Satellite configuration loaded with deployable antenna and satellite

Abstract

The invention provides a satellite configuration loaded with a deployable antenna and a satellite. The satellite configuration loaded with the deployable antenna comprises a platform cabin, a load cabin truss, the deployable antenna, a solar array, a data transmission antenna and a star sensor, the platform cabin is connected with the load cabin truss, and the solar array and the deployable antenna are installed on the two sides, in the same direction, of the platform cabin and the load cabin truss respectively. A data transmission antenna is installed on one side face, located between the solar arrays on the two sides, of the platform cabin, and the star sensor is installed on the load cabin truss. The unfolding direction of the deployable antenna is perpendicular to the longitudinal direction of the load cabin truss, that is, the longitudinal direction of the deployable antenna is parallel to the longitudinal direction of the satellite, the longitudinal size of the deployable antenna can be greatly increased, and therefore the deployable antenna can adapt to installation of a large-area deployable antenna in a one-dimensional folding and unfolding mode.

Description

Technical Field

[0001] The present invention relates to the field of satellite technology, and in particular to a satellite configuration and a satellite equipped with a deployable antenna. Background Art

[0002] According to the existing satellite configuration design, most of the main payloads, such as cameras on optical satellites and reflective antennas on communication satellites, have a small envelope and can be directly installed on the satellite platform. However, for satellites equipped with large-sized deployable antennas, their envelope when deployed in orbit far exceeds the available envelope of the launch vehicle, so they need to be folded and gathered on the satellite platform in the launch state, and after entering orbit, they are deployed to form a whole large-sized antenna.

[0003] Against the background of continuous development of satellite payload technology, the payload caliber of spacecraft such as microwave remote sensing satellites, data relay satellites, electronic reconnaissance satellites, and mobile communication satellites is getting larger and larger. However, due to the space constraints of the fairing of the launch vehicle, there are great restrictions on the size envelope and weight of satellites equipped with large-size deployable antennas. Therefore, there are high requirements for the satellite configuration of large space deployable antennas.

[0004] According to the prior art search, the Chinese patent with the authorization announcement number CN112009724B discloses a wedge-shaped satellite configuration and its assembly method. The wedge-shaped satellite configuration includes a configuration body, which is a heptahedral wedge configuration and a closed space formed by seven side panels. The configuration of this patent is not suitable for the installation of large deployable antennas.

[0005] The existing Chinese patent with authorization announcement number CN110450983B discloses an agile satellite configuration. This patent is suitable for satellites with compact layout and small inertia, but is not suitable for carrying large antennas.

[0006] The existing Chinese patent with authorization announcement number CN109927938B discloses a geostationary orbit real-aperture microwave detection satellite configuration in the field of satellite general technology. The satellite configuration of this patent is aimed at the geostationary orbit real-aperture microwave detection satellite, and its payload is a parabolic antenna.

[0007] The existing Chinese patent with authorization announcement number CN107600460B discloses a cryogenic optical satellite configuration suitable for the complex lighting conditions of a large elliptical frozen orbit. The invention aims to ensure that the satellite has a stable large-size sun-facing surface under the complex lighting conditions of a large elliptical frozen orbit to ensure the imaging temperature environment requirements of the cryogenic optical camera.

[0008] The existing Chinese patent with authorization announcement number CN109927937B discloses a separated solar observation satellite configuration that can achieve full physical space isolation between cabins and eliminate the micro-vibration effect of the satellite platform on the payload, but does not involve the description of large antenna loading. Summary of the invention

[0009] In view of the defects in the prior art, an object of the present invention is to provide a satellite configuration and a satellite equipped with a deployable antenna.

[0010] A satellite configuration equipped with a deployable antenna provided by the present invention includes: a platform cabin, a payload cabin truss, a deployable antenna, a solar array, a data transmission antenna, and a star sensor;

[0011] The platform cabin is connected to the payload cabin truss, and a solar array and a deployable antenna are respectively installed on both sides of the platform cabin and the payload cabin truss in the same direction;

[0012] The platform cabin is provided with a data transmission antenna on one side surface between the solar arrays on both sides, and the star sensor is provided on the payload cabin truss.

[0013] Preferably, the transverse surface of the load compartment truss is arranged to be trapezoidal and extends in the longitudinal direction, and the transverse surface of the load compartment truss is parallel to the end surface of the platform cabin connected to one end of the load compartment truss.

[0014] Preferably, the deployable antenna is composed of a plurality of deployable antenna array surfaces connected together, and the plurality of deployable antenna array surfaces are folded and installed on the payload bay truss.

