A configuration method of a high-efficiency compact serial satellite platform

Abstract

The application discloses a kind of configuration methods of high-efficient compact serial satellite platform, and serial satellite platform includes main force structure 1, load instrument cabin 2, platform instrument cabin 3, solar wing 4, thruster 5 and TT&C antenna 6, when the serial satellite platform is installed on launch vehicle and prepares to launch, it is in the state of retraction, the serial satellite platform is converted to the unfolded state after launch vehicle and main star transfer to specified orbit position in space.In the application, a kind of configuration layout method of serial launch satellite platform is designed to meet the interface requirements of current launch vehicle in China, the interface requirements of large satellite platform and the far point orbit transfer mode, which realizes a kind of configuration method of satellite platform for serial launch, which meets the interface requirements of current launch vehicle in China, the interface requirements of large communication satellite in China and the far point orbit transfer mode.

Description

TECHNICAL FIELD

[0001] The present application relates to the satellite overall configuration design technical field, and particularly relates to a configuration method of a high-efficiency compact serial satellite platform. BACKGROUND

[0002] Satellite configuration design is an important link in satellite overall design, and advanced configuration design can realize optimal design of satellite capacity by using existing space constraints.

[0003] The piggyback launch satellite is launched by using the capacity margin of the launch vehicle when launching the main load (hereinafter referred to as main satellite), and the launch vehicle pushes the main satellite and the piggyback launch satellite to the subscribed orbit in space.

[0004] The piggyback launch satellite platform is designed and produced for the convenience of piggybacking satellites, and the main structure form of the piggyback launch satellite, the power supply and distribution subsystem, the propulsion subsystem, the control subsystem, the measurement and control subsystem and the integrated electronic subsystem of the piggyback launch satellite are solidified. The installation interface of the satellite load is reserved. The configuration characteristics of the piggyback launch satellite platform can represent the configuration characteristics of the piggyback launch satellite based on the platform.

[0005] The serial piggyback launch satellite platform refers to the piggyback launch satellite, the launch vehicle and the main satellite arranged in series in the launch state. The present application is characterized by the configuration mode from bottom to top in the order of "launch vehicle -> piggyback launch satellite platform -> main satellite".

[0006] The "high-efficiency compact serial satellite platform" is a serial piggyback launch satellite platform suitable for the current launch capacity of China, which is based on the following requirements: the interface requirements / ambient space constraints of the mainstream large-thrust launch vehicle in China; the interface requirements / ambient space constraints of the mainstream large satellite platform (referring to the platform used by the main satellite) in China; the carrying capacity margin of the mainstream large-thrust launch vehicle launch capacity to the mainstream large satellite platform SUMMARY

[0007] The present application aims to solve the above problems, and a configuration layout method of a serial piggyback launch satellite platform matched with the interface requirements of the current launch vehicle in China, the interface requirements of the large satellite platform and the apogee orbit transfer mode is designed to meet the market demand of low-cost piggyback launch in China, and a configuration method of a high-efficiency compact serial satellite platform is proposed.

[0008] In order to achieve the above purpose, the present application adopts the following technical scheme:

[0009] The configuration method of the high-efficiency compact serial satellite platform comprises a main load-bearing structure, a load instrument cabin, a platform instrument cabin, a solar wing, a thruster and a measurement and control antenna.

[0010] Preferably, the satellite platform is in a retracted state when mounted on the launch vehicle, and is converted to an expanded state after the launch vehicle and the main satellite are transferred to a designated orbit position in space.

[0011] Preferably, the method comprises the following steps:

[0012] S1, a main load-bearing structure configuration mode of the satellite platform and a positional relationship with the launch vehicle and the main satellite;

[0013] S2, installation positions and configuration modes of a device cabin of the satellite platform, i.e., a platform instrument cabin and a load instrument cabin, on the platform;

[0014] S3, installation positions and configuration modes of solar wings of the satellite platform on the platform;

[0015] S4, installation positions and orientations of thrusters of the satellite platform on the platform;

[0016] S5, installation positions and orientations of TT&C antennas of the satellite platform on the platform.

