An active fault-tolerant and anti-interference trajectory tracking control method for an unmanned helicopter considering actuator failures

Abstract

The present invention discloses an active fault-tolerant anti-interference trajectory tracking control method for an unmanned helicopter considering actuator failure, comprising the following steps: establishing a full-loop disturbed dynamic model of the unmanned helicopter; respectively designing interference observers for a position control subsystem and an attitude control subsystem; realizing lumped interference estimation for a position loop and lumped interference estimation for an attitude loop; designing a composite dynamic inverse position controller and a composite dynamic inverse attitude controller; obtaining the total pitch of the control quantity and the inner loop attitude angle instruction through the composite dynamic inverse position controller; according to the inner loop attitude angle instruction, obtaining the tail rotor pitch, longitudinal cyclic pitch and lateral cyclic pitch of the unmanned helicopter through the composite dynamic inverse attitude controller. The present invention regards the influence of multi-source interference and actuator failure as lumped interference, effectively reduces the influence of multi-source interference and actuator failure on the system control accuracy, and significantly improves the trajectory tracking accuracy of the helicopter system.

Description

Technical Field

[0001] This invention belongs to the field of flight control technology, specifically relating to an active fault-tolerant and anti-interference trajectory tracking control method for unmanned helicopters that takes into account actuator failures. Background Technology

[0002] Unmanned helicopters are unmanned, vertically oriented rotorcraft capable of vertical takeoff and landing (VTOL) and spatial maneuvering. They offer advantages such as low cost and high flexibility, and are widely used in fields such as surveying, photography, and military reconnaissance. Compared to fixed-wing UAVs, unmanned helicopters do not require a wide, flat runway for takeoff and landing, enabling vertical takeoff and landing and hovering, and allowing them to operate in complex airspace with multiple obstacles. Compared to quadcopter UAVs, unmanned helicopters can achieve faster level flight and greater payload capacity. However, unmanned helicopters are complex nonlinear systems with underactuated characteristics and highly coupled control variables. Actuator failures significantly impact flight control performance. Furthermore, their flight is affected by external gusts, internal parameter perturbations, and unmodeled dynamics, making high-precision trajectory tracking control of unmanned helicopters extremely challenging.

[0003] In recent years, researchers both domestically and internationally have applied various control strategies and algorithms to address the trajectory tracking problem of unmanned helicopters, including linear PID control, robust control, backstepping control, and sliding mode control. However, these control methods mainly rely on the robustness of the algorithm itself to passively suppress the effects of faults and disturbances, sacrificing some performance indicators during the design and parameter tuning process. Therefore, there is a need to propose an active control method that can quickly suppress the adverse effects of actuator faults and multi-source disturbances. Summary of the Invention

[0004] Purpose of the invention: To address the adverse effects of actuator failures and multi-source disturbances on the control system of unmanned helicopters, this invention provides an active fault-tolerant and anti-interference trajectory tracking control method for unmanned helicopters that considers actuator failures. This method treats the effects of multi-source disturbances and actuator failures as lumped disturbances, and combines extended state observer technology with dynamic inverse control methods. While estimating and feedforward compensating for lumped disturbances, it retains the advantages of the dynamic inverse method, such as its simplicity and ease of engineering implementation. This effectively reduces the impact of multi-source disturbances and actuator failures on the system's control accuracy and significantly improves the trajectory tracking accuracy of the helicopter system.

[0005] Technical Solution: To achieve the above objectives, this invention provides an active fault-tolerant and anti-interference trajectory tracking control method for unmanned helicopters that considers actuator failures, comprising the following steps:

[0006] S1: Establish a full-loop dynamics model of an unmanned helicopter under disturbance;

[0007] S2: Based on the established full-loop disturbed dynamics model of the unmanned helicopter, disturbance observers for the position control subsystem and attitude control subsystem are designed respectively.

[0008] S3: The lumped disturbance estimation for the position loop and the lumped disturbance estimation for the attitude loop are realized by the disturbance observers of the position control subsystem and the attitude control subsystem, respectively.

[0009] S4: Combine the lumped disturbance estimation of the position loop to design a composite dynamic inverse position controller; combine the lumped disturbance estimation of the attitude loop to design a composite dynamic inverse attitude controller; this step realizes the feedforward compensation for lumped disturbance, enabling the system to have active anti-interference and fault tolerance capabilities.

