Thermal management method and device of aerospace vehicle and storage medium

Abstract

The invention discloses a thermal management method and device for an aerospace vehicle and a storage medium, relates to the technical field of aircrafts, and discloses the thermal management method for the aerospace vehicle, and the method comprises the steps: obtaining a first fuel coolant temperature and a second fuel coolant temperature in real time; determining a first target fuel coolant flow rate and a second target fuel coolant flow rate based on the first fuel coolant temperature and the second fuel coolant temperature; the first centrifugal pump is controlled based on the first target fuel coolant flow and the second centrifugal pump is controlled based on the second target fuel coolant flow. The first centrifugal pump and the second centrifugal pump are controlled through the first fuel coolant temperature and the second fuel coolant temperature so as to adjust the flow of the fuel coolant in the fuel pipeline, temperature controllability and temperature gradient arrangement of the whole aerospace vehicle are achieved, and then the temperature of airborne equipment and / or a part to be cooled is actively reduced. The passive thermal protection burden is effectively reduced, and the repeated utilization rate of the aerospace vehicle is improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aircraft, in particular to a space-aircraft thermal management method, device and storage medium. BACKGROUND

[0002] With the continuous development of aerospace science and technology, the flight environment of the new generation of aircraft has changed significantly, and is no longer limited to a single aviation space or space space, showing the characteristics of cross-domain flight. Compared with traditional aircraft or spacecraft, the flight environment of space-aircraft is complex and changeable, its flight envelope covers a large airspace of 0~100km and a wide speed domain of 0~30Ma, and the maximum heat flux density can reach MW / m 2 level, and the aerodynamic heating problem is obvious.

[0003] At present, passive thermal protection technologies such as ablative thermal protection and passive radiation thermal protection are mainly used to solve the problem of aerodynamic heating. Ablative thermal protection absorbs heat by ablating the surface of the material to form a thermal insulation layer, which plays a role in reducing temperature and insulation, but the aerodynamic shape of the aircraft changes after ablation, which is not conducive to the reuse of space-aircraft. Passive radiation thermal protection uses advanced heat-resistant materials such as carbon-carbon composites and ceramics to dissipate heat through thermal radiation. Heat-resistant materials also have a clear upper temperature limit and cannot cope with extreme heat flow environments alone. Carbon-carbon composites will oxidize and fail in high-temperature oxidation environments, and high-temperature resistant materials such as ceramics have high density and high brittleness, which reduces the effective payload capacity of the aircraft and is easily damaged by mechanical impact.

[0004] Therefore, how to reduce the passive thermal protection burden and improve the thermal protection capability and reuse rate of space-aircraft is a problem that needs to be solved at present. SUMMARY

[0005] The main purpose of the present application is to provide a space-aircraft thermal management method, device and storage medium, which aims to solve the technical problem of how to reduce the passive thermal protection burden and improve the thermal protection capability and reuse rate of space-aircraft.

[0006] To achieve the above purpose, the present application provides a space-aircraft thermal management method, which comprises an airborne device, a heat dissipation site, a fuel pipeline, a fuel tank, a first centrifugal pump, a second centrifugal pump and an engine combustion chamber; each airborne device is respectively provided with a first heat exchanger and a second heat exchanger; the heat dissipation site is provided with a microchannel 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, each first heat exchanger and each micro-channel structure are in communication with the first pipeline; one end of the second pipeline is in communication with the inlet of the fuel tank, and the other end is in communication with the inlet pipe of the engine combustion chamber; the first centrifugal pump is arranged on the third pipeline, one end of the third pipeline is in communication with the outlet of the fuel tank, and the other end is in communication with the inlet pipe of the engine combustion chamber; the inlet of the second centrifugal pump is in communication with the outlet of the first centrifugal pump through the first pipeline and the third pipeline; Each second heat exchanger is 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 thermal management method of the aerospace vehicle comprises: Real-time acquisition of the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each micro-channel structure; Based on the first fuel coolant temperature and the second fuel coolant temperature, 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 are determined; Based on the first target fuel coolant flow rate, the first centrifugal pump is controlled, and based on the second target fuel coolant flow rate, the second centrifugal pump is controlled.

[0007] In an embodiment, 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 heat dissipation site; 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.

[0008] In an 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 comprises: The first target temperature, the second target temperature, the first fuel coolant temperature and the second fuel coolant temperature are input into a state space equation, and the first target fuel coolant flow rate and the second target fuel coolant flow rate are determined through the output of the state space equation.

[0009] In an embodiment, the step of acquiring the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each micro-channel structure in real time comprises: acquiring the first fuel coolant temperature in real time through a temperature sensor at the outlet of each first heat exchanger; acquiring the second fuel coolant temperature in real time through a temperature sensor at the outlet of each micro-channel structure.

