Multi-duct casing complex load transient characteristic analysis method

Abstract

The invention discloses a multi-duct casing complex load transient characteristic analysis method, and belongs to the technical field of aero-engines. According to the method, the structure and the working process of the multi-duct variable-cycle engine are analyzed, and the multi-duct variable-cycle engine is divided into three turbofan modes and one turbojet mode; based on the aerodynamic thermodynamic configuration model, MATLAB software is used for part-level modeling, and part-level modeling data is formed; establishing a common working equation set by using modeling data, and performing iterative solution by adopting a Broyden quasi-Newton method to obtain a solution of the equation set; calculating the limit aerodynamic load of the casing system under the typical working condition; in combination with the flight envelope of the engine, 12 typical heights and Mach number states are selected, a limit working point is solved, and transient load characteristics are analyzed. According to the method, the transient response characteristics of the multi-duct casing system under the complex load condition can be accurately captured, the analysis precision and efficiency are improved, a theoretical basis is provided for design optimization and reliability evaluation of the casing system, and the method has important engineering value.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engines, and in particular to a method for analyzing transient characteristics of complex loads on a multi-duct casing. Background Art

[0002] Aircraft engines are the core components of aircraft, and their performance directly affects the thrust, fuel efficiency and safety of aircraft. Traditional turbojet engines and turbofan engines have certain advantages in their respective fields, but they show significant differences under different flight conditions: turbojet engines are more efficient in high-altitude supersonic flight, but have poor fuel efficiency in low-altitude subsonic flight; turbofan engines perform well in low-altitude subsonic flight, but their thrust efficiency drops significantly in high-altitude supersonic flight.

[0003] Multi-bypass variable cycle engines achieve the goal of maintaining high thrust output and reducing fuel consumption within a wide flight envelope by adjusting aerodynamic and thermal parameters, becoming an important direction of modern aviation engine technology. However, its complex airflow organization and mechanical structure will produce significant transient loads and thermal shocks during the switching of working modes (such as between turbofan mode and turbojet mode). This dynamic characteristic not only challenges the reliability and life of the engine casing system, but also puts higher requirements on the real-time and accuracy of the analysis method.

[0004] Most existing technologies focus on the study of aerodynamic characteristics under a single steady-state condition, which makes it difficult to effectively capture the complex transient response in mode switching, and lack efficient mathematical modeling and solution methods, and cannot meet the real-time analysis requirements under cross-air and cross-speed conditions. Therefore, developing an efficient method that can comprehensively analyze the dynamic characteristics of multi-bypass variable cycle engines has become a key research direction to improve their reliability and performance. Summary of the invention

[0005] The purpose of the present invention is to provide a method for analyzing transient characteristics of complex loads on a multi-duct variable cycle engine, thereby solving the problems of insufficient analysis accuracy, difficulty in capturing dynamic characteristics and low computational efficiency in the prior art by modeling and analyzing the structure, working process and dynamic characteristics of the engine casing system.

[0006] In order to achieve the above object, the present invention is implemented by the following technical solutions:

[0007] A method for analyzing complex load transient characteristics of a multi-duct casing, characterized by comprising the following steps:

[0008] Analyze the structure and working process of the multi-bypass variable cycle engine, determine the working mode, divide the working process into three turbofan modes and one turbojet mode, and form the corresponding working process analysis results;

[0009] Based on the work process analysis results, MATLAB software is used to perform component-level modeling of the multi-bypass variable cycle engine based on the aerodynamic thermodynamic configuration model to form component-level modeling data;

[0010] Using component-level modeling data, a common working equation group of a multi-bypass variable cycle engine is established, and a Broyden quasi-Newton method is used to iteratively solve the solution of the common working equation group to obtain a solution of the equation group;

[0011] Based on the solution of the equations, the ultimate aerodynamic load of the multi-duct casing system during operation is calculated;

[0012] By using the extreme aerodynamic loads and combining the flight envelope of the multi-bypass variable cycle engine, 12 typical altitudes and Mach number states are selected to solve the extreme operating points in turbofan mode and turbojet mode, and the transient load characteristics of the multi-bypass casing system are analyzed.

[0013] As a preferred solution of the present invention, the working modes include a subcruise turbofan mode, a supercruise turbofan mode, a high-speed turbofan mode and a high-speed turbojet mode, wherein:

[0014] The sub-cruise turbofan mode is suitable for flight speeds of 0-1.3 Ma and altitudes of 0-11000 m. The combustion chamber only opens the main combustion chamber and the mode selection valve is closed.

[0015] The supercruise turbofan mode is suitable for flight speeds of 1.3–3.5 Ma and altitudes of 11,000–23,000 m, with the main combustion chamber and the dual variable combustion chamber open, and the mode selection valve closed;

[0016] The high-speed turbofan mode is suitable for flight speeds of 3.5–4.0 Ma and altitudes of 23,000–27,000 m. The combustion chambers open the main combustion chamber, dual variable combustion chamber, and afterburner, and the mode selection valve is closed;

[0017] The high-speed turbojet mode is suitable for flight speeds of 3.7-5.0 Ma and altitudes of 24,000-30,000 m. The combustion chamber opens the dual-variable combustion chamber and the afterburner, and the mode selection valve is opened.

[0018] As a preferred solution of the present invention, the multi-bypass variable cycle engine includes a fan, a high-pressure compressor, a combustion chamber and a turbine, and component-level modeling is performed for these performance components, wherein:

[0019] The multi-duct variable cycle engine includes three fans, namely a front fan, a rear fan and a core fan. In turbofan mode, all three fans are involved in the work; in turbojet mode, the front fan and the rear fan are in a windmill state;

[0020] The multi-bypass variable cycle engine is a three-rotor structure, including a high-pressure turbine, an intermediate-pressure turbine and a low-pressure turbine, wherein the intermediate-pressure turbine guide vanes and the low-pressure turbine guide vanes are geometrically adjustable components, and the turbine performance characteristics are adjusted by adjusting the guide vane angle.

