A method for controlling the attitude of a tri-cowl aircraft with a heading rudder

Abstract

The application discloses a kind of three-cascade aircraft attitude control methods with rudder, and the method comprises the following steps: step one, establishing body coordinate system;Step two, establishing the control efficiency model of three-cascade aircraft with rudder;Step three, establishing the control allocation method of three-cascade aircraft with rudder;Step four, establishing the attitude control method of three-cascade aircraft with rudder;Step five, the expected value calculated in step four is substituted into the control allocation matrix designed in step three, to obtain the final control command.The control method proposed in the application can effectively control the attitude of three-cascade aircraft with rudder.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of control of multi-rotor aerial vehicles, and particularly to a three-cascade aircraft attitude control method with rudders. BACKGROUND

[0002] Ground transportation needs a large amount of and long time of infrastructure construction investment in the development process. In the city, the ground transportation is easily affected by various unexpected situations, and the operation efficiency is affected. For remote areas, in the face of some unexpected situations, it is also difficult to quickly develop ground transportation infrastructure construction in a short time, and the cost is relatively high. Therefore, low-altitude transportation has more prominent advantages in these situations. Especially, the low-altitude transportation can cope with the sudden transportation demand of the city and the remote area, such as the transportation of medical supplies, instruments, and wounded personnel, and the rapid mobilization of materials and rescue personnel in disaster areas.

[0003] At present, the application scene of low-altitude transportation is usually designed based on the configuration of four-rotor, six-rotor, eight-rotor and other rotor aircraft. However, with the increase of the number of rotors, the volume also increases under the same load. The excessive volume will limit the application scene of the low-altitude aircraft, and too many rotors will also limit the actual maximum flight distance. On the other hand, due to the configuration of the dual-rotor aircraft, the maneuverability is poor, and it is difficult to provide sufficient maneuverability during flight to cope with various disturbances that may occur in low altitude. Based on these reasons, the three-rotor aircraft can minimize the volume of the aircraft while considering the maneuverability, and balance the maneuverability and volume of the rotor aircraft. It is expected to be more widely used in future low-altitude transportation. The existing three-rotor aircraft is provided with two conventional rotors in front of the body and a tiltable rotor at the rear, so as to realize the control of yaw, but this configuration is more suitable for small unmanned aerial vehicles. In view of the large load demand required by low-altitude, the volume is directly increased based on the existing three-rotor, and a larger torque is required to control the larger tiltable rotor, which puts higher requirements on the control equipment. At the same time, the delay caused by the tilting of the large rotor is larger, and the control design is more restricted. SUMMARY

[0004] The purpose of the present application is to provide a three-cascade aircraft attitude control method with rudders, which realizes the attitude control of the three-cascade aircraft with rudders.

[0005] The technical solution for achieving the purpose of the present application is: a three-cascade aircraft attitude control method with rudders, comprising:

[0006] Step one, establishing a body coordinate system;

[0007] Step two, establishing a control efficiency model based on the geometric layout of the power system of the three-cascade aircraft;

[0008] Step three, based on the control efficiency model, a control allocation method for the three-cannula aircraft with rudder is established, and the input is the output of the attitude control;

[0009] Step four, based on the control theory, an attitude control method for the three-cannula aircraft with rudder is established, and the output is the input of the control allocation method;

[0010] Step five, based on the expected total thrust, expected roll moment, expected pitch moment and expected yaw moment calculated in step four, the control allocation matrix designed in step three is substituted to obtain the final control command.

[0011] An electronic device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, wherein the processor executes the program to realize the above-mentioned attitude control method for the three-cannula aircraft with rudder.

[0012] A computer readable storage medium having a computer program stored thereon, wherein the program is executed by a processor to realize the above-mentioned attitude control method for the three-cannula aircraft with rudder.

[0013] A computer program product comprising a computer program, wherein the computer program is executed by a processor to realize the above-mentioned attitude control method for the three-cannula aircraft with rudder.

[0014] Compared with the prior art, the present application has the following advantages: the present application can efficiently and quickly control the yaw channel of the three-cannula aircraft by using the tilting rudder surface below the two cannulas. The present application has a significant difference between the yaw moment generated by the yaw rudder surface and the yaw moment (also known as counter torque) generated by the rotor, and the counter torque of the three-cannula and the yaw moment generated by the yaw rudder surface are designed separately to realize more accurate control. BRIEF DESCRIPTION OF DRAWINGS

[0015] Figure 1 It is a geometric diagram of the three-cannula unmanned aerial vehicle with rudder.

