An Incremental Nonlinear Control Method for Damaged Aircraft with Performance Prediction

Abstract

The present invention provides an incremental nonlinear control method for damaged aircraft with performance prediction, including: establishing an angular velocity motion equation and deriving the incremental form of the motion equation; using the control surface as the control input and calculating the control effectiveness matrix at the previous moment according to the flight state; obtaining the angular acceleration signal at the previous moment through the difference method; setting the angular velocity performance boundary, selecting the corresponding performance boundary function and performing error transformation; designing the virtual control quantity of the angular velocity; designing an incremental nonlinear fault-tolerant control law according to the angular velocity increment equation, angular acceleration, control effectiveness matrix, rudder deflection signal and virtual control quantity to achieve fault-tolerant flight control of damaged aircraft. This method can effectively improve the robustness of the flight control system and solve the fault-tolerant control problem of aircraft with wing damage. In addition, this method realizes the unification of the transient performance and steady-state performance of damaged aircraft and ensures that the angular velocity dynamic response of damaged aircraft is within the expected performance boundary.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft damage-tolerant control, and particularly to an incremental nonlinear control method for damaged aircraft with performance prediction. Background Art

[0002] As a key component of an aircraft, the wing is the main component that generates lift for the aircraft. Wing damage will not only significantly reduce the lift of the aircraft, but also destroy the inherent balance, making the left and right lift of the aircraft asymmetric, thus generating an additional rolling moment, causing the damaged aircraft to continuously dive with its head down, and the degree increasing continuously. Therefore, wing damage is fatal to the aircraft and poses a severe challenge to the robustness of the flight controller.

[0003] Although the existing fault-tolerant control laws can improve the robustness of the flight control system, the dynamic response of the aircraft is often ignored during the fault-tolerant adjustment process, so the transient performance of the damaged aircraft cannot be guaranteed, thus reducing the flight quality of the aircraft. In addition, since the structure of the general fault-tolerant flight control system is relatively complex and usually requires a large amount of offline model data as support, this makes the design and implementation process of the fault-tolerant control law extremely complex. Therefore, there is an urgent need for a fault-tolerant control law to solve the fault-tolerant control problem of damaged aircraft and the problem of inconsistent transient and steady-state performance in general fault-tolerant control laws. This kind of fault-tolerant control is required to not only improve the robustness of the aircraft to damage interference, but also ensure that the dynamic performance of the aircraft meets the expected requirements after damage occurs. Summary of the Invention

[0004] To solve the above problems, the present invention provides the following technical solutions.

[0005] An incremental nonlinear control method for damaged aircraft with performance prediction, comprising the following steps:

[0006] Establish an angular velocity motion equation and derive an angular velocity increment equation;

[0007] Taking the elevator, aileron and rudder as control input quantities, calculate the control effectiveness matrix at the previous moment according to the flight state of the aircraft;

[0008] Obtain the angular acceleration signal and the control surface deflection signal at the previous moment;

[0009] Set the angular velocity performance boundary, select the corresponding performance boundary function and perform error transformation to obtain an error conversion function and a conversion error;

[0010] Design a virtual control quantity with performance preset according to the error conversion function and the conversion error;

[0011] According to the angular velocity increment equation, angular acceleration signal, control effectiveness matrix, rudder deflection signal and virtual control quantity with performance preset, an incremental nonlinear fault-tolerant control law is designed to realize flight control of damaged aircraft.

[0012] Preferably, establishing the angular velocity motion equation and deriving the angular velocity increment equation comprises the following steps:

[0013] Use CATIA software to draw a 3D aircraft model and calculate the normal aircraft mass m, wing area S, span b, and average aerodynamic chord length and moment of inertia J;

[0014] The Xflow software is used to simulate the wind tunnel to calculate the aircraft's control derivatives, including: aileron delta a Derivative of roll Rudder δ r Derivative of roll Elevator δ e Derivative of pitch Aileron derivative with respect to yaw and the rudder's derivative with respect to yaw

[0015] Establish the aircraft angular velocity motion equation:

[0016]

[0017] Where:

[0018]

[0019] in, and F = [X, Y, Z] T Indicates the moment acting on the aircraft and the force G b External aerodynamic force; ω = [p, q, r] T is the three-axis angular velocity; v b =[u,v,w] T Indicates the three-axis speed of the machine system; Δr=[Δx cg ,Δy cg ,Δz cg ] T Indicates the position of the center of gravity change;

[0020] The Taylor series expansion method is used to obtain the incremental form of the angular velocity dynamics. The expansion results are as follows:

[0021]

[0022] Where x represents the state quantity other than angular velocity, and ε represents the items related to the state increment and angular velocity increment, as follows:

[0023]

[0024] Preferably, taking the elevator, aileron and rudder as control input variables, calculating the control effectiveness matrix at the previous moment according to the flight state of the aircraft, includes the following steps:

[0025] Specify the elevator δ e , aileron δ a and rudder δ r as the control input variable δ, and calculate the aircraft control effectiveness matrix according to the obtained control derivatives as follows:

[0026]

[0027] where x0 represents the aircraft state variables at the previous moment, Q = 0.5ρ‖v b ‖ 2 represents the dynamic pressure.