[0010] S5: The collective control quantity and the inner loop attitude angle command are obtained through the composite dynamic inverse position controller;

[0011] S6: Based on the inner loop attitude angle command, the tail rotor pitch, longitudinal cyclic pitch, and lateral cyclic pitch of the unmanned helicopter are obtained through the composite dynamic inverse attitude controller.

[0012] Furthermore, the disturbed dynamics model of the unmanned helicopter in step S1 includes a disturbed unmanned helicopter position loop dynamics model with actuator failure, an attitude loop dynamics model affected by multi-source disturbances, and an actuator model affected by failure. The establishment process is as follows:

[0013] A1: Establish a dynamic model of the position loop of a disturbed unmanned helicopter that includes actuator failure:

[0014]

[0015] Where P = [x, y, z] T V = [v x ,v y ,v z ] T Let represent the position and velocity of the unmanned helicopter in the inertial frame of reference, respectively; m be the mass of the unmanned helicopter; g be the gravitational acceleration constant; and e = [0, 0, 1]. T Let Θ be a three-dimensional vector, Θ = [φ, θ, ψ]. T Let T be the helicopter attitude angles, φ, θ, and ψ be the roll angle, pitch angle, and yaw angle, respectively. m Main rotor lift, F d ∈R 3×1 For multi-source interference caused by aerodynamic parameter perturbation, external gusts, etc., R(Θ) is the transformation matrix from the body axis to the inertial frame;

[0016] A2: Establish a dynamic model of the attitude loop affected by multi-source disturbances:

[0017]

[0018] Where Ω = [p, q, r] T Let p, q, and r be the angular velocity of rotation about the fuselage axis, respectively, and let J = diag{I}. x ,I y ,I z Let Γ be the moment of inertia matrix, where Γ = [L, M, N]. T For the three-axis moments, L, M, and N are the rolling moment, pitching moment, and yaw moment, respectively, and D... A ∈R 3×1 For attitude loop multi-source interference, H(Θ) is the transformation matrix;

[0019] A3: Establish an actuator model affected by the fault:

[0020] Establish the relationship between main rotor lift and collective pitch:

[0021]

[0022] Where, δ col Main rotor collective pitch, Z w and Z col For coefficients, The collective pitch failure factor is given, and w is the helicopter's velocity along the z-axis of the fuselage. This is the nominal value of the main rotor lift when the collective pitch is fault-free. The main rotor lift caused by collective pitch failure is specifically expressed as follows:

[0023]

[0024] Establish the relationship between rolling torque L, pitching moment M, and periodic moment:

[0025]

[0026] Among them, L a ,L b M a M b A lon A lat B lon B lat Here are system parameters, and ε is the time constant. The fault coefficients for the longitudinal cycle distance and the transverse cycle distance are respectively, and for ease of writing, the following definitions are introduced:

[0027]

[0028] Then equation (4) can be rewritten as:

[0029]

[0030] Among them, L n M n L represents the nominal rolling torque and pitching moment during the period without failure. f M f Rolling torque and pitching moment caused by periodic pitch faults:

[0031]

[0032] Establish the relationship between yaw moment and tail rotor pitch and collective pitch:

[0033]

[0034] Where, N r N col N ped For system parameters, δ ped The tail rotor pitch, N is the tail rotor pitch failure factor. n The nominal yaw moment value is ignoring faults, N f Yaw moment caused by tail rotor failure:

[0035] N n =N r r+N col δ col +N ped δ ped ,

[0036] Furthermore, the design method for the disturbance observer of the position control subsystem in step S2 is as follows:

[0037] Define the position command tracking error as: e x =xx r ,e y =yy r ,e z =zz r , where x r ,y r ,z r Assuming the position command, by simultaneously solving equations (1) and (3), we can obtain the dynamic equation for the tracking error of the position subsystem:

[0038]

[0039] The lumped interference includes interference and fault information:

[0040]

[0041] The design method for the disturbance observer of the attitude control subsystem is as follows:

[0042] Define attitude angle tracking error:

[0043]

[0044] Based on equation (2), the system equation for the tracking error of the attitude angle subsystem is obtained:

