A-satellite platform configuration

Abstract

The application provides an A-shaped satellite platform configuration, which is centered on a payload, and provides an A-shaped satellite platform configuration, which designs a satellite platform as an A-shaped truss structure, comprises a central truss and a honeycomb structure plate, and realizes integrated design of the satellite platform and the payload. The bottom of the A-shaped platform is used for bearing axial and shearing loads when being connected with a multi-point connection interface of a carrier; by fully utilizing an envelope space, internal space design of the satellite configuration is optimized, main transmission path design is optimized; according to the size of the equipment envelope size, reasonable layout and placement are performed, and the cable direction, the plug space of the electrical connector are fully considered, so that the compact and reasonable structure space can be ensured, the expansion path requirement of the large-scale expandable antenna payload is met, and the area of the satellite solar cell array is ensured by fully utilizing the envelope space provided by the carrier rocket.

Description

Technical Field

[0001] This invention belongs to the field of satellite technology within aerospace technology, specifically relating to an A-shaped satellite platform configuration. More specifically, it relates to an A-shaped satellite platform configuration for mounting a large deployable antenna. Background Technology

[0002] As the performance indicators of satellite payloads continue to improve, the size of onboard antennas is also increasing, especially with the growing demand for satellites capable of carrying large deployable antennas. my country's existing satellite platforms generally feature a cubic configuration with dual solar panels on both sides, and an internal compartmentalized design consisting of a propulsion module, service module, and payload module. This configuration can meet the current need to keep payload components in operational condition by directly mounting them on the satellite platform or by folding the satellite into sections on its sides. However, for satellites carrying payloads with large deployable antennas, directly placing the payload on the satellite platform not only fails to meet the given radial envelope space and equipment installation and operation space requirements, but also fails to meet the given baseband frequency requirements.

[0003] The payload carried by the satellite platform of this invention is a large deployable antenna, which needs to be compressed in a folded state and can be unlocked and deployed in orbit. Existing satellite platforms cannot meet these requirements, which imposes stringent constraints on the satellite's configuration and layout design.

[0004] Patent document CN107600460A, entitled "A Low-Temperature Optical Satellite Configuration Suitable for Complex Illumination Conditions in a Highly Elliptical Frozen Orbit," discloses a low-temperature optical satellite configuration suitable for complex illumination conditions in a highly elliptical frozen orbit. The main purpose is to ensure the satellite has a stable, large-area shaded surface to meet the imaging temperature requirements of the low-temperature optical camera. This differs from the A-shaped satellite platform configuration proposed in this invention, which can be used to mount a large deployable antenna, and the method for installing a large-area planar antenna under limited constraints and ensuring the structure meets the dynamic environment requirements of the launch vehicle. These belong to different technical directions and application fields.

[0005] Patent document CN106742063A discloses an "Integrated Satellite Configuration". This patent document discloses an integrated satellite configuration, which is mainly for realizing the unlocking and separation of the payload compartment, non-contact high-precision and high-stability attitude control of the two compartments, and attitude maneuvering function under repeated locking of the two compartments. This is completely different from the A-shaped satellite platform configuration with a large deployable antenna proposed in this invention, which uses the method of point-type pyrotechnics with a common nodal circle design to achieve reliable separation of the satellite and rocket. Summary of the Invention

[0006] In view of the deficiencies in the prior art, the purpose of this invention is to provide an A-shaped satellite platform configuration.

[0007] According to the present invention, an A-shaped satellite platform configuration includes: a central truss 1, a base plate 201, a +Y upper side plate 202, a -Y upper side plate 206, a +X upper side plate 210, a +X lower side plate 211, a -X upper side plate 212, a -X lower side plate 213, a +Y lower side plate assembly, and a -Y lower side plate assembly;

[0008] Wherein, the origin O is the theoretical center of the point-type separation device at the bottom of the central truss 1 and the star-rocket separation surface; the Z-axis is perpendicular to the star-rocket separation surface along the origin and points to the satellite; the X-axis is inside the star-rocket separation surface, perpendicular to the Z-axis and pointing to the +X upper side plate 210 direction; the Y-axis is in a right-handed relationship with the X and Z axes.

[0009] Preferably, the +Y lower side panel assembly includes a first +Y lower side panel 203, a second +Y lower side panel 204, and a third +Y lower side panel 205 distributed in sequence; the -Y lower side panel assembly includes a first -Y lower side panel 207, a second -Y lower side panel 208, and a third -Y lower side panel 209 distributed in sequence.

[0010] Preferably, the angle of the cone at the top of the central truss 1 is 13°.

