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Space: Science & Technology / 2021 / Article

Review Article | Open Access

Volume 2021 |Article ID 5245136 | https://doi.org/10.34133/2021/5245136

Y. Xia, J. Li, R. Zhai, J. Wang, B. Lin, Q. Zhou, "Application Prospect of Fission-Powered Spacecraft in Solar System Exploration Missions", Space: Science & Technology, vol. 2021, Article ID 5245136, 15 pages, 2021. https://doi.org/10.34133/2021/5245136

Application Prospect of Fission-Powered Spacecraft in Solar System Exploration Missions

Received09 Jun 2020
Accepted29 Jul 2020
Published26 Feb 2021

Abstract

Fission power is a promising technology, and it has been proposed for several future space uses. It is being considered for high-power missions whose goal is to explore the solar system and even beyond. Space fission power has made great progress when NASA’s 1 kWe Kilowatt Reactor Using Stirling TechnologY (KRUSTY) prototype completed a full power scale nuclear test in 2018. Its success stimulated a new round of research competition among the major space countries. This article reviews the development of the Kilopower reactor and the KRUSTY system design. It summarizes the current missions that fission reactors are being considered as a power and/or propulsion source. These projects include visiting Jupiter and Saturn systems, Chiron, and Kuiper belt object; Neptune exploration missions; and lunar and Mars surface base missions. These studies suggest that the Fission Electric Propulsion (FEP)/Fission Power System (FPS) is better than the Radioisotope Electric Propulsion (REP)/Radioisotope Power System (RPS) in the aspect of cost for missions with a power level that reaches ~1 kWe, and when the power levels reaches ~8 kWe, it has the advantage of lower mass. For a mission that travels further than ~Saturn, REP with plutonium may not be cost acceptable, leaving FEP the only choice. Surface missions prefer the use of FPS because it satisfies the power level of 10’s kWe, and FPS vastly widens the choice of possible landing location. According to the current situation, we are expecting a flagship-level fission-powered space exploration mission in the next 1-2 decades.

1. Introduction

At present, chemical energy [13] and solar energy [46] are the main forms of energy supply for space applications. However, they have difficulty in meeting the energy demand of future space science and space applications [711]. Nuclear power, due to its high energy density, is believed to be one of the next-generation energy solutions. Radioisotope Electric Propulsion (REP) [1215] has been widely assessed for outer planet explorations as a means to provide sufficient power. The power of REP systems generally falls within the 1 kW range and offers negligible propulsive capability to nonsmall spacecraft. Fission Electric Propulsion (FEP) [1622] has also been proposed and expected to expand our space exploring capability by orders of magnitude. Typical FEP studies assume a system power of 100 kW or more and may require multiple launches with in-space assembly of the components. To make up the gap between the two technologies, NASA has made great efforts to make a 1-10 kWe space reactor of Kilowatt Reactor Using Stirling TechnologY (KRUSTY). The KRUSTY-based application may hopefully come into fly in this or the next decade. Fission power is promising and will play a role in the great power game. NASA supported it under the Game Changing Development (GCD) program [23]. The ESA judges that a successful FPS project realization is of global impact and comparable with the Apollo and International Space Station (ISS) project [24].

The United States, under the Systems for Nuclear Auxiliary Power (SNAP) program [25], launched the first and only spacecraft powered by a nuclear reactor in 1965. In the later four decades, some nuclear reactor-based projects have been supported by the US to promote space nuclear power. The projects included the Space Power Advanced Reactor (SPAR) program [26], the Space Prototype (SP-100) [27], the Prometheus, and the Fission Surface Power (FSP) [28, 29]. FSP is still in progress while the other projects have been finished [30]. In this century, the Los Alamos National Laboratory (LANL) and the Glenn Research Center (GRC) have made significant progress on KRUSTY. In 2018, its nonnuclear test verification has been completed, as well as its in-zone nuclear tests under steady-state, transient, and accident conditions [31]. NASA planned for it a fly demonstration or application for deep space exploration missions around 2025.

The former Soviet Union had successfully launched 33 BUK and 2 TOPAZ-I space reactor-powered missions. The BUK type adopted thermoelectric power generation technology, with an output power of 3 kWe and a lifetime of 135 days in orbit [32, 33]. TOPAZ-I used a thermionic stack, which weighted 1200 kg and had an output electric power of 6 kWe. It could operate in orbit for up to 342 days [32, 33]. The Russians continued its research and development of a Nuclear Gas Core Reactor (NGCR) for rockets since 1954 [34] and performed other research for weapons [35, 36]. In this century, Russia has revealed very little information on its progress but is believed to have been carrying out research on megawatt space reactor technology which is mainly based on Brayton thermoelectric transfer technology [37]. It is planning a fission-powered spacecraft travelling to Mars mission perhaps around 2025.

