**Radioisotope Power: A Key Technology for Deep Space Exploration**

George R. Schmidt1, Thomas J. Sutliff1 and Leonard A. Dudzinski2 *1NASA Glenn Research Center, 2NASA Headquarters USA* 

#### **1. Introduction**

Radioisotope Power Systems (RPS) generate electrical power by converting heat released from the nuclear decay of radioactive isotopes into electricity. Because all the units that have flown in space have employed thermoelectrics, a static process for heat-to-electrical energy conversion that employs no moving parts, the term, Radioisotope Thermoelectric Generator (RTG), has been more popularly associated with these devices. However, the advent of new generators based on dynamic energy conversion and alternative static conversion processes favors use of "RPS" as a more accurate term for this power technology. RPS were first used in space by the U.S. in 1961. Since that time, the U.S. has flown 41 RTGs, as a power source for 26 space systems on 25 missions. These applications have included Earthorbital weather and communication satellites, scientific stations on the Moon, robotic explorer spacecraft on Mars, and highly sophisticated deep space interplanetary missions to Jupiter, Saturn and beyond. The New Horizons mission to Pluto, which was launched in January 2006, represents the most recent use of an RTG. The former U.S.S.R. also employed RTGs on several of its early space missions. In addition to electrical power generation, the U.S. and former U.S.S.R. have used radioisotopes extensively for heating components and instrumentation.

RPS have consistently demonstrated unique capabilities over other types of space power systems. A comparison between RPS and other forms of space power is shown in Fig. 1, which maps the most suitable power technologies for different ranges of power level and mission duration. In general, RPS are best suited for applications involving long-duration use beyond several months and power levels up to one to 10 kilowatts.

It is important to recognize that solar power competes very well within this power level range, and offers much higher specific powers (power per unit system mass) for applications up to several Astronomical Units (AU) from the Sun. However, RPS offer the unique advantage of being able to operate continuously, regardless of distance and orientation with respect to the Sun. The flight history of RTGs has demonstrated that these systems are long-lived, rugged, compact, highly reliable, and relatively insensitive to radiation and other environmental effects. Thus, RTGs and the more capable RPS options of the future are ideally suited for missions at distances and extreme conditions where solarbased power generation becomes impractical. These include travel beyond the asteroid belt, operation within the radiation-intensive environments around Jupiter and close to the Sun,

Radioisotope Power: A Key Technology for Deep Space Exploration 421

**Launch** 

29 June 1961

15 Nov 1961

28 Sep 1963

5 Dec 1963

21 Apr 1964

18 May 1968

14 Apr 1969

14 Nov 1969

**Date Status** 

RTG operated for 15 yrs. Satellite now shutdown.

RTG operated for 9 yrs. Operation

RTG operated as planned. Non-RTG electrical problems on satellite caused failure after 9 months.

RTG operated for over 6 yrs. Satellite lost navigational capability after 1.5

Mission aborted because of launch vehicle failure. RTG burned up on reentry

as designed.

reused.

Mission aborted because of range safety destruct. RTG fuel recovered and

RTGs operated for over 2.5 yrs. No data taken after that.

RTG operated for about 8 years until station was shutdown.

yrs.

intermittent after 1962 high alt test. Last signal in 1971.

**Destination/ Application** 

Earth Orbit/ Navigation Sat

Earth Orbit/ Navigation Sat

Earth Orbit/ Navigation Sat

Earth Orbit/ Navigation Sat

Earth Orbit/ Navigation Sat

Earth Orbit/ Meteorology Sat

Earth Orbit/ Meteorology Sat

Lunar Surface/ Science Station

**Spacecraft/ System** 

1 Transit 4A SNAP-3B7

2 Transit 4B SNAP-3B8

3 Transit 5BN-1 SNAP-9A

4 Transit 5BN-2 SNAP-9A

5 Transit 5BN-3 SNAP-9A

6 Nimbus B-1 SNAP-19B2

7 Nimbus III SNAP-19B3

8 Apollo 12 SNAP-27

**Principal Energy Source (#)** 

RTG (1)

RTG (1)

RTG (1)

RTG (1)

RTG (1)

RTG (2)

RTG (2)

RTG (1)

extended operation within permanently shadowed and occulted areas on planetary surfaces, and general applications requiring robust, unattended operations.

Fig. 1. Suitability of space power system technologies.

Table 1 presents a chronological summary of the U.S. missions that have utilized radioisotopes for electrical power generation. Although three missions were aborted by launch vehicle or spacecraft failures, all of the RTGs that flew met or exceeded design expectations, and demonstrated the principles of safe and reliable operation, long life, high reliability, and versatility of operating in hostile environments. All of the RTGs flown by the U.S. comprise seven basic designs: SNAP-3/3B, SNAP-9A, SNAP-19/19B, SNAP-27, TRANSIT-RTG, MHW-RTG and GPHS-RTG. The first four types were developed by the Atomic Energy Commission (AEC) under the auspices of its Systems for Nuclear Auxilliary Power (SNAP) program. Although the original objective was to provide systems for space, the SNAP program also developed generators for non-space, terrestrial applications.

The GPHS-RTG is the most recently developed unit, and has been the workhorse on all RPS missions since 1989. A cutaway view of the unit is shown in Fig. 2. NASA and the Department of Energy (DOE) are looking beyond this capability, and are currently developing two new units: the Multi-Mission RTG (MMRTG), which draws on the design heritage of the SNAP-19, and the new Advanced Stirling Radioisotope Generator (ASRG) with its much more efficient dynamic conversion cycle.

extended operation within permanently shadowed and occulted areas on planetary

Table 1 presents a chronological summary of the U.S. missions that have utilized radioisotopes for electrical power generation. Although three missions were aborted by launch vehicle or spacecraft failures, all of the RTGs that flew met or exceeded design expectations, and demonstrated the principles of safe and reliable operation, long life, high reliability, and versatility of operating in hostile environments. All of the RTGs flown by the U.S. comprise seven basic designs: SNAP-3/3B, SNAP-9A, SNAP-19/19B, SNAP-27, TRANSIT-RTG, MHW-RTG and GPHS-RTG. The first four types were developed by the Atomic Energy Commission (AEC) under the auspices of its Systems for Nuclear Auxilliary Power (SNAP) program. Although the original objective was to provide systems for space,

the SNAP program also developed generators for non-space, terrestrial applications.

The GPHS-RTG is the most recently developed unit, and has been the workhorse on all RPS missions since 1989. A cutaway view of the unit is shown in Fig. 2. NASA and the Department of Energy (DOE) are looking beyond this capability, and are currently developing two new units: the Multi-Mission RTG (MMRTG), which draws on the design heritage of the SNAP-19, and the new Advanced Stirling Radioisotope Generator (ASRG)

surfaces, and general applications requiring robust, unattended operations.

