James Fincannon

Glenn Research Center, Cleveland, Ohio

Compilation of Trade Studies for the Constellation 

Program Extravehicular Activity Spacesuit 

Power System 

NASA/TM—2009-215803

November 2009
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James Fincannon

Glenn Research Center, Cleveland, Ohio

Compilation of Trade Studies for the Constellation 

Program Extravehicular Activity Spacesuit 

Power System 

NASA/TM—2009-215803

November 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2009-215803 iii 

Compilation of Trade Studies for the Constellation Program 

Extravehicular Activity Spacesuit Power System 

James Fincannon 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This compilation of trade studies performed from 2005 to 2006 addressed a number of power system 

design issues for the Constellation Program Extravehicular Activity Spacesuit. Spacesuits were required 

for spacewalks and in-space activities as well as lunar and Mars surface operations. The trades 

documented here considered whether solar power was feasible for spacesuits, whether spacesuit power 

generation should be a distributed or a centralized function, whether self-powered in-space spacesuits 

were better than umbilically powered ones, and whether the suit power system should be recharged in 

place or replaced. 
NASA/TM—2009-215803 iv 

 
NASA/TM—2009-215803 v 

Contents 

Abstract ........................................................................................................................................................ iii

1.0 Trade Study 1: High Efficiency Solar Power Augmentation for the Advanced EVA Suit ............. 1

1.1 Introduction ......................................................................................................................... 1

1.2 Solar Cell Assumptions ...................................................................................................... 1

1.3 Solar Panel Structure Assumptions ..................................................................................... 1

1.4 Solar Panel Design Assumptions ........................................................................................ 1

1.5 Configuration Assumptions ................................................................................................ 2

1.6 Projected Area Assumptions ............................................................................................... 6

1.7 Environment Assumptions .................................................................................................. 6

1.7.1 Shadow Factors ...................................................................................................... 6

1.7.2 Albedo Factor ........................................................................................................ 6

1.7.3 Pointing Factor ....................................................................................................... 7

1.8 Operational Assumptions .................................................................................................... 8

1.8.1 Electronics Assumptions ........................................................................................ 8

1.8.2 Other Masses ......................................................................................................... 8

1.9 Calculation Method ............................................................................................................. 9

1.10 Results ................................................................................................................................. 9

1.11 Conclusions ....................................................................................................................... 14

1.12 Additional Information: Tracking Solar Panel Concept ................................................... 14

1.13 Additional Information: Solar Fiber Concept ................................................................... 15

2.0 Trade Study 2: Comparison of Localized and Centralized Power System Options for the 

Advanced EVA Suit ....................................................................................................................... 16

2.1 Results Summary .............................................................................................................. 16

2.2 Introduction ....................................................................................................................... 16

2.3 Assumptions...................................................................................................................... 16

2.4 Commonality Assessment ................................................................................................. 18

2.5 Analysis Results ................................................................................................................ 19

2.6 Conclusion ........................................................................................................................ 25

3.0 Trade Study 3: Umbilical-Only Power Versus Suit-Mounted Power for the Advanced 

EVA Suit ........................................................................................................................................ 26

3.1 Results Summary .............................................................................................................. 26

3.2 Purpose ............................................................................................................................. 26

3.3 Past/Current Umbilical Data ............................................................................................. 26

3.3.1 Emergency Backup Power ................................................................................... 27

3.3.2 Contingency EVA ................................................................................................ 27

3.3.3 Applications of Umbilicals .................................................................................. 28

3.3.4 Anticipated Sortie Duration and Umbilical Length ............................................. 28

3.4 Analysis Method ...............................................................................................................28

3.5 Power System Assumptions .............................................................................................. 29

3.5.1 Analysis Limitations ............................................................................................ 30

3.5.2 Analysis Results ................................................................................................... 30

3.6 Other Aspects .................................................................................................................... 41

3.7 Conclusions ....................................................................................................................... 42

3.8 Future Modifications ......................................................................................................... 43

4.0 Trade Study 4: Lunar Surface EVA Suit Power Subsystem Rapid Recharge Versus 

Replace ........................................................................................................................................... 44

4.1 Results Summary .............................................................................................................. 44

4.2 Introduction ....................................................................................................................... 44

4.3 PLSS Power Subsystem Options ...................................................................................... 44
NASA/TM—2009-215803 vi 

4.4 PLSS Power Level Assumptions ...................................................................................... 45

4.4.1 PLSS Primary Fuel Cell Assumptions ................................................................. 46

4.4.2 Recharge Methods ............................................................................................... 46

4.4.3 Other PLSS Power Subsystem Sizing Assumptions ............................................ 47

4.4.4 PLSS Power Subsystem Redundancy Assumptions ............................................ 48

4.4.5 PLSS Power Subsystem Sizing Results ............................................................... 49

4.4.6 RRS Station Power Subsystem Requirements ..................................................... 54

4.5 RRS Station Power Options .............................................................................................. 54

4.6 RRS Station Redundancy Assumptions ............................................................................ 55

4.6.1 RRS Station Power Subsystem Mass Results ...................................................... 55

4.6.2 Combined RRS and 2 PLSS Power Subsystems Mass Results ........................... 62

4.6.3 Recharge Time Effect on RRS Station Mass ....................................................... 70

4.6.4 Quantitative and Qualitative Ranking Comparison of Options ........................... 70

4.7 Conclusions ....................................................................................................................... 71

4.8 Additional Information: Rapid Recharge Battery Data .................................................... 72

4.8.1 Charge Rates and Times ...................................................................................... 72

4.8.2 Charging Profiles ................................................................................................. 72

4.8.3 Rapid Recharge Battery Sizing Based on Charge Rates ...................................... 73

4.8.4 Thermal Effects.................................................................................................... 73

4.8.5 Safety ................................................................................................................... 73

4.8.6 Battery Life .......................................................................................................... 73

4.8.7 Rapid Recharge Vendors ..................................................................................... 73

4.8.8 Assumed Analysis Data ....................................................................................... 75

4.8.9 References ............................................................................................................ 75

4.9 Additional Information: Lunar Surface Recharging Methods .......................................... 76

4.9.1 References ............................................................................................................ 77

 
NASA/TM—2009-215803 1 

1.0 Trade Study 1: High Efficiency Solar Power Augmentation for 

the Advanced EVA Suit 

Date: June 08, 2005 

1.1 Introduction 

The purpose of this trade study is to evaluate the feasibility of using solar power to reduce the burden 

on the energy storage system for the advanced EVA suit. The goal is to reduce the overall mass of the 

power system by replacing some of the capacity of the energy storage through the use of solar generated 

power. 

1.2 Solar Cell Assumptions 

High efficiency solar cells were assumed for this application. Lower efficiency cells were not 

considered because the required coverage area is impractical. Two kinds of solar cells were examined, 

32 percent (reference efficiency) and 40 percent. The 32 percent efficiency is representative of a 

Lockheed Martin Marietta four junction GaAs/SiGe cell, selected because of its low cell density. Other 

SOA (state of the art) cells have efficiencies from 27 to 35 percent, but the selected one provides the 

highest SOA specific power (i.e., W/kg) due to its lower cell density. The 40 percent efficiency cell 

(currently at low TRL, technology readiness level, but being developed) was a hypothetical, optimistic 

projection (likely some variant of four junction cells) for far-term use. 

1.3 Solar Panel Structure Assumptions 

Two types of solar panel structure were assumed. One was the traditional “rigid” panel with a 

structural honeycomb mounting core and a thick facesheet. This is the standard method for solar panels 

attached to the surface of spacecraft or deployed in solar array “wings”. The second approach was a 

“flexible” panel. This is typically used in the Ultraflex design and essentially omits the core and facesheet 

by using a specialized method of attachment. The present maximum cell efficiency utilized for this 

approach is 20 percent. To be used (if possible) for higher efficiency 32 percent or projected 40 percent 

cells would require considerable development costs. Thus, it is likely flexible panels will only be for farterm use. 

1.4 Solar Panel Design Assumptions 

The packing factor for the solar panel was assumed to be 0.8. This is a typical value for small satellite 

solar panels. By reducing this value the solar panel area increases, but it enables reduction of the solar 

panel temperature using passive means. No thermal calculations to determine the best packing factor for 

the considered environments were attempted at this time. 

Since no thermal design calculations were performed, the operating temperature of the solar panel 

was assumed. It was unknown how the thermal load would be controlled, although passive means are 

desirable. If the solar cells (and panel) were 100 °C then 32 percent cells would decrease in efficiency to 

26.9 percent. All assessed cells were assumed to have the same temperature effects. 
NASA/TM—2009-215803 2 

The following table includes other assumptions used in the analysis. 

Packing factor ......................................................................................... 0.80 

Solar panel utilization factor ................................................................... 0.95 

Solar cell operating temperature ........................................................... 28 °C 

Solar panel cell mismatch factor ............................................................. 0.98 

Solar panel interconnect efficiency factor ............................................... 0.97 

Solar panel blocking diode drop ............................................................. 0.99 

Solar panel harness ................................................................................. 0.99 

Other losses due to solar panel lifetime..................................................... 0.0 

Solar intensity (near Mars plane) ................................................... 606 W/m2

Solar intensity (near Earth plane) ................................................. 1370 W/m2

The layers of the solar panel are listed below. Note that the rigid and flexible panels differ by the core 

and facesheet layers. The core thickness for the rigid panels was assumed. A more detailed structural 

analysis is required for this value. 

Coverglass (Ceria-doped microsheet) 2620 kg/m3 0.1016 mm 

Adhesive (DC93-500) 1080 kg/m3 0.0127 mm 

Cell 2330 kg/m3 0.1397 mm 

Back contact (copper) 8940 kg/m3 0.1016 mm 

Adhesive (DC93-500) 1080 kg/m3 0.00381 mm 

Facesheet (honeycomb K13C2U-3 ply) 1659 kg/m3 0.2032 mm (grid) 

0.0 mm (flexible) 

Adhesive (DC93-500) 1080 kg/m3 0.1016 mm 

Core (Al-honeycomb-3/8-5056-1.0) 16 kg/m3 10.0 mm (grid) 

0.0 mm (flexible) 

1.5 Configuration Assumptions 

Four different “fixed” solar panel layouts were considered. The first was a flat solar panel which was 

envisioned to be either mounted on the suit front, back, or by way of a pole attached to the suit portable 

life support system backpack (likely slightly above the helmet and centered). The second was a two sided 

solar panel with one on the suit front and one on the suit back, or, alternatively, attached by way of a pole 

to the suit backpack. The third layout was a cube of panels attached on a pole to the backpack. The fourth 

layout was sphere (or greater than 12 faces) of cells attached to the backpack. 

Other aspects of configuration include the ability to cant the flat panels, particularly for surface 

operations. It is assumed that, prior to the sortie, the astronaut sets the cant angle on the pole mounted 

solar panels (flat options) to maximize the incident solar energy. This depends on the latitude, longitude 

and time of year for either Mars or Moon location. For two sided panels, this canting can be in a tent-like 

formation. Therefore, at the North or South pole of the Moon or Mars, the astronaut would set the cant 

angle to 0 ° (aligned with local vertical) in order to maximize the incident solar energy for the flat panel 

options and at the equator when the Sun is overhead, he would set the cant angle to 90 °. 

No consideration was given to suit clearance issues. It was assumed the pole mounted panels were 

attached after egress. 

The following illustrations depict the variations of the solar panels. It should be noted that the layout 

of the suit mounted panels are not optimized for other considerations (equipment, switches, lines, access 

panels). The depicted image is more to give a sense of practical size limits. The depicted area is based on 

the projected area of 0.15 m2

. 
NASA/TM—2009-215803 3 

Suit Mounted Panel Option 

(front and/or back) 

Boom Mounted Solar Panel Option 

(one- or two-sided) 
NASA/TM—2009-215803 4 

Boom Mounted Solar Panel Option 

(surface operations, one-sided, 1 axis cantable prior to sortie) 

Boom Mounted Solar Panel Option 

(surface operations, two-sided, 1 axis cantable prior to sortie) 
NASA/TM—2009-215803 5 

Boom Mounted Cube Solar Panel Option 

Boom Mounted Sphere Solar Panel Option 
NASA/TM—2009-215803 6 

1.6 Projected Area Assumptions 

An assumed maximum projected area was used to size the solar panel. It is not known how much area 

is practical considering all the other uses of the surface. For all configurations, the maximum projected 

are for the solar panels was 0.15 m2

 (1.6 ft2). This value was selected because it was the maximum likely 

coverage area for the front of the suit. Note that this ignores any other considerations for usage of the area 

on the suit. The projected area for the cube and sphere are the maximum possible area that the objects can 

show to the Sun. Thus, for a sphere of radius R, the projected area is PI*R2

. 

For a cube, with side length L, the projected area is 1.5*L2

. 

Note that energy losses due to shallow solar incidence angles were ignored in this study. Such losses 

would reduce the power level by about 5 to 10 percent. 

1.7 Environment Assumptions 

Key factors in sizing the solar panel include shadowing, albedo and pointing angle. The study 

assumed suit operations from LEO to the lower level of the radiation belt, from the upper level of the 

Earth’s radiation belt to the Moon, on the Moon surface and on the Mars surface. Although it is 

recognized that these values vary continuously, in order to quantify these factors for sizing the solar 

panels, it is necessary to use an average estimate of these over likely averaging periods. 

1.7.1 Shadow Factors 

For LEO and Space operations, shadow factors are used to represent the amount of shadowing caused 

by the vehicle/suit passing into eclipse and the amount of shadowing due to proximity with the vehicle. 

For Moon and Mars operations, the shadowing is assumed to be caused by the terrain. Shadowing 

changes continuously based on many factors (astronaut location, orientation, time of year). It should be 

recognized that the shadow factor could vary from 0 to 1.0 if examined at points in time. 

For LEO operations, the average time a spacecraft is in planetary shadow through the entire year is 38 

percent of the time for the ISS altitude (400 km) but with 28° inclination (best for Kennedy launches and 

likely the value for CEV launches). The value goes down to 34 percent of the time for the lower level of 

the radiation belt (650 km). A conservative estimate for the amount of shadowing over an average sortie 

based only on eclipse shadowing is 38 percent of the time (a 0.38 shadow fraction). For some times of 

year and somewhat different inclinations, there are zero eclipse times, but it turns out that this is 

infrequent (the worse case is 39 percent). For shadowing due to hardware, it is assumed that the 

suit/astronaut is blocked from the Sun only about 25 percent of the time (of the insolation time period) 

due to the fact of being in close proximity to a large spacecraft (a 0.25 shadow fraction). This is 

somewhat optimistic but difficult to model. Statistical data on whether the astronaut suit is in insolation 

during an EVA is not available and the CEV vehicle orientation and configuration unknown. The total (a 

0.63 shadow fraction) provides the estimate for the LEO operations shadow factor. 

For In-Space/beyond LEO operations (i.e., translunar transit), the eclipse time would be negligible 

thus the only shadowing would still be due to the hardware (0.25). This same value could be used for InSpace use to Mars. 

For Moon and Mars operations, it is assumed that shadowing due to terrain (boulders) and hardware 

occurs for no more than 10 percent of the time on average, thus the shadow factor is 0.10. Implicit in this 

assumption is the exclusion of any analysis cases where the astronaut is located in eclipse exclusive 

environments (craters, dark side or night operations). 

1.7.2 Albedo Factor 

The albedo factor is a measure of the amount of solar energy that is diffusely reflected from 

surrounding surfaces near the astronaut. This term is an average albedo over time during insolation. 
NASA/TM—2009-215803 7 

For LEO operations, albedo energy comes from two sources: the Earth and the spacecraft. For the 

Earth this is assumed to be 0.2 at the ISS altitude (this is the fraction of albedo energy from the Earth 

surface albedo, 0.37, to reach a flat plate pointing at Earth at the ISS altitude). This must be adjusted 

during energy calculations using the pointing factor to account for random orientation of the astronaut 

(panel) with respect to the energy source. Note that for orbital in space EVA around the Moon, the albedo 

value of 0.03 is appropriate for 100km orbit, but since this is so small, it is assumed zero (equivalent to 

the albedo value outside the Earth’s radiation belt and between the Moon and Mars orbit). 

Spacecraft albedo energy depends on the thermal properties of the exterior of the vehicle. White paint 

would reflect 75 percent of the light and black paint 3 percent with a variety of materials ranging between 

these extremes. For this analysis 0.75 was assumed. 

For non-LEO In-Space operations, the albedo only comes from the spacecraft and has the value as 

above. For Moon surface operations, the assumed average albedo is 0.07 (obviously it is higher and lower 

depending upon location). For Mars operations, the accepted average albedo is 0.15. 

For surface operations, spheres are assumed to gather albedo from a hemispherical area and cubes 

gather this energy from three sides. 

As an application note, it should be noted that just as with solar energy, the orientation of the panel 

with respect to the Sun is critical in determining the amount of solar energy that can impinge upon the 

solar panel, the same applies to the albedo energy. However, to determine the average albedo energy onto 

the solar panels, both the random orientation of the panel with respect to the hardware/vehicle surface and 

the random orientation of the hardware to the Sun must be accounted for (such pointing factors are 

described in the next section). 

1.7.3 Pointing Factor 

Pointing factor is the average of the cosines of the solar panel unit normal vector from the Sun vector. 

It is the average projected area divided by the maximum projected area (i.e., the maximum possible area 

that can see the Sun if one re-orients the object). This is used to account for random orientation of the 

astronaut during nominal operations. For any in-space operation, the likely orientation of a flat panel with 

respect to the Sun is any combination of rotations about the X and Y astronaut axis. These cannot be 

known beforehand, thus averaging is required. For any surface operations, the astronaut is assumed to 

randomly rotate about his local vertical axis (although it is known some bending is performed, this is 

assumed to be infrequent). 

For in-space operations, the calculated pointing factor for one sided panels is 0.2. For two sided 

panels, it is 0.4. For a solar panel cube, it is 0.8. For a solar panel sphere, it is 1.0. 

For Moon surface operations, it is assumed that the likely rotation of the astronaut is about the local 

vertical axis and this axis depends on the latitude and longitude of the astronaut. The solar angle to the 

surface changes little during a sortie. The flat solar panel(s) are assumed to be manually cantable prior to 

the sortie to be at the best orientation. At lunar noon at the equator, the cant angle is 90° (with reference to 

the vertical axis of the astronaut suit). At lunar noon at the poles (or 90° anywhere from the equator case), 

this angle is 0° (parallel to the vertical axis of the suit). This angle is all that is needed to define the 

orientation of the Sun with respect to the Moon and the sortie location. Thus, it is possible to examine the 

entire range of sortie locations without defining them by time of month, latitude or longitude. For a cant 

angle of 90°, the one sided flat panel has a pointing factor of 1.0. At 0° cant angle, the other extreme is 

encountered (0.32 for the one-sided flat plate). The pointing factor at 0° cant is 0.64 for a two-sided flat 

plate and 0.85 for the solar panel cube. Regardless of location on the Moon, the pointing factor for the 

solar panel sphere is 1.0. 

For Mars surface operations, the conditions are similar to the Moon except that the solar angle to the 

surface changes considerably during a long sortie (~8 hr) on Mars (a “typical” day/night cycle versus the 

Moon’s month long day/night cycle). For these analyses, it is assumed that the solar angle or time of day 

does not change during the sortie. Therefore, the pointing factor will be the same as in the Moon cases. 
NASA/TM—2009-215803 8 

Other cant angles different from 0 and 90° were not presented in this report because 0 and 90 provide 

the extreme sizing points. Usually, 90° cant angle (by the equator) provide the most energy. Depending 

on the results of this analysis and determination of reasonable break point (minimum percentage of mass 

reduction by adding solar augmentation), future assessments at other cant angles can be used to map 

viable locations to use solar augmentation. 

1.8 Operational Assumptions 

The main trade study was performed assuming a nominal power level of 200 W were required for a 

time of 8 hr. Sensitivity analysis for a selected power system configuration was performed for 100 W for 

2 hr, 150 W for 6 hr and 200 W for 8 hr. 

The contingency power level was assumed to be 60 W for 30 min met by a primary energy storage 

system of 220 W-hr/kg. 

It was assumed that the EVA suit power/energy storage system was only used when separated from 

other power sources such as a rover. Thus, the energy storage time is referenced from that time of total 

separation. Otherwise, it can be assumed the EVA suit draws upon the power of the rover or other 

translation vehicle in order to extend the traverse capability. 

The main energy storage system was assumed to have an 80 percent depth of discharge regardless of 

whether it could possibly discharge further because realistic operational limits do not allow EVA to be 

extended to the last watt-hr of energy. 

1.8.1 Electronics Assumptions 

The battery discharge unit efficiency was sized for a regulated bus. The solar panel 

electronics/regulator was assumed to be a regulated direct energy transfer (results in lower than 

100 percent efficiency in utilizing the solar power because of inadequacy in matching the peak power 

point on the cell I-V curve). 

1.8.2 Other Masses 

Thermal masses were estimated to be 0.01 kg/ W for the energy storage part of the power system. For 

the boom mounted options, the solar panels are assumed to be thermally connected (e.g., cross straps) to 

permit thermal radiation from the noninsolation surface to help keep the cells at lower temperatures. For 

Solar Panel 

Energy 

Storage 

Loads 

28VDC 

200 W, 8 hr 

80 percent Max DoD 

50,100,150,200 W-hr/kg

Discharge Unit

No Load DC/DC converters assumed 

Cables: 99 percent Efficiency, .001 kg/ W (no detailed 

model) Assume no charge unit (switch out only) 

Regulator

Distribution Unit

90 percent 

Efficiency 

.02 kg/ W 

99 percent 

Efficiency 

.02 kg/ W 

95 percent Efficiency 

.025 kg/ W 

Regulated Direct Energy 

Transfer 

Cable 

Cable 

Cable 

Cable 

Cable 
NASA/TM—2009-215803 9 

suit mounted options, either the panel is thermally connected to the suit or additional solar panel area is 

needed for inter-cell radiative surfaces. The thermal mass for the solar part of the power system (related 

just to the solar panels) was assumed to be 20 percent of the solar panel mass. 

Structural masses for the energy storage system were assumed to be 10 percent of the mass of the 

entire power system. The structural mass (used to attach the solar panel to the suit or support the solar 

panel: the pole or mechanism) for the solar part of the power system was assumed to be 20 percent of the 

solar panel mass. 

