Inflatable Solar Arrays: Revolutionary Technology?

1999-01-2551 Inflatable Solar Arrays: Revolutionary Technology? Mark S. Grahne, David P. Cadogan ILC Dover, Inc. Copyright 1998 Society of Automotive Engineers, Inc ABSTRACT Recent technological advancements in space inflatable structures, in the areas of material rigidization and controlled deployment, have presented a new possibility to the space community with a low cost, lightweight alternative to mechanically deployed space structures. Space inflatable structures have many benefits and advantages over current mechanical systems. They are low in mass and can be packaged into small volumes, which can potentially reduce the overall program cost by reducing the launch vehicle size. Reduction in total system mass and deployment complexity can also increase system reliability. This new technology is fast becoming a reality, especially in the field of inflatable solar arrays and other applications for spacecraft components. Many solar array applications such as the Mars rover inflatable solar array, the JPL Deep Space Four (ST4) inflatable blanket solar array, and the Teledesic blanket solar array have been developed and prototypes have been built and tested. Several flight experiments are underway and will be flown in the very near future. INTRODUCTION Recently, ILC Dover has developed several methods of deploying and rigidizing inflatable structures in orbit or on planetary surfaces. These technologies enable a flexible composite laminate structure to be densely packaged for launch, deployed via inflation gas in space, and finally cured, or rigidized, in orbit (in situ) without further need of inflation pressure to maintain structural integrity. In conjunction with the recent enhancements made in the space inflatable structures technology, the advancements in the technology of flexible thin film solar cells have made the inflatable solar array very attractive to the next generation of spacecraft. The combination of these two technologies enable the further development of lightweight, low cost, and small packing volume solar arrays. In other words, spacecraft could potentially have larger and more powerful solar arrays without the undesirable mass and storage penalty of rigid array systems if these technologies were employed. This paper describes several types of inflatable solar arrays, and related applications for the technology that may benefit solar array technology in the future. A number of specific missions, which are under development, are also discussed. Recent developments in rigidization methods and controlled deployment are also presented to support the flight readiness of this technology. With the advent of new space inflatable structure technologies, many new applications have been envisioned. Applications of inflatable systems including Mars rover solar arrays, satellite blanket solar arrays, reflectarray radar antennas, communications antennas, synthetic aperture radar arrays, solar concentrators, structural members such as booms and trusses, impact attenuation devices, habitats, and sun shields, have been developed and enhanced through many pre-flight programs. Out of these applications, inflatable solar arrays are one of the most promising candidates thus far for further development to advance the state-of-the-art of satellite structures, especially in the areas of specific power and reducing system cost. 1 SPACE INFLATABLE TECHNOLOGY ENHANCEMENTS Two areas of enhancement that contributed directly to the advancement of the application of inflatable solar array technology are the development of structural materials with various rigidization methods and the development of controlled deployment mechanisms. In general, space inflatable structures are fabricated from flexible materials (thin films or coated fabrics) that are made structural by internal pressure. Environmental threats such as micrometeoroids and debris impact increase the probability of inflatable structure leakage over time. This given, in orbit rigidization of the structure is needed to extend system life. Pure inflatable structures are used only for missions with short operational life and where the

