VIII shows a comparison of ctors tested in the Rover programme.ear rocket programme, which cost $1.7 billion ($7 billion in ), was considered a technical success but it was terminated and mme has since used chemical propellants. However, nuclear

round experimental engine (XE-prime) installed in Engine Test Stand lear Rocket Development Station in Nevada. Source: NASA.

propulsion may be the option of choice for a manned mission to Mars according to recent conference announcements. Indeed, a 1990 assessment of the programme concluded that a Mars mission could be performed with NERVA technology [2].

5.1.1. Safe affordable fission engine (SAFE)

Another work is being Space Flight Cthermal powepower systems28%) flow via 100 kW(e) for 37

concept, SAFE, is a propulsive heat pipe power system on which done at the Los Alamos National Laboratory and the Marshall enter. SAFE-400 (see Fig. 23) is designed to provide 400 kW of r for more than ten years through two independent Brayton , the reactor heat being deposited into the gas (He 72%, Xe two independent heat pipe to gas heat exchangers. This provides a 25% efficiency.

SAFE-400 axial dimensions:

Fuelled length = 56 cmBeO axial refelector length = 4 cmFission gas plenum length = 5 cmTotal fuel pin length = 70 cm

Radial reflector length = 62 cmControl drum length = 56 cm

Heat exchanger plenum length = 5 cm (4x)

Heat exchanger heated length = 25 cm (2x)

Heat exchanger gaps = 2.5 cm (3x)Heat pipe length = 145 cm

FIG. 23. The SAFE-400 reactor. Source: Nuclear News.

38

The SAFE-400 reactor contains 127 identical modules made of niobiumzirconium (1 wt%) alloy. Each module contains a Nb1ZrNa heat pipe at its centre surrounded by three niobiumzirconium tubes each of which contains a rhenium clad uranium nitride fuel sleeve. The wick of the heat pipe is fabricated from Nb1Zr mesh; the 60% void being filled with sodium during operation. The heat pipes extend 75 cm outside the core. The fission power is transferred to the heat pipes at a vapour temperature of 1200 K and, thence, to the Brayton cy

The systeheating in exissince it can be to be passivelyexample, it is Moreover, it iprepare it for beryllium cont

The mas80 kg if changecases of reducbility, safety mreduced. Thus,reliability.

The technmodule SAFEmodule SAFEhas been replafabrication tec2003 in a new f

5.1.2. Heat p

Getting tperson to surApproximatelycapabilities of is impractical and geographiHOMER is sh

HOMERcally for prodcle heat exchangers.m uses existing technology and can be tested with electrical ting facilities, so development time is short. It is also flexible used with Stirling or Brayton cycles. Furthermore, it is designed safe in all credible launch or re-entry accident scenarios. For

subcritical even if fully immersed and surrounded by wet sand. s designed so that no operations are required after launch to startup. In operation, the reactor is controlled by Nb1Zr clad rol drums which have a boron carbide absorber layer.s of the reactor is 512 kg. This could be reduced to as little as s in design parameters and operations are made. However, in all ed dimensions and different control configurations, the relia-argins, ease of fabrication and ease of integration would be

the mass is a compromise between greater safety and increased

ology has been tested in heat pipe demonstrations and with a 12 -30 core. Also, the SAFE-400 has been tested through the 19 -100 programme, which is of similar design but where the Nb1Zr ced by stainless steel for economy. The programme also tested hniques. The integrated core and heat exchanger were tested in acility at the Marshall Space Flight Center.

ipe operated Mars exploration reactor (HOMER)

o Mars may be the attainment of a primary objective but for a vive on the surface a source of electrical energy is needed. 320 kW(e) are required, which just takes the task beyond the

an RTG because of the mass of plutonium required. Solar power because of the distance of Mars from the sun and for seasonal c sunlight issues. Thus, nuclear fission is the remaining option. own in Fig. 24. fulfils the need for a small power source. It is designed specifi-

ucing electricity on the surface of the planet. The low power

requirement mof power densthere is no sign

HOMERStirling engineconsidered a m

The reac316 stainless stby the 0.38g Mshield to a heengine.

