Combined spacecraft biaxial sun orientation control method

Abstract

The invention discloses a biaxial sun orientation control method for a combined spacecraft. The method comprises the following steps: determining a current configuration, a flight attitude and a working condition of the combined spacecraft; and selecting a sun-oriented control strategy matched with the current configuration, flight attitude and working condition of the combined spacecraft, and controlling the attitude of the biaxial sun-oriented system to realize biaxial sun-oriented control. According to the method, during the on-orbit building and operation period of the multi-cabin and multi-flight-attitude spacecraft, the generating capacity of the solar cell wing of each cabin reaches the optimal state, and stable power output is provided for each cabin of the combined spacecraft.

Description

Technical Field

[0001] The invention belongs to the technical field of power supply of a spacecraft power system, and in particular relates to a combined spacecraft dual-axis sun-orientation control method. Background Art

[0002] The power supply system provides a continuous supply of electricity for the flight of spacecraft and is an important part of the spacecraft. As the size of spacecraft continues to increase, spacecraft need to assemble and expand the cabins on orbit (combined spacecraft), the configuration is becoming more and more complex, and the flight attitude is becoming more and more variable. Under different flight states, the solar incidence angles of the solar panels of different cabins of the combined spacecraft are different, resulting in uneven power generation capacity of the solar panels of each cabin; therefore, the requirements for the solar panel's ability to orient to the sun are getting higher and higher. Under different cabin configurations and different flight attitudes, it is necessary to ensure that the solar panel has sufficient power generation capacity. Summary of the invention

[0003] The technology of the present invention solves the problem: Overcoming the shortcomings of the prior art, providing a dual-axis solar orientation control method for a combined spacecraft, aiming to achieve optimal power generation of the solar panels of each compartment during the in-orbit assembly and operation of a multi-compartment, multi-flight attitude spacecraft, and provide stable power output for each compartment of the combined spacecraft.

[0004] In order to solve the above technical problems, the present invention discloses a combined spacecraft dual-axis sun-orientation control method, comprising:

[0005] Determine the current configuration, flight attitude and operating conditions of the modular spacecraft;

[0006] A solar orientation control strategy that matches the current configuration, flight attitude and operating conditions of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control.

[0007] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, the combined spacecraft includes: an experimental module I, a core module and an experimental module II;

[0008] The configuration of the combined spacecraft includes: single-cabin configuration, two-cabin straight-line configuration, two-cabin L configuration and three-cabin T configuration; among them, single-cabin configuration: experimental cabin I or experimental cabin II; two-cabin straight-line configuration: experimental cabin I is coaxially docked with the core cabin to form a two-cabin straight-line combination; two-cabin L configuration: the experimental cabin I in the two-cabin straight-line combination is transferred to form a two-cabin L-type combination; three-cabin T configuration: the experimental cabin II is docked with the two-cabin L-type combination to form a three-cabin T-type combination;

[0009] The flight attitude of the combined spacecraft includes: single-cabin flight of experimental cabin I, single-cabin flight of experimental cabin II, flight of two-cabin straight-line combination, flight of two-cabin L-type combination and flight of three-cabin T-type combination; among which, single-cabin flight of experimental cabin I includes: single-cabin inertial flight of experimental cabin I and three-axis ground flight of single-cabin experimental cabin I; single-cabin flight of experimental cabin II includes: single-cabin inertial flight of experimental cabin II and three-axis ground flight of single-cabin experimental cabin II; flight of two-cabin straight-line combination includes: inertial flight of two-cabin straight-line combination and three-axis ground flight of two-cabin straight-line combination; flight of two-cabin L-type combination includes: three-axis ground flight of two-cabin L-type combination and docking preparation flight of two-cabin L-type combination; flight of three-cabin T-type combination includes: three-axis ground flight of three-cabin T-type combination, orbital flight of three-cabin T-type combination, radial docking preparation flight of three-cabin T-type combination and forward-backward docking preparation flight of three-cabin T-type combination;

[0010] The operating conditions of the combined spacecraft include: attitude control conditions, orbit control conditions, docking conditions, transfer conditions and evacuation conditions.

[0011] In the above-mentioned combined spacecraft dual-axis solar orientation control method, both experimental modules are equipped with a combined dual-degree-of-freedom solar orientation system; by controlling the attitude of the combined dual-degree-of-freedom solar orientation system, the solar cell wing of the experimental module works in the optimal solar orientation mode, so as to enhance the solar orientation directivity of the solar cell wing of the experimental module, make the solar incident angle reach the optimal state, and the solar cell wing converts solar energy to the maximum extent, thereby improving the power generation capacity of the solar cell wing and providing stable power output for the combined spacecraft.

[0012] In the above-mentioned combined spacecraft dual-axis solar orientation control method,

[0013] The combined dual-degree-of-freedom solar orientation system installed at the end of the experimental cabin I is recorded as a combined dual-degree-of-freedom solar orientation system I; wherein, the combined dual-degree-of-freedom solar orientation system I includes: an A-axis solar orientation subsystem I, a B-axis solar orientation subsystem I and a truss assembly I; the A-axis solar orientation subsystem I and the B-axis solar orientation subsystem I are orthogonally installed on the cabin body of the experimental cabin I through the truss assembly I; the A-axis solar orientation subsystem I is used to drive the solar cell wing of the experimental cabin I to rotate around the A1 axis, and the A1 axis is the cabin body axis of the experimental cabin I; the B-axis solar orientation subsystem I is used to rotate the solar cell wing of the experimental cabin I around the B1 axis, and the B1 axis is the axis of the solar cell wing of the experimental cabin I;

[0014] The combined dual-degree-of-freedom solar-directed system installed at the end of the experimental cabin II is recorded as the combined dual-degree-of-freedom solar-directed system II; wherein, the combined dual-degree-of-freedom solar-directed system II includes: an A-axis solar-directed orientation subsystem II, a B-axis solar-directed orientation subsystem II and a truss assembly II; the A-axis solar-directed orientation subsystem II and the B-axis solar-directed orientation subsystem II are orthogonally installed on the cabin body of the experimental cabin II through the truss assembly II; the A-axis solar-directed orientation subsystem II is used to drive the solar cell wing of the experimental cabin II to rotate around the A2 axis, and the A2 axis is the cabin body axis of the experimental cabin II; the B-axis solar-directed orientation subsystem II is used to rotate the solar cell wing of the experimental cabin II around the B2 axis, and the B2 axis is the axis of the solar cell wing of the experimental cabin II.

