Spacecraft thermal control subsystem design method and device based on mbse model and medium

Abstract

The application belongs to the technical field of spacecraft thermal control design, and provides a spacecraft thermal control subsystem design method, device and medium based on an MBSE model, which comprises: task analysis, converting a text requirement into a digital requirement model; scheme design, performing preliminary scheme design according to the requirement model; parameter modeling, establishing a simulation model to perform simulation verification and closed-loop iterative optimization on the scheme, and determining a final design scheme; and detailed design, generating production files according to the final scheme. The application realizes seamless data transmission and simulation driving optimization from the requirement to the product through the unified model throughout the whole process, can significantly shorten the development cycle, reduce the cost, and improve the design reusability and reliability, and solves the problems of design iteration difficulty, long cycle and high cost caused by information fragmentation in the spacecraft thermal control design.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft thermal control design technology, specifically, it relates to a design method, device and medium for a spacecraft thermal control subsystem based on the MBSE model. Background Technology

[0002] The spacecraft thermal control subsystem is one of the core subsystems ensuring the reliable operation of a spacecraft in orbit, and its design directly affects the spacecraft's performance and lifespan. Traditional thermal control subsystem design primarily relies on a document-centric development model. Information from each design phase, such as requirements analysis, scheme design, simulation verification, and detailed design, is scattered across different technical reports, design drawings, and interface protocols. This model has significant drawbacks: First, information transfer between phases via documents can easily lead to misunderstandings of requirements and inconsistent data, complex interface management, frequent design rework, and increased R&D risks and costs. Second, the rationality of the design scheme heavily depends on later physical prototype testing for verification, resulting in late problem discovery and long iteration cycles, making it difficult to meet the current requirements for short-cycle, low-cost spacecraft development. Finally, design knowledge and experience are solidified in unstructured documents, making it difficult to form standardized modules for reuse, leading to low efficiency in the development of new models.

[0003] To address these issues, the industry has begun to explore model-based systems engineering concepts, using unified models to manage design information. For example, some existing methods utilize models for requirements management or thermal simulation analysis, and even attempt to link task requirements with system configuration and simulation optimization. However, these methods mostly focus on a specific stage or part of the design process, failing to provide a systematic approach that seamlessly integrates and connects data across the entire process, including requirements analysis, solution design, parametric simulation verification, and detailed product design. In these methods, model breakpoints and data silos still exist between different stages of the design process, especially in the detailed design stage, from simulation-verified optimization to the generation of specific manufacturing documents, where effective connections are lacking.

[0004] Patent document "Systems and methods for describing, simulating and optimizing spaceborne systems and missions" (US11127102B2) discloses a system and method for evaluating and optimizing the hardware system configuration of spacecraft missions through simulation and verification configurations to ensure performance and resource compatibility, allowing the reuse of existing spacecraft platforms and payloads, thereby maximizing revenue and minimizing costs. However, it lacks a seamlessly integrated design methodology covering the entire process from requirements analysis to detailed product design and manufacturing preparation.

[0005] The patent document "A Structural Thermal Control Integrated Design Method for Micro and Nano Satellites" (CN119720386A) discloses a method for designing the organic integration of structural subsystems and thermal control subsystems by comprehensively considering load-bearing capacity, size, mass, mechanical interface, external heat flow and internal heat source. This method integrates thermally conductive materials, surface treatment of structural components, thermal insulation pads, MLI and thermally conductive fillers to improve thermal control performance.

[0006] Therefore, a spacecraft thermal control subsystem design method based on the MBSE model is needed to fundamentally solve the problem of coordination and iterative optimization throughout the entire design process. Summary of the Invention

[0007] To address the shortcomings of existing technologies, the present invention aims to provide a design method, system, and medium for spacecraft thermal control subsystems based on the MBSE model. This invention aims to solve the technical problems in existing spacecraft thermal control subsystem design methods, such as unclear requirement transmission, complex interface management, long R&D cycles, high costs, and poor reusability of design results, caused by information fragmentation at each design stage and reliance on document transmission.

[0008] According to the present invention, a spacecraft thermal control subsystem design method based on the MBSE model includes: Task analysis steps: Based on the development task specification of the spacecraft thermal control subsystem, the textual requirements are transformed into a structured digital requirement model, which includes the functions, performance, interfaces and constraints related to the thermal control subsystem. Scheme design steps: Based on the digital requirement model, conduct a preliminary scheme design for the spacecraft thermal control subsystem, and generate a design model corresponding to the preliminary scheme design; Parametric modeling steps: Based on the design model, establish a parametric simulation model for performance verification; by running the simulation, compare the simulation results with the performance requirements defined in the digital requirement model. If the simulation results do not meet the performance requirements, return to the scheme design steps to optimize and iterate the preliminary scheme design until the simulation results meet the performance requirements. The preliminary scheme design that meets the performance requirements at this time is determined as the final design scheme. Detailed design steps: Based on the three-dimensional thermal control model and technical requirements corresponding to the final design scheme, generate detailed technical documents for production and manufacturing.

[0009] Preferably, in the task analysis step, the digital requirement model is established using the SysML language.

