Method for generating digital prototype of aero-engine gas path system based on system modeling language

Abstract

The invention belongs to the technical field of digital prototypes of aero-engine gas path systems, and particularly relates to a method for generating a digital prototype of an aero-engine gas path system based on a system modeling language. Starting from demand definition of the aero-engine gas path system, demand, structure, behavior and parameter models of the system are established, a mechanism model of the established system is deduced, and a performance model of the system is established in combination with a digital prototype tool, so that the demand verification digital prototype of the aero-engine gas path system is directly generated. The virtual test is carried out in the simulation environment, the actual operation condition of the engine is simulated, closed-loop verification of requirements is effectively achieved, forward design of an overall scheme is completed, logic errors in the design are reduced, the research and development period is shortened, and the research and development cost is reduced.

Description

Technical Field

[0001] The invention belongs to the technical field of digital prototypes of aircraft engine gas path systems, and in particular relates to a method for generating a digital prototype of an aircraft engine gas path system based on a system modeling language. Background Art

[0002] Aircraft engines are highly complex and sophisticated thermal machines that need to work continuously in extreme environments, so there are strict requirements for their design and manufacturing level. The performance of the air path system of an aircraft engine is crucial to the normal operation, thrust generation and overall performance of the aircraft. The design of the air path system directly affects the efficiency and performance of the engine. Therefore, the development and design of the air path system of an aircraft engine is of great significance.

[0003] System engineering theory plays a leading role in the development of aerospace, and the development of aircraft engines is a typical complex system engineering. With the increase in demand for aircraft engine development and the complexity of tasks, traditional text-based system engineering methods have gradually shown their limitations, including inaccurate system description, poor relevance of design content, and disconnection between system verification and system design. For example, the literature "Li Chen, Wu Xin, Cui Lifeng, et al. Design method of aircraft engine control system based on MBSE thinking [J]. Aero Engines, 2021, 47(04): 123-130. DOI: 10.13477 / j.cnki.aeroengine.2021.04.015.".

[0004] Model-Based Systems Engineering (MBSE) provides a comprehensive system perspective that allows engineers to manage and analyze the entire system lifecycle, including requirements, design, verification, and maintenance. MBSE uses models for top-down analysis and bottom-up synthesis. Through model analysis and simulation, it helps to discover design problems at an early stage and optimize the overall design solution. At the same time, MBSE is the main tool for building a model-driven digital system and promoting digital transformation. It is based on the commonly used Systems Modeling Language (SysML) and follows the new R&D model of "digital design-virtual experiment-digital manufacturing-physical manufacturing" to ensure that products are fully verified in a virtual environment, improve R&D efficiency and quality, and reduce R&D costs and risks. Summary of the invention

[0005] The purpose of the present invention is to develop a method for generating a digital prototype of an aero-engine gas path system based on a system modeling language, starting from the demand definition of the aero-engine gas path system, establishing the demand, structure, behavior and parameter model of the system, deriving and building the mechanism model of the system, and combining the digital prototype tool to build the performance model of the system, thereby directly generating a digital prototype for demand verification of the aero-engine gas path system. Conduct virtual experiments in a simulation environment to simulate the actual operation of the engine, effectively realize closed-loop verification of the demand, complete the forward design of the overall plan, reduce logical errors in the design, shorten the R&D cycle, and reduce R&D costs.

[0006] The present invention adopts the following technical solution:

[0007] A method for generating a digital prototype of an aero-engine air path system based on a system modeling language comprises the following steps:

[0008] Step 1: Modeling the requirements and architecture of the aircraft engine gas path system.

[0009] Based on SysML, the aircraft engine gas path system is modeled from four aspects: requirements, behavior, structure, and parameters. Starting from the top-level requirements, the design purpose of the system is clarified, and then the requirements are refined into functional activity behavior analysis, static system structure composition, and system parameter constraint relationship to complete the architecture modeling of the overall solution.

[0010] Step 2: Infer the mechanism model of the air path system of the air-driving engine.

[0011] According to the workflow of the aircraft engine air path system obtained in step 1, the attributes and behavior analysis of the 11 components, including atmosphere, air inlet, fan, bypass, low-pressure compressor, high-pressure compressor, combustion chamber, high-pressure turbine, low-pressure turbine, inner duct nozzle, and outer duct nozzle, are carried out in turn to build the mechanism model of the aircraft engine. Among them, attribute analysis refers to analyzing the function of the component, describing the operating characteristics, component parameters, and intermediate variables of each component. Behavior analysis refers to using the corresponding thermodynamic equations to describe the dynamic behavior of the component.

[0012] Step 3: Performance simulation and verification of aircraft engine gas path system.

