Aero-engine direct thrust control method considering air inlet distortion

Abstract

The invention relates to the technical field of aero-engine control, in particular to an aero-engine direct thrust control method considering air inlet distortion, which comprises the following steps: constructing an aero-engine model with air inlet distortion; the real-time thrust of the air inlet distortion aero-engine model at each moment in the full-state test run process and related parameters related to the thrust are obtained; constructing a neural network model; and the output thrust of the aero-engine is obtained. According to the direct thrust control considering the air inlet distortion, the air inlet distortion of flight under the special working condition is considered, the control accuracy and the flight safety are improved, a traditional aero-engine control system achieves indirect thrust control by controlling parameters related to thrust, a large safety margin needs to be reserved, margin waste is reduced through direct thrust control, and the cost is reduced. The working potential of the aero-engine is fully explored, and meanwhile the engine efficiency is improved.

Description

Technical Field

[0001] This invention relates to the field of aero-engine control technology, and more specifically to a direct thrust control method for aero-engines that takes into account intake distortion. Background Technology

[0002] Intake distortion refers to the phenomenon of uneven or irregular airflow distribution entering the engine. In the field of aero-engines, intake distortion is usually caused by flight conditions (such as crosswinds, high-temperature exhaust gas intake, etc.) or intake duct design problems. Intake distortion is a common occurrence during aircraft flight, often happening under conditions such as high angle of attack, high sideslip angle, low-altitude flight, and the use of weapons in flight. Intake distortion causes uneven airflow entering the compressor, resulting in changes in the angle of attack of the airflow on the compressor blades. Severe intake distortion may cause the compressor operating point to deviate from the design condition, entering the surge boundary region. Surge is an unstable flow phenomenon that causes strong oscillations and backflows in the airflow within the compressor, drastically degrading compressor performance and potentially causing engine shutdown. Once surge occurs, engine thrust will drop significantly and cannot be stably output, rendering traditional thrust prediction methods based on stable operating conditions ineffective. To address this issue, air intake distortion tests are conducted during aero-engine design. Devices such as insert plates, simulation plates, and simulation nets are used to artificially create uneven airflow in order to test the engine's resistance to distortion.

[0003] The propulsion system is the power source of aircraft. Ensuring that the propulsion system provides stable thrust is a basic requirement of the propulsion control system. Currently, the data-based PI controller is widely used in the thrust control of aero engines. This controller can only provide a relatively conservative thrust control effect and cannot provide direct and stable thrust control under inlet distortion environment. Therefore, there is an urgent need for a more efficient thrust controller to meet the above pain points.

[0004] Traditional aero-engine thrust controllers indirectly estimate thrust by measuring parameters (such as high and low pressure rotor speeds and pressure ratios) using onboard sensors, and then perform thrust control. However, due to different operating points, operating states, performance degradation, and individual differences within the engine's flight envelope, the parameters measured by onboard sensors cannot accurately characterize the thrust. This forces designers to retain a large design margin or safety factor to ensure safe engine operation.

[0005] Currently, extensive research has been conducted abroad on thrust estimation and direct thrust control using methods such as neural network nonlinear mapping and Kalman filtering estimation. Improved turbofan engines and propfan engines have adopted direct thrust control, significantly enhancing aircraft flight performance. However, publicly available information indicates that neither domestic nor international research has considered the impact of inlet distortion on the direct thrust control of aero-engines.

[0006] Existing technologies suffer from several drawbacks: traditional direct thrust control systems fail to account for air intake distortion during flight under special operating conditions, resulting in low thrust control accuracy and potentially impacting flight safety. Furthermore, traditional aero-engine control systems achieve indirect thrust control by manipulating thrust-related parameters, requiring a substantial safety margin. Additionally, the nonlinear mathematical models of traditional aero-engine thrust estimators are limited by the nonlinearity of aero-engines.

[0007] Therefore, there is a need to provide a direct thrust control method for aero-engines that takes into account inlet distortion in order to solve the above problems. Summary of the Invention

[0008] To address the problems of existing technologies, this invention provides a direct thrust control method for aero-engines that takes into account inlet distortion, thereby resolving the existing issues.

[0009] The first aspect of this invention provides an aero-engine model based on inlet distortion according to parallel compressor theory, which adopts the following technical solution, including: Along the airflow path of the aero-engine, the following models are sequentially arranged between the inlet model and the low-pressure compressor model: distortion splitter model, parallel fan model, mixing chamber model, and inner and outer bypass splitter model. Among them, the distortion splitter model is used to split the airflow at the intake outlet into two parallel airflows; The parallel fan model includes a distortion-free sub-fan and a distortion-with sub-fan, and the inlets of the distortion-free sub-fan and the distortion-with sub-fan are connected to the corresponding outlets of the distortion splitter model; The mixing chamber model is used to combine the airflow from the undistorted sub-fan and the distorted sub-fan outlet of the parallel fan model into one stream; The inner and outer bypass splitter model is used to split the airflow at the outlet of the mixing chamber model into two streams according to the bypass ratio of the aero-engine. One stream enters the inner bypass and the other enters the outer bypass.

