Anthony J. Colozza

Vantage Partners, LLC, Brook Park, Ohio

Robert L. Cataldo

Glenn Research Center, Cleveland, Ohio

Applicability of STEM-RTG and High-Power

SRG Power Systems to the Discovery and Scout 

Mission Capabilities Expansion (DSMCE)

Study of ASRG-Based Missions

NASA/TM—2015-218885

October 2015
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Anthony J. Colozza

Vantage Partners, LLC, Brook Park, Ohio

Robert L. Cataldo

Glenn Research Center, Cleveland, Ohio

Applicability of STEM-RTG and High-Power

SRG Power Systems to the Discovery and Scout 

Mission Capabilities Expansion (DSMCE)

Study of ASRG-Based Missions

NASA/TM—2015-218885

October 2015

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Glenn Research Center

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NASA/TM—2015-218885 1 

Applicability of STEM-RTG and High-Power SRG Power Systems 

to the Discovery and Scout Mission Capabilities Expansion 

(DSMCE) Study of ASRG-Based Missions 

Anthony J. Colozza 

Vantage Partners, LLC 

Brook Park, Ohio 44142 

Robert L. Cataldo 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This study looks at the applicability of utilizing the Segmented Thermoelectric Modular Radioisotope 

Thermoelectric Generator (STEM-RTG) or a high-power radioisotope generator to replace the Advanced 

Stirling Radioisotope Generator (ASRG), which had been identified as the baseline power system for a 

number of planetary exploration mission studies. Nine different Discovery-Class missions were examined 

to determine the applicability of either the STEM-RTG or the high-power SRG power systems in 

replacing the ASRG. The nine missions covered exploration across the solar system and included orbiting 

spacecraft, landers and rovers. Based on the evaluation a ranking of the applicability of each alternate 

power system to the proposed missions was made. 

Study Purpose 

The objective of this study was to assess potential benefits that a new radioisotope power system 

concept might have on a Discovery-class NASA space science mission. In 2007, nine mission concept 

studies (discussed in more detail later) were awarded that focused on missions that would be enabled by 

the implementation of the Advanced Stirling Radioisotope Generator (ASRG), which was then in a late 

phase of development. The positive results of these concept studies led to NASA offering the ASRG for 

the Discovery 12 Announcement of Opportunity (AO) in 2010. 

Since that time, budgetary realignments within the NASA Science Mission Directorate (SMD) ended 

the ASRG flight project; however, a Stirling technology maturation effort continues to be funded, along 

with thermoelectric technology development and maturation activities. 

During the summer of 2014, a fresh look at the potential suite of SMD missions and corresponding 

power requirements was undertaken as part of the recent NASA Nuclear Power Assessment Study 

(NPAS) activity, also described later. In order to increase the database of radioisotope-enabled Discoveryclass mission concepts beyond those proposed for Discovery 12, a “look back” at the DSCME studies was 

undertaken to characterize the pros and cons of the system concepts developed during the NPAS as 

contrasted to the original ASRG mission concepts. 

Background 

In the summer of 2014 NASA completed a study evaluating the potential benefits of advanced 

radioisotope power systems (RPS) and small fission power systems to future NASA robotic science 

missions. During the execution of the Nuclear Power Assessment Study (NPAS), several concepts for 

new RPS were evaluated over a range of mission classes, including Discovery, New Frontiers and 

Flagship opportunities. In 2010, the Announcement of Opportunity (AO) for Discovery 12 allotted up to, 
NASA/TM—2015-218885 2 

two Advanced Stirling Radioisotope Generators (ASRG) as one of the four NASA incentivized 

technologies offered to proposers. 

In 2007, NASA SMD initiated a NASA Research Opportunity in Space and Earth Science (ROSES) 

solicitation seeking proposals for a Discovery and Scout Mission Capabilities Expansion (DSMCE) study. 

Nine proposals were downselected to assess the potential application of the ASRG and to evaluate the 

benefits to their proposed mission. The encouraging results of the studies paved way for the ASRG being 

offered in the Discovery 12 AO; two of the final three competing proposals that were awarded Phase A 

funding utilized two ASRGs each in their mission concepts. 

At the time of the studies, the predicted ASRG mass and mission power provided as guidance for the 

DSMCE studies were 24 kg and 145 We beginning of mission (BOM) power, respectively. As the ASRG 

design matured, the predicted BOM power level decreased to ~135 We and the mass increased to 32 kg. 

Since the ASRG flight project was subsequently cancelled, the opportunity emerged to revisit future 

mission power requirements for a wide spectrum of potential SMD missions, with a goal of ascertaining 

whether a single unit could provide sufficient power output to serve these missions. As part of the NPAS 

effort, the RPS Program’s Mission Analysis Team (MAT) assessed past Discovery, New Frontiers and 

Flagship missions as well as the National Research Council Visions and Voyages for Planetary Science in 

the Decade 2013 to 2022 decadal survey report for power requirements with hope of identifying a 

strategic power level that could be implemented on a host of missions using a single design. 

After compiling the power requirements of all the relevant mission concepts that where enabled by or 

could benefit from RPS, a candidate unit size of approximately 300 We emerged as a plausible solution. 

In the assessment of the three classes of SMD missions, the Discovery class fell within ~250 to 300 We; 

New Frontiers ~300 to 600 We, and Flagships between 900 to 1200 We. In addition, the number of power 

systems to be integrated on a spacecraft was limited to four due to typical available area on a spacecraft 

body and the challenges of integrating multiple units at the launch site. 

Thus if a new system were to be developed, a 300 We end-of-mission (EOM) power output would be 

desirable (based on a 14-year mission timeline), where a single unit would be used on Discovery class 

missions, 1 to 2 units on New Frontiers and 3 to 4 units on Flagship missions. Two concept designs were 

developed for the NPAS: a high-power Stirling Radioisotope Generator (SRG) and a Segmented 

Thermoelectric Modular Radioisotope Thermoelectric Generator (STEM-RTG). 

To better validate the attributes of these concepts and how they actually provide benefit to various 

missions, a task was initiated to quantify these benefits by taking a look at the representative DSMCE 

studies for Discovery that originally utilized the 145 We ASRGs and retrofit these concepts with 

~300 We SRG and the STEM-RTG. 

Introduction 

The Advanced Stirling Radioisotope Generator (ASRG) power system was the baseline power system 

identified for a number of potential space missions. However, its recent cancelation means that other 

potential power system will need to be identified that could meet the requirements of those (or similar) 

future mission. 

The ASRG system, illustrated in Figure 1, would generate electrical power through a Stirling engine 

operating off of two General Purpose Heat Source (GPHS) blocks that contain the system’s plutonium 

dioxide fuel. The system shown is based on the configuration established through the ASRG design 

process. The specifications of the ASRG are given in Table 1. 

The Stirling engine within the ASRG converts heat from the isotope blocks to provide electrical 

power. Each radioisotope block, GPHS, provides 250 W of thermal power at their beginning of life. The 

blocks are comprised of a number of plutonium dioxide pellets, as illustrated in Figure 2. Each block 

measures 9.948 by 9.32 by 5.82 cm and has a mass of 1.44 kg (Ref. 2). 
NASA/TM—2015-218885 3 

Figure 1.—Illustration of ASRG Arrangement and Components (Ref. 1). 