[0015] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0016] When the current flight attitude of the combined spacecraft is the single-cabin inertial system flight of the experimental cabin I, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin I are driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I;

[0017] When the current flight attitude of the combined spacecraft is the three-axis ground flight of the single cabin of the experimental cabin I, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin I are driven to rotate continuously around the B1 axis within the range of 0° to 360° through the B-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft;

[0018] When the current flight attitude of the combined spacecraft is the single-cabin inertial system flight of the experimental cabin II, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin II are driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II;

[0019] When the current flight attitude of the combined spacecraft is the three-axis ground flight of the single-cabin experimental module II, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental module II are driven by the B-axis solar orientation subsystem II to rotate continuously around the B2 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft.

[0020] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0021] When the current flight attitude of the combined spacecraft is the single-cabin flight of the experimental cabin I, and the current working condition is any one of the attitude control condition, orbit control condition, docking condition and evacuation condition, the A1 axis of the A-axis sun orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis sun orientation subsystem I is controlled to be in a horizontal zero position or a vertical zero position;

[0022] When the current flight attitude of the combined spacecraft is the single-cabin flight of the experimental cabin II, and the current working condition is any one of the attitude control condition, orbit control condition, docking condition and evacuation condition, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal zero position or a vertical zero position.

[0023] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0024] When the current flight attitude of the combined spacecraft is the inertial flight of the two-cabin combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ are driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem Ⅰ;

[0025] When the current flight attitude of the combined spacecraft is a three-axis ground-to-earth flight of a two-cabin in-line combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ are driven by the B-axis solar orientation subsystem Ⅰ to continuously rotate around the B1 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft.

[0026] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0027] When the current flight attitude of the combined spacecraft is a two-cabin in-line combination flight, and the current working condition is any one of the attitude control condition and the orbit control condition, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal zero position.

[0028] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0029] When the current flight attitude of the combined spacecraft is a three-axis ground-to-earth flight of a two-cabin L-shaped combination, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I;

[0030] When the current flight attitude of the combined spacecraft is that the two-cabin L-shaped combination is docked and ready for flight, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position.

[0031] In the above-mentioned dual-axis solar orientation control method for a combined spacecraft, a solar orientation control strategy matching the current configuration, flight attitude and working condition of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control, including:

[0032] When the current flight attitude of the combined spacecraft is a three-cabin T-type combination three-axis ground flight, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; the solar cell wing of the experimental cabin II is driven to rotate continuously around the A2 axis within the range of 0° to 360° through the A-axis solar orientation subsystem II, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I, and the solar cell wing of the experimental cabin II is driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II;

[0033] When the current flight attitude of the combined spacecraft is the three-cabin T-type combined orbital system flight, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; the solar cell wing of the experimental cabin II is driven to rotate continuously around the A2 axis within the range of 0° to 360° through the A-axis solar orientation subsystem II, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I, and the solar cell wing of the experimental cabin II is driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II;

[0034] When the current flight attitude of the combined spacecraft is the radial docking preparation for flight of the three-cabin T-type combination, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position;

[0035] When the current flight attitude of the combined spacecraft is that the three-cabin T-type combination is docked forward and backward in preparation for flight, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position.

[0036] The present invention has the following advantages:

[0037] The present invention discloses a method for dual-axis solar orientation control of a modular spacecraft, which is applicable to the occasions where a power supply system of a modular spacecraft is powered under multiple cabins and multiple flight attitudes. During the assembly and operation of the modular spacecraft, under different cabin configurations and different flight attitudes, the attitude of the dual-axis solar orientation system can be adjusted by adopting different solar orientation control strategies to achieve dual-axis solar orientation control, thereby ensuring that the power generation of the solar cell wings of each cabin is always in the optimal state under different cabin configurations and different flight attitudes, and providing stable power output for each cabin of the modular spacecraft. Under different cabin configurations and different flight attitudes, the fluctuation range of the power output of the power channel is reduced from 60% of the traditional solution to 7%, and the stability rate is greatly improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Figure 1 This is a flow chart of a combined spacecraft dual-axis sun-orientation control method according to an embodiment of the present invention;

[0039] Figure 2 This is a schematic diagram of the sun orientation mode of an experimental cabin I / II cabin during inertial flight in an embodiment of the present invention;

[0040] Figure 3 This is a schematic diagram of a solar orientation method for a single-cabin spacecraft during three-axis earth-to-earth flight in an embodiment of the present invention;

[0041] Figure 4 It is a schematic diagram of the sun orientation mode of a two-cabin inline combination during inertial flight in an embodiment of the present invention;

[0042] Figure 5 It is a schematic diagram of a sun-orientation method of a two-cabin in-line combination during three-axis ground-to-ground flight in an embodiment of the present invention;

[0043] Figure 6 It is a schematic diagram of the sun orientation mode of a two-cabin L-shaped combination during three-axis ground-to-ground flight in an embodiment of the present invention;

[0044] Figure 7 It is a schematic diagram of a sun-orientation method of a two-cabin L-shaped assembly docked and ready for flight in an embodiment of the present invention;

[0045] Figure 8 It is a schematic diagram of a sun-orientation method of a three-cabin T-shaped assembly during three-axis ground flight / three-cabin T-shaped assembly orbital flight in an embodiment of the present invention;

[0046] Fig. 9 This is a schematic diagram of a sun-orientation method for a three-cabin T-shaped assembly in radial docking and preparation for flight in an embodiment of the present invention;

[0047] Fig.10 This is a schematic diagram of a sun-orientation method of a three-cabin T-shaped assembly in an embodiment of the present invention when docking forward and backward to prepare for flight;

[0048] Fig.11 It is a structural schematic diagram of a combined dual-degree-of-freedom sun-aiming system in an embodiment of the present invention;

[0049] Fig.12 It is a block diagram of a combined dual-degree-of-freedom solar tracking system in an embodiment of the present invention;

[0050] Fig.13 It is a schematic diagram of equipment layout and connection relationship on one side of a truss assembly in an embodiment of the present invention;

[0051] Fig.14 is a schematic diagram of the equipment layout and connection relationship on the other side of a truss assembly in an embodiment of the present invention;

[0052] Fig.15 is a structural schematic diagram of a truss assembly in an embodiment of the present invention;

[0053] Fig.16 This is a schematic diagram of the structure of an A-axis sun orientation subsystem in an embodiment of the present invention;

[0054] Fig.17 It is a block diagram of the composition of a B-axis solar orientation subsystem in an embodiment of the present invention. DETAILED DESCRIPTION

[0055] In order to make the objectives, technical solutions and advantages of the present invention more clear, the embodiments disclosed in the present invention will be further described in detail below with reference to the accompanying drawings.

