SR-978

LIGHT-WEIGHT STRUCTURAL MATERIALS WITH INTEGRALRADIATION SHIELDING, THERMAL CONTROL AND ELECTRONICS

Prakash B. Joshi, Alan H. Gelb, Mark R. Malonson, Eric J. Lund, and B. David GreenPhysical Sciences Inc., Andover, MA

Edward SilvermanTRW, El Segundo, CA

Elizabeth T. ShinnAir Force Research Laboratory, Materials Laboratory, Wright-Patterson AFB, OH

Edward R. LongNASA Langley Research Center, Hampton, VA

Presented at the44th International SAMPE Symposium and Exhibition

May 23-27, 1999Long Beach, CA

Copyright © 1999 by Physical Sciences Inc.

Published by Society for the Advancement of Material and Process Engineering with permission.

Reprinted by permission from the Society for the Advancement of Material and ProcessEngineering (SAMPE). Please contact SAMPE regarding permission to copy or republish.

Copyright 1999 by Physical Sciences Inc. Published by Society for the Advancement of Material and Process Engineeringwith permission.

LIGHT-WEIGHT STRUCTURAL MATERIALS WITH INTEGRALRADIATION SHIELDING, THERMAL CONTROL AND ELECTRONICS

Prakash B. Joshi, Alan H. Gelb, Mark R. Malonson, Eric J. Lund, and B. David GreenPhysical Sciences Inc.

Andover, MA

Edward SilvermanTRW

El Segundo, CA

Elizabeth T. ShinnAir Force Research Laboratory, Materials Laboratory

Wright-Patterson AFB, OH

Edward R. LongNASA Langley Research Center

Hampton, VA

ABSTRACT

This paper describes a feasibility investigation for developing structural composite materialswith integrated active thermal control and control system electronics for space applications. Thematerials were also designed to provide mass-effective shielding against space radiation.Graphite-epoxy composites with integrated variable emissivity solid state electrochromic tilesand impregnated with materials for enhanced radiation shielding were fabricated. The tiles wereelectrically connected to a Kapton panel which was bonded to the composite and incorporatedconductor traces for electronics. Key emissivity, radiation (proton) attenuation, and mechanicalproperty measurements were made on the integrated composite structure. Limited variableemissivity performance was demonstrated in the laboratory. Proton attenuation through thematerials agreed well with design predictions. Preliminary measurements in flexure tests showhigh values of tensile strength and modulus for the integrated composite. The failure wasinitiated by cracking of the electrochromic tiles. The emissivity modulation range, mechanicalruggedness, and radiation attenuation per unit mass of the integrated composite can be improvedby using thin-film conductive polymer electrochromics currently under development.

KEY WORDS: Multifunctional Structures, Variable Emissivity Thermal Control, Light-Weight-Space Radiation Shielding

1. INTRODUCTION

1.1 Multifunctional Composite Structures In recent years the trend in spacecraft designhas been towards smaller, lighter, higher performance satellites with sophisticated payloads andinstrumentation. The satellites use smaller launch vehicles, so they need to be packaged in aweight and volume constrained design. Over the past 25 years, lightweight, composite materialshave been increasingly used to reduce the structural weight and volume of satellite components.However, concentrating solely on structural mass reduction does not lead to further loweringof spacecraft mass because the structure typically represents as little as 10 to 15% of the totalmass. Miniaturization of avionics per se further reduces mass, but not the large parasitic massassociated with avionics containers, cables, structural support of packaged avionics, orconnectors. These parasitic components can contribute as much as 50% of the mass of thespacecraft. Therefore, innovative electronics packaging concepts that are light-weight, reliable,and achieve smaller area and volume, must be developed.

One solution to the above problem is to design structural elements of a spacecraft with multiplefunctions, or multifunctional structures (MFS). It is envisioned that the subsystem functions andcomponents, such as thermal control, cabling associated with the command and data handlingsystem, and perhaps even solid state batteries, can be integrated into the load bearing structure.Printed circuits can also be laminated integrally into the structural composite face sheets orpanels. This approach, coupled with higher density electronics packaging technology, cansubstantially reduce the overall weight and volume of small satellites. The challenge is toaccomplish this integration of multiple functions/components into the spacecraft structure at anaffordable cost.