TABLE 1.—ASRG SPECIFICATIONS (REF. 1) 

Property Value 

Beginning of life output power 130 W 

Thermal to electrical conversion efficiency 26 percent 

Dimensions 0.76 by 0.46 by 39 m 

Mass 32 kg 

Figure 2.—GPHS Block Assembly Drawing (Ref. 2). 
NASA/TM—2015-218885 4 

Other radioisotope-based power systems can be considered as alternatives to mission studies that were 

designed around the ASRG system. Two potential systems that may be applicable are the STEM-RTG and 

the high power SRG. These systems are in the early stages of development but have projected characteristics 

similar to that of the ASTG and therefore may be applicable to missions that utilized the ASRG. 

Advanced Radioisotope Thermoelectric Generator 

The Jet Propulsion Laboratory and Aerojet-Rockedyne devised the STEM-RTG system as a potential 

follow-on to the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG), now in use on Mars 

by the Curiosity rover. The STEM-RTG concept is based on thermoelectric heat-to-power conversion 

technology and would provide increased performance over the MMRTG in both reduced mass and higher 

efficiency. This notional system is illustrated in Figure 3. 

The base number of GPHS blocks used as the heat source within this system can vary from 8 up to 18 

depending on the power needs of the mission concept. The estimated specifications and operational 

performance of the STEM-RTG are given in Table 2. 

Figure 3.—Illustration of the STEM-RTG Power System Conceptual Design (Ref. 3). 

TABLE 2.—STEM-RTG SPECIFICATIONS (REF. 3) 

Property Value 

Number of GPHS blocks 8 12 16 18 

Beginning of life power 217 W 334 W 454 W 513 W 

Beginning of mission power 207 W 319 W 432 W 489 W 

End of mission power (17 years) 166 W 255 W 346 W 391 W 

Dimensions length, max diameter 0.58 m, 0.52 m 0.83 m, 0.53 m 1.07 m, 0.54 m 1.19 m, 0.54 m 

Heat rejection temp. (sink temp. 4 K, 270 K) 459 K, 472 K 459 K, 472 K 460 K, 473 K 460 K, 473 K 

Beginning of life specific power 7.5 W/kg 8.0 W/kg 8.4 W/kg 8.6 W/kg 

Mass 29.2 kg 42.2 kg 54.2 kg 60.3 kg 

Beginning of life efficiency 10.9% 11.1% 11.3% 11.4% 
NASA/TM—2015-218885 5 

High Power SRG 

The high-power SRG would be a follow-on to the ASRG but with the prospect of operating at higher 

power levels, by utilizing more GPHS blocks. The notional SRG system could utilize from 2 to 8 GPHS 

blocks to provide heat to operate the Stirling engines. This conceptual power system is still in its early 

design stages of development. Initial designs borrow heavily from the ASRG in components and 

operating conditions. However, advanced materials and components are being considered to increase the 

performance by reducing mass and thereby increasing the specific power of the system. A conceptual 

design is illustrated in Figure 4. 

The estimated performance and specifications of the SRG conceptual design are given in Table 3. 

Figure 4.—High Power SRG Conceptual Illustration (Ref. 3). 

TABLE 3.—SRG SPECIFICATIONS (REF. 3) 

Property Value 

Number of GPHS blocks 2 4 6 8 

Beginning of life power 116 W 215 W 331 W 456 W 

Beginning of mission power 113 W 207 W 319 W 440 W 

End of mission power (17 years) 93 W 173 W 266 W 366 W 

Dimensions length, max diameter 0.19 m, 0.50 m 0.33 m, 0.45 m 0.33 m, 0.65 m 0.36 m, 0.95 m 

Heat rejection temp. (sink temp. 4 K, 270 K) 400 K, 440 K 428 K, 468 K 428 K, 468 K 408 K, 448 K 

Beginning of life specific power 7.5 W/kg 7.5 W/kg 7.9 W/kg 7.9 W/kg 

Mass 17.3 kg 32.0 kg 46.8 kg 64.6 kg 

Beginning of life efficiency 26.0% 24.0% 24.7% 25.5% 
NASA/TM—2015-218885 6 

Mission Options 

The predicted output power level range of the ASRG, as well as the STEM-RTG and SRG, matches 

well with Discovery–class missions. These missions are designed to be lower cost, highly focused science 

missions, selected on a short timescale every two to four years. This type of mission utilizes spacecraft 

and surface vehicles that generally utilize less than 1 kw of electrical power, with many in the sub-500 W 

level. Also the lifetimes of planetary science missions, which are on the order of multiple years, fit well 

with RPS. A listing of past Discovery mission and their power requirements is given in Table 4. 

All of the previous and scheduled Discovery missions (with InSight due to launch in 2016) utilized a 

photovoltaic array/battery system for electrical power. However, future missions are likely to continue to 

extend the Discovery program further out into the solar system where photovoltaic systems may not be 

practical. Therefore alternative power systems, such as RPS, can provide the ability to meet these mission 

needs. The ASRG power system was identified as a prime candidate for a number of potential future 

Discovery class missions. These missions are listed in Table 5. 

Since the ASRG flight development project has been stopped, the mission concepts that were based 

on utilizing it as a power system should be reexamined to see if other power systems would be applicable. 

Each of the missions listed in Table 5 could also potentially utilize either the STEM-RTG or SRG as a 

replacement for the ASRG. An initial evaluation of substituting these potential power systems for each of 

those missions is given in the following sections. 

TABLE 4.—PAST AND PLANNED DISCOVERY MISSIONS (REF. 4) 

Mission Approximate 

launch mass, 

kg 

Base power output Planned lifetime Power 

system 

Pathfinder 900 <25 W 90 days PV/Battery 

NEAR 500 1800 W (at 1 AU) 

880 W (at 1.5 AU) 

<100 W (S/C Cruise) 

6 years PV/Battery 

Lunar prospector 300 200 W 1 year PV/Battery 

Stardust 400 330 W 7 years PV/Battery 

Genesis 650 260 W 3 years PV/Battery 

CONTOUR 1000 200 W 6 years PV/Battery 

Messenger 1200 650 W 10 years PV/Battery 

Deep Impact 650 620 W (spacecraft) 6 years PV/Battery 

Dawn 1300 1300 W 10 years PV/Battery 

Kepler 900 1100 W 4 years PV/Battery 

GRAIL 600 800 W 1 year PV/Battery 

InSight (launch in 2016) 700 450 W 2 years PV/Battery 

TABLE 5.—FUTURE DISCOVER MISSION OPTIONS 

Mission Approximate

launch mass,

kg 

Base power output Planned 

lifetime, 

years 

DSMCE 1: Venus Atmospheric Long-Duration Observatory for 

in situ Research (VALOR) 

1500 140 W (on Balloon) 1 

DSMCE 2: Lunar Lander (EXOMOON) 1500 143 W 1 

DSMCE 3: Mars Polar Ice Borehole (Kuklos) 900 160.5 W Nominal 

600–900 W Drill 

3 

DSMCE 4: Lunar Polar Volatile Explorer (JEDI) 3600 1160 Max W, 

280 W Nominal 

1.5 

DSMCE 5: Io Volcano Observer (IVO) 2400 262 W 8 

DSMCE 6: Trojan Asteroid Lander (Ilion) 1500 256 W 11 

DSMCE 7: Comet Coma Rendezvous Sample Return (CCRSR) 1500 202 W 14 

DSMCE 8: Titan Mare Explorer (TiME) 1500 242 W 8 

DSMCE 9: Comet Hopper (CHopper) 1400 267 W 10 
NASA/TM—2015-218885 7 

DSMCE 1: Venus Atmospheric Long-Duration Observatory for In-Situ Research 

(VALOR) (Ref. 5) 

The VALOR mission would explore the upper atmosphere of Venus at a level of approximately 

55 km for a minimum of 30 Earth days. This would be accomplished by utilizing a super-pressure 

balloon, as illustrated in Figure 5, which would operate from the mid-latitudes to the pole. The science 

goals for the mission concept are to investigate the key processes, composition and dynamics of the 

Venus atmosphere. This included measuring reactive species within the atmosphere, Nobel gases and 

their isotopes and light gas species. The mission would have also provided meteorological measurements, 

lighting monitoring and global circulation mapping. To accomplish this the primary science instruments 

would have included: 

 Gas Chromatograph Mass Spectrometer 

 Atmospheric Structure Instrument 

 Radio Tracking 

 Lightning Detector 

Figure 5.—Venus Super-Pressure Balloon Concept Illustration (Ref. 5). 