[0015] Preferably, the deployment direction of the deployable antenna is perpendicular to the longitudinal direction of the load bay truss, and the longitudinal direction of the deployable antenna is parallel to the longitudinal direction of the load bay truss.

[0016] Preferably, the transverse surface of the load compartment truss is arranged to be an isosceles trapezoid.

[0017] Preferably, the solar array is composed of a plurality of solar cell arrays connected together, and the plurality of solar cell arrays are folded and installed on the side of the platform cabin.

[0018] Preferably, when folded, the patches of the solar cell array face outwards.

[0019] Preferably, a T-shaped rod is installed on the platform cabin, and the data transmission antenna extends outward along the T-shaped rod.

[0020] Preferably, the inner angle of the long side of the isosceles trapezoid is 82°.

[0021] Compared with the prior art, the present invention has the following beneficial effects:

[0022] 1. The deployment direction of the deployable antenna in the present application is perpendicular to the longitudinal direction of the payload cabin truss, that is, the longitudinal direction of the deployable antenna is parallel to the longitudinal direction of the satellite, which can greatly increase the longitudinal dimension of the deployable antenna, thereby being able to adapt to the installation of a large-area deployable antenna in a one-dimensional folding and deployment mode;

[0023] 2. In this application, the star sensor and the deployable antenna are installed on the truss at the same time. The star sensor is a measurement component of the satellite attitude. The pointing relationship between the star sensor and the antenna can be calibrated on the ground. When in orbit, this relationship will undergo unknown changes due to thermal deformation and other influences. Placing the antenna and the star sensor as close as possible on a carbon fiber truss with less thermal deformation can reduce this unknown change and the influence of the deformation on the load in orbit, thereby meeting the requirements of precise attitude determination.

[0024] 3. The cross section of the payload compartment truss is projected into an isosceles trapezoid in this application, which can effectively reduce the cantilever length when the antenna is folded, reduce the response during the launch phase, and reduce the envelope of the entire satellite when folded;

[0025] 4. The data transmission antenna of the present application extends out of the star body through a T-shaped rod to ensure that the beam of the data transmission antenna is not blocked by the large-area antenna. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0027] Figure 1 This is a schematic diagram of the satellite configuration in a folded state;

[0028] Figure 2 This is a schematic diagram of the satellite configuration deployment state;

[0029] Figure 3 This is a top view of the satellite configuration when it is folded;

[0030] As shown in the figure:

[0031] DETAILED DESCRIPTION

[0032] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0033] Example 1

[0034] like Figure 1-3As shown, this embodiment includes: a platform cabin 1, a payload cabin truss 2, a deployable antenna 3, a solar array 4, a digital transmission antenna 5 and a star sensor 6; the platform cabin 1 is connected to the payload cabin truss 2, and the solar array 4 and the deployable antenna 3 are respectively installed on both sides of the platform cabin 1 and the payload cabin truss 2 in the same direction; the digital transmission antenna 5 is installed on one side of the platform cabin 1 located between the solar arrays 4 on both sides, and a T-shaped rod is installed on the platform cabin 1, and the digital transmission antenna 5 extends outward along the T-shaped rod. The star sensor 6 is installed on the payload cabin truss 2.

[0035] The deployable antenna 3 is composed of a plurality of deployable antenna array surfaces connected together, which are folded and installed on the payload cabin truss 2. The solar array 4 is composed of a plurality of solar cell arrays connected together, which are folded and installed on the side of the platform cabin 1, with the patch surface of the solar cell array facing outward when folded.

[0036] In one embodiment, the transverse surface of the load compartment truss 2 is set to be an isosceles trapezoid and extends in the longitudinal direction, and the transverse surface of the load compartment truss 2 is parallel to the end surface of the platform cabin 1 connected to one end of the load compartment truss 2. The deployment direction of the deployable antenna 3 is perpendicular to the longitudinal direction of the load compartment truss 2, and the longitudinal direction of the deployable antenna 3 is parallel to the longitudinal direction of the load compartment truss 2.

[0037] In one embodiment, the interior angle of the long side of the isosceles trapezoid is 82°.

[0038] Example 2

[0039] Example 2 is a preferred example of Example 1.

[0040] like Figure 1 As shown, this embodiment includes: a satellite body, a large-sized deployable antenna 3, a deployment mechanism, a solar array 4, a data transmission antenna 5 and a star sensor 6.

