Aerial net recovery attitude control method for a rotary-wing aircraft

Abstract

The application discloses a rotorcraft air collision net recovery attitude control method and system. The method provided by the application uses a state observer to observe the attitude angular rate of the rotorcraft in real time, and inputs the attitude angular rate of the rotorcraft into a terminal sliding mode attitude control law, so as to realize the attitude control of the rotorcraft. The application can meet the demand of the attitude control of the rotorcraft in the air collision net recovery, and makes the rotorcraft be able to resist the interference of the damage of the power unit caused by the environment, airflow and collision at the end of the net recovery, so as to be beneficial to realizing the precise, rapid and efficient air collision net recovery.

Description

Technical Field

[0001] The present invention relates to the technical field of rotorcraft control, and in particular to a method and system for controlling the recovery attitude of a rotorcraft during an aerial net collision. Background Art

[0002] The rotorcraft has the characteristics of strong power, good maneuverability and fast command response, and has a certain anti-interference ability during the air net collision recovery process.

[0003] However, in actual engineering applications, the interference faced by rotorcraft may be stronger and come from multiple aspects:

[0004] (1) Aircraft airflow: The net used for net recovery is usually fixed to the bottom of the rotor-type carrier aircraft. The carrier aircraft itself has a large load-bearing capacity, so there is a strong downwash airflow near the net; or when a tail-towed cabin-type net structure is used, the wake (tail vortex) from the carrier aircraft will cause significant interference to the end of the net recovery. This effect often occurs at the end of the net recovery;

[0005] (2) Environment: Wind from the environment, especially lateral gusts, will have a significant impact on the recovery of the net. This impact will occur during the entire recovery process.

[0006] (3) Internal problems: These include the structure, power, and control of the rotorcraft. Power problems with rotorcraft generally occur during flight, including collisions during multi-machine recovery or collisions between the recovery end and the net, resulting in the loss of propellers. These problems need to be solved by improving the control performance of the rotorcraft.

[0007] Therefore, whether the rotorcraft maintains a stable attitude under the above-mentioned interference is related to whether the net collision recovery process is smooth and even to the flight safety of the rotorcraft itself. Summary of the Invention

[0008] In response to the above problems in the prior art, the present invention proposes a method and system for controlling the attitude of a rotorcraft during an aerial net collision recovery.

[0009] In order to achieve the above-mentioned objectives, in a first aspect, the present invention provides a method for controlling the attitude of a rotorcraft during an aerial net collision recovery, which utilizes a state observer to observe the attitude angular rate of the rotorcraft in real time, and inputs the attitude angular rate of the rotorcraft into a terminal sliding mode attitude control law to realize attitude control of the rotorcraft.

[0010] In a second aspect, the present invention provides a rotorcraft airborne net-impact recovery attitude control system. The system comprises:

[0011] A first establishing module is used to establish an attitude dynamics model of the rotorcraft based on the interference received by the rotorcraft;

[0012] a second establishing module, configured to establish a state observer according to the disturbance to which the rotorcraft is subjected;

[0013] The attitude control module is used to obtain the terminal sliding mode attitude control law and control the attitude of the rotorcraft based on the terminal sliding mode attitude control law.

[0014] The beneficial effects of the method and system for controlling the attitude of a rotorcraft recovering from a net collision in mid-air of the present invention include:

[0015] (1) The method provided by the present invention uses a state observer to accurately and quickly estimate the attitude angular rate of the rotorcraft when the rotorcraft is disturbed, and can achieve rapid convergence of the attitude angle error;

[0016] (2) The method provided by the present invention uses terminal sliding mode attitude control to achieve rapid convergence within a limited time, so that the entire attitude control of the rotorcraft has a high response speed; at the same time, when the rotorcraft is disturbed, it has strong anti-interference ability and can maintain attitude stability;

