Calibration of solar cells to be used as reference standards in simulator testing of cells and arrays was accomplished. Thirty-eight modules were carried to an altitude of about 36 kilometers during the solar cell calibration balloon flight

Seaman, C. H.

Weiss, R. S.

English

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IM
JPL PUBLICATION 80-31
--`"l
Results of the 1979
Balloon Flight solar
Calibration Program
C.H. Seaman
R.S. Weiss
NASA/JPL
Cell
(NASA-CR-163013) RESULTS OF THE 1979
NASA/JPL BALLOON FLIGHT SOLAR CELL
CALIBRATION PROGRAM (Jet p ropulsion Lab.)
17 p HC A02/MF A01	 CSCL 10A
N80--24752
Unclas
G:3/44 20943
^ 	 I
'^^sr/Fgb^D
May 1, 1980
National Aeronautics and
Space Administration
Jet Propulsion? Laboratory
California Institute of Technology
Pasadena, California
JPL PUBLICATION 80-31
il l)",, 	 tt
Results of the 1979
Balloon Flight Solar
Calibration Program
C.H. Seaman
R.S. Weiss
NASA/J P L
Cell
May 1, 1980
National Aeronautics and
Space Administration
Jet Propulsion Laboratory
California Institute of Technology
Pasadena, California
.4
The research described in this publication was carried out by the Jet Propulsion
Laboratory, California Institute of Technology, under NASA Contract No. NAS7-100.
s4
PREFACE
The work described in this report was performed by the Control and Energy
Conversion Division of the Jet Propulsion Laboratory. The flight was conduc-
ted with the cooperation of the National Scientific Balloon Facility, located
in Palestine, 'Texas. A summary of the data is presented.
ABSTRACT
r
	
	 The 1979 sOneduled solar cell calibration balloon flight was successfully
completed on Au "at 8, meeting all objectives of the program. thirty-eight
modules were carried to an altitude of about 36 kilometers. These calibrated
cells can be used as reference standards in simulator testing of cells and
arrays.
iii0
ACKNOWLEDGMENT
The authors wish to extend appreciation for the cooperation and support
provided by the entire staff of the National Scientific Balloon Facility.
Gratitude is also extended to assisting JPL personnel, especially B. E.
Anspaugh, R. G. Downing and L. B. Sidwell.
The cooperation and patience extended by all participating organizations
is greatly appreciated.
,!
iv
—CONTENTS
I. INTRODUCTION -------------------------------------------------- ------	 1
II. PROCEDURE ------------w---------------------------------------- ------	 3
III. SYSTEM DESCRIPTION -------------------------------------------- ------	 4
IV. DATA REDUCTION ------------------------------------------ ------------	 4
V. MONITOR CELLS ------------------------------------------------- ------	 7
VI. FLIGHT PERFORMANCE -------------------------------------------- ------	 8
VII. CONCLUSIONS --------------------------------------------------- ------	 9
REFERENCES ----------------------------------------------------------- ,------ 10
Tables
1. Cell Calibration Data ----•------------------------------------- ------	 6
2. Repeatability of Standard Module BFS-17A ---------------------- ------	 8
Figures
1. Extinction Optical Thickness ---------------------------------------- 2
2. Cell Spectral Re gponse ---------------------------------------------- 2
3. Error vs. Zenith Angle ---------------------------------------------- 3
4. Solar Module Payload ------------------------------------------------ 5
5. Balloon Mount ------------------------------------------------------- 5
6. Module Location Chart ----------------------------------------------- 7
v
n	 -	 s
SECTION I
INTRODUCTION
The primary source of electrical power for unmanned space vehicles is
the direct conversion of solar energy through the use of solar cells. As ad-
vancing cell technology continues to modify the spectral range of solar cells
to utilize more of the sun's spectrum, designers of solar arrays must have in-
formation detailing the impact of these modifications on cell conversion effi-
ciency to be able to confidently minimize the active cell area required and
hence the mass of the array structure.
Since laboratory simulaticr, of extra-atmospheric solar radiation has
not been accomplished on a prac:icol scale with sufficient fidelity, high
altitude exposure must be taken as the Lest representation of space itself.
The error due to residual atmosphere may be computed using
W	 co
Fd
 R^ T^ E^ A X
	(1-e)	 R^ E^ AX	 (1)
where	 RX = cell spectral response
EX
 = extra-atmospheric solar spectral irradiance (AMO)
T^ = sky path -pectral transmissivity
e	 fractional departure from true AMO response (error)
TX can be computed using
T X
 = Exp - (r X sec Z)
where	 TX = spectral extinction optical thickness (Reference 1)
Z	 = solar zenith angle
The actual limits in Eq. (1) will be determined by the cell spectral
response, RX.
The values of these summations have been computed for transmissivities
corresponding to an altitude of 36 kilometers (the altitude of this flight)
and four solar zenith angles using available data (References 1 and 2). The
spectral extinction optical thickness is plotted in Figure 1. Values of
percent error (100c) vs. solar zenith angle for two representative cell spec-
tral responses (given in Figure 2) are shown in Figure 3. Data on this flight
was taken for solar zenith angles between 25 degrees and 23 degrees. It is
seen that the flight altitude used in these calibrations is such that the
computed error due to residual atmosphere is negligible.
To reach and maintain the required altitude, the calibration program
makes use of balloons provided and launched by the National scientific Balloon
Facility, Palestine, Texas.
1
0.
51(
M
4,(
010 3.:
X
NNW
z 3.t
Q
u
 2„
aU
Z 2.1
_U
G
Z
1,1
WAVELENGTH, ,am
Figure 1. Extinction Optical Thickness
1,00
3 0,50
E
F;
wNZ
d 0,20
a
JQ
0.10
aN
w
5J
R 0,05
O-- SILICON
GaAIAs
L_	 I	 I	 I	 I	 I0.3	 0.4	 0.5	 0.6	 0,7	 0.8	 0.9	 1,0	 1,1
WAVELENGTH, µm
Figure 2. Cell Spectral Response
2
0,02
I	 GoA'IAs	 0CL1
/,0
35
!q
"^ 30
ui
a
C1
Z4
a 25
`Z.W
S
U 20
V1
15	 ^5 ERROR
vs
SOLAF ZENITH ANGLE AT
ALTITUDE , 36 KILOMETERS
5l	 tt	 ^
-0,20	
-0,15
	