[0041] The satellite body includes a platform cabin 1 and a payload cabin, and the payload cabin is provided with a payload cabin truss 2. The solar array 4 is installed on both sides of the platform cabin 1. The data transmission antenna 5 is installed on the ground of the platform cabin 1 through the data transmission mounting plate, and the data transmission antenna 5 extends out of the satellite body through a T-shaped rod to ensure that the beam of the data transmission antenna 5 is not blocked by the large-area deployable antenna 3. The star sensor 6 and the deployable antenna 3 are installed on the payload cabin truss 2. During the launch phase, the large-size deployable antenna 3, the solar array 4, and the data transmission antenna 5 are folded and pressed against the side of the satellite body to ensure that the satellite envelope meets the requirements of the lateral envelope interval 7 in the carrier fairing.

[0042] The deployable antenna 3 is folded and placed on both sides of the satellite payload compartment to reduce the envelope of the entire satellite during launch. After the deployable antenna 3 is deployed in orbit, it is connected to the satellite through the deployment mechanism. The normal direction of the deployable antenna 3 array surface always remains perpendicular to the longitudinal direction of the satellite before and after deployment. The distance direction (i.e. Figure 1The deployable antenna 3 is connected to the satellite through an over-constrained connection after deployment to ensure the position accuracy and deployment accuracy of the deployable antenna 3 after deployment.

[0043] like Figure 2 As shown, after entering orbit, the solar array 4, the deployable antenna 3 and the data transmission antenna 5 are deployed step by step in sequence. The solar array 4 is deployed first, and its deployment path does not interfere with other components outside the satellite; then the deployable antennas 3 folded on both sides are deployed to form a whole antenna array; finally, the data transmission antenna 5 is deployed to the ground, and after deployment, it can meet the field of view requirements of ground data transmission.

[0044] In one embodiment, the array surface of the deployable antenna 3 is divided into four parts and folded and pressed in the payload cabin in the satellite launch state. The solar array 4 is installed on both sides of the satellite platform cabin 1, unlocked and unfolded after the satellite and rocket are separated, and when folded, the patch surface of the solar array (the side that can absorb solar energy) faces outward.

[0045] like Figure 3 As shown, in one embodiment, the cross-section of the load compartment truss 2 is an isosceles trapezoid, and the inner angle of the long side of the isosceles trapezoid is 82°, which can effectively reduce the cantilever length of the deployable antenna 3 when it is folded.

[0046] In the description of the present application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present application.

[0047] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A satellite configuration equipped with a deployable antenna, characterized in that: include: Platform cabin (1), payload cabin truss (2), deployable antenna (3), solar array (4), data transmission antenna (5) and star sensor (6); The platform cabin (1) is connected to the payload cabin truss (2), and a solar array (4) and a deployable antenna (3) are respectively installed on both sides of the platform cabin (1) and the payload cabin truss (2) in the same direction; The platform cabin (1) is provided with a data transmission antenna (5) on one side surface between the two side solar arrays (4), and the star sensor (6) is installed on the payload cabin truss (2).

2. The satellite configuration equipped with a deployable antenna according to claim 1, characterized in that: The transverse surface of the load compartment truss (2) is arranged in a trapezoidal shape and extends in the longitudinal direction, and the transverse surface of the load compartment truss (2) is parallel to the end surface of the platform cabin (1) connected to one end of the load compartment truss (2).

3. The satellite configuration equipped with a deployable antenna according to claim 1, characterized in that: The deployable antenna (3) is composed of a plurality of deployable antenna array surfaces connected together, and the plurality of deployable antenna array surfaces are folded and installed on the load compartment truss (2).

4. The satellite configuration equipped with a deployable antenna according to claim 2, characterized in that: The deployment direction of the deployable antenna (3) is perpendicular to the longitudinal direction of the load compartment truss (2), and the longitudinal direction of the deployable antenna (3) is parallel to the longitudinal direction of the load compartment truss (2).

5. The satellite configuration equipped with a deployable antenna according to claim 2, characterized in that: The transverse surface of the load compartment truss (2) is arranged to be an isosceles trapezoid.

6. The satellite configuration equipped with a deployable antenna according to claim 1, characterized in that: The solar array (4) is composed of a plurality of solar cell arrays connected together, and the plurality of solar cell arrays are folded and installed on the side of the platform cabin (1).

7. The satellite configuration equipped with a deployable antenna according to claim 6, characterized in that: When the solar cell array is folded, the patch faces outwards.

8. The satellite configuration equipped with a deployable antenna according to claim 1, characterized in that: A T-shaped rod is installed on the platform cabin (1), and the data transmission antenna (5) extends outward along the T-shaped rod.

9. The satellite configuration equipped with a deployable antenna according to claim 5, characterized in that: The inner angle of the long side of the isosceles trapezoid is 82°.

10. A satellite, characterized in that: A satellite configuration equipped with a deployable antenna as described in any one of claims 1 to 9 is adopted.

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