Development History and Flight Performance of SERT 2 Solar Array

8/7/2019 Development History and Flight Performance of SERT 2 Solar Array

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1. Report No. I 2. Government Accession No. I 3. Recipient's Catalog No.NASA TM X-2246 I

4. Title and Subtitle 5. Report Date

6. Performing Organization CodeApril 1971DEVELQPMENT HISTORY AND FLIGHT P E R F O R W C E OF

SERT E SOLAR AI7. Author(s) 8. Performing Organization Report No.

Clifford E, Siegert E-603810. Work Unit No.

9. Performing Organization Name and AddressLewis Research CenterNational Aeronau tics and Space Administration 11. Contract or Grant No.Cleveland, Ohio 44135 13. Type of Report and Period Covered

12. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, ID. C . 20546 14. Sponsoring Agency Code

I15. Supplementary Notes

16. AbstractThe SERT 11solar ar ray consists of 33 000 2- by 2-em W on P silicon cell s configured to supply anominal electrical power at end of mission of 1100 W at 56 V for an ion thruster and 180 Wat 28 V fo r the spacecraf t9s command system, telemetry system, attitude control system, andsecondary experimen ts. The design and testing of the arr ay was limited to modificationsnecessary for the SERT I1 mission since similarly designed ar ra ys wer e flown successfully onsever al previous missions. The flight performance of the so lar a rr ay indicated that all re-quirements for the missi on have been met.

17. Key Words (Suggested by Author(s)) 18. Distribution StatementSolar arr ay Unclassified - unlimited2- by 2-ern cells

19. Security Classif. (o f this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price*Unclassified Un elass fied 33 $3.00

tFo r sale by the National Technical I nforma tion Service, Springfield, Virginia 22151


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NCE

i s ch CeY

The $E solar ar ra y consists of 33 000 2- by 2-centimeter Id onconfigured to supply a nominal electrical power at end of mission of 100 watts at 56 voltsfo r an ion thrus ter andlemetr y system, attitude control system, and secondary experimentstesting of the a rr ay was limited to modifications neces sar y fo r the Ssimilarly designed arr ay s were flown successfully on several previous missionsflight performance of the solar ar ra y indicated that all requirements f or the missionhave been met.

8 volts fo r the spacecraft's command system, te-

The

variety of missions have been performed in the N A space program using sol arcells as t h e source of el ect rical power, he designers for missions in the past mis-sions and miss ions in the future have had common concern, the selection of a solararray that will furnish sufficient power for the mission objectives to be accomplished.Futur e designers can exercis e greater car e in their selection of s ola r ar ra ys i f infor-mation is available about the design, development, test ing, and flight performance ofpreviously flown arrays mission was to endurance test an ion thrusterf o r a minimum of 6 months in a 1000-kilometer polar orbit. he sol ar ar ra y selectedto supply 1280 watts for the operation of the ion thruster experiment and spacecraftfunctions was an array that had been designed and flown successfully on previous AirFo rc e missions. The number of squ are feet of so la r cel ls and the power produced inorbit makes the SERT l2 so la r ar ray the largest ar ra y flown by NASA to date. The

The major objective of the SER


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solar array is the first ar ra y flown by NASA that was configured to supply morethan one potential. A section of I100 w a t t s for thru ster operation was configured f or anominal potential of 56 volts (*28 V), and the 180-watt section for spacec raft operationw a s configured for a nominal potential of 28 volts,

The purpose of th is r epor t is to present the requirements of the so la r arr ay for theT 1 mission, the overall design philosophy of the program9 the mechanical and the

electrical modifications necessary to adapt the existing designed ar ra y to the Emission, th e history of the limited development and testing progr ams , and the data ofthe sola r ar rays 's flight performance. The presentation of flight data will include theflight voltage and curre nt characteri stics , the operating temperatures of the arr ay andthe influence of or bi t position and geometric location, and an evaluation of the radiation-caused degradation of sho rt circuit current .

The SERT 11 spacecraft was to be launched by a horad/Agena D vehicle into anominal 1000-kilometer Sun synchronous circular orbit, inclined at 99-38'. A mercurybombardment ion thruster system w a s to operate for a minimum period of 6 months andcause an orbi t altitude variation of 100 kilometers. A f t e r a minimum period of 6 monthssunlight, the spacecr aft would en ter thewould last for 2 months and then normal spacecraf t operation and th ru st er operation wouldbe resumed.spacecraft, The ion thr ust ers are located on the Earth-facing end of the vehicle.

arth's shadow each orbit . This eclipsing

Figure 1 llu str ates the orb it of t he spacecra ft and the orientation of the Agena and

The sol ar ar ray was to consist of two deployable so la r arr ay wings. Each wingwas to be mounted to the aft rack of the Agena. For launch, the tw o arrays are foldedinside the envelope of the Agena booster adapter st ruc tur e. Once in orbi t the releasemechanism would activate and the deployment mechanism would extend the ar ra y. Fig-ur e 2 shows the fully extended so la r ar ray.

