47807330 Solar Sails System Development Trends

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Spacecraft Solar Array Technology Trends

P. Alan Jones & Brian R. SpenceAEC-Able Engineering Company, Inc.

93 Castilian Dr.Goleta, CA 93117

805-685-2262email: [email protected] & [email protected]

AbstractPhotovoltaic solar array systems are the mostcommon method for providing spacecraft power generation.The flexibility and variability of the many array types andconfigurations combine to accommodate a multitude ofmission applications and space environments. Solar arraytechnologies and their system configurations changeddramatically over the years as more aggressive anddemanding requirements were imposed. This paperaddresses the historical solar array technology trends. Theevolution of solar array technologies and the keyrequirements responsible for driving this evolution ispresented. Industry growth trends towards future systemsare identified. Evolutionary technological improvements inphotovoltaics, structural platforms, and deployment systemsare shown. Array selection criteria for a variety ofrequirements, applications, and environments are presented.Solar array technologies required to meet future missiontrends are shown.

TABLE OF CONTENTS

1. INTRODUCTION

2. SOLAR ARRAY HISTORY3. DRIVING REQUIREMENTS4. SOLAR ARRAY TECHNOLOGIES

- Photovoltaics- Structural Platforms

- Deployment Systems

- Mechanisms5. FUTURE SOLAR ARRAY

TECHNOLOGIES & TRENDS6. CONCLUSIONS7. REFERENCES8. BIOGRAPHIES

1. INTRODUCTION

Photovoltaic solar array systems are the most commonmethod for providing spacecraft power generation. In a timeperiod of less than four decades space solar arrays havegrown in size from less than 1 watt to systems over75,000 watts, such as the International Space Station Alpha(ISSA) solar array. GEO spacecraft power growth as afunction of time is shown in Figure 1.

0

5

10

15

20

25

30

35

40

45

1984 1986 1988 1990 1992 1994 1996 1998 2000 2002 2004 2006 2008 2010

Delivery Year

Power(

kW)

Simple exponent ia l model f i ts actual power

growth c urve from 1978-1999.

Figure 1. GEO Spacecraft Power vs. Time

The flexibility and variability of the many solar array typesand configurations combine to accommodate a variety ofmission applications and space environments. Electricalpower generation by photovoltaic conversion represents aclean and environmentally safe process for providing energyto a spacecraft system. These environmentally safe featuresled photovoltaic means to be considered as a politicallyacceptable replacement for nuclear radioisotope thermalgenerators (RTGs) for near planetary missions.

Solar array technologies changed dramatically over the yearsas more aggressive and demanding requirements werecontinuously b eing imposed by engineers who were creatingspacecraft systems with greater capabilities. As a result ofthese driving requirements and unique mission applications,innovative solar array technologies were developedthroughout the course of history. Primary solar arraytechnology developments included the optimization ofevolved structural platforms, lightweight substrates,

innovative deployment systems, and higher efficiencyphotovoltaics.

The development of solar array technologies is providingthe spacecraft engineer with a broad trade space of feasiblesolutions. The multitude of array subsystem and systemlevel solutions is allowing the engineer a greater ability tooptimize an array and specifically tailor it to a particularmission. As more exo tic future mission applications arise,which will impose even more stringent requirements, the


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solar array trade space will undoubtedly be furtherbroadened.

2. SOLARARRAY HISTORY

Historically, solar array technologies evolution was closelycoupled with expanding spacecraft power requirements.

Enhancements in spacecraft and launch vehicle systemstended to increase spacecraft capability which, in turn, drovethe development of more efficient solar array technologies.This trend, when coupled with the cost pressures inherent ina commercial industry, increased the demand for more costand mass efficient high power systems.

Array structural platforms evolved throughout the yearsfrom simple spacecraft body-mounted configurations, tosingle panel flip-out configurations, to kinematically complexmulti-panel deployable systems. This array configurationevolution is depicted schematically in Figure 2.

Graphic courtesy of JPL

Figure 2. Solar Array Configuration Evolution

The first solar array to fly in space was launched on17 March 1958 on the U.S. Vanguard I spacecraft[1]. TheVanguard I spacecraft is shown in Figure 3. This arrayconsisted of six simple body-mounted panels populated with10% efficient silicon solar cells. The total power output ofthis array was less than 1 watt[1]. The array on theVanguard I spacecraft quickly ushered in the spacephotovoltaic solar cell age as more advanced systems weredeveloped to meet ever-increasing power requirements.


