Spacecraft thermal management method and device and storage medium

Abstract

The application discloses a space-aircraft thermal management method, device and storage medium, relates to the field of aircraft technology, and discloses a space-aircraft thermal management method, which comprises the following steps: acquiring a first fuel coolant temperature and a second fuel coolant temperature in real time; determining a first target fuel coolant flow and a second target fuel coolant flow based on the first fuel coolant temperature and the second fuel coolant temperature; and controlling a first centrifugal pump based on the first target fuel coolant flow and controlling a second centrifugal pump based on the second target fuel coolant flow. The first centrifugal pump and the second centrifugal pump are controlled by the first fuel coolant temperature and the second fuel coolant temperature to adjust the fuel coolant flow in a fuel pipeline, so that the temperature of the space-aircraft whole machine is controllable and the temperature gradient is arranged, the temperature of on-board equipment and / or a part to be cooled is actively reduced, the passive thermal protection burden is effectively reduced, and the reusability of the space-aircraft is improved.

Description

Technical Field

[0001] This application relates to the field of aircraft technology, and in particular to a thermal management method, apparatus and storage medium for aerospace vehicles. Background Technology

[0002] With the continuous development of aerospace science and technology, the flight environment of the new generation of aircraft has changed significantly, no longer limited to a single airspace or spacespace, but exhibiting the characteristics of cross-domain flight. Compared with traditional aircraft or spacecraft, the flight environment of aerospace vehicles is complex and variable, with a flight envelope covering a large airspace of 0-100km and a wide speed range of 0-30Ma, and the maximum heat flux density can reach MW / m³. 2 At this level, the pneumatic heating problem is obvious.

[0003] Currently, passive thermal protection technologies such as ablation thermal protection and passive radiation thermal protection are mainly used to address aerodynamic heating problems. Ablation thermal protection absorbs heat through the ablation of the material surface, forming a heat insulation layer that provides cooling and insulation. However, the aerodynamic shape of the aircraft changes after ablation, which is detrimental to the reusability of aerospace vehicles. Passive radiation thermal protection uses advanced heat-resistant materials such as carbon-carbon composites and ceramics to dissipate heat through thermal radiation. However, heat-resistant materials also have a defined temperature limit and cannot cope with extreme heat flow environments on their own. Carbon-carbon composites will oxidize and fail in high-temperature oxidizing environments, and high-temperature resistant materials such as ceramics have high density and brittleness, reducing the payload capacity of the aircraft and making it susceptible to mechanical impact damage.

[0004] Therefore, how to reduce the burden of passive thermal protection and improve the thermal protection capability and reusability of aerospace vehicles is an urgent problem that needs to be solved. Summary of the Invention

[0005] The main objective of this application is to provide a thermal management method, device, and storage medium for aerospace vehicles, aiming to solve the technical problems of how to reduce the burden of passive thermal protection, improve the thermal protection capability of aerospace vehicles, and increase their reusability.

[0006] To achieve the above objectives, this application proposes a thermal management method for an aerospace vehicle, the aerospace vehicle including airborne equipment, a heat dissipation area, fuel lines, a fuel tank, a first centrifugal pump, a second centrifugal pump, and an engine combustion chamber; each airborne equipment is equipped with a first heat exchanger and a second heat exchanger; the heat dissipation area is equipped with a microchannel structure; the fuel lines include a first line, a second line, and a third line;

[0007] The second centrifugal pump is installed on the first pipeline, and each of the first heat exchangers and each of the microchannel structures are respectively connected to the first pipeline; one end of the second pipeline is connected to the inlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber; the first centrifugal pump is installed on the third pipeline, one end of the third pipeline is connected to the outlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber; the inlet of the second centrifugal pump is connected to the outlet of the first centrifugal pump through the first pipeline and the third pipeline;

[0008] Each second heat exchanger is adapted to a corresponding first heat exchanger. The coolant in the second heat exchanger is used to absorb the heat generated by the airborne equipment, and the fuel coolant in the first heat exchanger is used to absorb the heat from the coolant in the second heat exchanger.

[0009] The thermal management method for the aforementioned aerospace vehicle includes:

[0010] The temperature of the first fuel coolant corresponding to each heat exchanger and the temperature of the second fuel coolant corresponding to each microchannel structure are acquired in real time.

[0011] Based on the first fuel coolant temperature and the second fuel coolant temperature, determine the first target fuel coolant flow rate at the outlet of the first centrifugal pump and the second target fuel coolant flow rate at the outlet of the second centrifugal pump.

[0012] The first centrifugal pump is controlled based on the first target fuel coolant flow rate, and the second centrifugal pump is controlled based on the second target fuel coolant flow rate.

[0013] In one embodiment, the step of determining the first target fuel coolant flow rate at the outlet of the first centrifugal pump and the second target fuel coolant flow rate at the outlet of the second centrifugal pump based on the first fuel coolant temperature and the second fuel coolant temperature includes:

[0014] Obtain the first target temperature corresponding to each airborne device, and the second target temperature corresponding to each part to be cooled;

[0015] Based on the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature, the first target fuel coolant flow rate and the second target fuel coolant flow rate are determined.

[0016] In one embodiment, the step of determining the first target fuel coolant flow rate and the second target fuel coolant flow rate based on the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature includes:

[0017] The first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature are input into the state space equation, and the first target fuel coolant flow rate and the second target fuel coolant flow rate are determined by the output of the state space equation.

[0018] In one embodiment, the step of acquiring the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each microchannel structure in real time includes:

[0019] The temperature of the first fuel coolant is obtained in real time through the temperature sensors at the outlets of each first heat exchanger.

[0020] The temperature of the second fuel coolant is obtained in real time through temperature sensors at the outlets of each microchannel structure.

[0021] In one embodiment, the fuel line is provided with a multi-port connector; the first line, the second line, the third line and the feed line are respectively connected to the multi-port connector.

[0022] In one embodiment, the part to be cooled includes at least the leading edge of the aircraft and the leading edge of the wing.

[0023] In addition, to achieve the above objectives, this application also proposes a thermal management device for a spacecraft, the device comprising: a memory, a processor, and a computer program stored in the memory and executable on the processor, the computer program being configured to implement the steps of the aforementioned thermal management method for a spacecraft.

