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Thermal Management for Multifunctional Structures Suraj P. Rawal, David M. Barnett, and David E. Martin Abstract— Multifunctional structures (MFS) is an innovative concept that offers a new methodology for spacecraft design, eliminating chassis, cables and connectors, and integrating the electronics into the walls of the spacecraft. The MFS design consists of multilayer flexible circuit patches bonded onto a structural composite panel, and multichip modules (MCM’s) performing specific functions are bonded onto the circuit patches which are interconnected via flexible circuit jumpers. Incorporation of the high power density two-dimensional (2-D) and three-dimensional (3-D) MCM’s into smaller and more efficient packaging designs still has the fundamental requirement to maintain component temperatures within design limits. Higher component qualification temperatures, such as 393 K, can result in smaller spacecraft radiator areas that are consistent with efficient packaging schemes. During the MFS development effort, a structural radiator panel was fabricated using high thermal conductivity (Hi-K) composite facesheets, and several thermal management designs using combinations of Hi-K doublers (150–1500 W/m-K), Hi-K (150–700 W/m-K) corefill, and deployable radiators to maximize MCM’s heat rejection. Results of the thermal vacuum tests and details of the thermal design methodology are presented in this paper. Index Terms— Avionics, miniature spacecraft, multichip module (MCM), multifunctional structures (MFS), thermal control options, thermal management.
I. INTRODUCTION A. Spacecraft Trends and MFS Design
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PACECRAFT design is primarily driven by a trend to increase payload mass fraction and develop high performance, low-cost, small satellites. Future small spacecraft will achieve their low mass and size not only by reduction of the power source mass and volume but through miniaturization of the whole avionics system including the large parasitic mass associated with avionics containers, cables, structural support of packaged avionics, or connectors. These para-sitic components can contribute as much as 50% of the mass of space science spacecraft [1]–[3]. Rapid advances in large-scale integrated electronics packaging, lightweight composite structures, and high conductivity materials are enabling the development of an important new manufacturing and integration technology called multifunctional structures [1], [4], [5]. The overall MFS concept is to embed electronics assemblies (e.g., multi-chip modules or Manuscript received December 11, 1998; revised February 19, 1999. This work was presented in part at the International Conference and Exhibition on Multichip Modules and High Density Packaging, Denver, CO, April 15–17, 1998. This paper was supported in part by Lockheed Martin Astronautics Independent Research and Development project Lightweight Spacecraft Technology (D-90D). The authors are with Lockheed Martin Astronautics, Denver, CO 80201 USA (e-mail:
[email protected]). Publisher Item Identifier S 1521-3323(99)06394-7.
MCM’s), miniature sensors, and actuators into load carrying structures along with associated embedded cabling for power and data transmission (Fig. 1). This level of integration effectively eliminates traditional boards, boxes, large connectors, bulky cables, thermal baseplates, etc., thereby yielding major weight, volume and cost savings. Furthermore, each major improvement in miniaturization of components, sensors and devices becomes a candidate for integration into MFS. The MFS design inherently facilitates the increase in payload-mass fraction by maximizing the ratio of the volume of fundamental electronic parts to the total packaging volume. MCM’s offer considerable reduction in both the microelectronic packaging volume and interconnection signal delay. However, this integration results in higher heat fluxes at the first- and second-level packaging technologies. Therefore, thermal management of microelectronic devices is required for proper operation and acceptable reliability. The methods of connection for the signals and power from the MCM to the next level package (e.g., printed wiring board or 3-D MCM stack) can present a challenge for adequate heat removal. The MFS approach provides the system cabling and interconnection, the fourth-level package, as 2-D and/or 3-D MCM’s are directly mounted on the structural baseplate. Addressing both the electrical performance and thermal considerations, the MCM designers select appropriate substrates, die attachment materials, interconnect methods, and packaging materials. Recognizing the inherent thermal design of an MCM package, this paper presents the thermal management options which have been evaluated in the development of the MFS designs for miniature spacecraft. II. MFS THERMAL CONTROL ASPECTS A. Trends in Spacecraft Electronics Boxes Satellite electronic boxes requiring high-heat dissipation are usually mounted on Hi-K plate elements. The plate elements, often referred to as heat sink plates or thermal doublers, are bonded to honeycomb sandwich panels (the baseplate). The panels transfer heat energy from the satellite into space. A sandwich panel combined with such a plate element can be treated as a radiation fin. The greater the heat flux on a box/component footprint, the higher the required heat rejection capability of the fin to maintain the component’s temperature within the allowed range. With miniaturization in electronic components, the waste heat generated has remained nearly constant. Even in cases where the box power has decreased, the reduction has not kept pace with the footprint reductions, thus resulting in an increase in the overall Watt (W) density (power/footprint area). Fig. 2 shows this miniaturization progression for a typical command and data handling (C&DH) box.