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The power requirement for the sol ar a rr ay when in orbi t w a s a minimum of I100watts for the operation of the th ruster and a minimum of I80 watts fo r the operation ofthe spacecraft . These on-orbit power requirements, when converted to a standard ofair mass zero (AMO)nd 25' C, require the thruste r section of the ar ra y to furnishI425 watts at a minimumof 6 2 volts (*38V), and the housekeeping section to furnish286 watts at a minimum of 34 volts.

solar ar ra y consisted of 90 panels, Figure 3 illustrates the arrange-ment of the se panels. The fourteen panels that are shown shaded (seven pe r wing) werethe panels that supply the electrical power for the operation of the spacecraft . The re -maining 76 panels (38 pe r wing) supply the electr ica l power fo r the operation of thethruster , Note the polarity distribution of the panels used to supply to the ion th ruster s.This configuration w a s selec ted to provide minimum magnetic torques

The SE

A l s o shown in figure 3 is an individual panel with the arrangement of the 74 sub-ach submodule consists of five 2- by %centimeter cells connected in paral-

lel, The submodules ar(SCA). For 25' C andcurrent w a s 682 milli amperes and maximum power wasconstruction of a five- cel l submodule.

eries connected to for m a panel o r solar collector assemblyopen circuit voltage of the S w a s 43.2 volts, sho rt circut

igure 4 shows the

umeThe instrumentation requirements for the sola r ar ra y wer e six temperature se nsors ,

an open circuit voltage monitor, a short circuit current monitor, and a peak power moni-tor . The open circui t voltage monitor was a single 2- by %centimeter cell connected to a2. 5-kilohm load res istor. he peak power monitor was a single 2- by 2-centimeter cellconnected to a &ohm load res ist or. The sh or t circu it current monitor consisted of two

ne cell was located on the active side of the arr ay and a second cell w a s locatedon the backside almost directly behind the fro nt cell. The cell on the backside of thear ra y would aid in determining Sun-spacecraft angles i f a reacquistion of the spacecraftwere necessary. The tw o cells were connected in parallel to a I-ohm load resistor.igure 5 illust rates the location of the six temperature sen so rs and the three moni-to r cells. All of the se nsor s except temperature sen so r 3 we re located on the backsideof a sol ar collector assembly. Temperature sen sor 3 w a s located on the backside of theplate to which the monitor cells were mounted. The temperature sen sor is almostdirectly behind the peak power monitor.

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The solar cell base resistivity used on the simi larly designed a rr ay s previouslyflown w a s IO ohm-centimeter. When the flight data for previous programs were re-viewed and compared to the SEand voltage characteristics of r solar array were lower than theminimum requirements for th he cell configuration for the90 panels of the array could have been redesigned into a different series-parallel com-bination; however, this approach would cause an increase in development effort. Asecond approach w a s to investigate the voltage and power charac ter ist ics of cells o therthan IO ohm-centimeter. The result of theinvest igation was the selec tion of a 2-ohm-centimeter cell which would meet mission requirements without changing the design ofthe panels,

ission requirements, it was found that the power

he deployment mechanism consists of a mounting bed, an adjustable liquid-dampedspring actuator, a slider, and a support assembly. The design of the deploymentmechanism for previous flown missions canted and rotated the solar arrtrical power w a s available for a large variation in Sun angles. The SERrequired a deployment mechanism which orientated the solar array w i t h minimum inci-dence angle to the Sun and located the center of mass of the sol ar a r ra y along the center-line of the spacecraft . The center of ma ss alinement was important because the space-craft attitude w a s gravity gradient stabilized. Three different configurations of thesolar ar ra y and spacecraft with the associated location of center of m ass are shown infigure 6.