Fig. 1. Suitability of space power system technologies.

with its much more efficient dynamic conversion cycle.


Radioisotope Power: A Key Technology for Deep Space Exploration 423

**Launch** 

14 Mar 1976

20 Aug 1977

5 Sep 1977

18 Oct 1989

6 Oct 1990

15 Oct 1997

Jan 19 2006

**Date Status** 

Single launch with double payload. LES 8 shutdown in 2004. LES 9 RTG still operating.

RTGs still operating.

successfully operated to Jupiter, Saturn, Uranus, Neptune, and

RTGs still operating.

successfully operated to Jupiter, Saturn, and

RTGs continued to operate until 2003, when spacecraft was

deorbited into Jupiter

RTG continued to operate until 2008, when spacecraft was

RTGs continue to operate successfully. Scientific mission and operations still continue.

RTG continues to operate successfully. Spacecraft in transit to

Pluto.

intentionally

atmosphere.

deactivated.

Spacecraft

beyond.

Spacecraft

beyond.

**Destination/ Application** 

Earth Orbit/ Com Sats

Planetary/ Payload & Spacecraft

Planetary/ Payload & Spacecraft

Planetary/Payload & Spacecraft

Planetary/Payload & Spacecraft

Planetary/Payload & Spacecraft

Planetary/Payload & Spacecraft

**Spacecraft/ System** 

19 LES 8, LES 9 MHW-RTG

20 Voyager 2 MHW-RTG

21 Voyager 1 MHW-RTG

22 Galileo GPHS-RTG

23 Ulysses GPHS-RTG

24 Cassini GPHS-RTG

<sup>25</sup>New

Horizons

(4)

(3)

(3)

(2)

(1)

(3)

(1)

GPHS-RTG

Table 1. U.S. Missions Using Radioisotope Power Systems (RPS)

**Principal Energy Source (#)** 


**Launch** 

11 Apr 1970

31 Jan 1971

26 July 1971

2 Mar 1972

16 Apr 1972

2 Sep 1972

7 Dec 1972

5 Apr 1973

20 Aug 1975

9 Sep 1975

**Date Status** 

Mission aborted. RTG reentered intact with no release of Pu-238. Currently located at bottom of Tonga Trench in South Pacific Ocean.

RTG operated for over 6.5 years until

RTG operated for over 6 years until station was shutdown.

Last signal in 2003. Spacecraft now well beyond orbit of Pluto.

RTG operated for about 5.5 years until

RTG still operating as

station was shutdown.

of mid-1990s.

station was shutdown.

RTG operated for almost 5 years until

Last signal in 1995. Spacecraft now well beyond orbit of Pluto.

RTGs operated for over 6 years until lander was shutdown.

RTGs operated for over 4 years until relay link was lost.

station was shutdown.

**Destination/ Application** 

Lunar Surface/ Science Station

Lunar Surface/ Science Station

Lunar Surface/ Science Station

Planetary/Payload & Spacecraft

Lunar Surface/ Science Station

Earth Orbit/ Navigation Sat

Lunar Surface/ Science Station

Planetary/Payload & Spacecraft

Surf/Payload & Spacecraft

Surf/Payload & Spacecraft

Mars

Mars

**Spacecraft/ System** 

9 Apollo 13 SNAP-27

10 Apollo 14 SNAP-27

11 Apollo 15 SNAP-27

12 Pioneer 10 SNAP-19

13 Apollo 16 SNAP-27

14 Triad-01-1X Transit-

15 Apollo 17 SNAP-27

16 Pioneer 11 SNAP-19

17 Viking 1 SNAP-19

18 Viking 2 SNAP-19

**Principal Energy Source (#)** 

RTG (1)

RTG (1)

RTG (1)

RTG (4)

RTG (1)

RTG (1)

RTG (1)

RTG (4)

RTG (2)

RTG (2)



Radioisotope Power: A Key Technology for Deep Space Exploration 425

A variety of radioisotopes have been evaluated for space and terrestrial applications. The isotope initially selected for development was Cerium-144 (Ce-144), because it was one of the most plentiful fission products available from reprocessing defense reactor fuel at AEC's Hanford Site. Its short half-life (290 days) made Ce-144 compatible with the 6-month military reconnaissance satellite mission envisioned as the RPS application at that time. The cerium oxide fuel form and its heavy fuel capsule met all safety tests for intact containment of the fuel during potential launch abort fires, explosions, and terminal impacts. However, the high radiation field associated with the beta/gamma emission of Ce-144 complicated handling and caused problems with payload interaction, as well as safety issues upon reentry from orbit. Although Ce-144 was used to fuel SNAP-1, the first RTG, it was never

By the late 1950s, large amounts of Polonium-210 (Po-210) became available, also as a byproduct of the nuclear weapons program. Po-210 is an alpha emitter with a very high power density (~1,320 W/cm3) and low radiation emissions. It is made by neutron irradiation of Bismuth-209 targets in a nuclear reactor. It was used in polonium-beryllium neutron sources. Po-210 metal was used to fuel the small (5 We) SNAP-3 RTG in order to demonstrate RTG technology. It was first displayed at the White House in January 1959. Several SNAP-3 RTGs were fueled with Po-210 and used in various exhibits. However, the short 138-day half-life of Po-210 makes it suitable for only limited duration space power

In order to provide a longer-lived radioisotope fuel, Strontium-90 (Sr-90), an abundant fission product with a 28.6-year half-life, was recovered from defense wastes at Hanford. A very stable and insoluble fuel form, strontium-titanate, was developed and widely used in terrestrial power systems. Because Sr-90 and its daughter Yttrium-90 emit high-energy beta particles, they give off significant bremsstrahlung radiation and require heavy shielding. However, shield mass is not as critical for most terrestrial power systems as it is for space

By 1960, Plutonium-238 (Pu-238) had been identified as an attractive radioisotope fuel. It could be made by irradiating Neptunium-237 (Np-237) targets in defense production reactors. The availability of Pu-238 was extremely limited due to a shortage of Np-237 target material, which must be recovered from processing (and recycling) high burn-up, enriched uranium fuel. However, Pu-238 has all the desirable characteristics for a space power system fuel: long half-life (87.74 years), low radiation α-particle emissions, high power density and useful fuel forms (as the metal or the oxide form). Therefore, after flight qualification of its heat source, a Pu-238 fueled SNAP-3A RTG was launched on the Transit

The first Pu-238 heat sources used in space were relatively small and employed Pu-238 metal or plutonium-zirconium alloy fuel forms contained in tantalum-lined superalloy (Haynes-25) fuel capsules. These heat sources withstood all postulated launch pad accident and downrange impact environments, but they were designed to burn-up and disperse throughout the upper atmosphere in the event of reentry from space. This type of accident happened during the fifth launch of an RTG (SNAP-9A aboard Transit 5BN-3) when the spacecraft failed

Subsequent Pu-238 fueled space power systems were designed to use progressively higher temperature fuel forms and containment materials with a progressively higher degree of containment of the fuel under all postulated accident conditions (including reentry). As the

4A Navy navigation satellite in June 1961 – the first use of nuclear power in space.

to achieve orbit and the RTG burned up over the Indian Ocean in April 1964.

used in space.

applications.

power applications.

