L 

JPL PUBLICATION 86-42 

Results of the 1986 NASAlJPL 

Balloon Flight Solar Cell 

Calibration Program i 

B.E. Anspaugh k 

R.S. Weiss 

iNASA-CR- 1799 16) RESCL7S CF ‘IEE 1986 N87- 120 48 

NASA/JPL EALLCCS FLIGHT SCLAh CALIBBATXON 

PhOGErAM (Jet Prc&ulsica Lab.) 41 p CSCL 108 

Unclas 

G3/44 44933 

November 1, 1986 

NASA’ 

National Aeronautics and 

Space Administration 

Jet Propulsion Laboratory 

California Institute of Technology 

Pasadena, California 

\ 
JPL PUBLICATION 86-42 

Results of the 1986 NASAIJPL 

Balloon Flight Solar Cell 

Calibration Program 

B.E. Anspaugh 

R.S. Weiss 

November 1, 1986 

National Aeronautics and 

Space Administration 

Jet Propulsion Laboratory 

California Institute of Technology 

Pasadena, California 
The research described in this publication was carried out by the Jet Propulsion 

Laboratory, California Institute of Technology, under a contract with the National 

Aeronautics and Space Administration. 

Reference herein to any specific commercial product, process, or service by trade 

name, trademark, manufacturer, or otherwise, does not constitute or imply its 

endorsement by the United States Government or the Jet Propulsion Laboratory, 

California Institute of Technology. 
The 1986 solar cell calibration balloon flight was successfully 

completed on July 15, 1986, meeting all objectives of the program. Thirty 

modules were carried to an altitude of 118,000 ft (36.0 km). 

calibrated cells can now be used as reference statdards in simulator testing 

of cells and arrays. 

The 

iii 
Tb authors wish to extend 

- 

-reciation for the cooperation rmd support 

provided by the entire staff of the National Scientific Balloon pacility 

located in Palestine, Texas. Gratitude is also extended to assisting JPL 

personnel, especially Ed Mann for performing all the pre-flight and postflight solar cell measurements. The cooperation and patience extended by 

all particfpating organizations are greatly appreciated. 

iv 
CONTENTS 

1 . I~DUCTIONAND~IEW .................... 1 

2 . pRE-FL1mpRocEDUREs ...................... 3 

2.1 MOWTLEFABRICATION ..................... 3 

2.2 CELLMEASURFMENTS ..................... 3 

2.3 TEWERAW COEFFICI~ AND LEAST SQUARE FITS ....... 4 

2.4 PANELASSEMBLYANDCHECKMIT ................ 4 

2.5 PREXA~pEzocEMiREsATPALESTINE ............. 5 

3 . BALLooNsysTpM ......................... 10 

3.1 BALU30NDESCRIPTION .................... 10 

I 

3.2 TOP PAYLOAD ........................ 11 

3.3 BOTTOMPAYLOAD ....................... 13 

4 . FLIGHTSEQUENCX ......................... 18 

4.1 PRE-LAU" PREPARATIONS .................. 18 

4.2 FLIGHT ........................... 22 

4.3 FLIGHT TERMINATION ..................... 23 

5 . DATAANALYSIS .......................... 26 

5.1 coMp[ITwANALYSIS ..................... 26 

5.2 CALIBRATION RESULTS .................... 29 

I 

5.3 UPDATE OF BE'S-17A DATA ................... 29 

6 . CONCLUSIONS ........................... 33 

-E .............................. 34 

V 
Figures 

1 . Photograph of the 1986 Balloon Flight Solar Panel ..... 6 

2 . 1986MoctulemtianChart ................. 7 

3 . Aluminum Hoop Assembly with Tracker Mounted ........ 9 

4 . Block Diagram of Balloon Telemetry System ......... 14 

5 . Inflation of the Main Balloon ............... 20 

6 . BalloonLaunch ....................... 21 

7 . 1986 Balloon Flight Profile ................ 25 

Tables 

1 . Calibration Results: Balloon Flight No . 1425P. 

July15. 1986 ....................... 31 

2 . Repeatability of Standard Solar Cell BE'S-17A ........ 32 
SECTION 1 

INlRODUCTION AND OVERVIEW 

The primary source of electrical power for unmanned space vehicles is 

the direct conversion of solar energy through the use of solar cells. As 

advancing cell technology contimes to modify the spectral response of solar 

cells to utilize more of the sun's spectrum, designers of solar cells and 

arrays must have the capability of measuring these cells in a light beam 

that is a close match to the solar spectrum. The solar spectrum has been 

matched very closely by laboratory solar simulators. But the design of 

solar cells and the sizing of solar arrays requires such highly accurate 

measurements that the intensity of these simulators must be set very 

accurately. A small error in setting the simulator intensity can 

conceivably cause a disastrous mis-sizing of a solar panel, causirg either 

a premature shortfall in power or the launch of an oversized, overweight 

solar panel. 

