**5. Dinamic conversion energy**

For dynamic systems the conversion mechanism consists on that the thermal energy is partially transformed into mechanical work, moving an alternator to produce electric power. Rankine, Brayton and Stirling systems use this conversion mechanism. Typical cycle diagram is shown in Fig. 5. The isotope heats an inert gas working fluid which is expanded through a turbine. The high-efficiency Brayton cycle is capable to recuperate part of the energy. The turbine discharge gas is cooled, first in the recuperator, then in the radiator. The resulting low pressure gas is passed though the compressor, compressed to the highest cycle pressure and heated at essentially constant pressure in the recuperator before being returned to the heat source. The recuperator recovers a significant amount of heat, which would otherwise be dissipated through the radiator resulting in a higher cycle efficiency (Abelson, 2004; Benett, 2006; Lange, 2008).

Fig. 5. Dynamic Isotope Power System Cycle

The MMRTG will generate 120 W of power at launch from a Pu-238 heat source assembly containing a stack of 8 Step 2 GPHS modules, which are described in section 3. The MMRTG operates at a normal output voltage of 28 V-dc. Both diameter and length of MMRTG are 64 cm diameter and 66 cm, respectively. The central heat source cavity is separated from the thermoelectric converter by a helium isolation liner. The helium generated by the Pu-238 is dumped to the environment by diffusion through an elastomeric gasket seal. The thermoelectric converter cavity can operate in both atmospheric environment or space

The thermocouples are connected in a series/parallel electrical circuit to improve the efficiency up 6.8%. Waste heat is radiated from the eight radial fins. These fins are made of aluminum alloys coated with a high-emissivity to disintegrate and release the GPHS modules in the case of reentry into the Earth's atmosphere. The MMRTG is both lighter and

For dynamic systems the conversion mechanism consists on that the thermal energy is partially transformed into mechanical work, moving an alternator to produce electric power. Rankine, Brayton and Stirling systems use this conversion mechanism. Typical cycle diagram is shown in Fig. 5. The isotope heats an inert gas working fluid which is expanded through a turbine. The high-efficiency Brayton cycle is capable to recuperate part of the energy. The turbine discharge gas is cooled, first in the recuperator, then in the radiator. The resulting low pressure gas is passed though the compressor, compressed to the highest cycle pressure and heated at essentially constant pressure in the recuperator before being returned to the heat source. The recuperator recovers a significant amount of heat, which would otherwise be dissipated through the radiator resulting in a higher cycle efficiency

vacuum (Ritz, 2004; Lange, 2008).

**5. Dinamic conversion energy** 

(Abelson, 2004; Benett, 2006; Lange, 2008).

Fig. 5. Dynamic Isotope Power System Cycle

smaller than RTG system.

High efficiency in the RPS would both reduce system mass and fuel requirement, decreasing the total cost. Normally, RPS are tested at the ground in a vacuum chambers for a long time (>1000 hours). This would demonstrate that the design work also in the space with high power conversion efficiency.

Since the Brayton cycle is useless for power generation under 0.5 kW, missions with lower power requirement need an auxiliary electrical power generator.

The Stirling power system is based on a kinematics engine driving a three phase alternator. The initial design only worked during six months. It would not be appropriate for long-term missions. using a free piston combined with a linear alternator would be a promising technology for space power applications. The Stirling system, at difference of Brayton type, might provide low power. Further, Stirling engines operating in reverse have already used in space to provide cryogenic cooling for imaging sensors. The device has reciprocating pistons and displacers as it shown in Fig. 6. The motion of the components depends on physical springs or gas and on the cycle pressure swing of the engine. Typically, the engine contains only two moving parts.