Figure 3. U.S. Vanguard I Spacecraft

Simple body-mounted solar array configurations were mostprevalent in the early years. To accommodate increasingpower requirements the entire exterior surface area of thespacecraft was soon utilized for mounting solar cells. Aspower continued to grow, spacecraft were fitted with higherefficient photovoltaics and/or solar paddles in an effort to

extend the available solar array area. An example of aspacecraft configured with paddle type arrays is shown inFigure 4.


Figure 4. Spacecraft with early Paddle Arrays

Spacecraft designs and mission applications soon requiredeven more powerful solar arrays than what could beprovided by paddles. Solar panels attached to orientationdrives provided one solution and large cylindrically shapedspacecraft (accommodating the majority of the launch

vehicle volume) with body-mounted solar cells providedanother. An example of a cylindrically shaped spacecraftwith body-mounted solar cells is shown in Figure 5. Thesetwo solutions passivated power requirements for a shorttime but were limiting in power growth potential.

Photo courtesy of TRW

Figure 5. Cylindrically Shaped Spacecraft with

Body-Mounted Solar Cells

Spacecraft capability and mis sion applications continued torequire increased power. The trend was clear. As spacecraftcapability and mission applications drove array technologydevelopment, increased solar array power requirements alsoincreased to keep pace. Body-mounted array systems werelimited in capacity and what the industry needed was acompletely new and innovative solar array system which


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could provide power growth potential for futureapplications. This need provided the catalyst for thedevelopment of the large multi-panel sun-orientateddeployable solar array. An example of a large multi-paneldeployable solar array is shown in Figure 6.

Photo courtesy of FSS

Figure 6. Multi-Panel Deployable Solar Array

Today, the multi-panel deployable solar array is the mostcommon system utilized for high power applications. Itsability to provide accommodation for power growth (up to 15kW), high reliability, competitive weight (45 W/kg), and lowcost (


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process. Shortly after, the U.S. Air Force commenced amanufacturing technology program with Tecstar todemonstrate GaAs/Ge growth through a high volume metal-organic chemical vapor deposition (MOCVD) reactor. Highvolume production with the GaAs/Ge MOCVD process, orvariants of, is currently being performed at Spectrolab andTecstar. Today, single-junction GaAs/Ge solar cells accountfor over 50% of the total photovoltaic production.

According to many solar cell vendors, the demand forsingle-junction GaAs/Ge will soon replace silicon for mostfuture applications. To counter this demand for GaAs/Gephotovoltaics, various solar cell vendors developed a thinhigh-efficiency silicon solar cell which provides over 17%BOL efficiency. This cell is providing many commercialusers with an intermediate solution between standard siliconand GaAs/Ge, as it costs less than GaAs/Ge and producescomparable array level mass characteristics.

Recently, in the late 1990s, the basic single-junction 19%GaAs/Ge cell saw considerable evolution. The majorevolution of this basic cell platform was to grow additionalphotovoltaic active layers over the single Ga As/Ge junction.

This technique led to the development of multijunction solarcells. Multijunction GaInP2/GaAs/Ge solar cells, developedfor the U.S. Air Forces manufacturing technology program,produced efficiencies up to 24.2% for large-area cells. Thetriple junction variants have yielded efficiencies over25.5%[9]. As power requirements continue to grow, andmass and volume requirements remain constrained, the needfor even higher efficiency photovoltaics will be critical.Advanced multijunction solar cells will be a criticalcomponent of future solar array power growth.

The anticipated deployment of the many proposed low-earth-orbit satellite constellations has renewed interest in

low cost thin film photovoltaics. The most promising near-term thin film photovoltaics are amorphous silicon (a-Si),cadium telluride (CdTe), and copper indium galliumdiselenide (CIGS)[13]. These thin films have beensuccessfully employed in terrestrial applications but haveyet to be adapted to space in high volume quantities.Although the BOL efficiencies of thin film photovoltaics aresignificantly less than space standard photovoltaics theircosts savings potential is enormous. Transitioning thistechnology to the space sector will require a major massproduction development effort and a thorough qualificationtesting program. The anticipated benefits of thesetechnologies are promising, however they still awaitrepeatable space demonstration and commercial viability

[13].

Structural Platforms

Solar array structural platforms can be categorized as rigid,flexible, and concentrator systems[2]. An array platform canbe configured as a deployable or non-deployable system.

Most deployable solar arrays flown to date employed rigidhoneycomb panels interconnected with spring-driven hingesand electrical harnessing. These systems stow folded and

attach directly to the spacecraft sidewall. A picture of a rigidpanel solar array in its stowed configuration is depicted inFigure 7.

