The Earth Observing System, which is part of the International Mission to Planet Earth, is NASA's main contribution to the Global Change Research Program. Five large platforms are to be launched into polar orbit: two by NASA, two by the European Space Agency, and one by the Japanese. In such an orbit the radiation resistance of indium phosphide solar cells combined with the potential of utilizing 5 micron cell structures yields an increase of 10 percent in the payload capability. If further combined with the Advanced Photovoltaic Solar Array, the total additional payload capability approaches 12 percent

Bailey, Sheila G.

Flood, Dennis J.

Weinberg, Irving

English

NASA Technical Reports Server

NASA Technical Memorandum 105222
_ Advanced Power Systems for EOS
-(NASA-TM-105222) AOVANC_D POWER SYSTEMS FOR N91-31217
EOS (NASA) B p CSCL 21H
i '_ _i
Unc1 _,s
G3/20 0040120
Sheila G. Bailey, Irving Weinberg,
and Dennis J. Flood
Lewis Research Center
Cleveland, Ohio
Prepared for the ........................
...... 26th Intersociety Energy Conversion Engineering Conference
cosponsored by ANS, SAE, ACS, AIAA, AS/vIE, IEEE, and AIChE
Boston, Massachusetts, August 4-9, 1991
https://ntrs.nasa.gov/search.jsp?R=19910021903 2020-03-19T17:20:32+00:00Z