Fig. 6. Schematic diagram of the Stirling Isotope Power System

Two types of free pistons, linear alternator, Stirling machines are under development for mission in the range between 10 and 100 watts. The first type, under development by NASA. The power conversion efficiency for the Stirling producing 100 W is about 30%. The second type uses flexural spring to support the moving component to prevent friction and to provide enough axial springing for the free piston movement. The engine relies on the gap between the cylinder and the displacer to serve as regenerator for the system. As with

Radioisotope Power Systems for Space Applications 469

Several accidents can occur in a space missions. Typical phases for deep space exploration missions (interplanetary mission) consists on: phase 1, called as ascent, begins with litoff of the Space Shuttle vehicle from launch pad, and then continues until the Solid Rocket Boosters are jettisoned some time after; phase 2, Second stage. This phase includes the first burn of the Orbital Maneuvering System (OMS) engines. The Shuttle main engine cutoff is included in this phase; phase 3, on Orbit, starting with the first burn of the OMS (OMS-1) and ends when the payload are deployed form the Orbiter. The phase include the first and second burns of the OMS (OMS-1 and OMS-2) for following the correct orbit and circularization; phase 4, Payload deploy, when reach the Earth escape velocity; phase 5, Maneuvers. To make possible some outer missions, is needed Gravitational Assist Maneuver, to obtain an impulse on the Spacecraft using the rotation energy of the planet. Critical issue is an Earth Gravity Assist, because the SC come back to the Earth; and a possible reentry (phase 6), exclusively for missions which ends with an spacecraft on an

Various consequences could result from the accident environments that have been defined for the safety evaluation in the Final Safety Analysis Report (FSAR). In phase 1, the possible accidents resulting from Solid Rocket Booster (SRB) failures, either self induced or resulting from Range Safety destruct, can in certain instances lead to damaged GPHS modules with subsequent release of fuel due to: impact by SBR case fragments and subsequent impact agains ground surfaces or launch pad structures. In phase 2, vehicle breakup resulting from orbiter failures can result in reentry of the RTG and breakup of the GPHS modules on hard ground surfaces. In both phases 3 and 4, Shuttle failures can result in reentry of the SC (and RTGs) with subsequent breakup and release of the GPHS modules to impact on ground surfaces. In the case of the spacecraft should fail to reach escape velocity it would reenter into the Earth atmosphere. The heat of reentry would release the heat source from the generator and allow it to impact to the ground. The capsule would be exposed to reentry heating, Earth impact, and oxidation. If the heat shield were to fail, the unprotected capsule could fail in reentry and expose the bare fuel disks to the reentry and impact conditions (JPL, 1994; Richins, 2007). Additionally, in an Earth gravitational maneuver scenario, SC might reenter at very high velocity due to a spacecraft failure or a mission failure, such as

In regions on the space near Sun, NASA has historically used a few solar electric power systems such as solar panels. Several mission such as Mars Observer, the Viking Orbiters and Mariners missions were solar powered missions. For improving the systems efficiency,

For outer planet missions, NASA has used radioisotope thermoelectric generators for the Cassini spacecraft. High electrical power for mission science requirements in powering the instruments and communication systems makes the RTG systems better option than solar arrays. The low efficiency of the solar cells for distances beyond Jupiter is an important drawback. Further, the spacecraft must be as lighter as possible. The size of the theoretical arrays of solar panels to obtain the power required for all sciences systems would be very

As regards on the solar cell technology, the actual production efficiencies of advanced solar cells have historically lower than research findings. The high-efficiency ESA solar cell

the Mars Global Surveyor used solar power with gallium-arsenide cells (JPL, 1994).

puncture of the SC propellant tank by a micrometeoroid (space debris).

**7. RTGs versus solar arrays** 

large, increasing the spacecraft mass.

Earth reentry.

the Brayton cycle, heat regeneration is essential to achieve high efficiencies in Stirling engines. The design of a 10-W system has been tested, using fossile fuel combustion with a 20 % efficiency.

One of the problem of these systems is the attitude dynamic effects over the spacecraft. Both Brayton and Stirling systems have accumulated many time of testing. However, more tests would be required for outer planet missions that are expected to take more than 5 years.

## **6. RPS safety and accident evaluation**

The Department of Energy (DOE) has worked to improve the safety of the RPS under all accident conditions, including accidents occurring near the launch pad and for orbital reentry accidents. The Pu-238 fuel for was changed from a metal to a more stable pressed oxide (*PuO2*).