The top b o llustrations in figure 6 show configurations in which the center of masswould be located along the longitudinal axis of the vehicle and the lower configurationillustrates the offset cente r of mas s location used fo r the SE mission. The changein deployment mechanism design to achieve the configuration in the top illus tration wasvery extensive. The change in the deployment mechanism design to achieve the middleconfiguration w a s a moderate one; however, it required the manufacturing of a left andright wing, The configuration shown in the bottom illustration was achieved by a minorchange to the existing deployment. mechanism. This was the configuration chosen andthe resulting mechanism is shown in figure 7,

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ease MechanismThe release mechanism for previous p rograms is shown in figure 8 . The tie bar

is fastened to the outboard panel by a bracket and the tie rod is held at the ends by pintain a successful release. A study w a s conducted to redesign the release mechanism toincrease the reliability of the sys tem. The design concepts A , B and C in figures 9and 10 us e the concept of attaching one end of the tie rod to the release mechanismclamp and securing the tie rod to the other clamp by tw o pin pullers. These designs A ,

his type of configuration required both pin pull ers to operate in ord er to ob-

B, and C require the operation of both pin puller s fo r a successful release.The final design illust rated in the bottom of figure 10 has the capability of a suc-

cessful release with only one pin puller operating. The operation of only one pin pullerwill allow the swivel section on the tie rod end to rotate free from the second pin puller.

The thruste r ar ra y supplied power to a power conditioner. One of the speci ficat ionsimposed on the th ruster a r ray by the power conditioner design was a limitation of opencircuit voltage to less than 75 volts during periods of thruster operation. When thethruster is operating at 80 percent or greate r thrust level, the operating voltage of thethruster array was calculated to be between 55 and 65 volts. However, during init ialsta rtu p of the thruste r, the thr ust er draws approximately 1 ampere and, as a result,the thruster ar ra y is near open circuit voltage.array temperature is 59' C o r above. During the init ial selection of the 2 ohm-centimeter cells, the predicted temperature was at 59' C and open circuit voltage wasless than 75 volts.. A s thermal studies were completed, a revised temperature as lowas 45' C was predicted, and as the actual current-voltage charac ter ist ics of the flighthardware were obtained, it became apparent that the open circuit voltage early in themission could exceed the 75- volt requirement.design change, revised temperature predictions and final curr ent and voltage predictionswas to sho rt out six of the five-cell submodules. A typical panel of 74 submodules isshown in figure 11. All 74 submodules are connected in series, The s ix submodulesthat are shaded wer e the ones that were shorted out. Only the 76 panels fo r the thrus terar ray wer e modified. The 14 panels fo r the housekeeping ar ra y were not modified be-cause the output of this section was regulated by switching mode regulators, s o thehigher open c ircuit voltage was not a problem,

This condition of operat ing near open circuit voltage is not by itself a problem if the

The method selected to correc t the open c ircuit voltage problem based on minimum

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The predicted open circuit voltage at 45' C for the thruste r section before shortingof submodules w a s 79,4 volts; .the open ci rcu it voltage after the submodules wereshorted w a s 73.0 volts. The voltage at maximum power was changed from 61.2 to 56,2volts; the sho rt circuit current and current at maximum power were not effected.

Degaussing o YWhen orbital stability studies were performed a maximum magnetic moment of 3000

gauss- centimeter was assumed fo r the so la r array. Measurements performed on sev-eral test panels revealed an average magnetic intensity sufficient to produce torques fo rthe ent ire ar ra y above 3000 gauss-centimeterw a s traced to the use of a magnetic test fixture to hold the five-cell submodules duringtheir electrical check out after assembly.w a s created in a 7- by 7- by 3-foot volume by 300 turns of 12-gage wi re car rying8.14 amperes at 60 hertz. The ar ray was in a stowed configuration. The magneticfield was gradually increased by increasing the cur ren t to 8- 14 amperes in a 2-minuteperiod; the field was held constant for 2 minutes, and then gradually decreased in a%minute period. Measurements wer e taken, and the degaussing procedure was re-peated with different orientations of the ar ray until minimum readings we re obtained.Final magnetic measurements we re taken when the a rr ay was fully extended.centimeter

33 he so ur ce of the magnetic intensity

The solar arra y w a s degaussed by introducing each wing into a 10-gauss field which

The magnitude of the res idual magnetic moment of t he so la r ar ray was 2900 gauss-3 This amount was determined to be within lim it s fo r orbit stability.

I- Y

. The purpose of the deployment test was to verify the new design o the deploymentmechanism. The st ructur e and deployment mechanism are designed to be operated in a1-g field w a s important. The test fixture for this test, shown in figure 1 2 , supportedeach section of wing by cable in a low friction roller assembly and provided a simulatedzero-g environment for the solar array. The solar a rr ay struc ture shown is not theflight st ruc tur e but a model. The tests were recorded by high-speed motion picturecamera s. The film w a s used to analyze the motion of th e sol ar a r ray as it deployed.