[0013] Based on the mechanism model obtained in step 2, a digital prototype tool (MWORKS is used here, but not limited to this tool) is used to build a system performance model for simulation testing: First, the input and output interfaces of each component are defined and instantiated through text modeling. Then, the component models are encapsulated in sequence and connected according to the physical topology of the components to form a digital prototype model of the system. Finally, the simulation parameters and simulation variables are set, and the performance indicators of the system are simulated and calculated throughout the process, and the calculation results are intuitively displayed through the user interface, thereby verifying the top-level requirements of the system and completing the forward design process from concept to implementation.

[0014] Beneficial effects of the present invention:

[0015] (1) Requirements verification and design optimization: The system modeling language is used to build a digital prototype for requirements verification of the aircraft engine gas path system, which realizes closed-loop verification of requirements and completes the forward design of the overall solution. This helps to reduce logical errors in the design, improve R&D quality, and reduce R&D risks.

[0016] (2) Virtual simulation and performance prediction: Digital prototypes are directly generated from the system modeling language, replacing part of the physical prototype verification process. The engine is simulated and tested in a virtual environment, which can be fully verified and demonstrated before actual production. In addition, the engine performance can be verified under different design and operating conditions to ensure that it meets the expected goals and achieve performance prediction and verification. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 It is a design flow chart of the present invention;

[0018] Figure 2 is the demand graph of the present invention;

[0019] Figure 3 is a use case diagram of the present invention;

[0020] Figure 4 is an activity diagram of the present invention;

[0021] Figure 5 It is a sequence diagram of the present invention;

[0022] Figure 6 is a block definition diagram of the present invention;

[0023] Figure 7 is an internal block diagram of the present invention;

[0024] Figure 8 It is a simplified diagram of the gas path system of the present invention;

[0025] Fig. 9 It is a demonstration diagram of the Modelica digital prototype model of the present invention;

[0026] Fig.10 It is a performance model input parameter diagram of the present invention. DETAILED DESCRIPTION

[0027] The specific implementation of the present invention is further described below in conjunction with the accompanying drawings and technical solutions.

[0028] like Figure 1 The method of generating a digital prototype of an aeroengine gas path system based on a system modeling language of the present invention comprises:

[0029] Step 1: Modeling the requirements and architecture of the aircraft engine gas path system.

[0030] (1.1) Conduct demand analysis:

[0031] Analyze the design goals of the system, clarify the design purpose of the system, capture the top-level task requirements and subdivide the requirements, and generate a requirement diagram such as Figure 2 As shown. The top-level requirement for aircraft engine research and development is to provide the required thrust according to different flight conditions and missions (id="1", text="provide the required thrust according to the different flight conditions and flight missions of the aircraft"). In this process, certain parameters need to be introduced to calculate the thrust. These parameters include the intake total pressure recovery coefficient (id="1.1", text="sigma1=0.97"), low-pressure compressor pressure ratio (id="1.2", text="pi_lpc=3"), bypass pressure ratio (id="1.3", text="pi_bypass=2"), high-pressure compressor pressure ratio (id="1.4", text="pi_hpc=14"), combustion chamber total pressure recovery coefficient (id="1.5 ", text="sigma2="0.95"), high-pressure turbine pressure ratio (id="1.6", text="pi_hpt="2"), low-pressure turbine pressure ratio (id="1.7", text="pi_lpt="3.5"), internal nozzle total pressure recovery coefficient (id="1.8", text="sigma3="0.97") and external nozzle total pressure recovery coefficient (id="1.9", text="sigma4="0.97"), and these parameters are designed into 9 hierarchical requirements. The relationship between hierarchical requirements and top-level requirements is a "derivative" relationship. Each requirement has a unique identifier (id) and a corresponding text description (text). These requirements together constitute the detailed requirements in the design process of aircraft engine gas path system.

[0032] After that, the stakeholders of the aircraft engine gas path system are sorted out in detail, and the requirements are covered through use cases to generate a use case diagram such as Figure 3 As shown. Based on the requirements, the boundary of the aircraft engine is defined as the entire engine system. Subsequently, the scope of the use cases within the engine is considered, including takeoff, climb, cruise, landing, and providing thrust. Finally, based on the typical usage of aircraft engines, various participants that interact with the aircraft engine outside the system boundary are identified, including aircraft systems, atmospheric environment, and pilots, where all three participants are represented by "actor". The pilot manipulates the aircraft system, and there is an interaction relationship between the aircraft system and the five use cases of the engine. There is an interaction relationship between the atmospheric environment and the four use cases of takeoff, climb, cruise, and landing. All interaction relationships in the figure are represented by "communication path".

[0033] (1.2) Conduct behavioral analysis:

[0034] Determine the functions and logical relationships of the system and implement the definition and analysis of activities. First, convert the system requirements into several executable models and use activity diagrams (such as Figure 4 The process considered in the activity diagram is the startup process, and the research objects considered are the flight control system, auxiliary power unit, aircraft engine, atmospheric environment and fuel system. The transfer and convergence of the activity nodes in the diagram illustrate the functions and logical relationships of the system.