[0017] (3) The method provided by the present invention can meet the requirements of attitude control of the rotorcraft during the recovery of the net in the air, so that the rotorcraft can resist interference such as damage to its own power unit caused by the environment, airflow, and collision at the end of the recovery, which is conducive to achieving accurate, fast and efficient recovery of the net in the air. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 A schematic flow chart of a method for controlling the recovery posture of a rotorcraft during an aerial net collision according to the present invention is shown;

[0019] Figure 2 A schematic structural diagram of a rotorcraft airborne net-collision recovery attitude control system according to the present invention is shown;

[0020] Figure 3 A schematic diagram of interference estimation for three channels of roll, pitch and yaw using a state observer according to embodiment 1 of the present invention is shown;

[0021] Figure 4 FIG. 1 is a schematic diagram showing the estimation of attitude angular rate by the state observer according to Embodiment 1 of the present invention;

[0022] Figure 5 A schematic diagram showing the estimation error of the attitude angular rate by the state observer according to embodiment 1 of the present invention is shown;

[0023] Figure 6A schematic diagram showing a response curve of a quadrotor drone in response to a 25° attitude command according to Example 1 of the present invention is shown;

[0024] Figure 7 A schematic diagram of sinusoidal signal tracking under different interference conditions according to embodiment 1 of the present invention is shown;

[0025] Figure 8 A schematic diagram of attitude angle tracking error according to embodiment 1 of the present invention is shown. DETAILED DESCRIPTION

[0026] The preferred embodiments of the present invention are described in detail below with reference to the accompanying drawings so that the advantages and features of the present invention can be more easily understood by those skilled in the art, thereby making a clearer and more precise definition of the protection scope of the present invention.

[0027] It should be noted that, in this document, the terms "comprises," "includes," or any other variations thereof are intended to encompass non-exclusive inclusion, such that a process, method, article, or apparatus comprising a series of elements includes not only those elements but also other elements not explicitly listed, or elements inherent to such process, method, article, or apparatus. In the absence of further limitations, the elements defined by the phrase "comprises..." do not exclude the presence of other identical elements in the process, method, article, or apparatus comprising the elements.

[0028] The complex interference conditions encountered during the net-impact recovery process place high demands on the rotorcraft's control performance. Maintaining a stable attitude under these conditions is crucial to the success of the net-impact recovery process and even the flight safety of the rotorcraft itself.

[0029] To address the aforementioned issues, the present invention provides a method for controlling the attitude of a rotorcraft during a net-strike recovery. This method utilizes a state observer to observe the rotorcraft's attitude angular rate in real time and inputs this rate into a terminal sliding-mode attitude control law to achieve attitude control of the rotorcraft. This method enables the rotorcraft to withstand interference from the environment, airflow, and damage to its power unit caused by collisions at the end of net-strike recovery, achieving precise, rapid, and efficient net-strike recovery.

[0030] In the present invention, the dynamic characteristics of the optoelectronic pod and the rotorcraft itself are ignored. The rotorcraft is preferably a rotorcraft drone, more preferably a rotorcraft drone with four or more propellers.

[0031] Specifically, the present invention provides a method for controlling the attitude of a rotorcraft during airborne net collision recovery. Figure 1 As shown, the method mainly includes the following steps:

[0032] S101. Establishing an attitude dynamics model of the rotorcraft based on the interference to the rotorcraft.

[0033] Specifically, according to the running trajectory of the rotorcraft during net recovery, the interference received by the rotorcraft is determined, and the interference-based rotorcraft attitude dynamics model is obtained according to characteristic information affecting the attitude angle of the rotorcraft using the interference.

[0034] Preferably, the posture dynamics model can be expressed by formula 1:

[0035]

[0036] Where, Θ represents the attitude angle, which can be obtained by the sensor carried by the rotorcraft; τ b It represents the propeller lift moment, that is, the lift moment generated by the propeller, which can be calculated based on the propeller parameters. τ i represents propeller i torque, f i represents the propeller i lift, l represents the length of the arm, c T is the propeller lift coefficient, W i is the speed of propeller i, i=1,2,3...n, n is the number of propellers; It represents the attitude angular rate (attitude angle change rate), which can be obtained by taking the derivative of the attitude angle with respect to time;

[0037] B1 and B2 are two simplified coefficients.

