David R. Pantano, Frank Dottore, and E. Wayne Tobery

Lockheed Martin Space Systems Company, King of Prussia, Pennsylvania

Steven M. Geng and Jeffrey G. Schreiber

Glenn Research Center, Cleveland, Ohio

Joseph L. Palko

Connecticut Reserve Technologies, Inc., Cleveland, Ohio

Utilizing Radioisotope Power System Waste

Heat for Spacecraft Thermal Management

NASA/TM—2005-213990

October 2005

AIAA–2005–5548
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David R. Pantano, Frank Dottore, and E. Wayne Tobery

Lockheed Martin Space Systems Company, King of Prussia, Pennsylvania

Steven M. Geng and Jeffrey G. Schreiber

Glenn Research Center, Cleveland, Ohio

Joseph L. Palko

Connecticut Reserve Technologies, Inc., Cleveland, Ohio

Utilizing Radioisotope Power System Waste

Heat for Spacecraft Thermal Management

NASA/TM—2005-213990

October 2005

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

Third International Energy Conversion Engineering Conference

sponsored by the American Institute of Aeronautics and Astronautics

San Francisco, California, August 15–18, 2005

AIAA–2005–5548
Acknowledgments

This work was enabled through the Department of Energy (DOE) Stirling Radioisotope Generator (SRG110)

program. The authors wish to thank James Braun, Jaime Reyes, Robert Cockfield, Dan Tantino, Jack Chan, and

Joe Brouch, Lockheed Martin Space Systems Company, King of Prussia, for their support in preparing this paper.

Available from

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endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

Available electronically at http://gltrs.grc.nasa.gov
NASA/TM—2005-213990 1 

Utilizing Radioisotope Power System Waste Heat for 

Spacecraft Thermal Management 

David R. Pantano, Frank Dottore, and E. Wayne Tobery 

Lockheed Martin Space Systems Company 

King of Prussia, Pennsylvania 19406 

Steven M. Geng and Jeffrey G. Schreiber 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Joseph L. Palko 

Connecticut Reserve Technologies, Inc. 

Cleveland, Ohio 44136 

Abstract 

An advantage of using a Radioisotope Power System (RPS) for deep space or planetary surface 

missions is the readily available waste heat, which can be used for a number of beneficial purposes 

including: maintaining electronic components within a controlled temperature range, warming propulsion 

tanks and mobility actuators, and maintaining liquid propellants above their freezing temperature. Previous 

missions using Radioisotope Thermoelectric Generators (RTGs) dissipated large quantities of waste heat 

due to the low efficiency of the thermoelectric conversion technology. The next generation RPSs, such as 

the 110-Watt Stirling Radioisotope Generator (SRG110) will have higher conversion efficiencies, thereby 

rejecting less waste heat at a lower temperature and may require alternate approaches to transferring waste 

heat to the spacecraft. RTGs, with efficiencies of ~6 to 7 percent, reject their waste heat at the relatively 

high heat rejection temperature of 200 °C. This is an advantage when rejecting heat to space; however, 

transferring heat to the internal spacecraft components requires a large and heavy radiator heat exchanger. 

At the same time, sensitive spacecraft instruments must be shielded from the thermal radiation of the RTG. 

The SRG110, with an efficiency around 22 percent and 50 °C nominal housing surface temperature, can 

readily transfer the available waste heat directly via heat pipes, thermal straps, or fluid loops. The lower 

temperatures associated with the SRG110 avoid the chances of overheating other scientific components, 

eliminating the need for thermal shields. This provides the spacecraft designers more flexibility when 

locating the generator for a specific mission. A common misconception with high-efficiency systems is that 

there is not enough waste heat for spacecraft thermal management. This paper will dispel this 

misconception and investigate the use of a high-efficiency SRG110 for spacecraft thermal management and 

outline potential methods of waste heat utilization in several conceptual missions (Lunar Rover, Mars 

Rover, and Titan Lander). The advantages associated with the SRG110 as they relate to ease of assembly, 

less complex interfaces, and overall mass savings for a spacecraft will be highlighted. 

Nomenclature 

°C Degrees Celsius 

atm atmosphere 

BOM Beginning of Mission 

cm centimeters 

g gram 

kg kilogram 

kPa kilo Pascals (1000 N/m2

) 

L/min liters per minute 

m/s meters per second 

psia absolute pressure in pounds per square inch 

Pu238 Plutonium 238 

Wdc Watts electrical Direct Current 

W/kg Watts electrical per kilogram 
NASA/TM—2005-213990 2 

W/m2

 Watts per square meter 

Wt Watts thermal 

I. Introduction 

A. Radioisotope Power Systems 

A Radioisotope Power System (RPS) generates electricity from isotopic decay heat of a Plutonium-238 

(Pu238) heat source. This heat can be converted into electricity through a number of conversion technologies 

with the remainder of the thermal energy generally rejected to the environment as wasted energy. RPS 

systems are used on missions where solar panels would be ineffective e.g., locations with low solar flux at 

vast distances from the sun or in areas where there is little or no solar insolation for long durations, such as 

the lunar night or permanently shadowed craters. They are also capable of providing continuous power for 

very long durations, up to 25+ years, if necessary. An RPS is a perfect option to provide power for many 

types of missions including deep space missions, planetary or satellite surface missions, and missions to 

explore other planets where the spacecraft can take 7 to 10+ years to reach its destination. 

In the past, thermoelectrics have been the only conversion method used on RPSs for deep space and 

planetary surface missions, such as Pioneer, Voyager, Viking, and Cassini. (ref. 1). The technology of the 

past 40 years was not capable of efficiencies greater than 7 percent, requiring higher amounts of radioactive 

fuels to generate relatively small amounts of electrical power (ref. 2). Advancements in mechanical 

conversion technologies with power systems like the Stirling Radioisotope Generator (SRG110) have led to 

improved efficiencies, up to 22 percent, thus reducing the fuel required to generate the same amount of 

electrical power (refs. 3 and 4). 