[0021] As a preferred embodiment of the present invention, the geometrically adjustable characteristics of the intermediate-pressure turbine and low-pressure turbine guide vanes are used to adjust turbine performance parameters by the following method:

[0022] Based on the condition that different guide vane angles α are zero, the reference pressure ratio, reference efficiency and reference flow rate are obtained by combining turbine experimental data and two-dimensional interpolation method;

[0023] When the guide vane angle α is non-zero, the pressure ratio, efficiency and flow rate of the turbine are calculated according to the following formulas:

[0024] The turbine pressure ratio π is:

[0025] π=(π map -1)(1+α+k π )+1;

[0026] Where: π map is the reference pressure ratio; k π is the influence coefficient of guide vane angle on pressure ratio;

[0027] The turbine efficiency η is:

[0028]

[0029] The turbine flow rate W is:

[0030] W=W map (1+α+k w );

[0031] Where: W map is the reference flow rate; k w is the influence coefficient of guide vane angle on flow rate;

[0032] The meshgrid and interp2 functions of MATLAB are used to fit the turbine characteristic curve, and the total pressure, total temperature and turbine power at the turbine outlet are obtained by combining aerodynamic thermodynamic calculations.

[0033] As a preferred solution of the present invention, the common working equation group includes:

[0034] Based on the structure of a multi-bypass variable cycle engine, a mathematical model is established to describe the relationship between the airflow and energy transfer inside the engine. The mathematical model includes:

[0035] The flow balance equation is used to describe the conservation relationship of airflow mass between engine components;

[0036] The static pressure balance equation is used to describe the continuity of static pressure between engine components;

[0037] The power balance equation, which describes the energy transfer relationship between the high-pressure, intermediate-pressure, and low-pressure rotors;

[0038] The common working equation group consists of a set of 43 implicit nonlinear equations, wherein the change of each variable has a coupled effect on the overall residual of the equation group.

[0039] Furthermore, the common equations are the basis for the joint operation of multiple components. Based on the component-level modeling of the multi-bypass variable cycle engine mentioned above, the steady-state operation process of the multi-bypass variable cycle engine is further established, and the common equations of the performance model are finally solved. The core of the numerical calculation and solution of the component-level model is to solve the common working equations of the components. When the residual error of the common working equations is less than the specified convergence error after iterative solution, it can be considered that the aerodynamic thermodynamic parameters of each component calculated by the current model can represent the current real working conditions of the multi-bypass variable cycle engine.

[0040] First, seven initial variables are selected as auxiliary working lines, and the initial value matrix X is defined as:

[0041] X=[β1,β2,β3,β4,β5,β6,β7];

[0042] Where: β1, β2, β3 represent the auxiliary characteristics of the high-pressure turbine, the intermediate-pressure turbine and the low-pressure turbine respectively.

[0043] line; β4 represents the auxiliary lines of the characteristic diagram of the second and third ducts; β5 and β6 represent the front mixing

[0044] β7 represents the auxiliary lines of the characteristic diagram of the tail nozzle.

[0045] Based on the initial variables, seven sets of equations describing the engine airflow transfer and energy conservation were established, including flow balance and static pressure balance equations:

[0046] Flow balance equation:

[0047] The flow balance equation requires that the airflow between engine components maintains mass conservation, and its residual matrix E is expressed as:

[0048] E=[∈1,∈2,∈3,∈4,∈5,∈6,∈7];

[0049] Where: i represents the residual of each equilibrium equation;

[0050] The specific formula for flow balance is as follows:

[0051] High pressure turbine flow balance:

[0052] W HPT,in -W HPT,out =ε1;

[0053] Where: W HPT,in W represents the inlet air mass flow rate of the high-pressure turbine; HPT,out represents the outlet air mass flow rate of the high-pressure turbine; ε1 is the residual value of the high-pressure turbine flow balance;

[0054] Intermediate pressure turbine flow balance:

[0055] W MPT,in -W MPT,out =ε2;

[0056] Where: W MPT,in W represents the inlet air mass flow rate of the intermediate pressure turbine; MPT,out represents the outlet air mass flow rate of the intermediate pressure turbine; ε2 is the residual value of the intermediate pressure turbine flow balance;

[0057] Low pressure turbine flow balance:

[0058] W LPT,in -W LPT,out =ε3;

[0059] Where: W LPT,in W represents the inlet air mass flow rate of the low-pressure turbine; LPT,out represents the outlet air mass flow rate of the low-pressure turbine; ε3 is the residual value of the low-pressure turbine flow balance;

[0060] Bypass static pressure balance:

[0061] p s,second -p s,third =ε4;

[0062] Where: p s,second Indicates the static pressure value of the second section of the outer duct; p s,third It represents the static pressure value of the third section of the duct; ε4 is the residual value of the static pressure balance of the duct;

[0063] Static pressure balance of the front mixing chamber (static pressure at the outlet of the intermediate pressure turbine and the outlet of the core fan):

[0064] p s,MPT,out -p s,core,out =ε5;

[0065] Where: p s,MPT,out Indicates the static pressure value at the outlet of the intermediate pressure turbine; p s,core,out It represents the static pressure value at the core fan outlet; ε5 is the residual value of the static pressure balance of the front mixing chamber;

[0066] Static pressure balance of the rear mixing chamber (static pressure at the outlet of the front mixing chamber and the outlet of the low-pressure turbine):

[0067] p s,Fmix,out -p s,LPT,out =ε6;

[0068] Where: p s,Fmix,out Indicates the static pressure value at the outlet of the front mixing chamber; p s,LPT,out Represents the static pressure value at the outlet of the low-pressure turbine; ε6 is the residual value of the static pressure balance of the rear mixing chamber;

[0069] Tail nozzle flow balance:

[0070] W Nozzle,in -W Nozzle,out =ε7;