During the work presented in this paper, we investigated the integration of electrochromicdevices (tiles) and associated electronics for active thermal control into graphite-epoxy com-posites. Radiation shielding properties were incorporated by taking advantage of the attenuationthrough the electrochromic tile and the graphite composite.

1.2 Space Radiation Shielding The Earth’s radiation environment consists of Van Allenradiation belts composed of protons and electrons. These charged particles span a widespectrum of energies, from tens of KeV to tens of MeV for electrons and from hundreds of keVto hundreds of MeV for protons. The fluxes of these particles also vary greatly depending uponaltitude, latitude, and solar activity. The charged particles, especially protons, are also producedin great quantities during solar flare events. High energy protons and electrons penetratespacecraft materials and systems, depositing some or all of their energy, as the craft orbitsaround the Earth or traverses interplanetary space. This energy deposition can cause detrimentalshort-term and long-term effects on space system components - in particular, electronics.

Space electronics systems employ enclosures to shield sensitive components from spaceradiation. The purpose of shielding is to attenuate the energy of charged particles as they passthrough the shield material, such that the energy per unit mass (or dose) absorbed in silicon issufficiently below the maximum dose ratings of electronic components. The standard practicein space hardware is the use of aluminum as both a radiation shield and structural enclosure.Depending on mission altitude and inclination, and the dose rating of electronics, the thicknessof aluminum necessary for shielding can substantially exceed that required for structuralstrength, resulting in significant weight penalties.

To reduce the structural weight, satellite designers use composite materials which have higherstrength-to-weight ratios than aluminum. However, conventional graphite epoxy composites

are not as efficient shielding materials as aluminum because of their lower density, that is, forthe same mass, composites provide 30 to 40% less radiation attenuation than aluminum.Conversely, for the same radiation attenuation, the composites tend to be 30 to 40% thicker thanaluminum.

1.3 Space Systems Thermal Control All spacecraft components have prescribed operationaltemperature ranges. The temperature ranges can be rather wide, -25 to 60�C for passiveelectronics, or rather narrow, 5 to 40�C for batteries, depending on the component. Theobjective of thermal control is to maintain the temperature distribution in the interior of asystem, typically electronics, within certain limits to insure operation, survival, and long-termreliability. The designer selects coatings for the exposed surfaces with desired opticalproperties, resistance heaters, heat pipes to conduct away the heat, conductive couplings atmounting interfaces, and radiators to reject heat to space. Temperatures of exposed surfaces ofthe spacecraft and the heat rejected by radiator surfaces can be controlled using variableemissivity materials. Active thermal control with variable emissivity coatings also eliminatesthe need for bulky and heavy mechanical louvers for emissivity control, thus saving weight andvolume of spacecraft systems.

By reducing temperature swings, the variable emissivity coatings allow the use of commercialgrade components (which have a smaller operating temperature range) rather than military gradeparts in payloads. The severity of space qualification thermal testing which stresses andweakens the components prior to flight is substantially reduced. The result is reduction indevelopment costs, improved longevity and higher reliability of on-orbit operations.

1.4 Electrochromic Materials We describe briefly the basic features of electrochromic (EC)materials and devices which make them attractive for thermal control applications. ECs possessthe remarkable ability to change their optical properties (e.g., transmissivity) under the actionof a voltage pulse. This change is reversible, i.e., the direction of change can be reversed byreversing voltage polarity. Furthermore, the optical property change is accomplished in a"pulse-and-latch" manner, that is, the voltage (typically less than 5 V) needs to be appliedtransiently to change the state of the material and continuous application of potential to maintainthat state is not necessary. The power required for the change of state is minimal, on the orderof a few milliwatts, depending upon the time interval over which the change is desired.Normally, time intervals on the order of hundreds of seconds will be adequate for spacecraftthermal control applications.

There are several types of EC materials. We concern ourselves here with a tungsten oxide-basedEC, the so-called solid state EC (SSEC). An EC "material" in actuality consists of layers ofseveral materials. However, the optical properties are dominated by a single layer, e.g., tungstenoxide (WO3), in the example to follow. The EC layers can be created by several techniques suchas evaporation, sputter-deposition, etc. The total thickness of EC layers is typically on the orderof 1 to 2 µm. A schematic of a WO3-based EC device is shown in Figure 1.