[0035] The entire startup process starts from the "start switch" node in the "flight control system". After that, the "starter belt" in the "auxiliary power unit" starts to work, driving the engine to rotate. At the same time, the "atmospheric environment" supplies air to the engine. Under the action of the starter belt, the speed of the "aircraft engine" gradually increases until it reaches the specified speed value. At this time, the "fuel system" starts to supply fuel. In the "aircraft engine", the combustion chamber mixes the supplied fuel and air. After that, the "flight control system" controls the ignition operation to ignite the mixed fuel and air in the combustion chamber. After the ignition is successful, the "aircraft engine" is officially started, completing the entire starting process. Through the orderly implementation of these steps, the aircraft engine is successfully started from a stationary state and starts to operate.

[0036] Then, using the sequence diagram (such as Figure 5 The figure (shown) describes the sequential interaction between the various components that make up an aircraft engine. External air enters the engine through the air inlet and is divided into two paths at the fan outlet. One path is the outer duct airflow, which directly discharges air through the outer duct nozzle to generate thrust. The other path is the inner duct airflow. After the inner duct airflow is compressed by low-pressure air and high-pressure air, it is mixed with fuel in the main combustion chamber and burned to generate heat. The expanded high-pressure combustion gas works on the high-pressure turbine and the low-pressure turbine, extracting torque to the high-pressure shaft and the low-pressure shaft, providing kinetic energy for the low-pressure compressor and the high-pressure compressor, and finally accelerating the air discharge to generate thrust.

[0037] (1.3) Design structure composition:

[0038] To model the static structure of the system, first, the basic composition of the aircraft engine components is shown through the block definition diagram. The aircraft engine consists of an air inlet, a fan, an outer duct, a low-pressure compressor, a high-pressure compressor, a low-pressure rotor, a main combustion chamber, a high-pressure rotor, a low-pressure turbine, a high-pressure turbine, an inner nozzle, and an outer nozzle. Figure 6A component-level model of an aircraft engine is built, with each component represented as a "block". The three main types of relationships between blocks are generalization, association, and dependency. Here, the relationship between each sub-module and the aircraft engine module is "part association".

[0039] Afterwards, through the internal block diagram ( Figure 7 ) is used to design the component interfaces of the system, taking into account the gas path interface, mechanical interface and fuel interface. The fuel system of the aircraft engine ensures that the fuel is delivered to the main combustion chamber at an appropriate flow rate through the fuel interface. The high-pressure rotor shaft connects the high-pressure compressor and the high-pressure turbine through a mechanical interface, and the low-pressure rotor shaft connects the low-pressure compressor and the low-pressure turbine through a mechanical interface. The gas path system inside the engine is equipped with multiple channels to guide the air flow ("channel 1" to "channel 10"). The outside air enters the air intake through gas path channel 1 and enters the fan through gas path channel 2. After the fan, the air is divided into two paths. One path enters the outer duct through gas path channel 9, and then is directly discharged from the outer duct nozzle through gas path channel 10 to generate thrust; the other path enters the low-pressure compressor through gas path channel 3, enters the high-pressure compressor through gas path channel 4, enters the main combustion chamber through gas path channel 5, enters the high-pressure turbine through gas path channel 6, enters the low-pressure turbine through gas path channel 7, and is discharged from the inner nozzle through gas path channel 8 to generate thrust. These interface designs ensure the coordinated operation of airflow, fuel and mechanical components inside the engine, which together support the efficient operation of the engine.

[0040] (1.4) Establish parameter constraint relationship:

[0041] In the air path system of aircraft engines, there are complex constraints. The constraint relationships between the properties of the system and components are represented by parameter diagrams. Fig. 9As shown in the lower left corner, the center of the diagram is a constraint block, which contains formulas that represent the pressure relationship between components. Surrounding this central block are the eight component blocks specified in the requirements, including the inlet, low-pressure compressor, high-pressure compressor, main combustion chamber, high-pressure turbine, low-pressure turbine, inner nozzle and outer nozzle, each of which is configured with corresponding constraint attributes. Through the "equal" connection in the diagram, the design parameters of each component block are linked to the constraint block, thereby establishing a constraint relationship between the component attributes. Specifically, the intake total pressure recovery coefficient of the inlet duct is connected to sigma1 of the constraint block, the pressure ratio of the low-pressure compressor is connected to pi_lpc of the constraint block, the pressure ratio of the high-pressure compressor is connected to pi_hpc of the constraint block, the combustion chamber total pressure recovery coefficient of the main combustion chamber is connected to sigma2 of the constraint block, the pressure ratio of the high-pressure turbine is connected to pi_hpt of the constraint block, the pressure ratio of the low-pressure turbine is connected to pi_lpt of the constraint block, the endotube nozzle total pressure recovery coefficient is connected to sigma3 of the constraint block, and the endotube nozzle total pressure recovery coefficient and bypass pressure ratio are connected to sigma4 and pi_bypass of the constraint block respectively.

