A thermal management system model for aviation gas turbine engines

Abstract

The present application provides a model of a thermal management system for an aviation gas turbine engine, which belongs to the technical field of thermal management of aviation engines. The system model includes: an engine controller sub-model, which is used to calculate the current engine geometric adjustable variables and fuel flow rate according to the current atmospheric environmental conditions and engine operating instructions, and calculate the hot return oil flow rate according to the limiting temperature and the actual feedback temperature corresponding to the limiting temperature point determined according to the temperature bearing characteristics and safety requirements of the engine itself; an engine performance calculation sub-model, which is used to calculate the engine performance parameters required by the engine mechanical system sub-model and / or the engine fuel system sub-model; an engine mechanical system sub-model, which is used to calculate the heat generated by the engine bearing and the heat exchange with the air outside the bearing cavity; an engine fuel system sub-model, which is used to calculate the heat generated by the engine fuel system; and a radiator sub-model, which is used to calculate the temperature and pressure of the corresponding outlet after the heat exchange between the fuel and lubricating oil.

Description

Technical Field

[0001] The present application belongs to the field of aviation engine thermal management technology, and in particular, relates to an aviation gas turbine engine thermal management system model. Background Art

[0002] As the number of aircraft-borne equipment increases and the heat generation power continues to increase, traditional air circulation can no longer meet the cooling needs of aircraft-borne equipment. In addition, in order to ensure the reliable and high-speed operation of aviation gas turbine engines, the lubricating oil flowing through the bearings of each pivot also needs to be dissipated in time. In order to consume the excess heat generated by the aircraft and engine subsystems, some engines will establish an engine thermal management system with fuel as the heat transfer medium, and achieve heat management through automatic control of flow allocation, so as to avoid forcing pilots to intervene due to overheating and cooling by increasing fuel consumption.

[0003] However, traditional engine heat transfer modeling analysis has the following shortcomings:

[0004] 1) Difficulty in integrating engine components and subsystems for heat transfer analysis:

[0005] The analysis objects of the engine thermal management system include not only the complete machine components, such as fans, compressors, combustion chambers, high-pressure turbines, low-pressure turbines, and outer ducts, but also the engine mechanical systems, such as various pivot bearings, accessory casings, lubricant-fuel radiators, and lubricant pumps, as well as the engine fuel system, such as various fuel pumps and fuel regulators. The types of media for heat transfer between the complete machine components and the mechanical and fuel systems are not single, and include air, lubricant oil, and fuel. If a complete three-dimensional structural simulation model is established and analyzed with the help of the finite element method, in order to ensure the simulation accuracy, a large number of grids will inevitably make the iterative calculation extremely time-consuming and inefficient. Therefore, usually only simulation calculations of some typical characteristic points within the flight envelope are performed, and it is difficult to achieve continuous changes in the engine state or real-time state analysis.

[0006] 2) Only heat transfer analysis cannot be used to design the control algorithm of the thermal management system:

[0007] The purpose of the engine thermal management system is to meet the heat dissipation requirements of the entire aircraft by automatically adjusting the flow rate. Whether the thermal management is effective and whether it can avoid the pilot being forced to intervene and increase fuel consumption to achieve heat dissipation depends on the logic algorithm of the engine controller software. However, flow control itself is a continuous transient process. It is impossible to carry out the design of thermal management control algorithms by relying only on the heat transfer analysis of limited discrete steady-state points.

[0008] 3) It is impossible to achieve the integrated design of the thermal management system (aircraft-engine integration):

[0009] For an engine thermal management system that uses fuel as the core heat transfer medium, the core function of the system is how to make full use of the fuel tank fuel heat sink to meet the heat dissipation needs of the entire aircraft. The direct heat dissipation source of the aircraft, in addition to engine combustion, is only the aircraft ram air heat dissipation. Therefore, engine thermal management design and optimization require an integrated aircraft-engine design, while taking into account the heating and heat dissipation of the aircraft and engine subsystems. This cannot be achieved by heat transfer modeling and analysis of engine components and subsystems alone.

[0010] The above three problems limit the rationality and compliance verification of the thermal management system design. At the very least, it is impossible to achieve intelligent automatic regulation of heat and more effective utilization of the fuel heat sink in the fuel tank. At worst, the thermal management system cannot meet the heat dissipation needs of the entire aircraft, requiring pilots to intervene frequently and consume excess heat by increasing fuel, which seriously affects the flight range and flight time, and even causes flight safety issues when the engine cannot improve its status due to a malfunction. Summary of the invention

[0011] The purpose of the present application is to provide an aviation gas turbine engine thermal management system model to solve or alleviate at least one problem in the background technology.