[0071] Where: W Nozzle,in W represents the air mass flow rate at the tail nozzle inlet; Nozzle,out represents the mass flow rate of the airflow at the tail nozzle outlet; ε7 is the residual value of the tail nozzle flow balance;

[0072] In addition, in order to describe the energy transfer characteristics of the multi-pass variable cycle engine, it is also necessary to establish the rotor power balance equation. The details are as follows:

[0073] Rotor power balance equation:

[0074] The three rotors (high pressure rotor, intermediate pressure rotor and low pressure rotor) of the multi-pass variable cycle engine must satisfy the law of energy conservation, and their power balance equations are:

[0075] High pressure rotor power balance:

[0076]

[0077] Medium pressure rotor power balance:

[0078]

[0079] Low pressure rotor power balance:

[0080]

[0081] Where: N H ,N M ,N L are the speeds of the high-pressure, medium-pressure and low-pressure rotors respectively; η HPT ,η MPT ,η LPT are the mechanical efficiencies of the high-pressure, medium-pressure, and low-pressure rotors, respectively; P HPT ,P MPT ,P LPT are the power of high-pressure, medium-pressure and low-pressure turbines respectively; PHPC ,P core ,P front ,P behind are respectively the high pressure compressor power, core compressor power, front fan power and rear fan power; J H ,J M ,J L are the rotational inertia of the high pressure, medium pressure and low pressure rotors respectively;

[0082] The common working equations of multi-bypass variable cycle engines are essentially a set of implicit nonlinear equations. There are 43 equations that need to be solved. The change of each variable will affect the final residual result of the entire equation group, which places high demands on the convergence and real-time performance of solving implicit nonlinear equations. At present, there is no method to solve the variable solutions of implicit nonlinear equations by analytical methods. We can only solve these equations to obtain the numerical optimal solution under a certain error.

[0083] As a preferred solution of the present invention, the mathematical model is iteratively solved by Broyden quasi-Newton method, and the iterative calculation process includes the following steps:

[0084] Set the initial input variables and establish the computational framework for iterative solution;

[0085] Update the relationship between input variables and residuals, and gradually correct the variables;

[0086] The working state of the multi-bypass variable cycle engine is solved by iterative calculation until the result converges to the set error threshold.

[0087] As a preferred solution of the present invention, the iterative calculation process of the Broyden quasi-Newton method includes the following steps:

[0088] Initial state variable setting: Set the initial state variable vector X0, where X represents the key parameters describing the state of the multi-duct casing system, including: turbine pressure ratio, temperature, flow rate, and speed;

[0089] Calculation of initial residual matrix: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X) and calculate the initial residual matrix E0. The formula is as follows:

[0090] E = F(X);

[0091] Among them, E is the residual matrix, which represents the deviation between the current state parameters and the target parameters of the multi-duct casing system;

[0092] Jacobian matrix calculation: According to the aerodynamic thermodynamic model F(X), the initial Jacobian matrix J is calculated using the formula:

[0093]

[0094] Among them: E1, E2, ..., E7 are residual matrix components; X1, X2, ..., X7 are state variable components;

[0095] Iterative variable update: The initial state variables, initial residual matrix and Jacobian matrix are brought into the iterative variable update formula to correct the state variables. The calculation formula is as follows:

[0096]

[0097] Where: X i represents the state variable of the i-th iteration; J -1 is the inverse matrix of the Jacobian matrix; is the transpose of the residual matrix of the i-th iteration;

[0098] Residual matrix update: The updated state variables are input into the aerodynamic thermodynamic model again, and the update formula is:

[0099] E i+1 =F(X i+1 );

[0100] Where: E i+1 represents the residual matrix of the i+1th iteration;

[0101] Jacobian matrix correction: According to the state variable correction value and the residual matrix correction value, the Jacobian matrix is ​​corrected. The calculation formula is as follows:

[0102] Y=B i+1 -B i ;

[0103] Z=E i+1 -E i ;

[0104]

[0105] Among them: B i+1 is the Jacobian matrix approximation of the i+1th iteration; B i Jacobian matrix approximation for the current iteration;

[0106] Y is the change between the state variable correction values; Z is the change of the residual matrix; J i+1 is the modified Jacobian matrix; Y T is the transposed matrix of the change in the state variable correction value Y;

[0107] Repeat the iteration until the residual matrix E converges to the set threshold.

[0108] As a preferred embodiment of the present invention, the mathematical model further includes:

[0109] Determine the flight envelope based on the cross-airspace and cross-speed domain design requirements of the multi-bypass variable cycle engine;

[0110] Twelve extreme working points are selected from the flight envelope for simulation and solution, where:

[0111] The two working points are ground slow and ground takeoff states;

[0112] The six operating points are typical states in turbofan mode;

[0113] The four operating points are typical conditions in turbojet mode;

[0114] The limit operating point is simulated and solved to obtain the high-pressure rotor speed, the medium-pressure rotor speed and the low-pressure rotor speed;

[0115] The changes in the outlet temperatures of the dual-variable combustion chamber and afterburner combustion chamber during the mode switching of a multi-duct variable cycle engine are analyzed to determine the thermal shock damage characteristics of the multi-duct casing system caused by transient loads.

[0116] As a preferred solution of the present invention, the 12 extreme working points selected from the flight envelope include:

[0117] The six typical working points in turbofan mode correspond to the following working conditions:

[0118] Subsonic cruise mode;

[0119] Supersonic cruise mode;

[0120] High speed afterburner mode;

[0121] The four typical flight states in turbojet mode correspond to the following operating points:

[0122] D7(24,3.7), D8(27,4.0), D9(28,4.3) and D 10 (30,5.0).

