Intelligent multispectral load satellite system and operation control method thereof

Abstract

The embodiment of the invention provides an intelligent multispectral load satellite system and an operation control method thereof. The system comprises an energy subsystem, a comprehensive electronic subsystem, a communication subsystem and a load subsystem. Wherein a satellite-borne computer of the integrated electronic subsystem is provided with an attitude and orbit control module, a satellite service module and a power distribution module. The attitude and orbit control module can generate an adjustment instruction according to a current task stage in a set task mode and an attitude signal collected by the sensor. Meanwhile, the satellite service module can generate a device control instruction according to the current task stage. The power distribution module generates a power distribution instruction according to the current task stage, the adjustment instruction and the device control instruction. According to the system, the satellite system can be controlled by integrating the current task stage and the attitude signal through the adjustment instruction, the device control instruction and the power distribution instruction, so that the problem that the satellite system cannot fully utilize hardware resources is solved.

Description

Technical Field

[0001] The present application relates to the field of aerospace control technology, and in particular to an intelligent multi-spectral payload satellite system and an operation control method thereof. Background Art

[0002] A satellite system is a system that uses artificial satellites to orbit the Earth and achieve various functions through cooperation with ground equipment. Depending on the functions achieved, the satellite system can carry satellite payloads corresponding to the functions achieved. Among them, satellite payloads refer to various equipment and instruments installed on satellites to perform specific tasks and collect information from the Earth or space.

[0003] Taking remote sensing satellites as an example, remote sensing satellite systems can carry out optical monitoring of the earth's surface by carrying remote sensing information collection sensors such as spectral payload cameras, thereby obtaining surface spectral information. Remote sensing information collection sensors such as spectral payload cameras can have multiple working bands to obtain spectral information of different wavelengths, realizing a variety of applications such as surface vegetation monitoring, fire monitoring, and atmospheric composition monitoring.

[0004] During operation, the satellite system will have a variety of operating postures and device start-stop states according to the set mission mode. Accordingly, the satellite system has different data processing capacity requirements and power consumption in different operating postures and device start-stop states. However, since the satellite cannot be upgraded after being launched into the space environment, the satellite system needs to reserve a large amount of data processing capacity to cope with different working states. Therefore, the satellite system cannot fully utilize hardware resources and it is difficult to achieve miniaturization. Summary of the invention

[0005] In view of this, an embodiment of the present application provides an intelligent multi-spectral payload satellite system and an operation control method thereof to solve the problem that the satellite system cannot fully utilize hardware resources.

[0006] According to one aspect of the present application, there is provided an intelligent multi-spectral payload satellite system, the system comprising:

[0007] an energy subsystem configured to provide energy input;

[0008] An integrated electronic subsystem connected to the energy subsystem; the integrated electronic subsystem includes an onboard computer, and the onboard computer is equipped with an attitude and orbit control module, a satellite service module, and a power distribution module;

[0009] The attitude and track control module is connected to a plurality of attitude sensors; the plurality of attitude sensors are configured to collect multi-modal attitude signals, and send the attitude signals to the attitude and track control module; the attitude and track control module is configured to generate an adjustment instruction according to the current task stage and the attitude signal in a set task mode, and send the adjustment instruction to an adjustment target device, wherein the task mode includes a plurality of task stages divided according to the running time and / or running attitude; the adjustment target device is a device for adjusting the running attitude determined according to the current task stage and the attitude signal;

[0010] The satellite service module is configured to generate a device control instruction according to the current mission stage, and the device control instruction is used to start or shut down the associated device of the current mission stage; the power allocation module is configured to generate a power allocation instruction according to power demand information, and the power demand information includes one or more combinations of the current mission stage, the adjustment instruction and the device control instruction; the power allocation instruction is used to allocate power supply to the associated device of the current mission stage and / or the adjustment target device;

[0011] A communication subsystem, connecting the energy subsystem and the integrated electronic subsystem;

[0012] A payload subsystem is connected to the energy subsystem and the integrated electronic subsystem; the payload subsystem includes a multi-spectral payload camera, and the multi-spectral payload camera is configured to collect remote sensing data according to the current mission stage and send the remote sensing data to the integrated electronic subsystem.

[0013] Optionally, the system further includes:

[0014] The structural subsystem includes a load-bearing mechanism, an unfolding assembly, and a separating assembly; the load-bearing mechanism is a frame structure formed by connecting a first load-bearing member and a second load-bearing member through a plurality of brackets;

[0015] The unfolding assembly includes a flap and an unfolding hinge, the power source includes a solar panel, and the solar panel is arranged on the flap; the flap is rotatably connected to both sides of the load-bearing mechanism through the unfolding hinge;

[0016] The separation assembly includes a separation mechanism and a separation switch. The separation mechanism is a plug-in device that uses a memory alloy to achieve unlocking. The separation mechanism is connected to the power supply through the separation switch. The separation switch is connected to the satellite service module. The satellite service module is further configured as follows:

[0017] Generate separation instructions according to the set task mode;

[0018] The separation instruction is sent to the separation switch, and the separation instruction is used to control the separation switch to close, so that the separation mechanism is energized to perform a separation action.

[0019] Optionally, the system further includes:

[0020] The thermal control subsystem includes a temperature control component and a heat preservation component; the heat preservation component is arranged on the load-bearing mechanism, and the temperature control component is arranged in the internal space of the frame structure of the load-bearing mechanism; the load camera is temperature-insulated by the heat preservation component;

[0021] The onboard computer is also equipped with a thermal control module; the temperature control assembly includes a thin film temperature control sheet and a temperature sensor; the temperature sensor and the thin film heating sheet are connected to the thermal control module; the thermal control module is also configured as follows:

[0022] receiving the temperature inside the satellite detected by the temperature sensor;

[0023] When the temperature inside the star is not within a preset operating temperature range, a temperature control instruction is sent to the thin film temperature control sheet, wherein the temperature control instruction is used to control the thin film temperature control sheet to generate an adjusted temperature rise.

[0024] Optionally, the system further includes:

[0025] The electrical subsystem includes a cable assembly and a debugging separation switch; the cable assembly is an independent module integrating the cable and the connector in the satellite; the debugging separation switch is arranged on the separation assembly and is configured to detect the separation state of the separation assembly; the debugging separation switch is connected to the satellite service module, and the satellite service module is further configured to:

[0026] Receiving separation state information detected by the debugging separation switch;

[0027] If the separation status information is the first information, the next task of the separation task is executed according to the set task mode, and the first information is used to indicate that the separation component has executed the separation action;

[0028] If the separation status information is the second information, the separation task is repeatedly executed according to the set task mode, and the second information is used to indicate that the separation component has not performed the separation action.

[0029] Optionally, the system further includes:

[0030] A propulsion subsystem, comprising a vacuum arc thruster; the vacuum arc thruster is connected to the power supply through the power distribution module; the power distribution module is further configured to:

[0031] Obtaining the running track information within the preset detection period;

[0032] Comparing the operation track information with the planned operation path to obtain deviation data;

[0033] Setting working parameters of the vacuum arc thruster according to the deviation data, wherein the working parameters include thrust power, thrust time, thrust direction, and propellant ejection amount;

[0034] The operating parameters are sent to the vacuum arc thruster to control the vacuum arc thruster to perform thrusting according to the operating parameters.

[0035] Optionally, the plurality of attitude sensors include a navigation satellite assembly, an angular velocity sensor, and a magnetic sensor; and the attitude and trajectory control module is further configured as follows:

[0036] parsing a spatial positioning signal, an angular velocity signal and a magnetic signal from the multimodal attitude signal detected by the attitude sensor, wherein the spatial positioning signal is a signal detected by the navigation satellite component; the angular velocity signal is a signal detected by the angular velocity sensor; and the magnetic signal is a signal detected by the magnetic sensor;

[0037] Calculate current posture information according to the spatial positioning signal, the angular velocity signal and the magnetic signal;

[0038] According to the set task mode, an adjustment instruction is generated according to the current posture information, wherein the adjustment instruction includes an adjustment target device and an adjustment parameter, and the adjustment target device is a posture adjustment device associated with any one of the spatial positioning signal, the angular velocity signal, and the magnetic signal;

[0039] The adjustment instruction is sent to the adjustment target device, so that the adjustment target device operates according to the adjustment parameter.

