RU2196079C2 - Spacecraft - Google Patents

Abstract

FIELD: space engineering; designing spacecraft. SUBSTANCE: spacecraft has target equipment compartment, instrument compartment, power equipment compartment with engine plant, power supply system with solar battery, automation and voltage stabilization complex, storage batteries mounted on heat exchange units forming monoblocks, temperature control system with hydraulic loops and units for taking, supplying and removing heat in form of thermal plates and emitting external radiators included in loop of said temperature control system. Monoblocks are provided with thermal pipes located in parallel at preset distance relative to each other. End parts of these pipes form condensation areas; they are combined in separate thermal plates by means of plate made from heat-conducting material. Tubes of capillary structure secured to thermal plates through heat-conducting material perform function of capillary pump forming evaporation zone of thermal tube. Groups of thermal plates are interconnected in parallel and are connected to external emitting radiator by means of pipe lines, thus forming autonomous loop. EFFECT: increased service life; reduced mass; enhanced reliability. 1 dwg

Description

 The invention relates to space technology and can be used in the design of spacecraft (SC).

 In space technology, among others, the task is to increase the active life of the spacecraft. The most difficult technical problems along this path arise in the design of automatic spacecraft.

 As you know, a spacecraft is a technical device, consisting, as a rule, of target equipment and supporting systems (Spacecraft. Under the general editorship of KP Feoktistov, M., Military Publishing House, 1993). Optical, optoelectronic, radio engineering or other systems can be used as the target equipment, depending on the type and purpose of the spacecraft, directly performing the task. The number of supporting systems includes: propulsion system, power supply system (BOT), on-board control complex (BKU), temperature control system (CTP) and other systems.

 Among the systems of modern spacecraft, significantly affecting its period of active existence are, first of all, continuously functioning, in particular BOT and STR. In the power supply system, the weak link is the batteries (AB), and in the STR - mobile electromechanical devices (fans, pumps, regulators) and hydraulic lines.

To extend the service life (resource) of the battery, it is necessary to provide the required temperature conditions in a relatively narrow range during their operation. For low-orbit spacecraft, the optimal range of operating temperatures of ABs, characterized by relatively high charge and discharge currents, is, as a rule, 10-25 ^o С (Thermal control system and battery performance for Western European satellites. Astronautics and rocket dynamics, express information, 6 , pp. 23-29, 1989). In addition, the number of charge-discharge cycles significantly affects the battery life.

 The reliability of the CTP depends on many factors, including the length of the hydraulic lines, which determines the degree of loading of the pumping pumps. At the same time, it is extremely undesirable to disconnect the hydraulic lines during the spacecraft tests both at the plant and in the operating organization.

 Such operations can lead to latent defects of the STR: damage, wear of the sealing joints, chemical change in the coolant, contamination and corrosion of the hydraulic lines - this will lead to failures of the STR, manifested in the normal operation of the spacecraft.

 A known temperature control system for an artificial Earth satellite (US Patent 4880050, F 28 D 15/00, 1989, analog), which is equipped with thermal boards for efficient heat removal from the equipment to external radiation panels. In this technical device, the most important devices can be installed on thermal boards, through the channels of which the liquid coolant circulates. This method of cooling devices can significantly narrow the range of operating temperatures to optimal values and ensure the thermal regime of on-board equipment located in both sealed and non-sealed compartments, which positively affects the resource characteristics of the spacecraft.

 The disadvantage of the analogue is that the use of thermal circuits with a circulating coolant leads to an increase in the mass of the spacecraft, since they have a significant area and thickness. In addition, the use of thermal cards with connecting them to the STR line degrades the characteristics of the latter, since the length of the hydraulic line increases with a corresponding increase in hydraulic resistance, the reliability of the STR decreases, since the depressurization of any thermal card unambiguously destroys the STR as a whole.

 Another disadvantage of the analogue is that during ground tests of the spacecraft, its STR should be constantly turned on to cool the spacecraft systems. Prolonged ground tests of the spacecraft in this case lead to a limitation of the STR resource, and, therefore, to a decrease in the active life of the spacecraft.