Fig. 2. GPHS-RTG.

### **2. RPS design**

A typical RPS generator consists of two subsystems: a thermal source and an energy conversion system. The thermal source provides heat, which is produced by the decay process within the radioisotope fuel. This heat is partially transformed into electricity in the energy conversion system. Most of the remaining amount is rejected to space via radiators, although a small portion can be used to heat spacecraft components.

### **2.1 Thermal source**

The performance characteristics of an attractive fuel include: a long half-life (i.e., the time it takes for one-half of the original amount of fuel to decay) compared to the operational mission lifetime; low radiation emissions; high specific power and energy; and a stable fuel form with a high enough melting point. The fuel must be producible in useful quantities and at a reasonable cost (compared to its benefits). It must be capable of being produced and used safely, including in the event of potential launch accidents.

Thermal source designs have been driven by aerospace nuclear safety standards, which have evolved considerably over time. For example, the fuel form for the early SNAP-3B and 9A systems was designed to burn up in the event of an atmospheric reentry, and disperse at high altitudes. Later systems, such as SNAP-19, were designed for fuel containment in the event of reentry. A key design feature now is to immobilize the Pu-238 fuel during all nominal and potentially abnormal phases of the mission, including launch abort, reentry into Earth's atmosphere, and post-reentry impact.

Establishing a fuel production and a fuel form fabrication capability is a very costly and time-consuming endeavor. Flight qualification of a new fuel form requires considerable effort in terms of costs and schedule. In addition, there are only a limited number of radioisotope fuels that meet the requirements for half-life, radiation, power density, fuel form, and availability for use in space power applications.

A typical RPS generator consists of two subsystems: a thermal source and an energy conversion system. The thermal source provides heat, which is produced by the decay process within the radioisotope fuel. This heat is partially transformed into electricity in the energy conversion system. Most of the remaining amount is rejected to space via radiators,

The performance characteristics of an attractive fuel include: a long half-life (i.e., the time it takes for one-half of the original amount of fuel to decay) compared to the operational mission lifetime; low radiation emissions; high specific power and energy; and a stable fuel form with a high enough melting point. The fuel must be producible in useful quantities and at a reasonable cost (compared to its benefits). It must be capable of being produced

Thermal source designs have been driven by aerospace nuclear safety standards, which have evolved considerably over time. For example, the fuel form for the early SNAP-3B and 9A systems was designed to burn up in the event of an atmospheric reentry, and disperse at high altitudes. Later systems, such as SNAP-19, were designed for fuel containment in the event of reentry. A key design feature now is to immobilize the Pu-238 fuel during all nominal and potentially abnormal phases of the mission, including launch abort, reentry

Establishing a fuel production and a fuel form fabrication capability is a very costly and time-consuming endeavor. Flight qualification of a new fuel form requires considerable effort in terms of costs and schedule. In addition, there are only a limited number of radioisotope fuels that meet the requirements for half-life, radiation, power density, fuel

although a small portion can be used to heat spacecraft components.

and used safely, including in the event of potential launch accidents.

into Earth's atmosphere, and post-reentry impact.

form, and availability for use in space power applications.

Fig. 2. GPHS-RTG.

**2. RPS design** 

**2.1 Thermal source** 

A variety of radioisotopes have been evaluated for space and terrestrial applications. The isotope initially selected for development was Cerium-144 (Ce-144), because it was one of the most plentiful fission products available from reprocessing defense reactor fuel at AEC's Hanford Site. Its short half-life (290 days) made Ce-144 compatible with the 6-month military reconnaissance satellite mission envisioned as the RPS application at that time. The cerium oxide fuel form and its heavy fuel capsule met all safety tests for intact containment of the fuel during potential launch abort fires, explosions, and terminal impacts. However, the high radiation field associated with the beta/gamma emission of Ce-144 complicated handling and caused problems with payload interaction, as well as safety issues upon reentry from orbit. Although Ce-144 was used to fuel SNAP-1, the first RTG, it was never used in space.

By the late 1950s, large amounts of Polonium-210 (Po-210) became available, also as a byproduct of the nuclear weapons program. Po-210 is an alpha emitter with a very high power density (~1,320 W/cm3) and low radiation emissions. It is made by neutron irradiation of Bismuth-209 targets in a nuclear reactor. It was used in polonium-beryllium neutron sources. Po-210 metal was used to fuel the small (5 We) SNAP-3 RTG in order to demonstrate RTG technology. It was first displayed at the White House in January 1959. Several SNAP-3 RTGs were fueled with Po-210 and used in various exhibits. However, the short 138-day half-life of Po-210 makes it suitable for only limited duration space power applications.

In order to provide a longer-lived radioisotope fuel, Strontium-90 (Sr-90), an abundant fission product with a 28.6-year half-life, was recovered from defense wastes at Hanford. A very stable and insoluble fuel form, strontium-titanate, was developed and widely used in terrestrial power systems. Because Sr-90 and its daughter Yttrium-90 emit high-energy beta particles, they give off significant bremsstrahlung radiation and require heavy shielding. However, shield mass is not as critical for most terrestrial power systems as it is for space power applications.

By 1960, Plutonium-238 (Pu-238) had been identified as an attractive radioisotope fuel. It could be made by irradiating Neptunium-237 (Np-237) targets in defense production reactors. The availability of Pu-238 was extremely limited due to a shortage of Np-237 target material, which must be recovered from processing (and recycling) high burn-up, enriched uranium fuel. However, Pu-238 has all the desirable characteristics for a space power system fuel: long half-life (87.74 years), low radiation α-particle emissions, high power density and useful fuel forms (as the metal or the oxide form). Therefore, after flight qualification of its heat source, a Pu-238 fueled SNAP-3A RTG was launched on the Transit 4A Navy navigation satellite in June 1961 – the first use of nuclear power in space.

The first Pu-238 heat sources used in space were relatively small and employed Pu-238 metal or plutonium-zirconium alloy fuel forms contained in tantalum-lined superalloy (Haynes-25) fuel capsules. These heat sources withstood all postulated launch pad accident and downrange impact environments, but they were designed to burn-up and disperse throughout the upper atmosphere in the event of reentry from space. This type of accident happened during the fifth launch of an RTG (SNAP-9A aboard Transit 5BN-3) when the spacecraft failed to achieve orbit and the RTG burned up over the Indian Ocean in April 1964.