[0038] H1 and H2 are two simplified coefficients about the conversion matrix C and the moment of inertia J. H1=(C -1 ) T JC -1 , C represents the attitude angular rate and angular rate ω b =[ω x ω y ω z ] T The transformation matrix between them can be expressed as follows:

[0039]

[0040] in, θ and ψ represent the roll angle, pitch angle, and yaw angle, respectively, which can be obtained by sensors carried on the rotorcraft;

[0041] J represents the moment of inertia, which can be expressed as follows:

[0042]

[0043] Among them, J xx 、J yy 、J zz is the moment of inertia of the rotorcraft body on the three axes, which can be obtained through measurement;

[0044] Denotes the diagonal matrix ω b =diag{ω x ,ω y ,ω z} antisymmetric matrix, ω x 、ω y 、ω z They represent the roll, yaw, and pitch angular rates, i.e., the components of the rotorcraft's angular velocity on the three axes of the body, which can be obtained by sensors carried on the rotorcraft;

[0045] D represents the total impact of the disturbance on the system state vector (hereinafter referred to as total disturbance), τ r It represents the total disturbance torque, which is the sum of the internal disturbance torque of the rotorcraft and the external disturbance torque, τ r =τ 内 +τ 外 =τ c +τ f +τ d +(Λ-I)τ b , τ c represents the gyroscopic torque, which can be calculated based on the inherent parameters of the rotorcraft; τ f represents the damping torque of the motor and propeller, which can be calculated based on the inherent parameters of the rotorcraft; τ d represents the environmental interference torque, which can be calculated based on the impact of the environment on the rotorcraft; Λ represents the efficiency matrix of the preset motor and propeller, Λ=diag{Λ i}, i=1,2,3...n, n represents the number of propellers; I represents the unit matrix.

[0046] The robustness and accuracy of terminal sliding mode attitude control are affected by unknown bounded disturbances. Although increasing the parameters can reduce the error, this also increases the conservatism of the control to a certain extent. Therefore, to improve the adaptability to disturbances, the present invention adopts a state observer to estimate the disturbances.

[0047] S102: Establish a state observer according to the interference to the rotorcraft.

[0048] In the present invention, in order to simultaneously estimate the attitude angular rate and the total disturbance of the attitude control system, a finite-time convergent state observer is established.

[0049] Preferably, the state observer can be expressed by Formula 2:

[0050]

[0051] Among them, z1 and z2 are attitude angle Θ and attitude angular rate respectively. ; z3 is the estimated value of the total interference D; k1, k2, k3 represent setting parameters, and k1>0, k3>0, k2>k1 / k3; λ represents setting parameters, and 0<λ<1; m1, m2, m3 represent setting parameters, and 0<m1<1, m2=(2m1+1) / 3, m3=(m1+2) / 3.

[0052] The study found that when λ, k1, k2, k3, m1, m2, and m3 are within the above ranges, the state observer of Formula 2 is used, so that under different interference conditions, the estimation errors of the attitude angular rate and the total interference are small, that is, the state observer has high stability, thereby achieving accurate estimation of the attitude angular rate, total interference, etc.

[0053] By using the above-mentioned state observer, the attitude angular rate of the rotorcraft can be accurately and quickly estimated when the rotorcraft is disturbed; at the same time, when the interference conditions increase, the attitude angular rate error can be quickly converged.

[0054] S103: Acquire a terminal sliding mode attitude control law, and control the attitude of the rotorcraft based on the terminal sliding mode attitude control law.

[0055] In the present invention, the purpose of the terminal sliding mode attitude control law is to control the rotorcraft to quickly readjust to the desired attitude angle after being disturbed, so as to achieve accurate mid-air net recovery.

[0056] Specifically, step S103 may include the following steps:

[0057] S103-1. Obtain a sliding surface according to the posture dynamics model.