Photo courtesy of AEC-Able

Figure 7. Rigid Panel Solar Array in itsStowed Configuration

After release of the launch tiedowns these arrays deployoutward, in an accordion fashion, until each panel iscompletely flat. Deployment is, preferably, fully coordinatedby synchronization mechanisms and deployment rate isgoverned by a damping mechanism(s). Aluminumhoneycomb core and facesheet materials were the typicalconstruction of rigid panels in early systems.Fiberglass/epoxy composite facesheets were thenincorporated to reduce weight. Ply orientatedKevlar/cyanate ester and carbon/cyanate ester compositefacesheets replaced the fiberglass/epoxy systems to further

reduce weight and volatile outgassing. In some instances aKapton/polyimide film was used as a facesheet material.Additionally, most fabricators now employ an adhesivereticulation technique during the construction ofhoneycomb panels which minimized the facesheet-to-coreadhesive, further reducing weight. More advancedfacesheet materials being implemented today consist ofsingle-ply orthographically oriented open-weave materials,and extreme high-modulus, high-strength directionallyoriented carbon fiber laminates. Another rigid panelconstruction currently under consideration is the isogridreinforced structure. A typical rigid panel solar array withhoneycomb panels is shown in Figure 8.


Figure 8. INDOSTAR PUMA Rigid Panel Solar Array

Stowed Rigid Array


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Depending upon application, honeycomb panels can be alarge component of the overall array system level weight.For some missions which demand reduced weight and/orimproved power growth potential, honeycomb panels arebeing replaced in favor of a tensioned flexible blanketsystem. The flexible blanket solar array primarily consists ofa laminated blanket assembly with photovoltaics and a

structural deployment system. Blanket tensioning isachieved and maintained on-orbit by the deploymentmechanism. The deployment mechanisms are generallycoilable or articulated mast type systems, although tubularsystems have been employed and inflatable systems havebeen contemplated. The first flexible blanket solar arraydeveloped was for the Communications Technology Satellite(CTS). A drawing of the CTS spacecraft with its deployableflexible blanket solar array is shown in Figure 9.

Graphic courtesy of JPLFigure 9. CTS Spacecraft Flexible Blanket Solar Array

NASAs first major flexible blanket solar array, successfullyflown on the Solar Array Flight Experiment (SAFE) program,employed a coilable mast deployer system. The SAFEflexible blanket solar array is shown, partially deployed, inFigure 10.

Photo courtesy of LMSC

Figure 10. NASA SAFE Flexible Blanket Solar Array

Flexible blanket substrates are composed of a laminatedfiberglass or carbon fiber composite and Kapton/polyimidefilm substrate, or a single-ply composite material. As such,these substrates are very thin (>1mm) and flexible in nature.Special accommodations are required in the array packagingdesign to withstand the stowed launch and handling

environments. Two types of stowage configurations havebeen implemented with flexible blanket systems. The firsttype is a folded blanket configuration which deploysoutward in a similar accordion fashion as the rigid panelsystem. An open cell polyimide foam material is employedon each side of the stowed container or as an interleavematerial between blanket folds. When stowed, the foam ispre-loaded against the folded array to provide cushioningand protection for the delicate photovoltaic cells duringlaunch and handling. The SAFE solar array, shown in Figure10, is an accordion type flexible blanket system.

An alternate stowage configuration is the roll-up type inwhich the blanket is rolled up on a cylinder during launch

and unrolled by a deployment mechanism duringdeployment. A picture of the Hughes FRUSA roll-up solararray is shown in Figure 11. A compliant interleave materialof embossed Kapton or a separated sheet of polyimide foamis used to provide protection of the photovoltaics from thelaunch environment. The Hubble Space TelescopeSpacecraft employed a flexible roll-up array. A picture of thisarray is shown in Figure 12.


Figure 11. Hughes FRUSA Flexible Roll-Up Solar Array

Flexible blanket systems are becoming more popular and arebeing used on many spacecraft. The most notablespacecraft to implement flexible blanket solar arrays includethe Milstar, the Hubble Space Telescope, Olympus, CTS andERS-1[4]. New spacecraft which have yet to fly but willutilize thin flexible solar array technologies will includeNASAs EOS-AM[5] and the International Space StationAlpha (ISSA)[6]. A picture of the EOS-AM flexible blanketsolar array is shown in Figure 13. This innovative solar arrayproduces over 7 kW BOL and is the first NASA flexible


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blanket array to employ single-junction GaAs/Gephotovoltaics.