B. Advantages of High-Efficiency Systems 

There are four advantages of high-efficiency systems over low-efficiency systems: lower surface 

temperature, greater flexibility, lower mission cost, and decreased mass. In general, the low-efficiency 

systems must operate at high surface temperatures to reject the waste heat without prohibitively large 

radiator fins. Typical Radioisotope Thermoelectric Generator (RTG) housing surface temperatures, such as 

those on the General Purpose Heat Source Radioisotope Thermoelectric Generators (GPHS-RTGs) for the 

Ulysses, Galileo, and Cassini missions, approach 240 °C while rejecting roughly 4000 Wt of waste heat 

(ref. 2). The next generation RTG, the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) 

uses a different thermoelectric material and rejects roughly 1900 Wt of waste heat at a fin root temperature 

of approximately 200 °C (ref. 3). While high surface temperatures are an advantage when trying to reject 

large amounts of heat to the environment, the high surface temperatures present a challenge when the waste 

heat must be used in a controlled fashion as part of the overall spacecraft thermal management system. A 

preference for the low-efficiency MMRTG may be partially based on the misconception that there is 

insufficient waste heat available from high-efficiency systems. This paper addresses this misconception. 

1. Spacecraft Thermal Heating Loads 

Typical deep space or planetary spacecraft heat is provided from electrical heaters and/or Radioisotope 

Heater Units (RHUs). While Cassini used three RTGs and did receive ~140 Wt thermal radiation from the 

RTGs to warm components, it also included electrical heaters and 157 RHUs (35 on Huygens), which 

provide 1 Wt each (refs. 5 and 6). The Galileo mission used 2 RTGs, but did not utilize any of the RTG 

waste heat. Galileo relied on electrical heaters and 120 RHUs to provide thermal power to warm the 

spacecraft components (ref. 7). The Mars Science Laboratory (MSL), scheduled for launch in 2009, could 

require up to a maximum of 500 Wt makeup heat from an RPS to warm the internal components. From 

these examples, it is reasonable to estimate that most deep space or planetary spacecraft would use at most 

500 Wt of heat for thermal management. 

2. Thermal Advantages 

There are a number of practical ways to harness the waste heat from an RTG, including using radiation 

exchange from the hot housing surface, high-temperature heat pipes, or a high-temperature forced fluid 

loop. The high surface temperatures of the RTGs would result in high fluid temperatures if heat pipes or a 

forced fluid loop were implemented. This would lead to significant thermal design challenges to prevent 

overheating of the spacecraft components. A heat exchanger simplifies the thermal design, but effectively 

transferring the waste heat into the spacecraft thermal management system would require a large and heavy 
NASA/TM—2005-213990 3 

component. Conversely, the high-efficiency systems such as the SRG110 have housing surface 

temperatures of ~50 °C and are more conducive to direct attachment to the spacecraft thermal control 

system (ref. 4). 

Additionally, the high surface temperatures limit acceptable RTG locations and therefore the flexibility 

of the spacecraft design. At a 200 °C surface temperature, the radiative flux coming off the RTG surface is 

roughly 2000 W/m2

. The spacecraft may require thermal shields to protect sensitive instrumentation from 

these thermal radiation loads. A high-efficiency system with lower surface temperatures may allow the 

spacecraft designer to locate the power source almost anywhere. In essence, a high-efficiency system 

facilitates designing the spacecraft without much consideration of the power source, whereas a spacecraft 

with an RTG must basically be designed around the low-efficiency source. 

3. Reduction in Fuel 

High-efficiency power sources have an obvious advantage in requiring less fuel to generate the same 

amount of electrical power as compared to their low-efficiency counterparts. The SRG110, for example, 

offers a four-fold reduction in fuel over the MMRTG. The GPHS heat sources are heavy, 1.6 kg apiece. 

The total weight of the heat sources in the MMRTG is almost 13 kg, while only 3.2 kg for the SRG110. A 

decrease in mass is transferred directly to the mission in the way of lower mission costs. Furthermore, 

reducing the mass of the system may enable use of a lighter launch vehicle or increase the payload 

available for scientific purposes. 

4. Reduction in Required Hardware 

The benefits of a high-efficiency power source extend beyond the savings in fuel. Reductions in 

hardware as it relates to radiator heat exchangers and thermal shields for sensitive equipment are also 

possible, thereby reducing cost, design time, testing, and mass. This reduction in hardware can have an 

effect on the design of the overall system. 

In general, when using a low-efficiency power system, the expected environment would define the 

total area, embedded heat exchanger method, and structural support for the heat exchanger. For example, 

the heat exchanger would have to be larger for a mission in an extremely cold location such as Titan versus 

a warmer location like Mars. Due to the easier direct attachment of the thermal control system, such a large 

and heavy piece of hardware would be completely unnecessary in either location if a high-efficiency 

system was used instead. 

Typically, most thermal shields are carried on solar-powered spacecraft to protect sensitive components 

from the heat of the sun. While these solar shields would still be necessary in some high-efficiency RPSpowered spacecraft applications, thermal shields would not be necessary to protect onboard components 

from the heat from the spacecraft’s own power source. Cassini, for example, carried three thermal shields, 

one for each RTG, to protect the scientific payload. The lower surface temperatures of the high-efficiency 

power source would relax the requirements for these thermal shields, thus providing an opportunity for 

further program savings. 

C. Spacecraft Thermal Management 

The thermal management of present day spacecraft presents competing challenges between heating and 

cooling methods in an effort to maintain an acceptable thermal environment for sensitive avionics and 

instrumentation. Fortunately, to address these challenges, thermal engineers have a variety of thermal 

management tools at their disposal, including electric heating elements, conducting straps, louvers, heat 

pipes, fluid loops, etc. However, these challenges may differ depending on the type of mission. For 

example, thermal management of a spacecraft in Earth orbit may be handled differently than a spacecraft 

on a deep space mission. For the Earth orbiting spacecraft, options for available thermal energy may extend 

beyond the onboard source (e.g., solar energy), whereas for a deep space mission, these options become 

more limited. Because solar panel size must increase exponentially as a function of distance from the sun to 

generate the same power, when designing for a deep space or planetary surface mission, an alternate power 

source such as an RPS is generally more effective at generating electrical power than solar energy. 

