N O T I C E

THIS DOCUMENT HAS BEEN REPRODUCED FROM MICROFICHE. ALTHOUGH IT IS RECOGNIZED THAT

CERTAIN PORTIONS ARE ILLEGIBLE, IT IS BEING RELEASED IN THE INTEREST OF MAKING AVAILABLE AS MUCH

INFORMATION AS POSSIBLE

https://ntrs.nasa.gov/search.jsp?R=19800016259 2020-03-31T18:00:01+00:00Z

IM

JPL PUBLICATION 80-31

--`"l

Results of the 1979Balloon Flight solarCalibration ProgramC.H. SeamanR.S. Weiss

NASA/JPLCell

(NASA-CR-163013) RESULTS OF THE 1979NASA/JPL BALLOON FLIGHT SOLAR CELLCALIBRATION PROGRAM (Jet p ropulsion Lab.)

17 p HC A02/MF A01 CSCL 10A

N80--24752

UnclasG:3/44 20943

^ I

'^^sr/Fgb^D

May 1, 1980

National Aeronautics andSpace Administration

Jet Propulsion? LaboratoryCalifornia Institute of TechnologyPasadena, California

JPL PUBLICATION 80-31

il l)",, tt

Results of the 1979Balloon Flight SolarCalibration ProgramC.H. SeamanR.S. Weiss

NASA/J P LCell

May 1, 1980

National Aeronautics andSpace Administration

Jet Propulsion LaboratoryCalifornia Institute of TechnologyPasadena, California

.4

The research described in this publication was carried out by the Jet PropulsionLaboratory, California Institute of Technology, under NASA Contract No. NAS7-100.

s

4

PREFACE

The work described in this report was performed by the Control and EnergyConversion Division of the Jet Propulsion Laboratory. The flight was conduc-ted with the cooperation of the National Scientific Balloon Facility, locatedin Palestine, 'Texas. A summary of the data is presented.

ABSTRACT

r

The 1979 sOneduled solar cell calibration balloon flight was successfullycompleted on Au "at 8, meeting all objectives of the program. thirty-eightmodules were carried to an altitude of about 36 kilometers. These calibratedcells can be used as reference standards in simulator testing of cells andarrays.

iii0

ACKNOWLEDGMENT

The authors wish to extend appreciation for the cooperation and supportprovided 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 organizationsis greatly appreciated.

,!

iv

—

CONTENTS

I. INTRODUCTION -------------------------------------------------- ------ 1

II. PROCEDURE ------------w---------------------------------------- ------ 3III. 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 isthe direct conversion of solar energy through the use of solar cells. As ad-vancing cell technology continues to modify the spectral range of solar cellsto 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 andhence the mass of the array structure.

Since laboratory simulaticr, of extra-atmospheric solar radiation hasnot been accomplished on a prac:icol scale with sufficient fidelity, highaltitude exposure must be taken as the Lest representation of space itself.The error due to residual atmosphere may be computed using

W coFd 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 spectralresponse, RX.

The values of these summations have been computed for transmissivitiescorresponding to an altitude of 36 kilometers (the altitude of this flight)and four solar zenith angles using available data (References 1 and 2). Thespectral extinction optical thickness is plotted in Figure 1. Values ofpercent 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 flightwas taken for solar zenith angles between 25 degrees and 23 degrees. It isseen that the flight altitude used in these calibrations is such that thecomputed error due to residual atmosphere is negligible.

To reach and maintain the required altitude, the calibration programmakes use of balloons provided and launched by the National scientific BalloonFacility, Palestine, Texas.

1

0.

51(

M

4,(

010 3.:

XNNWz 3.t

Qu 2„

aUZ 2.1_U

GZ

1,1

WAVELENGTH, ,am

Figure 1. Extinction Optical Thickness

1,00

3 0,50E

F;

wNZd 0,20aJQ

0.10aNw5JR 0,05

O-- SILICONGaAIAs

L_ I I I I I

0.3 0.4 0.5 0.6 0,7 0.8 0.9 1,0 1,1WAVELENGTH, µm

Figure 2. Cell Spectral Response

2

0,02

I GoA'IAs 0CL1

/,0

35

!q

"^ 30uia

C1

Z4

a 25

`Z.W

SU 20V1

15 ^5 ERRORvs

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 componentsof the il?ght system, the cells are mounted by the participants on JPL-suppliedstandard modules according to directions in Reference 3 1 which details mater-ials, techniques, and workmanship standards for assembly. The JPL standardmodule is a machined copper block 3.7 cm x 4.8 cm x 0.3 cm thick, rimmed by0.3 cm thick fiberglass, painted a high reflectance white, with infulatedsolder 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, forexample, also loads the cell in its short circuit current condition.

The mounted cells are then subjected to preflight measurements in theJPL X25L solar simulator. This measurement when compared to a postflightmeasurement under the same conditions may be used to detect cell damage orinstabilities.

3

^ ^y.+L .wR4z'tz'.f y+w

Pr y or to shipment to the launch facility, the modules are mounted onthe sun tracker bed plate, Figure 4.

Upon arrival at the Palestine Facility, the tracker and module payloadare checked for proper operation, the data acquisition and Pulse Code Modula-tion telemetry systems are calibrated, and mounting of the assembly onto theballoon is then accomplished, figure 5.

At operating altitude the sun tracker bed plate is held pointed at thesun to within ±1 degree, The response of each module, temperatures ofrepresentative modules, sun lock information, and system calibration voltagesare sampled twice each second and telemetered to the ground station where theyare presented in teletype form for real time assessment and are also recordedon magnetic tape for later processing, Float altitude information is obtainedfrom data supplied by the balloon facility.