supply of make up gas does not pose a problem. Most of the applications discussed below require in situ rigidization to provide long term structural integrity. RIGIDIZABLE STRUCTURES Space rigidized structures are fabricated from flexible composite laminates that are rigidized in situ via some external influence. They can be fabricated into many different shapes such as toroids, spheres, dish structures, tubes, etc., which can be designed into various types of structures. This class of structures can be deployed in various orbits and in gravitational surface environments such as the Moon or Mars. The rigidized components are designed for typical operational lifetimes of seven to fifteen years without environmental concerns. Over the last forty years many rigidization methods have been investigated sporadically. However, with more intensive research and development in the recent years by the space inflatable industry, many reliable rigidized structural components have been produced through the use of advanced materials and design. Some of the most promising rigidization methods include: Heat Cured Thermoset Composite Laminates (Thermal Heating) Thin-walled Aluminum/polyimide Laminates Thermoplastic Composite Laminates (Passive Cooling) UV Curable Composite Laminates Foam Inflation Inflation Gas Reaction Laminates Out of these rigidization methods Thermal Heating and Thin-walled Aluminum are the most promising for space and planetary surface applications respectively. The Thermal Heating method is currently under further development at ILC Dover for flight experiment in the year 2000. The Thin-walled Aluminum System and the Thermal Heating System will be discussed in greater detail later. For further information on other methods of rigidization the reader is encouraged to consult the reference documents listed in the paper. Figure 1 - Composite Laminate Cross-Section The MLI heating blanket is designed and fabricated from various layers of vapor-deposited aluminum (VDA) Kapton, VDA Mylar, and spacers. The purpose of the MLI blanket is to keep out the harsh space environment and at the same time maintain the required curing temperature inside. The design of the MLI blanket is tailored to the specific mission environment and requirements. The support tube laminate typically consists of four separate layers: (1) The restraint layer that maintains the shape of the inflatable structure; (2) The heater assembly layer which provides the proper temperature for deployment and curing (in some designs the restraint layer and the heater layer are combined to become one assembly); (3) The rigidizable composite laminate layer, fabricated from prepreg materials such as epoxy/graphite, is the support structure when cured; And (4) The bladder layer, manufactured from black Kapton, keeps and maintains inflation pressure during deployment. The thermoplastic composite laminates, which use passive cooling as the curing method, has a similar construction. Thin-walled Aluminum Method In this approach of rigidization a laminate is fabricated from Kapton film and ductile aluminum, where the Kapton film is positioned on both sides of the aluminum as shown in Figure 2. Thermal Heating Method The composite laminate system, which consists of a thermoset matrix resin and a fiber reinforcement such as graphite, is cured or rigidized by heating. The thermoset resin hardens after being heated to a specified temperature and cure time. This rigidization method can be designed to cure from solar energy, or from the spacecraft power, or from a combination of both. The properties of the composite material are consistent with those used in today s spacecraft design. A typical composite laminate cross-section, as shown in Figure 1, consists of two multiple-layered components: (1) the MLI blanket, and (2) the support tube laminate. Figure 2 - Aluminum Laminate Cross-Section 2 To rigidize the cross-section, the structure is inflated to eliminate the wrinkles in the laminate and to a point of just yielding the aluminum. After yielding the aluminum the Kapton/aluminum laminate is rigidized and will maintain structural integrity. The inflation gas is then vented to space. A typical aluminum laminate cross-section

consists of two multiple-layered components: (1) the MLI blanket, and (2) the support tube laminate. The MLI blanket is similar to that used for thermal cured laminates, and it is tailored to mission environment and requirements. The support tube laminate is fabricated from layers of Kapton-adhesive-aluminum-adhesive- Kapton laminate. The thickness of each layer and the number of layers used can be tailored and designed to mission specific requirements. Controlled Deployment System In conjunction with the development of rigidization methods, ILC Dover has also developed a number of deployment techniques to ensure a proper and controlled deployment of inflatable structures. The purpose for the controlled deployment system is to (1) keep the deploying system within a known envelope, (2) improve structural reliability during deployment by avoiding entanglement with itself or other components, and (3) minimize shock or