Space station multi-truss system

Abstract

The application discloses a space station multi-truss system, characterized in that the system comprises a truss module and a sun-orienting mechanism a; the truss module comprises a truss structure, an unfolding mechanism, a flexible solar wing, a connecting unlocking mechanism, an electromechanical adapter interface and a sun-orienting mechanism b; a plurality of truss modules are connected in parallel to form the multi-truss system through the connecting unlocking mechanism; and the unfolding mechanism sequentially connects the truss structure with a spacecraft to form an axial series system during unfolding. The space station multi-truss system disclosed by the application fully and reasonably utilizes a launch envelope space, provides a solar wing with unlimited expansion area and installation space, and meets the demand of a large spacecraft for increasingly large on-orbit power generation and power supply.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of space station overall design, and particularly relates to a space station multi-truss system. BACKGROUND

[0002] Space stations or large spacecrafts, such as the "Peace" space station, the international space station, and the "Tiangong" space station of China, have super-large main structures, strict astronaut life support systems, and a large number of platform devices and experimental devices, and have great energy demand and device installation demand.

[0003] Solar energy is the most widely used clean energy for spacecrafts at present, and therefore large spacecrafts often need to be equipped with large-area solar wings. However, conventional spacecrafts can only carry and install two large solar wings at a time during one launch, and the launch efficiency is often low. For example, the international space station used multiple launches to gradually establish the current 2500m 2 scale solar wing total area. Therefore, it is of great significance to study how to reasonably use the launch envelope, improve the launch efficiency of solar wings at a time, and increase the area and number of solar wings. SUMMARY

[0004] To solve the above technical problems, the application provides a space station multi-truss system, which comprises a truss module and a sun-orientation mechanism a;

[0005] The truss module comprises a truss structure, an unfolding mechanism, a flexible solar wing, a connection unlocking mechanism, an electromechanical adapter interface, a sun-orientation mechanism b, and a plurality of truss modules are connected in parallel by the connection unlocking mechanism to form a multi-truss system;

[0006] The sun-orientation mechanism a is used for connecting the spacecraft and the truss structure;

[0007] The flexible solar wing is symmetrically installed on the front and rear sides of the truss structure and is connected to the root of the flexible solar wing through the sun-orientation mechanism b;

[0008] The connection unlocking mechanism is symmetrically installed on the left and right sides of the truss structure;

[0009] The unfolding mechanism is installed on the left and right sides of two adjacent truss structures, and when unfolded, the unfolding mechanism sequentially connects the truss structures and the spacecraft to form an axial series system;

[0010] The truss structure is provided with an electromechanical adapter interface on the top surface and the bottom surface, and after the unfolding mechanism is unfolded, each truss structure forms a series system through the electromechanical adapter interface.

[0011] Further, the truss structure is a polyhedron with six or more docking surfaces.

[0012] Further, the sun-orientation mechanism b corresponds to the flexible solar wing one by one, each truss module has two sun-orientation mechanisms b and two flexible solar wings.

[0013] Further, the number of truss modules increases with the increase of the size requirement of the flexible solar wing, the truss module fixedly connected with the sun-orientation mechanism a is truss module 1, the truss module 1 is a reference module, and the multiple truss system adopts a parallel configuration in a left-right symmetry around the truss module 1 when launched.

[0014] Further, the symmetrical truss modules are rearranged by using a space station mechanical arm after the truss modules are launched into orbit, and the deployment mechanism is used for deployment and docking adaptation.

[0015] Further, the deployment mechanism is arranged alternately at the right upper corner and the right lower corner of the truss structure starting from the truss module 1.

[0016] Further, the axial series connection system is composed of truss module rotation deployment, the rearranged truss modules use the deployment mechanism, the non-deployed module rotates 90° to the truss module 1, forms an inverted triangular offset configuration, and then the non-deployed module rotates 90° to the deployed part, forms a straight line series connection configuration, and the process is sequentially performed.

[0017] Further, the truss module can be installed with a platform and space experiment equipment inside and outside.

[0018] Further, the truss module can be installed at any part of the spacecraft or launched alone under the allowable carrying envelope.

[0019] Further, the truss module is transferred to the left side or the right side of the truss module 1 by using a mechanical arm after being launched into orbit, and then sequentially rotated and deployed by using the deployment mechanism to form a straight line series connection configuration.

[0020] The present application has the following advantages:

[0021] The present application provides a multiple truss system with scalable quantity and variable configuration, the total area of the solar wing power generation array and the number of trusses depend on the power demand of power generation and power supply in orbit, and the carrying envelope space during launch can be fully utilized, and the power generation, power supply and carrying capacity of the spacecraft can be doubled. BRIEF DESCRIPTION OF DRAWINGS

[0022] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or prior art description. Obviously, the drawings in the following description are some embodiments of the present application, and for those skilled in the art, other drawings can also be obtained from these drawings without creative labor. It should be noted that the drawings are not necessarily drawn to scale for clarity and ease of illustration.