[0040] Optionally, the navigation satellite assembly includes a front radio frequency processing unit, a high dynamic navigation unit, a microprocessor and an interface unit electrically connected in sequence;

[0041] The front radio frequency processing unit is connected to the navigation antenna of the communication subsystem; the front radio frequency processing unit is configured to receive satellite signals through the navigation antenna, perform filtering and amplification processing on the satellite signals, and transmit the processed satellite signals to the high dynamic navigation unit;

[0042] The high dynamic navigation unit is configured to perform signal conversion on the satellite signal according to a signal conversion item, and to resolve navigation data from the converted signal; the signal conversion item includes down-conversion, automatic gain control, and analog-to-digital conversion;

[0043] The microprocessor is configured to receive the navigation data from the high dynamic navigation unit, perform orbit determination calculation on the navigation data, obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit, wherein the navigation parameters include positioning data, orbit determination data and time information.

[0044] Optionally, the integrated electronic subsystem further includes a star sensor, the star sensor includes an optical lens and an image processing unit, the optical lens is configured to collect space optical signals, and send the space optical signals to the image processing unit; the image processing unit has a built-in field programmable gate array, and is configured to convert the optical signals into celestial body positioning images, and send the celestial body positioning images to the attitude and orbit control module; the attitude and orbit control module is further configured to:

[0045] Receiving an attitude angle signal detected by the angular velocity sensor;

[0046] Calculate three-axis attitude information according to the standard attitude angle signal and the celestial body positioning image;

[0047] Get the standard posture information of the current task mode;

[0048] Calculating the posture difference between the standard posture information and the three-axis posture information;

[0049] If the posture difference is greater than or equal to the posture difference threshold, the adjustment instruction is generated.

[0050] Optionally, the integrated electronic subsystem further includes a sun-sensitive sensor, which is connected to the attitude and orbit control module; the sun-sensitive sensor is configured to detect sun position information using a four-quadrant differential working principle, and send the sun position information to the attitude and orbit control module; the attitude and orbit control module is further configured to:

[0051] Acquiring a magnetic signal detected by the magnetic sensor;

[0052] Determine the geocentric orientation information according to the magnetic signal;

[0053] Performing dual-vector attitude determination according to the solar position information and the geocentric position information to generate attitude information;

[0054] The adjustment instruction is generated based on the posture information.

[0055] According to another aspect of the present application, a satellite system operation control method is provided, which is applied to the above-mentioned intelligent multi-spectral payload satellite system, and the method includes:

[0056] Acquire a multimodal posture signal and set a current task stage in a task mode, wherein the task mode includes a plurality of task stages divided according to an operating time and / or an operating posture;

[0057] Generate an adjustment instruction according to the current task stage and the posture signal, the adjustment instruction including an adjustment parameter and an adjustment target device, the adjustment parameter is set according to the posture signal; the adjustment target device is a device for adjusting the operating posture determined according to the current task stage and the posture signal;

[0058] Finding the associated components of the current task stage;

[0059] Generate a device control instruction according to the current task stage, wherein the device control instruction is used to start or shut down an associated device of the current task stage;

[0060] Generate a power allocation instruction according to the power demand information, wherein the power demand information includes one or more combinations of the current task stage, the adjustment instruction, and the device control instruction; the power allocation instruction is used to allocate power supply to the associated device of the current task stage and the adjustment target device;

[0061] The adjustment instruction and the power allocation instruction are sent to the adjustment target device, and the device control instruction and the power allocation instruction are sent to the associated device of the current task stage.

[0062] By means of the above technical scheme, the embodiment of the present application provides an intelligent multi-spectral payload satellite system and its operation control method, wherein the system includes an energy subsystem, an integrated electronic subsystem, a communication subsystem and a payload subsystem. Among them, the onboard computer of the integrated electronic subsystem is configured with an attitude and orbit control module, a satellite service module and a power distribution module. The attitude and orbit control module can generate adjustment instructions according to the current mission stage in the set mission mode and the attitude signal collected by the sensor. At the same time, the satellite service module can generate device control instructions according to the current mission stage. The power distribution module generates power distribution instructions according to the current mission stage, adjustment instructions and device control instructions. The system can control the satellite system by integrating the current mission stage and attitude signals through adjustment instructions, device control instructions and power distribution instructions to solve the problem that the satellite system cannot fully utilize hardware resources.

[0063] The above description is only an overview of the technical solution of the present application. In order to more clearly understand the technical means of the present application, it can be implemented in accordance with the contents of the specification. In order to make the above and other purposes, features and advantages of the present application more obvious and easy to understand, the specific implementation methods of the present application are listed below. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0065] Figure 1 A schematic diagram of the structure of an intelligent multi-spectral payload satellite system provided in an embodiment of the present application;

[0066] Figure 2 A schematic diagram of the architecture of the intelligent multi-spectral payload satellite system provided in the embodiment of the present application;

[0067] Figure 3 A schematic diagram of the structure of the integrated electronic subsystem provided in the embodiment of the present application;

[0068] Figure 4 A schematic diagram of the structure of a navigation satellite assembly provided in an embodiment of the present application;

[0069] Figure 5 A schematic diagram of the attitude and trajectory control process provided in an embodiment of the present application;

[0070] Figure 6 The attitude and orbit control principle diagram provided for the embodiment of the present application;

[0071] Figure 7 A schematic diagram of the structural subsystem provided for the embodiment of the present application;

[0072] Figure 8 A schematic diagram of a flow chart of generating adjustment instructions based on a celestial positioning image provided in an embodiment of the present application;

[0073] Fig. 9 A schematic diagram of a flow chart for generating adjustment instructions based on solar position information provided in an embodiment of the present application. DETAILED DESCRIPTION

[0074] The present application will be described in detail below with reference to the accompanying drawings and in combination with embodiments. It should be noted that the embodiments and features in the embodiments of the present application can be combined with each other without conflict.

[0075] In the embodiments of the present application, the satellite system refers to a system that uses artificial satellites to orbit the earth and cooperate with ground equipment to achieve various functions. Depending on the functions achieved, the satellite system can carry satellite payloads corresponding to the functions achieved. Among them, satellite payloads generally refer to various devices and instruments installed on satellites, which are used to perform specific tasks and collect information from the earth or space.

[0076] In some embodiments, the satellite system can carry out optical monitoring of the earth's surface by carrying remote sensing information collection sensors such as spectral payload cameras, thereby obtaining surface spectral information. Remote sensing information collection sensors such as spectral payload cameras can have multiple working bands for obtaining spectral information of different wavelengths.

[0077] During operation, the satellite system will have a variety of operating postures and device start-stop states according to the set mission mode. Accordingly, the satellite system has different data processing capacity requirements and power consumption in different operating postures and device start-stop states. However, since the satellite cannot be upgraded after being launched into the space environment, the satellite system needs to reserve a large amount of data processing capacity to cope with different working states. Therefore, the satellite system cannot fully utilize hardware resources.

[0078] In order to make full use of the hardware resources of the satellite system, an intelligent multi-spectral payload satellite system is provided in this embodiment. Figure 1 As shown, the intelligent multi-spectral payload satellite system may include: an energy subsystem, an integrated electronic subsystem, a communication subsystem, and a payload subsystem. In order to adapt to the on-orbit mission of the satellite system, the intelligent multi-spectral payload satellite system may also include a propulsion subsystem, a structural subsystem, a thermal control subsystem, an electrical subsystem, etc.

[0079] The energy subsystem is used to provide energy input for each device, mechanism, and device in the entire satellite system. In some embodiments, the energy subsystem may include a power supply and a power management module. The power supply may be a solar charging and discharging mechanism composed of a solar cell array and a lithium-ion battery pack. Figure 2 As shown, the power management module may include an EPS main control unit, which is used to control the charging and discharging of the power supply using a maximum power tracking algorithm.