The JPL solar cell calibration program was conceived to produce 

reference standards for the purpose of properly setting solar simulator 

intensities. The concept was to fly solar cells on a high-altitude balloon, 

measure their output at altitudes near 120,000 ft, recover the cells, and 

use them as reference standards. The reference 

call is placed in the simulator beam and the beam intensity is adjusted 

until the reference cell reads the same as it read on the balloon. As long 

as the reference cell has the same spectral response as the cells or panels 

to be measured, this is a very accurate method of setting the intensity. 

The procedure is simple. 

1 
But as solar cell technology changes, the spectral response of the solar 

cells charges also, and reference standards must be continually renewed. 

Until the summer of 1985, there had always been a question as to how 

much the atmosphere above the balloon modified the solar spectrum. If the 

modification was significant, the reference cells might not have the 

required accuracy. Solar cells made in recent years have increasingly 

higher blue responses and if the atmosphere has any effect at all, it would 

be expectedto modify the calibration of such blue cells much more than that 

of their predecessors. 

In late 1984, a collection of solar cells representing a wide cross 

section of solar cell technology was flown on the Discovery shuttle as a 

part of the Solar Cell Calibration Facility (SCCF) experiment. They were 

calibrated as reference cells on this flight using procedures similar to 

those used on the balloon flights. The same cells were then flown on the 

1985 balloon flight and remeasured. Since the two sets of measurements gave 

very nearly identical results (Ftef. l), the reference standards from balloon 

flights may continue to be used with high confidence. 
SECTION 2 

PFtE-FLI~ mzocEDuREs 

2.1 MODULE FABRICATION 

The cells were mounted by the participants onJPL-suppliedstandard 

modules accordirq to standard procedures developed for the construction of 

reference cells. The JPL standard module is a machined copper block, rimmed 

by a fiberglass circuit board with insulated solder posts. This assembly 

was painted a high reflectance white and was permanently provided with a 

load resistor. This resistor performs two tasks. First, it lcads the cells 

near short circuit current which is the cell parameter that varies in direct 

proportion to light intensity. Second, it scales the cell outputs to read 

near 100 mV during the flight, matching a constraint imposed by the 

telemetry electronics. Load resistance values are 0.5 ohm for a 2 x 2 cm 

cell, 0.25 ohm for a 2 x 4 cm cell, etc. The load resistors are precision 

(0.1 percent, 20 ppm/%) and have a resistance stability equal to or better 

than 9.002% Over a three year period. 

The 1986 flight incorporated 30 solar cell modules from 8 different 

vendors. Four GaAs cells and 

three Si cells which had been flown on earlier flights were reflown. JPL 

Module BFS-17A was included again, as it has been on every balloon flight 

since 1963. Its output continues to update our historical record of the 

calibration procedure repeatability. 

The cells included both Si and GaAs types. 

2.2 CELL MEASAfter the cells were mounted onto the copper blocks, the electrical 

Output of each cell module was measured under illumination by the JPL X25 

3 
Mark I1 solar simulator. For these measurements, the simulator intensity 

was set using only one reference cell--no attempt was made to match the 

spectral response of the reference standard to the individual cell modules. 

The absolute accuracy of these measurements is therefore unknown, but the 

measurements do allow checking of the modules for any unacceptable assembly 

losses or instabilities. After the balloon flight, the cells were measured 

in exactly the same way to check for any cell damage or instabilities that 

may have occurred as a result of the flight. 

2.3 TEMPERATURE COEFFICIENTS AND LEAST SQUARE FITS 

The temperature coefficients of the mounted cells were also measured 

before the flight. The modules were mounted on a temperature controlled 

block in a vacuum chamber. Outputs were measured at 0, 20, 40, 60, and 80 

deg C under illumination with the X25. The temperature coefficients of the 

cell modules were computed from these data by fitting the output vs 

temperature relationship with a linear least squares fit. 

2.4 PANEL ASSEMBLY AND CHECKOUT 

After the electrical measurements were completed, the modules were 

mounted on the solar panel and connected electrically. Figure 1 is a 

photograph of the modules after completion of these steps ard Figure 2 is a 

diagram which identifies the modules in the photograph by their serial 

numbers. After completion of the panel assembly, the panel and tracker 

together were given complete functional tests in terrestrial sunlight. The 

assembled tracker and panel were placed in sunlight on a clear, bright day, 

and checked for the tracker's ability to acquire and track the sun while 

each cell module was checked for electrical output. When these tests were 

4 
completed satisfactorily, the assembly was shipped to the National 

Scientific Balloon E'acility (NSBF) in Palestine, Texas, for flight. 