[0058] In the present invention, two state variables x1 and x2 of the attitude control system are set, where x1 represents the difference between the current attitude angle and the desired attitude angle, and x2 represents the current attitude angular rate;

[0059] Right now e=Θ-Θ d ,Θ d is the desired attitude angle;

[0060] Furthermore, the error dynamics model of the attitude angle can be expressed as follows:

[0061]

[0062] The present invention is based on sliding mode variable structure control. In sliding mode variable structure control, the selection of sliding mode surface has a great influence on the control effect.

[0063] The terminal sliding mode attitude control adopted in the present invention has the advantages of strong robustness, high precision, fast convergence speed and no singularity problem. Compared with the traditional terminal sliding mode control, it has a higher convergence rate and faster error convergence.

[0064] Specifically, the sliding surface can be expressed by formula 3:

[0065] s=x2+α|x1| β sgn(x1) Equation 3

[0066] Where s represents the sliding surface; α and β represent setting coefficients, α>0, 0<β<1.

[0067] In the present invention, the value of β actually introduces a nonlinear function into the terminal sliding mode attitude control, so that the sliding mode tracking error can converge to zero in a finite time.

[0068] The sliding surface shown in Equation 3 introduces nonlinear terms to improve the dynamics of the attitude control system and increase its convergence speed. Once the attitude control system enters the sliding mode state, it can converge within a finite time.

[0069] S103-2. Set the reaching law.

[0070] In the present invention, a unique convergence law is set to improve the convergence speed of the terminal sliding mode attitude control.

[0071] In order to achieve a finite time to reach the sliding surface and reduce the chattering phenomenon, preferably, the reaching law can be expressed by formula 4:

[0072]

[0073] in, represents the reaching law; σ, p, q are the setting coefficients, σ>0, p>0, 0≤q<1.

[0074] Research has found that the convergence law shown in Equation 4 shortens the sliding mode state approach process through the power term q, while controlling the speed of the attitude control system reaching the sliding surface to avoid being too large, thereby weakening vibration; moreover, the power term q can make the attitude control system converge within a finite time, which is beneficial to the response speed of attitude control.

[0075] S103-3. Obtain a terminal sliding mode attitude control law for the rotorcraft based on the sliding mode surface and the reaching law.

[0076] Specifically, the terminal sliding mode attitude control law of the rotorcraft can be expressed by Equation 5:

[0077]

[0078] From the above formula, we can see that when x1≈0, |x1| β-1 The term will produce a singularity, and |x1| β-1 Modifying it to the above form can avoid the occurrence of singularities.

[0079] in, When θ is a positive constant less than 1 and the value range of θ is 0 to 0.5, the generation of singularity can be avoided and the stability of the terminal sliding mode attitude control can be guaranteed.

[0080] In a second aspect, the present invention also provides a rotorcraft air net collision recovery attitude control system, which uses a state observer to observe the attitude angle of the rotorcraft in real time, and inputs the attitude angle of the rotorcraft into the terminal sliding mode attitude control law to achieve attitude control of the rotorcraft.

[0081] Specifically, if Figure 2 As shown, the system mainly includes:

[0082] A first establishing module 201 is configured to establish an attitude dynamics model of the rotorcraft based on the disturbance to which the rotorcraft is subjected;

[0083] A second establishing module 202 is configured to establish a state observer according to the disturbance to which the rotorcraft is subjected;

[0084] The attitude control module 203 is used to obtain a terminal sliding mode attitude control law and control the attitude of the rotorcraft based on the terminal sliding mode attitude control law.

[0085] The rotorcraft aerial net-collision recovery attitude control system provided by the present invention can be used to execute the rotorcraft aerial net-collision recovery attitude control method described in the first aspect above. Its implementation principle and technical effects are similar and will not be repeated here.

[0086] Preferably, each module in the rotary-wing aircraft air net recovery attitude control system of the present invention can be directly in hardware, in a software module executed by a processor, or in a combination of the two.

[0087] The software modules may reside in RAM memory, flash memory, ROM memory, EPROM memory, EEPROM memory, registers, hard disk, removable disk, CD-ROM, or any other form of storage medium known in the art. An exemplary storage medium is coupled to the processor such that the processor can read information from and write information to the storage medium.

