Lee S. Mason

Glenn Research Center, Cleveland, Ohio

David I. Poston

Los Alamos National Laboratory, Los Alamos, New Mexico

A Summary of NASA Architecture Studies Utilizing

Fission Surface Power Technology

NASA/TM—2011-216819

April 2011

AIAA–2010–6599
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question via the Internet to help@

sti.nasa.gov

• Fax your question to the NASA STI Help Desk 

at 443–757–5803

• Telephone the NASA STI Help Desk at

 443–757–5802

• Write to:

 NASA Center for AeroSpace Information (CASI)

 7115 Standard Drive

 Hanover, MD 21076–1320
Lee S. Mason

Glenn Research Center, Cleveland, Ohio

David I. Poston

Los Alamos National Laboratory, Los Alamos, New Mexico

A Summary of NASA Architecture Studies Utilizing

Fission Surface Power Technology

NASA/TM—2011-216819

April 2011

AIAA–2010–6599

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Eigth International Energy Conversion Engineering Conference (IECEC)

sponsored by the American Institute of Aeronautics and Astronautics

Nashville, Tennessee, July 25–28, 2010
Acknowledgments

The authors wish to acknowledge the contributions of the other Fission Surface Power (FSP) team members including John Warren 

(NASA Headquarters), Scott Harlow (DOE Headquarters), Mike Houts (MSFC), Jim Werner (INL), Lou Qualls (ORNL), and 

Ross Radel (SNL) as well as the other staff members at these locations. The FSP Project is part of the Exploration Technology 

Development Program (ETDP) managed by Frank Peri at Langley Research Center. ETDP is part of the Exploration Systems 

Mission Directorate at NASA Headquarters where Chris Moore serves as the Program Executive. The NASA architecture studies 

involve a large network of individuals across the NASA centers, which are currently coordinated by the Johnson Space Center 

Lunar Surface Systems Offi ce managed by Chris Culbert. The International Architecture Working Group (IAWG) Power Function 

Team is headed by Marc Haese from European Space Agency (ESA).

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2011-216819 1 

A Summary of NASA Architecture Studies Utilizing 

Fission Surface Power Technology 

Lee S. Mason 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

David I. Poston 

Los Alamos National Laboratory 

Los Alamos, New Mexico 87545 

Abstract 

Beginning with the Exploration Systems Architecture Study in 2005, NASA has conducted various 

mission architecture studies to evaluate implementation options for the U.S. Space Policy. Several of the 

studies examined the use of Fission Surface Power (FSP) systems for human missions to the lunar and 

Martian surface. This paper summarizes the FSP concepts developed under four different NASA-sponsored 

architecture studies: Lunar Architecture Team, Mars Architecture Team, Lunar Surface Systems/ 

Constellation Architecture Team, and International Architecture Working Group-Power Function Team. 

Introduction 

Under the NASA Exploration Technology Development Program (ETDP) and in partnership with the 

Department of Energy (DOE), NASA has an ongoing project to develop Fission Surface Power (FSP) 

technology. The primary goals of the project are 1) develop FSP concepts that meet expected surface power 

requirements at reasonable cost with added benefits over other options, 2) establish a hardware-based 

technical foundation for FSP design concepts and reduce overall development risk, 3) reduce the cost 

uncertainties for FSP and establish greater credibility for flight system cost estimates, and 4) generate the 

key products to allow NASA decision makers to consider FSP as a preferred option for flight development. 

The FSP project was initiated in 2006 as the NASA Prometheus Program and the Jupiter Icy Moons 

Orbiter (JIMO) mission were phased out. As a first step, NASA Headquarters commissioned the 

Affordable Fission Surface Power System Study (AFSPSS) to evaluate the potential for an affordable 

FSP development approach. With a cost-effective FSP strategy identified, the FSP team evaluated design 

options and selected a preliminary reference concept to guide technology development. Since then, the 

FSP preliminary reference concept has served as a point-of-departure for several NASA mission 

architecture studies examining the use of nuclear power and has provided the foundation for a series of 

“Pathfinder” hardware tests. The long-term technology goal is a Technology Demonstration Unit (TDU) 

integrated system test using full-scale components and a non-nuclear reactor simulator. 

The FSP team consists of Glenn Research Center (GRC), Marshall Space Flight Center (MSFC), and the 

DOE National Laboratories at Los Alamos (LANL), Idaho (INL), Oak Ridge (ORNL), and Sandia (SNL). The 

project is organized into two main elements: Concept Definition and Risk Reduction. Under Concept 

Definition, the team performs trade studies, develops analytical tools, and formulates system concepts. Under 

Risk Reduction the team develops hardware prototypes and conducts laboratory-based testing. 
NASA/TM—2011-216819 2 

Project Context 

One of the major challenges to the implementation of space fission power systems is development 

cost. In April 2006, NASA and DOE initiated the AFSPSS to determine the design features and expected 

costs of a representative FSP system. A government study team with members from several NASA field 

centers and DOE laboratories evaluated technology options and design variables and selected a reference 

concept based on affordability and risk. A low-risk approach was selected over other options that could 

offer higher system performance and/or lower mass. The team also defined a credible development 

schedule and generated a detailed Work Breakdown Structure (WBS)-based cost estimate. The results 

indicated that the initial FSP system could be developed, flight-qualified, and delivered to the lunar 

surface by 2020 for approximately $1.4B (2007 dollars) with follow-on systems costing about $215M 

each (Ref. 1). 