[0016] Figure 2 It is a diagram showing the included angle when observing the tilting rudder surface from the center of mass of the aircraft to the rudder surface.

[0017] Figure 3 It is a diagram of the attitude controller and the control allocation structure.

[0018] Figure 4 It is a pitch response curve diagram.

[0019] Figure 5 It is a roll response curve diagram.

[0020] Figure 6is a yaw response curve schematic diagram.

[0021] Figure 7 is a height response curve schematic diagram. DETAILED DESCRIPTION

[0022] The application proposes a three-cannula aircraft control method based on rudder with heading. The aircraft is based on three-rotor triangular layout, increases yaw multi-faces under the rotor to be responsible for yaw control, and increases the cannula to improve the overall efficiency of the aircraft power system. The method specifically includes the following steps:

[0023] Step one: establish the body coordinate system.

[0024] The coordinate system is fixed to the body, as shown in Figure 1 . The three rotors of the unmanned aerial vehicle are numbered as rotor 1, rotor 2, and rotor 3. The center of gravity of the body is taken as the origin o b , the forward direction is the positive direction of the x b axis, the right direction is the positive direction of the y b axis, and the downward direction is the positive direction of the z b axis. Two groups of rudders are installed under the rotor 1 and the rotor 2, and the plane of the rudder is just through the center of mass.

[0025] Step two: establish the control efficiency model of the three-cannula aircraft with rudder. According to Figure 1 , the control efficiency model is established as:

[0026]

[0027] Where, {F x F y F z} respectively represent the tension acting on the body {x b y b z b} axis; {τ x τ y τ z} respectively represent the moment acting on the three coordinate axes of the body coordinate system; {T1 T2 T3} respectively represent the thrust generated by the three rotors themselves, and the range is [0T max ], T max represents the maximum thrust that each rotor can generate; represents the angle between rotor 1 or rotor 2 and the x b axis of the body coordinate system; represents the distance of the motor mounting center of rotor 1 or rotor 2 relative to the x b axis of the body along the y b axis; The line connecting the motor mounting centers of rotors 1 and 2 is shown along the x-axis relative to the center of mass of the aircraft. b Distance between axes; This indicates the x-axis distance between the motor mounting center of rotor 3 and the aircraft's center of mass. b The distance; δ1 and δ2 represent the angles between the control surfaces of rotor 1 and rotor 2 and the vertical direction of the body coordinate system, respectively, such as Figure 2 As shown, the symbol definition of the vertical angle between the control surface corresponding to rotor 1 and the body coordinate system is illustrated; c represents the proportional coefficient of the thrust-to-torque generated by the rotor.

[0028] Step 3: Establish a control distribution method for a three-ducted aircraft with a rudder.

[0029] For a three-ducted rotorcraft with a turret, its pitch and roll control is similar to that of a common multi-rotor aircraft. The difference lies in the addition of an extra turret for directional control. Simultaneously, the counter-torque of each ducted rotor also affects the yaw motion of the aircraft. Considering the difference between the mechanism of rotor counter-torque's effect on directional control and the mechanism of the turret's effect on directional control, the control allocation matrix designed in this invention separates the turret and rotor counter-torque, allowing for separate setting of turret and counter-torque control parameters in the attitude controller. The control allocation method is represented by a linear matrix as follows:

[0030]

[0031] Among them, F z,d To act on the organism z b Force on the axis; {τ x,d τ y,d τ z,d} represent the expected torques of the three ducted rotors acting on the three axes of the airframe coordinate system; it should be noted that τ z,d It is the expected anti-torque achieved by the three rotors; τ z,δ,d The yaw moment is the yaw force generated by the desired rudder.

[0032] Step 4: Establish an attitude control method for a three-ducted aircraft with a yaw.

[0033] The desired input for attitude control is the desired roll angle φ. d Pitch angle θ d Yaw angle ψ d And the expected height p z,d The feedback inputs for attitude control are the current roll angle φ, pitch angle θ, yaw angle ψ, roll rate p, pitch rate q, yaw rate r of the three-ducted aircraft, and the current altitude p of the aircraft. z and speed v z{F x,d F y,d F z,d}, desired moments about the three axes of the body frame {τ x,d τ y,d τ z,d}, and desired yaw moment τ z,δ,d generated by the rudder. The outputs of the attitude controller are used as inputs to the control allocation. The structure of the controller and the control allocation is shown in Figure 3 . The specific steps to construct the attitude controller are as follows:

[0034] Step 4.1: Calculate the desired body angular rates {p d q d r d} based on the desired roll angle, pitch angle, yaw angle and the current roll angle, pitch angle, yaw angle of the trijet aircraft.

