Aircraft engine limit protection control method based on multi-mode command regulator

Abstract

The present invention discloses an aircraft engine limiting protection control method based on a multi-mode command regulator, including designing a main controller and establishing an inner loop control system prediction model; designing a conventional command regulator; designing a fast response command regulator and a limit value regulator; designing a mode switching logic; and digital simulation of a multi-mode command regulator limiting protection control method. The present invention establishes an inner loop control system prediction model, selects the command regulator working mode through a maneuvering flight signal and a limit parameter steady-state model, determines whether a conventional command regulator, a fast response command regulator, and a limit value regulator are working, and uses an online rolling optimization strategy to solve the command regulation problem with input and output constraints, thereby realizing multi-parameter fast tracking and multi-parameter limiting protection of aircraft engines in multiple modes, improving the response speed of aircraft engine thrust under maneuvering conditions, and realizing thrust increase control in emergency situations.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft engine control, and in particular relates to an aircraft engine limiting protection control method based on a multi-mode instruction regulator. Background Art

[0002] Aircraft engine limit protection control is an important means in aircraft engine control systems to ensure safe and reliable engine operation. Aircraft engine limit protection control methods use different limit protection control strategies in the engine steady-state control and transitional state control processes to keep the key variables of the engine within the allowable limit range and prevent the engine from entering abnormal operating conditions.

[0003] When the aircraft is in non-maneuverable flight state, the aircraft engine limit protection control method maintains the key variables within the allowable limit range and provides the minimum response speed that meets the design requirements. As advanced fighters have put forward higher requirements for super maneuverability and short takeoff and landing, in aircraft maneuvering flight or emergency state, the aircraft engine limit protection control method should not only maintain the key variables within the allowable limit range, but also, more importantly, the key variables should work as close to their limit boundaries as possible to give full play to the rapidity of the thrust or power response of advanced aircraft engines. In order to meet the aircraft engine control requirements under different aircraft states, it is very important to select different aircraft engine limit protection control methods to realize aircraft engine thrust control and limit protection functions according to different flight states.

[0004] At present, the aircraft engine limit protection control often adopts the low-selection-high-selection switching control method to realize the switching between the main controller and the limiter, ensuring that the various constraint outputs (such as temperature, pressure, surge margin, speed, etc.) of the engine do not exceed the limit during the control task. This method has problems such as high conservatism and affecting the performance of the main controller when applied to multivariable control, which greatly reduces the thrust response speed of the aircraft engine. To achieve different aircraft engine control requirements, the main controller and the limit controller gain need to be redesigned. The aircraft engine limit protection control method based on the multi-mode command regulator connects the command regulator in series with the main controller, without redesigning the main controller. By changing the optimization target and the limit matrix online, the limit protection control requirements in multiple modes can be achieved, which is easy to be applied to the aircraft engine limit protection control problem with multiple control requirements. Summary of the invention

[0005] Purpose of the invention: In order to improve the thrust control and limit protection capabilities of aircraft engines in multiple working modes, the present invention proposes an aircraft engine limit protection control method based on a multi-mode command regulator. On the basis of a main controller with good stability, robustness and tracking control response speed, a multi-mode command regulator is designed, and the mode switching logic is used to change the optimization target and limit matrix online, and the virtual command is optimized in real time, so as to realize multi-parameter fast tracking and multi-parameter limit protection in multiple working modes of aircraft engines.

[0006] To achieve the above object, an aircraft engine limit protection control method based on a multi-mode command regulator comprises the following steps:

[0007] Step 1: Design the main controller and establish the inner loop control system prediction model;

[0008] Step 2: Design a conventional instruction regulator;

[0009] Step 3: Design a fast response command regulator and a limit value regulator;

[0010] Step 4: Design mode switching logic;

[0011] Step 5: Digital simulation of the multi-mode command regulator limit protection control method.

