Enabling Future Spacecraft Missions through Isothermal Bus Thermal ...

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Enabling Future Spacecraft Missions Through Isothermal Bus Thermal Management Derek W. Hengeveld1 LoadPath, Albuquerque, NM 87106 Andrew D. Williams2 U.S. Air Force Research Laboratory, Kirtland Air Force Base, New Mexico 87117 and Brent S. Taft3 U.S. Air Force Research Laboratory, Kirtland Air Force Base, New Mexico 87117 The Isothermal Bus architecture is a novel idea focused on approaching a single thermal node S/C representation. In effect, heat is shared efficiently between cold and hot components and in combination with thermal control, spatial and temporal variations in temperature are minimized. The Isothermal Bus can enable higher capability systems along with providing schedule and cost savings. A nominal case study was provided that compared a traditional TCS to the Isothermal Bus concept with a 7:1 switching ratio. While the traditional TCS could accommodate 300 W of bus power, it would require survival heaters. However, the Isothermal Bus concept would enable a 167% increase in bus power while almost eliminating the need for survival heat.



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Assembly, Integration, and Test Active Thermal Tile High Thermal Conductance Integrated Thermal Energy Management System Low Earth Orbit Multilayer Insulation Plug-and-play Outgoing Longwave Radiation Spacecraft Satellite Modular and Reconfigurable Thermal System Thermal Control Panel Thermal Control Subsystem Variable Heat Transfer

II. Introduction

PACE exploitation provides tremendous opportunities. Since 1957, spacecraft (S/C) have been developed to take advantage of this new high ground by providing communication, scientific observation, weather monitoring, navigation, remote sensing, surveillance, and data-relay services.1 However, space presents extraordinary challenges. Consequently, S/C have become exceedingly complex and costly. Current S/C can take 1

Senior Engineer, AIAA Senior Member. Senior Mechanical Engineer, Space Vehicles Directorate, AIAA Senior Member. 3 Mechanical Engineer, Space Vehicles Directorate, AIAA Member. 1 American Institute of Aeronautics and Astronautics 2

from 3 to 7 years to deploy2 and cost from millions to billions of dollars.3 Increased complexity and cost can be traced to the growing importance and expectations of space. In addition, high launch costs push manufacturers to extend design life to reduce life-cycle costs.4 However, extended design life is accomplished through added redundancy that in turn increases complexity and cost. More flexible and even cheaper S/C can be realized through robust design approaches. Robust S/C are designed to meet a broad range of mission requirements (i.e. components, technologies, environments); consequently, they drastically reduce non-recurring engineering costs and greatly diminish design, development, and Assembly, Integration, and Test (AI&T) schedules. This is in stark contrast to traditional designs, which are optimized and intended for a specific set of requirements. In recent years, there has been several efforts to develop and demonstrate robust satellite architectures to reduce schedule and total cost while maintaining acceptable levels of performance, reliability, and lifetime. Examples include TacSat-1 through 5,5 PnP Sat,6 HexPak,7, 8 SMARTBus,9 MightySat,10 and CubeSats.11

TacSat-212 Figure 1. Example robust S/C architectures.

Plug-and-play Satellite (PnPSat)6

Robust S/C are required to handle a wide variety of missions/scenarios. As a result, the underlying thermal control subsystem (TCS) must be robust enough to maintain appropriate thermal requirements for satellite components under the most severe (hot and cold) environments. Achieving the advantages of a robust design will require a paradigm shift in how the space community presently operates. Of the numerous challenges, the S/C thermal control subsystem (TCS) presents a significant challenge because robust architectures require independence of functionality (i.e. modularity), well-defined interfaces, standardization, and flexibility.13 The traditional TCS approach is a fixed, customized, tightly-coupled system and generally does not yield itself easily to robust architectures; the problem is that the system cannot be de-coupled to the level required for a robust architecture without significant system overhead. Several works have focused on determining the best architecture for robust thermal control. Williams investigated three different architectures including high thermal conductance (HTC), thermally isolated, and variable heat transfer (VHT) architectures.14 Work by Young advanced the previous effort to include traditional, thermally isolated, HTC without VHT, modular HTC without VHT, thermally isolated with VHT, traditional with VHT, HTC with VHT, and modular HTC with VHT architectures.15 Results of this work showed that a modular HTC with VHT bus provided superior performance with respect to temperature gradient, thermal stability, thermal control, power savings, and dynamic range; this approach is the basis for the Isothermal Bus concept.