[0056] One of the core ideas of the present invention is that the power supply system provides energy guarantee for the modular spacecraft, which not only guarantees the power supply of each cabin during the autonomous flight, but also adapts to the changes in power demand brought about by the flight of various combinations of multiple cabins. On the one hand, the power supply system needs to provide basic cabin power supply for different cabin flight modes during the assembly process of the modular spacecraft. On the other hand, the power supply system needs to adapt to the changes in on-orbit loads for tasks such as orbit change, docking, transfer, and evacuation during the assembly process of the modular spacecraft. In this process, the flexible solar cell wing of the power supply system needs to adapt to the complex load conditions under different configurations of multiple cabins. According to the on-orbit assembly process and power supply requirements of large modular spacecraft, the present invention designs a dual-axis solar orientation control method for modular spacecraft, adopting a design scheme of decentralized layout and centralized control. The experimental cabin is equipped with a dual-axis solar orientation system, which realizes the dual-axis solar orientation of the solar cell wing through the A-axis solar orientation subsystem and the B-axis solar orientation subsystem. The dual-axis solar-directed directional control method of the combined spacecraft can meet the power supply demand of a single cabin of the experimental cabin during flight, and can also meet the power supply demand of each cabin section of the combined spacecraft under different configurations and flight attitudes during the assembly process of the combined spacecraft.

[0057] Reference Figure 1 In this embodiment, the combined spacecraft dual-axis sun-orientation control method includes:

[0058] Step 1: Determine the current configuration, flight attitude and operating conditions of the combined spacecraft.

[0059] In this embodiment, the modular spacecraft mainly includes: experimental cabin I, core cabin and experimental cabin II. Since the modular spacecraft is launched as a single cabin, assembled and operated on orbit, and has multiple and complex on-orbit combined configurations, the working mode of the power supply system must be able to adapt to all states from basic assembly to docking with manned spacecraft and cargo spacecraft. The assembly of the modular spacecraft is carried out by docking and assembly with the core cabin as the core.

[0060] Multiple configurations

[0061] The configurations of modular spacecraft mainly include: single-cabin configuration, two-cabin straight-line configuration, two-cabin L configuration and three-cabin T configuration.

[0062] Single cabin configuration: experimental cabin I or experimental cabin II.

[0063] Two-cabin in-line configuration: Experimental cabin I is coaxially docked with the core cabin to form a two-cabin in-line combination.

[0064] Two-cabin L-configuration: The experimental cabin I in the two-cabin straight-line combination is shifted to form a two-cabin L-shaped combination.

[0065] Three-cabin T-configuration: Experimental cabin II is docked with the two-cabin L-shaped combination to form a three-cabin T-shaped combination.

[0066] Multiple flight attitudes

[0067] The flight postures of the combined spacecraft mainly include: single-cabin flight of experimental cabin I, single-cabin flight of experimental cabin II, flight of two-cabin straight-line combination, flight of two-cabin L-shaped combination and flight of three-cabin T-shaped combination.

[0068] The single-cabin flight of Experimental Cabin I mainly includes: the single-cabin inertial system flight of Experimental Cabin I and the single-cabin three-axis ground flight of Experimental Cabin I.

[0069] The single-cabin flight of Experimental Cabin II mainly includes: the single-cabin inertial system flight of Experimental Cabin II and the single-cabin three-axis ground flight of Experimental Cabin II.

[0070] The flight of the two-cabin in-line combination mainly includes: the inertial system flight of the two-cabin in-line combination and the three-axis ground flight of the two-cabin in-line combination.

[0071] The flight of the two-cabin L-shaped combination mainly includes: three-axis ground flight of the two-cabin L-shaped combination and docking preparation flight of the two-cabin L-shaped combination.

[0072] The flight of the three-cabin T-type combination mainly includes: three-axis ground flight of the three-cabin T-type combination, orbital flight of the three-cabin T-type combination, radial docking preparation flight of the three-cabin T-type combination and forward and backward docking preparation flight of the three-cabin T-type combination.

[0073] Multiple working conditions

[0074] During the construction and operation of modular spacecraft, the cabin needs to complete various action tasks such as attitude control, orbit change, docking, transfer, and evacuation. Different cabin action tasks will produce different magnitudes and directions of impact loads on the solar panels. The working conditions of modular spacecraft mainly include: attitude control conditions, orbit control conditions, docking conditions, transfer conditions, and evacuation conditions. Among them:

[0075] Attitude control conditions: Factors affecting the load conditions of attitude control conditions include: cabin configuration, cabin attitude adjustment direction, GNC attitude adjustment mode, etc.

[0076] Orbit control conditions: Factors affecting the load conditions of orbit control conditions include: the position difference between the current orbit and the target orbit, orbit raising or lowering control, orbit control engine working time, cabin configuration, etc.

[0077] Docking conditions: Factors affecting the loading conditions of docking conditions include: docking port position, docking speed of the visiting aircraft, aerodynamic force of the visiting aircraft engine plume, and target aircraft cabin configuration.

[0078] Transposition condition: Factors affecting the load conditions of the transposition condition include: the flight attitude of the cabin during the transposition, the transposition acceleration of the transposition mechanism, etc.

[0079] Evacuation conditions: Factors affecting the load conditions of the transfer conditions include: docking position, evacuation speed of the evacuation aircraft, aerodynamic force of the evacuation aircraft engine plume, target aircraft cabin configuration, etc.

[0080] Combined dual-degree-of-freedom solar system

[0081] The power supply system uses a large-area flexible solar cell wing, which has the characteristics of large inertia, high flexibility, and dense modes. The dual-axis solar directional control must not only ensure smooth tracking of the sun under various cabin configurations, avoid resonance between the flexible solar cell wing and the aircraft attitude control, but also be able to withstand the load impact of various working conditions. In view of the complex load conditions of multiple cabins, multiple configurations, and multiple working conditions mentioned above, the method described in the present invention is based on a combined dual-degree-of-freedom solar system to achieve dual-axis solar directional control, and adopts three closed-loop controls of current, speed, and position to ensure smooth tracking of the sun and precise stopping. In order to adapt to the setting of the solar cell wing locking position under various flight conditions, the combined dual-degree-of-freedom solar system is designed with three control modes: horizontal zeroing, vertical zeroing, and angle servo, to achieve stop control at any position, and electrical locking and mechanical locking control modes are designed to provide different sizes of locking holding torques.