[0035]

[0036] where, denotes the numerical differentiation of ψ d , denotes the numerical differentiation of ψ are adjustable parameters.

[0037] Step 4.2: Calculate the desired moments based on the desired body angular rates {p d q d r d} and the current body angular rates {p q r}.

[0038]

[0039] where, denotes the numerical differentiation of p d , denotes the numerical differentiation of p denotes the numerical differentiation of q d , denotes the numerical differentiation of q are adjustable parameters.

[0040] Step 4.3: Calculate the desired lift force F z,d acting on the z z axis of the trijet aircraft based on the desired flight altitude p z and the current altitude p b and the vertical velocity v z,d .

[0041]

[0042] in, All are adjustable parameters, and sat represents the saturation function, defined as follows:

[0043]

[0044] Step 5: Calculate the expected value F obtained in Step 4. z,d , τ x,d , τ y,d , τ z,d , τ z,δ,d Substitute these values ​​into the control allocation matrix (2) designed in step three to obtain the final control commands T1, T2, T3, δ1, δ2.

[0045] The present invention will now be described in further detail with reference to the accompanying drawings.

[0046] Example

[0047] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. All other embodiments obtained by those skilled in the art based on the embodiments in this application without inventive effort are within the scope of protection of this application.

[0048] Step 1: Establish a rigid body simulation model

[0049]

[0050] Where, p = [p x p y p z ] T Indicates the position of the aircraft; v = [v x v y v z ] T The velocity of the aircraft is represented by g; g = 9.81 represents the acceleration due to gravity; e3 =

[001] T ;R eb Represents the rotation matrix from the body coordinate system to the inertial coordinate system; F = [F x F y F z ] T Θ represents the force vector acting on the aircraft's coordinate system; m represents the mass of the aircraft at takeoff; Θ = [φ θ ψ] T This represents a vector consisting of three Euler angles; ω = [pqr] T Indicates the angular velocity of the organism; These represent the differentials of the aircraft position, aircraft velocity, Euler angle, and aircraft angular rate, respectively.

[0051]

[0052] Represents the relationship between the body's angular velocity ω and the Euler angular velocity. The transformation relationship; τ=[τ x τ y τ z ] T This represents the torque vector acting on the body; J = diag([J xx J yy J zz ]) represents the inertial tensor of the aircraft.

[0053] Step 2: Based on the rigid body model (6) and the control efficiency model (1), a simulation environment for a three-ducted aircraft with a rudder was built using Simulink software. The model parameters were selected as follows: m = 180 kg, J = diag([554750]) kg·m 2 , Ψ=0.84rad, L1=0.898m, L2=0.383m, L3=0.825m, c=0.0389.

[0054] Step 3: Construct the controller based on controller equations (2), (3), (4), and (5), where the parameter values ​​are: k roll,p =10,k pitch,p =10,k yaw,p =20,k yaw,d =4,k rollrate,p =30,k rollrate,i =0.05, k rollrate,d =15,k pitchrate,p =40,k pitchrate,i =0.1, k pitchrate,d =15,k yawrate,p =0,k yawrate,i =0,k yawrate,p2 =30,k yawrate,i2 =0,k p,z =1,k v,z,p =800,k v,z,i =0.5.

[0055] Step 4: Given the desired pose φ d θ d , ψ d And the expected height p z,d Simulation begins. Results are as follows: Figure 4 , Figure 5 , Figure 6 , Figure 7As shown, the aircraft’s pitch, roll, yaw and altitude channels all achieved good tracking performance, and the designed controller can effectively and stably control the three-channel aircraft with a yaw rudder.

