Spacecraft heat dissipation component control method and computer equipment
By collecting multi-dimensional data through a distributed temperature sensor array and an environmental perception module, and combining it with preset temperature thresholds and external environmental data, the deployment attitude of the spacecraft's heat dissipation components is precisely controlled. This solves the problem of frequent opening and closing caused by single temperature feedback, reduces wear and energy consumption, and improves heat dissipation efficiency.
Patent Information
- Application Number
- CN202511867928.4
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-11
- Publication Date
- 2026-02-27
AI Technical Summary
When spacecraft are controlling heat dissipation in the complex space environment, a single temperature feedback causes the heat dissipation structure to open and close frequently, increasing wear and energy consumption, and it cannot effectively cope with transient thermal conditions.
By collecting multi-dimensional data through a distributed temperature sensor array and an environmental perception module, and combining it with preset temperature thresholds and external environmental data, the deployment posture of heat dissipation components can be precisely controlled to avoid frequent opening and closing actions.
It reduces mechanical wear and energy consumption of the heat dissipation structure, improves the heat dissipation efficiency of spacecraft, and adapts to the complex changes in the space environment.
Smart Images
Figure CN121573213A_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of heat dissipation control technology, and in particular to control methods, computer equipment, and computer-readable storage media for spacecraft heat dissipation components. Background Technology
[0002] In related technologies, spacecraft operating in the complex space environment often rely on single-feedback control based on temperature information. However, this can lead to excessively frequent opening and closing of the spacecraft's heat dissipation structure. This can not only increase wear and tear on the heat dissipation structure but also potentially increase the spacecraft's energy consumption. Summary of the Invention
[0003] This application provides a control method for a spacecraft heat dissipation component, a computer device, and a computer-readable storage medium.
[0004] This application provides a control method for a spacecraft heat dissipation component, the method comprising: Acquire the temperature information and external environmental data of the spacecraft's heat dissipation components; Based on the temperature information and the preset temperature threshold, the target state of the spacecraft heat dissipation component is determined; The deployment attitude of the spacecraft's heat dissipation components is controlled based on the target state, the temperature information, and the external environment data.
[0005] In this way, temperature information and external environmental data of the spacecraft's heat dissipation components are acquired. Next, based on the temperature information and a preset temperature threshold, the target state of the spacecraft's heat dissipation components is determined. Finally, based on the target state, temperature information, and external environmental data, the deployment attitude of the spacecraft's heat dissipation components is controlled. This method, by determining the target state of the heat dissipation components based on temperature information and external environmental data, and then precisely controlling the deployment attitude of the heat dissipation structure by combining these factors, allows for multi-dimensional control of the spacecraft's heat dissipation components' deployment attitude. This avoids frequent opening and closing actions caused by single temperature feedback, thereby reducing mechanical wear and energy consumption, and improving the spacecraft's heat dissipation efficiency.
[0006] In some embodiments, the external environment data includes illumination information and spacecraft operational attitude data, and the spacecraft heat dissipation component includes a multi-source sensing layer, which includes a distributed temperature sensing array and an environmental sensing module. The distributed temperature sensing array includes multiple temperature sensors, which are evenly distributed in the heat dissipation area of the spacecraft's heat dissipation component. The environmental perception module is used to collect the illumination information and the spacecraft's operational attitude data.
[0007] Thus, the external environment data includes illumination information and spacecraft operating attitude data, the spacecraft heat dissipation component includes a multi-source perception layer, and the multi-source perception layer includes a distributed temperature sensing array and an environment perception module. The distributed temperature sensing array includes a plurality of temperature sensors that are uniformly distributed in a heat dissipation region of the spacecraft heat dissipation component. The environment perception module is configured to collect the illumination information and the spacecraft operating attitude data. In this way, the distributed temperature sensing array, as the core of internal temperature collection, ensures that the temperature conditions at each position in the heat dissipation region can be captured through the uniform distribution of the plurality of temperature sensors, thereby avoiding the limitations of local data. The environment perception module focuses on external influencing factors and provides a basis for temperature change prediction by accurately collecting the illumination information and the spacecraft operating attitude data.
[0008] In some embodiments, the obtaining of the temperature information of the spacecraft heat dissipation component and external environment data includes: obtaining initial temperature information collected by each temperature sensor; determining an average value of the initial temperature information as the temperature information; obtaining the illumination information and the spacecraft operating attitude data collected by the environment perception module.
[0009] Thus, the initial temperature information collected by each temperature sensor is obtained. Then, the average value of the initial temperature information is determined as the temperature information. Then, the illumination information and the spacecraft operating attitude data collected by the environment perception module are obtained. In this way, by taking the average value of the initial temperature information as the temperature information, the influence of local temperature measurement errors and abnormal fluctuations can be eliminated, and the temperature information can truly and comprehensively reflect the overall heat state of the heat dissipation region, thereby providing accurate internal data basis for control decisions. Moreover, by capturing the dynamic changes of the external environment in real time, the control decisions are no longer limited to single internal temperature feedback, but can consider internal and external factors comprehensively, thereby effectively adapting to complex transient thermal working conditions and space environment changes in space.
[0010] In some embodiments, the spacecraft heat dissipation component includes a heat dissipation structure, and the heat dissipation structure includes a circular blade structure and a plurality of sector-shaped blade structures, wherein each sector-shaped blade structure is independently controllable.
[0011] Thus, the spacecraft heat dissipation component includes a heat dissipation structure, and the heat dissipation structure includes a circular blade structure and a plurality of sector-shaped blade structures, wherein each sector-shaped blade structure is independently controllable. In this way, the plurality of independently controllable sector-shaped blade structures enable the heat dissipation structure to have differentiated and precise adjustment capabilities, and enable targeted heat dissipation regulation according to temperature differences in different regions of the spacecraft and dynamic changes of the external environment.