[0082] Both experimental modules are equipped with a combined dual-degree-of-freedom solar orientation system. By controlling the attitude of the combined dual-degree-of-freedom solar orientation system, the solar panels of the experimental module are operated in the optimal solar orientation mode to enhance the solar orientation directivity of the solar panels of the experimental module, so that the solar incident angle reaches the optimal state, the solar panels convert solar energy to the maximum extent, and the power generation capacity of the solar panels is improved, providing stable power output for the combined spacecraft. In addition, the core module is equipped with a single-degree-of-freedom solar orientation system. Specifically: the combined dual-degree-of-freedom solar-directed system installed at the end of the experimental cabin I is recorded as the combined dual-degree-of-freedom solar-directed system I; wherein, the combined dual-degree-of-freedom solar-directed system I includes: an A-axis solar-directed orientation subsystem I, a B-axis solar-directed orientation subsystem I and a truss assembly I; the A-axis solar-directed orientation subsystem I and the B-axis solar-directed orientation subsystem I are orthogonally installed on the body of the experimental cabin I through the truss assembly I; the A-axis solar-directed orientation subsystem I is used to drive the solar cell wing of the experimental cabin I to rotate around the A1 axis, and the A1 axis is the axis of the cabin body of the experimental cabin I; the B-axis solar-directed orientation subsystem I is used to rotate the solar cell wing of the experimental cabin I around the B1 axis, and the B1 axis is the axis of the solar cell wing of the experimental cabin I. The combined dual-degree-of-freedom solar-directed system installed at the end of the experimental cabin II is recorded as the combined dual-degree-of-freedom solar-directed system II; wherein, the combined dual-degree-of-freedom solar-directed system II includes: an A-axis solar-directed orientation subsystem II, a B-axis solar-directed orientation subsystem II and a truss assembly II; the A-axis solar-directed orientation subsystem II and the B-axis solar-directed orientation subsystem II are orthogonally installed on the cabin body of the experimental cabin II through the truss assembly II; the A-axis solar-directed orientation subsystem II is used to drive the solar cell wing of the experimental cabin II to rotate around the A2 axis, and the A2 axis is the cabin body axis of the experimental cabin II; the B-axis solar-directed orientation subsystem II is used to rotate the solar cell wing of the experimental cabin II around the B2 axis, and the B2 axis is the axis of the solar cell wing of the experimental cabin II.

[0083] like Fig.11As shown, the combined dual-degree-of-freedom solar orientation system includes: an A-axis solar orientation subsystem 1, a B-axis solar orientation subsystem 2 and a truss assembly 3. Among them, the A-axis solar orientation subsystem 1 and the B-axis solar orientation subsystem 2 are orthogonally installed on the spacecraft cabin 5 (experiment cabin I or experiment cabin II) through the truss assembly 3; the A-axis solar orientation subsystem 1 is used to drive the solar cell wing 4 to rotate around the A-axis to realize the axial rotation control of the solar cell wing 4A, and the A-axis is the axis of the spacecraft cabin 5; the B-axis solar orientation subsystem is used to drive the solar cell wing 4 to rotate around the B-axis to realize the axial rotation control of the solar cell wing 4B, and the B-axis is the axis of the solar cell wing 4. That is, the combined dual-degree-of-freedom solar orientation system realizes dual-degree-of-freedom solar orientation through the A-axis solar orientation subsystem 1 and the B-axis solar orientation subsystem 2, so as to ensure that the solar incidence angle of the spacecraft reaches the optimal state under different flight orbits and flight attitude conditions, and the power generation of the solar cell wings reaches the maximum capacity, thereby providing stable power output for the spacecraft; at the same time, it provides solar orientation driving torque and stall locking torque for the solar cell wings with large inertia, and has high load-bearing performance.

[0084] A-axis sun orientation subsystem

[0085] like Fig.12 As shown, the A-axis solar orientation subsystem 1 mainly includes: an A-axis solar orientation mechanism 101, an A-axis main drive controller 102, an external temperature controller 103 and an A-axis backup drive controller 104.

[0086] The A-axis main drive controller 102 and the A-axis backup drive controller 104 are used to realize the drive control of the A-axis solar orientation mechanism 101. The A-axis main drive controller 102 is installed in the cabin of the spacecraft cabin 5, and the A-axis backup drive controller 104 is installed on the truss assembly 3.

[0087] The fixed end of the A-axis solar orientation mechanism 101 is connected to the spacecraft cabin 5, and the rotating end is connected to the truss assembly 3. The A-axis solar orientation mechanism 101 is used to drive the truss assembly 3, the B-axis solar orientation subsystem 2 and the solar cell wing 4 to rotate around the A-axis as a whole under the driving control of the A-axis main drive controller 102 or the A-axis backup drive controller 104, so as to realize the A-axis rotation control of the solar cell wing 4.

[0088] The outboard temperature controller 103 is installed on the truss assembly 3 and is used to control the temperature of the A-axis solar orientation mechanism 101 so that the A-axis solar orientation mechanism 101 is within the required operating temperature range.

[0089] like Fig.16As shown, the A-axis sun orientation mechanism 101 specifically includes: three sets of drive locking mechanisms, one set of TBA (Trundle Bearing Assembly) slewing support mechanism, four sets of switching locking mechanisms and one set of rotating electrical transmission device 1019. Among them:

[0090] The TBA slewing support mechanism is used to realize the slewing support function of the fixed end and the rotating end, and is mainly used for on-track load bearing, torque transmission and rotation speed transmission. The TBA slewing support mechanism specifically includes: TBA guide rail assembly, main mode slewing support assembly, cabin end flange 10183, standby mode slewing support assembly and truss end flange 10185. Among them, the TBA guide rail assembly includes: cabin side guide rail 10181 and truss side guide rail 10186, the cabin side guide rail 10181 and the truss side guide rail 10186 are arranged in parallel, and the inner side of the cabin side guide rail 10181 and the inner side of the truss side guide rail 10186 are both large gear rings. The main mode slewing support assembly is composed of 8 main mode TBA rolling units 10182, which are evenly distributed on the cabin side rails 10181 in the circumferential direction, and play a supporting and rotating role; the 8 main mode TBA rolling units 10182 are connected to the spacecraft cabin 5 through the cabin end flange 10183. The standby mode slewing support assembly is composed of 8 standby mode TBA rolling units 10184, which are evenly distributed on the truss side rails 10186 in the circumferential direction, and play a supporting and rotating role; the 8 standby mode TBA rolling units 10184 are connected to the truss assembly 3 through the truss end flange 10185.