[0123] As a preferred embodiment of the present invention, the method further comprises the following steps:

[0124] When the turbofan mode is switched to the turbojet mode, the mode regulating valve is opened, the main combustion chamber is closed, the core fan and the high-pressure compressor are in a windmill state, the high-pressure rotor and the intermediate-pressure rotor are stopped, the dual-variable combustion chamber and the afterburner are ignited, and the low-pressure rotor is started;

[0125] In the high-speed turbofan mode with a Mach number of 3.5 to 3.8 Ma, the mode switching is completed, and the total mode conversion time does not exceed 30 seconds;

[0126] The transient load characteristics of the multi-duct casing system during mode switching are analyzed, and the transient load characteristics include: thermal shock load caused by the increase in the combustion chamber outlet air flow temperature from 1700K of the dual-variable combustion chamber to 2200K of the afterburner combustion chamber; drastic changes in the low-pressure turbine inlet temperature and small changes in the low-pressure rotor speed.

[0127] Compared with the prior art, the beneficial effects of the present invention are as follows: the present invention can comprehensively analyze the structure and working process of a multi-bypass variable cycle engine, and by dividing the engine operation into three turbofan modes and one turbojet mode, it covers the working states under different flight conditions, solves the limitation of the prior art that only a single working condition can be analyzed, and provides an accurate basis for mode division for cross-airspace and cross-speed domain engine design. The method uses MATLAB software to perform component-level modeling based on the aerodynamic thermodynamic configuration model, comprehensively simulates the dynamic characteristics of the main components of the multi-bypass variable cycle engine such as the fan, combustion chamber, compressor and turbine, forms high-precision component-level modeling data, and provides a reliable data basis for the accurate calculation of the overall performance of the engine. The present invention establishes a common working equation group of a multi-bypass variable cycle engine, covering core equations such as flow balance, static pressure balance and power balance, and uses Broyden quasi-Newton method to iteratively solve the implicit nonlinear equation group, which greatly improves the calculation efficiency and convergence speed, and can quickly obtain the key performance parameters of the engine under different working conditions. At the same time, based on the solution of the common equation group, the limit aerodynamic load of the engine casing system in the typical working process is accurately calculated, and the transient response characteristics under mode switching and complex load conditions are effectively captured, providing a reliable basis for the reliability evaluation and structural optimization of the multi-duct casing system. The present invention combines the flight envelope of the engine, selects 12 typical altitudes and Mach number states, accurately calculates the limit working points under the turbofan mode and turbojet mode, and analyzes the transient load characteristics (such as aerodynamic load and thermal shock changes) during the mode switching process through simulation, breaking through the technical bottleneck of the prior art that is difficult to capture transient responses, and providing technical support for the dynamic optimization design of the multi-duct casing system. Through the above technical means, the present invention can adapt to the dynamic operation of the multi-duct variable cycle engine under high temperature, high speed and complex load conditions, significantly improve the performance analysis accuracy and adaptability of the engine, lay a solid foundation for the reliability design and life prediction of the multi-duct casing system, and solve the problems of insufficient analysis accuracy, difficulty in capturing dynamic characteristics and low calculation efficiency in the prior art, with significant technical progress and engineering practical value. BRIEF DESCRIPTION OF THE DRAWINGS

[0128] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the accompanying drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other accompanying drawings can be obtained based on these accompanying drawings without paying creative labor.

[0129] in:

[0130] Figure 1 is a flow chart of a method according to an embodiment of the present invention;

[0131] Figure 2 is a flow chart of a subcruise turbofan mode of a multi-bypass variable cycle engine in an embodiment of the present invention;

[0132] Figure 3 is a flow chart of a supercruise turbofan mode of a multi-bypass variable cycle engine in an embodiment of the present invention;

[0133] Figure 4 is a flow chart of a high-speed turbofan mode of a multi-pass variable cycle engine in an embodiment of the present invention;

[0134] Figure 5 is a flow chart of a high-speed turbojet mode of a multi-bypass variable cycle engine in an embodiment of the present invention;

[0135] Figure 6 1 is a schematic diagram of the flight envelope of a multi-bypass variable cycle engine in an embodiment of the present invention;

[0136] Figure 7 It is a schematic diagram of the change of the combustion chamber outlet temperature and the change of the low-pressure rotor speed under the mode transient in an embodiment of the present invention. DETAILED DESCRIPTION

[0137] In order to make the purpose, technical solution and advantages of the embodiment of the present invention clearer, the technical solution of the embodiment of the present invention will be clearly and completely described below in conjunction with the drawings of the embodiment of the present invention. Obviously, the described embodiment is a part of the embodiment of the present invention, not all of the embodiments. Based on the described embodiment of the present invention, all other embodiments obtained by ordinary technicians in this field belong to the scope of protection of the present invention.

[0138] like Figure 1 FIG. 1 is an embodiment of the present invention, which provides a method for analyzing transient characteristics of complex loads on a multi-duct casing, characterized in that it comprises the following steps:

[0139] Step 1: According to different flight missions, the mode selection valve and adjustable geometry can be adjusted to adjust the engine power cycle to achieve the required mission objectives. According to the airflow direction, combustion chamber usage and mode selection valve switch, it can be divided into four working modes: subcruise turbofan mode, supercruise turbofan mode, high-speed turbofan mode and high-speed turbojet mode. The mode flow path is as follows: Figure 2-Figure 5 shown.

[0140] like Figure 2 As shown, the working process of the Asian cruise turbofan mode is that the air intake duct sucks in gas, the inlet airflow passes through the air intake duct and enters the front fan. At this time, a small part of the airflow flows into the outermost third outer duct, which is recorded as airflow Q subsonic1 The rest of the airflow flows into the rear fan. After being compressed by the rear fan, the airflow is divided into two parts. One part of the airflow flows into the second duct and undergoes heat exchange through the heat exchanger, which is recorded as airflow Q. subsonic2 The remaining airflow flows into the core fan for compression. After passing through the core fan, the airflow is divided into two parts. One part of the airflow flows into the first outer duct and then enters the low-pressure turbine for cooling, which is recorded as Q subsonic3 The last part of the airflow flows into the high pressure compressor of the inner duct and enters the main combustion chamber, which is recorded as Q subsonic4 .

