RU2648520C2 - Space platform - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: invention relates to structure and arrangement of spacecrafts (SC), primarily space platforms (SP) combining service subsystems and providing operation of a payload module (PLM). SP contains an instrument compartment (IC) in the form of a rectangular parallelepiped with instruments and equipment (2), correction and orientation engines (3) (under the base of IC), rotary wings (5) of a solar panel and PLM assembly units. Some of the IC panels are radiators. The main power element for placing the SP and PLM is a central pipe (CP) (6), on which the IC is mounted. The ends of the CP (6) protrude beyond the plane of the IC panels. On the ends, there are docking points with a separation subsystem. Inside the CP, there are the tanks of the working fluid for the correction and orientation engines (3). Correction engines (3) can be located on the PLM panels (9).

EFFECT: weight reduction and increase in the density of the spacecraft arrangement created on the basis of the spacecraft, achievement of unification, possibility of simultaneous launch of several spacecrafts, and reduction of labour costs in course of the spacecraft creation.

21 cl, 5 dwg

Description

This invention relates to the rocket and space industry in the manufacture of spacecraft (SC).

The space platform (KP) is a structurally and functionally isolated module that combines service subsystems that provide the payload. In the process of creating a spacecraft, the gearbox is combined with the payload, which also represents a structurally and functionally separate module. The payload combines airborne transponders, antennas and all other elements that directly provide the solution to the target task of relaying information of a particular spacecraft in accordance with the necessary requirements. Since each particular spacecraft has its own frequency plan and service areas, and its repeater has its own scheme, the number of transponders and their energy characteristics, the payload for each spacecraft is unique. Solved KP problems are common for a number of spacecraft of the same class close in characteristics, therefore it is a unified element common to the whole series and having standard output characteristics and interfaces. Common to a number of the same type of spacecraft should be only the basic principles of constructing the payload and interfaces with the KP, which are the main condition for the possibility of application to create a spacecraft based on one of the unified platforms. The modular construction allows, within the framework of the standard gearbox design, to obtain various modifications that most fully satisfy the requirements of specific spacecraft created on the basis of the gearbox.

The unified structural basis of unpressurized platforms is a power structure made in the form of a central pipe (CT), with the following installed on it:

- rotary wings of the solar battery (BS);

- separation devices;

- storage tanks for the working fluid;

- Instrument compartment, with instruments and equipment for office systems, propulsion systems (Space platform [Electronic resource]. - Access mode: https://ru.wikipedia.org/wiki/Cosmic_platform, free. - Heading from the screen).

To ensure easy integration with various payload modules (MPN) corresponding to different spacecraft, the KP has simple and clearly defined unified interfaces, including:

- mechanical interface;

- electrical interface;

- thermal interface;

- information interface.

The construction and characteristics of the interfaces are universal and provide the ability to integrate various spacecraft with the MPN, which correspond to the range of platform interface requirements.

Spatially, all interfaces are located in the docking areas of the platform structures and the payload, and they are easily accessible at all stages of ground operation.

To install the SC on the launch vehicle (CB), the gearbox has a unified interface.

The interface with the SV is also used for docking with ground transportation and technological equipment during the assembly, integration and testing of the platform and the spacecraft as a whole, as well as transportation and preparation at the launch site.

The KP incorporates on-board systems capable of ensuring the functioning of the spacecraft in the area of launching into orbit, drift and installation at a given point in the orbit, and of fulfilling targets during the life of:

- general management of all subsystems and equipment and interaction with the ground control complex;

- transferring the platform from the starting configuration to the working one;

- orientation and stabilization of the spacecraft hull with the required accuracy;

- spacecraft retention at a given point in the orbit with the required accuracy;

- the formation of control forces and moments in the process of orientation, stabilization of the spacecraft and control of its movement;

- power supply to all subsystems of the platform and MPN in all operating modes;

- maintaining the temperature conditions of all elements of the KP and MPN within the specified limits;

- maintaining all the elements of the spacecraft in the required mutual position at all stages of operation and protection from external influences;

- ensuring the conduct of ground testing and testing of the spacecraft and its on-board systems, interaction with ground-based testing equipment.