[0091] The driving locking mechanism is used to control the rotation or electrical locking and stopping of the A-axis sun-orienting mechanism 101. Specifically, the driving locking mechanism realizes the rotation function through the motor, the reducer and the clutch, and realizes the electrical locking and stopping function through the brake and the clutch; the clutch of the driving locking mechanism in the working state is in the meshing state, and the clutch of the driving locking mechanism in the cold backup and non-working state is in the disengaged state. Among them, the three sets of driving locking mechanisms are: driving locking mechanism a1011, driving locking mechanism b1012 and driving locking mechanism c1013; the three sets of driving locking mechanisms are in a backup relationship with each other, and at the same time, one set of driving locking mechanisms is working, and the other two sets of driving locking mechanisms are in cold backup. The driving locking mechanism a1011 is located on the cabin side, including: brake a10111, motor a10112, reducer a10113, clutch a10114 and pinion a10115 connected in sequence, and the pinion a10115 is meshed with the inner large gear ring of the cabin side guide rail 10181. The driving locking mechanism b1012 is located on the cabin side, and includes: a brake b10121, a motor b10122, a reducer b10123, a clutch b10124 and a pinion b10125 connected in sequence, and the pinion b10125 is meshed with the inner large gear ring of the cabin side guide rail 10181. The driving locking mechanism c1013 is located on the truss side, and includes: a brake c10131, a motor c10132, a reducer c10133 and a pinion c10135 connected in sequence, and the pinion c10135 is meshed with the inner large gear ring of the truss side guide rail 10186.

[0092] The four sets of switching locking mechanisms are: switching locking mechanism a1014 and switching locking mechanism b1015 located on the cabin side, and switching locking mechanism c1016 and switching locking mechanism d1017 located on the truss side; wherein the four sets of switching locking mechanisms are used in combination to realize the switching between the main mode and the standby mode. When the main mode slewing support assembly is working, the switching locking mechanism a1014 and the switching locking mechanism b1015 are unlocked, and the switching locking mechanism c1016 and the switching locking mechanism d1017 are locked; when the standby mode slewing support assembly is working, the switching locking mechanism a1014 and the switching locking mechanism b1015 are locked, and the switching locking mechanism c1016 and the switching locking mechanism d1017 are unlocked.

[0093] The rotating electric transmission device 1019 is used to realize the transmission of power / electrical signals between the fixed end and the rotating end of the A-axis solar orientation mechanism 101. One end of the rotating electric transmission device 1019 is connected to the cabin end flange 10183, and the power / electrical signal is transmitted to the electrical connector of the cabin end flange 10183 through a wire; the other end of the rotating electric transmission device 1019 is connected to the truss end flange 10185, and the power / electrical signal is transmitted to the electrical connector of the truss end flange 10185 through a wire.

[0094] The A-axis main drive controller 102 controls the main mode operation; at this time, the A-axis main drive controller 102 first controls the switching locking mechanism a1014 and the switching locking mechanism b1015 to unlock; then, controls the driving locking mechanism a1011 or the driving locking mechanism b1012 to work, so that the A-axis solar orientation mechanism 101 drives the truss assembly 3, the B-axis solar orientation subsystem 2 and the solar cell wing 4 to rotate around the A-axis as a whole, thereby realizing the axial rotation control of the solar cell wing 4A.

[0095] The A-axis backup drive controller 104 controls the backup mode to work; at this time, the A-axis backup drive controller 104 first controls the switching locking mechanism c1016 and the switching locking mechanism d1017 to unlock; then, the control drives the locking mechanism c1013 to work, so that the A-axis sun-facing orientation mechanism 101 drives the truss assembly 3, the B-axis sun-facing orientation subsystem 2 and the solar cell wing 4 to rotate around the A-axis as a whole, thereby realizing the axial rotation control of the solar cell wing 4A.

[0096] The torque and speed output functions of the A-axis sun orientation mechanism 101 are jointly realized by the driving locking mechanism and the TBA slewing support mechanism: the motor output torque and speed in the driving locking mechanism are decelerated and the torque is amplified by the reducer, and then output by the small gear in the driving locking mechanism to drive the large gear ring on the inner side of the guide rail in the TBA slewing support mechanism, which is decelerated and the torque is amplified again by the large gear ring and then output to the cabin end flange / truss end flange.

[0097] B-axis sun orientation subsystem

[0098] like Fig.12 As shown, the B-axis solar orientation subsystem mainly includes: N B-axis solar orientation mechanisms 201 and N B-axis drive controllers 202. Among them, the number of B-axis solar orientation mechanisms 201 and B-axis drive controllers 202 is consistent with the number of solar cell wings 4, that is, each solar cell wing corresponds to one B-axis solar orientation mechanism and one B-axis drive controller, N≥2.

[0099] The B-axis driving controller 202 is installed on the truss assembly 3, and is used to realize the driving control of the B-axis solar orientation mechanism 201; and realize the temperature control of the B-axis solar orientation mechanism 201, so that the B-axis solar orientation mechanism 201 is within the required working temperature range.

[0100] The fixed end of the B-axis solar orientation mechanism 201 is connected to the truss assembly 3, and the rotating end is connected to the solar cell wing 4. The B-axis solar orientation mechanism 201 is used to drive the solar cell wing 4 to rotate around the B-axis under the driving control of the B-axis driving controller 202, so as to realize the B-axis rotation control of the solar cell wing 4.

[0101] like Fig.17As shown, the B-axis drive controller 202 mainly includes: a CPU module, a motor drive module, a temperature control module and a power module. Among them, the CPU module is used to parse the rotation control instructions sent by the spacecraft GNC subsystem and generate a drive signal. The motor drive module is used to realize the drive control of the B-axis solar orientation mechanism 201 based on the drive signal generated by the CPU module, drive the solar cell wing 4 to rotate around the B axis, and realize the B-axis rotation control of the solar cell wing 4. The temperature control module is used to realize the temperature control of the B-axis solar orientation mechanism 201, so that the B-axis solar orientation mechanism 201 is within the required operating temperature range. The power module is used to supply power to the CPU module, the motor drive module and the temperature control module.

