Small GEO communication satellite configuration

Abstract

The invention provides a small GEO communication satellite configuration. The small GEO communication satellite configuration comprises a satellite main body structure, a solar cell array, an antenna unit and a satellite-rocket separation mechanism, the satellite main body structure comprises structural components, the structural components are assembled into a trapezoidal hexahedral box-type star body, side plates are arranged in the X direction, the Y direction and the Z direction in a three-dimensional coordinate system OXYZ with the geometric center of the structural components as the original point O, the side plates comprise a + Z side plate, a-Z side plate, a + Y side plate, a-Y side plate, a + X side plate, a-X side plate and an OX side plate, and the + Z side plate, the-Z side plate, the + Y side plate, the-Y side plate, the + X side plate, the-X side plate and the OX side plate are arranged in the X direction, the Y direction and the Z direction respectively. The + Z side plate, the-Z side plate, the + Y side plate, the-Y side plate, the + X side plate and the-X side plate are connected to form a trapezoidal hexahedral box-type star, the OX side plate is mounted between the + Y side plate and the-Y side plate and abuts against the-Z side plate, and the OX side plate is located at the center line of the + Y side plate, the-Y side plate and the-Z side plate. The design of a traditional bearing cylinder is removed, the flexibility of the internal space of the satellite is improved, and the space utilization efficiency is improved.

Description

Technical Field

[0001] The present application relates to the technical field of spacecraft structures, and in particular to a small GEO communication satellite configuration. Background Art

[0002] Satellite configuration design is a crucial component of overall satellite design. Its fundamental mission is to determine the satellite's primary load-bearing structure, module composition, dimensions, and satellite-rocket interface within various constraints, and to complete the layout design of various satellite equipment. A suitable satellite configuration design can better meet payload requirements, reducing satellite envelope size and weight, and improving satellite and launch space utilization.

[0003] Traditional GEO (geostationary orbit) satellite structures mostly adopt a rectangular structure with a central load-bearing tube. This configuration is technically mature and has high mechanical strength. The cylindrical space of the load-bearing tube can be used to install chemical propulsion fuel tanks, such as the LS-1300 satellite platform of the US Laura Company, the SpaceBus4000 platform of the European TAS Company, the Eurostar3000 platform of the European ADS Company, and the DFH-4 satellite platform of China.

[0004] However, with the rapid growth in global demand for GEO (geostationary orbit) communication satellites, low-cost, high-efficiency satellite design has become crucial. In this context, GEO satellites will have greater advantages in being miniaturized, lightweight, and using launch methods such as "multiple satellites on one rocket." The traditional rectangular configuration of the central support tube is no longer suitable for miniaturization and lightweighting due to its complex structure. The rectangular configuration also has a relatively low space utilization rate in the cylindrical rocket fairing, and as higher specific impulse electric propulsion is increasingly used, the propulsion tank is significantly reduced, and the internal space of the cylindrical support tube will be vacant, making it difficult for other square equipment on the satellite to be used.

[0005] In order to reduce the envelope size of GEO satellites, reduce the weight of GEO satellites, improve the utilization efficiency of GEO satellite internal space, and improve the utilization efficiency of rocket fairing space, the satellite configuration must be optimized accordingly. Summary of the Invention

[0006] To achieve the above-mentioned purpose, this application proposes a small GEO communication satellite configuration, including: a satellite main structure, a solar cell array, an antenna unit and a satellite-rocket separation mechanism; the satellite main structure includes structural components, the structural components are assembled into a trapezoidal hexahedral box-type satellite body, and the solar cell array, the satellite-rocket separation mechanism and the antenna unit are respectively installed on the structural components.

[0007] Specifically, in a three-dimensional coordinate system OXYZ with the geometric center of the structural component as the origin O, side panels are respectively arranged along the X direction, the Y direction and the Z direction.

[0008] Specifically, the side panels include: +Z side panel, -Z side panel, +Y side panel, -Y side panel, +X side panel, -X side panel and OX side panel, wherein the +Z side panel, -Z side panel, +Y side panel, -Y side panel, +X side panel and -X side panel are connected to form a trapezoidal hexahedral box-type star body, the OX side panel is installed between the +Y side panel and the -Y side panel and abuts against the -Z side panel, and the OX side panel is located at the midline of the +Y side panel, -Y side panel and -Z side panel.