[0024] In addition, to achieve the above objectives, this application also proposes a storage medium, which is a computer-readable storage medium, on which a computer program is stored, and when the computer program is executed by a processor, it implements the steps of the aforementioned thermal management method for aerospace vehicles.

[0025] One or more technical solutions proposed in this application have at least the following technical effects:

[0026] The thermal management method for aerospace vehicles provided in this application forms a heat dissipation circuit using fuel as a coolant through a heat exchanger, microchannel structure, and fuel pipeline. The fuel coolant in the heat dissipation circuit dissipates heat from the airborne equipment and parts requiring cooling, achieving active thermal protection for the aerospace vehicle. Simultaneously, the first and second centrifugal pumps are controlled by the temperatures of the first and second fuel coolants to regulate the flow rate of the fuel coolant in the fuel pipeline, thereby actively cooling the high-temperature equipment in the airborne equipment and parts requiring cooling. In other words, the highest temperature equipment is cooled first by controlling the flow rate of the fuel coolant, achieving temperature controllability and temperature gradient arrangement of the entire aerospace vehicle, thereby actively reducing the temperature of the airborne equipment and / or parts requiring cooling, effectively reducing the burden of passive thermal protection, and improving the thermal protection capability and reusability of the aerospace vehicle. Attached Figure Description

[0027] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application.

[0028] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, for those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0029] Figure 1 A flowchart illustrating an embodiment of the thermal management method for aerospace vehicles according to this application;

[0030] Figure 2 A schematic diagram of the system architecture provided for an embodiment of the thermal management system of the aerospace vehicle of this application;

[0031] Figure 3 A schematic diagram of a microchannel structure provided for an embodiment of the thermal management system of the aerospace vehicle of this application;

[0032] Figure 4 This is a schematic diagram of the equipment structure of the hardware operating environment involved in the thermal management method of the aerospace vehicle in the embodiments of this application;

[0033] Figure 5 A simplified schematic diagram of a fluid model for a two-node fluid network;

[0034] Figure 6 This is a node diagram of a thermodynamic model in related technologies.

[0035] The purpose, features, and advantages of this application will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0036] It should be understood that the specific embodiments described herein are merely illustrative of the technical solutions of this application and are not intended to limit this application.

[0037] To better understand the technical solution of this application, a detailed description will be provided below in conjunction with the accompanying drawings and specific implementation methods.

[0038] The main solution of this application embodiment is as follows: An aerospace vehicle includes airborne equipment, a heat dissipation area, fuel pipelines, a fuel tank, a first centrifugal pump, a second centrifugal pump, and an engine combustion chamber; each airborne device is equipped with a first heat exchanger and a second heat exchanger; the heat dissipation area is equipped with a microchannel structure; the fuel pipeline includes a first pipeline, a second pipeline, and a third pipeline; the second centrifugal pump is disposed on the first pipeline, and each first heat exchanger and each microchannel structure is connected to the first pipeline; one end of the second pipeline is connected to the inlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber; the first centrifugal pump is disposed on the third pipeline, one end of the third pipeline is connected to the outlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber; the inlet of the second centrifugal pump... The first centrifugal pump outlet is connected via a first pipeline and a third pipeline; each second heat exchanger is adapted to a corresponding first heat exchanger, the coolant in the second heat exchanger is used to absorb the heat generated by the airborne equipment, and the fuel coolant in the first heat exchanger is used to absorb the heat from the coolant in the second heat exchanger; the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each microchannel structure are acquired in real time; based on the first fuel coolant temperature and the second fuel coolant temperature, the first target fuel coolant flow rate at the outlet of the first centrifugal pump and the second target fuel coolant flow rate at the outlet of the second centrifugal pump are determined; the first centrifugal pump is controlled based on the first target fuel coolant flow rate and the second centrifugal pump is controlled based on the second target fuel coolant flow rate.

[0039] In this embodiment, for ease of description, the following description will focus on the thermal management device for identifying aerospace vehicles.

[0040] With the continuous development of aerospace science and technology, the flight environment of the new generation of aircraft has changed significantly, no longer limited to a single airspace or spacespace, but exhibiting the characteristics of cross-domain flight. Compared with traditional aircraft or spacecraft, the flight environment of aerospace vehicles is complex and variable, with a flight envelope covering a large airspace of 0-100km and a wide speed range of 0-30Ma, and the maximum heat flux density can reach MW / m³. 2 At this level, the pneumatic heating problem is obvious.

[0041] Currently, passive thermal protection technologies such as ablation thermal protection and passive radiation thermal protection are mainly used to address aerodynamic heating problems. Ablation thermal protection absorbs heat through the ablation of the material surface, forming a heat insulation layer that provides cooling and insulation. However, the aerodynamic shape of the aircraft changes after ablation, which is detrimental to the reusability of aerospace vehicles. Passive radiation thermal protection uses advanced heat-resistant materials such as carbon-carbon composites and ceramics to dissipate heat through thermal radiation. However, heat-resistant materials also have a defined temperature limit and cannot cope with extreme heat flow environments on their own. Carbon-carbon composites will oxidize and fail in high-temperature oxidizing environments, and high-temperature resistant materials such as ceramics have high density and brittleness, reducing the payload capacity of the aircraft and making it susceptible to mechanical impact damage.

[0042] Therefore, how to reduce the burden of passive thermal protection and improve the thermal protection capability and reusability of aerospace vehicles is an urgent problem that needs to be solved.

[0043] This application provides a solution that forms a heat dissipation circuit using fuel as a coolant through a heat exchanger, a microchannel structure, and fuel pipelines. The fuel coolant in the heat dissipation circuit dissipates heat from the airborne equipment and parts requiring cooling of the spacecraft, achieving active thermal protection for the spacecraft. Simultaneously, the first and second centrifugal pumps are controlled by the temperatures of the first and second fuel coolants to regulate the flow rate of the fuel coolant in the fuel pipelines, thereby actively cooling the high-temperature equipment in the airborne equipment and parts requiring cooling. In other words, the highest temperature equipment is cooled first by controlling the flow rate of the fuel coolant, achieving temperature controllability and temperature gradient arrangement of the entire spacecraft, thereby actively reducing the temperature of the airborne equipment and / or parts requiring cooling, effectively reducing the burden of passive thermal protection, and improving the thermal protection capability and reusability of the spacecraft.