[0012] The technical solution of the present application is: a thermal management system model for an aviation gas turbine engine, the thermal management system model comprising:

[0013] An engine controller sub-model, the engine controller sub-model is used to calculate the current engine geometric adjustable variables and fuel flow rate according to the current atmospheric environmental conditions and engine operation instructions, and the engine controller sub-model is used to calculate the hot return oil flow rate according to the limit temperature and the actual feedback temperature corresponding to the limit temperature point determined by the engine's own temperature bearing characteristics and safety requirements;

[0014] An engine performance calculation sub-model, the engine performance calculation sub-model being used to calculate engine performance parameters required by an engine mechanical system sub-model and / or an engine fuel system sub-model;

[0015] An engine mechanical system submodel, wherein the engine mechanical system submodel is used to calculate the heat generated by the engine bearing and the heat exchange of the air outside the bearing cavity, wherein the heat generated by the bearing is calculated by using different friction heat generation calculation methods according to the actual bearing type;

[0016] An engine fuel system submodel, wherein the engine fuel system submodel is used to calculate the heat generated by the engine fuel system, and the heat generated by the engine fuel system includes the power loss of the system itself and the heat exchanged through the radiator;

[0017] The radiator sub-model is used to calculate the temperature and pressure of the corresponding outlet after the heat exchange between the fuel and the lubricating oil. The radiator sub-model needs to select the corresponding calculation method according to the different radiator types, and the calculation of the radiator flow resistance characteristics needs to be determined according to the different types of radiator structures.

[0018] Furthermore, the input parameters of the engine controller sub-model include:

[0019] Atmospheric conditions including flight altitude, Mach number, and indicated flight speed; and

[0020] Engine control commands including engine throttle lever angle;

[0021] The output parameters of the engine controller sub-model include:

[0022] Engine geometry adjustable variables including low pressure compressor adjustable guide vane angle, high pressure compressor adjustable guide vane angle, and nozzle exit area; and

[0023] The fuel flow rate includes the main combustion chamber fuel flow rate and the afterburner fuel flow rate.

[0024] Furthermore, the input parameters of the engine performance calculation sub-model include:

[0025] Atmospheric environmental conditions including flight altitude, Mach number, and indicated flight speed;

[0026] Engine geometry adjustable variables including low pressure compressor adjustable guide vane angle, high pressure compressor adjustable guide vane angle, and nozzle exit area; and

[0027] The fuel flow rate includes the fuel flow rate of the main combustion chamber and the fuel flow rate of the afterburner;

[0028] The output parameters of the engine performance calculation sub-model include:

[0029] The physical speed including the engine low-pressure rotor speed and high-pressure rotor speed; and

[0030] The cross-sectional temperature includes the fan inlet total temperature, compressor inlet total temperature, compressor outlet total temperature, and low-pressure turbine outlet total temperature.

[0031] Furthermore, the engine performance calculation sub-model is constructed using a component method, and the process includes:

[0032] Separating the engine into a plurality of independent sub-assemblies, the sub-assemblies including an inlet, a compression section, a main combustion chamber, a turbine, a mixing chamber, an afterburner, and a tail nozzle;

[0033] According to the aerodynamic thermodynamics principle followed during the operation of the engine, each sub-component is modeled separately using the variable specific heat calculation method;

[0034] Then, a nonlinear equation group is established according to the constraints of the common working conditions of the engines, and the nonlinear equation group is solved to obtain the common working point of the engines, and then the required output parameters are solved.

[0035] Furthermore, the input parameters of the engine mechanical system sub-model include:

[0036] Contains physical speeds for calculating the engine low-pressure rotor speed and high-pressure rotor speed for calculating speed quantities such as the circumferential speed of the roller bearing cage and the angular speed of the ball bearing;

[0037] Contains the cross-sectional temperatures of the fan inlet total temperature, compressor inlet total temperature, compressor outlet total temperature, and low-pressure turbine outlet total temperature used to calculate the temperatures of each bearing cavity of the engine; and

[0038] The temperature, pressure and flow rate of the lubricating oil after passing through each radiator;

[0039] The output parameters of the engine mechanical system sub-model include:

[0040] The temperature and pressure of the lubricating oil that is mixed and sent to the radiator after dissipating heat from each bearing cavity of the engine.