REACTOR CONTROLDRIVES

STIRLING ENGINE

VIBRATION DAMPER

LOOP HEAT PIPTO RADIATOR

HEAT EXCHANG

REACTOR AXIARADIATION SHI

REACTOR ANDHEAT PIPES

FI39

eans that the reactor operates within well-understood regimes ity, burnup and fission gas release. The fluence is so low that ificant irradiation damage to core materials.-15 is a 15 kW(th) reactor designed to couple with a 3 kW via heat pipes. Since the system is low power, it can be odule of a larger array should more power be needed.

tor uses 102 uranium nitride fuel pins, each 44 cm long, clad in eel, cooled by 19 stainless steelsodium heat pipes and assisted artian gravity. The heat pipes extend 40 cm beyond the core axial at exchanger. There, the heat is transmitted to a Stirling cycle

ES

ER

LELD

2.4 m

G. 24. The HOMER power source. Source: Nuclear News.

40

5.2. SOVIET/RUSSIAN DIRECTIONS

The 1950s were the years when individual groups of specialists in the former Soviet Union initiated innovative research to work on NTP.

It is possible to create a rocket with a nuclear engine. As shown earlier, this engine can provide a specific impulse 22.5 times higher than chemically fuelled engines. For this specific impulse and the needed mass and size charac-teristics, hydrotemperature o30 kW/cm3.

In the Rthe neutron mofuel elements casing formingThis preferencwas a fundame

The fuel In the assembmeet the speci

A typica(which can e

WORKINSUPPLY

CONTROL D

CONT

FUEL

COMBUSTIONBOTTOM

NOZZLE

FIG. 25. Rgen must be heated in the engine reactor to a mass average f 3000 K. The specific power flux in the reactor core would be

ussian programme, a heterogeneous reactor layout is used with derator located separately from the uranium fuel elements. The

are surrounded by heat insulation and enclosed in the metal the complete independent reactor unit the fuel assembly. e for a heterogeneous reactor and element by element testing ntal difference between the Russian and US programmes. assembly is the basic unit of a heterogeneous propulsion reactor. ly the working fluid is heated to the temperatures required to fic impulse of the design.l reactor design (Fig. 25) includes: the cooled power casing nd with a nozzle), the fuel assembly core itself, the high

G MEDIUM

RUM DRIVE

ROL DRUM

ASSEMBLY

CHAMBER

CHAMBER

TURBO-DRIVENPUMP ASSEMBLY

REACTOR SHIELDING

MODERATOR

REACTOR VESSEL

NOZZLE COOLINGMAIN LINE

NOZZLE

ussian NTP system (conceptual design). Source: Kurchatov Institute.

temperature heat insulation, the bearing unit, the inlet unit (which provides the uniform working fluid supply to the fuel assemblies and which can contain a temperature compensation device) and the end reflector and protection elements. The essence of the concept is a nuclear reactor with a heterogeneous core comprised of individual fuel assemblies (including fuel elements based on uranium, zirconium and niobium carbides) located in the zirconium hydride moderator body. The beryllium reflector, the shadow radiation shielding and the hydrogen manned missio

Fuel requ

(a) High den(b) High resi(c) High cor(d) A maxim(e) A maxim(f) Propertie

The mossolid solutionsdensity of ~2 heating the hy

The NTPconducting fuethe fuel assemb

Also, forin Sections 5.2research and dtional innovatifor terrestrial a

5.2.1. IGR r

The IGRassembly mateto be tested winside a waterassemblies of v

The purp41

loop surround the core. Radiation shielding is included for ns.irements include:

sity of uranium per unit volume of fuel;stance to radiation swelling; rosion resistance to the working fluid; um allowable temperature of the working fluid; um number of the heatingcooling cycles; s that provide passive safety.

t acceptable fuel compositions meeting these requirements are of carbides (UCZrC, UCNbC, UCTaC) with a uranium g/cm3. This fuel would provide a temperature of 3300 K for

drogen propellant. parameters were first verified experimentally in the course of l assembly simulator tests in the IGR reactor and subsequently lies and the reactor core units were tested in the IVG-1 reactor.

the Russian NTP programmes (i.e. all three concepts described .1, 5.2.2 and 5.2.3), it should be noted that this is an area of space evelopment that can be beneficial to various ongoing interna-ve reactor and fuel cycle research and development initiatives pplications (see Section 8.7).

eactor

reactor (1961) allowed research to be undertaken on fuel and rials under full-scale operational conditions. The fuel assembly as mounted in the central experimental channel of the reactor cooled ampoule type metal structure. Fuel elements and fuel arying design were tested in the reactor.oses of these tests in the IGR reactor were to:

