Attitude control method for three-duct aircraft with heading rudder

Abstract

The invention discloses an attitude control method for a three-duct aircraft with a heading rudder. The method comprises the following steps: step 1, establishing an aircraft body coordinate system; 2, establishing a control efficiency model of the three-duct aircraft with the heading rudder; step 3, establishing a control distribution method of the three-duct aircraft with the heading rudder; 4, establishing an attitude control method of the three-duct aircraft with the heading rudder; and 5, substituting the expected value calculated in the step 4 into the control distribution matrix designed in the step 3 to obtain a final control instruction. According to the control method provided by the invention, the attitude of the three-duct aircraft with the heading rudder can be effectively controlled.

Description

Technical Field

[0001] The invention relates to the field of multi-rotor aircraft control, in particular to a three-duct aircraft attitude control method with a heading rudder. Background Art

[0002] In the process of development, ground transportation requires a large amount of long-term investment in infrastructure construction. In cities, ground transportation is easily restricted by various unexpected situations, which affects the operating efficiency. In remote areas, in the face of some emergencies, it is impossible to quickly carry out the construction of ground transportation infrastructure in a short period of time, and the cost is relatively high. Therefore, low-altitude transportation has more prominent advantages in these situations. In particular, low-altitude transportation can cope with sudden transportation needs in cities and remote areas, such as the transportation of medical drugs, equipment, and wounded people, and the rapid mobilization of materials and rescue personnel in disaster-stricken areas.

[0003] At present, the application scenarios of low-altitude transportation are usually designed based on the basic configuration of rotorcraft such as quad-rotors, hexa-rotors, and octa-rotors. However, as the number of rotors increases, its volume increases under the same load. Too large a volume will limit the application scenarios of low-altitude aircraft, and too many rotors will also limit its actual maximum flight distance. On the other hand, due to its configuration, the dual-rotor aircraft has poor maneuverability and cannot provide sufficient maneuverability during flight to cope with various interferences that may occur in the low altitude. For these reasons, the three-rotor aircraft can minimize the volume of the aircraft while taking into account the maneuverability, and balance the maneuverability and volume of the rotorcraft. And it is expected to be more widely used in low-altitude transportation in the future. The existing three-rotor aircraft will have two conventional rotors in front of the fuselage and a tilting rotor at the rear to achieve yaw control, but this configuration is more suitable for micro-UAVs. In order to meet the heavy load demand at low altitude, the volume of the existing three-rotor is directly increased. In order to control the larger tilt-rotor, a larger torque is required, which puts higher requirements on the control equipment. At the same time, the tilt of the large rotor brings greater delay and imposes greater constraints on the control design. Summary of the invention

[0004] The object of the present invention is to provide a method for attitude control of a three-ducted aircraft with a rudder, so as to realize attitude control of the three-ducted aircraft with a rudder.

[0005] The technical solution to achieve the purpose of the present invention is: a three-ducted aircraft attitude control method with a heading rudder, comprising:

[0006] Step 1: Establish the body coordinate system;

[0007] Step 2: Establish a control efficiency model based on the geometric layout of the power system of the three-ducted aircraft;

[0008] Step 3: Based on the control efficiency model, a control allocation method for a three-ducted aircraft with a heading rudder is established, and its input is the output of attitude control;

[0009] Step 4: Based on control theory, an attitude control method for a three-ducted aircraft with a heading rudder is established, and its output is used as the input of the control allocation method;

[0010] Step 5: Substitute the expected total thrust, expected rolling moment, expected pitching moment, and expected yaw moment calculated in step 4 into the control allocation matrix designed in step 3 to obtain the final control command.

[0011] An electronic device comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor implements the above-mentioned attitude control method of a three-ducted aircraft with a heading rudder when executing the program.

[0012] A computer-readable storage medium stores a computer program, which, when executed by a processor, implements the above-mentioned attitude control method for a three-ducted aircraft with a heading rudder.

[0013] A computer program product comprises a computer program, which implements the above-mentioned attitude control method of a three-ducted aircraft with a heading rudder when executed by a processor.