.zero-gravity environment and not in a l-g field, so fixturing to deploy the a rr ay i n a

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The actuator assembly was instrumented with st ra in gages which provided informationabout the loads being applied to the assembly during deployment.ar m assembly. For each test the deployment t imes were recorded and the flatness be-tween panels on the so la r arr ay was measured. The fil m data, st ra in gage information,deployment times, and panel flatness measurements were used in the selection of thebest adjustment for the actuator a r m assembly.

The test results showed deployment times from 46.8 to 54 seconds and the anglebetween panels varied from l.79.0 to 180.0. he fac t that the deployment times andthe panel flatness were reproducible established confidence in the deployment mecha-

en deployment tests were performed with different adjustments on the actuator

nism.

The purpose c the release mechanism tests was to integrate it with L e other com-ponents of the sys tem, verify normal operation, and demonstra te mechanicTOverify normal operation, two pin puller s are actuated by four squibs. To demonstratemechanical redundancy, only one pin puller would be operated by one and two squibs.

The zero-g environment for the release mechanism test was provided by a differentfixture than w a s used f or the deployment tests. For the release mechanism tests, thesolar ar ray was mounted to a pedestal as shown in figure 13. The pedestal allowed arotational freedom of movement during deployment, When the wing w a s deployed inspace, it would not necessarily follow a straight line, as it was re strai ned to do duringground testing in the zero-g test fixture. The movement of the pedestal would compen-sate for the r es traint applied to the wing.

The release mechanism was actuated seven times, s ix times by squibs and oncepneumatically. For tw o of the release tests, both pin pulle rs were operated by foursquibs, two squibs for each pin puller. Two tests wer e performed with pin puller P notoperating and pin puller 2 being activated first by two squibs and then by a single squiThen tw o tests were performed with pin puller 2 not operating and pin puller 1 beingactivated first by two squibs and then by a single squib.pin pullers and four squibs, with one pin puller and two squibs, and with one pin pullerand one squib.

The results of the tests verified the operation of the release mechanism with two .


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ceThe solar a rray had previously been qualified for the launch environment of an

Atlas/Agena. The purpose of this test w a s to establish confidence that the solar arrayassembly could withstand the higher dynamic loading from a Thorad/Agena D aunch.The test solar array w a s not the flight struc ture but a struc ture very si milar to flighthardware.a deployment test before and after the vibration test, electrical continuity checks,harnes s insulation measurements, visual inspections, and post vibration check of elec-tric al characteri stics of a single so la r panel.

It w a s recognized that an overtest can occur on a shaker when a greater number ofcycles are introduced into the structure at resonant frequencies than were actually ex-perienced in flight. Two possible methods to prevent such an overtest are to programthe shaker with amplitudes equal to the sinusoidal components and vary the sweep rate,thereby duplicating the number of cycles, or program the shaker with a constant sweeprate and reduce the amplitudes of the sinusoidal components proportional to the dampingratio of the st ruc ture. The method of constant sweep rate w a s selected fo r the vibrationtesting of the solar array.accomplished in three steps. The first step w a s to transform the actual flight data intosinusoidal components which are reproducible on a shaker. Next, the damping ratio ofthe solar array w a s experimentally determined, The third step w a s to modify the am-plitudes of the sinusoidal components proportional to th e experimentally determineddamping ratioshown in figures 14, 15, and 16, respectively.so la r ar ra y mounted to the shaker.

The so la r arr ay assembly satisfactorily passed the visual inspection, the deploy-ment test, the electrical continuity check, and th e harness insulation measurements.The postvibration illumination test performed on th e single panel indicated no seriousdegradation due to the higher vibration levels.

The sol ar ar ra y confidence test consisted of a vibration test in the x, y9 and z-axes ,

The procedure for obtaining vibration test amplitudes from actual flight data was

The expected flight levels and the actual tes levels for the x-, y-, and z-axes areigure 317 shows a photograph of the

The purpose of the flight acceptance test w a s to perform the following on the flight(I)Visual check of the mechanical structure

arness continuity and insulation checkssolar array:

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ower output check on each of the 90 panels(4) Assembly of so la r ar ra y struct ure to mounting bed(5)Actuation of release mechanism(6) Three deployment tests on each wingThe flight sol ar ar ra y assembly completed the flight acceptance test without any

anomalies.

The purpose of the mechanical f i t check of the solar arra y to the Agena w a s toproperly aline the sol ar ar ra y mounting bed to the Agena aft rack, check fo r properclearances on the assembled unit, mechanically fit and se cu re the solar a rra y harness,and check fo r proper continuity from the sol ar ar ra y terminal board through the Agenavehicle. In fig ure 18 the solar array is shown in its stowed configuration mounted to theaft rack of the Agena vehicle. The solar a r ray fit check w a s completed w i t h only oneminor modification, the Agena booster adapter connector w a s relocated because of itsinterference with the Agena heat shield.

