RU2264954C2 - Spacecraft at regular orientation relative to sun - Google Patents

Abstract

FIELD: space engineering; spacecraft flying in geostationary or high-altitude elliptical orbits.

SUBSTANCE: proposed spacecraft has module case with projecting members. Two opposite faces of each module perform function of radiators with built-in thermal tubes. Arranged in modules are engine unit and some heat-loaded units and onboard devices (number n). Other units (number k) , for example metal-hydrogen storage batteries are secured to engine unit and are heat-insulated from first units. Units and devices are secured to engine unit by means of brackets through heat-insulating gaskets. Unit is made in form of three-layer honeycomb panel where thermal tubes with heaters are laid. Each of k-units has thermal contact with axial U-shaped thermal tube embracing them. This thermal tube is brought in contact with evaporator of loop thermal tube connected with radiator by means of vapor line which is communicated with loop thermal tube and its evaporator through condensate lines. Radiators are mounted beyond boundaries of projecting parts shading zones on side of extravehicular space.

EFFECT: increased cooling effect of spacecraft temperature control system; reduction of mass of this system.

2 cl, 4 dwg

Description

The invention relates to space technology and can be used in the development of spacecraft (SC), in which the orientation relative to the Sun is regular, for example, a spacecraft in a geostationary orbit (GSO), a spacecraft with a solar-terrestrial orientation in circular highly elliptical orbits, etc.

As an analogue, the spacecraft described in patent RU 2053937 can be adopted.

On this spacecraft, the basis of which is a sealed hull and which is equipped with a solar power installation, to increase the reliability of operation and improve the bulk-mass characteristics, additional non-tight containers are installed in the form of a parallelepiped with a small height, with batteries installed inside them, control and monitoring equipment for the solar power installation (SEU), means of ensuring thermal conditions. In this case, leaky containers are pivotally mounted on the spacecraft body and fixed in the unfolded position relative to the spacecraft body using clamps. The bases of the storage batteries (AB) and the SEU equipment are connected by heat transfer elements to the bases of the leaky container, which are emitters for radiative heat release.

The disadvantages of the considered analogue include:

AB and SEU are placed on a single base of an unpressurized container, and therefore, are combined in a heat ratio. Considering that the permissible operating temperature ranges of the battery and electronic components of the SEU equipment can vary significantly (the temperature on the basis of electronic devices can be maintained at + 40 ° C, while for comfortable charging and storage conditions, the temperature under the battery should not be exceed + 10 ° C), the radiation surface areas of leaky containers will be oversized, which will require additional energy costs for heating the installed equipment.

The presence of additional mechanisms for transferring leaky containers from the transport position to the working one, followed by the opening of solar panels (SB) hinged on the side surface of the leaky containers, complicates the design and control system of the spacecraft and introduces an additional element of unreliability.

As a prototype, as the closest in technical essence, a spacecraft with a regular orientation relative to the Sun can be adopted, protected by patent RU 2092398.

The spacecraft contains a body (container), consisting of modules installed along the longitudinal axis, which can be a U-, U- and H-shaped configuration. Inside the hull there are n (n = 1, 2 ...) heat-loaded nodes and devices of on-board multipurpose payload equipment and service systems. Two opposite sides of the spacecraft's hull (sides of the U- and U-shaped modules) with the nominal orbital orientation of the spacecraft on the GSO are deployed from the direction to the Sun and directed towards the north and south of the Earth.

These faces are radiator honeycomb panels with integrated heat pipes (TT). Heat-loaded n-th nodes and spacecraft devices are directly installed on the faces.

The spacecraft also contains an engine block with engines of the correction, orientation and stabilization system, fuel tanks and working fluid supply lines. In addition, the apparatus contains protruding relative to the housing structural elements. These include SB panels, antennas of on-board transponders of the "C-" and "Ku-" range. Moreover, the two wings of the folding panels of the solar battery are installed symmetrically to the longitudinal axis from the side of the "northern" and "southern" radiators and can rotate relative to the spacecraft hull.

One of the modules (in the prototype - the module of the H-shaped configuration) contains two additional radiator honeycomb panels with built-in heat pipes from the "east" and "west" sides falling into the sun lighting zone, enclosed in a single reversible heat circuit with the "western" and " eastern "honeycombs.