[0141] There is bleed air between the high-pressure compressor stages. Part of the airflow is used to cool the intermediate-pressure turbine, and part of the high-temperature airflow at the high-pressure compressor outlet flows into the heat exchanger pipe to exchange heat with the low-temperature airflow in the second duct before flowing into the high-pressure turbine. subsonic3 The air is mixed in the front mixing chamber, and then enters the dual variable combustion chamber and the low pressure turbine, and then mixes with the second and third bypass airflows Q in the rear mixing chamber. subsonic1 and Q subsonic2 Mixing occurs and finally the thrust is obtained by discharging the fuel through the tail nozzle. This mode ensures that the multi-bypass variable cycle engine can be in a mode of low fuel consumption and high thrust output.

[0142] like Figure 3 As shown in the figure, the working process of the supercruise turbofan mode is that the main combustion chamber outlet airflow is mixed with the second outer bypass airflow and then flows into the double variable combustion chamber for combustion again, and then driven by the low-pressure turbine to do work. The other parts are the same as the subcruise turbofan mode. The third outer bypass airflow is Q supersonic1 , the second bypass airflow is Q supersonic2 , the first bypass airflow is Q supersonic3 , the internal flow is Q supersonic4 This mode ensures that the multi-bypass variable cycle engine can be in a mode with low fuel consumption and high combat requirements.

[0143] like Figure 4As shown in the figure, in the high-speed turbofan mode, the low-pressure turbine outlet airflow mixes with the third outer bypass airflow and flows into the afterburner to burn, and then is discharged through the tail nozzle to generate thrust. The other parts are the same as the supercruise turbofan mode. Similarly, the airflow is recorded as Q Hi (i=1,2,3,4). In this mode, the multi-bypass variable cycle engine can be in a state of high-efficiency thrust output but high fuel consumption rate, as a short intermediate process of switching between turbofan and turbojet modes.

[0144] like Figure 5 As shown in the figure, in the high-speed turbojet mode, the multi-bypass variable cycle engine is in the high-speed turbojet mode, the first and second bypasses become inner bypasses, most of the inlet airflow flows along the inner bypass into the dual variable combustion chamber for combustion, a small part flows into the third outer bypass, and then the low-pressure turbine outlet airflow is mixed with the third outer bypass outlet airflow and flows into the afterburner for combustion, and finally discharged through the tail nozzle. The inner bypass airflow is denoted as Q turbojet1 , the airflow in the third bypass is Q turbojet2 By optimizing the distribution of core airflow and energy conversion efficiency, high thrust output and rapid response are achieved to meet the needs of hypersonic flight missions.

[0145] Step 2: Based on MATLAB software and the aerodynamic thermodynamic configuration scheme of the Basic Science Center, component-level modeling of the multi-bypass variable cycle engine was performed.

[0146] Preferably, the working modes include subcruise turbofan mode, supercruise turbofan mode, high-speed turbofan mode and high-speed turbojet mode, wherein:

[0147] The sub-cruise turbofan mode is suitable for flight speeds of 0-1.3 Ma and altitudes of 0-11000 m. The combustion chamber only opens the main combustion chamber and the mode selection valve is closed.

[0148] The supercruise turbofan mode is suitable for flight speeds of 1.3–3.5 Ma and altitudes of 11,000–23,000 m, with the main combustion chamber and the dual variable combustion chamber open, and the mode selection valve closed;

[0149] The high-speed turbofan mode is suitable for flight speeds of 3.5–4.0 Ma and altitudes of 23,000–27,000 m. The combustion chambers open the main combustion chamber, dual variable combustion chamber, and afterburner, and the mode selection valve is closed;

[0150] The high-speed turbojet mode is suitable for flight speeds of 3.7-5.0 Ma and altitudes of 24,000-30,000 m. The combustion chamber opens the dual-variable combustion chamber and the afterburner, and the mode selection valve is opened.

[0151] Preferably, the multi-bypass variable cycle engine includes a fan, a high-pressure compressor, a combustion chamber and a turbine, and component-level modeling is performed for these performance components, wherein:

[0152] The multi-duct variable cycle engine includes three fans, namely a front fan, a rear fan and a core fan. In turbofan mode, all three fans are involved in the work; in turbojet mode, the front fan and the rear fan are in a windmill state;

[0153] The multi-bypass variable cycle engine is a three-rotor structure, including a high-pressure turbine, an intermediate-pressure turbine and a low-pressure turbine, wherein the intermediate-pressure turbine guide vanes and the low-pressure turbine guide vanes are geometrically adjustable components, and the turbine performance characteristics are adjusted by adjusting the guide vane angle.

[0154] Preferably, the geometrically adjustable characteristics of the intermediate-pressure turbine and low-pressure turbine guide vanes are used to adjust turbine performance parameters by:

[0155] Based on the condition that different guide vane angles α are zero, the reference pressure ratio, reference efficiency and reference flow rate are obtained by combining turbine experimental data and two-dimensional interpolation method;

[0156] When the guide vane angle α is non-zero, the pressure ratio, efficiency and flow rate of the turbine are calculated according to the following formulas:

[0157] The turbine pressure ratio π is:

[0158] π=(π map -1)(1+α+k π )+1;

[0159] Where: π map is the reference pressure ratio; k π is the influence coefficient of guide vane angle on pressure ratio;

[0160] The turbine efficiency η is:

[0161]

[0162] The turbine flow rate W is:

[0163] W=W map (1+α+k w );

[0164] Where: W map is the reference flow rate; k w is the influence coefficient of guide vane angle on flow rate;

[0165] The meshgrid and interp2 functions of MATLAB are used to fit the turbine characteristic curve, and the total pressure, total temperature and turbine power at the turbine outlet are obtained by combining aerodynamic thermodynamic calculations.

