A method for analyzing the transient characteristics of complex loads on multi-duct casings

Abstract

This invention discloses a method for analyzing the transient characteristics of multi-ducted engine casings under complex loads, belonging to the field of aero-engine technology. This method analyzes the structure and operation of a multi-ducted variable cycle engine, classifying it into three turbofan modes and one turbojet mode. Based on an aerodynamic thermodynamic configuration model, component-level modeling is performed using MATLAB software, generating component-level modeling data. A common set of working equations is established using the modeling data, and the solution is obtained through iterative solving using the Broyden quasi-Newton method. The ultimate aerodynamic loads of the casing system under typical operating conditions are calculated. Combined with the engine flight envelope, 12 typical altitude and Mach number states are selected to solve for the ultimate operating points and analyze the transient load characteristics. This invention can accurately capture the transient response characteristics of multi-ducted engine casing systems under complex load conditions, improving analysis accuracy and efficiency, providing a theoretical basis for the design optimization and reliability assessment of casing systems, and has significant engineering value.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology, and more specifically to a method for analyzing the transient characteristics of complex loads on a multi-duct casing. Background Technology

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

[0003] Multi-bypass variable-cycle engines (MBOs) achieve the goal of maintaining high thrust output while reducing fuel consumption across a wide flight envelope by adjusting aerodynamic and thermodynamic parameters, becoming an important direction in modern aero-engine technology. However, their complex airflow organization and mechanical structure generate significant transient loads and thermal shocks during operating mode switching (such as between turbofan and turbojet modes). This dynamic characteristic not only challenges the reliability and lifespan of the engine casing system but also places higher demands on the real-time performance and accuracy of analytical methods.

[0004] Most existing technologies focus on aerodynamic characteristics under single steady-state conditions, making it difficult to effectively capture complex transient responses during mode switching. Furthermore, they lack efficient mathematical modeling and solution methods, failing to meet the real-time analysis requirements across airspace and velocity domains. Therefore, developing an efficient method capable of comprehensively analyzing the dynamic characteristics of multi-bypass variable-cycle engines has become a key research direction for improving their reliability and performance. Summary of the Invention

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

[0006] To achieve the above objectives, the present invention employs the following technical solution:

[0007] A method for analyzing the transient characteristics of complex loads on multi-duct casings, characterized by comprising the following steps:

[0008] The structure and working process of the multi-bypass variable cycle engine were analyzed, the working mode was determined, and the working process was divided into three turbofan modes and one turbojet mode, and corresponding working process analysis results were formed.

[0009] Based on the analysis results of the working process, using MATLAB software and based on the aero-thermodynamic configuration model, component-level modeling of the multi-duct variable cycle engine was carried out to form component-level modeling data.

[0010] Using component-level modeling data, a common working equation set for a multi-duct variable cycle engine is established, and the solution of the common working equation set is obtained by iteratively solving the solution of the equation set using the Broyden quasi-Newton method.

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

[0012] By utilizing the ultimate aerodynamic load and combining it with the flight envelope of the multi-ducted variable cycle engine, 12 typical altitude and Mach number states were selected to solve the ultimate operating points in turbofan and turbojet modes, and the transient load characteristics of the multi-ducted casing system were analyzed.

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

[0014] The sub-cruise turbofan mode is suitable for flight speeds of 0–1.3 Mach and altitudes of 0–11000 m. Only the main combustion chamber is open, 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. The main combustion chamber and dual variable combustion chamber are open, and the mode selection valve is closed.

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

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

[0018] As a preferred embodiment 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-bypass variable cycle engine includes three fans: a front fan, a rear fan, and a core fan. In turbofan mode, all three fans are engaged; in turbojet mode, the front fan and the rear fan operate in a windmill state.

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

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

[0022] Based on the condition that the guide vane angle α is zero for different purposes, the reference pressure ratio, reference efficiency, and reference flow rate are obtained by combining turbine experimental data and two-dimensional interpolation methods.

