Intelligent multispectral payload satellite system and its operation control method

Abstract

This application provides an intelligent multispectral payload satellite system and its operation control method. The system includes an energy subsystem, an integrated electronic subsystem, a communication subsystem, and a payload subsystem. The onboard computer of the integrated electronic subsystem is equipped with an attitude and orbit control module, a satellite service module, and a power allocation module. The attitude and orbit control module generates adjustment commands based on the current mission stage and attitude signals acquired by sensors under a set mission mode. Simultaneously, the satellite service module generates device control commands according to the current mission stage. The power allocation module generates power allocation commands based on the current mission stage, adjustment commands, and device control commands. The system can control the satellite system comprehensively based on the current mission stage and attitude signals through adjustment commands, device control commands, and power allocation commands, thereby solving the problem of satellite systems not fully utilizing hardware resources.

Description

Technical Field

[0001] This application relates to the field of aerospace control technology, and in particular to an intelligent multispectral payload satellite system and its operation control method. Background Technology

[0002] A satellite system is a system that utilizes artificial satellites operating in Earth orbit to perform various functions in conjunction with ground-based equipment. Depending on the function being performed, a satellite system can carry corresponding satellite payloads. Satellite payloads broadly refer to the various devices and instruments installed on a satellite for performing specific tasks and collecting information from Earth or outer space.

[0003] Taking remote sensing satellites as an example, remote sensing satellite systems can acquire surface spectral information by carrying remote sensing information acquisition sensors such as spectral payload cameras to conduct optical monitoring of the Earth's surface. These spectral payload cameras and other remote sensing information acquisition sensors can have multiple operating bands to acquire spectral information at different wavelengths, enabling various applications such as monitoring surface vegetation, fire monitoring, and atmospheric composition monitoring.

[0004] During operation, satellite systems exhibit various operational attitudes and device activation / deactivation states according to pre-defined mission modes. Correspondingly, these different operational attitudes and device activation / deactivation states result in varying data processing capacity requirements and power consumption. However, since satellites cannot undergo hardware upgrades after launch into the space environment, the satellite system needs to reserve significant data processing capacity to cope with different operational states. Consequently, the satellite system cannot fully utilize its hardware resources, hindering miniaturization. Summary of the Invention

[0005] In view of this, embodiments of this application provide an intelligent multispectral payload satellite system and its operation control method to solve the problem that satellite systems cannot fully utilize hardware resources.

[0006] According to one aspect of this application, a smart multispectral payload satellite system is provided, the system comprising:

[0007] The energy subsystem is configured to provide energy input;

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

[0009] The attitude control module is connected to multiple attitude sensors; the multiple attitude sensors are configured to acquire multimodal attitude signals and send the attitude signals to the attitude control module; the attitude control module is configured to generate adjustment instructions based on the current task stage in a set task mode and the attitude signals, and send the adjustment instructions to the adjustment target device; the task mode includes multiple task stages divided according to running time and / or running posture; the adjustment target device is a device determined based on the current task stage and the attitude signals for adjusting the running posture.

[0010] The satellite service module is configured to generate device control instructions according to the current mission phase. The device control instructions are used to start or stop associated devices in the current mission phase. The power allocation module is configured to generate power allocation instructions based on power demand information, which includes one or more combinations of the current mission phase, the adjustment instructions, and the device control instructions. The power allocation instructions are used to allocate power to associated devices and / or the adjustment target devices in the current mission phase.

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

[0012] The payload subsystem connects the energy subsystem and the integrated electronic subsystem; the payload subsystem includes a multispectral payload camera configured to acquire remote sensing data according to the current mission phase and to transmit the remote sensing data to the integrated electronic subsystem.

[0013] Optionally, the system further includes:

[0014] The structural subsystem includes a load-bearing mechanism, a deployment assembly, and a separation assembly; the load-bearing mechanism is a frame structure formed by connecting a first load-bearing member and a second load-bearing member through multiple supports;

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

[0016] The separation assembly includes a separation mechanism and a separation switch. The separation mechanism is a plug-in connector that uses a shape memory alloy for unlocking. The separation mechanism is connected to the power supply via the separation switch. The separation switch is connected to the space service module, which is further configured to:

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

[0018] The separation command is sent to the separation switch, which controls the separation switch to close, so that the separation mechanism is energized to perform the separation action.

[0019] Optionally, the system further includes:

[0020] The thermal control subsystem includes a temperature control component and a thermal insulation component; the thermal insulation component is installed on the load-bearing mechanism, and the temperature control component is installed in the internal space of the frame structure of the load-bearing mechanism; the load camera is temperature isolated through the thermal insulation component.

[0021] 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 further configured to:

[0022] Receive the internal temperature of the satellite detected by the temperature sensor;

[0023] When the temperature inside the satellite is not within the preset operating temperature range, a temperature control command is sent to the thin-film temperature control sheet. The temperature control command is used to control the thin-film temperature control sheet to generate an adjustable temperature rise.

[0024] Optionally, the system further includes:

[0025] The electrical subsystem includes a cable assembly and a test disconnect switch; the cable assembly is an independent module integrating onboard cables and connectors; the test disconnect switch is mounted on the disconnect assembly and configured to detect the disconnection status of the disconnect assembly; the test disconnect switch is connected to the satellite service module, which is further configured to:

[0026] Receive the separation status 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. The first information is used to indicate that the separation component has performed the separation action.

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

[0029] Optionally, the system further includes:

[0030] The propulsion subsystem includes a vacuum arc thruster; the vacuum arc thruster is connected to the power supply via the power distribution module; the power distribution module is further configured to:

[0031] Obtain the running trajectory information within a preset detection period;

[0032] The deviation data is obtained by comparing the operational trajectory information with the planned operational path.

[0033] The operating parameters of the vacuum arc thruster are set according to the deviation data, including thrust power, thrust time, thrust direction and propellant ejection rate;

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

[0035] Optionally, the plurality of attitude sensors include navigation satellite components, angular velocity sensors, and magnetic sensors; the attitude and orbit control module is further configured to:

[0036] Spatial positioning signals, angular velocity signals, and magnetic force signals are extracted from the multimodal attitude signals detected by the attitude sensors. The spatial positioning signals are signals detected by the navigation satellite components; the angular velocity signals are signals detected by the angular velocity sensors; and the magnetic force signals are signals detected by the magnetic force sensors.

[0037] Calculate the current attitude information based on the spatial positioning signal, the angular velocity signal, and the magnetic force signal;

[0038] According to the set task mode, an adjustment command is generated based on the current attitude information. The adjustment command includes the adjustment target device and the adjustment parameter. The adjustment target device is any one of the attitude adjustment devices associated with the spatial positioning signal, the angular velocity signal and the magnetic force signal.

[0039] The adjustment command is sent to the target device to cause the target device to operate according to the adjustment parameters.

[0040] Optionally, the navigation satellite assembly includes a front-end radio frequency processing unit, a high dynamic navigation unit, a microprocessor, and an interface unit that are connected in sequence.