[0041] Furthermore, the input parameters of the engine fuel system sub-model include:

[0042] The engine high pressure rotor speed used to calculate the speed of each pump;

[0043] The fuel flow rate includes the main combustion chamber fuel flow rate and the afterburner fuel flow rate; and

[0044] The hot oil return flow rate of each channel is calculated according to the thermal management control algorithm;

[0045] The output parameters of the engine fuel system submodel include:

[0046] The temperature and pressure of the fuel flowing to each radiator.

[0047] Furthermore, the sources of the system's own power loss include: power loss caused by the efficiency of the fuel pump, power loss caused by the pressure loss of the fuel along the way, and power loss caused by the flow loss of the fuel system; among them, the total loss is regarded as the difference between the input power of the fuel system and the effective power of the fuel system. The input power of the fuel system refers to the output power provided by the engine accessory casing to each fuel pump, and the effective power of the fuel system includes the effective power of supplying fuel to the combustion chamber and driving the servo mechanism.

[0048] Furthermore, the radiator sub-model includes one or more radiators, and the input parameters of each radiator include:

[0049] The temperature, pressure and flow rate of the fuel before heat exchange;

[0050] The temperature, pressure and flow rate of lubricating oil before heat exchange;

[0051] The output parameters for each heat sink include:

[0052] The temperature and pressure of the fuel after heat exchange;

[0053] The temperature and pressure of the lubricating oil after heat exchange.

[0054] The aviation gas turbine engine thermal management system model proposed in the present invention can realize the quasi-one-dimensional thermal simulation calculation of all engine components and fuel / lubricating oil systems, and can perform rapid thermal calculation analysis of any steady-state point or continuous transient process within the full flight envelope. The engine thermal management system model provides a thermal management control algorithm interface and a joint calculation interface for the integrated design of the aircraft and engine, which can better carry out the improvement and optimization design of the thermal management system, realize more adequate utilization of the fuel tank fuel heat sink, and more intelligent automatic heat regulation, meet the heat dissipation needs of the whole machine, and try to avoid relying on the pilot to consume excess heat by increasing fuel to ensure flight range, flight time and safety. In addition, the method has wide applicability, strong expansibility, and is easy to improve. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0056] Figure 1 This is a schematic diagram of the aviation gas turbine engine thermal management system model of the present application.

[0057] Figure 2 Schematic diagram of the input / output parameters of the engine controller sub-model in this application.

[0058] Figure 3 Schematic diagram of input / output parameters of the engine performance calculation sub-model in this application.

[0059] Figure 4 This is a schematic diagram of the input / output parameters of the engine mechanical system sub-model in this application.

[0060] Figure 5 Schematic diagram of the input / output parameters of the engine fuel system sub-model in this application.

[0061] Figure 6 Schematic diagram of input / output parameters of the heat sink sub-model in this application. DETAILED DESCRIPTION

[0062] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application.

[0063] The present application proposes a thermal management system model for an aviation gas turbine engine, aiming to solve the drawback that existing three-dimensional heat transfer modeling methods cannot integrate all the components or systems of the entire machine.

[0064] like Figure 1 As shown, the aviation gas turbine engine thermal management system model provided by this application includes the following parts:

[0065] The engine thermal management system model 10 is divided into five sub-models according to the functions, and the five sub-models include an engine controller sub-model 11, an engine performance calculation sub-model 12, an engine mechanical system sub-model 13, an engine fuel system sub-model 14 and a radiator sub-model 15;

[0066] 1) The engine controller sub-model is mainly used to simulate the operation of the onboard software in the digital electronic controller, that is, to calculate the current engine geometric adjustable variables and fuel flow according to the current atmospheric environmental conditions and engine control instructions. This part of the control process is formulated according to the engine's own control plan.

[0067] See also Figure 2 As shown in FIG. 1 , the input parameters of the engine controller sub-model include: atmospheric environmental conditions such as flight altitude (H), Mach number (Ma), flight speed indicator (Vb); and engine control instructions such as engine throttle lever angle (PLA).

[0068] The output parameters of the engine controller sub-model include: engine geometric adjustable variables such as the low-pressure compressor adjustable guide vane angle (αf), the high-pressure compressor adjustable guide vane angle (αc), and the nozzle outlet area (A8); and fuel flow rates such as the main combustion chamber fuel flow rate (Wfb) and the afterburner fuel flow rate (Wfab).