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III. Isothermal Bus Thermal Control Subsystem The Isothermal Bus architecture is a novel idea focused on approaching a single thermal node S/C representation. In effect, heat is shared efficiently between cold and hot components and in combination with thermal control, spatial and temporal variations in temperature are minimized. The Isothermal Bus, based on a modular HTC with VHT concept, is a robust TCS approach enabled through four key tenants. These include: 1) MODULARITY: A single thermal bus thermally couples all components on a S/C and provides heat load sharing. In addition, radiators are coupled only to the thermal bus. This approach provides functional separation and 16 interface decoupling needed for a Figure 2. Traditional versus modular approach. robust design (Figure 2). 2) REDUCED TEMPERATURE GRADIENTS: Thermal gradients (i.e. spatial isothermality) are minimized through high thermal conductance architectures and/or improved power distributions. However, a true ‘isothermal’ bus architecture is not realistic and therefore designs approaching that performance (i.e. quasiisothermal) should be considered. It is not clear what should characterize a quasi-isothermal bus and what these implications have on the performance and design envelope of the satellite. For example, is a 10 K temperature difference across the satellite good enough or would a more or less strict standard provide ‘good enough’ results. 3) ENVIRONMENTAL DECOUPLING: Exterior surfaces are insulated to minimize the influence of environmental heat loads. All environmental coupling is achieved through the radiator and therefore through the thermal bus. 4) THERMAL CONTROL: Thermal control is included in the thermal bus at the system (e.g. radiator) and/or component (e.g. heat switch) level. This provides the dynamic range of the thermal bus and provides temporal control (i.e. temporal isothermality). As with temperature gradients, the definition of a standard control authority along with the resulting design envelope needs to be addressed. Some of these traits can be verified by examining the results of Young.15 Candidate solutions for thermal control architectures were analyzed using high fidelity on-orbit S/C models. Design reference missions were developed to capture various combinations of equipment power and masses and mission orbital profiles. Figure 3 compares the results for the traditional, HTC, and Isothermal Bus architectures. Figure 3. Comparison of orbital simulation results of three The advantage of the Isothermal Bus architectures: traditional, high thermal conductance, and concept is demonstrated by examining the Isothermal Bus. range of temperatures over a single orbit. The traditional approach shows a temperature 3 American Institute of Aeronautics and Astronautics

range (i.e. maximum minus minimum temperature over a single orbit) of ~42°C. By improving bus thermal conductance in the HTC approach, spatial thermal gradients are reduced to a temperature range of ~18°C. However, temporal changes in temperature still exist. By including VHT in the Isothermal Bus concept, the temperature range is reduced to ~14°C. This example demonstrates a few of the advantages of the Isothermal Bus; however, there are several other advantages not captured. These include: 1) HIGHER CAPABILITY SYSTEMS: Improved isothermality allows for higher capability (i.e. higher power) systems. Consider a GaN device generating 600 W over a 0.40 cm2 area (i.e. 1,500 W/cm2) and a heat spreader that distributes this heat to a 0°C thermal sink with an average heat flux of 6.2 W/cm2 (i.e. 600 W at 96.7 cm2) and temperature drop of no more than 25°C. To better understand the problem, the heat equation in polar coordinates was solved. Assuming steady-state, no internal heat generation, constant thermal conductivity values, and no angular dependence, the governing equation simplifies to Figure 4 where r and z are radial and thru thickness coordinates while K is a ratio of in-plane to thru-thickness thermal conductivity.