From the prior art, the “Spacecraft of modular execution" is known (patent for invention RU No. 2092398, B64G 1/10, publication date 10/10/1997), which relates to space technology in which the orientation with respect to the Sun is regular, for example, SC in geostationary orbit, SC with Solar-Earth orientation in circular and highly elliptical orbits, etc. The inventive placement (partially or completely) on four faces of the parallelepiped (instrument container) of radiation surfaces allows the instrument container to be square in shape and, as a result, to obtain a compact (not elongated along the longitudinal axis) instrument container that fits optimally into the payload zone . An increase in the refrigerating capacity of the same exchange of the instrument container (up to 30%) due to previously unused external surfaces provides an additional opportunity to place heat-loaded on-board equipment (BA) in the internal volume of the instrument container, allows you to compose a larger number of heat-loaded instruments and components and optimally use (occupy ) the entire volume provided under the cylindrical radome of the launch vehicles. The design and layout scheme of the spacecraft, containing and integrating the main nodes and elements of the target equipment (payload) and service systems, is built on the block-modular principle of structurally detached blocks for functional purpose and taking into account the permissible temperature range of the BA.

The disadvantages of this technical solution are:

- layout restrictions imposed on the payload when applying this technical solution. In modern spacecraft, antennas with rigid reflectors up to 2.5 m in size are used, the location of the reflectors on the west and east sides (the surfaces by which an increase in cooling capacity is achieved) will reduce the effectiveness of radiation surfaces due to the shading of space open to radiators and additional heat gain during re-emission with warmer antenna reflectors;

- increased mass of the spacecraft due to the use of mainly honeycomb panels in the strength scheme to ensure rigidity, strength, geometric stability and thermoelasticity;

- the current level of miniaturization of equipment allows you to place the equipment of service systems on a smaller number of layout areas, while there will be no need to organize additional radiation panels.

The Space Platform is also known (patent for invention RU No. 2376212, B64G 1/00, B64G 1/22, publication date 10.12.2009), which relates to the design and layout of space technology products. The platform contains a parallelepiped-shaped frame with side, top and bottom panels mounted on it. A BS containing the root and end sections is pivotally mounted on the frame. The bar of the gravity device is installed on the top panel. Inside, the frame space is divided by an intermediate panel into the utility system compartment and the payload compartment. The electromagnetic devices of the orientation and stabilization system are mounted on the bottom panel from the side of the frame. The magnetometer is mounted on the top panel. The side panels are pivotally connected to the frame and provided with fastening elements mounted on them from the frame inside the payload compartment. BS rotation hinges are placed on the side of the side panels that do not contain frame connection nodes with the separation system. The root and end sections are equipped with limiters for their mutual rotation by an angle not exceeding 270 °. The total length of two pairs of root and end sections exceeds the total length of the end edges of the frame located on the side of the side panels that do not contain the specified connection nodes. Thin-film photoelectric converters are glued to the side panel containing the connection nodes of the frame with the separation system. The height of the root and end sections does not exceed the height of the compartment service systems. Side panels contain rotary fragments installed in the area of the payload compartment and equipped with opening drives. The technical result of the invention is to ensure constant charging of on-board chemical batteries in flight, reducing the mass of the structure and improving the accuracy of measuring the level of the magnetic field with a magnetometer.

The disadvantage of this technical solution is the lack of universality of the space platform. KP is not a complete assembly unit, as the payload is located in one frame with service systems, therefore, when creating spacecraft with various target equipment on its base, modernization (structural changes of the space platform when installing the target equipment of the spacecraft) of the platform will be required. This leads to an increase in ground experimental testing, an increase in the time and cost of creating spacecraft.