In 2013-2014, the European Disruptive technologies for space Power and Propulsion (DiPoP) project highlighted the main issues and recommendations that are necessary to develop space nuclear power [38]. In 2017, the European-Russian Megawatt Highly Efficient Technologies for Space Power and Propulsion Systems for Long-duration Exploration Missions (MEGAHIT) project [39] developed concepts of space, ground, and nuclear demonstrators. In 2018, its follow-up project, DEMOnstrators for Conversion, Reactor, radiator and Thrusters for Electric PrOpulsion Systems project (DEMOCRITOS) [40], succeeded in a ground-based test of a Russian nuclear reactor with 1 MWe of power plus the important thermal emission solution by the use of droplet radiators. The technical choice of test space applications preferred to start from 30 kWe to 200 kWe gas-cooled or liquid metal-closed cycle Brayton and their near-earth application [24]. Afterwards, the International Nuclear Power and Propulsion System (INPPS) flagship nonhuman (2020th) and human (2030th) Mars exploration missions [4144] were established to meet the celestial Earth-Mars-Earth-Jupiter/Europa trajectory in 2026-2031 for maximal space flight tests. The first flagship space flight is a complex, complete test mission for the second flagship towards Mars with humans. With a path choice different from the United States and Russia, Europe started directly at a MWe reactor power system. They are expecting the Mars-INPPS payload mass of up to 18 tons to allow the transport of three different payloads, scientific, pure commercial, and new media communication in one mission.

China [4548] and Japan [49] also have had a few early-staged concept designs for space reactors. Even though India, an emerging space power, is committing to growing its ground nuclear power capacity, it has not shown any public ambition for space nuclear power.

Space fission power can enable and benefit several mission objectives. The article only focuses on a part of the big picture. It is limited to space exploration purposes: namely, outer-planet missions, manned Mars/lunar surface missions, and interstellar precursor missions. Space fission power can also benefit national or planetary defense by supplying stable or pulsed high power, enhancing spacecraft’s mobility and serviceability, altering comet/asteroid orbit, and functioning along with airborne and terrestrial defense systems. It also shows commercial potential in space such as extending satellite use with much higher power, providing maintenance, mobility, retrieval, transition to prolong life span, and lower cost of one spacecraft. It can even enable lunar and deep space tourism [24].

This paper reviews the developments of Kilopower and KRUSTY’s fission power. It then further looks into fission powers’ potential in space sciences’ future. The paper explores plans for studying Jupiter and beyond, as well as lunar and Mars missions. This review discusses the possibility, advantages, and disadvantages of applying fission power systems within the framework or constraints of a few current missions. It is worth mentioning that, with the application of fission reactor, there will no longer be a power cap in the range of hundred watts to payloads, i.e., the power change will change the architecture of the whole payload planning and scientific achievement expectation. Specifically, space fission source is expected to upgrade missions by enabling orbiters to replace flybys, landers to replace orbiters, and multiple targets to replace a single target on bases that it can support higher power instruments, allow more instruments, increase the instruments’ duty cycles and download data with higher rates, support real-time operations and in situ analysis, enable electric propulsion to a lower mass, and improve flexibility. Yet due to limited discussion and proposals coming up so far, this frame changing advantages is unable to be covered in the scope of this article. This article refers mainly to NASA reports and projects. Thanks to NASA’s openness, scholars can follow and summarize these techniques and then form a reference for relevant researches.

2. Kilopower and KRUSTY Progress

NASA had partnered with the Department of Energy’s National Nuclear Security Administration, LANL, Nevada National Security Site (NNSS), and Y-12 National Security Complex by using their existing nuclear infrastructure. It has completed a quick prototype reactor and test. The Demonstration Using Flattop Fission (DUFF) test was completed by LANL using the existing “Flattop” criticality experiment [5052] at the NNSS to provide the nuclear heat source. This setup is shown in Figure 1. Many “firsts” were established in September 2012 when the DUFF experiment was completed [53, 54]. These milestones included the first Stirling engine driven by fission source and the first heat pipe to transfer heat from reactor to conversion system. DUFF’s success showed that a nuclear test of one small fission system can be accomplished with reasonable cost and a compact schedule [31, 55]. DUFF was envisioned as a simple step, to push space fission power forward and create confidence in space reactors’ use in future space missions. DUFF was the predecessor of KRUSTY. KRUSTY refers to a power system containing a reactor, heat transfer, Stirling power conversion, heat rejection control, and an overarching structure. KRUSTY now is developing Kilopower reactor concepts, which refers to only the fission reactors. After successful completion of the KRUSTY experiment in March 2018, the Kilopower project team began developing mission concepts for possible future flight demonstrations [56] to the moon or Mars. It is also conducting risk reduction research and improvements for better flight preparations.

Kilopower reactor concepts [5762] utilize a solid block of fuel and are constrained to use heat pipes to transfer energy from fission core. Kilopower reactors generally intend for simple, low power (1-10 kWe, with specific power of 2.5-6.5 W/kg). In addition, it aims for space and surface power systems with conceptual designs aiming specifically at a 1 kWe lunar demo mission and also a 10 kWe Mars In Situ Resource Utilization (ISRU) demo mission. Kilopower has limitations to reach power scales of 100 kWe. More recently, its fuel of Uranium-Molybdenum (UMo, provided by DOE) takes designs of two different 235U enrichments: 93% (HEU) and 19.75% (LEU) [63]. There are pros and cons for each choice, but overall, HEU is superior in performance. The only significant advantage of LEU is rooted in the fact that it is in accordance with the antiproliferation policy so that there is a simplified approval process. Figure 2 shows MCNP schematics of HEU and LEU enrichments combining 5 kWt (1 kWe) and 50 kWt (10 kWe) power levels. In the figure, fuel is in yellow and the beryllium oxide (BeO) reflector in blue. The developing group of LANL believes that a Kilopower reactor concept is the simplest among ever-proposed space power reactor concepts. Its simplicity leads to high reliability for two reasons. First, Kilopower systems can inherently survive any transients (e.g., loss of power conversion heat removal) at any power level without any control intervention needed. In addition, the design of the heat pipes assures the independent pipe function which avoids a single point failure and leads to inherent redundancy. They also claim that Kilopower reactors should survive launch and landing impacts because of the solid fuel and less fragile heat pipes.