Heat-loaded devices and components of the on-board equipment are located on the central transverse (“instrument”) cell block panel.

In the module of the H-shaped configuration, the CTs on the east and west honeycomb panels are connected to the unified network of CTs of the instrument panel honeycomb panels.

On the radiator honeycomb panels of the block of the H-shaped configuration, opening and compact folding mechanical drives are installed at the edges or in the middle of the radiators heat-insulating curtains.

Consider the main disadvantages of the prototype.

The entire thermal circuit for heat removal in the H-shaped unit can work only in alternating-periodic mode, since the external heat flux from the Sun will always affect one of the honeycomb panels - “eastern” or “western”. Moreover, so that the external heat flux from the illuminated honeycomb panel is not connected to the equipment located on the central honeycomb panel, the illuminated honeycomb panel must be periodically closed by curtains.

Thus, in terms of cooling capacity, approximately half of a typical TT network with two radiators is constantly operating in an H-shaped unit. The other half should be blocked.

With this in mind, as well as the presence on board of the mass of available two curtains, the following conclusion can be drawn: the required mass of cold-producing devices per unit of rated cold produced by them during the day for the "northern" and "southern" honeycomb panels will be more than 2 times less than for "eastern" and "western" honeycomb panels, subject to the use of the same panel design with TT. At the same time, we refer to heat-producing devices as heat pipes and curtains that protect the “western” and “eastern” honeycomb panels from overheating.

A significant disadvantage of the prototype is the fact that the thermal management system (COTP) is constructed in such a way that does not take into account the heat generation of individual blocks. So, in the prototype, all the heat pipes of each block are located in parallel planes with the same pitch (no more than 200 mm). At the same time, it is known that onboard equipment and individual components of spacecraft load spacecraft with heat differently. So, for example, power automation of regulation of the spacecraft power supply system, power equipment of stationary plasma engines during their operation, modules of metal-hydrogen batteries, etc. to a much greater extent they generate heat during their operation than electronic devices - digital computers, transceiver on-board equipment of the service control channel, etc. Therefore, it is necessary to differentiate to a certain extent the work of SOTR KA.

The objective of the invention is to reduce the mass of refrigeration units on board the spacecraft per unit of cold produced by them by increasing the cooling capacity of the spacecraft thermal control system.

The technical result is achieved in that in a spacecraft with a regular orientation relative to the Sun, containing a body with protruding structural elements, consisting of modules installed along the longitudinal axis, two opposite sides of each of which, which are radiators with integrated heat pipes, are deployed from the direction to the Sun, at the same time, in the modules there are n (n = 1, 2, ...) heat-loaded units and instruments of the on-board equipment, an engine block with engines of a correction, orientation and stabilization system , by fuel tanks and working fluid supply lines, in contrast to the known to the engine block k (k = 1, 2, ...) heat-loaded units and devices of on-board equipment, heat-insulated by screen-vacuum thermal insulation from n-heat-loaded units and devices on board, are attached equipment, and the indicated k-th nodes and devices of the on-board equipment are attached to the engine block by brackets installed on heat-insulating gaskets, while the engine block is made in the form of a three-layer panel consisting of two mounting plates a honeycomb filled filament with heat pipes with heaters inside, each of the kx components and devices of the on-board equipment is surrounded by an axial U-shaped heat pipe that has heat contact with their bodies, and each of the heat axial pipes, in turn, has heat contact with the evaporator of the loop heat pipe, the output of which through the steam line is connected to the input of the radiator-radiator, and the output of the radiator-radiator through the condensate pipe is connected to the input of the evaporator, while the radiator-radiators are installed in a flat bones parallel to the faces turned from the direction to the Sun, beyond the boundaries of the shading zones by the protruding structural elements of the space open to radiators-radiators.

In addition, in the specified spacecraft, the heat-loaded units are made in the form of metal-hydrogen storage batteries bodies of solid power plates of a honeycomb design, in the cylindrical openings of which metal-hydrogen batteries are installed.