[0044] It should be noted that the executing entity in this embodiment can be a computing service device with data processing, network communication, and program execution functions, such as a tablet computer, personal computer, or mobile phone, or an electronic device capable of performing the above functions, such as a thermal management device for a spacecraft. The following description uses a thermal management device for a spacecraft as an example to illustrate this embodiment and the subsequent embodiments.

[0045] Based on this, embodiments of this application provide a thermal management method for aerospace vehicles, referring to... Figure 1 , Figure 1 This is a flowchart illustrating an embodiment of the thermal management method for aerospace vehicles according to this application.

[0046] In this embodiment, the aerospace vehicle includes onboard equipment, heat dissipation components, fuel lines, a fuel tank, a first centrifugal pump, a second centrifugal pump, and an engine combustion chamber. For example... Figure 2As shown, the airborne equipment includes a first airborne equipment and a second airborne equipment, and the parts to be cooled include the leading edge of the wing and the leading edge of the aircraft. Each airborne equipment is equipped with a first heat exchanger and a second heat exchanger.

[0047] The airborne equipment refers to the equipment within the spacecraft that requires heat dissipation protection, and this airborne equipment is equipped with a heat exchanger. The areas requiring heat dissipation are the parts of the spacecraft that require heat dissipation protection. These areas are equipped with microchannel structures. Specifically, these areas include, but are not limited to, the leading edge (nose) and wing leading edges of the spacecraft. Multiple areas requiring heat dissipation can be located in locations within the spacecraft where aerodynamic heating is significant. The microchannel structure includes multiple fuel coolant circuits, and the shape of the microchannel structure is adaptively designed according to the shape of the areas requiring heat dissipation. Figure 3 As shown, Figure 3 The microchannel structure is ring-shaped and has multiple semi-circular fuel coolant circuits. One end of the circuit is the fuel coolant inlet, and the other end is the fuel coolant outlet.

[0048] The fuel line is equipped with a centrifugal pump, which includes a first centrifugal pump and a second centrifugal pump. The first centrifugal pump draws fuel coolant from the fuel tank into the fuel line. The flow rate of fuel coolant drawn into the fuel line can also be adjusted according to the temperature of the airborne equipment and the parts to be cooled.

[0049] The fuel pipeline includes a first pipeline, a second pipeline, and a third pipeline. A second centrifugal pump is disposed on the first pipeline, and each of the first heat exchangers and each microchannel structure is connected to the first pipeline. One end of the second pipeline is connected to the inlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber. The first centrifugal pump is disposed on the third pipeline, one end of which is connected to the outlet of the fuel tank, and the other end is connected to the inlet pipe of the engine combustion chamber. The inlet of the second centrifugal pump is connected to the outlet of the first centrifugal pump through the first and third pipelines.

[0050] Specifically, such as Figure 2 As shown, the inlet of the second centrifugal pump is connected to the outlet of the first centrifugal pump through the first pipeline and the third pipeline. The outlet of the second centrifugal pump is connected to the inlet of the first heat exchanger corresponding to the first airborne equipment. The outlet of the first heat exchanger corresponding to the first airborne equipment is connected to the inlet of the first heat exchanger corresponding to the second airborne equipment. The outlet of the first heat exchanger corresponding to the second airborne equipment is connected to the inlet of the microchannel structure corresponding to the leading edge of the wing. The inlet of the microchannel structure corresponding to the leading edge of the aircraft is connected to the outlet of the first heat exchanger corresponding to the first airborne equipment. The outlets of the microchannel structure corresponding to the leading edge of the aircraft and the microchannel structure corresponding to the leading edge of the wing are respectively connected to the feed pipe of the engine combustion chamber. The pipeline connecting the second centrifugal pump, the first heat exchanger and the microchannel structure is the second pipeline.

[0051] In the third pipeline, the fuel coolant (fuel) flows into the combustion chamber of the spacecraft's engine through the feed pipe. A portion of the fuel coolant in the second pipeline can also flow into the combustion chamber of the spacecraft's engine through the feed pipe to ensure the normal operation of the spacecraft. The remaining fuel coolant flows back to the fuel tank through the second pipeline.

[0052] Each second heat exchanger is adapted to a corresponding first heat exchanger. The coolant in the second heat exchanger is used to absorb the heat generated by the airborne equipment, and the fuel coolant in the first heat exchanger is used to absorb the heat from the coolant in the second heat exchanger.

[0053] In this embodiment, each airborne device is equipped with two heat exchangers, and the positions of the heat exchangers are rationally set according to the airborne device. Each first heat exchanger and each microchannel structure is connected to a first pipeline to form a heat dissipation loop. Specifically, the inlet and outlet of each first heat exchanger, and the inlet and outlet of each microchannel structure, are connected to the first pipeline. For instance, the inlet of one first heat exchanger is connected to the outlet of a centrifugal pump via the first pipeline, and its outlet is connected to the inlet of another heat exchanger; or the outlet of one first heat exchanger is connected to the inlet of a microchannel structure. Figure 2 As shown, the inlet of the first airborne equipment corresponding to the first heat exchanger is connected to the outlet of the second centrifugal pump through the first pipeline. The outlet is connected to the inlet of the second airborne equipment corresponding to the second heat exchanger and the inlet of the microchannel structure corresponding to the leading edge of the aircraft through the first pipeline. The outlet of the second airborne equipment corresponding to the second heat exchanger is connected to the inlet of the microchannel structure corresponding to the leading edge of the wing through the first pipeline. The outlet of the microchannel structure corresponding to the leading edge of the aircraft and the outlet of the microchannel structure corresponding to the leading edge of the wing are respectively connected to the feed pipe through the first pipeline.

[0054] It should be noted that each second heat exchanger is equipped with an independent coolant, which circulates independently within the second heat exchanger. The position of the second heat exchanger is adapted to the corresponding first heat exchanger. The coolant in the second heat exchanger absorbs the heat generated by the airborne equipment, and the fuel coolant in the first heat exchanger absorbs the heat from the coolant in the second heat exchanger, thereby achieving the cooling of the airborne equipment.