[0010] A second aspect of the present invention provides a method for matching components in the air path of an aero-engine model with inlet distortion based on parallel compressor theory, the method comprising: The test parameters for the aero-engine described in claim 1 are set, including: physical speed of the low-pressure rotor, physical speed of the high-pressure rotor, pressure ratio function value without distortion subfan, pressure ratio function value with distortion subfan, pressure ratio function value of the low-pressure compressor, pressure ratio function value of the high-pressure compressor, flow ratio function value of the high-pressure turbine, flow ratio function value of the low-pressure turbine, bypass ratio of the distortion splitter, bypass ratio of the inner and outer bypass splitters, airflow rate of the inlet, and total temperature at the outlet of the main combustion chamber. Based on the given parameters, residual equations are set to satisfy the balance conditions for the joint operation of component models in the aero-engine. These residual equations include: low-pressure rotor power balance equation, high-pressure rotor power balance equation, distortion-free subfan inlet flow balance equation, distortion-with subfan inlet flow balance equation, low-pressure compressor inlet flow balance equation, high-pressure compressor inlet flow balance equation, high-pressure turbine inlet flow balance equation, low-pressure turbine inlet flow balance equation, inner duct nozzle flow balance equation, outer bypass nozzle flow balance equation, and mixing chamber inlet static pressure balance equation. The target test parameters are obtained by assigning a preset increment to the initial value of each test parameter, and thermodynamic calculations are performed on the aero-engine to obtain the residual value. If the 2-norm of the residual value does not meet the error range, then a system of linear equations is constructed and solved to obtain a new set of trial parameters. This process continues until the 2-norm of the residual value meets the error range, thus completing the matching of each component in the aero-engine's air path.

[0011] A third aspect of the present invention provides a direct thrust control method for an aero-engine that takes into account inlet distortion, the method comprising: Based on the modeling mechanism of each component of a typical engine structure, an aero-engine model with intake distortion as shown in claim 1 is constructed. The real-time thrust and related parameters of the aero-engine with intake distortion were obtained at each moment during the full-state test of the aero-engine model. The related parameters are: high-pressure rotor speed, low-pressure rotor speed, bypass ratio, total pressure at the outlet of the high-pressure compressor, total pressure at the outlet of the low-pressure turbine, total temperature at the outlet of the high-pressure compressor, total temperature at the outlet of the low-pressure turbine, and engine fuel flow rate. Construct a neural network model, and use the relevant parameters at the current moment to predict the actual thrust of the engine. A fuzzy control model is constructed, and the error between the actual thrust and the preset target thrust, as well as the error change rate, are used as input parameters of the fuzzy control model. The output is the thrust control quantity of the aero-engine, and the output thrust of the aero-engine is controlled according to the thrust control quantity of the aero-engine.

[0012] A further technical solution of the present invention is that the expression for the real-time thrust of the inlet-distortion aero-engine is:

[0013] In the formula, This indicates the engine's non-installed thrust; This indicates the mass flow rate at the outlet of the internal tail nozzle; This indicates the velocity of the gas exiting the internal tail nozzle; Indicates the outlet pressure of the internal tail nozzle; Indicates the exit area of ​​the internal tail nozzle; Indicates the intrinsic import flow; This indicates the mass flow rate at the outlet of the bypass nozzle; Indicates the outlet pressure of the bypass nozzle; Indicates the outlet area of ​​the outer bypass duct nozzle; Indicates the velocity of the gas entering the intake manifold; This indicates the inlet pressure of the air intake.

[0014] A further technical solution of the present invention is that constructing a neural network model includes: Construct the initial neural network model; The relevant parameters at each time step are used as the input to the initial neural network model, and the engine thrust at each time step is used as the output of the initial neural network model. The neural network model is obtained by training the initial neural network model.

[0015] A further technical solution of the present invention is that the step of using the error between the actual thrust and the preset target thrust, as well as the rate of change of the error, as input parameters of the fuzzy control model, and outputting the thrust control quantity of the aero-engine is as follows: The fuzzy control model includes: a fuzzy controller and a PID controller; The error between the engine's actual thrust and the preset target thrust, as well as the rate of change of the error, are input into the fuzzy controller, which outputs the changes in the proportional parameter, integral parameter, and derivative parameter. Based on the changes in the proportional, integral, and derivative parameters, as well as the initial proportional, integral, and derivative parameters of the PID controller, the final proportional, integral, and derivative parameters are output. The thrust control quantity of the aero-engine is controlled based on the final proportional parameter, final integral parameter, and final derivative parameter.

[0016] A further technical solution of the present invention is that the membership function of the output parameters of the fuzzy controller is a triangular membership function.

[0017] A further technical solution of the present invention is that the membership function of the input parameters of the fuzzy controller is a Gaussian membership function.

[0018] A further technical solution of the present invention includes, after controlling the output thrust of the aero-engine with the thrust control quantity, the following: Obtain the target error and the rate of change of the target error when controlling the aero-engine with the thrust control quantity of the aero-engine. The target error and the rate of change of the target error are input into the fuzzy PID control module, which outputs the thrust control quantity of the aero-engine. The output thrust of the aero-engine is controlled according to the thrust control quantity of the aero-engine until the error between the output thrust of the aero-engine controlled by the thrust control quantity and the preset target thrust is 0. Then, the thrust when the error is 0 is taken as the output thrust of the aero-engine.

[0019] A further technical solution of the present invention is that the thrust control quantity of the aero-engine is the fuel flow rate of the aero-engine.

[0020] The beneficial effects of the present invention are: 1. Traditional direct thrust control systems do not consider intake distortion under special flight conditions, resulting in low thrust control accuracy and even affecting flight safety. The direct thrust control system for aero-engines in this invention considers intake distortion under special flight conditions, improving control accuracy and flight safety. Furthermore, traditional aero-engine control systems achieve indirect thrust control by controlling thrust-related parameters, requiring a large safety margin. Direct thrust control reduces this "margin waste," fully utilizing the aero-engine's operating potential while improving engine efficiency.