[0009] Through the above technical solution, not only the traditional load-bearing tube design is eliminated, the flexibility of the satellite's internal space is increased, and the space utilization efficiency of the satellite platform is improved; in addition, the difference from the existing trapezoidal hexahedron box structure is that the OX side panel is set up, which not only serves to separate functional spaces, but also can replace the load-bearing frame to support the side panel structure.

[0010] Specifically, the side panels are aluminum honeycomb sandwich panels.

[0011] Specifically, the solar cell array includes: ±Y solar wing deployment panels, ±Y solar wing root hinges, ±Y solar wing inter-panel hinges and ±Y solar wing compression and release devices, wherein the ±Y solar wing root hinges are respectively connected to the +Y side panel and the -Y side panel, the ±Y solar wing root hinges are respectively connected to the ±Y solar wing deployment panels, the ±Y solar wing span compression and release devices are installed on the ±Y solar wing deployment panels, and the ±Y solar wing deployment panels respectively include at least two solar wing panels, which are connected to each other through the ±Y solar wing inter-panel hinges.

[0012] Specifically, the antenna unit includes: phased array antenna, measurement and control antenna, beacon antenna, V antenna and QV feed antenna; the phased array antenna, measurement and control antenna and beacon antenna are installed on the +Z side panel facing the ground; the QV feed antenna is installed on the -Z side panel; and the V antenna is installed on the +X side panel.

[0013] Specifically, the satellite-rocket separation mechanism is set on the -X side panel.

[0014] Specifically, the QV feed antenna and the V antenna point to the +Z direction after the satellite enters orbit.

[0015] It also includes a phased array antenna, wherein the phased array antenna is installed on the +Z side panel.

[0016] Compared with the prior art, the present invention is beneficial in that:

[0017] 1. The traditional load-bearing tube design has been removed, which increases the flexibility of the satellite's internal space, improves the space utilization efficiency of the satellite platform, and adapts to the electric propulsion system.

[0018] 2. Compared with the conventional rectangular configuration, the trapezoidal configuration has two asymmetric sides. When the planar phased array antenna is arranged on the large trapezoidal surface, it has more antenna area, can obtain more antenna gain, lower satellite envelope size, and lower satellite weight.

[0019] 3. It can launch two satellites with one rocket. The cylindrical rocket fairing significantly improves the utilization efficiency of the carrier rocket and reduces the launch cost of a single satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] The accompanying drawings are included to provide a further understanding of the embodiments and are incorporated into and constitute a part of this specification. The accompanying drawings illustrate the embodiments and, together with the description, serve to explain the principles of the present application. Other embodiments and many of the expected advantages of the embodiments will be readily apparent as they become better understood by reference to the following detailed description. The elements of the drawings are not necessarily to scale with respect to each other. Like reference numerals designate corresponding similar parts.

[0021] Figure 1 1 is a schematic diagram of the overall structure of a small GEO communication satellite configuration according to an embodiment of the present application;

[0022] Figure 2 This is a structural diagram of structural components of a small GEO communication satellite configuration according to an embodiment of the present application.

[0023] The meaning of the numbers in the figure are: 1. Structural components; 1a. +X side panel; 1b. -X side panel; 1c. -Y side panel; 1d. +Y side panel; 1e. OX side panel; 1f. -Z side panel; 1g. +Z side panel; 2. Solar cell array; 3. Phased array antenna; 4. QV feed antenna; 5. Satellite-rocket separation mechanism; 6. V antenna; 7. Beacon antenna; 8. Tracking and control antenna. DETAILED DESCRIPTION

[0024] In the following detailed description, reference is made to the accompanying drawings, which form a part of the detailed description and are shown by illustrative specific embodiments in which the present application can be practiced. In this regard, directional terms, such as "top", "bottom", "left", "right", "up", "down", etc., are used with reference to the orientation of the figures being described. Because the components of the embodiments can be positioned in several different orientations, directional terms are used for illustrative purposes and are in no way limiting. It should be understood that other embodiments can be utilized or logical changes can be made without departing from the scope of the present application. Therefore, the following detailed description should not be taken in a limiting sense, and the scope of the present application is defined by the appended claims.