The well-known "Spacecraft" (patent for the invention RU No. 2463219, B64G 1/10, B64G 1/50, publication date 10/10/2012). The invention relates to space technology and for the design of automatic spacecraft for use in near-Earth orbits with instrument containers made of honeycomb panels using heat pipes. The spacecraft contains a multi-purpose payload, an unpressurized instrument box of a parallelepiped (prismatic) shape with integrated heat pipes in a honeycomb panel with installed heat-loaded devices. The edges of the container are radiator. Instrument modules with cross-shaped power honeycomb panels and transverse partitions are installed in an unpressurized instrument container. Cross-shaped honeycomb panels are interconnected by heat pipes. All honeycomb panels of the instrument container are connected into a single heat network by collector heat pipes equipped with electric heaters in the zone of each honeycomb panel with blocks of control temperature sensors. The cruciform honeycombs and brackets of the propulsion system are connected by heat pipes to the collector heat pipes of the instrument container. A thermoregulatory coating is applied to the outer surfaces of the honeycomb panels containing integrated heat pipes, the rest of the external surfaces of the honeycomb panels are thermally insulated. The instrument container is equipped with additional adjustable radiation heat exchangers with loop heat pipes, the evaporators of which are connected to the collector heat pipes of the instrument container through honeycomb panels with integrated heat pipes. Condensers of contour heat pipes are placed above the insulated parts of the honeycomb panels of the instrument container. Improved thermal stabilization of instruments and equipment is achieved with a uniform temperature field within each honeycomb panel and between honeycomb panels of the instrument container while increasing the density of the layout of the instrument container.

The disadvantage of this technical solution is that the payload equipment and the equipment of the service systems are combined in one instrument container. Changing the composition of the payload equipment will lead to the need to change (refine) the design and temperature control system of the instrument unit, as All honeycombs are connected to a single heat network. The need to refine the instrument unit with different target equipment leads to the fact that with the same set of service systems it is necessary to carry out the entire amount of ground experimental testing, which does not allow to reduce the time and cost of spacecraft.

Also known: "Modular core structure for dual launch" ("Modular core structure for dual-manifest spacecraft launch" patent for invention US 20140239125 A1, B64G 1/66, B64G 1/40, publication date 08/28/2014). The invention relates to the construction of a spacecraft. The spacecraft comprises an upper central structure or lower central structure. The upper central structure may include an upper cylinder to support the upper spacecraft in a dual launch configuration. The lower central structure may include a lower cylinder to support the lower spacecraft with the upper cylinder mounted on top of the lower cylinder. The upper cylinder may have an inner diameter of the upper cylinder similar to the inner diameter of the lower cylinder.

The disadvantage of this invention is that the invention mainly relates to the power structure of the spacecraft and is not a device functionally separate module that combines service subsystems that provide the payload. Also, the presented invention does not have in its description on-board systems capable of ensuring the functioning of the spacecraft and the performance of targets over the life of the aircraft.

The closest analogue is the "Multipurpose office platform for creating spacecraft" (patent for invention RU No. 2375267, B64G 1/10, B64G 1/22, 12/10/2009), the platform contains a module of office equipment in the form of a rectangular parallelepiped formed by an end plate and four side boards. Two intermediate boards are installed inside, dividing the module into three compartments for office equipment. The instruments of the orientation and stabilization system and antennas are mounted on the side board. On one of the boards, docking nodes with a separation system are mounted. The propulsion system is mounted in the region of the proposed center of mass. BS panels are mounted on brackets protruding beyond the module. The MPN installation units are located on the free ends of the module side boards and protruding brackets. Moreover, the devices of the target payload equipment are located in the space between the BS and the free zone of the module from the side of its open part.

A number of significant disadvantages characteristic of the prototype is as follows:

- all spacecraft created on the basis of a multi-purpose service platform should have the same dimensions, because the dimensions of the MPN are limited by the brackets of the BS installation, this imposes restrictions on the payload and reduces the versatility of the platform;

- the design of the platform does not allow the use of parabolic reflector antennas with rigid reflectors as part of the target equipment, which limits the capabilities of the target load.