[0010] Build a reusable baseline model library that includes at least one baseline model, and make differentiated modifications to the baseline models when performing new task designs.

[0011] The design process includes heat dissipation surface design, temperature uniformity design, thermal control compensation design, and multi-layer thermal insulation component design.

[0012] The temperature uniformity design includes automatic generation of heat pipe layout, following the principle of optimization: An isothermal heat network is formed under each individual unit mounting surface of the cabin, while satisfying the constraints of minimizing the number and weight of heat pipes.

[0013] The individual units in the layout model carry thermal model information, including the three-dimensional model, heat consumption, and installation method.

[0014] Preferably, in the parameter modeling step, a weighted Pareto front algorithm is used to automatically optimize the heat dissipation surface area, including: Construct the utility objective function:

[0015] in, Indicates the temperature compliance rate; This represents the normalized average high-temperature margin. This represents the normalized heater compensation power consumption; , , They represent , , The weight of A; A represents the heat dissipation area.

[0016] Based on the single-machine thermal interface data in the digital demand model, within a given maximum and minimum heat dissipation surface area... , Internal computation enables the utility objective function Maximize the optimal heat dissipation area .

[0017] Preferably, the parametric simulation model includes an external heat flow model of the spacecraft, a single-unit heat generation model inside the spacecraft, and a heat transfer model of the thermal control components.

[0018] The establishment of the parametric simulation model includes extracting the orbital parameters and single-machine thermal interface parameters from the SysML model, exporting the thermal analysis parameter table, and inputting it into the simulation software.

[0019] When the SysML model parameters are updated, the thermal simulation model is updated through the parameter table.

[0020] Preferably, in the parameter modeling step, typical operating conditions covering the entire life cycle of the spacecraft are set for simulation, including orbit insertion phase operating conditions, on-orbit normal operation operating conditions, and fault operating conditions.

[0021] Preferably, the detailed technical documents include a summary table of thermal control materials and components, a three-dimensional layout diagram, and technical requirements documents.

[0022] Preferably, after the detailed design step, the method further includes: Identify and archive common modules with preset standardized interfaces and configurable parameters from at least one model selected from the following model groups for reuse in the research and development of thermal control subsystems for different types of spacecraft.

[0023] The digital requirements model, the design model corresponding to the final design scheme, and the parametric simulation model.

[0024] Preferably, in the scheme design step, a data association is established between the digital requirement model and the generated design model.

[0025] In the parametric modeling step, a data association is established between the design model and the established parametric simulation model to achieve traceability from requirements to the final design solution.

[0026] A spacecraft thermal control subsystem design apparatus according to the present invention includes: at least one processor; and a memory storing computer-executable instructions thereon, wherein when the instructions are executed by the at least one processor, the apparatus performs the spacecraft thermal control subsystem design method based on the MBSE model.

[0027] According to the present invention, a computer-readable storage medium is provided thereon storing a computer program, which, when executed by a processor, implements the aforementioned design method for a spacecraft thermal control subsystem based on the MBSE model.

[0028] Compared with the prior art, the present invention has the following beneficial effects: 1. This invention, through comprehensive simulation verification and closed-loop iterative optimization of the design scheme during the parameter modeling stage, can identify and resolve potential design defects before product launch, effectively replacing some expensive and time-consuming physical tests, significantly reducing the number of physical prototype tests, thereby shortening the development cycle and reducing R&D costs.

[0029] 2. This invention constructs a requirement model during the task analysis phase and ensures its consistency and traceability in subsequent design, simulation, and verification stages, forming a complete traceability chain from requirements to design to verification. This ensures that all design activities revolve around the initial requirements, avoids distortion in requirement transmission, and improves forward design capabilities and design reliability.

[0030] 3. By constructing standardized requirements analysis models, functional analysis models, and design integration models, this invention enables the general modules to be directly reused in the research and development of different types of spacecraft. Only local adjustments and verifications are needed according to the new mission requirements to quickly generate new solutions, which significantly improves the reusability of research and development results and accelerates the development progress.

[0031] 4. This invention takes a unified model-based system engineering model as its core, integrates information from the entire process, including requirements, functions, design, simulation, and process, and realizes seamless connection and collaborative management of data in each stage. Any change in the model can be automatically or semi-automatically transmitted to related stages, effectively avoiding the data inconsistency problem in the traditional document-based mode, forming a closed loop of design-verification-optimization, and reducing design risks. Attached Figure Description

[0032] Other features, objects, and advantages of the invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings: Figure 1 A flowchart illustrating the design methodology for a spacecraft thermal control subsystem based on the MBSE model; Figure 2 This is a schematic diagram of the functional breakdown structure of the thermal control subsystem in an embodiment of the present invention; Figure 3 This is a schematic diagram of the three-dimensional heat pipe layout in an embodiment of the present invention; Figure 4 This is a schematic diagram of the temperature field distribution cloud map of the entire star in an embodiment of the present invention; Figure 5 This is a schematic diagram of the layout and circuit principle of the three-dimensional heater in an embodiment of the present invention. Detailed Implementation

[0033] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0034] The present invention provides a spacecraft thermal control subsystem design method based on the MBSE model, which can be executed by a spacecraft thermal control subsystem design device. This device can be one or more computer devices, such as workstations, servers, or distributed computing systems, which typically include at least one processor and memory. The memory stores computer-executable instructions, and when the processor executes these instructions, the aforementioned model-based systems engineering spacecraft thermal control subsystem design method can be implemented.

