Method and system for rapid response control of aircraft engine thrust based on redundant actuators

Abstract

The present invention relates to a method and system for rapid thrust response control of an aero-engine based on redundant actuators, and belongs to the field of aero-engines. The method comprises the following steps: constructing a transfer function for controlling thrust based on the main fuel flow rate and the tail nozzle throat area; calculating the transfer functions of the three modules for controlling thrust based on the main fuel flow rate and the tail nozzle throat area; solving the transfer functions for controlling thrust based on the main fuel flow rate and the tail nozzle throat area, and designing a valve position control (VPC) controller based on the transfer function L, thereby achieving rapid thrust response control. While increasing the thrust control bandwidth, the present invention avoids the problems of turbine inlet temperature exceeding the limit and compressor surge caused by the high bandwidth under traditional single-variable control.

Description

Technical Field

[0001] The present invention belongs to the field of aero-engines, and in particular relates to a method and system for rapid thrust response control of an aero-engine based on redundant actuators. Background Art

[0002] In low-speed flight scenarios, incoming air pressure is low, resulting in low control efficiency for control surfaces, which increases the pilot's operational burden. However, at low altitudes and low speeds, the propulsion system has ample available thrust, allowing it to act as a "virtual control surface" to participate in aircraft attitude control, effectively improving flight quality and trajectory control accuracy. Typical application scenarios include carrier-based aircraft landings and vertical / short takeoff and landing aircraft. In these scenarios, thrust needs to be able to quickly respond to changes in the pilot's throttle lever. However, traditional engine speed control modes have a low bandwidth and cannot meet control requirements, necessitating the exploration of thrust-based control modes.

[0003] In related fields, similar research has been conducted on thrust control, such as using nonlinear predictive model control methods to directly control engine thrust. This method can effectively improve the real-time performance of engine model predictive control. Although this method directly controls engine thrust, it does not consider achieving rapid thrust response.

[0004] Prior art methods for direct thrust control of aircraft engines employ nonlinear model predictive control (MPC). This method improves model accuracy and reduces sensitivity to training data noise. However, this invention targets response speed over a wide range, rather than a narrow range.

[0005] Existing technologies utilize a high-speed, slow-running control mode to achieve rapid response control of aircraft engines. This method improves engine acceleration performance and shortens the time it takes to accelerate from slow to maximum speed. However, this method is designed for the specific high-speed, slow-running state and does not adequately meet the requirements for rapid response control in other engine states. Summary of the Invention

[0006] Technical issues to be solved:

[0007] To address the shortcomings of existing technologies and mitigate the risks of overheating and surge caused by rapid engine thrust response in traditional fuel regulation schemes, the present invention provides a rapid thrust response control method for aircraft engines based on redundant actuators. This method improves the thrust control bandwidth while avoiding turbine inlet temperature overshoot and compressor surge associated with the high bandwidth of traditional single-variable control.

[0008] The technical solution of the present invention is: a method for rapid thrust response control of an aero-engine based on redundant actuators, the specific steps of which are as follows:

[0009] Step 1: Construct the main fuel flow W f and the tail nozzle throat area A8 controls the thrust F n The transfer function L is:

[0010] L = g c1 g a1 g1+g c2 g a2 g2

[0011] Among them, g c1 and g c2 is the controller module to be determined, g1 is ΔW f to ΔF n The transfer function, g2 is ΔA8 to ΔF n The transfer function, g a1 and g a2 W f and the transfer function of the A8 actuator, Δ represents the difference of the physical quantity relative to its steady-state value;

[0012] Step 2: Calculate g c1 and g c2 , the specific formula is as follows

[0013]

[0014]

[0015] Among them, K p It is an undetermined parameter with design freedom, and its size determines W f The relative contribution of K and A8 to thrust response; i The calculation formula is as follows:

[0016]

[0017] Among them, ω b is the desired thrust bandwidth, g a1 (0) is the initial value of the fuel actuator transfer function, g1(0) is the thrust to fuel flow W f The initial value of the response function;

[0018] Step 3: Controller module g obtained based on step 2 c1 and g c2 , W f and the transfer function g of the A8 actuator a1 and g a2 Set it as a first-order system for simulation, the specific form is as follows:

[0019]

[0020] Among them, ω a To adjust the countdown of time;

[0021] Step 4: Obtain g1 and g2 by performing system identification at the engine steady-state operating point;

[0022] Step 5: Based on the g obtained in steps 2-4 c1 and g c2 、g a1 and g a2 , g1 and g2 to solve the transfer function L, and then design the valve position control VPC controller according to the transfer function L, so as to achieve rapid response control of the thrust.