[0041] The front-end radio frequency processing unit is connected to the navigation antenna of the communication subsystem; the front-end 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 signal conversion terms, and to compute navigation data from the converted signal; the signal conversion terms include 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 calculations on the navigation data to obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit. The navigation parameters include positioning data, orbit determination data, and time information.

[0044] Optionally, 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 acquire space optical signals and transmit 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 positioning images and transmit the celestial positioning images to the attitude and orbit control module. The attitude and orbit control module is further configured to:

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

[0046] Calculate the three-axis attitude information based on the attitude angle signal and the celestial positioning image;

[0047] Obtain the standard pose information for the current task mode;

[0048] Calculate the attitude difference between the standard attitude information and the three-axis attitude information;

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

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

[0051] Acquire the magnetic signal detected by the magnetic sensor;

[0052] The geocentric orientation information is determined based on the magnetic signal;

[0053] Perform dual-vector attitude determination based on the solar azimuth information and the geocentric azimuth information to generate attitude information;

[0054] The adjustment command is generated based on the attitude information.

[0055] According to another aspect of this application, a satellite system operation control method is provided, applied to the aforementioned intelligent multispectral payload satellite system, the method comprising:

[0056] Acquire multimodal attitude signals and set the current task stage in the task mode, wherein the task mode includes multiple task stages divided according to running time and / or running attitude;

[0057] An adjustment command is generated based on the current task stage and the attitude signal. The adjustment command includes adjustment parameters and an adjustment target device. The adjustment parameters are set according to the attitude signal. The adjustment target device is a device used to adjust the operating attitude, determined according to the current task stage and the attitude signal.

[0058] Locate the associated devices for the current task phase;

[0059] Device control instructions are generated according to the current task stage, and the device control instructions are used to start or stop the associated devices in the current task stage;

[0060] A power allocation instruction is generated based on power demand information, which 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 to the associated devices and the adjustment target devices in the current task stage.

[0061] The adjustment command and the power allocation command are sent to the target device for adjustment, and the device control command and the power allocation command are sent to the associated device in the current task phase.

[0062] By employing the above technical solutions, this application provides an intelligent multispectral payload satellite system and its operation control method. The system includes an energy subsystem, an integrated electronic subsystem, a communication subsystem, and a payload subsystem. The onboard computer of the integrated electronic subsystem is equipped with an attitude and orbit control module, a satellite service module, and a power allocation module. The attitude and orbit control module generates adjustment commands based on the current mission stage and attitude signals acquired by sensors under a set mission mode. Simultaneously, the satellite service module generates device control commands according to the current mission stage. The power allocation module generates power allocation commands based on the current mission stage, adjustment commands, and device control commands. The system can control the satellite system comprehensively based on the current mission stage and attitude signals through adjustment commands, device control commands, and power allocation commands, thereby solving the problem of satellite systems not fully utilizing hardware resources.

[0063] The above description is only an overview of the technical solution of this application. In order to better understand the technical means of this application and to implement it in accordance with the contents of the specification, and to make the above and other objects, features and advantages of this application more obvious and understandable, the following are specific embodiments of this application. Attached Figure Description

[0064] The accompanying drawings, which are included to provide a further understanding of this application and form part of this application, illustrate exemplary embodiments and are used to explain this application, but do not constitute an undue limitation of this application. In the drawings:

[0065] Figure 1 A schematic diagram of an intelligent multispectral payload satellite system provided in this application embodiment;

[0066] Figure 2 This is a schematic diagram of the intelligent multispectral payload satellite system architecture provided in the embodiments of this application;

[0067] Figure 3 This is a schematic diagram of the integrated electronic subsystem structure provided in an embodiment of this application;

[0068] Figure 4 This is a schematic diagram of the navigation satellite component structure provided in an embodiment of this application;

[0069] Figure 5 This is a schematic diagram of the attitude control process provided in an embodiment of this application;

[0070] Figure 6 This is a schematic diagram of the attitude control principle provided in an embodiment of this application;

[0071] Figure 7 This is a schematic diagram of the structural subsystem provided in an embodiment of this application;

[0072] Figure 8 This is a schematic diagram of the process for generating adjustment instructions based on celestial positioning images, provided in an embodiment of this application.

[0073] Figure 9 This is a schematic diagram illustrating the process of generating adjustment instructions based on solar azimuth information, provided in an embodiment of this application. Detailed Implementation

[0074] The present application will be described in detail below with reference to the accompanying drawings and embodiments. It should be noted that, unless otherwise specified, the embodiments and features described in the embodiments of the present application can be combined with each other.

[0075] In this embodiment, the satellite system refers to a system that utilizes an artificial satellite operating in Earth orbit to perform various functions in conjunction with ground equipment. Depending on the function being performed, the satellite system can carry a satellite payload corresponding to that function. Here, satellite payload broadly refers to various devices and instruments installed on a satellite for performing specific tasks and collecting information from Earth or outer space.

[0076] In some embodiments, a satellite system can acquire surface spectral information by carrying remote sensing sensors such as spectral payload cameras to perform optical monitoring of the Earth's surface. These spectral payload cameras and other remote sensing sensors can have multiple operating bands to acquire spectral information at different wavelengths.

[0077] During operation, satellite systems exhibit various operational attitudes and device activation / deactivation states according to their pre-defined mission modes. Correspondingly, these different operational attitudes and device activation / deactivation states result in varying data processing capacity requirements and power consumption. However, since satellites cannot undergo hardware upgrades after launch into the space environment, the satellite system needs to reserve significant data processing capacity to cope with different operational states. Consequently, the satellite system cannot fully utilize its hardware resources.

[0078] To fully utilize the hardware resources of the satellite system, this embodiment provides an intelligent multispectral payload satellite system, such as... Figure 1 As shown, the intelligent multispectral payload satellite system may include: an energy subsystem, an integrated electronic subsystem, a communication subsystem, and a payload subsystem. To adapt to the satellite system's on-orbit missions, the intelligent multispectral payload satellite system may also include a propulsion subsystem, a structural subsystem, a thermal control subsystem, and an electrical subsystem.

[0079] The energy subsystem provides energy input to all equipment, mechanisms, and devices within 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 point tracking algorithm.

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

[0081] The integrated electronic subsystem connects 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, responsible for the satellite system's control and data processing. The on-board computer can perform functions such as spacecraft attitude control, orbit control, status monitoring, navigation and positioning, and fault diagnosis.

[0082] like Figure 3 As shown, the IOBC can provide a rich set of communication bus interfaces, such as CAN, RS232, RS422, RS485, IIC, SPI, etc. It can also provide three-phase brushless motor control, three-axis torque converter control, integrate Global Navigation Satellite System (GNSS), Inertial Measurement Unit (IMU), Magnetometer (MTM), and provide multiple payload power supply interfaces.