Figure 4. Illustration of thermal spreader problem Figure 5. Illustration of effective in-plane thermal conductivity requirements Using the details from above, a nominal thermal spreader of 1 cm thickness was modeled with thru-thickness thermal conductivity of 400 W/m-K (i.e. copper). The bottom of the thermal spreader was perfectly coupled to a thermal sink at 0°C (Figure 4). The temperature profile versus normalized distance from center is shown in Figure 5 for varying in-plane thermal conductivities. This figure illustrates how improving isothermality enables higher capability systems. A simple copper thermal spreader would have a hot spot of ~115°C, much greater than requirements. Only with in-plane thermal conductivities 25x that of copper (i.e. 10,000 W/m-K) or greater can a 25°C temperature drop be achieved and a higher power/capability system can be achieved. 2) REDUCED COST AND SCHEDULE: The Isothermal Bus architecture could allow a thermal engineer to treat the bus as a single node, thereby eliminating the need for complex thermal models. In addition, a modular architecture helps enable integration of new component and/or new component placements. For a traditional bus, this could easily force a system redesign. Finally, the Isothermal Bus could accommodate a wide range of missions, payloads, components, and orbits, which would reduce cost. One TCS for multiple missions eliminates the non-recurring engineering costs inherent to point designs. Although these advantages are great, they come with trade-offs. These include: added complexity of an active thermal control system and the relatively low technology readiness level (TRL) for this new and advanced approach. Mass penalties could also be incurred; however, these could be offset by reduced heater requirements (e.g. batteries, heaters, wiring, etc.) found in traditional cold-biased systems. 4 American Institute of Aeronautics and Astronautics

IV. Isothermal Bus Design Traditional TCS design methodologies provide highly optimized and capable systems that are strongly influenced by a specific set of mission requirements. These include S/C configuration (i.e. component and payload selection along with their placement and orientation) and orbit definition, which defines the external thermal environment. The sum of these requirements in turn dictates TCS design. Traditional, risk averse TCS are designed to accommodate a single combination of these design parameters and are not intended for ‘off-design’ conditions. The traditional TCS design process is extremely time intensive (Figure 6). Virtually every aspect of the mission, orbit, payload, bus, component, and operations must be known before the thermal design can be completed. The goal is to guarantee operation or survival during all aspects of the mission while minimizing system mass, design complexity, and heater power.

Traditional Isothermal Bus Figure 6. Comparison of the traditional thermal development approach and the Isothermal Bus approach. The first step is to determine worst-case environmental heat inputs, component temperature limits, and internal power dissipation. Using these inputs, external and internal heat sources are balanced with heat radiated to space. This is done with a simple first-order model where basic thermal control designs are evaluated and parametric studies are completed. As the design develops, model fidelity increases and the process becomes iterative. As the design of the satellite and the components change, the thermal model must be updated to evaluate its effect on the TCS and the overall design of the S/C. For many systems, semi-passive approaches are used where the hot case is cooled through conduction to and internal radiation exchange with the radiator surfaces, and heaters are used to ensure proper operation during the cold case. This ‘cold-biased’ design approach is very effective for point designs but not robust enough to manage changes within the system. Once the detailed design and thermal model are completed, they must be validated in a thermal balance test where the system is subjected to multiple operational phases including one or two hot operational cases, one cold operational case, and one cold nonoperational case. It occurs during the thermal vacuum test campaign and is typically conducted only on the lead vehicle of a series of S/C.3 Using the results from the thermal balance test, the 5 American Institute of Aeronautics and Astronautics

