Aeroengine thrust fast response redundant control method and system considering combustion delay

Abstract

The application discloses an aero-engine thrust fast response redundant control method and system considering combustion delay, and belongs to the technical field of aero-engine control. The method steps are as follows: an open-loop transfer function L of main fuel flow and tail nozzle throat area respectively controlling thrust is established based on a desired thrust closed-loop bandwidth; a decoupling condition between the variable tail nozzle throat area and the thrust is set to eliminate the mutual influence between them; four undetermined RTM controller functions in the transfer function L are solved; the actuator transfer functions of the main fuel flow and the tail nozzle throat area are set as first-order systems for solving; system identification is carried out at an engine steady-state working point to obtain a transfer function of a difference between the main fuel flow and the tail nozzle throat area relative to a steady-state value to a difference between the thrust and the steady-state value; the transfer function L is solved, and then the RTM control is completed according to the transfer function L, so that the fast response control of the thrust is realized. The application solves the key technical problems in the prior art.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine control technology, specifically relating to a method and system for rapid response redundancy control of aero-engine thrust considering combustion delay. Background Technology

[0002] In the context of integrated flight / thrust control requirements, rapid thrust response is one of the key technologies to ensure flight safety and performance.

[0003] Several patents and academic papers currently relate to engine thrust control. For example, existing technologies disclose the use of online sliding window deep neural networks as prediction models. While this model employs a deep learning structure, which improves accuracy, it may increase the complexity of the controller, directly impacting the solution of control variables. Other disclosed methods include online optimization and multivariable control design for aero-engines. These methods, under constraints, achieve control and online optimization of multiple aero-engine variables based on thrust, speed, and other requirements. However, this approach may require significant computational resources to solve linear optimization problems.

[0004] In terms of academic research, the paper "An Overview of the Development of Aero-engine Control Systems" published in *Measurement and Control Technology* provides a comprehensive review of thrust control technology, while "Direct Thrust Control of Engines Based on Neural Network Inverse Control" published in *Propulsion Technology* proposes a new control strategy. These studies provide a theoretical basis for engine thrust control, but in practical applications, how to achieve rapid thrust response under combustion delay conditions remains an urgent problem to be solved.

[0005] Regarding the issue of sluggish thrust response caused by combustion delay, the main drawback of existing technologies is the difficulty in achieving rapid and accurate thrust regulation when considering combustion delay. This is primarily because traditional thrust control methods rely on fuel flow regulation, which is affected by combustion delay, resulting in thrust regulation lag. Furthermore, existing thrust control methods often rely on complex control algorithms or require significant computational resources, further limiting the real-time performance and flexibility of thrust control.

[0006] In aero engines, traditional manipulated variables (especially fuel flow) have a significant impact on adjusting thrust (F) due to combustion delay. n There is a hysteresis phenomenon. In contrast, geometrically variable regions, such as the nozzle throat region (A8), have rapid response capabilities but limited thrust adjustment capabilities. Therefore, to address these issues, this invention proposes an innovative scheme combining redundant control inputs, becoming a pioneering solution for achieving RTM. Summary of the Invention

[0007] The technical problem to be solved:

[0008] To overcome the shortcomings of existing technologies, this invention provides a redundant control method for rapid thrust response in aero-engines that considers combustion delay. This method supplements traditional fuel flow control by introducing the exhaust nozzle throat area as a redundant actuator; and achieves high-precision attitude and trajectory control through Rapid Thrust Modulation (RTM). This invention enables more direct and rapid thrust changes while avoiding the combustion delay problems associated with fuel flow regulation. The control method of this invention improves the system's rapid thrust adjustment capability and enhances its robustness to combustion delay and its uncertainties, thereby solving key technical problems in existing technologies while ensuring rapid thrust response.