[0017] Preferably, a mechanical coordinate system of the satellite platform is as follows:

[0018] The origin of the coordinate system is defined at the center of a main load-bearing structure and a launch vehicle interface ring;

[0019] The Z-axis coincides with the axis of the main load-bearing structure and is directed to the center of a main satellite interface ring;

[0020] The X-axis is in the plane of the main load-bearing structure and the launch vehicle interface ring and is directed to the direction of an expandable TT&C antenna mounting surface.

[0021] Preferably, a flight coordinate system of the satellite platform is as follows:

[0022] The origin of the coordinate system is the same as that of the mechanical coordinate system of the satellite platform, and the plane in which the X-axis and the Y-axis are located is in the XOY plane of the mechanical coordinate system.

[0023] As described above, the application has the following advantages:

[0024] 1. The application is designed to meet the market demand for low-cost launch in China, and a configuration layout method of a serial satellite platform for launch is designed to meet the interface requirements of the current launch vehicle in China, the interface requirements of a large satellite platform, and the far point orbit transfer mode. A high-efficiency and compact serial satellite platform configuration method is proposed, and a configuration method of a satellite platform for serial launch is realized, which meets the interface requirements of the current launch vehicle in China, the interface requirements of a large communication satellite in China, and the far point orbit transfer mode. BRIEF DESCRIPTION OF DRAWINGS

[0025] Figure 1 Fig. 1 shows a schematic diagram of a general configuration of a serial satellite platform in a folded state according to an embodiment of the present application;

[0026] Figure 2 Fig. 2 shows a schematic diagram of a general configuration of a serial satellite platform in an unfolded state according to an embodiment of the present application;

[0027] Figure 3 Fig. 3 shows a schematic diagram of a position relationship between a satellite platform, a carrier rocket and a main satellite according to an embodiment of the present application;

[0028] Figure 4 Fig. 4 shows a schematic diagram of a relationship between a platform mechanical coordinate system and an on-orbit flight according to an embodiment of the present application;

[0029] Figure 5 Fig. 5 shows a schematic diagram of an installation position and configuration mode of a payload instrument cabin on a platform according to an embodiment of the present application;

[0030] Figure 6 Fig. 6 shows a schematic diagram of an installation position and configuration mode of a platform instrument cabin on a platform according to an embodiment of the present application;

[0031] Figure 7 Fig. 7 shows a schematic diagram of an installation position and configuration mode of a solar wing on a platform according to an embodiment of the present application;

[0032] Figure 8 Fig. 8 shows a schematic diagram of an installation position and pointing direction of a thruster on a platform according to an embodiment of the present application;

[0033] Figure 9 Fig. 9 shows a schematic diagram of an installation position and pointing direction of a TT&C antenna on a platform according to an embodiment of the present application.

[0034] LEGEND

[0035] 1, main load-bearing structure; 2, payload instrument cabin; 3, platform instrument cabin; 4, solar wing; 5, thruster; 6, TT&C antenna. DETAILED DESCRIPTION

[0036] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the protection scope of the present application.

[0037] Please refer to Figures 1-9The application provides a technical scheme:

[0038] A configuration method of a high-efficiency compact satellite platform in series, which comprises a main load-bearing structure 1, a load instrument cabin 2, a platform instrument cabin 3, a solar wing 4, a thruster 5 and a TT&C antenna 6.

[0039] Specifically, the satellite platform in series is in a folded state when it is installed on a launch vehicle for launching, and its form is as shown in Figure 1 After the launch vehicle and the main satellite are transferred to a designated orbit position in space, the satellite platform in series is converted into an unfolded state, and its form is as shown in Figure 2 Figure 1 Figure 2 The coordinate systems in the two figures are mechanical coordinate systems of the satellite platform in series;

[0040] A configuration method of a high-efficiency compact satellite platform in series, which comprises the following steps:

[0041] S1, a configuration mode of a main load-bearing structure 1 of the satellite platform in series and a positional relationship with a launch vehicle and a main satellite;

[0042] S2, installation positions and configuration modes of equipment cabins, i.e., the platform instrument cabin 3 and the load instrument cabin 2, on the platform;

[0043] S3, an installation position and a configuration mode of a solar wing 4 on the platform;

[0044] S4, an installation position and a pointing direction of a thruster 5 on the platform;

[0045] S5, an installation position and a pointing direction of a TT&C antenna 6 on the platform.