[0010] In an embodiment, the fuel pipeline is provided with a multi-way joint; the first pipeline, the second pipeline, the third pipeline and the inlet pipeline are in communication with the multi-way joint.

[0011] In an embodiment, the parts to be cooled down include at least the leading edge of the aircraft and the leading edge of the wing.

[0012] In addition, to achieve the above-mentioned purpose, the present application also provides a space-aircraft thermal management device, which comprises a memory, a processor and a computer program stored in the memory and executable on the processor, and the computer program is configured to implement the steps of the aforementioned space-aircraft thermal management method.

[0013] In addition, to achieve the above-mentioned purpose, the present application also provides a storage medium, which is a computer-readable storage medium, and the storage medium stores a computer program, and the computer program is executed by a processor to implement the steps of the aforementioned space-aircraft thermal management method.

[0014] The one or more technical solutions provided by the present application have at least the following technical effects: The space-aircraft thermal management method provided by the present application forms a fuel-cooled heat dissipation circuit through the heat exchanger, the micro-channel structure and the fuel pipeline, cools the airborne equipment and the parts to be cooled down by the fuel coolant in the heat dissipation circuit, realizes active thermal protection of the space-aircraft, controls the first centrifugal pump and the second centrifugal pump through the first fuel coolant temperature and the second fuel coolant temperature to adjust the fuel coolant flow in the fuel pipeline, actively cools the high-temperature equipment in urgent need of cooling among the airborne equipment and the parts to be cooled down, i.e. cools the equipment with the highest temperature first through the fuel coolant flow, realizes temperature-controllable and temperature-gradient arrangement of the space-aircraft, and further actively reduces the temperature of the airborne equipment and / or the parts to be cooled down, effectively reduces the passive thermal protection burden, and improves the thermal protection capability and reusability of the space-aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0015] The accompanying drawings, which are incorporated in and constitute a part of the specification, illustrate embodiments consistent with the present application and, together with the description, serve to explain the principles of the application.

[0016] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the accompanying drawings needed to be used in the embodiments or prior art description will be briefly introduced as follows. Obviously, the drawings can also provide further understanding of the prior art for those of ordinary skill in the art without any creative effort.

[0017] Figure 1 The flowchart provided for an embodiment of the thermal management method of the aerospace vehicle of the present application; Figure 2 The system architecture diagram provided for an embodiment of the thermal management system of the aerospace vehicle of the present application; Figure 3 The structure diagram of the microchannel structure provided for an embodiment of the thermal management system of the aerospace vehicle of the present application; Figure 4 The device structure diagram of the hardware running environment involved in the thermal management method of the aerospace vehicle in the embodiments of the present application; Figure 5 The simplified schematic diagram of the fluid model of the two-node fluid network; Figure 6 The thermodynamic model node diagram in the related art.

[0018] The purpose implementation, functional features and advantages of the present application will be further described with reference to the embodiments and the accompanying drawings. DETAILED DESCRIPTION

[0019] It should be understood that the specific embodiments described herein are only used to explain the technical solutions of the present application, and are not used to limit the present application.

[0020] In order to better understand the technical solutions of the present application, the following will be described in detail with reference to the drawings and specific embodiments in the specification.

[0021] The main solution of the embodiment of the application is that the aerospace 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, each first heat exchanger and each 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; an inlet of the second centrifugal pump is communicated with an 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, a coolant in the second heat exchanger is used to absorb heat generated by the airborne equipment, and a fuel coolant in the first heat exchanger is used to absorb heat of the coolant in the second heat exchanger; a first fuel coolant temperature corresponding to each heat exchanger and a second fuel coolant temperature corresponding to each micro-channel structure are acquired in real time; based on the first fuel coolant temperature and the second fuel coolant temperature, a first target fuel coolant flow of the outlet of the first centrifugal pump and a second target fuel coolant flow of the outlet of the second centrifugal pump are determined; the first centrifugal pump is controlled based on the first target fuel coolant flow, and the second centrifugal pump is controlled based on the second target fuel coolant flow.

[0022] In the embodiment, for convenience of description, the following describes an execution subject of identifying a thermal management device of an aerospace vehicle.

[0023] With the continuous development of aerospace science and technology, the flight environment of a new generation of aircraft has changed significantly, and is no longer limited to a single aviation space or space, and presents the characteristics of cross-domain flight. Compared with traditional aircraft or spacecraft, the flight environment of an aerospace vehicle is complex and changeable, the flight envelope covers a large airspace of 0-100 km and a wide speed domain of 0-30 Ma, and the maximum heat flux density can reach MW / m 2 The problem of aerodynamic heating is obvious.