[0009] The technical solution of this invention is: a method for rapid response redundancy control of aero-engine thrust considering combustion delay, the specific steps of which are as follows:

[0010] Based on the desired thrust closed-loop bandwidth, an open-loop transfer function L is established to control the thrust separately for the main fuel flow rate and the nozzle throat area. A decoupling control strategy is used to avoid mutual interference between the two variables. The transfer function L includes g c11 g c21 g c12 and g c22 Four undetermined RTM controller transfer functions, and actuator transfer functions g for main fuel flow and nozzle throat area. a1 and g a2 The transfer functions g1 and g2 are the difference between the main fuel flow rate and the tail nozzle throat area relative to their steady-state values ​​and the difference between the thrust and its steady-state values.

[0011] Set decoupling conditions between variable nozzle throat area and thrust to eliminate their mutual influence;

[0012] Solve for the four undetermined RTM controller functions g in the transfer function L. c11 g c21 g c12 and g c22 ;

[0013] The actuator transfer function g, which combines the main fuel flow rate and the area of ​​the tailpipe throat, is used to calculate the main fuel flow rate and the area of ​​the tailpipe throat. a1 and g a2 Solve as a first-order system;

[0014] By performing system identification at the engine's steady-state operating point, the transfer functions g1 and g2 are obtained from the difference between the main fuel flow rate and the tail nozzle throat area relative to their steady-state values ​​to the difference between the thrust and its steady-state values.

[0015] Based on the above-obtained g c11 gc21 g c12 g c22 g a1 g a2 Calculate g1 and g2, solve for the transfer function L, and then design variables based on the transfer function L to complete RTM control, thereby achieving rapid response control of thrust.

[0016] A further technical solution of the present invention is: the open-loop transfer function L is

[0017]

[0018] Where L = g c11 g a1 g1+g c21 g a2 g2+g c12 g a1 g1+g c22 g a2 g2, the integrator 1 / s ensures that the steady-state error of the step change in the thrust command is zero.

[0019] A further technical solution of the present invention is: the decoupling condition between the throat area of ​​the variable nozzle and the thrust is:

[0020] g c12 g a1 g1+g c22 g a2 g2 = 0

[0021] A further technical solution of the present invention is: the four undetermined RTM controller functions g c11 g c21 g c12 and g c22 The solution formula is as follows:

[0022]

[0023]

[0024]

[0025]

[0026] Where, ω b2 G1 represents the expected bandwidth of ΔA8. * K is the transfer function constructed by removing the delay term from g1. p It is an undetermined parameter with design freedom; K i The calculation formula is as follows:

[0027]

[0028] Where, ω b1 For ΔF n Expected bandwidth, g a1 (0) is the initial value of the transfer function of the main fuel flow actuator, and g1(0) is the thrust-to-main fuel flow rate W. f The initial value of the response function.

[0029] A further technical solution of the present invention is: the actuator transfer function g of the main fuel flow rate and the tailpipe throat area. a1 and g a2 The first-order system is:

[0030]

[0031] Where a and b are positive numbers, e -τs This is to delay combustion.

[0032] A further technical solution of the present invention is that the adjustment time of the actuator is one order of magnitude less than the adjustment time of the controller.

[0033] A further technical solution of the present invention is: g in the transfer function L c21 g a2 g2 is the control transfer function for the throat area of ​​the tailpipe, and its expression is as follows:

[0034]

[0035] Where, when K i =ω b When, g is satisfied c21 (0)g a2 (0) < ∞, which guarantees the neutrality of A8.

[0036] A fast-response redundant control system for aero-engine thrust that takes combustion delay into account includes two main controllers and two bypass controllers.

[0037] The inputs to the two main control systems are reference values ​​ΔF representing the difference between the thrust and its steady-state value. n，ref The RTM controller transfer function g is used respectively. c11 g c21 To achieve control transmission, g c11 via transfer function g a1 The difference ΔF between the output thrust of g1 and its steady-state value n g c21 via transfer function g a2 The difference ΔF between the output thrust of g2 and its steady-state value n ;

[0038] The two bypass controllers use the RTM controller transfer function g c12 and g c22 Implement control transmission and cancel it out to 0 in the control process.