[0021] 2. Traditional aero-engine thrust estimators suffer from limited accuracy due to the nonlinearity of aero-engines, resulting in nonlinear mathematical models. Neural network-based thrust estimators, however, require no prior knowledge and directly reflect the characteristics of the engine. They offer significant advantages and superior qualities in modeling and controlling highly nonlinear aero-engine systems, such as parallel processing, fault tolerance, distributed processing, and robustness. Attached Figure Description

[0022] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0023] Figure 1 This is a schematic diagram of the structure of an aero-engine considering intake distortion in an embodiment of the present invention; Figure 2 This invention relates to an aero-engine model with intake distortion based on parallel compressor theory. Figure 3 This is a flowchart of a matching method for components in the air path of an aero-engine model with intake distortion based on parallel compressor theory, according to the present invention. Figure 4 This is a schematic diagram illustrating the principle of the fuzzy PID controller in an embodiment of the present invention; Figure 5 This is a schematic diagram of the fuzzy PID control module in an embodiment of the present invention; Figure 6 This is a schematic diagram showing the control effect of the fuzzy PID controller in an embodiment of the present invention and the comparison results with traditional PID control; Figure 7 This is a negative feedback control system in an embodiment of the present invention.

[0024] In the diagram, 1 is the intake duct inlet; 2 is the intake duct outlet; 3 is the high-pressure compressor outlet; 4 is the combustion chamber outlet; 5 is the low-pressure turbine outlet; 6 is the inner duct nozzle inlet; 8 is the inner duct nozzle nozzle; 13 is the bypass duct inlet; 16 is the bypass duct outlet; 18 is the bypass duct nozzle nozzle; 21 is the inner duct inlet; 22 is the low-pressure compressor inlet; 24 is the low-pressure compressor outlet; 25 is the high-pressure compressor inlet; 41 is the high-pressure turbine inlet; 44 is the high-pressure turbine outlet; and 45 is the low-pressure turbine inlet. Detailed Implementation

[0025] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0026] Example 1 Example 1 provides an aero-engine model with inlet distortion based on parallel compressor theory, such as... Figure 1 As shown, in order to achieve precise control of the aircraft, this embodiment 1, based on a parallel compressor model, takes a high-bypass ratio twin-shaft turbofan engine as the research object and establishes an aero-engine model with intake distortion based on parallel compressor theory. Figure 1As shown, after the airflow is drawn into the engine, it first flows through the intake duct and fan, and then is discharged through the inner and outer bypass ducts. The inner bypass duct consists of a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, and an exhaust nozzle. The fan, low-pressure compressor, and low-pressure turbine are connected by a low-pressure shaft; the high-pressure compressor and high-pressure turbine are connected by a high-pressure shaft; there is no mechanical connection between the high-pressure shaft and the low-pressure shaft, only an air passage. When the inlet airflow is distorted, a distortion splitter model, a parallel fan model, and a mixing chamber model are introduced to circumferentially partition the non-uniform flow field, and a distortion-free sub-fan is used. With a fan with aberration Perform independent calculations. Figure 1 middle, Total temperature distortion index, The total pressure distortion index is defined as the temperature and pressure being higher in the distorted region compared to the undistorted region.

[0027] Therefore, as Figure 2 As shown, the aero-engine model based on parallel compressor theory with inlet distortion includes: an inlet model, a low-pressure compressor model, a high-pressure compressor model, a combustion chamber model, a high-pressure turbine model, a low-pressure turbine model, and a tail nozzle model. Specifically, in this embodiment 1, a distortion splitter model, a parallel fan model, a mixing chamber model, and an inner / outer bypass splitter model are sequentially arranged between the inlet model and the low-pressure compressor model along the aero-engine airflow path. Among them, the distortion splitter model is used to divide the airflow at the inlet outlet into two parallel airflows; the parallel fan model includes a distortion-free sub-fan and a distortion-with sub-fan, and the inlets of the distortion-free sub-fan and the distortion-with sub-fan are connected to the corresponding outlets of the distortion splitter model; the mixing chamber model is used to combine the airflows from the distortion-free sub-fan and the distortion-with sub-fan outlets of the parallel fan model into one stream; the inner / outer bypass splitter model is used to divide the airflow at the outlet of the mixing chamber model into two streams according to the bypass ratio of the aero-engine, one stream entering the inner bypass and the other entering the outer bypass.

[0028] It should be noted that, as Figure 2 As shown, in one specific embodiment, the parallel fan model uses two fan models to describe the distortion-free sub-fan. With a fan with aberration The inlet channel of the parallel fan model coincides with the outlet of the two fan-shaped regions of the distortion splitter model, and the inlet airflow parameters are equal to the airflow parameters of the airflow channels of the two fan-shaped outlet regions of the distortion splitter model.

[0029] The section numbers of the air system used in this model are as follows: Figure 3As shown, after inputting the initialization parameters required for each engine component provided by GasTurb, the design point calculation begins. Given atmospheric conditions, the inlet airflow at inlet 1 is assigned, and the inlet airflow parameters at outlet 2 are calculated using the called inlet model. Considering inlet distortion, the outlet airflow at the inlet is used to obtain the airflow parameters for the two fan-shaped sections a1 and b1 at the fan inlet using the called distortion splitter model; the inlet airflow at fan section a1 is used to calculate the outlet airflow parameters at fan section a2 using the called compressor model, and the inlet airflow at fan section b1 is used to calculate the outlet airflow parameters at fan section b2 using the called compressor model; the two airflows at fan outlet sections a2 and b2 are used to calculate the inlet airflow parameters for the inner and outer bypass splitters using the called mixing chamber model.

[0030] The following are the specific implementation steps for the distortion shunt model and the mixing chamber model: (as follows) Figure 2 As shown, in this embodiment, it is assumed that the engine's inlet airflow, after passing through the intake duct, splits into two parallel airflows at the fan inlet, which enter the parallel fan model, namely, the fan without distortion sub-fan a and the fan with distortion sub-fan b. It is assumed that there is no exchange of matter or energy between these two airflows, and they remix into one airflow at the fan outlet section. This airflow then splits into two more airflows, which enter the bypass duct and the low-pressure compressor respectively, and then flow into their respective corresponding engine components. The distortion splitter model utilizes the bypass ratio... The concept is used to achieve flow distribution in the two-fan-shaped outlet areas, and based on the total temperature distortion index. and total pressure distortion index The degree of distortion is defined by the definition.