[0028] Preferably, obtaining the angular acceleration signal at the previous moment includes the following steps:

[0029] Obtain the angular velocity information at the previous moment using a second-order filter H(s), and obtain the angular acceleration signal at the previous moment through differencing as shown in the following equation:

[0030]

[0031] Preferably, setting the angular velocity performance boundary, selecting the corresponding performance boundary function and performing error transformation, obtaining the error transformation function and transformation error, includes the following steps:

[0032] Design the performance boundary function ξ(t), which satisfies: when t1 > t2, 0 < ξ(t1) < ξ(t2), and is designed as:

[0033] ξ(t) = (ξ0 - ξ ∞ )e -at + ξ ∞ (7)

[0034] In the formula, and a > 0;

[0035] Restrict the angular velocity transient performance of the normal / damaged aircraft, that is, require the aircraft angular velocity error to always satisfy:

[0036] -M'ξ(t) < e(t) < N'ξ(t) (8)

[0037] In the formula, M', N' > 0, e(t) = ω(t) - ω cmd (t) is the conventional angular velocity error;

[0038] The error is converted, and the result is as follows:

[0039] e(t) = ξ(t)Ω(z(t)) (9)

[0040] Wherein, z(t) represents the conversion error, and the conversion function Ω(z(t)) is as follows:

[0041]

[0042] Preferably, designing a virtual control quantity with a performance preset according to the error conversion function and the conversion error includes the following steps:

[0043] According to the error conversion function, the virtual control quantity ν ω (t) is designed as:

[0044]

[0045] Wherein, k > 0 represents the angular velocity bandwidth.

[0046] Preferably, designing an incremental nonlinear fault-tolerant control law according to the angular velocity increment equation, the angular acceleration signal, the control effectiveness matrix, the rudder surface deflection signal, and the virtual control quantity with a performance preset includes the following steps:

[0047] According to the angular acceleration, the control effectiveness matrix, the rudder surface deflection signal, and the virtual control quantity with a performance preset, the incremental nonlinear angular velocity fault-tolerant control law is designed as:

[0048]

[0049] Wherein, δ0 represents the rudder surface deflection measured by the sensor at the previous moment.

[0050] Advantages of the present invention:

[0051] The incremental nonlinear angular velocity fault-tolerant control law with a preset performance designed by the present invention can effectively improve the robustness of the aircraft to faults and disturbances, and while ensuring the stability of the damaged aircraft, it can take into account the transient performance of the aircraft during the fault-tolerant adjustment process, realizing the unity of the transient performance and the steady-state performance of the damaged aircraft. In addition, the designed fault-tolerant control law has a simple structure, does not rely on an accurate system model, and does not require a complex parameter tuning process during the design process, so it is easy to be extended to different aircraft. Description of the Drawings

[0052] Figure 1 It is the overall structural framework of an incremental nonlinear control method for a damaged aircraft with performance prediction according to an embodiment of the present invention;

[0053] Figure 2It is the comparison chart of roll angular velocity response in the simulation experiment of the embodiment of the present invention;

[0054] Figure 3 It is the comparison chart of pitch angular velocity response in the simulation experiment of the embodiment of the present invention;

[0055] Figure 4 It is the comparison chart of yaw angular velocity response in the simulation experiment of the embodiment of the present invention;

[0056] Figure 5 It is the comparison and performance boundary of roll angular velocity error in the simulation experiment of the embodiment of the present invention;

[0057] Figure 6 It is the comparison and performance boundary of pitch angular velocity error in the simulation experiment of the embodiment of the present invention;

[0058] Figure 7 It is the comparison and performance boundary of yaw angular velocity error in the simulation experiment of the embodiment of the present invention;

[0059] Figure 8 It is the comparison chart of control rudder surface input in the simulation experiment of the embodiment of the present invention. Specific embodiments

[0060] In order to make the purpose, technical solutions and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

[0061] Embodiment 1

[0062] An incremental non - linear control method of the present invention for damaged aircraft with performance prediction. As Figure 1 shown:

[0063] S1: Establish the angular velocity motion equation and derive the incremental form of the angular velocity motion equation.