[0080] For example, the lithium-ion battery pack uses 21700-type ternary lithium-ion batteries, 4 series and 2 parallel, the nominal total capacity of the battery pack is 10AH, and the maximum system peak current is 20A. The solar cell array uses triple-junction gallium arsenide solar cells, with a 6-fold sailboard layout, each fold sailboard can be arranged with 6 solar cells, and can provide a peak charging power of 39W. The EPS main control unit adopts the MPPT maximum power tracking algorithm, with a rated output power consumption of ≥40W, a peak power consumption of 64W, a conversion efficiency BCR ≥90%, and a CAN communication interface.

[0081] The integrated electronic subsystem is connected to the energy subsystem to obtain power supply through the energy subsystem. The integrated electronic subsystem includes an on-board computer (IOBC). The on-board computer is a computer equipped on the satellite system, which is responsible for the control and data processing of the satellite system. The on-board computer can realize the functions of attitude control, orbit control, status monitoring, navigation positioning, fault diagnosis, etc. of the spacecraft.

[0082] like Figure 3 As shown, IOBC can provide a variety of communication bus interfaces, for example, it can provide CAN, RS232, RS422, RS485, IIC, SPI and other interfaces, and provide three-phase brushless motor control, three-axis magnetic torquer control, integrated Global Navigation Satellite System (GNSS), Inertial Measurement Unit (IMU), Magnetometer (MTM), and provide multiple load power supply interfaces.

[0083] In order to control the operation process of the satellite system, the attitude and orbit control module, the satellite service module and the power distribution module can be integrated in the integrated electronic subsystem. That is, the onboard computer is configured with the attitude and orbit control module, the satellite service module and the power distribution module. Among them, the attitude and orbit control module is used to adjust the attitude according to the current operation state of the satellite system. To this end, the attitude and orbit control module is connected to a plurality of attitude sensors; the plurality of attitude sensors are configured to collect multi-modal attitude signals and send the attitude signals to the attitude and orbit control module.

[0084] In some embodiments, the plurality of attitude sensors include a navigation satellite assembly, an angular velocity sensor, and a magnetic sensor. The navigation satellite assembly is used to detect a spatial positioning signal. For example, an onboard integrated modular GNSS can be used as a navigation satellite assembly. The navigation satellite assembly using the GNSS onboard integration method can achieve 30m positioning accuracy, 1ms time accuracy, 0.5m / s speed accuracy according to the UART communication method, and can operate stably for more than 3 years at an operating temperature of -40 to 85°C.

[0085] like Figure 4 As shown, in some embodiments, the navigation satellite assembly includes a pre-RF processing unit, a high-dynamic navigation unit, a microprocessor, and an interface unit that are electrically connected in sequence. The pre-RF processing unit is connected to the navigation antenna of the communication subsystem; the pre-RF processing unit is configured to receive satellite signals through the navigation antenna, perform filtering and amplification processing on the satellite signals, and transmit the processed satellite signals to the high-dynamic navigation unit.

[0086] The high dynamic navigation unit is configured to perform signal conversion on the satellite signal according to a signal conversion item, and to resolve navigation data from the converted signal; the signal conversion item includes down-conversion, automatic gain control, and analog-to-digital conversion;

[0087] The microprocessor is configured to receive the navigation data from the high dynamic navigation unit, perform orbit determination calculation on the navigation data, obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit, wherein the navigation parameters include positioning data, orbit determination data and time information.

[0088] For example, the RF input port of the navigation satellite component can be directly connected to the satellite navigation antenna, and the satellite signal is input to the front RF unit for filtering and amplification, and then output to the high dynamic navigation unit, where the signal is processed by down-conversion, AGC, AD conversion, etc., as well as the real-time capture, tracking and positioning of the 12-channel GPSL1CA code navigation signal and the 20-channel BDB1I code navigation signal. After the orbit determination calculation of the MCU, the data positioning, orbit determination data and time information are finally output through the serial port, and the 1PPS time synchronization signal is output. At the same time, the MCU also has the functions of abnormal state monitoring and processing, positioning data identification, interface control, etc.

[0089] Among them, the front RF processing unit, as a unit for processing BD and GPS navigation RF signals, has certain requirements for the sensitivity, anti-interference ability and reliability index of the reception. Therefore, the front RF processing unit can be selected with high-reliability devices that meet the use of space environment and have been verified by long-term flight. The front RF processing unit is connected to the passive antenna, performs out-of-band filtering, low-noise amplification and passive power division on the weak navigation signal received by the antenna, and provides RF signals for each high-dynamic receiver on the main and backup receivers.

[0090] The high-dynamic navigation unit is also called a high-dynamic baseband processing unit. To ensure its stability and reliability and to meet the use requirements of low-orbit satellites, each antenna is connected to an independent navigation module for positioning and solving the problem separately, and the results of independent calculations are judged and output to ensure its reliability. For example, in order to meet the requirements of high-precision positioning, the high-dynamic baseband processing unit can complete the rapid capture of RF signals, complete the mixing correlation of digital intermediate frequency signals, output the integral accumulation results for solving, and complete the loop processing of the signal, complete the positioning solution, output the positioning results and the original observation information to the orbit interface processing unit, and complete the generation of interrupts and second pulses.

[0091] The angular velocity sensor is used to detect angular velocity signals. For example, a MEMS gyroscope can be used as an angular velocity sensor. The MEMS gyroscope can be integrated on board and with dual gyroscope backup, communicate through SPI, achieve zero bias stability of 0.8° / h, better than a measurement range of ±50°, and a measurement accuracy of 0.01° / s, and can operate stably for more than 3 years at an operating temperature of -40 to 85°C.

[0092] The magnetic sensor is used to detect magnetic signals. The magnetic sensor is a magnetic sensor component, and the MTM can also be deployed in an onboard integrated manner and communicate through SPI. The range of the magnetic sensor can be ±8 gauss, and the accuracy is 1%. It is also required to operate stably for more than 3 years at an operating temperature of -40 to 85°C.

[0093] In some embodiments, in order to further miniaturize the satellite system, sensors for attitude and orbit control, such as gyroscopes, magnetometers, GNSS, magnetic torquers, flywheel controllers, etc., can be integrated into the IOBC. At the same time, the IOBC can directly power the star sensor and the sun sensor and obtain their data through the signal bus. The attitude control algorithm is also executed in the main control chip of the IOBC. Figure 5 As shown, the attitude and track control module is also configured as follows:

[0094] S101, generating an adjustment instruction according to the current task stage in the set task mode and the posture signal.

[0095] The mission mode includes multiple mission phases divided according to the operating time and / or operating attitude. For example, the mission mode can set multiple mission phases according to the operating process of the satellite after launch, including separation phase, initialization phase, self-test phase, entry phase, measurement and control phase, exit phase, etc. In different mission phases, the satellite system needs to start or shut down some components according to the requirements of the current mission phase.

[0096] In some embodiments, the device that needs to change the operating state in a certain mission stage can be called the associated device of the mission stage. Among them, changing the operating state means adjusting the device to adapt to the corresponding mission stage requirements when entering the mission stage. The change of operating state can include changing from unstarted to started, from started to unstarted, from standby to working state, etc. For example, when entering the initialization stage, the satellite system is powered on, and in addition to the IOBC main chip and battery management component working, the satellite system operates the sailboard to unfold and starts to initialize the various components of the system. Then the associated devices of the initialization stage may include the IOBC main chip, the battery management component, the sailboard, and the initialized system components.

[0097] Since different mission phases have different requirements for the attitude of the satellite, it is necessary to obtain the current mission phase in the set mission mode when adjusting the attitude. The current mission phase can be determined based on the current time, the current operating position (whether it is inbound), and the current attitude signal detected by the attitude sensor. For example, according to the on-orbit mission mode setting, in the [T1, T2] time interval, the satellite system performs measurement and control, and the current time T can be obtained. If T is within the [T1, T2] time interval, it can be determined that the current mission phase is the measurement and control phase.

[0098] In some embodiments, when multiple attitude sensors include navigation satellite components, angular velocity sensors, and magnetic sensors, the attitude and trajectory control module can parse the spatial positioning signal, angular velocity signal, and magnetic signal from the multimodal attitude signal detected by the attitude sensor. Then calculate the current attitude information based on the spatial positioning signal, the angular velocity signal, and the magnetic signal. Then, according to the set mission mode, generate an adjustment instruction based on the current attitude information. Wherein, the adjustment instruction includes an adjustment target device and an adjustment parameter. The adjustment target device is a device for adjusting the operating attitude determined according to the current mission stage and the attitude signal, that is, the adjustment target device is an attitude adjustment device associated with any one of the spatial positioning signal, the angular velocity signal, and the magnetic signal.