[0031] Define the following parameter: the flow rate at the intake outlet. Total temperature Total pressure Total enthalpy and oil-gas ratio ; Distortionless fan import flow Import total temperature, import total pressure, import total enthalpy, import oil-gas ratio ; with aberrant fan import flow Import total temperature, import total pressure, import total enthalpy, import oil-gas ratio ; Distortionless fan With a fan with aberration The ratio of the sector area (i.e., the bypass ratio of the distortion shunt model). Total temperature distortion index Total pressure distortion index .

[0032] Among them, the total temperature distortion index of the distortion splitter model , and the total pressure distortion index ; The air flow parameter distribution in the two fan-shaped regions of the distortion splitter model is that the air flow parameters at the inlets of the two fan-shaped regions in the distortion splitter model are equal to the air flow parameters at the outlet of the inlet duct model. The air flow parameters include: the flow rate at the outlet of the inlet duct model, the total temperature , the total pressure , the total enthalpy , and the fuel-air ratio ; The outlets of the two fan-shaped regions of the distortion splitter model are divided into the outlet corresponding to the undistorted sub-fan and the outlet with a distorted sub-fan . Among them: The air flow parameters at the outlet are: the outlet flow rate , the total temperature at the outlet, the total pressure at the outlet, the total enthalpy at the outlet, and the fuel-air ratio at the outlet are equal to the inlet values of the aeroengine model with inlet distortion; The outlet flow rate at the outlet, the total temperature at the outlet, the total pressure at the outlet, the total enthalpy at the outlet, and the fuel-air ratio at the

[0033] outlet are equal to the model inlet values. That is, the inlet air flow parameters of the undistorted sub-fan are: ; The inlet air flow parameters of the distorted sub-fan are: After obtaining the two sets of outlet air flow parameters of the distortion splitter model, assuming that each sub-fan operates according to the compressor characteristic line with uniform undistorted inlet air, the fan model can be called separately for independent calculation and then the mixing chamber model can be called. To describe the calculation process of the mixing chamber model, the following parameters are defined: the flow rate at the cross-section of the outlet of the undistorted sub-fan, the total temperature , the total pressure , the static pressure , the specific enthalpy , the area , and the velocity coefficient Flow rate at the cross-section of the fan outlet with distortion sub-fan Total temperature Total pressure static pressure enthalpy ,area speed coefficient Flow rate at the outlet section of the mixing chamber model Total temperature Total pressure enthalpy speed coefficient Assumptions are made for the mixing chamber model calculation: ① The two airflows flow in parallel and are uniformly distributed at their respective inlet sections; ② The static pressures of the two airflows are equal when they enter the mixing chamber, i.e. ③ Heat loss during mixing is not considered, and the outlet airflow parameters are uniformly distributed; ④ The mixing chamber is a tube with a constant cross-section.

[0034] According to the non-distorting sub-fan and fans with distortion By writing the continuity equation based on the outlet cross-sectional parameters, the outlet airflow of the mixing chamber model can be obtained as:

[0035] The specific enthalpy at the mixing chamber outlet can be obtained from the energy continuity equation. :

[0036] Seek Then, the total temperature of the gas flow at the outlet of the mixing chamber can be obtained iteratively using the bisection method. .

[0037] Next, the areas of the two inlet sectors of the mixing chamber model, i.e., the sub-fans, are calculated. Export area Kazuko Fan Export area The method is as follows: (1) Predict a sub-fan Export area Solve for the sub-fan according to the flow continuity equation. velocity coefficient of the outlet airflow ; (2) Then the fan outlet static pressure ; (3) According to the definition of bypass ratio in the above distortion shunt model, the sub-fan Export area is Similarly, we can find the sub-fan. outlet static pressure ; (4) Iterative calculation using the bisection method Until the sub-fan outlet static pressure and sub-fan The outlet static pressure is equal, that is... At this time, the sub-fan Export area Sub-fan Export area Repeat the above steps to obtain the sub-fan. And the fan outlet static pressure and .

[0038] Momentum balance equations for inlet and outlet airflow in the mixing chamber model:

[0039] Flow continuity equations for inlet and outlet airflow in the mixing chamber model:

[0040] Among them, the distortion-free sub-fan Flow rate at the outlet section Total temperature Total pressure static pressure enthalpy ,area speed coefficient ; with aberrant fan Flow rate at the outlet section Total temperature Total pressure static pressure enthalpy ,area speed coefficient Flow rate at the outlet section of the mixing chamber model Total temperature Total pressure enthalpy speed coefficient .

[0041] By simultaneously solving the momentum balance equations for the inlet and outlet airflows of the mixing chamber model and the flow continuity equations for the inlet and outlet airflows of the mixing chamber model, the total pressure at the outlet of the mixing chamber model can be obtained. and the outlet airflow velocity coefficient of the mixing chamber model Then the total temperature of the outlet airflow of the mixing chamber model can be obtained. Total pressure and traffic .