[0166] The set of equations working together include:

[0167] Based on the structure of a multi-bypass variable cycle engine, a mathematical model is established to describe the relationship between the airflow and energy transfer inside the engine. The mathematical model includes:

[0168] The flow balance equation is used to describe the conservation relationship of airflow mass between engine components;

[0169] The static pressure balance equation is used to describe the continuity of static pressure between engine components;

[0170] The power balance equation, which describes the energy transfer relationship between the high-pressure, intermediate-pressure, and low-pressure rotors;

[0171] The common working equation group consists of a set of 43 implicit nonlinear equations, wherein the change of each variable has a coupled effect on the overall residual of the equation group.

[0172] Preferably, the mathematical model is solved iteratively by Broyden quasi-Newton method, and the iterative calculation process includes the following steps:

[0173] Set the initial input variables and establish the computational framework for iterative solution;

[0174] Update the relationship between input variables and residuals, and gradually correct the variables;

[0175] The working state of the multi-bypass variable cycle engine is solved by iterative calculation until the result converges to the set error threshold.

[0176] Preferably, the iterative calculation process of the Broyden quasi-Newton method comprises the following steps:

[0177] Initial state variable setting: Set the initial state variable vector X0, where X represents the key parameters describing the state of the multi-duct casing system, including: turbine pressure ratio, temperature, flow rate, and speed;

[0178] Calculation of initial residual matrix: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X) and calculate the initial residual matrix E0. The formula is as follows:

[0179] E = F(X);

[0180] Among them, E is the residual matrix, which represents the deviation between the current state parameters and the target parameters of the multi-duct casing system;

[0181] Jacobian matrix calculation: According to the aerodynamic thermodynamic model F(X), the initial Jacobian matrix J is calculated using the formula:

[0182]

[0183] Among them: E1, E2, ..., E7 are residual matrix components; X1, X2, ..., X7 are state variable components;

[0184] Iterative variable update: The initial state variables, initial residual matrix and Jacobian matrix are brought into the iterative variable update formula to correct the state variables. The calculation formula is as follows:

[0185]

[0186] Where: X i represents the state variable of the i-th iteration; J -1 is the inverse matrix of the Jacobian matrix; is the transpose of the residual matrix of the i-th iteration;

[0187] Residual matrix update: The updated state variables are input into the aerodynamic thermodynamic model again, and the update formula is:

[0188] E i+1 =F(X i+1 );

[0189] Where: E i+1 represents the residual matrix of the i+1th iteration;

[0190] Jacobian matrix correction: According to the state variable correction value and the residual matrix correction value, the Jacobian matrix is ​​corrected. The calculation formula is as follows:

[0191] Y=B i+1 -B i ;

[0192] Z=E i+1 -E i ;

[0193]

[0194] Among them: B i+1 is the Jacobian matrix approximation of the i+1th iteration; B i Jacobian matrix approximation for the current iteration;

[0195] Y is the change between the state variable correction values; Z is the change of the residual matrix; J i+1 is the modified Jacobian matrix; Y T is the transposed matrix of the change in the state variable correction value Y;

[0196] Repeat the iteration until the residual matrix E converges to the set threshold.

[0197] Preferably, in step 3, a common working equation group of a multi-bypass variable cycle engine is established, and the Broyden quasi-Newton method is used to iteratively solve the common equation group.

[0198] At the design point H = 20000m and Ma = 2.35, the multi-bypass variable cycle engine is in the supercruise state, the main combustion chamber and the double variable combustion chamber are opened, and the fuel flow, rotor speed and initial guess matrix X are shown in Table 1:

[0199] Table 1 Initial parameter settings

[0200]

[0201] The iteratively solved steady-state simulation results are compared with the design point values, and the results are shown in Table 2:

[0202] Table 2 Simulation results

[0203]

[0204] The results show that the maximum error between the steady-state simulation results and the design point does not exceed 0.9%.

[0205] In step 4, the performance model is used to calculate the limit operating point of the multi-ducted casing system, and the transient load characteristics of the multi-ducted casing system are analyzed.

[0206] According to the design requirements of the new generation of multi-bypass variable cycle engines across airspace and speed ranges, the airspace and speed ranges of the multi-bypass variable cycle engine cover 0-33km and 0-5Ma, including turbofan mode and turbojet mode. The flight envelope diagram of the multi-bypass variable cycle engine is shown in the figure below: Figure 6 As shown in Table 3, H = 20000m and Ma = 2.35 are the design points, and 12 typical altitudes and Mach number states are selected for analysis. The highest turbofan mode reaches H = 24000m and Ma = 3.7, as shown in Table 3:

[0207] Table 3 Typical operating points of multi-bypass variable cycle engines

[0208]

[0209]

[0210] Except for the ground slow state, the remaining 10 points are the limit working points of the multi-bypass variable cycle engine under different flight missions. In turbofan mode, the aerodynamic load of the multi-bypass variable cycle engine comes from the high-pressure rotor, the intermediate-pressure rotor and the low-pressure rotor; in turbojet mode, the high-pressure rotor and the intermediate-pressure rotor are both in the windmill state, and the speed is very low, so there is no need to calculate the speed of the high-pressure rotor and the intermediate-pressure rotor, only the low-pressure rotor speed is calculated.

[0211] The limit point of the multi-bypass variable cycle engine turbofan mode is divided into three working conditions: subsonic cruise mode, supersonic cruise mode and high-speed afterburner mode. The limit point is denoted as D i(H i ,Ma i ), according to the common equation solving method of multi-bypass variable cycle engine, set the limit point input conditions, including the intermediate pressure turbine guide vane angle α M , low pressure turbine guide vane angle α L , Main combustion chamber fuel inlet flow W in,fb1 , Dual variable combustion chamber fuel inlet flow W in,fb2 and afterburner fuel inlet flow W in,fb3 , where the mode selection valve is in the closed state.

