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

AAAS Space: Science & Technology Volume 2021, Article ID 5245136, 15 pages https://doi.org/10.34133/2021/5245136

Y. Xia,1,2 J. Li,1 R. Zhai,1 J. Wang,1,2 B. Lin,1 and Q. Zhou3

1Beijing Institute of Spacecraft Environment Engineering, Beijing, China 2Science and Technology on Reliability and Environmental Engineering Laboratory, Beijing, China 3Institute of Nuclear and New Energy Technology of Tsinghua University, The Key Laboratory of Advanced Reactor Engineering and Safety, Ministry of Education, Beijing, China

Correspondence should be addressed to Q. Zhou; [email protected]

Received 9 June 2020; Accepted 29 July 2020; Published 26 February 2021

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 NASA has made great efforts to make a 1-10 kWe space reactor of Kilowatt Reactor Using Stirling TechnologY At present, chemical energy [1–3] and solar energy [4–6] are (KRUSTY). The KRUSTY-based application may hopefully the main forms of energy supply for space applications. How- come into fly in this or the next decade. Fission power is ever, they have difficulty in meeting the energy demand of promising and will play a role in the great power game. future space science and space applications [7–11]. Nuclear NASA supported it under the Game Changing Development power, due to its high energy density, is believed to be one (GCD) program [23]. The ESA judges that a successful FPS of the next-generation energy solutions. Radioisotope Elec- project realization is of global impact and comparable with tric Propulsion (REP) [12–15] has been widely assessed for the Apollo and International Space Station (ISS) project [24]. outer planet explorations as a means to provide sufficient The United States, under the Systems for Nuclear Auxil- power. The power of REP systems generally falls within the iary Power (SNAP) program [25], launched the first and only 1 kW range and offers negligible propulsive capability to spacecraft powered by a nuclear reactor in 1965. In the later nonsmall spacecraft. Fission Electric Propulsion (FEP) [16– four decades, some nuclear reactor-based projects have been 22] has also been proposed and expected to expand our space supported by the US to promote space nuclear power. The exploring capability by orders of magnitude. Typical FEP projects included the Space Power Advanced Reactor (SPAR) studies assume a system power of 100 kW or more and may program [26], the Space Prototype (SP-100) [27], the Prome- require multiple launches with in-space assembly of the com- theus, and the Fission Surface Power (FSP) [28, 29]. FSP is ponents. To make up the gap between the two technologies, still in progress while the other projects have been finished 2 Space: Science & Technology