impulse induced to the spacecraft during deployment. Some of the most promising controlled deployment devices include: Columnation Devices Roll-up/Reverse Roll-up Devices Internal Compartmentalization Break Cords/Peel Flaps Becket Loops Out of these controlled deployment devices the Roll-up design and the Columnation design are the most promising for the solar array application. The Roll-up design is currently under further development at ILC Dover for flight experiment in the near future. The Roll-up and the Columnation Devices will be discussed next in greater detail. The reader is encouraged to consult the reference documents for more detail. Roll-up Device One method of deployment control that has seen wide application in tubes and struts is the roll-up. This approach utilizes a rolled inflatable tube with an embedded mechanism to control its rate of unrolling when inflation gas is introduced. This is similar to a common party favor that unfurls when you blow into it. There are two classes of deployment control used in roll-up devices; 1)mechanisms embedded in the tube itself, and 2) mechanisms mounted at the end of the tube. Figure 3 - Roll-up Device (with Membrane) The second class of deployment control method involves utilizing a torque mechanism at the end of the tube. An ILC Dover proprietary Torque Mechanism is located in the end cap to provide resistance to unrolling so that interim beam stiffness during deployment is achieved. This approach is simple, reliable, compact, and the roll-up is a benign packing procedure for rigidizable laminates. Columnation Devices The columnation device, Figure 4, provides a deployment method that when inflated extends axially in a straight telescopic motion with some degree of beam stiffness. The inflatable tube is axially collapsed on a short mandrel in its packed state. The top of the mandrel has a tube compression feature (seal) that applies some resistance to axial motion of the tube. Inflation gas is introduced through the mandrel into the forward end of the tube. As plug load builds in the inflatable, it overcomes the frictional resistance and advances the tube. This method places axial load in the tube wall during deployment, allowing it to behave as an inflated beam and exhibit structural stiffness. Several variations of this device, including inflatable and collapsible mandrels that allow the packed beam to fit into very small volumes, have been developed and tested with good results. Embedded mechanisms include Velcro strips mounted longitudinally on the tube s exterior, or constant force springs mounted to the tube's interior in the same fashion. Introduction of gas into the tube would separate the Velcro due to shape change during inflation, or unroll the springs. Each system provides a calculable value of resistance, which dictates the internal pressure in the tube. This in turn dictates the interim beam stiffness of the tube during deployment, prior to rigidization. Figure 4 - Columnation Device 3

INFLATABLE SOLAR ARRAYS The advantages of using inflatable systems technology in designing a large solar array are reduced stowage volume and mass, increased specific power (greater than 100 W/kg), and reduced cost over current mechanically deployed solar arrays. The inflatable solar array is particularly attractive for missions that demand high power output with launch vehicle size restrictions. ILC is working or has developed a number of different solar array systems for planetary or space use. A few of these systems are described below. ST4 INFLATABLE SOLAR ARRAY The JPL ST4 spacecraft requires high power output arrays (2 wings at 6 kw per wing) to accomplish the mission of rendezvous with the Temple 1 comet and landing to perform scientific measurements. The ST4 inflatable solar array is a 3-meter by 15-meter solar array capable of producing 6 kw of power. The array is currently under development for a shuttle flight experiment in late 2000. ILC Dover is the prime contractor to JPL for this experiment, with AEC-Able developing the array blanket and L Garde developing the instrumentation. The basic configuration of the ST4 inflatable solar array, Figure 5, consists of four major subsystems: Solar Array Blanket Structural Support Components Controlled Deployment System Inflation System Top Panel (Full length) Inflatable/Rigidizable Tube Solar Panels 9.8 ft [3.00 m] Solar Array Blanket 48 ft [14.65 m] Bottom Panel (Split) Figure 5 - ST4 Inflatable Solar Array Inflation System The configuration of the ST4 solar array is a modular split blanket style with the deployment tube located on the array centerline. When stowed, the solar array modules are