The “affordable” design approach was considered representative of a number of potential system 

concepts. In order to determine a FSP reference concept, the team generated a comprehensive list of system 

design options and conducted screening studies that led to six plausible concepts for further study. All of the 

plausible concepts presumed the use of a low-temperature (<900 K) reactor heat source with conventional 

materials as a path toward achieving an affordable solution. The plausible concepts included a liquid-metalcooled reactor with Stirling, Brayton, thermoelectric, or organic-Rankine power conversion, a gas-cooled 

reactor with Brayton power conversion, and a heat pipe cooled reactor with Stirling power conversion. The 

concepts were evaluated for performance and relative cost against a common set of mission requirements 

and development constraints derived from the earlier affordable study. In 2008, a management review panel 

led by NASA Headquarters selected the liquid-metal reactor with Stirling power conversion as the FSP 

preliminary reference concept and recommended Brayton as a backup conversion option if unforeseen 

difficulties arise with the Stirling technology development efforts (Ref. 2). 

The resulting preliminary reference concept includes a liquid-metal-cooled, fast-spectrum reactor with 

Stirling power conversion and water-based heat rejection (Ref. 3). The reactor uses uranium oxide (UO2) 

fuel pins in a hexagonal core with an external radial reflector and control drums. Heat is transferred to the 

Stirling power convertors by a pumped sodium-potassium (NaK) reactor coolant loop. The core structure 

and coolant piping are constructed of stainless steel to reduce cost and development risk. The radial 

reflector is beryllium in a stainless-steel shell. The control drums are beryllium and boron carbide (B4C), 

also enclosed in stainless steel. The reactor is located at the bottom of an approximate 2-m-deep 

excavation. The lunar regolith limits radiation from the reactor to less than 5 rem/year at a 100 m radius. 

The Stirling convertors generate single-phase alternating current (AC) electric power that is converted to 

direct current (DC) for user loads. Stirling waste heat is removed by a pumped water coolant loop coupled 

to a series of two-sided, vertical radiator panels. The radiator panels are composed of titanium-water heat 

pipes in a composite facesheet sandwich. The FSP concept is designed to produce a net power of 40 kWe 

with a full-power service life of at least 8 yr. This same technology could be used for missions at 

essentially any location (equator to poles) on the lunar or Mars surface. 

Derived Requirements 

Table 1 presents a summary of the top-level requirements developed for the FSP system. The 

requirements are termed “derived” because they were predominantly defined by the FSP team in response 

to suggestions by NASA Headquarters and the various NASA architecture study teams. The FSP safetyrelated requirements were generated by the FSP team based on previous space fission system 

development projects, such as SP–100 and JIMO. These requirements will undoubtedly be reviewed (and 

perhaps expanded) by independent design experts once FSP reaches flight development status. For now, 

they provide a reasonable starting point to guide FSP concept definition and technology development. 

 
NASA/TM—2011-216819 3 

TABLE 1.—FSP-DERIVED REQUIREMENTS 

Requirement Rationale 

1. The FSPS shall be designed to produce no less than 

40 kWe net power output (after accounting for all power 

losses and auxiliary loads). 

Provides sufficient power for extended-stay crew habitation, ISRU 

production facilities, rover recharging, and science equipment, 

including margin. 

2. The FSPS shall be designed for use at any location on 

the lunar surface. Radiator sizing shall be based on 

worst-case surface temperatures and Sun angles. 

Provides maximum flexibility in locating the lunar outpost. 

(The preferred lunar outpost location has not been determined.) 

3. The FSPS shall be designed to operate for no less than 

8 yr at full power. 

Provides maximum service life without introducing excessive risk in 

FSP development and qualification. 

4. The FSPS shall be flight ready for an initial launch and 

deployment no later than 2022. 

Assures FSPS availability for initial outpost deployment based on 

current lunar emplacement schedules. 

5. The FSPS shall be designed to produce no less than 50% 

power output after the first credible component failure. 

Assures FSPS power availability to meet essential crew power 

requirements following a component failure. 

6. The FSPS shall be recoverable from all credible 

operational upsets and transients without adverse safety 

consequences to the crew or outpost. 

Assures FSPS power availability following an off-nominal event 

and a return to safe FSP operation. 

7. The radiation from the FSPS shall be less than 

5 rem/yr to an unshielded crew member located at the 

outpost. 

Provides a guideline for FSPS shield design that corresponds to 

10% of the astronaut annual dose limit. (The allowable crew dose 

from the FSPS has not been determined.) 

8. The reactor shall remain subcritical during all planned 

and credible unplanned mission events prior to FSPS 

startup. 

Assures that the FSPS does not present a radiological safety hazard 

before initial startup is commanded. 

9. At its end-of-life, the reactor shall be decommissioned in 

a safe shutdown condition. 