[0039] A further technical solution of the present invention is: the g c11 Main fuel flow W f The instruction ΔW is the difference between its steady-state value and its actual value. f，cmd The signal is sent to the flow actuator controller, which then outputs the main fuel flow W from the actuator. f The instruction ΔW is the difference between its steady-state value and its actual value. f The thrust difference ΔF relative to its steady-state value is obtained by sending the data to the flow transfer controller. n .

[0040] A further technical solution of the present invention is: the g c21 The instruction A is the difference between the throat area A8 of the tailpipe and its steady-state value. 8，cmd The signal is sent to the area actuation controller, which then sends the instruction ΔA8, representing the difference between the nozzle throat area A8 output by the actuator and its steady-state value, to the area transfer controller to obtain the thrust difference ΔF relative to its steady-state value. n .

[0041] Beneficial effects

[0042] The beneficial effects of this invention are as follows: This invention utilizes the main fuel flow rate W f The two variables, A8 and the nozzle throat area, achieve redundant thrust control. A8 compared to W... f It can change thrust more directly and quickly, but its adjustment range is limited. Therefore, after a rapid thrust adjustment, the A8 needs to be gradually returned to the "middle position" to ensure sufficient adjustment margin in both directions of increasing and decreasing thrust. To achieve this, by using the conventional fuel flow W... f An A8 circuit is added to the existing circuit, and a decoupling control strategy is designed. This allows the system to avoid the effects of combustion delay in the fuel circuit while continuously achieving rapid thrust adjustment, ensuring the independence between A8 and thrust.

[0043] Simulation verification was conducted based on a certain type of engine. The simulation results show that, considering the objective condition of combustion delay, the traditional operating variables of aero-engines, especially fuel flow rate, are affected by the combustion delay when adjusting F... n There is a hysteresis phenomenon. When adjusting A8, we should also consider how to achieve the ideal thrust bandwidth under the condition of combustion delay, that is, by adding g. c22 g c12 The bypass controller design takes into full account the existence of combustion delay.

[0044] Furthermore, to ensure neutrality, g should be satisfied. c21 (0)g a2 (0) < ∞, meaning it does not contain an integrator. Substituting the above parameters, we construct g. c21 g a2 g2 is shown below:

[0045]

[0046] It can be observed that when K i =ω b When, g is satisfied c21 (0)g a2 (0) < ∞, at this time A8 will return to its initial position or set value after stabilization, the steady-state error is 0, and it has returned to its "middle position", which ensures that it always has a certain usable range in both the increasing and decreasing directions; W f The thrust control meets the requirement of a large variable range. The overall main fuel quantity W... f By using two variables, A8 and the throat area of ​​the tail nozzle, redundant thrust control is achieved, which not only increases the thrust control bandwidth but also provides stronger robustness to combustion delay and its uncertainties.

[0047] This invention achieves high-precision attitude and trajectory adjustments through rapid thrust modulation, successfully improving the response speed and accuracy of aero-engine thrust. Simulation results show that this redundant control method effectively improves thrust bandwidth while ensuring flight parameter stability. Furthermore, when only A8 is adjusted, the rapid thrust modulation method effectively reduces thrust fluctuations; even with only a 5% step change in A8, the maximum thrust fluctuation is only 0.26%. Figure 7 As shown. This means that during the design and optimization of aero engines, the cross-sectional area of ​​the exhaust nozzle can be adjusted more freely to meet various performance requirements. Overall, this innovation enhances the robustness of the system and provides a safer and more stable thrust control scheme for aero engines. Attached Figure Description

[0048] Figure 1 Control structure diagram;

[0049] Figure 2 : Equivalent two-input two-output system diagram;

[0050] Figure 3 Schematic diagram of thrust steady-state characteristics (W) f,1 >W f,2 >W f,3 );

[0051] Figure 4 Steady-state characteristics of engine thrust;

[0052] Figure 5 : Thrust command step response curve;

[0053] Figure 6 Low-pressure compressor characteristic diagram;

[0054] Figure 7 Analysis of the coupling effect between A8 and thrust. Detailed Implementation

[0055] The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the invention, and should not be construed as limiting the invention.