ADVANCED POWER SYSTEMS FOR EOS
Sheila G. Bailey,IrvingWeinberg, and DennisJ. Flood
NationalAeronauticsand Space Administration
Lewis ResearchCenter
Cleveland,Ohio44135
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ABSTRACT
The Earth Observing System, which is part of the
International Mission to Planet Earth, is NASA's main
contribution to the Global Change Research Program. The
program opens a new era in international cooperation to
study the Earth's environment. Five large platforms are to
be launched into polar orbit, two by NASA, two by the
European Space Agency, and one by the Japanese. In
such an orbit the radiation resistance of indium phosphide
solar cells combined with the potential of utilizing 5 micron
cell structures yields an increase of 10% in the payload
capability. If further combined with the Advanced
Photovoltaic Solar Array, the total additional payload
capability approaches 12%.
INTRODUCTION
The international scientific community is organizing
research efforts to advance our knowledge of both natural
and human-induced global change. The U. S. Global
Change Research Program, a consensus interagency
plan, defines the U. S. element of those efforts. Mission
to Planet Earth, the central NASA contribution to the this
program includes two proposed initiatives: the Earth
Observing System (EOS) and Earth Probes. EOS
consists of a space-based observing system, a data and
information system, and a scientific research program.
This represents the initiation of a comprehensive, global
observing system with broad and high-resolution spectral
and spatial, as well as long-term temporal, coverage of the
Earth. The space component will consist of two series of
polar-orbiting platforms: EOS-A and EOS-B, with the
earliest possible launch of EOS-A in late 1997. EOS will
be supplemented by companion European and Japanese
platforms, as well as the continufng operational
environmental satellites. The 15-year observational
period will be achieved using three identical satellites per
series, each with a five-year design lifetime.
EOS-A DESCRIPTION
Figure 1 illustrates the EOS-A platform [Ref. 1]. EOS-A
has a 5-year design life, a payload mass of 3,000 to 3,500
kg, 6 kW average power for the platform itself and 3.2 kW
for the payload, 300 Mbps peak data rate, 30 to 50 Mbps
average data rate, and 0.5 terabyte of on-board data
storage. The present baseline calls for replacement of the
platform after 5 years; however, the platform is designed
to be serviceable if an effective means of on-orbit
servicing becomes available during the life of the mission.
The satellite will be launched into polar orbit from
Vandenberg Air Force Base on a Titan-IV rocket. The
need for global coverage every one to three days dictates
a sun-synchronous orbit with a quasi-two-day repeat; a
705 km altitude, 98.2-degree inclination orbit meets this
need. It will have a 13:30 local equatorial-crossing time.
The platform science objectives include global
quantification of the hydrologic cycle, monitoring of the
upper atmosphere and surface radiation flux, a
comprehensive study of clouds, the characterization of
biological activity and ecosystems through optical
measurements, an improved study of atmospheric
circulation, and the global determination of surface
mineralogy.
POWER SYSTEM WITH SILICON SOLAR
CELLS
A currently considered platform design allocates 1407 kg
to providing electrical power using large area (8 x 8) cm
silicon (Si) cells, 0.2 mm thick, covered with a 0.125 mm
thick ceria stabilized glass microsheet. Improvements in
the photovoltaic performance of the array would reduce
the mass and permit more of the spacecraft mass fraction
to be allocated to orbital communications, observations,
and other equipment packages.
The large area silicon cell, 10 ohm-cm base resistivity,
with dual anti-reflective coating and a back surface field
has an average efficiency of 14.2% at 28°C, beginning of
life (BOL). These cells are currently under production for
Space Station Freedom [Ref. 2].
The array would consist of approximately 58 pairs of
panels with each panel containing 200 series wired cells.
The panel pair is also wired in series to produce a nominal
160 Vdc output. The array would be designed to produce
12.2 kW at 5 years and 9.76 kW at 7.5 years. The payload
requires 3.2 kW average and 4.2 kW peak.
The silicon cells are mounted on a light flexible kapton
substrate to decrease the array mass, however, it also
leaves the rear surface of the cell much more vulnerable to
radiation. The equivalent shielding thickness for the rear
cell surface of a back surface field (BSF) cell .2 mm thick
is .05 mm [Ref. 3]. In an orbit of 98.2 ° at 705 km this
results in a 5-year total equivalent
Figure1. EOS-Asatellite
fluenceof1.39E14e'/cm2forprotonsand4.7E12e-/cm2
forelectrons[Ref.4](seeFigures2and3). Theresulting