[0042] Step 2: Infer the mechanism model of the air path system of the air-driving engine.

[0043] For the convenience of description, the gas system is simplified considering the characteristics of the engine with dual rotors and separate exhaust. Figure 8 As shown, the meaning of the work station number of each cross-sectional position of the aircraft engine in the figure is explained in detail through Table 1.

[0044] Table 1 Meaning of the workstation numbers of various parts of the engine

[0045]

[0046] The properties and behavior analysis of the 11 components, including atmosphere, air inlet, fan, bypass, low-pressure compressor, high-pressure compressor, combustion chamber, high-pressure turbine, low-pressure turbine, inner duct nozzle, and outer duct nozzle, are carried out in sequence. Among them: property analysis refers to analyzing the function of the component, describing the operating characteristics, component parameters, and intermediate variables of each component. Behavior analysis refers to using specific thermodynamic equations to describe the dynamic behavior of the component.

[0047] The formulas for atmospheric environment static temperature T0 and atmospheric environment static pressure P0 are as follows, where H is the flight altitude.

[0048]

[0049] The formulas for the total temperature T1 and the total pressure P1 at the inlet of the air duct are as follows, where k is the air isentropic ratio and Ma is the Mach number.

[0050]

[0051] The formulas for the total temperature T2 at the inlet duct outlet and the total pressure P2 at the inlet duct outlet are as follows, where σ1 is the recovery coefficient of the total inlet pressure.

[0052] T2=T1

[0053] P2=P1σ1

[0054] After the air flows through the fan, it is divided into two paths: one is the inner flow, which continues to be compressed by the compressor, and the other is the outer flow, which is directly discharged into the atmosphere through the outer duct to generate thrust. The ratio of the air flow through the outer duct and the inner duct is called the bypass ratio m. The fan inlet mass flow G2 and the fan outer duct outlet mass flow G 13 and the mass flow rate G at the fan outlet 21 The formula is as follows, where K m is a proportionality coefficient related to the gas properties, F2 is the cross-sectional area of ​​the flow channel, and q(λ) is the relative flux density.

[0055]

[0056]

[0057] Bypass outlet total temperature T 13 and bypass outlet total pressure P 13 The formula is as follows, where π bypass is the bypass pressure ratio.

[0058] T 13 =T2

[0059] P 13 =P2π bypass

[0060] Low pressure compressor work L 25 , low pressure compressor outlet total temperature T 25 , low pressure compressor outlet total pressure P 25 The formula is as follows, where C pa is the constant pressure specific heat of air, π lpc is the low pressure compressor pressure ratio, η lpc is the isentropic efficiency of the low-pressure compressor.

[0061] L 25 =C pa T2(π lpc (k-1) / k -1) / η lpc

[0062] T 25 =T2+L 25 / C pa

[0063] P 25=P2π lpc

[0064] The formulas for high-pressure compressor work L3, high-pressure compressor outlet total temperature T3, and high-pressure compressor outlet total pressure P3 are as follows, where π hpc is the high pressure compressor pressure ratio, η hpc is the isentropic efficiency of the high pressure compressor.

[0065] L3=C pa T 25 (π hpc (k-1) / k -1) / η hpc

[0066] T3=T 25 +L3 / C pa

[0067] P3=P 25 π hpc

[0068] The formulas for the total temperature T4 and total pressure P4 at the combustion chamber outlet are as follows: f is the fuel flow rate, G3 is the mass flow rate at the high pressure compressor outlet, H f is the calorific value of the fuel, η2 is the combustion efficiency, C pcg is the constant-pressure specific heat of aviation kerosene, and σ2 is the total pressure recovery coefficient of the combustion chamber.

[0069]

[0070] P4=P3σ2

[0071] High pressure turbine work L 42 , high pressure turbine outlet total temperature T 42 , High-pressure turbine outlet total pressure P 42 The formula is as follows, where π hpt is the high-pressure turbine pressure ratio, η hpt is the isentropic efficiency of the high-pressure turbine.

[0072]

[0073] T 42 =T4-L 42 / C pcg

[0074] P 42 =P4 / π hpt

[0075] The formulas for low-pressure turbine work L5, low-pressure turbine outlet total temperature T5, and low-pressure turbine outlet total pressure P5 are as follows, where π lpt is the low-pressure turbine pressure ratio, η lptis the isentropic efficiency of the low-pressure turbine.

[0076]

[0077] T5=T 42 -L5 / C pcg

[0078] P5=P 42 / π lpt

[0079] The formulas for the total temperature T9 and total pressure P9 at the outlet of the internal nozzle are as follows, where σ3 is the total pressure recovery coefficient of the internal nozzle.