[30]. In this century, the Los Alamos National Laboratory picture. It is limited to space exploration purposes: namely, (LANL) and the Glenn Research Center (GRC) have made outer-planet missions, manned Mars/lunar surface missions, significant progress on KRUSTY. In 2018, its nonnuclear test and interstellar precursor missions. Space fission power can verification has been completed, as well as its in-zone nuclear also benefit national or planetary defense by supplying stable tests under steady-state, transient, and accident conditions or pulsed high power, enhancing spacecraft’s mobility and [31]. NASA planned for it a fly demonstration or application serviceability, altering comet/asteroid orbit, and functioning for deep space exploration missions around 2025. along with airborne and terrestrial defense systems. It also The former Soviet Union had successfully launched 33 shows commercial potential in space such as extending satel- BUK and 2 TOPAZ-I space reactor-powered missions. The lite use with much higher power, providing maintenance, BUK type adopted thermoelectric power generation technol- mobility, retrieval, transition to prolong life span, and lower ogy, with an output power of 3 kWe and a lifetime of 135 days cost of one spacecraft. It can even enable lunar and deep in orbit [32, 33]. TOPAZ-I used a thermionic stack, which space tourism [24]. weighted 1200 kg and had an output electric power of This paper reviews the developments of Kilopower and 6 kWe. It could operate in orbit for up to 342 days [32, 33]. KRUSTY’s fission power. It then further looks into fission The Russians continued its research and development of a powers’ potential in space sciences’ future. The paper Nuclear Gas Core Reactor (NGCR) for rockets since 1954 explores plans for studying Jupiter and beyond, as well as [34] and performed other research for weapons [35, 36]. In lunar and Mars missions. This review discusses the possibil- this century, Russia has revealed very little information on ity, advantages, and disadvantages of applying fission power its progress but is believed to have been carrying out research systems within the framework or constraints of a few current on megawatt space reactor technology which is mainly based missions. It is worth mentioning that, with the application of on Brayton thermoelectric transfer technology [37]. It is fission reactor, there will no longer be a power cap in the planning a fission-powered spacecraft travelling to Mars range of hundred watts to payloads, i.e., the power change mission perhaps around 2025. will change the architecture of the whole payload planning In 2013-2014, the European Disruptive technologies for and scientific achievement expectation. Specifically, space fis- space Power and Propulsion (DiPoP) project highlighted sion source is expected to upgrade missions by enabling the main issues and recommendations that are necessary to orbiters to replace flybys, landers to replace orbiters, and develop space nuclear power [38]. In 2017, the European- multiple targets to replace a single target on bases that it Russian Megawatt Highly Efficient Technologies for Space can support higher power instruments, allow more instru- Power and Propulsion Systems for Long-duration Explora- ments, increase the instruments’ duty cycles and download tion Missions (MEGAHIT) project [39] developed concepts data with higher rates, support real-time operations and in of space, ground, and nuclear demonstrators. In 2018, its situ analysis, enable electric propulsion to a lower mass, follow-up project, DEMOnstrators for Conversion, Reactor, and improve flexibility. Yet due to limited discussion and radiator and Thrusters for Electric PrOpulsion Systems pro- proposals coming up so far, this frame changing advantages ject (DEMOCRITOS) [40], succeeded in a ground-based test is unable to be covered in the scope of this article. This article of a Russian nuclear reactor with 1 MWe of power plus the refers mainly to NASA reports and projects. Thanks to important thermal emission solution by the use of droplet NASA’s openness, scholars can follow and summarize these radiators. The technical choice of test space applications pre- techniques and then form a reference for relevant researches. ferred to start from 30 kWe to 200 kWe gas-cooled or liquid metal-closed cycle Brayton and their near-earth application 2. Kilopower and KRUSTY Progress [24]. Afterwards, the International Nuclear Power and Pro- pulsion System (INPPS) flagship nonhuman (2020th) and NASA had partnered with the Department of Energy’s human (2030th) Mars exploration missions [41–44] were National Nuclear Security Administration, LANL, Nevada established to meet the celestial Earth-Mars-Earth-Jupiter/- National Security Site (NNSS), and Y-12 National Security Europa trajectory in 2026-2031 for maximal space flight tests. Complex by using their existing nuclear infrastructure. It The first flagship space flight is a complex, complete test mis- has completed a quick prototype reactor and test. The sion for the second flagship towards Mars with humans. Demonstration Using Flattop Fission (DUFF) test was With a path choice different from the United States and Rus- completed by LANL using the existing “Flattop” criticality sia, Europe started directly at a MWe reactor power system. experiment [50–52] at the NNSS to provide the nuclear They are expecting the Mars-INPPS payload mass of up to heat source. This setup is shown in Figure 1. Many “firsts” 18 tons to allow the transport of three different payloads, were established in September 2012 when the DUFF scientific, pure commercial, and new media communication experiment was completed [53, 54]. These milestones in one mission. included the first Stirling engine driven by fission source China [45–48] and Japan [49] also have had a few early- and the first heat pipe to transfer heat from reactor to staged concept designs for space reactors. Even though India, conversion system. DUFF’s success showed that a nuclear an emerging space power, is committing to growing its test of one small fission system can be accomplished with ground nuclear power capacity, it has not shown any public reasonable cost and a compact schedule [31, 55]. DUFF ambition for space nuclear power. was envisioned as a simple step, to push space fission Space fission power can enable and benefit several mis- power forward and create confidence in space reactors’ sion objectives. The article only focuses on a part of the big use in future space missions. DUFF was the predecessor Space: Science & Technology 3

5 kWt HEU 50 kWt HEU Fuel: 33 kg, U235: 28 kg Fuel: 44 kg, U235: 38 kg Reactor: 161 kg Reactor: 268 kg

Flattop Heat pipe Stirling engine 33 cm 40 cm Figure 1: DUFF experiment setup [57]. (a) (b)

5 kWt LEU 50 kWt LEU Fuel: 348 kg, U235: 57 kg Fuel: 415 kg, U235: 68 kg of KRUSTY. KRUSTY refers to a power system containing Reactor: 832 kg Reactor: 973 kg 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 comple- tion of the KRUSTY experiment in March 2018, the Kilo- power 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 [57–62] utilize a solid block of fuel and are constrained to use heat pipes to 57 cm 58 cm transfer energy from fission core. Kilopower reactors (c) (d) generally intend for simple, low power (1-10 kWe, with specific power of 2.5-6.5 W/kg). In addition, it aims for Figure 2: Kilopower reactor MCNP schematics. (a) 5 kWt HEU space and surface power systems with conceptual designs design. (b) 50 kWt HEU design. (c) 5 kWt LEU design. (d) 50 kWt aiming specifically at a 1 kWe lunar demo mission and LEU design. also a 10 kWe Mars In Situ Resource Utilization (ISRU) demo mission. Kilopower has limitations to reach power KRUSTY, a system-leveled integration concept [23, 31, scales of 100 kWe. More recently, its fuel of Uranium- 64–67], was seen as a successful step towards future space Molybdenum (UMo, provided by DOE) takes designs of fission system deployment providing long-lived, robust, reli- two different 235U enrichments: 93% (HEU) and 19.75% able, 1-10 kWe power for space missions. The basic KRUSTY (LEU) [63]. There are pros and cons for each choice, but design is fuel, control, reflector, heat pipes, and shield. overall, HEU is superior in performance. The only signifi- Figure 3 shows the basic layout of a 1 kWe system. The cant advantage of LEU is rooted in the fact that it is in 1 kWe variation of Kilopower was selected for the KRUSTY accordance with the antiproliferation policy so that there nuclear demonstration, which consists of a HEU (UMo) is a simplified approval process. Figure 2 shows MCNP reactor core operating at 800°C, with the sodium heat pipes schematics of HEU and LEU enrichments combining delivering heat from the core to eight 125 W Stirling conver- 5 kWt (1 kWe) and 50 kWt (10 kWe) power levels. In the tors. It rejects waste heat with titanium-water heat pipes figure, fuel is in yellow and the beryllium oxide (BeO) coupled to aluminum radiator panels. The core is 4 kWt, reflector in blue. The developing group of LANL believes and the Stirlings produce a net output of 1 kWe. that a Kilopower reactor concept is the simplest among KRUSTY used an electrically heated, nonfissioning ever-proposed space power reactor concepts. Its simplicity depleted uranium reactor core to complete its nonnuclear leads to high reliability for two reasons. First, Kilopower system level tests. The tests were completed in 2017 at systems can inherently survive any transients (e.g., loss the GRC [68]. It succeeded its nuclear ground testing at of power conversion heat removal) at any power level the NNSS in 2018 [59]. The KRUSTY nuclear level tests, without any control intervention needed. In addition, the as its concept design, use a prototypic HEU (UMo) core design of the heat pipes assures the independent pipe and sodium heat pipes. Eight Stirling engines in its design function which avoids a single point failure and leads to are simulated with 6 thermal simulators to save cost and 2 inherent redundancy. They also claim that Kilopower Advanced Stirling Convertors (ASC-E2’s) for performance reactors should survive launch and landing impacts evaluation. Combining these is equivalent to the thermal because of the solid fuel and less fragile heat pipes. draw of conversion units during full power. The 4 Space: Science & Technology