[0014] Compared with the prior art, the present invention has the following significant advantages: the present invention can efficiently and quickly control the yaw channel of a three-ducted aircraft by utilizing the tilting control surfaces under the two ducts. The present invention has the characteristic that the way the yaw moment is generated by the yaw control surfaces is significantly different from the way the yaw moment (also called anti-torque) is generated by the rotor, and control channels are separately designed for the anti-torque of the three-ducted aircraft and the yaw moment generated by the yaw control surfaces, so as to achieve more precise control. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 This is a geometric diagram of a three-ducted UAV with a heading rudder.

[0016] Figure 2 It is a schematic diagram of the angle when observing the steering surface from the center of mass of the aircraft.

[0017] Figure 3 Schematic diagram of the attitude controller and control distribution structure.

[0018] Figure 4 Schematic diagram of the pitch response curve.

[0019] Figure 5 Schematic diagram of the roll response curve.

[0020] Figure 6Schematic diagram of yaw response curve.

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

[0022] The present invention proposes a control method for a three-ducted aircraft with a rudder. The aircraft is based on a three-rotor triangular layout, with a yaw multi-surface added under the rotor to be responsible for yaw control, and a duct is added to improve the overall efficiency of the aircraft power system. The method specifically includes the following steps:

[0023] Step 1: Establish the body coordinate system.

[0024] The coordinate system is fixed to the body, such as Figure 1 The three rotors of the drone are numbered as rotor 1, rotor 2, and rotor 3. The center of gravity of the drone is taken as the origin. b , forward is x b The positive direction of the axis, rightward is y b The positive direction of the axis, downward is z b The two sets of heading rudders are installed under rotor 1 and rotor 2 respectively, and the plane where the rudder surface is located just passes through the center of mass.

[0025] Step 2: Establish the control efficiency model of the three-ducted aircraft with rudder. Figure 1 As shown, the control efficiency model is established as:

[0026]

[0027] Among them, {F x F y F z} respectively represent the effects on the body {x b y b z b}Tension on the axis; {τ x τ y τ z} respectively represent the torque 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 Indicates the maximum thrust that each rotor can produce; Indicates the x coordinate system of rotor 1 or rotor 2 and the body b The angle formed by the axes; Indicates the motor installation center of rotor 1 or rotor 2 relative to the fuselage x b Along the y axis b The distance between the axes; The line connecting the motor installation centers of rotor 1 and rotor 2 relative to the center of mass of the aircraft is x b The distance between the axes; The x-axis represents the distance between the motor installation center of rotor 3 and the center of mass of the aircraft. b δ1 and δ2 represent the included angles between the control surfaces of rotor 1 and rotor 2 and the vertical direction of the body coordinate system, respectively. Figure 2 As shown, the symbol definition of the vertical angle between the control surface corresponding to the rotor 1 and the body coordinate system is shown; c represents the proportional coefficient of the pull-anti-torque moment generated by the rotor.

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

[0029] For a three-ducted aircraft with a heading rudder, the pitch and roll control is similar to the common multi-rotor aircraft principle. The difference is that the three-ducted aircraft adds an additional heading rudder to achieve heading control. At the same time, the counter-torque of each ducted rotor also affects the yaw motion of the aircraft. Considering that there is a difference between the action mechanism of the rotor counter-torque on the heading and the action mechanism of the heading rudder on the heading. Therefore, the control allocation matrix designed in the present invention separates the heading rudder from the rotor counter-torque, so that the control parameters of the heading rudder and the counter-torque can be set separately in the attitude controller. The control allocation method is expressed as a linear matrix:

[0030]

[0031] Among them, F z,d For the desired effect on the body b Force on the axis; {τ x,d τ y,d τ z,d} are the expected moments of the three ducted rotors acting on the three axes of the body coordinate system; it should be noted that τ z,d is the anti-torque expected to be achieved by the three rotors; τ z,δ,d is the yaw moment produced by the desired rudder.

[0032] Step 4: Establish the attitude control method of the three-ducted aircraft with yaw rudder.