1.9 Calculation Method 

To perform these calculations, standard solar panel and energy storage sizing relationships were used. 

The baseline cases assumed hypothetical energy storage systems of 50, 100, 150 and 200 W/kg. The 

reason for this is that the goal of this trade study is to find the point at which the solar panel mass is lower 

than the mass of the energy storage that is no longer needed because of the solar panels. It is recognized 

that some energy storage will be needed, but that solar arrays could reduce that need and that there is a 

trade-off in generation versus storage. 

Margins were not considered in this analysis. Typical values of margin for either power or mass 

estimates are 20 percent. 

1.10 Results 

Figure 1 shows the results for the EVA suit based on near term technology for use outside the 

spacecraft in low Earth orbit operation below the Earth radiation belt (i.e., minimal coverglass/shielding). 

These assume that the rigid panel, 32 percent solar cells are all that will be available in the time frame. 

This figure shows the relative mass savings from the total baseline non-solar power system mass estimate. 

The horizontal axis has the various assumed energy storage specific energy values. If one assumes the 

current EMU lithium (Li) polymer battery is the state of art value for W-hr/kg, then 100 W/kg and above 

should be the range of likely application. The best savings in total power system mass (using a spherical 

solar panel) is 8 percent. However, the near term projection for Li polymer batteries is 200 W-hr/kg 

(based on NASA’s, “Energy Storage Technology for Future Space Science Missions”). This reduces the 

maximum benefit to 4 percent. 

The four dashed lines are used to illustrate the benefit of simply improving the standalone baseline 

energy storage system by various percentages rather than invest resources into solar augmentation. The 

baseline energy storage specific energy (as shown on the horizontal axis) was increased by 5, 10, 15 and 

20 percent to create an improved energy storage system to which the baseline was compared. 

Examination of these lines shows that by improving the current EMU Li polymer battery (100 W-hr/kg) 

by about 13 percent in specific energy equals the benefit of adding the spherical solar panel. 

Figure 2 shows the parametric variation of projected solar panel surface area for the EVA suit based 

on 32 percent efficient cells with spherical, rigid solar panels for use outside the spacecraft in low Earth 

orbit operation below the Earth radiation belt (minimal coverglass/shielding). Although the maximum 

likely projected area is 0.15 m2

, this chart shows that increasing the area by about 50 percent improves by 

about 4 more percentage points from the baseline mass (for the 100 W-hr/kg energy storage). Higher 

energy storage specific energies provide even less improvement. 

Figure 3 shows the effect of the load power requirement and the duration of the load on the sizing of 

the solar augmented power system. Clearly, lower power levels are better suited for the solar 

augmentations than higher power levels and longer durations provide more benefit than short. The 100 

W-hr/kg energy storage mass improves 8 percentage points from the baseline when going from high 

power (200 W) to low power (100 W) sizing for 8 hr load duration. This improvement is only 5 

percentage points when the load duration is about 4 hr. For higher specific energy storage systems, the 

improvements for going to lower power level are less than 5 percentage points. 
NASA/TM—2009-215803 10 

Figure 4 shows the results for the EVA suit based on near term technology for use outside the 

spacecraft between geosynchronous orbit altitude and the lunar orbit (i.e., beyond the Earth’s radiation 

belt, thus requiring thicker coverglass/shielding). These cases assume the use of a rigid panel with 32 

percent solar cells. This figure shows the relative mass savings from the total baseline non-solar power 

system mass estimate. The horizontal axis has the various assumed energy storage specific energy values. 

If one assumes the current EMU Li polymer battery is the state of art value for W-hr/kg, then 100 W/kg 

and above should be the range of likely application. The results are improved beyond the LEO In-Space 

power system primarily due to eclipse shadowing considerations. The best savings in total power system 

mass (using a spherical solar panel) is 13 percent. However, for the near term projection for Li polymer 

batteries, 200 W-hr/kg, the maximum benefit is 9 percent. 

Figure 1: In-space/LEO: 1600 W-hr load energy requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

1 Sided/ 0.15m^2/ 32%/ Rigid

2-sided/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 32%/ Rigid

5% higher W-hr/kg

10% higher W-hr/kg

15% higher W-hr/kg

20% higher W-hr/kg

Figure 2: Projected Area Effect for Sphere, 32% Cell, Rigid, LEO (1600 W-hr load 

energy requirement)

80

85

90

95

100

0.05 0.1 0.15 0.2 0.25

Projected Area m^2

% of Baseline Mass 

50 W-hr/kg

100 W-hr/kg

150 W-hr/kg

200 W-hr/kg
NASA/TM—2009-215803 11 

The four dashed lines are used to illustrate the benefit of simply improving the standalone baseline 

energy storage system by various percentages rather than invest resources into solar augmentation. The 

baseline energy storage specific energy (as shown on the horizontal axis) was increased by 5, 10, 15 and 

20 percent to create an improved energy storage system to which the baseline was compared. 

Examination of these lines shows that improving the current EMU Li polymer battery (100 W-hr/kg) or 

an equivalent energy storage system by about 25 percent (by extrapolating the curves) in specific energy 

equals the benefit of adding the spherical solar panel. 

Figure 5 shows the results for far-term mass benefits of using advanced solar cell and solar panel 

technologies for low Earth orbit in-space application. State of the art (near term) data is provided for 

Figure 3: Load Level and Duration Effect on Spherical Solar Panel (32% cell, 

rigid, .15m^2 projected area, LEO)

75

80

85

90

95

100

105

110

115

12345678

Load duration (hrs)

% of Baseline Mass 

50 W-hr/kg, 200W load

100 W-hr/kg, 200W load

150 W-hr/kg, 200W load

200 W-hr/kg, 200W load

50 W-hr/kg, 100W load

100 W-hr/kg, 100W load

150 W-hr/kg, 100W load

200 W-hr/kg, 100W load

Figure 4: In-space/beyond GEO: 1600 W-hr load energy requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

1 Sided/ 0.15m^2/ 32%/ Rigid

2-sided/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 32%/ Rigid

5% higher W-hr/kg

10% higher W-hr/kg

15% higher W-hr/kg

20% higher W-hr/kg
NASA/TM—2009-215803 12 

comparison. Assuming the current EMU Li polymer battery (100 W-hr/kg), then maximum benefit is 11 

percent. For the 200 W-hr/kg energy storage, this benefit is 7 percent. 

Figure 6 shows the far-term mass benefits of using advanced solar cell and solar panel technologies 

for use outside the spacecraft between geosynchronous orbit altitude and the lunar orbit (i.e., beyond the 

Earth’s radiation belt, thus requiring thicker coverglass/shielding). State of the art (near term) data is 

shown for comparison. Assuming the current EMU Li polymer battery (100 W-hr/kg), then maximum 

benefit is 18 percent. For the 200 W-hr/kg energy storage, this benefit is 12 percent. 

Figure 5: In-space/LEO-far term: 1600 W-hr load energy requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

1 Sided/ 0.15m^2/ 32%/ Flexible

1 Sided/ 0.15m^2/ 40%/ Flexible

1 Sided/ 0.15m^2/ 32%/ Rigid

1 Sided/ 0.15m^2/ 40%/ Rigid

2-sided/ 0.15m^2/ 32%/ Rigid

2-sided/ 0.15m^2/ 40%/ Rigid

Cube/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 40%/ Rigid

Sphere/ 0.15m^2/ 32%/ Flexible

Sphere/ 0.15m^2/ 40%/ Flexible

Sphere/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 40%/ Rigid

Figure 6: In-space/beyond GEO-far term: 1600 W-hr load energy requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

1 Sided/ 0.15m^2/ 32%/ Flexible

1 Sided/ 0.15m^2/ 40%/ Flexible

1 Sided/ 0.15m^2/ 32%/ Rigid

1 Sided/ 0.15m^2/ 40%/ Rigid

2-sided/ 0.15m^2/ 32%/ Rigid

2-sided/ 0.15m^2/ 40%/ Rigid

Cube/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 40%/ Rigid

Sphere/ 0.15m^2/ 32%/ Flexible

Sphere/ 0.15m^2/ 40%/ Flexible

Sphere/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 40%/ Rigid
NASA/TM—2009-215803 13 

Figure 7 shows the Moon surface operations mass benefits of using advanced solar cell and solar 

panel technologies (as well as the state of art). Two cant angles (reflecting different locations on the 

Moon and times of year) are shown for the flat panel cases. A 90° cant is for noon at the equator and a 0° 

cant is for polar operation. Assuming the current EMU Li polymer battery (100 W-hr/kg), then maximum 

benefit is 18 percent. For the 200 W-hr/kg energy storage, this benefit is only 11 percent. 

Figure 8 shows the Mars surface operations mass benefits of using advanced solar cell and solar panel 

technologies (as well as the state of art). Two cant angles (reflecting different locations on the Mars and 

times of year) are shown for the flat panel cases. A 90° cant is for noon at the equator and a 0° cant is for 

polar operation. Assuming the current EMU Li polymer battery (100 W-hr/kg), then the maximum benefit 

is 8 percent. For the 200 W-hr/kg energy storage, this benefit is 5 percent. 

Figure 7: Moon surface: 1600 W-hr Load Energy Requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

Cube/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 40%/ Rigid

Sphere/ 0.15m^2/ 32%/ Flexible

Sphere/ 0.15m^2/ 40%/ Flexible

Sphere/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 40%/ Rigid

1 Sided/ 0.15m^2/ 32%/ Flexible/Cant=0deg

1 Sided/ 0.15m^2/ 40%/ Flexible/Cant=0deg

1 Sided/ 0.15m^2/ 32%/ Rigid/Cant=0deg

1 Sided/ 0.15m^2/ 40%/ Rigid/Cant=0deg

1 Sided/ 0.15m^2/ 32%/ Flexible/Cant=90deg

1 Sided/ 0.15m^2/ 40%/ Flexible/Cant=90deg

1 Sided/ 0.15m^2/ 32%/ Rigid/Cant=90deg

1 Sided/ 0.15m^2/ 40%/ Rigid/Cant=90deg

Figure 8: Mars Surface: 1600 W-hr Load Energy Requirement

75

80

85

90

95

100

105

50 100 150 200

Energy Storage W-hr/kg

% of Baseline Mass 

Cube/ 0.15m^2/ 32%/ Rigid

Cube/ 0.15m^2/ 40%/ Rigid

Sphere/ 0.15m^2/ 32%/ Flexible

Sphere/ 0.15m^2/ 40%/ Flexible

Sphere/ 0.15m^2/ 32%/ Rigid

Sphere/ 0.15m^2/ 40%/ Rigid

1 Sided/ 0.15m^2/ 32%/ Flexible/Cant=0deg

1 Sided/ 0.15m^2/ 40%/ Flexible/Cant=0deg

1 Sided/ 0.15m^2/ 32%/ Rigid/Cant=0deg

1 Sided/ 0.15m^2/ 40%/ Rigid/Cant=0deg

1 Sided/ 0.15m^2/ 32%/ Flexible/Cant=90deg

1 Sided/ 0.15m^2/ 40%/ Flexible/Cant=90deg

1 Sided/ 0.15m^2/ 32%/ Rigid/Cant=90deg

1 Sided/ 0.15m^2/ 40%/ Rigid/Cant=90deg
NASA/TM—2009-215803 14 

Note that all in-space operations were not considered. The suit power system will likely have 

requirements for Venus swing-by (near Venus orbit distance) and have higher solar energy possibilities. 

There may be operational restrictions on EVAs (shadow side of vehicle?). The data presented for the 

power system for in-space use are applicable only to the near Earth orbit (i.e., 1 AU). Other analyses are 

needed for Venus and Mars orbit operation extremes. 

The only metric considered at this time was overall power system mass because this is main driver in 

considering solar augmentation to the power system. 

1.11 Conclusions 

Solar augmentation may provide as much as 8 percent for LEO and 13 percent for beyond GEO InSpace suit power system mass savings over an all-energy storage power system (assuming a minimum of 

100 W-hr/kg energy storage is available for the near term). A far-term in-space suit has the potential of 7 

percent for LEO and 12 percent for beyond GEO/near 1 AU (assuming 200 W-hr/kg energy storage is 

available for the mid to far term). 

Solar augmentation can provide as much as 11 percent for the Moon surface suit and 5 percent for the 

Mars surface suit power system mass savings over an all-energy storage power system (assuming a 

minimum of 200 W-hr/kg energy storage is available, which is assumed possible for the mid to far term). 

Because the benefit of using solar augmentation relative to a likely all-energy storage power system is 

relatively low and since improvements in energy storage can likely outpace any benefits of using solar 

energy in this application, it is not considered feasible to use solar augmentation for the EVA suit for near 

term applications. If the benefit of using solar augmentation had been at least 25 percent, then it may have 

had merit for practical application. Although it is possible that some benefit exists for using solar 

augmentation in the mid to far term surface suit applications, the technology pathway for applicability to 

Mars operations is not viable (due primarily to the reduced solar illumination). 

In addition to these quantitative analyses, solar augmentation for the advanced EVA suit has 

numerous fundamental issues that discourage its use. 

1) Operation during Lunar or Mars night or for long periods in craters or where high shadowing 

occurs obviates the use of solar panels. 

2) The desire to be able to conduct mission planning with guaranteed power and energy levels is 

enhanced with fixed energy storage rather than fluctuating solar power. One cannot plan a mission on 

averages and estimates of solar power, one plans a mission based on fixed power. If the desire is to 

simply have “extra” energy beyond which you really need, then solar may be acceptable. 

3) Suit complexity is increased significantly when both energy storage and energy generation are 

used. This reduces reliability and increases number of spare parts required as well as net development and 

life cycle cost. 

4) Inability to have significant areas of solar cell coverage on the suit will reduce likely power levels 

from the optimistic projected area assumed in this trade study. 

5) Clearance and ergonomic issues with pole mounted options imply this approach is infeasible, even 

though desirable from a power viewpoint. 

6) High probability of damage of pole mounted options for in-space use likely prohibits their use 

(surface use only!). 

1.12 Additional Information: Tracking Solar Panel Concept 

After the main analysis of this report, another option was devised in which a one sided solar panel 

was deployed on a boom and articulated on two axes using gimbal drives. The concept used derivative 

Moog drives and required special gimbal pointing and control electronics. The concept assumed that 

perfect pointing was not needed (20° accuracy). The back of the panel would be used for solar panel heat 

rejection. 
NASA/TM—2009-215803 15 

Results revealed that the penalty of power usage by the gimbal drives (which had to continuously aim 

at the Sun even with rapid astronaut movement) and electronics combined with the extra mass of these 

items made it less competitive than “fixed” cube and sphere solar panel options. 

1.13 Additional Information: Solar Fiber Concept 

Initially, a concept using a new technology of solar energy generating fibers woven into the astronaut 

suit was omitted from this study as impractical due to 1) low inherent efficiency, 2) unknown fiber 

durability (since they would be in high wear areas), 3) unknown fiber resistance to space environment 

(not space rated or qualified), 4) dust effects on energy reduction (especially on lower extremities), 5) 

maintenance complications (cleaning suit without damaging solar fibers), 6) thermal effects (cannot cover 

whole suit with fibers without increasing thermal load) and 7) lack of viability for technology pathway to 

Mars surface operations (inadequate solar energy for low efficiency fibers and limited suit area). 

However, this option will be reconsidered in the future given adequate technology projections of fiber 

efficiency. 
NASA/TM—2009-215803 16 

2.0 Trade Study 2: Comparison of Localized and Centralized 

Power System Options for the Advanced EVA Suit 

Date: August 25, 2005 

2.1 Results Summary 

The conclusion of the trade study was that only small total power system mass benefits (<11 percent) 

were possible with the localized power system option. The mass actually increases for the localized power 

system option when these systems are required to have maximal commonality (to reduce mission logistics 

mass/cost). Mass and volumetric concerns (at hands and feet) also limit the localized option to lower 

power levels and shorter durations. Centralized power system/energy storage with power distribution 

cables is the best option for a combined metric of total power system mass, hand/foot volume and mass, 

and logistics. A recommendation is made of minimizing unique power distribution cables for the 

centralized power system option by connecting fixed segments (~0.5 m) to reduce spares inventory and 

logistic cost. 

2.2 Introduction 

The purpose of this trade study is to evaluate the benefits of distributing the power systems/energy 

storage throughout the suit to specific regional sites instead of using a centralized energy storage system 

with power distribution cables. 

This study does not consider the energy “redistribution” between the various localized energy sites 

(i.e., to enhance reliability and perform load sharing). The reason for this is that, qualitatively, it makes 

more sense to enhance reliability of a centralized energy storage site that runs the critical loads than 

depend upon complex switching and cabling needed to re-route energy throughout the suit. The other 

problem is that there would be too many fault propagation issues due to complex distribution topologies 

which would require detailed analysis to address. 

This study also does not consider the distribution of power to many sensors via wiring or RF or other 

means. It is assumed that such wiring/powering methods are handled downstream of the power 

system/energy storage units. 

2.3 Assumptions 

Table 1 shows a listing of the assumptions and variables used for this trade study. 

In order to perform this analysis, a DC power distribution cable design spreadsheet was modified to 

enable cable mass optimization/sizing of various power levels and cable lengths. Cabling and wires for 

instrumentation and control were excluded from consideration because of lack of definition of these 

elements. The power cables were designed with filler fibers to assist to strength, 

coatings/sheath/insulation to protect from shorting/wear/micrometeorites, and connectors. All distribution 

cables assume two separate power wire pairs (positive and ground) and a backup redundant wire pair 

(assumed manually switchable). Copper cable wires were assumed because copper optimizes to be a 

slightly lower mass than aluminum for this application. An insulator safety factor was assumed to assure 

adequate shorting protection. A cable power safety factor was applied to the peak power level to 

adequately size the wires. It was assumed that fuses activate past that maximum design power level to 

eliminate damage to the cable and loads. DC-DC converters used to provide low voltage regulated power 

for electronics applications (representing a small portion of site power) were included at each power load 

site. Note that specific powers (i.e.,W/kg) are applied to the input power of each component. 
NASA/TM—2009-215803 17 

TABLE 1.—ASSUMPTIONS AND VARIABLES USED FOR THIS TRADE STUDY 

Power System Parameters 

Voltage Type Direct Current 

Power System Voltage Level V 28 

Nominal Energy Storage Depth of Discharge percent 80 

Total Site Power Into DCDC Converter percent 20 

Nominal Energy Storage Specific Energies Considered W-hr/kg 50,100,150,200 

Energy Storage Durations Considered hr 2,8 

Total Load Levels Considered (Exact sized only) W 100,200 

Cable Parameters 

Cable Wire Material Copper 

Cable Wire Density gm/cm3 8.9 

Cable temperature K 279 

Cable Power Safety Factor 2 

Number of Redundant Cable Pairs/Transmission Line 1 

Cable Jacket Filler Density gm/cm3 1.7 

Insulator Density gm/cm3 1.7 

Cable Jacket Density gm/cm3 1.7 

Insulation Dielectric Strength Safety Factor 2 

Contingency Parameters 

Backpack Contingency Energy Storage Specific Energy W-hr/kg 220 

Backpack Contingency Energy Storage Load Level W 25 

Backpack Contingency Energy Storage Load Duration min 30 

Component Specific Powers 

DCDC Converter Specific Power W/kg 380 

Power Distribution Unit Specific Power W/kg 50 

Charge/discharge Regulator Unit Specific Power W/kg 50 

Thermal System Specific Power W/kg 100 

Nondistribution Cable Specific Power W/kg 1000 

Component Efficiencies 

DCDC Converter Efficiency percent 90 

Power Distribution Unit Efficiency percent 99 

Charge/discharge Regulator Unit Efficiency percent 90 

Mass Estimation Parameters 

Connector Mass (Percentage of Cable Mass) percent 10 

Energy Storage Structure (Percentage of Energy Storage Mass) percent 10 

Other Power System Structural Mass (Percentage of the DCDC and cable masses) percent 10 

The thermal energy system specific power (for its mass sizing) was based on the maximum power 

level encountered during the mission timeline. Cable masses internal to the power systems (e.g., between 

the distribution assembly and charge/discharge unit and between the charge/discharge regulator unit and 

the energy storage unit) which were not primary power distribution cables were assumed to have a nonoptimized mass based on a fixed specific power. The only emergency energy storage was for the suit back 

load location. 

Table 2 shows the power level and cable distance for each of the load sites. 

These sites were divided into the following regional locations: left hand, right hand, left foot, right 

foot, helmet (top), suit front (chest) and backpack. Thus, for a centralized power system, the power 

distribution cables would extend only to these particular locations. Loads (sensors, heaters, computers and 

lights) nearest a region were assumed to be powered by the same primary power distribution cable. 

However, the cables that go from the power system/energy source to each separate load were not 

considered since it is not clear at this time how many of these loads there are, what their exact power 

needs are and even if they may be powered without cables. This was assumed to have a negligible effect 

on a relative comparison because both centralized and localized power subsystems would be required to 

carry the same mass for powering these loads. Table 2 lists the assumed distances between the regional 

loads where the distance is from the center of the backpack to the furthest point for each load site. Also 

shown are the “maximally-common” and “reduced-commonality” power assumptions for each load site. 
NASA/TM—2009-215803 18 

The power values were selected based on estimated low and high values (the medium value is derived 

from these extremes). 

TABLE 2.—POWER LEVEL AND CABLE DISTANCE FOR EACH OF THE LOAD SITES 

 Cable 

distance 

(m) 

Maximally Common 

Energy Storage Units 

Reduced Commonality 

Energy Storage Units 

Total Power 

Level (W) 100 100 100 200 200 200 100 100 100 200 200 200 

Distribution 

Level Low Medium High Low Medium High Low Medium High Low Medium High 

Load Location 

Left hand 1 5 10 15 10 20 30 5 10 15 10 20 30 

Right hand 1 5 10 15 10 20 30 5 10 15 10 20 30 

Left foot 1.5 5 10 15 10 20 30 3.5 7.75 12 7 15.5 24 

Right foot 1.5 5 10 15 10 20 30 3.5 7.75 12 7 15.5 24 

Helmet 0.66 5 10 15 10 20 30 1 3.5 6 2 7 12 

Back 0.33 80 55 30 160 110 60 80 55 30 160 110 60 

Chest 0.5 5 10 15 10 20 30 2 6 10 4 12 20 

For “maximally common” remote energy storage/load sites, it was assumed that all sites other than 

the backpack had to have the same energy storage size/capacity (since they were likely to be similar 

anyway). Thus, only two unique energy storage sizes need to be inventoried. On the downside, the effect 

of this commonality is to oversize the local energy sources compared to what they require if “exactly” 

sized. For the nominal 100 W system, commonality forces the total power level to go to 110 to 120 W 

depending on the level of energy required at each regional site. For a nominal 200 W system, the range is 

220 to 240 W. 