[0099] S102: Send the adjustment instruction to the adjustment target device.

[0100] After generating the adjustment instruction, the attitude and trajectory control module may send the adjustment instruction to the adjustment target device so that the adjustment target device operates according to the adjustment parameters. Figure 6 As shown, when the satellite system enters the on-orbit measurement phase, the current operating attitude P of the satellite system can be determined through the attitude signal. If the operating attitude does not meet the attitude requirement P0 of the on-orbit measurement phase, the target device for attitude adjustment can be determined to be the momentum wheels according to the deviation ΔP between the current operating attitude P and the attitude P0 of the on-orbit measurement phase, and the adjustment parameter ΔP is generated, that is, the adjustment instruction is generated. Then the adjustment instruction is sent to the momentum wheels, so that the momentum wheels perform attitude control.

[0101] Among them, the momentum wheel set is a device that performs attitude control through momentum calculation. For example, the momentum wheel set can perform attitude control of the satellite system through a single-machine integration method of three orthogonal and one oblique installation, based on the FOC control algorithm, with a speed control control method. In order to meet the attitude control requirements, the momentum wheel set can have a speed range of -6000 to 6000. With an angular momentum better than 8mNms and a torque of 2mNm, it can operate for more than 3 years at an operating temperature of -40 to 85°C.

[0102] The satellite service module is responsible for satellite time management, on-orbit mission execution, whole satellite status management and monitoring, deployment component actions, etc. The satellite service module can be implemented by running mission-related applications on the onboard computer. In some embodiments, the satellite service module is implemented in the form of a finite state machine, that is, the satellite system is in a specific state and performs specific actions, such as turning on and off the power of a single machine, turning on and off the measurement and control data transmission, and executing tasks.

[0103] For example, the satellite service module can perform time management, storage management, subsystem management, flight attitude management, thermal control management, working mode control, satellite power consumption management, satellite status monitoring and maintenance, etc. Time management means determining and managing the on-board time reference based on the GNSS module time information or the time data recorded on the ground. When the satellite is unexpectedly powered off for a long time, it can also be calibrated based on the GNSS time.

[0104] Storage management is designed with multiple storage spaces in the satellite management system. The satellite management application is designed with storage management functions to uniformly store and manage some configuration information, task queues and other data that require non-volatile storage. Subsystem management means that the satellite module implements the preset flight mission through command control of the subsystem according to the design of the flight mission, cooperates with the payload adaptation unit to realize the control and management of the onboard payload and realize the preset satellite function.

[0105] Flight attitude management is the onboard computer acting as the controller of the attitude control subsystem, which realizes the attitude control of the satellite according to the different working modes and attitude control expectations of the satellite. Thermal control management is to perform thermal control management on the key nodes inside the entire satellite according to the operating temperature requirements of the camera and battery, so as to ensure that the key components work within the temperature range required by the design.

[0106] Working mode control is to design different working modes according to the different working stages of the satellite to achieve the satellite's flight mission. Satellite power consumption management is to monitor the energy of the entire satellite in real time during the satellite flight. When the satellite faces the risk of battery over-discharge, the satellite is energy-controlled, the entire satellite is controlled in a safe mode, and high-energy-consuming equipment is turned off to ensure the stability and normality of the satellite energy system.

[0107] Satellite status monitoring and maintenance is the process of regularly collecting and real-time monitoring the status of the entire satellite during satellite operation, and diagnosing and handling foreseeable faults. When an abnormal state occurs in the satellite operation, the fault can be automatically handled to ensure the stability of the satellite operation.

[0108] In order to perform corresponding mission control in different mission stages, the satellite service module is configured to generate device control instructions according to the current mission stage. Among them, the device control instructions are used to start or shut down the associated devices of the current mission stage. For example, when the satellite system enters the initialization mission stage, the separation switch needs to be in a separated state, and the system is powered on. In addition to the IOBC main chip and the battery management component working, the system operates the sailboard to unfold and starts to initialize the various components of the system. At this time, the associated devices of the current mission stage are the sailboard and the various system components that need to be initialized. Therefore, control instructions for unfolding the sailboard and for initializing the various system components can be generated. The control instructions are then sent to the sailboard and the various system components that need to be initialized to perform the system initialization task.

[0109] The power distribution module is used to provide energy for each part of the satellite system and dynamically adjust the energy supply of each part according to the overall power distribution of the current system. For example, after being configured into the IOBC, the power distribution module can provide energy for the communication system and the payload camera, and reserve 12V and 5V outputs to support subsequent expansion applications.

[0110] Therefore, the power allocation module is configured to generate a power allocation instruction based on power demand information, wherein the power demand information includes one or more combinations of the current task stage, the adjustment instruction and the device control instruction; the power allocation instruction is used to allocate power supply to the associated devices of the current task stage and / or the adjustment target device.

[0111] After the satellite system is launched, most of the energy supply is completed by the energy subsystem, that is, the solar cell array and lithium-ion battery pack in the energy subsystem provide power. However, the total amount of power provided by the solar cell array and lithium-ion battery pack is limited. Therefore, during the operation of the satellite system, the power allocation module can generate power allocation instructions based on the power demand information.

[0112] In some embodiments, the power allocation instruction can be generated according to the current mission stage, that is, after the satellite system enters a new mission stage, the power allocation module can control some devices to turn on and some devices to turn off according to the mission requirements of the current mission stage.

[0113] For example, the measurement mission phase of the satellite system is to perform spectrum acquisition on a specific area. After entering the measurement mission phase, the power allocation instruction can send a power allocation instruction to the multi-spectral payload camera in the payload subsystem, thereby controlling the multi-spectral payload camera to start running.

[0114] In some embodiments, the power allocation instruction can also be generated based on the adjustment instruction. That is, when the satellite system needs to make an attitude adjustment, since the attitude adjustment needs to be performed by the momentum wheel set, the three-axis magnetic rod and other actuators to generate an adjustment action, after the adjustment instruction is generated, the power allocation module can read the adjustment target device in the adjustment instruction and generate a power allocation instruction for the adjustment target device. Then, the power allocation instruction is sent to the adjustment target device, so that the adjustment target device obtains power supply and performs the adjustment action.

[0115] In some embodiments, the power allocation instruction can also be generated according to the device control instruction. That is, when the satellite system operates according to the predetermined mission mode, after the satellite service module generates the device control instruction according to the current mission stage, the power allocation module can read the associated devices of the current mission stage from the device control instruction and generate the power allocation instruction for the associated devices of the current mission stage. Then, the power allocation instruction is sent to the associated devices of the current mission stage, so that the associated devices of the current mission stage can obtain power supply and perform the corresponding tasks.

[0116] The communication subsystem connects the energy subsystem and the integrated electronic subsystem. The communication subsystem is used to establish a communication connection between the satellite system and the ground monitoring system, send data to the ground monitoring system, or receive data from the ground monitoring system. For example, the satellite orbit altitude is 500km, the minimum elevation angle of the ground station antenna is 5°, and the maximum distance between the satellite and the ground is about 2100km. For the downlink, when performing the ±50 degree telemetry mission, the operating frequency is 8350MHz. The corresponding on-board transmit power is 0dBW, the on-board transmit channel feed loss is 1dB, the on-board antenna gain is -3dBi, and the on-board transmit EIRP is -4dBW. The effective range is 2090km, the space loss is 177.37dB, the atmospheric loss is 3.5dB, the antenna pointing loss is 0.5dB, the total received signal power level of the ground station is -185.28dBW, the G / T value of the ground station is 24dB / k, the Boltzmann constant is -228.6dBW / HzK, the C / N0 received by the ground station is 67.32dBHz, the data rate is 8192bps, the received Eb / N0 is 28.18dB, and the theoretical value of Eb / N0 is 10.6dB. The coding gain is 0dB and the demodulation loss is 3dB. The communication subsystem can include an X-band communication machine, which is also called an X-band measurement, control and data transmission integrated machine, which is used to establish a satellite-to-ground wireless communication link. The X-band communication machine can achieve a communication capability of 8192bps, 20M data transmission and 32GB capacity.