Photo courtesy of NASA

Figure 12. Hubble Space Telescope Flexible

Roll-Up Solar Array


Figure 13. EOS-AM Flexible Blanket Solar Array

The solar arrays for the ISSA are enormous. The ISSA arrayconsists of six wings which combine to produce a power of

over 75 kW. Figure 14 depicts one half of the qualificationwing for the ISSA solar array.

Photo courtesy of LMSC

Figure 14. ISSA Solar Array Qualification Wing

(One Half of Wing Blanket Shown)

Flexible blanket systems with polyimide blanket compositionare inherently susceptible to thermal-snap phenomenabecause of their relatively low thermal mass and materialthermal expansion/contraction mismatches. This conditionprimarily occurs at eclipse exit under immediate solarillumination incidence. Thermal snap is caused by thermalexpansion/contraction mismatches between the tensionedblanket system and its supporting deployment structure. As

the flexible blanket system is rapidly heated it experiences anabrupt thermal mismatch between the blanket anddeployment structure. The abrupt mismatch results in asignificant rate of displacement along the arrays center ofgravity. The abrupt change in position results in a dynamicimpulse to the spacecraft which must be reacted by theattitude and control system. Thermal snap and jitteringeffects caused by these mismatches occur after every eclipseand can be detrimental to certain types of spacecraft thatrequire precision pointing.

Flexible arrays are generally more mass efficient as powerrequirements increase. A qualitative plot, shown inFigure 15, depicts array specific power trends for rigid and

flexible array technologies.

TYPICAL ARRA Y MASS EFFICIENCY TRENDS

0

10

20

30

40

50

60

70

80

0 1000 2000 3000 4000 5000 6000 7000 8000 9000

Array Power (W)

Spec

ificPower

(W/kg

)

Conventional Rigid Array

Conventional Flexible Array

Figure 15. Rigid & Flexible Array

Specific Power Trends

For power systems less than ~3kW, flexible blanket arraysbecome less weight competitive because of the largeparasitic masses of their supporting deployment system andstowage container. An exception to this trend is theUltraFlex solar array, shown in Figure 16[2]. The UltraFlex


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incorporates an innovative antennae-like structure totension the flexible blanket which eliminates the need forlarge deployment mechanisms and blanket boxes. Becauseof UltraFlexs low parasitic mass fraction this systemprovides specific powers of over 100 watt/kg for lower powerarrays.


Figure 16. UltraFlex Flexible Blanket Solar Array

Solar array platforms which concentrate sunlight onto asmaller area are being seriously considered for many nextgeneration systems in an effort to significantly reduce costand mass. In some applications, the cost of photovoltaicsand their laydown onto substrates represents as much as70% of the cost of a complete array system. Designs thatsignificantly reduce the amount of active photovoltaic arearepresent a realistic approach for cost reduction. NASAbegan the development of concentrators in the early 1980swith the emphasis placed on developing a system whichwould provide cost benefits for high power arrays, such asthe International Space Station. Early concentrator workproduced designs which required very precise sun

tracking/pointing tolerances (+ 0.1

0

) and precision optics,making these systems impractical for typical applications.The result of this initial concentrator work producedNASAs Solar Dynamic Array which relies on a largeparabolic reflector to concentrate light onto a heat engine. Apicture of NASAs solar dynamic array during a ground testis shown in Figure 17.

Graphic courtesy of NASA

Figure 17. NASA Solar Dynamic Array

The U.S. Department of Defense (DOD) was also interestedin concentrator arrays from a natural and hostile threatsurvivability perspective. Their requirements drove theconcentrator design to practical configurations that weresimpler to integrate and matched the performance ofconventional rigid panel systems. Many feasible

concentrator concepts evolved from the DOD sponsoredSUPER program. Under the USAF program a reflectiveconcentrator system was developed that resembled a seriesof venetian blinds. This technology was developed intoPhase B but then canceled. Due to the designs sensitivityto local slope errors and resultant high manufacturing costderiving from the systems reflective optics this technologyhas not been taken further. Another concentrator developedin the SUPER program by TRW relied on a number of smallparabolic reflective dishes integrated within each panelsubstrate in a mini-Cassagranian optical configuration. Apicture of the TRW SUPER array is shown in Figure 18.


Figure 18. TRW Super Concentrator Solar Array

In the late 1980s and early 1990s, NASA joined BMDO todevelop an entirely new mini-dome Fresnel lens lightconcentrating array. A test module of the mini-dome systemwas produced and successfully flown as part of the AirForces PASP+ experiment, and is shown in Figure 19. Themini-dome system still required moderate alpha and betapointing/alignment tolerances (+2

0) and this deficiency led to

the development of the SCARLET (Solar Concentrator Array

with Linear Refractive Element Technology) solar arraysystem.