2.5 PRE-LAU" F'FUX- AT PALESTINE 

The National Scientific Balloon Facility was established in 1963 at 

Palestine, Texas. This location was chosen because it has favorable weather 

conditions for balloon launching and a large number of clear days with light 

surface winds. The high-altitude winds in this part of the country take the 

balloons over sparsely populated areas so the descending payloads are 

unlikely to hit anyone. The JPL calibration flights have flown from the 

Palestine facility since 1973. The flights are scheduled to fly in July 

since the sun is high in the sky at that time of year, and the sunlight 

passes through a minimum depth of atmosphere before reaching the solar 

mdules . 

Upon arrival at Palestine, the tracker and module payload were again 

checked for proper operation. The data encoder was connected to the module 

payload and an end-to-end check was performed on the payload, telemetry, 

receiving and decoding systems. A self-contained voltage reference box was 

mounted to the tracker assembly and its four output voltage levels wired 

into the telemetry stream along with the module outputs. The analog to 

digital converter system was calibrated by recording these four voltage 

levels as they were irput to the system and as they were converted, decoded, 

and sent through the system as digital output values. The thermistor 

channels were calibrated by replacing one of the thermistors with a known 

resistor while the entire system was operating. The process was repeated 

using a set of nine carefully calibrated resistors which correspond to 

specific resistances on the resistance vs temperature calibration curve for 

5 
.. 

6 
(0 

El 

0 

7 
that thermistor. As for the reference voltage procedure, input vs output 

values were recorded. Then the procedure was repeated to calibrate all the 

thermistor channels. 

After all the checkouts d calibrations were performed, the tracker 

was mounted onto the aluminum tubular hoop assembly which will ride on the 

top portion (or apex) of the balloon. Figure 3 is a photograph of the 

tracker and solar panel after mounting onto the hoop. 

a 
1. 

1 Figure 3. Aluminum Hoop Assembly with Tracker Maunted 

9 
_'I' SETION 3 

BALIXKIN SYSTEM 

The main components of the balloon flight system were (1) The apexmounted hoop assembly which contains the experimental package, the data 

encoder, the recovery system and the camera package, (2) the balloon, and 

(3) the lower pyload which contained the telemetry and power systems. 

3.1 BAILDON DESCRIPTION 

The balloon employed for the JPL solar cell calibration high-altitude 

flights had a volume of 2.261 million ft3 (64,000 m3) and was made from 0.7 

mil (18 micron) thick stratofilm, a polyethylene film designed for balloon 

use. The balloon was designed to lift itself alay with the bottom ard top 

payloads to a \float altitude of 120,000 ft (36 km). At float altitude, the 

balloon will have a diameter of roughly 160 ft (50 m). A multiconductor 

cable to electrically connect the top and bottom payloads was hilt into the 

balloon duritq its manufacture. The balloon was built with an internal rip 

line designed to rip a hole in the side of the balloon for termination of 

the flight. A special structure was built into the top of the balloon for 

attaching the top payload. Two poppet valves incorporated into this 

mounting structure will open and release helium from the balloon at the end 

of the flight. The pappet valves act as a back-up to the rip line. 

Tryim to inflate and launch a bellom with a sizeable weight attached 

to its tap is like trying to balance a 30 ft broom harrdle on the end of your 

firger when the broom handle has a 50 lb weight on top. A tow balloon, tied 

to the top payload,was used during the inflation and launch phases to add 

stability and keep the top payload on top. This smaller balloon, about 

10 
31,000 ft3, is desi- to lift atmt 160 lb (73 kg). 

cut loose from tk top paylaad after launch when it reached an altitude of 

The tow balloon was 

about 2,000 ft(600 m). 

3.2 TOP PAYUXLI 

The top payload consists of the tracker, solar panel, voltage reference 

box, multiplexer, data encoder, single frame muvie camera, clock, descent 

parachute, and tracking beacon. All these items were mounted to the 

aluminum hoop assembly shown in Figure 3. The hoop assembly, with 

appropriately placed Styrofoam crush pds, served the following functions: 

Permitted the top-mounted payload to "float" on top of the 

balloon and minimized billowing of balloon material 

arcNlyi tk tap payload. 

Served as the mounting surface to the balloon's top end 

fittiq. 

Provided a convenient point for attaching the tow balloon and 

the descent parachute. 

Acted as a shock damper to protect and minimize damage to the 

top payload at taucMown. 

The complete apex-mounted hoap assembly, as flown, weighea approximately 

81 (37 kq) and descended as a unit by parachute at flight termination. 

The sun tracker, shown in Figure 3, is capable of orienting the solar 

panel toward the sun, compensating for the motion of the balloon by using 

two-axis tracking in both azimuth and elevation. The tracker has the 

capability to maintain its lock onto the sun to within 21 deg. To verify 

that the tracker was operating properly, the output of an on-sun indicator 

was constantly monitored during flight by feeding its output to the 

multiplexer and enteriq its signal into the telemetry stream. The cmindicator consists of a small, circular solar cell mounted at the bottom of 

11 
a collimator tube, 1 in. (2.54 cm) diameter, 4.5 in. (11.4 cm) long. It was 

attached to the solar parael so that it pointed at the sun when the panel was 

perpendicular to the sun. The autplt of the on- indicator falls off very 

rapidly as the collimator tube points away from the sun and provides a very 

sensitive irdication of proper tracker operation. 