 
NASA/TM—2015-218885 8 

The transit to Venus is illustrated in Figure 6. The transit shows that the closest approach to the Sun 

would have occurred at Venus. 

The ASRG power system was selected as the proposed balloon power system since it could provide 

long life power for the 30-day baseline mission under nighttime and low-Sun angle operation. The design 

used a single ASRG located on the gondola as shown in Figure 7. 

Figure 6.—Proposed Earth to Venus Transfer Orbit for the VALOR Mission (Ref. 5). 

Figure 7.—ASRG Location in VALOR 

Concept Aeroshell and Gondola (Ref. 5). 
NASA/TM—2015-218885 9 

The ASRG would provide up to 140 W of continuous power during operation within the Venus 

atmosphere. The total entry system mass was estimated to be 931 kg with a 23 percent contingency. The 

operating pressure within the Venus atmosphere would be 0.5 bar of CO2. The atmosphere also contains 

sulfuric acid (H2SO4) droplets, so any exposed surfaces of the power system would need to be resistant to 

this corrosive acid. Other reactive gasses that are expected to be encountered would include water vapor 

(H2O), carbon monoxide (CO) and carbonyl sulfide (OCS). 

The ASRG power system limited the entry g-load to 30 g’s. The ASRG also required a thermal 

interface to remove heat from the aeroshell during transit to Venus, and the aeroshell must be capable of 

absorbing the waste heat generated by the ASRG during entry until the aeroshell separates during descent. 

The waste heat provided by the ASRG was estimated to be 360 W with a heat rejection temperature of 

80 °C. The operating conditions at the design altitude for the balloon are 0.5 bar atmospheric pressure 

with a temperature of 30 °C. 

For the VALOR mission either the eight GPHS block STEM-RTG power system or the four GPHS 

block SRG power system would be a viable candidate to replace the ASRG. The eight-block STEM-RTG 

system and four-block SRG system have an estimated beginning of mission power of 207 W, which is 

greater than the required 140 W identified for the mission. Since the mission duration would be relatively 

short, the power output of either the STEM-RTG or SRG would be above or near the 200 W level 

throughout the mission. The increased power could provide opportunities for increased science 

operations. The estimated masses of the STEM-RTG and SRG power systems are 29.2 and 32 kg, 

respectively. However, the ASRG mass was not a significant driver for the mission design. Therefore, a 

change in the mass of the power system should not significantly impact the mission. A change in mass 

would affect the Aerostat sizing somewhat and could provide extra mass for science payload. Therefore, 

providing extra power could potentially enable additional science instruments to be carried onboard the 

gondola. 

Potential concerns with the STEM-RTG and SRG systems as a replacement for the ASRG would be 

in the design of the spacecraft and aeroshell during transit. The added waste heat generated from these 

systems would need to be accommodated to maintain the interior temperature of the spacecraft during 

transit within acceptable levels. Upon entry into the Venus atmosphere, the aeroshell would need to be 

sufficiently sized to absorb the additional waste heat. Within the Venus atmosphere the extra waste heat 

should not be a concern. Also, since only a single ASRG was being used, and there was sufficient volume 

available around it, the larger volume of these systems should not require the configuration within the 

aeroshell or gondola to be changed. 

DSMCE 2: Lunar Polar Volatiles Mobile Lander (ExoMoon) (Ref. 6) 

The ExoMoon mission concept would send a lander with limited mobility to explore a lunar polar 

crater. The lander design is illustrated in Figure 8. The crater would be selected to provide direct-to-Earth 

communications periodically over a two-week cycle. The lander would utilize a single ASRG for 

providing power. The main science goals are to determine the character, composition, physical state, 

abundance and spatial variability of volatiles within a permanently shadowed region on the moon. Also, 

the lander would be used to determine the history and sources of the volatiles, and the processes for how 

the volatiles were deposited. The main science instruments include; 

 Subsurface sampler 

 Mass Spectrometer 

 Neutron Spectrometer 

 Electric Field Instrument 

 Electrostatic Analyzer 

 Low Light Spectral Imager 
NASA/TM—2015-218885 10 

Figure 8.—ExoMoon Mobile Lander Concept Illustration (Ref. 6). 

The single ASRG power system would have supplied approximately 143 W of output power 

continuously throughout the mission with a mass estimate of 32 kg. The ASRG system was selected as 

the main power source since there would be no solar radiation available within the permanently shaded 

region of the crater where the lander would be operating within, and the mission duration of greater than 

1 year would not be feasible with a purely battery powered vehicle. 

The estimated power requirements throughout the ExoMoon mission are given in Table 6. 

For the lunar landing phase of the mission, the estimated power requirement would be greater than the 

output of the ASRG. To accommodate this power level a 20 A-hr battery is utilized to make up the power 

difference for this phase of the mission. The battery would then be recharged by the ASRG for use later in 

the mission if needed. 

The ASRG would be located on the top surface external to the vehicle, as show in Figure 8. Waste 

heat is radiated directly to space or the surroundings. No additional thermal interface with the lander is 

required. The mobile lander is installed within the Atlas V401 launch fairing in an upright orientation as 

shown in Figure 9. 

For the ExoMoon mobile lander, either the eight GPHS block STEM-RTG power system or the four 

GPHS block SRG power system could be a viable candidate to replace the ASRG. This mission duration 

is much shorter than the estimated 17-year operational time for the STEM-RTG or SRG. Therefore the 

EOM output power for both systems is close to the BOM output power. Both the eight-block STEM-RTG 

system and four GPHS block SRG system have an estimated end of the ExoMoon mission output power 

of ~200 W, with masses of 29.2 and 32 kg, respectively. Both of these power systems provide similar 

operating capabilities and mass as the single ASRG systems used in the spacecraft design. Both systems 

also provide greater output power providing greater power margin for the mission as well as potentially 

adding additional capabilities to the mission. The increased output power could be used to reduce the size 

of the battery needed during the landing phase of the mission. 
NASA/TM—2015-218885 11 

TABLE 6.—EXOMOON MISSION POWER REQUIREMENTS (REF. 6) 

Mission Launch Trans-Lunar 

coast 

Lunar landing Science operations 

Total required power 85.07 W 121.77 W 231.53 W 77.97 W 

ASRG power output 143.00 W 143.00 W 143.00 W 143.00 W 

ASRG power margin 41.1% 16.0% –57.7% 46.0% 

Figure 9.—The ExoMoon Mobile Lander Within the Launch Vehicle Fairing (Ref. 6). 

Since the ASRG would be mounted on the exterior of the lander, any change in volume or the 

increase in waste heat associated with utilizing either the SRG or STEM-RTG should not have a 

significant effect on the lander design or layout. Also, as shown in Figure 9, there is ample space 

surrounding the rover within the faring so any change in dimensions or volume of the power system 

should not have a significant effect on the launch vehicle. 