 Known spacecraft (RF patent 2144889 from 01/27/2000, prototype), containing a compartment with the target equipment, a sealed instrument compartment with an integrated propulsion system, a temperature control system with hydraulic circuits and devices for the selection, supply and discharge of heat, including those made in the form of thermal boards with hydraulic channels included in the circuit of the specified temperature control system, an electrical power system including a solar battery, an automation and stabilization complex installed in the instrument compartment of tension, as well as the batteries located in the aggregate compartment, along the length of each battery a monoblock forming a thermal plate is fixed with it, the monoblock fastening units to the construction of the aggregate compartment are made in the body of each battery, and additional hydraulic channels are connected to each other in the thermal plate of each monoblock using pipelines, while these channels and pipelines form an autonomous open line.

 This technical solution allows to partially eliminate the disadvantages of the analogue. During ground tests, the spacecraft can be thermostated without turning on the STR, thereby preserving the required STR resource for regular operation. In addition, it is possible to preserve battery life when using technological batteries during testing.

The disadvantage of the prototype is, first of all, that when replacing a technological battery with a standard battery or a failed battery with a working battery, it is absolutely necessary to disconnect the hydraulic line, followed by lengthy and costly leak tests of the CTP. This leads to significant inconvenience to ground operation of the spacecraft. The disadvantages characteristic of the analogue also remain. Indeed, the use of thermal boards leads to:
increase in the mass of the spacecraft due to the significant geometric dimensions of the design of the thermal board;
reduce the reliability of the STR due to the likelihood of depressurization of any thermal plate and the unique failure of the STR as a whole.

 The objective of the invention is to increase the reliability of the STR and reduce the mass of the spacecraft, ceteris paribus, as well as improving the convenience of ground-based operation of the spacecraft.

 This problem is solved by the fact that in the known spacecraft containing a compartment with the target equipment, a compartment with devices, an aggregate compartment with an integrated propulsion system, a power supply system including a solar battery, a complex of automation and voltage stabilization, rechargeable batteries installed on the corresponding heat exchange devices and combined with the latter in monoblocks, a temperature control system with hydraulic circuits and units for the selection, supply and discharge of heat, including in the form of thermal boards and radiation of external radiators included in the circuit of the specified temperature control system, each monoblock is equipped with heat pipes located parallel to a given distance from each other, the end parts of the heat pipes forming condensation zones are combined by a plate of heat-conducting material into separate thermal boards, with each thermal board through a heat-conducting the material is fixed to a tube with a capillary structure acting as a capillary pump and forming the evaporation zone of the contour heat pipe, each a group of thermal boards is connected in parallel with each other and connected to a radiating external radiator using pipelines, forming an autonomous circuit.

 The drawing shows the proposed device KA.

 It consists of a compartment with target equipment 1, an instrument compartment 2, a temperature control system 3, an aggregate compartment 4 with an integrated propulsion system 5 installed in it, a power supply system containing control equipment 6, a solar battery 7 and rechargeable batteries 8. Rechargeable batteries 8 and corresponding heat exchangers 9 form monoblocks, since the latter are attached to the AB with tightening screws (not shown). Monoblocks are installed on the construction of the aggregate compartment 4. The heat exchange device 9 is made of heat pipes 10, which are located in the same plane in parallel at a predetermined distance from each other. In this case, the heat pipes 10 have a flat surface for installing AB 8 on them. The surface of the heat pipes under the AB 8 forms a thermoplate-evaporator 11. Moreover, the ends of the heat pipes 10 are rigidly fixed on both sides and combined by a plate of heat-conducting material to form a separate thermoplate 12. A tube 13 with a capillary structure that acts as a capillary pump and forms the evaporation zone of the contour heat pipe 14 is fixed to each thermal plate 12 through a heat-conducting material, for example, paste (not shown). Each group moplat 12 connected in parallel with each other, forms an autonomous circuit, which also includes channels 15 of the radiating external radiator 16 and pipes 17. A contour heat pipe is formed by serial connection of the thermal card 12, pipelines 17 and channels 15 of the radiating external radiator 16. Heat exchange the device 9 for structural reasons may contain the transport zone of the heat pipes 10, if the thermal card-evaporator 11 and the thermal card 12 cannot be installed at a close distance from each other.

 Heat pipes 10 and pipes 13 are partially filled with liquid ammonia and sealed. When AB 8 releases a certain amount of heat, liquid ammonia, evaporating, takes heat from it and AB 8 cools.