A reflectlon shield was attached to the panel to prevent any stray 

reflected light from reachirg any of the modules. This shield was made of 

sheet aluminum and attached to three edges of the solar panel. It is the Ushaped, black object shown on the -1 in Figure 3. 

The solar cell modules were mounted onto the sun tracker platform with 

an interface of Dow Corning No. 340 silicone heat sink compourd ad held in 

place with four screws. The heat sink compound was used to achieve a highly 

conctucting thermal contact between the modules and the panel and to smooth 

aut the temperature distribution Over the solar panel as much as possible. 

The solar panel temperature was monitored using thermistors. Some of 

the solar cell modules were constructed with calibrated precision 

thermistors embedded in the copper substrate directly beneath the solar 

cell. Four of these modules were mounted on the solar panel at strategic 

locations so tbir temperature readings gave an accurate representation of 

panel temperature. Placement of these modules on the panel is shown in 

Figure 2. 

An ultra-wide angle, single frame movie camera mounted at the perimeter 

of the aluminum hoop provided visual documentation of tracker operation. A 

battery-powered timer activated the shutter at ten seccd intervals, so that 

50 ft of 8 mm movie film is sufficient to record the entire flight from 

launch to landing. A wind-up clock was placed in the camera's field-of-view 

12 L 
for correlation of tracker operation to the telemetered data. The pictures 

provide a complete record of ascent, tracker operation at float altitude, 

descent, toucMown, am3 post-touchdown events. 

3.3 BOTMM PAYUMD 

The bottom payload was entirely furnished by the NSBE'. It consists of 

a battery power supply, a ballast module for balloon control, and an 

electronics module known as the Ccmsolidated Instrument Pacwe (CIP). 

Power for operating most of the electrical and electronic equipment on 

the balloon was supplied by a high capacity complement of lithium batteries. 

This supply, furnishing 28 VDC regulated power and 36 VDC unregulated power, 

powered the sun tracker and all ths instruments in the CIP. Several other 

small battery sources were used at various locations on the balloon for 

instruments that require small amounts of power. For example, the trackiw 

beacons, the voltage reference barc, and the camera timer all had individual 

battery power supplies. All batteries were sized to supply power for at 

least twice the expected duration of a normal flight. 

High-altitude balloons tend to slowly lose helium during the course of 

the flight. As a consequence, a helium balloon will tend to reach float 

altitude and then begin a slow descent. To counteract this tendency, a 

ballast system was included as part of the bottom payload. It contained 

approximately 125 pounds (57 leg) of ballast in the form of very fine steel 

shot. The shot may be released in any desired amount by radio command. By 

proper use of this system float altitude may be maintained to within +2,000 

ft (+600 m). 

The telemetry system was contained in the Consolidated Instrument 

Package. A block diagram of the telemetry system is shown in Figure 4. The 

13 
A I RBORNE EQUIPMENT 

REFERENCE 

VOLTAGES 

SOLAR CELL PCM DATA SUBCARRIER 

VOLTAGES ENCODER e OSC I LLATOR I 

1 

SOLAR CELL 

TEMPERATURES c TRANSMITTER 

GROUND EQUIPMENT ’r b 

TELEMETRY - FWFM PCM 

RECEIVER DISCRIMINATORS DECODER 

COMPUTER 

SYSTEM 4 

a PRINTER 

INCREMENTAL 

MAGNETIC TAPE 

DATA REDUCTION AT JPL 

(UNIVAC 1100 COMPUTER) 

Figure 4. Block Diagram of Balloon Telemtry System 

14 
system sent all data transmissions concerning the flight over a common RF 

carrier. The CIP also contained a command system for serding commands to 

the balloon for controlling scientific payloads or for controlling the 

housekeeping functions on the balloon. Specifically, the CIP contained the 

following equipment: 

Pulse-code-modulated (PCM) data encoder. 

Programable Rsad Only Memory (PROM). 

Four Multiplexers (100 mV full scale). 

MKS pressure transducers. 

Qnega receiver. 

Subcarrier oscillators (as required) 

L-band EM transmitter. 

High-frequency tracking beacon transmitter. 

Transponder for air traffic control tracking. 

(10) FCM camnand receiver-decoder. 

The PCM data encoder amplified the analog signals from the solar cells, 

thermistors, on-sun indicator, and reference voltages, then performed an 

analog-to-digital conversion. It had a programmable control unit which was 

used to set bit rate, bits per word, parity, analog-to-digital conversion, 

and format. Four 32-channel 

multiplexers allow sampling of up to 128 data channels and amplify the low 

level signals from the experimental packqe. The amplifier was designed to 

process voltage signals at input levels up to 100 mV. The multiplexer 

stepped through the various channels at a rate of two scans per second, 

i.e., every data channel is read twice each second. 