[0045]

[0046] In the formula, Γ n The nominal values ​​of the torque around the fuselage shaft when there are no faults in the periodic pitch, collective pitch, and tail rotor pitch:

[0047] Γ n =[L n M n N n ] T ,

[0048] The matrix H(Θ) is the result of taking the derivative of each element of H(Θ) with respect to t, as follows:

[0049]

[0050] D in the formula LA ∈R 3×1 The lumped disturbance includes information on disturbances, faults, and the second derivative of the attitude angle command:

[0051]

[0052] Further, step S3 specifically includes:

[0053] Lumped disturbance estimation for location loop:

[0054] For the unmanned helicopter external loop position control subsystem of formula (1), an intermediate control variable is introduced to simplify the writing:

[0055]

[0056] The x-channel extended state observer is designed as follows:

[0057]

[0058] Design an expanded state observer for the y-channel:

[0059]

[0060] Design of the z-channel extended state observer:

[0061]

[0062] In the formula, These are the observer parameters, and their values ​​are all positive constants. These are the estimated values ​​of the lumped interference for the three channels x, y, and z, respectively.

[0063] Lumped disturbance estimation of the attitude loop:

[0064] For the attitude loop tracking error system of Equation (8), an extended state observer is designed to estimate the lumped disturbance:

[0065]

[0066] In the formula, the observer gain matrix Belongs to R 3×3 The diagonal array, and These are the derivative of the attitude angle command tracking error and the estimated value of the lumped disturbance, respectively.

[0067] Furthermore, the design method for the composite dynamic inverse position controller in step S4 is as follows:

[0068] For the external loop position subsystem, a composite dynamic inverse position controller is designed:

[0069]

[0070] In the above formula, k xp ,k xd ,k yp ,k yd ,k zp ,k zd These are controller parameters, and all their values ​​are positive integers.

[0071] The design method of the composite dynamic inverse attitude controller is as follows:

[0072] To achieve the three-channel decoupling of the attitude angle tracking error system in formula (8), the following definition is introduced:

[0073]

[0074] The attitude angle tracking error system can then be written as:

[0075]

[0076] By combining the lumped disturbance estimate of the observer, a composite dynamic inverse attitude controller is designed:

[0077]

[0078] In the above formula, k φd ,k φp ,k θd ,k θp,k ψd ,k ψp These are the parameters for the attitude controller, and all their values ​​are normal numbers.

[0079] Further, step S5 specifically includes:

[0080] The composite dynamic inverse position controller calculates the total control pitch δ from the intermediate control input. col And the inner loop attitude angle command:

[0081]

[0082] Further, step S6 specifically includes:

[0083] According to definition (16) and the composite dynamic inverse attitude controller (18), we can obtain:

[0084]

[0085] The longitudinal and lateral periodic distances of the unmanned helicopter can be obtained by solving equations (5) and (19):

[0086]

[0087] The tail rotor pitch of the unmanned helicopter can be obtained by solving equation (6) and combining equations (15) and (19):

[0088] δ ped =(N n -N r rN col δ col ) / N ped (twenty one)

[0089] δ ped The tail rotor pitch is δ. lon For longitudinal periodic pitch, δ lat It is a lateral periodic pitch variation.

[0090] Furthermore, in step A1, R(Θ) is the transformation matrix from the body axis to the inertial frame, specifically in the form of:

[0091]

[0092] Among them, cosφ=cφ, sinφ=sφ, cosθ=cθ, sinθ=sθ, cosψ=cψ, sinψ=sψ.

[0093] Furthermore, in step A2, H(Θ) is a transformation matrix, specifically represented as follows:

[0094]

[0095] Among them, sinφ=sφ, cosφ=cφ, tanθ=tθ, sinθ=sθ..

[0096] The method of this invention first establishes a full-loop disturbed dynamic model of an unmanned helicopter, including actuator failures, affected by multi-source disturbances. Second, based on extended state observer technology, disturbance observers are designed for the position control subsystem and attitude control subsystem respectively to estimate the lumped disturbance. Then, combining the disturbance estimation information and a benchmark dynamic inverse algorithm, a composite fault-tolerant anti-disturbance controller for the position and attitude loops is designed to obtain the required forces and torques. Finally, based on the required forces and torques, the actual control quantities of the helicopter—cyclic pitch and tail rotor pitch—are calculated.