[0024] Currently, the problem of aerodynamic heating is solved by using passive thermal protection technologies such as ablation thermal protection and passive radiation thermal protection. Ablation thermal protection absorbs heat by surface ablation of the material to form a thermal insulation layer, thereby playing a role in temperature reduction and thermal insulation. However, the aerodynamic shape of the aircraft changes after ablation, which is not conducive to the reuse 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. Heat-resistant materials also have a clear upper temperature limit and cannot cope with extreme heat flow environments alone. Carbon-carbon composites will oxidize and fail in high-temperature oxidation environments, and high-temperature-resistant materials such as ceramics have high density and high brittleness, which reduces the effective payload capacity of the aircraft and makes it vulnerable to mechanical impact damage.

[0025] Therefore, how to reduce the burden of passive thermal protection and improve the thermal protection capability and reuse rate of aerospace vehicles is a problem that needs to be solved at present.

[0026] The present application provides a solution to form a heat dissipation loop with fuel as the coolant through the heat exchanger, micro-channel structure and fuel pipeline, dissipate heat for the airborne equipment and the heat dissipation site of the aerospace vehicle through the fuel coolant in the heat dissipation loop, control the first centrifugal pump and the second centrifugal pump through the first fuel coolant temperature and the second fuel coolant temperature to adjust the fuel coolant flow in the fuel pipeline, and actively cool the high-temperature equipment in urgent need of cooling in the airborne equipment and the heat dissipation site, that is, the fuel coolant flow is used to cool the equipment with the highest temperature first, realize the temperature controllability and temperature gradient arrangement of the aerospace vehicle, and then actively reduce the temperature of the airborne equipment and / or the heat dissipation site, effectively reduce the burden of passive thermal protection, and improve the thermal protection capability and reuse rate of the aerospace vehicle.

[0027] It should be noted that the execution subject of the present embodiment can be a computing service device with data processing, network communication and program running functions, such as a tablet computer, a personal computer, a mobile phone, etc., or an electronic device capable of realizing the above functions, a thermal management device of an aerospace vehicle, etc. The thermal management device of the aerospace vehicle is taken as an example to illustrate the present embodiment and the following embodiments.

[0028] Based on this, the present application provides a thermal management method of an aerospace vehicle, which is described with reference to Figure 1 , Figure 1 is a flowchart of an embodiment of the thermal management method of the aerospace vehicle of the present application.

[0029] In the present embodiment, the aerospace vehicle includes airborne equipment, a heat dissipation site, a fuel pipeline, a fuel tank, a first centrifugal pump, a second centrifugal pump and an engine combustion chamber. As shown in Figure 2As shown, the airborne equipment includes a first airborne equipment and a second airborne equipment, and the parts to be cooled include a wing leading edge and an aircraft leading edge. Each airborne equipment is respectively provided with a first heat exchanger and a second heat exchanger.

[0030] The airborne equipment is a device in the aerospace vehicle that needs to be cooled and protected, and the airborne equipment is provided with a heat exchanger. The part to be cooled is a part in the aerospace vehicle that needs to be cooled and protected, and the part to be cooled is provided with a micro-channel structure. Specifically, the part to be cooled includes but is not limited to the aircraft leading edge (head) and the wing leading edge, and multiple parts to be cooled can be respectively arranged at positions where aerodynamic heating is obvious in the aerospace vehicle. The micro-channel structure includes multiple fuel coolant circuits, and the shape of the micro-channel structure is adaptively set according to the shape of the part to be cooled. Figure 3 As shown, Figure 3 The micro-channel structure in the above is annular, and is provided with multiple semicircular fuel coolant circuits, one end of which is an inlet of the fuel coolant, and the other end is an outlet of the fuel coolant.

[0031] The fuel pipeline is provided with a centrifugal pump, which includes a first centrifugal pump and a second centrifugal pump. The first centrifugal pump is used to extract the fuel coolant in the fuel tank to the fuel pipeline, and the flow of the fuel coolant extracted to the fuel pipeline can also be adjusted according to the temperature of the airborne equipment and the part to be cooled.

[0032] The fuel pipeline includes a first pipeline, a second pipeline and a third pipeline. The second centrifugal pump is arranged on the first pipeline, and each first heat exchanger and each micro-channel structure is in communication with the first pipeline. One end of the second pipeline is in communication with the inlet of the fuel tank, and the other end is in communication with the inlet pipe of the engine combustion chamber. The first centrifugal pump is arranged on the third pipeline, one end of the third pipeline is in communication with the outlet of the fuel tank, and the other end is in communication with the inlet pipe of the engine combustion chamber. The inlet of the second centrifugal pump is in communication with the outlet of the first centrifugal pump through the first pipeline and the third pipeline.