The EC device consists of five layers, shown schematically in the figure. The outer layers, TC1

and TC2 (typically, indium tin oxide), are transparent electron conductors and act as the cellelectrodes. The IC layer (such as LiNbO3) is an ionic conductor and electronic insulator. It isthe cell electrolyte. In some cases, the electrolyte can be a polymer. The EC2 layer (such asLiCoO2) stores the ions for doping the electrochromically active layer, EC1, composed of WO3.The cell functions as follows. A small voltage difference is imposed between the TC1 and TC2

layers. Current flows via ionic conduction through the cell. The electrolyte layer, IC, composedof LiNbO3, transports lithium ions between EC1 and EC2. The EC2 layer is a layered oxide suchas LiCoO2 which supplies the lithium ions. It is the injection and ejection of lithium ions to and

Figure 1. Basic features of an EC device with typical materials [courtesy Reference (1)].

from the EC1 layer that varies the optical properties. The EC1 layer is composed of WO3. In itsunlithiated state, it is highly transmitting at visible and infrared wavelengths. In its lithiatedstate, as LixWO3, it is highly reflecting at these wavelengths. Thus, by adjusting the trans-missivity of the EC cell against a high emittance background, one can construct a variableemissivity devive.

In our work, we employed tungsten oxide based ECs sandwiched between glass and germaniumsubstrates, with the germanium (which is transparent in LWIR) coated with low solarabsorptivity coatings facing space. The EC devices are electrically simple, with two electrodesacross which bias voltage is applied. In an active thermal control system comprising numerousEC devices, an intelligent electronic system will be required to sense temperature distributionon a surface and control voltages on individual EC devices. The electronics must be miniatureand integrable with the composite structure.

1.5 Electronics Embedding in Composite Structures Miniaturization and compact packag-ing of electronics is key to light-weight space avionics systems. Technologies such as ASICs,High Density Interconnects, hybridization, etc. have been developing at a rapid pace in recentyears. Miniaturization can be prohibitively expensive unless high production volumes areanticipated. One of the growing trends in modern electronics, both commercial and space is touse flexible circuit boards instead of the rigid G10 or similar materials. This allows the “board”to be flexed and shaped as desired to conform with other hardware. Flexible “ribbons” cables,accordian-like stacking of stiff boards connected by flex panels are quite common in electronicstoday. While embedding conductor traces between layers of flexible material such as kaptoncan be readily done, mounting of electronic components onto a flex panel is not, especially ona multilayer kapton flex board. For the work described in this paper, we selected to embedconductor traces within a kapton epoxy.

2. COMPOSITE MATERIAL DESIGN

2.1 Radiation Shielding Design: General Considerations This paper concerns the tech-nology of composite materials with integral space radiation shielding, active thermal controlelements, and their control electronics. Given the current trend toward commercial off-the-shelf(COTS) components, the orbital lifetime of space systems is substantially reduced since thesecomponents are not designed for high radiation tolerance. Therefore, systems with COTSelectronics require enclosures whose mass is determined by substantial shielding requirements,rather than structural requirements. Thus, significant weight penalties can result when shieldingspace systems that use COTS electronics and operate in high radiation environments. Ourshielding technology is particularly effective in minimizing these weight penalties.

Another attractive feature of our shielding techniques is its multifunctionality. The samemetallic components embedded in composites for radiation attentuation can also provide greaterthermal and electrical conductivity to the materials, thus providing thermal control and EMIshielding capabilities as well as reducing effecs of space systems charging on orbit. Thematerials can also readily incorporate surface treatments necessary for protecting against thespace environment; for example, atomic oxygen and solar ultraviolet radiation.

The processes of how charged particles interact with matter and lose their energy during passagethrough materials has been described by the classical theory of Bethe (2). From the theory it canbe shown that to minimize mass of the shield, low atomic number (low-z) elements are mosteffective on per unit mass basis. If it is necessary to reduce the thickness of the shield, then highatomic number (high-z) elements are most effective on per unit thickness basis. The recentpractice of replacing aluminum with composites for reducing structural mass fraction ofspacecraft offers the possibility of taking advantage of the composites’ inherently low-zcomposition (hydrogen, carbon, and other elements) for mass-efficient radiation shielding.Therefore, the key aspect of our shielding materials is to select the formulation of the matrix(e.g., epoxy in graphite [carbon fiber]-epoxy composites) such that low-z components aremaximized.