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Fig. 1. MFS panel including structural panel with integral thermal control, polyimide flex circuitry, MCM sockets, and formed cover provide electromagnetic shielding and protection.
Fig. 2. Trends in miniaturization of satellite electronic box mounted on composite panel.
B. Thermal Control Options The sandwich panel fin receives a heat flux from within a rectangular footprint on the inner skin and solar energy loads on the outer skin. The heat transfer area required to reject the waste heat from the electronic box is governed by the radiation equation:
where: is the rate of heat rejection to the environment is the heat transfer surface area; is the (energy/ time); is the emissivity of the radiator thermal control coating; W/m K ); is the Stephan-Boltzman Constant ( is the effective area weighted average radiator temperature; radiation environment. (This parameter takes into account the effects of all surfaces in the radiator hemispherical field of view including space and other spacecraft surfaces.) The area weighted radiator temperature takes into account temperature gradients in the radiator. The minimum radiator area occurs when there is no temperature gradient in the radiator. A 25 W box mounted to a radiator panel requires about 1 sq. ft. of 100% efficient radiator area to keep the box baseplate below its limit of 40 C with a 25 C radiation
sink. For a given thermal dissipation and environment, the minimum required radiation area is a constant, regardless of the electronic box footprint. For a traditional box, the required area may be very nearly the box footprint. For smaller footprint boxes, the radiator panel must have enough lateral conductance to spread the heat from the box baseplate footprint to the required radiator area. The lateral conductance is controlled by
where rate of heat flow (same as in the radiation equation); material thermal conductivity; cross-sectional area in the heat flow direction; baseplate temperature; average radiator temperature (same as the radiation equation); heat flow path length. For a traditional box with a footprint nearly the same size as the radiator, there is little spreading required ( is small), hence the thermal conductivity and the thickness of the radiator are not as important. An electronic box with a small footprint
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Fig. 3. Thermal control designs solutions for increasing thermal loads from electronic components.
and the same power depends on the heat spreading ability of the radiator material. This is illustrated by setting the radiation equation and the conduction equation equal to each other and solving for the baseplate tempera-ture. This can be done since the heat that is conducted into the radiator is the same heat that is radiated to the environment. The resulting equation takes the form Fig. 4. Three-dimensional MCM integrated on the MFS panel.
As the footprint gets smaller, the conduction length gets larger reducing the efficiency of the radiator and, hence, goes up (a typical box baseplate limit is 40 C to 50 C). To must equally increase, for example, counter this effect, by using HI-K facesheets. There is a practical limit to the material conductivity and cross-sectional area. When these parameters cannot be increased, a heat spreading device must be added, such as a thermal doubler or heat pipes. The effect of a heat spreader is to return the radiator to as close to an isothermal condition as possible. The isothermal radiator is the most efficient and smallest, which minimizes volume and mass requirements. Fig. 3 shows the potential thermal control options (including thermal doublers, heat pipes, deployable radiator fins, and capillary pumped loop) for the wide range of heat loads. A Hi-K facesheet such as K13C2U or K1100 fiber reinforced cyanate ester (CE) composite is often used to dissipate the moderate heat loads. As the heat load increases, specific system and component level thermal analyzes are performed to support thermal/structural/material design trade studies to obtain an optimum solution. For example, 3-D carbon-carbon 200 W/m-k and low modulus with has been successfully used as thermal doublers under the electronic boxes in spacecraft avionics subsystems. The HiK core fill with through-the-thickness conductivity of 400 to 800 W/m-K have been used in structural panels requiring ) through-the-thickness conducsignificantly increased ( tance. With increasing heat load, and driven by stringent mass and performance requirements, the next generation of small satellites could incorporate high conductivity panels with embedded heat pipes and also a deployable fin. In the case of a 2-D or 3-D MCM stack [for C&DH (Fig. 2, right side)], a heat spreading device is required, but since the chassis of the box has been eliminated, the limit temperature