[0033] Specifically, as shown, Figure 2 The inlet of the second centrifugal pump is in communication with the outlet of the first centrifugal pump through the first pipeline and the third pipeline, the outlet of the second centrifugal pump is in communication with 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 in communication with 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 in communication with the inlet of the micro-channel structure corresponding to the wing leading edge, the inlet of the micro-channel structure corresponding to the aircraft leading edge is in communication with the outlet of the first heat exchanger corresponding to the first airborne equipment, and the outlet of the micro-channel structure corresponding to the aircraft leading edge and the outlet of the micro-channel structure corresponding to the wing leading edge are respectively in communication with the inlet pipe of the engine combustion chamber. The pipeline connecting the second centrifugal pump, the first heat exchanger and the micro-channel structure is the second pipeline.

[0034] 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.

[0035] 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.

[0036] 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.

[0037] 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.

[0038] 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.

[0039] 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.

[0040] It should be noted that a controllable opening valve (controllable valve) can also be arranged at the fuel tank inlet to adjust the fuel flow rate flowing back to the fuel tank through the second pipeline, for example, when there is a part or airborne equipment in urgent need of cooling, by reducing the fuel flow rate flowing back to the fuel tank through the second pipeline, the temperature rise rate of the fuel in the fuel tank can be reduced to ensure that the temperature of the fuel in the fuel tank is lower than that of the part or airborne equipment in urgent need of cooling, so as to effectively cool the part or airborne equipment in urgent need of cooling.

[0041] The thermal management method of the aerospace vehicle includes steps S110-S130: 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; Step S120, determination of 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; Step S130, 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.

[0042] In the embodiments of the present application, during the flight of the aerospace vehicle, 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, wherein 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.

[0043] In a feasible implementation, step S110 can include steps A11-A12: Step A11, real-time acquisition of the first fuel coolant temperature by the temperature sensor at the outlet of each first heat exchanger; Step A12, real-time acquisition of the second fuel coolant temperature by the temperature sensor at the outlet of each microchannel structure.

[0044] In the embodiments, temperature sensors are arranged at the outlets of each first heat exchanger and at the outlets of each microchannel structure, the first fuel coolant temperature is acquired in real time by the temperature sensor at the outlet of each first heat exchanger, and the second fuel coolant temperature is acquired in real time by the temperature sensor at the outlet 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.

[0045] After the first fuel coolant temperature and the second fuel coolant temperature are obtained, 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 heat dissipation device, the first fuel coolant temperature and the second fuel coolant temperature, wherein the target temperatures of the airborne devices and the heat dissipation devices can be reasonably set according to actual working conditions.

[0046] Further, in a feasible implementation, the step S120 can include steps B11-B12: Step B11, obtaining the first target temperature corresponding to each airborne device and the second target temperature corresponding to each heat dissipation site; Step B12, 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.

[0047] In the embodiment, after the first fuel coolant temperature and the second fuel coolant temperature are obtained, the first target temperature corresponding to each airborne device and the second target temperature corresponding to each heat dissipation site are obtained. And the first target fuel coolant flow rate and the second target fuel coolant flow rate are determined based on the first target temperature, the second target temperature, the first fuel coolant temperature and the second fuel coolant temperature.

[0048] Further, in a feasible implementation, the step B12 can include step B121: Step B121, 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.

[0049] In the embodiment of the application, the target fuel coolant flow rate is calculated through 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 input as the input of the state space equation, and the first target fuel coolant flow rate and the second target fuel coolant flow rate are obtained through the output of the state space equation.

[0050] Wherein, the derivation process of the state space equation is as follows: First, the capacitive equation and the resistive equation are derived.

[0051] In the fluid network, the actuator of the system will first change the lift / flow resistance of a pipeline, and the flow of the pipeline will change under the condition that the node pressure is constant. The flow causes changes, which in turn affect the pressure of the node. The change in node pressure causes changes in the flow of other pipelines. Whether it is a two-node model or a multi-node model, the working principle of the fluid mechanics layer is consistent.

[0052] In order to qualitatively and quantitatively study the fluid network, two types of equations need to be established to describe the properties of the fluid, namely the capacitance equation and the resistance equation of the fluid. The capacitance equation is used to describe the relationship between the pressure change of the fluid node and the flow change. The resistance equation is used to describe the relationship between the pressure change and the flow in the pipeline. Through the two types of equations, the change process of the fluid can be completely described. For the sake of understanding, this embodiment takes a two-node model as an example to illustrate the pressure-flow dynamic characteristics.

[0053] Among them, the capacitance equation is derived as follows: The compression equation of the compressible fluid is: ; Among them, E is the bulk modulus of the compressible fluid, P is the pressure of the compressible fluid, V is the volume of the compressible fluid.