[0011] Preferably, the central truss 1 is divided into multiple regions by the base plate 201, the +Y upper side plate 202, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the -Y upper side plate 206, the first -Y lower side plate 207, the second -Y lower side plate 208, the third -Y lower side plate 209, the +X upper side plate 210, the +X lower side plate 211, the -X upper side plate 212, and the -X lower side plate 213.

[0012] Preferably, the inner surfaces of the base plate 201, the +Y upper side plate 202, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the -Y upper side plate 206, the first -Y lower side plate 207, the second -Y lower side plate 208, the third -Y lower side plate 209, the +X upper side plate 210, the +X lower side plate 211, the -X upper side plate 212, and the -X lower side plate 213 serve as the mounting base surfaces for the spaceborne equipment.

[0013] Preferably, it also includes: +Y-wing solar cell array 301 and -Y-wing solar cell array 302;

[0014] The +Y wing solar cell array 301 and -Y wing solar cell array 302 are respectively closed and pressed onto the +Y upper side plate 202 and the -Y upper side plate 206; solar cells are mounted on the surfaces of the +Y upper side plate 202, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the -Y upper side plate 206, the first -Y lower side plate 207, the second -Y lower side plate 208, and the third -Y lower side plate 209 facing the external space.

[0015] Preferably, the base plate 201, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the first -Y lower side plate 207, the second -Y lower side plate 208, the third -Y lower side plate 209, the +X lower side plate 211, and the -X lower side plate 213 form a closed three-dimensional structure, and the interior serves as the mounting base for spaceborne equipment, power subsystems, and hydrazine cylinders.

[0016] Preferably, it is suitable for large deployable antennas.

[0017] Preferably, the bottom surface of the eight truss legs on the central truss 1-Z surface is designed as the star-rocket separation surface. The center of the eight truss legs adopts a common pitch circle design with a diameter of Φ2200 and is evenly arranged. Each mounting leg is equipped with a point-type separation device to realize the connection and separation between the satellite and the launch vehicle.

[0018] Preferably, the eight thrusters 4 are mounted on the platform base plate, with two thrusters each arranged on the ±X side and the ±Y side, and the thrust direction is consistent with +Z, providing attitude and orbit maintenance and control functions for the satellite.

[0019] Compared with the prior art, the present invention has the following beneficial effects:

[0020] 1. Based on fully utilizing the envelope space of the launch vehicle satellite fairing and ensuring the area of ​​the satellite solar cells, this invention provides an A-shaped satellite platform configuration that meets the requirements for the folding and on-orbit deployment of large deployable antennas, provides installation space for instruments and equipment, and ensures the energy supply of the entire satellite.

[0021] 2. In this invention, during satellite assembly, the outer side panels on the +Y and -Y sides can be opened. Based on the base plate, central truss, and upper and lower top plates, the internal frame ensures the installation of onboard equipment and cable connections on the inner side panels on the +Y and -Y sides. After installation, the outer side panels on the +Y and -Y sides are then installed. This method not only ensures the assembly accuracy of the satellite structure but also guarantees sufficient space for satellite assembly operations, improving operational efficiency.

[0022] 3. The satellite platform configuration provided by this invention has a stable structure and can bear large-size and high-mass payloads. It is highly adaptable to large payloads that require high precision and are difficult to install, such as large deployable antennas.

[0023] 4. The configuration design method adopted in this invention can meet the on-orbit application requirements of flat panel antennas of different sizes, which is convenient for operation and implementation. In particular, it can be extended to the situation of parallel development of multiple satellites, which can effectively reduce the cycle of satellite launch and networking.

[0024] 5. The bottom of the A-shaped platform described in this invention is used to withstand axial and shear loads when connected to the multi-point connection interface of the launch vehicle; by making full use of the envelope space, the internal space design of the satellite configuration is optimized, and the main force transmission path design is optimized; according to the size of the equipment envelope, the layout and placement are reasonable, and the cable routing and the insertion and removal space of the electrical connectors are fully considered, which can ensure the compactness and rationality of the structural space, while meeting the deployment path requirements of the large deployable antenna payload, and ensuring the area of ​​the satellite solar cell array by making full use of the envelope space provided by the launch vehicle. Attached Figure Description

[0025] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0026] Figure 1 This is a schematic diagram of the satellite platform configuration of the present invention;

[0027] Figure 2 This is an exploded view of the satellite platform of the present invention.