[0025] like Figure 1 and Figure 2As shown, a small GEO communication satellite configuration includes: a satellite main structure, a solar cell array 2, an antenna unit and a satellite-rocket separation mechanism 5; the satellite main structure includes a structural component 1, which is assembled into a trapezoidal hexahedral box-type satellite body, and the solar cell array 2, the satellite-rocket separation mechanism 5 and the antenna unit are respectively installed on the structural component 1.

[0026] In a three-dimensional coordinate system OXYZ with the geometric center of the structural component 1 as the origin O, side panels are respectively arranged along the X direction, the Y direction and the Z direction.

[0027] The side panels include: +Z side panel 1g, -Z side panel 1f, +Y side panel 1d, -Y side panel 1c, +X side panel 1a, -X side panel 1b and OX side panel 1e, wherein the +Z side panel 1g, -Z side panel 1f, +Y side panel 1d, -Y side panel 1c, +X side panel 1a and -X side panel 1b are connected to form a trapezoidal hexahedral box-type star body, the OX side panel 1e is installed between the +Y side panel 1d and the -Y side panel 1c and abuts against the -Z side panel 1f, and the OX side panel 1e is located at the center line of the +Y side panel 1d, -Y side panel 1c and -Z side panel 1f.

[0028] The solar cell array 2 includes: ±Y solar wing deployment panels, ±Y solar wing root hinges, ±Y solar wing inter-panel hinges and ±Y solar wing compression and release devices, wherein the ±Y solar wing root hinges are respectively connected to the +Y side panel 1d and the -Y side panel 1c, the ±Y solar wing root hinges are respectively connected to the ±Y solar wing deployment panels, the ±Y solar wing span compression and release devices are installed on the ±Y solar wing deployment panels, and the ±Y solar wing deployment panels each include at least two solar wing panels, which are connected to each other through the ±Y solar wing inter-panel hinges.

[0029] The antenna unit includes: phased array antenna 3, measurement and control antenna 8, beacon antenna 7, V antenna 6 and QV feed antenna 4; the phased array antenna 3, measurement and control antenna 8 and beacon antenna 7 are installed on the +Z side panel 1g facing the ground; the QV feed antenna 4 is installed on the -Z side panel 1f; the V antenna 6 is installed on the +X side panel 1a.

[0030] The satellite-rocket separation mechanism 5 is provided on the -X side plate 1b.

[0031] The QV feeding antenna 4 and the V antenna 6 point to the +Z direction after the satellite enters orbit.

[0032] It also includes a phased array antenna 3, wherein the phased array antenna 3 is installed on the +Z side panel 1g.

[0033] In a specific embodiment, the satellite mechanical coordinate system is defined as follows: the coordinate origin O is located in the mechanical separation plane between the satellite-rocket separation mechanism 5 and the carrier rocket, and coincides with the geometric center of gravity of the satellite side;

[0034] OZ axis (satellite's yaw axis): parallel to the satellite-rocket separation plane, passing through the origin O, with the positive direction pointing to the direction of the satellite relative to the earth after entering orbit;

[0035] OX axis (satellite's roll axis): perpendicular to the separation plane of the rocket and spacecraft, passing through the origin O, with the positive direction pointing to the flight direction of the launch vehicle and the positive direction pointing to the flight direction of the satellite after entering orbit;

[0036] OY axis (satellite's pitch axis): forms a right-handed rectangular coordinate system with the OZ and OX axes;

[0037] The OXYZ coordinate system is a rectangular coordinate system fixed to the satellite. The OX axis is also the vertical upward direction when the satellite is parked vertically on the ground, and its axis passes through the center of the Earth. When the satellite is in orbit, under theoretical attitude conditions, the OX axis lies within the orbital plane of the satellite, with its positive direction pointing in the direction of the satellite's forward motion. The positive direction of the OZ axis points toward the center of the Earth, and the OY axis is aligned with the normal to the orbital plane.