[0212] Table 4 Status point input parameter settings

[0213]

[0214]

[0215] Table 5 Relative speed of each state point

[0216]

[0217] The working modes of multi-bypass variable cycle engines are divided into turbofan mode and turbojet mode. In the actual working process, these two working modes will switch back and forth, and the strong transient temperature and load characteristics caused by the working mode switching. Taking the aerodynamic load as an example, when the turbofan mode is switched to the turbojet mode, the mode regulating valve is opened, the main combustion chamber is closed, the core fan and the high-pressure compressor are in the windmill state, the high-pressure rotor and the intermediate-pressure rotor stop working, the dual-variable combustion chamber and the afterburner are ignited, the airflow temperature at the outlet of the dual-variable combustion chamber reaches about 1700K, the airflow temperature at the outlet of the afterburner reaches about 2200K, and the low-pressure rotor starts working.

[0218] When the multi-bypass variable cycle engine is in level flight, it is generally in high-speed turbofan mode with a Mach number of 3.5-3.8 Ma. The mode conversion starts from closing the main combustion chamber and opening the mode regulating valve, to high-altitude start-up, ignition of the dual variable combustion chamber and the afterburner, and the aircraft acceleration process, which does not exceed 30 seconds. The changes in the combustion chamber outlet temperature and the low-pressure rotor speed under the mode transient are as follows: Figure 7 shown.

[0219] In summary, the present invention can comprehensively analyze the transient characteristics of the multi-duct casing system under high temperature, high speed and complex load conditions by establishing a common working equation group of the multi-duct variable cycle engine and combining component-level modeling with an efficient iterative solution method. It not only solves the problems of insufficient analysis accuracy, difficulty in capturing dynamic characteristics and low calculation efficiency in the prior art, but also provides an important theoretical basis for the reliability design and optimization of the multi-duct casing system. The present invention is suitable for the design and development of a new generation of multi-duct variable cycle engines, has significant technological advancement and engineering practical value, and is of great significance to promoting the development of technology in the field of aviation engines.

[0220] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. Moreover, the specific features, structures, materials or characteristics described may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine different embodiments or examples described in this specification and the features of different embodiments or examples, unless they are contradictory.

[0221] Any process or method description in the flow chart or otherwise described herein can be understood to represent a module, fragment or portion of a code including one or more executable instructions for implementing the steps of a specific logical function or process. And the scope of the preferred embodiment of the present application includes other implementations, in which the functions may not be performed in the order shown or discussed, including in a substantially simultaneous manner or in a reverse order according to the functions involved.

[0222] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any technician familiar with the technical field can easily think of various changes or substitutions within the technical scope disclosed in the present application, which should be included in the protection scope of the present application. Therefore, the protection scope of the present application should be based on the protection scope of the claims.

Claims

1. A method for analyzing transient characteristics of complex loads on a multi-duct casing, characterized in that: The following steps are included: Analyze the structure and working process of the multi-bypass variable cycle engine, determine the working mode, divide the working process into three turbofan modes and one turbojet mode, and form the corresponding working process analysis results; Based on the work process analysis results, MATLAB software is used to perform component-level modeling of the multi-bypass variable cycle engine based on the aerodynamic thermodynamic configuration model to form component-level modeling data; Using component-level modeling data, a common working equation group of a multi-bypass variable cycle engine is established, and a Broyden quasi-Newton method is used to iteratively solve the solution of the common working equation group to obtain a solution of the equation group; Based on the solution of the equations, the ultimate aerodynamic load of the multi-duct casing system during operation is calculated; By using the extreme aerodynamic loads and combining the flight envelope of the multi-bypass variable cycle engine, 12 typical altitudes and Mach number states are selected to solve the extreme operating points in turbofan mode and turbojet mode, and the transient load characteristics of the multi-bypass casing system are analyzed.

2. A method for analyzing complex load transient characteristics of a multi-duct casing according to claim 1, characterized in that: The working modes include subcruise turbofan mode, supercruise turbofan mode, high-speed turbofan mode and high-speed turbojet mode, wherein: The sub-cruise turbofan mode is suitable for flight speeds of 0-1.3 Ma and altitudes of 0-11000 m. The combustion chamber only opens the main combustion chamber and the mode selection valve is closed. The supercruise turbofan mode is suitable for flight speeds of 1.3–3.5 Ma and altitudes of 11,000–23,000 m, with the main combustion chamber and the dual variable combustion chamber open, and the mode selection valve closed; The high-speed turbofan mode is suitable for flight speeds of 3.5–4.0 Ma and altitudes of 23,000–27,000 m. The combustion chambers open the main combustion chamber, dual variable combustion chamber, and afterburner, and the mode selection valve is closed; The high-speed turbojet mode is suitable for flight speeds of 3.7-5.0 Ma and altitudes of 24,000-30,000 m. The combustion chamber opens the dual-variable combustion chamber and the afterburner, and the mode selection valve is opened.

3. The method for analyzing complex load transient characteristics of a multi-duct casing according to claim 1, characterized in that: The multi-bypass variable cycle engine includes a fan, a high-pressure compressor, a combustion chamber and a turbine, and component-level modeling is performed for these performance components, wherein: The multi-duct variable cycle engine includes three fans, namely a front fan, a rear fan and a core fan. In turbofan mode, all three fans are involved in the work; in turbojet mode, the front fan and the rear fan are in a windmill state; The multi-bypass variable cycle engine is a three-rotor structure, including a high-pressure turbine, an intermediate-pressure turbine and a low-pressure turbine, wherein the intermediate-pressure turbine guide vanes and the low-pressure turbine guide vanes are geometrically adjustable components, and the turbine performance characteristics are adjusted by adjusting the guide vane angle.