Claims

1. A method for attitude control of a three-ducted aircraft with a rudder, characterized in that, include: Step 1: Establish the body coordinate system, as follows: The three rotors of the drone are numbered as Rotor 1, Rotor 2, and Rotor 3, with the drone's center of gravity as the origin. Forward The positive direction of the axis is to the right. The positive direction of the axis is downward. In the positive direction of the axis; two sets of rudders are installed below rotor 1 and rotor 2 respectively, and the plane of the rudder surface passes through the center of mass; Step 2: Based on the geometric layout of the propulsion system of the three-ducted aircraft, establish a control efficiency model; Step 3: Based on the control efficiency model, establish a control allocation method for a three-ducted aircraft with a yaw rudder, the input of which is the output of attitude control; The control allocation method for establishing a three-ducted aircraft with a rudder is as follows: The control allocation matrix separates the rudder and rotor counter-torque, allowing the attitude controller to set the control parameters for the rudder and counter-torque separately. The control allocation method is represented by a linear matrix as follows: ；（2） in, In order to act on the body Force on the axis; These represent the expected torques of the three ducted rotors acting on the three axes of the airframe coordinate system; The desired yaw moment generated by the rudder; These represent the thrust generated by the three rotors themselves, and their ranges are... , This indicates the maximum thrust that each rotor can generate; This indicates the motor mounting center of rotor 1 or rotor 2 relative to the fuselage. Along the axis Distance between axes; The line connecting the motor mounting centers of rotor 1 and rotor 2 is relative to the center of mass of the aircraft along... Distance between axes; This indicates the distance along the motor mounting center of rotor 3 relative to the aircraft's center of mass. The distance; , These represent the angles between the control surfaces of rotor 1 and rotor 2 and the vertical direction of the airframe coordinate system, respectively. Step 4: Based on control theory, establish an attitude control method for a three-ducted aircraft with a yaw rudder, and use its output as the input for the control allocation method; Step 5: Based on the expected total thrust, expected roll moment, expected pitch moment, and expected yaw moment calculated in Step 4, substitute them into the control allocation matrix designed in Step 3 to obtain the final control command.

2. The attitude control method for a three-ducted aircraft with a yaw as described in claim 1, characterized in that, Establish a control efficiency model for a three-ducted aircraft with a rudder: ； （1） in, These respectively represent the effects on the body Tension on the shaft; These represent the torques acting on the three coordinate axes of the body coordinate system; This indicates the coordinate system of rotor 1 or rotor 2 relative to the body coordinate system. The included angle formed by the axes; It represents the ratio of the thrust to the counter-torque generated by the rotor.

3. The attitude control method for a three-ducted aircraft with a yaw as described in claim 1, characterized in that, An attitude control method for a three-ducted aircraft with a yaw rudder is established, as follows: The desired input for attitude control is the desired roll angle. Pitch angle Yaw angle and the height of expectations The feedback input for attitude control is the current roll angle of the three-ducted aircraft. Pitch angle Yaw angle Roll rate Pitch rate yaw rate and the current altitude of the aircraft and speed The output of the attitude controller is the desired force acting on the three axes of the body coordinate system. The desired torques of the three rotors acting on the three axes of the airframe coordinate system. and the yaw moment expected to be generated by the rudder. The output of the attitude controller serves as the input for control assignment. The specific steps for constructing the attitude controller are as follows: Step 4.1: Calculate the desired airframe angular rate based on the desired roll angle, pitch angle, yaw angle, and the current roll angle, pitch angle, and yaw angle of the three-ducted aircraft. ; ；（3） in, express The numerical derivative, express The numerical derivative; , , , These are adjustable parameters; Step 4.2: Based on the desired body angular velocity and the current body angular velocity Calculate the desired torque; ；（4） in, express The numerical derivative, express The numerical derivative, express The numerical derivative, express The numerical derivative; , , , , , , , , , All parameters are adjustable; Step 4.3: Based on the desired flight altitude and current height and vertical velocity Calculations applied to a three-ducted aircraft Tension on the shaft : ；（5） in, , , All parameters are adjustable. The saturation function is defined as follows: 。 4. The attitude control method for a three-ducted aircraft with a yaw rudder according to claim 3, characterized in that, The expected value calculated in step four , , , , Substituting these values ​​into the control allocation matrix (2) designed in step three, we obtain the final control command. , , , , .

5. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the attitude control method for a three-ducted aircraft with a rudder as described in any one of claims 1-4.

6. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by the processor, the program implements the attitude control method for a three-ducted aircraft with a rudder as described in any one of claims 1-4.

7. A computer program product, comprising a computer program, characterized in that, When executed by a processor, the computer program implements the attitude control method for a three-ducted aircraft with a rudder as described in any one of claims 1-4.

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