Solar Power

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National Aeronautics and Space Administration
1
www.nasa.gov
Solar Power for Outer Planets Study
Solar Power for Outer Planets Study
Presentation to Outer Planets Assessment Group
November 8, 2007
Scott W. Benson/NASA Glenn Research Center
National Aeronautics and Space Administration
www.nasa.gov
National Aeronautics and Space Administration
2
www.nasa.gov
Background & Outline
Background & Outline


Alan Stern request: “…a quick look study for how we
could extend the Juno and Rosetta 5 AU-class
missions on solar arrays to enable solar array
missions at Saturn (10 AU) and Uranus (20 AU)”



Study Process
Study Process



Cell and Array Technology Findings
Cell and Array Technology Findings



Power System Sizing
Power System Sizing



Mission and System Integration Studies
Mission and System Integration Studies



Technology Planning
Technology Planning



Conclusions
Conclusions
National Aeronautics and Space Administration
3
www.nasa.gov
• •

Dawn Dawn


36.4 m
2
planar array area


10.3 kW at 1 AU


1.3 kW at 3 AU (-88 °C)


Triple Junction cells
• •

Juno Juno


Phase B design


45 m
2
planar array area


9.6 kW BOL at 1 AU


414 W at 5.5 AU (-130 °C)


Triple Junction cells
• •

Rosetta Rosetta


61.5 m
2
planar array area


7.1 kW BOL at 1 AU


400 W at 5.25 AU (-130 °C)


Silicon Hi-ETA cells
Most Distant Use of Solar Arrays
Most Distant Use of Solar Arrays
National Aeronautics and Space Administration
4
www.nasa.gov
Study Process
Study Process



Review prior studies and flight system publications
Review prior studies and flight system publications



Assess PV cell and array technologies
Assess PV cell and array technologies



Understand cell performance in outer planet applications
Understand cell performance in outer planet applications



Analyze power system performance
Analyze power system performance



Coordinate with technology, vendor and user community
Coordinate with technology, vendor and user community
through workshop at Space Photovoltaic Research &
through workshop at Space Photovoltaic Research &
Technology (SPRAT)
Technology (SPRAT)



Coordinate with Juno project
Coordinate with Juno project



Characterize system integration considerations
Characterize system integration considerations



Define technology paths
Define technology paths
National Aeronautics and Space Administration
5
www.nasa.gov
Solar Cell Technology Findings
Solar Cell Technology Findings
Solar Cell Capability
Solar Cell Capability


Nominal low intensity, low temperature
(LILT) state-of-the-art (SOA) cell
performance is viable at 5 AU and beyond


Cell efficiency increases with lower
temperature but decreases with lower
intensity


LILT Effect: off-nominal drop

in cell
performance, must be mitigated to
effectively use solar power in outer
solar system


Understood and mitigated on earlier silicon cells


Effect observed on SOA multi-junction (MJ) cells, cause not yet identified
Cell-to-cell variation


LILT Effect can be mitigated:
Cell screening, optimization or advanced concentrator technology


On-going advances in cell technology can provide improvements


NASA will need to adapt those to LILT conditions
GRC FY07 LILT IRAD testing results
Low Light Triple Junction Performance
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-180 -160 -140 -120 -100 -80 -60 -40 -20 0 20 40
Temperature ( C
1 AU
5.6 AU
22.1 AU
National Aeronautics and Space Administration
6
www.nasa.gov
Applicable Technologies
Applicable Technologies



Solar Cells
Solar Cells
• •

State State- -of of- -art performance at 1AU (AM0, 25C) art performance at 1AU (AM0, 25C)


Multi-junction III-V cells, triple-junction: 28 - 30%


Silicon: 16 - 19%


Thin-film: not space-qualified (6 - 10% currently)
• •

Expected advances in cell performance Expected advances in cell performance


Multi-junction: 30 - 33% in next 3 years
Development pursued by both cell vendors
Driven by military/commercial applications
35 - 40% cell design under development


Multi-junction: mass and cost reduction
Thinned substrate or no-substrate technology
to drastically reduce cell/array mass
Reusable substrates and improved manufacturing
to increase yield and reduce cost
• •

Advanced cell approaches Advanced cell approaches


Cells designed or optimized for outer solar system missions
Eliminate LILT Effect in future MJ cell generations
Optimize cells for bandgap narrowing at low temperatures