The “reduced commonality” values were based on the assumption that the only common local power 

system sizes were both hand sites and both feet sites. Besides those common elements, each load site was 

assumed to have different and unique power requirements. Thus, this results in five unique hardware 

types, with the back power level highest, then the hand, foot, chest and helmet. Although it is possible 

that the larger energy storage/power systems (backpack) can be composed of smaller sub-elements, this 

aspect was considered irrelevant to this trade study. 

2.4 Commonality Assessment 

The concept behind commonality is that common hardware unit types enable the reduction of spares 

quantities. For example, for low failure rates (<1 per “tour” or “stay” or logistics supply interval or 

deployment), it is still required that whole replacement units for each unique hardware type be carried 

(not by the astronaut, but in the outpost or vehicle). For common hardware units, the fractional number of 

spares can be combined to whole numbers. If a power source hardware unit has a 5 percent failure rate per 

tour, but there is only one such unit, a whole spare must be still be carried at the base. If there are 20 

different units with this failure rate, then 20 separate spares must be carried. However, if the 20 units are 

common rather than different, then only one spare (or perhaps up to three depending on probability 

models and failure distribution modes) needs to be carried at the exploration logistics supply base. Thus, 

launch mass and spares cost is minimized. 

For the nominal centralized energy storage with distribution cable approach, there is only one 

nominal energy storage system (although it could possibly be subdivided in some manner for lower level 

repairs, this is not considered here). However, two approaches for the power distribution cables must be 

considered. One approach has fixed length cables for going to the hand sites, other different fixed length 

cables to the foot sites and other separate unique length cables for the chest, helmet and back. This means 

one power system type and five unique cable types must be accommodated for spares (a total of 6 unique 

“items” where “items” are the highest level replacement unit). 
NASA/TM—2009-215803 19 

Alternatively, a higher commonality approach would be to select a fixed cable segment length which 

can be connected to other lengths to fit all these applications (except perhaps the backpack power cable). 

Thus, three .5 m cable segments could be connected to form the entire cable to a foot site, two .5 m cable 

segments could be connected to form the entire cable to the hand site and the remaining load sites could 

simply use one .5 m cable segments. This means one power system type and one unique cable type must 

be accommodated for spares (a total of two unique “items”). The drawback is that the “non-unique” 

power cables are not mass optimized for their intended application (power level) and also may have some 

cumbersome “slack” due to the standardized segment lengths. The former issue seems to be a minimal 

concern since cable mass did not vary significantly from low to high power distribution levels (<10 

percent of the cable mass). The later may be a concern for total power system mass because the slack is 

really unneeded mass, besides being an unnecessary volume burden to deal with. 

For the localized energy storage/power system sites, the reduced commonality option has five unique 

types (sizes) of power units and one backpack power distribution power is counted for this comparison (a 

total of 6 unique “items”). 

The maximally common option has two unique types of units and one backpack power distribution 

power is counted for this comparison (a total of three unique “items”). 

Assessing these commonality issues at a very preliminary level suggests that the two options, from a 

logistics point of view, which are unfavorable would be the centralized power system with unique power 

distribution cables (six “items”) or the reduced commonality localized power system option (six “items”). 

The two options with good logistics characteristics are the centralized power system with segmented 

power distribution cables (two “items”) and the maximally common localized power system (three 

“items”). 

2.5 Analysis Results 

Examination of results for the nominal centralized energy storage with power distribution cables 

option showed that the mass of the distribution cable (external to the backpack unit) make up for between 

1 to 10 percent of the total power system mass, depending upon the energy storage specific energy, power 

level, and load durations. Given this initial observation, it was highly unlikely that significant total power 

system mass savings were possible with the localized power system option (in which no such distribution 

cables were used). 

Figure 1 shows the relative total mass of the localized energy storage units/power systems compared 

to the centralized energy storage/cable power distribution system. Several generic energy storage systems 

with assumed specific energies are shown. It is apparent that some level of mass improvement was 

possible by eliminating the distribution cables if the localized energy sources are not driven to be as 

common as possible. As much as 9 percent total mass savings was possible for low total power level and 

duration. Higher power levels and durations have less total mass savings (<5 percent). The greatest 

savings occurs for higher energy storage specific energies (because the cable mass becomes a higher 

proportion of the total power system mass since the energy storage mass is reduced). Total savings of less 

than 20 percent indicate that it is likely that the benefit of attempting to localize the energy storage is not 

worth the development effort of localized power sources. 

If one considers the commonality aspect, then the localized option total mass savings disappear and 

actually become worse than the centralized approach. Because the localized energy units would need to 

be made the same size to enhance the commonality (reduce logistics mass), then clearly the mass of the 

total power system must be greater than if the exact power requirements are used for energy storage 

sizing. The lower power levels and durations are least affected by variations in the energy storage units 

sizing. Also, the higher the specific energy of the energy storage units the lower the adverse mass penalty. 

Still, the unavoidable result is a higher mass of from 1 to 18 percent from the nominal centralized/power 

cable distribution option. This definitely indicates that it is not “advanced EVA suit mass beneficial” to 

have the “maximally common” localized energy approach. 
NASA/TM—2009-215803 20 

Figures 2, 3 and 4, respectively, show the total power system mass at the hand, foot and helmet 

locations for the remotely located energy storage options. These values are preliminary estimates of mass 

based on many assumptions. Only the reduced commonality cases are presented because they provide the 

lowest mass levels (most beneficial for the advanced EVA suit mass). It is apparent that higher power 

levels and durations at the hand or foot sites increase the mass burden for the astronaut to carry 

significantly. 

Excessively cumbersome masses per hand or foot are assumed to be >3 kg for lunar operation 

(3kg*2.2lb_earth/kg*1.0lb_moon/6.0lb_earth=1.1lb_moon) and >1.5 kg for Mars operation 

(1.5*2.2lb_earth/kg*1.0lb_mars/3.0lb_earth=1.1lb_mars). Based on this, a 50 W-hr/kg energy storage 

generally is not feasible for remotely located energy storage units unless low durations and power levels 

are considered (for Mars or Moon). For the more likely available 100 W-hr/kg energy storage, it is not 

feasible for 8 hr duration and high site power level. For Mars operation, all high site power levels exceed 

the desirable site masses; only lower site power levels will meet the assumed mass limit. 

Figures 5 and 6, respectively, show the length of one side of a cube in centimeters of a theoretical 

power system at the hand and foot locations for the remotely located energy storage options. The various 

energy storage units (with specific energies of 50, 100, 150 and 200 W-hr/kg) have assumed specific 

volumes (100, 200, 300 and 300 W-hr/liter, respectively). Only the reduced commonality cases are 

presented because they provide the lowest mass levels and lower volume levels. The idea of the figures is 

to obtain a conceptual idea of the amount of movement impact of having the remotely located power units 

at the hand and foot sites. It is assumed that the exact shape of the units may be more conformal or thinner 

(if possible), a cube is the worst case shape for such a device attached to the hand or foot. The size of the 

power source is seen to get rather large for higher power levels and longer durations. A practical limit 

seems to be a volume of (9 cm3

) with rectangular or conformal shape (a cube would be infeasible due to 

clearances). 

Figure 1: Total Power System Mass Ratio (Remotely located energy 

storage/Centralized energy storage)

0.9

0.95

1

1.05

1.1

1.15

1.2

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

50W-hr/kg,Maximally common

100W-hr/kg,Maximally common

150W-hr/kg,Maximally common

200W-hr/kg,Maximally common

50W-hr/kg,Reduced commonality

100W-hr/kg,Reduced commonality

150W-hr/kg,Reduced commonality

200W-hr/kg,Reduced commonality
NASA/TM—2009-215803 21 

Figure 2: Localized Power Source Total Mass (Hand Site)

0

1

2

3

4

5

6

7

8

9

10

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

Hand Site Power System Mass (kg)

50w-hr/kg

100w-hr/kg

150w-hr/kg

200w-hr/kg

Figure 3: Localized Power Source Total Mass (Foot Site)

0

1

2

3

4

5

6

7

8

9

10

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

Foot Site Power System Mass (kg) 

50w-hr/kg

100w-hr/kg

150w-hr/kg

200w-hr/kg
NASA/TM—2009-215803 22 

Figure 4: Localized Power Source Total Mass (Helmet Site)

0

1

2

3

4

5

6

7

8

9

10

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

Helmet Site Power System Mass (kg) 

50w-hr/kg

100w-hr/kg

150w-hr/kg

200w-hr/kg

Figure 5: Localized Power Source Total Size (Hand Site)

0

2

4

6

8

10

12

14

16

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

Length on Cube side (cm) 

50w-hr/kg

100w-hr/kg

150w-hr/kg

200w-hr/kg
NASA/TM—2009-215803 23 

Figures 7, 8 and 9 shows a 3D model with three sizes of cube: 3, 9 and 16 cm. Clearly the size 

becomes cumbersome above (9 cm3

). Figure 9 is indicative of the volume for long duration, high power 

and low specific energy storage/specific volume. 

Examination of the sensitivity of the total power transmission cable mass to various power cable 

safety factors (i.e., oversizing the wire for much higher loads than planned on) showed that increasing the 

factor from 2 to 3 increased the total cable mass by about 10 percent. A power cable safety factor of 2 was 

felt to be justified given the current knowledge of the site load levels. 

During power distribution cable mass optimization, the wire gauge was allowed to be as small as 

AWG 38 (0.010 cm diameter wire) for very low power loads to AWG 19 (.091 cm diameter) for the 

higher power loads (typically the backpack). The average wire gauge for the power distribution cables to 

the hand and foot sites was AWG 30 (0.025 cm diameter). The wire mass optimizer can be set to have a 

minimum allowable wire gauge. When the limit to the wire diameter was set to AWG 22 (.0643 cm 

diameter), the total mass of the centralized power system with distribution cables increased 0.1 to 2.5 

percent (depending on the case). Figure 10 shows the total power system mass comparison when the wire 

diameter was limited (compare with Fig. 1). The primary difference is that it is clear that fixing the wire 

radius to AWG 22 causes a plateau of cable mass, thus the exact power level at the distribution site no 

longer affects the result (as long as the total power level is the same). Also, for the lowest distribution 

power/duration and high specific energy storage units (case 2-hr-100 W-low distributed power level), the 

high commonality remotely located option is lower than the centralized option. This indicates that larger 

minimum wire radius limits may make the remotely distributed energy option lower mass than the 

centralized energy option, although it will likely never save the groundruled amount of 20 percent of total 

power system mass (to make the approach worth the effort and risk). Even for the reduced commonality 

remote distributed energy option, which already are lower in mass than the centralized energy option, 

although there is a slight improvement in the total power system mass, the 20 percent total power system 

mass savings groundrule is not met (a maximum reduction of 11 percent is seen for low duration/power 

level options). 

Figure 6: Localized Power Source Total Size (Foot Site)

0

2

4

6

8

10

12

14

16

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

Length on Cube Side (cm) 

50w-hr/kg

100w-hr/kg

150w-hr/kg

200w-hr/kg
NASA/TM—2009-215803 24 

Figure 7.—3 cm cube 

Figure 8.—9 cm cube 

Figure 9.—16 cm cube 
NASA/TM—2009-215803 25 

2.6 Conclusion 

A centralized power system with power distribution cables is the best option of those considered. This 

option modified to enhance commonality in the distribution cables may have somewhat higher mass and 

volume but would have lower spares requirements. Localized power sources with reduced commonality 

may provide somewhat lower mass, but this benefit is outweighed by considering 1) commonality 

(spares/logistics), 2) power source volume/clearance issues, 3) localized appendage power source mass, 

4) recharge or swap out complexity of multiple units and 5) development cost of multiple power source 

types. Localized power source with maximal commonality lose the total mass benefits while retaining the 

other negative attributes (except for commonality and logistic considerations). 

The aspect that cannot be assessed by this trade study is the difficulty in integrating the cable into the 

suit. Cables are not exposed (outside the suit protective garment) so they must be fed within that layer in 

some manner. Qualitatively, it seems that feeding one cable from backpack to each hand and foot is not a 

complex matter; it is wiring the entire suit to each sensor that is complex. It is unlikely that any future 

suits would have woven “main” power cables since the desire to make these kinds of cables a replacement 

unit is high (rather than replace an entire “woven” garment or “woven” power thread). 

Power cables are likely to remain the only method of distributing power levels above 1 W. There may 

be some possibility to distribute power below that level using other methods (e.g., radio frequency, 

induction), but these must be addressed in detail based on the power needs of the sensors. However, if a 

power cable is used for hand and feet, then perhaps the same cable (and added RF/other power emitting 

methods/electronics) can be used to provide power to miscellaneous very low power sensors using these 

other methods without the complexity of individual wiring to them. 

Figure 10: Total Power System Mass Ratio for Wire Radius Limit=AWG22 

(Remotely located energy storage/Centralized energy storage)

0.85

0.9

0.95

1

1.05

1.1

1.15

1.2

2hr-100W-low distr

2hr-100W-med distr

2hr-100W-high distr

8hr-100W-low distr

8hr-100W-med distr

8hr-100W-high distr

2hr-200W-low distr

2hr-200W-med distr

2hr-200W-high distr

8hr-200W-low distr

8hr-200W-med distr

8hr-200W-high distr

50W-hr/kg,Maximally common

100W-hr/kg,Maximally common

150W-hr/kg,Maximally common

200W-hr/kg,Maximally common

50W-hr/kg,Reduced commonality

100W-hr/kg,Reduced commonality

150W-hr/kg,Reduced commonality

200W-hr/kg,Reduced commonality
NASA/TM—2009-215803 26 

3.0 Trade Study 3: Umbilical-Only Power Versus Suit-Mounted 

Power for the Advanced EVA Suit 

Date: September 23, 2005 

3.1 Results Summary 

Based on the Crew Exploration Vehicle design, umbilical lengths of up to 20 m can be expected. The 

maximum mass savings of a suit-mounted energy storage power system versus an umbilical-only power 

system (for a 20 m umbilical) is about 4 kg for one suit (includes power cable in umbilical and any suitrelated power system). The most favorable suit-mounted energy storage options are for low power <50 W 

and short duration (<1 hr) with recharging. The primary energy (non-rechargeable) storage system has the 

most mass savings although it is limited by being a onetime use (this may be acceptable depending upon 

the number of planned/groundruled EVAs per mission). If other umbilical resources are considered (O2, 

cooling, power, etc), the launch mass savings (based on a Shuttle EMU suit) increase to at least 40 kg for 

one suit (mass due to umbilical infrastructure/ mechanisms and O2/cooling fluids inside the umbilical are 

omitted here, but would increase the savings considerably). Suit-mounted energy storage has enhanced 

mobility and reduced encumberness as compared to umbilical-only power options. Umbilical-only power 

is the best selection for in-space pressurized (in-cabin) suits. Surface suits favor suit-mounted energy 

storage with short umbilicals for recharge (longer umbilicals are not favored for surface suits because of 

dust, wear, abrasion and severing concerns, as well as reduced mobility and increased encumberness). 

3.2 Purpose 

The purpose of this trade study is to evaluate the benefits of using suit-mounted energy storage 

systems for the advanced EVA suit as opposed to connecting the suit to a power source via an umbilical. 

3.3 Past/Current Umbilical Data 

Umbilicals can be used for in-space EVA suits as well as surface suits. Table 1 displays the typical 

uses and characteristics of past/current EVA suits with umbilicals. Tether data is also included. 

Umbilicals can provide various resources to an EVA suit including cooling, oxygen, 

communications/sensor linkups, mechanical linkage/support and power. Umbilicals may either provide 

nominal power to the suit during a sortie or be used to recharge a suit attached nominal power system. 

In designing a power umbilical, the maximum length is important to know. This is dependent upon 

the application. Which resources are carried in the umbilical is important since this affects size, flexibility 

and mass. Past anecdotes illustrate that very long (>18 m) umbilicals (which carry O2 and cooling) were 

too cumbersome. The Skylab umbilical designs were theoretically able to extend to 36 m (Skylab data is 

incomplete thus it is not clear what was the implemented umbilical length/characteristics). Such long 

umbilicals require that one astronaut is dedicated to perform the EVA and the other to handle (deploy or 

retract) the umbilical. One version of the Orlan suit had a power/communications-only umbilical of as 

much as 25 m. The Gemini suit had a 15 m umbilical which carried O2 and cooling primarily, but was 

considered very bulky and was difficult to return to the airlock. 

It is assumed that the umbilical is connected near one point (nominally the egress point or airlock) 

and the length of the umbilical is the limit to the astronaut travel. It is possible that multiple EVA 

umbilical external connection sites could be included in the CEV design (assuming a small EVA power 

system to handle movement between EVA sites and a moderate length umbilical to handle most of the 

sortie activity), but this is considered unlikely due to the added complexity and launch mass balanced 

against the infrequent use of such ports. For in-space, pressurized application, it is connected to one of 
NASA/TM—2009-215803 27 

possibly many ports on the inside. For surface operation, it is either connected to a port near the airlock or 

recharge station, or to a rover port. 

TABLE 1.—PAST UMBILICAL/TETHER DATA 

Program Type Location Purpose Length, 

m 

Note 

Voskhod Umbilical In-Space Comm/O2 5 Berkut suit 

Gemini Umbilical In-Space Comm/O2/cooling/sensors 8 to 15 Tangling problems at 15 m length, bounced off 

Agena 

Apollo surface Umbilical In-Space Cooling 3 Used to share resources between suits in an 

emergency 

Apollo surface Umbilical In-Space Power? ? Connection with Rover? 

Apollo Umbilical In-Space Comm/O2/cooling/sensors/mech 2.5 Apollo 15 

Skylab Umbilical In-Space Comm/O2/cooling/sensors/mech 18, 36 two 

in series 

Apollo pressure suit and astronaut life support 

assembly (emergency, O2, pressure control unit, life 

support umbilical) 

Skylab Umbilical In-Space ? 14 In-pressurized cabin use 

Shuttle EMU (IEU) Umbilical In-Space Comm/O2/cooling/sensors/power/fluids 7 Used in unpressurized 

Shuttle EMU (SCU) Umbilical In-Space Comm/O2/cooling/sensors/power/fluids 3.3 Used in airlock (pressurized) 

Shuttle (not suit) Umbilical In-Space Power/comm. 3, 10 To power satellites 

Orlan Umbilical In-Space Power/comm. 25 

Orlan Umbilical In-Space O2/fluids/cooling 3 

Gemini Tether In-Space Mechanical 8, 33, 47 

Apollo Tether In-Space Mechanical 9 

Shuttle EMU Tether In-Space Mechanical 1 to 18 

3.3.1 Emergency Backup Power 

Another consideration is whether the EVA suit has emergency backup power or not. Generally, in 

past and current EVA suits, if umbilicals are used, the EVA suit has an emergency backup O2 system to 

enable the astronaut to return a pressurized environment. However, only the most recent in-space suits 

such as the Orlan and Shuttle EMU have emergency backup power. The reason for this is that past suits 

seem to require less electronics and be more mechanical. It is not clear whether this can be considered 

true for future suits since communications at least is a critical function that seems to require some power 

if the umbilical fails. However, it may be groundruled that, for a specific EVA suit application, 

emergency backup power is not needed. 

3.3.2 Contingency EVA 

The in-space, external EVA suit has been previously groundruled (by the AEVA program) for 

contingency EVAs only (as opposed to assembly and nominal operations). A contingency EVA is 

performed as an emergency to save the spacecraft and/or its crew. Thus, transfer between orbiting 

vehicles or transfer between compartments in the same vehicle, emergency repair procedures on 

equipment (orbiting vehicle external solar array or radiator deployment, antenna deployment, thermal tile 

inspection/repair) are permitted but nominal assembly or scientific data acquisition/data retrieval are 

excluded. Table 2 describes typical contingency EVA applications (generic and historical). Duration, 

power level (i.e., may be lower than nominal EVAs), mobility requirements (e.g., may need to detach 

umbilical at location different from egress point) of each contingency EVA cover a wide range. 

Since contingency EVAs should not happen frequently (given a well designed vehicle), it is possible 

to consider one time use suit-mounted non-rechargeable energy storage systems. Replacement energy 

storage units (and other resources) could be carried for a certain threshold quantity of contingency EVAs 

(to be determined by analysis). 
NASA/TM—2009-215803 28 

TABLE 2.—EXAMPLES OF CONTINGENCY EVAS 

Salyut6 Remove the KRT-10 antenna from Salyut 6’s aft docking port 

Apollo 9 Demo of contingency EVA transfer between the Apollo Lunar Module and Command Module (in case IVA blocked) 

Skylab Failure of solar panel, loss of power and cooling, released remaining solar array, erected sunshade 

STS–6 Demo of payload bay door closure 

STS–37 EVA to deploy GRO high gain antenna which could not be deployed after the GRO was raised from the payload bay 

STS–57 Secure carrier antennae that would not latch on retraction 

Mir 1987 Removed obstruction (most likely a sheet) that was fouling the docking between Mir and Kvant 

Mir 1990 Fixed insulation blankets on the Soyuz landing capsule. Indentified the faulty hatch problems/features and forced it closed. 

ISS 2001 A contingency EVA was ordered to clear an obstruction in the docking port 

Example Adjust a displaced insulation blanket 

Example Adjust a displaced micrometeoroid shield 

Example Check the mating of a cable connector that might have malfunctioned during operations. 

3.3.3 Applications of Umbilicals 

The primary reason for a power umbilical is to provide the main power for an in-space, external EVA 

suit, eliminating the need for a suit-mounted energy storage system. 

A power umbilical can also be used to recharge the EVA suit energy storage (either for contingency 

or nominal operations). This can be used for in-space or surface suits (in pressurized or unpressurized 

environments). For surface suits, the same method as in-space recharging can be used (in an 

unpressurized airlock or external location or while in transit on a rover). These likely have short lengths 

(<3 m). 

Use of long umbilicals for powering or recharging surface suits is likely impractical due to dust 

adhesion, mechanical abrasion/wear, risk of severing/puncture due to sharp rocks, mass and mobility 

limitations. 

