Multi-source intelligent power supply system for near-earth large combined spacecraft

Abstract

The invention discloses a multi-source intelligent power supply system for a near-earth large combined spacecraft. The system comprises an intelligent closed-loop control system for stabilizing the temperature difference of a large sun orientation mechanism, a multi-area distributed power distribution system suitable for being restarted, a ground system suitable for a 360-degree rotary multi-cabin aircraft structure and a combined double-degree-of-freedom sun orientation system. According to the invention, the long-term stable operation of the near-earth large combined spacecraft is realized.

Description

Technical Field

[0001] The present invention belongs to the field of spacecraft energy technology, and in particular relates to a multi-source intelligent power supply system for a near-earth large-scale combined spacecraft. Background Art

[0002] The power supply system is an important platform system to ensure the normal operation of large spacecraft. With the development of my country's aerospace technology, the reliability, safety and long life of energy supply will provide important energy guarantee for the long-term and continuous development of manned space activities and space science research, space applications, technology experiments and other activities.

[0003] The existing space power system technology cannot meet the power requirements of 90KW high-power generation and transmission and the long-life operation requirements of more than 15 years. It is impossible to repair and upgrade faulty equipment, expand and maintain the system power during on-orbit flight, and cannot adapt to the changing needs of multiple configurations of spacecraft. It does not have the ability to intelligently identify faults on-orbit and to modify and repair faulty equipment on-orbit. Summary of the invention

[0004] The technical problem solved by the present invention is: to overcome the deficiencies of the prior art, to provide a multi-source intelligent power supply system for a near-Earth large-scale combined spacecraft, and to achieve long-term stable operation of the near-Earth large-scale combined spacecraft.

[0005] The object of the present invention is achieved through the following technical solutions: a multi-source intelligent power supply system for a near-earth large-scale combined spacecraft, comprising: an intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed orientation mechanism, a multi-region distributed power distribution system suitable for restarting, a system suitable for a 360° rotating multi-cabin aircraft structure, and a combined dual-degree-of-freedom sun-directed system; wherein the intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed orientation mechanism: collects the temperature parameters corresponding to each temperature measurement channel, obtains the temperature difference of each temperature measurement channel according to the temperature parameters corresponding to each temperature measurement channel and a preset temperature control reference voltage, obtains the switching signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measurement channel according to the switching signal of the heating circuit corresponding to each temperature measurement channel; the multi-region distributed power distribution system suitable for restarting: obtains a primary power supply from a system bus, converts the primary power supply into a secondary power supply in the cabin, and supplies power to the second-level cabin equipment of the power supply and distribution control circuit in the cabin area with the secondary power supply in the cabin; obtains a primary power supply from a system bus, converts a The secondary power source is converted into an extravehicular secondary power source, and the extravehicular secondary power source is used to power the second-level extravehicular equipment of the power supply and distribution control circuit in the extravehicular area; the command voltage is obtained from the system bus, and the command voltage is provided to the second-level in-cabin equipment of the power supply and distribution control circuit in the extravehicular area; the system suitable for the 360° rotating multi-cabin aircraft structure includes: a fixed-end structural module, a structural transmission module and a rotating-end structural module; wherein the fixed-end structural module is connected to the structural transmission module; the structural transmission module is connected to the rotating-end structural module The combined dual-degree-of-freedom solar orientation system comprises: an A-axis solar orientation subsystem, a B-axis solar orientation subsystem and a truss assembly; wherein the A-axis solar orientation subsystem and the B-axis solar orientation subsystem are orthogonally mounted on the spacecraft cabin through the truss assembly; the A-axis solar orientation subsystem is used to drive the solar cell wing to rotate around the A-axis to achieve the A-axis rotation control of the solar cell wing, and the A-axis is the axis of the spacecraft cabin; the B-axis solar orientation subsystem is used to drive the solar cell wing to rotate around the B-axis to achieve the B-axis rotation control of the solar cell wing, and the B-axis is the axis of the solar cell wing.

[0006] In the above-mentioned multi-source intelligent power supply system for large near-Earth combined spacecraft, the intelligent closed-loop control system for stabilizing the temperature difference of the large-scale solar-directed orientation mechanism includes: a temperature acquisition circuit, a heating control module and a heater loop module; wherein the temperature acquisition circuit: acquires the temperature parameters corresponding to each temperature measuring channel, and transmits the temperature parameters corresponding to each temperature measuring channel to the heating control module; the heating control module: receives the temperature parameters corresponding to each temperature measuring channel, obtains the temperature difference of each temperature measuring channel according to the temperature parameters corresponding to each temperature measuring channel and the preset temperature control reference voltage, obtains the switching signal of the heating loop corresponding to each temperature measuring channel according to the temperature difference of each temperature measuring channel, and transmits the switching signal of the heating loop corresponding to each temperature measuring channel to the heater loop module; the heater loop module: receives the switching signal of the heating loop corresponding to each temperature measuring channel, and controls the autonomous heating and disconnection of the heating loop corresponding to each temperature measuring channel according to the switching signal of the heating loop corresponding to each temperature measuring channel.

[0007] In the above-mentioned near-Earth large-scale combined spacecraft multi-source intelligent power supply system, the temperature difference of each temperature measuring channel is obtained according to the temperature parameter corresponding to each temperature measuring channel and the temperature control reference voltage, including: subtracting the temperature parameter corresponding to each temperature measuring channel from the temperature control reference voltage to obtain the temperature difference corresponding to each temperature measuring channel; the switch signal of the heating circuit corresponding to each temperature measuring channel is obtained according to the temperature difference of each temperature measuring channel, including: when the temperature difference of each temperature measuring channel is greater than 8°C, the switch signal of the two-stage heating circuit corresponding to each temperature measuring channel is turned on; when the temperature difference of each temperature measuring channel is not less than 4°C and not greater than 8°C, the switch signal of the first-stage heating circuit corresponding to each temperature measuring channel is turned on, and the switch signal of the first-stage heating circuit corresponding to each temperature measuring channel is turned off; when the temperature difference of each temperature measuring channel is ≤4°C, the switch signal of the two-stage heating circuit corresponding to each temperature measuring channel is turned off.

[0008] In the above-mentioned multi-source intelligent power supply system of the large near-Earth combined spacecraft, the heating control module includes an A / D converter and a CPU; wherein the A / D converter is connected to the CPU; the A / D converter: receives the temperature parameters corresponding to each temperature measurement channel, converts the temperature parameters corresponding to each temperature measurement channel into the temperature digital parameters corresponding to each temperature measurement channel, and transmits the temperature digital parameters corresponding to each temperature measurement channel to the CPU; the CPU: obtains the temperature difference of each temperature measurement channel according to the temperature digital parameters corresponding to each temperature measurement channel and the preset temperature control reference voltage, obtains the switching signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and transmits the switching signal of the heating circuit corresponding to each temperature measurement channel to the heater circuit module.

[0009] In the multi-source intelligent power supply system for the large near-Earth combined spacecraft, the multi-region distributed power distribution system applicable to restarting comprises: a power supply and distribution control circuit for the in-cabin area, a power supply and distribution control circuit for the out-cabin area, and a sun-orientation device; wherein the power supply and distribution control circuit for the in-cabin area is connected to the power supply and distribution control circuit for the out-cabin area through the sun-orientation device; the power supply and distribution control circuit for the in-cabin area obtains a primary power supply from the system bus, converts the primary power supply into a secondary power supply for the in-cabin area, and supplies the secondary power supply for the in-cabin area to the power supply and distribution control circuit for the in-cabin area. The first-level in-cabin equipment is powered; the second-level in-cabin equipment of the power supply and distribution control circuit in the in-cabin area obtains the command voltage from the system bus; the power supply and distribution control circuit in the outer area obtains the primary power from the system bus through the sun-facing device, converts the primary power into an outer-cabin secondary power, and uses the outer-cabin secondary power to power the second-level outer-cabin equipment of the power supply and distribution control circuit in the outer area; the power supply and distribution control circuit in the outer area obtains the command voltage from the system bus through the sun-facing device, and provides the command voltage to the second-level in-cabin equipment of the power supply and distribution control circuit in the outer area.