The waste heat from an RPS can be recovered by the spacecraft to provide additional thermal control to 

the avionics and instrumentation without resorting to additional electrical heaters. Using thermal power to 

generate electricity, only to then reconvert the electricity back into thermal energy may be a very inefficient 

use of the limited available spacecraft power. 
NASA/TM—2005-213990 4 

1. Heat Pipes 

Using waste heat from the RPS to directly warm electronics or other spacecraft hardware requires a heat 

transport device. One such device is a heat pipe. A heat pipe is a self-contained device used to transport 

heat efficiently across moderate distances. Heat transport is accomplished by the evaporation and 

condensation of an internal working fluid, which is then recirculated via the capillary action of an internal 

wick. Heat pipes have been used in a multitude of space applications and are efficient, lightweight, cost 

effective, and reliable devices. Four different types of heat pipes can be used to transfer and control the 

amount of heat that is transferred: constant conductance heat pipes (CCHP), variable conductance heat 

pipes (VCHP), loop heat pipes (LHP), and diode heat pipes. 

A CCHP consists of a pipe that is lined with a wick and capped off at both ends. The pipe is completely 

evacuated prior to being charged with a small amount of working fluid (water and ammonia are among the 

most efficient at 50 °C). When heat is introduced to any spot along the length of the CCHP the working 

fluid in that location evaporates and then condenses at a colder location along the pipe. Capillary action 

works to restore equilibrium (i.e., to uniformly saturate the wick) and transports the fluid to any location 

that has been depleted of its working fluid. CCHPs are often embedded in honeycomb panels as an 

effective method of isothermalizing the panel. They can be bent into a two or three-dimensional shapes and 

work best in a zero-g environment or where the effects of gravity are minimized. 

If heat were added at one end of a CCHP and removed from the other end, the vapor phase would 

sweep the entire length of the pipe before condensing and being returned as liquid by the wick. This 

phenomenon is used to the designers’ advantage in a VCHP. A VCHP is a CCHP with a small amount of 

non-condensable gas introduced along with the working fluid. As the vapor sweeps the entire length of the 

pipe, it sweeps the non-condensable gas along with it. This gas accumulates as a plug effectively blocking a 

portion of the condenser. The size of the plug is a function of many factors; however, with the addition of a 

gas reservoir located at the condenser end, a VCHP can be configured to maintain a specific temperature 

range in the component in need of temperature stability. Therefore, a VCHP is used when temperature 

regulation of the heat source is needed. 

CCHPs and VCHPs work well in the zero gravity environment of space. They are, however, sensitive to 

orientation when used in a gravitational field, working fine when completely horizontal and having 

reduced, or no, capacity in a vertical orientation. 

An LHP can best be described as a heat pipe with the functions separated and grouped together. The 

wick of a CCHP lines the entire length of the heat pipe, whereas an LHP wick is contained in an evaporator 

1 to 12 inches (2.5 to 30.5 cm) in length, which is then attached to the heat source. The wick is the dividing 

line between the liquid and vapor phases. One side of the wick is in contact with liquid and the other with 

vapor. Source heat is transferred into the wick and causes evaporation of the working fluid. The vapor 

flows into the wickless condenser tube, which can be as small as 0.125 inches (0.318 cm) in diameter and 

100 or more feet (30.5+m) in length. Heat removal can occur uniformly along the entire length of the 

condenser tube or at a specific location. Either way the vapor is condensed and pushed along by the vapor 

behind it until it reaches the wick at which point, the cycle is repeated. 

A diode heat pipe, as the name suggests, will transfer heat in only one direction. This can be 

accomplished in one of two ways. By utilizing a gas reservoir, similar to a VCHP but with a smaller 

amount of non-condensable gas, the heat pipe will function like a normal CCHP when operating in the 

forward direction. In this mode, the non-condensable gas resides completely within the reservoir. When 

operated in reverse, the non-condensable gas is swept out of the reservoir and completely fills the working 

length of the heat pipe blocking operation. A second method is to have a working-fluid trap (i.e., an 

unwicked reservoir) at the evaporator end. During forward operation no fluid will condense in the reservoir 

since it is at the same temperature as the evaporator. During reverse operation, the fluid will eventually end 

up in the reservoir, since there is no wick to bring the fluid back into circulation. When the temperature 

gradient is reestablished in the forward direction the working fluid evaporates out of the reservoir and the 

heat pipe is once again fully functional. 

Attachment of the heat pipe to the hot and cold surfaces can be accomplished by using a thermally 

conductive adhesive, press fit, or by clamping. A thermal interface material may be used between the two 

mating surfaces to enhance conductivity. In a typical RTG application, the heat pipes would be imbedded 

into the radiator heat exchanger and then routed into the spacecraft. For the higher efficiency systems, the 

heat pipes could be directly mounted onto the housing surface. 
NASA/TM—2005-213990 5 

2. Louvers and Insulation 

Typically, a spacecraft would be insulated around the primary structure and propulsion system, with 

penetrations for radiators for heat dissipating electronic components or for instruments. Spacecraft 

insulation typically consists of multi-layer foils, with the number of layers sized locally to achieve the 

required effective emissivity. The radiators are sized for maximum heat dissipation and worse case thermal 

environments. In situations where components are turned off or put in standby mode for power 

management reasons, louvers may be employed to reduce the effective radiator area by shielding all or part 

of the radiator surface. The louvers can be actuated passively by temperature changes or actively via 

electronic controls that sense temperature changes. Typical passive louvers consist of a bi-metal that 

expands or contracts as temperatures increase or decrease to move some form of a mechanical shield over 

the areas to be protected. 