[0055] In one feasible implementation, the fuel line is provided with a multi-port connector; the first line, the second line, the third line and the feed line are respectively connected to the multi-port connector.

[0056] In this embodiment, a multi-port connector is used to divert the fuel coolant in the first pipeline. A portion of the fuel coolant flows into the feed pipe through the multi-port connector and enters the combustion chamber of the aerospace vehicle's engine to ensure the normal operation of the aerospace vehicle. The remaining fuel coolant flows into the second pipeline through the multi-port connector and then flows back to the fuel tank via the second pipeline.

[0057] It should be noted that a controllable opening valve (controllable valve) can also be installed at the fuel tank inlet. This controllable opening valve can be used to regulate the flow rate of fuel returning to the fuel tank through the second pipeline. For example, when there are parts or airborne equipment that urgently need cooling, by reducing the flow rate of fuel returning to the fuel tank through the second pipeline, the rate of temperature rise of the fuel in the fuel tank can be reduced, ensuring that the temperature of the fuel in the fuel tank is lower than the temperature of the parts or airborne equipment that urgently need cooling, so as to effectively cool the parts or airborne equipment that urgently need cooling.

[0058] The thermal management method for the spacecraft includes steps S110~S130:

[0059] Step S110: Real-time acquisition of the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each microchannel structure;

[0060] Step S120: Based on the first fuel coolant temperature and the second fuel coolant temperature, determine the first target fuel coolant flow rate at the outlet of the first centrifugal pump and the second target fuel coolant flow rate at the outlet of the second centrifugal pump.

[0061] Step S130: Control the first centrifugal pump based on the first target fuel coolant flow rate and control the second centrifugal pump based on the second target fuel coolant flow rate.

[0062] In this embodiment of the application, during the flight of the spacecraft, the current first fuel coolant temperature of each heat exchanger and the current second fuel coolant temperature of each microchannel structure are acquired in real time. The first fuel coolant temperature can be the temperature at the outlet of each first heat exchanger, and the second fuel coolant temperature can be the temperature at the outlet of each microchannel structure.

[0063] In one feasible implementation, step S110 may include steps A11-A12:

[0064] Step A11: The temperature of the first fuel coolant is obtained in real time through the temperature sensors at the outlets of each first heat exchanger;

[0065] Step A12: The temperature of the second fuel coolant is obtained in real time through the temperature sensors at the outlets of each microchannel structure.

[0066] In this embodiment, temperature sensors are installed at the outlets of each first heat exchanger and at the outlets of each microchannel structure. The temperature of the first fuel coolant is obtained in real time through the temperature sensors at the outlets of each first heat exchanger, and the temperature of the second fuel coolant is obtained in real time through the temperature sensors at the outlets of each microchannel structure. That is, the first fuel coolant temperature is the temperature collected by the temperature sensor at the outlet of the first heat exchanger, and the second fuel coolant temperature is the temperature collected by the temperature sensor at the outlet of the microchannel structure.

[0067] After obtaining the first fuel coolant temperature and the second fuel coolant temperature, the first target fuel coolant flow rate at the outlet of the first centrifugal pump and the second target fuel coolant flow rate at the outlet of the second centrifugal pump are determined based on the first fuel coolant temperature and the second fuel coolant temperature. Specifically, the first target fuel coolant flow rate and the second target fuel coolant flow rate are determined according to the target temperature of each airborne device, the target temperature of each device to be cooled, the first fuel coolant temperature, and the second fuel coolant temperature. The target temperatures of the airborne devices and the devices to be cooled can be reasonably set according to the actual operating conditions.

[0068] Furthermore, in one feasible implementation, step S120 may include steps B11-B12:

[0069] Step B11: Obtain the first target temperature corresponding to each airborne device and the second target temperature corresponding to each part to be cooled.

[0070] Step B12: Based on the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature, determine the first target fuel coolant flow rate and the second target fuel coolant flow rate.

[0071] In this embodiment, after obtaining the first fuel coolant temperature and the second fuel coolant temperature, the first target temperature corresponding to each airborne device and the second target temperature corresponding to each part to be cooled are obtained. The first target fuel coolant flow rate and the second target fuel coolant flow rate are then determined based on the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature.

[0072] Furthermore, in one feasible implementation, step B12 may include step B121:

[0073] Step B121: Input the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature into the state space equation, and determine the first target fuel coolant flow rate and the second target fuel coolant flow rate through the output of the state space equation.

[0074] In this embodiment of the application, the target fuel coolant flow rate is calculated by a pre-set state-space equation, that is, the first target temperature, the second target temperature, the first fuel coolant temperature, and the second fuel coolant temperature are used as inputs to the state-space equation, and the first target fuel coolant flow rate and the second target fuel coolant flow rate are obtained by the output of the state-space equation.

[0075] The derivation of the state-space equations is as follows:

[0076] First, derive the capacitive and resistive equations.

[0077] In a fluid network, the system's actuators first change the head / flow resistance of a certain pipeline. With the node pressure remaining constant, the flow rate of the pipeline will change. The change in flow rate will then affect the pressure at the node, and the change in node pressure will then cause changes in the fluid flow rate of other pipelines. Whether it is a two-node model or a multi-node model, the working principle of the fluid dynamics layer is the same.

[0078] To qualitatively and quantitatively study fluid networks, two types of equations are needed to describe the fluid's properties: capacitive equations and resistive equations. Capacitive equations describe the relationship between pressure changes and flow rates at fluid nodes, while resistive equations describe the relationship between pressure changes and flow rates in a pipeline. These two types of equations can comprehensively describe the fluid's dynamic processes. For ease of understanding, this embodiment uses a two-node model as an example to illustrate the pressure-flow dynamic characteristics.

[0079] The derivation of the capacitive equation is as follows:

[0080] The compression equation for a compressible fluid is:

[0081] ;

[0082] in, E The bulk modulus of a compressible fluid. P The pressure of a compressible fluid. V The volume of the compressible fluid.

[0083] When there is Q=dV / dt When fuel of a certain volumetric flow rate passes through the cavity, the rate of change of cavity pressure is:

[0084] ;

[0085] in, Q This represents the volumetric flow rate of the fuel coolant.