[0012] In some embodiments, the preset temperature threshold comprises a first preset temperature threshold and a second preset temperature threshold, and the determining the target state of the spacecraft heat dissipation component according to the temperature information and the preset temperature threshold comprises: in a case where the temperature information is less than or equal to the first preset temperature threshold, determining that the target state is a first state, the first state being used to indicate that a difference between heat dissipated by the spacecraft heat dissipation component and heat generated by a spacecraft where the spacecraft heat dissipation component is located is greater than or equal to a first preset difference threshold; in a case where the temperature information is greater than the first preset temperature threshold and less than or equal to the second preset temperature threshold, determining that the target state is a second state, the second state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the first preset difference threshold and greater than or equal to a second preset difference threshold; in a case where the temperature information is greater than the second preset temperature threshold and less than or equal to a third preset temperature threshold, determining that the target state is a third state, the third state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the second preset difference threshold and greater than or equal to a third preset difference threshold; in a case where the temperature information is greater than the third preset temperature threshold, determining that the target state is a fourth state, the fourth state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the third preset difference threshold.
[0013] In this way, in a case where the temperature information is less than or equal to a first preset temperature threshold, the target state is determined as a first state, and the first state is used to indicate that a difference between heat dissipated by the spacecraft heat dissipation component and heat generated by the spacecraft in which the spacecraft heat dissipation component is located is greater than or equal to a first preset difference threshold. Then, in a case where the temperature information is greater than the first preset temperature threshold and less than or equal to a second preset temperature threshold, the target state is determined as a second state, and the second state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the first preset difference threshold and greater than or equal to a second preset difference threshold. Then, in a case where the temperature information is greater than the second preset temperature threshold and less than or equal to a third preset temperature threshold, the target state is determined as a third state, and the third state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the second preset difference threshold and greater than or equal to a third preset difference threshold. Finally, in a case where the temperature information is greater than the third preset temperature threshold, the target state is determined as a fourth state, and the fourth state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the third preset difference threshold. In this way, by dividing the temperature range into the first preset temperature threshold, the second preset temperature threshold, and the third preset temperature threshold, the heat dissipation requirement is accurately quantified into four states, which can avoid extreme adjustment caused by a single threshold and effectively reduce the over-response phenomenon, thereby reducing the mechanical wear and energy consumption.
[0014] In some embodiments, the controlling the deployment posture of the spacecraft heat dissipation component according to the target state, the temperature information, and the external environment data comprises: In a case where the target state is the first state, the heat dissipation structure is controlled to be completely closed.
[0015] In this way, in a case where the target state is the first state, the heat dissipation structure is controlled to be completely closed. In this way, the energy consumption can be effectively reduced, and the frequent start-stop of the heat dissipation structure in a low-temperature environment can be avoided, thereby reducing the mechanical wear caused by the opening and closing of the blades.
[0016] In some embodiments, the controlling the deployment posture of the spacecraft heat dissipation component according to the target state, the temperature information, and the external environment data comprises: In a case where the target state is the second state, the current deployment posture of the heat dissipation structure is maintained.
[0017] In this way, in a case where the target state is the second state, the current deployment posture of the heat dissipation structure is maintained. In this way, the mechanical friction caused by frequent adjustment can be avoided, the service life of the heat dissipation structure can be prolonged, and the energy consumption for driving the blades to move can be reduced.
[0018] In some embodiments, the external environment data comprises illumination information and spacecraft running attitude data, and the controlling, according to the target state, the temperature information and the external environment data, of the deployment attitude of the spacecraft heat dissipation component comprises: In a case where the target state is the third state, a target adjustment region is determined according to the illumination information and the spacecraft running attitude data, the target adjustment region being used to indicate a region in the heat dissipation structure that is not illuminated by sunlight; A target opening degree of the target adjustment region is calculated according to the temperature information; The deployment attitude of the heat dissipation structure is controlled according to the target adjustment region and the target opening degree.
[0019] In this way, in a case where the target state is the third state, a target adjustment region is determined according to the illumination information and the spacecraft running attitude data, the target adjustment region being used to indicate a region in the heat dissipation structure that is not illuminated by sunlight. Then, a target opening degree of the target adjustment region is calculated according to the temperature information. Finally, the deployment attitude of the heat dissipation structure is controlled according to the target adjustment region and the target opening degree. In this way, by avoiding the light-receiving region through the target adjustment region, heat dissipation and heat absorption can be avoided from occurring at the same time.
[0020] In some embodiments, the controlling, according to the target state, the temperature information and the external environment data, of the deployment attitude of the spacecraft heat dissipation component comprises: In a case where the target state is the fourth state, the heat dissipation structure is controlled to be completely deployed, and fault information is generated.
[0021] In this way, in a case where the target state is the fourth state, the heat dissipation structure is controlled to be completely deployed, and fault information is generated. In this way, temperature rise can be quickly suppressed, and equipment safety can be protected.
[0022] The embodiments of the present application provide a computer device, which comprises a processor and a memory, the memory storing a computer program, and the processor implements the steps of the above method when executing the computer program.
[0023] The embodiments of the present application provide a computer readable storage medium, which stores a computer program, and the computer program implements the steps of the above method when executed by a processor.
[0024] Additional aspects and advantages of the embodiments of the present application will be in part apparent and in part pointed out hereinafter. BRIEF DESCRIPTION OF DRAWINGS
[0025] The above and / or additional aspects and advantages of the present application will become apparent and more readily appreciated from the following description, taken in conjunction with the following drawings in which: Figure 1 is one of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 2 is another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 3 is still another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 4 is still another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 5 is still another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 6 is still another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application; Figure 7 is still another of flow diagrams of a control method of a spacecraft heat dissipation component according to some embodiments of the present application. DETAILED DESCRIPTION
[0026] Embodiments of the present application are described in detail below with reference to the attached drawing figures, wherein the same or like reference numerals and letters in different figures indicate the same or like components or elements throughout the figures. The implementation described below is an example only and is not intended to be limiting of the present application as the present application can have a variety of modifications and alternative forms. Specifically, the embodiments described below are provided purely for explanatory purposes and should not be construed as limiting the present application in any manner.