DSMCE 3: Mars Polar Ice Borehole (Kuklos) (Ref. 7) 

The Kuklos mission concept would explore the historical climate, water and atmosphere of Mars by 

burrowing a hole in the north polar ice of Mars and imaging the strata along the walls of the hole. The 

system would thermally drill a borehole by melting tens of meters into the north polar cap. During the 

boring process a camera would record the visible stratigraphy along the borehole walls. A pump would be 

used to extract the melt water as it is produced and send it to the surface, where additional analysis on its 

composition would be conducted. The lander was designated to be ASRG powered and operate year 

round in the Mars polar region. The lander, shown in Figure 10 is based off of the Phoenix lander 

platform and would utilize two ASRGs. 

The operating time on the surface of Mars would be 113+ days. The notional mission timeline for the 

mission is given in Figure 11. The ASRG was selected as the power system because it enabled continuous 

year-long drilling capability, and it would allow for low sun angle and wintertime operation at high 

latitudes. Throughout the mission from launch to operation the highest temperature the ASRG case would 

reach is 92 °C. This occurs while in transit while on the surface of Mars the maximum temperature is 

approximately 22 °C. The estimated ASRG case temperature during transit and while on the surface is 

shown in Figure 12. 
NASA/TM—2015-218885 12 

Figure 10.—Kuklos Polar Lander Concept (Ref. 7). 

Figure 11.—Kuklos Mission Concept Timeline (Ref. 7). 

Figure 12.—ASRG Case Temperature for Kuklos While in Transit and on the Surface of Mars (Ref. 7).
NASA/TM—2015-218885 13 

The science instruments chosen for the mission concept include two thermal drills with temperature 

sensors, an IceCam for visualizing the stratigraphy in high resolution, a laser spectrometer for water 

evaluation, inorganic chemistry analyzer for dust examination, meteorological package for measuring 

pressure, temperature, humidity and wind, surface imaging camera and a miniature seismometer. 

The thermal drill uses electrical power to heat a tip that melts the ice and then pumps up the water to 

the surface. The drill consumes the largest amount of power of all the science equipment. The best and 

worst case estimates range from approximately 600 to 900 W and vary with depth. The remaining science 

equipment and lander systems require a nominal power of 160.5 W. 

The power available from the two ASRGs would be 270 W while on the surface and 256 W in cruise. 

During cruise, the power available from the ASRGs would be augmented by a solar array/battery system 

to provide the total power needed by the vehicle for this phase of the mission, estimated to be 320 W. The 

base constant power requirement is 72 W. Additional power up to the design power of 270 W provides 

the ability to recharge the batteries used for the drilling process. The lander would not utilize the waste 

heat generated by the ASRGs. 

The ASRG mass with contingency was estimated to be 48 kg for the combination of both units. 

To meet the lander power requirements with two ASRG systems, batteries were also incorporated. 

These batteries would be charged by the ASRGs and then used to supplement the ASRG output to 

provide the full vehicle load power of over 1 kW during drilling. The drilling process would be done 

incrementally. Once the batteries are discharged, the drilling stops and the ASRGs are used to recharge 

the batteries for the next drilling sequence. 

Because of the volume requirements of the ASRGs, the Phoenix heritage aeroshell was increased in 

diameter from 2.65 to 3.25 m. 

For the Kuklos lander and mission concept, either the twelve GPHS block STEM-RTG power system 

or the six GPHS block SRG power system could be viable candidates to replace the ASRG. Both the 

twelve GPHS block STEM-RTG system and six GPHS block SRG system have an estimated beginning 

of mission output power of 319 W with a mass of 42.2 and 46.8 kg, respectively. Since the mission 

duration would be relatively short the output power from either power system will be near the beginning 

of mission power level and above the 270 W that would have been supplied by the ASRGs throughout the 

mission. The end of mission power levels listed in tables 2 and 3 are for long duration missions of 

17 years. For shorter duration mission, such as Kuklos the output power from either the STEM-RTG or 

SRG would be fairly constant throughout the mission. Both of these power systems provide similar 

operating capabilities and mass as the two ASRG systems used in the spacecraft design. Some concerns 

with the STEM-RTG and SRG systems as a replacement for the ASRG would be in the design of the 

spacecraft and aeroshell during transit. The added waste heat generated from these systems would need to 

be accommodated to maintain the interior temperature of the spacecraft during transit within acceptable 

levels. On the Mars surface, the extra waste heat should not be a concern. The larger volume of these 

systems may require their configuration within the lander to be changed so that it could fit within the 

aeroshell, or the aeroshell would need to be increased in size to accommodate their larger size. 

DSMCE 4: Lunar Polar Volatile Explorer (JEDI) (Ref. 8) 

The objective of the Journey to the land of Eternal Darkness and Ice (JEDI) mission concept would be 

to explore the lunar polar volatiles. This involves determining the distribution, concentration and origin of 

hydrogen within the permanently shadowed areas near the lunar poles. The JEDI mission would utilize a 

mobile platform or rover, as illustrated in Figure 13, to explore the permanently shadowed craters at the 

lunar pole. 

Due to operating within the shadowed regions, the rover was designed to be ASRG powered. It 

would, however, need to operate in both shadow and while sunlit with a driving range of 20 km. The 

process for exploring for volatiles would involve auguring to a depth of 1 m with a heated auger tip that 

can liberate the volatiles from the surface, which would then be collected and analyzed. The science 

instruments would include a neutron spectrometer, mass spectrometer; laser induced breakdown 
NASA/TM—2015-218885 14 

spectrometer and an alpha particle X-ray spectrometer. The total mission duration is estimated to be 16 

months from launch. The rover and relay satellite are launched together within an Atlas V401. The rover 

and relay satellite are released and held together for transfer to lunar orbit. Once in lunar orbit the relay is 

released and the rover descends to the lunar surface. The configuration during the mission timeline is 

illustrated in Figure 14. 

The JEDI rover utilizes a single ASRG to provide power and heat for its 9-month dark crater 

operations. Over this 9 month period 7 sites would be investigated, 4 in permanent shadow and 3 in 

partial or full sunlight. The ASRG would be mounted vertically on the rover main deck. This provides 

clear access for installation prior to launch. The carrier spacecraft through a conductive interface between 

the ASRG and the spacecraft utilize waste heat from the ASRG. The temperature within the permanently 

shadowed regions is estimated to be between 40 and 100 K. The effective sink temperature for the ASRG 

on the surface is estimated to be 209 and 238 K while in transit to the Moon. 

Figure 13.—JEDI Mission Concept Lunar Rover (Ref. 8). 

Figure 14.—Rover Configuration at Different Mission Stages (Ref. 8). 
NASA/TM—2015-218885 15 

TABLE 7.—POWER REQUIREMENTS THROUGHOUT THE JEDI MISSION (REF. 8) 

 Cruise Surface 

Mission Phase Launch Rx Tx LOI Burn Descent Idle Rove Comm Drill 

Power Req (W) 73 223 263 587 688 1160 118 274 174 280 

Operating Time (hr) 2.5 1.0 4.0 0.1 0.5 0.1 1.0 1.0 0.7 1.4 

ASRG Power (W) 71 143 143 143 143 143 145 145 145 145 

ASRG Energy (W-hr) 178 143 570 14 71 14 145 145 102 203 

Battery Energy (W-hr) 4 0 44 44 273 102 0 128 20 189 

Battery DoD 1% 0% 5% 5% 32% 12% 0% 15% 2% 23% 

The power system mass for the ASRG was estimated to be 39.1 kg including a 16 percent 

contingency. The end of life ASRG power estimate (11 months of operation on the surface) is 145.1 W. A 

28 A-hr battery is also utilized to supply supplemental power for the rover during operations. The power 

requirements throughout the mission are listed in Table 7. 