thermal engineer validates the model and finalizes the design. The final proof test for the system is the thermal vacuum test, but this tends to be a workmanship test and is required for every vehicle before flight. The result is a process that takes six months to three years to complete for small S/C. Isothermal Bus TCS design is drastically different from traditional approaches as it will be intended for a broad range of scenarios. It can handle variations in the number, power distribution, and placement of components, and allow for rapid integration of new technologies. In addition, an Isothermal Bus TCS could handle a wide-range of orbits and therefore thermal environment variations. In doing this, a single S/C could accommodate multiple missions, but worst-case mission environments must be known a priori. There are two primary approaches to defining the worst-case design environments. The first is a very conservative approach in which all of the worst-case conditions for direct solar, albedo, and Earth outgoing, longwave radiation (OLR) heat loads are summed. This approach would provide 100% mission assurance but would also result in an overdesigned thermal control system and most likely a severe heater power requirement. The second approach is to evaluate all possible LEO combinations and from these determine statistically relevant worst-case environmental conditions.17 Under this approach, orbital-averaged environmental conditions as a function of temporal variations of beta angle (i.e. the minimum angle between the orbit place and solar vector), orbit inclination, orbit altitude, and historical launch data were evaluated. From these variables, relationships of beta angle versus inclination and a historical inclination distribution were developed. These two relationships were combined to develop a statistical weighting matrix and ultimately viable weighting matrices at varying threshold levels. By applying the variable weighting matrix to orbital-averaged heat load models and increasing the threshold value, they determined a single hot- and cold-case design orbit for a wide-range of surface optical characteristics. For the hot case, a beta angle of 72°, inclination of 52°, and altitude of 350 km provide an appropriate environmental heat load for all cases with a maximum error of 1.6%. For the cold case, two conditions need to be evaluated depending on the predominant satellite surface type. For solar reflector surfaces, a beta angle of 0°, inclination of 65°, and altitude of 1000 km provide reasonable cold-case conditions, while a beta angle of 0°, inclination of 28°, and altitude of 350 km are appropriate for all other surface types. The advantage of Isothermal Bus architecture is that the majority of the design can be completed before mission initiation (Figure 6). The TCS can be designed for a range of orbits, internal component heat loads, and mission operations. In addition, a thermal balance test can be conducted to validate the thermal model and verify the design. The only steps that would have to be completed after mission initiation would be a simple go/no-go check to verify the missions bounds are within the robust TCS design bounds and then to rapidly assemble, integrate, and flightqualify the system, including thermal vacuum testing for proof of workmanship. However, the accelerated design process can provide significant problems to the thermal engineer due to the relatively large and not clearly defined design space created by combinations of payloads, bus components and orbits.

V. Enabling Technologies The Isothermal Bus, based on a modular HTC with VHT concept, is a robust TCS approach enabled through four key tenants including: modularity, reduced temperature gradients, environmental decoupling, and thermal control. Several efforts are focused on overcoming each of these challenging concepts. A. Modularity Several modular TCS approaches have been developed. Each relies on a underlying thermal bus to thermally couple components. Examples include: TherMMS,15, 16 SMARTS,18, 19 FACTS14 and development efforts including HexPak,7, 8 SMARTBus,20 and ITEMS.21 Figure 7 shows both the SMARTS and ITEMS concepts. The SMARTS (Satellite Modular and Reconfigurable Thermal System) concept adheres to four design rules: 1) modest radiator oversizing, 2) maximum external insulation, 3) internal isothermalization and 4) radiator heat flow modulation.18 The SMARTS TCS is based on a frame and panel satellite architecture consisting of multiple Al-isogrid or Alhoneycomb panels bolted together along common edges. An embedded heat pipe is placed around the edges of each panel, while C-shaped heat pipes are distributed throughout the center. The combination of these two heat pipes allows for heat spreading across each panel and attempts to approach isothermal conditions.

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SMARTS TCS Concept18 Figure 7. Example modular thermal control approaches.