[0046] A mechanical coordinate system of the satellite platform in series is as follows:

[0047] The origin of the coordinate system is defined at the center of a docking annular ring of the main load-bearing structure 1 and the launch vehicle;

[0048] The Z-axis coincides with the axis of the main load-bearing structure 1 and points to the center of the docking annular ring of the main satellite;

[0049] The X-axis is in the plane of the docking annular ring of the main load-bearing structure 1 and the launch vehicle and points to the direction of the installation surface of the TT&C antenna 6;

[0050] The Y-axis meets the right-hand rule with the X-axis and the Z-axis.The mechanical coordinate system of the satellite platform in series is parallel to the mechanical coordinate system of the main satellite, and the origin of the mechanical coordinate system of the satellite platform in series has coordinates (0, 0, -H) in the mechanical coordinate system of the main satellite (the definition of H is shown in Figure 3 ​), the definition method of the mechanical coordinate system is convenient for describing the space avoidance relationship between the product carried on the satellite platform, the main satellite and the launch vehicle;

[0051] The satellite platform is installed on the launch vehicle in the preparation stage of launching, and the installation mode is shown in Figure 3 . The satellite platform is installed on the instrument cabin of the launch vehicle, the main satellite is installed on the satellite platform, the Z axes of the launch vehicle, the satellite platform and the main satellite are coincident, and the satellite-rocket combination has a serial configuration mode.

[0052] The flight coordinate system of the satellite platform is as follows:

[0053] The origin of the coordinate system is the same as the origin of the mechanical coordinate system of the satellite platform, and the X axis and the Y axis are in the XOY plane of the mechanical coordinate system;

[0054] The X axis (X 在轨 ) is rotated clockwise by 30° relative to the X axis (X 搭载平台 ) of the mechanical coordinate system. The definition method of the flight coordinate system makes the X 在轨 point to the flight direction of the satellite, and the Z 在轨 point to the ground, which is convenient for describing the on-orbit attitude of the satellite.

[0055] The on-orbit flight attitude of the satellite is shown in Figure 4 . The sun wing 4 in the unfolded state is coincident with the geographic north direction in the -Y on-orbit direction of the satellite, and the TT&C antenna 6 in the unfolded state is approximately coincident with the geographic east direction in the +X on-orbit direction of the satellite;

[0056] The installation position and allowable space of the load instrument cabin 2 installed on the satellite platform are shown in Figure 5 . The radius of the docking ring of the satellite is R1, the inner diameter (radius) of the fairing of the launch vehicle is R2, and the height of the main load-bearing structure of the satellite is H. The installation position of the load instrument cabin 2 on the platform is as follows: the load instrument cabin 2 is installed on the -X side of the mechanical coordinate system, and the required installation range is in the form of a sector with an angle of 60°, which is symmetrical relative to the X axis, and has an outer ring radius C1, an inner ring radius D1, an inner height A1 and an outer height B1. Then, C1 = R2-50mm, D1 = R1+170mm, A1 = H-100mm, and B1 = A1+150mm.

[0057] The installation position and allowable space of the platform instrument cabin 3 installed on the satellite platform are shown in Figure 6 . The installation position of the platform instrument cabin 3 on the platform is as follows: the platform instrument cabin 3 is installed on the side of the mechanical coordinate system that is rotated clockwise by 60° relative to the X axis, and the required installation range is in the form of a sector with an angle of 60°, which has an outer ring radius C2, an inner ring radius D2, an inner height A2 and an outer height B2. Then, C2 = R2-50mm, D2 = R1+160mm, A2 = H-100mm, and B2 = A2+150mm.