The tasks to be solved by the claimed invention are aimed at selling a product that meets modern requirements for reducing dimensions and weight, increasing the density of the layout created by spacecraft based on this gearbox, reducing the amount of work on adapting spacecraft based on this gearbox to SV, increasing technical and operational characteristics, ensuring the simultaneous launch of several spacecraft made on the basis of this KP, as well as reducing the time and cost of creating a spacecraft based on this KP.

The tasks are solved due to the fact that the claimed control unit contains an instrument compartment, with instruments and equipment for service systems, correction engines, orientation engines, MPN installation units, docking units with a separation system, BS rotary wings. The main structural element of the structure is the power structure, made in the form of a central heating station, on which the instrument compartment is located, while the ends of the central heating protrude beyond the plane of the instrument compartment panels. At the ends of the central part there are docking nodes with a separation system, inside the central part there are storage tanks for the working fluid for correction engines and orientation engines. The instrument compartment is made in the form of a rectangular parallelepiped formed of panels, some of which are radiator panels. Inside and outside the instrument compartment are devices and equipment of service systems. Outside on the instrument compartment are located: correction engines, orientation motors, docking nodes with a payload, BS rotary wings, folding in the starting state, fixed symmetrically from two opposite sides of the instrument compartment using rods to rotation devices, and BS wings are made in the form of flat panels bonded together.

Preferably, the CT can be made in the form of a spatial mesh structure of composite materials.

The mesh structure can be formed by the intersection of spiral-shaped, longitudinal and annular elements.

A CT can consist of composite or metal skin with honeycomb in between.

CT can be made of a layered composite material.

CT can be made of solid fiber and a polymer binder.

Inside the DH from the side of the instrument compartment can be placed extension engines.

Extension engines are preferably electric rocket engines — plasma or ion.

Dilution motors preferably use gaseous xenon as the working fluid.

Extension motors can be mounted on the drive with the ability to change their angular position.

Correction engines can be located on the panels of the instrument compartment, the surfaces of which are facing the wings of the BS, and the correction motors are preferably placed at an angle to the surface of the instrument compartment on special brackets on the axis of symmetry on both sides, relative to the wings of the BS.

Correction engines can be located on the MPN KA, created on the basis of this gearbox.

Correction engines can be paired at an angle to the surface of the instrument compartment on special brackets on both sides of the instrument compartment on the sides perpendicular to the wings of the BS.

Correction motors can be mounted on the drive with the ability to change their angular position.

Correction engines are preferably electric rocket engines — plasma or ion.

Correction engines preferably use gaseous xenon as the working fluid.

The end face of the central heating facility, opposite the end with the instrument compartment, can be used for docking with a passing spacecraft during group launch.

Radiator panels can be made rectangular and are the side faces of the instrument compartment.

The panels of the instrument compartment are preferably made of composite material.

On the panel-base of the instrument compartment located near the end of the central heating center, eight orientation motors can be placed at an angle to the surface of the panel-base of the instrument compartment on special brackets at the corners of the base panel and at the periphery along the axes of symmetry.

BS wings can be made in the form of flat panels fastened together, which consist of a frame and photoelectric converters.

Technical results provided by the given set of features are:

- weight reduction due to the reduction in the dimensions of structural elements, as well as the use of central heating systems as a single power base on which the CP and MPN are located;

- increasing the density of the layout of the spacecraft created on the basis of KP, on the basis of the principle of optimal use of the volume of ZPG launch vehicle, while KP is placed in the base, and MPN - under the cone of the fairing ZPG;

- reduction in the volume and timing of work on adaptation of the SV carried out to ensure the launch of the spacecraft on the basis of this platform due to the unified mechanical interface and the unified electrical interface;

- increase of technical and operational characteristics, weight reduction due to the use of additional engines and optimal placement of correction engines;

- the ability to simultaneously launch several spacecraft on the basis of this gearbox, associated with the execution of the central heating circuit, at the ends of which there are interfaces with the separation system, allowing the spacecraft to dock with the upper stage and other spacecraft;

- reducing the time and cost of spacecraft, through the parallel development and creation of modules.