[0012] Furthermore, the specific steps in step 1 are as follows:

[0013] Step 1-1: Under the current flight altitude, Mach number and working state, establish the discrete state space model of the aircraft engine and normalize the model, as shown in formula (1):

[0014]

[0015] Where A, B, C, and D are dimensionally appropriate matrices, x(k), u(k), and y(k) are the state variables, input variables, and output variables of the aircraft engine, and the subscript e is the steady-state value of the corresponding variable. The state variables of the aircraft engine include the fan speed n L , compressor speed n H , the input includes the fuel flow W f , tail nozzle throat area A8, output includes fan speed, compressor speed, engine pressure ratio EPR, low-pressure turbine outlet temperature T6, compressor outlet pressure P3;

[0016] Step 1-2: Based on the discrete state space model of the aircraft engine, an augmented LQ tracking controller is designed as the main controller of the inner loop system. The controller form is shown in formula (2):

[0017] Δu(k)=K1Δx(k)+K2(y m(k)-r(k)) (2)

[0018] Where K1 is the dimensionally appropriate proportional coefficient matrix, K2 is the dimensionally appropriate integral coefficient matrix, Δu(k) is the first-order difference of the aircraft engine input, Δx(k) is the first-order difference of the aircraft engine state, and y m (k) is the controlled output of the aircraft engine, including the compressor speed and the engine pressure ratio, and r(k) is the aircraft engine tracking instruction;

[0019] Step 1-3: Under the same flight altitude, Mach number and working state, the discrete state space model of the inner loop control system is established by using the decoupled step response and fitting method, as shown in formula (3):

[0020]

[0021] Where A cl , B cl , C cl , D cl is a suitable-dimensional matrix, x cl , v(k) and y cl are the state quantity, input quantity and output quantity of the inner loop control system, and the subscript e is the steady-state value of the corresponding quantity. The state quantity of the inner loop control system of the aircraft engine includes the fan and compressor speed and their first-order differences, fuel flow rate, tail nozzle throat area, and the input quantity includes the compressor speed virtual instruction v nL and the engine pressure ratio virtual command v EPR The output includes compressor speed, low-pressure turbine outlet temperature, compressor outlet pressure, fuel flow, tail nozzle throat area, and fuel flow first-order difference ΔW f , the first-order difference of the tail nozzle throat area ΔA8.

[0022] Furthermore, the specific steps in step 2 are as follows:

[0023] Step 2-1: Consider the aircraft engine input and output constraints

[0024]

[0025] The subscripts max and min represent the maximum and minimum limits respectively. Based on the discrete state space model of the inner loop control system, the maximum output allowable set is solved by the incremental optimization algorithm.

[0026] O j ={(x cl (0),v(0))|y cl (k;(x cl (0),v(0)))∈Y,k=0,1,…,j} (5)

[0027] Where Y = {ycl (k)|Sy cl (k)≤s}, Finally determine the prediction time domain length j;

[0028] Step 2-2: Take the quadratic function of the difference between the compressor speed and the engine pressure ratio command and its virtual command as the optimization target Q is a positive definite matrix. According to the maximum output allowed set, the discrete state space model of the inner loop control system of the aircraft engine is transformed into the static and dynamic constraint matrices h, H x , H v Initial setting

[0029]

[0030] The superscript number indicates the number of recursions. The recursive algorithm establishes static and dynamic constraint matrices as the constraints of the optimization problem, as shown in formula (7).

[0031]

[0032] When t=j, the recursive algorithm ends.

[0033] Step 2-3: Under the current state of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the optimization problem established by the optimization objectives and constraints of step 2-2

[0034]

[0035] The virtual instructions v(k) of the compressor speed and engine pressure ratio at the current moment are obtained. If the optimization problem (8) fails to be solved, the virtual instructions at the previous moment are used as the virtual instructions at the current moment.

[0036] Furthermore, the specific steps in step 3 are as follows:

[0037] Step 3-1: Under the current state of the aircraft engine inner loop control system, when the fast instruction regulator is selected by the design mode switching logic in step 4, the quadratic programming algorithm is used to solve the optimization problem established by the following optimization objectives and constraints:

[0038]

[0039] Obtain the virtual command v(k) of the compressor speed and engine pressure ratio at the current moment. If the optimization problem (9) fails to be solved, use the virtual command at the previous moment as the virtual command at the current moment.

[0040] Step 3-2: After step 4, the design mode switching logic selects the fast instruction regulator and the limit value regulator to work together, then adjust the limit value

[0041] s(k)=s(k-1)+ρ(s max -s(k-1)) (10)

[0042] Where s max Regenerate the matrix h based on s(k) by limiting the maximum and minimum values ​​of the vector j+1 And solve the optimization problem (9) to obtain the virtual command v(k) of the compressor speed and engine pressure ratio at the current moment.