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

[0024] The turbine pressure ratio π is:

[0025] π=(π map -1)(1+α+k π +1;

[0026] Where: π map For reference pressure ratio; k π The coefficient representing the influence of the guide vane angle on the 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] Among them: W map For reference flow rate; k w The coefficient representing the influence of the guide vane angle on the flow rate;

[0032] The turbine characteristic curves were fitted using MATLAB's meshgrid and interp2 functions, and the total pressure, total temperature, and turbine power at the turbine outlet were obtained by combining aerodynamic thermodynamic calculations.

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

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

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

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

[0037] The power balance equation is used to describe the energy transfer relationship between high-pressure, medium-pressure, and low-pressure rotors;

[0038] The common working equation set consists of a set of 43 implicit nonlinear equations, in which changes in each variable have a coupled effect on the overall residual of the equation set.

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

[0041] X=[β1,β2,β3,β4,β5,β6,β7];

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

[0043] Lines; β4 represents auxiliary lines on the characteristic diagrams of the two and three bypass ducts, respectively; β5 and β6 represent the premixing lines, respectively.

[0044] Auxiliary lines for the characteristic diagrams of the mixing chamber and the after-mixing chamber; β7 represents the auxiliary lines for the characteristic diagram of the tail nozzle.

[0045] Based on the initial variables, seven sets of equations describing the airflow transfer and energy conservation in the engine 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: ε i Represents the residuals of each equilibrium equation;

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

[0051] High-pressure turbine flow balance:

[0052] W HPT,in -W HPT,out =ε1;

[0053] Among them: W HPT,in W represents the inlet gas mass flow rate of the high-pressure turbine; HPT,out ε represents the mass flow rate of the outlet gas from the high-pressure turbine; ε1 is the residual value of the flow balance of the high-pressure turbine.

[0054] Medium-pressure turbine flow balance:

[0055] W MPT,in -W MPT,out =ε2;

[0056] Among them: W MPT,in W represents the inlet air mass flow rate of the intermediate-pressure turbine; MPT,out ε1 represents the outlet airflow mass flow rate of the intermediate-pressure turbine; ε2 is the residual value of the flow balance of the intermediate-pressure turbine.

[0057] Low-pressure turbine flow balance:

[0058] W LPT,in -W LPT,out =ε3;

[0059] Among them: W LPT,in W represents the inlet gas mass flow rate of the low-pressure turbine; LPT,out ε3 represents the mass flow rate of the outlet airflow of the low-pressure turbine; ε3 is the residual value of the flow balance of the low-pressure turbine.

[0060] External bypass duct static pressure balance:

[0061] p s,second -p s,third =ε4;

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

[0063] Pre-mixing chamber static pressure balance (intermediate-pressure turbine outlet static pressure and core fan outlet static pressure):

[0064] p s,MPT,out -p s,core,out =ε5;

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

[0066] Post-mixing chamber static pressure balance (pre-mixing chamber outlet static pressure and low-pressure turbine outlet static pressure):

[0067] p s,Fmix,out -p s,LPT,out =ε6;

[0068] Where: p s,Fmix,out p represents the static pressure value at the outlet of the pre-mixing chamber. s,LPT,out ε represents the static pressure value at the low-pressure turbine outlet; ε6 is the residual value of the static pressure balance in the after-mixing chamber;

[0069] Tail nozzle flow balance:

[0070] W Nozzle,in -W Nozzle,out =ε7;

[0071] Among them: W Nozzle,in W represents the mass flow rate of the airflow at the nozzle inlet. Nozzle,out ε7 represents the mass flow rate of the airflow at the nozzle exit; ε7 is the residual value of the nozzle flow balance.