[0023] Figure 1 Schematic diagram of a truss module 1 of the present application;

[0024] Figure 2 Schematic diagram of a truss module 2 of the present application;

[0025] Figure 3 Schematic diagram of a truss module 3 of the present application;

[0026] Figure 4 Schematic diagram of a first stage of the present application;

[0027] Figure 5 Schematic diagram of a second stage of the present application;

[0028] Figure 6 Schematic diagram of a third stage of the present application;

[0029] Markings in the figures:

[0030] 1 - truss module 1, 101 - truss structure, 102 - electromechanical adapter interface, 103 - deployment mechanism 1, 104 - front sun-tracking orientation mechanism b, 105 - right connection unlocking mechanism, 106 - front solar wing, 107 - sun-tracking orientation mechanism a, 108 - left connection unlocking mechanism, 109 - rear solar wing, 110 - rear sun-tracking orientation mechanism b;

[0031] 2 - truss module 2, 201 - truss structure, 202 - upper electromechanical adapter interface, 203 - front sun-tracking orientation mechanism b, 204 - right connection unlocking mechanism, 205 - front solar wing, 206 - deployment mechanism 2, 207 - lower electromechanical adapter interface, 208 - left connection unlocking mechanism, 209 - rear solar wing, 210 - rear sun-tracking orientation mechanism b;

[0032] 3 - truss module 3, 301 - truss structure, 302 - upper electromechanical adapter interface, 303 - deployment mechanism 3, 304 - front sun-tracking orientation mechanism b, 305 - right connection unlocking mechanism, 306 - front solar wing, 307 - lower electromechanical adapter interface, 308 - left connection unlocking mechanism, 309 - rear solar wing, 310 - rear sun-tracking orientation mechanism b.

[0033] Specific implementation

[0034] The application will be further described below with reference to the drawings.

[0035] In this embodiment, the principle and content of the application are introduced in detail mainly by taking one launch carrying three truss modules as an example.

[0036] The structure of the truss module 1 is shown in Figure 1 The truss module 1 is defined as a reference module, which comprises a truss structure 101, a deployment mechanism 103, two connection unlocking mechanisms, i.e., a left connection unlocking mechanism 108 and a right connection unlocking mechanism 105, for connection and unlocking with other truss modules, a mechanical and electrical adaptation interface 102 for mechanical connection with other truss modules and connection and transmission of electrical power and signals, a sun-orientation mechanism a 107 for connecting the truss structure 101 with a spacecraft and realizing sun-orientation rotation, two flexible solar wings, i.e., a front solar wing 106 and a rear solar wing 109, and two sun-orientation mechanisms b, i.e., a front sun-orientation mechanism b 104 and a rear sun-orientation mechanism b 110.

[0037] The structure of the truss module 2 is shown in Figure 2 which comprises a truss structure 201, a deployment mechanism 206 located at the lower right corner of the module, two connection unlocking mechanisms, i.e., a left connection unlocking mechanism 208 and a right connection unlocking mechanism 204, two mechanical and electrical adaptation interfaces, i.e., an upper mechanical and electrical adaptation interface 202 and a lower mechanical and electrical adaptation interface 207, two flexible solar wings, i.e., a front solar wing 205 and a rear solar wing 209, and two sun-orientation mechanisms b, i.e., a front sun-orientation mechanism b 203 and a rear sun-orientation mechanism b 210.

[0038] The structure of the truss module 3 is shown in Figure 3 which comprises a truss structure 301, a deployment mechanism 303 located at the upper right corner of the module, two connection unlocking mechanisms, i.e., a left connection unlocking mechanism 308 and a right connection unlocking mechanism 305, two mechanical and electrical adaptation interfaces, i.e., an upper mechanical and electrical adaptation interface 302 and a lower mechanical and electrical adaptation interface 307, two flexible solar wings, i.e., a front solar wing 306 and a rear solar wing 309, and two sun-orientation mechanisms b, i.e., a front sun-orientation mechanism b 304 and a rear sun-orientation mechanism b 310.

[0039] The truss structures 101, 201 and 301 are all polyhedrons with six or more interfacing surfaces.

[0040] Preferably, the truss is in the shape of a cuboid, in which the front and rear sides are provided with solar wings, and the left and right sides are provided with other trusses through the connection unlocking mechanisms for parallel installation with other trusses during launch, and a single truss can be connected with two other trusses at most.

[0041] AsFigure 4 The first stage of the multi-truss system variable configuration process is shown in the schematic diagram. When launched, the truss module 1 is in the center, the truss module 2 is on the right, and the truss module 3 is on the left. The truss system is left-right symmetrical and parallel configuration. The serial number from left to right is 3-1-2. After entering the orbit, the mechanical arm grabs the left module 3. Then the connection unlocking mechanism between the truss module 1 and the truss module 3 is unlocked. The mechanical arm transfers the module 3 to the rightmost side. The connection unlocking mechanism between the modules 2 and 3 is locked. The deployment mechanism between the modules 2 and 3 is locked. The offset parallel configuration is formed, and the serial number from left to right is 1-2-3.

