WO2018147760A1

WO2018147760A1 - Micro class earth remote sensing spacecraft

Info

Publication number: WO2018147760A1
Application number: PCT/RU2017/000193
Authority: WO
Inventors: Александр Сергеевич МАЛИНИН; Пётр Викторович КУДРЯШОВ; Дмитрий Вадимович ДМИТРИЕВ; Владимир Евгеньевич ШМАГИН; Пётр Евгеньевич РОЗИН; Роман Николаевич АРХАНГЕЛЬСКИЙ; Александр Евгеньевич МИЛОВ; Сергей Владимирович ИОСИПЕНКО
Original assignee: Общество с ограниченной ответственностью "Даурия - спутниковые технологии"
Priority date: 2017-02-09
Filing date: 2017-03-31
Publication date: 2018-08-16
Also published as: RU2651309C1

Abstract

The invention relates to spacecraft of the platform type, developed in accordance with the CubeSat standard. The present spacecraft comprises a body in the shape of a parallelepiped and consisting of an upper panel and side panels, fastened to a service equipment frame, said frame dividing the body internally into two parts. Solar cells are mounted on the outer sides of the panels. The service equipment frame is in the form of a machined plate, on one side of which are mounted power supply and control units with built-in transceiving equipment, cables, and elements of an orientation and stabilization control system. Attached to the other side of the service equipment frame is an opto-electronic system, disposed on its own frame, the aperture of which is closed by a camera cap (with a solar cell thereon), and reaction wheels are disposed between the bearing ribs of the frame. VHF and GPS receiver antennae are disposed on the upper panel of the body of the spacecraft. The technical result of the invention is that of miniaturizing the structure and on-board systems of a platform within the framework of the CubeSat standard (with different form factors), while at the same time simplifying and expediting the process of producing, testing and assembling the spacecraft.

Description

Micro Class Earth Remote Sensing Spacecraft

FIELD OF THE INVENTION

The technical solution relates to the field of space technology, specifically to space platforms (KP) for the creation of micro-class spacecraft, developed in accordance with the CubeSat standard and used for remote sensing of the Earth (ERS).

State of the art

At the moment, the remote sensing market is shifting towards an increase in the following quantitative indicators of the quality of space systems:

· Increasing the spectral resolution of satellites.

• Increasing the spatial resolution of satellites.

• Improving the regularity of shooting a given region of the earth's surface.

At the same time, given the degree of miniaturization and microelectronics performance, a trend toward a change in the following spacecraft characteristics has clearly been outlined:

· Increasing the speed of transferring Earth survey products to ground-based points of reception of target information.

• Decrease in the total mass and size of the spacecraft and the mass of the ECO.

• Reducing the cost of manufacturing and launching the spacecraft.

It is worth noting that one of the main disadvantages of modern remote sensing systems is the high cost of developing and launching each spacecraft and, as a result, economic constraints when creating a large group, which in turn reduces the efficiency of shooting. Therefore, one of the main trends in the creation of remote sensing systems is the use of low-budget spacecraft of small and micro class.

The space platform is known from the prior art, which comprises a parallelepiped-shaped supporting body, equipped with flip modules connected with detachable hinge assemblies to the supporting body, rotary solar batteries mounted on the supporting body by electric drives, service system devices located inside the supporting body, fastening elements payload and nodes of the connection of the bearing body with the separation system. Hinged modules are equipped with swing mechanisms and fixation units for hinged modules to the bearing housing. Inside the hinged modules, payload fastening elements are placed. Additional solar panels are installed on the hinged modules (patent RU 2410294, B64G 1/10 12/30/2008).

The disadvantages of this space platform include the fact that the folding modules are unsuitable for placing in them a payload of remote sensing with large the aperture of the entrance pupil, in addition, when opening the folding modules, the accuracy of the orientation and stabilization of the space platform will significantly decrease, which will negatively affect (up to the complete absence of such an opportunity) the quality of the satellite imagery of the earth’s surface. In addition, the presence of protruding elements on the space platform does not allow it to be launched as part of a transport launch container (TPK), which can reduce shock loads on the optoelectronic system (OES) of a spacecraft during its separation.