[0010] In the multi-source intelligent power supply system for the large near-Earth combined spacecraft, the power supply and distribution control circuit of the in-cabin area includes a bus filter and a second-level in-cabin device; wherein the bus filter obtains a primary power supply from the system bus, converts the primary power supply into an in-cabin secondary power supply, and supplies power to the second-level in-cabin device with the in-cabin secondary power supply; the second-level in-cabin device obtains a command voltage from the system bus; the power supply and distribution control circuit of the extravehicular area includes a shunt regulator and a second-level extravehicular device; wherein the shunt regulator obtains a primary power supply from the system bus through the sun-directed device, and converts the primary power supply into an in-cabin secondary power supply. The primary power supply is converted into an off-board secondary power supply, and the off-board secondary power supply is used to supply power to the second-level off-board equipment of the power supply and distribution control circuit of the off-board area; the shunt regulator obtains the command voltage from the system bus through the sun-directed device, and provides the command voltage to the second-level on-board equipment of the power supply and distribution control circuit of the off-board area; the second-level on-board equipment includes a charge and discharge regulator a, a charge and discharge regulator b, a charge and discharge regulator c, a charge and discharge regulator d, an on-board drive controller and a power manager; wherein the primary power supply is 100V, and the on-board secondary power supply includes +5V, ± 12V and +28V; the in-cabin secondary power supply required by the charge and discharge regulator a is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator b is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator c is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator d is +5V and ±12V; the in-cabin secondary power supply required by the in-cabin drive controller is +5V, ±12V and +28V; the in-cabin secondary power supply required by the power manager is +5V and ±12V; The level 2 extravehicular equipment includes an integrated drive controller, an extravehicular drive controller, an extravehicular temperature controller and a Taimin controller; wherein, the primary power supply is 100V, and the extravehicular secondary power supplies include +5V, ±12V and +28V; the extravehicular secondary power supply required by the integrated drive controller is +5V, ±12V and +28V; the extravehicular secondary power supply required by the extravehicular drive controller is +5V, ±12V and +28V; the extravehicular secondary power supply required by the extravehicular temperature controller is +5V and ±12V; the extravehicular secondary power supply required by the Taimin controller is +5V.

[0011] In the above-mentioned multi-source intelligent power supply system for large near-Earth combined spacecraft, the fixed-end structural module includes a first power receiving device, a lithium-ion battery and a first structural ground bus module; wherein the structural ground bus module is connected to the first power receiving device; the lithium-ion battery is connected to the first structural ground transmission module; the first power receiving device includes an in-cabin drive controller, a bus filter, a charge and discharge regulator and a power manager; wherein the first structural ground bus module is respectively connected to the in-cabin drive controller, the bus filter, the charge and discharge regulator, the power manager and the lithium-ion battery; the structural ground transmission module includes an electric transmission component and an insulating layer; wherein the electric transmission component is arranged in the insulating layer; the first structural ground bus module is connected to the electric transmission component; the electric transmission component is connected to the rotating end structural module; the electric transmission component includes An empty ring, a first grounding ring and a second grounding ring; wherein the first structural ground confluence module is respectively connected to one end of the first grounding ring and one end of the second grounding ring; the other end of the first grounding ring and the other end of the second grounding ring are both connected to the rotating end structural ground module; the empty ring is connected to the rotating end structural ground module; the rotating end structural ground module includes a shunt regulator, a second structural ground confluence module, a power receiving device, a driving mechanism, a solar wing and an on-track installation device; wherein the driving mechanism is connected to the solar wing; the shunt regulator is connected to the empty ring; the shunt regulator is connected to the driving mechanism; the second structural ground confluence module is respectively connected to the other end of the first grounding ring and the other end of the second grounding ring; the second structural ground confluence module is respectively connected to the shunt regulator, the driving mechanism, the power receiving device and the on-track installation device.

[0012] In the above-mentioned near-earth large-scale combined spacecraft multi-source intelligent power supply system, the A-axis solar orientation subsystem includes: an A-axis solar orientation mechanism, an A-axis main drive controller, an extravehicular temperature controller and an A-axis backup drive controller; the A-axis main drive controller and the A-axis backup drive controller are used to realize the drive control of the A-axis solar orientation mechanism; wherein the A-axis main drive controller is installed in the cabin of the spacecraft cabin, and the A-axis backup drive controller is installed on the truss assembly; the fixed end of the A-axis solar orientation mechanism is connected to the spacecraft cabin, and the rotating end is connected to the truss assembly; wherein the A-axis solar orientation mechanism is used to drive the truss assembly, the B-axis solar orientation subsystem and the solar cell wing to rotate around the A-axis as a whole under the drive control of the A-axis main drive controller or the A-axis backup drive controller, so as to realize the A-axis rotation control of the solar cell wing; the extravehicular temperature controller is installed on the truss assembly, which is used to realize the temperature control of the A-axis solar orientation mechanism, so that the A-axis solar orientation mechanism is within the required working temperature range.

[0013] In the above-mentioned near-Earth large-scale combined spacecraft multi-source intelligent power supply system, the B-axis solar orientation subsystem includes: N B-axis solar orientation mechanisms and N B-axis drive controllers; the number of B-axis solar orientation mechanisms and B-axis drive controllers is consistent with the number of solar cell wings; wherein, each solar cell wing corresponds to a B-axis solar orientation mechanism and a B-axis drive controller; N≥2; the B-axis drive controller is installed on the truss assembly, and is used to realize the drive control of the B-axis solar orientation mechanism; and realize the temperature control of the B-axis solar orientation mechanism, so that the B-axis solar orientation mechanism is within the required operating temperature range; the fixed end of the B-axis solar orientation mechanism is connected to the truss assembly, and the rotating end is connected to the solar cell wing; wherein, the B-axis solar orientation mechanism is used to drive the solar cell wing to rotate around the B-axis under the drive control of the B-axis drive controller, so as to realize the B-axis rotation control of the solar cell wing.

[0014] In the above-mentioned multi-source intelligent power supply system for large near-Earth combined spacecraft, the truss assembly includes: a truss structure and a truss cable; wherein the truss structure is used to provide a fixed connection for the A-axis solar orientation mechanism and the B-axis solar orientation mechanism, and to provide an installation platform for the extravehicular temperature controller, the A-axis backup drive controller and the B-axis drive controller; the truss cable is laid and fixed on the truss structure, and is used to provide an electrical connection between the A-axis solar orientation mechanism and the A-axis main drive controller, the extravehicular temperature controller and the A-axis backup drive controller; and, to provide an electrical connection between the B-axis solar orientation mechanism and the B-axis drive controller.

[0015] Compared with the prior art, the present invention has the following beneficial effects:

[0016] (1) The present invention solves the problem of large temperature difference between the two ends of a large-scale solar orientation mechanism due to different exposure. The closed-loop control system reasonably controls the temperature difference between the two ends of the control mechanism to ensure the long-term operation of the large-scale solar orientation mechanism. The present invention can effectively identify the temperature conditions of various parts of a large-scale mechanism through temperature acquisition and effectively control the temperature difference range between the mechanisms. The present invention intelligently identifies and compares the information obtained through temperature acquisition and automatically triggers the on / off control of the heating system. When problems occur in the existing control strategy, the present invention can readjust the temperature control strategy of the system through the parameter injection interface to enhance the fault tolerance of the system.

[0017] (2) The present invention reasonably controls the working conditions of the secondary power supply and distribution system of each device in the power supply system, realizes the power-off maintenance, equipment replacement or function expansion capabilities of the secondary power supply and distribution system, and ensures the long-term stable operation of the secondary power supply and distribution system of each device; the present invention adopts a decentralized power distribution method, and sets a level 1 device to manage the power supply interface between general equipment and the system according to different layout areas; when a single device fails, the faulty device can be disconnected from the system through the level 1 device, and the system can reduce the spread of faults between devices; the present invention connects the level 1 device through the busbar, and when the busbar maintains a 100V voltage-stabilizing state, the system's power supply and distribution system can be established here to ensure that the system restarts again; the present invention is compatible with newly added expansion equipment, and enhances the fault tolerance of the system;

[0018] (3) The present invention realizes the connection between the rotating end structural module and the fixed end structural module through the structural transmission module. The rotating end structural module and the fixed end structural module are innovatively set at both ends of the 360° transmission channel as the common point for connecting the equipment in the two compartment space areas, and finally realizes that the equipment in each compartment can still realize effective structural connection when the aircraft is rotated 360°; the present invention realizes on-orbit maintainable operation through the equipment electrical connector transmission channel connection, and innovatively utilizes the performance of the electrical connector to disconnect the electrical connector of the equipment and disconnect the structural transmission path at the same time, and connect the electrical connector of the equipment and connect the structural transmission path at the same time;

[0019] (4) The present invention adopts a combined dual-degree-of-freedom solar-facing system, which can enhance the solar-facing directionality of the solar wing by driving the two axial degrees of freedom of the solar wing, so as to ensure that the solar incidence angle reaches the optimal state under different flight orbits and flight attitudes of the spacecraft, and the power generation of the solar wing reaches the maximum capacity, thereby providing stable power output for the spacecraft, improving the power generation capacity of the solar wing of the spacecraft power system under different flight orbits and flight attitudes, and ensuring the stability of the power output power of the power system; at the same time, it has extremely strong carrying capacity, and provides solar-facing directional driving torque and stall locking torque for the large-inertia solar wing in each flight phase. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0021] Figure 1 It is a structural block diagram of an intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed orientation mechanism provided by an embodiment of the present invention;

[0022] Figure 2This is a schematic diagram of the control and temperature measurement layout of a large-scale sun-directed mechanism provided in an embodiment of the present invention;

[0023] Figure 3 It is a control block diagram of a multi-area distributed power distribution system applicable to restarting provided by an embodiment of the present invention;

[0024] Figure 4 is a control block diagram of a power distribution system after adding equipment according to an embodiment of the present invention;