Subsequent Pu-238 fueled space power systems were designed to use progressively higher temperature fuel forms and containment materials with a progressively higher degree of containment of the fuel under all postulated accident conditions (including reentry). As the

Radioisotope Power: A Key Technology for Deep Space Exploration 427

Encapsulation is an important aspect of the design of the thermal source and consists of several elements, each of which serves one or more important functions in the safe handling and use of the fuel. The state-of-the-art in fuel encapsulation and containment is the

The GPHS is modular in design, thus allowing it to be stacked into variable thermal source configurations. Eighteen of these modules are stacked together to serve as the thermal source for the GPHS-RTG, shown in Fig. 2. It is being used in an 8-module stack for the recently developed MMRTG, and two individual GPHS will be used as the heat source for

Safety is the principal design driver for the GPHS. The main objective is to keep the fuel contained or immobilized to prevent inhalation or ingestion by humans. Each module is composed of five main elements: the fuel; the fuel cladding; the graphite impact shell (GIS); the carbon-bonded carbon fiber (CBCF) insulation; and the Fine Weave Pierced Fabric (FWPFTM) aeroshell. Each GPHS module contains four fuel pellets made of a high-temperature PuO2 ceramic with a thermal inventory of approximately 62.5 Wt (Watts-thermal) per pellet and 250.0 Wt per module. Each module has a total mass of

During its development program in the late 1970's, the GPHS went through a number of exacting engineering tests to assess its performance under operating conditions, including vibration and operating temperature. An extensive safety testing and analyses program was conducted to assess the GPHS performance under a range of postulated accident conditions such as launch pad explosions, projectile impacts, propellant fires, impacts, and atmospheric

**2.2 Fuel encapsulation and containment** 

Fig. 3. GPHS module assembly.

about 1.43 kg.

reentry.

the new ASRG, currently under development.

General Purpose Heat Source (GPHS) module shown in Fig. 3.

intact reentry heat source technology was developed, the fuel inventories (power levels) per launch also increased. A number of RTGs were launched on NASA and Navy missions with Pu-238 dioxide microsphere and plutonia-molybdenum-cermet (PMC) fuel forms in the late-1960s and early-1970s. Since the mid-1970s, pressed Pu-238 oxide fuel forms have been exclusively used in all RPS launched into space.

The amount of Pu-238 that could be produced has always been a limiting factor in its use in space missions. Therefore, several other radioisotopes have been thoroughly evaluated for space use over the years. Sr-90 and Po-210 fuels were considered for use in higher powered military satellite constellations for which there were insufficient quantities of Pu-238 available. These programs were cancelled before they were completed, so these fuels were never used in space by the U.S.

Curium-242 (Cm-242) was selected to fuel an isotope power system for the 90-day Surveyor mission to the Moon. Both the SNAP-11 RTG and SNAP-13 thermionic generators were developed for the Surveyor mission. Cm-242 is produced by reactor irradiation of Americium-241 (Am-241) targets. Cm-242 has a short half-life of 162 days, which is acceptable for a 90 day mission, and has a very high power density, which is necessary for a thermionic heat source. It also has a high melting point oxide fuel form capable of the high operating temperature necessary for thermionic energy conversion. A Cm-242 demonstration heat source was produced for the SNAP-13 engineering unit. However, it was decided that the Surveyor program would not use isotope power units, and Cm-242 fueled power systems have never been used in space. Due to its short half-life, Cm-242 is not suitable for the longer durations required by most space missions.

At one time, Curium-244 (Cm-244) was investigated as a potential alternative to Pu-238, because it was expected to become available in significant quantities from the U.S. program to develop breeder reactor fuel cycles. Cm-244 was considered an attractive space fuel because it has a relatively long half-life (18.2 years), a power density five times greater than that of Pu-238 and has a very stable, high temperature oxide fuel form. However, higher neutron and gamma emissions due to the higher rate of spontaneous fission of Cm-244 would increase shielding requirements for handling and for protection of spacecraft instrumentation. The increase weight of shielding and power flattening equipment required with Cm-244 makes it less desirable than Pu-238, especially for long duration missions. Cm-244 is also more difficult to produce, requiring successive neutron captures starting with Pu-239. Many years ago, several kilograms of Cm-244 were made as a target material for the Californium-252 program, but there is currently no practical production or processing capability for large quantities of Cm-244.

In the final analysis, Pu-238 is clearly superior to other radioisotope fuels for use in long duration space missions. The technology for producing and processing Pu-238 fuel forms has been refined over the past 50 years. Pu-238 fueled heat sources have been through rigorous flight qualification testing and have performed reliably in all of the RPS employed in the U.S. space program to date.

The most significant issue with Pu-238 is its limited availability. For the past 50 years the production and processing of Pu-238 fuel has been accomplished as a by-product of the production of materials for nuclear weapons. The discontinuation of this production in the 1990s eliminated the traditional means for producing Pu-238. During the 2000s, the U.S. began to purchase Pu-238 from Russia. However, this supply is also limited, so in the longterm, resumption of production is necessary.

#### **2.2 Fuel encapsulation and containment**

426 Radioisotopes – Applications in Physical Sciences

intact reentry heat source technology was developed, the fuel inventories (power levels) per launch also increased. A number of RTGs were launched on NASA and Navy missions with Pu-238 dioxide microsphere and plutonia-molybdenum-cermet (PMC) fuel forms in the late-1960s and early-1970s. Since the mid-1970s, pressed Pu-238 oxide fuel forms have been

The amount of Pu-238 that could be produced has always been a limiting factor in its use in space missions. Therefore, several other radioisotopes have been thoroughly evaluated for space use over the years. Sr-90 and Po-210 fuels were considered for use in higher powered military satellite constellations for which there were insufficient quantities of Pu-238 available. These programs were cancelled before they were completed, so these fuels were

Curium-242 (Cm-242) was selected to fuel an isotope power system for the 90-day Surveyor mission to the Moon. Both the SNAP-11 RTG and SNAP-13 thermionic generators were developed for the Surveyor mission. Cm-242 is produced by reactor irradiation of Americium-241 (Am-241) targets. Cm-242 has a short half-life of 162 days, which is acceptable for a 90 day mission, and has a very high power density, which is necessary for a thermionic heat source. It also has a high melting point oxide fuel form capable of the high operating temperature necessary for thermionic energy conversion. A Cm-242 demonstration heat source was produced for the SNAP-13 engineering unit. However, it was decided that the Surveyor program would not use isotope power units, and Cm-242 fueled power systems have never been used in space. Due to its short half-life, Cm-242 is