[0117] The X-band measurement and control digital transmission and measurement and control integrated machine can include a digital transmission transmitter, a measurement and control receiver, and a measurement and control transmitter. The digital transmission transmitter can receive and store satellite payload, satellite service and other data, replay the stored data when the satellite passes through the station, and perform AOS framing, scrambling, encoding and other processing. After processing, microwave modulation is performed, and the data is fed into the antenna through microwave channel circuits such as radio frequency amplifiers and transmitted to the ground digital transmission receiving station.

[0118] The TT&C receiver can receive uplink telemetry data from the ground station, upload update program data, etc. When receiving uplink data, it demodulates it in real time on orbit, and performs channel coding format and data identification processing, and stores the received data appropriately or directly sends it to the onboard computer for further data processing. The TT&C transmitter can receive telemetry data from the onboard computer in real time, and send the telemetry data on the satellite when passing through the station.

[0119] In order to meet the communication needs of the satellite system, the X-band measurement and control digital transmission and measurement and control integrated machine can realize payload communication with Cameralink according to the CAN control method, and the storage capacity is up to 32GB. In different mission modes, it can show different performance. For example, in remote control mode, the sensitivity of the communication system is 106dbm, and the frequency range is 7242Mhz~7250Mhz. According to the BPSK modulation method, the rate reaches 8192bps, and the bit error rate is less than 1×10 -6 Similarly, in telemetry mode, the power is 22dbm, the frequency is 8325Mhz~8375Mhz, and the BPSK modulation method is used to achieve a rate of 8192bps, with a bit error rate of less than 1×10 -6 In the digital transmission mode, the power is 30dbm and the frequency is 8325Mhz~8375Mhz. According to the LDPC, 7 / 8 modulation method, the rate can reach 5Mbps-20Mbps and the speed is less than 1×10 -6 bit error rate.

[0120] The communication subsystem may also include an antenna assembly, which may include a receiving antenna (uplink antenna) and a transmitting antenna (downlink antenna). For example, based on the overall requirements of the satellite system, the antenna indicators, system links and space use environment requirements are comprehensively evaluated. The receiving antenna may use a double-layer microstrip antenna. The gain of the double-layer microstrip antenna is 9dBi, ±60° and the Tsukuba ratio is ≤1.5. It can run stably for more than 3 years at an operating temperature of -90 to 90°C. The transmitting antenna may use a 2×2 array microstrip antenna, and the X-metering, control and digital transmission integrated machine measurement and control transmission and digital transmission transmission are time-division multiplexed. The gain of the transmitting antenna is 0dbi, ±50°, and the Tsukuba ratio is ≤1.5. It also needs to run stably for more than 3 years at an operating temperature of -90 to 90°C.

[0121] In some embodiments, the communication subsystem is configured to establish a communication link according to a preset frequency band. For example, in addition to the X-band communicator shown above, the communication subsystem may also include a UV-band communicator to establish a satellite-to-ground wireless communication link. The communication subsystem is also connected to the energy subsystem through the power distribution module, so that the energy subsystem obtains power supply during the communication process through the power distribution module.

[0122] The load subsystem is used to perform specific monitoring tasks according to the design requirements of the satellite system. The load subsystem can be connected to the energy subsystem to obtain power supply from the energy subsystem. The load subsystem also establishes a connection relationship with the integrated electronic subsystem to be controlled by the integrated electronic subsystem to perform corresponding tasks.

[0123] In some embodiments, the payload subsystem includes a multi-spectral payload camera, which may include an optical part and an electronic part. For example, the payload camera has a focal length f of 443.325 mm, an F / # of 5.1, a spectral range of 450 to 900 nm, a resolution of 3.45 m, a width of 10.9 km, an orbital height H of 500 km, and a pixel size α of 3.2 μm.

[0124] The multi-spectral payload camera is configured to collect remote sensing data according to the current mission stage, and send the remote sensing data to the integrated electronic subsystem. For example, according to a predetermined mission mode, when the satellite system enters the measurement stage, the satellite service module can generate a device control instruction according to the current mission stage. Among them, the device control instruction can specify that the associated device of the current mission stage is a payload camera, so that the device control instruction is sent to the payload camera. After receiving the device control instruction, the payload camera can respond to the device control instruction and perform spectral acquisition. During the spectral acquisition process, remote sensing spectral data can be generated along with the monitoring process, and the remote sensing spectral data can be sent to the integrated electronic subsystem. The integrated electronic subsystem can process the spectral data and send it to the ground monitoring system through the communication subsystem.

[0125] It can be seen that in the above embodiment, the system can control the satellite system by integrating the current mission phase and attitude signals through adjustment instructions, device control instructions and power allocation instructions to solve the problem that the satellite system cannot fully utilize hardware resources.

[0126] like Figure 7As shown, in some embodiments, the system also includes a structural subsystem. The structural subsystem involves mechanical structures such as support structures, stacks, flywheels, star sensors, stand-alone brackets, skins, solar panels, panel deployment mechanisms, antennas, cameras, etc. The structural subsystem may include a load-bearing mechanism 1, an deployment component 2, and a separation component 3. Among them, the load-bearing mechanism 1 is a frame structure formed by connecting a first load-bearing member 11 and a second load-bearing member 12 through a plurality of brackets 13. For example, the first load-bearing member 11 and the second load-bearing member 12 are two plate-like structural members arranged on both sides of the satellite system. The first load-bearing member 11 and the second load-bearing member 12 are connected by a plurality of brackets 13 to form a frame structure with an envelope size of 125mm (X) × 230mm (Y) × 420mm (Z).

[0127] The frame structure can form a structure layout of the sky part and the ground part on the satellite system. For example, the sky part can be equipped with an X-band oscillator antenna for measurement and control, a sun sensor, and a GPS antenna on the top surface. After the sailboard is unfolded, there are solar cells on the sky surface, which is designed as the main illuminated surface. The ground part includes an X-band conical spiral antenna, an X-band data transmission antenna, a camera, and two separation switches.

[0128] The first load-bearing member 11 can serve as the main load-bearing structure of the satellite system. The main load-bearing structure is the design and assembly reference of the entire satellite structure. There are four slide rail surfaces on the left and right sides of the first load-bearing member, which are in contact with the separation mechanism. They serve as bearing surfaces during transportation and launch. During assembly and measurement, the upper left is used as the reference origin of the entire satellite. The first load-bearing member is also provided with mounting holes and mounting surfaces for each single machine and bracket. The bracket mounting surface on the main load-bearing plate has positioning pin holes for accurate positioning of the intermediate bracket. The openings on the main load-bearing plate are for expanding the installation space, thermal control holes, and weight reduction settings.

[0129] The second load-bearing member 12 can be used as a secondary load-bearing structure of the satellite system. The secondary load-bearing structure has similar functions to the main load-bearing plate, but is not used as a reference for installation and measurement. The intermediate bracket mounting surface on the secondary load-bearing plate is designed with a positioning pin hole for positioning the intermediate bracket.

[0130] The bracket 13 may include a standard bracket and a special bracket. The standard bracket may specifically include a single-pole bracket, a single-pole bracket, a double-pole bracket, and a rectangular bracket. The standard bracket may be provided with screw holes and positioning pins connected to the main and secondary load-bearing structures, and there are fixing holes for the skin and cables on the side. The public bracket not only connects the main and secondary load-bearing plates, but also serves as a supporting bracket for various components. Some of the brackets can be used for the installation of Taimin, vibrator antennas, and GPS antennas. The conical antenna bracket can be used to install the conical antenna. The digital transmission antenna bracket is used to install the X digital transmission antenna. Depending on the support target, the public bracket may include a magnetic torquer bracket, a star sensor bracket, a flywheel bracket, etc., which are used to install a magnetic torquer, a star sensor, and a flywheel, respectively.