F in a l De p lo ym e n t Te stThe final deployment test w a s performed in conjunction with the thermal taping

modification necessary for a spr ing launch and the degaussing of the entire sola r arr ay .Thermal taping of the tie rod, which is on the shade side of the vehicle during ascent,is necessary t o prevent excess tension due to temperatu re differentials. The thermtaping had been applied in anticipation of a f a l l launch. When the launch w a s rescheto a spr ing launch, the thermal tape was removed from the tie rod of one wing and ap-plied to the tie rod of the other wing.cally modified by shorting out si x of the submodules on each of the thruand a shade shield w a s installed on the sho rt circuit cur rent monitor.to disassemble the solar array structure from the mounting bed and remove the associ-ated harness in or der to accomplish the electrical modification. The final deploymenttest revealed that a ser vic e loop in the power harness w a s not correct, and, as a result,the so la r ar ra y would not have fully extended in flight. The harnessing w a s reworkedw h i l e the array w a s in a fully extended configuration to ensure proper format ion of theser vice loop.

Two months prior to the final deployment test, the sol ar a rr ay had been electri-

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ange, Vandenburg Air F ~ r c e ase, Lompoc, C'iPPifom-ia, thefollowing tests were performed prior to launch:

(Before solar ar ra y was mated to vehicle)(I)Solar array continuity and insulation checks

Electr ica l compatibility check with vehicleElectr ic checkout of the squib circuitsher sol ar a r ray was mated to vehicle)

(4 ) Continuity and insulation checks, and instal lation of squibs(5)Polarity checks on power harnesses(6) Connection of sol ar ar ra y power harness to sp ac ec ra f support unit and to space-

er successfully completing these tests, the sola r arr ay was ready to support thecraft

ehe predicted current voltage characteristics fo r the thruster arr ay are presented

in figure 19, The I-V curves were drawn for the predicted temperature of 116' F(47' C ) with a *tP3' F *To C ) deviation. An effective solar flux, wM& includes space-craft angle and s o r flux variation due to the Earth's orbit, of 41 Qhour (129 mW/cm was used to calculate values for these curves,thruster ar ray for the first 6 months. The tw o data points for orbits 95 on 2/436 on 3/7/70 repre sent the thr ust er ar ra y open circuit voltage. Thr ee data points rep-resent the voltage and current characteristics of the thrus ter ar ra y during the opesationof thruster 2. The ata point fo r orb it 00 on 2/11/78 is the 38 percent thmst level ofthe thr ust er, the data point for orbit 109 on 2/11/70 is the 80 percent thrust Bevel, andthe data point for orbit 13 on 2/12/70 is the BOO percent thrust BevelB, The remainingfour data points rep resen t the voltage and cu rrent charac teris tics of the thruster arra yduring the operation of th ru st er 1 . The data point for orbit I40 on 2/64/70 represents30 percent thrust level, the point for o it 146 on 2/14/70 represents 80 percent thrustlevel, the data point for orbit 171 on 2 /70 represents 100 percent level, and the poinfor orbit 2200 on 7/15/70 represents 100 percent t h r u s t level I49 days later than orbit171.

Lu per square foot -There are nine flight data points plotted which represen t the performance of the

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The predicted current-voltage (I-V) characteristics for the houskeeping array arep r e s e n t 4 in figure 20. t he I-v curves were drawn f o r a predicted temperature of 116' F(47' C)with a -+IS' F -+ e ) eviation. An effective so lar flux of 410 Bki per squarefoot - hour (129 mw/cand 2208, The electric& loads for orbit 2 on 2/3/78 were the command system, datahandling, spacecraf t instrumentation, and two tape recorders The electrical loads fororbit 8 on 2/3/70 were the electrical loads of orbit 2 and two control moment gyros.The electrical loads for orbit 556 on 3/36/70 were the electrical loads of orbit 8, aswell as a miniature electrostatic accelerometer ( ESA), radiofrequency intewere the electrical loads of orbit 8 and a spa ce probe experiment.

w a s used to calculate the values for these curves.The four data points plotted on the curves in figure 20 are for orbits 2, 8, 556,