On April 21, 1964 the Transit-5-BN-3 mission was aborted because of a launch vehicle failure resulting in burn-up of the RTG during reentry, in keeping with the RTG design at the time. Some amount of the plutonium fuel was dropped in the upper atmosphere. The RTG design was changed to provide for survival of the fuel modules during orbital reentry.

A second accident occurred when the Nimbus B-1 was launched on May 18, 1968. It was aborted shortly after launch by a range destruction safety. The heat sources were recovered intact in about 90 meters under water in the California coast without release of plutonium. The fuel capsules were reworked and the fuel was used in a later mission (Abelson, 2004; Furlog, 1999).

The third incident occurred in April 1970, when the Apollo 13 mission to the moon was aborted following an oxygen tank explosion in the spacecraft service module. Upon return to Earth, the Apollo 13 lunar excursion module with a SNAP-27 RTG on board reentered the atmosphere and broke up above the south Pacific Ocean. The heat source module fell into the ocean. Atmospheric and oceanic monitoring showed no evidence of release of nuclear fuel.

The ceramic form covering plutonium-238 dioxide is heat-resistant and limits the rate of vaporization in fire or reentry conditions. The material also has low solubility in water. This material does not disperse though the environment.

More than 35 years have been researched in the engineering concepts and testing of RPS systems. Multiple layers of protective materials, including iridium capsules (or platiniumshodium capsules for RHUs) and high strength, heat-resistant graphite blocks are used to protect the radionuclide and prevent its release. Iridium is a strong, corrosion-resistant metal that is chemically compatible with plutonium dioxide. In addition, graphite is used because it is lightweight and highly heat-resistant. Several test for potential accident scenarios to know how RTG responses has been developed. Results of the failure mechanisms provide the basis for the determination of the source terms which are the characterization of plutonium releases including their quantity, location and particle size distribution. Recent large fragment tests in the GPHS safety test program have demonstrated in Solid Rocket Boosters (SRB) accident case, fragments impacting the full RTG system will not breach the fueled clads at velocities up to 0.12 km/s.

The multi-layer containment concept employed for the systems is designed to contain the radioisotope but even if the containment is breached, the ceramic pellet has been designed to limit dispersal of the material into the environment.

the Brayton cycle, heat regeneration is essential to achieve high efficiencies in Stirling engines. The design of a 10-W system has been tested, using fossile fuel combustion with a

One of the problem of these systems is the attitude dynamic effects over the spacecraft. Both Brayton and Stirling systems have accumulated many time of testing. However, more tests would be required for outer planet missions that are expected to take more than 5 years.

The Department of Energy (DOE) has worked to improve the safety of the RPS under all accident conditions, including accidents occurring near the launch pad and for orbital reentry accidents. The Pu-238 fuel for was changed from a metal to a more stable pressed

On April 21, 1964 the Transit-5-BN-3 mission was aborted because of a launch vehicle failure resulting in burn-up of the RTG during reentry, in keeping with the RTG design at the time. Some amount of the plutonium fuel was dropped in the upper atmosphere. The RTG design was changed to provide for survival of the fuel modules during orbital

A second accident occurred when the Nimbus B-1 was launched on May 18, 1968. It was aborted shortly after launch by a range destruction safety. The heat sources were recovered intact in about 90 meters under water in the California coast without release of plutonium. The fuel capsules were reworked and the fuel was used in a later mission (Abelson, 2004;

The third incident occurred in April 1970, when the Apollo 13 mission to the moon was aborted following an oxygen tank explosion in the spacecraft service module. Upon return to Earth, the Apollo 13 lunar excursion module with a SNAP-27 RTG on board reentered the atmosphere and broke up above the south Pacific Ocean. The heat source module fell into the ocean. Atmospheric and oceanic monitoring showed no evidence of release of nuclear

The ceramic form covering plutonium-238 dioxide is heat-resistant and limits the rate of vaporization in fire or reentry conditions. The material also has low solubility in water. This