Assures that the FSP does not present a human safety hazard after 

final shutdown is commanded. 

10. The FSPS mass shall be minimized and no greater than 

the current cargo lander down-mass capability of 

14 000 kg. 

Permits the FSPS to be delivered as a fully integrated package with 

available cargo mass to accommodate other payloads. 

11. The FSPS design shall be extensible to the Mars surface. 

All materials and design strategies shall be compatible 

with the Martian environment. 

Provides maximum return on FSPS technology investment by 

designating its applicability for both the Moon and Mars. 

12. The FSPS shall be designed for robotic deployment 

using teleoperation. 

Permits the FSPS to be installed at the lunar outpost without local 

human assistance, but does not preclude it. 

The key requirements that drive FSP system design are power level and service life. The 40 kWe 

power output is consistent with numerous studies that have estimated power requirements for the initial 

phase of a human lunar outpost dating back to the 1990’s Space Exploration Initiative and before. That 

power level is also well suited for an initial space reactor because it is large enough to demonstrate the 

mass effectiveness of nuclear fission, but not too large to over-complicate the design and development 

process. In actuality, the fission technology developed for the 40-kWe design is readily scalable between 

10 and 100 kWe. Below 10 kWe, the mass and cost advantages are not as compelling. Above 100 kWe, 

the reactor and power conversion technologies selected for FSP may need to be reevaluated. 

The 8-yr service life also represents a reasonable balance of performance and risk. It is long enough to 

accommodate most estimates for lunar and Mars surface mission duration. For longer missions, it would be 

prudent to utilize multiple FSP units and stagger their delivery to provide overlap. The 8-yr design life is 

also well within current technology projections for low-temperature liquid-metal reactors and dynamic 

power conversion. In addition, notional FSP development schedules indicate that sufficient qualification 

testing can be performed to demonstrate 8-yr life while still meeting the proposed launch date. 

System mass is another requirement that could influence FSP design. The current derived requirement 

is that the FSP system mass be less than the payload capacity of the lander. The current cargo lander 

concept is projected to deliver approximately 14 000 kg to the lunar surface. The 40-kWe FSP system can 

easily be accommodated within this mass constraint, and most estimates show the system to be less than 

one-half of the lander cargo capacity. The generous lander payload allocation eliminates system mass as a 

major FSP design driver and allows the system to utilize low-risk technology to minimize development 

cost and increase system reliability. Nevertheless, the FSP system design incorporates various mass 

saving features in order to maximize the mass available for other payloads. This also assures that the 

concept is relevant for future applications that may be more mass constrained. 
NASA/TM—2011-216819 4 

Design Summary 

The preliminary reference concept layout is shown in Figure 1. The reactor core is located at the 

bottom of an approximate 2-m-deep excavation with an upper plug shield to protect the equipment above 

from direct radiation. The NaK pumps, Stirling convertors, and water pumps are mounted on a 5-m-tall 

truss structure that attaches to the top face of the shield. Two symmetric radiator wings are deployed via a 

scissor mechanism from the truss. Each radiator wing is approximately 4 m tall by 16 m long and is 

suspended 1 m above the lunar surface. In its stowed configuration, the FSP system is approximately 

3 by 3 by 7 m tall. 

The buried configuration was selected for the preliminary reference concept because it minimizes the 

mass of radiation shielding that must be delivered from Earth. It also simplifies the Power Management 

and Distribution (PMAD) because the buried reactor can be located relatively close to the outpost to 

shorten transmission cable length. There are numerous other FSP installation options that could be 

developed depending on mission needs. The basic technology building blocks of the liquid-metal-cooled 

reactor, Stirling power conversion, and water-based heat rejection would be essentially the same. The 

decision on FSP configuration can easily be deferred until the flight program since most of the design 

challenges related to the configuration are engineering based rather than technology based. 

The preliminary reference concept schematic is shown in Figure 2. The use of redundant components 

and parallel fluid loops allows the system to produce partial power in the event of unexpected failures. The 

schematic shows the system energy balance and the anticipated temperatures, pressures, and flow rates at 

some of the key interfaces. The reactor (Rx) produces 186 kWt with a peak fuel pin clad temperature of 

860 K. It delivers heated NaK at 850 K to a pair of intermediate heat exchangers (IHX) using two fully 

redundant electromagnetic primary pumps (PP). The IHX is a NaK-to-NaK heat exchanger that provides a 

buffer between the primary NaK and the Stirling convertors, and a means to adjust the NaK flow rate and 

resulting temperature drop across the Stirling convertors separately from the reactor flow and temperature 

drop. Each intermediate NaK loop services two Stirling convertors at a supply temperature of 824 K. The 

effective Stirling hot-end cycle temperature is 778 K. The secondary NaK loops include an intermediate 

electromagnetic pump (IP) of similar design to the primary NaK pump. 

Figure 1.—FSP concept layout. 
NASA/TM—2011-216819 5 

Figure 2.—FSP concept schematic. 