The invention is illustrated by drawings, which depict:

- in FIG. 1 shows a general view (operating state of the gearbox in axonometric projection);

- in FIG. 2 shows a general view (starting state of the gearbox in axonometric projection);

- in FIG. 3 shows the placement of the storage tanks of the working fluid for correction engines, orientation engines;

- in FIG. 4 shows a bottom view (starting state of the gearbox in axonometric projection);

- in FIG. 5 depicts a general view (private version of the platform).

The control unit contains an instrument compartment 1, with instruments and equipment 2 for service systems, correction engines 3, orientation motors 4, MPN installation units (not shown in the drawings), docking units with a separation system (not shown in the drawings), 5 BS rotary wings. The main structural element of the structure is the power structure, made in the form of CT 6, on which the instrument compartment 1 is located, while the ends of the CT 6 protrude beyond the plane of the panels of the instrument compartment 1. At the ends of the CT 6 are docking nodes with a separation system, tanks are located inside the CT 6 7 storage of the working fluid for engines 3 correction, engines 4 orientation. The instrument compartment 1 is made in the form of a rectangular parallelepiped formed of panels, some of which are radiator panels. Inside and outside the instrument compartment 1 are devices and equipment 2 service systems. Outside on the instrument compartment 1 are located: correction engines 3, orientation motors 4, docking nodes with a payload, BS rotary wings 5, folding in the starting state, mounted symmetrically from two opposite sides of the instrument compartment 1 using rods to the rotation devices, with wings 5 BS made in the form of flat panels fastened together.

CT 6 is made in the form of a hollow composite cylindrical compartment for placement of structural panels, equipment and fastening of structural elements of the spacecraft. Preferably, the CT 6 can be made in the form of a spatial mesh structure of composite materials. The mesh structure can be formed by the intersection of spiral, longitudinal and annular elements, and can also consist of composite or metal cladding with honeycomb core between them. CT 6 can also be made of a layered composite material or can be made of solid fiber and a polymer binder. CT 6 has more bending stiffness than the reinforced panels, takes up loads at startup and transfers them to the interface with the CB adapter. Moreover, the placement of tanks 7 in the central heating center 6 coincides with the axis of symmetry of the spacecraft passing through the center of the tanks, which provides better control of the position of the center of gravity of the spacecraft even in case of uneven fuel consumption from the tanks.

The end face of CT 6, opposite the end with the instrument compartment 1, can be used for docking with a passing SC during group launch. At the ends of CT 6, docking nodes with a separation system are fixed. At the end of CT 6 from the side of the instrument compartment 1, the docking station is a unified mechanical and electrical interfaces. The unified docking unit allows to reduce the volume and terms of work on adaptation of the SV carried out to ensure the launch of the spacecraft on the basis of this platform, namely:

- the unified mechanical interface of the spacecraft / SV, which eliminates the fitting tests of the transition system of the spacecraft with SV for each spacecraft;

- a unified electrical interface allows the release of a single protocol of electrical pairing for all spacecraft on the basis of this platform from the SV and during adaptation work only to confirm its implementation;

- minimize joint tests of the spacecraft with the SV at the technical complex and the launch complex, check only newly formed electrical circuits and the passage of commands.

The docking node at the opposite end of CT 6 (without instrument compartment 1) is a mechanical interface that allows docking with a passing SC in case of a group launch. In the case of a single launch of a spacecraft based on this gearbox, elements of the MPN can dock to this interface. Also, in the area of the end face of the central heating zone, zones are provided for organizing the electrical interface, a passing SC / SV or a passing SC / SC based on the presented platform.

Inside DH 6 from the side of the instrument compartment 1 can be placed engines 8 additional. The additional engines 8 are preferably electric rocket engines, plasma or ion, which preferably use gaseous xenon as their working medium. The extension motors 8 can be mounted on the drive with the ability to change their angular position. The additional engines 8 are designed to put the SC into final orbit. Motivation engines 8 based on electric rocket engines can reduce the mass of the spacecraft, which leads to lower costs for launching this spacecraft, or to develop a more powerful spacecraft for this mass, which improves the technical characteristics of the spacecraft.