For the JEDI rover either the eight GPHS block STEM-RTG power system or the four GPHS block 

SRG power system could be a viable candidate to replace the ASRG. This mission duration is much 

shorter than the estimated 17-year operational time for the STEM-RTG or SRG. Therefore the end of 

mission output power for both systems is close to the beginning of mission output power. Both the eight 

block STEM-RTG system and four GPHS block SRG system have an estimated end of the JEDI mission 

output power of ~200 W, with masses of 29.2 and 32 kg, respectively. Both of these power systems 

provide similar operating capabilities and mass as the single ASRG systems used in the spacecraft design. 

Both systems also provide greater output power providing greater power margin for the mission as well as 

potentially adding additional capabilities to the mission. The increased output power could be used to 

eliminate one or more strings of the Trans-Lunar solar array and decrease the surface re-charge time and 

shortening the time needed to complete the baseline mission. 

Since the ASRG was mounted on the exterior of the rover any change in volume or the increase in 

waste heat associated with utilizing either the SRG or STEM-RTG would not have a significant effect on 

the rover design or layout. Also, as shown in Figure 14, there is ample space surrounding the rover within 

the faring so any change in dimensions or volume of the power system should not have a significant effect 

on the launch vehicle. 

DSMCE 5: Io Volcano Observer (IVO) (Ref. 9) 

The IVO mission concept observe and gather data on Jupiter’s moon Io through a spacecraft in orbit 

around Jupiter that provides a number of close flybys of Io, as illustrated in Figure 15. The science goals 

of the mission concept are: 

 Understand the eruption mechanisms 

 Determine the interior structure 

 Investigate the tidal heating mechanisms 

 Investigate the tectonics 

 Understand the atmosphere and ionosphere 

 Investigate the existence of a magnetic field 

 Investigate the surface chemistry 

 Improve the understanding of the Jupiter system 

The proposed mission baseline utilizes 3 Earth flybys and 1 Venus flyby to reach Jupiter, as shown in 

Figure 16. The transit time to Jupiter is approximately 6.5 years. Once at Jupiter there would be 7 flybys 

of Io over an 18-month period. The Venus flybys provide the hottest operating conditions for the 

spacecraft. 
NASA/TM—2015-218885 16 

Figure 15.—IVO Spacecraft Concept (Ref. 9). 

Figure 16.—IVO Orbital Trajectory for Transfer Between Earth and Jupiter (Ref. 9).

 
NASA/TM—2015-218885 17 

When out of range of Io the science instruments would be used to investigate Jupiter and Europa. The 

science instruments include a narrow-angle camera, thermal mapper, Ion and neutral mass spectrometer, 

fluxgate magnetometer and possibly a wide angle camera, second ion and neutral mas spectrometer, near 

IR imaging spectrometer and dust detector. Operating within the Jovian environment exposes the 

spacecraft to a significant radiation risk. A minimum of 100 kRad tolerant electronics should be utilized 

for all electrical components within the spacecraft. 

The ASRG systems would be utilized in conjunction with a rechargeable Lithium Ion battery to 

provide power to the spacecraft throughout the mission. The ASRG waste heat is also utilized to maintain 

the desired internal operating temperature of the spacecraft. The ASRG systems are located internal to the 

spacecraft structure, below the main propellant tank as illustrated in Figure 17. The mass of the power 

system was 98.4 kg, with the ASRG system mass estimated at 61.9 kg for 2 systems including a 

30 percent contingency. The external spacecraft thermal system is utilized to remove heat from the 

ASRGs. This system is sized to accommodate the heat load throughout the mission, including operation 

during the Venus flybys. 

Figure 17.—IVO Concept Spacecraft Internal Component Layout (Ref. 9). 

 
NASA/TM—2015-218885 18 

The ASRG systems provided 262 W of power throughout the mission. The Lithium Ion battery 

provided 2688 W-hr of energy to be utilized to supplement the ASRGs during the mission. The batteries 

were utilized during the following mission phases along with predicted depth-of-discharge (DoD). 

 Safe mode with the transmitter on (DoD, 91.9%) 

 Navigation prior to capture (DoD, 92.0%) 

 Maneuvering (DoD, 81.2%) 

 Science playback mode (DoD, 91.9%) 

The advantages of utilizing the ASRG system as the main power source for the spacecraft included 

the ability to provide a high data rate communications link, spacecraft pointing flexibility and long-term 

operation. Due to the low solar intensity at Jupiter, approximately 50 W/m2, a solar powered mission is 

difficult requiring a large solar array. 

For the IVO spacecraft, either the twelve GPHS block STEM-RTG power system or the six GPHS 

block SRG power system could be viable candidates to replace the ASRG. The twelve-block STEM-RTG 

system has an estimated end of mission output power of 255 W with a mass of 42.2 kg. This is lower than 

the estimated mission power of 262 W, however it is for an end of mission time of 17 years, so based on 

the IVO mission duration of eight years sufficient power should be available throughout the mission from 

this system. The six GPHS block SRG has a 266 W estimated end of mission output power and a mass of 

46.8 kg. Both of these power systems provide similar operating capabilities and mass as the two ASRG 

system used in the spacecraft design. Both systems also provide greater output power providing greater 

power margin for the mission. The main concern is that since the spacecraft design has the ASRG systems 

located internal to the spacecraft, any change in volume and waste heat generated by utilizing either the 

SRG or STEM-RTG in place of the ASRG could potentially require a redesign of the spacecraft layout 

and thermal control system. 

DSMCE 6: A DSMCE Trojan Asteroid Lander (Ilion) (Ref. 10) 

The Ilion mission concept would rendezvous and ultimately land on a Trojan asteroid and perform 

science tasks while in orbit and on the surface. The notioanal spacecraft is shown in Figure 18. The prime 

target of the mission would be asteroid Nestor. The science objective for the mission is to perform the 

following tasks: 

 Determine the mass, density, and interior structure and configuration of target. 

 Determine the surface properties and morphological features of target, including space 

weathering, maturity and lag deposits 

 Determine the mineralogical, elemental and isotopic composition and physical state of volatiles 

 Identify and characterize any out-gassing or atmosphere from target Trojan 

 Investigate dynamical evolution of target Trojan and search for satellites 

 Place Trojan asteroids in context of other bodies 

The science instruments that would be utilized to meet the mission goals include the following items. 

 Wide angle camera 

 Narrow-angle camera 

 LIDAR 

 Infrared Spectrometer 

 Gamma Ray Spectrometer 

 Alpha Proton X-Ray Spectrometer 

 Laser Desorption Mass Spectrometer 

 Neutron Spectrometer 
NASA/TM—2015-218885 19 

Figure 18.—Ilion Spacecraft/Lander Concept Illustration (Ref. 10). 

The total mission time is 11 years, which includes a 10-year cruise period. The spacecraft starts from 

Earth orbit with a direct transfer to Jupiter’s orbital distance, as illustrated in Figure 19. 