[0043] Furthermore, the specific steps in step 4 are as follows:

[0044] Step 4-1: If the maneuvering flight signal α=0, the multi-mode command regulator enters the normal mode, and only the normal command regulator works;

[0045] Step 4-2: If the maneuvering flight signal α=1, further determine the limit output maximum and minimum values ​​T 6max , T 6min , P 3max , P 3min The output steady-state value T of the limit on the reference command is tracked 6re , P 3re If the relationship T 6re ≥T 6max , T 6re ≤T 6min , P 3re ≥P 3max , P 3re ≤P 3min If none of them holds true, the multi-mode command regulator enters the maneuvering mode, and only the rapid response command regulator works. If one of the terms in the relationship holds true, the multi-mode command regulator enters the thrust increase mode, and the rapid response command regulator and the limit regulator work.

[0046] Furthermore, the specific steps in step 5 are as follows:

[0047] Step 5-1: Given the maximum, minimum, maximum, minimum and simulation duration k of the aircraft engine input and output limits, and initialize the simulation time t=0;

[0048] Step 5-2: Obtain the current flight altitude, Mach number, working status and maneuvering flight signal, and give appropriate multi-dimensional step tracking instructions;

[0049] Step 5-3: Obtain the prediction model and prediction time domain length of the inner loop control system at the current moment, make a judgment based on the mode switching logic in step 4, select the current working mode of the multi-mode command regulator, establish the optimization target, determine the restriction constraints, and establish the static and dynamic constraint matrices based on step 2 or 3;

[0050] Step 5-4: Obtain the current state quantity of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the instruction optimization problem, and obtain the virtual instruction at the current moment;

[0051] Step 5-5: The main controller uses the virtual instruction as a tracking instruction to calculate the fuel flow rate and tail nozzle throat area at the next moment, and calculates the state quantity and output quantity of the aircraft engine at the next moment;

[0052] Step 5-6: t=t+1, repeat steps 5-2 to 5-5 until t=k.

[0053] Beneficial effects: The invention provides an aircraft engine limit protection control method based on a multi-mode command regulator, which has the following technical effects compared with the prior art by using the above technical solution:

[0054] (1) The present invention adopts a multi-mode command regulator, which can adaptively select a suitable command regulator and limit value regulator to perform limit protection control according to the flight state and thrust demand, thereby realizing multi-parameter rapid tracking and multi-parameter limit protection in multiple working modes of the aircraft engine;

[0055] (2) The fast response command regulator used by the multi-mode command regulator of the present invention in the maneuvering mode can limit the parameters as close to the limit values ​​as possible while ensuring that multiple parameters do not exceed the limit, so that the inner loop control system has a faster dynamic response speed;

[0056] (3) The multi-mode command regulator of the present invention uses a fast response command regulator and a limit value regulator to work together in the thrust increase mode, thereby ensuring that multiple parameters do not exceed the limit and realizing engine thrust increase control in emergency situations. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Figure 1 It is the structural diagram of the aircraft engine limiting protection control system based on the multi-mode command regulator.

[0058] Figure 2 It is a digital simulation flow chart of aircraft engine limit protection control based on multi-mode command regulator.

[0059] Figure 3 It is a compressor speed response diagram comparing the switching method, normal mode, and maneuvering mode.

[0060] Figure 4 It is an engine pressure ratio response diagram comparing the switching method, normal mode, and maneuvering mode.

[0061] Figure 5 It is the low-pressure turbine outlet temperature response diagram comparing the switching method, conventional mode, and maneuvering mode.

[0062] Figure 6It is a compressor outlet pressure response diagram comparing the switching method, normal mode, and maneuvering mode.

[0063] Figure 7 This is a fuel flow chart comparing the switching method, normal mode, and maneuvering mode.

[0064] Figure 8 It is a diagram of the tail nozzle throat area comparing the switching method, conventional mode, and maneuvering mode.

[0065] Fig. 9 It is the first-order difference diagram of fuel flow rate for the comparison of switching method, conventional mode and maneuvering mode.

[0066] Fig.10 It is the first-order difference diagram of the tail nozzle throat area comparing the switching method, conventional mode, and maneuvering mode.

[0067] Fig.11 This is the speed response diagram of the compressor in thrust increase mode.