FQ9 and space probe experiment. The electrical loads for orbit 2200 on

The curves shown in figures 211 and 22 represent the temperature variation of onewing of the solar array as a function of its position around the Earth. The tim e spancovered is the first 98 days of the mission. As shown in figure 5, temperature sens or 4was located at the outside tip of the solar array, sensor 5 was located in the middle ofthe array, and sensor 6 was ocated on the inside section nea r the Agena vehicle.shown in the figures are thetur e variation for the various orbits.the warmest. The temperature gradient from sensor 4 to sensor 5 is sma lle r than theof the Agena vehicle is limialbedo on the temperature variation of the so la r array, In figure 21 the spacecraftangle w a s 21'the spacecratt angle was 6" and sensor 4 emperature variation was 4 , 5 O F (2, io c).similar reduction o f temperature variation can be noted for the other tw o sensors,tion of the array because the current drawn f r o m this array was nearly constant, Forfigures 2 1 and 28, the current being drawn was 86,4 mperes. The telemetry resoki-kion f o r Vgae thruster ar ra y voltage i s larger khan the varbation of the thruster array

mster array voltage tha t are associated with the tempera-The temper ature curves show that the temperature of the arr ay nearest the Agena is

te mpe ra br e gradient from sor 5 o sensor 6 - This indicates that the thermal effectto the half of the solar ar ra y wing nearest th e vehicle.

An examination of the curves wi%l eveal the effect of spacecraf t angle and EarthTsd ternperatwe variation S ~ Fensor 4 was I C B , T O P (5, ' c), figure 22

The thruster ar ra y voltage w a s selected fo r purposes of showing the voltage vasia-


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voltage, hence the variations are represented by one of two different voltages fo r a par-ticular orbit,

The three solar ar ray monitors were open circu it voltage, short circuit current,and peak power. The open circuit voltage monitor failed; however, this data could beobtained from the thruster section of the ar ra y whenever the thrus ter experiment isturned off.plotted in figure 23 for the beginning of mission, 30 days, 60 days, and 103 days afterlaunch.

The la rge variation in the value of the short circuit current is caused partially byspacecraft angle, Earth's albedo, and a shade shield. The position of the shade shieldw a s such that data, as shown, from the equator to the South Pole were affected becausethe sho rt circuit curr ent monitor w a s being shadowed. Data fr om the South Pole t o theNorth Pole were affected by the spacec raft angle and Earth's albedo. The curve forFebruary 4 has an abrupt change as the spacecraft w a s at the equator traveling towardthe North Pole. This change w a s caused when the shadow cast by the tie rod crossedthe sho rt circuit current monitor.has a similar trend in variation as the solar arr ay temperatures. The larger thespacecraft angle, the la rg er the amplitude variation in short circuit current. This am-plitude variation resulted from added flux of the Ear th' s albedo. Generally, the largerthe spacecraft angles th e larger the deviation of the spacecraft orbit from the Earth'sterminator. From the vernal equinox to autumnal equinox the spacecraft is on the sunside of the terminator at the North Pole and on the shade side of the terminator at the

ole. From autumnal to vernal equinox, the spacecraft's orbit is on the shadeside of terminator at the North ole and on the sun side of terminator at the South Pole.peak power monitor are shown in figur e 24. The data shown for the peak power monitorare taken from the sa me orbits as the sho rt circui t curren t monitor. . The abrupt changein value of peak power for the orbit onarray tie rod crossing the cell.larger the spacecraft angle, the arger the variation in the amplitude of the value fo rpeak power. It can be noted from the curves that the peak power value for February 4

Short circuit current. - The flight data for the sh or t circuit current monitor are

The variation of the sh or t circuit current which was not affected by the shade shield

Peak power monitor. - The flight data for a span of 60 days of the mission fo r theebruary 4 is caused by the shadow of the solar

he peak power data show a trend si mi la r to that of the short circuit current. The

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is higher than the value fo r April 2 despite the fact that on April 2 the spacecraf t anglewas less and the so la r flux was grea ter . This lower value is attributed to increasedtemperatures of the cell mounting plate and radiation degradation.

At the beginning of the miss ion, the temperature of the plate to which the cel lswere mounted was below the telemetry range and is estimated to be 90' F (32' C). A sthe mission progressed and the paint on the mounting plate degraded, the temperaturetained from the peak power monitor was the voltage reading across a 4-ohm resistor.This voltage is affected by the temperature of the cell, as well as radiation degradationeffects on both current and voltage.

Radiation degradation. - The data presented in f igure 25 show the relative percent-age of radiation degradation of the sh or t circuit cur rent monitor. The previous datapresented in figures 21, 22, 23, and 24 give an indication of the difficulty involved incomparing data points of the solar ar ra y monitors over a 6-month period because of thevariation of sol ar flux, spacecraft angle, Earth's albedo, and temperatures . The tech-nique used for comparing data over a 6-month span was to use the temperature of thesolar array as an indicator of the amount of effective so la r flux falling on the ar ray.thrus ter ar ra y would be indicated by a decreas e in thrus ter ar ra y voltage. Several datapoints taken on the thrus te r a r ray during open c ircui t voltage conditions showed no volt-age degradation.

