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Abstract. The advent of very small satellites, such as nano and microsatellites, logically leads to a requirement for smaller thermal control subsystems. In addition, the thermal control needs of the smaller spacecraft/instrument may well be different from more traditional situations. For example, power for traditional heaters may be very limited or unavailable, mass allocations may be severely limited, •and fleets of nano/microsatellites will require a generic thermal design as the cost of unique designs will be prohibitive. Some applications may require significantly increased power levels while others may require extremely low heat loss for extended periods. Small spacecraft will have low thermal capacitance thus subjecting them to large temperature swings when either the heat generation rate changes or the thermal sink temperature changes. This situation, combined with the need for tighter temperature control, will present a challenging situation during transient operation. The use of "off-the-shelf' commercial spacecraft buses for science instruments will also present challenges. Older thermal technology, such as heaters, thermostats, and heat pipes, will almost certainly not be sufficient to meet the requirements of these new spacecraft/instruments. They are generally too heavy, not scalable to very small sizes, and may consume inordinate amounts of power. Hence there is a strong driver to develop new technology to meet these emerging needs. Variable emittance coatings offer an exciting alternative to traditional control methodologies and are one of the technologies that will be flown on Space Technology 5, a mission of three microsatellites designed to validate "enabling" technologies. Several studies have identified variable emittance coatings as applicable to a wide range of spacecraft, and to potentially offer substantial savings in mass and/or power over traditional approaches. This paper discusses the development of the variable emittance thermal suite for ST-5. More specifically, it provides a description of and the infusion and validation plans for the variable emittance coatings.

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This paper describes the testing of the prototype loop heat pipe (LHP) for the Geoscience Laser Altimeter System (GLAS). The primary objective of the test program was to verify the loop's heat transport and temperature control capabilities under conditions pertinent to GLAS applications. Specifically, the LHP had to demonstrate a heat transport capability of 100 W, with the operating temperature maintained within +2K while the condenser sink was subjected to a temperature change between 273K and 283K. Test results showed that this loop heat pipe was more than capable of transporting the required heat load and that the operating temperature could be maintained within +_K. However, this particular integrated evaporator-compensation chamber design resulted in an exchange of energy between the two that affected the overall operation of the system. One effect was the high temperature the LHP was required to reach before nucleation would begin due to inability to control liquid distribution during ground testing. Another effect was that the loop had a low power start-up limitation of approximately 25 W. These issues may be a concern for other applications, although it is not expected that they will cause problems for GLAS under micro-gravity conditions. EEB.O_D.Ug.TJgBAs oart of the Earth Science Enterprise, the science objectives of the Geoscience Laser Altimeter System (GLAS) are to obtain ice sheet and ocean topography, and global profiles of land and vegetative canopy. In order to accomplish this, GLAS utilizes three lasers that dissipate approximately 120 W each when operating, but only one laser is needed at a time. The lasers must be maintained at a temperature of 293±2K.In order to transfer the heat and meet the temperature requirements, a heat pipe/loop heat pipe (LHP) system was proposed for therr_l control.Since GLAS was still in the conceptual design stage, it was desirable to build and test a thermal control system reflective of that propose1 for the mission. In the actual application, a heat pipe will be mounted to each of the I_sers through a thermal interface. The condenser section of the heat pipe will be connected to the evaporator section of the LHP. The co,-xlenser region of the LHP will be mounte¢: to a radiator panel. The front of the panel will then radiate to space. NASNGoddard Space Flight Center purchased a prototype LHP from D,vnatherm Corporation in 1997. The Thermal Engineering Branch at Goddard subsequently completed testing of this prototype.Testing for the GLAS LHP was divided into three pads. The first set of tests, which was designee to determine the temperature drop through the system and the thermal conduct_nces at various interfaces, was performed strictly using a heat pipe assembly. In these te_-;ts,a chill block was used to simulate -1- American Institute of Aeronautics and Astronautics AIAA-99-0473 Figure 1 GLAS LHP Figure 2 GLAS LHP Evaporator andCompensation Chamber the LHP evaporator and a heater block was used to simulate the laser box. The second set of tests was c...