Quantum dots, nanotechnology to increase efficiency
Far-term: efficiency increase through better utilization of solar spectrum
Projected Efficiency for MJ Solar Cells
10
15
20
25
30
35
1990 1995 2000 2005 2010
Year
Projected Efficiency for MJ Solar Cells
10
15
20
25
30
35
1990 1995 2000 2005 2010
Year
National Aeronautics and Space Administration
7
www.nasa.gov
Solar Cell LILT Effects
Solar Cell LILT Effects
Data collected at GRC
on SOA MJ production-
line cells


Cause of LILT Effect has not been identified for multi-junction III-V

cell technology


Occurs on a cell-to-cell basis


Magnitude of effect is not currently predictable


Little research has been done in this area, none to understand root cause of problem
0
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0 0.5 1 1.5 2 2.5 3 3.5
Voltage (V)
Blue Curve Data
FF = 64.1%
Efficiency ~ 24.8 %
Red Curve Data
FF = 77%
Efficiency ~ 29.8 %
Intensity = .02 Suns (7 AU)
Temperature = -150°C
LILT
Effect
National Aeronautics and Space Administration
8
www.nasa.gov
LILT Effect Mitigation
LILT Effect Mitigation
Cell Screening


Successful on
Dawn


In progress
for Juno


GRC and Juno data
indicate that effect
worsens in frequency
and magnitude
with lowering intensity


Cell screening may not
be applicable beyond
Jupiter
Cell Optimization


Silicon cells designed

for LILT on Rosetta


5.2 AU, -130 °C


Future cells could be optimized


To eliminate LILT Effect


To optimize cell performance and mass
for LILT conditions
Concentration


Maintains

intensity


Minimizes

LILT Effect


Reduces cell

count


Increased spacecraft system

effects (pointing requirements)
Low Light Triple Junction Performance
18
20
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28
30
32
34
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-180 -160 -140 -120 -100 -80 -60 -40 -20 0 20 40
Temperature ( C
1 AU
5.6 AU
22.1 AU
5% cell efficiency
benefit at Saturn with
8X concentration
National Aeronautics and Space Administration
9
www.nasa.gov
Array Technology Findings
Array Technology Findings
Advanced Solar Array Technology
Advanced Solar Array Technology


Multiple technical paths exist to extend photovoltaic power use towards
the outer solar system


UltraFlex


Near-term, high maturity


Baseline for Orion power
TRL6 by 2009 with subsequent qualification


SquareRigger


Mass competitive at large power levels


Rectangular bays offer better scaling characteristics


Compatible with planar and concentrator designs


Stretched Lens Array SquareRigger (SLASR)


Incorporates lightweight linear refractive concentrator
derived from Deep Space 1 SCARLET


SLA component flight demonstration on TACSAT-4


Can scale to very high power levels


Technology development is required:


To extend UltraFlex diameter beyond state-of-art size


To complete SquareRigger development at the array level
UltraFlex

Wing
SLASR Bay
2.5 x 5m
National Aeronautics and Space Administration
10
www.nasa.gov
0
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Distance (AU)
P
o
w
e
r

(
W
)
SLASR: 179 kg, 211m^2(~18bays)
SLASR: 89 kg, 105m^2(~10bays)
Four wing, 6 m diameter Ultraflex: 179 kg, 113 m^2
Planar Squarerigger: 179 kg, 86m^2(~8bays)
Two wing, 6 m diameter Ultraflex: 89 kg, 56 m^2
Planar Squarerigger: 89 kg, 43m^2(~4bays)
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Distance (AU)
P
o
w
e
r

(
W
)
SLASR: 179 kg, 211m^2(~18bays)
SLASR: 89 kg, 105m^2(~10bays)
Four wing, 6 m diameter Ultraflex: 179 kg, 113 m^2
Planar Squarerigger: 179 kg, 86m^2(~8bays)
Two wing, 6 m diameter Ultraflex: 89 kg, 56 m^2
Planar Squarerigger: 89 kg, 43m^2(~4bays)
Near Term Capability:
Near Term Capability:
Power from a Fixed Power System Mass
Power from a Fixed Power System Mass


Solar array options are sized to be the same mass as the two and

four wing Ultraflex
arrays.