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Photo courtesy of NASA

Figure 19. PASP+ Mini-Dome Concentrator Experiment

The SCARLET array employs a linear Fresnel lens elementand can accommodate generous off-pointing conditions (+30alpha and +240 beta). A schematic of a SCARLET moduleassembly is shown in Figure 20.

Fresnel Lens

Glass over Silicone

Lens Panel

Gr/CE

Frame Caps,

Delrin

Bypass Diodes (5)

Flex Circuit Bondedto Carbon-Carbon

Module Base

Solar Cells (5)

Multi-bandgap

GaInP2/GaAs/Ge

Blocking Diode

Graphic courtesy of AEC-Able

Figure 20. Schematic of SCARLET Module Assembly

Unlike other concentrator systems, the off-pointingcapability of SCARLET is compliant enough toaccommodate GEO spacecraft which employ standard singleaxis tracking systems. SCARLET was flight qualified in 1985for use on the METEOR spacecraft which ultimately was

destroyed during a launch vehicle failure. Figure 21 depictsthe SCARLET concentrator array flown on the METEORspacecraft.

Photo courtesy of AEC Able

Figure 21. SCARLET Solar Array for the

METEOR Spacecraft

An advanced SCARLET system is being configured forNASAs JPL New Millennium Deep Space One (DS1)spacecraft and is scheduled to launch in mid-1998. TheSCARLET DS-1 New Millennium solar array is shown inFigure 22.

In parallel with the NASA/BMDO program, the U.S. AirForce Phillips Laboratory and NASA sponsored thedevelopment of a low concentration ratio (LCR) planarreflective solar array. This array was conceived in the mid-1980s by Hughes Aircraft Company, as an approach forproviding hardenability from radiation and man-madeenvironments. This concept was further developed by anumber of companies since that time. The LCR reflective

planar array technology is sensitive to reflective surfacetolerances. When the slope errors of the reflective surfacesare not controlled properly, illumination variations occurwhich current-limit strings and significantly reduce power.


Figure 22. New Millennium DS1

SCARLET Solar Array

The cost savings potential achievable with LCR arrays is notas significant when compared with technologies that employhigher concentration levels. A LCR planar reflective panelsolar array is slated to fly aboard NASAs Small Satellite

Technology Initiative (SSTI) Clark spacecraft. A similarLCR array concept with a concentration ratio approaching~2.5X is being developed at the Naval ResearchLaboratory[3]. A picture depicting the NRL planar reflectivepanel concentrator solar array is shown in Figure 23.

Graphic courtesy of NRL

Figure 23. NRL Reflective Planar

Concentrator Solar Array


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Other concentrator concepts being developed includespectrum splitting using holographs to achieve up to 40%efficiency, light transmission through fiber optics to ashielded cell enclosure[7,8], and large Fresnel lens refractivesystems employing inflatable structures[2]. Theseaforementioned systems are currently in the conceptualstages and far from development. As a result, it is difficult topredict their feasibility and the ultimate impact they will have

on the spacecraft system.

Deployment Systems

Growing satellite power requirements have dictated themulti-panel deployable solar arrays to deploy larger areapanels and a higher number of them. For rigid panel typearrays the deployment system is integral with the panelassemblies. In a rigid panel system, each panel is attachedto an adjacent panel with hinges and deployment springs.The deployment force of this system is driven through acomplement of primary and redundant springs along eachpanel hinge line. To provide a controlled and coordinated

deployment each panel is generally linked to its adjacentpanel with a synchronization device. The panelsynchronization devices are coupled to one another andterminated at the base hinge/spacecraft interface. The rateof deployment is governed by a damping mechanism locatedat the base mechanism. The most common deploymentsynchronization method consists of cable/pulley systems orpantographically coupled panel systems. The solar array tobe flown on the next generation GPS IIF spacecraft willemploy a pantographic structural synchronization coupling.The GPS IIF solar array is depicted in Figure 23.

Graphic courtesy of BNA

Figure 23. PUMA Pantographically Coupled Array

for GPS IIF

Some multi-panel rigid arrays rely on other techniques forcontrolling and synchronizing deployment. One techniqueemploys individual dampers located at each panel hinge linewhich independently govern and control the deploymentrate of each individual panel. The challenge of this system isto match the deployment rates of each individual hinge line,precisely, such that the array deploys in a predictable andsafe path.