Depending on the severity of the charged particle environment, the dose rating of theelectronics, and the required degree of attenuation, the thickness of a purely graphite-epoxyshield can be 40% greater than aluminum. In some applications, this increase in thickness maynot be acceptable. Our shielding technique reduces the thickness by embedding high-z materialsinto the epoxy formulation. The physical form and atomic number range of the high-z materialis determined by the desired degree of attenuation and shield mass reduction relative toaluminum. The shielding design approach allows the composite materials to be tailored to thegiven charged environment and internal dose for the electronics. The design involves tradeoffsamong shield mass, thickness, internal dose, and mechanical and other properties. Severalformulations of graphite-epoxy composites were designed and fabricated into plates equivalentto 125 mils of aluminum in radiation attenuation. The designs were based on models developedat Physical Sciences Inc. (PSI) for attenuation of proton and electron energies as the particlestraversed and interacted with the shield materials.

The epoxy formulation used in graphite composites consists of a low outgassing, high-hydrogen-content resin with very low viscosity to allow higher possible loading of embeddingmaterials. Graphite fibers were introduced into the resin, 40% fibers to 60% epoxy by volume.Mid- to high-z elements or compounds were added to the epoxy. Percent loading is defined asthe weight of additive to the total weight of the additive plus epoxy mixture. Several 15.2 x15.2 cm plates of graphite-epoxy shields were fabricated.

2.2 Mechanical and Thermal Control Design We now discuss the mechanical and thermalcontrol design approach. For a structure concept, we selected a rectangular rib stiffened paneldue to its simplicity, ease of fabrication, and low cost. The rectangular grid was also compatiblewith the square EC tiles which were selected subsequently for thermal control due to their readyavailability. To develop an integrated material concept with the above attributes, we startedwith the graphite-epoxy composite which was then impregnated with appropriate additives toenhance radiation shielding as described earlier.

To impart variable emissivity to the space facing surface of this material, we seleted solid stateEC (SSEC) tiles being developed by EIC Laboratories, Norwood, MA. We envisioned theelectronics to control the tiles to be located on the back surface of the core composite, withelectrical conductors passing through holes drilled into the composite. The conductor traces

(a) Square rib-stiffenedstructure

(b) Integrated material

connecting various electronics are encapsulated within a thin kapton layer in our concept.Surface-mount electronic components are soldered to this flexible kapton panel and the panelis adhesively bonded to the back surface of the composite. The rib stiffened panel and theintegrated material concept are illustrated in Figure 2.

Figure 2. Integrated material and rib-stiffened structure concept.

The material concept shows one method of enhancing radiation shielding, in this case anexpanded copper mesh. The mesh allows the epoxy to seep through, providing good bondingfor shear strength. The SSEC tiles, as described later, incorporate fairly thick germanium andglass substrates which contribute to additional radiation shielding. The aluminum ribs providestiffness as well as radiation shielding between adjacent tiles. These multifunctional aspects ofour material/structure design concept are listed in Table 1.

Table 1. Functional Contributions of Various Layers/Components of Integrate Material

Layer/Component

Thick-ness*(mm)

Function

StructuralStrength/Stiffness

RadiationAttenuation

Thermal/Electrical

ConductivityEmissivity

ControlElectrical

Connectivity

& Graphite-epoxy

6 Yes Yes, mass-efficiency

In-plane - Yes, throughholes

& Cu meshlayer(s)

0.075 Bondingthrough openareas for shearstrength

Protons,electron

In-plane, cross-thicknessconductivitypossible withproper design

- -

& Kapton flexcircuit panelwith pins

0.25 - - - - Yes

& Aluminumribs

1.60 Stiffness Yes,betweentiles

& EC tiles 1.60 Yes, Ge andglass

- Yes

& Electricalcomponents

- - - Yes

*Total thickness 7.92 mm selected to transmit <5 krad dose in 2000 x 500 km orbit.

Figure 3. Normal Reflectance range of a variableemissivity device fabricated with polymer lithiumion conducting electrolyte (EIC Laboratoriesdata).

Figure 4. Design concept for active thermal control of ECs.

An example of the infrared reflectancerange of an SSEC variable emittancedevice is shown in Figure 3. In the 8 to12 µm range, the reflectance variesfrom ~15 to 45% for a switchingvoltage of ±1 V. Since transmittanceequals (1-reflectance), the data inFigure 3 would indicate emissivitychanges of about 0.85 to 0.6 inswitching from -1 V to +1 V.