[0023] A further technical solution of the present invention is: in step 1, ΔW f to ΔF n The control includes three control modules: the undetermined controller module, whose transfer function is unknown, set as g c1 ; The execution control module that receives the control command of the pending controller module has a transfer function of g a1 , for the engine W f Control; the control module of main fuel flow to thrust has a transfer function of g1.

[0024] A further technical solution of the present invention is: in step 1, the control from ΔA8 to ΔF includes three control modules: a pending controller module, whose transfer function is unknown and is set to g c2 ; The execution control module that receives the control command of the pending controller module has a transfer function of g a2 , controls the engine's A8; the control module of the tail nozzle throat area to thrust, its transfer function is g2.

[0025] A further technical solution of the present invention is: in step 2, K p The larger the W f The higher the degree of participation in the transition process, the range of its value is [0, +∞). The reason why it cannot be negative is to avoid the reduction of fuel flow during the thrust increase process.

[0026] A further technical solution of the present invention is: in step 3, the larger the bandwidth, the faster the response time, and the value range is 10 to 100 rad / s.

[0027] A further technical solution of the present invention is: in step 3, the adjustment time of the actuator is one order of magnitude shorter than the adjustment time of the controller.

[0028] A rapid response control system for aircraft engine thrust based on redundant actuators, including ΔW f to ΔF n The control route and ΔA8 to ΔFn The VPC controller is constructed in parallel with the control route, where ΔW f to ΔF n The control route includes a first pending controller module, a first execution control module, and a control module for main fuel flow to thrust; the control route from ΔA8 to ΔF includes a second pending controller module, a second execution control module, and a control module for tail nozzle throat area to thrust; a desired thrust control instruction is sent to the VPC controller, and the VPC control system can track the instruction to achieve a fast response control effect.

[0029] A further technical solution of the present invention is: the transfer function of the VPC controller is,

[0030] L = g c1 g a1 g1+g c2 g a2 g2.

[0031] An electronic device comprises at least one processor and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor to enable the at least one processor to execute the redundant actuator-based aircraft engine thrust rapid response control method.

[0032] A computer-readable digital storage medium stores computer instructions, which are used to enable a processor to implement the redundant actuator-based aero-engine thrust rapid response control method when executed.

[0033] Beneficial effects

[0034] The beneficial effects of the present invention are as follows: in the present invention, the control of the thrust by the tail nozzle throat area satisfies the characteristics of a small variable range and fast response; and after the thrust is quickly adjusted, it can return to the reference value after reaching stability. The principle is:

[0035]

[0036] Therefore, the final steady-state change value of A8 is 0, and it returns to its "middle position", ensuring that there is always a certain range of use in both the increasing and decreasing directions; the control of the thrust by the main fuel flow meets the characteristics of a large variable range. f The two variables of the thrust and the tail nozzle throat area A8 are used to perform redundant control on the thrust, thereby achieving the effect of improving the thrust control bandwidth while avoiding the excessive temperature before the turbine and compressor surge caused by the high bandwidth under traditional single-variable control.

[0037] Simulations conducted on a B2 bomber and an F118 engine demonstrated that, under near-ground landing conditions, a thrust bandwidth of 10 rad / s can be achieved without exceeding all parameters. The redundant control approach reduces fuel overshoot, minimizes fluctuations in turbine inlet temperature, and reduces the high-pressure compressor surge margin, effectively mitigating the risk of overheating and surge. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Figure 1 :Block diagram of thrust rapid response control system;

[0039] Figure 2 :Thrust steady-state characteristics diagram (W f,1 >W f,2 >W f,3 );

[0040] Figure 3 : Engine thrust steady-state characteristic diagram;

[0041] Figure 4 : Thrust command step response curve (incremental form);

[0042] Figure 5 : Low-pressure compressor surge margin. DETAILED DESCRIPTION

[0043] The embodiments described below with reference to the accompanying drawings are exemplary and are intended to explain the present invention, but should not be construed as limiting the present invention.