[0086] Since flow rate is divided into input flow rate and output flow rate, and mass flow rate is usually used, the capacitive equation is:

[0087] ;

[0088] in, q in For input traffic, q out To output flow, ρ This refers to the density of the fuel coolant.

[0089] The derivation of the resistive equation is based on the fact that the pressure difference across the pipe is equal to the flow resistance lost in the pipe. Specifically, the resistive equation can be expressed as:

[0090] ;

[0091] in, P The pressure difference of the fuel coolant in the pipeline, The flow resistance equation for the pipeline is as follows: It is related to the viscosity and friction of the fluid.

[0092] Then, a thermodynamic model is constructed based on the two-node fluid network, and the state-space equations of the thermodynamic model are derived based on the derived capacitive and resistive equations.

[0093] For fluid models of two-node fluid networks, such as Figure 5 As shown, Figure 5 This is a simplified schematic diagram of a fluid model for a two-node fluid network. The model includes a fluid source (fuel tank) Res1., a centrifugal pump, a first pipe with volume V1, and a second pipe with volume V2, all connected sequentially. The first and second pipes have different volumes, primarily to describe the pressure and flow rate changes in the pipes following the centrifugal pump. Fuel flows from the fluid source (fuel tank) Res1., passes through a centrifugal pump, enters the first pipe with volume V1, and then enters the second pipe with volume V2. Therefore, the capacitive equation for the two-node fluid network is:

[0094] ;

[0095] in, The inflow volume is V The rate of change of fuel coolant pressure in pipeline 1. The inflow volume is V The rate of change of fuel coolant pressure in pipeline 2, The inflow volume is V The change in the flow rate of the fuel coolant in pipeline 1. The inflow volume is V The change in the flow rate of the fuel coolant in pipeline 2. The engine's fuel consumption is also the outflow volume of... VThe change in the flow rate of fuel coolant in pipeline 2.

[0096] The state-space equation of the capacitive equation of a two-node fluid network can be specifically expressed as:

[0097] ;

[0098] The resistive equations of a two-node fluid network can be expressed as:

[0099] ;

[0100] Linearizing the resistance equations of a two-node fluid network, the resulting linearized resistance equations can be expressed as follows:

[0101] ;

[0102] in, P 0 represents the fuel tank pressure. P 1 represents the pressure at the first pipe. P 2 represents the pressure at the second pipe. P 0 represents the pressure increment at the centrifugal pump. P 1 represents the pressure increment at the first pipe. P 2 represents the pressure increment at the second pipe. n s For the pump speed increment, n s This refers to the rotational speed of the centrifugal pump. The flow rate of fuel coolant at the outlet of the centrifugal pump. The inflow volume is V The flow rate of fuel coolant in pipeline 2.

[0103] The polynomial fitting coefficients are the first pipeline resistance function. The polynomial fitting coefficients are the coefficients for the second pipeline resistance function. b 11 , b 12 , b 13 , b 14 , b 15 , b 16 These are the polynomial fitting coefficients for the centrifugal pump. The pipeline resistance function is obtained by data fitting, and its coefficients are determined by factors such as material type and surface smoothness.

[0104] Solving the linearized resistive equation, we get:

[0105] ;

[0106] Furthermore, the state-space equation of the resistive equation of a two-node fluid network can be expressed as:

[0107] ;

[0108] ;

[0109] in, β 11 , β 12 , β 21 All parameters are custom parameters.

[0110] By combining the state-space equations of the capacitive and resistive equations of a two-node fluid network, we can obtain the state-space equation of the two-node fluid network, the specific formula of which is as follows:

[0111] ;

[0112] The state-space equation of a two-node fluid network has two state variables, indicating that the number of state variables corresponds to the number of nodes or volumetric modules in the system.

[0113] The transfer function can be solved from the state-space equations of a two-node fluid network. After extracting the state variables and control variables and performing a Laplace transform, we obtain:

[0114] ;

[0115] The transfer functions obtained are as follows:

[0116] ;

[0117] ;

[0118] in, for P The transfer function of 1 for P The transfer function of 2.

[0119] when V When 1=0, it transforms into a first-order transfer function:

[0120] ;

[0121] ;

[0122] The controller parameters are designed based on the transfer function to control the state of each actuator and achieve pressure-flow regulation.

[0123] Next, combining the state-space equations of the resistive equations of a two-node fluid network... , , Using parameters such as [parameters], the state-space equations of the thermodynamic model are derived. For thermodynamic models, such as [examples of models]... Figure 6 As shown, Figure 6 In this model, CP1 represents any one or more of the first heat exchanger and the microchannel structure; W represents the heat generated, which can be heat generated by onboard equipment or the part to be cooled; the nodes of the two-node fluid network are the connections between the first pipe, the third pipe, and the feed pipe; the first branch is the connection between the third pipe and the first pipe; the second branch is the connection between the first pipe, the second pipe, and the feed pipe; and the controllable valve is located on the second pipe. The state-space equation of this thermodynamic model can be expressed as:

[0124]

[0125] Recorded as:

[0126] ;

[0127] in, T 0 represents the temperature of the fuel in the fuel tank. T 1 represents the temperature of the fuel at the first branch. T 2 represents the temperature of the fuel before it enters the engine. T 3 represents the temperature of the fuel at the first heat exchanger or microchannel structure. T 4 represents the temperature of the fuel at the second branch. T 0 represents the increase in fuel temperature in the fuel tank. T 1 represents the increase in fuel temperature at the first branch. T 2 represents the increase in fuel temperature before it enters the engine. T 3 represents the increment of fuel temperature at the first heat exchanger or microchannel structure. T 4 represents the increase in fuel temperature at the second branch. T w The temperature increase caused by heat generated by the equipment. The rate of change of fuel temperature increment in the fuel tank. The rate of change of fuel temperature increment at the first branch. The rate of change of fuel temperature increment before entering the engine. This represents the rate of change of fuel temperature increment at the first heat exchanger or microchannel structure. The rate of change of fuel temperature increment at the second branch. The rate of change of temperature increment caused by heat generated by the equipment. The rate of change of the increment in fuel mass flow rate returning to the fuel tank. The rate of change of the fuel mass flow rate increment exiting the first centrifugal pump. The rate of change of the increase in fuel mass flow rate entering the engine for combustion from the first branch. The rate of change of the fuel mass flow rate increment exiting the second centrifugal pump, The rate of change of the fuel mass flow rate increment entering the controllable valve from the second branch. The rate of change of the increase in fuel mass flow rate entering engine combustion from the second branch. The rate of change of the increment in fuel mass flow rate during engine combustion. This refers to the initial temperature of the fuel before it enters the engine. c p For isobaric specific heat capacity, This refers to the increase in heat generation caused by changes in operating conditions of the airborne system. K The heat transfer coefficient is... A For the heat transfer area, C t This is the specific heat capacity matrix under constant pressure. This is the matrix representing the rate of change of fuel temperature increment.