[0042] like Figure 2As shown, the airflow at the outlet of the mixing chamber model is re-divided into two streams, which enter the inner duct and the outer duct respectively. The cross-sectional airflow parameters of the inner duct inlet 21 and the outer duct inlet 13 are calculated based on the duct ratio BPR2 and the called inner and outer duct splitter models. The cross-sectional airflow parameters of the outer duct outlet 16 are calculated by the called duct loss model (since the gas will generate losses in the pipe, i.e., the total pressure loss coefficient, which is assumed to be a known value here). Then, the cross-sectional airflow parameters of the outer duct nozzle 18 are calculated by the called nozzle model. The static pressure of the outer duct nozzle 18 is equal to the static pressure of the incoming flow in the inlet duct. The inlet airflow at the duct inlet 21 is used to calculate the cross-sectional airflow parameters at the low-pressure compressor inlet 22 using the duct loss model. Then, the compressor model is used to calculate the cross-sectional airflow parameters at the low-pressure compressor outlet 24. The outlet airflow at the low-pressure compressor is used to calculate the cross-sectional airflow parameters at the high-pressure compressor inlet 25 using the duct loss model. Then, the compressor model is used to calculate the cross-sectional airflow parameters at the high-pressure compressor outlet 3. The outlet airflow at the high-pressure compressor is used to calculate the cross-sectional airflow parameters at the combustion chamber outlet 4 using the combustion chamber model. Due to the aforementioned adjustment rules, the air-fuel ratio (or fuel supply) is a given value. The outlet airflow at the combustion chamber is calculated using the duct loss model to obtain the high-pressure... The cross-sectional airflow parameters of turbine inlet 41 are calculated using the turbine model to obtain the cross-sectional airflow parameters of high-pressure turbine outlet 44. The cross-sectional airflow parameters of low-pressure turbine inlet 45 are calculated using the bypass loss model, and then the cross-sectional airflow parameters of low-pressure turbine outlet 5 are calculated using the turbine model. The high-pressure turbine inlet guide vanes and rotor blades draw cooling airflow from the cross-section of high-pressure compressor outlet 3. The cross-sectional airflow parameters of inner duct nozzle inlet 6 are calculated using the bypass loss model, and then the cross-sectional airflow parameters of inner duct nozzle nozzle 8 are calculated using the nozzle model. The static pressure of inner duct nozzle nozzle 8 is equal to the static pressure of the incoming airflow from the inlet duct.

[0043] Equations are established for the high-pressure rotor compressor torque and its output torque, the high-pressure rotor turbine torque and its output torque, and the physical speed of the high-pressure shaft and its speed. Similarly, equations are established for the low-pressure rotor compressor torque and its output torque (including the low-pressure compressor and fan), the low-pressure rotor turbine torque and its output torque, and the physical speed of the low-pressure shaft and its speed. This connects all components from beginning to end. To ensure smooth overall calculations, the following equations are adopted: , The regulatory pattern of W fGiven the fuel supply to the main combustion chamber and A8 as the throat area of ​​the tailpipe, the value of one of the trial parameters can be obtained by using the bisection method iteratively based on the given fuel supply to the main combustion chamber. The test equation is iteratively calculated based on the trial parameter until the common operating point of the engine, i.e., the steady-state operating point, is found.

[0044] Example 2 like Figure 3 As shown, this embodiment 2 provides a matching method for components in the air path of an aero-engine model with inlet distortion based on parallel compressor theory. During matching, a trial parameter x is set according to the aero-engine model with inlet distortion, and based on... Figure 2 The error verification equation y of the aero-engine model considering intake distortion is set to form a common working equation set for air path model matching:

[0045] The test condition is: when the value of x makes all y zeros, it indicates that the engine operation meets the balance condition; because the number of components in x is greater than the number of components in y. In this embodiment, a method for matching components in the air path of an aero-engine model with inlet distortion based on parallel compressor theory includes: setting trial parameters for the aero-engine with inlet distortion, the trial parameters including: the physical speed of the low-pressure rotor. Physical speed of high-voltage rotor The pressure ratio function value of the distortion-free sub-fan Pressure ratio function value with distorted sub-fan Low-pressure compressor pressure ratio function value High-pressure compressor pressure ratio function value High-pressure turbine flow ratio function value Low-pressure turbine flow ratio function value , distortion shunt bypass ratio Internal and external culvert splitter duct ratio Intake airflow Main combustion chamber outlet total temperature Based on the test parameters, residual equations are set to satisfy the balance conditions for the joint operation of component models in the aero-engine. These residual equations include: low-pressure rotor power balance equation, high-pressure rotor power balance equation, distortion-free subfan inlet flow balance equation, distortion-with subfan inlet flow balance equation, low-pressure compressor inlet flow balance equation, high-pressure compressor inlet flow balance equation, high-pressure turbine inlet flow balance equation, low-pressure turbine inlet flow balance equation, inner duct nozzle flow balance equation, outer bypass nozzle flow balance equation, and mixing chamber inlet static pressure balance equation. A preset increment is assigned to the initial value of each test parameter to obtain the target test parameter. Thermodynamic calculations are then performed on the aero-engine to obtain the residual value. If the 2-norm of the residual value does not meet the error range, a system of linear equations is constructed and solved to obtain a new set of test parameters. This process continues until the 2-norm of the residual value meets the error range, thus completing the matching of each component in the aero-engine's air path.

[0046] For example, in one specific embodiment, such as Figure 3 As shown, the physical speeds of the low-pressure rotor are sequentially increased. Physical speed of high-voltage rotor , distortion-free sub-fan Pressure ratio function value With a distorted fan Pressure ratio function value Low-pressure compressor pressure ratio function value High-pressure compressor pressure ratio function value High-pressure turbine flow ratio function value Low-pressure turbine flow ratio function value , distortion shunt bypass ratio Internal and external culvert splitter duct ratio Intake airflow An increment (Increment in this embodiment) The test parameters are set to 0.1% to 1% of the original test parameters, while keeping other test parameters unchanged from their initial values. A common method is to set the compressor physical speed n or the turbine inlet total temperature. In this embodiment, a given... Eleven engine thermodynamic calculations are performed to obtain eleven sets of residuals. If the 2-norm of these eleven sets of residuals does not meet the error limit, a linear equation system is constructed and solved to calculate a new set of trial parameter values. The above steps are repeated until the values ​​converge to the error range, thus completing the matching of the components in the aero-engine's air path.