[0035] Taking the thermal control subsystem design of a communication satellite operating in low Earth orbit as an example, by deeply integrating the system modeling language tool (SysML), 3D computer-aided design tools, and computer-aided engineering thermal simulation analysis tools in a unified, model-based design environment, we have achieved closed-loop optimization of the entire process from requirement definition to product manufacturing through model-based management and simulation-driven optimization.

[0036] by Figure 1 For example, the overall process of the method includes: task analysis steps, scheme design steps, parameter modeling steps, and detailed design steps that are connected before and after.

[0037] In the mission analysis step, the input is a text-formatted development mission specification for the low-Earth orbit environment satellite thermal control subsystem issued by the spacecraft overall design unit. This mission specification includes top-level requirements for the thermal control subsystem, such as: the satellite's operating orbit is a sun-synchronous orbit at an altitude of 700 km and an inclination of 98°; the total thermal power consumption of the satellite should not exceed 800W under the worst operating conditions, with the communication payload serving as the primary heat source, generating 400W of heat; the operating temperature of all onboard equipment must be maintained within the range of -10℃ to +50℃; the total weight of the thermal control subsystem must not exceed 60 kg; and the external power supply interface is a standard 28V DC power supply with an allowable fluctuation range of ±5%.

[0038] Specifically, in a model-based systems engineering platform, a system modeling language is used to deconstruct the aforementioned textual requirements into a structured and itemized model. For example, the performance requirement of "single-unit operating temperature -10℃ to +50℃" is transformed into a traceable performance requirement object with a unique ID, and then associated with all electronic components in the spacecraft model. Similarly, "total weight ≤ 60kg" is defined as a constraint requirement, while "power supply interface 28V±5%" is defined as an interface requirement.

[0039] In this way, all unstructured textual requirements are transformed into machine-readable, clearly defined digital requirement models.

[0040] In many preferred examples, the task analysis steps also include functional analysis and behavioral modeling to achieve a more comprehensive understanding of the system. Figure 2 Taking this as an example, the functional decomposition structure of the thermal control subsystem is illustrated. The top-level function is decomposed into sub-functions such as heat collection, heat transfer, heat dissipation, and heat compensation.

[0041] These functions are further mapped to specific physical behaviors. For example, heat transfer aims to achieve temperature equalization across the entire satellite, heat dissipation aims to cool critical areas, and thermal compensation is used to raise the temperature under cryogenic conditions. The analysis results of these functions and behaviors are also embedded in the model. For instance, activity diagrams or sequence diagrams from system modeling languages ​​are used to describe the execution flow of thermal control functions under different operating modes (such as orbit insertion mode, normal operation mode, and safety mode). Finally, the mission analysis steps output a series of model diagrams, including but not limited to requirement diagrams, function diagrams, interface diagrams, state machine diagrams, and behavior diagrams, which together constitute a unique and traceable design baseline for all subsequent design, simulation, and verification activities. These models collectively constitute the requirement analysis model and the functional analysis model.

[0042] In the aforementioned scheme design step, the design team, based on the digital requirement model and functional analysis model output from the mission analysis step, and referring to relevant national standards and industry specifications for spacecraft thermal control design, conducts a preliminary scheme design for the thermal control subsystem. This step aims to transform abstract functional requirements into concrete physical implementation schemes.

[0043] Specifically, it includes the following core design activities: Heat dissipation surface design: Based on the characteristics of low-orbit satellites rapidly flying over the Earth's illuminated and shadowed areas, and the partial shading of the outer surface by the solar panels, main heat dissipation surfaces are set on the side plates of the spacecraft in the +X and -X directions.

[0044] Preliminary calculations determined that the area of ​​each heat dissipation surface is 4.2 square meters. To achieve efficient heat dissipation, a polyimide aluminized film with excellent thermo-optical properties was selected as the thermal control coating for the heat dissipation surface, with the following technical requirements: solar absorptivity α. s =0.14, hemispherical emissivity ε=0.86.

[0045] Temperature uniformity design: To efficiently transfer the concentrated heat generated by high-heat-dissipation units such as communication payloads to the heat dissipation surface, the design scheme employs heat pipes (such as... Figure 2 As shown, heat is the primary means of heat transfer.

[0046] When designing for temperature uniformity, the system automatically generates the heat pipe layout using the optimization principle of heat pipe layout. The optimization principle includes: ensuring that heat pipes pass under each unit mounting surface on the deck to form an isothermal heat network, while satisfying the conditions of minimum number of heat pipes and minimum weight.