[0072] Furthermore, to describe the energy transfer characteristics of a multi-bypass variable cycle engine, a rotor power balance equation also needs to be established. Specifically:

[0073] Rotor power balance equation:

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

[0075] High-voltage rotor power balance:

[0076]

[0077] Medium-pressure rotor power balance:

[0078]

[0079] Low-voltage rotor power balance:

[0080]

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

[0082] The common operating equations of a multi-duct variable cycle engine are essentially a set of implicit nonlinear equations, requiring the solution of 43 equations. Changes to any variable affect the final residual result of the entire equation set, placing high demands on the convergence and real-time performance of solving these implicit nonlinear equations. Currently, there is no analytical method to find the variable solutions of these implicit nonlinear equations; only numerically optimal solutions satisfying certain errors can be found.

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

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

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

[0086] The solution for the operating state of the multi-duct variable cycle engine is completed by iterative calculation until the result converges to the set error threshold.

[0087] As a preferred embodiment 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-bypass casing system, including: turbine pressure ratio, temperature, flow rate, and speed;

[0089] Initial residual matrix calculation: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X), and calculate the initial residual matrix E0, as shown in the following formula:

[0090] E = F(X);

[0091] Where E is the residual matrix, representing the deviation between the current state parameters and the target parameters of the multi-duct casing system;

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

[0093]

[0094] Where: E1, E2, ..., E7 are the components of the residual matrix; X1, X2, ..., X7 are the components of the state variables;

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

[0096]

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

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

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

[0100] Among them: E i+1 Let represent the residual matrix of the (i+1)th iteration;

[0101] Jacobian matrix correction: The Jacobian matrix is ​​corrected based on the correction values ​​of the state variables and the correction values ​​of the residual matrix. 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 B is the approximate value of the Jacobian matrix in the (i+1)th iteration; i The approximate value of the Jacobian matrix in the current iteration;

[0106] Y represents the change between the state variable correction values; Z represents the change in the residual matrix; J i+1 Y is the corrected Jacobian matrix; T The transpose 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] The flight envelope is determined based on the cross-airspace and cross-speed domain design requirements of the multi-bypass variable cycle engine;

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

[0111] The two working points are ground idle and ground takeoff status;

[0112] The six operating points represent typical conditions under turbofan mode;

[0113] The four operating points represent typical conditions under turbojet mode;

[0114] The extreme operating points are simulated and solved to obtain the high-pressure rotor speed, medium-pressure rotor speed and low-pressure rotor speed;

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

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

[0117] The six typical operating points in turbofan mode correspond to the following operating states:

[0118] Subsonic cruise mode;

[0119] Supersonic cruise mode;

[0120] High-speed boost 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 includes the following steps:

[0124] When switching from turbofan mode to turbojet mode, open the mode adjustment valve, close the main combustion chamber, put the core fan and high-pressure compressor in the fan-mill state, stop the high-pressure rotor and intermediate-pressure rotor, ignite the dual-variable combustion chamber and afterburner, and start the low-pressure rotor.

[0125] In high-speed turbofan mode with Mach numbers of 3.5 to 3.8, the mode switching is completed in a total time of no more than 30 seconds;

