This article presents a methodological approach to designing a spacecraft thermal control system with coolant pumping with a cooling capacity of up to 7.0 kW. Two design options are considered. There is no airtight instrument container in the spacecraft layout scheme, and all heat-generating equipment is located directly on the power structure panels, so excess heat is removed from the spacecraft directly from the outside of the instrument panels.
With all the attractiveness of a two-phase circuit with a heat pipe circuit, its use in automatic spacecraft is complicated by the need to supply concentrated heat to the capillary evaporator of the circuit. To do this, it is necessary to collect heat from a large surface of the structure, on which a large number of heat sources are installed.
A schematic solution of the thermal control system is considered, in which the thermal power of the payload module is distributed between the panel of the service systems module and deployable radiators. At the same time, in the first version, the heat pipes of the payload module are connected to the heat pipes of the service systems module along the profile shelves, the contour heat pipes of the deployable radiator are connected only to the heat pipes of the service systems module. Ums, the heat load of the payload module is transferred to the heat pipes of the service systems module and then to the loop heat pipes.
The second option differs from the first one in that to equalize the temperatures of the panels of the payload module, each heat pipe of the North panel is connected to the heat pipes of the South panel.
From the results of a comparative analysis of the mass budget and energy efficiency, the conclusion follows: the DFC with capillary pumping is the most preferable option 2. which, with the same other characteristics, has a smaller mass. The specific mass-energy: characteristic of such a system is ~22.9 kg/kW.