[0102] like Fig.17 As shown, the B-axis sun-orientation mechanism 201 mainly includes: a stepper motor, a harmonic reducer, a zero position sensor, a drive shaft and a power / signal transmission device. In operation, the stepper motor drives the harmonic reducer to rotate, the harmonic reducer drives the drive shaft to rotate, and the drive shaft drives the solar cell wing 4 to rotate around the B axis to realize the B-axis rotation control of the solar cell wing 4; the zero position sensor is used to measure the rotation position of the drive shaft, and the measured rotation position of the drive shaft is fed back to the B-axis drive controller 202; the power / signal transmission device rotates with the solar cell wing 4, and is used to realize the power and signal transmission of the solar cell wing 4. It can be seen that the driving function of the B-axis sun-orientation mechanism 201 is mainly realized by the stepper motor and the harmonic reducer: the B-axis sun-orientation mechanism 201 operates continuously for a long time, and the torque output is provided by the stepper motor; after the stepper motor outputs the torque, the harmonic reducer is used to realize the deceleration and torque amplification, and it also has the function of improving the speed stability. The harmonic reducer is mainly affected by the inertia load torque of the solar cell wing and the internal friction torque of the B-axis solar orientation mechanism, and can withstand the impact load when the solar cell wing is unfolded, changed and docked; the torque output by the stepper motor is amplified by the harmonic reducer and transmitted and output through the drive shaft. The drive shaft base material is made of stainless steel, and the B-axis solar orientation mechanism adopts molybdenum disulfide solid lubrication. It has a wide operating temperature range, high load capacity, stable performance, long-term reliable operation, and can meet the working life requirements of more than 108 revolutions.

[0103] Control Mode

[0104] There are five control modes for the rotation of the solar cell wing 4A and B axes, including tracking mode, capture mode, zeroing mode, angle servo mode and stop mode. Among them:

[0105] The tracking mode includes: a normal tracking mode and an offset angle tracking mode. The normal tracking mode is used for full power generation of the solar cell wing, and the offset angle tracking mode is used for reducing the power generation of the solar cell wing. In the normal tracking mode, the solar incident angle is ensured to be 0°, one axis of the solar cell wing 4 is kept at 0°, and the other axis rotates at the flight orbit speed of the spacecraft; in the offset angle tracking mode, the solar incident angle is a fixed angle, which is any angle greater than 0° and not ±90°, one axis of the solar cell wing 4 is kept at the above fixed angle, and the other axis rotates at the flight orbit speed of the spacecraft.

[0106] In the capture mode, the solar panel quickly rotates from the positive or negative direction to a fixed angle of incidence of the sun, with a rotation speed of 0.3° / s.

[0107] In zeroing mode, the solar panel is controlled to quickly rotate from the forward or reverse direction to the required zero position, with a rotation rate of 0.3° / s; the required zero position includes: horizontal zero position and vertical zero position;

[0108] In the angle servo mode, the solar orientation mechanism rotates to the required specific position (determined by the position of the sun and the flight mission requirements, any position from 0° to 360°) and remains in a stopped state.

[0109] In the stop mode, the sun-directing mechanism stops at any position.

[0110] Truss components

[0111] like Fig.12 As shown, the truss assembly 3 mainly includes: a truss structure 301 and a truss cable.

[0112] The truss structure 301 is used to provide a fixed connection for the A-axis solar orientation mechanism 101 and the B-axis solar orientation mechanism 201, and to provide an installation platform for the extravehicular temperature controller 103, the A-axis backup drive controller 104 and the B-axis drive controller 202.

[0113] The truss cable is laid and fixed on the truss structure 301, and is used to provide electrical connection between the A-axis solar orientation mechanism 101 and the A-axis main drive controller 102, the extravehicular temperature controller 103 and the A-axis backup drive controller 104; and to provide electrical connection between the B-axis solar orientation mechanism 201 and the B-axis drive controller 202.

[0114] The truss structure 301 is assembled with rods and metal corner joints, and is covered with multiple layers of thermal control to insulate it from the external heat flow of the space environment.

[0115] like Figures 13-15As shown, the truss structure 301 includes: a large column section 3011 and a small column section 3012. Among them, the fixed end of the B-axis sun orientation mechanism 201 is connected to the large column section 3011. A number of instrument mounting plates are arranged on the small column section 3012, and the external temperature controller 103, the A-axis backup drive controller 104 and the B-axis drive controller 202 are respectively mounted on the corresponding instrument mounting plates; the instrument mounting plates are made of aluminum honeycomb panels.

[0116] It can be seen that the two axes of the combined dual-degree-of-freedom solar system realize orbital period solar tracking and solar altitude compensation respectively. One axis solar orientation control is used for orbital period solar tracking, and adopts continuous rotation control. The rotation speed of this axis is the orbital flight angular velocity of the spacecraft, that is, the time for the solar cell wing to rotate one circle is equal to the orbital period of the spacecraft; the other axis solar orientation control is used to compensate for the solar altitude, and adopts intermittent rotation control. For the low-orbit solar altitude, there is a change of about 5° every day. When the change of the solar altitude makes the solar cell wing's solar incidence angle accumulate to 14°, the axial solar orientation subsystem is controlled to rotate 20°. The axial solar incidence angle changes between -6° and 14°, so that the axial solar incidence angle is kept between 0 and 14°.

[0117] Step 2: Select a solar orientation control strategy that matches the current configuration, flight attitude and operating conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control.

[0118] Experimental cabin Ⅰ / Ⅱ single cabin flight

[0119] When the experimental module I / II is flying alone, it has two flight attitudes: inertial flight and three-axis ground flight. The inertial flight of the experimental module I / II is used for long-term flight missions with stable flight in orbit. The solar orientation is achieved by the entire module during the orbital period, and the rotation of the solar cell wings only needs to compensate for the change in the solar altitude angle. The three-axis ground flight of the experimental module I / II is used for short-term flight missions of rendezvous and docking.

[0120] Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including:

[0121] like Figure 2 As shown, when the current flight attitude of the combined spacecraft is the single-cabin inertial system flight of the experimental cabin Ⅰ / Ⅱ, the A1 / A2 axis of the A-axis solar orientation subsystem Ⅰ / Ⅱ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ / Ⅱ are driven to rotate intermittently around the B1 / B2 axis through the B-axis solar orientation subsystem Ⅰ / Ⅱ to compensate for the change in the solar altitude angle.

[0122] like Figure 3As shown, when the current flight attitude of the combined spacecraft is the three-axis ground flight of the single-cabin experimental module I, the A1 / A2 axis of the A-axis solar orientation subsystem I / II is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental module I / II are driven by the B-axis solar orientation subsystem I / II to continuously rotate around the B1 / B2 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft, so as to realize orbital period solar tracking.

[0123] When the current flight attitude of the combined spacecraft is single-cabin flight of the experimental cabin I / II, and the current working condition is any one of the attitude control condition, orbit control condition, docking condition and evacuation condition, the A1 / A2 axis of the A-axis solar orientation subsystem I / II is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 / B2 axis of the A-axis solar orientation subsystem I / II is controlled to be in a horizontal zero position or a vertical zero position.