Flexible blanket arrays rely on extendible boom systems fordeployment, and for providing and maintaining the correct

blanket pre-load tension. The deployment system mostcommonly used for flexible blanket systems is the coilablemast. The coilable mast deployer was used for the SAFE,Olympus, Milstar, and EOS-AM solar arrays. These systemsare extremely reliable, stow into a compact volume, havereasonable deployed stiffness, and are relatively low cost.The coilable mast system used for the EOS-AM solar array isshown in Figure 24.

For large flexible blanket solar arrays, such as theInternational Space Station, a mast system with a very highstiffness and strength capability is required. For theseapplications articulated mast deployer systems provide theneeded strength and stiffness characteristics. Thesesystems allow the structural element materials to be moreappropriately tailored to requirements. The articulated mastdeployer used for the International Space Station solar arrayis shown in Figure 25, in its stowed and deployedconfigurations.


Figure 24. Coilable Mast Deployer for the

EOS-AM Solar Array

Photos courtesy of AEC Able

Figure 25. ISSA Solar Array FASTMast Deployer


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Tubular boom deployer systems have also been used inunique applications. The CTS flexible blanket solar arrayemployed a centrally located tubular boom deployer. TheHubble Space Telescope roll-up flexible blanket array usedtwo tubular booms running the length of each blanket, asshown in Figure 12.

Inflatable boom deployers are a technology which promises

to provide systems with reduced mass and cost[14].Although these systems have yet to be successfully testedin spaceflight solar arrays, the inflatable technology isgetting considerable attention. An inflatable structure canbe composed of a B-staged pre -preg composite material thatcan be fully cured, in its deployed configuration, with amodest exposure to heat. Once completely cured, theinflatable boom becomes rigid to offer adequate deployedstrength and stiffness. These systems stow into anextremely compact launch volume, have minimal complexityand offer exceptionally low mass. The anticipated benefitsof inflatable technologies are promising, however they stillawait repeatable space demonstration and commercialviability.

Mechanisms

The major mechanisms within a solar array system areintegral spring/hinge assemblies (panel-to-panel assembliesand array-to-spacecraft assemblies), launch restraint/releasedevices, and sun-pointing orientation drive assemblies.

Most hinge assemblies are composed of pin and clevisconfigurations and include an integral spring along thehinge pin rotation axis. In some cases, flexible carpentertapes are substituted for conventional hinge devices. Thecarpenter tape approach is unique in that it combines both

hinge and spring features into one element. Flexible blanketsystems employ living flexure or pin type hinges which areintegral with the blanket assembly, or flexible carpenter tapetype hinges/springs which are laminated within the blanketassembly.

Notable developments have occurred with launchrestraint/release devices. Pyrotechnically actuated deviceshave historically been used to sever bolt, cable, and flexiblecord tiedowns. Pyrotechnique devices have had excellentreliability, but their inherent shock loading imposed on thearray upon release, and the stringent safety and qualityprovisions required for handling, have led to the

development of non-explosive actuators (NEA).

Non-explosive actuators are desirable, for some applications,to reduce the level of shock loading upon the array duringrelease. Non-explosive devices can be initiated with avariety of actuator mechanisms. Some non-explosiveactuators, such as the Starsys high-output paraffin system,includes features which allow the device to be resettableduring qualification and acceptance level testing. Theadvantages of these devices are that the units intended tobe flown are the same ones which successfully passedacceptance testing, and no refurbishing or replacement of

parts are required prior to launch. A picture of the Starsyshigh output paraffin actuator is shown in Figure 26.

Photo courtesy of Starsys

Figure 26. Starsys High Output Low Shock

Paraffin Actuator

Fokker Space Systems also developed a unique low shocklaunch restraint/release device. This system employs athermal knife which applies heat to cut through a fibrouscord tiedown. The system is space qualified and has flown

on numerous solar array systems. A picture of the Fokkerthermal knife launch restraint/release device is shown inFigure 27.

Graphics courtesy of FSS

Figure 27. Fokker Space Systems Thermal Knife Launch

Restraint/Release Device

Other NEA devices include a circumferentially burn wireinitiated device produced by G&H Technologies and quickreaction Nitonol shape memory alloy devices developed

initially by Lockheed Martin. These devices requirereplacement of components after each release. As with allNEAs, the releasing process is a slow reaction occurringover a long duration when compared with pyrotechniques.These releasing characteristics help produce very low shockloading into the array.