[0080] T9=T5

[0081] P9=P5σ3

[0082] Available nozzle pressure ratio π nc1 and critical pressure ratio π cr The formula is:

[0083]

[0084] When the available nozzle pressure ratio is less than the critical pressure ratio, the static pressure P s9 , air velocity V9 and Mach number Ma are:

[0085] P s9 =P0

[0086]

[0087] Ma=π nc1 (k-1) / k

[0088] When the available nozzle pressure ratio is greater than or equal to the critical pressure ratio, the static pressure P s9 , air velocity V9 and Mach number Ma are:

[0089] P s9 =P9 / π cr

[0090]

[0091] Ma=1

[0092] The formula for the outlet flow rate m9 of the inner tail nozzle is:

[0093]

[0094] Among them, q(Ma) is the relative flux density, and A9 is the cross-sectional area of ​​the end nozzle flow path.

[0095]

[0096] The thrust formula generated by the inner duct is as follows, where V0 is the atmospheric air velocity.

[0097] F9=m9(V9-V0)+A9(P s9 -P0)

[0098] Total temperature at the tail nozzle outlet of the duct T 19 , outlet total pressure P 19 The formula is as follows, where σ4 is the total pressure recovery coefficient of the outer nozzle.

[0099] T 19 =T 13

[0100] P 19 =P 13 σ4

[0101] Available nozzle pressure ratio π nc2 and critical pressure ratio π cr The formula is:

[0102]

[0103] When the available nozzle pressure ratio is less than the critical pressure ratio, the static pressure P s19 , air velocity V 19 And the Mach number Ma is:

[0104] P s19 =P0

[0105]

[0106] Ma=π nc2 (k-1) / k

[0107] When the available nozzle pressure ratio is greater than or equal to the critical pressure ratio, the static pressure P s19 , air velocity V 19 And the Mach number Ma is:

[0108] P s19 =P 19 / π cr

[0109]

[0110] Ma=1

[0111] External duct tail nozzle outlet flow rate m 19 The formula is:

[0112]

[0113] Among them, A 19 is the cross-sectional area of ​​the flow path of the external duct tail nozzle.

[0114] The thrust formula of the outer duct is as follows:

[0115] F 19 =m 19 (V 19 -V0)+A 19 (P s19 -P0)

[0116] The total engine thrust F formula is:

[0117] F=F9+F 19

[0118] Step 3: Performance simulation and verification of aircraft engine gas path system.

[0119] Based on the mechanism model in step 2, the digital prototype tool MWORKS is used for text modeling and programming to build a digital prototype model demonstration diagram of the system. Fig. 9 As shown in the figure. The model integrates and simulates the various components of the aircraft engine gas path system. The model consists of the core engine components such as the air inlet, fan, low-pressure compressor, high-pressure compressor, combustion chamber, high-pressure turbine, low-pressure turbine, internal tail nozzle and external tail nozzle. The main functions of the digital prototype include:

[0120] (1) Performance prediction: Calculate key performance indicators such as engine thrust and fuel consumption based on the input design parameters and operating conditions.

[0121] (2) Virtual testing and verification: Simulate the operation of the aircraft engine gas system in a virtual environment to verify whether the design meets the predetermined performance requirements and standards.

[0122] In order to ensure the accuracy of the model, the control and environmental variables are extracted from the public model data as the input data of the model. The input data covers the entire process of the aircraft taking off, ascending, high-altitude stable flight, and landing. Fig.10 Describes in detail the altitude H, Mach number Ma, fuel flow G fData changes during the entire flight process. After inputting the operating parameters of the aircraft engine gas path system, the digital prototype model can calculate the engine performance through the internal performance simulation model and display the operating results of the total thrust and fuel consumption rate. Here, the rationality of the model is analyzed by taking the engine thrust as an example. The curve of thrust F changing with time shows the change of the aircraft's thrust demand in different flight stages. During the takeoff and climb phases, the engine needs to provide sufficient thrust to overcome gravity. As the altitude increases, the thrust gradually decreases, but it still remains at a high level to maintain the flight speed. Entering the cruise phase, the thrust will remain at a relatively stable level. As the subsequent Mach number decreases, the thrust will increase to maintain the performance and stability of the aircraft. The Ft curve proves that the designed digital prototype model can provide the required thrust in different flight phases, realize accurate performance prediction, and effectively complete the closed-loop verification of the requirements, ensuring the forward design of the overall plan.