[0025] Figure 5 It is a structural block diagram of a system applicable to a 360° rotating multi-cabin aircraft structure provided by an embodiment of the present invention;

[0026] Figure 6 It is a structural schematic diagram of a combined dual-degree-of-freedom sun-aiming system in an embodiment of the present invention;

[0027] Figure 7 It is a block diagram of a combined dual-degree-of-freedom solar tracking system in an embodiment of the present invention;

[0028] Figure 8 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;

[0029] Fig. 9 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;

[0030] Fig.10 is a structural schematic diagram of a truss assembly in an embodiment of the present invention;

[0031] Fig.11 This is a schematic diagram of the structure of an A-axis sun orientation subsystem in an embodiment of the present invention;

[0032] Fig.12 This is a block diagram of a B-axis sun orientation subsystem in an embodiment of the present invention;

[0033] Fig.13 This is a flow chart of a combined spacecraft dual-axis sun-orientation control method according to an embodiment of the present invention;

[0034] Fig.14 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;

[0035] Fig.15 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;

[0036] Fig.16 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;

[0037] Fig.17 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;

[0038] Fig.18 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;

[0039] Fig.19 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;

[0040] Fig. 20 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;

[0041] Fig.21 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;

[0042] Fig. 22 It is a schematic diagram of the sun orientation mode of a three-cabin T-shaped assembly in an embodiment of the present invention when it is docked forward and backward in preparation for flight. DETAILED DESCRIPTION

[0043] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0044] This embodiment provides a multi-source intelligent power supply system for a large near-Earth combined spacecraft, which includes: an intelligent closed-loop control system for stabilizing the temperature difference of a large-scale solar-oriented mechanism, a multi-region distributed power distribution system suitable for restarting, a system suitable for a 360° rotating multi-cabin aircraft structure, and a combined dual-degree-of-freedom solar-oriented system; wherein,

[0045] The intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed directional mechanism: collects the temperature parameters corresponding to each temperature measurement channel, obtains the temperature difference of each temperature measurement channel according to the temperature parameters corresponding to each temperature measurement channel and a preset temperature control reference voltage, obtains the switch signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measurement channel according to the switch signal of the heating circuit corresponding to each temperature measurement channel;

[0046] The multi-area distributed power distribution system applicable to restarting: obtains primary power from the system bus, converts the primary power into a secondary power in the cabin, and supplies power to the second-level cabin equipment of the power supply and distribution control circuit in the cabin area with the secondary power in the cabin; obtains primary power from the system bus, converts the primary power into an extravehicular secondary power, and supplies power to the second-level extravehicular equipment of the power supply and distribution control circuit in the extravehicular area with the extravehicular secondary power; obtains command voltage from the system bus, and supplies the command voltage to the second-level cabin equipment of the power supply and distribution control circuit in the extravehicular area;

[0047] The system applicable to the 360° rotating multi-chamber aircraft structure comprises: a fixed end structural module, a structural transmission module and a rotating end structural module; wherein the fixed end structural module is connected to the structural transmission module; the structural transmission module is connected to the rotating end structural module;

[0048] 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; wherein the A-axis solar orientation subsystem 1 and the B-axis solar orientation subsystem 2 are orthogonally mounted on a spacecraft cabin 5 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.

[0049] Figure 1 1 is a block diagram of an intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed orientation mechanism provided by an embodiment of the present invention. Figure 1 As shown, the intelligent closed-loop control system for stabilizing the temperature difference of a large-scale solar-oriented mechanism includes: a temperature acquisition circuit, a heating control module and a heater loop module.

[0050] Temperature acquisition circuit: collects the temperature parameters corresponding to each temperature measurement channel, and transmits the temperature parameters corresponding to each temperature measurement channel to the heating control module; heating control module: receives the temperature parameters corresponding to each temperature measurement channel, obtains the temperature difference of each temperature measurement channel according to the temperature parameters corresponding to each temperature measurement channel and the preset temperature control reference voltage, obtains the switching signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and transmits the switching signal of the heating circuit corresponding to each temperature measurement channel to the heater circuit module; heater circuit module: receives the switching signal of the heating circuit corresponding to each temperature measurement channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measurement channel according to the switching signal of the heating circuit corresponding to each temperature measurement channel.

[0051] like Figure 1 As shown, the intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed mechanism also includes: a bus signal circuit and a bus processor; wherein the bus processor: when the temperature control reference voltage needs to be modified, receives external preset temperature control data, and transmits the preset temperature control data to the bus signal circuit; the bus signal circuit: receives the preset temperature control data, obtains a new temperature control reference voltage after processing the preset temperature control data, and injects the new temperature control reference voltage into the heating control module.

[0052] Obtaining the temperature difference of each temperature measurement channel according to the temperature parameter corresponding to each temperature measurement channel and the temperature control reference voltage includes: subtracting the temperature parameter corresponding to each temperature measurement channel from the temperature control reference voltage to obtain the temperature difference corresponding to each temperature measurement channel.

[0053] The switch signal of the heating circuit corresponding to each temperature measuring channel obtained according to the temperature difference of each temperature measuring channel includes: when the temperature difference of each temperature measuring channel is greater than 8°C, the switch signal of the two-stage heating circuit corresponding to each temperature measuring channel is turned on; when the temperature difference of each temperature measuring channel is not less than 4°C and not greater than 8°C, the switch signal of the first-stage heating circuit corresponding to each temperature measuring channel is turned on, and the switch signal of the first-stage heating circuit corresponding to each temperature measuring channel is turned off; when the temperature difference of each temperature measuring channel is ≤4°C, the switch signal of the two-stage heating circuit corresponding to each temperature measuring channel is turned off. It should be understood that each temperature measuring channel corresponds to two-stage heating circuits, the first-stage heating circuit and the second-stage heating circuit.

[0054] The heating control module includes an A / D converter and a CPU; wherein the A / D converter and the CPU are connected.

[0055] The A / D converter receives the temperature parameter corresponding to each temperature measurement channel, converts the temperature parameter corresponding to each temperature measurement channel into the temperature digital parameter corresponding to each temperature measurement channel, and transmits the temperature digital parameter corresponding to each temperature measurement channel to the CPU; the CPU: obtains the temperature difference of each temperature measurement channel according to the temperature digital parameter corresponding to each temperature measurement channel and the preset temperature control reference voltage, obtains the switch signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and transmits the switch signal of the heating circuit corresponding to each temperature measurement channel to the heater circuit module. Specifically, the CPU is an 8-bit single-chip microcomputer 80C32.

[0056] The heater circuit module includes: a power drive unit and a heating circuit; wherein each power drive unit receives a switch signal of the heating circuit corresponding to each temperature measurement channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measurement channel according to the switch signal of the heating circuit corresponding to each temperature measurement channel. It should be understood that each heating circuit corresponds to a power drive unit.

[0057] like Figure 1 As shown, the intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed orientation mechanism also includes: a power module; wherein the power module supplies power to the temperature acquisition circuit and the heating control module respectively.

[0058] The power module includes a surge protection circuit, an EMI filter and a DC / DC power module; wherein the surge protection circuit receives a primary power supply and transmits the primary power supply to the EMI filter; the EMI filter filters the primary power supply to obtain a filtered power supply and transmits the filtered power supply to the DC / DC power module; the DC / DC power module receives the filtered power supply and converts the filtered power supply into a preset power supply. Specifically, the primary power supply is 100V.

[0059] The temperature collection circuit uses thermistors Ti (i=1, 2, 3, indicating the first, second or third thermistor) to collect temperature parameters of the rotating end of the large sun-facing mechanism and the connecting bracket itself.

[0060] The heating control module uses the temperature parameters of the two end surfaces of the sun mechanism collected by the thermistor and the preset temperature control reference voltage Vref (i) After comparison, it is converted into a switch signal, i.e., a temperature control signal S (k) (k=1,2,3), through the temperature control signal S (k) Realize the control switch N of the power drive unit, that is, the heating circuit (n) Autonomous on and off control of the heating circuit control switch N (n) Using MOS tube, through the temperature control signal S (k) Realizes autonomous heating and disconnection control of the heater circuit.

[0061] like Figure 1 As shown, the heater circuit module also includes: a power supply control switch (K1, K2), the power supply control switch (K1, K2) is connected to the heating circuit, the power supply control switch (K1, K2) adopts a relay and remains in an on state, and can be controlled on and off by external instructions.

[0062] The temperature acquisition circuit transmits the temperature parameters corresponding to each temperature measurement channel to the bus signal circuit, the heating control module transmits the switch signal of the heating circuit corresponding to each temperature measurement channel to the bus signal circuit, and the bus signal circuit transmits the temperature parameters corresponding to each temperature measurement channel and the switch signal of the heating circuit corresponding to each temperature measurement channel to the bus processor. The bus processor judges the need to adjust the working reference of the heating control module based on the temperature parameters corresponding to each temperature measurement channel and the switch signal of the heating circuit corresponding to each temperature measurement channel. The ground sends the temperature control data to the bus processor, injects the bus signal into the bus signal circuit, and after signal processing, the new temperature control reference voltage Vref is (i) The modifications made realize the parameter adjustment of the heating control circuit and achieve the purpose of changing the system temperature control strategy.