[0056] To address the problems existing in the prior art, this invention provides a fast-response redundancy control method for aero-engine thrust considering combustion delay, comprising the following steps:

[0057] Step 1: Achieving ideal thrust bandwidth despite combustion delay. Using main fuel flow rate W... f The two variables, A8 and the area of ​​the nozzle throat, affect the thrust F respectively. n Redundancy control is required, therefore A8 and W must be considered. f The collaboration between the two variables is improved by introducing a decoupling control strategy to independently regulate them, avoiding their mutual influence and thus improving F. n The bandwidth.

[0058] One path is ΔW f To ΔF n Control, the two branches are ΔA8 to ΔF n The control involves two valves (operated variables) with different dynamic effects, coordinating inputs across different frequency bands. The symbols g1 and g2 represent the inputs from ΔW... f and ΔA8 to ΔF n The transfer functions, where the former includes a combustion delay. Their effects on thrust are assumed to be additive. The symbol g a1 and g a2 W f And the transfer function of the A8 actuator. Additionally, the symbol g... c11 g c21 g c12 and g c22 This is the transfer function of the undetermined RTM block. The subscript ref represents the reference value given by the pilot or flight control system, the subscript cmd represents the command from the controller to the actuator, and Δ represents the difference between the physical quantity and its steady-state value. The specific control structure diagram of RTM is shown below. Figure 1 .

[0059] The design begins with a reverse-engineering controller design process that specifies the open-loop transfer function L as follows:

[0060]

[0061] Where L = g c11 g a1 g1+g c21 g a2 g2+g c12 g a1 g1+g c22 g a2 g2, the integrator 1 / s ensures that the steady-state error of the thrust command step change is zero. This is achieved by selecting the appropriate closed-loop transfer function G. y From ΔF n,ref To ΔF n for

[0062]

[0063] Where ω b1 This is the desired thrust closed-loop bandwidth. It should be noted that there is no delay term in the thrust response because it is eliminated through the secondary loop described later. This characteristic clearly demonstrates the advantage of RTM over conventional fuel control; however, in equation (1), the delay term for g... c11 and g c21 The combination of constraints imposes a non-singular control effect, requiring the introduction of additional requirements to eliminate variability.

[0064] Step 2: Reset A8 to its optimal position to ensure continuous modulation capability and decouple A8 from thrust during flight. The above system can be equivalently described as follows: Figure 2 In the two-input, two-output system shown, to avoid the mutual interference between A8 and thrust, special consideration is given to the control module g in these two branches. c12 and g c22 Regarding the impact on A8, it is recommended that A... 8,ref Changes in F n The impact should be minimized to allow for free adjustment of A8. This feature allows for smooth transitions from different positions during flight, such as from RTM mode to normal flight operation mode, while ensuring that the aircraft's state is not affected. Decoupling is achieved when the following equation holds true.

[0065] g c12 g a1 g1+g c22 g a2 g2=0 (4)

[0066] In this case, g c22 A physical interpretation was obtained, as from ΔA 8,refThe transfer function to ΔA8 describes how changes in the reference position of A8 affect its actual position. At this point, A8 and thrust can be controlled independently, avoiding changes to A8 from affecting the aircraft's state. This allows for independent control of thrust and optimization of overall flight performance. This feature has potential application value in scenarios such as performance optimization and fuel consumption reduction.

[0067] Calculate g c11 g c21 g c22 and g c12 At that time, considering that the transfer function g1 includes a combustion delay, it is necessary to... c22 It contains a corresponding delay term to satisfy the equation. For simplicity, g is... c22 Defined as a first-order time-delay system, the transfer functions are solved according to the following formulas.