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

[0127] Compared with existing technologies, the beneficial effects of this invention are as follows: This invention can comprehensively analyze the structure and working process of a multi-bypass variable cycle engine. By dividing engine operation into three turbofan modes and one turbojet mode, it covers the working states under different flight conditions, overcoming the limitation of existing technologies that can only analyze a single operating condition. This provides a precise mode division basis for cross-airspace and cross-speed domain engine design. This method utilizes MATLAB software to perform component-level modeling based on an aerodynamic thermodynamic configuration model, comprehensively simulating the dynamic characteristics of major components of the multi-bypass variable cycle engine, such as the fan, combustion chamber, compressor, and turbine, forming high-precision component-level modeling data, providing a reliable data foundation for accurate calculation of the overall engine performance. This invention establishes a common set of working equations for the multi-bypass variable cycle engine, covering core equations such as flow balance, static pressure balance, and power balance, and uses the Broyden quasi-Newton method to iteratively solve the implicit nonlinear equations, significantly improving computational efficiency and convergence speed, enabling rapid acquisition of key performance parameters of the engine under different operating conditions. Simultaneously, based on the solution of the common equation set, the ultimate aerodynamic load of the engine casing system during typical operation was accurately calculated, effectively capturing the transient response characteristics under mode switching and complex load conditions, providing a reliable basis for the reliability assessment and structural optimization of the multi-duct casing system. This invention, combined with the engine's flight envelope, selected 12 typical altitude and Mach number states to accurately calculate the ultimate operating points in turbofan and turbojet modes. Through simulation analysis of transient load characteristics (such as aerodynamic load and thermal shock changes) during mode switching, it overcame the technical bottleneck of existing technologies that struggle to capture transient responses, providing technical support for the dynamic optimization design of multi-duct casing systems. Through these technical means, this invention can adapt to the dynamic operation of multi-duct variable cycle engines under high temperature, high speed, and complex load conditions, significantly improving the accuracy and adaptability of engine performance analysis. It lays a solid foundation for the reliability design and life prediction of multi-duct casing systems, solving problems such as insufficient analytical accuracy, difficulty in capturing dynamic characteristics, and low computational efficiency in existing technologies, demonstrating significant technological advancement and engineering practical value. Attached Figure Description

[0128] To more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the following description of the embodiments will be briefly introduced. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0129] in:

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

[0131] Figure 2 This is a flowchart of the subcruise turbofan mode of the multi-duct variable cycle engine in an embodiment of the present invention;

[0132] Figure 3 This is a flowchart of the supercruise turbofan mode of the multi-duct variable cycle engine in an embodiment of the present invention;

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

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

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

[0136] Figure 7 This is a schematic diagram of the changes in combustion chamber outlet temperature and low-pressure rotor speed under mode transients in an embodiment of the present invention. Detailed Implementation

[0137] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the described embodiments of the present invention are within the scope of protection of the present invention.

[0138] like Figure 1 As shown, this is an embodiment of the present invention, which provides a method for analyzing the transient characteristics of complex loads on multi-duct casings, characterized by comprising the following steps:

[0139] Step 1: Depending on the flight mission, the mode selection valve and adjustable geometry can be adjusted to regulate the engine's power cycle and achieve the required mission objectives. Based on airflow direction, combustion chamber usage, and mode selection valve operation, four operating modes can be defined: subcruise turbofan mode, supercruise turbofan mode, high-speed turbofan mode, and high-speed turbojet mode. The mode flow paths are as follows: Figures 2-5 As shown.

[0140] like Figure 2 As shown, the operation of the sub-cruise turbofan mode involves the intake duct drawing in gas. The inlet airflow passes through the intake duct and enters the front fan. At this time, a small portion of the airflow flows into the outermost third bypass duct, denoted as airflow Q. subsonic1 All other airflow flows into the rear fan. After compression by the rear fan, the airflow is divided into two parts. One part flows into the second bypass duct and undergoes heat exchange through the heat exchanger; this part is denoted as airflow Q. subsonic2 The remaining airflow flows into the core fan for compression. After passing through the core fan, the airflow splits into two parts. One part flows into the first outer bypass duct and then enters the low-pressure turbine for cooling, denoted as Q. subsonic3 The final portion of the airflow flows into the high-pressure compressor in the inner duct and enters the main combustion chamber, denoted as Q. subsonic4 .

[0141] There is bleed air between the high-pressure compressor stages. Part of the bleed air cools the intermediate-pressure turbine, while part of the high-temperature airflow from the high-pressure compressor outlet flows into the heat exchanger duct, where it exchanges heat with the low-temperature airflow in the second bypass duct before flowing back into the high-pressure turbine. subsonic3 The mixture is first mixed in the pre-mixing chamber, then enters the dual-variable combustion chamber and the low-pressure turbine, and finally mixes with the airflow from the second and third outer bypass ducts in the post-mixing chamber. subsonic1 and Q subsonic2 The mixture is produced and finally discharged through the tailpipe to generate thrust. This mode ensures that the multi-bypass variable cycle engine can operate in a mode with low fuel consumption and high thrust output.