Claims

1. A method for generating a digital prototype of an aeroengine air path system based on a system modeling language, characterized in that: The steps include: Step 1: Modeling the requirements and architecture of the aircraft engine gas path system; Based on SysML, the aircraft engine gas path system is modeled from four aspects: requirements, behavior, structure, and parameters. Starting from the top-level requirements, the design purpose of the system is clarified, and then the requirements are refined into functional activity behavior analysis, static system structure composition, and system parameter constraint relationship to complete the architecture modeling of the overall solution. Step 2: Reasoning the mechanism model of the air path system of the air-driving engine; According to the workflow of the aircraft engine air path system obtained in step 1, the attributes and behavior analysis of the atmosphere, air inlet, fan, bypass, low-pressure compressor, high-pressure compressor, combustion chamber, high-pressure turbine, low-pressure turbine, inner duct nozzle, and outer duct nozzle are performed in turn to build a mechanism model of the aircraft engine; wherein, attribute analysis refers to analyzing the function of the component, describing the operating characteristics, component parameters, and intermediate variables of each component; behavior analysis refers to using the corresponding thermodynamic equations to describe the dynamic behavior of the component; Step 3: Performance simulation and verification of aircraft engine gas path system; Based on the mechanism model obtained in step 2, a system performance model is built using a digital prototype tool for simulation testing: first, the input and output interfaces of each component are defined and instantiated through text modeling; then, the component models are encapsulated in sequence and connected according to the physical topology of the components to form a digital prototype model of the system; finally, the simulation parameters and simulation variables are set, the system performance indicators are simulated and calculated throughout the entire process, and the calculation results are intuitively displayed through the user interface, thereby verifying the top-level requirements of the system and completing the forward design process from concept to implementation.