[0131] The connection interface between structures is divided into through holes and threaded holes. The screws pass through the through holes of the structural bracket and are screwed into the threaded holes of the load-bearing plate connected to it to achieve the connection between the bracket and the load-bearing plate. The bracket is provided with two positioning pins at the top and bottom to ensure the installation accuracy. The single machine inside the star body is connected by the equipment and the bracket installation, and then fixed by the bracket and the load-bearing plate installation. The single machine bracket is used as the installation of the single machine equipment and also plays a supporting role in the star body. The connection between the bracket and the load-bearing plate is divided into through holes and threaded holes. The screws pass through the through holes of the load-bearing plate and are screwed into the threaded holes of the bracket connected to it to achieve single machine installation and structural connection.

[0132] The unfolding assembly 2 includes a flap 21 and an unfolding hinge 22 , and the power source includes a solar panel 4 , which is arranged on the flap 21 ; the flap 21 is rotatably connected to both sides of the load-bearing mechanism 1 through the unfolding hinge 22 .

[0133] For example, a satellite system may include two groups of solar panels, each of which is provided with a solar panel, and the two groups of panels are arranged in a mirror image on the left and right to form an energy subsystem including two solar panels. The size of the panel can be 178mm×320.5mm, and the panel has double-sided fabric. The solar panel has two states, a folded state and an unfolded state, and is supported and bound by hinges and burn-wire modules. The solar wing substrate can use a panel PCB, and the thickness of the panel substrate PCB is 0.8mm, and the panel has double-sided fabric. In space, the unfolding hinge can be heated by the resistor in the burn-wire module, and the rope that controls the folding of the solar wing can be melted, and the unfolding hinge can be unfolded to unfold the panel to the target position. The unfolding hinge can also add a locking design to better control the unfolding process.

[0134] In some embodiments, the separation component includes a separation mechanism and a separation switch. The separation mechanism is a plug and unplug device that uses a memory alloy to achieve unlocking; the separation mechanism is connected to the power supply through the separation switch. For example, the separation component uses a memory alloy pin puller to achieve unlocking, the electrical connector model is Y8C-4ZJBL, and the corresponding power supply plug model is Y8C-4TK. The total resistance of the memory alloy pin puller is about 1.1Ω, which is a pure impedance load. The power-on design is a dual-point dual-wire backup power supply, and the total current adaptation range is 2A to 30A. When used once on track, the total power supply energy is required to be 20 to 60J, the power supply current is 10A, and the power supply time is 200ms.

[0135] When the power supply current is less than 5A, the heat dissipation effect needs to be considered, and the input energy can be increased by more than 50%. The insulation resistance between the energized contacts and the shell of the memory alloy pin puller is not less than 20MΩ. When multiple independently unlocked separation seats are used to form a star-rocket connection separation component, the number, specifications and positions of the separation seats can be flexibly configured. The power supply requirements of each separation seat are the same as the above requirements, and multiple separation seats can be powered in series or in parallel according to the power supply capacity. To ensure the unlocking synchronization of multiple separation seats, the current obtained by each separation seat is consistent and greater than 5A.

[0136] In addition, the separation component can provide 2-way travel switch separation state measurement signals, which are closed before separation and open after separation. The carrier can choose to measure one or both of them. The installation position of the travel switch is symmetrical relative to the center of the separation component. The closed resistance of the travel switch is <0.2Ω, and the resistance after disconnection is ≥20MΩ. Each measurement signal is two swing wires with a length of about 200mm. The two wires are respectively equipped with CX-1A-5 pins and sockets. The CX-1A-5 pins and sockets are compatible and interchangeable with the CX-1B-5 pins and sockets.

[0137] In order to perform the separation task, the separation switch is connected to the satellite service module, and the satellite service module is also configured to generate a separation instruction according to the set task mode, and send the separation instruction to the separation switch. The separation instruction is used to control the separation switch to close, so that the separation mechanism is powered on to perform the separation action.

[0138] In some embodiments, the system further includes a thermal control subsystem for controlling the internal temperature of the satellite system so that the internal components of the satellite system can operate within the designed operating temperature range. The thermal control subsystem is mainly passive thermal control, supplemented by active thermal control. Active thermal control is achieved by a thin film heater, and passive thermal control can be achieved by materials such as thermal insulation materials and heat preservation materials. Therefore. The thermal control subsystem includes a temperature control component and a heat preservation component. The heat preservation component is arranged on the load-bearing mechanism, and the temperature control component is arranged in the internal space of the frame structure of the load-bearing mechanism; the payload camera is temperature-insulated by the heat preservation component.

[0139] For example, the thermal insulation design of the payload camera can be isolated from the satellite platform by a thermal insulation pad, so that the payload camera has one temperature measurement and two temperature control channels, and the IOBC controls the heating compensation. In addition, the battery components in the energy subsystem can also be designed with a heating circuit. The heat control of the satellite system can be achieved by covering the entire satellite with multiple layers and designing the heat dissipation for the parts that generate more heat.

[0140] The onboard computer is also equipped with a thermal control module; the temperature control component includes a thin film temperature control sheet and a temperature sensor; the temperature sensor and the thin film heating sheet are connected to the thermal control module; the thermal control module is also configured to receive the satellite temperature detected by the temperature sensor, and send a temperature control instruction to the thin film temperature control sheet when the satellite temperature is not within a preset operating temperature range. The temperature control instruction is used to control the thin film temperature control sheet to produce an adjusted temperature rise.

[0141] For example, in the thermal control subsystem, passive thermal control based on thermal insulation materials can be used for the entire satellite system. For the payload camera and battery components, one-way active heating compensation combined with passive thermal control can be used respectively, which can control the temperature inside the satellite within 10℃ ~ 45℃.

[0142] In some embodiments, the system further includes an electrical subsystem. The electrical subsystem includes a cable assembly and a debugging separation switch; the cable assembly is an independent module that integrates the satellite cable and the connector. Integration into an independent module can make the electrical subsystem an independent functional module, which is convenient for realizing fast cable connection and interface layout, and provides technical support for mass production. For example, the control interface between the rocket control system and the satellite is used to transmit the satellite-rocket separation command. The satellite triggers separation after receiving a pulse signal through a memory alloy pin puller. The pin puller has two bridge wires used in parallel, each with a resistance of about 2.2Ω, and a single total resistance of about 1.1Ω. It is a pure impedance load, and the power-on design is a dual-point dual-wire backup power supply, with a total current adaptation range of 2A to 30A. When the pin puller and the carrier equipment are electrically connected and powered on for unlocking test, in order to ensure the repeated service life of the pin puller, the power-on energy should be controlled at 10J to 15J. The power-on contacts of the memory alloy pin puller and the shell maintain an insulation resistance of ≥20MΩ.

[0143] Therefore, the debugging separation switch is provided on the separation component and is configured to detect the separation state of the separation component. The debugging separation switch is connected to the satellite service module. The satellite service module is also configured to receive the separation state information detected by the debugging separation switch, and if the separation state information is the first information, the next task of the separation task is executed according to the set task mode, wherein the first information is used to characterize that the separation component has executed the separation action, that is, after the separation switch detects the separation of the satellite and the rocket, the separation state information with the content of the first information can be generated. Similarly, if the separation state information is the second information, the separation task is repeated according to the set task mode, and the second information is used to characterize that the separation component has not executed the separation action.

[0144] By detecting the separation status information through the separation switch, the separation task can be repeated when it is detected that the satellite and rocket have not separated, so as to improve the mission processing capability and reliability of the satellite system.

[0145] In some embodiments, the system further comprises a propulsion subsystem. The propulsion subsystem can be used to perform orbit change or attitude adjustment. The propulsion subsystem can include a vacuum arc thruster (VAT); the vacuum arc thruster is connected to the power supply through the power distribution module.

[0146] For example, the working fluid of the vacuum arc thruster is titanium alloy metal. The vacuum arc thruster is an electromagnetic propulsion system in the electric propulsion system, and its cathode material is also used as a propellant working fluid. Using the principle of barrier dielectric discharge, an arc is instantly generated between the anode and the cathode. The cathode will be burned by the arc to form metal vapor and ionized into plasma. The instantaneous high-current arc will also generate a strong electromagnetic field. The ionized plasma will be ejected axially under the action of the electromagnetic field to generate thrust.