[0088] The processor may be a central processing unit (CPU), other general-purpose processors, digital signal processors (DSPs), application-specific integrated circuits (ASICs), field programmable gate arrays (FPGAs), or other programmable logic devices, discrete gate or transistor logic, discrete hardware components, or any combination thereof. A general-purpose processor may be a microprocessor, but in the alternative, the processor may be any conventional processor, controller, microcontroller, or state machine. The processor may also be implemented as a combination of computing devices, such as a combination of a DSP and a microprocessor, multiple microprocessors, one or more microprocessors combined with a DSP core, or any other such configuration. In the alternative, the storage medium may be integral to the processor. The processor and storage medium may reside in an ASIC. The ASIC may reside in a user terminal. In the alternative, the processor and storage medium may reside as discrete components in the user terminal.

[0089] Example

[0090] Example 1

[0091] The simulation experiment was carried out, and the two main considerations during the simulation process were the interference of the quadcopter's torque caused by gusts of wind and the interference of downwash airflow and power unit failure on the power. d =J[1+sin5t,1+cos5t,0.5(1+sin5t+cos5t)] T Nm,

[0092] J represents the moment of inertia, which is expressed as:

[0093] J xx =0.08698kg.m 2 , J yy =0.08698kg.m 2 , J zz =0.13652kg.m 2 At the 10th second of the simulation, the efficiency matrix of the motor and propeller in the drone is changed to Λ=diag{Λ i}, i = 1, 2, 3, 4, the power drops to 85% of the original power instantly during the net recovery process.

[0094] According to the interference to the quadrotor UAV, an attitude dynamics model of the quadrotor UAV based on interference is established as shown in Formula 1.

[0095]

[0096] Where, Θ represents the attitude angle, which is obtained by the sensor carried by the quadrotor drone; τ b represents the propeller lift torque, τ i represents propeller i torque, f i represents the propeller i lift, l represents the length of the arm, c T is the propeller lift coefficient, W i is the speed of propeller i, i = 1, 2, 3, 4;

[0097] B1 and B2 are two simplified coefficients.

[0098] H1 and H2 are two simplified coefficients about the conversion matrix C and the moment of inertia J. H1=(C -1 ) T JC -1 , C represents the attitude angular rate and angular rate ω b =[ω x ω y ω z ] T The transformation matrix between them is expressed as follows:

[0099]

[0100] in, θ and ψ represent the roll, pitch, and yaw angles, respectively, which are obtained by the sensors carried by the quadrotor drone;

[0101] J represents the moment of inertia, which is expressed as follows:

[0102]

[0103] Among them, J xx 、J yy 、J zz is the moment of inertia of the quadrotor drone on the three axes;

[0104] Denotes the diagonal matrix ω b =diag{ω x ,ω y ,ω z} antisymmetric matrix, ω x 、ω y、ω z denote the roll, yaw and pitch angular rates respectively;

[0105] D represents the total interference, τ r Represents the total interference torque, which is the sum of the internal interference torque and the external interference torque of the quadrotor drone, τ r =τ 内 +τ 外 =τ c +τ f +τ d +(Λ-I)τ b , τ c represents the gyroscopic torque; τ f represents the damping torque of the motor and propeller; τ d represents the environmental interference torque; Λ represents the efficiency matrix of the motor and propeller, Λ=diag{Λ i}; I represents the identity matrix.

[0106] According to the interference to the quadrotor UAV, the state observer shown in Equation 2 is established.

[0107]

[0108] Among them, z1 and z2 are attitude angle Θ and attitude angular rate respectively. is the estimated value of ; z3 is the estimated value of the total interference D; k1=8, k3=6, k2=12; λ=0.3; m1=0.6, m2=(2m1+1) / 3, m3=(m1+2) / 3.

[0109] The sliding surface is obtained according to the posture dynamics model.