Also, the space platform of remote sensing is known, which can be used to create small spacecraft remote sensing with a mass of 100-500 kg to work in low Earth orbits (patent for utility model Ns: 132422). The peculiarity of this space platform is that the body is made of interconnected self-supporting thermostabilized honeycomb panels with built-in frames and heat pipes; panels of the solar battery (SB) are configured to unfold the extreme wings perpendicularly with respect to the direction of unfoldment of the root parts of the wings of the SB; units and devices of service on-board systems that are resistant to radiation and ultraviolet radiation are installed on the outer surfaces of the hull, equipped with a propulsion system, a transition element and a fastening element of the payload module.

The disadvantages of this space platform include its high mass, which can reach from 100 to 500 kg, and bulkiness, which will lead to an increase in the production time and ground experimental testing of the spacecraft, which will entail additional labor costs and the cost of work. In addition, due to the high mass of the spacecraft based on this KP, the launch cost will be higher. The use of honeycomb panels with heat pipes also significantly increases the cost of this gearbox.

Also, a microsatellite for scientific research is known (patent RU 2572365, B64G 1/10, B64G 1/64 01/10/2016), put into orbit from a transport launch container, on the case of which the opening solar panels and antennas are installed in the attachment and rotation nodes, held by rotary levers of the case. The peculiarity of this microsatellite is that the microsatellite contains a rectangular parallelepiped-shaped housing, solar panels and antennas that are mounted on it in attachment and rotation units, and solar panels and antennas that are connected to it and a microsatellite separation system located on the lower end of the case, each attachment and rotation unit equipped with a spring mechanism, and on the microsatellite case mounted rotary levers that hold the solar panels and / or antennas in the transport position, and the housing and rotary levers equipped with rolling elements, which in the transport position rest on the inner surface of the transport and launch container of the microsatellite.

The disadvantages of this microsatellite are the uniqueness of its design and, as a result, the uniqueness of a cylindrical transport launch container in which this microsatellite is put into orbit, which makes it incompatible with the universal standard of CubeSat microsatellites and thereby complicates its adaptation to a wide range of launch vehicles. Also, the presence on the microsatellite design of elements protruding beyond the overall size of the housing leads to an increase in the volume and mass of the container bringing it into orbit.

In addition, the technical solutions of spacecraft developed by foreign companies for remote sensing tasks are known.

One such platform is the Dove space platform developed by the American company Planet (https: //directoi^.eoportal.orq/web/eoportal/satellite-missions/d/dove). This space platform is a satellite of the CubeSat 31 form factor, launched as part of a transport and launch container, which has solar panels that open along the long side, as well as a drop-down cover for the optoelectronic system, which includes a transmit-receive antenna.

The disadvantage of this space platform is its small size (the aperture of the entrance pupil of the optoelectronic system is not more than 10 cm), which does not allow to achieve high spatial resolution of the obtained images, and also limits the lens aperture.

Thus, to date, there are no existing groupings of high- and medium-resolution spacecraft that fully satisfy the requirements of customers at a new stage in the development of geo-information technologies.

The closest analogue of the claimed spacecraft is the space platform series Dove, manufactured by Planet. Spacecraft based on the Dove space platform are closest in manufacturing ideology (CubeSat standard), and also have similar mass and dimensional characteristics and spatial resolution characteristics. However, it is worth noting that the Dove series spacecraft are inferior to the declared Auriga spacecraft in spatial resolution (5 m to 2.5 m for an orbit altitude of 600 km) and the rate of discharge of target information (120 to 160 Mbit / s) SUMMARY OF THE INVENTION

The problem solved by the claimed invention is to provide remote sensing of the Earth with a high-resolution optoelectronic system and to ensure the transmission of satellite imagery data to ground-based reception means by means of a micro-class CubeSat spacecraft with form factor 16U, as well as simplification and acceleration of the manufacturing process, tests and installation of the spacecraft.