accordion-folded. Currently, Sharp s rigid highefficiency silicon photovoltaic assemblies are the basic building blocks for the modular blanket solar array. Flexible thin film solar cells using amorphous silicon, copper indium gallium diselenide or other materials hold great promise to provide even lower cost and lighter Gimbal weight photovoltaic modules in similar arrays in the future. listed in this paper. Structural Support Components This subsystem includes the inflatable beam, stowage panels, plume-offset panels, launch ties, and launch tie release mechanism. The inflatable beam can be fabricated from any of the aforementioned rigidization systems. However, the thermal heating method will be utilized for the flight experiment. The purpose of the inflatable beam is to provide a deployment mechanism and support structure for the solar array. It is similar in function to the mechanical deployment truss masts currently utilized on spacecraft. The beam is located on center of the split blanket and is stowed by rolling the tube, see Figure 6. 22 Solar Panels Roll-up Tube 118.11 [3000.00 mm] 52.70 [1338.58 mm] TYP. Launch Ties Top Panel 22.00 [558.80 mm] Inflation System 14.00 [355.60 mm] 4.50 [114.30 mm] Figure 6 - ST4 Inflatable Solar Array Stowed Configuration Controlled Deployment System The ILC roll-up device with the embedded torque mechanism is the current baseline design for the ST4 inflatable solar array deployment system. When gas is introduced in the base of the tube, the tube inflates and begins to unfurl. The torque mechanism provides resistance to the unrolling action and builds pressure in the tube at a uniform and constant level. The pressure in the tube yields a tensile stress in the tube wall, which provides interim beam stiffness during deployment. The torque mechanism is attached to the tip spreader plate on the array, which deploys the array blanket. Guide wires are also attached to both spreader plates and the blanket for control of the blanket during deployment. Inflation System Inflation systems for the space inflatable structures have included compressed gas, gas generators, and sublimation of powders and liquids into gases. Each of these approaches has its advantages and disadvantages depending on the specific application. In recent years, compressed gas systems have become very low mass and highly reliable, making them the most practical option for most space inflatable applications. A compressed gas system is currently baselined for the ST4 inflatable solar array inflation system. 4

MARS ROVER INFLATABLE SOLAR ARRAY TELEDESIC INFLATABLE SOLAR ARRAY The Teledesic program envisions a constellation of 288 satellites in low earth orbit to provide high data rate communication from anywhere on earth. Original concepts for the satellite included a pair of 6 kw solar array, see Figure 7, with a cost target of $100/W in production. Inflatable technology was considered to be a leading candidate in meeting this goal and was therefore investigated by the Teledesic team. JPL is developing several advanced rover vehicles for the exploration of the Martian surface. Several concepts call for large deployable solar arrays to meet the power needs of the rover. One such concept, see Figure 9, utilizes 1.5 meter deployable wheels and an inflatable solar array to cover vast surface areas in rapid times. Figure 9 - Inflatable Mars Rover Prototype Figure 7 - Teledesic Satellite ILC Dover, under contract to Boeing, designed an inflatable structure to support a 3-meter by 10-meter rectangular satellite solar array. A full-scale prototype demonstration unit was also fabricated and used in deployment trials, see Figure 8. The Mars Rover Solar Array prototype, Figure 10, is a working full-scale inflatable solar array for rover application. This prototype is a system that would be packaged in a small volume for launch and deployed in situ to collect solar energy. The array is parasol shaped and consists of four main components: (1) the canopy, which is the membrane that carries the solar modules; (2) the inflatable torus; (3) the inflatable column; and (4) sixteen solar modules. The sixteen (16) Kapton gores (the canopy) of the solar array are tensioned by a 8.0 cm (tube) diameter by 150 cm (major) diameter 16-sided inflatable torus, and supported by a 10.0 cm diameter inflatable column. Due to the requirement for multiple deployments, the torus of the prototype is constructed from a lightweight aluminized Mylar assembled by pressure sensitive adhesive tape of the same material. The prototype inflatable column is fabricated from a laminate of Kapton and aluminized Mylar. The inflatable torus and column of the future