KRUSTY, a system-leveled integration concept [23, 31, 6467], was seen as a successful step towards future space fission system deployment providing long-lived, robust, reliable, 1-10 kWe power for space missions. The basic KRUSTY design is fuel, control, reflector, heat pipes, and shield. Figure 3 shows the basic layout of a 1 kWe system. The 1 kWe variation of Kilopower was selected for the KRUSTY nuclear demonstration, which consists of a HEU (UMo) reactor core operating at 800°C, with the sodium heat pipes delivering heat from the core to eight 125 W Stirling convertors. It rejects waste heat with titanium-water heat pipes coupled to aluminum radiator panels. The core is 4 kWt, and the Stirlings produce a net output of 1 kWe.

KRUSTY used an electrically heated, nonfissioning depleted uranium reactor core to complete its nonnuclear system level tests. The tests were completed in 2017 at the GRC [68]. It succeeded its nuclear ground testing at the NNSS in 2018 [59]. The KRUSTY nuclear level tests, as its concept design, use a prototypic HEU (UMo) core and sodium heat pipes. Eight Stirling engines in its design are simulated with 6 thermal simulators to save cost and 2 Advanced Stirling Convertors (ASC-E2’s) for performance evaluation. Combining these is equivalent to the thermal draw of conversion units during full power. The conversion unit is gas cooled at this stage. The experiment setup is shown in Figure 4. It ran a 28-hour test with reactor startup, power ramp, full power steady operation, and shut down to simulate all scenarios that could happen in a mission.

The simplicity, the compactness, and the reliability on the existing infrastructure and nuclear material make KRUSTY a quick and affordable space reactor development. KRUSTY was tested as prototypic as possible, with a and a period of less than 3 years [63]. It also showed great safely features and a good possibility to be scalable, impact resistant, and long living.

Three major technological changes should be done to let Kilopower achieve significantly higher power levels [60]. The primary focus is to change a solid block core to a fuel rod/pellet core so that the fuel swelling is eliminated. The Kilopower team has not yet addressed this issue. The second technological switch is to change the Stirling to Brayton power conversion to adapt to to 100 kWe. The Stirling has its own technical challenges to reach >10 kWe power with only one single unit. The Brayton system is more complex, but the beauty of this switch is that heat pipes remove power from the fuel in the same manner. Thus, the Kilopower core keeps the load-following feature as the Stirling reactor. Finally, a direct-cycle Brayton has to be taken for >1 to 3 MWe power scene. In its design, the coolant flows directly through the holes in the core and drives the Brayton turbines in one single flow loop. This switch shall be examined carefully as its reactor dynamics will, unlike the first two, be affected by the coolant with periodic changing flow velocity, temperature, and pressure which changes with load in time. Also, a gas-cooled system faces difficulties with decay power removal and redundancy capability. The Kilopower team claimed they had promising results on these evolutionary concepts [60].

3. Possible Planetary Exploration Missions with Fission Power

3.1. Jupiter Icy Moon and Trojan Missions

NASA once had an ambitious mission, known as Jupiter Icy Moons Orbiter (JIMO) [7073]. It is aimed at exploring Jupiter and its three icy moons of Callisto, Ganymede, and Europa. The JIMO spacecraft was designed from the very beginning to be powered by a nuclear reactor and an associated electrical generation system, which is called the Reactor Power Plant (RPP). The baseline RPP consists of a reactor which is cooled by a liquid lithium loop that connects 4 dual-opposed Stirling convertors. In operation, two convertors operate full time to generate 100 kWe, while the other two convertors are held in reserve. All components of the system must function for a period of 12 years, and the reactor needs to run at full power for a period of 2 years. JIMO was canceled as the fission system that was needed was not developed.

The Jupiter-Europa Orbiter (JEO) [7476] comes next, and it is a predecessor to the Europa Clipper with a draft shown in Figure 5(a). The JEO mission would travel to Jupiter by a multiple-gravity-assist trajectory and then perform a study of the Jupiter system with a 30 months’ Jupiter system science phase and Europa with a Europa orbit phase of 9 months. JEO was originally designed with 5 Multimission Radioisotope Thermoelectric Generators (MMRTGs) which is equivalent to 1 Advanced Stirling Radioisotope Generator (ASRG) of 500 We. Considering a 1 kWe FPS was used to replace the RPS, it would double the available power which results in much higher data return rate and enable better data collection capabilities. It eliminates the need for Plutonium 238 in the RPS version, which remains very difficult to acquire. However, with FPS, the drawback is that it increases the power system mass at 260 kg for the RPS by a notably additional 360 kg [77].