2.3 Control System ElectronicsDesign The design concept for theactive control of ECs is shown inFigure 4. The operation of the controlsystem is quite simple. It incorporatesa microprocessor which receives acommand from the spacecraft to set asurface to a spectfic tempearature. Knowing the background (solar illumination, sun angle,radiative envrionment), the processor calculates the emissivity that the surface must attain.From the emissivity versus voltage database, and knowledge of the temperture from sensorsincorporated onto the surface, the processor then calculates the voltage that must be applied toindividual EC tiles. Via swiches, these control voltages are then applied to the ECs. Thecontroller continuously monitors the current drawn by each tile and its temperature. When thedesired emissivity state is achieved, the current drawn by the EC will drop significantly, and atpoint, the controller will “switch off” or remove the applied voltages, provided the surface hasattained the intended temperature. If not, the voltage is adjusted again to reach a new emissivityvalue, and the process repeats until the surface attains the commanded temperature.

The above control system design concept was applied to a 3 x 3, nine-tile configration. In actualfabrication, only the center tile was active. The other eight tiles were inactive, a choice dictatedby funding constraints.

2.4 Radiation Shielding Design of Integrated Composite As mentioned earlier and shownin Table 1, the EC tiles contribute to radiation shielding, in addition to the graphite-epoxycomposite. We also incorporated a copper mesh to demonstrate shielding design and fabricationtechnique in case SSEC tiles are not used in favor of other ECs which do not offer any shieldingthemselves, e.g., thin-film conductive polymer electrochromics (CPECs). The germanium(0.5 mm) and glass (1.1 mm) layers in the SSEC contribute most to shielding. The contributionsof other layers are negligible.

Figure 5. Transmitted proton doses through multifunctional composite and aluminum of thesame areal densities for 2000 × 500 km, 70 deg inclination orbit.

We applied our radiation shielding design methodology to the composite in Figure 2. Weselected a representative 2000 × 500 km elliptical orbit with 70 deg inclination for evaluatingproton shielding performance. The results obtained using PSI’s unidirectional flux protontransmission model are shown in Figure 5 in terms of transmitted annual dose (krad/yr) throughthe shield versus thickness of the graphite epoxy layer of the composite structure.

Calculations in Figure 5 show that for a graphite epoxy thickness of 0.6 cm (see Table 1), thetransmitted dose through the composite is lower by a small amount (about 5%) than aluminumshield of the same areal density (g/cm2). This reduction of transmitted proton dose through thecomposite over aluminum is small due to the presence of high density germanium and glasssubstrates in the electrochemical tile. This limitation is imposed by tungsten-oxide based ECs.The thin-film CPECs do not need the germanium window and the glass substrate. Therefore,the multifunctional structures employing CPEC for thermal control will be significantly moremass-efficient radiation shields than structures employing tungsten-oxide ECs.

We performed similar calculations for an electron rich orbit, 36,000 x 620 km, 7 deg inclination.Our conclusion from these calculations is that multifunctional structures made with tungsten-oxide based EC tiles are equal or slightly better radiation shields than aluminum of the samemass. Of course, the aluminum does not offer the integrated thermal coåntrol capability that themultifunctional composite offers.

3. COMPOSITE STRUCTURE FABRICATION AND ASSEMBLY

Figure 6 shows the mechanical drawing of the integrated multifunctional composite structurebased on the design concept of Figure 2. The structure incorporates nine EC tiles with thecenter tile active and the surroundg eight non-active. This panel measured 15.2 x 15.2 cm,although the actual fabricated panel measured 20.3 x 20.3 cm, with 2.54 cm of graphite epoxyextension all around the EC tiles for mounting fixtures, etc.

Each tile has two leads across which bias voltage is applied. Underneath each tile, bonded toslots in the top surface of the compsosite, a temperature sensor in placed. The sensor has twoleads. Thus, each tile location has four leads. A kapton flexible circuit panel which

Figure 6. Integrated multifunctional structure assembly.

encapsulates conductor traces for the electronics, is bonded to the back side of the composite.The flex ciruit incorporated a simplified version of the block diagram shown in Figure 4 forcontrolling the single active EC tile.