[0063] The temperature control range of the intelligent closed-loop control system that stabilizes the temperature difference of large-scale sun-directed mechanisms is set in grades:

[0064] Heater circuit N (a) (a is the odd-numbered heating circuit) The temperature control strategy is set as follows: if △T(a)>4℃, the heating switch is turned on; if △T(a)≤4℃, the heating switch is turned off. Heater circuit N (b) (b is the even number heating circuit) The temperature control range is set to: if △T(b)>8℃, the switch is turned on; if △T(b)≤8℃, the switch is turned off. △T(a) and △T(b) are: The highest temperature in the temperature measurement parameter T(n) is taken as the temperature control reference voltage Vref (i) , the temperature difference between the other temperatures and the reference voltage is obtained by comparing the two.

[0065] When the first-level odd-number heating circuit cannot meet the temperature control requirements, the second-level even-number heating circuit is started.

[0066] In order to effectively control the large temperature difference between the two ends of the large-scale solar directional mechanism due to different exposure, the intelligent closed-loop control system of the temperature difference of the large-scale solar directional mechanism is used to set up 47 temperature collection positions at the rotating end and connecting bracket of the large-scale solar directional mechanism, of which 8 channels T1~T8 are used for temperature control. Figure 2 shown.

[0067] By collecting temperature parameters at 8 locations, the temperature parameters are transmitted to the CPU control circuit through the temperature measurement channel, and the highest temperature is taken as the temperature control reference voltage Vref (i), compare other temperatures with the reference voltage. When the temperature difference between the two is greater than 4°C, the single heating circuit control switch N is properly controlled. (a) When the temperature difference between the two is greater than 8℃, the heating circuit control switch N is properly controlled. (b) The temperature difference in each part of the mechanism is controlled within 4℃, ensuring the long-term operation of the large-scale sun-directed mechanism.

[0068] The power supply control switch K1 in the heater circuit controls the 1st to 8th heating circuits, and K2 controls the 9th to 16th heating circuits; the power supply control switch K1 remains in the on state and can be controlled on and off by the external command i1; the power supply control switch K2 remains in the on state and can be controlled on and off by the external command i2.

[0069] 1st to 16th heating circuit start control switch N (1) ~N (16) , respectively through the temperature control signal S (1) ~S (16) Realize autonomous heating and disconnection control of each heater circuit.

[0070] The temperature parameters of T1~T47 are sent to the bus signal circuit through the temperature acquisition circuit. After the signal is processed into a bus signal, it is sent to the bus processor and then transmitted to the ground, so as to realize the ground monitoring of the voltage parameters and heating status corresponding to the temperature control point.

[0071] The ground monitors the working status of the large-scale sun-directing mechanism through the temperature parameters of T1~T47. When the temperature control reference voltage Vref is needed (i) When making changes, the temperature control data is sent to the bus processor, and the bus signal is injected into the bus signal circuit. After signal processing, the new temperature control reference voltage Vref (i) The parameters of the heating control circuit can be adjusted by injecting modifications into the CPU control circuit.

[0072] In order to improve the reliability of the intelligent closed-loop control system of temperature difference of large-scale sun-directed directional mechanism, the temperature control strategy adopts temperature range classification setting.

[0073] Heater circuit N (1) 、N (3) 、N (5) 、N (7) 、N (9) 、N (11) 、N (13)、 N (15) The temperature control temperature difference strategy is set as follows: if △T1>4℃, the switch is turned on; if △T1≤4℃, the switch is turned off.

[0074] Heater circuit N (2) 、N(4) 、N (6) 、N (8) 、N (10) 、N (12) 、N (14)、 N (16) The temperature control temperature difference range is set as follows: if △T2>8℃, the switch is turned on; if △T2≤8℃, the switch is turned off.

[0075] When the first-level odd-number heating circuit cannot meet the temperature control requirements, the second-level even-number heating circuit is started.

[0076] Figure 3 1 is a control block diagram of a multi-area distributed power distribution system applicable to restarting according to an embodiment of the present invention. Figure 3 As shown, the multi-area distributed power distribution system applicable for restarting includes: a power supply and distribution control circuit in the cabin area, a power supply and distribution control circuit in the outer cabin area, and a sun-directing device.

[0077] The power supply and distribution control circuit of the in-cabin area is connected to the power supply and distribution control circuit of the out-cabin area through the sun orientation device.

[0078] The power supply and distribution control circuit of the in-cabin area obtains primary power from the system bus, converts the primary power into in-cabin secondary power, and uses the in-cabin secondary power to power the level 2 in-cabin equipment of the power supply and distribution control circuit of the in-cabin area; the level 2 in-cabin equipment of the power supply and distribution control circuit of the in-cabin area obtains the command voltage from the system bus.

[0079] The power supply and distribution control circuit of the extravehicular area obtains primary power from the system bus through the sun-facing device, converts the primary power into extravehicular secondary power, and supplies power to the second-level extravehicular equipment of the power supply and distribution control circuit of the extravehicular area with the extravehicular secondary power; the power supply and distribution control circuit of the extravehicular area obtains command voltage from the system bus through the sun-facing device, and supplies the command voltage to the second-level intravehicular equipment of the power supply and distribution control circuit of the extravehicular area.

[0080] The power supply and distribution control circuit of the cabin area includes a bus filter and a second-level cabin device; wherein the bus filter obtains primary power from the system bus, converts the primary power into a cabin secondary power, and supplies power to the second-level cabin device from the cabin secondary power; the second-level cabin device obtains the command voltage from the system bus. Specifically, the primary power is 100V, and the cabin secondary power includes +5V, ±12V and +28V.

[0081] Specifically, the level 2 in-cabin equipment includes a charge and discharge regulator a, a charge and discharge regulator b, a charge and discharge regulator c, a charge and discharge regulator d, an in-cabin drive controller and a power manager.

[0082] The in-cabin secondary power supply required by the charge and discharge regulator a is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator b is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator c is +5V and ±12V; the in-cabin secondary power supply required by the charge and discharge regulator d is +5V and ±12V; the in-cabin secondary power supply required by the in-cabin drive controller is +5V, ±12V and +28V; the in-cabin secondary power supply required by the power manager is +5V and ±12V.

[0083] The power supply and distribution control circuit of the extravehicular area includes a shunt regulator and a second-level extravehicular device; wherein the shunt regulator obtains a primary power supply from the system bus through the sun-directed device, converts the primary power supply into an extravehicular secondary power supply, and supplies the extravehicular secondary power supply to the second-level extravehicular device of the power supply and distribution control circuit of the extravehicular area; the shunt regulator obtains a command voltage from the system bus through the sun-directed device, and supplies the command voltage to the second-level in-cabin device of the power supply and distribution control circuit of the extravehicular area. The primary power supply is 100V, and the extravehicular secondary power supply includes +5V, ±12V and +28V.

[0084] Level 2 extravehicular equipment includes an integrated drive controller, an extravehicular drive controller, an extravehicular temperature controller, and a Taimin controller. The extravehicular secondary power supply required by the integrated drive controller is +5V, ±12V, and +28V; the extravehicular secondary power supply required by the extravehicular drive controller is +5V, ±12V, and +28V; the extravehicular secondary power supply required by the extravehicular temperature controller is +5V and ±12V; the extravehicular secondary power supply required by the Taimin controller is +5V.

[0085] like Figure 4 As shown, the power supply and distribution control circuit of the extravehicular area also includes: an on-orbit expansion controller; wherein the shunt regulator obtains primary power from the system bus through the solar orientation device, converts the primary power into an extravehicular secondary power, and provides the extravehicular secondary power to the on-orbit expansion controller; the shunt regulator obtains a command voltage from the system bus through the solar orientation device, and provides the command voltage to the on-orbit expansion controller.

[0086] The system is divided into level 1 equipment and level 2 equipment. According to different layout areas, one device is set as level 1 equipment in one area. The level 1 equipment is responsible for the power supply interface between all level 2 equipment in the area and the system bus.

[0087] Level 1 equipment is the power supply and distribution management equipment within the system. It takes power from the system bus inside the product and directly provides secondary power supply to the product through the conversion circuit. The command bus within the system is uniformly provided by the level 1 equipment.

[0088] Level 2 equipment is other general working equipment in the system. According to the needs inside the product, it takes power from the system bus and uses conversion circuits such as "100V / +5V", "100V / ±12V" and "100V / +28V" to provide secondary power supply inside the product.

[0089] The system can meet the task of restarting after the system bus is powered off. The level 1 equipment is set to directly draw power from the power supply bus, and a 100V regulated input is provided through the regulated power supply interface to ensure the power supply bus voltage. A 28V command bus is formed through the conversion circuit in the level 1 equipment, and then uniformly distributed to each device in the system to ensure that each device can start the device power supply path through the 28V command bus and restore the system bus power supply. After the system bus power supply is restored, the 100V regulated input provided by the regulated power supply interface can be disconnected.