2. The method for generating a digital prototype of an aeroengine air path system based on a system modeling language according to claim 1, characterized in that: The step 1 is specifically as follows: (1.1) Conduct demand analysis: Analyze the design goals of the system, clarify the design purpose of the system, complete the capture of top-level task requirements and subdivide the requirements, and generate a requirement graph; the top-level requirement for aircraft engine development is to provide the required thrust according to different flight conditions and tasks, id="1", text="provide the required thrust according to different flight conditions and flight missions of the aircraft"; in this process, it is necessary to introduce parameters to calculate the thrust, including the total intake pressure recovery coefficient id="1.1", text="sigma1=0.97", low-pressure compressor pressure ratio id="1.2", text="pi_lpc=3", bypass pressure ratio id="1.3", text="pi_by pass=2”, high pressure compressor pressure ratio id="1.4", text="pi_hpc=14”, combustion chamber total pressure recovery coefficient id="1.5", text="sigma2=0.95”, high pressure turbine pressure ratio id="1.6", text="pi_hpt=2”, low pressure turbine pressure ratio id="1.7", text="pi_lpt=3.5”, internal nozzle total pressure recovery coefficient (d="1.8", text="sigma3=0.97” and external nozzle total pressure recovery coefficient id="1.9", text="sigma4=0.97”, these parameters are designed into 9 hierarchical requirements; The relationship between the hierarchical requirements and the top-level requirements is a "derivative" relationship; each requirement has a unique identifier id and a corresponding text description text, and these requirements together constitute the detailed requirements in the design process of the aircraft engine gas path system; After that, the stakeholders of the aircraft engine gas path system were sorted out, the requirements were covered through use cases, and a use case diagram was generated; Based on the requirements, the boundary of the aircraft engine is defined as the entire engine system; Then, consider the scope of the internal use cases of the engine, including takeoff, climb, cruise, landing, and providing thrust; finally, based on the typical use of aircraft engines, identify various participants that interact with aircraft engines outside the system boundary, including aircraft systems, atmospheric environment, and pilots. Among them, all three participants are represented by "actor"; the pilot operates the aircraft system, and the aircraft system has an interaction relationship with the five use cases of the engine. The atmospheric environment has an interaction relationship with the four use cases of takeoff, climb, cruise, and landing. All the interaction relationships in the figure are represented by "communication path"; (1.2) Conduct behavioral analysis: Determine the functions and logical relationships of the system, and implement the definition and analysis of activities; first, convert system requirements into several executable models, use activity diagrams to describe the system's behavior, and implement use case analysis; The process considered in the activity diagram is the startup process, and the research objects considered are the flight control system, auxiliary power unit, aircraft engine, atmospheric environment and fuel system; the transfer and convergence of activity nodes in the diagram illustrate the functions and logical relationships of the system; The entire starting process starts from the "start switch" node in the "flight control system"; then, the "starter belt" in the "auxiliary power unit" starts to work, driving the engine to rotate; at the same time, the "atmospheric environment" supplies air to the engine; under the action of the starter belt, the speed of the "aircraft engine" gradually increases until it reaches the specified speed value; at this time, the "fuel system" starts to supply fuel; in the "aircraft engine", the combustion chamber mixes the supplied fuel and air; then, the "flight control system" controls the ignition operation to ignite the mixed fuel and air in the combustion chamber; after successful ignition, the "aircraft engine" officially starts, completing the entire starting process; through the orderly implementation of these steps, the aircraft engine successfully starts from a stationary state and starts to operate; After that, a sequence diagram is used to describe the sequential interaction between the various components that make up an aircraft engine. External air enters the engine through the air inlet and is divided into two paths at the fan outlet. One path is the outer duct airflow, which directly discharges air through the outer duct nozzle to generate thrust. The other path is the inner duct airflow. After the inner duct airflow is compressed by low-pressure air and high-pressure air, it is mixed with fuel in the main combustion chamber and burned to generate heat. The expanded high-pressure combustion gas works on the high-pressure turbine and the low-pressure turbine, extracting torque to the high-pressure shaft and the low-pressure shaft, providing kinetic energy for the low-pressure compressor and the high-pressure compressor, and finally accelerating the air discharge to generate thrust. (1.3) Design structure composition: Model the static structure of the system. First, use the block definition diagram to show the basic composition of the aircraft engine components. The aircraft engine consists of an air inlet, a fan, a duct, a low-pressure compressor, a high-pressure compressor, a low-pressure rotor, a main combustion chamber, a high-pressure rotor, a low-pressure turbine, a high-pressure turbine, an inner nozzle and an outer nozzle. Build a component-level model of the aircraft engine, with each component represented as a block. The relationship between each sub-module and the aircraft engine module is "part association". Afterwards, the component interface design of the system was carried out through the internal block diagram, taking into account the gas path interface, mechanical interface and fuel interface; the fuel system of the aircraft engine ensures that the fuel is delivered to the main combustion chamber at an appropriate flow rate through the fuel interface; the high-pressure rotor shaft connects the high-pressure compressor and the high-pressure turbine through the mechanical interface, and the low-pressure rotor shaft connects the low-pressure compressor and the low-pressure turbine through the mechanical interface; the gas path system inside the engine is equipped with multiple channels to guide the air flow, including "channel 1" to "channel 10"; the outside air enters the air inlet through gas path channel 1 and enters the fan through gas path channel 2; the air after the fan is divided into two paths, one enters the outer duct through gas path channel 9, and then is directly discharged from the outer duct nozzle through gas path channel 10 to generate thrust; the other enters the low-pressure compressor through gas path channel 3, enters the high-pressure compressor through gas path channel 4, enters the main combustion chamber through gas path channel 5, enters the high-pressure turbine through gas path channel 6, enters the low-pressure turbine through gas path channel 7, and is discharged from the inner nozzle through gas path channel 8 to generate thrust; (1.4) Establish parameter constraint relationship: The constraint relationship between the properties of the system and components is represented by a parameter diagram. The center of the parameter diagram is a constraint block, which contains the formula representing the pressure relationship between the components. Surrounding the constraint block are the eight component blocks specified in the requirements, including the inlet, low-pressure compressor, high-pressure compressor, main combustion chamber, high-pressure turbine, low-pressure turbine, inner nozzle and outer nozzle. Each component block is configured with corresponding constraint attributes. Through the "equal" connection, the design parameters of each component block are linked to the constraint block, thereby establishing a constraint relationship between the component attributes. Specifically, the intake total pressure recovery coefficient of the inlet is related to the constraint. The pressure ratio of the low-pressure compressor is connected to pi_lpc of the constraint block, the pressure ratio of the high-pressure compressor is connected to pi_hpc of the constraint block, the total pressure recovery coefficient of the combustion chamber of the main combustion chamber is connected to sigma2 of the constraint block, the pressure ratio of the high-pressure turbine is connected to pi_hpt of the constraint block, the pressure ratio of the low-pressure turbine is connected to pi_lpt of the constraint block, the total pressure recovery coefficient of the endotube nozzle is connected to sigma3 of the constraint block, and the total pressure recovery coefficient of the endotube nozzle and the bypass pressure ratio are connected to sigma4 and pi_bypass of the constraint block respectively.