The technical result of the claimed invention is to miniaturize the design and onboard systems of the space platform to the dimensions of a CubeSat standard micro-satellite with a 16U form factor, provided that the target for remote sensing of the Earth is met by optical-electronic systems in high resolution and the data are transmitted to ground-based reception facilities, which leads to a simplification and acceleration of the manufacturing process, testing and installation of the spacecraft. The technical result of the claimed invention is achieved due to the fact that the micro-class Earth remote sensing spacecraft containing a body made in the form of a parallelepiped, consisting of side panels of the spacecraft body, mounted on a service equipment frame with the division of the internal volume of the body into two parts, and the top panel, while on the outer sides of the side panels and the top panel, side panels of the solar panels and the top panel of the solar panel are installed respectively The frame of service equipment is the main power element and is made in the form of a milled plate on which, on the one hand, a power supply and control unit with built-in GPS receiver and VHF transceiver, a high-speed Ka-frequency target information transmitter with a horn antenna are installed, electrical cables, and high-precision instruments of orientation and stabilization control systems mounted on the frame through the bracket, and on the other side of the frame of the service equipment by means of rods The optoelectronic system and three flywheel engines are fixed, while the frame of the optoelectronic system is made with power ribs, between which flywheel engines are located at a distance from the star sensors and the gyroscope, while the inlet of the optoelectronic system is closed by the camera cover , on the outer surface of which there is also a solar panel, on the upper panel of the spacecraft’s body are the antennas of the VHF band transceiver and GPS receiver, at the corners of the side panels on the side the top panel and the camera cover are 4 supporting projections. In the particular case of the implementation of the claimed technical solution, the instruments of the orientation and stabilization control system are made in the form of two star sensors, one microelectromechanical gyroscope, three flywheel engines, as well as magnetometers, solar sensors and magnetic actuators as part of solar panels,

In the particular case of the implementation of the claimed technical solution, the side panels are additionally interconnected.

In the particular case of the implementation of the claimed technical solution, the place of attachment of the side panels to each other from the side of the optoelectronic system is additionally reinforced at the corners with scarves.

In the particular case of the implementation of the claimed technical solution, the solar panel located on the outside of the cover is made on the basis of a multilayer printed circuit board with photoelectric converters.

In the particular case of the implementation of the claimed technical solution, the side and top solar panels are made on the basis of multilayer printed circuit boards with photoelectric converters and with built-in solar sensors, magnetometers, magnetic actuators and a microcontroller for control.

In the particular case of the implementation of the claimed technical solution, the chamber lid is equipped with torsion springs and is fixed in the closed state by a tension thread.

In the particular case of the implementation of the claimed technical solution, the opening of the lid under the influence of torsion springs is carried out by burning the tension thread.

In the particular case of the implementation of the claimed technical solution, the spacecraft is capable of launching, launching and separating as part of a transport launch container.

Brief Description of the Drawings

Details, features, and advantages of the present invention follow from the following description of embodiments of the claimed technical solution using the drawings, which show:

FIG. 1 - General view of the spacecraft in the working position (top view).

FIG. 2 - General view of the spacecraft in the working position (bottom view).

FIG. 3 - General view of the spacecraft in transport position. FIG. 4 - KA view without side panels + Y and -Y.

FIG. 5 - KA view without panels and covers.

FIG. 6 - Frame of service equipment with equipment (top view).

FIG. 7 - Frame of service equipment with equipment (bottom view).