The Jupiter Trojans, or simply Trojans, are the asteroid population located at the Sun-Jupiter Lagrange points L4 and L5 (i.e., 60-degree backwards and forward from Jupiter’s orbit). They are of great scientific interest. As early as 2009, the European scientific community proposed a flyby mission to the Jupiter Trojans [78, 79]. The Johns Hopkins University Applied Physics Laboratory carried out a study to reach the Jupiter Trojan asteroids to support NASA updating and extending the National Academies Space Studies Board’s current solar system exploration decadal survey (2013-2022) and evaluated the solar electric propulsion (SEP) and REP [80]. The concept draft is shown in Figure 5(b). When NASA investigated REP by using a common spacecraft design, a REP spacecraft to reach Jupiter’s leading Lagrange point and then orbit several Trojan asteroids is the baseline design [81]. In January 2017, NASA selected the Lucy [82, 83], a Discovery-class mission to explore Jupiter Trojan asteroids. Lucy is due to launch in October 2021; with Earth’s gravity assistance, it will travel 12 years to visit one Main Belt asteroid and 6 Trojans—7 different asteroids in total. These 6 particular Trojans were chosen because they have diverse spectral properties and thus are expected to represent well the remnants captured in the Sun-Jupiter Lagrangian points L4 and L5. The results will reveal primordial material that formed the outer planets as early as the solar system formation age. Some researchers have reevaluated the Trojan mission and believe that one small fission reactor could replace the original baseline REP configuration of ~60 We for the spacecraft functions and 750 We for propulsion [59].

3.2. Titan Saturn System Mission (TSSM)

The Titan Saturn System Mission (TSSM) [84, 85] was originated by NASA and ESA combining their independently selected missions. On the ESA side, it started from an L-class Titan and Enceladus mission (TandEM) mission in their Cosmic Vision 2015-2025 Call. On the NASA side, it got back to a 2007 NASA’s flagship-class Titan Explorer mission which was recommended by NASA’s 2006 Solar System Exploration Roadmap. Because of the many complex phenomena discovered from Saturn by Cassini [8688], the TSSM is to investigate Titan as a system. It is primarily to study its upper atmosphere, the neutral atmosphere, surface to interior, and its interactions with the magnetosphere. It will shine light on its origin, evolution, and astrobiological potential. TSSM secondarily observes Enceladus’ and Saturn’s magnetosphere. TSSM was proposed to launch in 2020 and should arrive at the Saturn system in 2029.

TSSM was designed with multiple propulsion technologies [85]. Reaching the geosynchronous transfer orbit, the SEP starts and uses the Earth-Venus gravity assists. The SEP stage was jettisoned 5 years later at the last Earth flyby for the heliocentric trans-Saturn cruise. Upon arriving at Saturn, a chemical system of bipropellant would perform the Saturn orbit inserting maneuvering and deceleration. The following 2-year Saturn tour phase would provide 16 Titan flybys and ≥7 close Enceladus flybys. Afterwards, it would reach the Titan orbit insertion phase with the main engine aerobraking while aerosampling. It would finally place the orbiter into a 1500 km, 85° inclination polar orbit cycling Titan approximately 5 times for every Earth day for a duration of 20 months. The orbiter will release montgolfière at the first Titan flyby and release a lake lander on the second Titan flyby as baseline.

The montgolfière was designed to be powered by US-supplied plutonium-fueled MMRTG (~1700 Wt and 500 We). In 2014, with the rapid progress of Kilopower and for US 2010 decadal survey, the GRC’s Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team reassessed TSSM with a 1 kWe fission reactor [89] in place of the 500 We MMRTG [90]. Here, the former and latter designs are referred to hereafter as the RPS version and the FPS version; a comparison is shown in Table 1.


SubsystemRPS 2008 ASRGFPS 1 kWe Stirling reactorFPS 1 kWe thermoelectric reactor

Science108 kg, 182 We, ~5 Tb data return108 kg, 182 We, ~9 Tb data return108 kg, 182 We, ~9 Tb data return
Mission~13 yr~15 yr (1 yr Earth spiral out)~16 yr (2 yr Earth spiral out
LauncherAtlas 551, short fairing; of 0.6 km2/s2 (6250 kg stage mass)Atlas 551, long fairing; of ~14.8 km2/s2 (8300 kg stage mass)Atlas 551, long fairing; of ~22 km2/s2 (9600 kg stage mass)
SEP stage~15 kWe, 500 kg Xe, 2+1 NEXT~19 kWe, 1400 kg Xe, 2+1 NEXT~19 kWe, 1800 kg Xe, 3+1 NEXT
Orbiter power system operation time~500 kg, ~7 yr operation (reactor activated ~7 yr after launch at the final Earth flyby)~700 kg, ~7 yr operation (reactor activated ~7 yr after launch at the final Earth flyby)
Aerobraking4-meter antenna for drag area, ballistic coefficient 77 kg/m2 (2 months of aerobraking campaign)4.5-meter drag flap plus a 2.25-meter antenna with the same 77 ballistic coefficients (2 months of aerobraking campaign)4.5-meter drag flap plus a 2.25-meter antenna 88 ballistic coefficient (~2.1 months of aerobraking campaign)
Communication4 m antenna, X/Ka, 25 We/35 We RF, 140 kbps2.25 m antenna, X/Ka, 70 We/250 We RF, 250 kbps2.25 m antenna, X/Ka, 70 We/250 We RF, 250 kbps
Attitude control system (Titan Ops)Four 25 Nms reaction wheels, LVLH around TitanFour 150 Nms reaction wheels, Gravity Gradient around TitanFour 150 Nms reaction wheels, Gravity Gradient around Titan
Total S/C dry mass (with margins)~3200 kg~4200 kg~5000 kg

The mission duration for the FEP version totals 15 years and 3 months. At the last Earth flyby, the SEP is jettisoned and then the reactor starts. The start at the last Earth flyby is for the fission power system life expansion and safety assurance by avoiding any radioactive leakage in a reentry failure scenario. The FEP version improved the mission by powering all the science instruments simultaneously. It therefore provides a higher communication data rate while it changes to a smaller antenna. The FPS version was favorable with an increase of 950 kg more while the mission time is 2 years more than the original RPS version. The study team boldly recommended using a 10 kWe FPS to incorporate a FEP system to take place of SEP stage. This change could simplify the spacecraft and reduce the chemical propellant and therefore could potentially reduce the total spacecraft mass. This 10 kWe version makes an all-sided attractive option than the RPS one. Figure 6 shows the FPS-powered TSSM spacecraft draft [91].