In a pressurized in-space cabin use, a powered suit is most likely to be powered via a short umbilical. 

3.3.4 Anticipated Sortie Duration and Umbilical Length 

Regarding sortie duration, for in-space, exterior use, Skylab assumed a nominal translation rate of 

about 0.3 ft/sec which means that 20 m can be covered in 1 min. At even 5 times this value, this implies 

significant time at the work site if the total sortie time is only 1 hr. Longer sortie times than 1 hr (and, 

especially, unlimited ones) would be favored. 

Regarding likely umbilical length, since the release of the Exploration Systems Architecture Study 

report, it is possible to view the likely maximum in-space distance which must be traversed in a 

contingency EVA. Assuming only the Crew Exploration Vehicle and Lander elements (i.e., the Departure 

Stage is jettisoned), the maximum umbilical distance is about 20 m. If the Departure Stage is included, the 

EVA traverse distance goes to up to as much as 60 m. This later case is considered unlikely in that 

contingency operations on the Departure Stage would be prohibited due to safety concerns (i.e., a mission 

abort would most likely occur). 

3.4 Analysis Method 

The primary goal of the analysis was to quantify the power system mass for both the suit-mounted 

energy storage option and the umbilical-only power cable option. This would enable a comparison of the 

suit mass, launch mass and provide data to gauge how encumbered the astronaut would be with each 

option. The same EVA suit power system model was utilized that was developed for the previous trade 

studies. This model was modified to include an umbilical (using cable sizing relationships from the suit 

power system). Various umbilical lengths (1, 5, 10, 15, and 20 m), power levels (25, 50, 100, 150, and 

200 W) and power durations (1, 2, and 4 hr) were analyzed. Various energy storage specific energies were 

considered (50, 100, 150, and 200 W-hr/kg). (Note that primary energy storage for either emergency or 

nominal usage were assumed to have a 220 W-hr/kg specific energy. Both primary (non-rechargeable) 
NASA/TM—2009-215803 29 

and secondary (rechargeable) energy storage were assessed for the suit-mounted energy storage options. 

Cases with and without emergency backup energy storage (using primary/non-rechargeable batteries) 

were considered. The recharge method was designed to be based on electrical recharge only (of batteries), 

not by recharging fuel cells via pressurized lines (i.e., fuel cells are ground ruled out for near-term AEVA 

in-space application). 

The basic method of generating mass data for the various options was to size the umbilical cables for 

the various power levels and lengths. Then, the nominal power system was sized for various power levels, 

durations and specific energies. The emergency energy storage system sized as appropriate. These values 

were combined to generate all the permutations of cases presented in the analysis results. 

3.5 Power System Assumptions 

Power system assumptions are listed in Table 3. For the primary/non-rechargeable battery energy 

storage options, a discharge regulator is used, however for rechargeable battery energy storage options, a 

charge/discharge regulator is used. Each component has a different efficiency which affects the sizing of 

the energy storage system. 

TABLE 3.—POWER SYSTEM ASSUMPTIONS 

Power System Parameters 

Voltage Type Direct Current 

Power System Voltage Level V 28 

Nominal Energy Storage Depth of Discharge percent 80 

Total Site Power Into DCDC Converter percent 20 

Cable Parameters 

Cable Wire Material Copper 

Cable Wire Density gm/cm3 8.9 

Cable temperature K 279 

Cable Power Safety Factor 2 

Number of Redundant Cable Pairs/Transmission Line 1 

Cable Jacket Filler Density gm/cm3 1.7 

Insulator Density gm/cm3 1.7 

Cable Jacket Density gm/cm3 1.7 

Insulation Dielectric Strength Safety Factor 2 

Contingency Parameters 

Backpack Contingency Energy Storage Specific Energy W-hr/kg 220 

Backpack Contingency Energy Storage Load Level W 25 

Backpack Contingency Energy Storage Load Duration min 30 

Component Specific Powers 

DCDC Converter Specific Power W/kg 80 

Power Distribution Unit Specific Power W/kg 50 

Charge/discharge Regulator Unit Specific Power W/kg 50 

Thermal System Specific Power W/kg 100 

Nondistribution Cable Specific Power W/kg 1000 

Component Efficiencies 

DCDC Converter Efficiency percent 95 

Power Distribution Unit Efficiency percent 99 

Charge/discharge Regulator Unit Efficiency (Secondary) percent 80 

Discharge Regulator Unit Efficiency (Primary/Emergency) percent 90 

Mass Estimation Parameters 

Connector Mass (Percentage of Cable Mass) percent 10 

Energy Storage Structure (Percentage of Energy Storage Mass) percent 10 

Other Power System Structural Mass (Percentage of the DCDC and cable masses) percent 10 

For the umbilical-only sizing, only DC-DC converters, cables and switchgear are sized. Note that the 

DC-DC converter specific power is based on its output power (i.e., input power divided by 5.0). 
NASA/TM—2009-215803 30 

The limiting wire gauge of AWG 22 is assumed for the internal electrical system. 

For the internal electrical system of the suit, an “average” distribution of power is assumed to size the 

cables (see Table 4). Local loads for total suit power levels of 25, 50, 100, 150 and 200 W are presented. 

TABLE 4.—DISTRIBUTION OF POWER 

 Power Level at Each Load Location 

 (W) 

Load location Cable distance (m) 25 50 100 150 200 

Left hand 1 2.5 5 10 15 20 

Right hand 1 2.5 5 10 15 20 

Left foot 1.5 1.94 3.875 7.75 11.625 15.5 

Right foot 1.5 1.94 3.875 7.75 11.625 15.5 

Helmet 0.66 0.88 1.75 3.5 5.25 7 

Back 0.33 13.74 27.5 55 82.5 110 

Chest 0.5 1.5 3 6 9 12 

3.5.1 Analysis Limitations 

Sizing of the emergency energy storage was fixed to a specific power level and duration for all 

options. No attempt was made to determine the transfer time of the astronaut from various distances to the 

airlock. Since the actual need for emergency power is unknown at this time, a fixed assumption is the best 

approach although it can be modified in the future. This likely has little effect on the results. 

Mass and sizing data on past umbilical systems is mostly unavailable. Better data would enhance this 

analysis. 

This analysis did not consider fuel cell recharge (via pressurized lines) or swap-out of hardware. This 

is anticipated to make little difference to the results. 

Enhanced umbilical protection was not included in the umbilical power cable model. It was assumed 

that the balance of the umbilical mass would have such a mass element (if needed). Charging such mass 

to the power cable portion would increase the mass benefits of suit-mounted energy storage options by a 

moderate amount. 

3.5.2 Analysis Results 

A basic ground rule carried from past analysis is that a total power system mass savings of at least 

20 percent is required for any alternative option to be considered viable. This has been done because of 

modeling uncertainty and the desire to have adequate savings to justify new development and hardware 

risk. 

It is difficult to determine the optimum design methodology for the umbilical power cable because of 

the many competing metrics. In order to minimize the number of cases presented in this trade study, two 

specific “typical” power umbilical designs were selected. Generally, these two types were based on 

minimum mass and minimum thermal load. 

Figure 1 shows the effect of sizing the power umbilical based on cable temperature. A thinner power 

cable provides more resistance and, thus, a higher voltage drop and parasitic power loss. This raises the 

temperature of the power cable. Conceivably, the cable temperature can be as high as 120 °C, but 

realistically, this is unacceptable for umbilicals for many obvious reasons. Although the lowest mass can 

be obtained for the umbilical for high temperatures, the increased temperature of the umbilical places a 

thermal load on the cooling loop (if present in the umbilical) and on the thermal design of a power only 

umbilical. In addition, the added required power (i.e., the parasitic power loss in addition to the nominal 

power) may be a burden to the vehicle (or rover). Since it is unknown what the temperature/thermal 

requirements are at this stage of EVA suit analysis, the temperature of 25 °C was selected as a typical, 

“reasonable” value. This was done because it was near the “knee” of the curves (which is near the plateau 

of cable mass benefit versus operating temperature). This case is considered the “ideal” but is likely to not 
NASA/TM—2009-215803 31 

be “real” since the goal to reduce parasitic power (a burden on any in-space vehicle which may be 

operating on limited battery storage only during a contingency operations event) and thermal load. 

Figure 2 shows the effect of sizing the power umbilical based on cable efficiency. As seen in the 

previous figure, lower efficiency cables are thinner, lighter, have a higher voltage drop, parasitic power 

loss and higher operating temperature. Higher efficiency cables are thicker, heavier, have very low 

voltage drop/parasitic losses/operating temperature. It is difficult to select an “ideal” cable efficiency. The 

typical method to do this is to examine the curves and select a point near the “knee” of the curve (which 

represents the plateau of cable mass benefits versus efficiency. This resulted in 95 percent efficiency 

being selected for the comparisons. This case, even though it is heavier than the previous one, is more 

likely to be closer to an actual implemented power cable for the umbilical because reduced power loads 

on the in-space vehicle are very important. 

Figure 3 shows the total power system mass (including power system umbilical) for various assumed 

power levels and durations. The emergency power system is included. This figure enables one to 

understand the relative magnitudes of the masses being compared. Masses range from 3.3 to 45 kg. 

Subsequent figures will provide a better method to compare the merits of umbilical only mass versus the 

on-suit energy storage options. 

Figure 4 shows the total power system mass (including power system umbilical) for various assumed 

power levels and durations. The emergency power system is not included. Some minor mass savings 

occur by excluding emergency power (which may be viable depending on the EVA suit design). Masses 

range from 2.2 to 38 kg. Subsequent figures enable a better relative comparison of the options. 

Figures 5 through 9 show the total power system mass including any suit mounted power system 

(including emergency energy storage) and power umbilical. These figures are designed to show the 

relative benefit of the various power system configurations relative to a baseline (proposed) alternative. 

Figure 5 has a suit mounted, primary (non-rechargeable) energy storage system as the baseline 

alternative. Using the 20 percent improvement in mass metric, then it can be seen that using the 

primary/non-rechargeable energy storage option is superior to a number of umbilical length, power level 

and power duration cases. The first observation is that the primary/non-rechargeable energy storage is 

superior in terms of mass as compared to other secondary energy storage options. The obvious drawback 

of the primary/non-rechargeable storage is that it is a onetime use only. If used for contingency EVA 

operation, the number of contingencies that must be planned for affects the logic of using such a system. 

Comparison with the umbilical power–only options show, in general, that lower power levels and 

durations enable the primary/non-rechargeable energy storage option to have superior mass. Primary/nonrechargeable energy storage provides at least 20 percent mass savings over the umbilical only power 

option with load durations of 1 hr and power level <50 W over umbilicals 10 m and longer (for the 

“realistic” 95 percent efficiency umbilical design case). For 2 hr and <50 W, the savings occur over 

umbilicals of about 12 m (visual interpolation) and longer. For 4 hr and <50 W, the savings occur over 

umbilicals of about 17 m (visual interpolation) and longer. For 1 hr and >100 W, the savings occur over 

umbilical of about 16 m (visual interpolation) and longer. For 2 hr and >100 W, the savings occur over 

umbilical of about 20 m and longer. For 4 hr and >100 W, there are no savings of >20 percent in mass 

over the umbilical. 

Figure 6 has a suit mounted, 50 W-hr/kg secondary energy storage system as the baseline alternative. 

This is typical of a near-term (NiMH battery). This option has only one application that is superior in 

terms of mass. This is the 1 hr load duration, 25 W power level case. Here, the secondary energy storage 

system saves at least 20 percent mass over umbilical-only power longer than 17 m (visual interpolation). 

Because of the limited applicability in terms of load level and duration, it seems that such an alternative is 

not viable. However, duration is not critical in that the 1 hr could be the EVA time between recharging. 
NASA/TM—2009-215803 32 

Figure 1: Minimum Mass Umbilical Power Cable Option (power rich/voltage flexible 

vehicle)

0

0.5

1

1.5

2

2.5

3

10 15 20 25 30 35 40 45 50

Power cable Temperature (C)

Total power cable mass (kg) 

1m length

5m length

10m length

15m length

20m length

Figure 2: Low Thermal Load Umbilical Power Cable Mass Option (power 

limited/voltage inflexible vehicle)

0

2

4

6

8

10

12

90 91 92 93 94 95 96 97 98

Transmission Efficiency (%)

Total power cable mass (kg)

1m length

5m length

10m length

15m length

20m length
NASA/TM—2009-215803 33 

Figure 3: Total EVA Suit Power System+Umbilical Mass (with emergency energy storage)

0

5

10

15

20

25

30

35

40

45

50

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Mass (kg) 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical

Figure 4: Total EVA Suit Power System+Umbilical Mass (no emergency energy storage)

0

5

10

15

20

25

30

35

40

45

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Mass (kg) .

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 34 

Figure 5: Ratio of Power-related Mass Relative to Primary Energy/No Tether Option (with 

emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical

Figure 6: Ratio of Power-related Mass Relative to 50W-hr/kg Secondary Energy/1m 

Tether Option (with emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 35 

Figure 7 has a suit mounted, 100 W-hr/kg secondary energy storage system as the baseline 

alternative. This option can be considered equivalent to a near-term Li polymer battery. For load 

durations of 1 hr and power levels <50 W, the 100 W-hr/kg energy storage options has at least 20 percent 

less mass than umbilical-only power options with lengths 15 m and longer. Even load durations of 2 hr 

and power levels <50 W have savings for umbilical lengths of 21 m (visual interpolation) and longer. 

Load durations of 4 hr have mass savings below the accepted limit. Therefore, there is some potential for 

this option to be feasible, as long as recharging occurs every 1 to 2 hr. 

Figure 8 has a suit mounted, 150 W-hr/kg secondary energy storage system as the baseline 

alternative. This energy storage system can be considered representative of a more advanced Li polymer 

type battery. For load durations of 1 hr and power levels of <50 W, this option saves at least 20 percent 

mass of umbilical-only cases of lengths from 13 m (visual interpolation) and longer. For 2 hr and <50 W, 

the savings occur for lengths from 17 m and longer. Load durations of 4 hr do not meet the mass savings 

criteria. For 1 hr and >100 W, savings occur for lengths >20 m. Therefore, for recharging intervals of 

about 1 hr and lower power levels, this option is feasible. Even for higher power levels, as long as the 

recharge interval is about 1 hr, it is feasible. 

Figure 9 has a suit mounted, 200 W-hr/kg secondary energy storage system as the baseline. This 

energy storage is representative of an advanced far-term Li polymer battery. This case is fairly similar to 

the primary/non-rechargeable energy storage option since they have similar W-hr/kg. The main difference 

is that the primary system is non-rechargeable and the electronics have different efficiencies. For 1 hr and 

<50 W, the (at least) 20 percent savings occur for umbilical lengths of 12 m (visual interpolation) and 

longer. For 2 hr and <50 W, the savings occur for 15 m and longer. For 4 hr and <50 W, the savings occur 

for 21 m (visual interpolation) and longer. For 1 hr and >100 W, the savings occur for 19 m (visual 

interpolation) and longer. Therefore, for low power levels, recharging at various intervals (1 to 4 hr) 

seems to make this option feasible. Even 1 hr recharge for higher power levels makes this option feasible. 

It is clearly more feasible than the primary battery (non-rechargeable) option from the standpoint of EVA 

duration due to the ability to recharge. Although this option seems viable, its drawback is the technology 

unavailability for the near and midterm. 

Figure 7: Ratio of Power-related Mass Relative to 100 W-hr/kg Secondary Energy-1m 

Tether Option (with emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 36 

Figure 8: Ratio of Power-related Mass Relative to 150W-hr/kg Secondary Energy/1m 

Tether Option (with emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical

Figure 9: Ratio of Power-related Mass Relative to 200W-hr/kg Secondary Energy/1m 

Tether Option (with emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 37 

Figures 10 through 14 show the total power system mass including any suit mounted power system 

(emergency power is not considered) and power umbilical. 

Figure 10 has a suit mounted, primary (non-rechargeable) energy storage system as the baseline 

alternative. Omitting the emergency storage worsens the mass savings of this option over the umbilical 

only cases. The reason for this is that an umbilical-only power system that must have extra mass on the 

EVA suit for emergency storage must add an entire power system mass, while for an existing EVA suit 

energy storage system with its related electronics/electrical system, some of which can be used by an 

emergency energy storage system. This means that using a nominal suit attached energy storage/power 

system provides some synergies and dual use hardware that the umbilical system would not have. 

Removing the emergency energy storage from both systems removes the advantage of the synergies and 

makes the umbilical power option lower in mass. Primary (non-rechargeable) energy storage provides at 

least 20 percent mass savings over the umbilical only power option with load durations of 1 hr and power 

level <50 W over umbilicals 18 m and longer. For 2 hr and <50 W, the savings occur over umbilicals of 

about 19 m (visual interpolation) and longer. For 4 hr and <50 W, the savings occur over umbilicals of 

about 23 m (visual interpolation) and longer. High power levels (>100 W) are not feasible from a mass 

savings perspective compared to the umbilical option (unless much longer umbilicals are considered). 

Low power levels may be feasible depending upon the travel distance requirements (to compare with the 

umbilical mass). 

Figure 11 has a suit mounted, 50 W-hr/kg secondary energy storage system as the baseline 

alternative. This is typical of a near-term (NiMH battery). No emergency energy storage is included. This 

option has only one application that is superior in terms of mass. This is the 1 hr load duration, 25 W 

power level case. Here, the secondary energy storage system saves at least 20 percent mass over 

umbilical-only power longer than 18 m (visual interpolation). Because of the limited applicability in 

terms of load level and duration, it seems that such an alternative is not viable. However, duration is not 

critical in that the 1 hr could be the EVA time between recharging. 

Figure 12 has a suit mounted, 100 W-hr/kg secondary energy storage system as the baseline 

alternative. This option can be considered equivalent to a near-term Li polymer battery. No emergency 

energy storage is included. For load durations of 1 hr and power levels <50 W, the 100 W-hr/kg energy 

storage options has at least 20 percent less mass than umbilical-only power options with lengths 20 m and 

longer. For 2 hr and power levels <25 W, savings exist for umbilical lengths of 18 m (visual 

interpolation) and longer. Load durations of 4 hr have no mass savings. Therefore, there is some potential 

for this option to be feasible, as long as recharging occurs every 1 to 2 hr for low power levels. 

Figure 13 has a suit mounted, 150 W-hr/kg secondary energy storage system as the baseline alternative. 

No emergency energy storage is included. This energy storage system can be considered representative of 

a more advanced Li polymer type battery. For load durations of 1 hr and power levels of <50 W, this 

option saves at least 20 percent mass of umbilical-only cases of lengths from 19 m (visual interpolation) 

and longer. For 2 hr and <50 W, the savings occur for lengths from 23 m (visual interpolation) and 

longer. Load durations of 4 hr do not meet the mass savings criteria. For 1 hr and >100 W, savings occur 

for lengths much greater than 20 m. Therefore, for recharging intervals of about 1 hr and lower power 

levels, this option is feasible. Even for some higher power levels, as long as the recharge interval is about 

1 hr, it is may be feasible from a mass savings perspective. 
NASA/TM—2009-215803 38 

 

Figure 10:Ratio of Power-related Mass Relative to Primary Energy/No Tether Option (no 

emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical

Figure 11:Ratio of Power-related Mass Relative to 50 W-hr/kg Secondary Energy/1m 

Tether Option (no emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass 

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 39 

 

Figure 14 has a suit mounted, 200 W-hr/kg secondary energy storage system as the baseline. No 

emergency energy storage is included. This energy storage is representative of an advanced far-term Li 

polymer battery. This case is fairly similar to the primary (non-rechargeable) energy storage option since 

they have similar W-hr/kg. The main difference is that the primary system is non-rechargeable and the 

electronics have different efficiencies. For 1 hr and <50 W, the savings of at least 20 percent occur for 

umbilical lengths of 18 m (visual interpolation) and longer. For 2 hr and <50 W, the savings occur for 20 

m and longer. For 4 hr and <25 W, the savings occur for 18 m (visual interpolation) and longer. No high 

power (>100 W) options seem feasible for the umbilical lengths considered (much longer than 20 m may 

be feasible). Therefore, for low to very low power levels, recharging at various intervals (1 to 4 hr) seems 

to make this option feasible. It is clearly more feasible than the primary battery option from the standpoint 

of EVA duration due to the ability to recharge. A drawback is the technology unavailability for the near 

and midterm. 

In the previous figures, only mass directly related to the power system was considered. Relatively 

speaking, the absolute magnitude of the mass savings is rather low. This analysis would be specifically 

applicable to a case where only a power umbilical was used (without other resources such as O2 and 

cooling). Detailed mass breakdowns of umbilicals are not available. However, some mass data is 

available for the Shuttle EMU and umbilical. 

Figure 15 was created to illustrate effect of the “balance of mass” contributors for the umbilical. On 

the left is the current Shuttle EMU suit mass. This would be representative of the suit-mounted energy 

storage cases described earlier in this report. On the right are cases for an umbilical only (power, cooling, 

O2 and communication) option. The masses are the sums of the space suit assembly (SSA), umbilical 

interface block/gauges/regulators, emergency O2 pack and the ISS EMU Umbilical-IEU (scaled for 

various lengths). The interface block is an estimate of the needed connections to connect solely with the 

suit (without the PLSS). The IEU is 3.3 m long and 13.6 kg (excluding the cooling fluid and O2 mass), 

thus 4 kg/m was derived for estimating longer umbilical lengths. The umbilical interface was estimated to 

be 1⁄4 of the PLSS mass. The figure illustrates that for umbilicals longer than 15 m, the non-umbilical 

options becomes competitive on a mass basis. The mass savings is 40 kg. Note that the figure does not 

include the umbilical cooling fluid or O2 mass which may increase the mass savings significantly (i.e., 

Figure 12:Ratio of Power-related Mass Relative to 100 W-hr/kg Secondary Energy/1m 

Tether Option (no emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 40 

one must add 20 to 40 kg to the 40 kg savings). For two astronaut EVAs, then two such umbilicals would 

be required, doubling the mass savings if the non-umbilical were used instead. Another aspect excluded 

from mass estimation is the required umbilical support/deployment/retraction hardware which is likely of 

considerable mass. 