[0142] like Figure 3 As shown, in the supercruise turbofan mode, the airflow at the main combustion chamber outlet mixes with the airflow from the second bypass duct and flows into the dual-variable combustion chamber for re-combustion. Then, it is driven by the low-pressure turbine to perform work. Other parts are the same as in the subcruise turbofan mode. Let the airflow from the third bypass duct be denoted as Q. supersonic1 The airflow in the second bypass duct is Q. supersonic2 The airflow in the first bypass duct is Q. supersonic3 The internal airflow is Q. supersonic4 This mode ensures that the multi-bypass variable cycle engine can operate in a mode with low fuel consumption and high operational requirements.

[0143] like Figure 4As shown, in the high-speed turbofan mode, the airflow from the low-pressure turbine outlet mixes with the airflow from the third bypass duct and flows into the afterburner for combustion. The mixture is then discharged through the exhaust nozzle to generate thrust. Other aspects are the same as in the supercruise turbofan mode. Similarly, the airflow is denoted as Q. Hi (i = 1, 2, 3, 4). In this mode, the multi-bypass variable cycle engine can achieve a high thrust output but a high fuel consumption rate, serving as a brief intermediate process between turbofan and turbojet modes.

[0144] like Figure 5 As shown, in high-speed turbojet mode, the multi-bypass variable cycle engine operates in high-speed turbojet mode. The first and second bypass ducts become inner ducts. Most of the inlet airflow flows into the dual-variable combustion chamber for combustion along the inner duct, while a small portion flows into the third outer bypass duct. Then, the airflow from the low-pressure turbine outlet mixes with the airflow from the third outer bypass duct before flowing into the afterburner for combustion, and finally exits through the exhaust nozzle. Let the airflow in the inner duct be denoted as Q. turbojet1 The airflow in the third bypass duct is Q. turbojet2 By optimizing the distribution of core airflow and energy conversion efficiency, high thrust output and rapid response were achieved, meeting the requirements of hypersonic flight missions.

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

[0146] Preferably, the operating 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 Mach and altitudes of 0–11000 m. Only the main combustion chamber is open, 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. The main combustion chamber and dual variable combustion chamber are open, and the mode selection valve is closed.

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

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

[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 has been performed for these performance components, wherein:

[0152] The multi-bypass variable cycle engine includes three fans: a front fan, a rear fan, and a core fan. In turbofan mode, all three fans are engaged; in turbojet mode, the front fan and the rear fan operate in a windmill state.

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

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

[0155] Based on the condition that the guide vane angle α is zero for different purposes, the reference pressure ratio, reference efficiency, and reference flow rate are obtained by combining turbine experimental data and two-dimensional interpolation methods.

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

[0157] The turbine pressure ratio π is:

[0158] π=(π map -1)(1+α+k π +1;

[0159] Where: π map For reference pressure ratio; k π The coefficient representing the influence of the guide vane angle on the 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] Among them: W map For reference flow rate; k w The coefficient representing the influence of the guide vane angle on the flow rate;

[0165] The turbine characteristic curves were fitted using MATLAB's meshgrid and interp2 functions, and the total pressure, total temperature, and turbine power at the turbine outlet were obtained by combining aerodynamic thermodynamic calculations.

[0166] The common working equation set includes:

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

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

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

[0170] The power balance equation is used to describe the energy transfer relationship between high-pressure, medium-pressure, and low-pressure rotors;

[0171] The common working equation set consists of a set of 43 implicit nonlinear equations, in which changes in each variable have a coupled effect on the overall residual of the equation set.

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

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

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

[0175] The solution for the operating state of the multi-duct variable cycle engine is completed by iterative calculation until the result converges to the set error threshold.