3.3. Chiron Orbiter

The 2060 Chiron or 1977 UB [92] is the closest primitive planet-crossing Centaur-type Cb objects with a diameter of and a period () of 50.794 years. Its characteristics are similar to both comets and asteroids and thus share its name as asteroid 2060 and also comet 95P/Chiron. Chiron is an unusual planetary body. It has a highly eccentric orbit with of 0.382 with a perihelion of 8 AU which is just inside Saturn’s orbit and the aphelion of 19 AU is in Uranus’ orbit. This minor planet is an ideal candidate for primitive body research as it is visible near perihelion. Unlike many other Centaur objects, it exhibits comet-like behavior.

In 2010, for the planetary science decadal survey committee and primitive bodies panel, NASA’s GRC’s COMPASS team and the Goddard Space Flight Center Architecture Design Laboratory (ADL) first cooperated on a Chiron Orbiter study [93, 94]. This study evaluated spacecraft strategies to put an 80 kg science instrument into Chiron’s orbit at a 10 m imaging resolution distance. Mission restrictions included a cost cap of $800 million (as a New Frontier class) and a launch window during the years 2015-2025. This study investigated chemical, chemical/SEP, and REP propulsion architectures. The latter REP options used either 6 standard 134 We ASRGs or 2 conceptual 550 We high-power (HP) ASRGs to power throughout the mission. The study concluded that REP propulsion was capable to deliver most science payloads to the Chiron orbit (6-ASRGs delivered a 72 kg payloads and 2-HP-ASRGs delivered 76 kg). However, neither REP choice fits the cost cap.

In 2012, COMPASS reevaluated the Chiron Orbiter with an added FEP option [95]. An 8 kWe FEP system was found equivalent to the REP baseline and capable to support science payloads. The FEP spacecraft maintained Atlas 551 launch vehicle and the 13-year time period by adding 7000 W ion engines. Figure 7 and Table 2 show the comparison of REP with the FEP Chiron Orbiter.


Power subsystemREPFEP

Science and mission duration44 kg CBEa/13 yr trip, 1 yr science44 kg CBE/13 yr trip, 1 yr science
Launch vehicleAtlas 551/Star 48Atlas 55
Launch mass (kg)13004000
Power level (EOLb)/mass αSix, 150 W ASRGc, 900 We/189 kg (4.7 We/kg)Single fast reactor, Stirling convertors 8000 We/1142 kg (7 We/kg)
Electric propulsion trust/weightThree 600 W Hall, ~450 kg XeThree 7000 W Ion, ~1600 kg Xe
Size (m)
 Deployed2.216
 Launch4 (includes Star 48)7
Nuclear material~6 kg, 239Pu~75 kg, 93% HEUd
Initial radioactivity (Ci)918404.8

aCurrent best estimate. bEnd of life. cAdvanced Stirling Radioisotope Generator. dHighly enriched uranium.

The FEP version (4000 kg launch mass) is particularly heavier than the REP one (1300 kg launch mass). It had to scale the reactor power (900 We to 8000 We) and electric thrusters three 600 W Hall with ~450 kg Xe to three 7000 W Ion with ~1600 kg Xe) up to offset the introduced extra mass. This increased propulsion capacity allowed the spacecraft to escape Earth without using the original Star 48 payload assist module. Once in the orbit of Chiron, the extra power benefit becomes more appealing. The power easily supported all the instruments and high communication rate as in the HP-ASRG REP version. There was much extra power that might likely influence the payload architecture. Also, as Chiron lacks substantial gravity, FEP enabled orbiting it (in previous choices, only REP could do the same).

3.4. Kuiper Belt Object Orbiter (KBOO)

There may be hundreds of millions of small icy objects in the Kuiper Belt [15] ring region beyond the orbit of Neptune. These objects are presumed to be leftovers from the formation of the outer planets and thought to be the source of most of the short-period comets that can be observed on Earth. They are assumed to be composed of frozen methane, ammonia, and water, which have presumably never thawed. They are collectively called Kuiper belt objects (KBOs) and also trans-Neptunian objects (TNOs). The KBOO mission picked trans-Neptunian Kuiper Belt Object 2001 XH255 as a target [15]. It has a slight eccentricity () orbit with a perihelion of 32.28 AU and a semimajor axis of 34.81 AU.

In 2011, GRC’s COMPASS team studied a REP spacecraft for a KBOO mission. The Jet Propulsion Laboratory’s (JPL’s) Team X similarly evaluated a Neptune Flagship Orbiter with REP. Both studies led to the KBOO mission. KBOO was scheduled to occur in 2030 and would take 16 years for a trip and conduct 1 year of science exploration. This concept used either 11 of 420 W ARTGs or 9 of 550 W HP-ASRGs. The original REP spacecraft has 3180 kg mass and will launch on a Delta IV Heavy rocket with a Star 63F upper stage to speed up to . The electric propulsion operates continuously for 7 years from launch until a Jupiter gravity assist. The spacecraft then goes for a 9-year long coast as it approaches JBO 2001 XH255.