[0147] Vacuum arc thrusters can combine the supply of non-toxic propellant with the thruster body into one module, eliminating the need for complex propellant storage and supply systems. They are suitable for space application needs of microsatellites and for performing propulsion tasks with high control precision.

[0148] In order to perform the propulsion task, the power distribution module is also configured to first obtain the operation track information within the preset detection period, and then compare the operation track information with the planned operation path to obtain deviation data. Then, the working parameters of the vacuum arc thruster are set according to the deviation data. Among them, the working parameters include propulsion power, propulsion time, propulsion direction and propellant ejection amount. By sending the working parameters to the vacuum arc thruster, the vacuum arc thruster is controlled to perform propulsion according to the working parameters.

[0149] Furthermore, as a refinement and expansion of the specific implementation methods of the above-mentioned embodiments, in order to fully illustrate the specific implementation process of this embodiment, some embodiments of the present application also provide an intelligent multi-spectral payload satellite system, wherein the integrated electronic subsystem of the intelligent multi-spectral payload satellite system also includes a star-sensitive sensor.

[0150] The star sensor includes an optical lens and an image processing unit. For example, the star sensor is composed of a sunshade, a lens, and an image processing unit. In order to achieve the miniaturization of the system, the electronics system adopts an assembly method of stacking two processing boards, namely a CMOS imaging system and an FPGA image processing and communication system. According to the working characteristics of the star sensor, the aberrations that affect the star point and measurement accuracy can be corrected first based on the optical signal collected by the optical lens, and the tolerances of other aberrations can be appropriately relaxed to meet the measurement accuracy requirements.

[0151] The hood design adopts the finite element analysis algorithm. The initial structure and initial values ​​of the surface reflection characteristics are given according to the design parameters. The extinction capacity is recalculated through finite element analysis, and the design parameters are adjusted until the extinction capacity meets the requirements. The lens structure adopts an integrated design of light, machine and heat. The optics, structure, material matching, processing and assembly methods are determined through simulation analysis and engineering analysis to ensure that the total defocus of the lens within a certain temperature range is within the range allowed by the measurement accuracy. The circuit part corresponding to the image processing unit can adopt an ultra-simplified, high-reliability circuit design architecture, and a high-precision design concept of combining rigidity and flexibility with the integration of core components. The system aerospace adaptability design principle of radiation-resistant components and single-particle-resistant circuits realizes the interface circuit board structure.

[0152] The optical lens is configured to collect space optical signals and send the space optical signals to the image processing unit; the image processing unit has a built-in field programmable gate array and is configured to convert the optical signals into celestial body positioning images and send the celestial body positioning images to the attitude and orbit control module. Figure 8 As shown, the attitude and track control module is also configured as follows:

[0153] S201, receiving an attitude angle signal detected by the angular velocity sensor;

[0154] S202, calculating three-axis attitude information according to the standard attitude angle signal and the celestial body positioning image;

[0155] S203, obtaining standard posture information of the current task mode;

[0156] S204, calculating the posture difference between the standard posture information and the three-axis posture information;

[0157] S205: If the posture difference is greater than or equal to the posture difference threshold, generate the adjustment instruction.

[0158] It can be seen that the attitude and orbit control module can accurately calculate the three-axis attitude information by acquiring the attitude angle signal and the celestial body positioning image. Then the standard attitude information of the current task mode is obtained, and the three-axis attitude information is compared with the standard attitude information to calculate the attitude difference between the standard attitude information and the three-axis attitude information. When the attitude difference is greater than or equal to the attitude difference threshold, an adjustment instruction is generated according to the attitude difference to achieve attitude adjustment.

[0159] In some embodiments, the integrated electronic subsystem also includes a sun-sensitive sensor. The sun-sensitive sensor can be used to detect the sun's position, so as to perform attitude control according to the sun's position, or to adapt to a specific mission phase. For example, the predetermined mission mode of the satellite system may include a sun capture mode, which is used to realize functions such as rate damping, sun capture (including pitch search and roll search), and solar cruise after the separation of the satellite and rocket. The sun capture mode can be used as a fail-safe mode for the transfer orbit and a safety mode for the worst case of the synchronous orbit. When the satellite and rocket are separated, or when a fault occurs during the operation of the satellite, the sun capture mode can be activated. In the sun capture mode, the satellite can be rate damped first, and the solar search can be started after the angular rates of each axis are less than 0.5° / s and last for a period of time. In the sun capture mode, the sun-sensitive sensor can measure the roll attitude and pitch angle information, and when the sun-sensitive sensor is facing the sun, it switches to the solar cruise mode.

[0160] The sun-sensitive sensor is connected to the attitude and orbit control module; the sun-sensitive sensor is configured to detect the sun's position information using a four-quadrant differential working principle, and send the sun's position information to the attitude and orbit control module. Fig. 9 As shown, the attitude and track control module is also configured as follows:

[0161] S301, obtaining a magnetic signal detected by the magnetic sensor;

[0162] S302, determining the geocentric orientation information according to the magnetic signal;

[0163] S303, performing dual-vector attitude determination according to the solar position information and the geocentric position information to generate attitude information;

[0164] S304: Generate the adjustment instruction based on the posture information.

[0165] It can be seen that through the dual-vector attitude determination of the sun sensor and magnetometer, X1 and X2 can be replaced by the sun vector and the geomagnetic vector respectively. These two vectors can be expressed in the inertial system by orbit calculation, and can also be expressed in the satellite body by the sun sensor and magnetometer. Therefore, using the above dual-vector attitude determination principle, the attitude conversion matrix from the system to the inertial system can be obtained, thereby achieving more precise attitude adjustment.

[0166] Based on the intelligent multi-spectral payload satellite system provided in the above embodiments, some embodiments of the present application also provide a satellite system operation control method, which is applied to the above intelligent multi-spectral payload satellite system, and the method includes:

[0167] Acquire a multimodal posture signal and set a current task stage in a task mode, wherein the task mode includes a plurality of task stages divided according to an operating time and / or an operating posture;

[0168] Generate an adjustment instruction according to the current task stage and the posture signal, the adjustment instruction including an adjustment parameter and an adjustment target device, the adjustment parameter is set according to the posture signal; the adjustment target device is a device for adjusting the operating posture determined according to the current task stage and the posture signal;

[0169] Finding the associated components of the current task stage;

[0170] Generate a device control instruction according to the current task stage, wherein the device control instruction is used to start or shut down an associated device of the current task stage;

[0171] Generate a power allocation instruction according to the power demand information, wherein the power demand information includes one or more combinations of the current task stage, the adjustment instruction, and the device control instruction; the power allocation instruction is used to allocate power supply to the associated device of the current task stage and the adjustment target device;

[0172] The adjustment instruction and the power allocation instruction are sent to the adjustment target device, and the device control instruction and the power allocation instruction are sent to the associated device of the current task stage.

[0173] By applying the technical solution of this embodiment, the method can generate adjustment instructions according to the current mission stage in the set mission mode and the attitude signal collected by the sensor through the attitude and orbit control module of the intelligent multi-spectral payload satellite system. At the same time, the device control instructions can be generated according to the current mission stage through the satellite service module. The power allocation module generates power allocation instructions according to the current mission stage, adjustment instructions and device control instructions. The method can control the satellite system by integrating the current mission stage and attitude signals through adjustment instructions, device control instructions and power allocation instructions to solve the problem that the satellite system cannot fully utilize hardware resources.

[0174] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0175] The above-described embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be construed as limiting the scope of the present application. It should be noted that, for a person of ordinary skill in the art, several modifications and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the attached claims.