[0097] Beneficial effects: Compared with the prior art, the present invention has the following advantages:

[0098] 1. This invention fully analyzes the nonlinear coupling characteristics of unmanned helicopters, and reduces the difficulty of adjusting the controller parameters by feedback cancellation of known nonlinear parts, thus ensuring the flight dynamic characteristics of unmanned helicopters.

[0099] 2. This invention treats the effects of actuator failure and multi-source interference as lumped interference. By estimating the lumped interference and feedforward compensation, the system has the ability to actively resist interference and tolerate faults.

[0100] 3. This invention significantly improves the response speed and tracking accuracy of unmanned helicopters to complex trajectories in the event of actuator failure. Its control strategy can also be used in a variety of similar nonlinear systems, and has broad engineering application prospects. Attached Figure Description

[0101] Figure 1 This is a block diagram of the full-loop active fault-tolerant anti-interference trajectory tracking control method used in this invention;

[0102] Figure 2 The image shows the effect of an unmanned helicopter tracking a cylindrical helical trajectory command in three-dimensional space under the action of the baseline dynamic inverse controller (NIDC) and the proposed composite dynamic inverse controller (CNDIC).

[0103] Figure 3 The response curves of the position three-channel and heading angle channel of the unmanned helicopter to commands under the action of the baseline dynamic inverse controller (NIDC) and the proposed composite dynamic inverse controller (CNDIC);

[0104] Figure 4 The command tracking error curves of the position three-channel and heading angle channel of the unmanned helicopter under the action of the baseline dynamic inverse controller (NIDC) and the proposed composite dynamic inverse controller (CNDIC);

[0105] Figure 5 The control input response curves of the unmanned helicopter under the action of the baseline dynamic inverse controller (NIDC) and the proposed composite dynamic inverse controller (CNDIC) are shown. Detailed Implementation

[0106] The present invention will be further illustrated below with reference to the accompanying drawings and specific embodiments. It should be understood that these embodiments are for illustrative purposes only and are not intended to limit the scope of the invention. After reading this invention, any modifications of the invention in various equivalent forms by those skilled in the art will fall within the scope defined by the appended claims.

[0107] This invention provides an active fault-tolerant and anti-interference trajectory tracking control method for unmanned helicopters that considers actuator failures, such as... Figure 1 As shown, it includes the following steps:

[0108] S1: Establish a full-loop disturbed dynamic model for the unmanned helicopter. The establishment process is as follows:

[0109] A1: Establish a dynamic model of the position loop of a disturbed unmanned helicopter that includes actuator failure:

[0110]

[0111] Where P = [x, y, z] T V = [v x ,v y ,v z ] T Let m and g represent the position and velocity of the unmanned helicopter in the inertial frame of reference, respectively. Let m be the mass of the unmanned helicopter and g be the gravitational acceleration constant. In this embodiment, m = 6 kg and g = 9.8 m / s². 2 ;

[0112] e = [0, 0, 1] T Let Θ be a three-dimensional vector, Θ = [φ, θ, ψ]. T Let T be the helicopter attitude angles, φ, θ, and ψ be the roll angle, pitch angle, and yaw angle, respectively. m Main rotor lift, F d ∈R 3×1 For multi-source interference caused by aerodynamic parameter perturbations, external gusts, etc., R(Θ) is the transformation matrix from the body axis to the inertial frame, specifically in the form of:

[0113]

[0114] Among them, cosφ=cφ, sinφ=sφ, cosθ=cθ, sinθ=sθ, cosψ=cψ, sinψ=sψ.;

[0115] A2: Establish a dynamic model of the attitude loop affected by multi-source disturbances:

[0116]

[0117] Where Ω = [p, q, r] T Let p, q, and r be the angular velocity of rotation about the fuselage axis, respectively, and let J = diag{I}. x ,I y ,I z} represents the rotational inertia matrix, I in this embodiment x =0.3; I y =0.68,I z =0.787;