A PROM was used to provide format control. 

The altitude of the balloon was measured with a capacitance-type 

electronic transducer which read pressure within the range of 1020 to 0.4 

15 
millibars (102,000 to 40 Newtons/m2) with an accuracy of 0.05%. The 

transducer produced a DC level which was encoded as PCM data and decoded at 

the receiving station into pressure, and then the altitude was calculated 

from the pressure reading. 

The Omega navigation system was used for flight tracKlng. An on-board 

receiver was used to receive the Omega signal for retransmission to the 

processor in the ground station. This system can provide position data to 

an uncertainty of less than 2 mi (3.2 km). The Omega signal was also 

multiplexed into the telemetry stream and updated every eight seconds. 

As previously mentioned, all the telemetry data was sent to the ground 

in the form of pulse code modulation. A UHF L-band transmitter in the CIP 

was used to generate the RF carrier. The L-band carrier was modulated by 

the pulse code and sent to the receiving station at Palestine. 

Several tracking beacons were used on the balloon. A low frequency 

transmitting beacon, attached to the lower payload, was used by the 

automatic direction finding (ADl?) system in the recovery airplane to track 

and locate the balloon during flight and recovery. When available, a high 

frequency transmitting beacon is attached to the top payload to help find 

the payload during recovery. An aircraft type transponder was flown so that 

Air Traffic Control could read the balloon's location on their radar systems 

throughout the flight. 

The purpose of the PCM commana system is to send commands to the 

balloon, i.e. to turn the tracker on or off, terminate the flight, etc. It 

was designed to reject false commardS and was highly reliable in operation. 

The data was encoded on a frequency-shift-keyed audio carrier. This signal 

was then decoded into data and timing control. Each command consisted of a 

16 
I 

double transmission of the data word. Both words must be decoded and pass a 

bit-by-bit comparison before a command is executed. Commands may be sent to 

the balldon from either the ground station at Palestine or from one of the 

recovery airplanes. 

The top end of a 8.5 m diameter parachute was fastened to the bottom of 

the balloon. The lower payload, containing the CIP, the battery power 

supply, and the ballast, was attached to the bottom of the parachute. 

Appropriate electrical cables and break-away connectors were rigged in 

parallel with the mechanical connections. The whole bottom assembly was 

designed to break away from the balloon and fall to earth suspended from the 

parachute at termination of the flight. 

17 
SECTION 4 

FLIGHT SEQUENCE 

4.1 PRE-LAUNCH PREPARATIONS 

The balloon launching pad at the NSBF is a large circular area, 2,000 

ft (600 m) in diameter. In the center of this large circle is another 

circular area, solidly paved, measuring 1,000 ft (300 m) in diameter. Graqs 

is planted in the area between the two circles, and a paved road surrounds 

the large circle. Paved spokes feed the center area from the outer road. 

When all pre-launch preparations had been completed and the staff 

meteorolcgist had predicted favorable weather and winds at Palestine and for 

some 300 mi (480 km) down-range, the equipment was taken to the launch site. 

The main balloon, protected by a plastic sheath, was laid out full-length on 

the circular paved area. It was aligned with the direction of the wind, 

with the top of the balloon upwind. The top end of the balloon was passed 

under, behind, and over the top of a large, smooth, horizontal cylinder 

mounted on the front end of a launch vehicle. One end of this launching 

cylinder was hinged to the launch vehicle. The other end of the cylinder 

had a latch which could be released by a trigger mechanism. After the 

balloon was passed Over the cylinder, the cylinder was pushed back to engage 

the latch so that the cylinder trapped the balloon. The top 10 m or so of 

the balloon was pllled forward from the cylinder, allowing the top payload 

to rest on the ground. It is this top 10 m of balloon that will later 

receive the helium gas during inflation. The helium will form a kind of 

bubble in this part of the balloon above the launching cylinder. After the 

launching cylinder was latched, final preparations of the top payload were 

18 
begun. The tow balloon was attached to the hoop with nylon lines, the 

clock was wound, the camera was energized, and a final checkout of the 

tracker ard data encoder was performed. 

Preparations at the bottom end of the balloon proceeded in parallel. 

The balloon was connected to the top center of the parachute canopy, and the 

parachute's shroud lines were connected to the bottom payload using a 

special fitting. All the connecting wires, lines and cables associated with 

the bottom payload cables were passed up and over a high, platform-like 

structure on the top of another launch vehicle. On the forward part of this 

structure, a fixed pin pointed forward. A hole in the special shroud line 

fitting was passed over this pin, and a quick release mechanism latched the 

fitting to the pin. The lower payload was thus suspended from the pin in 

front of the vehicle in full view of the vehicle's driver. After completion 

of all the mechanical and electrical hookups at both top and bottom 

payloads, a final checkout of all the CIP and ballast systems was performed. 