[0047] It should be noted that, It is a system of 11 equations. There are 12 components, if The value makes All zeros indicate that the engine is operating at equilibrium, meaning that at this time... The value of is the solution to the system of common working equations. This system of common working equations... There is no specific expression; it is composed of all the formulas for varying specific heats calculated from the aerodynamic and thermodynamic processes from the atmosphere and the inlet to the exhaust nozzle. Therefore, a certain physical quantity in the given parameter x needs to be given before using the multivariate Newton method, i.e., the Newton-Raphson method, to achieve matching of components in the air path. There are 11 error equations and 12 independent variables. The equation system is not "closed," meaning it cannot be solved. Therefore, an independent variable needs to be determined in advance to make the equation system "closed," and then the equation system can be solved. A common method is to give the physical speed of the compressor. Or the total temperature at the turbine inlet In this embodiment, a given... .

[0048] After completing the component-level modeling of the aero-engine considering intake distortion using the above method, the model can be used to perform engine thermodynamic calculations to obtain the engine's performance parameters such as thrust and fuel consumption rate, as well as the airflow, total temperature, total pressure, and total enthalpy cross-sectional parameters in the air system.

[0049] Example 3 Example 3 provides a direct thrust control method for aero-engines that considers inlet distortion, including: S1. Construct an aero-engine model with intake distortion; Based on the modeling mechanism of each component of a typical engine structure, an aero-engine model with intake distortion is constructed. S2. Obtain the real-time thrust and related parameters at each moment during the full-state test of the inlet-distorted aero-engine model; For example, in one specific embodiment, the steps for obtaining the real-time thrust and related parameters of the inlet-distorted aero-engine model at each moment during full-state testing are as follows: obtaining test data of engine thrust and related engine parameters during ground testing of the aero-engine; obtaining simulation data of engine thrust and related engine parameters during simulation of the inlet-distorted aero-engine model in component-level modeling software; and using the engine thrust and related engine parameters corresponding to the simulation data and test data as the real-time thrust and related parameters of the aero-engine during operation.

[0050] For example, in one specific embodiment, the relevant parameters are: high-pressure rotor speed, low-pressure rotor speed, bypass ratio, total pressure at the outlet of the high-pressure compressor, total pressure at the outlet of the low-pressure turbine, total temperature at the outlet of the high-pressure compressor, total temperature at the outlet of the low-pressure turbine, and engine fuel flow rate. For example, in one specific embodiment, the expression for the real-time thrust of the inlet-distortion aero-engine is:

[0051] In the formula, This indicates the engine's non-installed thrust; This indicates the mass flow rate at the outlet of the internal tail nozzle; This indicates the velocity of the gas exiting the internal tail nozzle; Indicates the outlet pressure of the internal tail nozzle; Indicates the exit area of ​​the internal tail nozzle; Indicates the intrinsic import flow; This indicates the mass flow rate at the outlet of the bypass nozzle; Indicates the outlet pressure of the bypass nozzle; Indicates the outlet area of ​​the outer bypass duct nozzle; Indicates the velocity of the gas entering the intake manifold; This indicates the inlet pressure of the air intake.

[0052] S3. Construct a neural network model; Specifically, a neural network model is constructed, and the actual thrust of the engine is predicted based on the relevant parameters at the current moment.

[0053] For example, in one specific embodiment, constructing a neural network model includes: constructing an initial neural network model; using the relevant parameters corresponding to each time step as the input of the initial neural network model, and the engine thrust corresponding to each time step as the output of the initial neural network model; and training the initial neural network model to obtain a neural network model. S4. Obtain the output thrust of the aircraft engine; Specifically, a fuzzy control model is constructed, and the error between the actual thrust and the preset target thrust, as well as the rate of change of the error, are used as input parameters of the fuzzy control model. The output is the thrust control quantity of the aero-engine, and the output thrust of the aero-engine is controlled according to the thrust control quantity of the aero-engine.

[0054] For example, in one specific embodiment, the step of using the error between the actual thrust and the preset target thrust, as well as the rate of change of the error, as input parameters of the fuzzy control model to output the thrust control quantity of the aero-engine is as follows: wherein the fuzzy control model includes: a fuzzy controller and a PID controller; the error between the actual thrust of the engine and the preset target thrust, as well as the rate of change of the error, are input to the fuzzy controller, and the changes corresponding to the proportional parameter, integral parameter, and derivative parameter are output; based on the changes corresponding to the proportional parameter, integral parameter, and derivative parameter, and the initial proportional parameter, initial integral parameter, and initial derivative parameter of the PID controller, the final proportional parameter, final integral parameter, and final derivative parameter are output; the thrust control quantity of the aero-engine is controlled based on the final proportional parameter, final integral parameter, and final derivative parameter.

[0055] For example, in one specific embodiment, such as in this embodiment... Figure 4 The principle of the fuzzy PID controller is shown, and the setup is as follows. Figure 5 The structure of the fuzzy PID control module shown is that, in the Simulink toolbox of Matlab software, the Fuzzylogiccontroller module (fuzzy controller) is first built.

[0056] The specific steps for establishing a fuzzy controller are as follows: Define the universe of discourse: In MATLAB, call the fuzzy function to normalize the input and output parameters of the fuzzy controller. Define the universe of discourse of the normalized input and output parameters on the fuzzy set as {-3, -2, -1, 0, 1, 2, 3}.

[0057] Define linguistic variables: Define 7 linguistic variables to describe their magnitude: NB (negative, large), NM (negative, moderate), NS (negative, small), ZO (zero), PS (positive, small), PM (positive, moderate), PB (positive, large), and select -3, -2, -1, 0, 1, 2, 3 as the central elements of the fuzzy subsets on the 7 universes of discourse, that is, when x = -3, -2, -1, 0, 1, 2, 3, the membership function value u(x) = 1.