The following positions are indicated in the figures:

1a, 1b, 1c, 1d - side panels of the SB (SB panels -X, -U, + Y, + X, respectively); 2 - the upper panel SB (panel SB -Z); 3 - the bottom panel of the SB (panel SB + Z); 4a, 46 — VHF antenna (in the open and folded state, respectively); 5 - GPS antenna; 6 - camera cover; 7 - corner scarves; 8 - supporting guides; 9 - optoelectronic system; 10 - supporting protrusions -Z; 11 - supporting projections + Z; 12

- ECO cover opening springs; 13 - star sensors; 14 - Ka-band transmitter; 15 - power supply and control unit; 16 - frame ECO; 17 - frame service equipment; 18a, 18g — spacecraft side panels (-X, + X panels, respectively, -U and + Y panels are not represented); 19 - the upper panel of the spacecraft (panel -Z); 20 - gyroscope; 21

- flywheel engines; 22 - bracket

Disclosure of invention

The spacecraft of the Auriga micro class is made in the form of a parallelepiped, consisting of the main power-receiving structural element in the form of a frame (17) of service equipment, while the frame (17) is a milled rectangular plate with four posts in the corners on which the main service equipment is installed , namely:

Power supply and control unit (15) with built-in GPS receiver and VHF transceiver, high-speed Ka-frequency target information transmitter with horn antenna (14), three flywheel engines (21) (DM), as well as an arm (22) with high-precision instruments installed on it, an orientation and stabilization control system (SUOS), made in the form of two star sensors (13) (ZD) and one microelectromechanical (MEMS) gyroscope (20); in this case, the DMs (21) are placed on the frame (17) from below, away from the ZD (13) and the gyroscope (20), in order to minimize vibration effects on the sensors and the gyroscope, as well as to prevent electric cables from entering the rotating parts of the DM (21) (laying of all electric cables is carried out from the upper side of the official equipment frame). On the lower side of the frame (17) of the office equipment, the threaded holes for mounting the ECO (9) through the frame (16) of the ECO (9) are arranged on the racks, while the flywheel engines (21) are located in such a way that they are located in the free space between the power ribs of the frame (16) ECO (9). Two threaded holes for connecting to four side panels (18a, 18b, 18b, 18g) (panels + X, -X, + Y, -Y) forming the outer spacecraft body are made on the sides of the frame (17) of the office equipment; in this case, the side panels (18a, 18b, 18c, 18g) are interconnected in the corners by screws (lap), the corner ribs at the joints of the side panels play the role of supporting guides (8), which the spacecraft rests on the rails inside the TPK in the transverse directions and slides (on rails TPK) in the process of separation in orbit; at the same time, the place of attachment of the side panels (18a, 18b, 18c, 18g) to each other from the ECO side (9) is additionally reinforced at the corners with 4 scarves. On the outer sides of the side panels (18a, 18b, 18c, 18g), side panels (1a, 1b, 1c, 1g) of solar panels (SB) are installed, made on the basis of multilayer printed circuit boards and being separate electronic devices with built-in them solar sensors (DM), magnetometers, magnetic actuators (MIO) and microcontroller control.

The upper side of the spacecraft is closed by the upper panel (19) (panel -Z); at the same time, the antennas of the VHF band transceiver and GPS receiver are located on the upper panel, as well as the upper panel (2) of the SB, which is also an electronic device based on a multilayer printed circuit board.

The lower side of the spacecraft, where the inlet of the optical scheme of the ECO (9) is located, is closed by a lid (6) at the stage of removal (panel + Z); at the same time, the lower (3) SB panel is located on the cover. After the stage of removing and separating the spacecraft from the TPK, the cover (6) opens under the action of torsion springs, ensuring the opening of the inlet of the ECO (9). At the corners of the side panels (18a, 18g) (+ X, -X panels) from the side of the upper (2) panel and the lower (3) panel, 4 support projections (10) are made to support the spacecraft in the longitudinal direction against the pusher and the TPK cover. The axes of the associated spacecraft coordinate system materialize in such a way that the Z axis

The spacecraft is directed along the optical axis of the OES (9) in the direction from the inlet of the OES (9) to the frame (17) of the service equipment; the spacecraft's Y axis is directed along the axis of the horn antenna of the high-speed Ka-band transmitter (14) in the direction of the antenna radiation; the X axis of the spacecraft complements the spacecraft coordinate system to the right coordinate triple. The proposed spacecraft is supposed to be used at low (up to 800 km high) circular solar-synchronous orbits for remote sensing tasks.