[0069] In addition, the engine controller sub-model provides an interface for the thermal management system control algorithm, such as determining the temperature point that needs to be limited according to the engine's own temperature characteristics and safety requirements, and setting the preset limit temperature and the actual feedback temperature (Tlimit_1 ～ Tlimit_n), the hot return oil flow rate (Wrhy_1~Wrhy_n) is calculated through the control algorithm.

[0070] 2) The engine performance calculation submodel is used to calculate the engine speed, cross-sectional temperature and other performance parameters required by other submodels (engine mechanical system submodel and engine fuel system submodel). Among them, the engine performance calculation submodel is built using the component method, that is, the engine is divided into multiple independent subcomponents, including the intake duct, compression component, main combustion chamber, turbine, mixing chamber, afterburner and tail nozzle, etc. According to the aerodynamic thermodynamics principle followed during the engine operation, each subcomponent is modeled separately using the variable specific heat calculation method; then, a nonlinear equation group is established based on the constraints of the common operation of the engine; finally, the nonlinear equation group is solved to find the common operating point of the engine, and then the required output parameters are solved.

[0071] like Figure 3 As shown in the figure, the input parameters of the engine performance calculation sub-model include: atmospheric environmental conditions such as flight altitude (H), Mach number (Ma), flight speed (Vb), etc.; engine geometric adjustable variables such as low-pressure compressor adjustable guide vane angle (αf), high-pressure compressor adjustable guide vane angle (αc), nozzle outlet area (A8), etc.; fuel flow rates such as main combustion chamber fuel flow rate (Wfb) and afterburner fuel flow rate (Wfab).

[0072] The output parameters of the engine performance calculation sub-model include: engine low-pressure rotor speed (N1), high-pressure rotor speed (N2); fan inlet total temperature (T2), compressor inlet total temperature (T25), compressor outlet total temperature (T3), low-pressure turbine outlet total temperature (T6) and other cross-sectional temperatures.

[0073] 3) The engine mechanical system submodel is mainly used to calculate the heat generated by the engine bearings and the heat exchange of the air outside the bearing cavity. The heat generated by the bearings needs to be calculated using different friction heat generation calculation methods according to the actual bearing type (such as roller bearings, ball bearings). The calculation methods for friction heat generation of each bearing type can be found in the references of the corresponding bearings, and will not be repeated here. The geometric structure of the outer wall of the front, middle and rear bearing cavities of the engine can be simplified and approximated by cylinders, truncated cones, etc. The outer wall temperature of the bearing cavity is the mixed temperature of the cross-section of each related component.

[0074] like Figure 4 As shown in the figure, the input parameters of the engine mechanical system sub-model include: the engine low-pressure rotor speed (N1) and high-pressure rotor speed (N2) for calculating the circumferential speed of the roller bearing cage, the angular velocity of the ball bearing and other speed quantities; the fan inlet total temperature (T2), compressor inlet total temperature (T25), compressor outlet total temperature (T3), low-pressure turbine outlet total temperature (T6) and other cross-sectional temperature parameters for calculating the temperature of each bearing cavity of the engine; in addition, it also includes the temperature (Toil_after heat dissipation), pressure (Poil_after heat dissipation) and flow rate (Qoil) of the lubricating oil after flowing through each radiator.

[0075] The output parameters of the engine mechanical system sub-model include: the temperature (Toil_before heat dissipation) and pressure (Poil_before heat dissipation) of the lubricating oil mixed and sent to the radiator after cooling each bearing cavity of the engine.

[0076] 4) The engine fuel system submodel is used to calculate the heat generated by the engine fuel system. The heat sources are two parts, including the system's own power loss and heat exchange through the radiator. The sources of the system's own power loss include: power loss caused by the efficiency of the fuel pump, power loss caused by the pressure loss of the fuel along the way, and power loss caused by the flow loss of the fuel system (including gear pump oil return and system oil return, etc.). The total loss can be regarded as the difference between the input power (provided by the rotation of the engine rotor and the transmission gear of the casing) and the effective power of the fuel system. The input power of the fuel system refers to the output power provided by the engine accessory casing to each fuel pump; the effective power of the fuel system includes the effective power of supplying fuel to the combustion chamber and driving the servo mechanism.

[0077] like Figure 5 As shown in the figure, the input parameters of the engine fuel system sub-model include: the engine high-pressure rotor speed (N2) used to calculate the speed of each pump; the fuel flow rates such as the main combustion chamber fuel flow (Wfb) and the afterburner fuel flow (Wfab); and the various hot return oil flows (Wrhy_1~Wrhy_n) calculated according to the thermal management control algorithm.