[0068]

[0069]

[0070]

[0071]

[0072] Among them g1 * It is the transfer function constructed by removing the delay term from g1. From formula (6), we can obtain that ΔA8 will affect ΔA. 8,ref A delayed response is made as the cost of eliminating the delay in the inference response. And with ΔF n,ref In comparison, ΔA 8,ref The changes do not occur too frequently, making the delay negligible.

[0073] Step 3: The controller module g was obtained in Step 2. c11 g c21 g c22 and g c12 Because of g a1 and g a2 W respectively f The transfer functions of the A8 actuator and the A8 actuator can be simulated as a first-order system, specifically in the form of...

[0074]

[0075] Where a and b are positive numbers, e -τs This is for combustion delay. The specific value depends on the established g. c11 and g c21 The adjustment time is determined by the control mechanism (generally, the adjustment time of the actuator is an order of magnitude less than that of the controller).

[0076] The two transfer functions, g1 and g2, can be obtained through system identification at the engine's steady-state operating point. Since obtaining the two transfer functions g1 and g2 through system identification is not a property of this patent, readers can consult relevant literature to understand the specific method.

[0077] Step 4: Based on g obtained in steps 2 and 3 c11 g c21 g c22 and g c12 g a1 and g a2 The transfer function L is solved by g1 and g2, and then the valve position control RTM controller is designed based on the transfer function L to achieve rapid response control of thrust.

[0078] The specific execution steps can also be summarized as shown in Table 1, which illustrates the process of synthesizing the RTM mode.

[0079]

[0080]

[0081] Figure 3 This demonstrates the steady-state thrust characteristics of a hybrid turbofan engine. (By...) Figure 3 It can be seen that the thrust F n The thrust varies with the nozzle area A8, and there exists a maximum thrust line. The figure also distinguishes between minimum phase and non-minimum phase regions, indicating that the engine's dynamic characteristics differ under different operating conditions. The RTM operating point in the figure represents the engine's real-time modeling (RTM) operating point, while the normal operating point represents the engine's operating point under normal conditions.

[0082] By increasing W f (Fuel flow) or reducing A8 can achieve higher thrust, but these two operations have different effects on engine parameters such as low-pressure rotor speed (NL), turbine inlet temperature (T41), and high-pressure compressor surge margin (SMHPC). Figure 3 The simulation also simulated the paths to enter and exit RTM, which are related to the engine response control strategy or the transition to different operating modes.

[0083] The above technical solution will be further explained below with reference to specific embodiments:

[0084] The engine design parameters used in this simulation are based on the F118 engine: ground conditions, turbine inlet temperature T. 41=1700K, overall pressure ratio 35, bypass ratio 0.8, equivalent airflow 125kg / s, and a convergent nozzle configuration, the design thrust is calculated to be 73.87kN. Based on these cyclic parameters, a component-level aerodynamic-thermodynamic model is established in the MATLAB / Simulink environment. Under near-ground landing conditions (Ma = 0.27, altitude 0km), W... f By varying the A8 within 35% to 100% of the design value and within 100% to 180% of the minimum area, the steady-state characteristics of the engine parameters are obtained. Figure 3 As shown. The B-2 bomber is equipped with four F118 engines. Based on thrust-drag balance and aircraft aerodynamic coefficients, calculations show that under landing conditions, each engine needs to provide approximately 20 kN of thrust. A working point of 35 kN thrust at 55% fuel flow is selected as the operating point. This operating point is to the right of the maximum thrust line and falls within the algorithm's applicable region. Figure 4 As indicated by the triangle marking. Parameter overruns mainly occur during engine acceleration; observe the increase in W at the operating point. f The trend of reducing the influence of A8 on the parameters shows that both have the same direction of action on thrust, verifying the use of main fuel quantity W in this invention. f The thrust control based on the nozzle throat area A8 is entirely feasible and can solve the problems caused by traditional thrust control.