[0128] in, , , , , h t , r t , T head and l t All parameters are custom parameters.

[0129] The state-space equations of the thermodynamic model can be obtained as follows: The formula is:

[0130] ;

[0131] The standard form of the state-space equations of a thermodynamic model can be expressed as:

[0132] ;

[0133] in,

[0134] ;

[0135] in, It is the flow rate of the fluid network. y A is the increment of the system output. T B T , , All parameters are custom parameters.

[0136] In thermodynamic layer calculations, only the static gain of the fluid needs to be considered. It can be represented as:

[0137] ;

[0138] Among them, K q_u This is the gain matrix of flow rate versus pressure. K q_e This is the gain matrix of flow rate on engine fuel consumption rate. M is the hydrodynamic layer characteristic parameter matrix. h Let A be the energy adjacency matrix. p Let B be the state matrix of the pressure layer. p The input correlation matrix of the pressure layer, d h E is the perturbation matrix of the thermodynamic layer. p This is the extended input matrix for the pressure layer.

[0139] The standard form of the state-space equations of the thermodynamic model and After merging, the state-space equation for the control quantity versus system temperature is obtained, which is the state-space equation required in this application. Furthermore, the first target fuel coolant flow rate and the second target fuel coolant flow rate can be calculated using this state-space equation. The standard form of the state-space equation from the thermodynamic model is... It can be seen that the control quantity u h It affects the flow distribution in the system, which in turn affects the temperature distribution.

[0140] By collecting the temperatures of the first airborne equipment, the second airborne equipment, the leading edge of the wing, and the nose (leading edge) of the aircraft, the controller determines the control parameters based on the corresponding target temperatures and the collected temperatures. The first target fuel coolant flow rate and the second target fuel coolant flow rate are obtained by integrating the control parameters through state-space equations.

[0141] Specifically, the temperature of the first fuel coolant and the temperature of the second fuel coolant are used as... T 3. Calculate the corresponding temperature of the first heat exchanger based on the first fuel coolant temperature and the first target temperature. T3. Calculate the microchannel structure corresponding to the second fuel coolant temperature and the second target temperature respectively. T 3. Put each T 3 and T 3, as the input to the state-space equation, will be obtained through the output of the state-space equation. and .

[0142] The result The state-space equation of the resistive equation for a two-node fluid network Substituting the state-space equations of the resistive equations of the two-node fluid network, the first target rotational speed of the first centrifugal pump is obtained. n s1 , set the first target speed n s1 Substituting the values ​​into the flow rate versus rotational speed function yields the first target fuel coolant flow rate corresponding to the first centrifugal pump. The formula for the flow rate versus rotational speed function is:

[0143] ;

[0144] in, c 11 , c 12 , c 13 , c 14 , c 15 , c 16 For the polynomial fitting coefficients of the centrifugal pump, The first target fuel coolant flow rate is the flow rate of the fuel coolant at the outlet of the first centrifugal pump, specifically determined by a relational function. The first target fuel coolant flow rate was obtained by performing polynomial fitting. .

[0145] The result The state-space equation of the resistive equation for a two-node fluid network Substituting the state-space equations of the resistive equations of the two-node fluid network, the second target rotational speed of the second centrifugal pump is obtained. n s2 The second target rotation speed n s2 Substituting the values ​​into the flow rate versus rotational speed function yields the second target fuel coolant flow rate corresponding to the second centrifugal pump. The formula for the flow rate versus rotational speed function is:

[0146] ;

[0147] in, The second target fuel coolant flow rate is the flow rate of the fuel coolant at the outlet of the second centrifugal pump, specifically determined by a relational function. The second target fuel coolant flow rate was obtained by performing polynomial fitting. .

[0148] Subsequently, the rotational speed of the first centrifugal pump is controlled based on the first target fuel coolant flow rate to adjust the fuel coolant flow rate at its outlet to achieve the first target fuel coolant flow rate, and the rotational speed of the second centrifugal pump is controlled based on the second target fuel coolant flow rate to adjust the fuel coolant flow rate at its outlet to achieve the second target fuel coolant flow rate.

[0149] The thermal management method for aerospace vehicles provided in this application forms a heat dissipation circuit using fuel as a coolant through a heat exchanger, microchannel structure, and fuel pipeline. The fuel coolant in the heat dissipation circuit dissipates heat from the airborne equipment and parts requiring cooling, achieving active thermal protection for the aerospace vehicle. Simultaneously, the centrifugal pump is controlled by the temperatures of the first and second fuel coolants to regulate the flow rate of the fuel coolant in the fuel pipeline, actively cooling the high-temperature equipment in the airborne equipment and parts requiring cooling. That is, the highest temperature equipment is cooled first by the flow rate of the fuel coolant, achieving controllable temperature and temperature gradient arrangement of the entire aerospace vehicle, thereby actively reducing the temperature of the airborne equipment and / or parts requiring cooling, effectively reducing the burden of passive thermal protection, and improving the thermal protection capability and reusability of the aerospace vehicle.

[0150] It should be noted that the above examples are only for understanding this application and do not constitute a limitation on the thermal management method of the aerospace vehicle of this application. Any simple modifications based on this technical concept are within the protection scope of this application.

[0151] This application provides a thermal management device for a spacecraft, comprising: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores instructions executable by the at least one processor, the instructions being executed by the at least one processor to enable the at least one processor to perform the thermal management method for the spacecraft described in Embodiment 1 above.