[0118] Γ=[L,M,N] T For the three-axis moments, L, M, and N are the rolling moment, pitching moment, and yaw moment, respectively, and D... A ∈R 3×1 For attitude loop multi-source interference, H(Θ) is the transformation matrix, specifically expressed as:

[0119]

[0120] Among them, sinφ=sφ, cosφ=cφ, tanθ=tθ, sinθ=sθ.;

[0121] A3: Establish an actuator model affected by the fault:

[0122] Establish the relationship between main rotor lift and collective pitch:

[0123]

[0124] Where, δ col Main rotor collective pitch, Z w and Z col For coefficients, Z is the total distance fault factor in this embodiment. w = -0.878, Z col = -5.71;

[0125] w is the velocity of the helicopter along the z-axis of the fuselage. This is the nominal value of the main rotor lift when the collective pitch is fault-free. The main rotor lift caused by collective pitch failure is specifically expressed as follows:

[0126]

[0127] Establish the relationship between rolling torque L, pitching moment M, and periodic moment:

[0128]

[0129] Among them, L a ,L b M a M b A lon A lat B lon B lat Here are system parameters, and ε is the time constant. The fault coefficients for the longitudinal periodic distance and the transverse periodic distance are respectively. In this embodiment, L a =83.88,L b =745.66,

[0130] M a =555.50,M b =11.02; A lat =0.296,A lon =1.945,B lat =2.12,B lon = -0.38; ε = -22.3.

[0131] For ease of writing, the following definition is introduced:

[0132]

[0133] Then equation (4) can be rewritten as:

[0134]

[0135] Among them, L n M n L represents the nominal rolling torque and pitching moment during the period without failure. f M f Rolling torque and pitching moment caused by periodic pitch faults:

[0136]

[0137] Establish the relationship between yaw moment and tail rotor pitch and collective pitch:

[0138]

[0139] Where, N r N col N ped For system parameters, N in this embodiment r =-23.9,N col =8.89,N ped =113.65.δ ped The tail rotor pitch, N is the tail rotor pitch failure factor. nThe nominal yaw moment value is ignoring faults, N f Yaw moment caused by tail rotor failure:

[0140] N n =N r r+N col δ col +N ped δ ped ,

[0141] S2: Based on the established full-loop disturbed dynamics model of the unmanned helicopter, disturbance observers are designed for the position control subsystem and attitude control subsystem respectively:

[0142] The design method for the disturbance observer of the position control subsystem is as follows:

[0143] Define the position command tracking error as: e x =xx r ,e y =yy r ,e z =zz r , where x r ,y r ,z r Assuming the position command, by simultaneously solving equations (1) and (3), we can obtain the dynamic equation for the tracking error of the position subsystem:

[0144]

[0145] The lumped interference includes interference and fault information:

[0146]

[0147] The design method for the disturbance observer of the attitude control subsystem is as follows:

[0148] Define attitude angle tracking error:

[0149]

[0150] Based on equation (2), the system equation for the tracking error of the attitude angle subsystem is obtained:

[0151]

[0152] In the formula, Γ n The nominal values ​​of the torque around the fuselage shaft when there are no faults in the periodic pitch, collective pitch, and tail rotor pitch:

[0153] Γ n =[L n M n N n ]T ,

[0154] The matrix H(Θ) is the result of taking the derivative of each element of H(Θ) with respect to t, as follows:

[0155]

[0156] D in the formula LA ∈R 3×1 The lumped disturbance includes information on disturbances, faults, and the second derivative of the attitude angle command:

[0157] Γ f =[L f M f N f ] T ,

[0158] S3: Lumped disturbance estimation for the position loop and lumped disturbance estimation for the attitude loop are achieved respectively through disturbance observers in the position control subsystem and attitude control subsystem.

[0159] Lumped disturbance estimation for location loop:

[0160] For the external loop position control subsystem (1) of the unmanned helicopter, an intermediate control variable is introduced:

[0161]

[0162] The x-channel extended state observer is designed as follows:

[0163]

[0164] Design an expanded state observer for the y-channel:

[0165]

[0166] Design of the z-channel extended state observer:

[0167]

[0168] In the formula, For the observer parameters, the specific values ​​in this embodiment are: These are the estimated values ​​of the lumped interference for the three channels x, y, and z, respectively.