[0064] Draw a three - dimensional model of the aircraft through CATIA software, and calculate the mass m, wing area S, span b, mean aerodynamic chord length and moment of inertia J of the normal aircraft;

[0065] Calculate the control derivatives of the aircraft by simulating the wind tunnel through Xflow software, specifically including: aileron δ a derivative with respect to roll rudder δ r derivative with respect to roll elevator δ e derivative with respect to pitch derivative of aileron with respect to yaw and derivative of rudder with respect to yaw

[0066] Establish the aircraft angular velocity motion equation:

[0067]

[0068] Where:

[0069]

[0070] Among them, and F = [X, Y, Z] T represents the moment acting on the aircraft and the aerodynamic force except the gravity G b ω = [p, q, r] T is the three-axis angular velocity; v b = [u, v, w] T represents the three-axis velocity in the body coordinate system; Δr = [Δx cg , Δy cg , Δz cg T represents the position change of the center of gravity. [[ID=??]]

[0071] Adopt the Taylor series expansion method to obtain the incremental form of the angular velocity dynamics, and the expansion result is as follows:

[0072]

[0073] Where x represents the state quantity other than the angular velocity, and ε represents the terms related to the state increment and the angular velocity increment, specifically as follows:

[0074]

[0075] S2: Take the control surface as the control input and calculate the aircraft control effectiveness matrix.

[0076] Specify the elevator δ e , aileron δ a and rudder δ r as the control input quantity δ. Calculate the aircraft control effectiveness matrix according to the control derivative obtained in S1, specifically as follows:

[0077]

[0078] Among them, x0 represents the aircraft state quantity at the previous moment, and Q = 0.5ρ||v b || 2 represents the dynamic pressure.

[0079] S3: Obtain the angular acceleration signal at the previous moment by means of difference.

[0080] Note: There seems to be a typo in the original text at line 37 where "??" is shown. It should likely be something else. This has been left as is in the translation for the sake of following the instructions.The angular velocity information at the previous moment is obtained by using a second-order filter H(s). The angular acceleration signal at the previous moment is obtained by the difference method. Specifically, it is shown as the following formula:

[0081]

[0082] S4: Set the angular velocity performance boundary according to the angular velocity expectation dynamics, and select the corresponding performance boundary function.

[0083] Design the performance boundary function ξ(t), which satisfies: when t1 > t2, 0 < ξ(t1) < ξ(t2), and it is designed as:

[0084] ξ(t) = (ξ0 - ξ ∞ )e -at + ξ ∞ (7)

[0085] In the formula, and a > 0;

[0086] Constrain the angular velocity transient performance of the normal / damaged aircraft, that is, it is required that the angular velocity error of the aircraft always satisfies:

[0087] -M'ξ(t) < e(t) < N'ξ(t) (8)

[0088] In the formula, M', N' > 0, e(t) = ω(t) - ω cmd (t) is the conventional angular velocity error;

[0089] Convert the error, and the result is as follows:

[0090] e(t) = ξ(t)Ω(z(t)) (9)

[0091] In the formula, z(t) represents the conversion error, and the conversion function Ω(z(t)) is as follows:

[0092]

[0093] S5: Design the virtual control quantity of the angular velocity according to the error conversion function and the conversion error.

[0094] According to the error conversion function in S4, the virtual control quantity ν ω (t) is designed as:

[0095]

[0096] In the formula, k > 0 represents the angular velocity bandwidth.

[0097] S6: Design an incremental nonlinear fault-tolerant control law based on the incremental form of the angular velocity equation of motion, angular acceleration, control effectiveness matrix, rudder deflection signal, and virtual control variables with performance presets, ultimately achieving flight control of the damaged aircraft. This includes the following steps:

[0098] According to the angular velocity increment equation, angular acceleration, control effectiveness matrix, rudder deflection signal and virtual control quantity with performance preset obtained from S1-S6, the incremental nonlinear angular velocity fault-tolerant control law is designed as follows:

[0099]

[0100] Among them, δ0 represents the deflection of the rudder surface measured by the sensor at the previous moment.

[0101] In this embodiment, under the condition of wing damage fault, the angular velocity fault-tolerant control law with incremental nonlinearity and preset performance is simulated and evaluated.

[0102] The aircraft is at an altitude of H = 3500m and a speed of V a =120m / s in straight and level flight, the initial state of the aircraft is θ0=α0=1.68°, The sampling time of the fault-tolerant controller is Δt = 0.02s, the controller parameter is selected as k = 10, and the preset error boundary is set as follows:

[0103]

[0104] Assuming that the left wing of the aircraft is damaged by 19.1% during flight, an angular velocity command is given under this damage fault. The incremental nonlinear control law with preset performance designed in this patent is compared with the general fault-tolerant control law. The simulation comparison results are as follows: Figures 2 - 8 shown.