[0047] Specifically, with Figure 3 For example, six high thermal conductivity heat pipes with a diameter of 8 mm and using ammonia as the working fluid were designed. These heat pipes were embedded or pasted onto the substrate of the spacecraft. Their layout was carefully designed, with one end in close contact with the main heat-generating unit and the other end extending to the heat dissipation surface area, thus forming an efficient heat transfer path.

[0048] Thermal compensation design: To cope with the low-temperature environment that may occur when the satellite enters the Earth's shadow region or during winter orbit, a thermal compensation scheme needs to be designed. Specifically, based on thermodynamic analysis, eight independent heating loops are configured, each with a rated power of 30W, and a total compensation power limit of 240W. Polyimide film type electric heaters (such as...) are selected. Figure 2 As shown), Figure 5 As shown, this heater has the advantages of being lightweight, flexible, and easy to install.

[0049] Multi-layer thermal insulation component design: All external surfaces of the spacecraft, except for the heat dissipation surface, as well as parts that need to be thermally isolated from the internal structure, are covered with multi-layer thermal insulation components to minimize the leakage of external heat flow and the unnecessary loss of internal heat.

[0050] In more preferred embodiments, a multilayer thermal insulation component with aluminized polyester film supported by 12 layers of embossed dots is designed.

[0051] After completing the preliminary design, these design schemes are parameterized and updated into a unified system model, forming a design model corresponding to the preliminary scheme design. During this process, each unit in the layout model carries complete thermal model information. When the layout or thermal interface parameters of a unit are updated, this information is synchronously saved in the comprehensive design model.

[0052] For example, generating a heat dissipation surface design chart that includes the orientation, precise area, and thermal control coating model, and drawing it in a 3D computer-aided design model, such as... Figure 3 The system displays a 3D heat pipe layout diagram and generates a heater design diagram that includes the number of heater paths, power, and preliminary layout. Simultaneously, the functional diagrams, interface diagrams, and component diagrams in the system model are also updated to reflect these newly added physical components and their connections. For example, the interface diagram clearly defines the electrical interfaces between the heaters and the power supply subsystem.

[0053] This series of computer-aided design models and related design documents together constitute the comprehensive design model. Furthermore, during the scheme design phase, a clear data link is established between the digital requirements model and the comprehensive design model through a model-based systems engineering platform. For example, the requirement of "temperature control accuracy ±2℃" is linked to the design parameters of the thermistor and temperature controller, achieving initial traceability from requirements to design.

[0054] The core objective of the parametric modeling step is to conduct comprehensive performance simulation verification of the preliminary scheme generated in the scheme design step. This step takes the spacecraft's three-dimensional computer-aided design model, which includes thermal design schemes, and related design parameters output from the scheme design stage as input to establish an integrated, parametric thermal simulation analysis model.

[0055] Specifically, the modeling process includes: Model creation: Import the three-dimensional computer-aided design model of the entire satellite, including the layout of thermal control components such as heat pipes and heaters, into professional spacecraft thermal analysis software (such as Thermal Desktop).

[0056] Model parameter definition: In this software, a comprehensive parametric simulation model is established.

[0057] Specifically, the system extracts the necessary input conditions for thermal analysis integrated into the SysML model (such as orbital parameters, annual maximum / minimum illumination angles, and thermal interface information for each unit, including heat consumption, weight, surface condition, and installation contact area), and exports a thermal analysis parameter table with one click. This table is then input into the thermal simulation software, which automatically generates the model's thermophysical parameters without requiring manual configuration. When parameters in the SysML model are updated, the thermal simulation model can be updated synchronously with one click, achieving a dynamic association between the SysML model and the thermal simulation model.

[0058] The model includes at least: an external heat flow model for the spacecraft, which accurately calculates the transient heat flow to the spacecraft's outer surface from direct solar radiation, Earth's albedo, and Earth's infrared radiation based on the satellite's orbital parameters (700km altitude, 98° inclination) and attitude control law; an internal heat generation model for individual components of the spacecraft, which sets each component's power consumption (e.g., 400W for the communication payload) as a constant or time-varying heat source based on the mission specifications; and a heat transfer model for thermal control components, which accurately simulates the equivalent thermal conductivity of heat pipes, the equivalent thermal conductivity of multilayer insulation components, and the thermo-optical properties of thermal control coatings.

[0059] Simulation Conditions: To comprehensively evaluate the performance of the design scheme, multiple typical operating conditions covering the entire life cycle of the spacecraft were set up for simulation analysis. These conditions include, but are not limited to: the orbital insertion phase after launch, when the satellite has not yet fully deployed and some equipment is not powered on; the normal operation in orbit, further subdivided into the summer highest orbital temperature condition (maximum Beta angle) and the winter lowest orbital temperature condition (minimum Beta angle), considering both the illuminated and shadowed areas; and possible fault conditions, such as abnormal shutdown of a high-power unit or attitude instability leading to direct sunlight on non-heat-dissipating surfaces.

[0060] After setting up, run the simulation analysis. After the simulation is complete, the software outputs detailed analysis results, such as... Figure 4 The image shows the temperature field distribution cloud map of the entire star. Figure 4 The 3D model of the spacecraft displays different colors on its surface, and the temperature graph on the left provides a clear visual representation of the temperature distribution of each component. Furthermore, it generates temperature curves of key individual units over time, a thermal balance calculation table for the entire satellite, and a summary table of individual unit temperatures.