[0169] Lumped disturbance estimation of the attitude loop:

[0170] For the attitude loop tracking error system of Equation (8), an extended state observer is designed to estimate the lumped disturbance:

[0171]

[0172] In the formula, the observer gain matrix Belongs to R 3×3 The diagonal array, and These represent the derivative of the attitude angle command tracking error and the estimated value of the lumped disturbance, respectively. In this example, the specific values ​​are...

[0173] S4: Design a composite dynamic inverse position controller:

[0174] For the external loop position subsystem, a composite dynamic inverse position controller is designed:

[0175]

[0176] In the above formula, k xp ,k xd ,k yp ,k yd ,k zp ,k zd For controller parameters, the specific value in this embodiment is k. xd =k yd =k zd =4; k xp =k yp =k zp =4.;

[0177] The design method of the composite dynamic inverse attitude controller is as follows:

[0178] To achieve the three-channel decoupling of the attitude angle tracking error system in formula (8), the following definition is introduced:

[0179]

[0180] The attitude angle tracking error system can then be written as:

[0181]

[0182] By combining the lumped disturbance estimate of the observer, a composite dynamic inverse attitude controller is designed:

[0183]

[0184] In the above formula, k φd ,k φp ,k θd ,k θp ,k ψd ,k ψp For the attitude controller parameters, the specific value in this embodiment is: k φd =k θd =k ψd=10; k φp =k θp =k ψp =25.

[0185] S5: The collective control torque and inner loop attitude angle commands are obtained through the composite dynamic inverse position controller.

[0186] The composite dynamic inverse position controller calculates the total control pitch δ from the intermediate control input. col And the inner loop attitude angle command:

[0187]

[0188] S6: Based on the inner loop attitude angle command, the tail rotor pitch, longitudinal cyclic pitch, and lateral cyclic pitch of the unmanned helicopter are obtained through the composite dynamic inverse attitude controller.

[0189] According to definition (16) and the composite dynamic inverse attitude controller (18), we can obtain:

[0190]

[0191] The longitudinal and lateral periodic distances of the unmanned helicopter can be obtained by solving equations (5) and (19):

[0192]

[0193] The tail rotor pitch of the unmanned helicopter can be obtained by solving equation (6) and combining equations (15) and (19):

[0194] δ ped =(N n -N r rN col δ col ) / N ped (twenty one)

[0195] δ ped The tail rotor pitch is δ. lon For longitudinal periodic pitch, δ lat It is a lateral periodic pitch variation.

[0196] Thus, the four actual control variables of the unmanned helicopter are obtained: collective pitch δ col Tail rotor pitch δ ped Longitudinal periodic pitch δ lon Lateral periodic pitch δ lat .

[0197] Based on the above scheme, in order to verify the anti-interference performance of the composite controller proposed in this invention, the proposed method is simulated and verified using MATLAB / Simulink. The Simulink toolbox in the simulation software MATLAB R2018b is used to build a full-loop model of the unmanned helicopter considering actuator failure, and then simulation and experiment are carried out.

[0198] Taking into full account actuator failures and external time-varying disturbances, a baseline Nonlinear Dynamic Inverse Controller (NDIC) was selected as a comparative verification against the Composite Nonlinear Dynamic Inverse Controller (CNDIC) proposed in this invention.

[0199] Initial state of the unmanned helicopter:

[0200] x(0)=0, y(0)=0, z(0)=0

[0201] v x (0)=0,v y (0)=0,v z (0)=0

[0202] φ(0)=0,θ(0)=0,ψ(0)=0

[0203] p(0)=0, q(0)=0;r(0)=0

[0204] Position commands and yaw angle commands:

[0205] 0s < t < 20s:

[0206] x r (t)=1+sin(0.2πt),y r (t)=cos(0.2πt),z r (t)=-0.5-0.5t,ψ r =0;

[0207] 20s < t < 40s:

[0208] x r (t)=1+sin(0.2πt),y r (t)=cos(0.2πt),z r (t)=-10.5,ψ r =0.