ITEMS TCS Concept21

ITEMS (Integrated Thermal Energy Management System) developed by JPL includes heat load sharing, heat rejection modulation, and minimized survival heater power.21 The ITEMS approach (Figure 7) utilizes a cooling loop to thermally integrate all S/C subsystems in conjunction with thermal switches and valves. The heat rejected from one subsystem is transferred to another subsystem where the heat is needed to maintain its minimum temperature. Any excess heat generated in the S/C above what is needed is rejected at a deployable radiator system that uses variable emittance devices on its surface. In addition, a phase change material is used for battery thermal control. This type of architecture provides the needed flexibility and accommodates low-cost overall design and implementation. B. Reduced Temperature Gradients Reduced temperature gradients are typically achieved through hardware. Examples include: heat pipes/loop heat pipes18, annealed pyrolytic graphite panels23, and relatively new oscillating heat pipe technologies.24, 25 One such technology is the Thermal Control Panel (TCP) developed by Thermal Management Technologies (Figure 8). This technology provides isothermal structural panels and low thermal impedance joints for use in an isothermal S/C structure. Each light weight panel provides the S/C structure while simultaneously exhibiting an effective thermal conductivity/mass ratio much greater than traditional S/C materials. In various forms the TCP can be used to load spread heat, create nearly isothermal S/C structures, Figure 8. Thermal Control Panel with 56W single 22 22 at less than 5K per panel temperature gradient. and to provide highly efficient space radiators. In addition to hardware approaches to isothermality, optimization approaches have advantages. In several studies thermally optimized component placements were investigated both globally and locally.26, 27 On average, optimized component distributions reduced maximum temperatures, increased minimum temperatures, and reduced maximum temperature differences by 5.4 K, 7.1 K, and 12.6 K, respectively, over evenly distributed components. The largest and smallest maximum temperature difference reductions were 17.8 K (at 400 W) and 9.5 K (at 100 W), respectively. C. Environmental Decoupling Used to conserve thermal energy and protect S/C from external radiation, insulation is a well-established and extensively utilized S/C TCS technology. Multilayer Insulation (MLI), the most common S/C insulation technology, 7 American Institute of Aeronautics and Astronautics

provides an effective method for insulating S/C. MLI consists of up to 25 layers of thermal control materials to obtain the desired optical and insulating properties.28 Combined with a well-established flight history, MLI blankets are effective and highly reliable. However, MLI requires a tedious design and installation process due to its inherent fragility and alternative forms of insulation are being sought. Trifu, Gould, Qassim, and Clark evaluated aerogel composite blankets as a potential replacement for MLI. Aerogel composite blankets had similar insulating performance to MLI.29 Compared to MLI, aerogels have limited flight history although recent missions have placed this technology in space. Most notably, the Mars rovers used carbon-opacified silica aerogel to insulate the most sensitive electronic components during cold Martian nights.28 Work continues to refine aerogel aerospace applications including the development of aerogel insulation panels.30 D. Thermal Control Thermal control can be achieved in numerous ways including at the bus- (e.g. radiator) and/or component-levels. Bus level thermal control has been investigated including electrochromic, variable-emissivity devices31 and loop heat pipe variable control.19 Component-level control includes paraffin-based technologies and variable thermal layers using thermoelectric devices.32, 33 The Space Test Program-Houston 4-Active Thermal Tile (STP-H4-ATT) mission investigated variable conductance thermal tiles that serve as a quick-insert thermal management device for satellite components (Figure 9). The tiles contain thermoelectric devices capable of operating in heating, cooling and neutral or “off” modes.34-36