The launch sequence began by inflating the tow balloon with helium. 

The mainballoonwastheninflatedby passing a pre-determined volume of 

helium through two long flll-tubes and into the balloon as shown in Figure 

5. The balloon was launched by triggering the latch on the launching 

cylinder. When the latch released, a stout spring caused the free end of 

the cylinder to fly forward, rotating about the hinge. This released the 

balloon. As the balloon rose, the launch vehicle at the lower end of the 

balloon began to move forward (downwind). When the driver of the launch 

vehicle had positioned the vehicle directly below the balloon and had his 

vehicle going along at the same speed as the balloon, he released the latch 

on the pin and the lower payload was released. Figure 6 shows the 

19 
Figure 5. Inflation of the Main Balloon 

20 
PAGE IS 

- OF'POORQtJALITY 

Figure 6. Balloon Launch 

21 
balloon system and the launch vehicle a few seconds after release of the 

launching cylinder and just as the downwind launch vehicle began to move. 

When the balloon reached 2,000 ft (600 m) , an aneroid switch triggered an 

explosive charge which released the tow balloon and the launch sequence was 

canplete. 

4.2 FLIGHT 

The balloon normally ascended at a rate of approximately 900 ft/min 

(4.6 m/s) and reached float altitude in slightly over two hours. During the 

ascent, the flight controller at Palestine maintained a constant contact 

with Air Traffic Control (ATC). Data from the on-board navigational system 

was continuously given to ATC so that air traffic in the area could be 

vectored around the balloon. 

After the balloon had been launched, solar cell voltages, interspersed 

with reference calibration voltages and thermistor voltages, were fed into 

the telemetry system. PCM and were 

transmitted to the NSBF ground station along with the navigational, 

altitude, and other information from the CIP. At the ground station they 

were decoded, recorded, and displayed in real time for monitoring of the 

flight. The flight was timed to reach float altitude approximately two haurs 

before solar noon, so when the telemetry showed that float altitude had been 

reached, the tracker was turned on by telemetry command. Tracker operation 

was monitored by observing the telemetered values for the current drawn by 

the tracker motors, the output of the on-sun indicator, and the outputs of 

certain selected solar cell modules. Data was recorded continuously from two 

hours before solar noon until one hour past solar noon. 

These voltages were converted to 

22 
4.3 FLIGHT TERMINATION 

Shortly after launch ard while the balloon was ascending, the reccwery 

airplane took off from Palestine with the recovery crew aboard. This 

airplane was equipped with a radio system which allowed the crew to maintain 

constant communication with the balloon base either by direct transmission 

or by radio relay from the balloon. The airplane also had a full command 

system so that it could send commands to the balloon. During July, the 

winds at altitudes above 8O,OOO ft (24 km) blow from East to West at speeds 

of about 50 knots (25 m/sec), so the airplane had to fly about 300 mi (180 

km) West of Palestine to be in position for recovery. The pilot can fly 

directly toward the balloon at any time by using the ADF to track the low 

frequency beacon. 

The pllot of the recovery airplane is responsible for the termination 

phase. Before leaving Palestine, the pilot had been provided a set of 

descent vectors bv the staff meteorologist. The descent vectors are an 

estimate of the trajectories the payloads will follow as they come down on 

their parachutes. Upon receiving word from Palestine that the experimenter 

had all the data he needed, the pilot flew der the balloon and established 

its position accurately by visual observation. Using the descent vectors he 

then plotted where the payloads should comt down. He also established 

contact with Air Traffic Control. When ATC advised that the payloads would 

not endanger air traffic, ard when the descent vector plots showed that the 

payloads would not come down in an inhablted area, the pilot sent the 

commands to the balloon for termination of the flight. 

The termination sequence consisted of first sending a command which cut 

the electrical cable running from the bottom payload to the top payload. 

23 
This command simultaneous ly cut the cables holding the top payload onto the 

top of the balloon and opened the poppet valves on the top of the balloon. 

A second command released the bottom parachute from the balloon, allowing 

the bottom payload to fall away and caused the balloon to become top heavy. 

As the bottom payload fell, a ripcord attached to the top of the parachute 

opened a large section in the balloon. The balloon collapsed, the top 

payload fell off the balloon, its parachute opened and they all began their 

descent. 

It requires about an hour for the two parachutes and the balloon to 

fall from float altitude at the descent rate of roughly 1250 ft/min (6.4 

m/s). During this time, the pilot monitored the position of the bottom 

payload using his ADF. He also monitored the altitude information from the 

pressure transducer and relayed this information along to ATC. After 

reaching the ground, all three items have to be found. The low frequency 

beacon on the bottom payload usually aids in locating it rather quickly. 

The other items are usually located by flying search patterns centered 

around the impact zones calculated from the descent vectors. Once the 

impact locations were established from the air, the recovery crew drew 

detailed maps and later returned with ground vehicles to recover the 

payloads. 