degradationfactor(efficiencyEOL/efficiencyBOL)is .82
after five years.
POWER SYSTEM WITH INDIUM PHOSPHIDE
SOLAR CELLS
Indium Phosphide (InP) solar cells are excellent
candidates for space power generation. Total area
efficiencies over 19%, air mass 0 (AM0), are now routinely
achieved for cells 4 cm 2 in area [Ref.5]. Figures 4 and 5
illustrate the superior radiation resistant properties of InP
[Refs. 6,7, and 8]. In addition a small module containing
InP solar cells has shown no degradation after four years
in space [Ref. 9]. A large-scale commercial process has
been developed to produce high efficiency diffused
junction InP solar cells and one thousand of these cells
are the power source for a lunar orbiter on board the
Japanese ISAS scientific satellite "MUSES-A", which was
launched January, 1990. However there was no telemetry
data from this satellite after one month.
Recent data have shown that a fluence of 10 MeV protons
produces damage in InP that can be approximated by a
fluence of 1 MeV electrons, about 650 times that of the 10
MeV proton fluence [Ref. 8]. This compares to 3500 for
silicon and 1000 for gallium arsenide [Ref. 10]. Utilizing
this relative damage coefficient and the silicon equivalent
fluences yields a degradation factor of .997 for InP cells
with .05 mm coverglass after 5 years in orbit. The total
annual equivalent 1 Mev electron fluence from trapped
protons and electrons is 2.3 E13 e'/cm 2. However, InP
cells, on board the LIPS III satellite with 12 mil coverglass,
in a nearly circular orbit of 1100 kilometers altitude with an
inclination of 60 +0 for four years have a radiation
environment of approximately 1.8 times the total fluence
of five years in the EOS orbit (after adjusting for the
differences in coverglasses) and have exhibited no
measurable degradation [Ref. 9]. Annealing under
continuous minority carrier injection by sunlight may
explain the apparent lack of degradation for these cells
[Ref.7].
Figure 2.
Figure 3.
70 . . . .
uJ 60
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40
30 I , I I I
0 2 6 8
, I ,
4
Shield Thickness (mil)
W 2OO
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0 • I • | . i . I • I _ I i
0 1 2 3 4 5 6 7
Shield Thickness (mil)
Annual equivalent 1 MeV electron fluence for silicon from trapped protons (Pmax, Voc),
705 km altitude, 100 ° inclination, as a function of shield thickness
Annual equivalent 1 MeV electron fluence for silicon from trapped electrons, 705 km
altitude, 100 ° inclination, as a function of shield thickness
300 , • , • ",'_ . , .. , , ,
CI
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E
O
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0.9
0.8
0.7 "
0.6
,-, Si (Ref. 3)
• InP (Refs. 7 & 8)
0+5 • • ,,,,,i • .,,,,,1
ioI:_ lo13 lOli
• .,•=..| • •lo1 ;o
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1 MeV Electron Fluence, (e/cm)
017
Figure 4. Normalized efficiency after 1 MeV electron irradiation
¢-
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0o7"
0.6
10 10
Figure 5.
• InP-1 [Ref. 7]
'b
• InP-2 [Ref. 8]
1011 1012 1013 1014
2
10 MeV Proton Fluence, (e/cm)
Normalized efficiency after 10 MeV proton irradiation
Therecent achievement of 19.1% AMO efficiencies with 4
cm2 InP cells [Ref. 5] indicates that increased efficiencies
and larger areas are indeed feasible. However, one
impediment to the development of InP space cells has
been the relatively high cost of InP wafers. Two potential
solutions to this problem are either heteroepitaxial InP
solar cell structures on silicon substrates [Ref. 11] or ultra
thin InP solar cells produced by the CLEFT [Ref. 12] or
peeled film techniques [Ref. 13] and bonded to a silicon
substrate. Modeling calculations indicate that efficiencies
of 22.5% are feasible in the laboratory [14]. Production
capability is therefore assumed to be able to achieve a
20% efficient, 5 micron InP solar cell structure on a 55
micron silicon substrate. This cell will be used to
demonstrate the mass saving on both a Space Station
Freedom array described earlier and a retractable version
of the Advanced Photovoltaic Solar Array (APSA) [Ref.
13] under development at NASA's Jet Propulsion
Laboratory.
COMPARISON AND CONCLUSION
In a relatively high radiation environment InP has a clear
advantage in radiation tolerance. Figure 6 compares the
degradation in normalized efficiency as a function of the
number of years in an EOS orbit for a silicon solar cell with
a 5 rail cover glass and an InP solar cell with a 2 mil cover
glass. The 5-year degradation factor of silicon is .82
compared to .997 in indium phosphide for the EOS orbit
under consideration. A comparison of temperature
coefficients, -0.042%/°C for InP and -0,066%/°C for Si,
1,0
Figure 6.
and assuming a 60°C operating temperature, a 1.47%
decrease is expected in InP and a 2.1% decrease in Si. In
addition the expected increase in temperature due to