[0033] The desired input for attitude control is the desired roll angle φ d , pitch angle θ d , yaw angle ψ d and the desired height p z,d The feedback input of attitude control is the current three-ducted aircraft's roll angle φ, pitch angle θ, yaw angle ψ, roll angle rate p, pitch angle rate q, yaw angle rate r, and the current aircraft's altitude p z and speed v z; The output of the attitude controller is the force {F x,d F y,d F z,d}, the torque {τ x,d τ y,d τ z,d}, and the yaw moment τ expected to be generated by the rudder z,δ,d The output of the attitude controller is used as the input of the control distribution. The structural relationship between the controller and the control distribution is as follows: Figure 3 The specific steps to construct the attitude controller are as follows:

[0034] Step 4.1: Calculate the expected body angular rate {p according to the expected roll angle, pitch angle, yaw angle and the current roll angle, pitch angle, yaw angle of the three-ducted aircraft d q d r d}.

[0035]

[0036] in, Represents ψ d The numerical differentiation of represents the numerical differentiation of ψ; It is an adjustable parameter.

[0037] Step 4.2: According to the desired body angular velocity {p d q d r d} and the current body angular rate {pqr} to calculate the desired torque.

[0038]

[0039] in, Indicates p d The numerical differentiation of represents the numerical differentiation of p, Indicates q d The numerical differentiation of represents the numerical differentiation of q; All are adjustable parameters.

[0040] Step 4.3: According to the desired flight height p z,d and the current height p z and vertical velocity v z , calculate the effect on the three-ducted aircraft z b Tension F on the shaft z,d .

[0041]

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

[0043]

[0044] Step 5: Substitute the expected value F calculated in step 4 z,d , τ x,d , τ y,d , τ z,d , τ z,δ,d , substitute into the control allocation matrix (2) designed in step 3 to obtain the final control instructions T1, T2, T3, δ1, δ2.

[0045] The present invention is further described in detail below in conjunction with the accompanying drawings.

[0046] Example

[0047] In order to make the purpose, technical scheme and advantages of the implementation of this application clearer, the technical scheme in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. Based on the implementation of this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application.

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

[0049]

[0050] Where p = [p x p y p z ] T represents the position of the aircraft; v = [v x v y v z ] T represents the speed of the aircraft; 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 represents the force vector acting on the body coordinate system; m represents the mass of the aircraft at takeoff; Θ = [φ θ ψ] T Represents a vector of three Euler angles; ω = [pqr] T represents the body angular velocity; It represents the differential of the aircraft position, aircraft velocity, Euler angle, and body angular rate in turn;

[0051]

[0052] Represents the body angular velocity ω and the Euler angular velocity The conversion relationship; τ = [τ x τ y τ z ] T Represents the torque vector acting on the body; J = diag ([J xx J yy J zz ]) represents the inertia tensor of the aircraft.

[0053] Step 2: Based on the rigid body model (6) and the control efficiency model (1), a three-ducted aircraft simulation environment with a rudder is constructed using Simulink software. The model parameters are: 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: Build a controller based on the controller equations (2), (3), (4), and (5), where the parameter values ​​involved 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 desired height p z,d Start the simulation. The result is as follows Figure 4 , Figure 5 , Figure 6 , Figure 7As shown in the figure, the pitch angle, roll angle, yaw angle and altitude channel of the aircraft respectively achieve good tracking effects. The designed controller can effectively and stably control the three-ducted aircraft with heading 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; Step 2: Establish a control efficiency model based on the geometric layout of the power system of the three-ducted aircraft; Step 3: Based on the control efficiency model, a control allocation method for a three-ducted aircraft with a heading rudder is established, and its input is the output of attitude control; Step 4: Based on control theory, an attitude control method for a three-ducted aircraft with a heading rudder is established, and its output is used as the input of the control allocation method; Step 5: Substitute the expected total thrust, expected rolling moment, expected pitching moment, and expected yaw moment calculated in step 4 into the control allocation matrix designed in step 3 to obtain the final control command.

2. The attitude control method of a three-ducted aircraft with a yaw rudder according to claim 1, characterized in that: 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 center of gravity of the body as the origin. b , forward is x b The positive direction of the axis, rightward is y b The positive direction of the axis, downward is z b The two sets of heading rudders are installed under rotor 1 and rotor 2 respectively, and the plane where the rudder surface is located passes through the center of mass.