TABLE 2.—FSP MASS SUMMARY 

1.0 Fission surface power system (kg) 5820 

1.1 Reactor module (kg) 1440 

1.2 Power conversion module (kg) 411 

1.3 Heat rejection module (kg) 767 

1.4 Power management and distribution module (kg) 1071 

1.5 Radiation shield module (kg) 2080 

1.6 Integration structure (kg) 51 

Each Stirling convertor (Stir) is composed of two axially opposed Stirling heat engines and two linear 

alternators. Power conversion thermal-to-electric efficiency is estimated at 26 percent. The alternators 

deliver 6 kWe each at 400 Vac rms and 60 Hz to the PMAD. A Local Power Controller (LPC), located 

approximately 100 m from the reactor, converts the 400 Vac to 120 Vdc for distribution to the Electrical 

Load Interface (ELI). The 48 kWe gross Stirling output power provides sufficient capacity to account for 

electrical losses (~3 kWe) and system parasitic loads (~5 kWe) and still delivers 40 kWe net for user 

loads. A Parasitic Load Radiator (PLR) dissipates electric power that is not required by the user loads and 

allows the system to be operated at constant power thus eliminating the complexity of thermal system 

load following. The ELI serves as the primary power bus and system interface for commands and 

telemetry. A 5-kWe photovoltaic (PV) array and a 30-kW-hr battery are included with the FSP PMAD for 

startup and backup power. 

The heat rejection subsystem is composed of four water heat transport loops and two radiator (Rad) 

wings (two loops per wing). The radiator wings receive heated water at 420 K from the Stirling convertors 

and return the water at 390 K using a mechanical radiator pump (RP), while rejecting approximately 35 kWt 

per loop. The resulting Stirling cold-end cycle temperature is 425 K. The total heat load is approximately 

140 kWt and the total two-sided FSP system radiator area is 185 m2

 assuming a 250 K effective sink 

temperature and 10 percent area margin. Each radiator wing includes 10 subpanels, each measuring 

approximately 2.7 m wide by 1.7 m tall. The preliminary reference concept mass summary for the buried 

reactor configuration is shown in Table 2. The total system mass without margin is 5820 kg. 
NASA/TM—2011-216819 6 

Lunar and Mars Architecture Studies 

Beginning with the Exploration Systems Architecture Study (ESAS) in 2005, NASA has conducted 

various mission architecture studies to evaluate implementation options for the U.S. Space Policy 

(formerly the Vision for Space Exploration). Several of the studies examined the use of fission power 

systems for human missions to the lunar and Martian surface. The FSP team contributed by supplying 

FSP design characteristics, developing mission-compatible configuration options, and defining a conceptof-operations consistent with the mission objectives. 

Lunar Architecture Team 

In 2007, the second phase of the Lunar Architecture Team (LAT2) developed an FSP-based 

architecture known as Option 6 for a polar lunar outpost at the Shackleton Crater site. The nuclear-based 

architecture was proposed to accelerate outpost buildup, achieve earlier 180-day crewed surface missions, 

and maximize the total number of crew days on the surface over a 10 year lunar campaign. Option 6 uses 

the buried reactor concept delivered on a cargo lander and installed by a combination of robots and crew, 

as shown in Figure 3. An earlier lander delivers a small solar array and battery to supply initial power, the 

FSP PMAD equipment, and a bladed-rover that prepares the site for the reactor. Once installed, the FSP 

system provides a robust power capability of 40 kWe resulting in substantial power margin for early 

outpost buildup and operations. It also provides capacity for power increases associated with the initial 

surface elements and the potential for expanded science and resource utilization. System trades comparing 

Option 6 to similar LAT2 architectures with solar PV arrays and regenerative fuel cells (RFCs) showed 

the FSP-based architecture to offer significantly more power with less power system mass and 

comparable cost despite the favorable conditions for solar power at Shackleton. 

A key question raised about the FSP installation was the feasibility of excavating the reactor hole. 

Independent studies were conducted by the in situ resource utilization (ISRU) team during LAT2 to evaluate 

methods for excavating a 2 m-deep hole. The ISRU study evaluated various digging methods and developed 

analytical models to predict the mass and power requirements for the machinery. It was determined that the 

process could use the same semi-automated regolith-moving equipment planned for the ISRU oxygen 

production plant. The recommended approach was to prepare an oversized hole with a ramp that could 

accommodate ingress/egress of a bladed rover. Preliminary estimates indicated the need to move about 

24 m3

 of regolith, including the final backfilling of the ramp, over a time period of 41 to 50 days. 

Figure 3.—FSP system for LAT2 Option 6. 
NASA/TM—2011-216819 7 

Mars Architecture Team 

During the same time period, the Mars Architecture Team (MAT) was reviewing power system 

options for a crewed mission to Mars. The basic architecture was derived from previous Mars mission 

concepts in which an initial cargo lander delivers a power system and ISRU plant to locally produce the 

return propellant before the crew ever leaves Earth. A nuclear system allows the propellant production to 

be completed faster and more efficiently through continuous day/night operations. The power 

requirements for the nuclear power option were about 30 kWe during the pre-crew deployment phase and 

about 20 kWe after the crew arrives. The 30-kWe power level was similar enough to the reference 

40 kWe lunar concept that no power system design changes were required. The MAT-based FSP concept 

assumed the reactor on a mobile cart with integral shielding that is robotically deployed from the lander, 

as shown in Figure 4. The above-grade reactor configuration was chosen for this application because the 

MAT wanted to avoid digging operations. FSP was selected as the baseline power system for MAT based 

on advantages in system mass, operational flexibility, and environmental robustness as compared to solar 

power systems with energy storage (Ref. 4). 