 

Figure 13:Ratio of Power-related Mass Relative to 150 W-hr/kg Secondary Energy/1m 

Tether Option (no emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical

Figure 14:Ratio of Power-related Mass Relative to 200 W-hr/kg Secondary Energy/1m 

Tether Option (no emergency backup)

0.4

0.6

0.8

1

1.2

1.4

1.6

1 hr, 25W

1 hr, 50W

1 hr, 100W

1 hr, 150W

1 hr, 200W

2 hr, 25W

2 hr, 50W

2 hr, 100W

2 hr, 150W

2 hr, 200W

4 hr, 25W

4 hr, 50W

4 hr, 100W

4 hr, 150W

4 hr, 200W

Normalized Total Power-related Mass

No storage, 1m umbilical

No storage, 5m/25C umbilical

No storage, 10m/25C umbilical

No storage, 15m/25C umbilical

No storage, 20m/25C umbilical

No storage, 5m/95% eff umbilical

No storage, 10m/95% eff umbilical

No storage, 15m/95% eff umbilical

No storage, 20m/95% eff umbilical

50W-hr/kg storage, 1m/95% eff umbilical

100W-hr/kg storage, 1m/95% eff umbilical

150W-hr/kg storage, 1m/95% eff umbilical

200W-hr/kg storage, 1m/95% eff umbilical

Primary storage, No umbilical
NASA/TM—2009-215803 41 

For reference purposes, results from a past report are presented here. In April 1973, Hamilton 

Standard published a report called “Space Shuttle EVA/IVA Support Equipment Requirements Study”. It 

included a section comparing self-contained versus umbilical systems. Various options were considered 

including closed, semi-closed, semi-open and open loop for both types of systems. Various suit operating 

pressures were examined. The pressure suit by itself weighed from 29.5 kg (for 4 psi) to 34 kg (for 

10 psi). Umbilical length of 30 m was reported. The mass of the umbilical was estimated at between 10 to 

13 kg. Thus the mass per unit length was .3 kg/m, much smaller than a value based on the existing Shuttle 

IEU. This report seemed to consider only the aspects affected by pressure level, thus the cooling 

lines/communications lines/sensors and power cables were excluded. Cooling was done by sublimation 

and thus no cooling lines are assumed in the umbilical. The water weight, tankage and portion of the 

thermal control system required for humidity control and equipment loads was included in the basic PLSS 

weights. Although it is interesting to consider the options other than closed, the problem with those 

options is that total system mass (including consumables) becomes too high. The report deletes from 

consideration the closed loop umbilical on the grounds that “The closed loop umbilical system is slightly 

larger and heavier than the self-contained closed loop system and also has the disadvantage of umbilical 

management.” Figures show that the closed loop umbilical system is twice as heavy as the self-contained 

closed loop system. For a 10 psi/closed loop umbilical system, the PLSS and Umbilical mass was 

estimated at 8.8 kg but for the self-contained closed loop system it was estimated to be 4.9kg (note that 

the pressurized suit mass of 32 kg and ECLSS mass of 6kg were the same for both systems). No mobility 

or “management” data was included in the report. The stowed volume for a 30 m umbilical was estimated 

at 0.2 m3

. 

3.6 Other Aspects 

The mobility that is provided by eliminating the use of an umbilical is difficult to quantify. A fixed 

length umbilical almost always implies a maximum distance the astronaut may travel from the airlock (or 

umbilical connection site). Given that the absolute distance requirement for any possible use is unknown, 

it is desirable to make this distance as long as possible to account for all uses. Qualitatively, the suitattached energy storage options offer the maximum mobility while umbilical options offer least mobility. 

Figure 15: Estimated Total EVA Suit Mass Comparison

0

20

40

60

80

100

120

140

160

180

Shuttle EMU Umbilical

poweronly,5m

Umbilical

poweronly,10m

Umbilical

poweronly,15m

Umbilical

poweronly,20m

Total Mass (kg) 

Umbilical

Umbilical Interface

Emergency O2

PLSS

Suit
NASA/TM—2009-215803 42 

Another quantitative problem regards “encumberness”. This metric is a measure of how difficult it is 

to move around or operate for a specific suit configuration. For in-space suits, the mass of either the suit 

or umbilical should not be a concern for the astronaut. However, suit volume, inertia and accelerations are 

affected by having an energy storage system on the suit. Umbilicals have a problem in that the longer they 

are (and long ones are usually needed for unknown EVA situations) the more difficult they are to 

accommodate, manipulate, deploy, situate, retract and stow. They must be routed to minimize damage 

risk to the spacecraft or to the umbilical. This encumberness metric is some combination of mass and 

these other aspects. Qualitatively, the encumberness of umbilicals is greater than the self-contained 

energy system options. 

Risk is another measure that must be qualitatively addressed. Long umbilicals have the risk of 

puncture or severing. However, since such risks can be addressed with emergency backups (for either 

umbilical-only or suit-mounted power options), it is not necessary to be concerned with such measures at 

this time. 

3.7 Conclusions 

Depending upon the requirements of the EVA suit, there are specific mass beneficial niches for the 

suit-mounted nominal energy storage (either primary/non-rechargeable or secondary/rechargeable) over 

the umbilical-only power option. These niches are typically low power (<50 W) and low duration (<1 hr). 

Groundruling that suit-mounted emergency energy storage be used enhances the mass benefits of using 

nominal suit mounted energy storage. While the total mass benefits are relatively modest (<4 kg for one 

suit) for typical likely lengths (20 m), the combination of mass benefits with enhanced mobility and 

reduced encumberness indicates that suit-mounted energy storage options should be considered for 

contingency in-space EVA use. 

Although suit-mounted primary battery energy storage has the best mass benefit, a secondary energy 

storage is beneficial because it is rechargeable (thus providing multiple uses) and more forwardly 

compatible to future uses (surface suit). Any use of secondary energy storage would require limited 

duration sorties to enable recharge of the batteries (and other resources). Even the relatively poor specific 

energy option (50 W-hr/kg) is superior to some umbilical power-only options. 

These power system mass savings are relatively modest, especially in terms of an in-space suit, since 

zero gravity makes mass somewhat less important (except as it contributes to inertia and acceleration 

effects). If launch mass is a concern, these savings are negligible. However, if the remaining umbilical 

mass elements are considered, sizable launch mass savings are created by using a suit-mounted system. 

Including these other umbilical mass elements (cooling, O2, etc) shows that umbilicals with a length of 

20 m may weigh as much as 40kg more (and likely much more due to mass of umbilical 

infrastructure/mechanisms and O2/cooling fluids inside the umbilical) per suit than suit-mounted energy 

options. 

Umbilical-only power is the best selection for in-space (pressurized/in-cabin) suits. There is no 

justification to use suit-mounted energy storage for such suit. Surface suits favor suit-mounted energy 

storage with short umbilicals for recharge (longer umbilicals are not favored for surface suits because of 

dust, wear, abrasion and severing concerns, as well as reduced mobility and increased encumberness). 

Non-rechargeable energy options have the benefit of maximum mobility (because they are not tied to 

a power recharge port but other resources may still need an umbilical) although the EVA duration has a 

relatively short “hard” time limit which requires swap out of the energy storage to accommodate. 

Umbilical power-only options (with our without emergency backup) have “unlimited” EVA duration 

and are fairly reusable (potential for numerous separate EVAs are possible with same hardware) although 

the mobility (range of travel and encumberness) is limited and mass is likely a problem for longer lengths. 
NASA/TM—2009-215803 43 

3.8 Future Modifications: 

1) Adjust the minimum wire gauge from AWG 22 (which may be excessive) to 24 to 26 AWG for 

power levels below 50 W. Base gauge selections on standards such as “JSC-49894 Electronic Part 

Selection and Design Guidelines for Low Criticality Space Flight Payloads” and “NASA-STD-8739.4 

Requirements for Interconnecting Cables, Harnesses and Wiring, Crimping, Interconnecting Cables, 

Harnesses and Wiring”. 

2) Include reliability and availability analysis on power system design topology (e.g., number of 

redundant wire pairs, battery cell connectivity). 

3) Consider zinc-oxygen primary batteries. The 1973 Hamilton Standard report mentioned that 

although these zinc-air (oxygen) primary batteries had the highest specific energy, they were ground ruled 

out for the self contained EVA suit because of an only 7 day activation life and it requires an oxygen flow 

of at least 48 cc/min to meet the requirements and thus poses an interface within the PLSS not otherwise 

present. Assuming a pure O2 instead of air feed into the battery could address the activation life issue. 

Assuming a clean sheet PLSS, the O2 interface should be permissible. The O2 mass needed to supply the 

system must be accounted (which it is not clear that they did in the trade study) and added 

interconnects/pressure lines for the oxygen must be considered. Currently, a company called Electric Fuel 

has zinc-air batteries rated at 350 W-hr/kg and a 800 W-hr battery pack based on 30 A-hr cells (12 to 

24 V). Such batteries have even been proposed to be “rechargeable” on the anode level via “anode 

cassettes” 

4) Examine effects of multiple recharge/umbilical attach points on CEV (assuming small suitmounted translation power system) 

 
NASA/TM—2009-215803 44 

4.0 Trade Study 4: Lunar Surface EVA Suit Power Subsystem 

Rapid Recharge Versus Replace 

Date: May 26, 2006 

4.1 Results Summary 

This trade study assessed both the PLSS and Rapid Recharge System (RRS) Station power subsystem 

masses for supporting lunar surface EVA operations. Options considered included 8 hr PLSS and 4 hr 

PLSS with recharging, refilling or replacing of the energy storage subsystem. The results of the trade 

study showed that primary fuel cell and battery (with specific energy >=200 W-hr/kg) energy storage 

options provide the lowest PLSS power subsystem mass. Given the low PLSS mass savings involved 

(<3 kg for >=100 W PLSS options), PLSS power subsystem rapid replacement, refilling or recharging is 

not justifiable. However, if low PLSS mass savings are desired, PLSS energy storage replacement or 

refilling can be performed with a low to moderate RRS Station mass (5 to 42 kg for a 100 W PLSS). For 

>=265 W-hr/kg battery PLSS options, recharge can be performed with ‘heavy’ RRS Stations (>160 kg for 

a 100 W PLSS). Solar energy reduces RRS Station mass modestly, but inductive charging increases both 

the PLSS and RRS Station mass. Recharging both suits simultaneously during 15 min has large thermal 

and electronics impacts thus either sequential recharging or longer recharge times should be considered. 

4.2 Introduction 

The goal of this trade study is to identify power subsystem options which minimize the mass of the 

lunar surface EVA suit PLSS and EVA PLSS support equipment (called the RRS Station). This report is 

divided into sections that 1) describe the options considered and their assumptions, 2) size the PLSS 

power subsystem options, 3) use the PLSS sizing results to help size the RRS Station options, 4) combine 

the PLSS and RRS Station power subsystem masses for a ‘net’ mass metric and 4) generate a ranking 

table to compare the relative benefits of all options. 

4.3 PLSS Power Subsystem Options 

Table 1 lists the various PLSS power subsystem options considered for this trade study. 

TABLE 1.—PLSS POWER SYSTEM OPTIONS 

Energy Storage Option Energy Storage type Battery 

Specific

Energy,

W-hr/kg

Max 

charge 

rate 

15 min 

recharge from 

this depth of 

discharge, 

percent 

Vendor/ 

8 hr 4 hr comment 

Replace 

4 hr 

Refill 

4 hr 

Recharge 

X X X Li ion battery 100 1C 25 Note 1 

X X X Li ion battery 150 1C 25 Note 1 

X X X Li polymer battery 200 1C 25 Note 1 

X X X Li sulfur battery 265 3C 75 Sion claimed present capability 

X X X Li sulfur battery 318 6C 80 Sion claimed present capability 

X X X Primary fuel cell 

(H2/O2) 

----- --- --- Note 1 

 X Nickel metal hydride 65 7C 80 Electroenergy claimed present 

capability 

 X Li ion battery 85 7C 80 A123 claimed present capability 

 X Li ion battery 100 2C 50 Note 1 

 X Li ion battery 150 2C 50 Note 1 

 X Li polymer battery 200 2C 50 Note 1 

The left four columns indicate whether the particular battery or fuel cell option (as defined by the 

right four columns) was considered for the various ‘Energy Storage Options’. 
NASA/TM—2009-215803 45 

The ‘8 hr’ energy storage option has the energy storage sized for a full 8 hr EVA, thus, has no in situ 

recharge/replacement/refilling and no RRS Station is needed (at least for the power subsystem). 

The ‘4 hr Replace/Refill’ energy storage options have the PLSS power subsystem sized for a 4 hr 

EVA. The PLSS is designed with replaceable battery modules or fuel cell tank modules (PLSS 

H2/O2/H2O tanks clustered together for simultaneous removal). Such replacements add to the design 

complexity and increase risk. Replacement requires either two astronauts (since the energy storage has 

traditionally been on the back of the astronaut) or the energy storage be located more forwardly 

accessible. For primary fuel cell refilling, high pressure recharge of the PLSS fuel cell tanks via a hose 

connection to the RRS Station is required. The pressurization level is based on safe flow rates to 

accommodate 15 min refilling in a non-optimum (dusty) environment. The RRS Station is designed to 

store and maintain either module type (or the high pressure refilling equipment and tanks) for a 7 day 

period prior to the EVA activity. After this 7 day period, a single, 4 hr EVA replace/refill capability for 

each of two astronauts is needed at the RRS Station. 

The ‘4 hr Recharge’ energy storage option has the PLSS power subsystem sized for a 4 hr EVA but 

has the PLSS power subsystem recharged within 15 min by the RRS Station. The electrical recharge can 

be either done conductively or inductively. The RRS Station must simultaneously recharge two 

astronaut’s PLSS power subsystems within 15 min to provide the ability to perform another 4 hr EVA. 

The column ‘Energy Storage Type’ lists those energy storage types considered to be most likely to be 

used in the lunar surface EVA application. Section 4.8 describes the various battery options in more detail 

including references to literature in which the characteristics are discussed. Primary (non-rechargeable) 

batteries were not considered due to their nonrenewable nature. Supercapacitors were not considered due 

to their very low specific energy. Also, it was assumed that EVA suit-mounted regenerative fuel cells 

would be impractical for this time frame. Supercritical and cryogenic storage of fuel for primary fuel cell 

subsystems were not considered at this time due to the added hardware complexities. 

The column ‘Battery Specific Energy W-hr/kg’ is the assumed value for this battery type. 

The column ‘Max Charge Rate’ is the assumed maximum rate at which the battery can safely be 

recharged. ‘1C’ indicates the battery may be recharged (from a completely discharged state) in 1 hr. The 

‘8 hr’ options may be assumed to be ‘slowly’ recharged/refilled at the base (not at the EVA sortie site). 

‘Slow’ (traditional) battery charge methods (=<1C) are battery life/safety friendly. Section 4.8 discusses 

rapid charging batteries in more detail. 

The column ‘15 Minute Recharge From This Depth of Discharge’ is the maximum amount that the 

battery can be discharged in order to be recharged at the maximum charge rate within 15 min. Thus, for a 

1C battery which can be fully recharged from a completely discharge state in 1 hr, it may only be 

discharged 25 percent in order to recharge the battery completely at 1C within 15 min. The effect of this 

partial discharging to accommodate the charge rate limitation is that the battery must be oversized to 

provide the energy requirement within 15 min. For all battery options with less than 4C charge rate, the 

battery must be oversized. For safety and life reasons, the maximum allowable battery depth of discharge 

is 80 percent. 

Under ‘Vendor/Comment’, some details relating to type of energy storage characteristics are listed. 

Vendors provide current test data as well as projected capabilities for their products. 'Note 1' indicates that 

the listed set of metrics were based on a possible range of capabilities for TRL6 by 2010. The 

methodology of how all of these battery metrics were determined is discussed in detail in Section 4.8. 

Note that the battery ‘roundtrip’ efficiency (used in the determination of the amount of energy the 

RRS Station must provide the PLSS for recharging) was assumed to be 80 percent for NiMH battery 

types and 90 percent for Li battery types. 

4.4 PLSS Power Level Assumptions 

Examination of the schematic study results (NASA Crew Robotics And Vehicle Equipment Contract 

Delivery Order #1–DO-CRAVE-EC5-001) for one EVA suit for a 4 hr EVA showed the required user 

energy to be 251 W-hr to 2490 W-hr (for a cold environment) and 133 W-hr to 1920 W-hr (for a 
NASA/TM—2009-215803 46 

moderate environment). Assuming only Schematics 3 (the minimum mass option) and 4 (low power in a 

moderate environment option), the required user energy is ~135 W-hr (for a moderate environment) and 

~282 W-hr (for a cold environment). The derived power levels are, thus, ~34 W (for a moderate 

environment) and ~70 W (for a cold environment). This trade study assumed a range of power levels (25, 

50, 100, 150, and 200 W) (as measured at the user electrical load after PLSS power subsystem losses). 

4.4.1 PLSS Primary Fuel Cell Assumptions 

Three primary fuel cell options were considered: size the primary fuel cell for 8 hr, size it for 4 hr and 

replace the fuel cell tank cluster and size it for 4 hr and refill the tank cluster using high pressure lines. 

Refilling the fuel cell tank cluster at the RSS site has some risk due to the required fluid connection in a 

dusty environment. However, since recharging of the EVA suit breathing tanks is likely to be required, 

then similar technologies can be assumed to be used for oxygen connections and recharging. This 

applicability does not exist for hydrogen recharging, therefore operational and technology development is 

needed for this gas. For the fuel cell tank cluster replacement option, only the oxygen, hydrogen and 

water tanks were assumed to be replaced at the RRS Station. The astronaut must be able to access the 

tanks (by placing them on the front/side or another astronaut must perform the replacement) in a dusty 

environment. Any quick disconnect/connect valves and attachments must be resistant to dust. It is 

difficult to estimate replacement times but a 15 min replacement seems feasible. One tank cluster would 

be used to power during the replacement period of the other cluster. 

A model for the primary fuel cell based power subsystem was created based on the data and 

algorithms listed in the report entitled ‘Small Portable PEM Fuel Cell Systems for NASA Exploration 

Missions’ (Kenneth A. Burke, NASA/TM—2005-213994, December 2005, AIAA–2005–5680). The key 

assumptions used in this trade study were mostly derived from this reference. These include the following 

assumptions. The fuel cell energy conversion efficiency was assumed to be a conservative 60 percent. 

The fuel cell active area per unit mass was assumed to be 2000 cm2

/kg. The fuel cell discharge current 

density was assumed to be 0.398 A/cm2

. The cell voltage was assumed to be .8V. The high pressure 

H2/O2 tanks burst pressure was assumed to be 573 atm (this limit is based on 15 min fill rate). A fuel cell 

stack multiplier (number of independent full power stacks for redundancy in case something should go 

wrong with one stack) was assumed to be 3.0. A tank cluster multiplier (number of tanks to split the fuel 

into for redundancy in case something should go wrong with one tank cluster) was assumed to be 2.0. 

Ancillary Equipment specific power was assumed to be 0.002 kg/ W. The other fuel cell 

packaging/insulation mass fraction was assumed to be 25 percent of all the other fuel cell mass. 

4.4.2 Recharge Methods 

There are two methods to provide electricity from the RRS Station to the PLSS power subsystem. 

Section 4.9 goes into more detail on these and includes various references to technical literature. These 

methods are summarized here. 

4.4.2.1 Conductive 

This method relies upon a traditional direct electrical connection (wires/cable harness) from RRS 

Station to PLSS. In order to reduce the mass carried on the EVA suit, this method assumes that any high 

power charge electronics (i.e., battery charge regulator, DC-DC converter) are located in the RRS Station. 

4.4.2.2 Inductive 

This method uses a noncontact method to transfer the electricity from the RRS Station to the PLSS 

power subsystem. DC power is converted at the RRS Station to high frequency AC power. The power is 

then transferred over a narrow gap (which depends upon the design) via separate halves of a transformer 

(one on RRS Station side and one on PLSS side). The AC power is converted back to DC in the PLSS 

using electronics which then charge the batteries. This method has been used in underwater vehicles and 
NASA/TM—2009-215803 47 

electric cars and the main reason it is considered for lunar operational use is because of the potential 

reduced impact of dust on power connections (i.e., enhanced power/current safety). 

Sizing characteristics of inductive chargers include the inductor specific power (one ‘half’ only) of 

370 W/kg with an inductor efficiency of 90 percent. Also, high power PLSS-side recharge electronics 

must be accounted for (i.e., AC-DC converter, battery charge regulator). Transfer of data is possible 

inductively (done with underwater vehicles) and may be needed to control the battery charging. 

4.4.3 Other PLSS Power Subsystem Sizing Assumptions 

Numerous other assumptions must be made in order to size the PLSS power subsystem. 

For thermal sizing, the mass of the power-related thermal subsystem is based on specific power where 

the power level considered is the load power. For the 8-hr battery and 4-hr battery replacement options, 

the thermal loads are ‘nominal’ and a specific power of 0.01kg/ W was assumed to estimate the mass. 

However, higher thermal loads are assumed for all primary fuel cell options (due to primary fuel cell 

60 percent efficiency) and all recharge options (15 min recharge effect and battery efficiency impact). For 

these the thermal subsystem specific power was assumed to be double the nominal value (0.02 kg/W). All 

thermal mass estimates are very rough and require more detailed analysis. 

For structure sizing, the mass is estimated as a fraction of the mass the structure must hold or contain. 

Two classes of structures are energy storage structure and electronics structure. The ‘nominal’ level of 

structure (all energy storage options except replacement options) assume that the energy storage structure 

mass is 10 percent of the energy storage mass. All replacement options (batteries or primary fuel cells) 

require extra structure for mechanisms that can be used in a dusty environment by gloved astronauts, thus, 

the energy storage structure mass is assumed to be 20 percent of energy storage mass. Also, for 

replacement, an extra 10 percent is added to the nominal mass of power cables/interconnects to account 

for enhanced dust resistance and astronaut handling/durability/aids. 

For electronics sizing, it was assumed that the power was delivered to the user at 28 Vdc. Electronics 

components were sized based on specific powers derived from spacecraft historical data. The battery 

regulator was assumed to be 0.02 kg/ W with an efficiency of 90 percent. The power distribution unit was 

assumed to be 0.02 kg/ W with an efficiency of 99 percent. And the DC-DC converters were assumed to 

be 0.0125 kg/ W with 95 percent efficiency. 

In order to size the distribution cables connecting the centralized PLSS power subsystem to the loads 

throughout the EVA suit, a distribution of the user power to various sites had to be assumed (Table 2). 

Because of the low power levels for distribution, cable sizing is based on minimum mass down to the 

minimum wire radius (.032cm). At these wire radii, the efficiencies are very high and losses low. 