[0058] Determine the membership function: In the membership function editor of FuzzyToolbox, select Gaussian (gaussmf) as the membership function for the input parameters (error e between actual thrust of the aero-engine and preset target thrust) and (rate of change of error between actual thrust of the aero-engine and preset target thrust). The output parameter (the corresponding change in the proportional parameter) should also be selected. Kp, the change in the integration parameter. The changes in Ki and the corresponding differential parameters The membership function of Kd is triangle (trimf).

[0059] Setting fuzzy rules: Based on the influence of the proportional parameter Kp, integral parameter Ki, and derivative parameter Kd on the system output characteristics, we can summarize the changes in the controlled parameter (the corresponding proportional parameter) for different errors e and error change rates ec. Kp, the change in the integration parameter. The changes in Ki and the corresponding differential parameters The self-tuning requirement of Kd is as follows: When the fuzzy error E is large, regardless of the magnitude of the fuzzy error change rate EC, in order to enhance the system's fast tracking performance and avoid large overshoot in the system response, a larger proportional parameter Kp and an integral parameter Ki=0 should be selected. In addition, since the deviation between the actual thrust and the preset target thrust increases instantaneously at the beginning, it is easy to cause differential oversaturation and cause the control quantity to exceed the normal range. Therefore, a smaller differential parameter Kd should be selected, or the differential parameter Kd=0 can be selected.

[0060] When the fuzzy error E is of moderate magnitude, in order to reduce system overshoot, appropriate proportional parameters Kp, Ki, and Kd should be selected to ensure the system's response speed.

[0061] When the fuzzed error E is small, a larger proportional parameter Kp and integral parameter Ki should be chosen to ensure good dynamic performance of the system and eliminate static error. At the same time, to avoid oscillations around the given value and to enhance the system's anti-interference capability, when the fuzzed error rate EC is large, a smaller differential parameter Kd should be chosen, or the differential parameter Kd=0 should be chosen. When the fuzzed error rate EC is small, a moderately sized Kd can be chosen.

[0062] Based on the corresponding change of the above proportional parameters Kp, the change in the integration parameter. The changes in Ki and the corresponding differential parameters Kd parameter tuning rules.

[0063] In the Simulink environment of Matlab software, according to Figure 5 Design the simulation block diagram of the system, establish a Fuzzylogiccontroller module, and add quantization factors Ke, Kec, K1, K2, and K3 on this basis; where Ke and Kec are fuzzification factors, and K1, K2, and K3 are defuzzification factors; and encapsulate the fuzzy controller and PID controller together to form a fuzzy PID controller, which realizes the relationship between the changes in the proportional parameter Kp, integral parameter Ki, and derivative parameter Kd of the PID controller and the corresponding changes in the proportional parameter output by the fuzzy controller. Kp, the change in the integration parameter. The changes in Ki and the corresponding differential parameters Kd is added together to obtain the final proportional parameter, final integral parameter, and final derivative parameter of the fuzzy PID controller output. The final proportional parameter, final integral parameter, and final derivative parameter can control the thrust control quantity of the aero-engine. The thrust control quantity of the aero-engine is the fuel flow rate Wf of the aero-engine. According to the characteristic equation of the membership function, through multiple bench tests using the expert experience method, the fuzzification factors Ke=0.9, Kec=0.1, the defuzzification factors K1=3, K2=1.2, K3=0.01 are obtained. The initial proportional parameter of the PID controller is Kp0=5, the initial integral parameter is Ki0=2, the initial derivative parameter is Kd0=1.2, and the sampling period is T=0.01s.

[0064] like Figure 6 As shown, the control effect of the fuzzy PID controller is evaluated by giving a unit step input and compared with the traditional PID control. In the comparison results, the overshoot of the fuzzy PID controller in this embodiment is reduced by 15%, the number of oscillations is reduced to one, and the control effect is significantly improved.

[0065] like Figure 7 As shown, the fuzzy control model, thrust estimator (trained neural network), and engine model (aero-engine simulation module) are connected to form a negative feedback control system.

[0066] For example, in one specific embodiment, the membership function of the input parameters of the fuzzy controller is a Gaussian membership function, and the membership function of the output parameters of the fuzzy controller is a triangular membership function.

[0067] For example, in one specific embodiment, after the thrust control quantity of the aero-engine controls the output thrust of the aero-engine, the method further includes: obtaining the target error and the target error change rate between the output thrust and the preset target thrust when the aero-engine is controlled according to the thrust control quantity of the aero-engine; inputting the target error and the target error change rate into the fuzzy PID control module, outputting the thrust control quantity of the aero-engine, and controlling the output thrust of the aero-engine according to the thrust control quantity of the aero-engine until the error between the output thrust and the preset target thrust when the thrust control quantity of the aero-engine controls the aero-engine is 0, and then taking the thrust when the error is 0 as the output thrust of the aero-engine.

[0068] For example, in one specific embodiment, the thrust control amount of the aircraft engine is the fuel flow rate of the aircraft engine.

[0069] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. An aero-engine model for inlet distortion based on parallel compressor theory, characterized in that, include: Along the airflow path of the aero-engine, the following models are sequentially arranged between the inlet model and the low-pressure compressor model: distortion splitter model, parallel fan model, mixing chamber model, and inner and outer bypass splitter model. Among them, the distortion splitter model is used to split the airflow at the intake outlet into two parallel airflows; The parallel fan model includes a distortion-free sub-fan and a distortion-with sub-fan, and the inlets of the distortion-free sub-fan and the distortion-with sub-fan are connected to the corresponding outlets of the distortion splitter model; The mixing chamber model is used to combine the airflow from the undistorted sub-fan and the distorted sub-fan outlet of the parallel fan model into one stream; The inner and outer bypass splitter model is used to split the airflow at the outlet of the mixing chamber model into two streams according to the bypass ratio of the aero-engine. One stream enters the inner bypass and the other enters the outer bypass.