Lunar Surface Systems/Constellation Architecture Team 

In 2008, Lunar Surface Systems (LSS) and the Constellation Architecture Team developed an FSPbased architecture known as Scenario 5. Two basic FSP options were investigated including the typical 

off-loaded and buried system and a new concept where the FSP system remained on the lander, as shown 

in Figure 5. In either case, the FSP system was to be delivered on the first cargo lander to provide a 

power-rich environment for early outpost buildup. Both systems also assumed a central power 

distribution node at the outpost. This provided an easy-access power bus for outpost loads such as 

habitats, ISRU equipment, rover recharging, and science experiments. It also placed the FSP system’s 

power and control electronics at a location that was readily accessible should maintenance be required. A 

small solar array (5 kWe) and battery (30 kW-hr) was included with the FSP PMAD for startup and 

emergency backup. A follow-on architecture evaluated by LSS, referred to as Scenario 12, included an 

FSP system that was delivered later in the lunar campaign using a similar design approach. 

 

Figure 4.—Mars FSP concept. 

 

Figure 5.—Lander-integrated FSP system for LSS 

Scenario 5. 
NASA/TM—2011-216819 8 

The FSP team did an extensive evaluation of radiation shielding options to support the Scenario 5 

architecture definition (Ref. 5). Figure 6 shows graphical representations of the Monte Carlo N-Particle 

(MCNP) transport code models developed for the four shielding approaches that were examined including 

A) FSP system off-loaded and reactor buried, B) FSP system off-loaded and placed on surface with 

surrounding regolith berm, C) FSP stays on the lander as delivered from Earth, and D) FSP system stays 

on the lander with regolith shielding augmentation. 

All options assumed a 3 mrem/hr (26.3 rem/yr) reactor dose rate to an unshielded astronaut at a 

specified separation distance. This is higher than the 5 rem/yr dose mentioned previously to account for 

more realistic crew routines and schedules. Crew length-of-stay is expected to be no greater than 180 days 

with the majority of time spent in shielded habitats and rovers away from the hypothetical reactor 

boundary. Given reasonable assumptions for crew operations, the total radiation to a crew member from 

the reactor based on the 3 mrem/hr dose rate at the specified distance is expected to total much less than 

5 rem per year of duty. The actual allowable astronaut radiation dose is not defined yet and will depend 

on many factors including natural radiation levels, proximity to nuclear sources (such as FSP), crew 

shielding, length of mission, and Extra-Vehicular Activity (EVA) duty cycle. The FSP-related dose is 

expected to be a small percentage of the total received by crew members during their lunar stay. The FSP 

shielding must also protect its own components located on the truss above the shield. The assumed dose 

limits for truss-mounted equipment above the shield were 5 Mrad (gamma) and 2.5×1014 neutrons/cm2

. 

Recent testing of Stirling convertor components and materials at SNL, ORNL, and Texas A&M 

University suggest that these dose limits could be increased. In most cases, the FSP equipment was the 

limiting factor in determining the required FSP shield mass. 

Figure 6.—Scenario 5 shielding options. 

 
NASA/TM—2011-216819 9 

The buried reactor case (A) resulted in a 2080 kg delivered shield, predominantly B4C, with the 

reactor core buried to a 2.3 m depth. This approach offered the shortest separation distance among the 

options at 100 m. It also offers the lowest delivered shield mass, and the potential for shield mass 

reduction by using water rather than B4C. The berm shield case (B) assumed a 2-m-tall regolith berm 

surrounding the reactor. It resulted in a 2660-kg delivered shield using water and depleted uranium (DU) 

and a 200 m separation distance. The landed shield cases assumed that the reactor remained in the central 

lander cavity between the propellant tanks at a height of approximately 4 m above the lunar surface. The 

“as-delivered” lander case (C) required a shaped shield of water and DU that was thicker in the direction 

of the outpost. It was still the heaviest delivered shield at 2980 kg and required a separation distance of 

1000 m. The regolith-augmented lander case (D) resulted in a 2250 kg delivered shield of water and DU 

supplemented with 0.8-m-thick regolith-filled annulus surrounding the water vessel and a 400 m 

separation distance. 

The separation distances for the various shield options were determined in conjunction with power 

transmission cable mass estimates. Generally, there is an optimum distance that balances shield mass and 

cable mass. Table 3 provides a summary of the power transmission assumptions and resulting cable 

masses for the four cases. Larger distances require more complex power transmission approaches. In all 

cases, the power distribution node was assumed to be located at the specified separation distance and a 

25 percent margin was added for cable length. The cable bundle was assumed to include a main power 

cable, auxiliary power cable (for FSP parasitic loads), and a data transmission cable (for FSP 

instrumentation signals). The main power cable includes parallel channels for each of the eight Stirling 

alternators. The auxiliary power cable is assumed to carry a total of 5 kWe via 10 parallel channels. 

