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Microsatellite BIRD (Bispectral InfraRed Detection), mass 92 kg, sizes 550×610×620 mm was put on October 22, 2001 in a sun-synchronous orbit. The passive thermal control system (TCS) provided a temperature range of −10…+30 °C for a payload. It is assembled from precision optical instruments and housekeeping equipment with average power about 35 W. In the observation mode a power consumption peak of 200 W is occurred during 10-20 min. The TCS ensured a thermal stable design of the payload structure and is realised by heat transfer elements (conductors and grooved heat pipes), which thermally connected the satellite segments, two radiators, multilayer insulation and low-conductive stand-offs. Three years in space have brought an enormous volume of telemetric data about thermal performance of the TCS, based on information from temperature sensors, power consumption, attitude relative to Sun and Earth.

Micro Satellites are one of promising instruments for near earth space research programs. The strong restrictions to mass and power budget of satellite subsystems and payload lead to choice of mainly passive Thermal Control Systems (TCS), often with heat pipe (HP) integration. New tendencies in micro satellite thermal requirements as multi temperature level payload instruments, thermal stability of mounting structures for precision optical devices cause the corresponding adequate modifications in thermal concept and in hardware realisation. Intended aim of this paper is to present the experience collected by authors during thermal design and preflight thermal tests of Micro Satellite BIRD, developed under the German small satellite program.

Heat pipes efficiency for space thermal engineering is widely recognized and illustrated. Among majority of heat transfer tasks the heat pipe employment for high precision optical devices' thermal control is interesting and challenging. The main heat pipe functions are associated with a) heat removal tasks in the wide temperature range (90 ÷ 300 K), b) heat transfer and redistribution inside/outside of system, c) providing of isothermality of mounting plate (seats) of devices; d) regulating of temperature in selected device or group of devices. The brief description of technical realization example of heat pipe application in peculiarities of small satellite design is subject of this report.

For the prediction of the transient behavior of thermal nodes which are interacting within a Thermal Mathematical Model (TMM) it is necessary to know the heat capacity of each node. For instance this is actual for components of opto-electronic devices for space exploration. Other assignment is to define the thermal properties of new structure materials and their combinations. Often the base for the correction of the TMM is the comparison of the calculated node temperatures with the node temperatures measured on a Thermal Engineering Model (TEM) during a Thermal Vacuum Test. The TEM has to be very similar to the flight hardware from the thermal point of view. But very expensive flight components are replaced in the TEM by thermal equivalent dummies. This makes it possible to use all components of the TEM for an unusual but simple experimental determination of their heat capacity as well.

The DLR (German Aerospace Center) plans to launch the microsatellite BIRD (Bi Spectral Infrared Detection) in 1999 as part of a Earth remote sensing mission with hot spot detection as matter of priority. This project represents the begin of a line of small satellite missions with ambitious scientific and technological objectives by application of new technology and respecting the limitations of microsatellites. The spacecraft bus design is based on the proposed orbit and the payload requirements. The scientific payload is a novel multi-spectral sensor system, consisting of two cooled infrared sensor arrays and the Wide Angle Opto-electronic Stereo Scanner (WAOSS). A serious constraint of the satellite design is the required compatibility to a piggyback launch. The concept of the satellite bus fits to the requirements with the satellite dimensions of about 550x610x620 mm3 and a total mass of approx. 85kg.

For the determination of the solar absorptance αs and the thermal emittance ε optical methods are preferred. These optical measuring procedures are relatively exact. They also deliver information about any potential spectral selective behaviour of surfaces. However, special measuring equipment is required. Particularly if space simulation facilities are already available, the method presented in [1] appears as a suitable enrichment of the measurement methods for the determination of the thermooptical properties αs and ε. It is possible with slight technical expense to expand the usage potential of such facilities. For this paper the same test rig as in [1] is used. In a common space simulation chamber (vacuum; cryogenic shroud; solar simulator) a test item (target) is arranged. In different measuring phases the target is heated up by an integrated electrical heater or by solar irradiation. The arrangement (target / shroud) is regarded as a two node model.

For the determination of the solar absorptance αs and the thermal emittance ε preferably optical methods are utilized. To these methods belong spectral reflectance αs and ε measurements and / or emission measurements (only ε), whereby for αs must be measured in the common range 0,3 μm to 2,5 μm and ε in the range 5 μm to 35 μm with an essentially better resolution than 10 nm. Besides so-called integrated reflectance measurements are known, for which special detectors are used, in the above indicated spectral ranges. These optical measuring procedures are relatively exact. They deliver also information about a perhaps spectral selective behavior of surfaces, however, special measuring equipment is required.

An approach to the creation of passive radiative cooling system ensuring temperature levels less than 220K for the optical sensor of scientific space equipment elements is considered. The system is intended for the arbitrary orientation-in-space function under solar radiation. Theoretical analysis of the application field of this system, using heat pipes with constant and variable thermal resistance in a range of solar constant variation (500…2700)W/m2 is given. Experimental results on system models, in which two engaged in parallel thermodiodes with freon-22 and ammonium were used, showed the possibility to attain device temperature levels less than 220 K at the solar constant magnitude 1400 W/m2 and device heat release under (1…2) W.