[0083] To control the operation of the satellite system, an attitude and orbit control module, a satellite service module, and a power allocation module can be integrated into the integrated electronic subsystem. That is, the onboard computer is equipped with these modules. The attitude and orbit control module is used to adjust the attitude of the satellite system based on its current operating status. For this purpose, the attitude and orbit control module is connected to multiple attitude sensors; these sensors are configured to acquire multimodal attitude signals and transmit these 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 magnetometer. The navigation satellite assembly is used to detect spatial positioning signals. For example, an onboard integrated modular GNSS can be used as the navigation satellite assembly. A navigation satellite assembly using an onboard integrated GNSS can achieve a positioning accuracy of 30m, a time accuracy of 1ms, and a velocity accuracy of 0.5m / s via UART communication, and can operate stably for more than 3 years in an operating temperature range of -40 to 85°C.

[0085] like Figure 4 As shown, in some embodiments, the navigation satellite assembly includes a front-end radio frequency processing unit, a high-dynamic navigation unit, a microprocessor, and an interface unit connected in sequence. The front-end radio frequency processing unit is connected to the navigation antenna of the communication subsystem; the front-end 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.

[0086] The high dynamic navigation unit is configured to perform signal conversion on the satellite signal according to signal conversion terms, and to compute navigation data from the converted signal; the signal conversion terms include 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 calculations on the navigation data to obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit. 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, inputting satellite signals to the preamplifier RF unit for filtering and amplification, and then outputting them to the high dynamic range navigation unit. This unit performs signal down-conversion, AGC, and AD conversion, as well as real-time acquisition, tracking, and positioning calculation of 12-channel GPS L1CA code navigation signals and 20-channel BDB1I code navigation signals. After orbit determination calculation by the MCU, the positioning, orbit determination data, and time information are finally output via serial port, along with a 1PPS time synchronization signal. The MCU also features abnormal state monitoring and handling, positioning data discrimination, and interface control functions.

[0089] The pre-amplifier radio frequency (RF) processing unit, which processes BD and GPS navigation RF signals, has certain requirements for receiver sensitivity, anti-interference capability, and reliability. Therefore, the RF processing unit can be selected from highly reliable devices that meet the requirements of the space environment and have undergone long-term flight verification. The RF processing unit is connected to the passive antenna and performs out-of-band filtering, low-noise amplification, and passive power division on the weak navigation signals received by the antenna, providing RF signals to the various high-dynamic receivers on the primary and backup receivers.

[0090] The high-dynamic navigation unit, also known as the high-dynamic baseband processing unit, is designed to ensure stability, reliability, and adaptability to low-Earth orbit (LEO) satellite applications. Each antenna is connected to an independent navigation module for positioning calculations, and the calculated results are evaluated before output to guarantee reliability. For example, to meet high-precision positioning requirements, the high-dynamic baseband processing unit can rapidly acquire radio frequency (RF) signals, perform mixing and correlation of digital intermediate frequency (IF) signals, output the integral and accumulated results for calculation, perform loop processing of the signal, complete the positioning calculation, output the positioning results and raw observation information to the LEO interface processing unit, and simultaneously generate 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 the angular velocity sensor. The MEMS gyroscope can be integrated on a board with dual gyroscope backup, communicate via SPI, achieve zero-bias stability of 0.8° / h, a measurement range better than ±50°, a measurement accuracy of 0.01° / s, and can operate stably for more than 3 years in an operating temperature range of -40 to 85°C.

[0092] The magnetic sensor is used to detect magnetic signals. It is a magnetic sensor component that can also be deployed on-board via an integrated MTM and communicates via SPI. The magnetic sensor has a measurement range of ±8 gauss and an accuracy of 1%. It also needs to operate stably for more than 3 years within an operating temperature range of -40 to 85°C.

[0093] In some embodiments, to further miniaturize the satellite system, sensors used for attitude and orbit control, such as gyroscopes, magnetometers, GNSS sensors, torque converters, and flywheel controllers, can be integrated into the IOBC. Simultaneously, the IOBC can directly power the star sensor and sun sensor and acquire their data via a signal bus. The attitude control algorithm is also executed in the IOBC's main control chip, i.e., Figure 5 As shown, the attitude control module is further configured as follows:

[0094] S101. Generate adjustment instructions based on the current task stage and the attitude signal in the set task mode.

[0095] The mission mode includes multiple mission phases divided according to operating time and / or operating attitude. For example, the mission mode can be set with multiple mission phases based on the satellite's post-launch operation, including separation phase, initialization phase, self-test phase, entry phase, telemetry and control phase, and exit phase. In different mission phases, the satellite system needs to activate or deactivate some components according to the requirements of the current mission phase.

[0096] In some embodiments, devices that require a change in operating state during a certain task phase can be referred to as associated devices for that task phase. Changing the operating state refers to adjusting the device to adapt to the requirements of that task phase upon entering that phase. Operating state changes can include changing from not started to started, from started to not started, and switching from standby to operating state. For example, upon entering the initialization phase, the satellite system powers on. Besides the IOBC main chip and battery management component operating, the satellite system deploys its solar panels and begins initializing the various system components. In this case, the associated devices during the initialization phase can include the IOBC main chip, battery management component, solar panels, and the system components being initialized.

[0097] Because different mission phases have different requirements for satellite attitude, attitude adjustment requires obtaining the current mission phase within the set mission mode. The current mission phase can be determined based on the current time, current operating position (whether it has entered orbit), and the current attitude signals detected by attitude sensors. For example, according to the on-orbit mission mode setting, if the satellite system performs telemetry, tracking, and command (TT&C) within the time interval [T1, T2], then the current time T can be obtained. If T is within the time interval [T1, T2], then the current mission phase can be determined as the TT&C phase.

[0098] In some embodiments, when multiple attitude sensors include navigation satellite components, angular velocity sensors, and magnetic sensors, the attitude and orbit control module can parse space positioning signals, angular velocity signals, and magnetic signals from the multimodal attitude signals detected by the attitude sensors. It then calculates the current attitude information based on the space positioning signals, angular velocity signals, and magnetic signals. Next, according to a set mission mode, it generates adjustment instructions based on the current attitude information. These adjustment instructions include a target device for adjustment and adjustment parameters. The target device is a device determined based on the current mission stage and the attitude signals for adjusting the operating attitude; specifically, the target device is any one of the associated attitude adjustment devices among the space positioning signals, angular velocity signals, and magnetic signals.

[0099] S102. Send the adjustment command to the adjustment target device.

[0100] After generating the adjustment command, the attitude control module can send the adjustment command to the target device so that the target device operates according to the adjustment parameters. For example, such as Figure 6 As shown, once the satellite system enters the on-orbit measurement phase, its current operating attitude P can be determined using attitude signals. If the operating attitude does not meet the attitude requirements P0 of the on-orbit measurement phase, the target device for attitude adjustment, momentum wheels, can be identified based on the deviation ΔP between the current operating attitude P and the attitude P0 of the on-orbit measurement phase. Adjustment parameters ΔP are then generated, i.e., adjustment commands are generated. These adjustment commands are then sent to the momentum wheels, enabling them to perform attitude control.