[0061] Subsequently, a compliance analysis is conducted. The simulation output, including the overall satellite temperature field distribution and individual unit temperature summary table, is automatically or semi-automatically compared with the performance requirements in the digital requirement model defined in the mission analysis steps.

[0062] In the initial simulation of the embodiment, the results showed that under the worst summer high-temperature conditions, the highest temperature of the communication load was 45℃; under the worst winter low-temperature conditions, the lowest temperature of a certain edge unit was -5℃, at which point the total compensation power of the automatically activated heater was 120W. Comparing these results with the required specifications (single unit temperature -10℃ to +50℃, compensation power budget ≤ 150W), it was found that all specifications were within the required range and had a certain margin.

[0063] Accordingly, the conclusion of the compliance analysis is "compliant". The system determines that there is no need to return to the design steps for optimization and iteration, and the current design scheme is confirmed as the final design scheme. In the model-based systems engineering platform, the state of this design scheme is officially locked and associated with the successful simulation results, forming a complete "design-verification" evidence chain.

[0064] In the final detailed design step, it should be noted that the thermal simulation model is mainly used to calculate temperature distribution and heat loss compensation to verify the compliance of performance indicators; it does not directly generate technical requirement documents. After the indicators are verified, the confirmed thermal control scheme parameters (such as heat pipe layout, heat dissipation surface distribution, and measuring point distribution) will be finalized in the 3D thermal control model. Based on this finalized 3D thermal control model and the accompanying technical requirement specifications, the system extracts key data, aiming to transform the design parameters into detailed technical documents that can be used for manufacturing, assembly, and testing.

[0065] The specific output includes: Summary Table of Thermal Control Materials and Components: This table details all thermal control products that need to be purchased or manufactured, including the specific grade and supplier information of the thermal control coating; the precise dimensions, working fluid, and bending radius requirements of the heat pipe; the model of the heater (e.g., selecting the H-30-28 polyimide film heater from supplier B), resistance value, and dimensions; the model of the thermistor (e.g., selecting the NTC-10K type, requiring a temperature measurement accuracy of ±0.5℃) and the supplier; and the material layers and density of the multilayer thermal insulation components.

[0066] 3D Layout: Outputs a 3D model and 2D drawings containing precise coordinates and installation tolerances. For example, ... Figure 3 The 3D heat pipe layout diagram shown will be accompanied by detailed dimension annotations, clearly indicating the precise route of each heat pipe on the substrate and the position of the mounting holes, with the installation tolerance required to be controlled within ±0.1mm.

[0067] like Figure 5 The 3D heater layout diagram shown precisely defines the mounting position and orientation of each heater on the equipment mounting plate. Compensating heating units A and B each have four heating circuits: A1 main, A2 main, A3 main, A4 main, and B1 main, B2 main, B3 main, B4 main. Additionally, a 3D thermistor layout diagram and a 3D multi-layer insulation assembly outline and opening definition diagram are included.

[0068] Technical Requirements Documents: Detailed technical requirements documents are prepared for key products. For example, the technical requirements document for thermal control coatings will specify that the coating adhesion is not less than 5 MPa, the high and low temperature resistance range is -196℃ to +150℃, and clearly define the acceptance test methods. The technical requirements document for heaters will specify their power deviation at rated voltage, insulation resistance not less than 100 MΩ, etc.

[0069] In more preferred embodiments, the thermal control subsystem design of this low-Earth orbit communication satellite demonstrated significant advantages through the aforementioned method. According to project statistics, the entire design cycle was shortened from the traditional 12 months to 8 months, and R&D costs were reduced by approximately 40%. This was primarily due to the replacement of at least two physical thermal balance tests with simulation in the parameter modeling step, greatly saving time and money. During the final product's on-orbit operation, telemetry data showed that the temperature of all individual units remained stable within the expected range of -5℃ to +45℃, highly consistent with the simulation predictions, fully validating the effectiveness and high reliability of the design method provided by this invention.

[0070] Compared to low-Earth orbit satellites, geostationary orbit satellites have a more stable thermal environment, but they also experience prolonged exposure to sunlight on one side of their surface and shadow periods of up to 72 minutes. In addition, their payloads typically consume more power and have more stringent requirements for temperature control precision.

[0071] In more advanced examples, this technology is applied to the thermal control subsystem design of a high-power communication satellite in geostationary orbit. During the mission analysis phase, the input development task specification requires the satellite to operate in a geostationary orbit at a specified east longitude, with a total thermal power consumption of up to 2000W. The core communication payload module consumes 1500W, and its internal temperature must be kept stable within a narrow range of 20℃±5℃. Similarly, a system modeling language is used to transform these more stringent requirements into a digital requirement model.