The ASRG was selected as the baseline power source for the spacecraft due to the distance from the 

sun and the need to land the spacecraft onto the surface of the asteroid. A solar array would be 

prohibitively large, on the order of 90 m2

 in surface area, especially for meeting the mission requirement 

to land on the asteroid surface. As shown in Figure 6.1, two ASRGs would be utilized. They are mounted 

on the exterior of the spacecraft 180° apart. From this orientation, they would be capable of radiating 

directly to space and would not need any additional thermal interface to reject their waste heat. 

The power system design is based on the New Horizons power subsystem with the GPHS-RTG 

replaced by the ASRG. A diagram of the power system layout is shown in Figure 20. The spacecraft 

would require 256 W of power at the end of the mission. Therefore, two ASRG systems were utilized to 

meet this requirement. The power system mass is estimated at 78 kg for both ASRG systems. 

For the Ilion spacecraft and lander, either the twelve GPHS block STEM-RTG power system or the 

six GPHS block SRG power system could be viable candidates to replace the ASRG. The twelve-block 

STEM-RTG system has an estimated end of mission output power of 255 W with a mass of 42.2 kg and 

the six GPHS block SRG has a 266 W estimated end of mission output power and a mass of 46.8 kg. This 

EOM power is sufficient to meet the mission requirements of 256 W EOL. Both of these power systems 

provide similar operating capabilities and mass as the two ASRG systems used in the spacecraft design. 

Both systems also provide greater output power at BOM, thereby providing greater power margin for 

the majority of the mission. The larger volume and waste heat load of the systems should not have a 

significant effect on the spacecraft design. Since the majority of the mission occurs in deep space, the 

added heat rejection would not be a concern. The ASRGs would be mounted on the exterior of the 

spacecraft and radiate directly to space. This arrangement can also be utilized with either the SRG or the 

STEM-RTG. Since both the STEM-RTG and SRG are single units that would be replacing two ASRGs, 

there may be a need to reconfiguration of the layout of the spacecraft to accommodate the new power 

system geometry. Adjustments to accommodate installation into the launch vehicle should be minimal. 
NASA/TM—2015-218885 20 

Figure 19.—Ilion Orbital Trajectory (Ref. 10). 

Figure 20.—Ilion Power System Layout Diagram (Ref. 10). 
NASA/TM—2015-218885 21 

DSMCE 7: Comet Coma Rendezvous Sample Return (CCRSR) (Ref. 11) 

The CCRSR mission concept would rendezvous with comet 46P/Wirtanen. The spacecraft conceptual 

design for this mission is shown in Figure 21. During the rendezvous it would collect dust grains, measure 

the gases released by the nucleus and photograph the comet from close range. Once the data collection 

phase is completed, the spacecraft would return to Earth with the collected samples. This mission would 

be somewhat of a follow-on to the Stardust mission, with the ability to collect more than 1000 times the 

amount of dust. The ASRG power system was selected for this mission because it enabled long-duration 

operation within the coma of the comet and does not have pointing requirements, thereby allowing the 

spacecraft to remain pointed at the comet nucleus throughout the mission. To meet the power 

requirements of this mission, two ASRGs were required. 

The mission was initially projected to launch in November 2015 and return the comet sample to Earth 

in December 2029, 14 year 1 month mission duration. The notional mission timeline is illustrated in 

Figure 22. 

Figure 21.—CCRSR Mission Spacecraft Conceptual Design (Ref. 11). 

Figure 22.—CCRSR Mission Concept Timeline (Ref. 11). 
NASA/TM—2015-218885 22 

Figure 23.—Orbit Throughout the Mission (Ref. 11). 

The proposed mission trajectory from Earth launch to Earth Return is shown in Figure 23. The orbital 

trajectory does takes the spacecraft out past the orbit of Mars and then through an Earth flyby then onto 

the comet which extends out to nearly the orbit of Jupiter. After co-orbiting with the comet for over three 

years, it then returns to Earth. Throughout the mission the spacecraft does not come closer to the sun than 

Earth’s orbit (1 AU). 

The ASRG system would be disposed of in a non-Earth crossing orbit prior to the sample return 

portion of the mission. The main science instruments included on the notional spacecraft include a Dust 

Sample Collector, Neutral-Ion Mass Spectrometer, Narrow and wide angle cameras and dust flux 

monitor. 

The spacecraft would carry both a solar array/battery and ASRG. The ASRG power system is a 

critical aspect of the mission. It would provide power when the spacecraft is orbiting in close proximity to 

the comet. During this portion of the mission the solar arrays are not applicable because the gas emanated 

from the comet would impinge the large surface area of the solar array, causing structural loading, drag 

and instability on the spacecraft. Also, this dusty environment could degrade the performance of the array 

and possibly damage it. 

For the three phases of the mission (checkout, cruise, sampling), the two ASRG power systems were 

estimated to provide sufficient power to operate the spacecraft loads with margin, as given in Table 8. 

The ASRG systems would be mounted external to the primary spacecraft bus, 180° apart, as illustrated in 

Figure 21. The solar array/battery controller would handle the ASRG shunt power requirements through 

the solar array/battery control card. If the power load of the spacecraft exceeds the ASRG output 

capabilities, power would be drawn from the batteries and then subsequently recharged by the ASRG. 

The ASRG would provide waste heat to the spacecraft to help maintain the internal operating 

temperature. This waste heat is transferred to the spacecraft through heat pipes. The installation of the 

ASRG system would occur between 30 days prior to launch to one day before launch (checkout to on-pad 

radioisotope installation). 
NASA/TM—2015-218885 23 

TABLE 8.—POWER LOAD & ASRG OUTPUT FOR THE THREE PHASES OF 

THE CCRSR MISSION CONCEPT (REF. 11) 

Mission Phase Power Contingency Total Power ASRG Output Power Power Margin 

Checkout w Telecom 176 W 15% 202 W 267.9 W 32.6% 

Cruise 118 W 14% 135 W 239 W 77% 

Sampling 173 W 15% 199 W 253.4 W 27.3% 

The allotted electrical power system mass was estimated to be 107.17 kg with a 14.1 percent 

contingency for a total mass of 122.25 kg. 

For the CCRSR spacecraft and mission concept, either the twelve GPHS block STEM-RTG power 

system or the six GPHS block SRG power system could be viable candidates to replace the ASRG. The 

twelve GPHS block STEM-RTG system has an estimated end of mission output power of 255 W with a 

mass of 42.2 kg and the six GPHS block SRG has a 266 W estimated end of mission output power and a 

mass of 46.8 kg. Both of these power systems provide similar operating capabilities and mass as the two 

ASRG systems used in the spacecraft design. Both systems also provide greater output power providing 

greater power margin for the mission. The larger volume and waste heat load of the systems should not 

have a significant effect on the spacecraft design. Since the majority of the mission occurs in deep space, 

the added heat rejection would not be a concern. Since both the STEM-RTG and SRG are single units that 

would be replacing two ASRGs, there may be a need to reconfiguration of the layout of the spacecraft to 

accommodate the new power system geometry. Adjustments to accommodate installation into the launch 

vehicle should be minimal. 

DSMCE 8: Titan Mare Explorer (TiME) (Ref. 12) 

The TiME mission concept would place a floating platform onto the surface of a Titan methane lake 

to provide science data on the Titan environment. The science objectives included looking at the methane 

lake chemistry, determine the lake depth, determine the physical properties of the lake, examine the local 

meteorology and it’s on the methane cycle and examine the morphology of the lake surface and shoreline. 