[0028] The diagram shows:

[0029] Detailed Implementation

[0030] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0031] like Figure 1 The diagram shown is an exploded view of the satellite platform configuration of this invention. Considering the characteristics of the aforementioned A-shaped satellite platform configuration, for ease of description, a satellite layout coordinate system O-XYZ is first established, defined as follows:

[0032] Origin of coordinates: The theoretical center of the point-type separation device at the bottom of the central truss 1 and the star-rocket separation surface;

[0033] Z-axis: perpendicular to the star-rocket separation surface along the origin of the coordinate system, pointing towards the satellite;

[0034] X-axis: Inside the star-rocket separation surface, perpendicular to the Z-axis, pointing towards the +X upper side plate 210 direction;

[0035] Y-axis: forms a right-handed relationship with the X and Z axes.

[0036] The satellite carries a large deployable antenna that needs to be compressed in a folded state and unlocked and deployed in orbit. To meet the launch envelope requirements and ensure the satellite's mechanical performance during launch and the space for the antenna to move in orbit, while also maximizing the area of ​​the solar array, the satellite platform is constructed in an A-shape configuration. The A-shape configuration design and internal space design are optimized. To fully utilize the envelope space provided by the launch vehicle and maximize the area of ​​the satellite's solar cells, solar cells are attached to the outer sides of the +Y upper side plate 202, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, and the outer sides of the -Y upper side plate 206, the first -Y lower side plate 207, the second -Y lower side plate 208, and the third -Y lower side plate 209, as needed.

[0037] The main structure of the satellite platform consists of a central truss 1, a base plate 201, an upper +Y side plate 202, a first lower +Y side plate 203, a second lower +Y side plate 204, a third lower +Y side plate 205, an upper -Y side plate 206, a first lower -Y side plate 207, a second lower -Y side plate 208, a third lower -Y side plate 209, an upper +X side plate 210, a lower +X side plate 211, an upper -X side plate 212, and a lower -X side plate 213, which together form a closed platform configuration with an A-shaped cross section.

[0038] The central truss 1 is divided into multiple regions by the base plate 201, the +Y upper side plate 202, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the -Y upper side plate 206, the first -Y lower side plate 207, the second -Y lower side plate 208, the third -Y lower side plate 209, the +X upper side plate 210, the +X lower side plate 211, the -X upper side plate 212, and the -X lower side plate 213, effectively improving the rigidity of the platform structure and providing a better load-bearing capacity. It provides the mounting base for the clamping seat and the space for unlocking and unfolding. The inner surfaces of the base plate 201, +Y upper side plate 202, first +Y lower side plate 203, second +Y lower side plate 204, third +Y lower side plate 205, -Y upper side plate 206, first -Y lower side plate 207, second -Y lower side plate 208, third -Y lower side plate 209, +X upper side plate 210, +X lower side plate 211, -X upper side plate 212, and -X lower side plate 213 serve as the mounting base for the spaceborne equipment. The base plate 201, the first +Y lower side plate 203, the second +Y lower side plate 204, the third +Y lower side plate 205, the first -Y lower side plate 207, the second -Y lower side plate 208, the third -Y lower side plate 209, the +X lower side plate 211, and the -X lower side plate 213 constitute a closed three-dimensional structure, which serves as the mounting base for spaceborne equipment, power subsystems, and hydrazine cylinders.

[0039] The satellite platform configuration of this invention optimizes the main force transmission path design while limiting the envelope space and the equipment installation space. Eight point-type separation devices with a shared pitch circle design on the central truss 1 and the base plate 201 serve as the main load-bearing structure for the force transmission path. Impact and vibration loads from the vehicle are transmitted to the honeycomb structure plates, such as the side plates, through the central truss 1.

[0040] During the overall satellite assembly, the following panels are opened: +Y upper side panel 202, first +Y lower side panel 203, second +Y lower side panel 204, third +Y lower side panel 205, -Y upper side panel 206, first -Y lower side panel 207, second -Y lower side panel 208, third -Y lower side panel 209, +X upper side panel 210, +X lower side panel 211, -X upper side panel 212, and -X lower side panel 213. Based on the base plate 201 and the central truss 1, the installation of onboard equipment and cable connections can be ensured through internal structural panels. This method not only guarantees the satellite's overall assembly operation space, reduces the assembly difficulty, and improves operational efficiency, but also avoids blind spots during the satellite assembly process.

[0041] The bottom surface of the eight truss legs on the Z-plane of the central truss is designed as the separation surface between the satellite and the launch vehicle. The center of the eight truss legs adopts a common pitch circle design with a diameter of Φ2200, which can be evenly arranged. Each mounting leg can be equipped with a point separation device to realize the connection and separation between the satellite and the launch vehicle.