Figure 7 is the fiighf profile for the 1986 balloon flight (No. 1425P 

in the nomenclature of the NSBF). It is a plot of altitude vs distance from 

the launch site from the time of launch until touchdown. The points are 

plotted on 10 minute time intervals. As the Figure shows, the touchdown 

site was approximately 340 mi (547 km) from the launch site and the total 

flight duration was just under seven hours. 

24 
x.. 

\ 

0 

2g-o 0 

mz 0 

OOL 1 

009 1 

00s 1 

OOP 1 

0081 0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

m 

-1 

0 

0 

00 

0 

h 

W 

IZ 

W 

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25 
SECTION 5 

DATA ANALYSIS 

The computer analysis was performed at JPL using the Univac 1100 

computer. Theprogram read the raw data from the magnetic tape produced 

during the flight, then corrected the cell data for temperature and sunearth distance according to the formula: 

VT,R = measured module output voltage at temperature T ard distance R. 

R = sun-earth distance in astronomical units (AU) 

A 

T = module output tgnperature coefficient. = module temperature in deg C. 

The remainder of this section describes the details of performing the above 

corrections and computing calibration values for the cells. 

5.1 c.xwuTER ANALYSIS 

The computer program read data from the magnetic tape a record at a 

time. Each record contained 16 scans of data plus a ground frame. Each 

scan of data consisted of two synch words followed by a reading of each of 

the 64 data channels. The ground frame contained the date and time of the 

scan, followed by the latitude, longitude and pressure as measured by 

instruments on the balloon. The program first checked the ground frame to 

see whether the record fell within the allowed data analysis time. (The 

allowed data analysis time window is an input to the program set by the 

parameters MINTIM and MAXTIM). The computer rejected the entire frame if 

the time of the current record fell outside the time window or if it could 

26 
read the time properly. The computer read records within the time window 

until it had accumulated 200 scans. Each set of 200 scans is called a pass. 

At this point, the data was in pulse code modulation (EM) counts. The PCM 

data for each channel was averaged, then a screening procedure was used to 

reject scans containing questionable data. For example, scans were rejected 

if the on-sun indicator was lower than a threshold value (input parameter 

OSMIN). Scans were also rejected if the PCM count for a data channel was 

not within the allowable PCM count range, if there was too large a count 

deviation on any channel from one scan to the next, or if there was an 

unrecoverable problem in reading the tape. 

The FCM data was next converted to engineerirq units. The program has 

provision for doing this in any of four different ways. The simplest and 

most commonly used method will be described here. During the calibration of 

the telemetry system, the output of the telemetry system vs the input from 

the voltage reference box was recorded to produce a table of mV input vs PCM 

count output. Similarly for the thermistor channels, a series of resistors 

was connected one by one across a thermistor channel ard the resulting PCM 

count for that channel recorded. Since the temperature corresponding to 

each resistance value was known, the construction of a PCM count vs 

temperature table was possible. This procedure was repeated for all 

thermistor channels and a calibration table was constructed for each 

thermistor channel. All these calibration values were input to t-he computer 

program. During conversion of solar cell data to mV, the computer performed 

a linear interpolation in the PCM vs mV table. At the completion of the 

initial computer analysis run, the output values of each channel 

corresponding to voltage reference box levels were ckecked. If they were 

found to have held constant during the flight, (they normally read constant 

27 
to within - + one PCM count out of 1,000) the use of the simple linear 

interpolation scheme was continued for the final data analysis. Since the 

relationship between thermistor resistance vs temperature is non-linear, a 

third degree polynomial was used for the interpolation of the temperature 

values. 

At the end of each pass, four averages were computed for each channel: 

(1) an initial average based on all acceptable scans, (2) a corrected 

average, using all data falling within a specified fractional deviation 

(input parameter ADEV) of the initial averages, (3) the corrected average 

multiplied by the square of the earth-sun radius vector (in AU) for the day 

of the flight (input parameter RV), and (4) a final average with all the 

above corrections plus a temperature correction to 28 deg C. This final 

correction used the values of the temperature coefficients measured in the 

laboratory ( input parameters TPCOEF). 

This completes the data analysis for one pass of data. The entire 

process was repeated by returning to the tape reading routine and reading 

another pass (200 scans) of data. All the screening and averaging routines 

were performed on this pass also, and this procedure was continued until 20 

passes of data had been analyzed or until MAXTIM was exceeded. 

In addition to the averages taken after each pass, an overall summary 

matrix was constructed which contained the fully corrected averages for each 

channel after each pass. A row of entries was added to this matrix after 

each pass. After all passes had been completed, an overall average and 

standard deviation were computed for each channel. These overall averages 

of approximately 4000 readings are the reported calibration values for the 

modules. 