These objectives are accomplished through a series of science instruments, which include; mass 

spectrometer, Sonar, meteorological instruments (Temperature, Pressure, Humidity, Wind Speed), 

composition instruments (refractive index, dielectric constant) and imaging cameras (descent and 

surface). The power consumption of the science instruments was estimated to be 91 W, including 

contingency power. There is some concern that the heat rejected from the ASRG systems would affect the 

measurements of the science instruments. Additional study and modeling would be needed to confirm this 

and correct for any influence when processing the data. 

The TiME lander would be 3.0 m in diameter and is a 2X scale of the Genesis design. The TiME 

lander concept is illustrated in Figure 24. The backshell of the lander has a hatch that would allow the 

ASRG to be installed just prior to launch. The ASRG would be installed into the ASRG Environment 

Enclosure (AEE) within the lander as illustrated in Figure 25. This enclosure provides a buffer between 

the waste heat generated by the ASRG and the interior of the lander thereby maintaining the desired 

temperature distribution within the lander. 
NASA/TM—2015-218885 24 

Figure 24.—TiME Mission Lander Concept (Ref. 12). 

Figure 25.—ASRG Location Within the Lander (Ref. 12). 

 
NASA/TM—2015-218885 25 

The mission duration from launch to end of mission was approximately eight years. The notional 

mission timeline is given in Figure 26. The orbital trajectory utilizes an Earth flyby in transit to Saturn as 

shown in Figure 27. 

The ASRG power system was a critical element to the floating lander design and mission feasibility. 

Due to the low solar intensity on the surface of Titan, approximately 1.5 W/m2

, a radioisotope system 

would be required as the power system for a long duration mission in this environment. Radioisotope 

power also eliminates landing orientation requirements and changes in output power with attitude 

changes. Two ASRG systems are utilized in the design to provide power to the lander. The ASRGs would 

be located at the top of the lander, as illustrated in Figure 8, underneath a radiator panel used for rejecting 

the waste heat during both transit in space and while on the Titan surface. Heat would be moved from the 

ASRG to the radiator through constant conductance heat pipes. Other options to this arrangement are also 

under consideration, such as a combined fin, heat-pipe radiation combination or a fin only heat rejection. 

These options may provide increased thermal and power system performance margins. In addition to 

enabling easy waste heat rejection, the ASRG mounted location in the center of the lake lander was done 

to optimize its center of mass. During transit there are 2 additional radiators that are utilized to reject heat 

from the ASRGs as well as 2 shunt radiators to shunt any excess power generated and reject it as heat. 

The ASRGs operate in conjunction with a battery. If power above the output of the ASRGs would be 

required, the batteries supply the difference, such as during communications transmission. And when 

power requirements are below what can be supplied by the ASRG, the excess power would be used to 

charge the batteries. The EOM output power requirement for the ASRG systems is 242.6 W with 

contingency and margin. 

Figure 26.—TiME Mission Schedule (Ref. 12). 

Figure 27.—TiME Mission Concept Orbital Trajectory 

(Ref. 12). 

 
NASA/TM—2015-218885 26 

For the TiME lander and mission, either the twelve GPHS block STEM-RTG power system or the six 

GPHS block SRG power system could be viable candidates to replace the ASRG. The twelve-block 

STEM-RTG system has an estimated end of mission output power of 255 W with a mass of 42.2 kg and 

the six GPHS block SRG has a 266 W estimated end of mission output power and a mass of 46.8 kg. Both 

of these power systems provide similar operating capabilities and mass as the two ASRG systems used in 

the spacecraft design. Both systems also provide greater output power providing greater power margin for 

the mission. The main concern with this system would be the added waste heat generated by the increased 

number of GPHS blocks. This increase in waste heat would need to be accounted for in a revised thermal 

design. This would increase the size of the radiators needed both during curse as well as on the surface of 

Titan. It would also require additional internal insulation surrounding the AEE to maintain the correct 

operating temperature within the lander. The effects of the added waste heat as well as the larger size of 

the STEM-RTG or SRG compared to the ASRG would likely cause an increase in the size of the lander or 

a repackaging of the components within the lander to accommodate the larger size such as a transverse 

mounting scheme. Since the mass between the systems is similar, the positioning of the alternate systems 

would be similar to that for the ASRG. 

DSMCE 9: Comet Hopper (CHopper) (Ref. 13) 

The Comet Hopper (CHopper) mission concept would rendezvous with comet 46P/Wirtanen and 

enter a co-orbit with it around the sun. The comet orbit varies from approximately 5 AU to 1 AU, as 

illustrated in Figure 28. 

The mission would have two phases; a remote mapping portion and a surface science portion. Once in 

co-orbit with the comet, the spacecraft would perform its remote science potion of the mission. Once this 

portion is completed, it would then land on the surface to begin the surface science portion of the mission. 

The lander would have the capability to move or hop to multiple locations on the surface. Two hops 

would be performed to explore a total of three surface locations. 

The main science goals of the CHopper mission would be to determine the composition and 

morphologic heterogeneity of the comet, determine the processes that drive comet evolution and variation 

in comet types and examine the physical structure of the comet. The main science instruments include; 

 Gas Mass Spectrometer 

 Multi-Spectrometer Imager 

 IR Spectrometer 

 APXS, 

 Thermal Probe 

 Panoramic Cameras 

The mission duration would be 8.5 years including a 6.5-year cruise to the comet location and 2-year 

science data-gathering portion. 

The ASRG was selected for this mission concept due to the need to operate at the 5 AU distance from 

the sun. At this distance, solar power would not be a viable power source for the spacecraft. Also, the 

requirement to land on the surface and hop to other locations would put significant structural loading on a 

solar array, as well as require it to operate within a dusty environment. 

The CHopper spacecraft utilized two ASRGs to provide power throughout the mission. A set of 

redundant 30 A-Hr batteries would also be utilized to provide power to meet peak demand loads. The 

power requirements throughout the mission are listed in Table 9. 
NASA/TM—2015-218885 27 

Figure 28.—CHopper Mission Concept Orbit (Ref. 13). 

TABLE 9.—CHOPPER MISSION CONCEPT POWER REQUIREMENTS (REF. 13) 

Mission Post comet 

arrival 

Mapping-MEV 

(Science/Overall) 

Sortie 1B-MEV 

(Science/Overall) 

Total Required Power 187.5 W 207.6/226.4 W 172.8/206.4 W 

ASRG Power Output 246.7 W 244.5 W 240.9 W 

Battery DoD 22% 34% 32% 

Power Margin 46% 33% 53% 

Although the average power consumption during the phases (listed below) is below that available 

from the ASRGs, the predicted power consumption profile had power levels above the ASRG output 

level; this therefore required batteries to make up this difference. The power profiles are shown in Figures 

29 and 30 for the mapping phase and sortie, respectively. 

Waste heat from the ASRG’s control units would be utilized to heat components within the spacecraft 

such as the hydrazine tanks. The two ASRGs would be positioned on the exterior of the spacecraft, as 

shown in Figure 31, where they can radiate directly to space without the need for a thermal interface 

system to reject their waste heat. 

For the CHopper spacecraft and mission concept, either the twelve GPHS block STEM-RTG power 

system or the six GPHS block SRG power system could be viable candidates to replace the ASRG. The 

twelve-block STEM-RTG system has an estimated end of mission output power of 255 W with a mass of 

42.2 kg and the six GPHS block SRG has a 266 W estimated end of mission output power and a mass of 

46.8 kg. Both of these power systems provide similar operating capabilities and mass as the two ASRG 

systems used in the spacecraft design. Both systems also provide greater output power providing greater 

power margin for the mission or reduced battery usage. Since the ASRG systems are mounted on the 

exterior of the spacecraft the use of either the STEM-RTG or SRG should not require significant 

modifications to the thermal system design. The added size of these systems would need to be accounted 

for, but based on the layout shown in Figure 31, this should not be a significant impact. Also the mass 

between the systems is similar, therefore the positioning of the alternate systems would be similar to that 

for the ASRG. 
NASA/TM—2015-218885 28 

Figure 29.—CHopper Mapping Phase Power Profile (Ref. 13).