A more detailed description of the design of the spacecraft micro class "Auriga" is presented below. The proposed spacecraft in the transport position is a parallelepiped (Fig.Z.), where the main power-receiving structural element is the frame (17) of the service equipment on which the service equipment is located as part of the power supply and control unit (15), high-speed transmitter (14) of target information Ka-band, three flywheel engines (21), as well as a precision bracket (22), on which two star sensors (13) and a gyroscope (20) of the spacecraft orientation and stabilization system are installed. Moreover, the Ka-band transmitter is placed on the frame (17) so that its horn antenna exits to the outer side of the KA in the direction of the + Y side of the KA, the power and control unit is located next to the Ka-band transmitter and occupies the space along the -X side on the frame SC, three flywheel engines (21) are located on the opposite side of the frame (17) of the service equipment (from the side -Z of the SC) in such a way as to enable the SC to turn in three mutually orthogonal planes, and star sensors (13) are located on the precision bracket (22) in such a way that their sighting axes are aligned with the axes of the spacecraft + X and -Y, respectively, and form a 90 ° angle between them, and the gyroscope (20) is placed on the bracket in the free zone between two ZDs (13).

The precision bracket (22) minimizes the error between the measuring coordinate systems of the instruments installed on it and the spacecraft construction coordinate system. Minimizing this error makes it possible to increase the accuracy of the spacecraft orientation during the operation of the target equipment. The use of a single base part in the form of a frame (17) of office equipment allows to realize the greatest possible accuracy of the relative installation of equipment, the manufacturability of the assembly process; provide the best heat transfer conditions for optimizing the temperature conditions of the spacecraft onboard equipment. The power and control unit (15), in addition to the on-board computer and power supply and control elements of the power supply system, includes lithium-ion rechargeable batteries as a secondary power source for the spacecraft, as well as a GPS signal receiver and an VHF range transceiver.

Such a combination of various on-board devices and systems as part of a single unit (15) allows to reduce the overall dimensions and weight of this unit in due to the use of a single enclosure and the exclusion of the onboard cable network for connecting devices and systems included in its composition.

The flywheel engines (21) are located on the frame (17) of the service equipment at a distance from the stellar sensors (13) and the gyroscope (20), in order to minimize the vibration effect on them.

Through the threaded holes of the rack frames of the service equipment (17), the frame (16) of the UES (9) is attached to it, which is the main supporting element of the payload of the spacecraft in the form of an optical-electronic system (9). At the same time, the frame (16) of the OES (9) has power ribs, in the intervals between which the housings of the flywheel engines (21) are placed, which also allows to reduce the overall size of the spacecraft.

The ECO (9) is a telescope in the shape of a rectangular parallelepiped, where the main power element is the frame (16) of the ECO on which the optical-mechanical and optical elements of the ECO are installed, the telescope's sighting axis coinciding with the Z axis of the spacecraft, and the entrance hole of the ECO optical system located on the side of -Z KA. OES (9) is made in the form of a mirror-lens axial telescope with an aperture of 241 mm and a focal length of 745 mm, which can significantly reduce the overall size and mass of the spacecraft, while ensuring high quality images with a resolution of up to 2.5 m per pixel in panchromatic mode at shooting into a nadir from an orbit 600 km high. On the sides of the frame (17) of the service equipment, 2 threaded holes were made for connecting to the four side panels KA + X, -X, + Y, -Y (18). These holes are located close to the midpoints of the sides of the frame (17) in order to reduce the mutual influence of temperature deformations of the panels (18) and the frame (17). Moreover, the side panels (18) are interconnected in the corners by overlapping screws. Corner ribs at the joints play the role of supporting guides (8), by which the spacecraft rests against the rails inside the TPK in the transverse directions and slides (along the TPK rails) during separation. From the side of the inlet of the ECO (9), the side panels (18) are additionally supported by four corner scarves (7).