[0152] The following is for reference. Figure 4This document illustrates a structural schematic diagram of a thermal management device suitable for implementing the embodiments of this application for a spacecraft. The thermal management device for the spacecraft in the embodiments of this application may include, but is not limited to, mobile terminals such as mobile phones, laptops, digital broadcast receivers, PDAs (Personal Digital Assistants), PADs (Portable Application Description), PMPs (Portable Media Players), in-vehicle terminals (e.g., in-vehicle navigation terminals), and fixed terminals such as digital TVs and desktop computers. Figure 4 The thermal management device for the aerospace vehicle shown is merely an example and should not impose any limitations on the functionality and scope of use of the embodiments of this application.

[0153] like Figure 4 As shown, the thermal management device of the spacecraft may include a processing unit 1001 (e.g., a central processing unit, a graphics processing unit, etc.), which can perform various appropriate actions and processes according to a program stored in a read-only memory (ROM) 1002 or a program loaded from a storage device 1003 into a random access memory (RAM) 1004. The RAM 1004 also stores various programs and data required for the operation of the spacecraft's thermal management device. The processing unit 1001, ROM 1002, and RAM 1004 are interconnected via a bus 1005. An input / output (I / O) interface 1006 is also connected to the bus. Typically, the following systems can be connected to I / O interface 1006: input devices 1007 including, for example, touchscreens, touchpads, keyboards, mice, image sensors, microphones, accelerometers, gyroscopes, etc.; output devices 1008 including, for example, liquid crystal displays (LCDs), speakers, vibrators, etc.; storage devices 1003 including, for example, magnetic tapes, hard disks, etc.; and communication devices 1009. Communication device 1009 allows the thermal management device of the spacecraft to communicate wirelessly or wiredly with other devices to exchange data. Although the figure shows a thermal management device for a spacecraft with various systems, it should be understood that it is not required to implement or possess all the systems shown. More or fewer systems may be implemented alternatively.

[0154] Specifically, according to the embodiments disclosed in this application, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments disclosed in this application include a computer program product comprising a computer program carried on a computer-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via a communication device, or installed from storage device 1003, or installed from ROM 1002. When the computer program is executed by processing device 1001, it performs the functions defined in the methods of the embodiments disclosed in this application.

[0155] The thermal management device for aerospace vehicles provided in this application, employing the thermal management method for aerospace vehicles described in the above embodiments, can solve the technical problems of how to reduce the passive thermal protection burden, improve the thermal protection capability of aerospace vehicles, and increase their reusability. Compared with the prior art, the beneficial effects of the thermal management device for aerospace vehicles provided in this application are the same as those of the thermal management method for aerospace vehicles provided in the above embodiments, and other technical features in this thermal management device for aerospace vehicles are the same as those disclosed in the method of the previous embodiment, and will not be repeated here.

[0156] It should be understood that the various parts disclosed in this application can be implemented using hardware, software, firmware, or a combination thereof. In the description of the above embodiments, specific features, structures, materials, or characteristics can be combined in any suitable manner in one or more embodiments or examples.

[0157] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

[0158] This application provides a computer-readable storage medium having computer-readable program instructions (i.e., a computer program) stored thereon, the computer-readable program instructions being used to execute the thermal management method for a spacecraft in the above embodiments.

[0159] The computer-readable storage medium provided in this application may be, for example, a USB flash drive, but is not limited to, electrical, magnetic, optical, electromagnetic, infrared, or semiconductor systems, devices, or any combination thereof. More specific examples of computer-readable storage media may include, but are not limited to: electrical connections having one or more wires, portable computer disks, hard disks, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage devices, magnetic storage devices, or any suitable combination thereof. In this embodiment, the computer-readable storage medium may be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system, system, or device. The program code contained on the computer-readable storage medium may be transmitted using any suitable medium, including but not limited to: wires, optical cables, RF (Radio Frequency), etc., or any suitable combination thereof.

[0160] The aforementioned computer-readable storage medium may be included in the thermal management device of the spacecraft; or it may exist independently and not be installed in the thermal management device of the spacecraft.

[0161] The aforementioned computer-readable storage medium carries one or more programs that, when executed by the thermal management device of the spacecraft, cause the thermal management device to: acquire in real time the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each microchannel structure; determine, based on the first fuel coolant temperature and the second fuel coolant temperature, a first target fuel coolant flow rate at the outlet of the first centrifugal pump and a second target fuel coolant flow rate at the outlet of the second centrifugal pump; control the first centrifugal pump based on the first target fuel coolant flow rate and control the second centrifugal pump based on the second target fuel coolant flow rate.

[0162] Computer program code for performing the operations of this application can be written in one or more programming languages ​​or a combination thereof, including object-oriented programming languages ​​such as Java, Smalltalk, and C++, and conventional procedural programming languages ​​such as the "C" language or similar programming languages. The program code can be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving remote computers, the remote computer can be connected to the user's computer via any type of network—including a Local Area Network (LAN) or a Wide Area Network (WAN)—or can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0163] The flowcharts and block diagrams in the accompanying drawings illustrate the architecture, functionality, and operation of possible implementations of systems, methods, and computer program products according to various embodiments of this application. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing a specified logical function. It should also be noted that in some alternative implementations, the functions indicated in the blocks may occur in a different order than those indicated in the drawings. For example, two consecutively indicated blocks may actually be executed substantially in parallel, and they may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, can be implemented using a dedicated hardware-based system that performs the specified function or operation, or using a combination of dedicated hardware and computer instructions.

[0164] The modules described in the embodiments of this application can be implemented in software or hardware. The names of the modules do not necessarily limit the functionality of the unit itself.

[0165] The readable storage medium provided in this application is a computer-readable storage medium that stores computer-readable program instructions (i.e., a computer program) for executing the thermal management method for the aforementioned spacecraft. This solves the technical problem of how to reduce the passive thermal protection burden, improve the thermal protection capability of spacecraft, and increase its reusability. Compared with the prior art, the beneficial effects of the computer-readable storage medium provided in this application are the same as those of the thermal management method for spacecraft provided in the above embodiments, and will not be repeated here.