[0176] Preferably, the iterative calculation process of the Broyden quasi-Newton method includes 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-bypass casing system, including: turbine pressure ratio, temperature, flow rate, and speed;

[0178] Initial residual matrix calculation: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X), and calculate the initial residual matrix E0, as shown in the following formula:

[0179] E = F(X);

[0180] Where E is the residual matrix, representing the deviation between the current state parameters and the target parameters of the multi-duct casing system;

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

[0182]

[0183] Where: E1, E2, ..., E7 are the components of the residual matrix; X1, X2, ..., X7 are the components of the state variables;

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

[0185]

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

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

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

[0189] Among them: E i+1 Let represent the residual matrix of the (i+1)th iteration;

[0190] Jacobian matrix correction: The Jacobian matrix is ​​corrected based on the correction values ​​of the state variables and the correction values ​​of the residual matrix. 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 B is the approximate value of the Jacobian matrix in the (i+1)th iteration; i The approximate value of the Jacobian matrix in the current iteration;

[0195] Y represents the change between the state variable correction values; Z represents the change in the residual matrix; J i+1 Y is the corrected Jacobian matrix; T The transpose 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 set for the multi-duct variable cycle engine is established, and the Broyden quasi-Newton method is used to iteratively solve the common equation set.

[0198] At the design points H = 20000m and Ma = 2.35, the multi-bypass variable cycle engine is in overcrowding mode with the main combustion chamber and dual variable combustion chambers open. The fuel flow rate, rotor speed, and initial prediction matrix X are shown in Table 1.

[0199] Table 1 Initial Parameter Settings

[0200]

[0201] The steady-state simulation results obtained through iterative solutions 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] Step 4: The performance model was used to calculate the extreme operating point of the multi-duct casing system, and the transient load characteristics of the multi-duct casing system were analyzed.

[0206] Based on the cross-airspace and cross-velocity design requirements of the new generation multi-bypass variable cycle engine, the multi-bypass variable cycle engine covers an airspace and velocity range of 0–33 km and 0–5 Ma, respectively, including turbofan and turbojet modes. The flight envelope of this multi-bypass variable cycle engine is shown below. Figure 6 As shown in Table 3, H = 20000m and Ma = 2.35 are the design points. Twelve typical altitudes and Mach numbers were selected for analysis. The turbofan model reached a maximum of 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] Excluding the ground idle state, the remaining 10 points represent the extreme operating 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 originates from the high-pressure rotor, intermediate-pressure rotor, and low-pressure rotor; in turbojet mode, the high-pressure rotor and intermediate-pressure rotor are in a windmill state with very low speeds, so it is not necessary to calculate the speeds of the high-pressure rotor and intermediate-pressure rotor, only the speed of the low-pressure rotor needs to be calculated.

[0211] The turbofan engine's extreme points are divided into three operating conditions: subsonic cruise mode, supersonic cruise mode, and high-speed afterburner mode. The extreme point is denoted as D. i(H i Ma i Based on the solution method for the common equations of multi-bypass variable cycle engines, limit point input conditions are set, including the intermediate-pressure turbine guide vane angle α. M Low-pressure turbine guide vane angle α L , Main combustion chamber fuel inlet flow rate W in,fb1 Dual-variable combustion chamber fuel inlet flow rate W in,fb2 and the fuel inlet flow rate W of the afterburner in,fb3 The mode selection valve is in the closed state.

[0212] Table 4 Status Point Input Parameter Settings

[0213]

[0214]

[0215] Table 5 Relative rotational speeds at each state point

[0216]

[0217] Multi-bypass variable cycle engines operate in two modes: turbofan mode and turbojet mode. During actual operation, these two modes switch back and forth, resulting in strong transient temperature and load characteristics. Taking aerodynamic load as an example, when switching from turbofan mode to turbojet mode, the mode control valve opens, the main combustion chamber closes, the core fan and high-pressure compressor operate in a fan-milled state, the high-pressure rotor and intermediate-pressure rotor stop working, the dual-variable combustion chamber and afterburner ignite, the outlet gas temperature of the dual-variable combustion chamber reaches approximately 1700K, and the outlet gas temperature of the afterburner reaches approximately 2200K, at which point the low-pressure rotor begins to operate.