[0085] Taking a typical twin-shaft, low-bypass turbofan engine as an example, neglecting the cavity effect and the gas temperature change caused by the heat absorption and release of the metallic material, the thrust response function g2 has the following second-order proper fractional form:

[0086]

[0087] A simulation comparison experiment was conducted using a certain type of engine. Simulation comparison graphs of RTM and KQ methods were plotted for three combustion delays: τ = 0s, τ = 0.1s, and τ = 0.2s. Figure 5 The thrust command step response curve is given; a working point of 55% fuel flow and 35kN thrust is selected. This point is to the right of the maximum thrust line and is located in the applicable region of the algorithm. The thrust pair W is identified at this working point. f The response functions g1 and g2 of A8 are as follows:

[0088]

[0089]

[0090] The thrust command is increased by 5% at 0.5s. Figure 5 The following is given in W f The KQ method and RTM control effect with A8 as the control variable, W fThe curves showing the changes in A8 and thrust are shown in the figure. Observing the parameters, we can see that: First, as the combustion delay increases, the KQ system oscillates, but the RTM control system is not significantly affected. Based on the estimation of a first-order system, the closed-loop bandwidth design target of 10 rad / s is achieved, demonstrating the robustness of the RTM method to combustion delay and its uncertainties. Second, A8 returns to its baseline value after the thrust stabilizes, achieving the design objective of return to center and providing room for variation for the next rapid thrust response.

[0091] Based on the above system, plot the characteristic trajectory of the low-pressure compressor under the RTM control system, as follows: Figure 6 As shown in the figure, it can be seen that the engine's operating trajectory is always within the compressor characteristic diagram. This indicates that the RTM control system effectively maintains the stable operation of the engine within the low-pressure compressor characteristic range, further verifying the system's reliability and conforming to objective physical laws.

[0092] Modify the RTM control loop in the MATLAB / Simulink environment and add A to the input. 8,ref The command perturbation. At 1s and 6s, the A8 command step is reduced and increased by 5% respectively, and the original thrust command step is removed. W f The curves showing the changes in thrust and low-pressure rotor speed are as follows: Figure 7 As shown in the image.

[0093] At 1s and 6s, the A8 command step decreases and increases by 5%, respectively. Under traditional control methods, the thrust would change significantly at these times. However, the rapid thrust modulation method avoids large thrust fluctuations, allowing W to... f As A8 changes, F... n The maximum fluctuation was 0.26%. Simulation results show that changing A... 8,ref For F n No impact. This means that when designing and optimizing aero engines, the cross-sectional area of ​​the exhaust nozzle can be adjusted more flexibly to meet more performance requirements, such as reducing noise and improving fuel efficiency, without worrying about adverse effects on thrust. This provides greater flexibility for the online optimization of aero engines.

[0094] RTM also used W f The thrust is controlled by A8, which returns to its reference value after the thrust stabilizes, satisfying the design requirement of return to neutrality and providing adjustment space for thrust control. Compared with the KQ algorithm, RTM control is more robust to combustion delay and uncertainty.