[0124] Two-cabin flight

[0125] When the two-cabin combination flies, it has two flight attitudes: inertial flight and three-axis ground flight. The inertial flight of the two-cabin combination is used for long-term flight missions with stable flight in orbit. The solar orientation is achieved by the entire cabin during the orbital period, and the rotation of the solar cell wings only needs to compensate for the change in the solar altitude angle. The three-axis ground flight of the two-cabin combination is used for short-term flight missions of rendezvous and docking.

[0126] Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including:

[0127] like Figure 4 As shown, when the current flight attitude of the combined spacecraft is the inertial system flight of the two-cabin inline combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ are driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem Ⅰ to compensate for the change in the solar altitude angle.

[0128] like Figure 5 As shown, when the current flight attitude of the combined spacecraft is a three-axis ground-to-earth flight of a two-cabin inline combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wing of the experimental cabin Ⅰ is driven by the B-axis solar orientation subsystem Ⅰ to continuously rotate around the B1 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft, so as to realize orbital period solar tracking.

[0129] When the current flight attitude of the combined spacecraft is a two-cabin in-line combination flight, and the current working condition is any one of the attitude control condition and the orbit control condition, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal zero position.

[0130] Two-cabin L-shaped combination flight

[0131] The flight of the two-cabin L-shaped combination mainly includes: three-axis ground flight of the two-cabin L-shaped combination and docking preparation flight of the two-cabin L-shaped combination.

[0132] Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including:

[0133] like Figure 6 As shown, when the current flight attitude of the combined spacecraft is a three-axis ground flight of a two-cabin L-shaped combination, the solar cell wing of the experimental cabin Ⅰ is driven by the A-axis solar orientation subsystem Ⅰ to continuously rotate around the A1 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft to achieve orbital periodic solar tracking; at the same time, the solar cell wing of the experimental cabin Ⅰ is driven by the B-axis solar orientation subsystem Ⅰ to intermittently rotate around the B1 axis to compensate for the change in the solar altitude angle.

[0134] like Figure 7 As shown, when the current flight attitude of the combined spacecraft is that the two-cabin L-shaped combination is docking and preparing for flight, in order to avoid the forward docking channel, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a vertical position.

[0135] Three-cabin T-type combination flight

[0136] The flight of the three-cabin T-type combination mainly includes: three-axis ground flight of the three-cabin T-type combination, orbital flight of the three-cabin T-type combination, radial docking preparation flight of the three-cabin T-type combination and forward and backward docking preparation flight of the three-cabin T-type combination.

[0137] Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including:

[0138] like Figure 8As shown, when the current flight attitude of the combined spacecraft is three-cabin T-type combination three-axis ground flight / three-cabin T-type combination orbital flight, the solar cell wing of the experimental cabin I is driven by the A-axis solar orientation subsystem I to continuously rotate around the A1 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft, so as to achieve orbital period solar tracking; the solar cell wing of the experimental cabin II is driven by the A-axis solar orientation subsystem II to continuously rotate around the A2 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft, so as to achieve orbital period solar tracking; at the same time, the solar cell wing of the experimental cabin I is driven by the B-axis solar orientation subsystem I to intermittently rotate around the B1 axis to compensate for the change of the solar altitude angle; the solar cell wing of the experimental cabin II is driven by the B-axis solar orientation subsystem II to intermittently rotate around the B2 axis to compensate for the change of the solar altitude angle.

[0139] like Fig. 9 As shown, when the current flight attitude of the combined spacecraft is the radial docking preparation for flight of the three-cabin T-type combination, in order to avoid the radial docking channel, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position.

[0140] like Fig.10 As shown, when the current flight attitude of the combined spacecraft is that the three-cabin T-type combination is in the forward and backward docking preparation for flight, in order to avoid the forward and backward docking channel, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position.

[0141] In summary, under different flight states, the solar incident angles of the solar panels of different compartments of the spacecraft are different, resulting in uneven power generation capabilities of the solar panels of each compartment. Through the combined spacecraft dual-axis solar orientation control method described in the present invention, the solar panels of each compartment are always operated in a better solar orientation mode, the solar orientation directivity of the solar panels is enhanced, the solar incident angle reaches the optimal state, the solar panels convert solar energy to the maximum extent, the power generation capacity of the solar panels of each compartment of the spacecraft is improved, and the power output power stability of the power supply system is ensured, which plays a key role in solving different compartment configurations and different flight attitudes, improving the power generation capacity of the solar panels, and ensuring the power output capacity of the source system.

[0142] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0143] The contents not described in detail in the specification of the present invention belong to the common knowledge of the professionals in this field.

Claims

1. A method for dual-axis solar orientation control of a combined spacecraft, characterized in that: include: Determine the current configuration, flight attitude and operating conditions of the modular spacecraft; A solar orientation control strategy that matches the current configuration, flight attitude and operating conditions of the combined spacecraft is selected to control the attitude of the dual-axis solar orientation system to achieve dual-axis solar orientation control.

2. The combined spacecraft dual-axis solar orientation control method according to claim 1, characterized in that: Combined spacecraft, including: Experimental Module I, Core Module and Experimental Module II; The configuration of the combined spacecraft includes: single-cabin configuration, two-cabin straight-line configuration, two-cabin L configuration and three-cabin T configuration; among them, single-cabin configuration: experimental cabin I or experimental cabin II; two-cabin straight-line configuration: experimental cabin I is coaxially docked with the core cabin to form a two-cabin straight-line combination; two-cabin L configuration: the experimental cabin I in the two-cabin straight-line combination is transferred to form a two-cabin L-type combination; three-cabin T configuration: the experimental cabin II is docked with the two-cabin L-type combination to form a three-cabin T-type combination; The flight attitude of the combined spacecraft includes: single-cabin flight of experimental cabin I, single-cabin flight of experimental cabin II, flight of two-cabin straight-line combination, flight of two-cabin L-type combination and flight of three-cabin T-type combination; among which, single-cabin flight of experimental cabin I includes: single-cabin inertial flight of experimental cabin I and three-axis ground flight of single-cabin experimental cabin I; single-cabin flight of experimental cabin II includes: single-cabin inertial flight of experimental cabin II and three-axis ground flight of single-cabin experimental cabin II; flight of two-cabin straight-line combination includes: inertial flight of two-cabin straight-line combination and three-axis ground flight of two-cabin straight-line combination; flight of two-cabin L-type combination includes: three-axis ground flight of two-cabin L-type combination and docking preparation flight of two-cabin L-type combination; flight of three-cabin T-type combination includes: three-axis ground flight of three-cabin T-type combination, orbital flight of three-cabin T-type combination, radial docking preparation flight of three-cabin T-type combination and forward-backward docking preparation flight of three-cabin T-type combination; The operating conditions of the combined spacecraft include: attitude control conditions, orbit control conditions, docking conditions, transfer conditions and evacuation conditions.