flight unit will be constructed from thin-walled aluminum/kapton laminates or UV cured laminates. Figure 8 - Teledesic Inflatable Solar Array Due to the limited packing volume of the array, the mast system is designed with a deployment tube, whose only function is to deploy the array, and two composite rigidizable tubes to provide structural support after deployment. This design was also constrained by a requirement for limiting frontal area in any orientation to reduce drag during orbital ascent. This requirement forced the use of a three-tube system rather than a single multifunctional tube structure. The mast structure is located on the backside of the array. The deployment tube uses an inflatable columnation device for controlled deployment. The two structural tubes are z-folded into a small compact stowage volume and pulled out during deployment. Solar cells selected for the prototype array are manufactured by Iowa Thin Film Technologies Inc. These modules consist of tandem amorphous silicon (a-si) solar cell material deposited on 2 mil thick polyimide. Each solar module is made of three, 12 volt, 50 ma, solar panels connected in parallel with a blocking diode. All three panels and the blocking diode are enclosed inside 5 mil polyester encapsulant. Each module is capable of generating 12 volt and 150 ma at 1000-watts/m 2 light intensity with 5% cell efficiency. The Mars solar array prototype is therefore capable of generating approximately 20 watts of power in a terrestrial environment or 12 watts in a Martian environment. Further optimization in cell population pattern, cell efficiency and mass will further increase specific power to weight ratio. 5

SYNTHETIC APERTURE RADAR ARRAY (SAR) Figure 10 - Mars Rover Inflatable Solar Array POWER SPHERE ILC Dover together with the Aerospace Corporation developed concepts for a Power Sphere solar array, which would enable the use of micro and nano satellites. The design of the power sphere allows for constant power generation independent of the sphere s orientation. The sphere could be packaged in very small stowage volume and would deploy and inflate using sublimating powder or compressed gas. The power sphere would generate in the neighborhood of 5watss of power with a system mass of 600grams. ILC Dover fabricated a proof of concept unit for demonstration purposes, Figure 11. The SAR program was a development effort performed by ILC Dover for NASA s JPL to demonstrate the feasibility of a low mass inflatable Synthetic Aperture Radar Array Antenna. The prototype system, see Figure 12, was a flat multiple-layered metalized thin film microstrip array (a three-membrane assembly) tensioned and supported by an inflated frame that can be rigidized after deployment. The size of the full-scale array was a 10-m x 3.3-m structure. The system was stowed in a rolled-up configuration on the spacecraft bus, and was deployed in a controlled manner via inflation gas. The objective of this program was to develop a functional subscale system that was less than 2 kg/m 2 of radiating area when projected to full scale. The subscale prototype met the requirements and demonstrated the feasibility of the low mass SAR concept. Figure 12 - Synthetic Aperture Radar Array The deployment method and supporting frame concept used in the SAR could also be used for a flexible blanket type solar array in future developments. That is, the membrane assembly of the SAR could be replaced with a flexible thin film solar cell blanket and the system would then act as an inflatable solar array. JPL 3-METER KA-BAND INFLATABLE REFLECTARRAY Figure 11 Power Sphere RELATED APPLICATIONS Developments in related applications are producing more packaging and deployment concepts that are applicable for deploying future inflatable solar arrays with flexible thin-film array blankets. Investigations in these projects are pushing the inflatable technologies forward both in rigidizable structures and controlled deployment systems. In conjunction with improvements made in the inflatable technologies, further advancements in the flexible thinfilm photovoltaic cells will make these concepts prime candidates for future inflatable solar arrays. The concepts discussed below also employed many of the latest advancements in the inflatable technology. The 3-Meter Ka-Band Reflectarray program is another development effort performed by ILC Dover for JPL to demonstrate the feasibility of super low mass telecommunication and spaceborne SAR concepts. This 3-Meter prototype, see Figure 13, consists of four major subassemblies: (1) the membrane assembly; (2) the rigid frame assembly; (3) the inflatable frame assembly; and (4) the suspension system. 6