The technical result of the invention is to increase the cooling capacity of the spacecraft thermoregulation system, achieved by placing TT radiators that remove heat from the most heat-loaded nodes and devices in planes parallel to the faces of the spacecraft deployed from the direction to the Sun. In addition, this effect is also expected to be achieved by using the area of radiators necessary and sufficient to solve the problem of heat removal from selected specific components and assemblies of the spacecraft.

To ensure the essence of the invention are further provided by figures 1-4.

Figure 1, 2 shows a spacecraft with a regular orientation relative to the Sun, while in a working geostationary orbit.

Figure 3 shows the structural block-modular diagram of the spacecraft.

Figure 4 shows the installation location on the spacecraft bodies of metal-hydrogen storage batteries.

Figure 1, 2 shows the spacecraft communications and broadcasting in a working geostationary orbit. The following notation is introduced:

1 - spacecraft;

2 - solar batteries (SB) KA;

3 - antenna repeaters;

4, 5 - body panels-radiators of the spacecraft, "north" and "south", respectively;

6 - radiators, emitters.

SB2 are installed along the longitudinal axis OZ of the spacecraft from the side of the "southern" and "northern" honeycomb panels of radiators, while the "+" OZ is directed towards the south pole of the Earth. Spacecraft repeater antennas are placed along the longitudinal axis of the spacecraft OY from the "eastern" and "western" side of the spacecraft panels. In addition, figure 2 conventionally shows the angular range of incidence of solar radiation relative to the GSO plane ± 23.5 °. As can be seen from the figures, radiators-radiators are installed in planes parallel to the planes of location of the "northern" and "southern" radiators. In the particular case, they can be in the same planes without obscuring the radiator panels themselves. The location of the radiators-emitters should be determined so that they are outside the shading zones by the SB panels and the antennas of outer space repeaters. In cases of installation on the spacecraft not of repeaters, but of another payload, shading of radiators-radiators by other structural elements protruding relative to the hull should not occur.

When these conditions are met, the refrigerating capacity of TT refrigerant condensers located inside the radiators-radiators will be maximum.

This follows from the fact that, depending on the time of the year, the angle between the plane of the orbit and the direction to the Sun varies within ± 23.5 °, and at angles ± 8.7 ° the orbit passes through the “shadow” of the Earth. The orbit is 24 hours. The orientation of the spacecraft during the flight is with the “X” axis along the radius vector to the Earth, while the “Y” axis is in the orbit plane and directed along the velocity vector, and the “Z” axis is perpendicular to it and directed towards the South Pole.

Thus, radiators-emitters will be constantly in the areas of minimum solar illumination, the least affected by the external heat flux of solar radiation.

Figure 3 shows the structural block-modular diagram of the spacecraft. In addition to the previously adopted, figure 3 introduced new notation:

7 - payload module (MPN);

8 - service systems module (MSS);

9 - sotopanel engine block (SDB);

10 - housing heat-loaded nodes and devices;

11 - additional brackets for fixing the cases of heat-loaded units and devices.

As can be seen from figure 3, the spacecraft is made in the form of a rectangular parallelepiped, docked in two parts. One part is MLN 7, which mainly houses the equipment of the airborne relay complex, the second - MCC 8.

The spacecraft body is formed by the "southern" and "northern" (respectively, along the "+ Z" and "-Z" axes) MCC and MPN panels, the "eastern" and "western" (respectively, along the "+ U" and "- U axes) ") panels, the upper end panel along the axis" + X "of the arrester and the lower end panel along the axis" -X "MCC. At the junction between the MPN and MSS is located SDB 9 with installed on it engines of orientation correction and stabilization systems, fuel tanks and supply lines of the working fluid.

The main heat-generating equipment is located on the “northern” and “southern” panels of MPN 7 and MSS 8. The MPN 7 contains spacecraft payload equipment (for example, an onboard relay complex), partially the apparatus of the motion control system (SUD) of the device (for example, external orientation sensors SC, etc.) MCC 8 houses an on-board digital computer complex, inertial executive bodies of the satellite orientation support system, logical-converting devices of the onboard equipment control system, and other devices and assembly s service systems of the spacecraft.