is governed by the electronics parts which may be able to go above 80 C (353 K). At 80 C, only 0.5 sq ft of radiator area is required to reject 25 W, compared to 1 sq ft required at 40 C. Fig. 4 shows a 3-D MCM stack using a thermal strap of copper or composite to transfer the heat to the radiator surface. This is required, since the conductance through the stack is poor. Another option is a conductive interface under each layer of the MFS to bring the heat to a small diameter (3 mm diameter) heat pipe. The heat pipe then transfers the heat from the stack to the radiator and spreads the heat as well. III. THERMAL ANALYSIS AND VACUUM TESTING OF MFS PANELS When incorporating high-power density MCM’s into smaller and more efficient packaging designs, it is a fundamental requirement that the component temperatures be maintained within given limits. Higher component qualification temperatures, such as 80 C to 120 C (393 K), can result in smaller spacecraft radiator areas that are consistent with efficient packaging schemes. A brief description of the thermal vacuum tests of the MFS panels [6] with three thermal designs is presented below. IV. MFS PANEL CONFIGURATION Three composite panels ( cm cm) with K13C2U/CE facesheets and aluminum honeycomb core and three different thermal designs (Panel A, B and C) were fabricated (Fig. 5). Panel A is a standard design panel that uses Hi-K K13C2U/CE panel facesheets. Panel B has thicker W/m-k) filler under the facesheets and a Hi-K ( MCM. Panel C is the same as Panel B, but also uses a 0.1-in.-thick isotropic carbon-carbon thermal doubler under the MCM. The thermal doubler reduces the thermal gradient
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Fig. 5. MFS panels with three thermal designs to reject heat generated by an MCM.
Fig. 6. Heater power density profile of the MFS panel with heat simulator MCM during the thermal vacuum test.
within the MCM ceramic low-temperature co-fired alumina base. RTV thermal interface material is used under the MCM for all three panels to reduce thermal contact conductance. These three MFS panels were tested under identical thermal vacuum conditions. The MCM was mounted in the center of each panel. The five test heaters and four heaters on the panel perimeter and a heater on the MCM test lid were used to represent various operational conditions. The panel was insulated on one side with multilayer insulation (MLI), and painted black on the opposite side that becomes the radiator. The MCM’s included several internal film heaters to represent circuit elements, which were operated at about 323 K and 393 K. The heater power levels were adjusted to achieve the test temperature. Higher power levels indicate better MCM heat rejection. V. TEST RESULTS AND MODEL CORRELATION The heater power density (ranging from 5–14 W/cm ) results (Fig. 6) indicated that the Panel C consistently exhibited the best performance. The through-the-panel temperature gradient for one of the test cases and analysis are shown in Fig. 7(a) and (b), respectively. Model predictions show the expected behavior of reduced thermal gradient for Panel C
[Fig. 7(b)]. However, the test results show that Panel A and Panel B exhibited similar temperature distributions in all test cases [Fig. 7(a)]. Recall that Panel A had no thermal core fill under the MCM, while Panel B had the Hi-K ( 700 W/m-k) filler. Panel C also had the same filler as Panel B and exhibited a slightly lower peak temperature at the panel midplane. These test results suggest that there was poor bonding between the thermal filler and the panel facesheets. Posttest ultrasonic inspection verified filler debonding for both Panel B and C. For Panel C, Fig. 7(b) shows that the model analysis agreed better with the test when the filler was eliminated. Thermal performance testing and the model correlation efforts indicated that the C-C thermal doublers with RTV offered a good approach to enhance MCM heat rejection. The high conductivity thermal corefill in the honeycomb core under the MCM can enhance heat rejection but care must be taken to ensure full bonding of the corefill to the facesheets. Subsequent process development efforts have been successful to ensure intimate bonding between the composite facesheets and to obtain nearly a 3 increase in the through-the-thickness panel conductance. For example, the through-the-thickness panel conductance increased from 0.15 W/ C for the unfilled core to 0.53 W/ C for the Hi-K corefilled region [7].
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REFERENCES
(a)
(b) Fig. 7. Through-the-panel temperature gradients (for case 12): (a) test and (b) analysis.