[0090] All new devices added to the system are Level 2 devices. The system reserves the power supply interface and 28V command interface of the system bus for the subsequent new devices. After the new devices are installed, they are connected to the system through cables and the reserved interfaces of the system. Finally, the power supply switch is controlled to close inside the Level 1 device, so that the new devices can be connected to the system bus for power supply.

[0091] (1) Power supply and distribution control circuit

[0092] The power supply and distribution control circuit of the system consists of two relatively independent and dispersed parts: the in-cabin area and the out-cabin area.

[0093] The power supply and distribution control circuit in the cabin area consists of a bus filter, charge and discharge regulators a~d, a cabin drive controller and a power manager.

[0094] The bus filter is a Class 1 device that draws power from the 100V power supply bus. The product uses a secondary power distribution method through a "100V / +28V" conversion circuit inside the product to provide power to the command relay coils of all equipment in the power supply system, with an output capacity of no less than 1A; and provides the remaining Class 2 products with a 100V power supply bus access through a secondary power distribution method.

[0095] The bus filter is also equipped with a 100V voltage-stabilized power supply interface for receiving 100V input from an external circuit. The voltage-stabilized power supply interface is equipped with an isolation relay. The voltage-stabilized power supply relay wire package in the single machine is powered by the 28V command bus of the external circuit.

[0096] The charge and discharge regulators a~d, the cabin drive controller and the power manager are Level 2 devices. They take power from the primary bus as required inside the product and use secondary power distribution through conversion circuits to provide their own working needs.

[0097] The power supply and distribution control circuit of the extravehicular area is composed of a shunt regulator, an integrated drive controller, an extravehicular drive controller, an extravehicular temperature controller and a Taimin controller.

[0098] The shunt regulator is a Class 1 device that draws power from the 100V power supply bus. The product uses a secondary power distribution method through the "100V / +5V" and "100V / ±12V" conversion circuits to meet its own working needs; and the 100V power supply bus access for the remaining Class 2 products is provided through the secondary power distribution method.

[0099] Taking into account the need for command bus management of all equipment in the power system, and optimizing the number of slip rings between the extravehicular equipment and the solar orientation device between the extravehicular areas, the 28V command bus of all extravehicular equipment is transferred inside the shunt regulator.

[0100] The integrated drive controller, extravehicular drive controller, extravehicular temperature controller and Taimin controller are Level 2 devices. According to the needs inside the product, they take power from the primary bus and use secondary power distribution through the conversion circuit to provide their own working needs.

[0101] (2) Voltage stabilizing starting circuit

[0102] The system can meet the task of restarting after the system bus is powered off. After the 100V power supply bus is powered off, the system will disconnect all power switches to ensure safety, and the bus voltage will be 0V.

[0103] When the system needs to be restarted, the external circuit inputs 100V through the bus filter voltage stabilization power supply interface.

[0104] When the external 28V command bus controls the voltage-stabilized power supply relay package to close the power supply switch, the voltage-stabilized power supply interface inputs 100V to maintain the system power supply bus voltage boosted to 100V.

[0105] The bus filter forms a 28V command bus through the "100V / 28V" conversion circuit, which is uniformly distributed to the charge and discharge regulators a~d, the cabin drive controller and the power manager cabin level 2 equipment, and at the same time transmitted to the shunt regulator of the level 1 equipment in the outer cabin area through the sun-directed device, and then transferred by the shunt regulator to the integrated drive controller, the outer cabin drive controller, the outer cabin temperature controller and the Taimin controller outer cabin level 2 equipment. Ensure that each device can start the device power supply path through the 28V command bus and restore the system bus power supply.

[0106] After the system bus power supply is restored, the external 100V regulated voltage input provided by the regulated power supply interface can be disconnected.

[0107] (3) New equipment

[0108] The on-orbit expansion controller is a new device added to the system, so it is a level 2 device. The shunt regulator in the extravehicular area reserves the power supply interface and command bus interface for the subsequent new equipment.

[0109] After installation, the on-orbit expansion controller is connected to the power supply interface and command bus interface reserved by the shunt regulator through a cable. The shunt regulator controls the reserved power supply path switch to close, so that the on-orbit expansion controller is connected to the system bus for power supply.

[0110] Figure 5 1 is a structural block diagram of a system applicable to a 360° rotating multi-cabin aircraft structure provided by an embodiment of the present invention. Figure 5 As shown, the system suitable for a 360° rotating multi-chamber aircraft structure includes: a fixed-end structural module, a structural transmission module and a rotating-end structural module; wherein the fixed-end structural module is connected to the structural transmission module; and the structural transmission module is connected to the rotating-end structural module.

[0111] The fixed-end structural module includes a first power receiving device, a lithium-ion storage battery and a first structural ground confluence module; wherein the structural ground confluence module is connected to the first power receiving device; and the lithium-ion storage battery is connected to the first structural ground transmission module.

[0112] The first power receiving device includes an in-cabin drive controller, a bus filter, a charge and discharge regulator and a power manager; wherein the first structure ground bus module is respectively connected to the in-cabin drive controller, the bus filter, the charge and discharge regulator, the power manager and the lithium-ion battery.

[0113] The structural ground transmission module includes an electric transmission component and an insulating layer; wherein the electric transmission component is arranged in the insulating layer; the first structural ground confluence module is connected to the electric transmission component; and the electric transmission component is connected to the rotating end structural ground module.

[0114] The electrical transmission component includes an empty ring, a first grounding ring and a second grounding ring; wherein the first structural ground confluence module is respectively connected to one end of the first grounding ring and one end of the second grounding ring; the other end of the first grounding ring and the other end of the second grounding ring are both connected to the rotating end structural ground module; the empty ring is connected to the rotating end structural ground module.

[0115] The rotating end structure ground module includes a shunt regulator, a second structure ground confluence module, a power receiving device, a driving mechanism, a solar wing and an on-track installation device; wherein the driving mechanism is connected to the solar wing; the shunt regulator is connected to the empty ring; the shunt regulator is connected to the driving mechanism; the second structure ground confluence module is respectively connected to the other end of the first grounding ring and the other end of the second grounding ring; the second structure ground confluence module is respectively connected to the shunt regulator, the driving mechanism, the power receiving device and the on-track installation device.

[0116] The powered equipment includes an integrated drive controller, a thermistor converter, an outboard drive controller and an outboard temperature controller; wherein the integrated drive controller, the thermistor converter, the outboard drive controller and the outboard temperature controller are all connected to the second structural ground confluence module.

[0117] The on-track installation device includes an extended shunt regulator and an extension and retraction controller; wherein the extended shunt regulator and the extension and retraction controller are both connected to the second structure ground confluence module.

[0118] The structural ground transmission module also includes an outboard flange drive assembly switching assembly; wherein the second structural ground confluence module is connected to the outboard flange drive assembly switching assembly.

[0119] The structural ground transmission module also includes a switching component of the driving component of the main bearing frame of the flange in the cabin; wherein the second structural ground confluence module is connected to the switching component of the driving component of the main bearing frame of the flange in the cabin.

[0120] The fixed end structural ground module is composed of overlapping electronic equipment, high resistance grounding equipment and the first structural ground confluence module;

[0121] Inside the system, powered devices such as "cabin drive controller", "bus filter", "charge and discharge regulator" and "power manager" are used as connecting electronic devices for equipment connection design.

[0122] "Lithium-ion battery" is a power supply device. The grounding terminal of the lithium-ion battery shell is designed with high-resistance grounding to prevent the power supply from leaking and still ensure the safety of the system through the high-resistance grounding path.

[0123] The "structural ground confluence module" is a fixed-end electronic device. The structural ground confluence module provides the overall confluence function of the structure, realizes a unified structural ground at the fixed end, and is finally connected to the structural ground confluence module at the rotating end through the transmission module.

[0124] The structural ground transmission module consists of a transmission slip ring and a transmission path. The structural ground transmission module is divided into a rotating end transmission path and a fixed end transmission path through the transmission slip ring. The transmission slip ring can realize the effective connection of the transmission paths when the transmission paths rotate 360°. The transmission slip ring adopts a high-resistance grounding design as a whole to prevent the power supply from leaking and still ensure the safety of the system through the high-resistance grounding path.

[0125] The rotating end structure ground module is composed of bonding electronic equipment, high-resistance grounding equipment and the second structure ground confluence module. The "integrated drive controller", "too sensitive converter", "outboard drive controller" and "outboard temperature controller" are used as power receiving equipment in the system, and are used as bonding electronic equipment for equipment bonding design.

[0126] The extended shunt regulator and the expansion and retraction controller are on-orbit installation equipment. As overlapping electronic equipment, they adopt the equipment overlapping design. In order to realize the maintainable operation on-orbit, the equipment structure is connected through the equipment electrical connector transmission channel, which is different from the direct equipment casing overlapping method at the fixed end.