TABLE 2.—USER POWER DISTRIBUTION 

Load Location Cable 

distance 

(m) 

Local Load Power Level (W) 

Left hand 1 2.5 5.0 10.0 15.0 20.0 

Right hand 1 2.5 5.0 10.0 15.0 20.0 

Left foot 1.5 1.9 3.9 7.7 11.6 15.5 

Right foot 1.5 1.9 3.9 7.7 11.6 15.5 

Helmet 0.66 0.9 1.7 3.5 5.2 7.0 

Back 0.33 13.7 27.5 55.0 82.5 110.0 

Chest 0.5 1.5 3.0 6.0 9.0 12.0 

Total user power level, W 25 50 100 150 200 

DC-DC converters were assumed to be needed at each load site to converter power for computer 

electronics. It was assumed that 20 percent of the allocated electrical load at each load site would be 

converted for this use. 
NASA/TM—2009-215803 48 

Some power cables (between the battery modules and discharge regulator, between the discharge 

regulator and power distribution unit, and between the power distribution unit and a DC-DC converter) 

were sized using a simplified method (e.g., 99.9 percent efficiency and .001 kg/W_net_suit power). 

High current charging cables for the various recharge options had to be sized using cable design 

algorithms. PLSS side recharging cable mass had to be sized to not exceed a temperature of 293 K while 

being able to handle the high currents that occur for 15 min recharging. In order to generate a minimum 

PLSS-side mass as well as reduce electrical losses/thermal loads, the PLSS-side electrical recharging 

cables are assumed to be the shortest practical length. For this trade study, the length was assumed to be 

0.15 m from the connector (into which external power is input into the PLSS) to the battery (e.g., if the 

battery is on the backpack, then the connector is on the back). The implication of this approach is that 

either two astronauts are needed to connect the RRS Station recharging line to the back of the PLSS, or 

the PLSS energy storage components are front or side EVA-suit mounted, or a specialized ‘backup-into’ 

self-connection assembly is utilized at the RRS Station. Li batteries, because they have smaller amp-hour 

sizes, require more, separate cables if the charge regulation unit is assumed to be on the RRS Station. The 

mass of such charge cables for Li batteries are about three times the mass of NiMH battery charge cables 

(1 A-hr versus 6 A-hr). Note that the charging power cables require two backup power cable pairs to 

assure that the batteries can be charged even with two faults in the cables. For NiMH battery types, the 

recharge cable mass ranges from 0.03 kg (1 cable for 25 W PLSS/excluding redundancy) to 0.24 kg (six 

cables for 200 W PLSS). For Li battery types, the recharge cable mass is from 0.13kg (5 cables for 25 W 

PLSS) to 0.89kg (35 cables for 200 W PLSS). 

For the PLSS-side refill high pressure lines, it was assumed that they were the shortest practical 

length in order to minimum mass and enhance safety. The implication of this approach is that either two 

astronauts are needed to connect the RRS Station refilling line to the back of the PLSS, or the PLSS 

energy storage components are front or side EVA-suit-mounted, or a specialized ‘backup-into’ selfconnection assembly is utilized at the RRS Station. 

It was assumed that reliable, safe and practical designs operated by a suited astronaut in a dusty lunar 

environment can be implemented for low and high power/current electrical connections (for data and 

power), high pressure H2/O2 connections, low pressure H2O connections and mechanisms/hinges. 

4.4.4 PLSS Power Subsystem Redundancy Assumptions 

The PLSS power subsystem must include consideration of redundancy to enhance reliability and 

astronaut safety. Each energy storage options as well as other aspects of the power subsystem must 

address redundancy. 

For the various battery energy storage options, it was assumed that it would be designed such that no 

extra mass would be required to achieve redundancy (i.e., no extra battery modules or extra cells within 

each battery). This was felt to be feasible since the quantity of battery modules were likely greater than 

two and each module likely composed of multiple cells. Thus, in case of a module level failure, the PLSS 

could operate at reduced power levels with the remaining battery module until the failed battery module is 

replaced. It is groundruled that the design would have more than one battery module and that they are 

operated to permit safe return to a replacement station. 

For the primary fuel cell option, the fuel cell stack was assumed to have three full power backups. It 

was assumed the fuel and storage was separated into two separate fuel cell tank clusters. In the unlikely 

event that there are two total failures of the fuel cell stack, the fuel cell can still provide the full nominal 

power level. In case one fuel cell tank cluster fails for any reason, the remaining tank cluster can be used 

to provide the nominal power level for a reduced time (or a reduced power level for a longer time). 

For all PLSS power subsystem options, it was assumed that power cables for nonessential loads (i.e., 

all loads except those on the back of the EVA suit) had one backup cable pair and that power cables for 

essential loads (i.e., only one assumed to be located on the back of EVA suit) had two backup cable pairs. 

All electronics are assumed to have internal redundancy. In general, all power subsystem components 

were assumed to have inherent ‘high’ (yet currently undefined) reliability. 
NASA/TM—2009-215803 49 

Despite these considerations and assumptions for redundancy, the mass required for the contingency 

power subsystem was also included within the total PLSS power subsystem mass. This was assumed 

because of historical safety considerations on Shuttle suits and because assurance of astronaut safety is of 

paramount importance. It is assumed that powered hardware is an absolute requirement for safe suit 

operation, thus a backup power system is needed in case of an emergency condition. Such a contingency 

power subsystem may be used instead of relying on redundancy or in addition to redundancy but is 

nevertheless a groundruled requirement. The depth of discharge for the contingency primary battery is 

assumed to be 90 percent. 

Other PLSS power subsystem sizing assumptions are included in Table 3. 

TABLE 3.—PLSS POWER SUBSYSTEM SIZING ASSUMPTIONS 

Power Subsystem Parameters 

Voltage Type Direct Current 

Voltage Level V 28 

Total Site Power Into DCDC Converter percent 20 

Total Load Levels Considered W 25, 50,100,150,200 

Cable Parameters 

Cable Wire Material Copper 

Cable Wire Density gm/cm3 8.9 

Cable temperature K 279 

Cable Power Safety Factor 2 

Number of Redundant Cable Pairs/Transmission Line 1 

Cable Jacket Filler Density gm/cm3 1.7 

Insulator Density gm/cm3 1.7 

Cable Jacket Density gm/cm3 1.7 

Insulation Dielectric Strength Safety Factor 2 

Contingency Parameters 

Contingency Energy Storage Specific Energy W-hr/kg 220 

Contingency Energy Storage Load Level W 25 

Contingency Energy Storage Load Duration min 30 

Component Specific Powers 

DCDC Converter Specific Power kg/W 0.0125 

Power Distribution Unit Specific Power kg/W 0.02 

Discharge Regulator Unit Specific Power kg/W 0.02 

Nondistribution Cable Specific Power kg/W 1000 

Component Efficiencies 

DCDC Converter Efficiency percent 95 

Power Distribution Unit Efficiency percent 99 

Discharge Regulator Unit Efficiency percent 90 

Mass Estimation Parameters 

Connector Mass (Percentage of Cable Mass) percent 10 

Other Power System Structural Mass (Percentage of the DCDC and cable masses) percent 10 

4.4.5 PLSS Power Subsystem Sizing Results 

Figures 1 through 8 depict the PLSS power subsystem sizing results. The first three figures show the 

PLSS power subsystem mass for each of the options for user power levels of 25, 50 and 100 W. The 8 hr 

Li battery (>=150 W-hr/kg) options, fuel cell (replace and refill) options, replace options and Li battery 

conductive recharge options (with >=2C charge rate excluding 100 W-hr/kg Li ion) all have 

competitively low mass. However, rapid charge (e.g., 7C charge rate), NiMH/Li ion options, 100 W-hr/kg 

Li ion option, all 1C recharge options and inductive charging options have higher masses. 

The net mass savings between replacing, recharging and 8 hr PLSS are small enough (i.e., comparing 

cases A.3/A.4/A.5 with C.3/C.4/C.5 show <3kg savings for 100 W PLSS) to indicate that the added 

complexities of replacement and recharging are not justified. 
NASA/TM—2009-215803 50 

Figure 4 depicts the same data as the previous 3 figures except now it is possible for the reader to 

select user power levels up to 200 W different from the discrete values of 25, 50 and 100 W. To assist in 

selecting the correct PLSS power subsystem options (which is complicated due to the number of 

lines/symbols), the legend has been ranked from highest line (top) to lowest line (bottom). 

3456789

A.1:Lith. Ion-100Whr/kg,8hr,1C-None

A.2:Lith. Ion-150Whr/kg,8hr,1C-None

A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

B.1:Fuel cell-,8hr,None

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

E.1:NiMH-65Whr/kg,4hr,7C-Cond. Rech.

E.2:Lith. Ion-85Whr/kg,4hr,7C-Cond. Rech.

E.3:Lith. Sulfur-265Whr/kg,4hr,3C-Cond. Rech.

E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

E.5:Lith. Ion-100Whr/kg,4hr,2C-Cond. Rech.

E.6:Lith. Ion-150Whr/kg,4hr,2C-Cond. Rech.

E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

E.8:Lith. Ion-100Whr/kg,4hr,1C-Cond. Rech.

E.9:Lith. Ion-150Whr/kg,4hr,1C-Cond. Rech.

E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

F.3:Lith. Sulfur-265Whr/kg,4hr,3C-Ind. Rech.

F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS Power Subsystem Mass (kg) 

Figure 1: 25W PLSS Power Requirement

5 6 7 8 9 10 11 12 13 14 15 16

A.1:Lith. Ion-100Whr/kg,8hr,1C-None

A.2:Lith. Ion-150Whr/kg,8hr,1C-None

A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

B.1:Fuel cell-,8hr,None

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

E.1:NiMH-65Whr/kg,4hr,7C-Cond. Rech.

E.2:Lith. Ion-85Whr/kg,4hr,7C-Cond. Rech.

E.3:Lith. Sulfur-265Whr/kg,4hr,3C-Cond. Rech.

E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

E.5:Lith. Ion-100Whr/kg,4hr,2C-Cond. Rech.

E.6:Lith. Ion-150Whr/kg,4hr,2C-Cond. Rech.

E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

E.8:Lith. Ion-100Whr/kg,4hr,1C-Cond. Rech.

E.9:Lith. Ion-150Whr/kg,4hr,1C-Cond. Rech.

E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

F.3:Lith. Sulfur-265Whr/kg,4hr,3C-Ind. Rech.

F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS Power Subsystem Mass (kg) 

Figure 2: 50W PLSS Power Requirement
NASA/TM—2009-215803 51 

Figure 5 summaries the replacement mass for battery module or fuel cell tank clusters and the fuel 

cell tank refilling mass for 25, 50 and 100 W PLSS user power levels. For user power levels <=50 W, the 

replace/refill masses are low. Fuel cell tank replace/refill masses have the lowest required mass. 

Figure 6 depicts the data in Figure 5 such that the reader may derive the PLSS power subsystem 

replace/refill mass for subsystems with user power levels other than 25, 50 and 100 W. 

Figure 7 shows the maximum charging power required to recharge the PLSS power subsystem 

batteries in 15 min. This data is used to size charging cables, charging related electronics and thermal 

8 10 12 14 16 18 20 22 24 26 28 30

A.1:Lith. Ion-100Whr/kg,8hr,1C-None

A.2:Lith. Ion-150Whr/kg,8hr,1C-None

A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

B.1:Fuel cell-,8hr,None

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

E.1:NiMH-65Whr/kg,4hr,7C-Cond. Rech.

E.2:Lith. Ion-85Whr/kg,4hr,7C-Cond. Rech.

E.3:Lith. Sulfur-265Whr/kg,4hr,3C-Cond. Rech.

E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

E.5:Lith. Ion-100Whr/kg,4hr,2C-Cond. Rech.

E.6:Lith. Ion-150Whr/kg,4hr,2C-Cond. Rech.

E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

E.8:Lith. Ion-100Whr/kg,4hr,1C-Cond. Rech.

E.9:Lith. Ion-150Whr/kg,4hr,1C-Cond. Rech.

E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

F.3:Lith. Sulfur-265Whr/kg,4hr,3C-Ind. Rech.

F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS Power Subsystem Mass (kg) 

Figure 3: 100W PLSS Power Requirement

0

10

20

30

40

50

60

25 50 75 100 125 150 175 200

PLSS Power Susbsystem Mass (kg) .

PLSS User Power Level (W)

Figure 4: PLSS Power Subsystem Power vs Mass

E.8:Lith. Ion-100Whr/kg,4hr,1C-Cond. Rech.

F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

E.9:Lith. Ion-150Whr/kg,4hr,1C-Cond. Rech.

F.3:Lith. Sulfur-265Whr/kg,4hr,3C-Ind. Rech.

F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

A.1:Lith. Ion-100Whr/kg,8hr,1C-None

E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

E.5:Lith. Ion-100Whr/kg,4hr,2C-Cond. Rech.

E.1:NiMH-65Whr/kg,4hr,7C-Cond. Rech.

E.2:Lith. Ion-85Whr/kg,4hr,7C-Cond. Rech.

A.2:Lith. Ion-150Whr/kg,8hr,1C-None

E.6:Lith. Ion-150Whr/kg,4hr,2C-Cond. Rech.

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

B.1:Fuel cell-,8hr,None

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill

E.3:Lith. Sulfur-265Whr/kg,4hr,3C-Cond. Rech.

E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace
NASA/TM—2009-215803 52 

subsystem and affects the RRS Station sizing. In addition to affecting mass, high power levels (and 

currents) may affect safety. 

Figure 8 shows the derived thermal load for conductive and inductive recharging of NiMH and Li 

energy storage types. These thermal loads are due to electronics, battery and charge cable efficiencies. 

High power/current charging (mandatory for a 15 min recharge) result in high thermal loads which 

require thermal management to keep the PLSS cool during this period. Such transient loads are likely 

beyond the capability of the PLSS thermal subsystem unless the user power levels are low or Li 

conductive recharging PLSS options selected. To address the recharging thermal issues, design options 

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 5.5 6 6.5 7

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

C.3:Lith. Polymer-200Whr/kg,4hr,1CReplace

C.4:Lith. Sulfur-265Whr/kg,4hr,3CReplace

C.5:Lith. Sulfur-318Whr/kg,4hr,6CReplace

Swap-out/Refill Mass (kg) 

Figure 5: Swapout/refill mass for 1 PLSS Power Subsystem (25W,50W,100W)

100W

50W

25W

0

2

4

6

8

10

12

14

25 50 75 100 125 150 175 200

Swap-out/Refill Mass (kg) .

PLSS User Power Level (W)

Figure 6: PLSS Power Subsystem Swap-out/Refill Mass vs Power 

C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

C.3:Lith. Polymer-200Whr/kg,4hr,1CReplace

C.4:Lith. Sulfur-265Whr/kg,4hr,3CReplace

C.5:Lith. Sulfur-318Whr/kg,4hr,6CReplace

B.2:Fuel cell-,4hr,Replace

B.3:Fuel cell-,4hr,Refill
NASA/TM—2009-215803 53 

require: 1) larger PLSS thermal control system mass, 2) coolant loop connection to PLSS from RRS 

Station or 3) a ‘cold plate’ attachment from RRS Station onto PLSS thermal system. Each of these 

options add mass and complexity to the PLSS. Due to the sizing method used, thermal effects/mass must 

be analyzed in more detail. 

0

1000

2000

3000

4000

5000

6000

25 50 75 100 125 150 175 200

Max.Charge Power Level-15 min (W) . .

PLSS User Power Level (W)

Figure 7: PLSS Power Subsystem Max Charge Power vs User Power

NiMH inductive recharge

Lithium-types inductive recharge

NiMH conductive recharge

Lithium-types conductive recharge

0

500

1000

1500

2000

2500

25 50 75 100 125 150 175 200

Max.Charge Thermal Load-15 min (W) . .

PLSS User Power Level (W)

Figure 8: PLSS Power Subsystem Max Thermal Load vs User Power NiMH inductive recharge

Lithium-types inductive 

recharge

NiMH conductive recharge

Lithium-types conductive 

recharge
NASA/TM—2009-215803 54 

4.4.6 RRS Station Power Subsystem Requirements 

The RRS Station is used to supply recharging, replacement or refilling resources (in addition to 

power) for one 4 hr EVA for two astronauts. For PLSS options in which the power subsystem is sized for 

a full 8 hr EVA, then the RRS Station must still be sized for power to support the recharge/replacing/ 

refilling of other PLSS subsystems. The baseline approach to the RRS Station activity is to assume both 

astronauts can perform the activity simultaneously and within 15 min. The RRS Station is assumed to be 

deployed prior to an EVA and the Station maintains the recharging or replacement capabilities both 

thermally and electrically until needed. Since the Station may be deployed 7 days prior to an EVA 

activity, the Station must provide its capabilities at the end of the 7 day period. Deployment and operation 

in either illuminated or shadowed lunar conditions must be considered. The continuous RRS Station 

housekeeping power is assumed to be an average of 20 W total during the entire 7 day period. This power 

level is for sensors, electronics and heaters. Power for communications and other power needs have not 

been considered at this time. Very low power levels are needed for thermal maintenance and/or battery 

trickle/top off charging. 

4.5 RRS Station Power Options 

A number of different options were considered for the RRS Station power subsystem. Table 4 lists 

these options. Separate power sources were considered for both the continuous ‘housekeeping’ power and 

for the ‘rapid recharge’ power. ‘Replace/refill’ (and also 8 hr PLSS power subsystem) options require 

only the RRS Station to provide housekeeping power. 

TABLE 4.—OPTIONS CONSIDERED 

Option type Housekeeping power source Rapid recharge power source 

Replace/refill None None 

Replace/refill Battery (200 W-hr/kg) Li polymer None 

Replace/refill Battery (318 W-hr/kg) Li sulfur None 

Replace/refill Solar None 

Replace/refill Primary fuel cell None 

Replace/refill Radioisotope (Stirling) None 

Recharge Solar Battery (200 W-hr/kg) Li polymer 

Recharge Solar Battery (318 W-hr/kg) Li sulfur 

Recharge Solar Primary fuel cell 

Recharge Solar Solar 

Recharge Radioisotope (Stirling) Flywheel 

Battery and primary fuel cell RRS Station components and their assumptions are similar to those 

discussed in the PLSS power subsystem sizing section and are not duplicated here. 

The RRS Station solar power subsystem option was assumed to be used when the Station was 

deployed in illuminated locations. It was assumed to be a one sided, non-tracking solar panel composed of 

27 percent efficiency/3-junction gallium arsenide solar cells. The solar panel was assumed to be fixed by 

the astronaut to the optimum solar pointing for the particular 7 day deployment. This resulted in a 

pointing factor of 91 percent (100 percent=perfect continuous pointing, 50 percent= pointed perfectly at 

the Sun half the time and away from the Sun the other half). The solar panel was sized including 7 percent 

albedo energy reflected from the surrounding lunar surface and the solar panel was assumed to operate at 

100C. The solar array regulator efficiency was assumed to have 95 percent efficiency and a specific 

power of 50 W/kg. Typical solar RRS Station sizing results show that the solar array by itself has a 

specific power of ~170 W/kg but when all other power subsystem masses are included (electronics, 

thermal, structure, harness/cables), this value goes down to ~10 W/kg. The total surface area of the solar 

panel to provide a continuous 20 W ‘housekeeping’ power is 0.13 m2

. For the total surface area of the 

‘rapid recharge’ solar panel, the areal power is ~200 W/m2

 where the power used is the peak charge 

power. For a 15 min conductive recharge of two 25 W PLSS power subsystems, the total area needed is 
NASA/TM—2009-215803 55 

~5 m2 and this goes to ~11 m2 for two 50 W PLSS units. Values above 10 m2 seem unrealistic for this 

application. 

The RRS Station Radioisotope power subsystem (RPS) option was considered because of its 

usefulness during continuous non-illuminated (lunar night/shadowed) operations. The two primary 

methods of converting thermal heat to electricity are advanced Stirling conversion (37 percent converter 

efficiency and 30 percent system efficiency) and thermoelectric (6 percent system efficiency). Because of 

the scarcity and expense of the radioisotope fuel, advanced Stirling, whose high conversion efficiency 

reduces the amount of fuel, was baselined (e.g., for every Stirling 1 fuel unit, 5 thermoelectric fuel units 

are needed). Although Stirling power subsystems have a limiting external operational temperature of 

100 °C (lunar max. is 120 °C), this limitation may be mitigated through judicious thermal 

orientation/management. RPS was not considered for high power recharge due to large amount of fuel 

required. Also, since the smallest RPS size is sized for 1 fuel unit, the minimum power level provided by 

Stirling is 65 We while for thermoelectric the 20 W required power is feasible using 2 fuel units. 

Although new development efforts could be performed to design smaller Stirling units, this may not be an 

issue since it is likely that the RRS Station power needs will exceed 20 W in any event. 

The RRS Station Flywheel energy storage subsystem was considered due to its capability for fast 

charge and discharge and relatively high energy density. Such subsystems are typically designed for 

60 min charge and 30 min discharge and are used commercially for uninterruptible power supplies. The 

concept for using this kind of energy storage would be to pre-deploy the RRS station with the flywheel 

not spinning. Then, either the solar or RPS power subsystem would be used to spin up the flywheel unit 

over a period of about 1 day. At that point the energy may be discharged according to the required power 

level. Note that technology development is required for this kind of flywheel charge/discharge cycle due 

to modifications needed to the motor/generator. Sizing assumptions based on projected capabilities for 

the flywheel subsystem show a specific energy of 50 W-hr/kg (with no specialized containment) and 

33 W-hr/kg (with containment, i.e., 50 percent of flywheel mass) is possible with adequate funding to 

TRL6 by 2010. Included in this projection are low power flywheel electronics (10 W startup/operating 

power per flywheel). 

4.6 RRS Station Redundancy Assumptions 

Despite the fact that the power subsystem is inherently designed for high reliability, since the RSS 

Station power subsystem is a critical system, redundancy is included. For the replacement options, it is 

assumed that one extra module (i.e., PLSS battery module and fuel cell tank cluster) is stored in RRS 

Station. For the refill (PLSS fuel cell tanks) options, it is assumed that 50 percent more fuel is stored in 

the RRS Station (only one tank cluster needed on the RRS Station to store the fuel). For the RRS Station 

battery options, 25 percent more mass is added to account for backup battery modules, cables, and 

electronics. The lower redundancy percentage is because it is a distributed system (i.e., has numerous 

battery modules) and the loss of a few would not affect the mission. For the RRS Station solar options, 

25 percent more mass is added to account for backup solar cells, interconnects, electronics, and cables. 