[0027] In the related art, when a spacecraft performs an on-orbit task in a complex and changeable space environment, its thermal control system often only relies on temperature information collected by a single temperature detection point to carry out simple closed-loop feedback control. This control mode lacks comprehensive consideration of multidimensional data and only uses the real-time temperature of a single monitoring point as the only basis for the action of the heat dissipation structure, resulting in obvious limitations in the control logic.
[0028] The particularity of the space environment makes the temperature of the spacecraft extremely susceptible to instantaneous fluctuations. On the one hand, the spacecraft frequently enters and exits the Earth's shadow zone during orbital operation, and the temperature changes from a high-temperature environment directly irradiated by the sun to a low-temperature environment without light in an instant, and the temperature change rate can reach several degrees Celsius per second. On the other hand, sudden conditions such as solar flare eruption, orbital height deviation, and change of solar incidence angle can cause the solar irradiance received by the spacecraft to change sharply, further exacerbating the instability of the temperature. A single feedback control mode cannot distinguish between these instantaneous fluctuations and persistent temperature changes, and as soon as the temperature deviates from the preset threshold, the opening and closing action of the heat dissipation structure will be triggered immediately.
[0029] In this way, the opening and closing action of the spacecraft heat dissipation structure will inevitably be too frequent: the heat dissipation blades may need to complete multiple cycles of expansion and closure in a short period of time. Such high-frequency mechanical action puts the core components such as the blade shaft and driving mechanism of the heat dissipation structure in a state of repeated stress for a long time, which is prone to mechanical fatigue, increased clearance, and other problems, not only reducing the control accuracy of the heat dissipation structure, but also possibly causing premature failure of the components, shortening the service life of the entire thermal control system. At the same time, frequent opening and closing actions require the start and braking of the driving motor, and these processes consume a large amount of energy. Moreover, the spacecraft carries a limited energy reserve, and this meaningless energy waste will squeeze the energy allocation of other equipment, which may affect the endurance capability of the spacecraft in orbit.
[0030] Based on the above problems, please refer to Figure 1 The embodiment of the present application provides a control method of a spacecraft heat dissipation component, which comprises the following steps: 01: obtaining temperature information and external environment data of the spacecraft heat dissipation component; 02: determining a target state of the spacecraft heat dissipation component according to the temperature information and a preset temperature threshold; 03: controlling the expansion posture of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data.
[0031] The embodiment of the present application also provides a computer device comprising a memory and a processor. The control method of the spacecraft heat dissipation component of the embodiment of the present application can be realized by the computer device of the embodiment of the present application. Specifically, the memory stores a computer program, and the processor is used to obtain temperature information and external environment data of the spacecraft heat dissipation component. And according to the temperature information and a preset temperature threshold, a target state of the spacecraft heat dissipation component is determined. And according to the target state, the temperature information and the external environment data, the expansion posture of the spacecraft heat dissipation component is controlled.
[0032] The application embodiment further provides a heat dissipation component posture control device. The spacecraft heat dissipation component control method of the application embodiment can be implemented by the heat dissipation component posture control device of the application embodiment. Specifically, the heat dissipation component posture control device comprises an acquisition module, a determination module and a control module. The acquisition module is configured to acquire temperature information of the spacecraft heat dissipation component and external environment data. The determination module is configured to determine a target state of the spacecraft heat dissipation component according to the temperature information and a preset temperature threshold. The control module is configured to control a deployment posture of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data.
[0033] Specifically, the spacecraft heat dissipation component refers to a heat dissipation device for a spacecraft, which can realize the exchange of heat inside the spacecraft and the external environment by adjusting the posture of the structure itself, so as to ensure that the spacecraft equipment operates in an appropriate temperature range.
[0034] The temperature information refers to data reflecting the heat level of the heat dissipation component and the surrounding area after being collected and processed by a temperature sensing device on the spacecraft heat dissipation component.
[0035] The external environment data refers to external space environment parameters affecting the heat dissipation effect of the spacecraft, mainly including illumination information and spacecraft operation posture data. The illumination information includes the radiation intensity of celestial bodies such as the sun on the spacecraft. The spacecraft operation posture data includes parameters such as the position of the spacecraft on the orbit and the attitude angle.
[0036] The preset temperature threshold refers to a temperature determination standard preset for realizing heat control of the spacecraft, including a first preset temperature threshold, a second preset temperature threshold and a third preset temperature threshold.
[0037] The target state refers to the working state of the spacecraft heat dissipation component determined according to the comparison result of the temperature information and the preset temperature threshold, including four states of a first state, a second state, a third state and a fourth state, which correspond to different heat dissipation levels.
[0038] The deployment posture refers to the opening position and the opening degree of the heat dissipation structure in the spacecraft heat dissipation component, which directly affects the heat dissipation area and the heat dissipation efficiency, and different heat dissipation effects can be achieved by adjusting the deployment posture.
[0039] Firstly, the initial temperature information of different heat dissipation areas of the heat dissipation component is collected by a plurality of temperature sensors in the distributed temperature sensing array, and then the mean value of all the initial temperature information is calculated to obtain temperature information reflecting the overall temperature condition of the heat dissipation component, so as to ensure the accuracy and comprehensiveness of the temperature data. At the same time, the illumination information data and the spacecraft operation posture data are collected by the environment perception module to complete the collection of the external environment data.
[0040] Subsequently, the processed temperature information is compared with a preset temperature threshold to determine a target state of the spacecraft heat dissipation component.
[0041] Finally, according to the determined target state, the temperature information and the external environment data are combined to precisely control the deployment posture of the heat dissipation structure.
[0042] In summary, for the spacecraft heat dissipation component control method and computer equipment provided by the embodiments of the present application, temperature information and external environment data of the spacecraft heat dissipation component are obtained. Then, according to the temperature information and a preset temperature threshold, a target state of the spacecraft heat dissipation component is determined. Finally, according to the target state, the temperature information and the external environment data, the deployment posture of the spacecraft heat dissipation component is controlled. In this way, by means of temperature information and external environment data, the target state of the heat dissipation component is determined according to the preset temperature threshold, and then the deployment posture of the heat dissipation structure is precisely controlled in combination with the target state, the temperature information and the external environment data. The deployment posture of the spacecraft heat dissipation component is controlled from multiple dimensions, which can avoid frequent opening and closing actions caused by single temperature feedback, thereby reducing the wear and tear of the mechanism and energy consumption, and improving the heat dissipation efficiency of the spacecraft.
