Павел Григорьевич Гакал
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Геннадий Александрович Горбенко
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Эдем Русланович Решитов
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Рустем Юсуфович Турна
The world trend in the development of space vehicles is the expansion of their functionality, which leads to an increase in the power consumption, most of which is allocated in the elements of spacecraft equipment in the form of heat. To remove heat from the equipment elements, transfer it to the heat sink subsystem with subsequent removal to outer space, and also to maintain the required temperature mode of the equipment operation, thermal control systems are used. The increase in the power-to-weight ratio and linear dimensions of new spacecraft in conditions of severe design and weight-and-size limitations leads to a complication and growth of the mass of the system of thermal control of space vehicles. At present, thermal control systems for space vehicles based on single-phase fluid heat transfer loops are used. For space vehicles with an energy consumption of more than 10 kW, thermal control systems based on two-phase heat transfer loops are the most promising. They have a number of advantages in comparison with single-phase thermal control systems: two-phase heat transfer loops can transfer much more heat per unit of flow; the use of heat transfer during boiling allows to maintain the temperature of objects practically on the whole extent of the circuit close to the saturation temperature; the mass of the thermal control system with a two-phase coolant is substantially less than with a single-phase coolant , and the energy consumption of the pump for pumping the coolant is negligible. In this paper, a two-phase heat transfer loop performances are analyzed. The process of increasing the thermal power up to the maximum under conditions of full filling of the accumulator is considered. The study was carried out on an experimental two-phase heat transfer loop with an ammonia. Transient processes associated with an increase in the thermal load from 73 % to 100 % are considered. The obtained data correlate well with the results of the calculation. Based on the results of the analysis, conclusions were made on the operability and stability of the spacecraft thermal control system under these conditions, and recommendations on the choice of the volume of the accumulator are given.
Design and Analysis of the Location of Cold and Heat Sources for the Payload’s Active Thermal Control System on the Lunar Surface