[0042] In a variation, a large deployable antenna can be mounted on the +X upper side plate 210 and the -X upper side plate 212.

[0043] This invention creates a new design method for satellite configuration technology that will subsequently carry large deployable antennas in my country. In particular, for the installation of large deployable antennas on satellites, the configuration design method of folding and unfolding combination has become the first choice for radar and remote sensing satellite configuration designers. This invention will have a wide range of applications in the field.

[0044] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0045] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. An A-shaped satellite platform configuration, characterized in that, include: Together they form a closed A-shaped platform configuration, consisting of a central truss (1), a base plate (201), a +Y upper side plate (202), a -Y upper side plate (206), a +X upper side plate (210), a +X lower side plate (211), a -X upper side plate (212), a -X lower side plate (213), a +Y lower side plate assembly, and a -Y lower side plate assembly; Among them, the origin O is the theoretical center of the point-type separation device at the bottom of the central truss (1) and the star-rocket separation surface; the Z-axis is perpendicular to the star-rocket separation surface along the origin and points to the satellite; the X-axis is inside the star-rocket separation surface, perpendicular to the Z-axis and points to the direction of the +X upper side plate (210); the Y-axis is in a right-handed relationship with the X and Z axes. The +Y lower side panel assembly includes a first +Y lower side panel (203), a second +Y lower side panel (204), and a third +Y lower side panel (205) distributed in sequence; the -Y lower side panel assembly includes a first -Y lower side panel (207), a second -Y lower side panel (208), and a third -Y lower side panel (209) distributed in sequence. The central truss (1) is divided into multiple areas by the base plate (201), the +Y upper side plate (202), the first +Y lower side plate (203), the second +Y lower side plate (204), the third +Y lower side plate (205), the -Y upper side plate (206), the first -Y lower side plate (207), the second -Y lower side plate (208), the third -Y lower side plate (209), the +X upper side plate (210), the +X lower side plate (211), the -X upper side plate (212), and the -X lower side plate (213); The inner surfaces of the base plate (201), the +Y upper side plate (202), the first +Y lower side plate (203), the second +Y lower side plate (204), the third +Y lower side plate (205), the -Y upper side plate (206), the first -Y lower side plate (207), the second -Y lower side plate (208), the third -Y lower side plate (209), the +X upper side plate (210), the +X lower side plate (211), the -X upper side plate (212), and the -X lower side plate (213) serve as the mounting base surfaces for the spaceborne equipment.

2. The A-shaped satellite platform configuration according to claim 1, characterized in that, The angle of the cone at the top of the central truss (1) is 13°.

3. The A-shaped satellite platform configuration according to claim 1, characterized in that, Also includes: +Y-wing solar cell array (301), -Y-wing solar cell array (302); The +Y wing solar cell array (301) and -Y wing solar cell array (302) are respectively closed and pressed onto the +Y upper side plate (202) and the -Y upper side plate (206); solar cells are mounted on the surfaces of the +Y upper side plate (202), the first +Y lower side plate (203), the second +Y lower side plate (204), the third +Y lower side plate (205), the -Y upper side plate (206), the first -Y lower side plate (207), the second -Y lower side plate (208), and the third -Y lower side plate (209) facing the external space.

4. The A-shaped satellite platform configuration according to claim 1, characterized in that, The base plate (201), the first +Y lower side plate (203), the second +Y lower side plate (204), the third +Y lower side plate (205), the first -Y lower side plate (207), the second -Y lower side plate (208), the third -Y lower side plate (209), the +X lower side plate (211), and the -X lower side plate (213) constitute a closed three-dimensional structure, which serves as the mounting base for spaceborne equipment, power subsystems, and hydrazine cylinders.

5. The A-shaped satellite platform configuration according to claim 1, characterized in that, Suitable for large deployable antennas.

6. The A-shaped satellite platform configuration according to claim 1, characterized in that, The bottom surface of the eight truss legs of the central truss (1)-Z surface is designed as the star-rocket separation surface. The center of the eight truss legs adopts a common pitch circle design with a diameter of Φ2200 and is evenly arranged. Each mounting leg is equipped with a point separation device to realize the connection and separation between the satellite and the launch vehicle.

7. The A-shaped satellite platform configuration according to claim 1, characterized in that, Eight thrusters (4) are installed on the platform base plate, with two on the ±X side and two on the ±Y side. The thrust direction is consistent with +Z, providing attitude and orbit maintenance and control functions for the satellite.

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