At one time, Curium-244 (Cm-244) was investigated as a potential alternative to Pu-238, because it was expected to become available in significant quantities from the U.S. program to develop breeder reactor fuel cycles. Cm-244 was considered an attractive space fuel because it has a relatively long half-life (18.2 years), a power density five times greater than that of Pu-238 and has a very stable, high temperature oxide fuel form. However, higher neutron and gamma emissions due to the higher rate of spontaneous fission of Cm-244 would increase shielding requirements for handling and for protection of spacecraft instrumentation. The increase weight of shielding and power flattening equipment required with Cm-244 makes it less desirable than Pu-238, especially for long duration missions. Cm-244 is also more difficult to produce, requiring successive neutron captures starting with Pu-239. Many years ago, several kilograms of Cm-244 were made as a target material for the Californium-252 program, but there is currently no practical production or processing

In the final analysis, Pu-238 is clearly superior to other radioisotope fuels for use in long duration space missions. The technology for producing and processing Pu-238 fuel forms has been refined over the past 50 years. Pu-238 fueled heat sources have been through rigorous flight qualification testing and have performed reliably in all of the RPS employed

The most significant issue with Pu-238 is its limited availability. For the past 50 years the production and processing of Pu-238 fuel has been accomplished as a by-product of the production of materials for nuclear weapons. The discontinuation of this production in the 1990s eliminated the traditional means for producing Pu-238. During the 2000s, the U.S. began to purchase Pu-238 from Russia. However, this supply is also limited, so in the long-

not suitable for the longer durations required by most space missions.

exclusively used in all RPS launched into space.

never used in space by the U.S.

capability for large quantities of Cm-244.

term, resumption of production is necessary.

in the U.S. space program to date.

Encapsulation is an important aspect of the design of the thermal source and consists of several elements, each of which serves one or more important functions in the safe handling and use of the fuel. The state-of-the-art in fuel encapsulation and containment is the General Purpose Heat Source (GPHS) module shown in Fig. 3.

Fig. 3. GPHS module assembly.

The GPHS is modular in design, thus allowing it to be stacked into variable thermal source configurations. Eighteen of these modules are stacked together to serve as the thermal source for the GPHS-RTG, shown in Fig. 2. It is being used in an 8-module stack for the recently developed MMRTG, and two individual GPHS will be used as the heat source for the new ASRG, currently under development.

Safety is the principal design driver for the GPHS. The main objective is to keep the fuel contained or immobilized to prevent inhalation or ingestion by humans. Each module is composed of five main elements: the fuel; the fuel cladding; the graphite impact shell (GIS); the carbon-bonded carbon fiber (CBCF) insulation; and the Fine Weave Pierced Fabric (FWPFTM) aeroshell. Each GPHS module contains four fuel pellets made of a high-temperature PuO2 ceramic with a thermal inventory of approximately 62.5 Wt (Watts-thermal) per pellet and 250.0 Wt per module. Each module has a total mass of about 1.43 kg.

During its development program in the late 1970's, the GPHS went through a number of exacting engineering tests to assess its performance under operating conditions, including vibration and operating temperature. An extensive safety testing and analyses program was conducted to assess the GPHS performance under a range of postulated accident conditions such as launch pad explosions, projectile impacts, propellant fires, impacts, and atmospheric reentry.

Radioisotope Power: A Key Technology for Deep Space Exploration 429

After passing through the energy conversion system, the unconverted waste heat must be rejected to the environment at lower temperatures. For space power systems some of the waste heat can be utilized to control the temperature of the spacecraft equipment, but

Thus, the operating temperatures for an RPS are set on the hot side by the heat source and conversion system material limitations (Thot) and on the cold side by the size, weight, and heat sink conditions of the radiator (Tcold). The overall efficiency of the energy conversion system is limited to something less than the Carnot efficiency of (Thot – Tcold)/Thot. Higher efficiencies can significantly reduce fuel usage, which has many implications for

Conversion system reliability is another important consideration. Since mission success depends on having sufficient electrical power over the life of the mission, conversion system selection must be consistent with mission power levels and lifetimes. For instance, it makes little sense to combine an unreliable or short-lived energy conversion unit with a 100% reliable, long-lived isotope heat source. Graceful power degradation over the life of a

Other important considerations in selecting a system include mass, size, ruggedness to withstand shock and vibration loads, survivability in hostile particle and radiation environments, scalability in power levels, flexibility in integration with various types of spacecraft (and launch vehicles), and versatility to operate in the vacuum of deep space or

All of the RPS units flown in space have utilized thermoelectric energy conversion. Thermoelectric converters are useful over a very wide range of power levels (from milliwatts to kilowatts) and their operating temperatures are ideally suited for radioisotope heat sources. Thermoelectric converters are reliable over operational lifetimes of several decades, compact, rugged, radiation resistant, easily adapted to a wide range of applications, and produce no noise, vibration or torque during operation. Thermoelectric converters require no start-up devices to operate, and begin producing electrical power (direct current and voltage) as soon as the heat source is installed. Power output is easily regulated at design level by maintaining a matched resistive load on the converter. The only disadvantage of thermoelectrics is their relatively low conversion efficiencies , which is

Thermoelectric materials, when operating over a temperature gradient, produce a voltage due to the Seebeck effect. When connected in series with a load, the internally generated voltage causes a current to flow through the load producing useful power. The Seebeck effect was discovered in 1825, but had little practical use, except in measuring temperatures with dissimilar metal thermocouples. With the advent of semiconductor materials in the

Power is produced in a thermoelectric element by placing it between a heat source and a heat sink. Good thermoelectric semiconductor materials have large Seebeck voltages in combination with a relatively high electrical conductivity and low thermal conductivity (in contrast to most metals). By proper doping, n and p type elements can be formed so that current will flow in the same or opposite directions as the heat. By electrically joining the n and p elements through a hot shoe, a thermocouple is formed which can be connected to other

ultimately the waste heat must be radiated to the space vacuum environment.

cost, availability, size, weight, and safety.

mission is acceptable as long as it is within predictable limits.

1950s, application of thermoelectrics has expanded dramatically.

on planetary surfaces with or without solar energy.

**2.4 Thermoelectric energy conversion** 

typically less than 10%.

The fuel pellets, one of which is shown in Fig. 4, are individually encapsulated in a welded iridium alloy clad.

Fig. 4. GPHS PuO2 Fuel Pellet.

The alloy is capable of resisting oxidation in a hypothetical post-impact environment while also being chemically compatible with the fuel and graphitic components during hightemperature operation and postulated accident environments.

Two fueled clads are encased in a cylindrical graphite impact shell (GIS) made of FWPFTM, a carbon-carbon composite material. The GIS is designed to provide protection to the fueled clads for postulated impact. Two of these GIS assemblies, each containing two fueled clads, are located in each FWPFTM aeroshell. A carbon-bonded carbon fiber (CBCF) insulator surrounds each GIS within the aeroshell to limit the peak temperature of the fueled clad during inadvertent reentry and to maintain a sufficiently high temperature to ensure its ductility upon the subsequently postulated impact.