Solar array mass includes hardware outboard of gimbal


SquareRigger standard bay size of 2.5 m by 5 m


2.5 yr gravity assist period + 2 years at AU location + 2 AU/year transit time


Heliocentric space
(no planetary radiation
effects)


Useful power will be
lower based on

planetary eclipses/
radiation degradation


SOA cell efficiency of
30% at 1 AU


LILT Effect-free cells


8X concentration with
SLASR
*
Juno
-

SOA cells
-

planar array
-

mass not
normalized
National Aeronautics and Space Administration
11
www.nasa.gov
System Integration Considerations
System Integration Considerations
• •

Mass impacts of carrying the solar array into deep space Mass impacts of carrying the solar array into deep space


Additional/larger systems: solar array, batteries, power conditioning systems, pointing
systems (larger reaction wheels)


Heavier thermal systems (lack of RPS waste heat)


Structures/mechanisms to attach the solar arrays
(impact from capture propulsion system)


Net impact is reduced payload compared to RPS systems
• •

Launch vehicle integration Launch vehicle integration


Volume constraints in packaging stowed arrays
• •

Spacecraft integration and operations in mission orbit Spacecraft integration and operations in mission orbit


Multiple subsystem requirements for pointing and slew


Possible incompatibilities with science objectives
• •

Power system design Power system design


Maintaining power through eclipse periods


Radiation tolerant design


Managing power in inner solar system, when generated
power from array can be 10’s - 100’s of kW
National Aeronautics and Space Administration
12
www.nasa.gov
Mission Applications
Mission Applications
• •

A range of missions were considered to encompass power system si A range of missions were considered to encompass power system sizing zing
and spacecraft integration drivers, including: and spacecraft integration drivers, including:
– –

Heliocentric distance: 5 Heliocentric distance: 5 - - 20 AU 20 AU
– –

Operations concept & power management: moon orbiters Operations concept & power management: moon orbiters
– –

Radiation: Radiation: Jovian Jovian moon orbiter moon orbiter
– –

Simplest missions: flybys Simplest missions: flybys
• •

Flagship Flagship- -class class


Saturn Orbiter, Titan Orbiter


Uranus Orbiter


Ganymede or Europa Orbiter
• •

PI PI- -led led


Saturn Flyby


Centaur Flyby or Rendezvous
• •

Saturn Orbiter analyzed in COMPASS team study Saturn Orbiter analyzed in COMPASS team study


Used GSFC Enceladus architecture option Saturn-OL as reference
• •

Europa Europa, Centaur and Uranus missions assessed analytically , Centaur and Uranus missions assessed analytically


Representative point analyses performed with selected mission power
National Aeronautics and Space Administration
13
www.nasa.gov
Example: Saturn Orbiter Mission
Example: Saturn Orbiter Mission
• •

Mission assumptions: Mission assumptions:


Titan/Enceladus cycling orbit


335 W continuous nominal power (per Enceladus study)


11.5 yr VVEEGA voyage to Saturn


Saturn and rings eclipse periods


Total radiation degradation of 15%
• •

Power system design options Power system design options


SOA cells/array:
Nine SLASR bays at 237 kg
Twelve Planar Squarerigger bays at 470 kg
Four, 7.2 m diameter Ultraflex arrays at 415 kg


Projected cells/array:
Eight SLASR bays at 205 kg
Ten Planar Squarerigger bays at 321 kg
Four, 6.7m diameter Ultraflex arrays at 268 kg
• •

System System- -level drivers level drivers


COMPASS study performed to assess system drivers,
details follow
• •

Technology feasibility Technology feasibility


Target power level can be achieved with near-term PV technology
Near-SOA Ultraflex

Arrays
SOA MJ Cell Performance
48 kW BOL at 1 AU
Interplanetary voyage
377 W at Saturn arrival
-

Planetary radiation
- Energy storage for

eclipse periods
335 W EOL power
40 m
2
National Aeronautics and Space Administration
14
www.nasa.gov
Other Missions
Other Missions
System System- -level drivers level drivers


Radiation and eclipse power level


Orbital operations and pointing
requirements
Technology feasibility Technology feasibility


SLASR and UltraFlex provide
feasible paths


Four, 7.0 m diameter Ultraflex

arrays at 513 kg


Requires detailed radiation
degradation trade study
Advanced SLASR Arrays Advanced SLASR Arrays
SOA MJ Cell Performance SOA MJ Cell Performance
45 kW BOL at 1 AU, 362 kg
Interplanetary voyage
1400 W at Jupiter arrival
-

Planetary radiation, 30%

total degradation
- Energy storage for

eclipse periods
720 W EOL at Europa
50 m
2
Europa Orbiter Europa Orbiter
SOA SOA Ultraflex Ultraflex Arrays Arrays
SOA MJ Cell Performance SOA MJ Cell Performance
33 kW BOL at 1 AU, 287 kg
Interplanetary voyage
300 W EOL at Echeclus
27 m
2
System System- -level drivers level drivers