Figure 30.—Sortie Phase Power Profile (Ref. 13). 
NASA/TM—2015-218885 29 

Figure 31.—Position of the ASRGs on the Exterior of the CHopper Spacecraft (Ref. 13). 

Summary 

A overview of the nine missions identified as candidates for utilizing the ASRG power system was 

presented in the previous sections. By considering the operational environment for each mission concept 

as well as the power level, vehicle configuration, and operational requirements, an assessment was made 

as to the potential applicability of the STEM-RTG or SRG to provide power for these missions. Versions 

of the STEM-RTG and SRG systems have similar output power levels to either single or multiple ASRG 

power systems. All of the mission concepts considered were limited to a maximum of utilizing two 

ASRG power systems, although some missions only utilized one ASRG system. The twelve block 

STEM-RTG system and the six block SRG system provided comparable output power to a two ASRG 

system, whereas an eight block STEM-RTG system and a four block SRG system was comparable to a 

single ASRG system. 

The utilization of an STEM-RTG or SRG system in the identified mission concepts depends on how 

easily they could be substituted for the ASRG system(s). This generally means minimizing any changes to the 

vehicle configuration or design or mission goals and objectives. Some of the main items to consider include: 

 Ease of waste heat rejection: Since the STEM-RTG and SRG systems would utilize more GPHS 

blocks then the ASRG system there would be additional heat to reject and also supply additional 

heat if required by the mission. 

 Installation mass: Since a single STEM-RTG or SRG would be replacing two ASRGs, in a 

number of the mission, the placement and mass of the STEM-RTG or SRG has to be consider. 

Total mass as well as center of mass changes to the spacecraft has to be considered. 

 
NASA/TM—2015-218885 30 

 Installation Volume: Since a single STEM-RTG or SRG would be larger than a single ASRG this 

increased volume must also be considered for vehicles utilizing a single ASRG. 

 The output power of the identified STEM-RTG and SRG systems would be higher throughout 

their mission lifetime then the comparable ASRG system. Therefore the ability to meet the output 

power requirements for the missions with the substituted systems should not be a problem. 

To aid in evaluating the applicability of either the STEM-RTG or SRG systems in replacing the 

ASRG power system for the nine mission concepts examined, a 1 to 10 scale (1 corresponding to no 

impact to 10 corresponding to major impact) is utilized. The scale identifies the perceived impact on the 

spacecraft of substituting the STEM-RTG or SRG for the ASRG regarding the identified areas of thermal 

control, mass distribution and volume constraint. This rating is given in Table 10. 

Based on the ratings shown in Table 10, the two lunar surface missions, DSMCE 2: EXOMOON and 

DSMCE 4: JEDI, would be the most applicable to integrating the STEM-RTG or SRG power systems. 

These mission concepts minimize the integration concerns with utilizing these power systems in place of 

the ASRG. This is because for those missions only one ASRG system would have been utilized. Therefor 

using either the STEM-RTG or SRG would be close to a 1 to 1 replacement. For the other missions that 

utilized two ASRGs, these would be replaced by a single STEM-RTG or SRG. This can cause mass 

distribution concerns as well as volume concerns with the power system shifted from two locations to a 

single location. Another advantage to these lunar missions is that the power system would be located on 

the exterior of the vehicle. This allows it to reject heat directly from its radiators. This arrangement would 

also allow the STEM-RTG or SRG to also radiate directly to space to reject their waste heat without any 

modification to the spacecraft or power system. The final advantage is that the power system would be 

exposed to space during transit to the moon. This also allows the system to reject heat without any 

modification to the spacecraft. 

The least applicable mission concepts were ones that utilized multiple ASRG systems and had them 

located within the spacecraft or surface vehicle. This configuration could likely require significant 

modification to the design to accommodate the increased waste heat load of the STEM-RTG or SRG. 

TABLE 10.—STEM-RTG & SRG IMPACT OF REPLACING ASRG POWER SYSTEM 

Mission System Thermal 

impact 

Mass and 

distribution 

impact 

Volume 

impact 

Total 

impact 

DSMCE 1: Venus Atmospheric LongDuration Observatory for in situ 

Research (VALOR) 

STEM-RTG 

SRG 

7 

5 

1 

1 

2 

2 

10 

8 

DSMCE 2: Lunar Lander 

(EXOMOON) 

STEM-RTG 

SRG 

2 

2 

1 

1 

2 

2 

5 

5 

DSMCE 3: Mars Polar Ice Borehole 

(Kuklos) 

STEM-RTG 

SRG 

2 

2 

7 

8 

5 

4 

14 

14 

DSMCE 4: Lunar Polar Volatile 

Explorer (JEDI) 

STEM-RTG 

SRG 

2 

2 

1 

1 

2 

2 

5 

5 

DSMCE 5: Io Volcano Observer 

(IVO) 

STEM-RTG 

SRG 

9 

7 

2 

2 

3 

5 

14 

14 

DSMCE 6: Trojan Asteroid Lander 

(Ilion) 

STEM-RTG 

SRG 

2 

2 

7 

8 

5 

4 

14 

14 

DSMCE 7: Comet Coma Rendezvous 

Sample Return (CCRSR) 

STEM-RTG 

SRG 

2 

2 

4 

5 

5 

4 

11 

11 

DSMCE 8: Titan Mare Explorer 

(TiME) 

STEM-RTG 

SRG 

8 

6 

2 

2 

4 

3 

14 

11 

DSMCE 9: Comet Hopper (CHopper) STEM-RTG 

SRG 

2 

2 

4 

5 

3 

3 

9 

10 
NASA/TM—2015-218885 31 

Conclusions 

One observation of this assessment of the applicability of the new RPS designs to the nine DSMCE 

missions is that, at a minimum, these missions could utilize a new RPS with a designed EOM power 

output in the 300 We range. And to a varying degree, some sort of spacecraft accommodation could be 

required to implement any of the new concepts. Furthermore, the more tightly coupled or integrated the 

power system would be with other spacecraft systems, e.g., thermal, the more difficult this would 

become. Without a redesign starting with a “clean sheet of paper,” it is difficult to determine to what 

degree a new RPS would benefit the design. It is not unreasonable to expect that a spacecraft design could 

make use of a single higher power unit as its power system option. Unfortunately, a redesign effort with 

each of the nine DSMCE teams is not practical or prudent at this time. 

NPAS assessed Discovery, New Frontiers and Flagship missions for their power needs and 

determined that if a new RPS development were to occur, a system producing about 300 We after 

~15 years (three years storage, 12-year mission) to support an outer planet mission’s typical trip time, 

could support all three SMD-type missions with a single or multiple number of units. 

Recommendations 

The power range common for New Frontiers missions is on the order of 500 to 600 We EOM. The 

use of small, 100 to 130 We units to supply this power level creates potential issues with integration with 

other spacecraft systems and launch vehicle integration logistics at the launch site. A follow-on study is 

recommended for a similar effort done here for Discovery class missions to be performed for the 

appropriate missions enabled by RPS, that have been identified in the two Decadal surveys and, in 

particular, the mission concept set announced for the NF-4 AO. 

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