[0095] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. A fast-response redundant control method for aero-engine thrust considering combustion delay, characterized in that... The specific steps are as follows: Based on the desired thrust closed-loop bandwidth, open-loop transfer functions are established to control the thrust separately for the main fuel flow rate and the nozzle throat area. L Furthermore, a decoupling control strategy is employed to avoid mutual interference between the two variables; the transfer function L Includes four pending RTM controller transfer functions g c11 , g c21 , g c12 and g c22 Actuator transfer function of main fuel flow rate and tailpipe throat area g a1 and g a2 The transfer function of the difference between the main fuel flow rate and the nozzle throat area relative to their steady-state values ​​to the difference between the thrust and its steady-state values. g 1 and g 2; The open-loop transfer function L for in, The integrator 1 / s ensures that the steady-state error of the thrust command step change is zero; The four undetermined RTM controller functions g c11 , g c21 , g c12 and g c22 The solution formula is as follows: in, ω b2 Indicates Δ A 8 expected bandwidth, It is by g 1. The transfer function constructed by removing the delay term. To delay combustion, K p These are undetermined parameters with design freedom. K i The calculation formula is as follows: in, ω b1 For Δ F n Expected bandwidth The initial value of the transfer function of the main fuel flow actuator. For thrust to main fuel flow W f The initial value of the response function; Set decoupling conditions between the variable nozzle throat area and thrust to eliminate their mutual influence; Solving the transfer function L Four undetermined RTM controller functions g c11 , g c21 , g c12 and g c22 ; The actuator transfer function that combines the main fuel flow rate and the area of ​​the tailpipe throat. g a1 and g a2 Solve as a first-order system; By performing system identification at the engine's steady-state operating point, the transfer function is obtained from the difference between the main fuel flow rate and the nozzle throat area relative to their steady-state values ​​to the difference between the thrust and their steady-state values. g 1 and g 2; Based on the above... g c11 , g c21 , g c12 , g c22 , g a1 , g a2 , g 1. g 2. Solve for the transfer function L Then, based on the transfer function L The design variables complete the RTM control, thereby achieving rapid response control of thrust.

2. The fast-response redundancy control method for aero-engine thrust considering combustion delay according to claim 1, characterized in that: The decoupling condition between the throat area of ​​the variable nozzle and the thrust is: 。 3. The fast-response redundancy control method for aero-engine thrust considering combustion delay according to claim 1, characterized in that: The actuator transfer function of the main fuel flow rate and the tailpipe throat area g a1 and g a2 The first-order system is: in, If it is a positive number, This is to delay combustion.

4. The fast-response redundancy control method for aero-engine thrust considering combustion delay according to claim 3, characterized in that: The settling time of the actuator is an order of magnitude shorter than that of the controller.

5. The fast-response redundancy control method for aero-engine thrust considering combustion delay according to claim 3, characterized in that: The transfer function L In g c21 g a2 g 2 represents the control transfer function for the throat area of ​​the tailpipe, and its expression is as follows: Among them, when When, g is satisfied c21 (0)g a2 (0) < ∞, which guarantees that A 8 is neutral.

6. A fast-response redundant control system for aero-engine thrust considering combustion delay, characterized in that: It includes two main path controllers and two bypass controllers; used to implement the fast response redundancy control method for aero-engine thrust considering combustion delay as described in any one of claims 1-5; The inputs to the two main path controllers are reference values ​​Δ, representing the difference between the thrust and its steady-state value. F n,ref The RTM controller transfer function is used respectively. g c11 , g c21 To achieve control transmission, g c11 sequentially through the transfer function g a1 , g 1. The difference Δ between the output thrust and its steady-state value F n ; g c21 sequentially through the transfer function g a2 , g 2. The difference Δ between the output thrust and its steady-state value F n ; The two bypass controllers use RTM controller transfer functions. g c12 and g c22 Implement control transmission and cancel it out to 0 in the control process.

7. The aero-engine thrust fast response redundancy control system considering combustion delay according to claim 6, characterized in that: The g c11 Main fuel flow W f The instruction Δ is the difference between its steady-state value and its steady-state value. W f,cmd The main fuel flow rate output by the actuator is then sent to the flow actuator controller. W f The instruction Δ is the difference between its steady-state value and its steady-state value. W f The thrust difference Δ is sent to the flow transfer controller to obtain the thrust relative to its steady-state value. F n .

8. The aero-engine thrust fast response redundancy control system considering combustion delay according to claim 7, characterized in that: The g c21 The area of ​​the tail nozzle throat A 8. Instructions for the difference relative to its steady-state value A 8,cmd The data is sent to the area actuator controller, which then outputs the nozzle throat area data from the actuator. A 8. The instruction Δ for the difference between its steady-state value and its steady-state value A 8. The thrust is sent to the area transfer controller to obtain the difference Δ between the thrust and its steady-state value. F n .

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