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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.

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.

Description of design principles of thermocontrol system for small subsatellites Magion- 4, 5 and thermal behavior during two years of flight exploitation are presented. Passive type of thermocontrol system is intended to support the temperature level (-10…50) °C in inner volume of equipment compartment at solar subsatellite orientation and in temporary Earth shadow. Heat energy balance of the object at required temperature range is provided by utilization of radiating surfaces with defined areas, by thermal coupling of hot and cold subsatellite parts and by adaptation of constructive subsatellite elements to solve thermal tasks. Survey of telemetric date concerns temperature state of the main major parts of satellite are under consideration as well.

Description of design principles of thermocontrol system for small subsatellites “Magion - 4, 5” is presented. Passive type of system is intended to support the temperature level (-10…50) °C in inner volume of equipment compartment at solar subsatellite orientation and in temporary Earth shadow. Heat energy balance of object at necessary temperature range is provided by utilization of radiating surfaces with defined areas, by thermal connection of hot and cold subsatellite parts and by adaptation of constructive subsatellite elements to solve thermal tasks. Stages of heat pipe design, its development and testing separately and installed into object, thermovacuum test of subsatellite are presented. Initial time range of flight operation of thermocontrol system is analyzed as well.

The approach of finite element method (FEM) for simulation of thermal state of space apparatus and heat pipes integrated is presented. The practical tool discussed is 2D finite software package “HEAT'90”. Questions dealing with modeling of 2D construction, accuracy of simplification of 3D objects to pseudo 3D are under consideration. Several practically used cases of heat pipes connections with surrounding are analyzed: heat pipes are installed onto surface of device, heat pipes are inset into honeycomb panels, heat pipe crossing. Samples of simulation of heat pipe network for honeycomb panels illustrate the range of possibilities to forecast the thermal state of complicated ramified systems by using non traditional numerical FEM analyzer.

The questions of the gas filled heat pipes' application for thermal control systems of scientific equipment are discussed. It is analyzed different extents of electronic components' integration: creating of thermal stability mounting places of devices; creating of cooled planes and surfaces on device's body; providing of thermal stability of internal components. It is proposed design decisions providing of compensation some variable parameters such as device heat flows, external heat influences that are typical for near-earth orbits. Tests' results have shown the principal ability to construct schemes for thermal control of electronic components.

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.