On the side of the frame of the service equipment (17) on the upper faces of the side panels of the spacecraft (18) through the threaded connections, the upper panel (19) of the spacecraft is mounted on which an active GPS antenna (5) and a drop-down antenna of the VHF radio service complex (4) are installed a quarter-wave vibrator. In addition, on the top panel (19) of the spacecraft there are two contacts of the compartment (not shown conditionally) that are activated when the spacecraft is separated from the TPK and, closing the power circuit, power is switched to the spacecraft onboard systems. The side panels of the upper panel (19) form the outer spacecraft body.

On the outer sides of the side panels (18) and the upper panel (19) on the threaded connection, the side panels (1) of the solar panels and the upper panel (2) of the SB are installed. These panels (1, 2) of the SB are separate terminal electronic devices based on multilayer printed circuit boards, which, in addition to photoelectric converters (PEC) of the power supply system, include such EMS systems as miniature solar sensors, magnetometers, and magnetic actuators. All of these peripherals are controlled by a microcontroller integrated in each SB panel, which allows each SB panel (1, 2) to be considered as a separate device on the I2C bus. The use of SB smart panels (1, 2) as electronic devices based on multilayer printed circuit boards can significantly reduce the size and mass of the spacecraft.

From the side of the inlet of the ECO (9), a camera opening cover (6) is located, which is in the closed position until the angular velocity of the spacecraft is damped after separation from the TPK and the spacecraft is stabilized in standby mode (the optical axis of the ECO is directed to the nadir). Opening the lid (6) of the chamber is carried out by burning out the thread holding it, after which by means of two springs of opening the lid of the ECO (12) it is put into the open position and held in it. On the outside of the cover (6) of the camera through the threaded connection is installed the lower panel SB (3). It, like other SB panels (1, 2) as part of the spacecraft, is made on the basis of a multilayer printed circuit board, but without placing the SUOS devices in its composition.

While the spacecraft is inside the TPK, in the longitudinal direction (along the spacecraft optical axis), the spacecraft is fixed on both sides (sides -Z and + Z, respectively) through the support protrusions (11, 12) milled from the end sides of the side panels of the spacecraft + X and - X (18a, 18g), respectively.

The assembly of the spacecraft begins with the installation on the frame (17) of the service equipment of the on-board systems instruments and the bracket (22) with the SUOS devices (Fig. 6, Fig. 7). At the next stage of assembly, the frame of the service equipment is assembled from the ECO (9) through the ECO frame (16) (Figure 5.). The assembly is carried out in a vertical position, when the inlet of the ECO is directed upwards, and the frame of the service equipment (17) is mounted on the assembly table with rigging equipment. At this and previous stages of spacecraft assembly, the onboard cable network is traced. Next, the KA + X and -X side panels (18a, 18g), respectively, are installed on the assembly of the UES (9) and the frame of the service equipment (17) through threaded holes in the frame. At the same time, a GPS antenna (5) and an VHF antenna in an inset form (46) are installed on the upper panel of the spacecraft (19), after which the upper panel of the spacecraft (19) is installed on the upper end side of the side panels KA + X and -X ( 18a, 18g) (Fig. 4.).

At the next assembly stage, the KA + Y and -Y side panels (not shown conventionally) are attached to the official equipment frame (17), which are connected to the KA + X and -X panels (18a, 18g) with an overlap, and from the side of the inlet of the ECO fastened with four corner scarves (7). At the same stage, on the end of the side panel -X (18g) from the inlet side of the ECO, the chamber cover (6) is attached.

The side panels of the SB (1a, 1b, 1v, 1g), the upper SB panel (2) and the lower SB panel (3), the camera cover (6) are put into the closed state and fixed with a thread (Fig. C) .). In this state, the spacecraft can be installed on the supporting protrusions (11) or (12) for further work on it.