3. Possible Planetary Exploration Missions with Fission Power

Ti/H2O heat pipe radiator 3.1. Jupiter Icy Moon and Trojan Missions. NASA once had (truncated in picture) an ambitious mission, known as Jupiter Icy Moons Orbiter (JIMO) [70–73]. It is aimed at exploring Jupiter and its three icy moons of Callisto, Ganymede, and Europa. The JIMO Stirling power spacecraft was designed from the very beginning to be pow- conversion system ered by a nuclear reactor and an associated electrical generation system, which is called the Reactor Power Plant (RPP). Na heat pipes 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 LiH/W shielding ~1.5 m to generate 100 kWe, while the other two convertors are held in reserve. All components of the system must function for a BeO refector 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 UMo metal fuel system that was needed was not developed. The Jupiter-Europa Orbiter (JEO) [74–76] comes next,

B4C absorber rod and it is a predecessor to the Europa Clipper with a draft shown in Figure 5(a). The JEO mission would travel to Figure 3: 1 kWe KRUSTY system layout for spacecraft. Jupiter by a multiple-gravity-assist trajectory and then perform a study of the Jupiter system with a 30 months’ Jupi- ter system science phase and Europa with a Europa orbit conversion unit is gas cooled at this stage. The experiment phase of 9 months. JEO was originally designed with 5 setup is shown in Figure 4. It ran a 28-hour test with reac- Multimission Radioisotope Thermoelectric Generators tor startup, power ramp, full power steady operation, and (MMRTGs) which is equivalent to 1 Advanced Stirling shut down to simulate all scenarios that could happen in a Radioisotope Generator (ASRG) of 500 We. Considering mission. a 1 kWe FPS was used to replace the RPS, it would double The simplicity, the compactness, and the reliability on the the available power which results in much higher data existing infrastructure and nuclear material make KRUSTY a return rate and enable better data collection capabilities. quick and affordable space reactor development. KRUSTY It eliminates the need for Plutonium 238 in the RPS was tested as prototypic as possible, with a cost < $20M and version, which remains very difficult to acquire. However, a period of less than 3 years [63]. It also showed great safely with FPS, the drawback is that it increases the power features and a good possibility to be scalable, impact system mass at 260 kg for the RPS by a notably additional resistant, and long living. 360 kg [77]. Three major technological changes should be done to let The Jupiter Trojans, or simply Trojans, are the asteroid Kilopower achieve significantly higher power levels [60]. The population located at the Sun-Jupiter Lagrange points L4 primary focus is to change a solid block core to a fuel rod/pel- and L5 (i.e., 60-degree backwards and forward from Jupiter’s let core so that the fuel swelling is eliminated. The Kilopower orbit). They are of great scientific interest. As early as 2009, team has not yet addressed this issue. The second technolog- the European scientific community proposed a flyby mission ical switch is to change the Stirling to Brayton power conver- to the Jupiter Trojans [78, 79]. The Johns Hopkins University sion to adapt to powers > 50 to 100 kWe. The Stirling has its Applied Physics Laboratory carried out a study to reach the own technical challenges to reach >10 kWe power with only Jupiter Trojan asteroids to support NASA updating and one single unit. The Brayton system is more complex, but extending the National Academies Space Studies Board’s cur- the beauty of this switch is that heat pipes remove power rent solar system exploration decadal survey (2013-2022) from the fuel in the same manner. Thus, the Kilopower core and evaluated the solar electric propulsion (SEP) and REP keeps the load-following feature as the Stirling reactor. [80]. The concept draft is shown in Figure 5(b). When NASA Finally, a direct-cycle Brayton has to be taken for >1to investigated REP by using a common spacecraft design, a 3 MWe power scene. In its design, the coolant flows directly REP spacecraft to reach Jupiter’s leading Lagrange point through the holes in the core and drives the Brayton and then orbit several Trojan asteroids is the baseline design turbines in one single flow loop. This switch shall be exam- [81]. In January 2017, NASA selected the Lucy [82, 83], a ined carefully as its reactor dynamics will, unlike the first Discovery-class mission to explore Jupiter Trojan asteroids. two, be affected by the coolant with periodic changing flow Lucy is due to launch in October 2021; with Earth’s gravity velocity, temperature, and pressure which changes with load assistance, it will travel 12 years to visit one Main Belt aster- in time. Also, a gas-cooled system faces difficulties with decay oid and 6 Trojans—7different asteroids in total. These 6 par- power removal and redundancy capability. The Kilopower ticular Trojans were chosen because they have diverse team claimed they had promising results on these evolution- spectral properties and thus are expected to represent well ary concepts [60]. the remnants captured in the Sun-Jupiter Lagrangian points Space: Science & Technology 5