[0218] During level flight maneuvers, multi-bypass variable cycle engines typically operate in high-speed turbofan mode, with Mach numbers between 3.5 and 3.8. The mode transition, from the closure of the main combustion chamber's mode control valve to high-altitude start-up, ignition of the dual variable combustion chamber and afterburner, and aircraft acceleration, lasts no more than 30 seconds. The changes in combustion chamber outlet temperature and low-pressure rotor speed during this transient mode transition are as follows: Figure 7 As shown.

[0219] In summary, this invention, by establishing a common set of working equations for a multi-duct variable cycle engine and combining component-level modeling with efficient iterative solution methods, can comprehensively analyze the transient characteristics of the multi-duct casing system under high temperature, high speed, and complex load conditions. This not only solves the problems of insufficient analytical accuracy, difficulty in capturing dynamic characteristics, and low computational efficiency in existing technologies, but also provides important theoretical basis for the reliability design and optimization of multi-duct casing systems. This invention is applicable to the design and development of next-generation multi-duct variable cycle engines, possessing significant technological advancements and engineering practical value, and is of great importance to promoting the development of technology in the field of aero-engines.

[0220] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of this application. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of those different embodiments or examples.

[0221] Any process or method description in the flowchart or otherwise herein can be understood as representing a module, segment, or portion of code comprising one or more executable instructions for implementing a particular logical function or process. Furthermore, the scope of the preferred embodiments of this application includes additional implementations in which functions may be performed not in the order shown or discussed, including substantially simultaneously or in reverse order depending on the functionality involved.

[0222] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any person skilled in the art can easily conceive of various variations or substitutions within the technical scope disclosed in this application, and these should all be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for analyzing the transient characteristics of complex loads on multi-duct casings, characterized in that, It includes the following steps: The structure and working process of the multi-bypass variable cycle engine were analyzed, the working mode was determined, and the working process was divided into three turbofan modes and one turbojet mode, and corresponding working process analysis results were formed. Based on the analysis results of the working process, using MATLAB software and based on the aero-thermodynamic configuration model, component-level modeling of the multi-duct variable cycle engine was carried out to form component-level modeling data. Using component-level modeling data, a common working equation set for a multi-duct variable cycle engine is established, and the solution of the common working equation set is obtained by iteratively solving the solution of the equation set using the Broyden quasi-Newton method. Based on the solution of the equation system, the ultimate aerodynamic load of the multi-duct casing system during operation is calculated. By utilizing the ultimate aerodynamic load and combining it with the flight envelope of the multi-ducted variable cycle engine, 12 typical altitude and Mach number states were selected to solve the ultimate operating points in turbofan and turbojet modes, and the transient load characteristics of the multi-ducted casing system were analyzed.

2. The method for analyzing the transient characteristics of complex loads on a multi-duct casing according to claim 1, characterized in that, The operating 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 Mach and altitudes of 0–11000 m. Only the main combustion chamber is open, 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. The main combustion chamber and dual variable combustion chamber are open, and the mode selection valve is closed. The high-speed turbofan mode is suitable for flight speeds of 3.5–4.0 Mach and altitudes of 23,000–27,000 m. The combustion chamber is open, including the main combustion chamber, dual variable combustion chamber, and afterburner. The mode selection valve is closed. The high-speed turbojet mode is suitable for flight speeds of 3.7–5.0 Mach and altitudes of 24,000–30,000 m. The combustion chamber is open with dual variable combustion chamber and afterburner, and the mode selection valve is open.