Orientation drive units are responsible for positioning thesolar array normal to the sun for extracting maximum power.These mechanisms consist of a one- or two-axis gimbal, anda motor/controller drive system. The drive motors areconfigured to provide sufficient torque to autonomously


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track the sun during the arrays life as well as to reach a fasterslew rate when desired. The gimbal mechanisms mustprovide the required strength and stiffness to accommodateall on-orbit array loads. Drive motors for these devices haveevolved throughout the years. Initially, DC brush motorswere employed in these systems. As mission life extendedand electronic controllers became more powerful, reliable andminiaturized, brushless DC motors were considered as an

advantageous replacement. Brushless DC motors provideenhanced controllability, longer life, and lower mechanicalnoise output, which are sometimes desirable characteristicsfor particular missions. DC stepper motors are alsocommonly used today for drive systems. The DC steppermotor incorporates simplistic drive electronics whichgenerally results in a lower system cost.

5. FUTURE SOLARARRAY TECHNOLOGIES

AND TRENDS

Historically, as solar array power requirements continued togrow, the demand for higher power systems that were lowerin cost and weight became more crucial. To meet nextgeneration industry needs solar array systems will also haveto provide power growth potential, low cost and lightweight.

Today, the transition of space from the government sector tocommercial purposes is driving solar array technologieseven further. Most of the commercial applications beingproposed today consist of high power geostationary earthorbit (GEO) based systems, and low to medium earth orbit(LEO to MEO) constellations composed of many spacecraft.Recent projections from leading GEO spacecraftmanufacturers indicate that solar array power requirements

will climb from the 8 to 10 kW ranges of today to 15 to 20 kWwithin 5 years, and up to 30 kW within the next decade[10].Meeting these new aggressive GEO applications will requireconsideration of alternative array technologies whichminimize performance/cost impacts at the spacecraft systemlevel and allow significant cost savings through economiesof scale. High concentration ratio (HCR) solar arrays appearto be the best candidate for meeting the high power GEOrequirements. The HCR arrays are able to significantlyreduce cost while maximizing power growth potential byflying high efficient solar cells more economically than othersystems. HCR arrays provide the most cost-effectiveplatform to utilize next generation photovoltaics as theycome on line. Projections from various LEO and MEO

constellations indicate that within the next decade over 500solar array systems will be required for project completion.A number of these proposed LEO and MEO applicationsemploy large production volumes and operate in highradiation environments. These applications will demandinnovative designs which deliver low system cost andemploy novel radiation hardening techniques. HCR arrayswhich can more mass efficiently provide radiation protectionand low cost, and extremely low cost thin film photovoltaicsare promising candidates for emerging LEO and MEO

missions. Cost and performance characteristics of variousarray technologies is shown in Figure 28.

10 15 20 25 30

Cell Conversion Efficiency (Absolute %)

ArrayCost

BETTER

VALUE

LOWER

VALUE

PLANAR

GaAs

PLANAR

S i

PLANAR

THIN FILM

HCR

GaAs

HCR MJ

GaAs

PLANAR

MJ GaAs

HCR Growth Path

10 $/W per % Efficiency

Planar Growth Path

188 $/W per % Efficiency

Figure 28. Cost/Performance Landscape

of Array Technologies

Concentrator systems which employ innovative cost and

mass reduction features, and can easily be adapted to aspacecraft with negligible impact, offer an attractive solutionfor meeting the cost and performance requirements of thefuture. High efficient photovoltaics will become moreprominent in future systems and will provide a much neededgrowth potential for existing arrays. Thin film photovoltaicsrepresent a cost-efficient technique for significantlyreducing solar array cost for low power and high volumeapplications.

The SCARLET concentrator solar array may represent suchan advanced HCR array technology for meeting future arrayrequirements. SCARLET is currently the most mature andpractical HCR array developed to date. It has been reported

that a SCARLET 15-kW array system will provide over a$10M cost savings when compared to an equivalentconventional rigid panel planar array[11]. Additionally,SCARLET systems provide a marginal mass savings whencompared to conventional arrays, which provides furthersavings at the spacecraft system level. Innovative systemswhich provide higher power, lower weight, and significantcost savings, such as SCARLET, will evolve solar arraytechnologies even further.

6. CONCLUSIONS

As in any vibrant technical arena solar array technologyevolution responds to the give and take of applicationrequirements. A historical review illustrates how solar cellefficiencies increased by 260%, how installed power levelshave grown by three orders of magnitude, and how a diversefamily of structural-mechanism solutions were developed.

There now exist various technical solutions tailored for aspectrum of missions, from interplanetary to low earthorbiting applications. The available technology base isbroad and proper systems optimization requires a goodunderstanding of the relevant drivers. For some missions


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like interplanetary landers absolute minimum mass andstowage volume will be paramount. For other applicationslike multi-unit commercial communications constellationsabsolute minimum cost will be critical to support a viablebusiness model.