(a) (b) (c)

Figure 4: Conﬁguration of the 1 kWe KRUSTY nuclear demonstration. (a) Design sketch with the vacuum tank removed. (b) Real setup with the vacuum tank removed. (c) Design sketch with the vacuum tank on [69].

map. Because of the many complex phenomena discovered from Saturn by Cassini [86–88], 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 sec- ondarily 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 Jupiter Europa Trojan Tour SEP starts and uses the Earth-Venus gravity assists. The Orbiter ~600 We ~800 W SEP stage was jettisoned 5 years later at the last Earth flyby 6 MMRTGs 8 MMRTGs for the heliocentric trans-Saturn cruise. Upon arriving at Sat- 4 ASRGs 6 ASRGs urn, a chemical system of bipropellant would perform the 1 small fssion system 1 small fssion system Saturn orbit inserting maneuvering and deceleration. The (a) (b) following 2-year Saturn tour phase would provide 16 Titan flybys and ≥7 close Enceladus flybys. Afterwards, it would Figure 5: Jupiter Europa Orbiter and Trojan Tour power choice reach the Titan orbit insertion phase with the main engine between REP and FEP. (a) Jupiter Europa Orbiter. (b) Trojan Tour. aerobraking while aerosampling. It would finally place the orbiter into a 1500 km, 85° inclination polar orbit cycling L4 and L5. The results will reveal primordial material that Titan approximately 5 times for every Earth day for a dura- formed the outer planets as early as the solar system formation of 20 months. The orbiter will release montgolfière at tion age. Some researchers have reevaluated the Trojan mis- the first Titan flyby and release a lake lander on the second sion and believe that one small fission reactor could replace fl fi Titan yby as baseline. the original baseline REP con guration of ~60 We for the The montgolfière was designed to be powered by US- spacecraft functions and 750 We for propulsion [59]. supplied plutonium-fueled MMRTG (~1700 Wt and 500 We). In 2014, with the rapid progress of Kilopower and 3.2. Titan Saturn System Mission (TSSM). The Titan Saturn for US 2010 decadal survey, the GRC’s Collaborative Model- System Mission (TSSM) [84, 85] was originated by NASA ing for Parametric Assessment of Space Systems (COM- and ESA combining their independently selected missions. PASS) team reassessed TSSM with a 1 kWe fission reactor On the ESA side, it started from an L-class Titan and Encela- [89] in place of the 500 We MMRTG [90]. Here, the former dus mission (TandEM) mission in their Cosmic Vision 2015- and latter designs are referred to hereafter as the RPS version 2025 Call. On the NASA side, it got back to a 2007 NASA’s and the FPS version; a comparison is shown in Table 1. flagship-class Titan Explorer mission which was recom- The mission duration for the FEP version totals 15 years mended by NASA’s 2006 Solar System Exploration Road- and 3 months. At the last Earth flyby, the SEP is jettisoned 6 Space: Science & Technology

Table 1: Titan Saturn system mission comparison of Radioisotope Power System (RPS) and Fission Power System (FPS).