Claims

1. An intelligent multi-spectral payload satellite system, characterized in that: The system comprises: an energy subsystem configured to provide energy input; An integrated electronic subsystem connected to the energy subsystem; the integrated electronic subsystem includes an onboard computer, and the onboard computer is equipped with an attitude and orbit control module, a satellite service module, and a power distribution module; The attitude and track control module is connected to a plurality of attitude sensors; the plurality of attitude sensors are configured to collect multi-modal attitude signals, and send the attitude signals to the attitude and track control module; the attitude and track control module is configured to generate an adjustment instruction according to the current task stage and the attitude signal in a set task mode, and send the adjustment instruction to an adjustment target device, wherein the task mode includes a plurality of task stages divided according to the running time and / or running attitude; the adjustment target device is a device for adjusting the running attitude determined according to the current task stage and the attitude signal; The satellite service module is configured to generate a device control instruction according to the current mission stage, and the device control instruction is used to start or shut down the associated device of the current mission stage; the power allocation module is configured to generate a power allocation instruction according to power demand information, and the power demand information includes one or more combinations of the current mission stage, the adjustment instruction and the device control instruction; the power allocation instruction is used to allocate power supply to the associated device of the current mission stage and / or the adjustment target device; A communication subsystem, connecting the energy subsystem and the integrated electronic subsystem; A payload subsystem is connected to the energy subsystem and the integrated electronic subsystem; the payload subsystem includes a multi-spectral payload camera, and the multi-spectral payload camera is configured to collect remote sensing data according to the current mission stage and send the remote sensing data to the integrated electronic subsystem.

2. The system according to claim 1, characterized in that The system further comprises: The structural subsystem includes a load-bearing mechanism, an unfolding assembly, and a separating assembly; the load-bearing mechanism is a frame structure formed by connecting a first load-bearing member and a second load-bearing member through a plurality of brackets; The unfolding assembly includes a flap and an unfolding hinge, the power source includes a solar panel, and the solar panel is arranged on the flap; the flap is rotatably connected to both sides of the load-bearing mechanism through the unfolding hinge; The separation assembly includes a separation mechanism and a separation switch. The separation mechanism is a plug-in device that uses a memory alloy to achieve unlocking. The separation mechanism is connected to the power supply through the separation switch. The separation switch is connected to the satellite service module. The satellite service module is further configured as follows: Generate separation instructions according to the set task mode; The separation instruction is sent to the separation switch, and the separation instruction is used to control the separation switch to close, so that the separation mechanism is energized to perform a separation action.

3. The system according to claim 2, characterized in that The system further comprises: The thermal control subsystem includes a temperature control component and a heat preservation component; the heat preservation component is arranged on the load-bearing mechanism, and the temperature control component is arranged in the internal space of the frame structure of the load-bearing mechanism; the load camera is temperature-insulated by the heat preservation component; The onboard computer is also equipped with a thermal control module; the temperature control assembly includes a thin film temperature control sheet and a temperature sensor; the temperature sensor and the thin film heating sheet are connected to the thermal control module; the thermal control module is also configured as follows: receiving the temperature inside the satellite detected by the temperature sensor; When the temperature inside the star is not within a preset operating temperature range, a temperature control instruction is sent to the thin film temperature control sheet, wherein the temperature control instruction is used to control the thin film temperature control sheet to generate an adjusted temperature rise.

4. The system according to claim 1, characterized in that The system further comprises: The electrical subsystem includes a cable assembly and a debugging separation switch; the cable assembly is an independent module integrating the cable and the connector in the satellite; the debugging separation switch is arranged on the separation assembly and is configured to detect the separation state of the separation assembly; the debugging separation switch is connected to the satellite service module, and the satellite service module is further configured to: Receiving separation state information detected by the debugging separation switch; If the separation status information is the first information, the next task of the separation task is executed according to the set task mode, and the first information is used to indicate that the separation component has executed the separation action; If the separation status information is the second information, the separation task is repeatedly executed according to the set task mode, and the second information is used to indicate that the separation component has not performed the separation action.

5. The system according to claim 1, characterized in that The system further comprises: A propulsion subsystem, comprising a vacuum arc thruster; the vacuum arc thruster is connected to the power supply through the power distribution module; the power distribution module is further configured to: Obtaining the running track information within the preset detection period; Comparing the operation track information with the planned operation path to obtain deviation data; Setting working parameters of the vacuum arc thruster according to the deviation data, wherein the working parameters include thrust power, thrust time, thrust direction, and propellant ejection amount; The operating parameters are sent to the vacuum arc thruster to control the vacuum arc thruster to perform thrusting according to the operating parameters.

6. The system according to claim 1, characterized in that The plurality of attitude sensors include a navigation satellite assembly, an angular velocity sensor, and a magnetic sensor; the attitude and trajectory control module is further configured as follows: parsing a spatial positioning signal, an angular velocity signal and a magnetic signal from the multimodal attitude signal detected by the attitude sensor, wherein the spatial positioning signal is a signal detected by the navigation satellite component; the angular velocity signal is a signal detected by the angular velocity sensor; and the magnetic signal is a signal detected by the magnetic sensor; Calculate current posture information according to the spatial positioning signal, the angular velocity signal and the magnetic signal; According to the set task mode, an adjustment instruction is generated according to the current posture information, wherein the adjustment instruction includes an adjustment target device and an adjustment parameter, and the adjustment target device is a posture adjustment device associated with any one of the spatial positioning signal, the angular velocity signal, and the magnetic signal; The adjustment instruction is sent to the adjustment target device so that the adjustment target device operates according to the adjustment parameters.

7. The system according to claim 6, characterized in that The navigation satellite assembly includes a front radio frequency processing unit, a high dynamic navigation unit, a microprocessor and an interface unit which are electrically connected in sequence; The front radio frequency processing unit is connected to the navigation antenna of the communication subsystem; The front radio frequency processing unit is configured to receive satellite signals through the navigation antenna, perform filtering and amplification processing on the satellite signals, and transmit the processed satellite signals to the high dynamic navigation unit; The high dynamic navigation unit is configured to perform signal conversion on the satellite signal according to a signal conversion item, and to resolve navigation data from the converted signal; the signal conversion item includes down-conversion, automatic gain control, and analog-to-digital conversion; The microprocessor is configured to receive the navigation data from the high dynamic navigation unit, perform orbit determination calculation on the navigation data, obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit, wherein the navigation parameters include positioning data, orbit determination data and time information.

8. The system according to claim 6, characterized in that The integrated electronic subsystem further includes a star sensor, which includes an optical lens and an image processing unit. The optical lens is configured to collect space optical signals and send the space optical signals to the image processing unit. The image processing unit has a built-in field programmable gate array and is configured to convert the optical signals into celestial body positioning images and send the celestial body positioning images to the attitude and orbit control module. The attitude and orbit control module is further configured to: Receiving an attitude angle signal detected by the angular velocity sensor; Calculate three-axis attitude information according to the standard attitude angle signal and the celestial body positioning image; Get the standard posture information of the current task mode; Calculating the posture difference between the standard posture information and the three-axis posture information; If the posture difference is greater than or equal to the posture difference threshold, the adjustment instruction is generated.

9. The system according to claim 6, characterized in that The integrated electronic subsystem further includes a sun-sensitive sensor, which is connected to the attitude and orbit control module; the sun-sensitive sensor is configured to detect sun position information using a four-quadrant differential working principle, and send the sun position information to the attitude and orbit control module; the attitude and orbit control module is further configured to: Acquiring a magnetic signal detected by the magnetic sensor; Determine the geocentric orientation information according to the magnetic signal; Performing dual-vector attitude determination according to the solar position information and the geocentric position information to generate attitude information; The adjustment instruction is generated based on the posture information.

10. A satellite system operation control method, characterized in that: Applied to the intelligent multi-spectral payload satellite system according to any one of claims 1 to 9, the method comprising: Acquire a multimodal posture signal and set a current task stage in a task mode, wherein the task mode includes a plurality of task stages divided according to an operating time and / or an operating posture; Generate an adjustment instruction according to the current task stage and the posture signal, the adjustment instruction including an adjustment parameter and an adjustment target device, the adjustment parameter is set according to the posture signal; the adjustment target device is a device for adjusting the operating posture determined according to the current task stage and the posture signal; Finding the associated components of the current task stage; Generate a device control instruction according to the current task stage, wherein the device control instruction is used to start or shut down an associated device of the current task stage; Generate a power allocation instruction according to the power demand information, wherein the power demand information includes one or more combinations of the current task stage, the adjustment instruction, and the device control instruction; the power allocation instruction is used to allocate power supply to the associated device of the current task stage and the adjustment target device; The adjustment instruction and the power allocation instruction are sent to the adjustment target device, and the device control instruction and the power allocation instruction are sent to the associated device of the current task stage.

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