-0.10	 -0,05
ERROR, °b
Figure 3. Error vs. Zenith Angle
SECTION II
PROCEDURE
To insure electrical and mechanical compatibility with other components
of the il?ght system, the cells are mounted by the participants on JPL-supplied
standard modules according to directions in Reference 3 1 which details mater-
ials, techniques, and workmanship standards for assembly. The JPL standard
module is a machined copper block 3.7 cm x 4.8 cm x 0.3 cm thick, rimmed by
0.3 cm thick fiberglass, painted a high reflectance white, with infulated
solder posts and is permanently provided with a precision (0.1 percent, 20
PPm/ OC) load resistor appropriate for scaling the cell output to the tele-
metry constraints. This load resistor, 0.5 ohm for a 2 cm x 2 cm cell, for
example, also loads the cell in its short circuit current condition.
The mounted cells are then subjected to preflight measurements in the
JPL X25L solar simulator. This measurement when compared to a postflight
measurement under the same conditions may be used to detect cell damage or
instabilities.
3
^	 ^y.+L .wR4z'tz'.f y+w
Pr y or to shipment to the launch facility, the modules are mounted on
the sun tracker bed plate, Figure 4.
Upon arrival at the Palestine Facility, the tracker and module payload
are checked for proper operation, the data acquisition and Pulse Code Modula-
tion telemetry systems are calibrated, and mounting of the assembly onto the
balloon is then accomplished, figure 5.
At operating altitude the sun tracker bed plate is held pointed at the
sun to within ±1 degree, The response of each module, temperatures of
representative modules, sun lock information, and system calibration voltages
are sampled twice each second and telemetered to the ground station where they
are presented in teletype form for real time assessment and are also recorded
on magnetic tape for later processing, Float altitude information is obtained
from data supplied by the balloon facility.
SECTION III
SYSTEM DESCRIPTION
A solar tracker mounted in a frame on top of the balloon carries the
module payload while the transmitter of the data link is located in the lower
Gondola along with batteries for power and ballast for balloon control. At
completion of the experiment, the upper payload and lower gondola are returned
by parachutes and recovered, A more complete description of the system in-
cluding the sun tracker can be found in Reference 4.
SECTION IV
DATA REDUCTION
The raw data as taken from the tape is corrected for temperature and
sun-Earth distance according to the formula (Reference 5)
V28,1 = VT R (R2 ) - a (T-28)
where VTrR = measured module output voltage at temperature T and distance R
R = sun-Earth distance in astronomical units
a = module output voltage temperature coefficient
T = module temperature in oC
The value of a is supplied by the participant.
4
h
rFigurt .. Typo. "I Soiar Module Payloci, GF /c/,v AR qQ
Figure 5. Balloon Mount
5
The calibration value is taken to be the average of 200 consecutive
data points taken around the time of nol.ar noon after indicated temperature
stability.
The flight data were thus reduced and modules with their data and calm
bration values were returned to the participants. Thin information in col-
lected in 'Fable 1. The placement of modules on the field of the tracker bed
for the 1919 flight is shown in Figure 6.
Table 1. Cell Calibration Data
BALLOON FLIGHT 79 - 1 DATE 8-0 . 79 ALTITUCE 1179000 AVa1.01392
CHANNEL MOCULC ORGANIZATION TEMP. INTENSITY 	 STANDARD	 AM09SOLAR SIM.	 COMPARISON9SOLAR
NUMBER NUMBER	 CODE	 ADJ. AVERAGE	 DEVIATION	 1 AUm 28 OCG.0	 SIMULATOR A FLT
	