[0058] The mounting position of the solar wing 4 on the mounting platform is shown in Figure 7 The mounting position of the solar wing 4 on the mounting platform is shown in

[0059] The mounting position and pointing direction of the thruster 5 on the mounting platform are shown in Figure 8 The mounting position and pointing direction of the thruster 5 on the mounting platform are shown in

[0060] The mounting position and pointing direction of the thruster 5 on the mounting platform are shown in Figure 9 The mounting position and pointing direction of the thruster 5 on the mounting platform are shown in

[0061] The capabilities of the satellite launching platform developed according to the application are shown in Table 1:

[0062] Table 1 Capabilities of the satellite launching platform

[0063]

[0064] The application is aimed at the market demand for low-cost satellite launching in China, and a serial satellite launching platform configuration layout method is designed, which matches the current launch vehicle interface requirements in China, the large satellite platform interface requirements, and the far point orbit transfer working mode. A high-efficiency and compact serial satellite platform configuration method is proposed, and a satellite platform configuration method for serial satellite launching is realized, which is adapted to the current launch vehicle interface requirements in China, the large communication satellite interface requirements in China, and the far point orbit transfer working mode.

[0065] The above description of the embodiments enables those skilled in the art to implement or use the application. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the application. Therefore, the application will not be limited to these embodiments shown herein, but will conform to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A configuration method for a highly efficient and compact serial satellite platform, characterized in that, The serial satellite platform includes a main load-bearing structure (1), a payload instrument compartment (2), a platform instrument compartment (3), a solar array (4), a thruster (5), and a telemetry and control antenna (6); The payload instrument compartment (2) is installed in the 60° sector area on the -X side of the main load-bearing structure (1), and its installation space satisfies the following: the outer ring radius is equal to the inner diameter radius of the launch vehicle fairing minus 50mm, the inner ring radius is equal to the docking ring radius of the main load-bearing structure plus 170mm, the inner height is equal to the height of the main load-bearing structure minus 100mm, and the outer height is equal to the inner height of the payload instrument compartment (2) plus 150mm; The platform instrument compartment (3) is installed in the fan-shaped area on the side of the main load-bearing structure (1) rotated 60° clockwise around the X-axis, and its installation space satisfies the following: the outer ring radius is equal to the inner diameter radius of the launch vehicle fairing minus 50mm, the inner ring radius is equal to the docking ring radius of the main load-bearing structure plus 160mm, the inner height is equal to the height of the main load-bearing structure minus 100mm, and the outer height is equal to the inner height of the platform instrument compartment (3) plus 150mm.

2. The configuration method for a highly efficient and compact serial satellite platform according to claim 1, characterized in that, When the serial satellite platform is installed on the launch vehicle in preparation for launch, it is in a retracted state. After the launch vehicle and the host satellite are transferred to the designated orbital position in space, the serial satellite platform is converted into an deployed state.

3. The configuration method for a highly efficient and compact serial satellite platform according to claim 2, characterized in that, Includes the following steps: S1. Configuration of the main load-bearing structure of the satellite platform and its positional relationship with the launch vehicle and the main satellite; S2. The installation location and configuration of the equipment compartment carrying the satellite platform, namely the platform instrument compartment (3) and the payload instrument compartment (2) on the platform; S3. The installation location and configuration of the solar array (4) on the satellite platform; S4. The installation position and orientation of the thruster (5) mounted on the satellite platform on the platform; S5. The installation position and orientation of the telemetry and control antenna (6) mounted on the satellite platform on the platform.

4. The configuration method for a highly efficient and compact serial satellite platform according to claim 3, characterized in that, The mechanical coordinate system of the satellite platform is: The origin of the coordinate system is defined at the center of the docking ring between the main load-bearing structure (1) and the launch vehicle; The Z-axis coincides with the axis of the main load-bearing structure (1) and points towards the center of the main star docking ring; The X-axis is located within the plane of the main load-bearing structure (1) and the docking ring of the launch vehicle, pointing towards the mounting surface of the deployable telemetry and control antenna (6).

5. The configuration method for a highly efficient and compact serial satellite platform according to claim 4, characterized in that, The flight coordinate system of the satellite platform is as follows: The origin of the coordinate system is the same as the origin of the mechanical coordinate system on the satellite platform, and the planes containing the X and Y axes lie in the XOY plane of the mechanical coordinate system.

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