2. A method for matching components in the air path of an aero-engine model with inlet distortion based on parallel compressor theory, characterized in that, include: The test parameters for the aero-engine described in claim 1 are set, including: physical speed of the low-pressure rotor, physical speed of the high-pressure rotor, pressure ratio function value without distortion subfan, pressure ratio function value with distortion subfan, pressure ratio function value of the low-pressure compressor, pressure ratio function value of the high-pressure compressor, flow ratio function value of the high-pressure turbine, flow ratio function value of the low-pressure turbine, bypass ratio of the distortion splitter, bypass ratio of the inner and outer bypass splitters, airflow rate of the inlet, and total temperature at the outlet of the main combustion chamber. Based on the given parameters, residual equations are set to satisfy the balance conditions for the joint operation of component models in the aero-engine. These residual equations include: low-pressure rotor power balance equation, high-pressure rotor power balance equation, distortion-free subfan inlet flow balance equation, distortion-with subfan inlet flow balance equation, low-pressure compressor inlet flow balance equation, high-pressure compressor inlet flow balance equation, high-pressure turbine inlet flow balance equation, low-pressure turbine inlet flow balance equation, inner duct nozzle flow balance equation, outer bypass nozzle flow balance equation, and mixing chamber inlet static pressure balance equation. The target test parameters are obtained by assigning a preset increment to the initial value of each test parameter, and thermodynamic calculations are performed on the aero-engine to obtain the residual value. If the 2-norm of the residual value does not meet the error range, then a system of linear equations is constructed and solved to obtain a new set of trial parameters. This process continues until the 2-norm of the residual value meets the error range, thus completing the matching of each component in the aero-engine's air path.

3. A direct thrust control method for aero-engines considering inlet distortion, characterized in that, include: Based on the modeling mechanism of each component of a typical engine structure, an aero-engine model with intake distortion as shown in claim 1 is constructed. The real-time thrust and related parameters of the aero-engine with intake distortion were obtained at each moment during the full-state test of the aero-engine model. The related parameters are: high-pressure rotor speed, low-pressure rotor speed, bypass ratio, total pressure at the outlet of the high-pressure compressor, total pressure at the outlet of the low-pressure turbine, total temperature at the outlet of the high-pressure compressor, total temperature at the outlet of the low-pressure turbine, and engine fuel flow rate. Construct a neural network model, and use the relevant parameters at the current moment to predict the actual thrust of the engine. A fuzzy control model is constructed, and the error between the actual thrust and the preset target thrust, as well as the error change rate, are used as input parameters of the fuzzy control model. The output is the thrust control quantity of the aero-engine, and the output thrust of the aero-engine is controlled according to the thrust control quantity of the aero-engine.

4. The direct thrust control method for aero-engines considering inlet distortion according to claim 3, characterized in that, The expression for the real-time thrust of an inlet-distorted aero-engine is: In the formula, This indicates the engine's non-installed thrust; This indicates the mass flow rate at the outlet of the internal tail nozzle; This indicates the velocity of the gas exiting the internal tail nozzle; Indicates the outlet pressure of the internal tail nozzle; Indicates the exit area of ​​the internal tail nozzle; Indicates the intrinsic import flow; This indicates the mass flow rate at the outlet of the bypass nozzle; Indicates the outlet pressure of the bypass nozzle; Indicates the outlet area of ​​the outer bypass duct nozzle; Indicates the velocity of the gas entering the intake manifold; This indicates the inlet pressure of the air intake.

5. The direct thrust control method for aero-engines considering inlet distortion according to claim 3, characterized in that, Building a neural network model includes: Construct the initial neural network model; The relevant parameters at each time step are used as the input to the initial neural network model, and the engine thrust at each time step is used as the output of the initial neural network model. The neural network model is obtained by training the initial neural network model.

6. The direct thrust control method for aero-engines considering inlet distortion according to claim 3, characterized in that, The steps for using the error between the actual thrust and the preset target thrust, as well as the rate of change of the error, as input parameters to the fuzzy control model, and outputting the thrust control quantity of the aero-engine are as follows: The fuzzy control model includes: a fuzzy controller and a PID controller; The error between the engine's actual thrust and the preset target thrust, as well as the rate of change of the error, are input into the fuzzy controller, which outputs the changes in the proportional parameter, integral parameter, and derivative parameter. Based on the changes in the proportional, integral, and derivative parameters, as well as the initial proportional, integral, and derivative parameters of the PID controller, the final proportional, integral, and derivative parameters are output. The thrust control quantity of the aero-engine is controlled based on the final proportional parameter, final integral parameter, and final derivative parameter.

7. The direct thrust control method for aero-engines considering inlet distortion according to claim 6, characterized in that, The membership function of the output parameters of the fuzzy controller is a triangular membership function.

8. A direct thrust control method for an aero-engine considering inlet distortion according to claim 6, characterized in that, The membership function of the input parameters of the fuzzy controller is a Gaussian membership function.

9. A direct thrust control method for an aero-engine considering inlet distortion according to claim 3, characterized in that, After controlling the output thrust of an aero-engine, the thrust control parameters also include: Obtain the target error and the rate of change of the target error when controlling the aero-engine with the thrust control quantity of the aero-engine. The target error and the rate of change of the target error are input into the fuzzy PID control module, which outputs the thrust control quantity of the aero-engine. The output thrust of the aero-engine is controlled according to the thrust control quantity of the aero-engine until the error between the output thrust of the aero-engine controlled by the thrust control quantity and the preset target thrust is 0. Then, the thrust when the error is 0 is taken as the output thrust of the aero-engine.

10. A direct thrust control method for an aero-engine considering inlet distortion according to claim 3, characterized in that, The thrust control quantity of an aero-engine is the fuel flow rate of the aero-engine.

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