During the first 6 months it w a s noted that, for the sa me ar ra y temperature, thethrus ter ar ra y voltage decreased twice. The first voltage decrea se occurred 36 daysafter launch; the second occurred 161 days after launch. The selection of these twodata points allows the thruster voltage to be determined more accurately than teleme tryresolution normally permits. The decrea se in thruster a r r a y voltage represents a

percent degradation in current. The curve in figure 27 for short circuit currentshows that the percentage change in degradation from day 36 to day 161 is 10 percent.The estimated degradation fro m launch to 16 days after launch was 14 percent.

of the mounting plate increased into the telemetry range. The telemetry reading ob-

The thruster a rr ay had nearly a constant current load. Any degradation of the

The resu lt s of the design, development, testing, and performance of the sola r ar ra yare as follows:

l e A previously designed s olar ar ra y was successfully modified to meet the requi re-ments of the SE 'I'nr mission, and flight qualified status was maintained with a minimumof confidence testing.

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2. The existence of two different potentials on the so la r ar ra y has not presented anyhe flight temperatures for the solar array were ne are r the upper limit of the

in-flight problemspredicted temperature range.degradation of the peak power.

4. The degradation of the shor t circui t current followed the curve of the predicted

Lewis Research Center,National Aeronautics and Space Administration,

Cleveland, Ohio, December 23, 1370,904- 3

NCSolar Array. Rep. kMSC/A941440-Rev., ockheed Missil es and

Space Co. (NASA C 72706), May 4, 1970.2. Falconer, John C. Final Extrapolated I-V Curves fo r SERT ar Array Sub-

system. Rep. k C/TM-69-22-107, Lockheed Missiles and Space C O , 1968,3. Anon.: SE Solar Arr ay Release Mechanism Engineering valuation Test,

C/A937499, Lockheed Missiles and Space eo . , 1968.4, Anon.: SE T I Solar Array Module Confidence Test.

Lockheed Missiles and Space Co. 1968.Solar Arr ay Engineiles and Space Co. ,


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//// \/ \/ \\/ \

\\\/

S //\ /\ /\ /o l a r\ / c i r c u l a r o r b it\ //,\

C0-9017- 9

F i gu r e 1. - SERT I I in orbit viewed from Sun.


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Figure 2. - Solar array in extended posit ion mounted to Agena in Earth orbit.


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+ Pan els with positive potential, negative groun d- Panels with negative potential, positive gro und

Thrus t or a r ra y:76 panels6 8 submodules1425W , 25' C, AM0

(a) 45 panels per wing arranged in 15 by 3 configuration,

(b) One of 90 panels for SERT II.

Housekeeping array :14 panels74 submodules286 W, 25" C, AM0

Figure 3. - Panel arrangement for solar array and submodule arrangement for a panel.


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2 0 M I L 7940 f us ed s i l i c acover s l ides--._

T unne l ov en ooe r a t i on2-by 2-cm photovol ta ic

So l de r p r e fo r m -,

BaseplateI n s u l a t i n g s h e e t-

G r i d bas e: magnes i um c as t i ng- Bac k si de : w h i t e r e f l ec t i v e pa i n t

F i g u r e 4. - Assembly of f i ve-ce l l submodule.

S o l a r a r r a y t e m p e r a t u r e s e n s o r4 7 6 7 r 3

D i r e c t i on o f s pac ec r a ft t r av e l/o n i t o r

F i g u r e 5. - Loc a ti on o f s i x t empe r a tu r e s ens o r s , open c i r c u i t vo lt age mon i t o r , s ho r t c i r c u i t c u r r e n tmon i t o r , and peak pow er mon i t o r .


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1S u n(a ) Pos i t i on l .+-

CG OrbitfS u nIb) Pos i t i on 2.

1Sun( c ) Pos i t i on 3.F i gu r e 6 - A l t e r n a t e s o l a r a r r a y c o n f i g u r a t i o nc ons i de r ed f o r SERT I1 mi s s i on .


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a

21


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#

a,VIma,-L0-VI.-VI

VWVInzaa,SUm.-

22


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23


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6l- POGO

.011 2 4 6 8 10 20 40 60 80100 200 400Frequency, HzF i g u r e 14. - Sin uso idal tes t leve ls fo r X-ax is exc i ta t ion.