Figure 9. Active thermal tile experiment.35, 36

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VI. Case Study To demonstrate the advantages of Isothermal Bus over a traditional approach, a case study was developed and evaluated. A nominal hexagonal S/C with typical frame and panel construction was modeled in Thermal Desktop® as shown in Figure 10. The 14,840 node model consists of 6 side panels, 1 top/payload deck, and 1 bottom deck. Each side panel measures ~0.71 m x 0.65 m while the top/bottom decks are 1.29 m across the longest dimension. The six side panels include body-mounted radiators while the remaining surfaces are fully insulated. To simplify the analysis, a single 15 cm x 15 cm component was placed on the inside face of each of eight panels. For each analysis, temperatures of the eight components were tracked over two complete geosynchronous orbits. The second orbit provides quasi-steady state results. Temperature results included maximum temperature (i.e. maximum temperature that any component sees during an entire, quasi-steady state orbit) and minimum temperature (i.e. minimum temperature that any component sees during an entire, Figure 10. Nominal hexagonal spacecraft with quasi-steady state orbit). Temperature results were then transparency to show internal components. a function of four variables. These include: 1) BUS POWER: The overall bus power (i.e. heat Table 1. Typical spacecraft component operating load) on the structure was varied from 0 W up to temperatures.37 1000 W. Power was distributed to the payload Component Operational Survival component (25%), radiator components (65%), Temperature Temperature and the bottom deck (10%). 2) SPATIAL ISOTHERMALITY: Structural panel Batteries 273 – 288 K 263 – 298 K transverse thermal conductivities were varied from a low-value of 20 up to a high value of 300 Electronics 253 – 323 K 233 – 348 K W/m-K to simulate a traditional honeycomb Reaction wheels 263 – 323 K 253 – 333 K architecture to a more modern Isothermal Bus (e.g. heat pipe) architecture. In addition, panelAntenna gimbals 233 – 353 K 223 – 363 K to-panel conductance was varied from 10 W/K Antennas 173 – 373 K 153 – 393 K to 20 W/K to model traditional and Isothermal Bus approaches. Solar arrays 123 – 383 K 73 – 403 K 3) POINTING: S/C pointing was adjusted to Payloads Varies Varies examine the effect of solar loading on the results. A variety of pointing conditions were run including nadir and sun-facing along with several S/C rotations. 4) RADIATOR EMISSIVITY: The traditional TCS utilized a radiator emissivity of 0.92. Control was simulated in the Isothermal Bus by varying radiator emissivity from a low value of 0.1 up to 0.7. This provided a switching ratio of 7 to 1. Typical operational and survival temperature limits for components were identified and summarized in Table 1. Based on these values, a nominal operating temperature regime for internal electronic components was assumed to be 273 to 313 K (0 to 40°C). Figure 11 summarizes the results of these simulations by plotting minimum and maximum component temperatures versus total bus power for both the traditional and Isothermal Bus. Both minimum and maximum temperature results show some scatter in results due to pointing effects. As expected, the traditional bus shows increasing component temperatures versus power. Since no control is provided with this system (i.e. thermal switching), survival heat is necessary to maintain minimum component temperatures above 273 K. However, at a power value of ~300 W, maximum component temperatures exceed 313 9 American Institute of Aeronautics and Astronautics

K. Because of this, maximum power for this system is 300 W. In addition, this system requires survival heat and associated system components to maintain minimum temperatures above requirements.

Figure 11. Demonstration of Isothermal Bus approach versus a traditional bus. Isothermal Bus results show both minimum and maximum temperatures. Isothermal Bus minimum temperatures are based on minimum component temperatures for a quasi-steady state orbit at a radiator emissivity of 0.7 (i.e. maximum thermal dissipation). Isothermal Bus maximum temperatures are based on maximum component temperatures for a quasi-steady state orbit at a radiator emissivity of 0.1 (i.e. minimum thermal dissipation). These results show the advantage of the Isothermal Bus concept. First, at relatively low power levels, an emissivity of 0.1 would keep the bus warm enough with only a modest amount of survival heat required. However, at relatively high powers, an emissivity of 0.7 would keep the bus cool enough up to ~800 W of bus power. As a result, the Isothermal Bus concept would enable a 167% increase in bus power while almost eliminating the need for survival heat.

VII. Conclusions The Isothermal Bus architecture is a novel idea focused on approaching a single thermal node S/C representation. In effect, heat is shared efficiently between cold and hot components and in combination with thermal control, spatial and temporal variations in temperature are minimized. The Isothermal Bus can enable higher capability systems along with providing schedule and cost savings. A nominal case study was provided that compared a traditional TCS to the Isothermal Bus concept with a 7:1 switching ratio. While the traditional TCS could accommodate 300 W of bus power, it would require survival heaters. However, the Isothermal Bus concept would enable a 167% increase in bus power while almost eliminating the need for survival heat.