Subsystem RPS 2008 ASRG FPS 1 kWe Stirling reactor FPS 1 kWe thermoelectric reactor Science 108 kg, 182 We, ~5 Tb data return 108 kg, 182 We, ~9 Tb data return 108 kg, 182 We, ~9 Tb data return Mission ~13 yr ~15 yr (1 yr Earth spiral out) ~16 yr (2 yr Earth spiral out Atlas 551, short fairing; C of Atlas 551, long fairing; C of ~14.8 km2/s2 Atlas 551, long fairing; C of Launcher 3 3 3 0.6 km2/s2 (6250 kg stage mass) (8300 kg stage mass) ~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 ~700 kg, ~7 yr operation (reactor Orbiter power ~500 kg, ~7 yr operation (reactor activated 171 kg > 13 yr operation time activated ~7 yr after launch at the final system ~7 yr after launch at the final Earth flyby) Earth flyby) 4.5-meter drag flap plus a 2.25-meter 4-meter antenna for drag area, 4.5-meter drag flap plus a 2.25-meter antenna with the same 77 ballistic Aerobraking ballistic coefficient 77 kg/m2 (2 antenna 88 ballistic coefficient (~2.1 coefficients (2 months of aerobraking months of aerobraking campaign) months of aerobraking campaign) campaign) 4 m antenna, X/Ka, 25 We/35 We 2.25 m antenna, X/Ka, 70 We/250 We RF, 2.25 m antenna, X/Ka, 70 We/250 We Communication RF, 140 kbps 250 kbps RF, 250 kbps Attitude control Four 25 Nms reaction wheels, Four 150 Nms reaction wheels, Gravity Four 150 Nms reaction wheels, Gravity system (Titan LVLH around Titan Gradient around Titan Gradient around Titan Ops) Total S/C dry mass (with ~3200 kg ~4200 kg ~5000 kg margins) and then the reactor starts. The start at the last Earth flyby is for the fission power system life expansion and safety assur- ance 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 132 ± 142 km and a period (P) 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 ðeÞ 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 Figure 6: Titan Saturn System Mission (TSSM) with 1 kWe FEP comet-like behavior. replacing MMRTGs [77, 91]. In 2010, for the planetary science decadal survey commit- tee and primitive bodies panel, NASA’s GRC’s COMPASS egies to put an 80 kg science instrument into Chiron’s orbit at team and the Goddard Space Flight Center Architecture a 10 m imaging resolution distance. Mission restrictions Design Laboratory (ADL) ﬁrst cooperated on a Chiron included a cost cap of $800 million (as a New Frontier class) Orbiter study [93, 94]. This study evaluated spacecraft strat- and a launch window during the years 2015-2025. This study Space: Science & Technology 7

water, which have presumably never thawed. They are collec- tively 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 (e =0:07) 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 C : 2 2 upper stage to speed up to 3 =6956 km /s . The electric propulsion operates continuously for 7 years from launch (a) (b) (c) until a Jupiter gravity assist. The spacecraft then goes for a Figure 7: Chiron Orbiter spacecraft [91]. (a) REP version. (b) 9-year long coast as it approaches JBO 2001 XH255. 8 kWe FEP version in launch vehicle. (c) 8 kWe FEP version In 2012, the COMPASS team revisited KBOO with an deployed. added FEP choice [15, 96, 97]. The study found that, constrained to the same launch vehicle, trip time, and science investigated chemical, chemical/SEP, and REP propulsion payloads, an 8 kWe FEP could replace the 4 kWe REP system. architectures. The latter REP options used either 6 standard Table 3 and Figure 8 show KBOO with a FEP concept in 134 We ASRGs or 2 conceptual 550 We high-power (HP) comparison. ASRGs to power throughout the mission. The study con- A REP system keeps almost constant specific power cluded that REP propulsion was capable to deliver most around 5 We/kg. The FEP system grows specific power as science payloads to the Chiron orbit (6-ASRGs delivered a power levels increase and reach 7 We/kg @ 8 kWe power 72 kg payloads and 2-HP-ASRGs delivered 76 kg). However, level. The KBOO mission with FEP doubles the power com- neither REP choice fits the cost cap. pared to its REP counterpart, with a cost of <20% extra mass. In 2012, COMPASS reevaluated the Chiron Orbiter with Initial radioactivity provides the other key difference, with an added FEP option [95]. An 8 kWe FEP system was found the REP system having 413260 Ci of radioactivity at launch equivalent to the REP baseline and capable to support science and FEP system having only 5 Ci. This reduced radioactivity payloads. The FEP spacecraft maintained Atlas 551 launch is much safer in scenarios that have a launch failure. 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. 4. Possible Human Exploration Missions with The FEP version (4000 kg launch mass) is particularly Fission Power heavier than the REP one (1300 kg launch mass). It had to scale the reactor power (900 We to 8000 We) and electric 4.1. Human Lunar Surface Mission. On 9th April 2019, NASA thrusters three 600 W Hall with ~450 kg Xe to three 7000 W released a timeline for mission Artemis. This mission will Ion with ~1600 kg Xe) up to offset the introduced extra mass. send men and the US’s first woman onto the moon to the This increased propulsion capacity allowed the spacecraft to south pole region by 2024 [98]. The Artemis plan includes escape Earth without using the original Star 48 payload assist two phases of rapidly landing astronauts on the moon by module. Once in the orbit of Chiron, the extra power benefit 2025 and realizing a sustained human presence on and becomes more appealing. The power easily supported all the around the moon by 2028. The latter includes a Gateway out- instruments and high communication rate as in the HP- post orbiting the moon for access and transfer. This ambi- ASRG REP version. There was much extra power that might tious outpost will enable access of the moon more than ever likely influence the payload architecture. Also, as Chiron before. lacks substantial gravity, FEP enabled orbiting it (in previous NASA is planning a “multipart” landing process to trans- choices, only REP could do the same). port astronauts to and from the moon [99]. The astronauts start from the Gateway outpost, ride on a “transfer element” 3.4. Kuiper Belt Object Orbiter (KBOO). There may be hun- spacecraft down to the low-lunar orbit, and then ride on a dreds of millions of small icy objects in the Kuiper Belt [15] “descent element” spacecraft down to the lunar surface. To ring region beyond the orbit of Neptune. These objects are return, they use an “ascent element” to reach the Gateway. presumed to be leftovers from the formation of the outer NASA is planning to make these elements both reusable planets and thought to be the source of most of the short- and refuelable. The Gateway shall have great power and pro- period comets that can be observed on Earth. They are pulsion to fulfill its goal. On 23rd May 2019, NASA assigned assumed to be composed of frozen methane, ammonia, and Maxar Technologies (formerly SSL) to develop a “power and 8 Space: Science & Technology

Table 2: Chiron Orbiter mission comparison of Radioisotope Electric Propulsion (REP) and Fission Electric Propulsion (FEP).