3. The attitude control method of a three-ducted aircraft with a yaw rudder according to claim 2, characterized in that: Establish a control efficiency model for a three-ducted aircraft with a rudder: Among them, {F x F y F z } respectively represent the effects on the body {x b y b z b }Tension on the axis; {τ x τ y τ z } respectively represent the torque 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 Indicates the maximum thrust that each rotor can produce; Indicates the x coordinate system of rotor 1 or rotor 2 and the body b The angle formed by the axes; Indicates the motor installation center of rotor 1 or rotor 2 relative to the fuselage x b Along the y axis b The distance between the axes; The line connecting the motor installation centers of rotor 1 and rotor 2 relative to the center of mass of the aircraft is x b The distance between the axes; The x-axis represents the distance between the motor installation center of rotor 3 and the center of mass of the aircraft. b ; δ1 and δ2 represent the angles between the control surfaces of rotor 1 and rotor 2 and the vertical direction of the fuselage coordinate system, respectively; c represents the proportional coefficient of the pull-counter torque generated by the rotor.

4. The attitude control method of a three-ducted aircraft with a yaw rudder according to claim 3, characterized in that: The control allocation method of a three-ducted aircraft with a heading rudder is established as follows: The control allocation matrix separates the rudder and the rotor anti-torque so that the control parameters of the rudder and the anti-torque can be set separately in the attitude controller; the control allocation method is expressed as a linear matrix: Among them, F z,d For the desired effect on the body b Force on the axis; {τ x,d τ y,d τ z,d } are the expected moments of the three ducted rotors acting on the three axes of the body coordinate system; τ z,d is the anti-torque expected to be achieved by the three rotors; τ z,δ,d is the yaw moment produced by the desired rudder.

5. The attitude control method of a three-ducted aircraft with a yaw rudder according to claim 4, characterized in that: The attitude control method of a three-ducted aircraft with a heading rudder is established as follows: The desired input for attitude control is the desired roll angle φ d , pitch angle θ d , yaw angle ψ d and the desired height p z,d The feedback input of attitude control is the current three-ducted aircraft's roll angle φ, pitch angle θ, yaw angle ψ, roll angle rate p, pitch angle rate q, yaw angle rate r, and the current aircraft's altitude p z and speed v z ; The output of the attitude controller is the force {F x,d F y,d F z,d }, the torque {τ x,d τ y,d τ z,d }, and the yaw moment τ expected to be generated by the rudder z,δ,d ; The output of the attitude controller is used as the input of the control distribution; The specific steps of constructing the attitude controller are as follows: Step 4.1: Calculate the expected body angular rate {p according to the expected roll angle, pitch angle, yaw angle and the current roll angle, pitch angle, yaw angle of the three-ducted aircraft d q d r d }; in, Represents ψ d The numerical differentiation of represents the numerical differentiation of ψ; is an adjustable parameter; Step 4.2: According to the desired body angular velocity {p d q d r d } and the current body angular rate {pqr} to calculate the desired torque; in, Indicates p d The numerical differentiation of represents the numerical differentiation of p, Indicates q d The numerical differentiation of represents the numerical differentiation of q; All are adjustable parameters; Step 4.3: According to the desired flight height p z,d and the current height p z and vertical velocity v z , calculate the effect on the three-ducted aircraft z b Tension force F on the shaft z,d : in, All are adjustable parameters, sat represents the saturation function, which is defined as follows:

6. The attitude control method of a three-ducted aircraft with a yaw rudder according to claim 5, characterized in that: The expected value F calculated in step 4 z,d , τ x,d , τ y,d , τ z,d , τ z,δ,d , substitute into the control allocation matrix (2) designed in step 3 to obtain the final control instructions T1, T2, T3, δ1, δ2.

7. 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, the attitude control method of a three-ducted aircraft with a rudder as described in any one of claims 1-6 is implemented.

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

9. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the attitude control method of a three-ducted aircraft with a rudder as described in any one of claims 1-6 is implemented.

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