3. The method for generating a digital prototype of an aeroengine air path system based on a system modeling language according to claim 1, characterized in that: The step 2 is specifically as follows: The formulas for atmospheric environment static temperature T0 and atmospheric environment static pressure P0 are as follows, where H is the flight altitude; The formulas for the total temperature T1 and total pressure P1 at the inlet of the air duct are as follows, where k is the air isentropic ratio and Ma is the Mach number; The formulas for the total temperature T2 and total pressure P2 at the inlet duct outlet are as follows, where σ1 is the recovery coefficient of the inlet total pressure; T2=T1 P2=P1σ1 After the air flows through the fan, it is divided into two paths: one is the inner flow, which continues to be compressed by the compressor, and the other is the outer flow, which is directly discharged into the atmosphere through the outer duct to generate thrust; the ratio of the air flow through the outer duct and the inner duct is called the bypass ratio m; the fan inlet mass flow G2, the fan outer duct outlet mass flow G 13 and the mass flow rate G at the fan outlet 21 The formula is as follows, where K m is a proportionality coefficient related to the gas properties, F2 is the cross-sectional area of ​​the flow channel, and q(λ) is the relative flux density; Bypass outlet total temperature T 13 and bypass outlet total pressure P 13 The formula is as follows, where π bypass is the bypass total pressure recovery coefficient; T 13 =T2 P 13 =P2π bypass Low pressure compressor work L 25 , low pressure compressor outlet total temperature T 25 , low pressure compressor outlet total pressure P 25 The formula is as follows, where C pa is the constant pressure specific heat of air, π lpc is the low pressure compressor pressure ratio, η lpc is the isentropic efficiency of the low-pressure compressor; L 25 =C pa T2(π lpc (k-1) / k -1) / h lpc T 25 =T2+L 25 / C pa P 25 =P2π lpc The formulas for high-pressure compressor work L3, high-pressure compressor outlet total temperature T3, and high-pressure compressor outlet total pressure P3 are as follows, where π hpc is the high pressure compressor pressure ratio, η hpc is the isentropic efficiency of the high pressure compressor; L3=C pa T 25 (p hpc (k-1) / k -1) / h hpc <h2 style=";text-align:left;direction:ltr">T3=T<h2 style=";text-align:left;direction:ltr"> 25 <h2 style=";text-align:left;direction:ltr"> +L3 / C<h2 style=";text-align:left;direction:ltr"> pa P3=P 25 p hpc The formulas for the total temperature T4 and total pressure P4 at the combustion chamber outlet are as follows: f is the fuel flow rate, G3 is the mass flow rate at the high pressure compressor outlet, H f is the calorific value of the fuel, η2 is the combustion efficiency, C pcg is the constant pressure specific heat of aviation kerosene, σ2 is the total pressure recovery coefficient of the combustion chamber; P4=P3σ2 High pressure turbine work L 42 , high pressure turbine outlet total temperature T 42 , High-pressure turbine outlet total pressure P 42 The formula is as follows, where π hpt is the high-pressure turbine pressure ratio, η hpt is the isentropic efficiency of the high-pressure turbine; T 42 =T4-L 42 / C pcg P 42 =P4 / π hpt The formulas for low-pressure turbine work L5, low-pressure turbine outlet total temperature T5, and low-pressure turbine outlet total pressure P5 are as follows, where π lpt is the low-pressure turbine pressure ratio, η lpt is the isentropic efficiency of the low-pressure turbine; T5=T 42 -L5 / C pcg P5=P 42 / π lpt The formulas for the total temperature T9 and total pressure P9 at the outlet of the internal nozzle are as follows: σ3 is the total pressure recovery coefficient of the internal nozzle; T9=T5 P9=P5σ3 Available nozzle pressure ratio π nc1 and critical pressure ratio π cr The formula is: When the available nozzle pressure ratio is less than the critical pressure ratio, the static pressure P s9 , air velocity V9 and Mach number Ma are: P s9 =P0 Ma=π nc1 (k-1) / k When the available nozzle pressure ratio is greater than or equal to the critical pressure ratio, the static pressure P s9 , air velocity V9 and Mach number Ma are: P s9 =P9 / π cr Ma=1 The formula for the outlet flow rate m9 of the inner tail nozzle is: Among them, q(Ma) is the relative flux density, A9 is the cross-sectional area of ​​the inner nozzle flow path; The thrust formula generated by the inner channel is as follows, where V0 is the atmospheric air velocity; <h2 style=";text-align:left;direction:ltr">F9 = m9(V9-V0) + A9(P<h2 style=";text-align:left;direction:ltr"> s9 <h2 style=";text-align:left;direction:ltr"> -P0) Total temperature at the tail nozzle outlet of the duct T 19 , outlet total pressure P 19 The formula is as follows, where σ4 is the total pressure recovery coefficient of the culvert nozzle; T 19 =T 13 P 19 =P 13 σ4 Available nozzle pressure ratio π nc2 and critical pressure ratio π cr The formula is: When the available nozzle pressure ratio is less than the critical pressure ratio, the static pressure P s19 , air velocity V 19 And the Mach number Ma is: P s19 =P0 Ma=π nc2 (k-1) / k When the available nozzle pressure ratio is greater than or equal to the critical pressure ratio, the static pressure P s19 , air velocity V 19 And the Mach number Ma is: P s19 =P 19 / p cr Ma=1 External duct tail nozzle outlet flow rate m 19 The formula is: Among them, A 19 is the cross-sectional area of ​​the flow path of the outer duct tail nozzle; The thrust formula of the outer duct is as follows: F 19 =m 19 (IN 19 -V0)+A 19 (P s19 -P0) The total engine thrust F formula is: F=F9+F 19 。

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