[0078] The output parameters of the engine fuel system submodel include: the temperature (Tfuel_before cooling) and pressure (Pfuel_before cooling) of the fuel flowing to each radiator.

[0079] The initial inlet pressure, temperature, flow rate of the engine fuel system, as well as the return oil pressure, temperature, and flow rate of the engine to the aircraft, can provide an interface for the integrated joint calculation of flight thermal management.

[0080] 5) The radiator sub-model is used to calculate the temperature and pressure of the corresponding outlet after the heat exchange between the fuel and lubricating oil. It is the heat exchange medium between the engine fuel system and the mechanical system. Among them, the radiator sub-model needs to select the corresponding calculation method according to different radiator types (such as fixed tube plate shell and tube radiator, U-tube shell and tube radiator, etc., in which the cold end fuel flows in the carrier tube and the hot end lubricating oil flows in the tube shell). The calculation of the radiator flow resistance characteristics needs to be specifically analyzed according to different types of radiator structures.

[0081] Taking the shell and tube radiator as an example, the flow resistance on the shell side and the tube side needs to be analyzed separately. The shell side resistance can be calculated by Bell's method: the shell side lubricating oil pressure loss of the heat exchanger with the circular baffle installed is equal to the sum of the pressure loss when the cross flow flows with the tube bundle in the middle of the baffle, the pressure loss when the cross flow flows with the tube bundle at the inlet / outlet baffle, and the pressure loss when the flow is parallel to the tube axis through the circular part of the baffle. The tube side resistance is mainly due to the flow contraction at the inlet and outlet of the heat transfer tube core body when the fuel flows through, and the flow expansion at the outlet. The fuel pressure loss corresponding to this sudden flow contraction and expansion, as well as the friction loss when the fuel flows through the heat transfer tube, and the additional pressure loss when flowing through the turn or U-shaped section of the tube box, together constitute the total fuel pressure drop.

[0082] like Figure 6 As shown, the radiator submodel includes one or more radiators, and the input parameters of each radiator include: the temperature (Tfuel_Before heat dissipation), pressure (Pfuel_Before heat dissipation) and flow rate (Qfuel) of the fuel before heat exchange; the temperature (Toil_Before heat dissipation), pressure (Poil_Before heat dissipation) and flow rate (Qoil) of the lubricating oil before heat exchange.

[0083] The output parameters of each radiator include: the temperature (Tfuel_after cooling) and pressure (Pfuel_after cooling) of the fuel after heat exchange; the temperature (Toil_after cooling) and pressure (Poil_after cooling) of the lubricating oil after heat exchange.

[0084] The thermal management system model constructed by the above method of the present application has a thermal management control algorithm interface and an interface for joint thermal simulation calculation with the aircraft, which can carry out advanced algorithm design and simulation verification, as well as joint calculation with the aircraft system to realize the integrated design of the aircraft and the engine. The aviation gas turbine engine thermal management system model constructed by this method is easy to improve and has wide applicability. The thermal management system optimized by using this model can more safely ensure that the whole system does not overheat and make more full use of the fuel heat sink, ensuring flight safety while realizing the unconstrained operation of the engine by the pilot.

[0085] The aviation gas turbine engine thermal management system model proposed in the present invention can realize the quasi-one-dimensional thermal simulation calculation of all engine components and fuel / lubricating oil systems, and can perform rapid thermal calculation analysis of any steady-state point or continuous transient process within the full flight envelope. The engine thermal management system model provides a thermal management control algorithm interface and a joint calculation interface for the integrated design of the aircraft and engine, which can better carry out the improvement and optimization design of the thermal management system, realize more adequate utilization of the fuel tank fuel heat sink, and more intelligent automatic heat regulation, meet the heat dissipation needs of the whole machine, and try to avoid relying on the pilot to consume excess heat by increasing fuel to ensure flight range, flight time and safety. In addition, the method has wide applicability, strong expansibility, and is easy to improve.