[0043] In some embodiments, the external environment data includes illumination information and spacecraft running posture data, and the spacecraft heat dissipation component includes a multi-source perception layer, which includes a distributed temperature sensing array and an environment perception module. The distributed temperature sensing array includes a plurality of temperature sensors, which are uniformly distributed in the heat dissipation region of the spacecraft heat dissipation component. The environment perception module is used to collect the illumination information and the spacecraft running posture data.
[0044] Specifically, the illumination information refers to the radiation intensity of celestial bodies such as the sun in the space environment where the spacecraft is located, which is a key environmental factor that determines how much heat the spacecraft receives, and its change will affect the heat dissipation demand of the heat dissipation component.
[0045] The spacecraft running posture data refers to parameters such as position and attitude angle of the spacecraft during on-orbit operation, reflecting the relative position relationship between the spacecraft and heat sources such as the sun, and thus affecting the heating situation in different regions.
[0046] The multi-source perception layer refers to a functional module in the spacecraft heat dissipation component responsible for comprehensive and accurate collection of temperature data and external environment data, including various sensing devices such as a distributed temperature sensing array and an environment perception module, providing comprehensive and reliable data support for subsequent intelligent decision-making.
[0047] The distributed temperature sensing array refers to a component in the multi-source perception layer that collects temperature information, which is formed by a plurality of temperature sensors arranged in a specific manner, and can realize omnidirectional monitoring of the temperature of the heat dissipation region.
[0048] The temperature sensor refers to a sensing element for directly detecting the temperature of the heat dissipation area, and can convert the temperature physical quantity into an electrical signal or a data signal that can be transmitted and processed, and is a basic unit of temperature data acquisition.
[0049] The heat dissipation area refers to an area in the spacecraft heat dissipation component that directly participates in heat dissipation, and is the part with the most concentrated temperature change and the most reflection of heat dissipation demand.
[0050] The environment perception module refers to a component in the multi-source perception layer that collects external environment data, integrates sensing devices suitable for space environment, and is used to capture dynamic changes of illumination information and spacecraft running attitude data.
[0051] The multi-source perception layer is the main module of data acquisition, including a distributed temperature sensing array and an environment perception module, which respectively undertake the functions of internal temperature collection and external environment collection, realize the data coverage of the internal-external two dimensions, and make up for the defects of traditional single temperature dimension collection. Among them, the distributed temperature sensing array is the core of internal temperature collection, through the uniform layout of multiple temperature sensors, it can ensure that the temperature of each position of the heat dissipation area can be captured, and the limitation of local data can be avoided. The environment perception module focuses on external influencing factors, and accurately collects illumination information and spacecraft running attitude data to provide basis for temperature change prediction.
[0052] In this way, the external environment data includes illumination information and spacecraft running attitude data, the spacecraft heat dissipation component includes a multi-source perception layer, and the multi-source perception layer includes a distributed temperature sensing array and an environment perception module. Among them, the distributed temperature sensing array includes multiple temperature sensors, and the multiple temperature sensors are uniformly distributed in the heat dissipation area of the spacecraft heat dissipation component. The environment perception module is used to collect illumination information and spacecraft running attitude data. In this way, the distributed temperature sensing array is the core of internal temperature collection, through the uniform layout of multiple temperature sensors, it can ensure that the temperature of each position of the heat dissipation area can be captured, thereby avoiding the limitation of local data. The environment perception module focuses on external influencing factors, and accurately collects illumination information and spacecraft running attitude data to provide basis for temperature change prediction.
[0053] Please refer to Figure 2 In some embodiments, step 01 (obtaining temperature information of the spacecraft heat dissipation component and external environment data) includes: 011: obtaining initial temperature information collected by each temperature sensor; 012: determining the average value of the initial temperature information as the temperature information; 013: obtaining illumination information and spacecraft running attitude data collected by the environment perception module.
[0054] In some embodiments, the processor is further configured to acquire initial temperature information collected by each temperature sensor, and determine an average of the initial temperature information as the temperature information.
[0055] In some embodiments, the processor is further configured to acquire initial temperature information collected by each temperature sensor, and determine an average of the initial temperature information as the temperature information.
[0056] Specifically, the distributed temperature sensor array is an acquisition unit of the initial temperature information. The plurality of temperature sensors are uniformly distributed in the heat dissipation region of the spacecraft heat dissipation component, which can comprehensively cover the key parts of heat dissipation and avoid local temperature omission. Each temperature sensor works independently to convert the temperature physical quantity at the location into initial temperature information in the form of an electrical signal. Moreover, considering the possible measurement error or local temperature abnormal fluctuation of a single sensor, the system does not directly use the data of a single sensor as the basis for decision-making, but integrates and processes the initial temperature information collected by all temperature sensors, eliminates the influence of local errors and abnormal fluctuations by calculating the arithmetic mean, and finally obtains temperature information that can reflect the overall heat level of the heat dissipation region.
[0057] The environment perception module is a core module for acquiring external environment data, which integrates special sensing devices such as a sun sensor and a spacecraft attitude / orbit parameter receiver. The sun sensor is responsible for accurately capturing light information, including solar radiation intensity, light angle, etc., which can monitor the mutation of light amplitude in real time, such as the sudden increase of radiation intensity caused by solar flares. The spacecraft attitude / orbit parameter receiver is linked with the spacecraft control system to acquire the running attitude data of the spacecraft in real time, including orbit height, attitude angle, running speed, etc., and the relative position relationship between the spacecraft and the sun.