[0110] Among them, two state variables x1 and x2 of the attitude control system are set, where x1 represents the difference between the current attitude angle and the desired attitude angle, and x2 represents the current attitude angular rate;

[0111] Right now e=Θ-Θ d ,Θ d is the desired attitude angle;

[0112] The error dynamics model of the attitude angle is expressed as follows:

[0113]

[0114] The sliding surface s is expressed by formula 3:

[0115] s=x2+α|x1|βsgn(x1) Equation 3

[0116] Among them, α=1, β=0.5.

[0117] Convergence Law It is expressed by formula 4:

[0118]

[0119] Among them, σ=0.5, p=110, q=0.1.

[0120] According to the sliding mode surface and the reaching law, the terminal sliding mode attitude control law of the quadrotor UAV is obtained as shown in Equation 5.

[0121]

[0122] in, θ=0.1.

[0123] The specific simulation results of the quadrotor drone are as follows: Figures 3 to 8 shown.

[0124] Example 2

[0125] The simulation process is similar to that of Example 1, with the only difference being that the power drops to 70% of the original power during the net recovery process. Figures 3 to 8 shown.

[0126] from Figure 3 It can be seen from the figure that under different disturbance conditions, the state observer of the present invention can achieve a better estimation of the disturbance torque.

[0127] from Figure 4 、 5 It can be seen that the state observer of the present invention can accurately and quickly estimate the attitude angular rate of the quadrotor drone when it is disturbed. As the disturbance increases, the error in the attitude angular rate in the state observer increases. In particular, at the instant of power drop, when the power drops to 70%, the peak error in the state observer's roll angular rate reaches 0.04 rad / s, and the peak error in the pitch angular rate reaches 0.2 rad / s. However, the state observer converges within 1 second, and the error remains within 0.01 rad / s.

[0128] Figure 6 In the above equation, the attitude response is divided into two parts: the response to the sinusoidal signal and the response to the step signal. The expected sinusoidal signal is θ d =25cos(0.2πt)°, ψ d =15sin(0.2πt)°; the step signal expected to respond to the three channels is 25°. Figure 6As can be seen, the response convergence times for the quadrotor drone of the present invention in the roll, pitch, and yaw channels are 1.3s, 1.3s, and 1.5s, respectively. The quadrotor's attitude converges quickly within a limited timeframe, thus offering a high response speed. Furthermore, when the quadrotor is subject to interference such as downwash at the net recovery terminal, it maintains strong anti-interference capabilities and maintains attitude stability even when power is reduced to 85% and 70%.

[0129] from Figure 7 As can be seen, the present invention has a good effect on the attitude control of the quadrotor drone, tracking the desired terminal sliding mode attitude control law within 1.5 seconds. When the quadrotor drone power is instantly reduced to 85% and 70% of the original, the present invention demonstrates strong anti-interference ability.

[0130] Combine Figure 8 It can be seen that at the 10th second, when the power instantaneously dropped to 70% of the original, the tracking errors of the attitude angles showed large differences, among which a large tracking error of about 1.6° appeared in the pitch channel. However, under the method of the present invention, the tracking error of the attitude angle can achieve rapid convergence.

[0131] The present invention has been described in detail above with reference to specific embodiments and exemplary examples. However, these descriptions are not to be construed as limiting the present invention. Those skilled in the art will appreciate that, without departing from the spirit and scope of the present invention, various equivalent substitutions, modifications, or improvements may be made to the technical solutions and implementations of the present invention, all of which fall within the scope of the present invention.