The aeroshell serves as the primary structural member of the GPHS module as it is stacked inside the RPS unit. The aeroshell is designed to contain the two graphite impact shell assemblies under a wide range of postulated reentry conditions and to provide additional protection against postulated impacts on hard surfaces at terminal velocity. FWPFTM was selected because its composite structure gave it a high margin of safety against the thermal stresses associated with postulated atmospheric reentries. The aeroshell also provides protection for the fueled clads from postulated launch vehicle explosion overpressures and fragment impacts and it can provide protection in the event of a propellant fire.

#### **2.3 Power conversion systems**

A portion of the heat generated from the thermal source is converted to useful electrical energy in the power conversion system. There are two general classes of energy conversion systems: static and dynamic. Static systems include thermoelectric, thermionic, and thermophotovoltaic conversion devices which can convert heat to electricity directly with no moving parts. Dynamic systems involve heat engines with working fluids that transform heat to mechanical energy which in turn is used to generate electricity. Dynamic systems include Stirling, Brayton and Rankine cycle engines that operate with various types of working fluids.

The fuel pellets, one of which is shown in Fig. 4, are individually encapsulated in a welded

The alloy is capable of resisting oxidation in a hypothetical post-impact environment while also being chemically compatible with the fuel and graphitic components during high-

Two fueled clads are encased in a cylindrical graphite impact shell (GIS) made of FWPFTM, a carbon-carbon composite material. The GIS is designed to provide protection to the fueled clads for postulated impact. Two of these GIS assemblies, each containing two fueled clads, are located in each FWPFTM aeroshell. A carbon-bonded carbon fiber (CBCF) insulator surrounds each GIS within the aeroshell to limit the peak temperature of the fueled clad during inadvertent reentry and to maintain a sufficiently high temperature to ensure its

The aeroshell serves as the primary structural member of the GPHS module as it is stacked inside the RPS unit. The aeroshell is designed to contain the two graphite impact shell assemblies under a wide range of postulated reentry conditions and to provide additional protection against postulated impacts on hard surfaces at terminal velocity. FWPFTM was selected because its composite structure gave it a high margin of safety against the thermal stresses associated with postulated atmospheric reentries. The aeroshell also provides protection for the fueled clads from postulated launch vehicle explosion overpressures and

A portion of the heat generated from the thermal source is converted to useful electrical energy in the power conversion system. There are two general classes of energy conversion systems: static and dynamic. Static systems include thermoelectric, thermionic, and thermophotovoltaic conversion devices which can convert heat to electricity directly with no moving parts. Dynamic systems involve heat engines with working fluids that transform heat to mechanical energy which in turn is used to generate electricity. Dynamic systems include Stirling, Brayton and Rankine cycle engines that operate with various types of

fragment impacts and it can provide protection in the event of a propellant fire.

temperature operation and postulated accident environments.

ductility upon the subsequently postulated impact.

iridium alloy clad.

Fig. 4. GPHS PuO2 Fuel Pellet.

**2.3 Power conversion systems** 

working fluids.

After passing through the energy conversion system, the unconverted waste heat must be rejected to the environment at lower temperatures. For space power systems some of the waste heat can be utilized to control the temperature of the spacecraft equipment, but ultimately the waste heat must be radiated to the space vacuum environment.

Thus, the operating temperatures for an RPS are set on the hot side by the heat source and conversion system material limitations (Thot) and on the cold side by the size, weight, and heat sink conditions of the radiator (Tcold). The overall efficiency of the energy conversion system is limited to something less than the Carnot efficiency of (Thot – Tcold)/Thot. Higher efficiencies can significantly reduce fuel usage, which has many implications for cost, availability, size, weight, and safety.

Conversion system reliability is another important consideration. Since mission success depends on having sufficient electrical power over the life of the mission, conversion system selection must be consistent with mission power levels and lifetimes. For instance, it makes little sense to combine an unreliable or short-lived energy conversion unit with a 100% reliable, long-lived isotope heat source. Graceful power degradation over the life of a mission is acceptable as long as it is within predictable limits.

Other important considerations in selecting a system include mass, size, ruggedness to withstand shock and vibration loads, survivability in hostile particle and radiation environments, scalability in power levels, flexibility in integration with various types of spacecraft (and launch vehicles), and versatility to operate in the vacuum of deep space or on planetary surfaces with or without solar energy.

#### **2.4 Thermoelectric energy conversion**

All of the RPS units flown in space have utilized thermoelectric energy conversion. Thermoelectric converters are useful over a very wide range of power levels (from milliwatts to kilowatts) and their operating temperatures are ideally suited for radioisotope heat sources. Thermoelectric converters are reliable over operational lifetimes of several decades, compact, rugged, radiation resistant, easily adapted to a wide range of applications, and produce no noise, vibration or torque during operation. Thermoelectric converters require no start-up devices to operate, and begin producing electrical power (direct current and voltage) as soon as the heat source is installed. Power output is easily regulated at design level by maintaining a matched resistive load on the converter. The only disadvantage of thermoelectrics is their relatively low conversion efficiencies , which is typically less than 10%.

Thermoelectric materials, when operating over a temperature gradient, produce a voltage due to the Seebeck effect. When connected in series with a load, the internally generated voltage causes a current to flow through the load producing useful power. The Seebeck effect was discovered in 1825, but had little practical use, except in measuring temperatures with dissimilar metal thermocouples. With the advent of semiconductor materials in the 1950s, application of thermoelectrics has expanded dramatically.

Power is produced in a thermoelectric element by placing it between a heat source and a heat sink. Good thermoelectric semiconductor materials have large Seebeck voltages in combination with a relatively high electrical conductivity and low thermal conductivity (in contrast to most metals). By proper doping, n and p type elements can be formed so that current will flow in the same or opposite directions as the heat. By electrically joining the n and p elements through a hot shoe, a thermocouple is formed which can be connected to other

Radioisotope Power: A Key Technology for Deep Space Exploration 431

Stirling cycle engines use a light working gas that expands by absorption of heat on the hot side and contracts by rejection of heat on the cold side causing rapidly changing pressure cycles across a piston forcing it to move in a reciprocating fashion. The movement of the

Traditional Stirling engines use a rhombic drive mechanism to convert the reciprocating motion into a rotary motion that drives an ordinary rotating alternator. This requires lubrication of a gear-box and seals to separate the working gas from the lubricating oil. The engine housing cannot be hermetically sealed because of the penetration of the rotary power