[0044] Based on the problems of overheating and surge caused by the rapid response of engine thrust in traditional fuel regulation schemes in the prior art, in order to reduce the risk of overheating and surge, the present invention provides a method for rapid response thrust control of an aircraft engine based on redundant actuators. When controlling thrust, the present invention uses two variables, the main fuel quantity Wf and the tail nozzle throat area A8, to perform redundant control of thrust. Compared with Wf, A8 can change thrust more directly, but A8 has a small variable range and is less capable of significantly changing thrust. Therefore, it is hoped that A8 will return to its "middle position" after participating in the rapid thrust adjustment, thereby ensuring that there is always a certain available range in both the increasing and decreasing directions. Therefore, the present invention adopts valve position control (VPC) to improve the thrust control bandwidth while avoiding the problems of turbine inlet temperature overrun and compressor surge caused by the high bandwidth under traditional single variable control.

[0045] To achieve the above technical effects, the technical solution of this embodiment includes the following steps:

[0046] Step 1: Construct the control function L of thrust Fn;

[0047] Using the main fuel flow Wf The two variables of thrust F and tail nozzle throat area A8 n Redundant control is performed. That is, dual-input-single-output redundant control scheme: (W f , A8)→F n , and establish a parallel VPC (valve position control) architecture.

[0048] One branch is ΔW f to ΔF n The other branch is ΔA8 to ΔF n Among them, ΔW f to ΔF n The control consists of three control modules: Design a control module (the transfer function is unknown, set as g c1 ), through which the control module transmits the control command to the next control module (actuator, with the transfer function g a1 ), thus affecting the engine's W f Control is performed, and finally it is the control module for the main fuel flow to thrust (the transfer function is g1).

[0049] ΔA8 to ΔF n The control consists of three control modules: Design a control module (the transfer function is unknown, set as g c2 ), through which the control module transmits the control command to the next control module (actuator, with the transfer function g a2 ) controls the engine's A8, and finally the tail nozzle throat area's thrust control module (the transfer function is g2).

[0050] The entire parallel architecture can be viewed as a control, and its transfer function can be expressed as

[0051] L = g c1 g a1 g1+g c2 g a2 g2 (1)

[0052] The specific control structure diagram is as follows Figure 1 . Among them g c11 and g c21 The controller module is pending. Figure 1 In the equation, g1 is ΔW f to ΔF n The transfer function, g2 is ΔA8 to ΔF n The transfer function, g a1 and g a2 W fThe transfer function of the A8 actuator is shown in Figure 2. The subscript ref represents the reference value given by the pilot or the flight control system, the subscript cmd represents the instruction given by the controller to the actuator, and Δ represents the difference of the physical quantity relative to its steady-state value.

[0053] Step 2: Calculate g c1 and g c2 (The transfer function of the controller module is to be determined), and g is calculated according to the following two formulas respectively c1 and g c2

[0054]

[0055]

[0056] Among them, K p It is an undetermined parameter with design freedom, and its size determines W f The relative contribution of A8 to thrust response - K p The larger the W f The higher the degree of participation in the transition process, the range of its value is [0, +∞). The reason why it cannot be negative is to avoid the reduction of fuel flow during the thrust increase process. i According to the formula

[0057]

[0058] To calculate; where ω b is the desired thrust bandwidth, g a1 (0) is the initial value of the fuel actuator transfer function, g1(0) is the thrust to fuel flow W f The initial value of the response function.

[0059] Step 3: Obtain the controller module g in step 2 c1 and g c2 , because g a1 and g a2 W f The transfer functions of the A8 actuator can be simulated as first-order systems.

[0060]

[0061] Among them, ω a It is the inverse of the adjustment time. The larger the bandwidth, the faster the response time. The general value range is 10 to 100 rad / s (the adjustment time of the actuator is generally one order of magnitude smaller than the adjustment time of the controller).

[0062] Step 4: The two transfer functions g1 and g2 can be obtained by performing system identification at the engine's steady-state operating point. (Using different identification methods will only affect the accuracy of the transfer functions; significant differences in the transfer functions obtained using different methods will not occur.) Since obtaining the two transfer functions g1 and g2 through system identification is not covered by this patent, readers are encouraged to consult relevant literature for specific methods.