In 2012, the COMPASS team revisited KBOO with an added FEP choice [15, 96, 97]. The study found that, constrained to the same launch vehicle, trip time, and science payloads, an 8 kWe FEP could replace the 4 kWe REP system. Table 3 and Figure 8 show KBOO with a FEP concept in comparison.


Power systemREPFEP

Science and mission duration100 kg CBEa/16 yr trip, 1 yr science100 kg CBE/16 yr trip, 1 yr science
Launch vehicleDelta IV Heavy/Star 63FDelta IV Heavy/Star 63F
Launch mass (kg)31003700
Power level (EOLb)/mass αNine, 550 W ASRGc, 4000 We/782 kg (5 We/kg)Single fast reactor, Stirling convertors 8000 We/1162 kg (7 We/kg)
Electric propulsion trust/weight1+1 3000 W NEXTd Ion, ~1200 kg Xe1+1 7000 W NEXTd Ion, direct drive, ~1200 kg Xe
Height (m)
 Deployed616
 Launch37
Nuclear material~27 kg, 238Pu~75 kg, 93% HEUe
Initial radioactivity (Ci)4132604.8

aCurrent best estimate. bEnd of life. cAdvanced Stirling Radioisotope Generator. dNASA’s Evolutionary Xenon Thruster. eHighly enriched uranium.

A REP system keeps almost constant specific power around 5 We/kg. The FEP system grows specific power as power levels increase and reach 7 We/kg @ 8 kWe power level. The KBOO mission with FEP doubles the power compared to its REP counterpart, with a cost of <20% extra mass. Initial radioactivity provides the other key difference, with the REP system having 413260 Ci of radioactivity at launch and FEP system having only 5 Ci. This reduced radioactivity is much safer in scenarios that have a launch failure.

4. Possible Human Exploration Missions with Fission Power

4.1. Human Lunar Surface Mission

On 9th April 2019, NASA released a timeline for mission Artemis. This mission will send men and the US’s first woman onto the moon to the south pole region by 2024 [98]. The Artemis plan includes two phases of rapidly landing astronauts on the moon by 2025 and realizing a sustained human presence on and around the moon by 2028. The latter includes a Gateway outpost orbiting the moon for access and transfer. This ambitious outpost will enable access of the moon more than ever before.

NASA is planning a “multipart” landing process to transport astronauts to and from the moon [99]. The astronauts start from the Gateway outpost, ride on a “transfer element” spacecraft down to the low-lunar orbit, and then ride on a “descent element” spacecraft down to the lunar surface. To return, they use an “ascent element” to reach the Gateway. NASA is planning to make these elements both reusable and refuelable. The Gateway shall have great power and propulsion to fulfill its goal. On 23rd May 2019, NASA assigned Maxar Technologies (formerly SSL) to develop a “power and propulsion element” as the first element of the Gateway. It targeted a launch to demonstrate its capability in late 2022.

For now, NASA announces to fly the “power and propulsion element” by means of SEP, but with three times more power than any SEP previously flown. Heretofore, mainly due to the power range being close to the SEP’s limit, the scientific community believes that FEP will be a strong competitor of SEP in human moon missions. In 2018, NASA and DOE stated that the test success of KRUSTY might lead to a mid-sized lunar lander flight demonstration in the mid-2020s. Meanwhile, we have noticed that the KRUSTY group is developing a 1 kWe lunar demo mission conceptual design with the same reactor core as Kilopower in a similar timeline. The concept 1 kWe FPS system with a HEU and a LEU demo is shown in Figure 9 [63]. This figure only shows the reactor core and the necessary reflector and shield which are believed to be intended to be buried under the moon surface. Above the surface are heat pipes to conduct heat to 4 pairs of dual-opposed Stirling engines to generate electricity and a deployed radiator, as shown in Figure 10 [100]. For now, they are believed to likely support a search for resources in the permanently shadowed craters. How SEP, REP/RPS, and FEP/FPS would be included in Artemis is yet too early to be known.

4.2. Human Mars Surface Mission

The next great step planned for human exploration after the moon is to send humans to Mars. NASA is trying to reach this goal first in the following 20 years and has released a Mars Design Reference Architecture (DRA 5.0) [101]. DRA has baselined FPS as the primary power source and the 10 kWe Kilopower reactor as the enabling technology.

The two main phases of the Mars program both require novel power technology. Phase I requires power for an ISRU plant’s autonomously deploy and supply. The plant will separate oxygen from the Martian atmosphere and cryogenically store it as “ascent element” propellant. After collection of the necessary propellant comes Phase II. In this phase, the power system provides support for the human crew and the science stage. The power level is determined by the number of astronauts. In early Mars missions with a crew of 4-6 astronauts, the DRA plans on using 40 kWe.

In 2016, the COMPASS team was commissioned by the Human Exploration and Operations Mission Directorate (HEOMD) to compare fission versus solar systems for both of these phases of the Martian missions [95]. Rucker et al. published the study [102], and the following is a brief summary.

4.2.1. Phase I: In Situ Resource Utilization (ISRU) Demonstrator

Phase I is designed on constrains of Delta IV Heavy as the launch vehicle, 7500 kg of the payload mass delivered to the Jezero crater (18°5118N, 77°3108E) on Mars. An ISRU plant will produce 4400 kg of liquid oxygen propellant for a 1/5 scale demonstration. The 1/5 scale demonstrator favors in terms of mass but requires more time to produce oxygen.