TABLE 3.—SCENARIO 5 POWER TRANSMISSION ANALYSIS 

Shield option A B C D 

Separation distance, m 100 200 1000 400 

Cable length, m 125 250 1250 500 

Transmission, Vac 400 400 2000 400 

Auxiliary bus, Vdc 120 120 120 270 

Auxiliary bus location Outpost Outpost FSP Outpost 

Main power cable, kg 48 128 688 400 

Auxiliary power cable, kg 120 450 120 300 

Data cable, kg 12 32 12 100 

High-voltage transformers, kg -------- -------- 144 -------- 

Total transmission mass, kg 180 610 964 800 

Cases A and B assumed direct 400-Vac-power cabling from the Stirling alternators to the power 

distribution node where the 400 Vac was converted to 120 Vdc for the user load bus. The same 120 Vdc 

bus was used to power FSP parasitic loads, such as pumps and motors, via a power cable from the 

distribution node back to the FSP system. The larger separation distance for Case C required the addition 

of high-voltage transformers near the FSP system to boost the transmission voltage to 2000 Vac. The 

2000 Vac provides a reasonable compromise on cable mass, development risk, and operational 

complexity. The 120-Vdc auxiliary power bus and FSP data bus was assumed to be co-located with the 

transformers at a 100 m distance from the FSP system. Case D assumed direct 400-Vac transmission, a 

120-Vdc user load bus, and a 270-Vdc auxiliary power bus and return cable. 

The LSS and Constellation Architecture Team settled on two FSP configurations for Scenario 5. The 

two systems used the same reactor, power conversion, heat rejection, and PMAD electronics. The offloaded configuration assumed the use of the “ATHLETE” utility rover for excavating a hole, removing 

the FSP system from the lander, transporting it to the site, and positioning it in the hole. The total FSP 

system mass was about 5800 kg including shielding and cabling. The landed configuration assumed the 

regolith-augmented shield with the lander cavity filled using a crane that scoops regolith collected near 

the lander by a bladed rover. The total FSP system mass was about 6600 kg with shielding and cabling. 
NASA/TM—2011-216819 10 

International Architecture Working Group—Power Function Team 

Beginning in April 2009, a Power Function Team was formed to support the International Architecture 

Working Group (IAWG) and develop power system concepts for a Global-Point-of-Departure human lunar 

mission. The team consisted of members from NASA, the European Space Agency (ESA), and the Japanese 

Space Agency (JAXA). The proposed IAWG mission architecture assumed an initial human mission to the 

lunar south pole (Shackleton crater) followed by robotic relocation of the initial surface assets to 

accommodate subsequent human missions to nonpolar sites such as Malapert Mountain and Schrodinger 

Basin. Once these initial areas were explored, a single site could be selected for a follow-on long-duration 

mission phase. Solar-based power systems with RFCs were baselined to include adequate energy storage 

capacity for 5-day eclipse durations at Shackleton. However, nighttime power requirements may exceed 

available energy storage capacity for missions beyond Shackleton. Two nuclear-based options were 

analyzed including a 2-kWe Large-Scale Stirling Radioisotope Generator (LSRG) and a 10-kWe Mobile 

Fission Power System (MFPS). The LSRG-based architecture enabled 11-day eclipse missions at Malapert, 

while the MFPS-based architecture provided sufficient capacity to enable both the 11-day eclipse missions 

at Malapert and 15-day eclipse missions at Schrodinger. 

Mobile Fission Power System 

The objective for the MFPS was to provide a small fission power system that could be easily 

deployed and moved if necessary. The basic concept was a scaled-version of the buried 40-kWe FSP 

system that would be off-loaded from the lander, placed on the lunar surface, and shielded with a regolith 

berm as shown in Figure 7. The system would be designed to produce 10 kWe output, continuously 

during the lunar daylight and nighttime periods. The reactor would be oversized, based on the larger 

40-kWe system, to provide growth capacity and simplify the qualification of possible follow-on systems. 

The system would utilize a single 12-kWe dual-opposed Stirling convertor, instead of the four units used 

on the 40-kWe system. Heat rejection would be provided by a simple two-panel deployable radiator, as 

opposed to the 10 radiator assemblies required on the 40-kWe system. A 200-m transmission cable with 

remote electrical controls and 120-Vdc load bus would provide the interface to the mission power loads. 

In addition to the regolith berm, a supplemental water shield would surround the reactor to limit radiation 

to 3 mrem/hr at the 200 m power hub. The key feature that distinguishes this concept from others is the 

capability to be shut down and moved to a new location if required. 

The MFPS mass summary with Current Best Estimate (CBE) mass values is provided in Table 4. The 

system could be delivered in as many as three separate packages. The water could be delivered 

specifically for this use or scavenged from lunar sources. The mass of regolith-moving equipment to 

create a surrounding berm is not included and assumed to be an available asset, if needed. 