[0068] Fig.12 This is the thrust increase mode engine pressure ratio response diagram.

[0069] Fig.13 This is the low-pressure turbine outlet temperature response diagram in thrust increase mode.

[0070] Fig.14 This is the pressure response diagram of the compressor outlet in the thrust increase mode. DETAILED DESCRIPTION

[0071] The technical solution of the present invention is further described in detail below 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 creative work are within the scope of protection of the present invention.

[0072] In order to improve the thrust control and limit protection capabilities of aircraft engines in multiple working modes, the present invention proposes an aircraft engine limit protection control method based on a multi-mode command regulator. Based on a main controller with good stability, robustness and tracking control response speed, a multi-mode command regulator is designed, and the mode switching logic is used to change the optimization target and limit matrix online, and the virtual command is optimized in real time, so as to realize multi-parameter fast tracking and multi-parameter limit protection in multiple working modes of aircraft engines.

[0073] Figure 1The structure diagram of the aircraft engine limit protection control system based on the multi-mode command regulator used by the method of the present invention is shown. The specific implementation of the present invention takes the design of an aircraft engine limit protection control method based on a multi-mode command regulator of a certain type of turbofan engine as an example. The design of the aircraft engine limit protection control method based on the multi-mode command regulator includes the following steps:

[0074] Step 1: Design the main controller and establish the inner loop control system prediction model;

[0075] Step 2: Design a conventional instruction regulator;

[0076] Step 3: Design a fast response command regulator and a limit value regulator;

[0077] Step 4: Design mode switching logic;

[0078] Step 5: Digital simulation of the multi-mode command regulator limit protection control method.

[0079] Taking the working state of 100% compressor speed and 100% engine pressure ratio of a certain type of turbofan engine at the ground point as an example, further, the specific steps in step 1 are as follows:

[0080] Step 1-1: Under the conditions of flight altitude of 0 km, Mach number of 0, compressor speed of 100%, and engine pressure ratio of 100%, a discrete state space model of the aircraft engine is established and the model is normalized, as shown in formula (1):

[0081]

[0082] Where A cl , B cl , C cl , D cl is a suitable-dimensional matrix, A, B, C, and D are suitable-dimensional matrices, x(k), u(k), and y(k) are the state variables, input variables, and output variables of the aircraft engine, and the subscript e is the steady-state value of the corresponding variable. The state variables of the aircraft engine include the fan speed n L , compressor speed n H , the input includes the fuel flow W f , tail nozzle throat area A8, output includes fan speed, compressor speed, engine pressure ratio EPR, low-pressure turbine outlet temperature T6, compressor outlet pressure P3, the coefficient matrix is ​​shown in formula (2)

[0083]

[0084]

[0085] Step 1-2: Design an augmented LQ tracking controller as the main controller of the inner loop system based on the discrete state space model of the aircraft engine. The controller form is shown in formula (3):

[0086] Δu(k)=K1Δx(k)+K2(y m (k)-r(k)) (3)

[0087] Where K1 is the dimensionally appropriate proportional coefficient matrix, K2 is the dimensionally appropriate integral coefficient matrix, Δu(k) is the first-order difference of the aircraft engine input, Δx(k) is the first-order difference of the aircraft engine state, and y m (k) is the controlled output of the aircraft engine, including the compressor speed and the engine pressure ratio, r(k) is the aircraft engine tracking instruction, and the controller parameters are shown in formula (4):

[0088]

[0089] Step 1-3: When the flight altitude is 0 km, the Mach number is 0, the compressor speed is 100%, and the engine pressure ratio is 100%, the discrete state space model of the inner loop control system is established by using the decoupled step response and fitting method, as shown in formula (5):

[0090]

[0091] Where x cl , v(k) and y cl are the state, input and output of the inner control system of the aircraft engine. The subscript e is the steady-state value of the corresponding quantity. The state of the inner control system of the aircraft engine includes the fan and compressor speeds and their first-order differences, fuel flow, and tail nozzle throat area. The input includes the compressor speed virtual instruction and the engine pressure ratio virtual command v EPR The output includes compressor speed, low-pressure turbine outlet temperature, compressor outlet pressure, fuel flow, tail nozzle throat area, and fuel flow first-order difference ΔW f , the first-order difference of the tail nozzle throat area ΔA8, and the coefficient matrix is ​​shown in formula (6)

[0092]

[0093] Furthermore, the specific steps in step 2 are as follows:

[0094] Step 2-1: Consider the aircraft engine input and output constraints

[0095]

[0096] The subscripts max and min represent the maximum and minimum limits, respectively. Based on the discrete state space model of the inner-loop control system of an aircraft engine, the maximum output allowable set is solved by an incremental optimization algorithm.