4. A method for analyzing complex load transient characteristics of a multi-duct casing according to claim 3, characterized in that: The geometrically adjustable characteristics of the intermediate-pressure turbine and low-pressure turbine guide vanes are used to adjust turbine performance parameters by: based on the condition that different guide vane angles α are zero, a reference pressure ratio, a reference efficiency and a reference flow rate are obtained by combining turbine experimental data and a two-dimensional interpolation method; When the guide vane angle α is non-zero, the pressure ratio, efficiency and flow rate of the turbine are calculated according to the following formulas: The turbine pressure ratio π is: π＝(π map -1)(1+α+k π )+1； Where: π map is the reference pressure ratio; k π is the influence coefficient of guide vane angle on pressure ratio; The turbine efficiency η is: The turbine flow rate W is: W=W map (1+α+k w ); Where: W map is the reference flow rate; k w is the influence coefficient of guide vane angle on flow rate; The meshgrid and interp2 functions of MATLAB are used to fit the turbine characteristic curve, and the total pressure, total temperature and turbine power at the turbine outlet are obtained by combining aerodynamic thermodynamic calculations.

5. The method for analyzing complex load transient characteristics of a multi-duct casing according to claim 1, characterized in that: The set of equations working together include: Based on the structure of a multi-bypass variable cycle engine, a mathematical model is established to describe the relationship between the airflow and energy transfer inside the engine. The mathematical model includes: The flow balance equation is used to describe the conservation relationship of airflow mass between engine components; The static pressure balance equation is used to describe the continuity of static pressure between engine components; The power balance equation, which describes the energy transfer relationship between the high-pressure, intermediate-pressure, and low-pressure rotors; The common working equation group consists of a set of 43 implicit nonlinear equations, wherein the change of each variable has a coupled effect on the overall residual of the equation group.

6. A method for analyzing complex load transient characteristics of a multi-duct casing according to claim 5, characterized in that: The mathematical model is solved iteratively by Broyden quasi-Newton method, and the iterative calculation process includes the following steps: Set the initial input variables and establish the computational framework for iterative solution; Update the relationship between input variables and residuals, and gradually correct the variables; The working state of the multi-bypass variable cycle engine is solved by iterative calculation until the result converges to the set error threshold.

7. A method for analyzing complex load transient characteristics of a multi-duct casing according to claim 6, characterized in that: The iterative calculation process of the Broyden quasi-Newton method includes the following steps: Initial state variable setting: Set the initial state variable vector X0, where X represents the key parameters describing the state of the multi-duct casing system, including: turbine pressure ratio, temperature, flow rate, and speed; Calculation of initial residual matrix: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X) and calculate the initial residual matrix E0. The formula is as follows: E = F(X); Among them, E is the residual matrix, which represents the deviation between the current state parameters and the target parameters of the multi-duct casing system; Jacobian matrix calculation: According to the aerodynamic thermodynamic model F(X), the initial Jacobian matrix J is calculated using the formula: Among them: E1, E2, ..., E7 are residual matrix components; X1, X2, ..., X7 are state variable components; Iterative variable update: The initial state variables, initial residual matrix and Jacobian matrix are brought into the iterative variable update formula to correct the state variables. The calculation formula is as follows: Where: X i represents the state variable of the i-th iteration; J -1 is the inverse matrix of the Jacobian matrix; is the transpose of the residual matrix of the i-th iteration; Residual matrix update: The updated state variables are input into the aerodynamic thermodynamic model again, and the update formula is: E i+1 =F(X i+1 ); Where: E i+1 represents the residual matrix of the i+1th iteration; Jacobian matrix correction: According to the state variable correction value and the residual matrix correction value, the Jacobian matrix is ​​corrected. The calculation formula is as follows: Y = B i+1 -B i ; Z=E i+1 -E i ; Among them: B i+1 is the Jacobian matrix approximation of the i+1th iteration; B i Jacobian matrix approximation for the current iteration; Y is the change between the state variable correction values; Z is the change of the residual matrix; J i+1 is the modified Jacobian matrix; Y T is the transposed matrix of the change in the state variable correction value Y; Repeat the iteration until the residual matrix E converges to the set threshold.

8. The method for analyzing complex load transient characteristics of a multi-duct casing according to claim 5, characterized in that: The mathematical model also includes: Determine the flight envelope based on the cross-airspace and cross-speed domain design requirements of the multi-bypass variable cycle engine; Twelve extreme working points are selected from the flight envelope for simulation and solution, where: The two working points are ground slow and ground takeoff states; The six operating points are typical states in turbofan mode; The four operating points are typical conditions in turbojet mode; The limit operating point is simulated and solved to obtain the high-pressure rotor speed, the medium-pressure rotor speed and the low-pressure rotor speed; The changes in the outlet temperatures of the dual-variable combustion chamber and afterburner combustion chamber during the mode switching of a multi-duct variable cycle engine are analyzed to determine the thermal shock damage characteristics of the multi-duct casing system caused by transient loads.

9. A method for analyzing complex load transient characteristics of a multi-duct casing according to claim 8, characterized in that: The 12 extreme operating points selected from the flight envelope include: The six typical working points in turbofan mode correspond to the following working conditions: Subsonic cruise mode; Supersonic cruise mode; High speed afterburner mode; The four typical flight states in turbojet mode correspond to the following operating points: D7(24,3.7), D8(27,4.0), D9(28,4.3) and D 10 (30,5.0).

10. The method for analyzing complex load transient characteristics of a multi-duct casing according to claim 1, characterized in that: The method further comprises the following steps: When the turbofan mode is switched to the turbojet mode, the mode regulating valve is opened, the main combustion chamber is closed, the core fan and the high-pressure compressor are in a windmill state, the high-pressure rotor and the intermediate-pressure rotor are stopped, the dual-variable combustion chamber and the afterburner are ignited, and the low-pressure rotor is started; In the high-speed turbofan mode with a Mach number of 3.5 to 3.8 Ma, the mode switching is completed, and the total mode conversion time does not exceed 30 seconds; The transient load characteristics of the multi-duct casing system during mode switching are analyzed, and the transient load characteristics include: thermal shock load caused by the increase in the combustion chamber outlet air flow temperature from 1700K of the dual-variable combustion chamber to 2200K of the afterburner combustion chamber; drastic changes in the low-pressure turbine inlet temperature and small changes in the low-pressure rotor speed.

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