[0054] When the fuel with a volume flow rate of Q = dV / dt passes through the cavity, the rate of change of the cavity pressure is: ; Among them, Q is the volume flow rate of the fuel coolant.

[0055] Since the flow is divided into input flow and output flow, and mass flow is usually used, the capacitance equation is: ; Among them, q in is the input flow, q out is the output flow, p is the density of the fuel coolant.

[0056] For the derivation of the resistance equation, the pressure difference at both ends of the resistance equation pipeline is equal to the flow resistance loss in the pipeline. The resistance equation can be specifically expressed as: ; Among them, P is the pressure difference of the fuel coolant in the pipeline, is the flow resistance equation of the pipeline, It is related to the viscosity and friction of the fluid.

[0057] 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.

[0058] 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: ; 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... V The change in the flow rate of fuel coolant in pipeline 2.

[0059] The state-space equation of the capacitive equation of a two-node fluid network can be specifically expressed as: ; The resistive equations of a two-node fluid network can be expressed as: ; Linearizing the resistance equations of a two-node fluid network, the resulting linearized resistance equations can be expressed as follows: ; 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 is the pressure increment at the first pipe, P 2 is the pressure increment at the second pipe, n s is the increment of the rotational speed of the pump, n s is the rotational speed of the centrifugal pump. is the flow rate of the fuel coolant at the outlet of the centrifugal pump, is the flow rate of the fuel coolant flowing into the pipe with a volume of V 2.

[0060] is the polynomial fitting coefficient of the first pipe resistance function, is the polynomial fitting coefficient of the second pipe resistance function, b 11 , b 12 , b 13 , b 14 , b 15 , b 16 is the polynomial fitting coefficient of the centrifugal pump. The pipe resistance function is fitted by data, and the coefficient is determined by factors such as the type of material and the degree of surface smoothness.

[0061] Solving the linearized resistance equation, we get: ; Further, the state space equation of the resistance equation of the two-node fluid network can be expressed as: ; ; wherein, β 11 , β 12 , β 21 are self-defined parameters.

[0062] Combining the state space equation of the capacitance equation of the two-node fluid network with the state space equation of the resistance equation, the state space equation of the two-node fluid network is obtained, and the specific formula is: ; In the state space equation of the two-node fluid network, there are two state variables, which indicates that there are several nodes or several volume modules in the system, and there are several state variables.

[0063] The transfer function can be solved by the state space equation of the two-node fluid network, and the part of the state variable and the control variable is taken out, and the Laplace transform is carried out to obtain: ; The transfer function is solved as follows: ; ; Among them, is the transfer function of 1, P is the transfer function of 2. P When 1=0, it becomes a first-order transfer function:

[0064] V ; ; ; Based on the transfer function, the controller parameters are designed to control the state of each actuator and realize the regulation of pressure-flow.

[0065] Then, combined with the parameters such as , , in the state space equation of the resistive equation of the two-node fluid network, the state space equation of the thermodynamic model is derived. For the thermodynamic model, as shown in Figure 6 , Figure 6 CP1 in is the first heat exchanger and any one or more of the microchannel structure, W is the heat generated by the airborne equipment or the part to be cooled, the node of the two-node fluid network is the connection between the first pipeline, the third pipeline and the inlet pipe, the first branch is the connection between the third pipeline and the first pipeline, the second branch is the connection between the first pipeline, the second pipeline and the inlet pipe, and the controllable valve is arranged on the second pipeline. The state space equation of the thermodynamic model can be expressed as:

[0066] Denoted as: ; Among them, T 0 is the temperature of the fuel in the fuel tank, T 1 is the temperature of the fuel at the first branch, T 2 is the temperature of the fuel before entering the engine, T 3 is the temperature of the fuel at the first heat exchanger or the microchannel structure, T 4 is the temperature of the fuel at the second branch, T 0 is the increment of the temperature of the fuel in the fuel tank, ​T 1 is the increment of fuel temperature at the first branch, T 2 is the increment of fuel temperature before entering the engine, T 3 is the increment of fuel temperature at the first heat exchanger or microchannel structure, T 4 is the increment of fuel temperature at the second branch, T w is the increment of temperature due to heat generated by the equipment, is the rate of change of the increment of fuel temperature in the tank, is the rate of change of the increment of fuel temperature at the first branch, is the rate of change of the increment of fuel temperature before entering the engine, is the rate of change of the increment of fuel temperature at the first heat exchanger or microchannel structure, is the rate of change of the increment of fuel temperature at the second branch, is the rate of change of the increment of temperature due to heat generated by the equipment, is the rate of change of the increment of fuel mass flow rate flowing back to the tank, is the rate of change of the increment of fuel mass flow rate flowing out of the first centrifugal pump, is the rate of change of the increment of fuel mass flow rate entering the engine for combustion at the first branch, is the rate of change of the increment of fuel mass flow rate flowing out of the second centrifugal pump, is the rate of change of the increment of fuel mass flow rate entering the controllable valve at the second branch, is the rate of change of the increment of fuel mass flow rate entering the engine for combustion at the second branch, is the rate of change of the increment of fuel mass flow rate entering the engine for combustion, is the initial temperature of the fuel before entering the engine, c p is the constant pressure specific heat capacity, is the increment of heat generated by the on-board system due to changes in operating conditions, K is the heat transfer coefficient, A is the heat transfer area, C t is the constant pressure specific heat capacity matrix, is the rate of change of the increment of fuel temperature matrix.