[0058] In this way, the initial temperature information collected by each temperature sensor is acquired. Then, the average of the initial temperature information is determined as the temperature information. Then, the light information and the spacecraft running attitude data collected by the environment perception module are acquired. In this way, by taking the average of each initial temperature information as the temperature information, the influence of local temperature measurement error and abnormal fluctuation can be eliminated, and the temperature information can truly and comprehensively reflect the overall heat state of the heat dissipation region, providing accurate internal data basis for control decision-making. Moreover, by capturing the dynamic changes of the external environment in real time, the control decision can no longer be limited to single internal temperature feedback, but can consider internal and external factors comprehensively, effectively adapting to the complex transient thermal working conditions and space environment changes in space.
[0059] In some embodiments, the spacecraft heat dissipation component includes a heat dissipation structure, the heat dissipation structure includes a circular blade structure and a plurality of sector blade structures, wherein each sector blade structure is independently controllable.
[0060] Specifically, the heat dissipation structure refers to the mechanical structure directly involved in heat dissipation in the spacecraft heat dissipation component, which changes the heat dissipation area and path by changing its own form, and is the execution unit for precise heat dissipation control.
[0061] The circular blade structure refers to the basic component of the heat dissipation structure, which is arranged in a ring shape and serves as the core bearing unit and basic heat dissipation area of the heat dissipation structure, and itself participates in basic heat dissipation.
[0062] The sector blade structure refers to the adjustable heat dissipation unit of the heat dissipation structure, which is distributed in a sector shape around or inside the circular blade structure, and realizes flexible adjustment of the heat dissipation area by changing the expansion angle, and is the core component for realizing hierarchical heat dissipation.
[0063] Independent control refers to that each sector blade structure is equipped with an independent driving mechanism and a control signal receiving unit, and can independently adjust the expansion angle and switching state according to the control instruction, and is not affected by the action of other sector blade structures, thereby realizing targeted heat dissipation regulation.
[0064] The sector blade structure is evenly distributed in a circle around the circular blade structure, forming a complete heat dissipation surface. Each sector blade structure is equipped with an independent micro driving mechanism and a control module, and the control module can receive a separate control instruction from the decision engine to drive the sector blade structure to expand or close around the mounting shaft.
[0065] In this way, the spacecraft heat dissipation component includes a heat dissipation structure, the heat dissipation structure includes a circular blade structure and a plurality of sector blade structures, wherein each sector blade structure is independently controllable. In this way, the plurality of independently controllable sector blade structures can enable the heat dissipation structure to have differentiated and precise regulation capability, and can realize targeted heat dissipation regulation according to the temperature difference of different regions of the spacecraft and the dynamic change of the external environment.
[0066] Please refer to Figure 3 In some embodiments, the preset temperature threshold includes a first preset temperature threshold and a second preset temperature threshold, and step 02 (determining the target state of the spacecraft heat dissipation component according to the temperature information and the preset temperature threshold) includes: 021: In the case that the temperature information is less than or equal to the first preset temperature threshold, the target state is determined as the first state; 022: In the case that the temperature information is greater than the first preset temperature threshold and less than or equal to the second preset temperature threshold, the target state is determined as the second state; 023: determining the target state as the third state when the temperature information is greater than a second preset temperature threshold and less than or equal to a third preset temperature threshold; 024: determining the target state as the fourth state when the temperature information is greater than the third preset temperature threshold.
[0067] In some embodiments, the determining module is further configured to determine the target state as the first state when the temperature information is less than or equal to a first preset temperature threshold, and determine the target state as the second state when the temperature information is greater than the first preset temperature threshold and less than or equal to a second preset temperature threshold. The determining module is further configured to determine the target state as the third state when the temperature information is greater than the second preset temperature threshold and less than or equal to a third preset temperature threshold, and determine the target state as the fourth state when the temperature information is greater than the third preset temperature threshold.
[0068] In some embodiments, the processor is further configured to determine the target state as the first state when the temperature information is less than or equal to a first preset temperature threshold, and determine the target state as the second state when the temperature information is greater than the first preset temperature threshold and less than or equal to a second preset temperature threshold. The processor is further configured to determine the target state as the third state when the temperature information is greater than the second preset temperature threshold and less than or equal to a third preset temperature threshold, and determine the target state as the fourth state when the temperature information is greater than the third preset temperature threshold.
[0069] Specifically, the first preset temperature threshold refers to a critical temperature for the heat dissipation system to enter a heat preservation or closed state, and when the temperature information is lower than or equal to the threshold, it indicates that the spacecraft has a very low heat dissipation demand.
[0070] The second preset temperature threshold refers to a critical temperature for the start and maintenance of a regular heat dissipation mode, and is between the first and third preset temperature thresholds, corresponding to a medium heat dissipation demand scenario.
[0071] The third preset temperature threshold refers to a critical temperature for fault warning and emergency heat dissipation, and exceeding the threshold indicates that the heat dissipation demand is extremely urgent and emergency measures need to be started.
[0072] The first state refers to a state in which the heat dissipation amount of the heat dissipation component is much greater than the heat generation amount of the spacecraft, and no additional heat dissipation is needed.
[0073] The second state refers to a state in which the heat dissipation amount of the heat dissipation component matches the heat generation amount of the spacecraft, and the current heat dissipation state can meet the demand.
[0074] The third state refers to a state in which the heat dissipation amount of the heat dissipation component is insufficient, and targeted heat dissipation enhancement is needed.
[0075] The fourth state refers to a state in which the heat dissipation amount of the heat dissipation component is severely insufficient, and there is a risk of overheating, which needs to be handled urgently.
[0076] The preset difference threshold refers to a criterion for quantifying the matching degree of the heat dissipation amount of the heat dissipation component and the heat generation amount of the spacecraft, and includes a first preset difference threshold, a second preset difference threshold, and a third preset difference threshold, corresponding to different heat balance requirements of target states.
[0077] It should be noted that the preset temperature threshold is set after system calibration, and the suitable working temperature range of the spacecraft equipment, the thermal characteristics of the space environment, and the adjustment capability of the heat dissipation structure are fully considered.