Similar to Saturn Orbiter study
Technology feasibility Technology feasibility


Target power level can be
achieved with near-term PV
technology
Advanced SLASR Arrays Advanced SLASR Arrays
Projected MJ Cell Performance Projected MJ Cell Performance
95 kW BOL at 1 AU, 327 kg
Interplanetary voyage
200 W EOL at Chiron
100 m
2
Centaur: Centaur: Echeclus Echeclus Centaur: Chiron Centaur: Chiron
System System- -level drivers level drivers


Array size amplifies all spacecraft
integration considerations
Technology feasibility Technology feasibility


Technology is achievable


Significant changes in
spacecraft concept required
Advanced SLASR Arrays
Projected MJ Cell Performance
232 kW BOL at 1 AU, 781 kg
Interplanetary voyage
400 W EOL at Uranus
250 m
2
Uranus Orbiter Uranus Orbiter
System System- -level drivers level drivers


Array size amplifies all spacecraft
integration considerations
Technology feasibility Technology feasibility


Significant changes in
spacecraft concept required
National Aeronautics and Space Administration
15
www.nasa.gov
Solar Saturn Probe Design Study
Solar Saturn Probe Design Study
• •

Reference ASRG Reference ASRG- -powered spacecraft powered spacecraft


Power: 335 W


3 ASRGs


Science payload: ~1000 kg, includes lander
• •

Solar Solar- -powered spacecraft powered spacecraft


48 kW solar arrays at 1AU


Science payload: ~550 kg
Solar Powered
Subsystem
Mass Change Compared
to RPS probe
Cause
Payload -

450 kg Increase in bus subsystems mass
Power + 340 kg Solar arrays, mechanisms, PMAD
ACS + 30 kg Heavier wheels (ACS propellant increased)
C&DH/Comm + 15 kg Increase in pointing, more complex spacecraft operations
Thermal + 30 kg Additional blankets, heaters, RHUs

due to lack of waste heat for RPS
Structures + 30 kg Solar array booms
National Aeronautics and Space Administration
16
www.nasa.gov
Power at Saturn:


Interplanetary-only (no
eclipses, planetary radiation)


Fixed mass power system
Key Technology: Key Technology:

Cell Improvements Cell Improvements

+5% eff., lower mass

+58% Power
SLASR/SOA Cells SLASR/SOA Cells
360 W 360 W, 233 kg, , 233 kg,
9 bays (2.5 m x 5 m) 9 bays (2.5 m x 5 m)
SLASR/ SLASR/

Projected Cells Projected Cells
420 W 420 W, 233 kg, , 233 kg,
9 bays (2.5 m x 5 m) 9 bays (2.5 m x 5 m)
Key Technology: Key Technology:

Lightweight Arrays Lightweight Arrays

+80% Power
+110% Power
Power improvements
achievable through
technology investment

Power improvements
achievable through
technology investment
Technology Leverage Summary
Technology Leverage Summary
Underlying Development and Technologies Underlying Development and Technologies


Qualify UltraFlex for low temperature application


LILT Effect Evaluation for MJ Cells


Blanket Technologies for Low Temperature Conditions
UltraFlex UltraFlex/ /
SOA Cells SOA Cells
200 W 200 W, 233 kg, , 233 kg,
4 4 wings, wings,
5. 5.4 m diameter 4 m diameter
UltraFlex/

Projected Cells
316 W, 233 kg,
4 wings,

6.3 m diameter
National Aeronautics and Space Administration
17
www.nasa.gov
Conclusions
Conclusions
• •

Near Near- -term term Ultraflex Ultraflex

arrays and SOA multi arrays and SOA multi- -junction cells can provide junction cells can provide
capability to perform low power (200 capability to perform low power (200- -300 W) missions out to 10 AU 300 W) missions out to 10 AU
– –

300 W mission to an inner Centaur appears achievable 300 W mission to an inner Centaur appears achievable
• •

Further investigation of LILT Effect is warranted if PV power is Further investigation of LILT Effect is warranted if PV power is

to be to be
considered for more demanding outer planet missions considered for more demanding outer planet missions
– –

LILT Effect can be avoided through multiple approaches LILT Effect can be avoided through multiple approaches
• •

Advanced cell and array technologies would extend the practical Advanced cell and array technologies would extend the practical
application of PV power through mass and efficiency benefits application of PV power through mass and efficiency benefits
– –

Clear technology paths exist to enhance PV application to outer Clear technology paths exist to enhance PV application to outer planet planet
missions missions
• •

Implementation of PV power will decrease payload mass Implementation of PV power will decrease payload mass
• •

Feasibility of PV use critically depends on mission and spacecra Feasibility of PV use critically depends on mission and spacecraft ft
concept concept

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