The spacecraft in the transport position (Fig.Z.) is installed by sliding the support rails (8) along the TPK rails (16U form factor) inside the transport launch container, is fixed in the longitudinal direction (along the optical axis) in the TPK structure through the support protrusions (11 ) and (12). At the same time, the inlet of the ECO should be directed towards the TPK opening door. In this position, the switched off spacecraft inside the TPK is delivered by means of launch into the calculated orbit, where it is separated from the TPK.

At the moment of separation of the spacecraft from the TPK, the contacts of the compartment are activated, which close the power switching circuit of the spacecraft onboard systems. A few minutes after the separation of the spacecraft from the TPK, the VHF range antenna (4a) is automatically opened and the spacecraft switches to the damping mode of the angular velocities obtained at the time of separation from the TPK by means of magnetic actuators as part of the SB panels. After stabilization, the spacecraft enters the standby mode of operation using stellar sensors and a gyroscope as measuring instruments and flywheel engines as executive bodies. In this mode, a three-axis orientation is built when the optical axis of the spacecraft is directed in the direction of the nadir. Unloading the accumulated kinetic moment on the flywheel engines is carried out by means of MIO.

Having completed all the preparatory and verification operations with the on-board onboard systems, the spacecraft can proceed to fulfill the target task by remote Earth sounding by means of ECO. A command is issued to open the lid of the chamber (6), through which the holding thread is burned, and the lid, under the action of the ECO lid opening springs (12), made in the form of torsion springs, goes into the open state (Figure 1, Figure 2.) . In this state, the lid remains for the entire spacecraft lifetime.

In the space survey mode, the OES of the spacecraft (9) takes pictures of the earth's surface specified in the flight mission and saves the survey data in a ROM of up to 1 Tb. At the same time, the space platform of the device allows you to shoot the earth's surface with a deviation from the nadir in the angle of heel and pitch by ± 30 °. The speed of re-targeting the spacecraft between objects of satellite imagery can be about 1.5 deg s. An operating mode is implemented when the ECO camera monitors a given region of the Earth to obtain stereo images.

In the mode of conducting a communication session with a ground-based point of reception of target information, the spacecraft is guided to the ground-level point with the horn antenna of the Ka-band transmitter (14) and holds the antenna orientation in the direction of the ground station throughout the entire session of transmitting satellite imagery data. In this case, the direction to the Sun is automatically tracked, and the spacecraft is guided by the horn antenna to a ground-based reception point so that it does not fall into the field of view of the ECO camera. After the communication session is completed, the spacecraft returns to the standby mode when the target axis of the ECO is directed to the nadir.

The novelty of the proposed technical solution is to miniaturize the design and onboard systems of the space platform to the size of a CubeSat standard micro-satellite with a 16U form factor to solve high-resolution remote sensing tasks.

The problem to which the invention is directed is achieved by:

- placing on-board service equipment on a single base part (17) allows to realize the greatest possible accuracy of the relative installation of equipment, manufacturability of the assembly process; provide the best heat transfer conditions for optimizing the temperature conditions of the spacecraft onboard equipment, as well as achieve miniaturization of the spacecraft.

- the use of miniature instruments of onboard systems, such as star sensors (13), flywheel engines (21), gyroscope (20), which allow reducing the overall size and mass of the spacecraft. - placement of several on-board devices and systems as part of a single power supply and control unit (15), which allows to reduce the area occupied by these devices and systems, as well as to reduce the weight and volume of the on-board cable network for connecting them together. - the use of a compact Ka-band transmitter (14), which provides higher energy of the communication channel and throughput per unit mass in comparison with X-band transmitters used on analogue satellites. The use of high-frequency components operating at the Ka-band (26.8 GG) frequencies as a part of the transmitter determines its compactness in combination with high-speed data transfer characteristics. Also, the compactness is added by the placement of a horn antenna in the transmitter. The static placement of the horn antenna on the design of the transmitter compares favorably with the antenna on the rotary device in terms of mass and dimensional characteristics, and also does not require laying onboard cable network design. - the use of SB panels (1a, 16, 1 in, 1g, 2) based on multilayer printed circuit boards with built-in microminiature SUOS devices, which can significantly reduce the mass and size of the spacecraft due to the fact that these SUOS devices do not occupy additional space in the spacecraft and do not require laying onboard cable network to them. - placement of flywheel engines (21) between the power ribs of the frame

OES (16), which makes it possible to reduce the size of the spacecraft along the optical axis.