Similarly, heat sources inside of a spacecraft can vary the amount of heat rejected into the spacecraft for 

thermal management by using a system of insulation and louvers. When the spacecraft requires more heat 

for temperature control, the louvers can close retaining more heat inside the spacecraft. Conversely, the 

louvers can open and reject more heat to the environment when not required. 

Louvers and insulation would not be a viable thermal management technique for a low-efficiency RPS 

due to the amount of waste heat. In most cases, the amount of waste heat generated by a low-efficiency 

RPS necessitates a cantilever mount from the spacecraft to provide the maximum area possible for heat 

rejection. However, the high-efficiency systems like the SRG110 could be mounted inside the spacecraft to 

provide the highest possible waste heat utilization. In this case, variable area radiators and insulation could 

be used together to effectively maintain components within their allowable flight temperature limits. This 

flexibility represents one of the distinct advantages with high-efficiency RPS systems. 

3. Forced Fluid Loop 

Another effective means for transferring thermal energy is with a pumped fluid loop. A pump provides 

the driving force necessary to create fluid flow over a hot surface transferring heat to the fluid and 

subsequently over another, colder surface that accepts heat from the fluid. A variable speed pump can be 

included to throttle flow based on temperature requirements of the system. With this type of system, flow is 

reduced to limit additional heat flow to the hardware in a hot environment, while in a cold environment the 

flow can be increased to transfer more heat. 

A forced fluid loop usually includes a redundant pump system with thin-walled aluminum tubing and, 

depending on predicted temperatures, a working fluid such as water or a refrigerant. The pump is mounted 

to the spacecraft and the fluid lines are soldered or brazed onto the hot and cold surfaces. To date, pumped 

fluid loops have been flown on several missions including the Mars Pathfinder and Mars Exploration 

Rovers (MER) (refs. 8 to 10). In these cases the working fluid was R–11 pressurized to 100 psia (690 kPa) 

to preclude formation of a vapor phase. During the cruise phase while in transit to Mars, the fluid carried 

the waste heat generated by the internal electronics of the lander (which was enclosed in an aeroshell) to an 

external radiator system to prevent overheating. The currently planned Mars Science Laboratory (MSL) 

mission will use a similar pumped loop. 

A forced fluid loop would be implemented similar to heat pipes. A low-efficiency system would have 

the tubing installed inside of a radiation heat exchanger, while a high-efficiency system would have the 

tubing mounted directly to the RPS fins or housing. 

II. Mission Requirements 

Thermal requirements of the various missions can vary depending on the mission profile and objectives. 

A system designed specifically to maintain acceptable temperatures in the darkness of the interstellar void 

would overheat if exposed to the 127 °C daytime temperatures of the lunar surface. The environmental 

boundary conditions throughout the mission, desired spacecraft size and type, and the special needs of 

scientific instruments, electronics, augmentation batteries, and propulsion elements all contribute to the 

mission requirements of the thermal management system. Through use of any one or a combination of the 

available thermal management techniques described previously, the RPS can be used to maintain the 

spacecraft components within allowable temperature limits. 
NASA/TM—2005-213990 6 

Passive Fins (8)

Fluid Manifold

Primary Loop

Backup Loop

Fluid Loop Fins (8)

Figure 1.—SRG110 Shown with Reconfigurable Fins. 

III. Interfaces 

The specific spacecraft design and operational environments will dictate the necessary thermal 

interfaces from the RPS to the spacecraft. A multi-mission power source must be readily adaptable to any 

spacecraft configuration and not force limitations on the spacecraft design. High-efficiency systems like the 

SRG110 can be designed with simple and robust interfaces for ease of assembly and integration with the 

spacecraft. The reconfigurable nature of the detachable SRG110 fins, as shown in figure 1, allow for 

flexibility in meeting mission requirements. Different fins could be attached for different missions, e.g., 

radiation fins, fins with forced fluid lines, heat pipes, conduction straps, or fins with a mechanical mounting 

provision to provide direct conduction to spacecraft components. 

Similar techniques could be used for the MMRTG, but implementation would be more challenging. 

Because of the high temperatures associated with this low-efficiency power source, the MMRTG must use 

a cantilever mount to locate the power source away from the spacecraft. Due to the quantity of waste heat 

generated, it would not be possible to completely enclose the MMRTG inside the spacecraft, thereby 

limiting its positioning flexibility. The high surface temperature would also require that the interfaces, 

either heat pipes, fluid loops, conduction straps, etc., be capable of withstanding the high temperatures. 

Most electronic components cannot survive contact with a 200 °C heat transport device, so considerations 

would have to be made to material selection, fluid properties, and contact interfaces to prevent the 

electronics from overheating. 

IV. Analysis 

To demonstrate the advantages and flexibility of a reconfigurable, high-efficiency power source, several 

possible missions were considered where the RPS was used for electrical power and thermal management 

for spacecraft components. For each example, the SRG110 served as the power source. The assumptions 

related to the operating conditions and design requirements of the SRG110 appear in table 1. The 

conditions and design requirements of the MMRTG are also listed in table 1 for comparison (refs. 3, 4, and 

12). The information listed in table 1 summarizes the operational performance of each system from 

presently available public references and does not reflect performance enhancements that may be possible 

through currently on-going design evolutions. 