Correction engines 3 can be located on the panels of the instrument compartment 1, the surfaces of which are turned to the wings 5 of the BS (Fig. 2), and the correction motors 3 are preferably placed at an angle to the surface of the instrument compartment 1 on special brackets on the axis of symmetry on both sides, relative to the wings 5 BS. Correction engines 3 can be located on panels 9 MPN (in Fig. 5 shows a variant of the placement of correction engines on MPN) KA, created on the basis of this KP. Correction engines 3 can be paired at an angle to the surface of the instrument compartment 1 on special brackets on both sides of the instrument compartment 1 on the sides perpendicular to the wings of the BS 5. Correction motors 3 can be mounted on the drive with the ability to change their angular position. Correction engines 3 are preferably electric rocket engines — plasma or ion. Correction engines 3 preferably use gaseous xenon as the working fluid. Variable placement of correction engines 3 allows you to choose the optimal layout with minimal impact on the spacecraft control system of disturbing moments from the action of the engine jets on the structural elements of the spacecraft. The minimum disturbing moments from the action of the engine jets allow reducing the mass of the spacecraft control system (for example, reducing the mass of gyroscopic power stabilizers (flywheels), and reducing the reserves of the working fluid necessary to maintain the orientation of the spacecraft). In special cases of implementation, each correction motor 3 is mounted on the drive and is able to change its angular position, i.e. change traction vector. Using these capabilities allows during operations to control the center of mass of the spacecraft to quench the kinetic moment (unloading) of inertial spacecraft controls, for example, gyroscopic power stabilizers (flywheels).

The instrument compartment 1 of the gearbox is made in the form of a rectangular parallelepiped formed of panels, some of which may be panels-radiators having a rectangular shape (side faces of the instrument compartment 1). The panels of the instrument compartment 1 are preferably made of composite material. Inside and outside the instrument compartment 1 are devices and equipment 2 service systems. In modern spacecraft, the technology of planar installation is used, as well as the methods of designing instruments and equipment of 2 spacecraft associated with this technology. The use of the instrument compartment 1 of a rectangular shape allows the use of many flat surfaces for mounting devices and equipment 2, and the use of flat panels allows you to mount devices and equipment 2 from two sides. The density of the layout and tracing of cable routes is achieved due to the minimum gaps between the devices and equipment. The use of mounting panels as radiators allows to reduce the number of structural elements (without having to provide a separate structure for placing equipment with a separate radiator), thereby reducing the dimensions and weight of the spacecraft. One of the requirements for the gearbox is the high density of the layout of the spacecraft created on the basis of this gearbox. The implementation of the instrument compartment 1 KP of minimum dimensions near the junction with the NE creates minimal restrictions on the layout of the target equipment. The location of the gearbox in the BCP allows you to place the target equipment in the remaining part of the BCP (can be used, including the conical part of the BCP). Eight orientation motors 4 (Fig. 4) can be placed on the base panel of the instrument compartment 1, located near the end of the CT 6, mounted at an angle to the surface of the base panel of the instrument compartment 1 on special brackets at the corners of the base panel and at the periphery along axis of symmetry.

BS rotary wings 5, folding in the starting state, are fixed symmetrically from two opposite sides of the instrument compartment 1 with the help of rods to rotation devices, and BS wings 5 are made in the form of flat panels fastened together, consisting of a frame and photoelectric converters. After the wings of the 5 BS are opened, using the rotation device, their rotation is ensured in order to be oriented to the Sun to ensure maximum efficiency.