[0067] wherein, , , , , h t , r t , Thead and l t are self-defined parameters.

[0068] The formula of the state space equation of the thermodynamic model can be obtained by The formula of the state space equation of the thermodynamic model can be obtained by ; The standard form of the state space equation of the thermodynamic model can be expressed as: ; wherein, ; wherein, is the flow of the fluid network, y is the increment of the system output, A T , B T , , are self-defined parameters.

[0069] In the calculation of the thermodynamic layer, only the static gain of the fluid can be considered, and then can be expressed as: ; wherein, K q_u is the gain matrix of the flow to the pressure, K q_e is the gain matrix of the flow to the fuel consumption rate of the engine, is the fluid mechanics layer characteristic parameter matrix, M h is the energy adjacency matrix, A p is the state matrix of the pressure layer, B p is the input associated matrix of the pressure layer, d h is the disturbance matrix of the thermodynamic layer, E p is the extended input matrix of the pressure layer.

[0070] After the standard form of the state space equation of the thermodynamic model is combined with , the state space equation of the control quantity-system temperature can be obtained, that is, the state space equation required by the present application, and then the first target fuel coolant flow and the second target fuel coolant flow can be calculated through the state space equation. Through the standard form of the state space equation of the thermodynamic model and , it can be known that the control quantity u h is the flow distribution that affects the temperature distribution in the system.

[0071] The controller determines the control parameters according to the corresponding target temperature and the collected temperature, and obtains the first target fuel coolant flow and the second target fuel coolant flow through integration of the control parameters by the state space equation.

[0072] Specifically, the first fuel coolant temperature and the second fuel coolant temperature are taken as T 3, the first target temperature and the first fuel coolant temperature are used to calculate the corresponding T 3, the second target temperature and the second fuel coolant temperature are used to calculate the corresponding T 3, the respective T 3 and T 3 are taken as the input of the state space equation of the resistive equation of the two-node fluid network, and the output obtained by the state space equation of the resistive equation of the two-node fluid network is and .

[0073] The obtained is taken as the input of the state space equation of the resistive equation of the two-node fluid network , and the first target rotating speed of the first centrifugal pump is obtained by substituting the state space equation of the resistive equation of the two-node fluid network n s1 The first target rotating speed n s1 is substituted into the flow rate-rotating speed relationship function to obtain the first target fuel coolant flow corresponding to the first centrifugal pump. The formula of the flow rate-rotating speed relationship function is: ; wherein, c 11 , c 12 , c 13 , c 14 , c 15 , c 16 is a polynomial fitting coefficient of the centrifugal pump, is the flow rate of the fuel coolant at the outlet of the first centrifugal pump, i.e., the first target fuel coolant flow, which is obtained by polynomial fitting through the relationship function . .

[0074] The obtained is taken as the input of the state space equation of the resistive equation of the two-node fluid network , the state space equation of the resistive equation of the two-node fluid network is substituted into the second target rotating speed of the second centrifugal pump n s2 , the second target rotating speed n s2 is substituted into the flow rate and rotating speed relationship function to obtain the second target fuel coolant flow rate corresponding to the second centrifugal pump. The formula of the flow rate and rotating speed relationship function is: ; wherein, is the flow rate of the fuel coolant at the outlet of the second centrifugal pump, i.e., the second target fuel coolant flow rate, which is specifically obtained by polynomial fitting of the relationship function . .

[0075] Subsequently, the rotating 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 the outlet of the first centrifugal pump to the first target fuel coolant flow rate, and the rotating 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 the outlet of the second centrifugal pump to the second target fuel coolant flow rate.