[0078] The system first acquires temperature information processed by the distributed temperature sensing array, and then compares the temperature information with the first preset temperature threshold, the second preset temperature threshold, and the third preset temperature threshold one by one.
[0079] When the temperature information is less than or equal to the first preset temperature threshold, it is determined to be the first state, indicating that the current heat dissipation amount is excessive, and no additional heat dissipation is needed. When the temperature information is greater than the first preset temperature threshold and less than or equal to the second preset temperature threshold, it is determined to be the second state, indicating that the heat dissipation and heat generation are basically balanced, and no adjustment is needed. When the temperature information is greater than the second preset temperature threshold and less than or equal to the third preset temperature threshold, it is determined to be the third state, indicating that the heat dissipation amount is insufficient, and the heat dissipation needs to be enhanced. When the temperature information is greater than the third preset temperature threshold, it is determined to be the fourth state, indicating that the heat dissipation is severely insufficient, and emergency heat dissipation is needed.
[0080] In this way, in a case where the temperature information is less than or equal to the first preset temperature threshold, the target state is determined as the first state, and the first state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft in which the spacecraft heat dissipation component is located is greater than or equal to the first preset difference threshold. Then, in a case where the temperature information is greater than the first preset temperature threshold and less than or equal to the second preset temperature threshold, the target state is determined as the second state, and the second state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the first preset difference threshold and greater than or equal to the second preset difference threshold. Then, in a case where the temperature information is greater than the second preset temperature threshold and less than or equal to the third preset temperature threshold, the target state is determined as the third state, and the third state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the second preset difference threshold and greater than or equal to the third preset difference threshold. Finally, in a case where the temperature information is greater than the third preset temperature threshold, the target state is determined as the fourth state, and the fourth state is used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the third preset difference threshold. In this way, by dividing the temperature range by the first preset temperature threshold, the second preset temperature threshold, and the third preset temperature threshold, the heat dissipation requirement is accurately quantified into four states, which can avoid extreme adjustment caused by a single threshold and effectively reduce the over-response phenomenon, thereby reducing the wear and tear of the mechanism and energy consumption.
[0081] Referring to Figure 4 In some embodiments, step 03 (controlling the deployment posture of the spacecraft heat dissipation component according to the target state, the temperature information, and the external environment data) comprises: 031: in a case where the target state is the first state, controlling the heat dissipation structure to be completely closed.
[0082] In some embodiments, the determining module is further configured to, in a case where the target state is the first state, control the heat dissipation structure to be completely closed.
[0083] In some embodiments, the processor is further configured to, in a case where the target state is the first state, control the heat dissipation structure to be completely closed.
[0084] Specifically, completely closed refers to one of the control states of the heat dissipation structure, that is, all the blades of the heat dissipation structure are in a closed state, and the heat dissipation area is completely blocked, which can minimize heat dissipation and achieve heat preservation function.
[0085] The processed temperature information is compared with a preset first temperature threshold. If the temperature information is less than or equal to the first temperature threshold, it is determined that the target state is the first state, i.e., the spacecraft is currently at a low temperature, and the heat preservation control needs to be started. Subsequently, the system sends a complete closing instruction to the heat dissipation structure: control all the circular blades and the fan-shaped blades to close synchronously, so that the heat dissipation area is completely blocked, the main path of heat dissipation to the outside is cut off, the internal heat of the spacecraft is maximally retained, and the heat preservation function is realized.
[0086] In some embodiments, when it is determined that the target state is the first state based on the comparison of the temperature information with the preset first temperature threshold, to avoid misjudgment caused by instantaneous temperature fluctuation, the system further verifies the accuracy of the first state determination in combination with external environment data: if the light intensity detection result shows that the current irradiance is extremely low (for example, the spacecraft is in the shadow area of the Earth), and the spacecraft running attitude data indicates that there is no trend of entering the light area in the short term, it can be confirmed that the low temperature is a persistent state, and the target state is the first state.
[0087] In this way, when the target state is the first state, the heat dissipation structure is completely closed. In this way, the energy consumption can be effectively reduced, and the frequent start and stop of the heat dissipation structure in a low temperature environment can be avoided, so that mechanical wear caused by the opening and closing of the blades can be reduced.
[0088] Please refer to Figure 5 In some embodiments, step 03 (controlling the deployment attitude of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data) comprises: 032: In the case where the target state is the second state, the current deployment attitude of the heat dissipation structure is maintained.
[0089] In some embodiments, the determination module is further configured to maintain the current deployment attitude of the heat dissipation structure in the case where the target state is the second state.
[0090] In some embodiments, the processor is further configured to maintain the current deployment attitude of the heat dissipation structure in the case where the target state is the second state.
[0091] Specifically, maintaining the current deployment attitude means keeping the current opening angle and distribution state of the blades unchanged, continuing the current heat dissipation efficiency, and avoiding unnecessary action adjustment.
[0092] The processed temperature information is compared with the preset first temperature threshold and second temperature threshold. If the temperature information is greater than the first temperature threshold and less than or equal to the second temperature threshold, it is determined that the target state is the second state, that is, the current temperature of the spacecraft is in the appropriate interval, and there is no need to adjust the heat dissipation posture. Subsequently, the system sends a current deployment posture maintaining instruction to the heat dissipation structure: control all the blades to maintain the current opening angle and distribution state, continue the current heat dissipation efficiency, and do not perform any additional deployment or closing action.
[0093] In some embodiments, when it is determined that the target state is the second state, to avoid misjudgment caused by instantaneous temperature fluctuation, the system further verifies the accuracy of the second state determination in combination with external environment data: if the light intensity detection result shows that the current irradiation intensity is stable, there is no sudden increase or decrease, and the spacecraft running posture data shows that there is no large attitude adjustment in the short term and the orbit position does not change significantly, it can be confirmed that the current appropriate temperature state has stability, and the target state is the second state.
[0094] In this way, in the case that the target state is the second state, the current deployment posture of the heat dissipation structure is maintained. In this way, mechanical friction caused by frequent adjustment can be avoided, the service life of the heat dissipation structure is prolonged, and energy consumption for driving the blades to move is reduced.