The lower redundancy percentage is because it is a distributed system (with numerous cells). For the RRS 

Station primary fuel cells option, there are 3 separate stacks and 3 separate tank clusters, each sized for 

full power/energy. For the RRS Station flywheel option, 3 separate flywheel assemblies (each at full 

power/energy) were assumed. Flywheels are inherently designed for safety and reliability by way of 

operation at lower speeds. For the RRS Station radioisotope power option, 3 separate 

converters/electronics units were assumed (each converter was oversized for this application at 60 W). 

4.6.1 RRS Station Power Subsystem Mass Results 

Figures 9 through 19 show the mass for each RRS Station option with each of the PLSS power 

subsystem recharge/replace options. Figure 9 shows the RRS Station mass for the replace/refill PLSS 

power subsystem options for a 100 W PLSS user power. 
NASA/TM—2009-215803 56 

In general, RRS Station options which are used to replace/refill PLSS power subsystems rather than 

perform a PLSS power subsystem recharge have the lowest mass. This is mainly due to the peak 

recharging power sizing effects on electronics and thermal subsystems. Solar replace/refill RRS Stations 

have the lowest mass but are limited to illuminated operation. Radioisotope, fuel cell and 318 W-hr/kg 

battery RRS Station options are moderate mass options which can operate in non-illuminated conditions. 

The 200 W-hr/kg battery-only RRS Station option has a relatively high mass. Fuel cell replace/refill 

PLSS options have the lowest mass RRS Station mass with 318, 265 and 200 W-hr/kg battery replace 

PLSS options having next lowest mass RRS Stations. For these RRS Station options, < 40 kg RRS Station 

(for the power subsystem only) seems feasible. 

Figure 10 shows the RRS Station mass for the recharge PLSS power subsystem options for a 100 W 

PLSS user power. The fuel cell and flywheel RRS Station options have the lowest mass. Combining solar 

for housekeeping power with other energy storage types for recharging provides modest mass savings. 

Battery-only RRS Station options have close to the highest masses. The solar-only RRS Station option 

has the largest mass due to the requirement of sizing for the peak charging power. The solar-only option 

has the potential to have a large area (based on 200 peak_chargingpowerwatts/m^2). Also, inductive 

charging increases the mass of the RRS Station by 60-200 kg due to inefficiencies of the method. 

0 10 20 30 40 50 60 70 80

PLSS: B.2:Fuel cell-

,4hr,Replace

PLSS: B.3:Fuel cell-

,4hr,Refill

PLSS: C.1:Lith. Ion100Whr/kg,4hr,1C-Replace

PLSS: C.2:Lith. Ion150Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer200Whr/kg,4hr,1C-Replace

PLSS: C.4:Lith. Sulfur265Whr/kg,4hr,3C-Replace

PLSS: C.5:Lith. Sulfur318Whr/kg,4hr,6C-Replace

RRS Station Mass (kg) 

Figure 9: RRS Station Masses:Replace/Refill Options (100W PLSS)

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Fuel Cell-only RRS Station

Radioisotope-only RRS Station

Solar-only RRS Station
NASA/TM—2009-215803 57 

Figures 11 through 19 show the RRS Station mass from the previous figures in a different format to 

permit the reader to select different PLSS user powers from the 25 W, 50 W and 100 W values. Each 

figure represents a different RRS Station power subsystem options as reflected in its title. The legend for 

each figure lists the various PLSS power subsystem options. The order of the items in the legend is based 

0 200 400 600 800 1000

PLSS: E.1:NiMH Battery 

Types:Conduct.Rech.

PLSS: E.2-E.10:Lithium Battery 

Types: Conduct. Rech.

PLSS: F.1:NiMH Battery 

Types:Induct.Rech.

PLSS: F.2-F.4:Lithium Battery 

Types: Induct. Rech.

RRS Station Mass (kg) 

Figure 10: RRS Station Masses:Recharge Options (100W PLSS)

Solar-only RRS Station

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Battery(200 W-hr/kg)/Solar RRS Station

Battery(318W-hr/kg)/Solar RRS Station

Flywheel/Radioisotope RRS Station

Flywheel/Solar RRS Station

Fuel Cell-only RRS Station

Fuel cell/Solar RRS Station

0

100

200

300

400

500

600

700

800

900

1000

1100

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 11: Battery-only RRS Mass vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options
NASA/TM—2009-215803 58 

on the magnitude of the line in the figure. Recharge RRS Station options have much steeper slopes and 

relatively higher magnitudes compared to replace/refill RRS Station options. 

0

100

200

300

400

500

600

700

800

900

1000

1100

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 12: Battery-only RRS Mass vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options

0

200

400

600

800

1000

1200

1400

1600

1800

2000

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 13: Solar-only RRS Mass vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options
NASA/TM—2009-215803 59 

0

100

200

300

400

500

600

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 14: Fuel Cell-only RRS Mass vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options

0

100

200

300

400

500

600

700

800

900

1000

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 15: Solar+Battery (200W-hr/kg) RRS Mass 

vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options
NASA/TM—2009-215803 60 

0

100

200

300

400

500

600

700

800

900

1000

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 16: Solar+Battery (318W-hr/kg) RRS Mass 

vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options

0

100

200

300

400

500

600

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 17: Solar+Fuel Cell RRS Mass 

vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options
NASA/TM—2009-215803 61 

0

100

200

300

400

500

600

700

800

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 18: Solar+Flywheel RRS Mass 

vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options

0

100

200

300

400

500

600

700

800

900

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 19: Radioisotope+Flywheel RRS Mass 

vs. User Power

PLSS:F.1-NiMH Inductive Recharge Options

PLSS:F.2-F.4:Lithium Battery Inductive Recharge Options

PLSS:E.1:NiMH Conductive Recharge Options

PLSS:E.2-E.10:Lithium Battery Conductive Recharge Options

PLSS:C.1:Lithium Battery Replace Option

PLSS:C.5:Lithium Batteries Replace Options

PLSS:B.1-B.2:Fuel Cell Refill/Replace Options
NASA/TM—2009-215803 62 

4.6.2 Combined RRS and 2 PLSS Power Subsystems Mass Results 

Figures 20 through 34 show the combined masses for each RRS Station option with each of the PLSS 

power subsystem recharge/replace/refill options. Two PLSS recharges/replacements/refillings (plus 

redundancy) are included in the RRS Station mass and two PLSS power subsystem masses are included 

in the combined mass. The results are similar to the previous sections with PLSS power subsystem mass 

only and RRS Station mass only. Figures 20 through 22 show the combined replace/refill/8 hr option 

PLSS power subsystem mass and RRS Station mass for three PLSS user power levels. Figures 23 through 

25 show the combined recharge option PLSS power subsystem mass and RRS Station mass for three 

PLSS user power levels. 

In general these figures show that the 8 hr PLSS power subsystem option (except for 8 hr PLSS 

100 W-hr/kg battery option) has the lowest combined mass. Replace, refill and 8 hr PLSS power 

subsystem options have moderate combined mass. For these PLSS options, the solar-only RRS Station 

has the lowest mass in the group but is limited to illuminated operation. Other methods (except RRS 

Station 200 W-hr/kg battery-only option) are next lowest in overall mass, are fairly competitive with each 

other and are not limited to in-Sun operation. The rechargeable PLSS power subsystem options have 

much higher mass compared to replace/refill/8hr options. For rechargeable PLSS options, fuel cell and 

flywheel RRS Station options result in lower combined mass whereas battery with solar options have 

moderate mass and solar-only and battery-only options have generally the highest combined mass. 

0 10 20 30 40 50 60 70 80

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.2:Lith. Ion-150Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

PLSS: B.1:Fuel cell-,8hr,None

PLSS: B.2:Fuel cell-,4hr,Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

RRS Station Mass (kg) 

Figure 20: 2 PLSS Power Subsystem(25W)+RRS Station Masses

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Fuel Cell-only RRS Station

Radioisotope RRS Station

Solar-only RRS Station
NASA/TM—2009-215803 63 

0 10 20 30 40 50 60 70 80 90

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.2:Lith. Ion-150Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

PLSS: B.1:Fuel cell-,8hr,None

PLSS: B.2:Fuel cell-,4hr,Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

RRS Station Mass (kg) 

Figure 21: 2 PLSS Power (50W) Subsystem+RRS Station Masses

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Fuel Cell-only RRS Station

Radioisotope RRS Station

Solar-only RRS Station

0 20 40 60 80 100 120

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.2:Lith. Ion-150Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.4:Lith. Sulfur-265Whr/kg,8hr,3C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

PLSS: B.1:Fuel cell-,8hr,None

PLSS: B.2:Fuel cell-,4hr,Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.2:Lith. Ion-150Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.4:Lith. Sulfur-265Whr/kg,4hr,3C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

RRS Station Mass (kg) 

Figure 22: 2 PLSS Power (100W) Subsystem+RRS Station Masses

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Fuel Cell-only RRS Station

Radioisotope RRS Station

Solar-only RRS Station
NASA/TM—2009-215803 64 

0 50 100 150 200 250

PLSS: E.4:Lith. Sulfur318Whr/kg,4hr,6C-Cond. Rech.

PLSS: E.7:Lith. Polymer200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.10:Lith. Polymer200Whr/kg,4hr,1C-Cond. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6CInd. Rech.

RRS Station Mass (kg) 

Figure 23: 2 PLSS Power (25W) Power Subsystem+RRS Station Masses

Solar-only RRS Station

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Battery(200 W-hr/kg)/Solar RRS Station

Battery(318W-hr/kg)/Solar RRS Station

Flywheel/Radioisotope RRS Station

Flywheel/Solar RRS Station

Fuel Cell-only RRS Station

Fuel cell/Solar RRS Station

0 100 200 300 400 500

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6CCond. Rech.

PLSS: E.7:Lith. Polymer200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.10:Lith. Polymer200Whr/kg,4hr,1C-Cond. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6CInd. Rech.

RRS Station Mass (kg) 

Figure 24: 2 PLSS Power (50W) Power Subsystem+RRS Station Masses

Solar-only RRS Station

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Battery(200 W-hr/kg)/Solar RRS Station

Battery(318W-hr/kg)/Solar RRS Station

Flywheel/Radioisotope RRS Station

Flywheel/Solar RRS Station

Fuel Cell-only RRS Station

Fuel cell/Solar RRS Station
NASA/TM—2009-215803 65 

Figures 26 through 34 show the combined results from the previous figures in a different format to 

permit the reader to select different PLSS user powers from the 25, 50 and 100 W values. Each figure 

represents a different RRS Station power subsystem options as reflected in its title. The legend for each 

figure lists the various PLSS power subsystem options. The order of the items in the legend is based on 

the magnitude of the line in the figure. Recharge RRS Station options have much steeper slopes and 

relatively higher magnitudes compared to replace/refill RRS Station options. 

0 200 400 600 800 1000

PLSS: E.4:Lith. Sulfur318Whr/kg,4hr,6C-Cond. Rech.

PLSS: E.7:Lith. Polymer200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.10:Lith. Polymer200Whr/kg,4hr,1C-Cond. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6CInd. Rech.

RRS Station Mass (kg) 

Figure 25: 2 PLSS Power (100W) Power Subsystem+RRS Station Masses

Solar-only RRS Station

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Battery(200 W-hr/kg)/Solar RRS Station

Battery(318W-hr/kg)/Solar RRS Station

Flywheel/Radioisotope RRS Station

Flywheel/Solar RRS Station

Fuel Cell-only RRS Station

Fuel cell/Solar RRS Station

0

200

400

600

800

1000

1200

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 26: Battery-only RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None
NASA/TM—2009-215803 66 

0

200

400

600

800

1000

1200

1400

1600

1800

2000

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 27: Solar-only RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

0

100

200

300

400

500

600

700

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 28: Fuel Cell-only RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None
NASA/TM—2009-215803 67 

0

200

400

600

800

1000

1200

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 29: Solar/Battery (200W-hr/kg) RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

0

100

200

300

400

500

600

700

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 30: Solar/Fuel Cell RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None
NASA/TM—2009-215803 68 

0

100

200

300

400

500

600

700

800

900

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 31: Solar/Flywheel RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

0

200

400

600

800

1000

1200

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 32: Battery-only (318W-hr/kg) RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None
NASA/TM—2009-215803 69 

0

100

200

300

400

500

600

700

800

900

1000

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 33: Radioisotope/Flywheel RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None

0

200

400

600

800

1000

1200

25 50 75 100 125 150 175 200

RRS Mass (kg) .

User Power (W)

Figure 34: Solar/Battery-only (318W-hr/kg) RRS +2 PLSS 

Power Subsystem Mass vs. Power

PLSS: F.1:NiMH-65Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.2:Lith. Ion-85Whr/kg,4hr,7C-Ind. Rech.

PLSS: F.4:Lith. Sulfur-318Whr/kg,4hr,6C-Ind. Rech.

PLSS: E.10:Lith. Polymer-200Whr/kg,4hr,1C-Cond. Rech.

PLSS: E.7:Lith. Polymer-200Whr/kg,4hr,2C-Cond. Rech.

PLSS: E.4:Lith. Sulfur-318Whr/kg,4hr,6C-Cond. Rech.

PLSS: C.1:Lith. Ion-100Whr/kg,4hr,1C-Replace

PLSS: C.3:Lith. Polymer-200Whr/kg,4hr,1C-Replace

PLSS: C.5:Lith. Sulfur-318Whr/kg,4hr,6C-Replace

PLSS: B.3:Fuel cell-,4hr,Refill

PLSS: A.1:Lith. Ion-100Whr/kg,8hr,1C-None

PLSS: A.3:Lith. Polymer-200Whr/kg,8hr,1C-None

PLSS: A.5:Lith. Sulfur-318Whr/kg,8hr,6C-None
NASA/TM—2009-215803 70 

4.6.3 Recharge Time Effect on RRS Station Mass 

15 min simultaneous recharging has a significant effect on RRS Station mass due to thermal and 

electronics sizing for peak powers that must be transferred to the PLSS in the short time period. 

Consideration should be given to extending this time period or removing the requirement for 

simultaneous recharging. 

4.6.4 Quantitative and Qualitative Ranking Comparison of Options 

A ranking table (Table 5) was created to numerically compare a combination of PLSS and RRS 

Station power subsystem qualitative and quantitative attributes for each option. Along the top of the table 

going from left to right are the various PLSS power subsystem options. A.1-A.5 and B.1 are the 8 hr 

PLSS power subsystem options. B.2-B.3 and C.1-C.5 are the replace and refill PLSS power subsystem 

options. E.1-E.10 are the conductive recharge PLSS power subsystem options. F.1-F.4 are the inductive 

recharge PLSS power subsystem options. For purposes of this table, only the best combined solar 

(housekeeping power)/battery (recharge power, if applicable) RRS Station option was considered as it 

applies to each PLSS power subsystem option. The list of questions in the left column are either 

answerable with a ‘Yes’ (given a point value of 1) and ‘No’ (given a point value of 0). The question 

‘Below median PLSS mass?’ provided a simple way to compare PLSS power subsystem masses. Those 

PLSS power subsystems below the median were given a ‘Yes’ (1). Similarly, the question ‘Combined 

mass of solar+battery RRS AND two suit power systems are below median RRS mass?’ relatively 

compares the total power subsystem mass metric (‘Yes’ for a combined mass less than the median). The 

other nine questions were arbitrarily selected as pertinent to the lunar surface EVA suit. They address 

dust, thermal, complexity, safety and technology readiness. 

Based on this table’s equal weighting of questions, it is clear from the ‘Sum’ row at the bottom, that 

most ‘8 hr’ PLSS options are superior to the other options. Also, most ‘replace’ options are superior to 

any recharge options. Different weighting systems were tried to examine their impact on the results. 

When normalized PLSS and RSS Station power subsystem masses were used instead of rounded values, 

the PLSS mass weighting factor increased to 10.0 and the RRS Station mass weighting factor increased to 

0

100

200

300

400

500

600

700

15 30 45 60

RRS Station Mass (kg) .

Total Recharge Time (minutes)

Figure 35: RSS Station Mass (100W PLSS, lithium battery 

conductive recharge) vs. Recharge Time

Solar-only RRS Station

Battery(200 W-hr/kg)-only RRS Station

Battery(318W-hr/kg)-only RRS Station

Battery(200 W-hr/kg)/Solar RRS Station

Battery(318W-hr/kg)/Solar RRS Station

Flywheel/Radioisotope RRS Station

Flywheel/Solar RRS Station

Fuel Cell-only RRS Station

Fuel cell/Solar RRS Station
NASA/TM—2009-215803 71 

5.0, the results were fairly close to this table except that replace options (C.4-C.5) moved closer to the 8 

hr PLSS options (A.2-A.5), especially for lower PLSS user power levels. 

TABLE 5.—METRIC COMPARISONS/RANKING OF OPTIONS 

Metrics 

(equally weighted) 

A.1:Lith. Ion-100 W-hr/kg,8 hr,1C-None A.2:Lith. Ion-150 W-hr/kg,8 hr,1C-None A.3:Lith. Polymer-200 W-hr/kg,8 hr,1C-None A.4:Lith. Sulfur-265W-hr/kg,8 hr,3C-None A.5:Lith. Sulfur-318W-hr/kg,8 hr,6C-None B.1:Fuel cell-,8 hr, None B.2:Fuel cell-,4 hr, Replace B.3:Fuel cell-,4 hr, Refill C.1:Lith. Ion-100 W-hr/kg,4 hr,1C-Replace C.2:Lith. Ion-150 W-hr/kg,4 hr,1C-Replace C.3:Lith. Polymer-200 W-hr/kg,4 hr,1C-Replace C.4:Lith. Sulfur-265 W-hr/kg,4 hr,3C-Replace C.5:Lith. Sulfur-318 W-hr/kg,4 hr,6C-Replace E.1:NiMH-65 W-hr/kg,4 hr,7C-Cond. Rech. E.2:Lith. Ion-85 W-hr/kg,4 hr,7C-Cond. Rech. E.3:Lith. Sulfur-265 W-hr/kg,4 hr,3C-Cond. Rech. E.4:Lith. Sulfur-318 W-hr/kg,4 hr,6C-Cond. Rech. E.5:Lith. Ion-100W-hr/kg,4 hr,2C-Cond. Rech. E.6:Lith. Ion-150W-hr/kg,4 hr,2C-Cond. Rech. E.7:Lith. Polymer-200 W-hr/kg,4 hr,2C-Cond. Rech E.8:Lith. Ion-100 W-hr/kg,4 hr,1C-Cond. Rech. E.9:Lith. Ion-150 W-hr/kg,4 hr,1C-Cond. Rech. E.10:Lith. Polymer-200 W-hr/kg,4hr,1C-Cond. Rec F.1:NiMH-65 W-hr/kg,4 hr,7C-Ind. Rech. F.2:Lith. Ion-85 W-hr/kg,4 hr,7C-Ind. Rech. F.3:Lith. Sulfur-265 W-hr/kg,4 hr,3C-Ind. Rech. F.4:Lith. Sulfur-318 W-hr/kg,4 hr,6C-Ind. Rech. 

Below median PLSS mass? 0 1 1 1 1 1 1 1 1 1 1 1 1 0 1 1 1 0 1 1 0 0 0 0 0 0 0

Combined mass of solar+battery RRS AND two suit power 

systems are below median RRS mass? 1 1 1 1 1 1 1 1 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0

No dust environment electrical power connections(assume 

sensors are RF or also carried inductively)? 1 1 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 1 1 1

No dust environment gas/fluid connections? 1 1 1 1 1 1 0 0 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0

No dust environment mechanisms? 1 1 1 1 1 1 0 1 0 0 0 0 0 1 1 1 1 1 1 1 1 1 1 1 1 1 1

No pressurized gas safety issue? 1 1 1 1 1 0 0 0 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1

No high current charging safety issue? 1 1 1 1 1 1 1 1 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0

No high thermal loads in charging/nominal operation? 1 1 1 1 1 0 0 0 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0

Low complexity in mechanisms? 1 1 1 1 1 1 0 1 0 0 0 0 0 1 1 1 1 1 1 1 1 1 1 1 1 1 1

Low thermal complexity in charging? 1 1 1 1 1 1 1 1 1 1 1 1 1 0 0 0 0 0 0 0 0 0 0 0 0 0 0

Available sooner than later? 1 1 0 1 0 0 0 0 1 1 0 1 0 1 1 1 0 1 1 0 1 1 0 1 1 1 1

Sum 10 11 10 11 10 8 5 6 8 8 7 8 7 4 5 5 4 4 5 4 4 4 3 5 5 5 5

4.7 Conclusions 

The most important finding was that, because of the relatively low PLSS power subsystem masses 

involved, an ‘8 hr’ PLSS power subsystem is superior to 4 hr PLSS recharging/ refilling/replacing for 

PLSS power levels =<100 W. Replacement of battery modules or fuel cell tank clusters and refilling of 

fuel cell tanks seems to be the next most practical solution in that although the PLSS may be modestly 

less heavy, including the RRS Station in consideration makes the options less competitive with the ‘8 hr’ 

PLSS options (which does not ‘need’ a RRS Station and thus the RRS Station mass contribution is 

essentially zero). 

For the replace/refill RRS Station options, the mass of its power subsystem is modest for 

housekeeping power. Solar power, which provides the lowest RRS Station mass, can be used to handle 

the housekeeping power needs if the Station is used in illuminated areas. Other alternatives are available 

that can operate in un-illuminated conditions at a moderate mass penalty. Any replace or refill PLSS 

options requires dust mitigation for electrical/fluid connections and mechanical parts. Refilling of high 

pressure tanks and handling of high pressure tanks may present a safety concern. Battery modules must be 

placed in a thermally controlled RRS Station container to maintain life until the battery is used in the 

EVA and may need specialized charging prior to use to maintain battery readiness (i.e., some batteries 

discharge slightly over time in a stored state). In addition, the RRS Station should provide the same kind 

of container to store ‘spent’ battery modules in a thermally safe environment until return to ‘Base’. 

Specialized charging may also be needed for this time period. In a related spin-off benefit, using 
NASA/TM—2009-215803 72 

traditional (battery life friendly, safe) recharging rates to recharge battery modules at the RRS Station 

could increase the potential number of EVA sorties. Since no requirements existed for this potential use, it 

was not considered in this trade study. 