[0097] O j ={(x cl (0),v(0))|y cl (k;(x cl (0),v(0)))∈Y,k=0,1,…,j} (8)

[0098] Where Y = {y cl (k)|Sy cl (k)≤s}, Finally, the prediction time domain length j=64 is determined;

[0099] Step 2-2: Take the quadratic function of the difference between the compressor speed and the engine pressure ratio command and its virtual command as the optimization target Q is a positive definite matrix. According to the maximum output allowed set, the discrete state space model of the inner loop control system of the aircraft engine is transformed into the static and dynamic constraint matrices h, H x , H v Initial setting

[0100]

[0101] The superscript number indicates the number of recursions. The recursive algorithm establishes static and dynamic constraint matrices as the constraints of the optimization problem, as shown in formula (10).

[0102]

[0103] Step 2-3: Under the current state of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the optimization problem established by the optimization objectives and constraints of step 2-2

[0104]

[0105] The virtual instructions v(k) of the compressor speed and engine pressure ratio at the current moment are obtained. If the optimization problem (11) fails to be solved, the virtual instructions at the previous moment are used as the virtual instructions at the current moment.

[0106] Furthermore, the specific steps in step 3 are as follows:

[0107] Step 3-1: Under the current state of the aircraft engine inner loop control system, when the fast instruction regulator is selected by the design mode switching logic in step 4, the quadratic programming algorithm is used to solve the optimization problem established by the following optimization objectives and constraints:

[0108]

[0109] Obtain the virtual command v(k) of the compressor speed and engine pressure ratio at the current moment. If the optimization problem (12) fails to be solved, use the virtual command at the previous moment as the virtual command at the current moment.

[0110] Step 3-2: After step 4, the mode switching logic is designed to select the fast command regulator and the limit value regulator to work together. In this example, in order to meet the thrust increase mode switching logic, the output constraint is

[0111]

[0112] Adjustment limit value

[0113] s(k)=s(k-1)+ρ(s max -s(k-1)) (14)

[0114] Where s max The vector consisting of the maximum and minimum limits, ρ = 0.9. Regenerate the matrix h according to s(k) j+1 And solve the optimization problem (12) to obtain the compressor speed and engine pressure ratio virtual instruction v(k) at the current moment.

[0115] Furthermore, the specific steps in step 4 are as follows:

[0116] Step 4-1: If the maneuvering flight signal α=0, the multi-mode command regulator enters the normal mode, and only the normal command regulator works;

[0117] Step 4-2: If the maneuvering flight signal α=1, further determine the limit output maximum value T 6max , P 3max The output steady-state value T of the limit on the reference command is tracked 6re , P 3re If the relationship T 6re ≥T 6max , P 3re ≥P 3max If none of them holds true, the multi-mode command regulator enters the maneuvering mode, and only the rapid response command regulator works. If one of the terms in the relationship holds true, the multi-mode command regulator enters the thrust increase mode, and the rapid response command regulator and the limit regulator work.

[0118] Figure 2 The figure shows a digital simulation flow chart of aircraft engine limit protection control based on a multi-mode command regulator. Furthermore, the specific steps in step 5 are as follows:

[0119] Step 5-1: Given the maximum, minimum, maximum, minimum and simulation duration k of the aircraft engine input and output limits, and initialize the simulation time t=0;

[0120] Step 5-2: Obtain the current flight altitude, Mach number, working status and maneuvering flight signal, and give appropriate multi-dimensional step tracking instructions

[0121] Step 5-3: Obtain the prediction model and prediction time domain length of the inner loop control system at the current moment, make a judgment based on the mode switching logic in step 4, select the current working mode of the multi-mode command regulator, establish the optimization target, determine the restriction constraints, and establish the static and dynamic constraint matrices based on step 2 or 3;

[0122] Step 5-4: Obtain the current state quantity of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the instruction optimization problem, and obtain the virtual instruction at the current moment;

[0123] Step 5-5: The main controller uses the virtual instruction as a tracking instruction to calculate the fuel flow rate and tail nozzle throat area at the next moment, and calculates the state quantity and output quantity of the aircraft engine at the next moment;

[0124] Step 5-6: t=t+1, repeat steps 5-2 to 5-5 until t=k.