The combined performance and low cost characteristics ofHCR systems should appeal to various high power missions

in the future. The transition of thin film photovoltaics to thespace environment will likely be attractive to the multi-spacecraft constellation applications. And the uniquerequirements of interplanetary missions ought to motivatethe industry to conceptualize even more unique systemsthan weve seen in the past.

7. REFERENCES

[1] Jet Propulsion Laboratory, California Institute ofTechnology, Solar Cell Array Design Handbook, Chapter1, October 1976.

[2] D. Allen, A Survey of Next Generation Solar Arrays,35th Aerospace Sciences Meeting & Exhibit, January 1997.

[3] M. Brown, I. Sokolsky, NRL Thin Film SolarConcentrator, 1997 Space Power Workshop.

[4] K.P. Bogus, Europes Space Photovoltaics Programme,Proceedings of the XII Space Photovoltaic Research &Technology Conference, NASA, 1994.

[5] M.J. Herriage, R.M. Kurland, C.D. Faust, E.M. Gaddy,and D.J. Keys, EOS AM-1 GaAs/Ge Flexible Blanket SolarArray, Proceedings of the 30th Intersociety Energy

Conversion Engineering Conference, ASME, 1995.

[6] R. Hill, C. Lu, J. Hartung, and J. Friefred, Current Status,Architecture, and Future Directions for the InternationalSpace Station Electric Power System, Proceedings of the30th Intersociety Energy Conversion EngineeringConference, ASME, 1995.

[7] U. Ortabasi, A Hardened Solar Concentrator System forSpace Power Generation: Photovoltaic Cavity Converter(PVCC), Space Technology 13, 1993.

[8] T. Nakamura and B. Irvin, Development of OpticalWaveguide for Survivable Solar Space Power Systems,

USAF Report PL-TR-92-3006, 1993.

[9] Capt. D.N. Keener & Dr. D. Marvin, Progress in theMultijunction Solar Cell Mantech Program, SpacePhotovoltaic Research & Technology Conference, 1997.

[10] M. McVey, Commercial Space System Practices,Presentation at the 15th Annual Space Power Workshop,1997.

[11] B.R. Spence, P.A. Jones, M.I. Eskenazi & D.M. Murphy,The SCARLET Solar Array for High Power GEO Satellites,IEEE Photovoltaic Specialis ts Conference, 1997.

[12] G.T. Crotty, P.J. Verlinden, M. Cudzinovic, R. M.Swanson, 18.3% Efficient Silicon Solar Cells for SpaceApplications, IEEE Photovoltaics Specialists Conference,1997.

[13] E.S. Fairbanks, M.T. Gates, Adaptation of Thin-FilmPhotovoltaic Technology for use in Space, IEEEPhotovoltaic Specialists Conference, 1997.

[14] M.J. ONeill, M. F. Piszczor, Inflatable Lenses forSpace Photovoltaic Concentrator Arrays, IEEE PhotovoltaicSpecialists Conference, 1997.

8. BIOGRAPHIES

P. Alan Jones:After graduation from the University of California at

Santa Barbara with a B.S.M.E. Mr. Jones performedvarious assignments for Able Engineering, includingmanaging the Advanced Photovoltaic Solar Arrayprogram, the Tethered Satellite System Strengtheningprogram, and the SUPER Solar Array program. Mr.Jones also managed the research and developmenteffort that produced four separatesolar array products for AbleEngineering. Mr. Jones holds fivepatents in the array technologyfield and has authored numeroustechnical papers. Mr. Jones nowmanages the Able Solar Array

Product Group and oversees theefforts in five direct programs aswell as IR&D and businessdevelopment matters.

Brian R. Spence:Mr. Spence received his B.S.M.E. from the University ofCalifornia, Santa Barbara, in 1986. With over 11 yearsexperience in aerospace, Mr. Spence has been involvedin the design, development and analysis of space-baseddeployable structures, mechanisms and solar arraysystems, in project engineering and managementcapacities. Mr. Spences most notable projects includethe ISSA Mobile Transporter, Mars Pathfinder

Deployable Ramp, Phillips Laboratory/U.S. Air ForceAstro -Edge solar array, NASA SSTI Astro-Edge solararray, HS702 planar reflective low concentrator solararray, and the Advanced SCARLET solar array. Mr.Spence is currently a member of thesolar array systems group at AEC-ABLE Engineering, and is primarilyinvolved in business developmentand advance IR&D activities. Mr.Spence has authored numeroustechnical papers and is a registered


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professional engineer in the state of California.

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