3. The combined spacecraft dual-axis solar orientation control method according to claim 2, characterized in that: Both experimental modules are equipped with a combined dual-degree-of-freedom solar-facing system. By controlling the attitude of the combined dual-degree-of-freedom solar-facing system, the solar panels of the experimental module can be operated in the optimal solar-facing mode to enhance the solar-facing directivity of the solar panels of the experimental module and make the solar incident angle reach the optimal state. The solar panels can convert solar energy to the maximum extent, improve the power generation capacity of the solar panels, and provide stable power output for the combined spacecraft.

4. The combined spacecraft dual-axis solar orientation control method according to claim 3, characterized in that: The combined dual-degree-of-freedom solar orientation system installed at the end of the experimental cabin I is recorded as a combined dual-degree-of-freedom solar orientation system I; wherein, the combined dual-degree-of-freedom solar orientation system I includes: an A-axis solar orientation subsystem I, a B-axis solar orientation subsystem I and a truss assembly I; the A-axis solar orientation subsystem I and the B-axis solar orientation subsystem I are orthogonally installed on the cabin body of the experimental cabin I through the truss assembly I; the A-axis solar orientation subsystem I is used to drive the solar cell wing of the experimental cabin I to rotate around the A1 axis, and the A1 axis is the cabin body axis of the experimental cabin I; the B-axis solar orientation subsystem I is used to rotate the solar cell wing of the experimental cabin I around the B1 axis, and the B1 axis is the axis of the solar cell wing of the experimental cabin I; The combined dual-degree-of-freedom solar-directed system installed at the end of the experimental cabin II is recorded as the combined dual-degree-of-freedom solar-directed system II; wherein, the combined dual-degree-of-freedom solar-directed system II includes: an A-axis solar-directed orientation subsystem II, a B-axis solar-directed orientation subsystem II and a truss assembly II; the A-axis solar-directed orientation subsystem II and the B-axis solar-directed orientation subsystem II are orthogonally installed on the cabin body of the experimental cabin II through the truss assembly II; the A-axis solar-directed orientation subsystem II is used to drive the solar cell wing of the experimental cabin II to rotate around the A2 axis, and the A2 axis is the cabin body axis of the experimental cabin II; the B-axis solar-directed orientation subsystem II is used to rotate the solar cell wing of the experimental cabin II around the B2 axis, and the B2 axis is the axis of the solar cell wing of the experimental cabin II.

5. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is the single-cabin inertial system flight of the experimental cabin I, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin I are driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I; When the current flight attitude of the combined spacecraft is the three-axis ground flight of the single cabin of the experimental cabin I, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin I are driven to rotate continuously around the B1 axis within the range of 0° to 360° through the B-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; When the current flight attitude of the combined spacecraft is the single-cabin inertial system flight of the experimental cabin II, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin II are driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II; When the current flight attitude of the combined spacecraft is the three-axis ground flight of the single-cabin experimental module II, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental module II are driven by the B-axis solar orientation subsystem II to rotate continuously around the B2 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft.

6. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is the single-cabin flight of the experimental cabin I, and the current working condition is any one of the attitude control condition, orbit control condition, docking condition and evacuation condition, the A1 axis of the A-axis sun orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis sun orientation subsystem I is controlled to be in a horizontal zero position or a vertical zero position; When the current flight attitude of the combined spacecraft is the single-cabin flight of the experimental cabin II, and the current working condition is any one of the attitude control condition, orbit control condition, docking condition and evacuation condition, the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal zero position or a vertical zero position.

7. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is the inertial flight of the two-cabin combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ are driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem Ⅰ; When the current flight attitude of the combined spacecraft is a three-axis ground-to-earth flight of a two-cabin in-line combination, the A1 axis of the A-axis solar orientation subsystem Ⅰ is controlled to be in a horizontal position and mechanically locked; at the same time, the solar cell wings of the experimental cabin Ⅰ are driven by the B-axis solar orientation subsystem Ⅰ to continuously rotate around the B1 axis in the range of 0° to 360°, and the rotation speed is the current flight speed of the combined spacecraft.

8. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is a two-cabin in-line combination flight, and the current working condition is any one of the attitude control condition and the orbit control condition, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal zero position.

9. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is a three-axis ground-to-earth flight of a two-cabin L-shaped combination, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I; When the current flight attitude of the combined spacecraft is that the two-cabin L-shaped combination is docked and ready for flight, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position.

10. The combined spacecraft dual-axis solar orientation control method according to claim 4, characterized in that: Select a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, control the attitude of the dual-axis solar orientation system, and realize dual-axis solar orientation control, including: When the current flight attitude of the combined spacecraft is a three-cabin T-type combination three-axis ground flight, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; the solar cell wing of the experimental cabin II is driven to rotate continuously around the A2 axis within the range of 0° to 360° through the A-axis solar orientation subsystem II, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I, and the solar cell wing of the experimental cabin II is driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II; When the current flight attitude of the combined spacecraft is the three-cabin T-type combined orbital system flight, the solar cell wing of the experimental cabin I is driven to rotate continuously around the A1 axis within the range of 0° to 360° through the A-axis solar orientation subsystem I, and the rotation speed is the current flight speed of the combined spacecraft; the solar cell wing of the experimental cabin II is driven to rotate continuously around the A2 axis within the range of 0° to 360° through the A-axis solar orientation subsystem II, and the rotation speed is the current flight speed of the combined spacecraft; at the same time, the solar cell wing of the experimental cabin I is driven to rotate intermittently around the B1 axis through the B-axis solar orientation subsystem I, and the solar cell wing of the experimental cabin II is driven to rotate intermittently around the B2 axis through the B-axis solar orientation subsystem II; When the current flight attitude of the combined spacecraft is the radial docking preparation for flight of the three-cabin T-type combination, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a horizontal position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a horizontal position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position; When the current flight attitude of the combined spacecraft is that the three-cabin T-type combination is docked forward and backward in preparation for flight, the A1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position and mechanically locked; the A2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position and mechanically locked; at the same time, the B1 axis of the A-axis solar orientation subsystem I is controlled to be in a vertical position, and the B2 axis of the A-axis solar orientation subsystem II is controlled to be in a vertical position.

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