The 20.3 x 20.3 cm graphite-epoxy panel with embedded copper mesh midway through thethickness was fabricated by Textron. Holes were drilled into the finished composite to passelectrical connections from the EC tiles to electronics. The kapton flex panel was fabricated byMinco. We assembled the multifunctional structure by integrating the EC tiles and the kaptonflex circuit panel on to the graphite epoxy structural panel, Figure 6. The pins of the flex circuitpanel passed through mating holes in the composite structural panel. The flex was bonded tothe panel using a low outgassing pressure sensitive adhesive. The holes in the compositesurrounding the pins were filled with a low outgassing, thermally conductive, electricallyinsulating epoxy. Next, nine thin-film resistance temperature detectors (RTDs) were bondedinto slots milled out on the front side of the composite panel and the leads from these RTD weresoldered to the appropriate pins of the flex circuit. These RTDs were located behind each of theEC tiles in the final assembly and used to monitor the tile temperatures. We then bonded thenine EC tiles to the front surface of the panel and soldered the electrode leads to the appropriatepins from the flex circuit. The EC tiles and RTDs were bonded to the composite panel usingthermally conductive epoxy. Finally, the aluminum rib structure was bonded to the front surfaceto fill the areas between the tiles. This rib structure had sections milled out to accommodate thepins from the flex circuit. The ribs were designed to provide added stiffness to the panel andradiation shielding between tiles equal to the shielding available through the tiles.

Figure 7 shows the photographs of the front and back surfaces of the assembled panel. As partof the assembly, we did not install electronic components on the kapton flex circuit panel.Rather, we used the pads provided at the edge of the kapton panel, as shown in Figure 7(b), toconnect external electronics to operate the active EC tile and to measure temperatures of alltiles. This approach was considered sufficient for the proof-of-concept demonstration purposes.

(a) Space facing side with EC tiles (b) Back side with kapton flex circuit panel

Figure 7. Photographs of assembled integrated multifunctional structural panel.

Figure 8. Infrared camera test setup.

4. INTEGRATED MULTIFUNCTIONAL MATERIALPERFORMANCE MEASUREMENTS

4.1 Emissivity Performance Measurements We subjected the integrated panel to a varietyof tests to evaluate its thermal control (variable emissivity) and radiation shielding performanceand its mechanical strength. An IR camera test was performed to measure the performance ofthe functional EC tile in the assembled system, in terms of emissivity variation versus voltage.We also measured the spatial uniformity of the tile’s emissivity as a function of EC state, andqualitatively determined the EC switching time. The test setup is shown schematically inFigure 8. Wire leads were soldered to the pads on the back of the flex circuit to make con-nections to the RTDs and EC electrodes. On the back side of the panel, we mounted a 15.2 x15.2 cm silicone heating pad, which had a maximum heat dissipation of 0.775 W/cm2. The padwas sandwiched between the panel and a thin aluminum plate, and was held on by two clampsalong the edge of the panel. The power dissipation of the heater pad was controlled via avariable transformer. The leads from the nine RTDs in the panel were connected to a digitalreadout through a 10-position rotary switch, so that each RTD temperature could be read in turn.

Figure 9. (a) IR (8 to 12 µm) image of integrated EC panel at high emissivity;(b) IR (8 to 12 µm) image of integrated EC panel at low emissivity.

The EC electrode leads were connected to a potentiostat, which set the EC voltage andmonitored the EC current.

The panel was mounted vertically on the bench top, and an imaging microbolometer infraredcamera was be pointed at it. This camera measures infrared radiance in the range of 8 to12 microns of wavelength. It has a measurement precision of about 32 µW/cm2. Because thetiles (both active and non-active) are moderately reflective in the infrared and are highlyspecular, we placed a blackened copper plate of known temperature in front of the panel at thespecular reflectance point (relative to the camera). This plate gave us a known, constantreflected IR signal which we could correct for in the data. Next to the panel, we placed acalibrated blackbody of known high emissivity, which we used to calibrate the radiancemeasurement of the camera.

The panel was heated to approximately 50�C, and the EC voltage was set to +2.5 V to put theEC tile into its high emissivity state. The blackbody was also heated to 50 C. We waited forthe temperature of the panel to stabilize and for the current draw of the EC to drop to a very lowlevel (<20 microamps). Images of both the panel and the calibrated blackbody were capturedwith the camera, and recorded the temperatures of the nine RTDs in the panel. Then, weswitched the voltage on the EC to -2.5 V to place it in its low emissivity state. We again waitedfor the EC current draw to drop to <20 microamps. We once again captured images of the paneland the blackbody, and recorded the RTD temperatures. It took about 1 hour for the current ofthe EC to drop to less than 20 microamps after the voltage was switched from +2.5V to -2.5V;however, the most of the change in radiance of the tile appeared to occur within the first 15 min,as determined qualitatively by viewing the image. The captured panel images are shown inFigures 9(a) and 9(b).