Figure 7.—Mobile fission power system concept. 
NASA/TM—2011-216819 11 

TABLE 4.—MOBILE FISSION POWER SYSTEM MASS SUMMARY 

Mobile FPS subsystem Mass, 

kg 

Includes 

Power plant 1615 Reactor, water shield vessel, power 

conversion, radiator, and 

truss/structure. 

Power Management and Distribution 

(PMAD) 

415 Cabling, electrical controls, and 120-

Vdc load interface bus. 

Water (for shield) 1310 Liquid water for filling shield vessel 

prior to reactor startup. 

Total current best estimate 3340 

The MFPS can provide various benefits for the relocation and long-duration mission phases. During 

the relocation phase, the MFPS provides a continuous 10-kWe day/night power source for stationary 

applications that can be shut down and relocated as needed. During the long-duration phase, the 10-kWe 

MFPS could be moved to the long-duration site and reused, or the design could be scaled to 40 kWe and a 

follow-on system could be delivered to the site. The higher power capacity could provide expanded 

outpost operations to allow closed loop life support, high-production rate ISRU, and larger crew size if 

desired. The 8-yr design life assures long-term campaign use for the entire mission phase. The same 

technology could be used on Mars, making this option very well aligned with “Mars Forward” goals. 

MFPS Concept-of-Operations 

A preliminary assessment was made to determine the activities associated with launching, installing, 

and operating a MFPS. The main activities are summarized in the paragraphs below. 

Launch and Delivery 

The reactor is launched cold and subcritical. The reactor presents essentially no radiological hazard 

during lander and launch vehicle integration. The power plant will be delivered to the launch site as a 

fully integrated package. The MFPS system can be delivered as a single payload or in a series of packages 

depending on the available capacity of the lander. Prior to delivery, the MFPS system would be 

acceptance tested at a DOE nuclear facility. The reactor contains Special Nuclear Material that requires a 

specific class of security and safeguards. This is the primary difference between the MFPS payload and 

other lunar surface payloads. The reactor is designed to preclude inadvertent criticality under potential 

launch accident scenarios. During launch and delivery, the MFPS system may require periodic electrical 

power to exercise fluid pumps and monitor instrumentation sensors. 

Off-Loading and Setup 

A suitable site for the power plant must be determined that is approximately 200 m away from the 

outpost. If the local terrain offers natural outcroppings, these could provide shielding benefits and 

possibly eliminate the need for the water shield and regolith berm. The PMAD (415 kg) would be offloaded from the lander. A single power electronics pallet (1 by 1 by 1 m, 175 kg) would be setup near the 

outpost with easy access for the electric power users. The pallet provides the interface for power loads 

and data communications. A transmission cable bundle (<1 m diam., 240 kg) would be connected to the 

electronics pallet and unfurled via a spool to the power plant site. A robotic rover could deploy the cable. 

The transmission cable includes the main power cable, data signal cable, and return power cable for 

power plant auxiliary loads. The power plant (3 by 1.5 by 7 m, 1615 kg) would be off-loaded from the 

lander and transported via a cargo rover to the installation site. The power plant would be set directly on a 

flat regolith surface. The center of mass is very low and centered. Support bracing could be added if 

desired. A bladed rover would be used to move regolith to form a 2-m-high berm around the power plant, 

in lieu of using natural topography. 
NASA/TM—2011-216819 12 

System Startup 

Startup could be performed by local crew members or remotely from Earth. After connection of the 

transmission cable, initial communications would be established via telemetry signals to the electronics 

pallet. Computers and instrumentation would be started at 100 W total input power. Trace heating of the 

one reactor coolant (NaK) loop and two radiator coolant (H2O) loops would be started with 100 W input 

power on each loop. The shield vessel would be filled with liquid water (1310 kg). The one reactor NaK 

pump would be started with 250 W input power. The NaK loop trace heaters would be turned off. The 

two radiator loops would be charged with H2O coolant from a heated storage tank. The two radiator H2O 

pumps would each be started with 100 W input power. The H2O loop trace heaters would be turned off. 

Three of the six reactor control drums would be adjusted to the reactor operating position. The fourth, 

fifth, and sixth control drums would be adjusted in succession resulting in initial reactor criticality. The 

control drums would be adjusted to reach 600 K NaK outlet temperature. The Stirling convertor would be 

started resulting in approximately 1 kWe power output. The two radiator panels would be deployed. The 

control drums would be adjusted in steps and the Stirling convertor stroke would be gradually increased 

with hold-points at 2.5, 5, 7.5, and 10 kWe. The regulated 120-Vdc bus would be enabled at 5 kWe. The 

total startup energy for this process is estimated at 10 kW hr. 

Water for Reactor Shield 

The combination of the water shield and regolith berm reduces the reactor radiation to 3 mrem/hr at 

200 m. The resulting dose to an unshielded astronaut at the 200 m boundary during a typical 30-day 

mission is about 2.2 rem. The OSHA limit for nuclear power workers is 5 rem/yr and the NASA guideline 

for astronauts is 50 rem annually. The use of natural topography could reduce or eliminate the need for 

the water and/or regolith berm. The 1310 kg of liquid water for the water shield could be delivered from 

Earth, or acquired from lunar sources. The water could also be scavenged from other lunar surface assets, 

such as fuel cells or propellant tanks. 