MILLIVOLTS	 PRE-FLT POS-FLT	 PRC-FLT	 FLIS.HT
V9.	 VS.
POS-FLT	 PAC-FLT
(PCRCCNTI (PERCENT)
79-001 JPL 07.77 .09925 86070 86.00 -091 1623
79 . 126 OCLI 01.49 008063 80.90 81.10 .25 075
79 . 113 HUGHC5 56043 .00701 36.00 55.9 8 -009 076
79-101 HUGHES 29039 .08709 29075 28690 -2.b6 -1.23
14-205 JPL 90. 76 .04927 P ft. 30 88.30 .23 3902
79-115 HUGHES 56.71 .08353 56.70 5600 000 .13
79-132 TRY 13054 000357 73085 73083 -003 -042
79-002 JPL 84.30 .08978 03060 83.20 -048 483
78-105 HUGHES 02.20 009530 80095 81600 .06 I.54
79-134 TRY 75.07 000434 75.10 74680 -040 -004
79-107 HU61105 49042 .08959 31050 30070 -2.54 56089
15-183 JPL 66.96 .09882 67690 68000 015 -1039
79-103 HUGHES 24033 000670 26.10 25.70 -1053 -6079
9FS-17A JPL 6004 .00071 60.40 61000 .99 -043
79-127 OCLI 82.01 .09412 01075 81630 -.55 040
T8-107 HUGHES 09.50 +10645 60005 80000 -.06 1065
79-003 JPL 82.45 .07081 81685 81.60 -031 673
79-122 OCLI 84.31 .10594 83065 83.33 -.38 .86
7 3 - I a 2 JPL 67.83 009426 60.65 68065 .29 -1019
10-.110 HUGHES 95.02 .07924 93095 93075 -.21 1014
79-133 TRhi 73.5A .24430 73.85 73040 -061 -.36
79 . 109 HUGHES 21.04 .07746 20.90 20.90 .DO 065
79-136 TRY 75.22 .00212 75.10 75.20 .13 015
19-004 JPL 83.45 .08756 83.05 02.75 -.36 046
79-105 HUGHES 28053 .08429 37.70 37.41 -077 -24.33
76 . 001 JPL 67.32 .07523 68.40 68.65 037 -1.58
79-005 JPL 76639 .07775 7600 75.70 -09 .12
79-006 JPL 79.30 +11971 79.55 79.06 -0S2 -.31
79-007 JPL 83689 .07629 83055 03.55 .00 041
79-000 JPL 81.25 .09327 80090 80070 -.25 .44
79-009 JPL 79.33 .07772 78095 78.90 -.06 048
79-111 HUGHES 63.55 .10558 63.50 62.45 -1.65 -08
79-010 JPL 55.14 .07319 55.35 55.27 -.14 -.35
79 . 124A OCLI 18.20 .09198 77.35 77.70 .45 1.10
79-1240 OCLI 77.93 .08796 76.90 77.37 .61 1.34
79-014 JPL 59.23 .09271 59060 59.20 -.67 -065
79-012 JPL 7248 .07837 72.95 72.65 -01 -.10
11-007 JAL 61.01 .09520 60.90 60.96 .06 .18
AVERAGE TEMPERATUPE (OEG.C) At FLOAT ALIITUDC a 45.95
6
i^
>ESHUSHES JPL HUAHES TROY JPL
[UG"
?4
FILTER OIACk
79.001
	 794126
JPL	 O ^Lf
HU^rHES
8h
SUN
(^10^
79-1,q4 Y
Taw
79-107
HUGHES
FILTER
73.103
Tb e4r,j
JPL
79.103
HUGHES
FILTER
CF •17A
JPL
79.127
QCLI
70•1C'7
HUGHES
993003
JPL
(10+ 01) (1 (13 A^ Ii6 Ib 19
P9.122 73.12 70.110 79.133 79-104 79.13b 74.00b 0(il 74•
I nU^iHEs TRW JPL HUGHES
a
JPL
FILIER FILTER
LT'
(232 b2ti
79.00b
JPL
^^20 F
I
74.012
	 7700
JPL	 JPL
LP_2LQI^ ^
79-
[FIUL
79.12b
PLS JPL	 CLI
 .
1--i—
	