[0042] As shown in Figure 5 The second stage of the multi-truss system variable configuration process is shown in the schematic diagram. The connection unlocking mechanism between the truss module 1 and the truss module 2 is unlocked. The deployment mechanism 1 is used to rotate the combination of the truss module 2 and the truss module 3 around the truss module 1 by 90° until 180°. The mechanical and electrical adaptation interfaces of the modules 1 and 2 are docked and locked. The inverted triangular offset configuration is formed.

[0043] As shown in Figure 6 The third stage of the multi-truss system variable configuration process is shown in the schematic diagram. The connection unlocking mechanism between the truss module 2 and the truss module 3 is unlocked. The deployment mechanism 2 is used to rotate the truss module 3 around the truss module 2 by 90° until 180°. The mechanical and electrical adaptation interfaces of the modules 2 and 3 are docked and locked. The final form of the straight-line series configuration is formed.

[0044] Preferably, the truss module can continue to install truss modules 4 and 5 to the required number on the left and right sides under the premise of allowing the carrying envelope. After entering the orbit, the mechanical arm is used to transfer all the truss modules to one side (left or right) of the truss module 1. Then the deployment mechanism is used to rotate and deploy in sequence to form a straight-line series configuration.

[0045] Preferably, the truss module can be installed to other parts of the spacecraft, including but not limited to the outside and inside of the spacecraft, the outside and inside of other spacecraft, and separate launch of the truss module. After entering the orbit, the mechanical arm is used to transfer all the truss modules to the left or right side of the truss module 1. Then the deployment mechanism is used to rotate and deploy in sequence to form a straight-line series configuration.

[0046] Preferably, the variable configuration process can be simplified to directly use the principle of building blocks, and the mechanical arm, on-orbit robot, or astronaut is used to install and dock to form a straight-line series configuration.

[0047] Although the present application has been disclosed with reference to the preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application without departing from the spirit and scope of the present application by using the disclosed methods and technical contents. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application without departing from the technical solutions of the present application shall fall within the protection scope of the present application. The contents not described in detail in the present application specification are the known technology of the person skilled in the art.

Claims

1. A multi-truss system for a space station, characterized in that: The system includes a truss module and a solar orientation mechanism a. The truss module includes a truss structure, a deployment mechanism, a flexible solar array, a connection and unlocking mechanism, an electromechanical adapter interface, and a solar orientation mechanism b. Multiple truss modules are connected in parallel through the connection and unlocking mechanism to form a multi-truss system. The solar orientation mechanism a is used for the connection between the spacecraft and the truss structure; The flexible solar array is symmetrically installed on the front and rear sides of the truss structure and is connected to the root of the flexible solar array through the solar orientation mechanism b. The connection unlocking mechanism is symmetrically installed on the left and right sides of the truss structure; The deployment mechanism is installed on the left and right sides of two adjacent truss structures. When deployed, the deployment mechanism sequentially connects the truss structures with the spacecraft to form an axial series system. The truss structure is provided with electromechanical adapter interfaces on its top and bottom surfaces. After the deployment mechanism is deployed, the truss structures form a series system through the electromechanical adapter interfaces. The number of truss modules increases with the increase in the size requirements of the flexible solar array. The truss module fixedly connected to the solar orientation mechanism a is the first truss module (1), which is the reference module. When the multi-truss system is launched, the truss modules adopt a left-right symmetrical parallel configuration arranged with the first truss module (1) as the center. After the truss modules are launched into orbit, the space station's robotic arm rearranges the symmetrical truss modules, and the deployment mechanism unfolds and docks them for adaptation.

2. The space station multi-truss system as described in claim 1, characterized in that: The truss structure is a polyhedron with at least six mating surfaces.

3. A space station multi-truss system as described in claim 1, characterized in that: The solar orientation mechanism b corresponds one-to-one with the flexible solar array, and each truss module has two solar orientation mechanisms b and two flexible solar arrays.

4. A space station multi-truss system as described in claim 1, characterized in that: The deployment mechanism is arranged alternately at the upper right corner and lower right corner of the truss structure, starting from the first truss module (1).

5. A space station multi-truss system as described in claim 1, characterized in that: The multi-truss system is composed of truss modules that are rotated and unfolded. The rearranged truss modules utilize an unfolding mechanism. The un-unfolded module rotates 90° relative to the first truss module (1) to form an inverted triangular offset configuration. Then, the un-unfolded module rotates 90° relative to the unfolded part to form a straight-line serial configuration. This process is repeated sequentially.

6. A space station multi-truss system as described in claim 1, characterized in that: Platforms and space experimental equipment can be installed inside and outside the truss module.

7. A space station multi-truss system as described in claim 1, characterized in that: The truss module can be installed anywhere on the spacecraft or launched separately, provided the payload envelope allows.

8. A space station multi-truss system as described in claim 1 or 7, characterized in that: After the truss module is put into the track, the robotic arm is used to transfer all the truss modules to the left or right side of the first truss module (1), and then the unfolding mechanism is used to rotate and unfold them in sequence to form a linear series configuration.

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