[0101] Among them, the momentum wheel assembly is a device that performs attitude control through momentum calculation. For example, the momentum wheel assembly can be integrated into a single unit using a three-orthogonal-one-skewed mounting configuration, based on the FOC control algorithm, to perform attitude control of a satellite system using a speed control method. To meet attitude control requirements, the momentum wheel assembly can operate within a speed range of -6000 to 6000 RPM. It boasts an angular momentum better than 8 mN / ms and a torque of 2 mN / m, enabling it to operate for more than 3 years at a working temperature range of -40 to 85°C.

[0102] The satellite service module is responsible for satellite time management, on-orbit mission execution, overall satellite status management and monitoring, and deployment of component actions. 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 as a finite state machine, meaning that the satellite system performs specific actions in a specific state, such as turning on and off individual power supplies, turning on and off telemetry and data transmission, and executing missions.

[0103] For example, the satellite management module can perform time management, storage management, subsystem management, flight attitude management, thermal control management, operating mode control, satellite power consumption management, and satellite status monitoring and maintenance. Time management involves determining and managing the onboard time reference based on GNSS module time information or ground-based time data. It can also be used for calibration based on GNSS time in the event of an unexpected and prolonged power outage.

[0104] Storage management is a feature of the satellite mission management system, which incorporates various storage spaces. The satellite mission management application includes storage management functions to uniformly manage non-volatile data such as configuration information and task queues. Subsystem management refers to the satellite mission module controlling subsystems according to the flight mission design to achieve preset flight missions. In conjunction with the payload adaptation unit, it controls and manages onboard payloads to realize preset satellite functions.

[0105] Flight attitude management involves the onboard computer acting as the controller of the attitude control subsystem, controlling the satellite's attitude according to different operating modes and attitude control requirements. Thermal management involves managing the thermal temperature of key nodes within the entire satellite based on the operating temperature requirements of cameras and batteries, ensuring that critical components operate within the designed temperature range.

[0106] Operating mode control involves designing different operating modes based on the different operational phases of the satellite to achieve its flight mission. Satellite power management involves real-time monitoring of the satellite's overall energy during flight. When the satellite faces the risk of battery over-discharge, energy control is implemented to put the entire satellite into a safe mode, shutting down high-energy-consuming equipment to ensure the stability and normal operation of the satellite's energy system.

[0107] Satellite status monitoring and maintenance involves periodically collecting and monitoring the overall status of the satellite during its operation, diagnosing and handling foreseeable faults, and automatically handling faults when abnormal conditions occur in satellite operation to ensure the stability of satellite operation.

[0108] To execute corresponding mission control at different mission phases, the satellite service module is configured to generate device control commands according to the current mission phase. These device control commands are used to activate or deactivate associated devices in the current mission phase. For example, when the satellite system enters the initialization mission phase, the disconnect switch needs to be in the disconnected state, and the system needs to be powered on. In addition to the IOBC main chip and battery management components operating, the system operates the solar panels to deploy and begins initializing various system components. At this time, the associated devices in the current mission phase are the solar panels and the various system components that need to be initialized. Therefore, control commands can be generated for deploying the solar panels and initializing the various system components. These control commands are then sent to the solar panels and the various system components that need to be initialized to execute the system initialization task.

[0109] The power distribution module is used to provide power to various parts of the satellite system and dynamically adjusts the power supply to each part according to the overall power distribution of the current system. For example, after being configured with the IOBC, the power distribution module can provide power to the communication system and the payload camera, while reserving 12V and 5V outputs to support future expansion applications.

[0110] Therefore, the power allocation module is configured to generate a power allocation instruction based on power demand information, which 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 to the associated device and / or the adjustment target device in the current task stage.

[0111] After launch, the satellite system relies heavily on its energy subsystem, specifically the solar arrays and lithium-ion battery packs, for power. Since the total power output from these components is limited, the power distribution module generates power allocation commands based on power demand information during satellite operation.

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

[0113] For example, if the measurement mission phase of a satellite system involves spectral acquisition of a specific area, then upon entering the measurement mission phase, a power allocation command can be sent to the multispectral payload camera in the payload subsystem, thereby controlling the multispectral payload camera to start operation.

[0114] In some embodiments, power allocation commands can also be generated based on adjustment commands. That is, when the satellite system needs to perform attitude adjustments, since these adjustments require actuators such as momentum wheel sets and three-axis magnetic rods to generate adjustment actions, after generating the adjustment command, the power allocation module can read the target device in the adjustment command and generate a power allocation command for that target device. The power allocation command is then sent to the target device, thereby enabling it to receive power and perform the adjustment action.

[0115] In some embodiments, power allocation instructions can also be generated based on device control instructions. That is, when the satellite system is operating according to a predetermined mission mode, after the satellite service module generates device control instructions based on the current mission phase, the power allocation module can read the associated devices for the current mission phase from the device control instructions and generate power allocation instructions for the associated devices in the current mission phase. The power allocation instructions are then sent to the associated devices in the current mission phase, enabling the associated devices in the current mission phase to receive power 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, sending data to or receiving data from the ground monitoring system. For example, if the satellite orbital 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 approximately 2100km, then for the downlink, when performing ±50° telemetry missions, 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, with a spatial loss of 177.37dB, an atmospheric loss of 3.5dB, an antenna pointing loss of 0.5dB, a total received signal power level of -185.28dBW at the ground station, a ground station G / T value of 24dB / k, a Boltzmann constant of -228.6dBW / HzK, a received C / N0 of 67.32dBHz, a data rate of 8192bps, a received Eb / N0 of 28.18dB (the theoretical value of Eb / N0 is 10.6dB), a coding gain of 0dB, and a demodulation loss of 3dB. The communication subsystem can include an X-band communication unit, also known as an X-band telemetry and data transmission integrated unit, used to establish a satellite-to-ground wireless communication link. The X-band communication unit can achieve a communication capability of 8192bps rate, 20Mbps data transmission, and 32GB capacity.

[0117] The X-band telemetry, tracking, and command (TT&C) integrated system can include a data transmitter, a TT&C receiver, and a TT&C transmitter. The data transmitter can receive and store data such as satellite payloads and satellite operations. When the satellite passes through a station, the stored data is played back and processed by AOS framing, scrambling, and encoding. After processing, it is microwave modulated and fed down to the ground data receiving station through microwave channel circuits such as RF amplifiers.

[0118] The telemetry, tracking, and command (TT&C) receiver can receive uplink telemetry data and uploaded update program data from ground stations. When receiving uplink data, it performs real-time demodulation in orbit, and performs channel encoding format decoding and data identification processing. The received data is then appropriately stored or directly sent to the onboard computer for further data processing. The TT&C transmitter can receive telemetry data from the onboard computer in real time and transmit the telemetry data from the satellite during station passes.