[0095] Please refer to Figure 6 In some embodiments, the external environment data includes light information and spacecraft running posture data, and step 03 (controlling the deployment posture of the spacecraft heat dissipation component according to the target state, temperature information and external environment data) includes: 033: In the case that the target state is the third state, determining a target adjustment region according to the light information and the spacecraft running posture data; 034: Calculating a target opening degree of the target adjustment region according to the temperature information; 035: Controlling the deployment posture of the heat dissipation structure according to the target adjustment region and the target opening degree.
[0096] In some embodiments, the determination module is further configured to, in the case that the target state is the third state, determine a target adjustment region according to the light information and the spacecraft running posture data, calculate a target opening degree of the target adjustment region according to the temperature information, and control the deployment posture of the heat dissipation structure according to the target adjustment region and the target opening degree.
[0097] In some embodiments, the processor is further configured to, in the case that the target state is the third state, determine a target adjustment region according to the light information and the spacecraft running posture data, calculate a target opening degree of the target adjustment region according to the temperature information, and control the deployment posture of the heat dissipation structure according to the target adjustment region and the target opening degree.
[0098] Specifically, the target adjustment region refers to a region in the heat dissipation structure that is not irradiated by sunlight, and the region is free from additional heat absorption interference.
[0099] The target opening degree refers to an angle at which the blades in the target adjustment region should be expanded according to the temperature exceeding amplitude, and can balance the heat dissipation demand and energy consumption control.
[0100] The processed temperature information is compared with the preset second temperature threshold and third temperature threshold. If the temperature information is greater than the second temperature threshold and less than or equal to the third temperature threshold, it is determined that the target state is the third state, that is, the spacecraft has an overheating condition and needs local heat dissipation.
[0101] Subsequently, the sunlight irradiation direction and intensity are analyzed in combination with the light information, the current attitude angle and orbit position of the spacecraft are determined according to the spacecraft running attitude data, and the region in the heat dissipation structure that is not irradiated by sunlight is accurately determined through spatial geometric operation, and the region is marked as the target adjustment region, so as to ensure that the heat dissipation process is free from additional heat absorption interference.
[0102] Then, the required heat dissipation area is calculated through a heat conduction model according to the difference between the temperature information and the second temperature threshold and in combination with the heat dissipation efficiency parameter of the heat dissipation structure. The target opening degree required for the target adjustment region to be expanded is converted according to the total area of the target adjustment region, so as to balance the heat dissipation speed and energy consumption control.
[0103] Finally, a control instruction is sent to the sector-shaped blades in the target adjustment region to drive the blades to expand at the target opening degree, and the blades in other regions remain in the original state.
[0104] In some embodiments, when it is determined that the target state is the third state, to avoid misjudgment caused by instantaneous temperature fluctuation, the system further verifies the accuracy of the third state determination in combination with external environment data: if the light intensity is stable and there is no sudden increase or decrease according to the light amplitude detection result, and the spacecraft running attitude data indicates that there is no large attitude adjustment in the short term and the orbit position does not change significantly, it is determined that the current appropriate temperature state has stability, and it is determined that the target state is the third state.
[0105] In this way, when the target state is the third state, the target adjustment region is determined according to the light information and the spacecraft running attitude data, and the target adjustment region is used to indicate a region in the heat dissipation structure that is not irradiated by sunlight. Then, the target opening degree of the target adjustment region is calculated according to the temperature information. Finally, the expansion attitude of the heat dissipation structure is controlled according to the target adjustment region and the target opening degree. In this way, the target adjustment region avoids the light-irradiated region, and heat dissipation and heat absorption can be avoided from occurring simultaneously.
[0106] Please refer to Figure 7In some embodiments, step 03 (controlling the deployment posture of the spacecraft heat dissipation component according to the target state, temperature information and external environment data) comprises: 036: In the case where the target state is the fourth state, the heat dissipation structure is controlled to be fully deployed, and failure information is generated.
[0107] In some embodiments, the determining module is further configured to, in the case where the target state is the fourth state, control the heat dissipation structure to be fully deployed, and generate failure information.
[0108] In some embodiments, the processor is further configured to, in the case where the target state is the fourth state, control the heat dissipation structure to be fully deployed, and generate failure information.
[0109] Specifically, fully deployed means that all the blades are at the maximum opening angle, the heat dissipation area is fully exposed, the heat dissipation area is maximized, and the strongest heat dissipation effect is achieved.
[0110] The failure information refers to a data set generated by the system after detecting that the temperature is seriously out of range, which contains key information such as failure type, temperature data, failure occurrence time and external environment parameters, and is used for ground personnel diagnosis and disposal.
[0111] The processed temperature information is compared with the preset third temperature threshold value. If the temperature information is greater than the third temperature threshold value, it is determined that the target state is the fourth state, i.e. the spacecraft temperature is seriously out of range, which has reached the failure level, and needs to be cooled urgently and alarmed. Subsequently, the system starts the emergency disposal process: sends a fully deployed command to the heat dissipation structure, drives all the circular blades and fan-shaped blades to reach the maximum opening angle synchronously, exposes the heat dissipation area completely, maximizes the heat dissipation area, quickly conducts and dissipates a large amount of heat inside the spacecraft, and suppresses the continuous rise of temperature.
[0112] In addition, failure information is generated and fed back: real-time records of temperature data, external environment parameters, failure occurrence time, heat dissipation structure action state and other information when the failure occurs are integrated to generate standardized failure information. Subsequently, through the spacecraft communication system, the failure information is transmitted to the ground control center in real time, and an audible and visual alarm is triggered inside the spacecraft to remind the on-orbit personnel to pay attention to the failure state.