Solar power architecture used ATK Ultraflex™ arrays and lithium ion batteries. The arrays operate at 120 VDC with a conversion efficiency of 33%. They were mounted on a gimbal that tracked the sun and sloping to an angle of 45° to mitigate dust. Lithium ion batteries store energy with a density of 165 Wh/kg. Three different powering architectures were compared. These were a plant operating only at daylight at 1/5 production rate (1A), a plant operating at 1/5 production full time (1B), and a plant operating at 2/5 production rate only in daylight (1C). The arrays and batteries were sized along with these options. For instance, a 120 d global dust storm was modeled and in addition the effect of 10 h/sol average daylight.

The fission architecture was modeled with a 10 kWe Kilopower system. A permanent radiator for the reactor was attached to the lander. The power conversion used 8-1.25 kWe Stirling engines in the dual opposed manner. The fission option had no power interruptions or loss in the dust storm conditions. Also, the choice of nuclear power provided flexible choices on landing locations. The fission system worked at 65% capacity (6.5 kWe) full time during these studies.

Most lander subsystems were almost identical between the SPS and FPS choices, except for the thermal controls. The comparison of these Phase I options are shown in Table 4. The conceptual drafts are shown in Figures 11 and 12.


Power systemSPS 1A: 1/5 rate daytime onlySPS 1B: 1/5 rate full timeSPS 1C: 2/5 rate daytime onlyFPS 2: 1/5 rate full time

Total payload mass (including growth) (kg)1128242515312751
Electrical subsystem mass (kg)45517336391804
ISRU subsystem mass (kg)192192335192
Power (kW)~8 at daylight~8 continuous~16 at daylight~6.5 continuous
Solar arrays4-5.6 m diam. each4-7.5 m diam. each4-7.5 m diam. eachNone
Night productionNoYesNoYes
Liquid oxygen production (kg/sol)4.510.89.010.8
Production time of 4400 kg oxygen, including dust storm outage (sol)1098527609407
Radiation tolerance100 krad for electronics and ISRU300 krad for electronics, 10 Mrad for ISRU

Among the SPS options, the 1C architecture balanced best between mass and oxygen production time, but its energy storage capability did not address the follow-on requirements to cycle on and off every day for the crewed stage. Thus, option 1B is the final choice as it has the minimal start and stop cycle times and satisfies the needs for both phases. SPS 1B and FPS are not significantly different in terms of mass and production performance.

4.2.2. Phase II: Crewed Mission

Phase II is designed on the constraints of a Space Launch System as the launch vehicle with mass has not yet been determined. This mission intends to send 4-6 crew members to the Jezero crater (18°5118N, 77°3108E) and the Columbus crater (29.8°S, 166.1°W) in 2038. The ISRU plant will produce 23000 kg of liquid oxygen propellant.

Referring to DRA 5.0 [101], four to six astronauts will stay on Mars for ~500 days and carry through 3 expeditions at 3 different locations. Each expedition has a premission that sends a larger payload as cargo landers with a lower energy trajectory. The lander houses the power unit, propellant production unit, and the Mars Ascent Vehicle (MAV). They convert the atmosphere CO2 into O2 and store it in the MAV. After the adequate propellant production and storage has been confirmed, the crew travels to a low-Earth orbit (LEO), rendezvouses with the Mars Transfer Vehicle (MTV), and then rides on it for a 175-225 d quick Mars transit. Arriving at Mars, the MTV will rendezvous and enter the landing vehicle, and then, the crew will start the expedition.

Rucker et al. [102] extended the COMPASS study to accommodate the logistics of the crew phase. The FPS concept is as shown in Figure 13 and Table 5 which gives a brief comparison between SPS and FPS. The 50 kWe FPS is composed of four 10 kWe Kilopower units plus one backup unit. It has a 12 years’ design life. It is delivered and deployed with the premission and powers through all three expeditions.


Crew expeditionPower generation/storage mass (kg)
FPSSPS
Jezero craterColumbus crater

Expedition 191541171312679
 Lander 1915456115909
 Lander 2020342704a
 Lander 3020342033
 Lander 4020342033
Expedition 2061026770
 Lander 1020342704a
 Lander 2020342033
 Lander 3020342033
Expedition 3000
 Lander 1000
 Lander 2000
 Lander 3000
Three expeditions’ total (kg)91541781519449

aThe noted mass include additional strings from ISRU to MAV.

The major difference in the crewed phase is that the power system has to provide energy overnight and during the global dust storms. In Table 5, the first lander of each expedition is heavier for it as it has the energy storage and power management addition. The results found that for crewed expeditions, FPS is around half the mass of its counterpart SPS, at both craters’ latitudes.

The FPS architecture on Mars has three rarely debated advantages. First, it is tolerant to dust storms. Second, it allows the landing location to be at any point on Mars. Last, it produces power for a long time and has potential to well expand these mission designs. There are two disadvantages of FPS. First, it produces radiation; the astronauts must keep away, and radiation safety protocols must be strictly regulated all time. Second, it is inferior in terms of simplicity and redundancy and has an unproven flight heritage compared to SPS. The key challenge for SPS is that it accumulates dust on its solar panels which reduced the access to sunlight. This very reason supports Spirit, Opportunity and such Martian rovers to change from SPS to MMRTG-based RPS.