More than 35 years have been researched in the engineering concepts and testing of RPS systems. Multiple layers of protective materials, including iridium capsules (or platiniumshodium capsules for RHUs) and high strength, heat-resistant graphite blocks are used to protect the radionuclide and prevent its release. Iridium is a strong, corrosion-resistant metal that is chemically compatible with plutonium dioxide. In addition, graphite is used because it is lightweight and highly heat-resistant. Several test for potential accident scenarios to know how RTG responses has been developed. Results of the failure mechanisms provide the basis for the determination of the source terms which are the characterization of plutonium releases including their quantity, location and particle size distribution. Recent large fragment tests in the GPHS safety test program have demonstrated in Solid Rocket Boosters (SRB) accident case, fragments impacting the full

The multi-layer containment concept employed for the systems is designed to contain the radioisotope but even if the containment is breached, the ceramic pellet has been designed

RTG system will not breach the fueled clads at velocities up to 0.12 km/s.

20 % efficiency.

oxide (*PuO2*).

reentry.

Furlog, 1999).

fuel.

**6. RPS safety and accident evaluation** 

material does not disperse though the environment.

to limit dispersal of the material into the environment.

Several accidents can occur in a space missions. Typical phases for deep space exploration missions (interplanetary mission) consists on: phase 1, called as ascent, begins with litoff of the Space Shuttle vehicle from launch pad, and then continues until the Solid Rocket Boosters are jettisoned some time after; phase 2, Second stage. This phase includes the first burn of the Orbital Maneuvering System (OMS) engines. The Shuttle main engine cutoff is included in this phase; phase 3, on Orbit, starting with the first burn of the OMS (OMS-1) and ends when the payload are deployed form the Orbiter. The phase include the first and second burns of the OMS (OMS-1 and OMS-2) for following the correct orbit and circularization; phase 4, Payload deploy, when reach the Earth escape velocity; phase 5, Maneuvers. To make possible some outer missions, is needed Gravitational Assist Maneuver, to obtain an impulse on the Spacecraft using the rotation energy of the planet. Critical issue is an Earth Gravity Assist, because the SC come back to the Earth; and a possible reentry (phase 6), exclusively for missions which ends with an spacecraft on an Earth reentry.

Various consequences could result from the accident environments that have been defined for the safety evaluation in the Final Safety Analysis Report (FSAR). In phase 1, the possible accidents resulting from Solid Rocket Booster (SRB) failures, either self induced or resulting from Range Safety destruct, can in certain instances lead to damaged GPHS modules with subsequent release of fuel due to: impact by SBR case fragments and subsequent impact agains ground surfaces or launch pad structures. In phase 2, vehicle breakup resulting from orbiter failures can result in reentry of the RTG and breakup of the GPHS modules on hard ground surfaces. In both phases 3 and 4, Shuttle failures can result in reentry of the SC (and RTGs) with subsequent breakup and release of the GPHS modules to impact on ground surfaces. In the case of the spacecraft should fail to reach escape velocity it would reenter into the Earth atmosphere. The heat of reentry would release the heat source from the generator and allow it to impact to the ground. The capsule would be exposed to reentry heating, Earth impact, and oxidation. If the heat shield were to fail, the unprotected capsule could fail in reentry and expose the bare fuel disks to the reentry and impact conditions (JPL, 1994; Richins, 2007). Additionally, in an Earth gravitational maneuver scenario, SC might reenter at very high velocity due to a spacecraft failure or a mission failure, such as puncture of the SC propellant tank by a micrometeoroid (space debris).

#### **7. RTGs versus solar arrays**

In regions on the space near Sun, NASA has historically used a few solar electric power systems such as solar panels. Several mission such as Mars Observer, the Viking Orbiters and Mariners missions were solar powered missions. For improving the systems efficiency, the Mars Global Surveyor used solar power with gallium-arsenide cells (JPL, 1994).

For outer planet missions, NASA has used radioisotope thermoelectric generators for the Cassini spacecraft. High electrical power for mission science requirements in powering the instruments and communication systems makes the RTG systems better option than solar arrays. The low efficiency of the solar cells for distances beyond Jupiter is an important drawback. Further, the spacecraft must be as lighter as possible. The size of the theoretical arrays of solar panels to obtain the power required for all sciences systems would be very large, increasing the spacecraft mass.