Recharging options, from both the PLSS power subsystem and RRS Station perspective, generally 

had the highest mass. Solar power is not viable for recharge power due to the large deployed area and 

high mass. Batteries or fuel cells are required, but have relatively high mass. The high mass is the result 

of high charging power level which is used to size electronics and thermal components. Another 

important finding of this trade study is that the RRS Station size can be reduced significantly with 

sequential instead of simultaneous recharging. 

Li battery types and primary fuel cells have advantageous mass characteristics assuming continued 

improvement of specific energy. Technology development may be needed for some battery types to 

increase battery charge/discharge rates from 1C to 2C with minimal impact on battery life. 

Inductive charging is not suitable for rapid charging due to high power levels (and thus mass) and 

reduced efficiency. Low power inductive applications may still be viable elsewhere in lunar activities 

(such as lunar robotic/manned rovers). Conductive charging is preferred due primarily to reduced losses 

during high power charging. High thermal rates due to rapid recharging over a 15 min period may require 

thermal cooling loops to be connected to the PLSS from the RRS Station during recharging, adding to the 

overall mass and complexity of the charging option. In addition, safety concerns from high 

currents/thermal loads during charging and dust effects on conductive/fluid connections make recharging 

options much less attractive. 

4.8 Additional Information: Rapid Recharge Battery Data 

A literature search regarding the state of the art of rapid recharge batteries was performed. The 

sources of data examined included technical papers, manufacturer presentations, patent applications, press 

releases and web sites. It is apparent that current state-of-the-art rapid recharge battery data is somewhat 

difficult to obtain due to the competitive market pressures. Such batteries have tremendous potential in 

the electric vehicle and consumer electronics markets. Space applications are hardly a consideration for 

such manufacturers. 

4.8.1 Charge Rates and Times 

Secondary (i.e., rechargeable) batteries have different kinds of charging times. These can be grouped 

as ‘slow’ (>3 hr), ‘fast’ (30 min-3 hr) and ‘rapid’ (~15 min). Each type corresponds to a charge rate and 

charging profile which differ based on battery type. 

A charge rate of ‘1C’ means the entire Amp-hour capacity of a battery can be recharged in 1 hr. Thus, 

a 10 A-hr battery can be recharged at 1C (i.e., at 10 A over a period of 1 hr). Also, a 10 A-hr battery can 

be recharged as 2C (i.e., at 20 A over a period of 30 min). Typical charge rates for Li or NiMH batteries 

are 1C although 1.5C and 2C have been indicated to be practical limits. 

4.8.2 Charging Profiles 

Charging profiles can either be constant current (for NiMH battery types) or constant current-constant 

voltage (for Li battery types). Constant current (CC) means that the battery is charged at a specific current 

until the battery is fully charged. Constant current-constant voltage (CCCV) means that most of the 

battery is charged at a constant current until a specific voltage is reached and then the remainder of the 

charging period is performed at this voltage (with a gradually reducing, linear current drop-off). For this 

reason, charging of Li batteries can take longer since the constant current portion (the highest current 

level) can charge from 0 to 80 percent state of charge, but time must be taken for constant voltage 

charging at lower current levels for going from 80 to 100 percent state of charge. 
NASA/TM—2009-215803 73 

4.8.3 Rapid Recharge Battery Sizing Based on Charge Rates 

There are two ways to handle the rapid recharge requirement: oversize the battery (effectively lowers 

the depth-of-discharge from 100 to 50 percent for twice the capacity for instance) or use rapid charge 

batteries. 

With oversizing, nonrapid recharge type batteries (e.g., which recharge >30 min) are used, yet they 

are recharged only for 15 min. This method implies longer battery life and enhanced battery safety due to 

slower recharging rates but this occurs at the price that the battery can only be partially recharged within 

the allotted time of 15 min. In order to still provide the required user energy, more batteries must be 

stored in the PLSS. Thus, effectively the batteries must be oversized to provide the same energy as one 

could obtain if the batteries were allowed to completely be discharged. 

Another aspect of the oversizing method, as applied to Li batteries, is that most of the charging can be 

performed in the constant current part of the charge profile while the constant voltage portion is neglected 

to save time. Of course, this requires life testing to understand the ramifications of such partial charging. 

It is possible that reconditioning methods could be developed. 

Using rapid recharge batteries is the other option. Although these can easily meet the 15 min recharge 

requirements, many questions exist regarding the limitations and applicability of these batteries. More 

testing is required to verify vendor data especially as it applies to a space-rated lunar EVA suit 

application. 

4.8.4 Thermal Effects 

Rapid recharge increases the battery temperature and the entire thermal load on the EVA suit. 

4.8.5 Safety 

Fast or rapid recharging requires sophisticated monitoring and control to ensure battery safety and 

battery life. This includes temperature, pressure and charge monitoring. Hardware designs must address 

these aspects. For example, rapid recharge NiMH batteries may require a pressure vessel to withstand at 

least 150 psi (10.2 atm) while Li type batteries do not seem to have such a requirement. These 

considerations are critical since recharging is proposed to occur with the astronaut inside the suit adjacent 

to the PLSS. Although it is likely that adequate hardware and software designs can alleviate this concern, 

a 15 min recharge is extremely aggressive and will require a great deal of testing to verify its safety and 

effect on battery life. 

4.8.6 Battery Life 

Fast or rapid recharging of batteries has a direct impact on battery capacity reduction over time. 

Research and development is usually focused on reducing such life effects. In addition, partial recharging 

of batteries may have long term battery life effects. It is unknown whether battery reconditioning methods 

can be developed to restore battery capacity to NiMH, Li ion and Li sulfur battery types. Research and 

development is required in this area. 

4.8.7 Rapid Recharge Vendors 

Three primary rapid recharge battery vendors were identified: Electronenergy for NiMH batteries, 

A123 for Li ion batteries, and Sion for Li sulfur batteries. 

4.8.7.3 Electroenergy NiMH batteries 

A technical presentation (Ref. A.1) for the Electroenergy NiMH battery in Nov 2004 shows that their 

6.5 A-hr/1.25V cell with 65 W-hr/kg specific energy can completely recharge in 8.5 min or 7 °C (a 13 Ahr cell in 12.3 min). Details about the life effects of such fast recharging are not available and independent 

confirmation of the test data is not available. Also, it is not clear whether there was battery pressure 

venting or whether the battery was completely sealed. For instance, for a 6.6 A-hr rapid charge cell that 
NASA/TM—2009-215803 74 

recharges in 8.5 min has a pressure rise of 15 psi and a 13 A-hr cell that recharges in 12.3 min has a 

pressure rise of 10 psi. NiMH batteries can vent hydrogen if they are overcharged. 

4.8.7.4 A123 Li Ion 

A patent application for the A123 Li battery (Ref. A.2) in October 2005 shows that their 0.72 A-hr 

cell with 85 W-hr/kg specific energy can be recharged from 0 to 80 percent state of charge in a 6 min 

(constant current) and 80 to 100 percent in 9 min. The A123 patent application covers many versions 

of their product but only one specific energy is listed. It is possible that each version has a different 

specific energy. The A123 website presents data that can be inferred to mean that the cell is either 30 or 

43 W-hr/kg. It is not specifically stated in the patent whether the cell must be completely recharged and 

how partial recharging would affect cell life. It is generally accepted that batteries must be fully recharged 

prior to use. A123 states they can recharge most of the battery in 5 min although the data shows they also 

follow a standard charge cycle of up to 10 min. Pressurization is not required for Li ion battery types. 

4.8.7.5 Sion Li Sulfur 

Technical papers of 2004 by Sion discuss Li sulfur battery data (Ref. A.3, A.4) and indicate that 

prototype Li sulfur cells, operating at 250 to 300 W-hr/kg, demonstrated a >3C charge or discharge rate 

and a temperature range of –60° to 60 °C. Also, 300 to 350 W-hr/kg cells were being tested in 2004. The 

reports show that cycle life (capacity over time) drops off significantly for the high specific energy cells 

(30 percent capacity reduction over 100 cycles for tested 2004 cells, 10 percent capacity reduction over 

100 cycles for estimated 2004 new cells). Cells were used to make batteries for ‘pen tablet’-type 

computers. The projected battery specific energy was 265 W-hr/kg (assuming a cell specific energy of 

350 W-hr/kg contributing 75 percent of the mass). The projected battery volumetric energy was 320 Whr/liter (assuming a cell volumetric energy of 400 W-hr/liter contributing 80 percent of the volume). 

Pressurization is not required for Li sulfur battery types. Independent replication of Li sulfur battery test 

data is required. 

4.8.7.6 Other References for Rapid Recharging of Batteries 

1) Technical papers (Refs. A.5, A.6, and A.7) from 2000 to 2002 indicate that an 85 A-hr NiMH 

battery (11 cells) with 60 W-hr/kg specific energy can be recharged from 20 percent to 70 percent in 

16 min (2.0C). This had a pressure charge control of 150 psi. The implication from the paper is that it is 

all right to partially recharge the battery (it only has a constant charge phase, unlike the Li battery which 

has a constant current and then constant voltage part). However, the authors indicate that more testing is 

needed to determine the battery life impacts of partial recharge. 

2) A technical paper (Ref. A.8) describes the temperature behavior of 1.6 A-hr NiMH batteries. A 

charge input of 80 percent at 3C increases cell temperature from 15 to 60 °C, for 2C it is from 15 to 

40 °C, and for 1C it is from 15 to 30 °C. 

3) A technical paper (Ref. A.9) indicates that a 0.7 A-hr Li battery can be rapidly recharged from 20 

to 100 percent state of charge at 1.5 °C in 34 min. The charging requires constant current and constant 

voltage portions and states that it is critical to monitor temperature and voltage to prevent damage to the 

battery. 

4) A technical paper (Ref. A.10) explores the rapid charging profile using computer-designed 

algorithms and determined that a 2C charge rate can be attained for a 0.93 A-hr Li battery. 

5) A technical paper (Ref. A.11) discusses a rapid charging method (bootstrap charging) for Li 

batteries which applies high currents for a short period at the beginning of charge, but continuing with the 

standard constant current-constant voltage method afterwards. The best that this method can achieve is 

30 min full recharge from starting at 20 percent state of charge (for a 1.1 A-hr battery). Data show that 

life is relatively unaffected by this method (assuming a mandatory standard Li type charge profile after 

the bootstrap). The method works best for a very low starting battery state of charge. Partial charging 

using this method is not considered in the report and testing of this method is required to determine life 
NASA/TM—2009-215803 75 

effects. The report indicates recharge of 20 to 59 percent SOC in 15 min (if partial recharging is allowed 

then the battery must be oversized by 41 percent). In comparison, for a standard Li 15 min charging 

method for the same battery, the recharge would only go from 20 to 45 percent SOC (thus if partial 

recharging is allowed, the batteries must be oversized by 55 percent). 

6) Incomplete and contradictory press release information (Ref. A.12-A.18) from companies called 

Altairnano and ABAT describe a Li battery with specific energies from 40 to 112 W-hr/kg and recharge 

times from 3 to 34 min. Altairnano demonstrated a 3 to 6 min recharge but the Li battery was 40 W-hr/kg. 

Altairnano provided materials to a Li battery manufacturer called ABAT, which has demonstrated a 30 to 

45 min recharge time (although Altair claims better electrode design can reduce this time). Altair claimed 

that the energy density is 3.75 times that of sealed lead-acid battery (30 W-hr/kg) or 100 W-hr/kg. No 

technical data is available and the properties cannot be confirmed. These vendor’s claims are not 

considered in this trade study. 

7) Older Li sulfur battery research and development by Moltech Corporation (aka Sion Power 

Corporation) is discussed (Ref.A.19). The specific energy of a prototype 0.8 A-hr cell is 180 W-hr/kg 

with a charge rate of 4C. The 2003 projected values are 300 W-hr/kg with a charge rate of 2C. In 

addition, the presentation says the future capability may range from 300 to 450 W-hr/kg. In another 

reference for that same year (Ref.A.20), the company projects 300 to 500 W-hr/kg as the future 

capability. 

4.8.8 Assumed Analysis Data 

Because of the broad range of possible values of specific energy, three groups of battery 

characteristics were analyzed in this trade study. The first group is the currently reported battery 

characteristics that are derived from testing by the battery vendors. For NiMH batteries, this is 

Electroenergyinc with a 15 min recharge and 65 W-hr/kg. For Li ion batteries, this is A123 with a 15 min 

recharge and 85 W-hr/kg. For Li sulfur batteries, this is Sion with a 3C recharge rate and 265 W-hr/kg. 

Note that these values have not been independently verified. The second group includes selected vendor 

future projections of their battery technology. The only member of this group is Sion with a 15 min 

recharge rate and 318 W-hr/kg. Other vendors were excluded from this group due to their lower specific 

energies. The final group includes technology projections of traditional Li ion and Li polymer battery 

types. Official NASA projections of these battery characteristics (Ref. A.21) are typically based on low 

charging rates (<2C). It is difficult to project the impact of higher charge rates on specific energy so the 

approach taken was to assume a range of likely specific energies (100, 150 and 200 W-hr/kg) and two 

maximum charge rates (1C and 2C). NiMH batteries were not included in this ‘projected’ group because 

the NiMH characteristics are similar to Li types (Ref. A.22). 

4.8.9 References 

A.1) “High Power Nickel Metal Hydride Battery,” DOE Energy Storage Systems Research Annual 

Peer Review Nov 10, 2004, Washington, D.C., James Landi/ Electroenergyinc. 

A.2) “Lithium Secondary Cell with High Charge and Discharge Rate Capability,” US Patent 

Application, US-20050233220A1, Oct, 20, 2005. 

A.3) “Lithium-Sulfur Rechargeable Batteries: Characteristics, State of Development, and Applicability 

to Powering Portable Electronics,” F. Tudron, J. Akridge, V. Puglisi, Sion Power Corporation, 

Power Sources 2004. 

A.4) “Lithium Sulfur Rechargeable Battery Technology for Computer Application,” Windows 

Hardware Engineering Conference 2004, June 17, 2004. 

A.5) “Fast charging nickel-metal hydride traction batteries,” Xiao Guang Yang, Journal of Power 

Sources 101 (2001), pg. 158-166. 

A.6) “Reliable fast recharge of nickel metal hydride cells,” Bor Yann Liaw, Solid State Ionics (2002), 

pp. 217–225. 
NASA/TM—2009-215803 76 

A.7) “Thermal Behavior of Small Nickel/Metal Hydride Battery during Rapid Charge and Discharge 

Cycles,” Takuto Araki, 2005, Electrochemical Society. 

A.8) “Limiting mechanism on rapid charging NiMH batteries,” Bor Yann Liaw, Electrochimica Acta 

47 (2001), pg. 875–884. 

A.9) “Fast-Charge in Lithium-Ion Batteries for Portable Applications,” J. Lopez, 2004, IEEE. 

A.10) “Search for an Optimal Rapid Charging Pattern for Lithium–Ion Batteries Using Ant Colony 

System Algorithm,” Yi-Hwa Liu, IEEE Transactions on Industrial Electronics, Vol. 52, No. 5, 

Oct. 2005. 

A.11) “Boostcharging Li-ion batteries: A challenging new charging concept,” P.H.L. Notten, Journal of 

Power Sources 145 (2005), pg. 89–94. 

A.12) http://www.greencarcongress.com/electric_battery/

A.13) http://www.greencarcongress.com/2005/06/electrode_nanom.html

A.14) http://www.b2i.us/profiles/investor/ResLibraryView.asp?ResLibraryID=6313&Go 

Topage=2&BzID=546&Category=564

A.15) http://www.b2i.us/Profiles/Investor/ResLibraryView.asp?ResLibraryID=8990&Go 

Topage=1&BzID=546&Category=24

A.16) http://www.b2i.us/profiles/investor/ResLibraryView.asp?ResLibraryID=9144&Go 

Topage=1&BzID=546&Category=24

A.17) “Li4Ti5O12/poly(methyl)thiophene asymmetric hybrid electrochemical device,” A. Du Pasquier, 

Journal of Power Sources 125 (2004), pg. 95–102. 

A.18) “Power-ion battery: bridging the gap between Li-ion and supercapacitor chemistries,” A. Du 

Pasquier, Journal of Power Sources 136 (2004), pg. 160–170. 

A.19) “Lithium-Sulfur Technology, Current Performance and Future Potential,” PO WER 2001, 

Moltech Corporation, October 2001. 

A.20) “Lithium Sulfur Rechargeable Battery Safety,” J.R. Akridge, Moltech Corporation, Battery 

Power Products & Technology, October 2001. 

A.21) “Energy Storage Technology for Future Space Science Missions,” JPL D-30268, Nov 2004. 

A.22) “Electrochemical Energy Storage Technology Paths--Batteries,” Tom Miller, NASA GRC, 

December 17, 1999. 

4.9 Additional Information: Lunar Surface Recharging Methods 

For battery-type energy storage, there are two methods to recharge the modules in place. First is a 

direct electrical hookup. This method requires some level of dust mitigation, high levels of current (for 

rapid recharge) and high heat loads in the batteries, charge regulators and cables. The second method is 

called inductive charging. It relies upon half of a transformer coil placed on the suit and the other half of 

the coil at the RRS Station. The each coil is surrounded with a nonconductive shell which is designed to 

self-align when the recharging procedure is to begin. This method has been designed for underwater 

vehicle recharging (Ref. B.1, B.2) where fouling of electrical connector ports is a problem. This system 

converts a DC power source (as would be generated by the RRS Station battery, fuel cell, solar array or 

radioisotope power system source) to 65 to 80 kHz sine wave AC power using an inverter so that the 

power can be transferred using a transformer. When the coil halves are placed in close proximity to each 

other, the power is transferred across the RRS Station coil to the PLSS power subsystem-side coil. A 

rectifier converts the AC power back to DC power. Finally, a power conditioner (dc-dc converter) is 

needed to provide the PLSS power subsystem battery charger with regulated voltage and current. For the 

underwater system (which was built and tested successfully), the efficiency of the power transfer was 

80 percent for power levels around 200 W. 

The following information lists the details of the referenced underwater system. The coil design used 

a type 78 ferrite material because of its low-loss, high-frequency properties. The windings of each coil 

used 28 bi-filar turns of #22 Heavy Poly-Thermaleze insulated magnet wire. The winding conformation 

had an aspect ratio of 2 to 1 and was chosen to minimize the profile of the winding. The input voltage 
NASA/TM—2009-215803 77 

level was 48 to 60 V. The protective housing of each coil was made of Lexan. The maximum energy 

transfer capability was rated at 200 W-hr/hr (Ref. B.3). The weight of the system was 11 lb and 6.5 in. in 

diameter (Ref. B.4). 

A high power design (Ref. B.4) is described to be a 5 in.3

 battery-side transformer weighing less than 

5 lb (same weight and size for the charger side) with the potential to provide 1000 W using four separate 

transformer couplings (at 250 W each). Since this system seems to be the most applicable to the EVA suit 

recharging since it can be scaled to the appropriate power level or 1000 W/5 lb (.0027 kg/w). Such a unit 

has not been built, but seems practical. 

Note that inductive charging is also currently in use with terrestrial electric vehicles (Ref. B.5) in lieu 

of direct electrical connection. The reason for considering inductive charging is that direct electrical 

connections may have safety issues (incorrect plugging, shock hazard, etc). Although several brands of 

inductive electric vehicle chargers have been produced (Delco Electronics is the most popular), direct 

electrical connection seems more popular. The power level of such chargers are high (6500 W) but they 

are for high charge voltages (200 to 500 V). The efficiency of these inductive charging systems is around 

90 percent. The weight of the entire inductive charge system is not clear from the literature but they are 

not weight optimized since most of the electronics are not on the vehicle, only the inductive coupler is 

likely to be low weight. Charge times are typically multi-hour. 

The most obvious drawback for inductive charging is that since there are inherent losses in such a 

system, an inductive-type RRS Station must be larger than one using direct electrical connections. 

However, this drawback must be weighed against the enhanced safety due to the reduced dust interfaces 

impacts of the inductive system. 

Inductive charging is not a valid approach for rapid recharging at high power levels since the mass of 

the transformer couplings on the PLSS-side is too high. For lower power levels below 250 W, the mass is 

likely acceptable, but this results in much longer recharge times. The most likely application may be for 

power transfer from a rover seat transformer coupler during long drives or perhaps for manned or robotic 

rover charging outside the manned base. 

4.9.1 References 

B.1) “Power Systems for Autonomous Underwater Vehicles,” Albert M. Bradley, IEEE Journal of Oceanic 

Engineering, Vol. 26, No. 4, Oct. 2001. 

B.2) “An Interface System for Autonomous Undersea Vehicles,” Michael D. Feezor, Vol. 26, No. 4, Oct. 

2001. 

B.3) “Sensor Power and Telemetry Requirements for DEOS Buoy Observatory,” 

http://obslab.whoi.edu/InstrPowerTelemetry.pdf. 

B.4) “Multi-User Autonomous Underwater Vehicle Docking Systems,” Edgar An, Grant #: N00014–98–

1–0861, http://www.onr.navy.mil/oas/reports/docs/om/03/oman2.pdf. 

B.5) “Design Considerations For Power Converters Supplying The SAE 5-1773 Electric Vehicle Inductive 

Coupler,” Nasser H. Kutkut, 1997, IEEE. 
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4. TITLE AND SUBTITLE

Compilation of Trade Studies for the Constellation Program Extravehicular Activity 

Spacesuit Power System 

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6. AUTHOR(S)

Fincannon, James 

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WBS 731384.06.04.01.31.06.50 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

This compilation of trade studies performed from 2005 to 2006 addressed a number of power system design issues for the Constellation 

Program Extravehicular Activity Spacesuit. Spacesuits were required for spacewalks and in-space activities as well as lunar and Mars 

surface operations. The trades documented here considered whether solar power was feasible for spacesuits, whether spacesuit power 

generation should be a distributed or a centralized function, whether self-powered in-space spacesuits were better than umbilically powered 

ones, and whether the suit power system should be recharged in place or replaced. 

15. SUBJECT TERMS

Extravehicular activity; Spacesuit; Power system; Solar power; Battery; Energy storage; Fuel cell; Solar arrays; Umbilical power; 

Power distribution; Centralized versus distributed power; Inductive versus direct power coupling; Flywheel; Radioisotope 

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