[0076] The space flight vehicle thermal management method provided by the present application forms a fuel-cooled heat dissipation circuit through the heat exchanger, the micro-channel structure and the fuel pipeline, and the fuel coolant in the heat dissipation circuit is used to dissipate heat for the on-board equipment and the parts to be cooled of the space flight vehicle, thereby achieving active thermal protection of the space flight vehicle. Meanwhile, the centrifugal pump is controlled by the first fuel coolant temperature and the second fuel coolant temperature to adjust the fuel coolant flow rate in the fuel pipeline, so as to actively cool the high-temperature equipment in urgent need of cooling among the on-board equipment and the parts to be cooled, i.e., the fuel coolant flow rate is used to cool the equipment with the highest temperature first, so as to realize controllable temperature and temperature gradient arrangement of the whole space flight vehicle, thereby actively reducing the temperature of the on-board equipment and / or the parts to be cooled, effectively reducing the passive thermal protection burden, and improving the thermal protection capability and the reusability of the space flight vehicle.

[0077] It should be noted that the above examples are only used for understanding the present application and do not constitute a limitation on the space flight vehicle thermal management method of the present application. More forms of simple transformation based on this technical concept are within the protection scope of the present application.

[0078] The present application provides a space flight vehicle thermal management device, which comprises at least one processor and a memory in communication connection with the at least one processor. The memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor to enable the at least one processor to execute the space flight vehicle thermal management method in Embodiment I.

[0079] Reference will now be made to Figure 4 , which shows a structural diagram of a space vehicle thermal management device suitable for use in implementing embodiments of the present application. The space vehicle thermal management device in embodiments of the present application can include, but is not limited to, mobile terminals such as mobile phones, notebook computers, digital broadcast receivers, PDAs (Personal Digital Assistants), PADs (Portable Application Descriptions), PMPs (Portable Media Players), car terminals (e.g., car navigation terminals), and the like, as well as stationary terminals such as digital TVs, desktop computers, and the like. Figure 4 The illustrated space vehicle thermal management device is merely an example and should not impose any limitations on the functions and use range of embodiments of the present application.

[0080] As Figure 4 shown, the space vehicle thermal management device can include a processing device 1001 (e.g., a central processing unit, a graphic processing unit, or the like) that can perform various appropriate actions and processes according to programs stored in a read-only memory (ROM) 1002 or loaded from a storage device 1003 into a random access memory (RAM) 1004. Various programs and data required for operation of the space vehicle thermal management device are also stored in the RAM 1004. The processing device 1001, the ROM 1002, and the RAM 1004 are connected to each other through a bus 1005. An input / output (I / O) interface 1006 is also connected to the bus. In general, the following systems can be connected to the I / O interface 1006: input devices 1007 including, for example, a touch screen, a touch pad, a keyboard, a mouse, an image sensor, a microphone, an accelerometer, a gyroscope, and the like; output devices 1008 including, for example, a liquid crystal display (LCD), a speaker, a vibrator, and the like; the storage device 1003 including, for example, a magnetic tape, a hard disk, and the like; and a communication device 1009. The communication device 1009 can allow the space vehicle thermal management device to communicate wirelessly or by wire with other devices to exchange data. Although the space vehicle thermal management device having various systems is shown in the drawing, it should be understood that all of the illustrated systems are not required to be implemented or provided. More or fewer systems can be alternatively implemented or provided.

[0081] In particular, according to the embodiments disclosed in the present application, the process described above with reference to the flowchart can be implemented as a computer software program. For example, the embodiments disclosed in the present application include a computer program product comprising a computer program carried on a computer readable medium, the computer program containing program codes for executing the method shown in the flowchart. In such embodiments, the computer program can be downloaded and installed from a network through a communication device, or installed from the storage device 1003, or installed from the ROM 1002. When the computer program is executed by the processing device 1001, the above-mentioned functions defined in the method of the embodiments disclosed in the present application are executed.

[0082] The space vehicle thermal management device provided by the present application adopts the space vehicle thermal management method in the above-mentioned embodiments, and can solve the technical problem of how to reduce the passive thermal protection burden and improve the thermal protection capability and reusability of the space vehicle. Compared with the prior art, the space vehicle thermal management device provided by the present application has the same beneficial effects as the space vehicle thermal management method provided by the above-mentioned embodiments, and other technical features in the space vehicle thermal management device are the same as the features disclosed in the previous embodiment method, which will not be repeated here.

[0083] It should be understood that various parts of the present application can be realized by hardware, software, firmware or a combination thereof. In the description of the above-mentioned embodiments, specific features, structures, materials or characteristics can be combined in any one or more embodiments or examples in a suitable manner.

[0084] The above is merely specific embodiments of the present application, but the protection scope of the present application is not limited thereto, and any person skilled in the art can easily think of changes or replacements within the technical scope disclosed in the present application, which should be covered within the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

[0085] The present application provides a computer readable storage medium having stored thereon computer readable program instructions (i.e. computer program) for executing the space vehicle thermal management method in the above-mentioned embodiments.