We reduced the data from the images, and, based on the average radiance measurements of thecenter (active) tile, we solved for the tile emissivity at the extreme voltage states. The resultswere an average emissivity of 0.77 at +2.5 V and 0.66 at -2.5 V. While the radiance of thecenter tile in Figure 9(a) (high emissivity state) is very uniform, the radiance of the tile inFigure 9(b) (low emissivity state) is not very uniform. We analyzed the data in the brightest anddarkest areas of this second image (low emissivity state) and found that the emissivity of the tilevaries from approximately 0.59 to 0.74. We also solved for the emissivity of the inactive tiles,and found it to be 0.57, which is close to the theoretical value (based on the refractive index ofGe and the emissivity of glass) of 0.59.

Figure 10. Experimental setup for protontransmission tests.

Figure 11. Measured proton transmission through multifunctional panel.

4.2 Proton Transmission Measurements With the panel designed for a proton rich orbitalenvironment, it was appropriate to perform shielding effectiveness tests in a proton beam. Theproton shielding effectiveness tests were performed at Harvard University’s CyclotronLaboratory in Cambridge, MA. This facility delivers precisely controlled proton fluxes on theorder of 108 p/cm2/s in a 1.9 cm diameter beam. The energy of the beam can be controlled bythe use of lexan attenuators to specific energies from 29 MeV to 148 MeV. For each protonenergy, we delivered a known fluence(approximately 3 x 108 p/cm2) of protonsto the panel, and measured the transmitteddose with a PIN diode dosimeter locatedbehind the panel. A photograph of thepanel mounted in the proton beam facilityis shown in Figure 10. We measured thedose transmitted through four locations onthe panel: at the centers of two of the non-functioning tiles, at the center of the ECtile, and at the corner junction betweenfour of the tiles (i.e., through the aluminumrib). Since the shielding effectiveness ofthe panel was theoretically equivalent to0.56 cm of aluminum, we also tested analuminum plate of that thickness forcomparison.

The results of these tests are shown in Figure 11. The plot shows the measurements of trans-mitted dose versus energy (normalized to a fluence of 6 x 109 protons/cm2) at various locationson the panel. Comparison with transmitted dose through a 0.56 cm aluminum plate (equal tothe integrated composite in g/cm2) and the model predictions for aluminum are also shown. Thepanel data matches the aluminum data and model prediction very well. The data from the centerof the non-functioning and EC tiles appears to be shifted to slightly higher energies comparedto aluminum, while the data from the tile junction (the rib) is shifted to slightly lower energies.This means that the tiles are providing shielding that is somewhat more effective than the

Figure 12. Flexure test setup.

Figure 13. Stress versus strain for multifunctional composite with integrated EC tiles.

0.56 cm aluminum (i.e., equivalent to a thicker aluminum), while the rib location is providingless effective shielding. This result is not surprising, since the rib density is somewhat lowerthan the tile density, and since there were slight gaps between the ribs and the tiles which mayhave allowed some proton penetration.

4.3 Mechanical Property Measurements A flexure test was performed on a multifunctionalmaterial sample (3 x 1 tiles) to predict the strength and modulus of the integrated panel. A20.3 x 5.7 cm piece of base composite material (the same material used in the multifunctionalpanel in Figure 7) was prepared, and three dummy tiles of glass/germanium were bonded to itusing epoxy. The spacing of the tiles was identical to the spacing on the actual multifunctionalpanel. The test specimen was placed on the flexure test apparatus, supported at the ends at aspan distance of 17.8 cm. It was loaded from the center of the back side of the specimen(opposite the center tile), and deflection of the surface of the center tile was measured using alinear variable differential transformer (LVDT). Load versus deflection was recorded on a chartrecorder. Stress and strain were calculated from the measured load and deflection. The testsetup is shown schematically in Figure 12. The resulting stress strain curve is shown inFigure 13.