System Operations 

Once the MFPS system reaches full power it can operate with minimal human intervention requiring 

only periodic health monitoring from a data console at the outpost or Earth. The MFPS system is 

designed to automatically respond to lunar day/night transients, electric load changes, and recoverable 

power plant faults without human intervention. The MFPS system would produce electric power 

continuously and shunt excess power not required by loads via a parasitic load radiator. A periodic 

control drum adjustment (perhaps twice per Earth month) may be performed to maintain reactor coolant 

temperature within a nominal band. The electronics pallet includes a power switch panel that allows users 

to connect electric loads as needed. This would be the power interface for habitats, landers, ISRU plants, 

rover recharging, science experiments, etc. The electronics pallet is located near the crew for 

maintenance, if required. The power plant is designed for an 8-yr service life without maintenance. The 

design life would be demonstrated as part of the ground verification testing. Performance degradation 

over the design life should be negligible. Radiator area margin is included to account for environmental 

degradation such as ultraviolet radiation and dust. The redundancy in the reactor control drums provides 

fault tolerance for reactivity adjustments. 

Shutdown and Relocation 

At end-of-life, the power plant would be shut down and could remain at the installation site. Within 

several weeks, the radiation will decrease below natural background levels and the plant should not pose 

any particular safety concern. The plant shutdown would include an adjustment of the control drums to 

make the reactor subcritical, and a commanded stop to the Stirling convertor, reactor pump, and radiator 

pumps. The power plant and PMAD could also be shut down and moved to another location. Robotic 

access would be permitted within hours; human access may be permitted within days. To prepare for 

relocation, the radiator panels would be retracted and the transmission cable would be disconnected. The 
NASA/TM—2011-216819 13 

water shield could be drained, if desired, to reduce the mass for transport. The MFPS system cannot 

provide power while it is being moved. At the new location, the setup and startup would be repeated as 

described previously. Subsequent startups should be easier and faster. After setup, the system could be 

operational within ~12 hr. The MFPS system could be relocated multiple times recognizing that each 

shutdown, movement, and restart increases the probability of a system malfunction. 

Conclusions 

Fission Surface Power (FSP) systems are currently being studied by NASA as an option for future 

human exploration missions to the Moon and Mars. NASA and the Department of Energy (DOE) have 

partnered to help mature FSP technology so that it may be considered for future flight development. The 

NASA/DOE team has generated a preliminary FSP reference concept to help guide technology 

development and provide a possible starting point for future flight systems. A portion of the FSP project 

is focused on concept definition and integration studies to evaluate the use of FSP technology for various 

mission architectures. This paper discusses some of the recent architecture studies and potential methods 

for utilizing FSP systems. The FSP team has supported these architecture studies by supplying FSP 

design characteristics, developing mission-compatible configuration options, and defining a concept-ofoperations consistent with the mission objectives. The technology is adaptable to meet a wide range of 

mission needs. Configuration options include buried-reactor systems, lander-integrated systems, surfacemounted systems, and systems which are deployed via a wheeled cart. Power levels can range from 

several kilowatts up to about 100 kWe without a significant change to the basic technology. The primary 

building blocks of the FSP system: reactor, power conversion, and heat rejection can be common to many 

design configurations making the technology extremely robust and versatile. 

References 

1. Affordable Fission Surface Power System Study Final Report, Exploration Systems Mission 

Directorate Memorandum For The Record, 5 October 2007. 

2. Executive Summary of Reference Fission Surface Power System Concept Selection, Exploration 

Systems Mission Directorate Memorandum For The Record, 23 July 2008. 

3. Lee Mason, David Poston, and Louis Qualls, “System Concepts for Affordable Fission Surface 

Power,” NASA/TM—2008-215166, January 2008. 

4. Human Exploration of Mars: Design Reference Architecture 5.0, Bret G. Drake, editor, NASA/SP—

2009-566, July 2009. 

5. David I. Poston, Lee S. Mason, and Michael G. Houts, “Radiation Shielding Architecture Studies for 

NASA’s Lunar Fission Surface Power System,” Proceedings of Nuclear and Emerging Technologies 

for Space (NETS–2009), Atlanta GA, 14–19 June 2009, Paper 208590. 
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-04-2011 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

A Summary of NASA Architecture Studies Utilizing Fission Surface Power Technology 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Mason, Lee, S.; Poston, David, I. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 463169.02.03.01.01 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-17386 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2011-216819 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 20 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Beginning with the Exploration Systems Architecture Study in 2005, NASA has conducted various mission architecture studies to evaluate 

implementation options for the U.S. Space Policy. Several of the studies examined the use of Fission Surface Power (FSP) systems for 

human missions to the lunar and Martian surface. This paper summarizes the FSP concepts developed under four different NASA-sponsored 

architecture studies: Lunar Architecture Team, Mars Architecture Team, Lunar Surface Systems/Constellation Architecture Team, and 

International Architecture Working Group-Power Function Team.

15. SUBJECT TERMS

Nuclear electric power generation 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

19 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18