When creating modern spacecraft for various purposes (communications, geodesy, navigation, etc.), it is envisaged to implement a comprehensive program of ground-based experimental testing of the spacecraft and its subsystems for all types of external influences determined by operating conditions in the spacecraft. The implementation of the KP in the form of a separate assembly with certain interfaces with the MPN and SV allows ground-based experimental testing of the KP without waiting for the production of MPN, which reduces the time required to manufacture the spacecraft. The implementation of KP in the form of a separate assembly makes it possible to carry out parallel production and assembly of MPN and KP and, as a result, shorten the spacecraft manufacturing time.

Claims (21)

1. A space platform containing an instrument compartment with instruments and equipment of service systems, correction engines, orientation engines, installation units of the payload module, docking units with a separation system, rotary wings of the solar battery (BS), characterized in that the main structural element is the power structure, made in the form of a central pipe (CT), on which the instrument compartment is located, while the ends of the CT protrude beyond the plane of the panels of the instrument compartment, the nodes of the docking system are placed on the ends of the CT my department, inside the central heating storage tanks are located for correction engines, orientation engines, the instrument compartment is made in the form of a rectangular parallelepiped made up of panels, some of which are radiator panels, inside and outside the instrument compartment there are devices and equipment of service systems, moreover, on the outside in the instrument compartment are correction engines, orientation engines, docking nodes with a payload, BS rotary wings, folding in the starting state, lennye symmetrically on two opposing sides of the instrument compartment via the rods to the pivot device, wherein the BS wings are in the form of flat panels connected to each other. 2. The space platform according to claim 1, characterized in that the CT is made in the form of a spatial mesh structure of composite materials. 3. The space platform according to claim 2, characterized in that the mesh structure is formed by the intersection of spiral-shaped, longitudinal and annular elements. 4. The space platform according to claim 1, characterized in that the CT consists of composite or metal skin with a honeycomb core between them. 5. The space platform according to claim 1, characterized in that the CT is made of a layered composite material. 6. The space platform according to claim 1, characterized in that the CT is made of solid fiber and a polymer binder. 7. The space platform according to p. 1, characterized in that in the central heating center from the side of the instrument compartment there are additional engines. 8. The space platform according to claim 7, characterized in that the extension engines are electric rocket engines - plasma or ion. 9. The space platform according to claim 8, characterized in that the exhaust engines use gaseous xenon as a working fluid. 10. The space platform according to claim 7, characterized in that the extension motors are mounted on the drive and are able to change their angular position. 11. The space platform according to claim 1, characterized in that the correction engines are located on the panels of the instrument compartment, the surfaces of which are facing the wings of the BS, and the correction engines are placed at an angle to the surface of the instrument compartment on special brackets on the axis of symmetry on both sides relative to the wings of the BS . 12. The space platform according to claim 1, characterized in that the correction engines are located on the payload module of the spacecraft (SC), created on the basis of this space platform. 13. The space platform according to claim 1, characterized in that the correction engines are arranged in pairs at an angle to the surface of the instrument compartment on special brackets on both sides of the instrument compartment on sides perpendicular to the wings of the BS. 14. The space platform according to any one of paragraphs. 11-13, characterized in that the correction motors are mounted on the drive and are able to change their angular position. 15. The space platform according to claim 1, characterized in that the correction engines are electric rocket engines - plasma or ion. 16. The space platform according to p. 15, characterized in that the correction engines use gaseous xenon as a working fluid. 17. The space platform according to claim 1, characterized in that the end face of the CT opposite the end face with the instrument compartment can be used for docking with a passing spacecraft during group launch. 18. The space platform according to claim 1, characterized in that the radiator panels are made in a rectangular shape and are lateral faces of the instrument compartment. 19. The space platform according to claim 1, characterized in that the panels of the instrument compartment are made of composite material. 20. The space platform according to p. 1, characterized in that on the panel-base of the instrument compartment, located near the end of the central heating center, eight orientation motors are installed at an angle to the surface of the panel-base of the instrument compartment on special brackets at the corners of the base panel and on peripheries along the axes of symmetry. 21. The space platform according to claim 1, characterized in that the BS wings are made in the form of flat panels fastened together, which consist of a frame and photoelectric converters.

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