[0119] To meet the communication requirements of satellite systems, the X-band telemetry, tracking, and command (TT&C) integrated unit can achieve payload communication via Cameralink using CAN control, with a storage capacity of 32GB. It can exhibit different performance characteristics in different mission modes. For example, in remote control mode, the communication system has a sensitivity of 106 dBm and a frequency range of 7242 MHz to 7250 MHz. Using BPSK modulation, it achieves a data rate of 8192 bps with a bit error rate of less than 1 × 10⁻⁶. -6 Similarly, in telemetry mode, with a power of 22 dBm and a frequency of 8325 MHz to 8375 MHz, BPSK modulation achieves a data rate of 8192 bps with a bit error rate of less than 1 × 10⁻⁶. -6 In data transmission mode, the power is 30 dBm, and the frequency is 8325 MHz to 8375 MHz. Using LDPC modulation in a 7 / 8 configuration, a data rate of 5 Mbps to 20 Mbps and a bandwidth of less than 1 × 10⁻⁶ Mbps can be achieved. -6 The bit error rate.

[0120] The communication subsystem may also include antenna assemblies, which can include receiving antennas (uplink antennas) and transmitting antennas (downlink antennas). For example, based on the overall requirements of the satellite system, antenna specifications, system links, and space environment requirements are comprehensively evaluated. The receiving antenna can be a dual-layer microstrip antenna. The dual-layer microstrip antenna has a gain of 9dBi, ±60°, and a build-up ratio ≤1.5. It must operate stably for more than 3 years in an operating temperature range of -90 to 90°C. The transmitting antenna can be a 2×2 array microstrip antenna, with time-division multiplexing of telemetry and data transmission using the X-band telemetry and data transmission integrated machine. The transmitting antenna has a gain of 0dBi, ±50°, and a build-up ratio ≤1.5, and also needs to operate stably for more than 3 years in an operating temperature range 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 communication unit shown above, the communication subsystem may also include a UV-band communication unit for establishing a satellite-to-ground wireless communication link. The communication subsystem is also connected to the energy subsystem through the power distribution module, thereby obtaining power supply from the energy subsystem during the communication process through the power distribution module.

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

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

[0124] The multispectral payload camera is configured to acquire remote sensing data according to the current mission phase and to transmit the remote sensing data to the integrated electronic subsystem. For example, according to a predetermined mission mode, when the satellite system enters the measurement phase, the satellite service module can generate device control commands based on the current mission phase. These device control commands can specify the associated device for the current mission phase as the payload camera, thereby sending the device control commands to the payload camera. Upon receiving the device control commands, the payload camera can respond by performing spectral acquisition. During spectral acquisition, remote sensing spectral data can be generated along with the monitoring process and transmitted to the integrated electronic subsystem. The integrated electronic subsystem can process the spectral data and transmit it to the ground monitoring system via the communication subsystem.

[0125] As can be seen, in the above embodiments, the system can control the satellite system by combining the current mission stage and attitude signals through adjustment commands, device control commands, and power allocation commands, so as to solve the problem that the satellite system cannot fully utilize hardware resources.

[0126] like Figure 7As shown, in some embodiments, the system further includes a structural subsystem. The structural subsystem involves mechanical structures such as support structures, stacks, flywheels, star sensors, single-unit brackets, skin, solar panels, panel deployment mechanisms, antennas, and cameras. The structural subsystem may include a load-bearing mechanism 1, a deployment assembly 2, and a separation assembly 3. 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 via multiple supports 13. For example, the first load-bearing member 11 and the second load-bearing member 12 are two plate-like structural members located on both sides of the satellite system. The first load-bearing member 11 and the second load-bearing member 12 are connected by multiple supports 13 to form a frame structure with an envelope size of 125mm(X)×230mm(Y)×420mm(Z).

[0127] A frame structure allows for the arrangement of space-facing and ground-facing components within a satellite system. For example, the space-facing component could include an X-band dipole antenna for telemetry and control, a sun sensor, and a GPS antenna on the top surface. Once deployed, the solar panels on the space-facing side are solar panels, designed as the primary surface receiving sunlight. The ground-facing component includes one X-band conical antenna, one X-band data transmission antenna, one camera, and two disconnect switches.

[0128] The first load-bearing component 11 serves as the main load-bearing structure of the satellite system. This main load-bearing structure is the design and assembly benchmark for the entire satellite structure. It has four sliding rail surfaces on each of its left and right sides, contacting the separation mechanism; these surfaces act as load-bearing surfaces during transportation and launch; and during assembly and measurement, the upper left side serves as the overall satellite reference origin. The first load-bearing component also has mounting holes and surfaces for various individual units and supports. The support mounting surfaces on the main load-bearing plate have locating pin holes for accurate positioning of intermediate supports. Openings on the main load-bearing plate are for expanding installation space, providing thermal control vias, and weight reduction features.

[0129] The second load-bearing component 12 can serve as a secondary load-bearing structure for the satellite system. The secondary load-bearing structure functions similarly to the main load-bearing plate, but it is not used as an installation or measurement reference. The intermediate support mounting surface on the secondary load-bearing plate is designed with positioning pin holes for positioning the intermediate support.

[0130] Support bracket 13 may include standard brackets and dedicated brackets. Standard brackets may specifically include single-pole brackets, single-pole brackets, double-pole brackets, and rectangular brackets. Standard brackets may have screw holes and locating pins for connection with the main and auxiliary load-bearing structures, and have mounting holes for the skin and cables on the sides. Common brackets, while connecting the main and auxiliary load-bearing plates, also serve as support brackets for various components. Some brackets can be used for mounting antennas such as tethered sensors, dipole antennas, and GPS antennas. Conical antenna brackets can be used to mount conical antennas. Data transmission antenna brackets are used to mount X-ray data transmission antennas. Depending on the supported target, common brackets may include magnetic torque generator brackets, star sensor brackets, flywheel brackets, etc., for mounting magnetic torque generators, star sensors, and flywheels, respectively.

[0131] The structural connection interfaces are divided into through holes and threaded holes. Screws pass through the through holes of the structural brackets and are screwed into the threaded holes of the connecting load-bearing plates to achieve the connection between the brackets and the load-bearing plates. The brackets are equipped with two locating pins at the top and bottom to ensure installation accuracy. Individual units inside the satellite are installed and connected to the brackets, which are then fixed by the brackets to the load-bearing plates. The individual unit brackets serve both as the mounting point for the individual unit and as a support within the satellite. The connection between the brackets and the load-bearing plates is divided into through holes and threaded holes. Screws pass through the through holes of the load-bearing plates and are screwed into the threaded holes of the connecting brackets to achieve the installation of individual units and structural connection.

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

[0133] For example, a satellite system may include two sets of solar panels, each with one solar cell. The two sets of panels are arranged side-by-side to form an energy subsystem containing two solar panels. The panels can be 178mm × 320.5mm in size, with double-sided fabric. The solar panels have two states: retracted and deployed, supported and secured by hinges and a wire-burning module. The solar panel substrate can use a solar panel PCB, with a thickness of 0.8mm and double-sided fabric. In space, the deployment hinge heats up by energizing a resistor in the wire-burning module, melting the tether controlling the solar panel's retraction, thus deploying the panel to the target position. A locking design can also be added to the deployment hinge for better control of the deployment process.