Claims

1. A method for controlling the recovery attitude of a rotorcraft after hitting a net in the air, characterized in that: The state observer is used to observe the attitude angular rate of the rotorcraft in real time, and the attitude angular rate of the rotorcraft is input into the terminal sliding mode attitude control law to realize the attitude control of the rotorcraft. The method comprises the following steps: S101, establishing an attitude dynamics model of the rotorcraft based on the interference to the rotorcraft; S102. Establish a state observer based on the interference to the rotorcraft; S103, obtaining a terminal sliding mode attitude control law, and controlling the attitude of the rotorcraft based on the terminal sliding mode attitude control law, In step S102, the state observer is expressed by formula 2: Wherein, z1 and z2 are the estimated values ​​of attitude angle and attitude angular rate respectively; z3 is the estimated value of total interference; λ, k1, k2, k3, m1, m2, and m3 are all setting parameters and are constants; Θ represents the attitude angle; τ b represents the propeller lift moment; Indicates attitude angular rate; B1 and B2 are two simplified coefficients, where H1 and H2 are two simplified coefficients about the transformation matrix and the moment of inertia, where H1 = (C -1 ) T JC -1 , Denotes the diagonal matrix ω b =diag{ω x ,ω y ,ω z } antisymmetric matrix, ω x 、ω y 、ω z They represent the roll, yaw and pitch angular rates respectively; C represents the conversion matrix between attitude angular rate and angular rate; J represents the moment of inertia; In step S103, the terminal sliding mode attitude control law is expressed by equation 5: Formula 5 in, is a positive constant less than 1; s represents the sliding surface; α and β represent the setting coefficients; x1 and x2 represent the two state variables of the attitude control system; σ, p, and q are setting coefficients.

2. The method for controlling the recovery attitude of a rotorcraft after hitting a net in the air according to claim 1, characterized in that: In step S101, the posture dynamics model is expressed by formula 1: Where, Θ represents the attitude angle; τ b represents the propeller lift moment; Indicates attitude angular rate; B1 and B2 are two simplified coefficients, where H1 and H2 are two simplified coefficients about the transformation matrix and the moment of inertia, where H1 = (C -1 ) T JC -1 , Denotes the diagonal matrix ω b =diag{ω x ,ω y ,ω z } antisymmetric matrix, ω x 、ω y 、ω z They represent the roll, yaw and pitch angular rates respectively; C represents the conversion matrix between attitude angular rate and angular rate; J represents the moment of inertia; D represents the total interference, where τ r It represents the total disturbance torque, which is the sum of the internal disturbance torque of the rotorcraft and the external disturbance torque, τ r =τ 内 +τ 外 =τ c +τ f +τ d +(Λ-I)τ b , τ c represents the gyroscopic torque; τ f represents the damping torque of the motor and propeller; τ d represents the environmental interference torque; Λ represents the efficiency matrix of the motor and propeller; I represents the unit matrix.

3. The method for controlling the recovery attitude of a rotorcraft after hitting a net in the air according to claim 1, characterized in that: k1>0, k3>0, k2>k1 / k3; and 0<m1<1 ， m2=(2m1+1) / 3,m3=(m1+2) / 3。 4. The method for controlling the recovery attitude of a rotorcraft after hitting a net in the air according to claim 1, wherein: In step S103, the sliding surface is expressed by equation 3: s=x2+α|x1| β sgn(x1) Equation 3 Where s represents the sliding surface; α and β represent setting coefficients, α>0, 0<β<1; x1 and x2 represent the two state variables of the attitude control system.

5. The method for controlling the recovery attitude of a rotorcraft after hitting a net in the air according to claim 4, characterized in that: x1 and x2 are expressed by the following formula: e=Θ-Θ d ,Θ d is the desired attitude angle.

6. The method for controlling the recovery attitude of a rotorcraft after hitting a net in the air according to claim 4, characterized in that: In step S103, the reaching law is expressed by equation 4: in, represents the reaching law; σ, p, q are the setting coefficients, σ>0, p>0, 0≤q<1.

7. A rotorcraft air strike net recovery attitude control system, which implements the rotorcraft air strike net recovery attitude control method according to any one of claims 1 to 6, characterized in that: include: A first establishing module is used to establish an attitude dynamics model of the rotorcraft based on the interference received by the rotorcraft; a second establishing module, configured to establish a state observer according to the disturbance to which the rotorcraft is subjected; The attitude control module is used to obtain the terminal sliding mode attitude control law and control the attitude of the rotorcraft based on the terminal sliding mode attitude control law.

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