The ISRU demonstration due in the mid-2020s has not yet been determined on SPS versus FPS. We can guess for now that the two technologies will likely combine and add redundancy in the near-term Mars missions with SPS taking the main role in the equatorial regions and FPS in the polar regions.

5. Discussion on Mission Applicability of Fission Power

Beyond these space missions that have just been discussed, fission power can also facilitate other applications, such as putting into NEO orbits for earth threatening deflection/destruction, “dead” spacecraft/space debris removal, survey, mining, outpost as fueling/maintenance station, and high-power ground radar, laser, microwave generator, particle beam, high data rate long distance communications, and others high-power usages [24]. Fission power may also enable new conceptual missions and instruments that have not yet be considered outside the scope of its current power capabilities.

Space reactors play an alternative addition to radioisotope ones in higher power missions especially for those with electric propulsion. The transition power level between RPS and FPS shall be concluded as a future space exploration mission design reference value. The NASA Nuclear Power Assessment Study (NPAS) discussed this transition for three factors: mass, cost, and plutonium fuel availability. The mass and cost of FPS is larger than RPS at lower power levels until mass breaking even in the range 8 kWe and 10 kWe and cost breaking even below 1 kWe. The NPAS study concluded ~1 kWe is a prudent transition point [103]. Learning from NASA’s past missions and future estimations, a “Discovery” mission class needs power in the range from 130 We to 267 We, “New Frontiers” mission class needs power in the range from 170 We to 750 We. “Flagship” missions require 150 We to 1000 We, and manned lunar and Mars missions fall in the range of a few 10’s to 100 kWe. Specifically, the range is 2-10 kWe for a crewed mission’s habitat. For long-range mobility, the range is 10-35 kWe for exploration science, and 15-30 kWe for ISRU. Based on this analysis and only discussing on the perspective of space science missions, we expect that RPS and FPS will stay on the propulsion source shelves as alternatives. In the near future, FPS are more likely to be used as a flagship-class multiple targeted primary body exploration mission or a Mars or lunar stage mission for the final crewed objective. NASA is almost certain to be the first agency to achieve this milestone.

Since last century’s SNAP and TOPAZ projects, space reactors have not made much progress, while Kilopower has been a huge success in a short time and has attracted worldwide attention. This gives us some ideas. Kilopower insists system simplicity. Simplicity brings both low cost and high reliability. A complex new technology gains benefit from a simple system to show a rapid success in the development. With the development of modern technologies, the most existing concept of space reactors has changed from static thermoelectric conversion to dynamic thermoelectric conversion. Dynamic thermoelectric conversion maintains the advantages of high thermoelectric conversion efficiency by at least a factor of 2, while eliminating the previous disadvantage of short life. It is expecting to have a 20-year practical life span. Space reactors may have different reactor and/or thermoelectric conversion technology choices at different power levels. Developers of Kilopower cores and KRUSTY systems also admit that there are many technical difficulties to upgrade to the power level beyond 100 kWe. This observation makes the Brayton cycle a backup choice at the MWe level. Brayton for MWe has converged to be a choice with MEGAHIT and many international and domestic peers’ concepts. Another consensus that is coming into acceptance is that space reactors should be developed and applied gradually from ~10 KW, ~100 kW, to ~MW. Ultrahigh power level cannot be accomplished overnight due to cost, technology readiness, and faith in customers.

6. Conclusion

Space reactor power or propulsion is a near-future game-changing technology, which all great powers are examining for their projects. Space reactors could benefit civil and military missions for a great range of both space locations and power levels. This article focuses only on space exploration missions and discusses fission comparison to its original radioisotope or solar power-based solutions.

Targeting on the solar system opportunities from near to far, Jupiter missions can replace 600 We REP with 1 kWe FEP, Trojan missions can replace 800 We REP with 1 kWe FEP, Saturn system missions can replace 500 We REP with 1 kWe FEP, Chiron and Neptune missions could replace 900 We REP with 8 kWe FEP, and the Kuiper belt object missions could replace 4 kWe REP with 8 kWe FEP. As in object surface missions, reactor-powered lunar Artemis missions are under consideration but it is too early to have a definite opinion. Nevertheless, fission reactors for both Mars’ ISRU and crewed phase are coming into plans, with 10 kWe almost equaling the performance of solar arrays and 50 kWe a clear advantage. Generally, FEP increased the launch mass and therefore requires to update the propulsion level and to extend transition time for outer space usage. On the other hand, it also gives tremendous extra power that may enable more scientific instruments to be used with more cycles. It can enable higher rate communication. It also provides the possibility, unseen before, to update mission. For example, it can update a flyby mission to an orbiter. As a whole, FEP’s use is sure to enrich the scientific output.

This mission-by-mission comparison agrees with NPAS’s general comparison that concluded that FEP/FPS wins REP/RPS when power levels reach ~1 kWe in aspect of cost and when power levels reach ~8 kWe in aspect of mass. Beyond ~Saturn REP may not give a cost acceptable solution and a plutonium affordable solution leaves FEP’s as the only choice. Objective surface missions prefer fission powers more than the solar counterpart as it widens the location ranges to almost whole surfaces with far less constraints on assuring weather conditions and access to the sun. As long as humans explore space, fission power will come into use sooner or later. According to current progress, we are expecting a flagship-level fission-powered space exploration mission in the next 1-2 decades.

Conflicts of Interest

All authors declare no possible conflicts of interests.

Authors’ Contributions

The authors were listed in order of contribution to this paper by providing literature research and summary, academic discussion, and insights.

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