[0063] Step 5: After completing these three steps, the VPC controller can be designed based on the thrust rapid response control transfer function L, thereby achieving rapid thrust response control. As shown in the block diagram of the thrust rapid response control system, a desired thrust control command (typically a thrust step control command) is sent to the VPC controller, and the VPC control system can track this command, achieving a rapid response control effect.

[0064] The above technical solution is verified through specific examples below.

[0065] The engine design parameters used in the simulation refer to the F118 engine: ground conditions, turbine inlet temperature T 41 =1700K, total pressure ratio 35, bypass ratio 0.8, converted air flow 125kg / s, and nozzle adopting convergent form. According to the above parameters, the design point thrust is about 70kN. Based on the above cycle parameters, a component-level aerodynamic thermodynamic model is established in the MATLAB / Simulink environment. Under the condition of near-ground landing (Ma=0.27, altitude 0km), W f The steady-state characteristics of the engine parameters are obtained by varying between 30% and 100% of the design value and between 100% and 200% of the minimum area of ​​A8. Figure 3 As shown. The B2 bomber is equipped with four F118 engines. According to the thrust-drag balance and the aircraft aerodynamic coefficient, it can be calculated that under landing conditions, a single engine needs to provide about 30kN thrust. 50% fuel flow and 30kN thrust are selected as the working point. This working point is to the right of the maximum thrust line and is located in the algorithm applicable area, as shown in Figure 3 The parameter overrun mainly occurs during the engine acceleration process, and the increase of W at the operating point is observed. f And the trend of reducing the influence of A8 on the parameters, we can see that the direction of the thrust is the same, but for N L 、T 41 and SM HPC The effect is the opposite (only SM HPC Basically not affected by the changes of A8), so during the thrust increase process, W f The effects of A8 cancel each other out, thus reducing parameter fluctuations, which verifies that the present invention adopts the main fuel oil quantity W fThe thrust control based on the tail nozzle throat area A8 is completely feasible and can solve the problems caused by traditional thrust control.

[0066] Use a certain model of engine to do simulation experiments, Figure 3 The engine thrust characteristic diagram is given; 40% fuel flow and 20kN thrust are selected as the working point. This point is on the right side of the maximum thrust line and is located in the applicable area of ​​the algorithm. At this working point, the thrust dependence W is obtained based on the least squares system identification. f And the response functions g1 and g2 of A8 are as follows:

[0067]

[0068]

[0069] Depend on Figure 3 It can be seen that the increase of W at the observation working point f From the trend of reducing the influence of A8 on the parameters, it can be seen that the directions of their effects on thrust are the same.

[0070] According to the proposed method, the VPC controller is designed for the above working point. The design conditions are: the desired thrust bandwidth ω b =10rad / s, W f The A8 actuator is a first-order inertia link, with a bandwidth of ω a1 =ω a2 =20rad / s, K p =1, we can get g c1 、g c2 、g a1 and g a2 for:

[0071]

[0072]

[0073]

[0074] In the MATLAB / Simulink environment, a VPC control system was built based on the component-level aerodynamic thermodynamic model. The simulation conditions were Ma = 0.27, altitude 0km, thrust command step 5%, W f 、F n , the incremental change curve of A8 is as follows Figure 4 As shown in . Figure 4 It can be seen that:

[0075] The thrust under VPC control reaches 95% of the set value in 0.3s, with no steady-state error or obvious overshoot, and has significant first-order link characteristics. The estimated bandwidth based on the first-order link is 9.7rad / s, which basically meets the expected design target of 10rad / s. The error mainly comes from the nonlinearity of the component-level model.

[0076] VPC also uses W f The A8 controls thrust, and after stabilizing, the A8 returns to its baseline value, meeting the design requirement of return-to-neutral and providing additional adjustment margin for thrust control. Compared to IMC, which uses a single fuel flow rate, VPC utilizes the A8 for regulation, resulting in smoother fuel flow changes and no overshoot. This further demonstrates the reliability of this invention.

[0077] Although the embodiments of the present invention have been shown and described above, it will be understood that the above embodiments are illustrative and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, replace and modify the above embodiments within the scope of the present invention without departing from the principles and purpose of the present invention.