[0086] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A thermal management system model for an aviation gas turbine engine, characterized in that: The thermal management system model includes: An engine controller submodel, the engine controller submodel is used to calculate the current engine geometric adjustable variables and fuel flow rate according to the current atmospheric environmental conditions and engine control instructions, and the engine controller submodel calculates the hot return oil flow rate according to the limiting temperature and the actual feedback temperature corresponding to the limiting temperature point determined by the engine's own temperature bearing characteristics and safety requirements, wherein the input parameters of the engine controller submodel include: atmospheric environmental conditions including flight altitude, Mach number, and flight speed, and engine control instructions including engine throttle lever angle; the output parameters of the engine controller submodel include: engine geometric adjustable variables including low-pressure compressor adjustable guide vane angle, high-pressure compressor adjustable guide vane angle, and nozzle outlet area, and fuel flow including main combustion chamber fuel flow and afterburner fuel flow; An engine performance calculation submodel, the engine performance calculation submodel is used to calculate the engine performance parameters required by the engine mechanical system submodel and / or the engine fuel system submodel, wherein the input parameters of the engine performance calculation submodel include: atmospheric environmental conditions including flight altitude, Mach number, and flight speed, engine geometric adjustable variables including low-pressure compressor adjustable guide vane angle, high-pressure compressor adjustable guide vane angle, and nozzle outlet area, and equal fuel flow including main combustion chamber fuel flow and afterburner fuel flow; the output parameters of the engine performance calculation submodel include: physical speed including engine low-pressure rotor speed and high-pressure rotor speed, and cross-sectional temperatures including fan inlet total temperature, compressor inlet total temperature, compressor outlet total temperature, and low-pressure turbine outlet total temperature; The engine mechanical system submodel is used to calculate the heat generated by the engine bearing and the heat exchange of the air outside the bearing cavity, wherein the heat generated by the bearing adopts different friction heat generation calculation methods according to the actual bearing type, wherein the input parameters of the engine mechanical system submodel include: the physical speed of the engine low-pressure rotor speed and high-pressure rotor speed for calculating the circumferential speed of the roller bearing cage and the angular velocity of the ball bearing, the cross-sectional temperature of the fan inlet total temperature, the compressor inlet total temperature, the compressor outlet total temperature, the low-pressure turbine outlet total temperature for calculating the temperature of each bearing cavity of the engine, and the temperature, pressure and flow rate of the lubricating oil after passing through each radiator for heat dissipation; the output parameters of the engine mechanical system submodel include: the temperature and pressure corresponding to the lubricating oil mixed and led to the radiator after heat dissipation in each bearing cavity of the engine; An engine fuel system submodel, the engine fuel system submodel is used to calculate the heat generated by the engine fuel system, the heat generated by the engine fuel system includes the power loss of the system itself and the heat exchanged through the radiator, wherein the input parameters of the engine fuel system submodel include: the engine high-pressure rotor speed used to calculate the speed of each pump, the fuel flow including the main combustion chamber fuel flow, the afterburner fuel flow and the various hot return oil flows calculated according to the thermal management control algorithm; the output parameters of the engine fuel system submodel include: the temperature and pressure of the fuel flowing to each radiator; the sources of the system's own power loss include: the power loss caused by the efficiency of the fuel pump, the power loss caused by the pressure loss of the fuel along the way and the power loss caused by the flow loss of the fuel system; wherein the total loss is regarded as the difference between the input power of the fuel system and the effective power of the fuel system, the input power of the fuel system refers to the output power provided by the engine accessory casing to each fuel pump, and the effective power of the fuel system includes the effective power of supplying fuel to the combustion chamber and driving the servo mechanism; A radiator sub-model, wherein the radiator sub-model is used to calculate the temperature and pressure of the corresponding outlets after the heat exchange between the fuel and the lubricating oil, wherein the radiator sub-model needs to select the corresponding calculation method according to different radiator types, and the calculation of the radiator flow resistance characteristics needs to be determined according to different types of radiator structures, wherein the radiator sub-model includes one or more radiators, and the input parameters of each radiator include: the temperature, pressure and flow rate of the fuel before the heat exchange and the temperature, pressure and flow rate of the lubricating oil before the heat exchange; the output parameters of each radiator include: the temperature and pressure of the fuel after the heat exchange and the temperature and pressure of the lubricating oil after the heat exchange.

2. The aviation gas turbine engine thermal management system model according to claim 1, characterized in that: The engine performance calculation sub-model is built using the component method, and the process includes: Separating the engine into a plurality of independent sub-assemblies, the sub-assemblies including an inlet, a compression section, a main combustion chamber, a turbine, a mixing chamber, an afterburner, and a tail nozzle; According to the aerodynamic thermodynamics principle followed during the operation of the engine, each sub-component is modeled separately using the variable specific heat calculation method; Then, a nonlinear equation group is established according to the constraints of the common working conditions of the engines, and the nonlinear equation group is solved to obtain the common working point of the engines, and then the required output parameters are solved.

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