[0105] It can be seen from the simulation results that after wing damage occurs, the fault-tolerant controller designed based on the method of this patent can quickly and stably damage the aircraft, and the control performance after damage does not decrease due to wing damage. Compared with general fault-tolerant controllers, the angular velocity response under the action of the fault-tolerant controller designed in this patent is faster than that of general fault-tolerant controllers, and the angular velocity error is always kept within the preset performance boundary. However, the peak value of the angular velocity error under the general fault-tolerant controller is greater than that of the fault-tolerant controller designed in this patent, and the transient performance does not meet the expected performance. Therefore, the incremental nonlinear fault-tolerant controller with preset performance designed in this patent can not only improve the robustness of the aircraft to fault interference, but also ensure the transient performance of the aircraft during the fault-tolerant adjustment process, achieving the goal of unifying transient performance and steady-state performance.

[0106] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included within the protection scope of the present invention.

Claims

1. An incremental non-linear control method for damaged aircraft with performance prediction, characterized in that Including the following steps: Establish the angular velocity motion equation and derive the angular velocity increment equation; Taking the elevator, aileron and rudder as control input quantities, calculate the control effectiveness matrix at the previous moment according to the aircraft flight state; Obtain the angular acceleration signal and the control surface deflection signal at the previous moment; Set the angular velocity performance boundary, select the corresponding performance boundary function and perform error transformation to obtain the error transformation function and the transformation error; Design a virtual control quantity with performance preset according to the error transformation function and the transformation error; Design an incremental nonlinear fault-tolerant control law according to the angular velocity increment equation, the angular acceleration signal, the control effectiveness matrix, the control surface deflection signal and the virtual control quantity with performance preset to achieve the flight control of the damaged aircraft; The setting of the angular velocity performance boundary, selecting the corresponding performance boundary function and performing error transformation to obtain the error transformation function and the transformation error includes the following steps: Design performance boundary function such that the boundary function satisfies: when then is designed as: wherein, , , and ; Constraining the angular velocity transient performance of the normal / damaged aircraft, that is, requiring that the aircraft angular velocity error always satisfies: In the formula, , is the conventional angular velocity error; Perform transformation on the error, and the result is as follows: In the formula, represents the conversion error, and the conversion function is as follows: ， 。 2. The incremental non-linear control method for damaged aircraft with performance prediction according to claim 1, characterized in that, The establishment of the angular velocity motion equation and the derivation of the angular velocity increment equation include the following steps: Draw the model of a three-dimensional aircraft using CATIA software and calculate the mass of a normal aircraft , wing area , span , mean aerodynamic chord and moment of inertia J ; Calculating the control derivatives of an aircraft by simulating a wind tunnel with Xflow software, specifically including: ailerons Derivatives with respect to roll , rudders Derivatives with respect to roll , elevators Derivatives with respect to pitch , derivatives of ailerons with respect to yaw , and derivatives of rudders with respect to yaw ; Establish the aircraft angular velocity motion equation: Where: Among them, and represent the moment acting on the aircraft and the aerodynamic force other than gravity outside; is the three-axis angular velocity; represents the three-axis velocity in the body coordinate system; represents the changing position of the center of gravity; Using the Taylor series expansion method, obtain the incremental form of the angular velocity dynamics, and the expansion result is as follows: In the formula, represents a state quantity other than the angular velocity, represents terms related to the state increment and the angular velocity increment, specifically as follows: 。 3. The incremental nonlinear control method for damaged aircraft with performance prediction according to claim 2, characterized in that, Taking the elevator, aileron and rudder as control input quantities, calculating the control effectiveness matrix at the previous moment according to the aircraft flight state includes the following steps: Elevator , Aileron and Rudder are used as control input quantities . According to the obtained handling derivatives , , , , , the aircraft control effectiveness matrix is calculated as follows: Among them, represents the aircraft state quantity at the previous moment, represents the dynamic pressure.

4. The incremental nonlinear control method for damaged aircraft with performance prediction according to claim 3, characterized in that, The obtaining of the angular acceleration signal at the previous moment includes the following steps: Adopt a second-order filter Obtain the angular velocity information at the previous moment, and obtain the angular acceleration signal at the previous moment through the differential method , as shown in the following formula 。 5. The incremental non-linear control method for damaged aircraft with performance prediction according to claim 4, characterized in that The design of a virtual control quantity with performance preset according to the error transformation function and the transformation error includes the following steps: According to the error conversion function, the virtual control quantity is designed as: In the formula, represents the angular velocity bandwidth.

6. The incremental nonlinear control method for damaged aircraft with performance prediction according to claim 5, wherein The design of an incremental nonlinear fault-tolerant control law according to the angular velocity increment equation, the angular acceleration signal, the control effectiveness matrix, the control surface deflection signal and the virtual control quantity with performance preset includes the following steps: According to the angular acceleration, the control effectiveness matrix, the control surface deflection signal and the virtual control quantity with performance preset, the incremental nonlinear angular velocity fault-tolerant control law is designed as: Among them, represents the rudder surface deflection measured by the sensor at the previous moment.

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