Figure 13 - JPL 3-Meter Ka-Band Inflatable Microstrip Reflectarray The reflectarray membrane, supplied by JPL and assembled by ILC, is suspended by a horseshoe shaped structure assembled from a straight rigid frame assembly and a U-shaped inflatable frame assembly. The rigid frame assembly is made from graphite epoxy with aluminum end caps. The inflatable frame assembly is fabricated from urethane coated Kevlar to simulate a rigidizable material in the final application. The membrane support frame is 25 cm in diameter and the feed horn support torus is 7.5 cm in diameter. The feed horn torus is supported by three vertical struts grouped within 90 degrees of each other. The struts are tapered in diameter, from the larger diameter of the membrane support tube to the smaller diameter of the feed horn torus, to minimize material in the RF path as well as system mass. One of the critical requirements for this program is the flatness of the membrane. For this reason many packaging and deployment concepts were considered, with the current configuration being the best for this particular application. Again, similar to the rectangular frame used for the SAR array, the horseshoe frame concept could also be used for solar array deployment. NGST (ISIS) SUNSHIELD ILC Dover is working with JPL and GSFC on the NGST (ISIS) Sunshield program. The goal of the program is to demonstrate a 1/3-scale model of the Next Generation Space Telescope (NGST) sunshield. A ½ scale model has already been demonstrated in 1G and the current program calls for the 1/3 scale model to be demonstrated on the shuttle in 2000. The sunshield is a diamond shaped membrane structure that measures approximately 5 x 11 x 1 meters when deployed. The system consists of 4 layers of membranes that are deployed and supported by 4 inflatable beams. The membranes, when stowed, are accordion-folded or Z folded into a small packing volume around the bus structure. The roll-up method for controlled deployment and the heat cured thermoset composite laminates are baselined for this demonstration. The system shown in Figure 14 is a half scale model deployed at ILC with a gravity negation system attached to the sunshield. Figure 14 - NGST Half Scale Sunshield CONCLUSIONS Inflatable solar array technology continues to mature and expand the possibilities for large-scale satellite solar arrays as well as planetary surface arrays for rover use. Technology advancements in inflatable rigidizable structures and controlled deployment systems, in conjunction with advancements in solar cell technology have lead to increases in specific power, reductions in system cost, and minimization of package volume for solar arrays. Lab and thermal vacuum chamber demonstrations of this technology have shown the viability of the approach. The flight demonstration of the ST4 solar array will provide the necessary flight heritage to allow the use of this technology for future deep space and commercial missions. CONTACT Mark S. Grahne ILC Dover, Inc. One Moonwalker Road Frederica, DE 19946 Grahnm@ilcdover.com 302.335.3911x462 302.335.1320 fax www.ilcdover.com REFERENCES 1. D.P. Cadogan, M. Grahne, and M. Mikulas, Inflatable Space Structures: A New Paradigm For Space Structure Design, IAF-98-I.1.02, 49th International Astronautical Congress, Melbourne, Australia, September 28 - October 2, 1998. 2. C. M. Satter and R. E. Freeland, Inflatable Structures Technology Applications and Requirements, AIAA 95-3737, AIAA 1995 Space Programs and Technologies Conference, Huntsville, AL, September 26-28, 1995. 7

3. D.P. Cadogan, R.W. Lingo, and C.R. Sandy, Proof- Of-Concept Demonstration Model of the Inflatable Roll-Up Synthetic Aperture Radar Array Antenna, ILC Dover Final Report for JPL Contract No. 96836, September 1997. 4. M.C. Lou, V.A. Feria, and J. Huang, Development of An Inflatable Space Synthetic Aperture Radar, AIAA 98-2103, 1998. 5. V.A. Feria, M.C. Lou, J. Huang, and S.E. Speer, Lightweight Deployable Space Radar Arrays, AIAA 98-1933, 1998 6. J. K. Lin and D. P. Cadogan, Concept Development, Design, and Fabrication of an Inflatable Solar Array for a Mars Rover Application, ILC Dover Final Report for JPL, June 1998. 7. E.S. Fairbanks and M.T. Gates, Adaptation of Thin- Film Photovoltaic Technology For Use in Space, 26th IEEE PVSC, Anaheim, California, 1997. 8. S. Guha, et. al., Recent Progress in Amorphous Silicon Alloy Leading to 13% Stable Cell Efficiency, 26th IEEE PVSC, Anaheim, California, 1997. 9. F. Jeffrey, et. al., Lightweight, Flexible, Monolithic Thin-Film Amorphous Silicon Modules on Continuous Polymer Substrates, Int. J. Solar Energy, 1996, Vol. 18, pp. 205-212. 8