[0127] The solar wing is a power supply device that requires a high-resistance grounding design, and the "shunt regulator" is a power transmission device that also uses a high-resistance grounding design. Both prevent the power supply from having a leakage fault and ensure the safety of the system through the high-resistance grounding path.

[0128] The second structural confluence module provides an overall structural confluence function for the rotating end electronic equipment, realizes a unified structural confluence at the rotating end, and is finally connected to the structural confluence module at the fixed end through the transmission module.

[0129] Aircraft power receiving equipment: Since the product draws power from the busbar, when the equipment power supply fails, the equipment fault can be isolated by disconnecting the power switch. Therefore, a direct bonding installation method is adopted in the grounding system.

[0130] Aircraft power supply equipment: According to product characteristics, power supply equipment is the power supply of aircraft. When the equipment fails, the equipment cannot be disconnected from the power supply path. Therefore, such equipment is installed in an insulated manner and grounded with high resistance in the grounding system. When the equipment fails, a small current can be released through the high resistance grounding path to protect the power supply system from working normally. Such as batteries and solar wings.

[0131] Aircraft on-orbit installation equipment: According to the characteristics of on-orbit installation of products, it is necessary to consider the establishment of a static electricity accumulation discharge circuit after the equipment is installed. Therefore, such equipment is directly overlapped and installed in the grounding system, and a structural ground is set inside the equipment and then connected to the cabin structure ground through wires. Such as extended shunt regulators and extended expansion and retraction controllers.

[0132] Aircraft power transmission equipment: According to the product characteristics, it needs to meet the requirements of 360° rotation and power transmission. The power transmission equipment is a single point link in the power supply path. When the equipment fails, the equipment cannot be disconnected from the power supply path. Therefore, such equipment is installed in an insulated manner and grounded with high resistance in the grounding system. When the equipment fails, a small current can be released through the high-resistance grounding path to protect the power supply system from working normally. Such as drive mechanism and solar orientation device.

[0133] This embodiment constructs a comprehensive safety structural ground system through technical measures such as overlapping, high-resistance grounding, and slip ring grounding.

[0134] The rotating multiple compartments of the aircraft are the working compartment, truss compartment and solar wing;

[0135] The rotating mechanism of the aircraft cabin wall includes a solar orientation device and a driving mechanism, wherein the solar orientation device connects the working cabin and the truss cabin, and the driving mechanism connects the truss cabin and the solar wing.

[0136] The equipment in the aircraft working compartment includes drive controller, bus filter, power manager, charge and discharge regulator, and battery;

[0137] The aircraft truss cabin equipment includes shunt regulator, integrated controller, deployment controller, outboard drive controller, outboard temperature controller, extended shunt regulator, and extended deployment controller;

[0138] The aircraft working cabin equipment drive controller, bus filter, power manager, and charge and discharge regulator are all powered devices, and are directly connected to the cabin structure for installation and overlap.

[0139] The battery of the aircraft working cabin equipment is the power supply equipment. It is insulated from the cabin and installed with high resistance connection measures. A 50KΩ high resistance is set inside the battery to connect to the grounding pile inside the equipment, and then led out to the working cabin structure ground through the connector.

[0140] The sun-directing device is directly overlapped with the installation surfaces of the truss cabin and the working cabin.

[0141] Grounding piles are set for the cabin flange, main load-bearing frame, and driving locking assembly and switching locking assembly installed on the cabin flange of the sun orientation device, and then the wires are led out to connect with the structural ground in the working cabin.

[0142] The outer flange of the solar orientation device and the driving locking assembly and the switching locking assembly installed on the flange are connected by grounding piles, and then the wires are led out to communicate with the truss cabin structure.

[0143] The electric transmission component of the solar orientation device is the main power transmission equipment, and the shell is insulated from the main structure of the solar orientation device. A grounding stake is set on the shell of the electric transmission component, and an 80KΩ resistor is connected in series, and then connected to the grounding stake of the flange outside the solar orientation device cabin for collection.

[0144] The aircraft truss cabin is equipped with a structural ground in quadrants II and IV for grounding of each product. The structural grounds in quadrants II and IV are connected through truss cables, and then connected to the two grounding slip rings of the solar orientation device by the truss structural ground, and then connected to the cabin structure ground in the working cabin through the slip rings. The connection between the truss structural ground and the cabin structure ground in the working cabin is realized.

[0145] The thermal control multi-layer structure coated outside the aircraft truss cabin is connected to the structures of the truss quadrants II and IV through cables.

[0146] The aircraft truss cabin equipment shunt regulator, integrated controller, extension and retraction controller, extravehicular drive controller, and extravehicular temperature controller are all powered devices. The internal structure of the equipment is used to lead out two structural ground contacts through connectors, and they are respectively connected to the structural ground set in quadrants II and IV of the truss cabin through truss cables.

[0147] The driving mechanism is a power transmission device, which is insulated from the truss cabin and directly overlapped with the solar wing mounting surface.

[0148] The solar wing is a power supply device, with a high resistance of 80KΩ connected in series on its shell structure, and connected to the two grounding rings in the drive mechanism through connectors. The grounding ring of the drive mechanism is connected to the structural ground of the truss cabin through the truss cable.

[0149] The empty ring of the driving mechanism and the empty ring of the sun-directing device are connected to the 100KΩ high resistance set inside the shunt regulator through a connector, the high resistance is connected to the shunt regulator structure, and then led out to the truss structure through the connector.

[0150] The extended shunt regulator and extended expansion and retraction controller are products to be installed on-orbit later. The equipment is a power-receiving device, so a grounding pile is set inside the equipment, and two structural ground contacts are led out through connectors. After the installation on-orbit and the cabin is completed, the equipment structural ground is connected to the truss cabin structural ground through cables.

[0151] The extended solar wing is a power supply device, with a high resistance of 80KΩ connected in series on its shell. After being installed on the track, it is connected to the structural ground of the extended shunt regulator through a cable. The structural ground of the extended shunt is connected to the structural ground of the equipment with the structural ground of the truss cabin through a connector to lead out a cable.

[0152] This embodiment realizes the connection between the rotating end structural ground module and the fixed end structural ground module through the structural ground transmission module. The rotating end structural ground and the fixed end structural ground are innovatively set at both ends of the 360° transmission channel as the common point of the connection ground of the equipment in the two compartment space areas, and finally realizes that the equipment in each compartment can still realize effective structural connection when the aircraft is rotated 360°; this embodiment realizes on-orbit maintainable operation through the equipment electrical connector transmission channel connection, and innovatively utilizes the performance of the electrical connector to realize that the structural ground transmission path is disconnected when the equipment is disconnected from the electrical connector, and the structural ground transmission path of the equipment is connected when the electrical connector is plugged in.

[0153] Reference Figure 6In this embodiment, 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. The A-axis solar orientation subsystem 1 and the B-axis solar orientation subsystem 2 are orthogonally mounted on the spacecraft cabin 5 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 achieve 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 achieve 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.

[0154] A-axis sun orientation subsystem

[0155] In this embodiment, if Figure 7 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.

[0156] 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.

[0157] 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.

[0158] 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.

[0159] like Fig.11As 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:

[0160] 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.

[0161] 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.

[0162] 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.

[0163] 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.

[0164] 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.

[0165] 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.

[0166] 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.

[0167] B-axis sun orientation subsystem

[0168] In this embodiment, if Figure 7 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.

[0169] 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.

[0170] 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.

[0171] like Fig.12As 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.

[0172] like Fig.12 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.

[0173] Control Mode

[0174] In this embodiment, the solar cell wing 4A axis and B axis rotation have five control modes, including: tracking mode, capture mode, zeroing mode, angle servo mode and stop mode. Among them:

[0175] 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.

[0176] 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.

[0177] In zeroing mode, the solar panel is controlled to quickly rotate from the forward or reverse direction to the desired zero position, with a rotation rate of 0.3° / s; the desired zero position includes: horizontal zero position and vertical zero position;

[0178] 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.

[0179] In the stop mode, the sun-directing mechanism stops at any position.

[0180] Truss components

[0181] In this embodiment, if Figure 7 As shown, the truss assembly 3 mainly includes: a truss structure 301 and a truss cable.

[0182] 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.

[0183] 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.

[0184] 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.

[0185] like Figure 8 , Fig. 9 and Fig.10As 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.

[0186] In this embodiment, the combined dual-degree-of-freedom solar orientation system is usually installed at the tail of the experimental cabin of the spacecraft; the experimental cabin includes: a cabin body and two sets of solar cell wings. Specifically: the fixed end of the A-axis solar orientation subsystem is connected to the cabin body of the experimental cabin, and the rotating end is connected to the truss structure. As mentioned above, the number of B-axis solar orientation mechanisms and B-axis drive controllers in the B-axis solar orientation subsystem is consistent with the number of solar cell wings: the fixed end of one B-axis solar orientation mechanism is connected to the truss structure, and the rotating end is connected to the root of one set of solar cell wings; the fixed end of the other B-axis solar orientation mechanism is connected to the truss structure, and the rotating end is connected to the root of the other set of solar cell wings; the two B-axis drive controllers are respectively installed on the instrument mounting plates on both sides of the truss structure.