[0038] Based on the OXYZ coordinate system constructed above, the satellite configuration adopts a trapezoidal hexahedron box design, providing installation and load-bearing support for other subsystems, forming a complete spacecraft. To accommodate the entire satellite equipment, maximize the ground layout space for the mission payload, and fully utilize the launch envelope space, the satellite body is designed as a trapezoidal box structure, composed of seven aluminum honeycomb sandwich panels: ±Z, ±Y, ±X, and OX.

[0039] The satellite configuration in the above embodiment takes into account the requirement of launching two satellites with one rocket. By adopting the above satellite configuration, the satellites are mainly distributed in the cylindrical section inside the fairing. This design can be adapted to a 4200mm fairing, where the diameter of the available space of the fairing is 3850mm. Preliminary analysis of the three-dimensional simulation assembly shows that the static envelope requirements inside the fairing can be met.

[0040] Obviously, those skilled in the art can make various modifications and changes to the embodiments of the present application without departing from the spirit and scope of the present application. In this way, if these modifications and changes are within the scope of the claims of the present application and their equivalents, the present application is also intended to cover these modifications and changes. The word "comprising" does not exclude the presence of other elements or steps not listed in the claims. The simple fact that certain measures are recited in mutually different dependent claims does not indicate that the combination of these measures cannot be used to advantage. Any reference signs in the claims should not be considered as limiting the scope.

Claims

1. A small GEO communication satellite configuration, characterized in that: include: A satellite main structure, a solar cell array, an antenna unit, and a satellite-rocket separation mechanism; the satellite main structure includes structural components, which are assembled into a trapezoidal hexahedron box-type satellite body. The solar cell array, the satellite-rocket separation mechanism, and the antenna unit are respectively mounted on the structural components. In a three-dimensional coordinate system OXYZ with the geometric center of the structural component as the origin O, side panels are respectively arranged along the X direction, the Y direction, and the Z direction. The side panels include: a +Z side panel, a -Z side panel, a +Y side panel, a -Y side panel, a +X side panel, a -X side panel, and an OX side panel. The +Z side panel, -Z side panel, +Y side panel, -Y side panel, +X side panel, and -X side panel are connected to form the trapezoidal hexahedron box-type satellite body. The OX side panel is mounted between the +Y side panel and the -Y side panel and abuts against the -Z side panel, and the OX side panel is located at the midline of the +Y side panel, the -Y side panel, and the -Z side panel.

2. A small GEO communication satellite configuration according to claim 1, characterized in that: The side panels are aluminum honeycomb sandwich panels.

3. A small GEO communication satellite configuration according to claim 2, characterized in that: The solar cell array includes: ±Y solar wing deployment panels, ±Y solar wing root hinges, ±Y solar wing inter-panel hinges and ±Y solar wing compression and release devices, wherein the ±Y solar wing root hinges are respectively connected to the +Y side panel and the -Y side panel, the ±Y solar wing root hinges are respectively connected to the ±Y solar wing deployment panels, the ±Y solar wing expansion compression and release devices are installed on the ±Y solar wing deployment panels, the ±Y solar wing deployment panels each include at least two solar wing panels, and the solar wing panels are connected to each other through the ±Y solar wing inter-panel hinges.

4. The small GEO communication satellite configuration according to claim 2, characterized in that: The antenna unit includes: a phased array antenna, a measurement and control antenna, a beacon antenna, a V antenna and a QV feed antenna; the phased array antenna, the measurement and control antenna and the beacon antenna are installed on the +Z side panel facing the ground; the QV feed antenna is installed on the -Z side panel; and the V antenna is installed on the +X side panel.

5. The small GEO communication satellite configuration according to claim 2, characterized in that: The satellite-rocket separation mechanism is arranged on the -X side plate.

6. A small GEO communication satellite configuration according to claim 4, characterized in that: The QV feeding antenna and the V antenna point to the +Z direction after the satellite enters orbit.

7. The small GEO communication satellite configuration according to claim 2, characterized in that: It also includes a phased array antenna, wherein the phased array antenna is installed on the +Z side panel.

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