Three separate and varied missions were selected for the analyses. The missions were selected based on 

the realistic nature of scientific interest, the amount of environmental information available for each 

mission, and the fact that together, the environments for the various missions represented a wide spectrum 

of potential operational conditions for the RPS. The missions included: Lunar Rover, Mars Rover, and 

Titan Lander. For each of these cases, the spacecraft power requirements at the beginning of mission 

(BOM) remained the same: 200 Wdc of electrical power and a variable thermal management heat load of 

approximately 500 Wt. Achieving this amount of electrical power would require two of the current state-orthe-art RPSs, either MMRTG or SRG110. The environmental factors, and therefore the thermal 

management requirements, were mission dependent. It should be noted that these three mission profiles 

represent only a small subset of the capabilities of an SRG110-powered spacecraft. 
NASA/TM—2005-213990 7 

Table 1. SRG110 and MMRTG Design Requirements and 

Operational Condition Assumptions

Mass (kg) 34 44

3.4

104 x 38 x 29

Multi-Mission

50

500

116

Stirling Engine

22%

SRG110

Envelope (cm) 65 x 63

Specific Power (W/kg) 2.8

Possible Mission Environments Multi-Mission

125 Nominal Output Power – BOM Deep 

Space Mission (Wdc)

Initial Fuel Loading (W 2000 t

)

Typical Surface Temperature (°C) 200

PbTe/TAGS 

thermoelectric

Conversion Technology

Typical Efficiency 6%

MMRTG

Mass (kg) 34 44

3.4

104 x 38 x 29

Multi-Mission

50

500

116

Stirling Engine

22%

SRG110

Envelope (cm) 65 x 63

Specific Power (W/kg) 2.8

Possible Mission Environments Multi-Mission

125 Nominal Output Power – BOM Deep 

Space Mission (Wdc)

Initial Fuel Loading (W 2000 t

)

Typical Surface Temperature (°C) 200

PbTe/TAGS 

thermoelectric

Conversion Technology

Typical Efficiency 6%

MMRTG

-269

0.9

0.9

127

1400

Lunar Surface 

– Direct Sun

Hot Case

Sky Sink Temperature (°C) -269

Surface Temperature (°C) -190

Surface Emissivity 0.9

Lunar Dust Absorptance 0.9

Incident Solar Flux (W/m 0 2)

Permanently 

Dark Crater

Location

Condition Cold Case

-269

0.9

0.9

127

1400

Lunar Surface 

– Direct Sun

Hot Case

Sky Sink Temperature (°C) -269

Surface Temperature (°C) -190

Surface Emissivity 0.9

Lunar Dust Absorptance 0.9

Incident Solar Flux (W/m 0 2)

Permanently 

Dark Crater

Location

Condition Cold Case

Table 2. Lunar Surface Assumptions

D. Lunar Rover 

The Lunar Rover concept considered here was based on two studies by NASA Jet Propulsion 

Laboratory (JPL) (refs. 11 and 12). These references provided detailed information regarding proposed 

architecture to enable a successful mission on the lunar surface, with either manned or unmanned vehicles. 

For the purposes of this investigation, two cases were studied; a hot case, which included direct solar 

insolation, and a cold case, which represented operation during lunar night or in a permanently shadowed 

crater. The assumed lunar conditions for these two cases are summarized in table 2. 

1. Lunar Rover Configuration 

Two SRG110s would be mounted internal to the lunar rover with VCHPs, diode heat pipes, or a 

pumped fluid loop running from the base of the SRG110 fins to the internal components of the rover. A 

variable heat transport mechanism would be required to support rover operations in both hot and cold 

conditions. For this investigation it was assumed that VCHPs would be used to transport the heat from the 

SRG110s to the rover radiators. 
NASA/TM—2005-213990 8 

2. Lunar Rover Thermal Model 

A thermal model was constructed using Thermal Desktop® (ref. 13) to predict the resulting 

temperatures and heat transferred to the rover under both environmental conditions. It was assumed that the 

rover required at least 500 Wt during the cold case and as little as possible during the hot cases. The surface 

of the Lunar Rover was assumed to be coated with a fine layer of lunar dust. Some form of dust mitigation 

was assumed, making the optical properties of the exposed surfaces have an emissivity of 0.9 and an 

absorptivity of 0.5 (ε = 0.9 and α = 0.5). 

3. Thermal Analysis Results for Lunar Rover Mission 

The thermal analysis results showed that the temperature difference between the two cases resulted in a 

challenging thermal management scenario. In the cold case, the analysis showed that the SRG110s could 

operate within expected design limits and still provide the required thermal and electrical power to the 

rover through all mission phases. The lunar hot case demonstrated the challenges inherent to lunar 

operations. In order to develop a common system for both hot and cold operations, compromises were 

made to optimize the performance for the cold case at the expense of the performance for the hot case. The 

temperature predictions for the SRG110 and the representation of the rover are shown in figure 2 and 

summarized in table 3. 

94

91.3

88.7

86.1

83.5

80.9

78.3

75.7

73.0

70.4

67.8

Temperature (°C)

20.9

17.7

14.4

11.2

7.9

4.6

1.4

-1.8

-5.1

-8.3

-11.6

Temperature (°C)

Lunar 

Hot 

Lunar 

Cold

94

91.3

88.7

86.1

83.5

80.9

78.3

75.7

73.0

70.4

67.8

Temperature (°C)

20.9

17.7

14.4

11.2

7.9

4.6

1.4

-1.8

-5.1

-8.3

-11.6

Temperature (°C)

94

91.3

88.7

86.1

83.5

80.9

78.3

75.7

73.0

70.4

67.8

Temperature (°C)

94

91.3

88.7

86.1

83.5

80.9

78.3

75.7

73.0

70.4

67.8

Temperature (°C)

20.9

17.7

14.4

11.2

7.9

4.6

1.4

-1.8

-5.1

-8.3

-11.6

Temperature (°C)

20.9

17.7

14.4

11.2

7.9

4.6

1.4

-1.8

-5.1

-8.3

-11.6

Temperature (°C)

Lunar 

Hot 

Lunar 

Cold

Figure 2.—Temperature Profiles for the SRG110 Lunar Rover Configuration in 

Hot and Cold Environments. 