A more recent development is the Free Piston Stirling engine which requires no lubricating fluids and produces electricity by means of a linear alternator within the hermetically sealed engine housing. The piston moves back and forth at a resonant frequency on a cushion of working gas between it and the surrounding cylinder wall. Piston displacement is controlled by gas pressure across the piston. A permanent magnet is attached to the power piston and produces electrical currents in surrounding alternator coils as it vibrates back and forth. Since the reciprocating motion of the piston would cause unbalanced vibration loads, these Stirling engines usually are designed in pairs with dynamically opposed pistons

Heat is also exchanged between the hot and cold gas flowing from one side of the piston to the other to enhance the conversion efficiency. Due to the limited volume of working gas within the Stirling engine, heat transfer between the heat source and the heater head of the engine, between the hot and cold gas, and between the cold gas and a radiator system are the most challenging requirements for an optimum engine design. The Stirling cycle provides the highest conversion efficiencies of any dynamic cycles at the same cycle temperatures. Therefore, efficiencies of 30% or more are possible at operating temperatures achievable with isotope heat sources and oxidation-resistant superalloy structural materials. The Stirling engine also promises to retain its high performance characteristics at lower power levels compared to the other dynamic systems, which is also attractive for

Research and development of other energy conversion technologies has been an important aspect of RPS programs in the past. Although thermoelectrics and Stirling have received the most attention, there are several other technologies that could achieve higher heat-to-electric

One of these is Thermophotovoltaics (TPV), which is another static form of electrical power conversion. A thermophotovoltaic (TPV) converter transforms the energy from infrared photons emitted by a hot surface into electricity using photovoltaic (PV) cells. TPV converters use advanced PV cells, spectrally-tuned to optimize conversion of the emitted photon energy. Controlling the frequency of photon energy impinging on the PV cells by means of selective emitters, PV cell materials and filter properties are key to achieving high performance. Studies in the past have suggested the possibility of achieving efficiencies of

On the dynamic side, the Brayton is another thermodynamic cycle consisting of a turbine/alternator, compressor and heat exchangers. An additional recuperator heat exchanger is often used to transfer heat within the cycle and improve cycle efficiency. An inert gas working fluid, typically a mixture of helium and xenon, is sequentially heated in

conversion efficiencies and considerably lower masses than the systems in use today.

piston can drive a linear alternator to produce electricity.

so that no net load is transmitted to the engine mounts.

radioisotope power systems.

up to 20% with TPV.

**2.6 Other energy conversion technologies** 

shaft. Such Stirling engines have been widely used throughout the world.

thermocouples at the cold shoe to form a converter with the desired output voltage and current. Thermocouples can be connected in a series-parallel arrangement to enhance reliability by minimizing the effect on total power due to an open circuit or short circuit failure in a single thermocouple. Typically, thermoelectric couples are low voltage, high current devices so a number of them must be connected in series to produce normal load voltages.

The most widely used thermoelectric materials in order of increasing temperature capability, are: Bismuth Telluride (BiTe); Lead Telluride (PbTe); Tellurides of Antimony, Germanium and Silver (TAGS); Lead Tin Telluride (PbSnTe); and Silicon Germanium (SiGe). All except BiTe have been used in space RTG applications. Many more materials have been, and are still being, investigated in hopes of finding that ideal thermoelectric material from which to produce higher efficiency, lower mass, and more stable performance over longer operating lifetimes.

The telluride materials are limited to a maximum hot junction temperature of 550 C. Due to the deleterious effects of oxygen on these materials and their high vapor pressure, the tellurides must be operated in a sealed generator with an inert cover gas to retard sublimation and vapor phase transport within the converter. Bulk-type, fibrous thermal insulation must be used due to the presence of the cover gas. Buildup of helium gas from αparticle emission must be controlled by using a separate container around the heat source or permeable seals in the generator design. Gas management considerations in the generator housing design and the use of bulk insulation materials increase the size and weight of the generator. However, this type of RTG is equally useful for space vacuum or for planetary atmospheric applications.

SiGe materials can be operated at hot junction temperatures up to 1,000 C. Their sublimation rates and oxidation effects, even at these higher temperatures, can be controlled by use of sublimation barriers around the elements and an inert cover gas within the generator during ground operation. A pressure release device, designed to open upon reaching orbitral altitude, opens the generator to space vacuum for operation on deep space missions. This allows use of multifoil thermal insulation and also vents the helium to space as it is generated. A SiGe RTG is usually smaller and lighter than is a telluride RTG of similar power level.

The overall efficiency of the two types of thermoelectric generators are comparable. Although the tellurides have a higher material efficiency than SiGe, the SiGe operates over a larger temperature gradient. Cold junction temperatures are determined more by radiator weight than by efficiency considerations for space RTGs and are normally in the range of 200-300 C. Although various convectively cooled radiator systems have been developed (e.g., heat pipes), conductively coupled finned radiators attached to the generator housing are normally more weight efficient for low-powered RTGs of up to 300 We.

#### **2.5 Stirling energy conversion**

For higher power levels of 100 We and above, the more efficient dynamic power conversion technologies enable better use of the limited radioisotope fuel, offer systems with a higher power-to-weight ratio and make it easier to integrate the radioisotope power system with the spacecraft compared to the number of RTGs required to produce kilowatts of power. Dynamic heat-to-electricity conversion efficiencies of 25% and more are achievable, which reduce the radioisotope inventory by at least one-quarter of that for RTGs. This reduces mass, cost, and potential safety risks for higher-powered radioisotope systems.

thermocouples at the cold shoe to form a converter with the desired output voltage and current. Thermocouples can be connected in a series-parallel arrangement to enhance reliability by minimizing the effect on total power due to an open circuit or short circuit failure in a single thermocouple. Typically, thermoelectric couples are low voltage, high current devices so a number of them must be connected in series to produce normal load voltages. The most widely used thermoelectric materials in order of increasing temperature capability, are: Bismuth Telluride (BiTe); Lead Telluride (PbTe); Tellurides of Antimony, Germanium and Silver (TAGS); Lead Tin Telluride (PbSnTe); and Silicon Germanium (SiGe). All except BiTe have been used in space RTG applications. Many more materials have been, and are still being, investigated in hopes of finding that ideal thermoelectric material from which to produce higher efficiency, lower mass, and more stable performance

The telluride materials are limited to a maximum hot junction temperature of 550 C. Due to the deleterious effects of oxygen on these materials and their high vapor pressure, the tellurides must be operated in a sealed generator with an inert cover gas to retard sublimation and vapor phase transport within the converter. Bulk-type, fibrous thermal insulation must be used due to the presence of the cover gas. Buildup of helium gas from αparticle emission must be controlled by using a separate container around the heat source or permeable seals in the generator design. Gas management considerations in the generator housing design and the use of bulk insulation materials increase the size and weight of the generator. However, this type of RTG is equally useful for space vacuum or for planetary