[0072] In the design process described, an adaptive design was implemented specifically for the characteristics of geostationary orbit satellites. Due to the stable thermal environment, the heat dissipation surfaces were designed as fixed radiators facing north and south, which are almost completely shielded from direct sunlight. To address the high heat dissipation of 2000W and the high temperature uniformity requirements of the payload compartment, the temperature homogenization design no longer uses ordinary heat pipes, but instead selects a loop heat pipe system with stronger heat transfer capacity and longer heat transfer distance. The evaporator of this loop heat pipe system is integrated with the mounting plate of the payload compartment, while the condenser is connected to the north and south radiators, forming a highly efficient heat "transfer" loop. In terms of thermal compensation design, based on heat loss calculations during a 72-minute shadow period, the total power of the heaters was increased, and their layout in key components and pipelines was optimized to achieve precise heat preservation during the "dormant" period.

[0073] In the parameter modeling step, the external heat flow model was set as a geosynchronous orbit model, which included accurate calculations of long shadow periods and seasonal variations in solar incidence angle (summer / winter solstice). The simulation analysis focused not only on verifying the overall satellite temperature distribution but also on verifying whether the startup characteristics, operational stability, and maximum heat transfer capacity of the loop heat pipe system met the design requirements. Simultaneously, a long shadow period operating condition was specifically set up to verify whether heater compensation alone could maintain all individual units above their survival temperature.

[0074] Simulation results show that under the summer solstice conditions, sunlight shines obliquely onto the north-facing radiator at a certain angle, causing the radiator temperature to rise, which in turn reduces the efficiency of the condenser section of the connected loop heat pipe, and the load chamber temperature exceeds the upper limit of 25°C.

[0075] Based on these simulation results, the design process was revisited for optimization. As an alternative approach, instead of simply increasing the heatsink area (as this would increase weight and complicate mechanical design), the thermal control coating distribution on the north-facing heatsink was optimized in the overall design model: in the edge region near the oblique solar radiation, the original single thermal control coating was replaced with a special coating (such as a quartz secondary surface mirror) with lower solar absorptivity / higher emissivity in some areas. After modifying the design model, the parametric modeling process was repeated for iterative verification.

[0076] New simulation results show that, after coating optimization, the maximum temperature of the load chamber under summer solstice conditions dropped to 24.5℃, meeting the requirement of 20℃±5℃. Thus, the final design scheme was determined.

[0077] The detailed design steps, in which the output detailed technical documents highlight the specific requirements of the loop heat pipe system. For example, the 3D layout diagram not only includes the precise routing of the pipes but also defines in detail the flatness and thermal resistance requirements of the contact surface between the evaporator and the equipment mounting plate, as well as the connection method between the condenser and the radiator. The technical requirements document also includes the electrical interface definition and control logic description for the loop heat pipe drive controller (such as the liquid receiver temperature control unit).

[0078] This design method can effectively address the complex thermal control design requirements of high-power, long-life geostationary orbit satellites. It not only supports the design of conventional thermal control technologies, but also models, simulates, and iteratively optimizes more complex thermal control technologies such as loop heat pipes, thereby ensuring that the design scheme reaches a high degree of maturity and reliability before investing in expensive physical manufacturing and testing.

[0079] In many preferred examples, the entire process is led by engineers and involves interaction with computer-aided design models and thermal simulation software.

[0080] Initial Simulation and Problem Exposure: A simulation of the most severe high-temperature operating conditions was run in thermal simulation software, based on a computer-aided design model including a 3mm thick heat spreader. After the simulation, the software returned the results. Engineers viewed the temperature contour plots and data reports generated by the simulation to... Figure 4 For example, it was found that the highest local temperature at the communication payload installation point reached 58℃, which exceeded the upper temperature limit of 50℃ specified in the task specification. At this time, the result of the indicator compliance analysis was "non-compliant".

[0081] Problem Analysis and Localization: Utilizing the powerful post-processing capabilities of the simulation software, an in-depth analysis of the simulation model was conducted. Observation of the heat flux density diagram revealed a significant "thermal resistance bottleneck" in the path of the 400W heat generated by the load to the heat pipe. In other words, the 3mm thick heat spreader plate lacked sufficient thermal conductivity, causing heat to accumulate in the area directly beneath the load and preventing rapid and uniform diffusion to the area connected to the heat pipe, thus creating localized hotspots.

[0082] Solution Optimization and Decision-Making: Based on the problem identification, return to the solution design steps and modify the comprehensive design model. At this point, the system can assist engineers in proposing multiple optimization ideas: Option A: Enhance local thermal conductivity. Increase the thickness of the heat spreader plate under the communication load from 3mm to 5mm. This is a direct improvement, but the downside is increased structural weight.

[0083] Option B: Increase the heat transfer path. While keeping the heat spreader plate thickness unchanged at 3mm, add an auxiliary heat pipe near the load mounting point. This heat pipe directly connects the load mounting area and the main heat dissipation surface.

[0084] Iterative verification and scheme validation: Two branch models corresponding to scheme A and scheme B were created in the computer-aided design model. Subsequently, the parametric modeling step was restarted to perform parametric modeling and simulation verification on the two optimized models again.

[0085] Simulation results for Scheme A show that the maximum payload temperature drops to 48℃, meeting the requirement of being below 50℃. Calculations indicate that this scheme results in a 0.5kg increase in satellite weight.