[0134] In some embodiments, the separation assembly includes a separation mechanism and a separation switch. The separation mechanism is a plug-in connector using shape memory alloy for unlocking; the separation mechanism is connected to the power supply via the separation switch. For example, the separation assembly uses a shape memory alloy plug-in for unlocking, with an electrical connector model of Y8C-4ZJBL and a corresponding power supply plug model of Y8C-4TK. The shape memory alloy plug-in has a total resistance of approximately 1.1Ω, making it a purely impedance load. The power supply design is for dual-point, dual-wire backup power supply, with a total current range of 2A to 30A. For single-use in-orbit operation, the required total power supply energy is 20–60J, the supply current is 10A, and the power supply time is 200ms.

[0135] When the supply current is less than 5A, the impact of heat dissipation must be considered, which can increase the input energy by more than 50%. The insulation resistance between the energized contacts of the shape memory alloy puller and the casing should not be less than 20MΩ. When multiple independently unlocking release seats are used to form a star-rocket connection release assembly, the number, specifications, and position of the release seats can be flexibly configured. The power supply requirements for each release seat are the same as those mentioned above, and multiple release seats can be connected in series or parallel according to the power supply capacity. To ensure the unlocking synchronization of multiple release seats, the current received by each release seat should be consistent and greater than 5A.

[0136] Furthermore, the separation assembly provides two channels of limit switch separation status measurement signals: closed before separation and open after separation. The device can select to measure one or both channels. The limit switch installation position is symmetrical relative to the center of the separation assembly. The limit switch closing resistance is <0.2Ω, and the opening resistance is ≥20MΩ. Each measurement signal consists of two stranded wires, approximately 200mm in length. Each wire has a CX-1A-5 pin and socket, which are compatible and interchangeable with the CX-1B-5 pins and sockets.

[0137] To perform the separation task, the separation switch is connected to the space service module. The space service module is also configured to generate a separation command according to a set mission mode and send the separation command to the separation switch. The separation command controls the separation switch to close, thereby energizing the separation mechanism 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, ensuring that the internal components of the satellite system operate within their designed operating temperature range. The thermal control subsystem primarily employs passive thermal control, supplemented by active thermal control. Active thermal control is achieved using thin-film heating elements, while passive thermal control can be implemented using materials such as heat-insulating materials. Therefore, the thermal control subsystem includes a temperature control component and a heat-insulating component. The heat-insulating component is mounted on the load-bearing mechanism, while the temperature control component is disposed within the internal space of the load-bearing mechanism's frame structure; the payload camera is temperature-isolated through the heat-insulating component.

[0139] For example, the thermal insulation design of the payload camera can be achieved by isolating it from the satellite platform using thermal pads, enabling the payload camera to have one temperature measurement channel and two temperature control channels, with heating compensation controlled by the IOBC (Integrated Electrical Control Unit). Furthermore, the battery components in the energy subsystem can also be designed with one heating circuit. By encasing the entire satellite in multiple layers and implementing heat dissipation designs for areas with high heat generation, thermal control of the satellite system can be achieved.

[0140] The onboard computer is also equipped with a thermal control module; the temperature control component includes a thin-film temperature control element and a temperature sensor; the temperature sensor and the thin-film heating element are connected to the thermal control module; the thermal control module is further configured to receive the onboard temperature detected by the temperature sensor, and when the onboard temperature is not within a preset operating temperature range, send a temperature control command to the thin-film temperature control element. The temperature control command is used to control the thin-film temperature control element to generate a 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 active heating compensation combined with passive thermal control can be used to control the internal temperature of the satellite between 10°C and 45°C.

[0142] In some embodiments, the system further includes an electrical subsystem. The electrical subsystem includes cable assemblies and a test disconnect switch; the cable assembly is an independent module integrating the satellite's internal cables and connectors. Integrating it as an independent module allows the electrical subsystem to function as a separate module, facilitating rapid cable connections and interface layout, and providing technical support for mass production. For example, the control interface between the rocket control system and the satellite is used to transmit satellite-rocket separation commands. The satellite receives a pulse signal via a shape memory alloy puller, triggering separation. The puller has two parallel bridge wires, each with a resistance of approximately 2.2Ω, resulting in a total resistance of approximately 1.1Ω. It is a purely impedance load, designed for dual-point, dual-wire backup power supply, with a total current range of 2A to 30A. When the puller is tested for ground electrical connection and unlocking with the launch vehicle equipment, the energizing energy should be controlled between 10J and 15J to ensure the puller's repeated service life. The insulation resistance between the energizing contacts and the outer casing of the shape memory alloy puller should be ≥20MΩ.

[0143] Therefore, the debugging separation switch is installed on the separation component and configured to detect the separation status of the separation component. The debugging separation switch is connected to the space service module. The space service module is also configured to receive separation status information detected by the debugging separation switch. If the separation status information is a first type of information, the next task of the separation task is executed according to a set task mode. The first type of information indicates that the separation component has performed a separation action; that is, after the separation switch detects the separation of the satellite and rocket, it can generate separation status information containing the first type of information. Similarly, if the separation status information is a second type of information, the separation task is repeated according to a set task mode. The second type of information indicates that the separation component has not performed a separation action.

[0144] By detecting separation status information through the separation switch, the separation task can be repeated even if the satellite and rocket have not separated, thereby improving the mission processing capability and reliability of the satellite system.

[0145] In some embodiments, the system further includes a propulsion subsystem. The propulsion subsystem can be used to perform trajectory changes or attitude adjustments. The propulsion subsystem may include a vacuum arc thruster (VAT); the vacuum arc thruster is connected to the power supply via the power distribution module.

[0146] For example, the working fluid in a vacuum arc thruster is a titanium alloy. A vacuum arc thruster is an electromagnetic propulsion system within an electric propulsion system, where the cathode material also serves as the propellant. Utilizing the principle of barrier dielectric discharge, an electric arc is instantaneously generated between the anode and cathode. The cathode is ablated by the arc, forming metal vapor, which is then ionized into plasma. The instantaneous high-current arc simultaneously generates a strong electromagnetic field, and the ionized plasma, under the influence of this electromagnetic field, is ejected axially to generate thrust.

[0147] Vacuum arc thrusters can combine the supply of non-toxic propellants with the thruster body into a single module, eliminating the need for complex propellant storage and supply systems. This makes them suitable for the space applications of microsatellites and for performing propulsion tasks requiring high control precision.

[0148] To perform the propulsion task, the power distribution module is further configured to first acquire the running trajectory information within a preset detection period, then compare the running trajectory information with the planned running path to obtain deviation data. The operating parameters of the vacuum arc thruster are then set based on the deviation data. These operating parameters include propulsion power, propulsion time, propulsion direction, and propellant ejection rate. By sending these operating parameters to the vacuum arc thruster, the module controls the vacuum arc thruster to perform propulsion according to these parameters.