SiGe materials can be operated at hot junction temperatures up to 1,000 C. Their sublimation rates and oxidation effects, even at these higher temperatures, can be controlled by use of sublimation barriers around the elements and an inert cover gas within the generator during ground operation. A pressure release device, designed to open upon reaching orbitral altitude, opens the generator to space vacuum for operation on deep space missions. This allows use of multifoil thermal insulation and also vents the helium to space as it is generated. A SiGe RTG is usually smaller and lighter than is a telluride RTG of similar

The overall efficiency of the two types of thermoelectric generators are comparable. Although the tellurides have a higher material efficiency than SiGe, the SiGe operates over a larger temperature gradient. Cold junction temperatures are determined more by radiator weight than by efficiency considerations for space RTGs and are normally in the range of 200-300 C. Although various convectively cooled radiator systems have been developed (e.g., heat pipes), conductively coupled finned radiators attached to the generator housing

For higher power levels of 100 We and above, the more efficient dynamic power conversion technologies enable better use of the limited radioisotope fuel, offer systems with a higher power-to-weight ratio and make it easier to integrate the radioisotope power system with the spacecraft compared to the number of RTGs required to produce kilowatts of power. Dynamic heat-to-electricity conversion efficiencies of 25% and more are achievable, which reduce the radioisotope inventory by at least one-quarter of that for RTGs. This reduces

are normally more weight efficient for low-powered RTGs of up to 300 We.

mass, cost, and potential safety risks for higher-powered radioisotope systems.

over longer operating lifetimes.

atmospheric applications.

**2.5 Stirling energy conversion** 

power level.

Stirling cycle engines use a light working gas that expands by absorption of heat on the hot side and contracts by rejection of heat on the cold side causing rapidly changing pressure cycles across a piston forcing it to move in a reciprocating fashion. The movement of the piston can drive a linear alternator to produce electricity.

Traditional Stirling engines use a rhombic drive mechanism to convert the reciprocating motion into a rotary motion that drives an ordinary rotating alternator. This requires lubrication of a gear-box and seals to separate the working gas from the lubricating oil. The engine housing cannot be hermetically sealed because of the penetration of the rotary power shaft. Such Stirling engines have been widely used throughout the world.

A more recent development is the Free Piston Stirling engine which requires no lubricating fluids and produces electricity by means of a linear alternator within the hermetically sealed engine housing. The piston moves back and forth at a resonant frequency on a cushion of working gas between it and the surrounding cylinder wall. Piston displacement is controlled by gas pressure across the piston. A permanent magnet is attached to the power piston and produces electrical currents in surrounding alternator coils as it vibrates back and forth. Since the reciprocating motion of the piston would cause unbalanced vibration loads, these Stirling engines usually are designed in pairs with dynamically opposed pistons so that no net load is transmitted to the engine mounts.

Heat is also exchanged between the hot and cold gas flowing from one side of the piston to the other to enhance the conversion efficiency. Due to the limited volume of working gas within the Stirling engine, heat transfer between the heat source and the heater head of the engine, between the hot and cold gas, and between the cold gas and a radiator system are the most challenging requirements for an optimum engine design. The Stirling cycle provides the highest conversion efficiencies of any dynamic cycles at the same cycle temperatures. Therefore, efficiencies of 30% or more are possible at operating temperatures achievable with isotope heat sources and oxidation-resistant superalloy structural materials. The Stirling engine also promises to retain its high performance characteristics at lower power levels compared to the other dynamic systems, which is also attractive for radioisotope power systems.

#### **2.6 Other energy conversion technologies**

Research and development of other energy conversion technologies has been an important aspect of RPS programs in the past. Although thermoelectrics and Stirling have received the most attention, there are several other technologies that could achieve higher heat-to-electric conversion efficiencies and considerably lower masses than the systems in use today.

One of these is Thermophotovoltaics (TPV), which is another static form of electrical power conversion. A thermophotovoltaic (TPV) converter transforms the energy from infrared photons emitted by a hot surface into electricity using photovoltaic (PV) cells. TPV converters use advanced PV cells, spectrally-tuned to optimize conversion of the emitted photon energy. Controlling the frequency of photon energy impinging on the PV cells by means of selective emitters, PV cell materials and filter properties are key to achieving high performance. Studies in the past have suggested the possibility of achieving efficiencies of up to 20% with TPV.

On the dynamic side, the Brayton is another thermodynamic cycle consisting of a turbine/alternator, compressor and heat exchangers. An additional recuperator heat exchanger is often used to transfer heat within the cycle and improve cycle efficiency. An inert gas working fluid, typically a mixture of helium and xenon, is sequentially heated in

Radioisotope Power: A Key Technology for Deep Space Exploration 433

That SNAP-3 actually never flew in space, but it became an invaluable showpiece for RPS and the SNAP program. President Eisenhower, who had been keenly interested in developing nuclear power for U.S. surveillance satellites, was shown this breakthrough device in January 1959, when the SNAP-3 was displayed on his desk in the Oval Office (Fig. 5). Eisenhower used the opportunity to emphasize his view of "peaceful uses" of nuclear technology, and it afforded him an opportunity to issue a challenge to NASA to develop missions that could exploit the device's potential. The SNAP-3 continued its marketing role, and was shown at several foreign capitals as part of the U.S.'s "Atoms for Peace" exhibits.

The first successful use of RTGs in space took place with the U.S. Navy's Transit satellite program. Also known as the NNS (Navy Navigation Satellite), the Transit system was used by the Navy to provide accurate location information to its ships. It was also used for general navigation by the Navy, as well as hydrographic and geodetic surveying, and was the first such system to be used operationally. The Johns Hopkins Applied Physics Laboratory (APL) developed the system, starting in 1957. Many of the technologies developed under the Transit program are now in use on the Global Positioning System

Several of the Transit developers had been considering the use of RPS since the beginning of the program. Although solar cells and batteries had powered the first six Transit satellites, there was concern that the battery hermetic seals would not meet the five-year mission requirement. Thus, APL accepted an offer from the AEC to include an auxiliary nuclear power source on the satellite. At that time, however, the radioisotope fuel of choice, Plutonium-238 (Pu-238), was unavailable due to AEC restrictions, and APL refused to use beta-decaying Strontium-90 because of the excessive weight associated with its necessary shielding. The AEC eventually acquiesced and agreed to provide the Pu-238 fuel. The SNAP-3 was converted from use of Po-210 to Pu-238, and acquired the new designation, SNAP-3B. The SNAP-3B RTGs on board these spacecraft supplemented solar cell arrays

and demonstrated operation of nuclear systems for space power applications.

Fig. 5. SNAP-3 presentation to President Eisenhower.

**4. Flight systems** 

**4.1 SNAP-3B** 

(GPS).

one heat exchanger, expanded through the turbine, passed through a gas cooler and pressurized by the compressor thus completing the cycle. A rotary alternator attached to the turbine shaft produces alternating current (AC) electrical power.