[0187] In this embodiment, a specific application scenario is taken as an example for description.

[0188] The combined dual-degree-of-freedom solar orientation system described in the embodiment of the present invention can be applied to the combined spacecraft dual-axis solar orientation control. Specifically:

[0189] The modular spacecraft mainly includes: Experimental Module I, Core Module and Experimental Module II. Since the modular spacecraft is launched as a single module, assembled and operated on orbit, and has multiple and complex configurations on orbit, the working mode of the power 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 based on the core module for docking and assembly.

[0190] Both the experimental cabin I and the experimental cabin II 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 wings of the experimental cabin are operated in the optimal solar orientation mode to enhance the solar orientation directivity of the solar cell wings of the experimental cabin, so that the solar incident angle reaches the optimal state, the solar cell wings convert solar energy to the maximum extent, and the power generation capacity of the solar cell wings is improved, providing stable power output for the combined spacecraft. In addition, the core cabin is equipped with a single-degree-of-freedom solar orientation system.

[0191] 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 cabin 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 cabin body axis 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.

[0192] 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 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 axis of the cabin body 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.

[0193] Based on the above, Fig.13 As shown in the figure, the dual-axis solar orientation control process of the combined spacecraft is as follows:

[0194] S1, determine the current configuration, flight attitude and working conditions of the combined spacecraft.

[0195] Multiple configurations

[0196] The configurations of modular spacecraft mainly include: single-cabin configuration, two-cabin straight-line configuration, two-cabin L configuration and three-cabin T configuration.

[0197] Single cabin configuration: experimental cabin I or experimental cabin II.

[0198] Two-cabin in-line configuration: Experimental cabin I is coaxially docked with the core cabin to form a two-cabin in-line combination.

[0199] 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.

[0200] Three-cabin T-configuration: Experimental cabin II is docked with the two-cabin L-shaped combination to form a three-cabin T-shaped combination.

[0201] Multiple flight attitudes

[0202] 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.

[0203] 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.

[0204] 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.

[0205] 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.

[0206] 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.

[0207] 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.

[0208] Multiple working conditions

[0209] 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:

[0210] Attitude control conditions: Factors affecting the load conditions of attitude control conditions include: cabin configuration, cabin attitude adjustment direction, GNC attitude adjustment mode, etc.

[0211] 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.

[0212] 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.

[0213] 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.

[0214] 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.

[0215] S2, selects a solar orientation control strategy that matches the current configuration, flight attitude and working conditions of the combined spacecraft, controls the attitude of the dual-axis solar orientation system, and realizes dual-axis solar orientation control.

[0216] Experimental cabin Ⅰ / Ⅱ single cabin flight

[0217] 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.

[0218] 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:

[0219] like Fig.14 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.

[0220] like Fig.15 As 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.

[0221] 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.

[0222] Two-cabin flight

[0223] 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.

[0224] 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:

[0225] like Fig.16 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.

[0226] like Fig.17 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.

[0227] 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.

[0228] Two-cabin L-shaped combination flight

[0229] 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.

[0230] 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:

[0231] like Fig.18As 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.

[0232] like Fig.19 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.

[0233] Three-cabin T-type combination flight

[0234] 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.

[0235] 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:

[0236] like Fig. 20 As 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.

[0237] like Fig.21As 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.

[0238] like Fig. 22 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.

[0239] This embodiment uses a combined dual-degree-of-freedom solar orientation system, which has the advantages of high pointing accuracy and strong carrying capacity, and is applied to the assembly and operation of large aircraft, effectively improving the power generation of solar cell wings. During the assembly and operation of large aircraft, under different flight phases and flight attitudes, under the orbital conditions of the solar altitude angle of ±66°, it can ensure that the incident angle of the solar cell wing is between 0° and 14°, so that the power generation of the solar cell wing is in a relatively stable maximum output state. Compared with the single-degree-of-freedom solar orientation system, the average power generation of the solar cell wing is increased by 15%, and the extreme power generation is increased by 140% (when the solar altitude angle is 66°).

[0240] Under different flight states, the solar incident angles of the solar panels in different sections of the spacecraft are different, resulting in uneven power generation capabilities of the solar panels in each section. Through the combined spacecraft dual-axis solar orientation control method of this embodiment, the solar panels in each section always work in a better solar orientation mode, which enhances the solar orientation directivity of the solar panels, makes the solar incident angle reach the optimal state, and maximizes the conversion of solar energy by the solar panels. The power generation capacity of the solar panels in each section of the spacecraft is improved, and the power output power stability of the power supply system is guaranteed. It plays a key role in solving the problem of improving the power generation capacity of the solar panels in different section configurations and different flight attitudes and ensuring the power output capacity of the source system.

[0241] This embodiment reasonably controls the working conditions of the secondary power supply and distribution system of each device in the power supply system, realizes the ability of the secondary power supply and distribution system to be powered off for maintenance, equipment replacement or function expansion, and ensures the long-term stable operation of the secondary power supply and distribution system of each device; this embodiment adopts a decentralized power distribution method, and sets up level 1 equipment to manage the power supply interface between general equipment and the system according to different layout areas; when a single device fails, the faulty device can be disconnected from the system through the level 1 device, and the system can reduce the spread of faults between devices; this embodiment connects the level 1 equipment through the bus, and when the bus maintains a 100V voltage-stabilizing state, the system's power supply and distribution system can be established here to ensure that the system is restarted again; this embodiment is compatible with newly added expansion equipment and enhances the fault tolerance of the system.

[0242] This embodiment solves the problem of large temperature difference at both ends of a large-scale solar-directing mechanism due to different exposure to sunlight, and reasonably controls the temperature difference at both ends of the mechanism through a closed-loop control system to ensure the long-term operation of the large-scale solar-directing mechanism; this embodiment can effectively identify the temperature conditions of various parts of a large-scale mechanism through temperature collection, and effectively control the temperature difference range between mechanisms; this embodiment intelligently identifies and compares the information obtained through temperature collection, and automatically triggers the on and off control of the heating system; this embodiment can readjust the temperature control strategy of the system through a parameter injection interface when problems occur in the existing control strategy, thereby enhancing the fault tolerance of the system.

[0243] 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.

Claims

1. A near-Earth large-scale combined spacecraft multi-source intelligent power supply system, characterized in that include: An intelligent closed-loop control system for stabilizing the temperature difference of large-scale solar-directed directional structures, a multi-region distributed power distribution system suitable for restarting, a system suitable for 360° rotating multi-cabin aircraft structures, and a combined dual-degree-of-freedom solar-directed system; among them, The intelligent closed-loop control system for stabilizing the temperature difference of a large-scale sun-directed directional mechanism: collects the temperature parameters corresponding to each temperature measurement channel, obtains the temperature difference of each temperature measurement channel according to the temperature parameters corresponding to each temperature measurement channel and a preset temperature control reference voltage, obtains the switch signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measurement channel according to the switch signal of the heating circuit corresponding to each temperature measurement channel; The multi-area distributed power distribution system applicable to restarting: obtains primary power from the system bus, converts the primary power into a secondary power in the cabin, and supplies power to the second-level cabin equipment of the power supply and distribution control circuit in the cabin area with the secondary power in the cabin; obtains primary power from the system bus, converts the primary power into an extravehicular secondary power, and supplies power to the second-level extravehicular equipment of the power supply and distribution control circuit in the extravehicular area with the extravehicular secondary power; obtains command voltage from the system bus, and supplies the command voltage to the second-level cabin equipment of the power supply and distribution control circuit in the extravehicular area; The system applicable to the 360° rotating multi-chamber aircraft structure comprises: a fixed end structural module, a structural transmission module and a rotating end structural module; wherein the fixed end structural module is connected to the structural transmission module; the structural transmission module is connected to the rotating end structural module; The combined dual-degree-of-freedom solar orientation system comprises: an A-axis solar orientation subsystem (1), a B-axis solar orientation subsystem (2) and a truss assembly (3); wherein the A-axis solar orientation subsystem (1) and the B-axis solar orientation subsystem (2) are orthogonally mounted on a spacecraft cabin (5) via 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 achieve A-axis rotation control of the solar cell wing (4), and the A-axis is the axis of the spacecraft cabin (5); and the B-axis solar orientation subsystem is used to drive the solar cell wing (4) to rotate around the B-axis to achieve B-axis rotation control of the solar cell wing (4), and the B-axis is the axis of the solar cell wing (4).

2. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 1 is characterized in that: The intelligent closed-loop control system for stabilizing the temperature difference of a large solar-oriented mechanism comprises: a temperature acquisition circuit, a heating control module and a heater loop module; wherein, The temperature acquisition circuit is used to acquire the temperature parameters corresponding to each temperature measurement channel, and transmit the temperature parameters corresponding to each temperature measurement channel to the heating control module; The heating control module receives the temperature parameter corresponding to each temperature measurement channel, obtains the temperature difference of each temperature measurement channel according to the temperature parameter corresponding to each temperature measurement channel and the preset temperature control reference voltage, obtains the switch signal of the heating circuit corresponding to each temperature measurement channel according to the temperature difference of each temperature measurement channel, and transmits the switch signal of the heating circuit corresponding to each temperature measurement channel to the heater circuit module; The heater circuit module receives a switch signal of the heating circuit corresponding to each temperature measuring channel, and controls the autonomous heating and disconnection of the heating circuit corresponding to each temperature measuring channel according to the switch signal of the heating circuit corresponding to each temperature measuring channel.

3. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 2 is characterized in that: Obtaining the temperature difference of each temperature measurement channel according to the temperature parameter corresponding to each temperature measurement channel and the temperature control reference voltage includes: subtracting the temperature parameter corresponding to each temperature measurement channel from the temperature control reference voltage to obtain the temperature difference corresponding to each temperature measurement channel; The switching signal of the heating circuit corresponding to each temperature measuring channel obtained according to the temperature difference of each temperature measuring channel includes: when the temperature difference of each temperature measuring channel is greater than 8°C, the switching signal of the two-stage heating circuit corresponding to each temperature measuring channel is on; when the temperature difference of each temperature measuring channel is not less than 4°C and not greater than 8°C, the switching signal of the first-stage heating circuit corresponding to each temperature measuring channel is on, and the switching signal of the first-stage heating circuit corresponding to each temperature measuring channel is off; when the temperature difference of each temperature measuring channel is ≤4°C, the switching signal of the two-stage heating circuit corresponding to each temperature measuring channel is off.

4. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 3 is characterized in that: The heating control module includes an A / D converter and a CPU; wherein, The A / D converter is connected to the CPU; The A / D converter receives the temperature parameter corresponding to each temperature measurement channel, converts the temperature parameter corresponding to each temperature measurement channel into a digital temperature parameter corresponding to each temperature measurement channel, and transmits the digital temperature parameter corresponding to each temperature measurement channel to the CPU; The CPU: obtains the temperature difference of each temperature measuring channel according to the temperature digital parameter corresponding to each temperature measuring channel and the preset temperature control reference voltage, obtains the switching signal of the heating circuit corresponding to each temperature measuring channel according to the temperature difference of each temperature measuring channel, and transmits the switching signal of the heating circuit corresponding to each temperature measuring channel to the heater circuit module.

5. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 1 is characterized in that: The multi-area distributed power distribution system applicable for restarting includes: a power supply and distribution control circuit in the cabin area, a power supply and distribution control circuit in the cabin area, and a sun-directing device; wherein, The power supply and distribution control circuit of the cabin area is connected to the power supply and distribution control circuit of the outer cabin area through the sun orientation device; The power supply and distribution control circuit in the cabin area obtains primary power from the system bus, converts the primary power into a secondary power supply in the cabin, and uses the secondary power supply in the cabin to supply power to the level 2 cabin equipment of the power supply and distribution control circuit in the cabin area; The second-level cabin equipment of the power supply and distribution control circuit in the cabin area obtains the command voltage from the system bus; The power supply and distribution control circuit of the extravehicular area obtains primary power from the system bus through the sun-directing device, converts the primary power into extravehicular secondary power, and uses the extravehicular secondary power to supply power to the second-level extravehicular equipment of the power supply and distribution control circuit of the extravehicular area; The power supply and distribution control circuit of the outer area obtains a command voltage from the system bus through the solar orientation device, and provides the command voltage to the second-level cabin equipment of the power supply and distribution control circuit of the outer area.

6. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 5 is characterized in that: The power supply and distribution control circuit of the cabin area includes a bus filter and a second-level cabin device; wherein the bus filter obtains primary power from the system bus, converts the primary power into a cabin secondary power, and supplies power to the second-level cabin device with the cabin secondary power; the second-level cabin device obtains a command voltage from the system bus; The power supply and distribution control circuit of the extravehicular area includes a shunt regulator and a second-level extravehicular device; wherein the shunt regulator obtains a primary power supply from the system bus through the sun-directed device, converts the primary power supply into an extravehicular secondary power supply, and supplies the extravehicular secondary power supply to the second-level extravehicular device of the power supply and distribution control circuit of the extravehicular area; the shunt regulator obtains a command voltage from the system bus through the sun-directed device, and supplies the command voltage to the second-level intravehicular device of the power supply and distribution control circuit of the extravehicular area; The level 2 cabin equipment includes charge and discharge regulator a, charge and discharge regulator b, charge and discharge regulator c, charge and discharge regulator d, a cabin drive controller and a power manager; wherein the primary power supply is 100V, and the cabin secondary power supply includes +5V, ±12V and +28V; the cabin secondary power supply required by the charge and discharge regulator a is +5V and ±12V; the cabin secondary power supply required by the charge and discharge regulator b is +5V and ±12V; the cabin secondary power supply required by the charge and discharge regulator c is +5V and ±12V; the cabin secondary power supply required by the charge and discharge regulator d is +5V and ±12V; the cabin secondary power supply required by the cabin drive controller is +5V, ±12V and +28V; the cabin secondary power supply required by the power manager is +5V and ±12V; The level 2 extravehicular equipment includes an integrated drive controller, an extravehicular drive controller, an extravehicular temperature controller and a Taimin controller; wherein the primary power supply is 100V, and the extravehicular secondary power supplies include +5V, ±12V and +28V; the extravehicular secondary power supply required by the integrated drive controller is +5V, ±12V and +28V; the extravehicular secondary power supply required by the extravehicular drive controller is +5V, ±12V and +28V; the extravehicular secondary power supply required by the extravehicular temperature controller is +5V and ±12V; the extravehicular secondary power supply required by the Taimin controller is +5V.

7. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 1 is characterized in that: The fixed end structure module includes a first power receiving device, a lithium ion battery and a first structure ground bus module; wherein the structure ground bus module is connected to the first power receiving device; the lithium ion battery is connected to the first structure ground transmission module; the first power receiving device includes an in-cabin drive controller, a bus filter, a charge and discharge regulator and a power manager; wherein the first structure ground bus module is respectively connected to the in-cabin drive controller, the bus filter, the charge and discharge regulator, the power manager and the lithium ion battery; The structural transmission module includes an electric transmission component and an insulating layer; wherein the electric transmission component is arranged in the insulating layer; the first structural confluence module is connected to the electric transmission component; the electric transmission component is connected to the rotating end structural module; The electric transmission assembly comprises an empty ring, a first grounding ring and a second grounding ring; wherein the first structural ground confluence module is respectively connected to one end of the first grounding ring and one end of the second grounding ring; the other end of the first grounding ring and the other end of the second grounding ring are both connected to the rotating end structural ground module; the empty ring is connected to the rotating end structural ground module; The rotating end structure ground module includes a shunt regulator, a second structure ground confluence module, a power receiving device, a driving mechanism, a solar wing and an on-track installation device; wherein the driving mechanism is connected to the solar wing; the shunt regulator is connected to the empty ring; the shunt regulator is connected to the driving mechanism; the second structure ground confluence module is respectively connected to the other end of the first grounding ring and the other end of the second grounding ring; the second structure ground confluence module is respectively connected to the shunt regulator, the driving mechanism, the power receiving device and the on-track installation device.

8. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 1 is characterized in that: The A-axis solar orientation subsystem (1) comprises: 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); An A-axis main drive controller (102) and an A-axis backup drive controller (104) are used to realize drive control of the A-axis sun orientation mechanism (101); wherein the A-axis main drive controller (102) is installed in a cabin of a spacecraft cabin (5), and the A-axis backup drive controller (104) is installed on the truss assembly (3); 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); wherein 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); 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 a required operating temperature range.

9. The multi-source intelligent power supply system for near-earth large-scale combined spacecraft according to claim 8 is characterized in that: The B-axis solar orientation subsystem (2) comprises: N B-axis solar orientation mechanisms (201) and N B-axis drive controllers (202); 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); wherein each solar cell wing corresponds to one B-axis solar orientation mechanism and one B-axis drive controller; N≥2; The B-axis direction driving controller (202) is installed on the truss assembly (3) and is used to realize driving control of the B-axis direction solar orientation mechanism (201); and realize temperature control of the B-axis direction solar orientation mechanism (201) so that the B-axis direction solar orientation mechanism (201) is within a required operating temperature range; 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); wherein the B-axis solar orientation mechanism (201) is used to drive the solar cell wing (4) to rotate around the B-axis under the drive control of the B-axis drive controller (202), thereby realizing the B-axis rotation control of the solar cell wing (4).

10. The multi-source intelligent power supply system for near-earth large combined spacecraft according to claim 9, characterized in that: The truss assembly (3) comprises: a truss structure (301) and a truss cable; wherein: The truss structure (301) is used to provide a fixed support connection for the A-axis solar orientation mechanism (101) and the B-axis solar orientation mechanism (201), and to provide a mounting platform for an external temperature control instrument (103), an A-axis backup drive controller (104), and a B-axis drive controller (202); 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 external temperature control device (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).

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