Average SRG110 Surface 77 5

Temperature (°C)

Heat Pipe Average Temperature (°C) 68 -12

Waste Heat to Rover – 2 SRG110s 76* 610**

BOM (Wt

)

Total Electrical Power – 2 SRG110s 201 234

BOM (Wdc)

Condition Hot Case Cold Case

Average SRG110 Surface 77 5

Temperature (°C)

Heat Pipe Average Temperature (°C) 68 -12

Waste Heat to Rover – 2 SRG110s 76* 610**

BOM (Wt

)

Total Electrical Power – 2 SRG110s 201 234

BOM (Wdc)

Condition Hot Case Cold Case

Table 3. Thermal Analysis Results for the SRG110 Lunar 

Rover Configuration

Notes: * Includes radiation only

** Includes heat transported by VCHPs only
NASA/TM—2005-213990 9 

E. Mars Rover 

In this simulation, SRG110s were applied to a Mars Rover concept. A typical Mars mission can involve 

wide temperature ranges depending on many factors, including local environmental conditions and landing 

locations of the spacecraft. By evaluating the SRG110 Mars Rover concept under both hot and cold 

conditions, missions anywhere within ±60° latitude on the Mars surface were covered. Table 4 summarizes 

the assumed conditions for each case. 

Table 4.—Mars Surface Conditions Assumed 

Condition Hot Case Cold Case 

Location 15° South Latitude 60° North Latitude 

Season Martian Summer Martian Winter 

Direct Solar Flux (W/m2) 580 0 

Diffuse Solar Flux (W/m2) 104 0 

Surface Temperature (°C) 31 –123 

Surface Emissivity 0.8 0.8 

Atmosphere Temperature (°C) 5 –123 

Sky Sink Temperature (°C) –101 –150 

Atmosphere Pressure (kPa) 0.67 1.34 

Surface Wind Speed (m/s) 0 30 

1. Mars Rover Configuration 

For the Mars rover mission, it was assumed that two SRG110s were mounted internally on the rover 

vehicle and coupled with a pumped fluid loop running from the base of the SRG110 fins to the internal 

components of the rover. It was also assumed that water was the working fluid in the pumped fluid loop at 

a pressure of 100 psia (690 kPa) with a maximum flow rate of 0.65 L/min. By mounting the SRG110s 

internally, the waste heat available for thermal management of the onboard electronics would be 

maximized, and the SRG110s would be shielded from the Martian wind when exposed to the cold 

environment. The SRG110s were protected on five sides by the insulated exterior shell of the rover and the 

sixth side was open to the rover electronics cavity. Not only would this allow easy installation, but also the 

open side provided a line of sight for thermal radiation to the internal components of the rover. The use of 

water as the working fluid provided the best heat transfer option when considering heat transport from the 

SRG110s to the external rover radiators while the rover was exposed to the hot environment. The Mars 

Rover configuration, based on the Mars Science Laboratory (MSL) design, with the SRG110s is shown in 

figure 3. 

Ground Boundary Temperature

Sky Boundary

Direct Solar 

and Diffuse IR

Rover

SRG110s Sun Track 

Enclosure 

Air Temperature (Wind)

Ground Boundary Temperature

Sky Boundary

Direct Solar 

and Diffuse IR

Rover

SRG110s Sun Track 

Enclosure 

Air Temperature (Wind)

Figure 3.—Mars Rover Configuration with SRG110s. 
NASA/TM—2005-213990 10 

2. Mars Rover Thermal Model 

The thermal model was again constructed using Thermal Desktop® to predict the resulting 

temperatures and heat transferred to the rover. The different environmental conditions presented different 

modeling challenges. When exposed to the cold conditions, it was assumed that the rover required 

approximately 500 Wt for onboard thermal management. When exposed to the hot environment, it was 

assumed that all of the waste heat in the pumped fluid loop bypassed the rover internal components and 

was rejected to the Martian atmosphere. The surface of the rover was assumed to be coated with a fine 

layer of Martian dust making the optical properties equal to the Martian surface, ε = 0.8 and α = 0.8. 

3. Thermal Analysis Results for Mars Rover Mission 

The thermal analysis results showed that the SRG110s operated within allowable temperature limits and 

still provided the required thermal and electrical power to the rover through all mission phases. 

Furthermore, it was demonstrated that the rover fluid cooling system varies by only 32 °C between the 

worst-case environmental conditions, demonstrating thermal control capability over the environmental 

range. The temperature predictions for the SRG110s are shown in figure 4 and summarized in table 5. The 

total amount of heat transferred to the rover in the cold case was defined as the heat transferred into the 

working fluid. For the hot case, the heat transmitted to the rover only included the radiative heat transmitted 

into the rover electronics cavity, as it was assumed that the pumped loop bypassed the rover internal 

components and transferred the heat instead to the external radiators. 

77.7

74.8

72.0

69.1

66.2

63.3

60.4

57.6

54.7

51.8

48.9

Temperature (°C)

44.4

41.9

39.5

37.0

34.6

32.1

29.7

27.2

24.8

22.3

19.9

Temperature (°C)

Mars 

Hot 

Mars 

Cold 

77.7

74.8

72.0

69.1

66.2

63.3

60.4

57.6

54.7

51.8

48.9

Temperature (°C)

77.7

74.8

72.0

69.1

66.2

63.3

60.4

57.6

54.7

51.8

48.9

Temperature (°C)

44.4

41.9

39.5

37.0

34.6

32.1

29.7

27.2

24.8

22.3

19.9

Temperature (°C)

44.4

41.9

39.5

37.0

34.6

32.1

29.7

27.2

24.8

22.3

19.9

Temperature (°C)

Mars 

Hot 

Mars 

Cold 

Figure 4.—Temperature Profiles for the SRG110 Mars Rover Configuration 

in Hot and Cold Environments. 

Average SRG110 Surface 62 30

Temperature (°C)

Fluid Temperature into Rover (°C) 54 22

Waste Heat to Rover – 2 SRG110s 58* 512**

BOM (Wt

)

Total Electrical Power – 2 SRG110s 206 216

BOM (Wdc)

Condition Hot Case Cold Case

Average SRG110 Surface 62 30

Temperature (°C)

Fluid Temperature into Rover (°C) 54 22

Waste Heat to Rover – 2 SRG110s 58* 512**

BOM (Wt

)

Total Electrical Power – 2 SRG110s 206 216

BOM (Wdc)

Condition Hot Case Cold Case

Table 5. Thermal Analysis Results for the SRG110 Mars 

Rover Configuration

Notes: * Includes radiation only

** Includes heat transported by pumped fluid loop only
NASA/TM—2005-213990 11 

F. Titan Lander 

As the final example for a possible mission, the Titan Lander concept was considered for the SRG110. 