VI. CCONCLUSION The MFS is a revolutionary design approach which integrates the electrical, thermal and structural functions on a spacecraft bus panel. By using copper/polyimide flex interconnect architecture, bulky and heavy cables/harnesses and connectors are eliminated. The flex circuit interconnects provide all power and signal distribution functions directly on the composite bus panel. The 2-D and 3-D MCM’s are also directly mounted on the structural panel via the flex circuit patches, virtually eliminating the need for large enclosures or circuit boards. Thermal management aspects of the MFS panel were evaluated by using three different design configurations to reject heat generated by a heat simulator MCM. Results of these tests showed that the combination of high conductivity facesheet, thermal doubler, high conductivity core fill provided adequate solution to dissipate 14 W/cm heater power density. The paper suggests the candidate thermal design options which can be realistically implemented to dissipate high heat fluxes on MFS panels. ACKNOWLEDGMENT The authors would like to thank Dr. A. Das, AFRL/PRS, for his technical guidance and encouragement for the MFS development effort and Dr. J. Chapter for his support in thermal test and analysis of the MFS panel.
[1] M. Obal and J. Sater, “Multifunctional structures: The future of spacecraft design?,” in Proc. 5th Int. Conf. Adaptive Structures, Sendai, Japan, Dec. 5–7, 1994. [2] C. Nunez, “JPL spacecraft mass trends,” Interoffice Memo 3132-93-076, Jet Propulsion Lab., Pasadena, CA, Feb. 19, 1993. [3] B. Muirhead, “Technology thrust areas for mass constrained spacecraft,” in Proc. AIAA/DARPA Conf. Lightweight Satellite Syst., Naval Postgraduate School, Monterey, CA, Aug. 4–6, 1987. [4] J. Sercel et al., “Modular and multifunctional systems in the new millennium program,” in Proc. AIAA-96-0702, 36th Aerosp. Sci. Meeting, Reno, NV, Jan. 15–18, 1996. [5] D. M. Barnett and S. Rawal, “Multifunctional structures technology experiment on deep space 1 mission,” in Proc. 16th Digital Avionics Syst. Conf., 1997, pp. 2.3-1–2.3-7. [6] J. Chapter, “Honeycomb panel spacecraft radiator with multichip module thermal analysis and vacuum testing,” in Proc. Space Technol. Appl. Int. Forum, pp. 801–812, 1997. [7] S. Rawal, “Thermal structural materials solutions for spacecraft applications,” Monthly Lett. Rep. 18, submitted to AFRL/Wright Lab, Mar. 15, 1997.
Suraj P. Rawal received the B.S. and M.S. degrees in metallurgical engineering from the Indian Institute of Technology, and the Ph.D. degree in materials science and engineering from Brown University, Providence, RI. He is the Manager of the Advanced Structures, Materials, and Controls Group, Lockheed Martin Astronautics (LMA) Flight Systems, Denver, CO. He has been active in the development of new materials, fabrication technologies, and hardware insertions for spacecraft applications. He is the Program Manager for the technical efforts; “Spacecraft Integrated Electronics Structures (SIES),” “Thermal Structural Materials Solutions for Spacecraft,” LMA IR&D D-90D project, “Lightweight Spacecraft Technologies,” and other multifunctional spacecraft technology programs. He has been instrumental in the development of lightweight composite radiators, electronic packages, and thermal straps using innovative designs and processing approaches. In addition, he has developed a thermal-mechanical property database of all advanced composite materials. He has authored over 35 articles on advanced composites, thermal management, and structure/property relationships Dr. Rawal has received numerous Lockheed Martin awards, including inventor, author, and technical achievement.
David M. Barnett received the B.S. degree in electrical engineering from Kansas State University, Manhattan, and the M.S. degree in electrical engineering from the University of Michigan, Ann Arbor. He is a Senior Staff Engineer in the Advanced Structures, Materials, and Controls Group, LMA Flight Systems, Denver, CO. He is the Prinicpal Investigator for several LMA programs including SIES (multifunctional structure electronic circuit design), black box design in support of CIS solar array development and follow-on demonstrations, and SMA driver design. He has considerable experience in the design, development, and fabrication of innovative electronic devices that are integrated into spacecraft application. His previous experience includes principal integrity engineer (PIE) for electronic hardware, avionics, sensors, data acquisition systems, and command and data handling (C&DH) hardware on several major aerospace programs. He has also authored several articles in peer-reviewed journals. Mr. Barnett received the Lockheed Martin inventor award, technical achievement awards, and commendations. He is a senior member of AIAA.
David E. Martin received the B.S. degree in mechanical engineering from Colorado State University, Fort Collins. He is a Staff Engineer in the Thermal Group, LMA Flight Systems, Denver, CO. He has considerable experience in the design and development of active and passive thermal control systems for spacecraft application. His previous experience includes product integrity engineer (PIE) for two-phase thermal control hardware and system thermal analyst on several aerospace programs. Mr. Martin is a member of ASME.