The maximum composite material stress reached 31.5 MPa, and the strain in the tile reached0.000635 mm/mm, before minor cracks in the germanium tile occurred. The stress was70.3 MPa and the strain was 0.00127 mm/mm before the first major crack occurred. The testwas run up to a maximum stress of 480.6 MPa (360 kg load) and strain of 0.0139 mm/mm(9.45 mm center deflection). The only damage that occurred was cracking, and some flaking,of the germanium tiles. Throughout the test, the tiles remained bonded to the composite, andthe base composite material remained intact.

As is apparent from shape of the curve, the tiles initially add some stiffness to the structure untilthe point that they begin to crack, at which point the overall stiffness is reduced. Note that theslope of the stress-strain curve is steeper before the major cracking occurred than it is after. Theflexural modulus, based on the slope of the curve after the first cracks, is 34.5 MPa. (Thismodulus is based on strain calculated as above, using the overall thickness of the material.However, once the tiles crack, the strain in the composite should be calculated based on thethickness of the composite alone. Using that strain calculation, the modulus is 45.5 MPa.) Theyield stress of the material, where the first cracking event occurred, was 31.5 MPa.

5. SUMMARY OF RESULTS

The development of light-weight structural materials with integrated thermal control, electronicsand radiation shielding was investigated. We designed, fabricated, and measured key per-formance parameters of a graphite-epoxy structural composite with integrated EC devices andkapton-encapsulated electronic circuitry. Variable emissivity properties of the integratedthermal control structure were demonstrated in the laboratory. Improvements over the measuredemissivity modulation range and the minimum emissivity value are expected with furtherdevelopments in EC technologies. Preliminary measurements in flexure tests show high valuesof tensile strength and modulus for the integrated composite. The failure is initiated due tocracking of the germanium substrate of the solid state EC tiles. Thin-film CPECs, currentlyunder development, are mechanically rugged and will not be subject to such failure.Measurements of proton transmission through the integrated composite verified our shieldingdesign methodology for embedded materials. The integrated composite was found to provideproton shielding effectiveness per unit mass slightly better than aluminum. The weight savingsover aluminum will increase substantially as lighter weight EC devices become available.

Several technical issues have been identified during our work and are summarized in Table 2along with the potential solution approaches and recommendations.

Table 2. Technical Issues and Solution Approaches

Technical Issue Solution Approach (Recommendations)

& Emissivity modulation range and minimumemissivity of ECs— û0 of 0.15 and 0min of 0.6 was achieved.

Need 0min <0.3, û0 > 0.6.

Use all solid-state EC materials with mostrecent advances. Substitute with CPECswhen developed to sufficient maturity.

& Uniformity of emissivity across entiresurface of EC device

Same as above

& Light weighting:— Thick germanium and glass substrates

are used in available solid state ECdevices

Use next generation of solid state EC devicesemploying thinner glass substrate and siliconsubstrate or, use thin-film CPECs.

& Mechanical fragility of solid state EC devices Same as above, plus better bonding of devicesto the structure with compliant adhesives

& Electronics for fully autonomous control ofseveral EC devices— Preliminary design of electronics was

developed to control nine devices, butimplemented for one device to showfeasibility. Control electronics wasexternal to the panel.

Develop complete control electronics tooperate all nine devices. Embed electroniccomponents onto kapton flex panel, bonded tostructural panel.

& Database on mechanical, thermal, and otherproperties of integrated composites forstructural/ thermal design and spacequalification— Only limited mechanical property data

has been obtained

Measure strength/stiffness in tension,compression and shear before and afterthermal/vacuum cycling. Obtain data onthermal/electrical conductivity, outgassing,electron transmission

6. REFERENCES

1. Grangvist, C-G., “Electrochromic Tungsten-Oxide-Based Thin Films: Physics, Chemistry,and Technology, Mechanical and Dielectric Properties of Thin Films and Smart Materials,”Phys. Rev. Letters, 301-369 (1993).

2. Knowles, G.F., Radiation Detection and Measurement, J. Wiley, 1989.

7. ACKNOWLEDGEMENTS

This work was supported by NASA Langley Research Center under Contract NAS1-98046 withMr. Edward Long as the Technical Monitor and by the Air Force Research Laboratory,Materials Laboratory, WPFAB, Dayton, under Contract F33615-97-C-5014 with Ms. ElizabethT. Shinn as the Technical Monitor.