[0113] In some embodiments, when it is determined that the target state is the fourth state, in order to avoid misjudgment caused by instantaneous temperature fluctuation, the system further verifies the accuracy of the second state determination in combination with external environment data: if the light intensity detection result shows that the current irradiance is stable, there is no sudden increase or decrease, and the spacecraft running posture data shows that there is no large attitude adjustment in the short term and the orbit position does not change significantly, it can be confirmed that the current appropriate temperature state has stability, and the target state is the fourth state.
[0114] Thus, the second target screening value is determined according to the ratio of the second target attenuation rate and the second attenuation rate characteristic value. In this way, by determining the second target screening value according to the ratio of the second target attenuation rate and the second attenuation rate characteristic value, the target battery can be associated with the battery pack, so as to eliminate the interference caused by the individual differences of different batteries in the battery pack and environmental factors, and further improve the accuracy of target battery abnormality detection.
[0115] The application further provides a computer readable storage medium, which stores a computer program. When the computer program is executed by a processor, the steps of the control method of the spacecraft heat dissipation component are implemented.
[0116] It can be understood that the computer program includes computer program code. The computer program code can be in the form of source code, object code, executable files or some intermediate forms. The computer readable storage medium can include any entity or device capable of carrying computer program code, recording medium, U disk, mobile hard disk, magnetic disk, optical disk, computer memory, read-only memory (ROM), random access memory (RAM) and software distribution medium.
[0117] The application embodiment further provides a computer program product, which includes computer program / instruction, and the computer program / instruction is executed by a processor to implement the above method.
[0118] In the description of the present application, the description of the terms "specifically", "further", "particularly", "understandably" and the like means that the specific features, structures, materials or characteristics described in connection with the embodiments or examples are included in at least one embodiment or example of the present application. In the present application, the illustrative description of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, the different embodiments or examples described in the present application and the features of the different embodiments or examples can be combined and combined by those skilled in the art without contradiction.
[0119] Any processes or methods described in the flow charts or otherwise described herein can be understood as representing one or more modules, segments, or portions of code that include executable instructions for performing specific logical functions or steps, and the preferred embodiments of the application include additional implementations that can not be described in detail in order to not obscure the application. The scope of the application should be determined by the claims.
[0120] Although the embodiments of the present application have been shown and described above, it should be understood by those skilled in the art that the above-described embodiments are exemplary only, and should not be understood as limiting the present application, and those skilled in the art can make changes, modifications, substitutions, and variations to the above-described embodiments within the scope of the present application.
Claims
1. A method of controlling a spacecraft heat dissipation component, characterized by, The method comprises: acquiring temperature information of the spacecraft heat dissipation component and external environment data; determining a target state of the spacecraft heat dissipation component according to the temperature information and a preset temperature threshold; controlling an unfolding posture of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data.
2. The method of claim 1, wherein, The external environment data comprises illumination information and spacecraft running posture data, and the spacecraft heat dissipation component comprises a multi-source perception layer, the multi-source perception layer comprising a distributed temperature sensing array and an environment perception module; The distributed temperature sensing array comprises a plurality of temperature sensors, which are uniformly distributed in a heat dissipation area of the spacecraft heat dissipation component; The environment perception module is used to collect the illumination information and the spacecraft running posture data.
3. The method of claim 2, wherein, The acquisition of the temperature information of the spacecraft heat dissipation component and the external environment data comprises: acquiring initial temperature information collected by each temperature sensor; determining an average value of the initial temperature information as the temperature information; acquiring the illumination information and the spacecraft running posture data collected by the environment perception module.
4. The method of claim 1, wherein, The spacecraft heat dissipation component comprises a heat dissipation structure, the heat dissipation structure comprising a circular blade structure and a plurality of sector-shaped blade structures, wherein each sector-shaped blade structure can be independently controlled.
5. The method of claim 4, wherein, The preset temperature threshold comprises a first preset temperature threshold and a second preset temperature threshold, and the determination of the target state of the spacecraft heat dissipation component according to the temperature information and the preset temperature threshold comprises: in a case where the temperature information is less than or equal to the first preset temperature threshold, determining that the target state is a first state, the first state being used to indicate that a difference between heat dissipated by the spacecraft heat dissipation component and heat generated by a spacecraft where the spacecraft heat dissipation component is located is greater than or equal to a first preset difference threshold; in a case where the temperature information is greater than the first preset temperature threshold and less than or equal to the second preset temperature threshold, determining that the target state is a second state, the second state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the first preset difference threshold and greater than or equal to a second preset difference threshold; in a case where the temperature information is greater than the second preset temperature threshold and less than or equal to a third preset temperature threshold, determining that the target state is a third state, the third state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the second preset difference threshold and greater than or equal to a third preset difference threshold; in a case where the temperature information is greater than the third preset temperature threshold, determining that the target state is a fourth state, the fourth state being used to indicate that the difference between the heat dissipated by the spacecraft heat dissipation component and the heat generated by the spacecraft is less than the third preset difference threshold.
6. The method of claim 5, wherein, The control of the unfolding posture of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data comprises: in a case where the target state is the first state, controlling the heat dissipation structure to be completely closed.
7. The method of claim 5, wherein, The control of the deployment attitude of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data comprises: In the case where the target state is the second state, the current deployment attitude of the heat dissipation structure is maintained.
8. The method of claim 5, wherein, The external environment data comprises illumination information and spacecraft operation attitude data, and the control of the deployment attitude of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data comprises: In the case where the target state is the third state, a target adjustment region is determined according to the illumination information and the spacecraft operation attitude data, the target adjustment region being used to indicate a region of the heat dissipation structure that is not illuminated by sunlight; A target opening degree of the target adjustment region is calculated according to the temperature information; The deployment attitude of the heat dissipation structure is controlled according to the target adjustment region and the target opening degree.
9. The method of claim 5, wherein, The control of the deployment attitude of the spacecraft heat dissipation component according to the target state, the temperature information and the external environment data comprises: In the case where the target state is the fourth state, the heat dissipation structure is controlled to be completely deployed, and a fault information is generated.
10. A computer device, comprising: The device comprises a processor and a memory, the memory storing a computer program, and the processor implements the steps of the method of any one of claims 1-9 when executing the computer program.