Power subsystem REP FEP Science and mission duration 44 kg CBEa/13 yr trip, 1 yr science 44 kg CBE/13 yr trip, 1 yr science Launch vehicle Atlas 551/Star 48 Atlas 55 Launch mass (kg) 1300 4000 Six, 150 W ASRGc, 900 We/189 kg Single fast reactor, Stirling convertors 8000 We/1142 kg Power level (EOLb)/mass α (4.7 We/kg) (7 We/kg) Electric propulsion Three 600 W Hall, ~450 kg Xe Three 7000 W Ion, ~1600 kg Xe trust/weight Size (m) Deployed 2.2 16 Launch 4 (includes Star 48) 7 Nuclear material ~6 kg, 239Pu ~75 kg, 93% HEUd Initial radioactivity (Ci) 91840 4.8 aCurrent best estimate. bEnd of life. cAdvanced Stirling Radioisotope Generator. dHighly enriched uranium.

Table 3: Kuiper Belt Object Orbiter (KBOO) comparison of Radioisotope Electric Propulsion (REP) and Fission Electric Propulsion (FEP).

Power system REP FEP Science and mission duration 100 kg CBEa/16 yr trip, 1 yr science 100 kg CBE/16 yr trip, 1 yr science Launch vehicle Delta IV Heavy/Star 63F Delta IV Heavy/Star 63F Launch mass (kg) 3100 3700 Nine, 550 W ASRGc, 4000 We/782 kg Single fast reactor, Stirling convertors 8000 We/1162 kg Power level (EOLb)/mass α (5 We/kg) (7 We/kg) Electric propulsion 1+1 3000 W NEXTd Ion, ~1200 kg Xe 1+1 7000 W NEXTd Ion, direct drive, ~1200 kg Xe trust/weight Height (m) Deployed 6 16 Launch 3 7 Nuclear material ~27 kg, 238Pu ~75 kg, 93% HEUe Initial radioactivity (Ci) 413260 4.8 aCurrent best estimate. bEnd of life. cAdvanced Stirling Radioisotope Generator. dNASA’s Evolutionary Xenon Thruster. eHighly enriched uranium.

(a) (b) (c)

Figure 8: Kuiper Belt Object Orbiter (KBOO) spacecraft [91]. (a) 4 kWe REP version. (b) 8 kWe FEP version. (c) 8 kWe FEP version deployed. Space: Science & Technology 9

1.5 m

1 kWe HEU 1 kWe LEU 10 kWe HEU 10 kWe LEU lunar demo lunar demo Mars ISRU Mars ISRU (a) (b) (c) (d)

Figure 9: Lunar demo underneath the moon surface with reactor core, reflector, shield, etc.: (a) 1 kWe HEU lunar demo, (b) 1 kWe LEU lunar demo, (c) 10 kWe HEU Mars ISRU, and (d) 10 kWe LEU Mars ISRU. 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 flythe“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 competi- tor 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. Figure 10: Lunar demo above the moon surface with Stirling Meanwhile, we have noticed that the KRUSTY group is devel- engine, radiator, and other components. oping a 1 kWe lunar demo mission conceptual design with the same reactor core as Kilopower in a similar timeline. The con- stage. The power level is determined by the number of astro- cept 1 kWe FPS system with a HEU and a LEU demo is shown nauts. In early Mars missions with a crew of 4-6 astronauts, in Figure 9 [63]. This figure only shows the reactor core and the DRA plans on using 40 kWe. the necessary reflector and shield which are believed to be In 2016, the COMPASS team was commissioned by the Human Exploration and Operations Mission Directorate intended to be buried under the moon surface. Above the sur- fi face are heat pipes to conduct heat to 4 pairs of dual-opposed (HEOMD) to compare ssion versus solar systems for both Stirling engines to generate electricity and a deployed radiator, of these phases of the Martian missions [95]. Rucker et al. as shown in Figure 10 [100]. For now, they are believed to published the study [102], and the following is a brief likely support a search for resources in the permanently summary. shadowed craters. How SEP, REP/RPS, and FEP/FPS would be included in Artemis is yet too early to be known. 4.2.1. Phase I: In Situ Resource Utilization (ISRU) Demonstrator. Phase I is designed on constrains of Delta IV 4.2. Human Mars Surface Mission. The next great step Heavy as the launch vehicle, 7500 kg of the payload mass deliv- planned for human exploration after the moon is to send ered to the Jezero crater (18°51′18″N, 77°31′08″E) on Mars. An humans to Mars. NASA is trying to reach this goal first in ISRU plant will produce 4400 kg of liquid oxygen propellant for the following 20 years and has released a Mars Design Refer- a 1/5 scale demonstration. The 1/5 scale demonstrator favors in ence Architecture (DRA 5.0) [101]. DRA has baselined FPS terms of mass but requires more time to produce oxygen. as the primary power source and the 10 kWe Kilopower Solar power architecture used ATK Ultraflex™ arrays reactor as the enabling technology. and lithium ion batteries. The arrays operate at 120 The two main phases of the Mars program both require VDC with a conversion efficiency of 33%. They were novel power technology. Phase I requires power for an ISRU mounted on a gimbal that tracked the sun and sloping plant’s autonomously deploy and supply. The plant will sep- to an angle of 45° to mitigate dust. Lithium ion batteries arate oxygen from the Martian atmosphere and cryogenically store energy with a density of 165 Wh/kg. Three different store it as “ascent element” propellant. After collection of the powering architectures were compared. These were a plant necessary propellant comes Phase II. In this phase, the power operating only at daylight at 1/5 production rate (1A), a system provides support for the human crew and the science plant operating at 1/5 production full time (1B), and a 10 Space: Science & Technology