.01 I 1 1 1 I I I I I4 6 8 10 20 40 60 80100 200 400Frequency, HzF i g u r e 15. - Sinus oidal tes t leve ls for Y-ax is exc i ta t ion.

24


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4

2F.01 I 1 1 1 I I I I I l l I4 6 8 10 20 40 60 80100 200Frequency, Hz

F i g u r e 16. - S i nus o i da l t es t l ev e ls f o r Z - ax i s ex c i t a ti on .

F i g u r e 17. - S o l a r a r r a y in s tow ed c on f i gu r a t i on mou n ted to shake table.

25


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Figure 18. - Solar ar ray in stowed position mounted to Agena in horizontal position.

80 63.1 13.94 125 (51.6)30 68. 5 7.4 127 (52. )a 113 2 100 6L3 17.22 128 (53.3)80 64.9 13.94 123 (50.5)30 68.5 7.4 125 (51 )0 122 (50.0)I I I I I I I I I I I I I I MU0 5 10 15 20 25 30 35 40 45 50 55 60 65 70" 75Potential, V

Figure 19. -Thr uste r array characteri tics for spring launch. Solar input, 410Btuper squ are foot - hour (129.2m W h ; array temperature, ll6*13" F (47*7"C).

26


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10987

a 6L 5v 4

321

+-L3

O r b i t P o t e n t i a l , C u r r e n t , T e F p e r a t u r e ,

0 2.5 5.0 7.5 10.0 12.5 15.0 17.5 20.0 22.5 25.0 27.5 30.0 32.5 35.0 37.5 40.0 42.5Potent ia l , VFigu re 20. - H ous ek eepi ng a r r ay c ha r a c te r i s t i c s f o r s p r i ng launc h . So l a r i npu t , 410 B tu pe rs qua r e f oo t - h o ur (129.2 mW/crnZ) ; ar ra y temperature, 116+13" F (47*7" C).

15 0

13 0

11 0

U

6 0cL

55alcc

L-0l

L

0l45

-

1 4 0 , ~ ~

-0

120()-0

T emper a tu r es ens o rA 61 7 50 4A A A A A AA A A AA A A A A n A n n

-5:%&i2 6 0Equa to r Sou th Po l e Equa to r N o r th Po l eO r b i t p o s i ti o n

F ig ure 21. - So l a r a r r a y t emper a tu r es and th r us te r a r r ay v o lt age fo r F eb r ua r y 18 , 1970. Spac ec r af tangle, 21"; so lar f lux , 418.5 B tu pe r squa re foot - ho ur (131.9 mW/cm2L

27


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Y0

EccPL

L

0nm0l

45

n n

150 rT emper a tu r es e n s o r

A 60 50 4A A A A A A A A A AA A A A A A

110 I

2 2 %S k r "-6o83qua to r Sou th Po l e Equa to r N o r th Po leOr b i t pos i t i on

F igu re 22. - So l a r a r r a y t emper a tu r es and th r u s te r a r r ay v o l t a e f o r Ap r i l 28 , 1970. Spac ec r a ft ang le ,6"; solar flux, 430 B tu pe r s qu a r e f oo t - hour (135.5 mW/cm9

Date Spacecraf t angle, So lar f lu x ,d eg B t u / f t 2 - h r ( m W h n 2 )0 May 27 , 1970 12. 5 416.5 (131.3)A Ap ri l 2, 1970 5.8 436 (137.4)D M a r c h 4, 1970 17. 8 428.5 (135.0)Feb rua rv4 . 1970 27.5 406 (127.9)

ID5 1 I I

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Date Spacecraf t angle, Sola r f lux , Cell t emper a tu r e ,deg B tu / f t 2 - h r ( mW/c m2) "F ("C)0 May27, 1970 12.5 416.5 (13L3) 107 (4 1 7)A A p r i l 2, 1970 5. 8 436 (l37. 4) 1U3 (39.4)D M a r c h 4, 1970 17. 8 428. 5 (U5.0) 99 (37.2)

6o r F e b r u a r v 4 . 1970 27.5 406 (127.9) 90 (32.2)

I I N or th Po l e30Equator Sou th Pole EquatorOr b i t pos i t ionF i g u r e 24. - Peak pow er p l o t ted aga i nst o r b i t pos i t i on f o r pe r i od i c o r b i t s t h r oug hou tmiss ion.

F i g u r e 25. - Percentage of d e gr a da ti on f o r s h o r t c i r c u i t c u r r e n t m o n i t o r f o rSERT I1 miss ion.

NASA-Langley, 1971- 1 E-6038 29


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Development History and Flight Performance of SERT 2 Solar Array

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