[0125] Digital simulation of aircraft engine limit protection control based on multi-mode command regulator Figures 3 to 14 As shown in the figure, it can be seen that the aircraft engine limit protection control method based on the multi-mode command regulator can realize multi-parameter fast tracking and multi-parameter limit protection in multiple working modes of the aircraft engine. In the conventional mode, the conventional command regulator realizes multi-parameter tracking and multi-parameter limit protection of the aircraft engine. In the maneuvering mode, the fast response command regulator obtains a superior fast response capability through more aggressive virtual commands. Its speed tracking root mean square error is reduced by 1.7% compared with the conventional command regulator, and 26.1% compared with the switching method. The pressure ratio tracking root mean square error is reduced by 18.2% compared with the conventional command regulator, and 52.2% compared with the switching method. The fast command regulator is the most active in adjusting the engine input and its rate, making the engine output closest to the limit value, giving full play to the engine performance, and reducing the conservatism of the limit protection control. In the thrust increase mode, through the adjustment of the limit regulator's multiple limit values ​​and the optimization of the command by the fast response command regulator, the engine tracks the upper steady-state limited reference command to achieve engine thrust increase control.

[0126] The above shows and describes the basic principles, main features and advantages of the present invention. It should be understood by those skilled in the art that the present invention is not limited to the above embodiments. The above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, which fall within the scope of the present invention. The scope of protection of the present invention is defined by the attached claims and their equivalents.