[0086] Simulation results for scheme B show that the maximum load temperature is reduced to 46℃, which is better than scheme A. Calculations show that this scheme only adds 0.2kg of weight (the weight of a thin heat pipe).

[0087] When iterating through the design process, to avoid manual trial and error, in a preferred embodiment, automatic optimization is performed based on the weighted Pareto front algorithm. Taking the search for the optimal heat dissipation surface area as an example, the mathematical modeling process includes: Define the design variable as heat dissipation area A (which can be a single variable or a combination of multiple heat dissipation surfaces).

[0088] Three objective functions are constructed based on the simulation output: Temperature compliance rate The higher this value, the better; Average value of high temperature margin The higher this value, the better; Heater compensation power consumption The smaller this value, the better.

[0089] Where N represents the total number of hot interfaces; i represents the hot interface ordinal number; Indicates the highest temperature of thermal interface i; This indicates the actual temperature of the thermal interface i.

[0090] Based on design experience, the above target values ​​are assigned weights respectively. , , and to and Perform maximum and minimum value normalization:

[0091]

[0092] in, , These represent the normalized average high-temperature margin and the heater compensation power consumption, respectively. , These represent the minimum and maximum values ​​of the average high-temperature margin, respectively. , These represent the minimum and maximum power consumption of the heater compensation, respectively.

[0093] No normalization is needed. The theoretical range can be calculated based on the single-machine thermal interface summary table (temperature index) in the SysML model. The maximum and minimum values ​​can be obtained through thermal design conditions. After normalization, all three indicators meet the standard that the larger the number, the better.

[0094] Then, a utility objective function is constructed. :

[0095] The system uses a given maximum and minimum heat dissipation surface area. , Based on MBSE system model data, the solution is automatically optimized:

[0096] This allows for the rapid attainment of the optimal heat dissipation surface area. Design parameters, etc.

[0097] After comprehensively weighing performance margins and weight costs, Scheme B was ultimately selected as the final design. This cyclical process of "modifying design parameters, running simulations, obtaining results, and determining whether the requirements are met" clearly demonstrates the closed loop of design-verification-optimization.

[0098] Entering the detailed design stage: Based on the final confirmed solution B, the design process smoothly enters the detailed design stage, generating various documents for production and manufacturing.

[0099] By conducting rapid simulation iterations in a purely digital model environment early in the design process, potential design performance bottlenecks can be identified and resolved at extremely low cost. This effectively avoids the huge costs and time delays caused by discovering problems only after a physical prototype has been manufactured, through expensive and time-consuming physical thermal equilibrium tests, which necessitates major design changes or even restarting production.

[0100] In more preferred examples, after the development of a low-orbit satellite has been completed, a low-orbit satellite with similar configuration but carrying different mission payloads is developed. The same satellite platform is used, but its core payload is changed from a communication payload to a high-power optical camera, its heat dissipation power is changed to 500W, and a data processing unit is added.

[0101] Compared to traditional design methods that require starting most of the design work from scratch, this approach significantly improves the development efficiency of serialized and platform-based spacecraft by building a reusable model library, supporting the technical characteristics of model reuse. Specifically, this includes: Model Inheritance and Baseline Establishment: A reusable baseline model library with specific inheritance and variant mechanisms is established. When a new task arises, the entire set of designed system models (including requirements analysis models, functional analysis models, and design synthesis models) is completely replicated within the model-based systems engineering platform to form a new design baseline. This baseline model already includes a validated thermal control design scheme for the satellite platform. Subsequent modifications only require instantiation of this baseline model, enabling rapid reuse.

[0102] Differentiation Analysis and Requirements Update: In the task analysis step, the new development task specification is input into the model-based systems engineering platform. The platform can automatically or semi-automatically compare the requirements items in the new task specification with the requirements models in the baseline model. The system will highlight all requirements items with differences, such as "core payload heat dissipation power" changing from 400W to 500W, or "new single unit: data processing unit, power consumption 50W". Simultaneously, since these requirements have already established traceability relationships in the completed design model, the system can automatically identify all downstream functional modules, physical components, and design parameters affected by these requirement changes.

[0103] Local Modifications and Model Evolution: Instead of redesigning the entire model, only the highlighted differences need to be modified locally. The following operations were performed during the scheme design phase: In the computer-aided design model for designing the comprehensive model, a three-dimensional model of the data processing unit was added, and its installation location was planned.

[0104] As the core load heats up, the layout of the heat pipes connected to the new load is adjusted, for example, by increasing the number of main heat pipes from two to three, or by using heat pipes with a larger diameter.

[0105] Add an independent heating circuit to the newly added data processing unit and update the relevant information in the heater design diagram.

[0106] Most of the satellite platform's design, such as the heat dissipation surface design, most structural components, and multi-layer thermal insulation components, can be directly inherited and reused without any modification, as the model is not affected by changes in requirements.