28 
5.2 CALIBRATION RESULTS 

Table 1 reports the calibration vaiues of all ti ceils calibrated on 

the 1986 balloon flight, corrected to 28 deg C and to 1 AU. The Table also 

reports the standard deviation of the 4,000 measurements, the pre-flight and 

post-flight readirqs of each module in the X-25 simulator, a comparison of 

the pre-flight with the post-flight simulator readings, and a comparison of 

the pre-f light simulator readings with the balloon calibration readings. 

The Table also reports the temperature coefficients measured for each module 

in the laboratory. 

The data in Table 1 indicate that we did not have as god agreement as 

usual in the comparison of the pre-f light data with the post-f light data. 

This discrepancy was introduced because a malfunction in the X-25 required 

replacing its internal optical mirrors just before the post-flight readings 

were taken. The higher blue content of the simulator with its new mirrors, 

coupled with the fact that the simulator intensity was set with an old 

standard cell with low blue spectral response, caused the difference between 

the pre- ard post-flight readings. In general, the post-flight readings of 

the silicon cells were siightly higher .than pre-flight readings, while the 

post-f light GaAs cell readings were slightly lower than the pre--f light 

readings. 

5.3 UPDATE OF' BE'S-17A DATA 

Several standard modules have been flown repeatedly over the 24-year 

period of Calibration flights. The record of the one with the longest 

history, BFS-1?A, appears in Table 2 alorg with some computed statistical 

parameters. The standard deviation of 0.26 (0.43%) and a maximum deviation 

29 
of 0.74 (1.23%) from the mean show that the BFS-17A readings are very 

tightly grouped. In addition to giving a measure of the consistency of the 

year-to-year measurements, BFS-17A also provides some insight into the 

quality of the solar irradiance falling on the solar panel with regard to 

uniformity, shadowing, or reflections. Since this cell has been mcxmtcd 

in various locations on the panel over the years and its readj.nqs have been 

quite consistent, there does not seem to be any evidence of nonirriiformity 

over the panel. 

30 
R 

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4 

If 

T 

dddd 

.... 

31 
Table 2. Repeatability of Standard Solar Cell BFs-17A 

(38 Flights Over a 24-year Period) 

Flight 

Date 

9/ 5/63 

8/ 3/64 

7/28/65 

8/ 9/65 

8/13/65 

7/29/65 

8/ 4/66 

8/12/66 

8/26/66 

7/14/67 

7/25/67 

8/10/67 

7/19/68 

7/29/68 

0/26/69 

9/ 8/69 

7/28/70 

81 8/64 

a/ 4/67 

Calibration 

Value, mV 

60.07 

60.43 

60.11 

59.90 

59.90 

59.93 

60.67 

60.25 

60.15 

60.02 

60.06 

60.02 

59.83 

60.02 

60.31 

60.20 

60.37 

60.17 

60.42 

Mean 

Standard Deviation 

Maximum Deviation 

Flight Calibraticn 

Date Value, mV 

8/ 5/70 

4/ 5/74 

4/23/74 

5/ 8/74 

10/12/74 

10/24/74 

6/ 6/75 

6/27/75 

6/ 10/77 

8/11/77 

7/20/78 

7/24/00 

7/25/81 

7/21/82 

7/12/83 

7/19/84 

7/12/85 

7/15/86 

a/ 8/29 

60.18 

0.26 

0.74 

60.32 

60.37 

60.37 

60.36 

60.80 

60.56 

60.20 

60.21 

60.35 

60.46 

60.49 

60.14 

60.05 

60.07 

59.86 

60.10 

59.04 

59.99 

59.44 

9/5/63 to 8/5/70 Each data point is an average of 20 to 30 

points per flight 

4/5/74 to 7/1/75 Each data point is an average of 100 or mre 

points per flight 

6/10/77 to 7/21/82 Each data point is an average of 200 data 

points 

7/12/83 to Present Each data pint is an awrage of 4000 data 

points 
SECTION 6 

CONCLUSIONS 

The 1986 balloon flight was a successful flight. The new, larger solar 

panel was flown for the second time arvl it again performed in a satisfactory 

manner. BFS-17A again returned readings which are nearly identical to the 

readings it returned on its previous 37 flights. Three silicon reference 

cells which had flown in 1984 and 1985 were reflown this year. The 1986 

calibration values derived for these cells were within 1% of their previous 

values. One of these cells 

had been flown in 1979 and three in 1982. The repeatability of the GaAs 

cells was not as good. The calibration values had all decreased between 2.9 

and 4.1%. The exact nature of this apparent instability is currently under 

investigation. 

Four GaAs reference cells were reflown also. 

33 
1. B.E. Anspaugh, R.G. Downing and L.B. Sidwell, Solar Cell 

Calibration Facility Validation of Balloon Flight Data: 

A Comparison of Shuttle and Balloon Flight Results, 

cation No. 85-78, Jet Propulsion Laboratory, Pasadena, Calif., 

October 15, 1985. 

JPL Azbli34