Information for this scenario was taken from the recently accomplished Huygens probe mission by the 

European Space Agency (ESA) and NASA JPL (ref. 6). The Huygens probe was released by the Cassini 

spacecraft on December 25, 2004, and landed on the Titan surface on January 14, 2005, where it survived 

for roughly 2 hours before its limited battery power was exhausted. Had the Huygens probe included an 

RPS system such as the SRG110, its mission could have been extended well beyond 2 hours and much 

more scientific information could have been retrieved. Titan represents an unusual environment that may 

include liquid methane and the possibility of hydrocarbon rain. The atmosphere is very thick and for this 

analysis it was assumed to effectively attenuate the extremely weak solar flux at the surface. The wind 

speeds in the upper atmosphere have been estimated at approximately 120 m/s; however, no information 

was available for the estimated wind speed at the surface. For the purposes of this study, it was assumed 

that the surface wind speed would be approximately 10 percent of the upper atmosphere wind. The Titan 

surface conditions assumed for this study are summarized in table 6. 

Diffuse Solar Flux (W/m 0 2)

Surface Temperature (°C) -180

Atmospheric Pressure (kPa) 152

Surface Wind Speed (m/s) 12

Direct Solar Flux (W/m 0 2)

Location Surface

Condition Cold Case

Diffuse Solar Flux (W/m 0 2)

Surface Temperature (°C) -180

Atmospheric Pressure (kPa) 152

Surface Wind Speed (m/s) 12

Direct Solar Flux (W/m 0 2)

Location Surface

Condition Cold Case

Table 6. Titan Surface Conditions Assumed

1. Titan Lander Configuration 

Like the Mars Rover, the Titan Lander would require an internal mount for two SRG110s. Constant 

conductance heat pipes (CCHPs) running from the base of the SRG110 fins to the internal components 

served as the thermal management system for the onboard components. Since it was expected that there 

was only a “cold” condition on Titan, constant conductance heat pipes or loop heat pipes could be used 

since the lander’s demand for waste heat would be constant. For this investigation it was assumed that the 

CCHPs would be aluminum with ammonia as the working fluid. Additionally, the SRG110 radiator fins 

would be as small as possible to minimize the heat transfer to the environment. 

2. Titan Lander Thermal Model 

A thermal model was constructed using Thermal Desktop® to predict the resulting temperatures and 

heat transferred to the lander. It was assumed that the lander requires approximately 500 Wt to survive 

these extremely cold conditions. 

3. Thermal Analysis Results for Titan Lander Mission 

The thermal analysis results showed that the SRG110s operated within allowable temperature limits and 

still provided the required thermal and electrical power to the lander. The temperature predictions for the 

SRG110s and the representation of the rover are shown in figure 5 and summarized in table 7. 
NASA/TM—2005-213990 12 

25.0

22.0

19.0

16.0

13.0

10.0

7.0

4.0

1.0

-2.0

-5.0

Temperature (°C)

25.0

22.0

19.0

16.0

13.0

10.0

7.0

4.0

1.0

-2.0

-5.0

Temperature (°C)

25.0

22.0

19.0

16.0

13.0

10.0

7.0

4.0

1.0

-2.0

-5.0

Temperature (°C)

25.0

22.0

19.0

16.0

13.0

10.0

7.0

4.0

1.0

-2.0

-5.0

Temperature (°C)

Figure 5.—Temperature Profiles for the SRG110 Titan 

Lander Configuration. 

Average SRG110 Surface Temperature (°C) 10

Heat Pipe Average Temperature (°C) -30

Waste Heat to Lander – 2 SRG110s BOM (W 482* t

)

Total Electrical Power – 2 SRG110s BOM (W 228 dc)

Condition Cold Case

Average SRG110 Surface Temperature (°C) 10

Heat Pipe Average Temperature (°C) -30

Waste Heat to Lander – 2 SRG110s BOM (W 482* t

)

Total Electrical Power – 2 SRG110s BOM (W 228 dc)

Condition Cold Case

Table 7. Thermal Analysis Results for the SRG110 Titan 

Lander Configuration

Notes: * Includes heat transported by CCHPs only

V. Summary and Conclusions 

High-efficiency RPSs, such as the SRG110, can be very flexible in their configuration and provide 

operation as both a power source and heat source in a range of environments. This paper outlined several 

methods that can be employed to utilize the waste heat from a high-efficiency RPS. In all of the examples, 

the amount of waste heat recovered from the high-efficiency RPS was typical of the usable waste heat 

recovered from a low-efficiency RTG. The advantages of using a high-efficiency power source extend past 

the benefit of reduced fuel content. The thermal advantages of having a lower surface temperature allow for 

greater flexibility and more direct use of the waste heat for spacecraft thermal management, including 

elimination of the need for additional radiation heat exchangers. Additionally, the lower surface 

temperatures of the high-efficiency RPS make many types of component thermal shields unnecessary. The 

cost savings for high-efficiency sources include not only the fuel material costs, but also all the costs 

associated with design, safety verification, and heat source processing. Mass savings are also realized by 

high-efficiency RPSs through reduction in hardware and lower heat source mass; the benefit of which could 

be realized by using smaller launch vehicles or increasing the scientific payload. 

Examples of extreme mission environments were investigated, and it was shown that the higherefficiency systems, like the SRG110, can meet the same electrical power and usable waste heat 

requirements as an RTG for less cost and lower mass, while providing more flexibility for spacecraft design 

and easier integration with the spacecraft. 
NASA/TM—2005-213990 13 

References 

1. Planning & Human Systems, Inc., Atomic Power in Space: A History, Department of Energy, 

Washington, DC, 1987. 

