RU2436716C2 - Method of producing spaceship thermal control system and device to this end - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to methods of devices of spaceship thermal control. Proposed method and device proceeds from exploiting the difference in coefficients of linear thermal expansion of spaceship body 1 (from aluminium alloy) and metal rod 3 (for example, from invar alloy). Spaceship is provided with driven shutter flaps 6 with heat isolation 2. Flaps drive comprises drums 8 fitted on flap axles 7 whereon thread 9 is wound, rod 3 secured on body 1 and converter 5. The latter has its one end coupled with said thread 9 and its opposite end coupled with rod 3. It serves to convert linear difference in coefficients of thermal expansion of spaceship body 1 and metal rod 3 into tensioning or loosening of thread 9 during rotation of drum 8 in opening or closing flaps 6. Device 5 may made up of pulley block suspended to bracket 4, or of a system of bell cranks.

EFFECT: simple and reliable system, power saving and reduced weight.

2 cl, 6 dwg

Description

This technical solution relates to the temperature control of spacecraft (SC) and can be used in rocket and space technology.

A known method of thermoregulation of spacecraft (see information on spacecraft under the program "Hurricane", the journal "Cosmonautics News" No. 02, 1999, the article "In flight three" Hurricanes "and the abstract" Satellite navigation systems GPS and Glonass ", l.19, 20), in which the aluminum case of the sealed container, containing the service and scientific equipment, is protected from the outside by thermal insulation, provide it with shutter flaps protected by thermal insulation, and rotationally driven to close-open by electric drives at the commands of the control unit of the system lovogo mode (BUSTR) sealed container by signals from the temperature sensors.

During operation of office and scientific equipment, it is required to maintain a certain thermal regime, therefore, when the temperature of the gas container’s atmosphere increases from the radiation of operating devices with temperature sensors, a command is sent to the control unit of the thermal mode of the pressure container to turn on the electric drives that open the shutter shutters and the thermal radiation surfaces of the cover of the pressure container provide unprotected thermal insulation emission of excess heat by radiation. When the temperature of the gas environment of the pressurized container is decreased by temperature sensors, the thermal control unit of the pressurized container receives a command to rotate the electric drives in the opposite direction, which closes the shutters, the heat release by unprotected thermal insulation by the thermal radiation surfaces of the pressurized container shell stops.

This method of thermoregulation of the spacecraft has significant disadvantages:

- spacecraft requires equipment to control the temperature of the gas environment of the pressure container and to control the electric drives;

- a large number of temperature sensors are needed;

- electric drives are needed;

- additional power supply is required for electric drives in the form of solar batteries and additional backup power supply in the form of rechargeable batteries;

- a large mass of components of the spacecraft thermal control system;

- a low degree of reliability of the spacecraft thermal control system due to the large number of its elements and the possibility of failure of one of them;

- the great complexity of tuning the spacecraft thermal control system;

- irrational use of radiation energy from operating spacecraft devices.

The objectives of this technical solution are:

- improving the reliability of the spacecraft temperature control system by reducing the number of elements of the spacecraft temperature control system and by increasing the reliability of the elements of this system;

- weight reduction of the spacecraft thermal control system;

- exclusion from the thermal control system of spacecraft electric drives;

- exclusion of power supply elements from the spacecraft thermal control system;

- providing automatic thermal control of the spacecraft;

- rational use of radiation energy of working devices.

It is known that the spacecraft body, usually made of aluminum or its alloys, has the largest constant thermal coefficient of linear expansion and when the spacecraft body temperature changes, its thermal deformation occurs. This cyclic thermal deformation of the expansion of the spacecraft body, obtained when it is heated from the radiation of the working spacecraft devices, and the thermal deformation of reducing the size of the spacecraft body by cooling it when the shutters are opened, are not used rationally in the overall balance of power consumption of the spacecraft.

To eliminate these shortcomings in the use of parasitic energy of heat radiation by instruments, it is proposed to install a rod made of a material with a significantly lower thermal coefficient of linear expansion, in contact with its thermo-radiation surface along its larger size, by attaching one end of the rod to one of the end walls of the spacecraft, and attaching the other end of the rod to a device that converts the thermal linear deformation of the spacecraft body into increased linear displacements to control the sash E shutters fixed to the opposite end wall of the spacecraft body. It is most rational to apply the method of thermoregulation of the spacecraft using the Invar brand H36 as a material for the alloy core, whose thermal expansion coefficient is not more than α = 1.5 × 0.000001 / deg in the temperature range from minus 60 ° С to plus 100 ° С.

Figure 1 shows the thermostatic device of the spacecraft according to the proposed method in the initial position with the shutters of the blinds closed (view of the spacecraft from above).

Figure 2 shows the thermostatic device of the spacecraft according to the proposed method in the initial position with the open shutters of the blinds (view of the spacecraft from above).

Figure 3 shows the thermostatic control device KA in section AA (section AA in figure 1).

Figure 4 shows the graph of the node B (a variant of the device that converts the thermal linear deformation of the spacecraft into increased linear displacements, for example, a chain hoist with elements of a thermal control device).