 In the heat release zone of AB 8 there is a zone of evaporation of heat pipes, the distance between the heat pipes being determined by the density of the heat flux from AB 8 to the evaporator thermal plate 11. The condensation zones of the heat pipes 10 of each thermal plate 12 are connected by a plate of heat-conducting material. To ensure guaranteed cooling of the radiator 16, certain optical coefficients are provided on its surface.

 The principle of operation of the means for ensuring the thermal regime of AB 8 is as follows. The heat energy generated by AB 8 is transferred to the heat pipes 10. The heat carrier in the form of liquid ammonia circulating inside the heat pipes 10, sensing this heat energy, partially or completely evaporates while maintaining the set temperature of the installation site AB 8. The evaporated heat carrier enters the condensation zone of the heat pipes 10, where the condensation of the coolant occurs with the transfer of thermal energy to the heat-conducting plate. Further, due to conductive heat transfer, thermal energy is transferred to the evaporation zone of the contour heat pipe to the capillary pump, where the heat carrier of the contour heat pipe evaporates. The capillary pump is a tube 13 with a capillary structure (see USSR author's certificate 1779648) and allows you to overcome high hydraulic resistance during movement of the coolant, thereby spreading the evaporation and condensation zones of the contour heat pipe over long distances. The coolant vaporized in the evaporation zone of the contour heat pipe enters the channels 15 of the external radiating radiator 16, where it is condensed and heat is released into outer space. Condensed coolant due to the discharge pressure developed by the capillary pump flows through the tube back into the evaporation zone of the loop heat pipe. Then the process is repeated again.

 The proposed design of the spacecraft allows you to fully and effectively solve the problem.

 A decrease in the mass of the spacecraft is achieved through the use of relatively thin and light heat pipes 10 as the heat transfer element of the heat exchange device 9 of the batteries 8. No additional devices are required for cooling the AB 8 during ground tests. Since the batteries 8 are fixed along the wall of the aggregate compartment, the heat pipes 10 are almost upright when testing the spacecraft, and their evaporation zone will be lower than the condensation zone of one of the edges of the heat pipes 10. This means that the liquid phase of ammonia is below, in contact with the thermal conductivity of the battery 8, gaseous - at the top, in contact with the thermal conductivity of the thermal plate 12. Thus, in terrestrial conditions, the heat exchanger 9 works in the same way tively, as its heat pipes 10 functions as a thermosyphon.

 An increase in the reliability of the STR, and hence an increase in its resource, is achieved by reducing the length of the STR line, since the heat pipes 10 and duplicated contour heat pipes, including channels of the external radiating radiators 15 (not shown), are not hydraulically connected to the STR lines, and also due to the absence of an effect on the operability of the CTP of the fact of depressurization of a certain number of heat pipes 10 or even one of the channels 15 of the external radiating radiator 16.

 The increase in battery life 8 can be achieved through the use of technological AB 8 during testing, since the installation and dismantling of AB 8 do not require disassembly and depressurization of the STR hydraulic circuits. This fact helps to increase both the reliability of the STR and the convenience of the surface operation of the spacecraft.

 Thus, the use of the proposed spacecraft will significantly improve the reliability of the STR and the spacecraft as a whole and reduce the mass, as well as improve the convenience of ground-based operation of the spacecraft.

Claims (1)

 A spacecraft containing a compartment with target equipment, a compartment with devices, an aggregate compartment with an integrated propulsion system, a power supply system that includes a solar battery, a complex of automation and voltage stabilization, batteries installed on the respective heat exchange devices combined with the latter in monoblocks, thermal control system with hydraulic circuits and units for the selection, supply and discharge of heat, including in the form of thermal boards and radiating external radiators, including included in the circuit of the specified temperature control system, characterized in that each monoblock is equipped with heat pipes located parallel to a given distance from one another, the end parts of the heat pipes forming the condensation zones are combined by a plate of heat-conducting material into separate thermal boards, with each thermal board through a heat-conducting the material is fixed to a tube with a capillary structure that acts as a capillary pump and forms the evaporation zone of the contour heat pipe, each group termoplat connected in parallel to each other and connected to an external radiator radiative via pipelines, forming an autonomous circuit.