2. Lockheed Martin Astronautics Valley Forge Operations, “Final Technical Report: GPHS-RTGs in 

Support of the Cassini Mission,” Document No. RR18, US Department of Energy, Oakland, CA, 1998. 

3. National Aeronautics and Space Administration (NASA), “New Horizons AO, Radioisotope Power 

System (RPS) Information Summary,” Washington, DC, October 2003. 

4. Department of Energy, “Space Radioisotope Power Systems; Stirling Radioisotope Generator,” April, 

2002, URL: http://www.ne.doe.gov/pdf/stirling.pdf [cited April 25, 2005]. 

5. Mauzy, A. (ed.), “NMT’s Contributions to the Cassini Saturn Mission Follow Division’s Space 

Exploration Tradition,” The Actinide Research Quarterly, US Department of Energy, Los Alamos 

National Laboratory, Fall 1996. pp. 1–10. 

6. Leberton, J.P., and Matson, D.L., “Huygens: Science, Payload & Mission”, European Space Agency 

(ESA), ESA-SP-1177, Noordwijk, The Netherlands, 1997. 

7. Department of Energy, “Nuclear Power in Space,” DOE/NE-0071, Department of Energy Office of 

Nuclear Energy, Science and Technology, Washington, DC. 

8. Birur, G.C., Bhandari, P., and Gram, M.B., “Mechanical Pumped Cooling Loop for Spacecraft 

Thermal Control,” SAE Technical Paper No. 961488, 26th International Conference on Environmental 

Systems, Monterey, CA, July 8–11, 1996. 

9. Birur, G.C., and Bhandari, P., “Mars Pathfinder Active Heat Rejection System: Successful Flight 

Demonstration of a Mechanically Pumped Cooling Loop,” SAE Technical Paper No. 981684, 28th 

International Conference on Environmental Systems, Danvers, MA, July 13–16, 1998. 

10. Ganapathi, G.B., Birur, G.C., Tsuyuki, G.T., McGrath, P.L., and Patzold, J. D., “Active Heat Rejection 

System on Mars Exploration Rover—Design Changes from Mars Pathfinder,” Space Technology and 

Applications International Forum (STAIF-2003) edited by M. El-Genk, AIP Conference Proceedings 

654, Melville, New York, 2003, pp. 206–217. 

11. Abelson, R.D. (ed.), “Enabling Exploration with Small Radioisotope Power Systems,” NASA Jet 

Propulsion Laboratory, JPL Pub 04–10, September, 2004. 

12. Abelson, R.D. (ed.), “Expanding Frontiers with Standard Radioisotope Power Systems,” NASA Jet 

Propulsion Laboratory, JPL D-28902, January, 2004. 

13. Panczak, T.D., Ring, S.G., Welch, M.J., and Johnson, D., “Thermal Desktop® Users Manual, Version 

4.7,” C&R Technologies, Littleton, CO, 2004.
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AIAA–2005–5548

E–15314

WBS–22–972–20–01

19

Utilizing Radioisotope Power System Waste Heat for Spacecraft

Thermal Management

David R. Pantano, Frank Dottore, E. Wayne Tobery, Steven M. Geng,

Jeffrey G. Schreiber, and Joseph L. Palko

Spacecraft temperature; Temperature control; Heat transfer; Stirling cycle

Unclassified -Unlimited

Subject Category: 34

Prepared for the Third International Energy Conversion Engineering Conference sponsored by the American Institute of Aeronautics

and Astronautics, San Francisco, California, August 15–18, 2005. David R. Pantano, Frank Dottore, and E. Wayne Tobery, Lockheed

Martin Space Systems Company, 230 Mall Boulevard, King of Prussia, Pennsylvania 19406; Steven M. Geng and Jeffrey G.

Schreiber, NASA Glenn Research Center; and Joseph L. Palko, Connecticut Reserve Technologies, Inc., 10030 Greenwich Drive,

Strongsville, Ohio 44136. Responsible person, Steven M. Geng, organization code RPT, 216–433–6145.

An advantage of using a Radioisotope Power System (RPS) for deep space or planetary surface missions is the readily available waste heat, which can

be used for a number of beneficial purposes including: maintaining electronic components within a controlled temperature range, warming propulsion

tanks and mobility actuators, and maintaining liquid propellants above their freezing temperature. Previous missions using Radioisotope Thermoelectric

Generators (RTGs) dissipated large quantities of waste heat due to the low efficiency of the thermoelectric conversion technology. The next generation

RPSs, such as the 110-Watt Stirling Radioisotope Generator (SRG110) will have higher conversion efficiencies, thereby rejecting less waste heat at a

lower temperature and may require alternate approaches to transferring waste heat to the spacecraft. RTGs, with efficiencies of ~6 to 7 percent, reject

their waste heat at the relatively high heat rejection temperature of 200 °C. This is an advantage when rejecting heat to space; however, transferring heat

to the internal spacecraft components requires a large and heavy radiator heat exchanger. At the same time, sensitive spacecraft instruments must be

shielded from the thermal radiation of the RTG. The SRG110, with an efficiency around 22 percent and 50 °C nominal housing surface temperature,

can readily transfer the available waste heat directly via heat pipes, thermal straps, or fluid loops. The lower temperatures associated with the SRG110

avoid the chances of overheating other scientific components, eliminating the need for thermal shields. This provides the spacecraft designers more

flexibility when locating the generator for a specific mission. A common misconception with high-efficiency systems is that there is not enough waste

heat for spacecraft thermal management. This paper will dispel this misconception and investigate the use of a high-efficiency SRG110 for spacecraft

thermal management and outline potential methods of waste heat utilization in several conceptual missions (Lunar Rover, Mars Rover, and Titan

Lander). The advantages associated with the SRG110 as they relate to ease of assembly, less complex interfaces, and overall mass savings for a

spacecraft will be highlighted.