[0086] The computer readable storage medium provided in the present application may, for example, be a U disk, but is not limited to an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, system, or device, or any combination of the above. More specific examples of the computer readable storage medium can include, but are not limited to, an electrical connection having one or more conductive wires, a portable computer disk, a hard disk, a random access memory (RAM), a read only memory (ROM), an erasable programmable read only memory (EPROM or flash memory), an optical fiber, a portable compact disk read only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present embodiment, the computer readable storage medium can 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 can be transmitted by any suitable medium, including but not limited to an electrical wire, an optical cable, an RF (Radio Frequency), and the like, or any suitable combination of the above.

[0087] The above computer readable storage medium can be contained in the thermal management device of the aerospace vehicle, or can exist separately without being assembled into the thermal management device of the aerospace vehicle.

[0088] The above computer readable storage medium carries one or more programs, when the one or more programs are executed by the thermal management device of the aerospace vehicle, the thermal management device of the aerospace vehicle: 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; based on the first fuel coolant temperature and the second fuel coolant temperature, determine 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; 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.

[0089] Computer program code for carrying out operations of the present application can be written in any combination of one or more programming languages, including an object oriented programming language such as Java, Smalltalk, C++ or the like and conventional procedural programming languages, such as the "C" programming language or similar programming languages. The program code can execute entirely on the user's computer, partly on the user's computer, as a stand-alone software package, partly on the user's computer and partly on a remote computer or entirely on the remote computer or server. In the latter scenario, the remote computer can be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or the connection can be made to an external computer (for example, through the Internet using an Internet Service Provider).

[0090] The flow diagrams and the block diagrams in the drawings are illustrations of architectures, functionalities, and operations of possible implementations of systems, methods, and computer program products according to various embodiments of the present application. In this regard, each block in the flow diagrams or block diagrams can represent a module, a segment, or a portion of code, which comprises one or more executable instructions for implementing the specified logical function(s). It should also be noted that in some alternative implementations, the functions noted in the blocks can occur out of the order noted in the figures. For example, two blocks shown in succession may, in fact, be executed substantially concurrently or the blocks may

[0091] The modules involved in the embodiments of the present application can be implemented in the form of software or in the form of hardware. In some cases, the name of the module does not constitute a limitation on the module itself.

[0092] The readable storage medium provided by the present application is a computer readable storage medium, which stores computer readable program instructions (i.e., a computer program) for executing the above-mentioned space flight vehicle thermal management method, and can solve the technical problem of how to reduce the passive thermal protection burden, improve the thermal protection capability of the space flight vehicle, and improve the reusability. Compared with the prior art, the computer readable storage medium provided by the present application has the same beneficial effects as the space flight vehicle thermal management method provided by the above-mentioned embodiments, and will not be described here.

[0093] The application also provides a computer program product comprising a computer program which, when executed by a processor, implements the steps of the method for thermal management of a space-aircraft as described above.

[0094] The computer program product provided by the application can solve the technical problem of how to reduce the passive thermal protection burden and improve the thermal protection capability and reusability of the space-aircraft. Compared with the prior art, the beneficial effects of the computer program product provided by the application are the same as those of the method for thermal management of a space-aircraft provided by the above-mentioned embodiments, and are not described here.

[0095] The above only describes some embodiments of the application, and does not limit the patent scope of the application. Any equivalent structural transformation made by using the content of the specification and drawings, or direct / indirect application in other related technical fields under the technical concept of the application is included in the patent protection scope of the application.

Claims

1. A method of thermal management of an aerospace vehicle, characterized by, The space-aircraft 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 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 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-aircraft comprises: Real-time acquisition of the first fuel coolant temperature corresponding to each 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.

2. The method of thermal management of a spaceplane as recited in claim 1, wherein, 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.

3. The method of thermal management of a spaceplane as recited in claim 2, wherein, 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: Input of 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 determination of the first target fuel coolant flow rate and the second target fuel coolant flow rate through the output of the state space equation.

4. The method of thermal management of a space vehicle of claim 1, wherein, The step of real-time acquisition of the first fuel coolant temperature corresponding to each heat exchanger and the second fuel coolant temperature corresponding to each micro-channel structure comprises: Real-time acquisition of the first fuel coolant temperature through a temperature sensor at the outlet of each first heat exchanger; The second fuel coolant temperature is obtained in real time by a temperature sensor at the outlet of each microchannel structure.

5. The method of thermal management of a space vehicle of claim 4, 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 communicated with the multi-way joint.

6. The method of thermal management of a space vehicle of any of claims 1 to 5, wherein, The part to be cooled at least includes the leading edge of the aircraft and the leading edge of the wing.

7. 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 space thermal management method of the space vehicle according to any one of claims 1 to 6.

8. 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 space thermal management method of the space vehicle according to any one of claims 1 to 6.

Images