Designing passive thermal control system with a capacity of up to 3 kW by heat pipes and active heating elements for a spacecraft

Aeronautical and Space-Rocket Engineering


DOI: 10.34759/vst-2022-1-67-80

Аuthors

Shilkin O. V.1*, Kolesnikov A. P.1**, Kishkin A. A.2***, Zuev A. A.2****, Delkov A. V.2*****

1. Compani «Information satellite systems of academician M.F. Reshetnev», 52, Lenin str., Zheleznogorsk, Krasnoyarsk region, 662972, Russia
2. Siberian State University of Science and Technology named after academician M.F. Reshetnev, 31, Krasnoyarsky Rabochy av., Krasnoyarsk, 660014, Russia

*e-mail: shilkin@iss-reshetnev.ru
**e-mail: kolesnikov@iss-reshetnev.ru
***e-mail: spsp99@mail.ru
****e-mail: dla2011@inbox.ru
*****e-mail: delkov-mx01@mail.ru

Abstract

The thermal control system (TCS) is intended for maintaining the required thermal conditions of all spacecraft elements and onboard equipment.

The spacecraft TCS designing is a significant part of the spacecraft general engineering. This is due to the fact that the TCS is a deeply integrated spacecraft system interrelated with main onboard systems, environment, structural elements and flight tasks.

It is necessary to account for the thermal loads from the onboard equipment, radiation and re-radiation from the Sun and planets, and many other factors while designing a spacecraft. With relatively small thermal capacities, the spacecraft has a leaky design and the TSR is being designed on passive means of thermo stating. Application of thermal models with lumped parameters is widespread in the design of spacecraft onboard equipment. This approach appropriateness is confirmed by the practice of various units of a spacecraft TSR electronic equipment designing, analyzing and testing. The presence of telemetry parameters creates the possibility and directions for techniques optimization for the spacecraft TSR with improved qualitative mass-energy characteristics design.

The most common liquid TSRs display the essential fault in terms of specific mass-energy characteristics due to the greater mass of a coolant fueling, employing only heat-capacitive heat accumulation, as a consequence of the vapor phase inadmissibility at the contour centrifugal pump, though both models and heat balances of such systems are elaborated enough.

The presented article deals with an approach to the design of structural schemes for the spacecraft thermal control system with passive coolant pumping with of at least 3 kW of thermal power productivity. Three options were considered herewith.

The first option studies application of the thermal control system based on heat pipes, installed on the radiating panels. The heat-emitting devices herewith is installed on the backside of the radiating surface, and heat pipes distribute the heat along the panels’ surface transferring heat from one panel to the other.

The second option suggests the device in the form of the central heat bus, in which the heat-emitting devices are located on the common cooling panel, and uncontrolled heat pipes are embedded into the board being cooled and carry the heat from the electronic equipment to the passive heat transfer device in the form of the capillary pump.

The heat transfer unit of the third option does not contain flexible pipelines, and connects the electronic equipment board with the emitting radiator by the rigid pipelines. To provide the possibility for temperature control of the board being cooled, the heat pipes’ condensing zones of the cooled board and emitting radiators are connected by the gas-regulated heat pipes.

As far as the system with passive coolant pumping is under consideration, such criteria as energy consumption, operability range, control accuracy and reliability for all options are practically the same, and dominant evaluation criterion is the mass, which computing for all three options is presented. The computational results revealed the first option advantage, for witch specific mass-energy characteristic was ~33 kg/kW (without considering the ration of a certain part of the mass to the load-bearing structure mass).

The results of the performed comparative analysis allow drawing a conclusion that at the spacecraft equipment thermal load up to 3 kW, the most optimal is the thermal control system, which design scheme is based on application of the exclusively axial heat pipes.

Keywords:

heat output, thermal control system, heat pipes, capillary pump, two-phase circuit

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