VIII. References 1

Donabedian, M. and Gilmore, D. G., Spacecraft thermal control handbook: Aerospace Press, 2003.

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2 Saleh, J. H. and Dubos, G., "Responsive space: Concept analysis, critical review, and theoretical framework," in AIAA Space, 2007. 3 Williams, A. D. and Palo, S. E., "Issues and Implications of the Thermal Control System on Responsive Space Missions," DTIC Document2006. 4 Noel, J., Escorpizo, R., and Jones, E., "Transforming the National Spacelift Architecture," ed, 2004. 5 Chaplain, C., Space Acquisitions: DoD Is Making Progress to Rapidly Deliver Low Cost Space Capabilities, but Challenges Remain: DIANE Publishing, 2008. 6 Fronterhouse, D., Lyke, J., and Achramowicz, S., "Plug-and-play satellite (PnPSat)," AIAA [email protected] Aerospace, 2007. 7 Hicks, M., Enoch, M., and Capots, L., "HEXPAK–A FLEXIBLE, SCALABLE ARCHITECTURE FOR RESPONSIVE SPACECRAFT," in Paper No. RS3-2005-3006, 3rd Responsive Space Conference, Los Angeles, CA, 2005. 8 Hicks, M. T., Hashemi, A., and Capots, L. 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A., and Williams, A. D., "Optimal Distribution of Electronic Components to Balance Environmental Fluxes," Journal of Spacecraft and Rockets, vol. 48, pp. 694-697, 2011. 28 Donabedian, M., Gilmore, D., Stultz, J., Tsuyuki, G., and Lin, E., "Insulation," Spacecraft thermal control handbook, vol. 1, pp. 161-205, 2002. 29 Trifu, R. M., Gould, G. L., Qassim, K., and Clark, J. L., "Ultra-lightweight aerogel superinsulation as an MLI replacement," in Engineering, Construction, and Operations in Challenging [email protected] sEarth and Space 2004, 2004, pp. 976-982. 30 White, S. O., Zafiropoulos, N. A., and Clark, J. L., "Aerogel insulation panels and manufacturing thereof," ed: Google Patents, 2014. 31 Shannon III, K. C., Sheets, J., Groger, H., and Williams, A., "Thermal management integration using plug-and-play variable emissivity devices," in SPIE Defense, Security, and Sensing, 2009, pp. 73300F-73300F-9.

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32 Hafer, W. T., Vitale, N., Macris, C., Ebel, R., McCullough, J., and Williams, A., "Design of a variable thermal layer (VTL) for a generic satellite component interface," in 49th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 16th AIAA/ASME/AHS Adaptive Structures Conference, 10th AIAA Non-Deterministic Approaches Conference, 9th AIAA Gossamer Spacecraft Forum, 4th AIAA Multidisciplinary Design Optimization Specialists Conference, 2008, p. 2259. 33 Hafer, W. T., Vitale, N. G., Macris, C., Ebel, R., McCullough, J., and Williams, A., "Design and use of a variable thermal layer (VTL) for rapid satellite component integration," in Proceedings of the 6th Responsive Space Conference, Los Angeles, CA, 2008, pp. 1-10. 34 NASA. (2016, 4/24/2016). Space Test Program-Houston 4-Active Thermal Tile (STP-H4-ATT). Available: www.nasa.gov/mission_pages/station/research/experiments/871.html 35 Delaney, R. K., "Exploring the Performance of Active Thermal Tiles for Space Applications," 2014. 36 Delaney, R. K. and Dumm, H.-P., "Ground Testing of Active Thermal Tiles," in 53rd AIAA Aerospace Sciences Meeting, 2015, p. 0375. 37 Larson, W. J. and Wertz, J. R., "Space mission analysis and design," Microcosm, Inc., Torrance, CA (US)1992.

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