Claims

1. A method for rapid thrust response control of an aero-engine based on redundant actuators, characterized in that The specific steps are as follows: Step 1: Construct the main fuel flow W f and the tail nozzle throat area A8 controls the thrust F n The transfer function L is: L=g c1 g a1 g1+g c2 g a2 g2 Among them, g c1 and g c2 is the controller module to be determined, g1 is ΔW f to ΔF n The transfer function, g2 is ΔA8 to ΔF n The transfer function, g a1 and g a2 W f and the transfer function of the A8 actuator, Δ represents the difference of the physical quantity relative to its steady-state value; Step 2: Calculate g c1 and g c2 , the specific formula is as follows Among them, K p It is an undetermined parameter with design freedom, and its size determines W f The relative contribution of K and A8 to thrust response; i The calculation formula is as follows: Among them, ω b is the desired thrust bandwidth, g a1 (0) is the initial value of the fuel actuator transfer function, g1(0) is the thrust to fuel flow W f The initial value of the response function; Step 3: Controller module g obtained in step 2 c1 and g c2 , W f and the transfer function g of the A8 actuator a1 and g a2 Set it as a first-order system for simulation, the specific form is as follows: Among them, ω a To adjust the countdown of time; Step 4: Obtain g1 and g2 by performing system identification at the engine steady-state operating point; Step 5: Based on the g obtained in steps 2-4 c1 and g c2 、g a1 and g a2 , g1 and g2 to solve the transfer function L, and then design the valve position control VPC controller according to the transfer function L, so as to achieve rapid response control of the thrust.

2. The method for rapid thrust response control of an aero-engine based on redundant actuators according to claim 1, characterized in that: In step 1, ΔW f to ΔF n The control includes three control modules: the undetermined controller module, whose transfer function is unknown, set as g c1 ; The execution control module that receives the control command of the pending controller module has a transfer function of g a1 , for the engine W f Control; the control module of main fuel flow to thrust has a transfer function of g1.

3. The method for rapid thrust response control of an aircraft engine based on redundant actuators according to claim 1, characterized in that: In step 1, ΔA8 to ΔF n The control includes three control modules: the undetermined controller module, whose transfer function is unknown, set as g c2 ; The execution control module that receives the control command of the pending controller module has a transfer function of g a2 , controls the engine's A8; the control module of the tail nozzle throat area to thrust, its transfer function is g2.

4. The method for rapid thrust response control of an aero-engine based on redundant actuators according to claim 1, characterized in that: In step 2, K p The larger the W f The higher the degree of participation in the transition process, the range of its value is [0, +∞). The reason why it cannot be negative is to avoid the reduction of fuel flow during the thrust increase process.

5. The method for rapid thrust response control of an aero-engine based on redundant actuators according to claim 1, characterized in that: In step 3, the larger the bandwidth, the faster the response time, and the value range is 10 to 100 rad / s.

6. The method for rapid thrust response control of an aero-engine based on redundant actuators according to claim 1, characterized in that: In step 3, the adjustment time of the actuator is one order of magnitude shorter than the adjustment time of the controller.

7. A system for executing the method for rapid thrust response control of an aircraft engine based on redundant actuators according to any one of claims 1 to 6, characterized in that: Including ΔW f to ΔF n The control route and ΔA8 to ΔF n The VPC controller is constructed in parallel with the control route, where ΔW f to ΔF n The control route includes a first pending controller module, a first execution control module, and a control module for main fuel flow to thrust; the control route from ΔA8 to ΔF includes a second pending controller module, a second execution control module, and a control module for tail nozzle throat area to thrust; a desired thrust control instruction is sent to the VPC controller, and the VPC control system can track the instruction to achieve a fast response control effect.

8. The system according to claim 7, characterized in that: The transfer function of the VPC controller is, L=g c1 g a1 g1+g c2 g a2 g2。 9. An electronic device, characterized in that: The invention comprises at least one processor and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can execute the method for rapid thrust response control of an aircraft engine based on a redundant actuator as described in any one of claims 1 to 6.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores computer instructions, and the computer instructions are used to enable a processor to implement the aircraft engine thrust rapid response control method based on redundant actuators as described in any one of claims 1 to 6 when executed.

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