The spacecraft body, consisting of modules installed along the longitudinal axis, is oriented in space using the SUD, which uses the sensors of the spacecraft's external orientation for control, and jet engines or power gyroscopes as executive bodies. This ensures a regular orientation of the apparatus relative to the Sun (see figure 2), in which you can always distinguish two opposite faces, turned away from it (for GSO it is the "northern" and "southern" faces), and two faces regularly illuminated by the Sun ( for GSO, this is the “western” and “eastern” faces). Based on the indicated orientation, the design of spacecraft modules is also carried out. The faces of the modules, turned away from the Sun, are radiators for discharging heat from heat-loaded nodes and spacecraft devices. On SDB 9 there is an engine block with engines of the correction, orientation and stabilization system, fuel tanks, supply lines for the working fluid to the jet engines. The heat-generating equipment for controlling the engine block is mainly located in MCC 8. Therefore, the heat exchangers located in the SDB 9 have, as a rule, a heating zone with the aim of supplying heat to individual units and assemblies of the engine block to prevent their overcooling and even possible freezing.

These panels are made three-layer - two sheets of aluminum alloy, between which a honeycomb is installed. Three-layer is also made SDB 9.

The “western” and “eastern” panels, the lower and upper end panels are power frames, covered by single-layer sheets of aluminum alloy.

In contrast to the prototype ke (k = 1, 2, ...), the bodies of the most heat-loaded units and assemblies 10 are attached to the SDB 9 (bottom, see figure 3) through additional brackets 11. In order to exclude the influence of heat fluxes from the indicated buildings for equipment of n nodes and devices of on-board equipment, where n = 1, 2, ..., located inside the spacecraft, the indicated housings ke, where k = 1, 2, ..., are insulated from surrounding devices and structural elements when help multilayer screen-vacuum thermal insulation (EVTI) (not shown conditionally in figure 3), as well as heat-insulating gaskets s established, in particular, under the brackets 11.

The applied modular structure of the spacecraft with a regular orientation allows the separation of its structure into power and thermal parts. All heat-loaded payload devices and service systems are arranged on two radiation panels of the "northern" and "southern" faces of these modules. The proposed technical solution in the existing structure of modular division combines both of the above functions, since on the one side of the case of heat-loaded units and devices installed on the brackets of SDB 9, they are used as structural elements of the structure (for example, enhancing its rigidity), on the other hand, they is the heat transfer elements of the COTP apparatus.

Separation of individual heat-loaded components and devices from all spacecraft equipment and their special placement in separate hulls makes it possible to create “special working conditions” for them, providing increased heat dissipation, and to increase the locally cooling capacity of the thermal management system (COTP).

As one of the most typical, we can consider an example of ensuring the thermal regime of modules of metal-hydrogen storage batteries (MVAB).

So, in accordance with the technical conditions for operation when operating as a part of the spacecraft, the temperature of the MBAB hulls should be in the range from 5 to 25 ° C for charging and not more than 45 ° C for discharge.

Figure 4 shows how the thermal mode of operation of heat-loaded buildings, in particular MVAB, is provided, while the following additional notation is introduced:

12 - MVAB corps;

13 - axial heat pipes (ATT);

14 - evaporators of loop heat pipes (ICTT);

15 - contour heat pipes (CTT).

In addition, dashed lines indicate the areas of equipment placement on the "northern" and "southern" panels of the spacecraft.

ATT 13 is proposed to be made in a U-shape, covering the MVAB 12 housings and having thermal contact with ICTT 14, the KTT 15 capacitors are placed in the radiators-radiators 6. In turn, the installation of radiators-radiators in the above manner allows for maximum cooling capacity of the KTT 15 capacitors. Thus, the temperature regime is ensured for the two MVAB 12 buildings. Moreover, each of the buildings is made in the form of a solid power plate made of MAG-1 magnesium alloy. The plate itself is a honeycomb structure with cylindrical holes for installing metal-hydrogen batteries, while U-shaped ATTs are attached between the MVAB 12 and IKTT 14 casings using the adhesive sealant "Elastosil 137-242".

To regulate heat fluxes from heat-loaded spacecraft hulls, for example, COTP made on the basis of various types of heat pipes (transport and diode), configuration, internal structure, and profile can be used.

ATT 13 is used to collect excess heat from MVAB 12 and its transportation to ICTT 14.

In the lower part of the U-shaped ATT 13 during ground preparation, a technological electric heater is placed, which ensures the start of the ATT.