As regards on the solar cell technology, the actual production efficiencies of advanced solar cells have historically lower than research findings. The high-efficiency ESA solar cell

Radioisotope Power Systems for Space Applications 471

events. For using less plutonium than required, RPS efficiency must improve. Using lowconductivity materials and high thermoelectric rating, *Z*, RPS efficiency would improve. A high-efficiency Stirling-type system would give an apparent mass/power benefit, as well as using less plutonium for a similar power output. If we want to continue using RPS with Plutonium-238 as fuelling, we have to develop more high-efficiency systems, avoiding vibrations on the attitude on the spacecraft, as itself occurs with dynamic-conversion system. The current RPS power conversion efficiency is not too high. It is also required

Tethers might be used as alternative to solve the severe power generation problem. An electrodynamic tether, which is a very long wire capable to generate the suggested power, might radiate waves to satisfy communication requirements itself (Sanchez-Torres et al., 2010). The large electromotive force produced by the tether moving in some plasma ambient near the planet generate induced current and then electric power (Sanmartin et al., 1993). Tethers might be very useful for generating electric power both in Low Earth Orbit (high plasma density and moderate magnetic field) and in Jovian conditions (low plasma

This work was supported by the Ministry of Science and Innovation of Spain (BES-2009-

Abelson, R. et al. (2004). Enabling Exploration with Small Radioisotope Power Systems*, JPL* 

Bennett, G. et al. (June 2006). Mission of Daring: The General-Purpose Heat Source

Brown, C. (2001). Elements of Spacecraft Design, *AIAA Education Series*, Reston, Virginia,

Furlog, R. & Wahlquist, E. (1999). U.S. Space Missions Using Radioisotope Power Systems,

Griffin, M. & French, J. (2004). Space Vehicle Design, Second Edition, *AIAA Education Series*,

Hastings, D. & Garrett, H. (2004). Spacecraft-Environment Interactions, *Cambridge University* 

JPL (July 1994). Cassini Program Environmental Impact Statement Supporting Study.

Lange, R. & Carrol, W. (2008). Review of Recent Advances of Radioisotope Power Systems*,* 

*Energy Conversion and Management*, 49, pp. 393-401, ISSN: 0196-8904

Volume 2: Alternate Mission and Power Study, *JPL Publication No. D-11777. Cassini* 

http://www.ans.org/pubs/magazines/nn/pdfs/1999-4-2.pdf

http://saturn.jpl.nasa.gov/spacecraft/safety/eisss2.pdf

*Engineering Conference and Exhibit*, San Diego, California, Avalible from

Radioisotope Thermoelectric Generator, *4th International Energy Conversion* 

lower cost power systems.

density and high magnetic field).

*Pub 04-10,* Avalible from

ISBN 1-56347-524-3

http://hdl.handle.net/2014/40856

http://www.fas.org/nuke/space/gphs.pdf

*Nuclear News*, pp. 26-34, Avalible from

Reston, Virginia, ISBN 1-56347-539-1

*Press,* Cambridge, UK, ISBN 0 521 60756 6

*Document No. 699-070-2*, Avalible from

**9. Acknowledgment** 

013319 FPI Grant).

**10. References** 

devices are relative thick and heavy compared to the usual solar cells. Further, these advanced cells would be radiation sensitive. Solar-powered Juno mission will be launched in August on 2011, to study Jupiter. The spacecraft avoid the intense radiation belts using a innovative polar orbit, obtaining a great visibility for both the solar light arriving from the sun and communications.

Large solar arrays would severely impact the design, mass and operation of the spacecraft. This structure would have to be deployable, i.e. it could fit inside the rocket payload, and then unfold once the SC reached the outer planet. The mechanical components to fold and unfold the arrays would increase notably the size and mass on the SC. The long solar arrays would also severely complicate the stability on the trajectory and the attitude for scientific observations and data transmission to the Earth. Large spacecraft size, indeed, would make the maneuvers slower, which is critical for scientific data collection.

The electrical power requirements of the spacecraft for science instruments and telecommunications, lunch mass, and mission lifetime are all of critical concern in choosing the electrical power source.