42

(a) Check the reliability of selected materials and protective coatings of fuel elements in hydrogen at temperatures between 3000 and 3300 K in a high neutron and irradiation environment;

(b) Substantiate optimum steady state operational temperatures of the fuel elements;

(c) Check fuel assembly structural elements and units, as well as the methods of construction and assembly of heat insulation materials;

(d) Obtain dspecific t

(e) Obtain dmodes of

(f) Investigathe amou

Followinreactor under

5.2.2. IVG-1

Apart frocharacteristics,bench prototyp

The desifuel elements power (thrust)each fuel assemof a water moddimensions angroups with a thrusts of up to

The IVGmoderator andsections. The psections includeand the centralchannel units oberyllium displtwice the averathe fuel assemb

Figure 26sectional view,ata on specific fuel assembly parameters, in particular, the hrust impulse;ata on the dynamic characteristics of fuel assemblies and optimal control;te fuel assembly operating peculiarities, in particular, to determine nt of the uranium and fission fragments released into the hydrogen.

g the loop tests, full-scale fuel assembly tests were conducted in a steady state operation.

experimental bench reactor

m testing fuel assemblies and reactor core elements with various the IVG-1 experimental reactor was destined to function as a e of a medium power (200400 kN of thrust) NTP unit.

gn of the 720 MW IVG-1 reactor is such that full-scale tests of and fuel assemblies of various types with a varying range of can be made. An individual supply of the gaseous hydrogen to

bly allows tests at rated gas outlet temperatures, while the use erator makes it possible to vary the fuel assembly cross-section

d power. The IVG-1 reactor has allowed testing of fuel assembly thrust of up to 40 t and of a single fuel assembly in a loop at 200 t.-1 reactor is a heterogeneous gas cooled reactor using a water a beryllium reflector. It consists of both permanent and replaceable ermanent sections are shown in Fig. 26. The replaceable core the central assembly which comprises a set of thirty test channels channel. The test fuel assemblies can be placed in either the test r in the central channel. As the central channel is embedded in the acer, the thermal neutron flux in the channel can be approximately ge cross-sectional flux of thermal neutrons, making it possible to test ly mounted in the central channel to failure. ((a) and (b)) shows the IVG-1 reactor cross-section and the

respectively.

FIG. 26(43

FIG. 26(a). The IVG-1 reactor cross-section.

b). The IVG-1 reactor sectional view. Source: Kurchatov Institute.

44

The IVG-1 reactor made it possible to:

(a) Confirm the selection of the structural materials used for the fuel elements and assemblies early on;

(b) Confirm the reliability of the fuel assembly designs in hydrogen within specified limits;

(c) Study the physics and thermalphysical characteristics of the fuel assemblie

(d) Study the

Successfumade it possibreactor (see Fi

FIG. 27. Startutemperature: 300s and core elements; dynamic properties of the fuel elements and assemblies.

l test performance of the fuel assemblies in the IVG-1 reactor le to start the next test stage conducting off-line tests of the g. 27). In bench tests the reactor generated a thrust of 36 kN; the

p of the IVG-1 experimental reactor (thermal power: 225 MW; hydrogen 0 K). Source: Kurchatov Institute.

tests being carried out using the IRGIT reactor, the specially designed prototype of the NTP reactor.

5.2.3. IRGIT reactor

The IRGIT reactor tests included the following stages: reactor physical startup; cold gas dynamic tests; physical startup check; cold hydrodynamic tests; powered

The phystest facility at at the Baikal-1Transferring thmechanical eff

For reasocarried out usi

Poweredpower level thmedium to a Figure 28 show

The maihensive check engineering anthe fuel fired t

(a) The therelementshydrogenassemblieof the conafter the

(b) Neutron margin, effects an

(c) The quanassemblie(on the b

(d) Operatio45

startup; fuel fired tests and post-test research.ical startup was undertaken in two stages: the first at the Strela the Institute for Physics and Power Engineering and the second test bench complex (now situated at Kurchatov in Kazakhstan). e system between two sites gave a valuable insight into the

ects of the first set of tests.ns of safety and economy, the cold hydrodynamic tests were

ng a substitute for the working fluid (nitrogen). startup, as distinct from physical startup, brings the reactor to a at is sufficient to heat the reactor structure and the working value close to the nominal design value (or a little lower). s the IRGIT reactor at the Keldysh Center Test Facility Berth.

n purposes of the fuel fired tests were to provide a compre-of the serviceability of the reactor and its