Figure 5 shows the graph of the node B (a variant of the device that converts the thermal linear deformation of the spacecraft into increased linear displacements, for example, as a device of a two-arm leverage system with elements of a temperature-regulating device). The device is shown in the position of the final phase of the spacecraft cooling.

Figure 6 shows the graph of the node B (a variant of the device that converts the thermal linear deformation of the spacecraft into increased linear displacements, for example, in the form of a device of a two-arm lever system with elements of a temperature-regulating device). The device is shown in the position of the initial phase of the spacecraft cooling.

The spacecraft thermal control device formed by the proposed method is shown in FIG. The spacecraft thermal control device consists of a spacecraft body 1, heat insulation 2 of the body, rod 3 with the lowest thermal coefficient of linear expansion, an arm 4 of the rod 3 fastening to the spacecraft body 1, device 5, which converts the space heat deformation of the spacecraft into increased linear displacements used to rotate the shutters , the shutters of the shutters 6, the axis 7 of the shutters of the shutters 6, the drum 8, rigidly fixed to the axis 7, the thread 9, transmitting and converting linear movements into rotational ones from the device 5 to the drum 8 and springing gear 10.

The temperature control device of the spacecraft in the initial position is shown in figure 1. The shutters of the shutters 6 with thermal insulation elements 2 close the space surface of the spacecraft and the threads 9, mounted on the drums 8, hold the shutters 6 with a spring 10 in such an equilibrium position, while the thermal regime of the casing 1, which is closed on the outside by thermal insulation 2, is in a predetermined regulated temperature range of temperatures . When heated, the spacecraft body 1 and the rod 3 in contact with it extend, but since the aluminum body of the spacecraft lengthens much more than the rod 3 of Invar 36H alloy, the difference in thermal deformation of the body of the spacecraft 1 and rod 3 stretches the device 5 (for example, the chain block ) for the rigging units, and the thread 9 “retracting” and distributed on the pulley blocks rotates the drum 8 together with the shutter 6 on the axis 7, opening the thermo-radiation surface of the spacecraft 1 and stretching the spring 10. The thermo-radiation surface of the body open by the shutters 6 SC 1 and resets the heat radiation and the temperature of the spacecraft body 1 decreases.

With decreasing temperature, the body of the spacecraft 1 and the rod 3 being in contact with it decrease in size due to the release of heat by radiation. Since the aluminum body of the spacecraft decreases in size larger, the thermal voltage to the device 5 weakens, the tension of the thread 9 decreases and the spring 10, compressing, pulls the thread 9 from the drum 8. The shutters 6 close the thermal radiation surface of the spacecraft body 1. Heat release stops. A cycle is heating the spacecraft body and then cooling it is automatically completed.

As a device for increasing thermal linear displacement, you can use a two-arm lever, but it will have large dimensions, because To convert small thermal deformations into increased linear displacements, a large ratio of the lever arm ratio is needed. To reduce the size of the lever device, it is necessary to use a system of two-shouldered levers in it. Figures 5 and 6 show a diagram of a lever device with a two-arm leverage system.

Using the method of thermal control of the spacecraft's body due to its thermal deformation allows:

- reduce the mass of components of the spacecraft thermal control system;

- exclude electric drives from the spacecraft thermal control system;

- to exclude power elements from the spacecraft thermal control system;

- provide automatic thermal control of the spacecraft;

- rationally use the radiation energy of working devices.

Claims (2)

1. The method of thermal control of the spacecraft (SC), which consists in the fact that the spacecraft body is protected by thermal insulation, it is supplied with shutters that are protected by insulation and rotationally actuated for closing-opening drives, characterized in that the rotary drives of the shutters of the shutters, providing their rotational movement for closing-opening is performed in the form of a device with torque-driven drums rigidly fixed to the axes of the shutters of the blinds, and with a thread wound on the drums, which provide from the drums together with the shutters of the blinds due to its tension-weakening device, which converts the difference in linear thermal deformation of the aluminum spacecraft body and the metal rod with a lower coefficient of linear expansion into increased linear displacements, which transmit the specified difference in linear thermal deformation of the spacecraft body and the metal rod, and the device , converting this difference into increased linear displacements, is performed, for example, in the form of a chain-link or a two-arm leverage system. 2. A spacecraft thermal control device comprising a spacecraft body, thermal insulation of the outer surface, a heat-radiation surface covered by shutters of shutters with a heat-insulating coating, and a drive for opening and closing shutters of shutters, characterized in that the drive of opening and closing of shutters is a device using the difference in temperature deformation of aluminum spacecraft body and metal rod with a lower coefficient of linear expansion, attached at one end to the spacecraft body, and the other end to the device, increasing the linear difference between the thermal deformation of the spacecraft body and the metal rod, and from this device the increased linear difference of thermal deformation by a flexible connection is transmitted to the drums, each of which is rigidly coaxially connected with the involuntary, for example, spring, axis of the shutter leaf, which when heating-cooling the spacecraft body and metal rod, respectively, opens and closes, providing regulation of the thermal regime of the spacecraft, and a device that increases the linear difference in thermal deformation of the spacecraft’s body and metal of the rod is, for example, chain hoist system or two-armed levers.

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