- the use of a compact OES (9) in the form of a mirror-lens axial telescope with an aperture of 241 mm and a focal length of 745 mm, which can significantly reduce the overall size and mass of the spacecraft while ensuring high quality images with a resolution of up to 2.5 m per pixel in panchromatic mode when shooting in nadir from an orbit 600 km high.

- development of the spacecraft according to the CubeSat standard, which allows launching the spacecraft as part of the TPK, which simplifies the process of adapting the spacecraft to the launch vehicle, and also reduces the shock loads on the spacecraft design when it is separated in comparison with the separation method using pyrotechnic means.

- the use of TPK to launch the spacecraft, which allows to unify the process of adaptation of the spacecraft to almost any launch vehicle.

- minimizing the number of spacecraft structural elements, the use of miniature on-board instruments and systems, low weight (21 kg) and size (250 x 250 x 450 mm in transport position (Fig.Z.)) spacecraft, which allows to simplify and speed up the process of installation, testing and adaptation of the spacecraft.

- implementation of SB panels (1, 2, 3) based on photovoltaic converters from three-stage gallium arsenide, which allowed reducing the dimensions and weight of SB without reducing the efficiency of power supply of on-board systems.

Claims

Claim

1. A microclass Earth remote sensing spacecraft comprising a parallelepiped-shaped housing consisting of side panels of a spacecraft housing mounted on a service equipment frame with the internal volume of the housing divided into two parts, as well as the upper panel, while on the outer sides the side panels and the top panel are installed side panels of the solar panels and the top panel of the solar battery, respectively,

and the official equipment frame is the main power element and is made in the form of a milled plate, on which, on the one hand, a power supply and control unit with integrated GPS receiver and VHF transceiver, a high-speed Ka-frequency target information transmitter with a horn antenna, electrical cables are installed , and high-precision instruments of the orientation and stabilization control system installed on the frame through the bracket, and on the other hand, the frame of the service equipment through the frame for the optoelectronic system is mounted, and three flywheel engines, while the frame of the optoelectronic system is made with power ribs, between which flywheel engines are located at a distance from the star sensors and the gyroscope,

at the same time, the inlet of the optoelectronic system is closed by a camera cover, on the outer surface of which there is also a solar panel, on the upper panel of the spacecraft’s body there are antennas of the VHF range transceiver and GPS receiver, at the corners of the side panels from the side of the upper panel and the camera cover 4 support ledges are made.

2. The spacecraft according to claim 1, characterized in that the instruments of the orientation and stabilization control system are made in the form of two star sensors, one microelectromechanical gyroscope, three flywheel engines, as well as magnetometers, solar sensors and magnetic actuators as part of solar panels ,

3. The spacecraft according to claim 1, characterized in that the side panels are additionally interconnected.

4. The spacecraft according to claim 1, characterized in that the place of attachment of the side panels to each other from the side of the optoelectronic system is additionally reinforced at the corners with scarves.

5. The spacecraft according to claim 1, characterized in that the solar panel located on the outside of the cover is made on the basis of a multilayer printed circuit board with photoelectric converters.

6. The spacecraft according to claim 1, characterized in that the side and top solar panels are made on the basis of multilayer printed circuit boards with photoelectric converters and with built-in solar sensors, magnetometers, magnetic actuators and a microcontroller.

7. The spacecraft according to claim 1, characterized in that the lid of the chamber is equipped with torsion springs and is fixed in the closed state by a tension thread.

8. The spacecraft according to claim 1, characterized in that the opening of the lid under the influence of torsion springs is carried out by burning the tension thread.

9. The spacecraft according to claim 1, characterized in that the spacecraft is arranged to launch, launch and separate as part of a transport launch container.