radiation degradation gives approximately a .07%
advantage to InP at the 5-year design point.
A major advantage in the utilization of InP occurs because
the increased optical absorption of InP permits a cell
design requiring only 5 microns of solar cell material
without a loss of efficiency. A 55 micron silicon substrate
and 50 micron coverglass were chosen to be compatible
with the APSA program design. Without considering the
impact of a reduced array size on other components of the
electrical power system, simply changing the cells of the
proposed silicon array to indium phosphide without altering
any other structural components saves 298 kg. This
would represent a 10% increase in payload capability.
The APSA wing is an ultra-iightweight flexible foldout
blanket with a weight optimized deployment mechanism
consisting of a fiberglass triiongeron lattice mast deployed
from a cylindrical aluminum cannister and deployment
actuator [Ref. 15]. Using both the APSA array and InP
cells secures a gain of 352 kg, which is approximately an
11.7% increase in payload capability.
The preceding results provide a compelling reason to
pursue the use of InP solar cells for space applications.
Todate the space use of these cells has been limited but
the data obtained does reflect the excellent radiation
resistance exhibited in the laboratory.
O
t-
.__ 0.9.
O
,w¢:
uJ
"1oG>
.N
03
0
z 0.8 [] InP (2 mil coverglass)
• Si (5 mil coverglass)
_.--_ =7
0.7 .... , I ' I I I I
0 2 4 6 8 10 12
Years in Orbit
Normalized efficiency as a function of the number of years in a 705 km polar orbit
6. REFERENCES
[1]EOS Reference Handbook, Goddard Space Flight
Center, 1990.
[2] G. F. Turner and S. C. Debrock, Space Power, Vol. 8,
Nos. 1/2, pp. 11-22,1989.
[3] B. E. Anspaugh, Solar Cell Radiation Handbook,
Addendum 1: 1982-1988.
[4] H. Y. Tada, J. R. Carter, Jr., B. E. Anspaugh, and R.
G. Downing, Solar Cell Radiation Handbook, Third Edition,
1982.
[5] C. J. Keavney, V. E. Haven, and S. M. Vernon,
Proceedings 2nd Int'l Conf. on Indium Phosphide and
Related Materials, pp. 435-438, 1990.
[6] A. Yamamoto, M. Yamaguchi, and C. Uemura, Appl.
Phy. Lett., 44, p. 611, 1984.
[7] I. Weinberg, to be published in Solar Ceils, 1991.
[8] T. Takamoto, H. Okazaki, H. Takamura, M. Ura, M.
Ohmod, M. Yamaguchi, S. Ikegami, H. Arai, T. Hayashi, A.
Ushirokawa, K. Takahashi, M. Koubata, and A. Ohnishi,
Int'l PVSEC-5, pp. 547-550, 1990.
[9] D. J. Brinker and I. Weinberg, Proceedings 21st IEEE
Photovoltaic Specialists Conf., pp. 1167-1171, 1990.
[10] B. E. Anspaugh and R. G. Downing, Proceedings
17th IEEE Photovoltaic Specialists Conf., pp. 23-30,
1984.
[11] C.J. Keavney, S. M. Vernon, V. E. Haven and S. J.
Wojtczuk, App/. Phys. Lett. 54, p. 1139, 1989.
[12] J. C. C. Fan, C. O. Bozler, and R. W. Mclelland,
Proceedings 15th IEEE Photovoltaic Specialists Conf., p.
666, 1981.
[13] E. Yablanovitch, T. Gmitter, J. P. Harbison, and R. H.
Batt, Appl. Phys. Lett. 51, p. 2222, 1987.
[14] C. Goradia, W. Thesling, !. Weinberg, to be published
inthe Proceedings of the XI Space Photovoltaic Research
and Technology Conference, 1991.
[15] R. M. Kurland and P. Stella, 9th Space Photovoltaic
Research and Technology Conference, pp. 122-137,
1988.
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4. TITLE AND SUBTITLE
Advanced Power Systems for EOS
6. AUTHOR(S)
Sheila G. Bailey, Irving Weinberg, and Dennis J. Flood
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio 44135- 3191
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National Aeronautics and Space Administration
Washington, D.C. 20546- 0001
Technical Memorandum
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11. SUPPLEMENTARY NOTES
Prepared for the 26th Intersociety Energy Conversion Engineering Conference cosponsored by ANS, SAE, ACS,
AIAA, ASME, IEEE, and AIChE, Boston, Massachusetts, August 4-9, 1991. Responsible person, Sheila G. Bailey,
(216) 433-2228.
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13. ABSTRACT (Maximum 200 words)
The Earth Observing System, which is part of the International Mission to Planet Earth, is NASA's main contribution
to the Global Change Research Program. The program opens a new era in international cooperation to study the
Earth's environment. Five large platforms are to be launched into polar orbit, two by NASA, two by the European
Space Agency, and one by the Japanese. In such an orbit the radiation resistance of indium phosphide solar cells
combined with the potential of utilizing 5 micron cell structures yields an increase of 10% in the payload capability. If
further combined with the Advanced Photovoltaic Solar Array, the total additional payload capability approaches 12%.
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Photovoltaics; EOS; Indium phosphide; Solar array
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