3. The method for analyzing the transient characteristics of complex loads on 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 was performed for these performance components, including: The multi-bypass variable cycle engine includes three fans: a front fan, a rear fan, and a core fan. In turbofan mode, all three fans are engaged; in turbojet mode, the front fan and the rear fan operate in a windmill state. The multi-bypass variable cycle engine has a three-rotor structure, including a high-pressure turbine, an intermediate-pressure turbine, and a low-pressure turbine. The guide vanes of the intermediate-pressure turbine and the guide vanes of the low-pressure turbine are geometrically adjustable components, and the turbine performance characteristics can be adjusted by adjusting the guide vane angle.

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

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

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

7. The method for analyzing the transient characteristics of complex loads on 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-bypass casing system, including: turbine pressure ratio, temperature, flow rate, and speed; Initial residual matrix calculation: Input the initial state variable X0 into the aerodynamic thermodynamic model F(X), and calculate the initial residual matrix E0, as shown in the following formula: E = F(X); Where E is the residual matrix, representing the deviation between the current state parameters and the target parameters of the multi-duct casing system; Jacobian matrix calculation: Based on the aerodynamic thermodynamic model F(X), the initial Jacobian matrix J is calculated using the following formula: Where: E1, E2, ..., E7 are the components of the residual matrix; X1, X2, ..., X7 are the components of the state variables; Iterative variable update: Substitute the initial state variables, initial residual matrix, and Jacobian matrix into the iterative variable update formula to correct the state variables. The calculation formula is as follows: Where: X i J represents the state variable in the i-th iteration; -1 It is the inverse of the Jacobian matrix; This is the transpose of the residual matrix of the i-th iteration; Residual matrix update: The updated state variables are input back into the aerodynamic thermodynamic model, and the update formula is: E i+1 =F(X i+1 ); Among them: E i+1 Let represent the residual matrix of the (i+1)th iteration; Jacobian matrix correction: Based on the correction values ​​of the state variables and the correction values ​​of the residual matrix, 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 B is the approximate value of the Jacobian matrix in the (i+1)th iteration; i The approximate value of the Jacobian matrix in the current iteration; Y represents the change between the state variable correction values; Z represents the change in the residual matrix; J i+1 Y is the corrected Jacobian matrix; T The transpose 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 the transient characteristics of complex loads on a multi-duct casing according to claim 5, characterized in that, The mathematical model also includes: The flight envelope is determined based on the cross-airspace and cross-speed domain design requirements of the multi-bypass variable cycle engine; Twelve extreme operating points were selected from the flight envelope for simulation and solution, where: The two working points are ground idle and ground takeoff status; The six operating points represent typical conditions under turbofan mode; The four operating points represent typical conditions under turbojet mode; The extreme operating points are simulated and solved to obtain the high-pressure rotor speed, medium-pressure rotor speed and low-pressure rotor speed; The changes in outlet temperatures of the dual variable combustion chamber and afterburner during the mode switching process of a multi-duct variable cycle engine were analyzed to determine the thermal shock damage characteristics of the multi-duct casing system under transient loads.

9. The method for analyzing the transient characteristics of complex loads on 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 operating points in turbofan mode correspond to the following operating states: Subsonic cruise mode; Supersonic cruise mode; High-speed boost 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 the transient characteristics of complex loads on a multi-duct casing according to claim 1, characterized in that, The method further includes the following steps: When switching from turbofan mode to turbojet mode, open the mode adjustment valve, close the main combustion chamber, put the core fan and high-pressure compressor in the fan-mill state, stop the high-pressure rotor and intermediate-pressure rotor, ignite the dual-variable combustion chamber and afterburner, and start the low-pressure rotor. In high-speed turbofan mode with Mach numbers of 3.5 to 3.8, the mode switching is completed in a total time of no more than 30 seconds; The transient load characteristics of the multi-duct casing system during the analysis mode switching process include: thermal shock load caused by the combustion chamber outlet airflow temperature rising from 1700K in the dual-variable combustion chamber to 2200K in the afterburner; drastic changes in the low-pressure turbine inlet temperature; and small changes in the low-pressure rotor speed.

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