Each CTT is a sealed device consisting of:

a porous nickel evaporator with an average pore diameter of ~ 1.2 μm with longitudinal vapor channels;

a steam line made of a thin-walled stainless tube serving as a channel for transporting steam from the evaporator to the condenser;

bypass line with valve;

a condenser made in the form of a radiator-emitter, which serves to condense the coolant vapor and transfer heat to the radiating surface;

a condensate conduit made of a thin-walled stainless tube, which serves to transport the liquid coolant from the condenser to the evaporator.

The principle of operation of KTT 15 is based on the use of phase transformations of the coolant (evaporation ↔ condensation). The heat generated by the heat source (MVAB 12) enters the KTT evaporator and is spent on the evaporation of the working fluid. The resulting steam enters through a steam line into a condenser-radiator, where it condenses. The heat released during condensation is removed to the environment. The mover of the coolant in the CTT is a capillary pump, which sucks the condensed coolant into the evaporator. The amount of heat discharged is controlled by bypass lines with valves.

The head, which modern capillary pumps are capable of developing, reaches 0.6 kg / cm ² , which makes it possible to raise the heat carrier by 6 m against gravity. Thus, CTTs can operate in terrestrial conditions with an arbitrary arrangement of the panel and the evaporator relative to each other.

Compared with the prototype, the proposed design of the spacecraft has a number of significant advantages.

The optimal placement of radiation panels (radiators) is provided for the discharge of excess heat (which are simultaneously CTT condensers) in relation to external heat fluxes from the Sun and the Earth. The indicated arrangement of radiators allows minimizing the required mass of spacecraft structures, excluding that part of it that would require the installation of additional CTs and protective curtains on the radiators.

The effective refrigerating capacity of the system, obtained due to the proposed arrangement of radiators, allows for the release of heat from the most heat-stressed n-nodes and assemblies. Moreover, the areas of these panels can be different, necessary and sufficient to ensure the thermal regime of a particular device or unit. In the prototype, COTP should ensure the operation of the spacecraft with the maximum excess heat flux from the units and assemblies installed on the panels. In the case when not all devices are turned on (not all heat sources are involved), additional energy consumption may be required to heat the “instrument” and other panels indicated in the prototype. Otherwise, it is possible to subcool these devices and assemblies.

Ensuring the placement of cases of heat-loaded devices and assemblies (in particular, MBAU cases) on the power panel of the engine block allows optimizing the layout of the spacecraft compartments, while minimizing the mass of additional structures for installing these bodies and strengthening the strength of the spacecraft body itself.

Claims (2)

1. A spacecraft with a regular orientation relative to the Sun, containing a housing with protruding structural elements, consisting of modules installed along the longitudinal axis, two opposite faces of each of which, which are radiators with integrated heat pipes, are deployed from the direction to the Sun, while the modules are located n (n = 1, 2, ...) of heat-loaded units and instruments of on-board equipment, an engine block with engines of a correction, orientation and stabilization system, fuel tanks and supply lines body, characterized in that k (k = 1, 2, ...) of heat-loaded units and devices of on-board equipment insulated by screen-vacuum heat insulation from n-heat-loaded units and devices of on-board equipment are attached to the engine block, and ke and on-board equipment instruments are attached to the engine block with brackets mounted on heat-insulating gaskets, and the engine block is made in the form of a three-layer panel consisting of two mounting planes with honeycomb core, inside of which there is a heat pipes with heaters are provided, each of the kx components and devices of the on-board equipment is covered by an axial U-shaped heat pipe having thermal contact with their bodies, while each of the axial heat pipes, in turn, has thermal contact with the evaporator of the contour heat pipe, the output of which through the steam line is connected to the input of the radiator-emitter, and the output of the radiator-radiator through the condensate line of the loop heat pipe is connected to the input of the evaporator of this heat pipe, while the radiator-emitters are installed in loskostyah parallel faces deployed from the direction to the sun, outside the occluded areas projecting elements for the construction of open space emitters radiators. 2. The spacecraft according to claim 1, characterized in that said heat-loaded nodes are made in the form of metal-hydrogen battery packs of solid power plates of a honeycomb design, in which cylindrical openings metal-hydrogen batteries are installed.

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