[0209] Actuator fault settings:

[0210] 0s<t≤10s:

[0211] 10s<t≤30s:

[0212] 30s<t≤40s:

[0213] External gust interference settings:

[0214]

[0215]

[0216] The specific simulation results and analysis are as follows:

[0217] Figures 2 to 4 The figures show the tracking effect of the unmanned helicopter on the command trajectory in three-dimensional space under the action of the composite dynamic inverse controller (CNDIC) and the reference dynamic inverse controller (NDIC), the response curves of the position three-channel and heading angle channel commands, and the command tracking error curve. It can be seen that the composite dynamic inverse control method proposed in this invention significantly reduces the impact of multi-source interference and actuator failure on control accuracy, and the anti-interference performance is significantly improved compared with the traditional dynamic inverse control method.

[0218] Figure 5 The control input response curves of the unmanned helicopter under the action of the composite dynamic inverse controller (CNDIC) and the reference dynamic inverse controller (NDIC) are presented. It can be seen that when the actuator efficiency decreases, the amplitude of the control quantity of the composite dynamic inverse control method proposed in this invention is correspondingly increased, compensating for the impact of the fault.

[0219] In summary, this invention can ensure that unmanned helicopters can more accurately track complex trajectories even under the influence of actuator failure and multi-source interference.