Table 4: Solar versus ﬁssion for In Situ Resource Utilization (ISRU) demonstration mission on Mars surface.

SPS 1A: 1/5 rate SPS 1B: 1/5 rate SPS 1C: 2/5 rate Power system FPS 2: 1/5 rate full time daytime only full time daytime only Total payload mass (including growth) (kg) 1128 2425 1531 2751 Electrical subsystem mass (kg) 455 1733 639 1804 ISRU subsystem mass (kg) 192 192 335 192 Power (kW) ~8 at daylight ~8 continuous ~16 at daylight ~6.5 continuous 4-7.5 m diam. Solar arrays 4-5.6 m diam. each 4-7.5 m diam. each None each Night production No Yes No Yes Liquid oxygen production (kg/sol) 4.5 10.8 9.0 10.8 Production time of 4400 kg oxygen, including 1098 527 609 407 dust storm outage (sol) 300 krad for electronics, Radiation tolerance 100 krad for electronics and ISRU 10 Mrad for ISRU

(a) (b) (a) (b) Figure 11: Mars SPS-ISRU lander concept [91]: (a) at launch; (b) Figure 12: Mars FPS-ISRU lander concept [77, 91]: (a) at launch; deployed on the surface. (b) deployed on the surface. plant operating at 2/5 production rate only in daylight Among the SPS options, the 1C architecture balanced (1C). The arrays and batteries were sized along with these best between mass and oxygen production time, but its options. For instance, a 120 d global dust storm was mod- energy storage capability did not address the follow-on ff requirements to cycle on and off every day for the crewed eled and in addition the e ect of 10 h/sol average daylight. fi The fission architecture was modeled with a 10 kWe Kilo- stage. Thus, option 1B is the nal choice as it has the minimal start and stop cycle times and satisfies the needs for both powersystem.Apermanentradiatorforthereactorwas fi ff attached to the lander. The power conversion used 8- phases. SPS 1B and FPS are not signi cantly di erent in 1.25 kWe Stirling engines in the dual opposed manner. The fis- terms of mass and production performance. sion option had no power interruptions or loss in the dust storm conditions. Also, the choice of nuclear power provided 4.2.2. Phase II: Crewed Mission. Phase II is designed on the flexible choices on landing locations. The fission system worked constraints of a Space Launch System as the launch vehicle at 65% capacity (6.5 kWe) full time during these studies. with mass has not yet been determined. This mission intends Most lander subsystems were almost identical between the to send 4-6 crew members to the Jezero crater (18°51′18″N, SPS and FPS choices, except for the thermal controls. The 77°31′08″E) and the Columbus crater (29.8°S, 166.1°W) in comparison of these Phase I options are shown in Table 4. 2038. The ISRU plant will produce 23000 kg of liquid oxygen The conceptual drafts are shown in Figures 11 and 12. propellant. Space: Science & Technology 11