[0166] This application also provides a computer program product, including a computer program that, when executed by a processor, implements the steps of the thermal management method for a spacecraft as described above.

[0167] The computer program product provided in this application can solve the technical problem of how to reduce the passive thermal protection burden, improve the thermal protection capability of aerospace vehicles, and increase their reusability. Compared with the prior art, the beneficial effects of the computer program product provided in this application are the same as those of the thermal management method for aerospace vehicles provided in the above embodiments, and will not be repeated here.

[0168] The above description is only a part of the embodiments of this application and does not limit the patent scope of this application. All equivalent structural transformations made under the technical concept of this application and using the contents of the specification and drawings of this application, or direct / indirect applications in other related technical fields, are included in the patent protection scope of this application.

Claims

1. A method of thermal management of an aerospace vehicle, characterized by, The space flight vehicle comprises airborne equipment, a part to be cooled, a fuel pipeline, a fuel tank, a first centrifugal pump, a second centrifugal pump and an engine combustion chamber; each airborne equipment is respectively provided with a first heat exchanger and a second heat exchanger; the part to be cooled is provided with a micro-channel structure; the fuel pipeline comprises a first pipeline, a second pipeline and a third pipeline; The second centrifugal pump is arranged on the first pipeline, and the first heat exchanger and the micro-channel structure are respectively communicated with the first pipeline; One end of the second pipeline is communicated with a feeding port of the fuel tank, and the other end is communicated with a feeding pipe of the engine combustion chamber; the first centrifugal pump is arranged on the third pipeline, one end of the third pipeline is communicated with a discharging port of the fuel tank, and the other end is communicated with the feeding pipe of the engine combustion chamber; the inlet of the second centrifugal pump is communicated with the outlet of the first centrifugal pump through the first pipeline and the third pipeline; Each second heat exchanger is respectively arranged in correspondence with the corresponding first heat exchanger, the coolant in the second heat exchanger is used for absorbing the heat generated by the airborne equipment, and the fuel coolant in the first heat exchanger is used for absorbing the heat of the coolant in the second heat exchanger; The heat management method of the space flight vehicle comprises: Real-time acquisition of the first fuel coolant temperature corresponding to each first heat exchanger and the second fuel coolant temperature corresponding to each micro-channel structure; Determination of the first target fuel coolant flow rate of the outlet of the first centrifugal pump and the second target fuel coolant flow rate of the outlet of the second centrifugal pump based on the first fuel coolant temperature and the second fuel coolant temperature; Control of the first centrifugal pump based on the first target fuel coolant flow rate and control of the second centrifugal pump based on the second target fuel coolant flow rate; The step of determining the first target fuel coolant flow rate of the outlet of the first centrifugal pump and the second target fuel coolant flow rate of the outlet of the second centrifugal pump based on the first fuel coolant temperature and the second fuel coolant temperature comprises: Acquisition of the first target temperature corresponding to each airborne equipment and the second target temperature corresponding to each part to be cooled; Determination of the first target fuel coolant flow rate and the second target fuel coolant flow rate based on the first target temperature, the second target temperature, the first fuel coolant temperature and the second fuel coolant temperature; The step of determining the first target fuel coolant flow rate and the second target fuel coolant flow rate based on the first target temperature, the second target temperature, the first fuel coolant temperature and the second fuel coolant temperature comprises: Inputting the first target temperature, the second target temperature, the first fuel coolant temperature and the second fuel coolant temperature into a state space equation, and determining the first target fuel coolant flow rate and the second target fuel coolant flow rate through the output of the state space equation, The formula of the state space equation is: ； ； Wherein, ； ； wherein, is the rate of change of the fuel temperature increment, T [ T 0, T 1, T 2, T 3, T 4, T w ], T 0 is the increment of the fuel temperature in the tank, T 1 is the increment of the fuel temperature at the first branch, T 2 is the increment of the fuel temperature before entering the engine, T 3 is the increment of the fuel temperature at the first heat exchanger or microchannel structure, T 4 is the increment of the fuel temperature at the second branch, T w is the temperature increment caused by the heat generated by the equipment, is the increment of the heat generated by the on-board systems caused by the change in operating conditions, is the flow of the fluid network, C t is the specific heat capacity at constant pressure matrix, y is the increment of the system output, K is the heat transfer coefficient, A is the heat transfer area, K q_u is the gain matrix of the flow with respect to the pressure, K q_e is the gain matrix of the flow with respect to the specific fuel consumption of the engine, is the matrix of the characteristic parameters of the fluid mechanics layer, M h is the energy adjacency matrix, A p is the state matrix of the pressure layer, B p is the input incidence matrix of the pressure layer, is the disturbance matrix of the thermodynamic layer, E p is the extended input matrix of the pressure layer, is the rate of change of the increment of the fuel mass flow of the engine combustion, is the specific heat capacity at constant pressure, is the control variable, , , 、 、 、 、 T head 、 、A T 、B T 、 、 are all custom parameters.

2. The method of thermal management of a spaceplane as recited in claim 1, wherein, The step of real-time acquisition of the first fuel coolant temperature corresponding to each first heat exchanger and the second fuel coolant temperature corresponding to each micro-channel structure comprises: The first fuel coolant temperature is acquired in real time through a temperature sensor at the outlet of each first heat exchanger; The second fuel coolant temperature is acquired in real time through a temperature sensor at the outlet of each micro-channel structure.

3. The method of thermal management of a spaceplane as recited in claim 2, wherein, The fuel pipeline is provided with a multi-way joint; the first pipeline, the second pipeline, the third pipeline and the feeding pipeline are in communication with the multi-way joint.

4. The method of thermal management of a spaceplane of any one of claims 1 to 3, wherein, The parts to be cooled down include at least the leading edge of the aircraft and the leading edge of the wing.

5. A thermal management device for an aerospace vehicle, comprising: The device comprises a memory, a processor and a computer program stored on the memory and executable on the processor, and the computer program is configured to implement the steps of the method for thermal management of the aerospace vehicle according to any one of claims 1 to 4.

6. A storage medium, characterized by The storage medium is a computer readable storage medium, and the storage medium stores a computer program, and the computer program is executed by the processor to implement the steps of the method for thermal management of the aerospace vehicle according to any one of claims 1 to 4.

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