[0107] Rapid Verification and Validation: In the parametric modeling step, since most thermal models can be directly reused, only the changed parts need to be updated based on the completed parametric simulation model: update the heating power value of the core load to 500W, and add the thermal model of the newly added data processing unit and its 50W heat output. Rerun the simulation analysis covering typical operating conditions. The simulation results may show new temperature distributions, based on which the heat pipe layout or heater control strategy can be fine-tuned. After one or two rapid iterations, a final design scheme that meets all requirements can be obtained.

[0108] By using this model reuse approach, the design cycle of the new thermal control subsystem is expected to be shortened by more than 50% compared to designing from scratch. The method provided by this invention is not merely an optimization of the design process for a single model, but a systems engineering practice capable of establishing an accumulative, evolving, and reusable design knowledge base.

[0109] By solidifying proven design solutions into standardized and modular models, the development process of subsequent spacecraft on the same platform or in series can be greatly accelerated, enabling knowledge transfer and significantly increasing efficiency. For example, standardized heater temperature control logic models and equivalent heat transfer models of multi-layer insulation components are general-purpose modules with preset standardized interfaces and configurable parameters. These modules can be easily archived and reused in the development of thermal control subsystems for different types of spacecraft.

[0110] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, making the system and its various devices, modules, and units of this invention function as logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0111] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A design method for a spacecraft thermal control subsystem based on the MBSE model, characterized in that, include: Task analysis steps: Based on the development task specification of the spacecraft thermal control subsystem, the textual requirements are transformed into a structured digital requirement model, which includes the functions, performance, interfaces and constraints related to the thermal control subsystem. Scheme design steps: Based on the digital requirement model, conduct a preliminary scheme design for the spacecraft thermal control subsystem, and generate a design model corresponding to the preliminary scheme design; Parametric modeling steps: Based on the design model, establish a parametric simulation model for performance verification; By running the simulation, the simulation results are compared with the performance requirements defined in the digital demand model. If the simulation results do not meet the performance requirements, the process returns to the scheme design step to optimize and iterate the preliminary scheme design until the simulation results meet the performance requirements. The preliminary scheme design that meets the performance requirements at this point is then determined as the final design scheme. Detailed design steps: Based on the three-dimensional thermal control model and technical requirements corresponding to the final design scheme, generate detailed technical documents for production and manufacturing.

2. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, In the task analysis step, the digital requirement model is established using the SysML language; Build a reusable baseline model library that includes at least one baseline model, and make differentiated modifications to the baseline models when performing new task designs; The design steps of the scheme include heat dissipation surface design, temperature uniformity design, thermal control compensation design, and multi-layer thermal insulation component design. The temperature uniformity design includes automatic generation of heat pipe layout, following the principle of optimization: An isothermal heat network is formed under each individual unit mounting surface of the cabin, while satisfying the constraints of minimum number of heat pipes and minimum weight. The individual units in the layout model carry thermal model information, including the three-dimensional model, heat consumption, and installation method.

3. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, In the parameter modeling step, a weighted Pareto front algorithm is used to automatically optimize the heat dissipation surface area, including: Construct the utility objective function: in, Indicates the temperature compliance rate; This represents the normalized average high-temperature margin. This represents the normalized heater compensation power consumption; , , They represent , , The weights; A represents the heat dissipation area; Based on the single-machine thermal interface data in the digital demand model, within a given maximum and minimum heat dissipation surface area... , Internal computation enables the utility objective function Maximize the optimal heat dissipation area .

4. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, The parametric simulation model includes the external heat flow model of the spacecraft, the internal heat generation model of individual components of the spacecraft, and the heat transfer model of the thermal control components. The establishment of the parameterized simulation model includes extracting the orbital parameters and single-machine thermal interface parameters from the SysML model, exporting the thermal analysis parameter table and inputting it into the simulation software; When the SysML model parameters are updated, the thermal simulation model is updated through the parameter table.

5. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1 or claim 4, characterized in that, In the parameter modeling step, typical operating conditions covering the entire life cycle of the spacecraft are set for simulation. The typical operating conditions include the orbit insertion phase, the normal on-orbit operation, and the fault condition.

6. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, The detailed technical documents include a summary table of thermal control materials and components, a three-dimensional layout diagram, and technical requirements documents.

7. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, Following the detailed design steps, the method further includes: Identify and archive common modules with preset standardized interfaces and configurable parameters from at least one model selected from the following model groups for reuse in the research and development of thermal control subsystems for different types of spacecraft; The digital requirements model, the design model corresponding to the final design scheme, and the parametric simulation model.

8. The spacecraft thermal control subsystem design method based on the MBSE model according to claim 1, characterized in that, In the solution design step, a data association is established between the digital requirements model and the generated design model; In the parametric modeling step, a data association is established between the design model and the established parametric simulation model to achieve traceability from requirements to the final design solution.

9. A design device for a spacecraft thermal control subsystem, characterized in that, include: At least one processor; as well as A memory having computer-executable instructions thereon, which, when executed by the at least one processor, cause the device to perform the spacecraft thermal control subsystem design method based on the MBSE model as described in any one of claims 1 to 8.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the computer program is executed by the processor, it implements the spacecraft thermal control subsystem design method based on the MBSE model as described in any one of claims 1 to 8.

Images