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 Figure 13: NASA concept of fission reactors (4 × 10 kWe) on Mars. (MTV), and then rides on it for a 175-225 d quick Mars tran- sit. Arriving at Mars, the MTV will rendezvous and enter the Table landing vehicle, and then, the crew will start the expedition. 5: SPS versus FPS mass comparison in the crewed phase of the Mars mission. Rucker et al. [102] extended the COMPASS study to accommodate the logistics of the crew phase. The FPS con- Power generation/storage mass (kg) cept is as shown in Figure 13 and Table 5 which gives a brief SPS Crew expedition comparison between SPS and FPS. The 50 kWe FPS is com- FPS Jezero Columbus posed of four 10 kWe Kilopower units plus one backup unit. crater crater ’ It has a 12 years design life. It is delivered and deployed with Expedition 1 9154 11713 12679 the premission and powers through all three expeditions. Lander 1 9154 5611 5909 The major difference in the crewed phase is that the Lander 2 0 2034 2704a power system has to provide energy overnight and during the global dust storms. In Table 5, the first lander of each Lander 3 0 2034 2033 expedition is heavier for it as it has the energy storage and Lander 4 0 2034 2033 power management addition. The results found that for Expedition 2 0 6102 6770 crewed expeditions, FPS is around half the mass of its coun- Lander 1 0 2034 2704a terpart SPS, at both craters’ latitudes. Lander 2 0 2034 2033 The FPS architecture on Mars has three rarely debated Lander 3 0 2034 2033 advantages. First, it is tolerant to dust storms. Second, it Expedition 3 0 0 0 allows the landing location to be at any point on Mars. Last, it produces power for a long time and has potential to well Lander 1 0 0 0 expand these mission designs. There are two disadvantages Lander 2 0 0 0 of FPS. First, it produces radiation; the astronauts must keep Lander 3 0 0 0 away, and radiation safety protocols must be strictly regu- Three expeditions’ total 9154 17815 19449 lated all time. Second, it is inferior in terms of simplicity (kg) fl a and redundancy and has an unproven ight heritage com- The noted mass include additional strings from ISRU to MAV. pared 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 Mar- not yet be considered outside the scope of its current power tian rovers to change from SPS to MMRTG-based RPS. capabilities. The ISRU demonstration due in the mid-2020s has not Space reactors play an alternative addition to radioiso- yet been determined on SPS versus FPS. We can guess for tope ones in higher power missions especially for those with now that the two technologies will likely combine and add electric propulsion. The transition power level between RPS redundancy in the near-term Mars missions with SPS taking and FPS shall be concluded as a future space exploration mis- the main role in the equatorial regions and FPS in the polar sion design reference value. The NASA Nuclear Power regions. Assessment Study (NPAS) discussed this transition for three factors: mass, cost, and plutonium fuel availability. The mass 5. Discussion on Mission Applicability of and cost of FPS is larger than RPS at lower power levels until Fission Power mass breaking even in the range 8 kWe and 10 kWe and cost breaking even below 1 kWe. The NPAS study concluded Beyond these space missions that have just been discussed, ~1 kWe is a prudent transition point [103]. Learning from fission power can also facilitate other applications, such as NASA’s past missions and future estimations, a “Discovery” putting into NEO orbits for earth threatening deflection/des- mission class needs power in the range from 130 We to truction, “dead” spacecraft/space debris removal, survey, 267 We, “New Frontiers” mission class needs power in the mining, outpost as fueling/maintenance station, and high- range from 170 We to 750 We. “Flagship” missions require power ground radar, laser, microwave generator, particle 150 We to 1000 We, and manned lunar and Mars missions beam, high data rate long distance communications, and fall in the range of a few 10’s to 100 kWe. Specifically, the others high-power usages [24]. Fission power may also range is 2-10 kWe for a crewed mission’s habitat. For long- enable new conceptual missions and instruments that have range mobility, the range is 10-35 kWe for exploration 12 Space: Science & Technology science, and 15-30 kWe for ISRU. Based on this analysis and extend transition time for outer space usage. On the other only discussing on the perspective of space science missions, hand, it also gives tremendous extra power that may enable we expect that RPS and FPS will stay on the propulsion more scientific instruments to be used with more cycles. It source shelves as alternatives. In the near future, FPS are can enable higher rate communication. It also provides the more likely to be used as a flagship-class multiple targeted possibility, unseen before, to update mission. For example, primary body exploration mission or a Mars or lunar stage it can update a flyby mission to an orbiter. As a whole, FEP’s mission for the final crewed objective. NASA is almost use is sure to enrich the scientific output. certain to be the first agency to achieve this milestone. This mission-by-mission comparison agrees with NPAS’s Since last century’s SNAP and TOPAZ projects, space general comparison that concluded that FEP/FPS wins reactors have not made much progress, while Kilopower REP/RPS when power levels reach ~1 kWe in aspect of cost has been a huge success in a short time and has attracted and when power levels reach ~8 kWe in aspect of mass. worldwide attention. This gives us some ideas. Kilopower Beyond ~Saturn REP may not give a cost acceptable solution insists system simplicity. Simplicity brings both low cost and a plutonium affordable solution leaves FEP’sastheonly and high reliability. A complex new technology gains benefit choice. Objective surface missions prefer fission powers more from a simple system to show a rapid success in the develop- than the solar counterpart as it widens the location ranges to ment. With the development of modern technologies, the almost whole surfaces with far less constraints on assuring most existing concept of space reactors has changed from weather conditions and access to the sun. As long as humans static thermoelectric conversion to dynamic thermoelectric explore space, fission power will come into use sooner or later. conversion. Dynamic thermoelectric conversion maintains According to current progress, we are expecting a flagship- the advantages of high thermoelectric conversion efficiency level fission-powered space exploration mission in the next by at least a factor of 2, while eliminating the previous disad- 1-2 decades. vantage 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 Conflicts of Interest power levels. Developers of Kilopower cores and KRUSTY All authors declare no possible conflicts of interests. systems also admit that there are many technical difficulties to upgrade to the power level beyond 100 kWe. This observa- tion makes the Brayton cycle a backup choice at the MWe Authors’ Contributions level. Brayton for MWe has converged to be a choice with MEGAHIT and many international and domestic peers’ con- The authors were listed in order of contribution to this paper cepts. Another consensus that is coming into acceptance is by providing literature research and summary, academic that space reactors should be developed and applied gradu- discussion, and insights. ally from ~10 KW, ~100 kW, to ~MW. Ultrahigh power level cannot be accomplished overnight due to cost, technology References readiness, and faith in customers. [1] B. A. Palaszewski, M. L. Meyer, L. Johnson, D. M. Goebel, 6. Conclusion H. White, and D. J. 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