Claims

1. An active fault-tolerant and anti-interference trajectory tracking control method for an unmanned helicopter considering actuator failures, characterized in that: The steps include: S1: Establish a full-loop disturbed dynamic model of an unmanned helicopter; S2: Based on the established full-loop disturbed dynamics model of the unmanned helicopter, the disturbance observers of the position control subsystem and the attitude control subsystem are designed respectively; S3: realizing the lumped interference estimation of the position loop and the lumped interference estimation of the attitude loop respectively through the interference observers of the position control subsystem and the attitude control subsystem; S4: Combined with the lumped disturbance estimation of the position loop, a composite dynamic inverse position controller is designed; Combined with the lumped disturbance estimation of the attitude loop, a composite dynamic inverse attitude controller is designed; S5: The control quantity total distance and inner loop attitude angle command are obtained through the composite dynamic reverse position controller; S6: According to the inner loop attitude angle command, the tail rotor pitch, longitudinal cyclic pitch and lateral cyclic pitch of the unmanned helicopter are obtained through the composite dynamic inverse attitude controller; The disturbed full-loop dynamic model of the unmanned helicopter in step S1 includes a disturbed position loop dynamic model of the unmanned helicopter including an actuator fault, an attitude loop dynamic model affected by multi-source interference, and an actuator model affected by the fault, and the establishment process is as follows: A1: Establish a dynamic model of the disturbed unmanned helicopter position loop including actuator failure: Where P = [x, y, z] T , V=[v x ,v y ,v z ] T Respectively represent the position and velocity of the unmanned helicopter in the inertial system, m is the mass of the unmanned helicopter, g is the gravitational acceleration constant, and e=[0,0,1] T is a three-dimensional vector, Θ=[φ,θ,ψ] T is the helicopter attitude angle, φ, θ, ψ are the roll angle, pitch angle and yaw angle respectively, T m is the main rotor lift, F d ∈R 3×1 is multi-source interference, R(Θ) is the transformation matrix from the body axis system to the inertial system; A2: Establishment of attitude loop dynamics model affected by multiple sources of interference: Where Ω = [p, q, r] T is the angular velocity of rotation around the body axis, p, q, r are the roll, pitch, and yaw angular rates respectively, J = diag{I x ,I y ,I z } is the moment of inertia matrix, Γ=[L,M,N] T is the three-axis moment, L, M, N are rolling moment, pitch moment and yaw moment respectively, D A ∈R 3×1 is the multi-source interference of the attitude loop, H(Θ) is the transformation matrix; A3: Model the actuators affected by the fault: Establish the relationship between main rotor lift and collective pitch: Among them, δ col is the main rotor collective pitch, Z w and Z col is the coefficient, is the total distance failure coefficient, w is the speed of the helicopter along the z-axis of the fuselage, is the nominal value of the main rotor lift when there is no collective pitch fault, is the main rotor lift caused by the collective pitch failure, and the specific expression is: Establish the relationship between rolling torque L, pitching moment M and periodic moment: Among them, L a ,L b ,M a ,M b ,A lon ,A lat ,B lon ,B lat is the system parameter, ε is the time constant, are the fault coefficients of the longitudinal period distance and the transverse period distance respectively, and the following definitions are introduced: Then equation (4) can be rewritten as: Among them, L n ,M n is the nominal rolling torque and pitching moment when the cyclic distance is fault-free, L f ,M f Roll torque and pitch moment caused by cyclic fault: Establish the relationship between yaw moment, tail rotor pitch and total pitch: Among them, N r ,N col ,N ped is the system parameter, δ ped is the tail rotor pitch, is the tail rotor pitch failure coefficient, N n is the nominal value of the yaw moment without considering the fault, N f Yaw moment caused by tail rotor failure: The design method of the disturbance observer of the position control subsystem in step S2 is: Define the position command tracking error as: e x =xx r ,e y =yy r ,e z =zz r , where x r ,y r ,z r is the position command, and the tracking error dynamic equation of the position subsystem can be obtained by combining equations (1) and (3): The aggregate interference includes interference and fault information: The design method of the disturbance observer of the attitude control subsystem is: Define attitude angle tracking error: According to equation (2), the tracking error system equation of the attitude angle subsystem is obtained: In the formula, Γ n is the nominal value of the torque around the fuselage axis when there is no fault in the cyclic pitch, collective pitch and tail rotor pitch: C n =[L n M n N n ] T is the matrix after the derivative of each element of the H(Θ) matrix with respect to t, as follows: D in the formula LA ∈R 3×1 is the aggregate disturbance, including the disturbance, fault and the second-order derivative information of the attitude angle command: The step S3 is specifically as follows: Aggregate interference estimation of the position loop: For the unmanned helicopter outer loop position control subsystem of formula (1), in order to simplify the writing, the intermediate control quantity is introduced: The x-channel extended state observer is designed as: Design the y-channel expanded state observer: Expand the z-channel state observer design: In the formula, are observer parameters, are the estimated values ​​of the aggregate interference of three channels x, y, and z respectively; Aggregate interference estimation of attitude loop: For the attitude loop tracking error system of formula (8), an extended state observer is designed to estimate the lumped disturbance: Where the observer gain matrix For R 3×3 The diagonal matrix of and are respectively the derivative of the attitude angle command tracking error and the estimated value of the lumped disturbance; The design method of the composite dynamic reverse position controller in step S4 is: For the outer loop position subsystem, a composite dynamic inverse position controller is designed: In the above formula, k xp ,k xd ,k yp ,k yd ,k zp ,k zd is the controller parameter; The design method of the composite dynamic inverse attitude controller is: In order to realize the three-channel decoupling of the attitude angle tracking error system in formula (8), the following definition is introduced: Then the attitude angle tracking error system can be written as: Combined with the lumped disturbance estimator of the observer, a composite dynamic inverse attitude controller is designed: In the above formula, k φd ,k φp ,k θd ,k θp ,k ψd ,k ψp are the attitude controller parameters; The step S5 is specifically as follows: The composite dynamic inverse position controller calculates the total distance δ of the control quantity through the intermediate control quantity col And the inner loop attitude angle command: The step S6 is specifically as follows: According to definition (16) and the composite dynamic inverse attitude controller (18), we can obtain: According to equation (5) and equation (19), the longitudinal period and lateral period of the unmanned helicopter are obtained: According to equation (6) and combined with equations (15) and (19), the tail rotor pitch of the unmanned helicopter is obtained: d ped =(N n -N r rN col d col ) / N ped (21) δ ped is the tail rotor pitch, δ lon is the longitudinal periodic pitch, δ lat It is the lateral periodic pitch variation.

2. The active fault-tolerant and anti-interference trajectory tracking control method for an unmanned helicopter considering actuator failure according to claim 1 is characterized in that: In step A1, R(θ) is the transformation matrix from the body axis system to the inertial system, and its specific form is: Among them, cosφ=cφ, sinφ=sφ, cosθ=cθ, sinθ=sθ, cosψ=cψ, sinψ=sψ.

3. The active fault-tolerant and anti-interference trajectory tracking control method for an unmanned helicopter considering actuator failure according to claim 1 is characterized in that: In step A2, H(θ) is a transformation matrix, which is specifically expressed as: Among them, sinφ=sφ, cosφ=cφ, tanθ=tθ, sinθ=sθ.

Images