SECTION III

SYSTEM DESCRIPTION

A solar tracker mounted in a frame on top of the balloon carries themodule payload while the transmitter of the data link is located in the lowerGondola along with batteries for power and ballast for balloon control. Atcompletion of the experiment, the upper payload and lower gondola are returnedby 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 andsun-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

r

Figurt .. 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 consecutivedata points taken around the time of nol.ar noon after indicated temperaturestability.

The flight data were thus reduced and modules with their data and calmbration values were returned to the participants. Thin information in col-lected in 'Fable 1. The placement of modules on the field of the tracker bedfor 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. COMPARISON9SOLARNUMBER NUMBER CODE ADJ. AVERAGE DEVIATION 1 AUm 28 OCG.0 SIMULATOR A FLT

MILLIVOLTS PRE-FLT POS-FLT PRC-FLT FLIS.HTV9. VS.

POS-FLT PAC-FLT(PCRCCNTI (PERCENT)

79-001 JPL 07.77 .09925 86070 86.00 -091 162379 . 126 OCLI 01.49 008063 80.90 81.10 .25 07579 . 113 HUGHC5 56043 .00701 36.00 55.9 8 -009 07679-101 HUGHES 29039 .08709 29075 28690 -2.b6 -1.2314-205 JPL 90. 76 .04927 P ft. 30 88.30 .23 390279-115 HUGHES 56.71 .08353 56.70 5600 000 .1379-132 TRY 13054 000357 73085 73083 -003 -04279-002 JPL 84.30 .08978 03060 83.20 -048 48378-105 HUGHES 02.20 009530 80095 81600 .06 I.5479-134 TRY 75.07 000434 75.10 74680 -040 -00479-107 HU61105 49042 .08959 31050 30070 -2.54 5608915-183 JPL 66.96 .09882 67690 68000 015 -103979-103 HUGHES 24033 000670 26.10 25.70 -1053 -6079

9FS-17A JPL 6004 .00071 60.40 61000 .99 -04379-127 OCLI 82.01 .09412 01075 81630 -.55 040

T8-107 HUGHES 09.50 +10645 60005 80000 -.06 106579-003 JPL 82.45 .07081 81685 81.60 -031 67379-122 OCLI 84.31 .10594 83065 83.33 -.38 .867 3 - I a 2 JPL 67.83 009426 60.65 68065 .29 -101910-.110 HUGHES 95.02 .07924 93095 93075 -.21 101479-133 TRhi 73.5A .24430 73.85 73040 -061 -.36

79 . 109 HUGHES 21.04 .07746 20.90 20.90 .DO 06579-136 TRY 75.22 .00212 75.10 75.20 .13 01519-004 JPL 83.45 .08756 83.05 02.75 -.36 04679-105 HUGHES 28053 .08429 37.70 37.41 -077 -24.3376 . 001 JPL 67.32 .07523 68.40 68.65 037 -1.5879-005 JPL 76639 .07775 7600 75.70 -09 .1279-006 JPL 79.30 +11971 79.55 79.06 -0S2 -.3179-007 JPL 83689 .07629 83055 03.55 .00 04179-000 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 -.3579 . 124A OCLI 18.20 .09198 77.35 77.70 .45 1.1079-1240 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 794126JPL O ^Lf

HU^rHES

8hSUN

(^10^

79-1,q4 Y

Taw79-107HUGHESFILTER

73.103Tb e4r,jJPL

79.103HUGHESFILTER

CF •17AJPL

79.127QCLI

70•1C'7HUGHES

993003JPL

(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

aJPL

FILIER FILTERLT'

(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 ^4L

T3 STD CELL

Tb STD CELL @6H

T.R VOLTAGE P,EF °OO 407

Figure 6. Module Location Chart

A detailed discussion of data reduction and an analysis of system errormay be found in Reference 4.

SECTION V

MONITOR CELLS

Several standard modules have been flown repeatedly over the 16-yearperiod 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 of0.39 percent and a maximum deviation of 0.92 percent from the mean.

In addition, the uniformity of the solar irradiance (i.e. no spuriousreflections, shadowing) over the field of the modules has been demonstratedsince the Location of this module was changed in that field from flight toflight.

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 60.328/3/64 60.43 4/5/74 60.378/8/64 60.17 4/23/74 60.377/28/65 59.90 5/8/74 60.368/9/65 59.90 10/12/74 60.80"/13/65 59.93 10/24/74 60.567/29/65 60.67 6/6/75 60.208/4/66 60.25 6/27/75 60.218/12/66 60.15 6/10/77 60.358/26/66 60.02 8/11/77 60.467/14/67 60.06 7/20/78 60.497/25/67 60.02 7/8/79 60.14

8/4/6; 59.83 i4ean 60.25

8/10/67 60.027/19./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 forperiod 9/5/63 to 8/5/70.

For flights an 4/5/74 through 7/1/75 each data point is an averageof 100 or more flight data points.

For flights ^tarting in September 1975, each data point is an averageof 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, atan altitude of 35.7 kilometers with sun-lock occurring within 2 minutes. Datawas taken for solar zenith angles between 25 degrees and 23 degrees. Theflight was terminated at 1145 hours, CST. Although the parachute did notfully deploy and the tracking beacon failed, the payload was recovered thefollowing 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 forlong periods of time and may be used as standards with confidence.

2. The calculated error due to residual atmosphere on this flight isless 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 spectralresponse, continued calibration of solar cells under AMO conditions is re-quired to assure that solar panel performance with all. its ramifications canbe accurately predicted.

9

{

REFERENCES

1

1. Handbook of Geophysics and Space Environments, AFCRL, S. L. ValleyEd., Chapter 7, 1965.

2. Drummond, A. J. and Thekaekara, M. P., The Fxtraterrestial SolarS 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 Report329 16D0, Jet Propulsion Laboratory, Pasadena, California, November1, 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