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

[0150] The star sensor includes an optical lens and an image processing unit. For example, a star sensor consists of a light shield, a lens, and an image processing unit. To achieve system miniaturization, the electronics system uses a stacked assembly of two processing boards: a CMOS imaging system and an FPGA image processing and communication system. Based on the operating characteristics of the star sensor, aberrations affecting star points and measurement accuracy can be preferentially corrected based on the optical signal acquired by the optical lens, while other aberration tolerances can be appropriately relaxed to meet measurement accuracy requirements.

[0151] The lens hood design employs a finite element method (FEM) to recalculate the extinction capability based on design parameters. Initial values ​​for the structure and surface reflectivity are provided, and the extinction capability is recalculated using FEM. Design parameters are then adjusted until the extinction capability meets the requirements. The lens structure adopts an integrated optical, mechanical, and thermal design. Simulation and engineering analyses are used to determine the matching of optics, structure, and materials, as well as processing and assembly methods, to ensure that the total defocusing amount of the lens remains within the allowable measurement accuracy range over a certain temperature range. The circuitry corresponding to the image processing unit utilizes an ultra-simplified, highly reliable circuit design architecture, incorporating a high-precision design approach that combines rigidity and flexibility in the integration of core components. Radiation-resistant components and single-particle-resistant circuitry are incorporated into the system's aerospace-adaptive design principles to achieve the interface circuit board structure.

[0152] The optical lens is configured to acquire space optical signals and transmit 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 positioning images and transmit the celestial positioning images to the attitude and orbit control module. Correspondingly, as... Figure 8 As shown, the attitude control module is further configured as follows:

[0153] S201. Receive the attitude angle signal detected by the angular velocity sensor;

[0154] S202. Calculate the three-axis attitude information based on the target attitude angle signal and the celestial positioning image;

[0155] S203. Obtain the standard attitude information of the current task mode;

[0156] S204. Calculate the attitude difference between the standard attitude information and the three-axis attitude information;

[0157] S205. If the attitude difference is greater than or equal to the attitude difference threshold, generate the adjustment command.

[0158] As can be seen, the attitude control module can accurately calculate the three-axis attitude information by acquiring attitude angle signals and celestial positioning images. It then acquires the standard attitude information for the current mission mode and compares the three-axis attitude information with the standard attitude information to calculate the attitude difference between the two. When the attitude difference is greater than or equal to an attitude difference threshold, an adjustment command is generated based on the attitude difference to achieve attitude adjustment.

[0159] In some embodiments, the integrated electronic subsystem further includes a solar sensor. The solar sensor can be used to detect the sun's position, thereby enabling attitude control based on the sun's position or adapting to specific mission phases. For example, the satellite system's predetermined mission modes may include a solar capture mode, used to achieve functions such as rate damping after satellite-launch separation, solar capture (including pitch search and roll search), and sun-tracking cruise. The solar capture mode can serve as a fail-safe mode for transfer orbits and a worst-case safety mode for geostationary orbits. The solar capture mode can be activated after satellite-launch separation or in the event of a malfunction during satellite operation. In solar capture mode, rate damping of the satellite can be applied first, and sun search can begin once the angular rates of all axes are less than 0.5° / s and remain so for a period of time. In solar capture mode, the solar sensor can measure roll attitude and pitch angle information; when the solar sensor is directly facing the sun, the system switches to sun-tracking cruise mode.

[0160] The solar sensor is connected to the attitude and orbit control module; the solar sensor is configured to detect solar azimuth information using a four-quadrant differential operating principle and to send the solar azimuth information to the attitude and orbit control module. Correspondingly, as... Figure 9 As shown, the attitude control module is further configured as follows:

[0161] S301. Obtain the magnetic signal detected by the magnetic sensor;

[0162] S302. Determine the geocentric orientation information based on the magnetic signal;

[0163] S303. Perform dual-vector attitude determination based on the solar azimuth information and the geocentric azimuth information to generate attitude information;

[0164] S304. Generate the adjustment command based on the attitude information.

[0165] As can be seen, by using dual-vector attitude determination with the sun sensor and magnetometer, X1 and X2 can be replaced with the solar vector and geomagnetic vector respectively. These two vectors can be expressed in the inertial frame by orbit calculation, or in the satellite body by the sun sensor and magnetometer. Therefore, by using the above dual-vector attitude determination principle, the attitude conversion matrix from the local system to the inertial frame can be obtained, thereby achieving more accurate attitude adjustment.

[0166] Based on the intelligent multispectral payload satellite system provided in the above embodiments, some embodiments of this application also provide a satellite system operation control method, applied to the above-mentioned intelligent multispectral payload satellite system, the method comprising:

[0167] Acquire multimodal attitude signals and set the current task stage in the task mode, wherein the task mode includes multiple task stages divided according to running time and / or running attitude;

[0168] An adjustment command is generated based on the current task stage and the attitude signal. The adjustment command includes adjustment parameters and an adjustment target device. The adjustment parameters are set according to the attitude signal. The adjustment target device is a device used to adjust the operating attitude, determined according to the current task stage and the attitude signal.

[0169] Locate the associated devices for the current task phase;

[0170] Device control instructions are generated according to the current task stage, and the device control instructions are used to start or stop the associated devices in the current task stage;

[0171] A power allocation instruction is generated based on power demand information, which 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 to the associated devices and the adjustment target devices in the current task stage.

[0172] The adjustment command and the power allocation command are sent to the target device for adjustment, and the device control command and the power allocation command are sent to the associated device in the current task phase.

[0173] By applying the technical solution of this embodiment, the method can generate adjustment commands through the attitude and orbit control module of the intelligent multispectral payload satellite system, based on the current mission stage and attitude signals collected by sensors under a set mission mode. Simultaneously, the satellite service module can generate device control commands according to the current mission stage. The power allocation module then generates power allocation commands based on the current mission stage, adjustment commands, and device control commands. This method can comprehensively control the satellite system based on the current mission stage and attitude signals through adjustment commands, device control commands, and power allocation commands, thereby solving the problem of satellite systems not fully utilizing hardware resources.

[0174] The technical features of the above embodiments can be combined in any way. For the sake of brevity, 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 embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are specific and detailed, they should not be construed as limiting the scope of this patent application. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this application should be determined by the appended claims.