Claims

1. An aircraft engine limit protection control method based on a multi-mode command regulator, characterized in that: The following steps are involved: Step 1: Design the main controller and establish the inner loop control system prediction model; Step 2: Design a conventional instruction regulator; Step 3: Design a fast response command regulator and a limit value regulator; Step 4: Design mode switching logic; Step 5: Digital simulation of the multi-mode command regulator limit protection control method; The specific steps of designing the main controller and establishing the inner loop control system prediction model in step 1 are as follows: Step 1-1: Under the current flight altitude, Mach number and working state, establish the discrete state space model of the aircraft engine and normalize the model, as shown in formula (1): Where A, B, C, and D are dimensionally appropriate matrices, x(k), u(k), and y(k) are the state variables, input variables, and output variables of the aircraft engine, and the subscript e is the steady-state value of the corresponding variable. The state variables of the aircraft engine include the fan speed n L , compressor speed n H , the input includes the fuel flow W f , tail nozzle throat area A8, output includes fan speed, compressor speed, engine pressure ratio EPR, low-pressure turbine outlet temperature T6, compressor outlet pressure P3; Step 1-2: Design an augmented LQ tracking controller as the main controller of the inner loop system based on the discrete state space model of the aircraft engine. The controller form is shown in formula (2): Δu(k)=K1Δx(k)+K2(y m (k)-r(k)) (2) Where K1 is the dimensionally appropriate proportional coefficient matrix, K2 is the dimensionally appropriate integral coefficient matrix, Δu(k) is the first-order difference of the aircraft engine input, Δx(k) is the first-order difference of the aircraft engine state, and y m (k) is the controlled output of the aircraft engine, including the compressor speed and the engine pressure ratio, and r(k) is the aircraft engine tracking instruction; Step 1-3: Under the same flight altitude, Mach number and working state, the discrete state space model of the inner loop control system is established by using the decoupled step response and fitting method, as shown in formula (3): Where A cl , B cl , C cl , D cl is a suitable-dimensional matrix, x cl , v(k) and y cl are the state quantity, input quantity and output quantity of the inner loop control system, and the subscript e is the steady-state value of the corresponding quantity. The state quantity of the inner loop control system of the aircraft engine includes the fan and compressor speed and their first-order differences, fuel flow rate, tail nozzle throat area, and the input quantity includes the compressor speed virtual instruction v nL and the engine pressure ratio virtual command v EPR The output includes compressor speed, low-pressure turbine outlet temperature, compressor outlet pressure, fuel flow, tail nozzle throat area, and fuel flow first-order difference ΔW f , the first-order difference of the tail nozzle throat area ΔA8; The specific steps of designing the conventional instruction regulator in step 2 are as follows: Step 2-1: Consider the aircraft engine input and output constraints The subscripts max and min represent the maximum and minimum limits respectively. Based on the discrete state space model of the inner loop control system, the maximum output allowable set is solved by the incremental optimization algorithm. O j ={(x cl (0),v(0))|y cl (k;(x cl (0),v(0)))∈Y,k=0,1,…,j} (5) Where Y = {y cl (k)|Sy cl (k)≤s}, Finally determine the prediction time domain length j; Step 2-2: Take the quadratic function of the difference between the compressor speed and the engine pressure ratio command and its virtual command as the optimization target Q is a positive definite matrix. According to the maximum output allowed set, the discrete state space model of the inner loop control system of the aircraft engine is transformed into the static and dynamic constraint matrices h, H x , H v Initial setting The superscript number indicates the number of recursions. The static and dynamic constraint matrices are established by the recursive algorithm as the constraints of the optimization problem, as shown in formula (7): When t=j, the recursive algorithm ends; Step 2-3: Under the current state of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the optimization problem established by the optimization objectives and constraints of step 2-2 Obtain the virtual command v(k) of the compressor speed and engine pressure ratio at the current moment. If the optimization problem (8) fails to be solved, use the virtual command at the previous moment as the virtual command at the current moment. The specific steps of designing the fast response command regulator and the limit value regulator in step 3 are as follows: Step 3-1: Under the current state of the aircraft engine inner loop control system, when the fast instruction regulator is selected by the design mode switching logic in step 4, the quadratic programming algorithm is used to solve the optimization problem established by the following optimization objectives and constraints: Obtain the virtual command v(k) of the compressor speed and engine pressure ratio at the current moment. If the optimization problem (9) fails to be solved, use the virtual command at the previous moment as the virtual command at the current moment. Step 3-2: After step 4, the design mode switching logic selects the fast instruction regulator and the limit value regulator to work together, then adjust the limit value s(k)=s(k-1)+ρ(s max -s(k-1)) (10) Where s max Regenerate the matrix h based on s(k) by limiting the maximum and minimum values ​​of the vector j+1 And solve the optimization problem (9) to obtain the current compressor speed and engine pressure ratio virtual instruction v(k); The specific steps of the design mode switching logic in step 4 are as follows: Step 4-1: If the maneuvering flight signal α=0, the multi-mode command regulator enters the normal mode, and only the normal command regulator works; Step 4-2: If the maneuvering flight signal α=1, further determine the limit output maximum and minimum values ​​T 6max , T 6min , P 3max , P 3min The output steady-state value T of the limit on the reference command is tracked 6re , P 3re If the relationship T 6re ≥T 6max , T 6re ≤T 6min , P 3re ≥P 3max , P 3re ≤P 3min If neither of the above conditions holds true, the multi-mode command regulator enters the maneuvering mode, and only the quick response command regulator works. If one of the conditions holds true, the multi-mode command regulator enters the thrust increase mode, and the quick response command regulator and the limit regulator work.

2. The aircraft engine limit protection control method based on the multi-mode command regulator according to claim 1 is characterized in that: The specific steps of digital simulation of the multi-mode command regulator limiting protection control method in step 5 are as follows: Step 5-1: Given the maximum, minimum, maximum, minimum and simulation duration k of the aircraft engine input and output limits, and initialize the simulation time t=0; Step 5-2: Obtain the current flight altitude, Mach number, working status and maneuvering flight signal, and give appropriate multi-dimensional step tracking instructions; Step 5-3: Obtain the prediction model and prediction time domain length of the inner loop control system at the current moment, make a judgment based on the mode switching logic in step 4, select the current working mode of the multi-mode command regulator, establish the optimization target, determine the restriction constraints, and establish the static and dynamic constraint matrices based on step 2 or 3; Step 5-4: Obtain the current state quantity of the aircraft engine inner loop control system, use the quadratic programming algorithm to solve the instruction optimization problem, and obtain the virtual instruction at the current moment; Step 5-5: The main controller uses the virtual instruction as a tracking instruction to calculate the fuel flow rate and tail nozzle throat area at the next moment, and calculates the state quantity and output quantity of the aircraft engine at the next moment; Step 5-6: t=t+1, repeat steps 5-2 to 5-5 until t=k.

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