= 30	 t 31 	 (39^	 A11a 0l/ Lbj
) INDICATES CHANNEL NUMBER
H - TI STD CELL 43
T2 TRACKER ELEC ^4
L
T3 STD CELL
Tb STD CELL @6
H
T.R VOLTAGE P,EF °OO 407
Figure 6. Module Location Chart
A detailed discussion of data reduction and an analysis of system error
may be found in Reference 4.
SECTION V
MONITOR CELLS
Several standard modules have been flown repeatedly over the 16-year
period of calibration flights. The record of the one with the longest his-
tory, BFS-17A, appears in Table 2. This data shows a standard deviation of
0.39 percent and a maximum deviation of 0.92 percent from the mean.
In addition, the uniformity of the solar irradiance (i.e. no spurious
reflections, shadowing) over the field of the modules has been demonstrated
since the Location of this module was changed in that field from flight to
flight.
7
,!!W
Table 2. Repeatability of Standard Module BFS-17A
(31 Flights over a 17-Year Period)
Flight date
	
Output, mV
	
Flight data
	
Output, mV
9/5/63 60.07 8/5/70 60.32
8/3/64 60.43 4/5/74 60.37
8/8/64 60.17 4/23/74 60.37
7/28/65 59.90 5/8/74 60.36
8/9/65 59.90 10/12/74 60.80
"/13/65 59.93 10/24/74 60.56
7/29/65 60.67 6/6/75 60.20
8/4/66 60.25 6/27/75 60.21
8/12/66 60.15 6/10/77 60.35
8/26/66 60.02 8/11/77 60.46
7/14/67 60.06 7/20/78 60.49
7/25/67 60.02 7/8/79 60.14
8/4/6; 59.83 i4ean 60.25
8/10/67 60.02
7/19./68 60.31 Std. Deviation 0.24
7/29/68 60.20
8/26/69 60.37 Maximum deviation 0.55
9/8/69 60.17
7128/70 60.42
Each data point is an average of 20 to 30 points per flight for
period 9/5/63 to 8/5/70.
For flights an 4/5/74 through 7/1/75 each data point is an average
of 100 or more flight data points.
For flights ^tarting in September 1975, each data point is an average
of 200 data ;tints.
SECTION VI
FLIGHT PERFORMANCE
The launch at 0700 hours, CST, on August 8 was accomplished without inci-
dent as was the float phase. The tracker was energized at 1045 hours, CST, at
an altitude of 35.7 kilometers with sun-lock occurring within 2 minutes. Data
was taken for solar zenith angles between 25 degrees and 23 degrees. The
flight was terminated at 1145 hours, CST. Although the parachute did not
fully deploy and the tracking beacon failed, the payload was recovered the
following afternoon undamaged.
8
8
SECTION VII
CONCLUSIONS
1. As emphasized by the history of repeatability of cell BF'S-17A, viz
±1 percent (see Table 2), silicon cells when properly cared for are stable for
long periods of time and may be used as standards with confidence.
2. The calculated error due to residual atmosphere on this flight is
less than 0.10 percent for silicon cells and less than 0.13 percent for gal-
lium arsenide cells.
3. As advancing technology continues to favorably modify cell spectral
response, continued calibration of solar cells under AMO conditions is re-
quired to assure that solar panel performance with all. its ramifications can
be accurately predicted.
9
{REFERENCES
	
1
1. Handbook of Geophysics and Space Environments, AFCRL, S. L. Valley
Ed., Chapter 7, 1965.
2. Drummond, A. J. and Thekaekara, M. P., The Fxtraterrestial Solar
S vt um, Inst. of Env. Sciences, Mount Prospect, Illinois, 1973.
3. Greenwood, R. F., "Solar Cell Modules Balloon Flight Standard,
Fabrication of," Procedure No. EP504443, Revision C, JPI. Pasadena,
California, June 11, 1974 OPL Internal Document).
4. Yasui, R. K. and Greenwood, R. F., Results of the 1973 NASA/JPI.
Balloon Flight Solar Cell Calibration p ogns , Technical Report
329 16D0, Jet Propulsion Laboratory, Pasadena, California, November
1, 1975.
9.	 Solar Cell Array Design Handbook, 3.6°2, .IPL SP 43-38, Vol. 1, 1976.
10	 NASA M VWH t A Card
1


Results of the 1979 NASA/JPL balloon flight solar cell calibration program

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