Claims

1. A smart multispectral payload satellite system, characterized in that, The system includes: The energy subsystem is configured to provide energy input; An integrated electronic subsystem is connected to the energy subsystem; the integrated electronic subsystem includes an onboard computer, which is equipped with an attitude and orbit control module, a satellite service module, and a power distribution module. The attitude control module is connected to multiple attitude sensors; the multiple attitude sensors are configured to acquire multimodal attitude signals and send the attitude signals to the attitude control module; the attitude control module is configured to generate adjustment instructions based on the current task stage in a set task mode and the attitude signals, and send the adjustment instructions to the adjustment target device; the task mode includes multiple task stages divided according to running time and / or running posture; the adjustment target device is a device determined based on the current task stage and the attitude signals for adjusting the running posture. The satellite service module is configured to generate device control instructions according to the current mission phase. The device control instructions are used to start or stop associated devices in the current mission phase. The power allocation module is configured to generate power allocation instructions based on power demand information, which includes one or more combinations of the current mission phase, the adjustment instructions, and the device control instructions. The power allocation instructions are used to allocate power to associated devices and / or the adjustment target devices in the current mission phase. A communication subsystem connects the energy subsystem and the integrated electronic subsystem; The payload subsystem connects the energy subsystem and the integrated electronic subsystem; the payload subsystem includes a multispectral payload camera configured to acquire remote sensing data according to the current mission phase and to transmit the remote sensing data to the integrated electronic subsystem.

2. The system according to claim 1, characterized in that, The system also includes: The structural subsystem includes a load-bearing mechanism, a deployment assembly, and a separation assembly; the load-bearing mechanism is a frame structure formed by connecting a first load-bearing member and a second load-bearing member through multiple supports; The unfolding assembly includes a flap and an unfolding hinge; the energy subsystem includes a power source, which includes a solar panel, which is disposed on the flap; the flap is rotatably connected to both sides of the load-bearing mechanism via the unfolding hinge. The separation assembly includes a separation mechanism and a separation switch. The separation mechanism is a plug-in connector that uses a shape memory alloy for unlocking. The separation mechanism is connected to the power supply via the separation switch. The separation switch is connected to the space service module, which is further configured to: Generate separation instructions according to the set task mode; The separation command is sent to the separation switch, which controls the separation switch to close, so that the separation mechanism is energized to perform the separation action.

3. The system according to claim 2, characterized in that, The system also includes: The thermal control subsystem includes a temperature control component and a thermal insulation component; the thermal insulation component is mounted on the load-bearing mechanism, and the temperature control component is located within the internal space of the frame structure of the load-bearing mechanism; the multispectral payload camera is temperature-isolated through the thermal insulation component. 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 temperature control sheet are connected to the thermal control module; the thermal control module is further configured to: Receive the internal temperature of the satellite detected by the temperature sensor; When the temperature inside the satellite is not within the preset operating temperature range, a temperature control command is sent to the thin-film temperature control sheet. The temperature control command is used to control the thin-film temperature control sheet to generate an adjustable temperature rise.

4. The system according to claim 2, characterized in that, The system also includes: The electrical subsystem includes a cable assembly and a test disconnect switch; the cable assembly is an independent module integrating onboard cables and connectors; the test disconnect switch is mounted on the disconnect assembly and configured to detect the disconnection status of the disconnect assembly; the test disconnect switch is connected to the satellite service module, which is further configured to: Receive the separation status 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. The first information is used to indicate that the separation component has performed the separation action. If the separation status information is the second information, the separation task is repeated according to the set task mode. The second information is used to indicate that the separation component has not performed a separation action.

5. The system according to claim 2, characterized in that, The system also includes: The propulsion subsystem includes a vacuum arc thruster; the vacuum arc thruster is connected to the power supply via the power distribution module; the power distribution module is further configured to: Obtain the running trajectory information within a preset detection period; The deviation data is obtained by comparing the operational trajectory information with the planned operational path. The operating parameters of the vacuum arc thruster are set according to the deviation data, including thrust power, thrust time, thrust direction and propellant ejection rate; The operating parameters are sent to the vacuum arc thruster to control the vacuum arc thruster to perform propulsion according to the operating parameters.

6. The system according to claim 1, characterized in that, The plurality of attitude sensors include navigation satellite components, angular velocity sensors, and magnetometers; the attitude and orbit control module is further configured to: Spatial positioning signals, angular velocity signals, and magnetic force signals are extracted from the multimodal attitude signals detected by the attitude sensors. The spatial positioning signals are signals detected by the navigation satellite components; the angular velocity signals are signals detected by the angular velocity sensors; and the magnetic force signals are signals detected by the magnetic force sensors. Calculate the current attitude information based on the spatial positioning signal, the angular velocity signal, and the magnetic force signal; According to the set task mode, an adjustment command is generated based on the current attitude information. The adjustment command includes the adjustment target device and the adjustment parameter. The adjustment target device is any one of the attitude adjustment devices associated with the spatial positioning signal, the angular velocity signal and the magnetic force signal. The adjustment command is sent to the target device to cause the target device to operate according to the adjustment parameters.

7. The system according to claim 6, characterized in that, The navigation satellite assembly includes a front-end radio frequency processing unit, a high dynamic navigation unit, a microprocessor, and an interface unit that are connected in sequence. The front-end radio frequency processing unit is connected to the navigation antenna of the communication subsystem; The front-end 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 signal conversion terms, and to compute navigation data from the converted signal; the signal conversion terms include 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 calculations on the navigation data to obtain navigation parameters, and send the navigation parameters to the attitude and orbit control module through the interface unit. 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 also includes a star sensor, which comprises an optical lens and an image processing unit. The optical lens is configured to acquire space optical signals and transmit 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 positioning images and transmit the celestial positioning images to the attitude and orbit control module. The attitude and orbit control module is further configured to: Receive the attitude angle signal detected by the angular velocity sensor; Calculate the three-axis attitude information based on the attitude angle signal and the celestial positioning image; Obtain the standard pose information for the current task mode; Calculate the attitude difference between the standard attitude information and the three-axis attitude information; If the attitude difference is greater than or equal to the attitude difference threshold, the adjustment command is generated.

9. The system according to claim 6, characterized in that, The integrated electronic subsystem also includes a solar sensor connected to the attitude and orbit control module. The solar sensor is configured to detect solar azimuth information using a four-quadrant differential operating principle and to send the solar azimuth information to the attitude and orbit control module. The attitude and orbit control module is further configured to: Acquire the magnetic signal detected by the magnetic sensor; The geocentric orientation information is determined based on the magnetic signal; Perform dual-vector attitude determination based on the solar azimuth information and the geocentric azimuth information to generate attitude information; The adjustment command is generated based on the attitude information.

10. A satellite system operation control method, characterized in that, Applied to the intelligent multispectral payload satellite system according to any one of claims 1-9, the method comprises: Acquire multimodal attitude signals and set the current task stage in the task mode, wherein the task mode includes multiple task stages divided according to running time and / or running attitude; An adjustment command is generated based on the current task stage and the attitude signal. The adjustment command includes adjustment parameters and an adjustment target device. The adjustment parameters are set according to the attitude signal. The adjustment target device is a device used to adjust the operating attitude, determined according to the current task stage and the attitude signal. Locate the associated devices for the current task phase; Device control instructions are generated according to the current task stage, and the device control instructions are used to start or stop the associated devices in the current task stage; A power allocation instruction is generated based on power demand information, which 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 to the associated devices and the adjustment target devices in the current task stage. The adjustment command and the power allocation command are sent to the target device for adjustment, and the device control command and the power allocation command are sent to the associated device in the current task phase.

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