CN110254741B - Design method of flight control system - Google Patents
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Abstract
The invention discloses a design method of a flight control system, and belongs to the technical field of unmanned aerial vehicles. From the flight mechanism of the spinning aircraft, an equivalent theory is established, a kinematic model and a dynamic model control model of the aircraft are deduced, a control system of the spinning aircraft is further established, and six-degree-of-freedom complete control of the aircraft is realized through periodic transformation and feedback control. The invention provides a novel method for controlling a low-degree-of-freedom aircraft, and solves the problem of stable control of an under-actuated aircraft.
Description
Technical Field
The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to a design method of a flight control system.
Background
Most flight control systems in use today use euler angles as the basis for control. The three channels of rolling, pitching and yawing in the remote control directly correspond to the rolling angle, the pitch angle and the yawing angle for controlling the attitude of the aircraft, and the movement of the aircraft in space is indirectly controlled through the three channels. In addition, the stability enhancement function of the flight control is based on the three Euler angles, and the rolling angle, the pitch angle and the yaw angle are regulated through PID, so that the expected value of the stability tends to be achieved, and the stability of the aircraft is realized.
The cured control model also limits the development of new aircraft. The aircraft dynamics equations established based on the body coordinate system all comprise six degrees of freedom, and the aircraft designed based on the aircraft dynamics equations are stable in six degrees of freedom. Low-degree-of-freedom aircraft can only stay in the conceptual or experimental stage due to lack of suitable and control models, and application is difficult to achieve.
The spinning aircraft always maintains a high-speed spinning state around the central axis of the machine body in the take-off and later flying processes, so that better mechanical stability and aerodynamic characteristics are obtained. However, the three euler angles of a spinning vehicle are unstable, and therefore a spinning vehicle is not compatible with conventional flight control.
The flying of the spinning aircraft is dynamic balance, is different from the requirement of the traditional aircraft on the stability of three attitude angles, reduces the requirement on control, does not control the angle and the angular speed of the spinning of the aircraft body, and reserves more control resources for the remaining two directions, so that the aircraft obtains better maneuvering performance and adapts to more complex environments. Under the condition of the same performance requirement, the requirements of the spinning aircraft on hardware configuration are reduced, so that the production cost is reduced. However, the stress analysis method, the mathematical model and the flight attitude control of the spinning aircraft are obviously different from those of the traditional aircraft, the control of the spinning aircraft cannot be realized by the traditional flight control system, and a set of brand new control model is required to be established.
Disclosure of Invention
In order to solve the problems mentioned in the background art, and solve the problem of unstable control of a spinning aircraft caused by limitations of an existing aircraft control model, the invention aims to provide a design method of a flight control system.
The invention provides a design method of a flight control system, which comprises the following steps:
(1) Analyzing a flight mechanism of the spinning aircraft, establishing an equivalent model according to a plane of a fuselage of the spinning aircraft, and respectively determining a plurality of degrees of freedom of the aircraft based on an equivalent coordinate system;
(2) And establishing a mathematical model of the spinning aircraft based on the equivalent coordinate system. The method comprises the steps of deducing a dynamics model, a kinematic model and a state equation model of an organism aiming at the flight mechanism and spin characteristics of the organism;
(3) A control system for the spinning vehicle is established. The control of five degrees of freedom of the aircraft is realized through periodic transformation. The PID algorithm is used for controlling three loops of the position, the speed and the gesture of the machine body, and the flying is controlled through real-time feedback.
In the design method of the flight control system provided by the invention, the equivalent model is an imaginary model established according to the flight mechanism of the spinning aircraft.
In the design method of the flight control system provided by the invention, the equivalent model is used for controlling the aircraft instead of the aircraft body. When the aircraft is stably spinning in the air, the space through which the aircraft rotates is equivalent to a disk-shaped aircraft. The motion state of the disc-shaped aircraft in space is similar to that of a common aircraft, and can be regarded as a special stability. The aircraft system using the disc can equivalently reflect the flight state of the aircraft.
In the design method of the flight control system provided by the invention, the equivalent coordinate system refers to the machine system of the equivalent model.
In the method for designing a flight control system provided by the invention, the degrees of freedom of the aircraft in the step (1) include inclination angle, direction angle and cycle parameters.
The tilting angle of the aircraft is the same as that of the equivalent model, namely the tilting degree of the aircraft or the equivalent model relative to the horizontal direction;
the direction angle of the aircraft is the same as the direction angle of the equivalent model, namely the tilting direction of the aircraft or the equivalent model;
And the periodic parameter is used for describing the relative relation between the aircraft body and the equivalent model.
In the design method of the flight control system provided by the invention, the important property that the airframe spin of the equivalent coordinate system does not cause the coordinate system change provides an important basis for the gesture representation and control model of the spin aircraft and other low-freedom aircraft, and the control model is simplified based on the establishment of the equivalent coordinate coefficient model, so that an effective coordinate system is provided for the stable control of the low-freedom aircraft.
The rotor of the spinning aircraft is not fixed relative to the fuselage in azimuth from the user coordinate system, so that the practical situation of the external force applied to the spinning aircraft is directly acted on the equivalent model is quite complex. Aiming at the characteristic of high-speed spin of a spinning aircraft, the invention provides an equivalent stress method.
For any one azimuth, three rotors will continue to pass through in one spin cycle. When the rotating speed reaches a certain degree, the rotor wings which provide lift force can be uniformly distributed on the circumference taking the length of the horn as the radius in the time of one period with a certain probability density. And after the lift force of the upper horn is overlapped, the radius of the equivalent circumference is contracted inwards, and the lift force density is further increased.
The effect of superposition of all rotor wing lifting force and lifting force generated by aircraft spinning is equivalent to a lifting force circumference at the periphery of the airframe. Each point on the circumference corresponds to a lift density f (γ):
Wherein, F i is the lift force of each rotor wing passing through the position in one period, F 0 is the lift force generated by the spinning of the fuselage, and r is the equivalent radius.
For any one flight state, there is a corresponding f (γ) for representing the stress situation of the spinning vehicle. Since rotor thrust F i is continuously variable, F (γ) is a continuously-derivable function.
In the design method of the flight control system, the stress equivalent method further simplifies the mechanical model of the low-freedom-degree aircraft, is a simplified analysis method for the symmetrical aircraft, and provides convenience for dynamic modeling analysis of the aircraft.
The design method of the five-degree-of-freedom aircraft control system provided by the invention fully utilizes the characteristics of the altitude symmetry and the spin of the spinning aircraft to carry out simplified analysis when analyzing the flight mechanism of the spinning aircraft. Based on the characteristics of the spin of the spinning vehicle, the degree of freedom of the vehicle in the direction of the z-axis of the machine system is not considered. Thus, the x-axis and y-axis of a spinning aircraft architecture are constantly changing in space and are equivalent.
In the design method of the flight control system provided by the invention, the inclination angle and the direction angle of the equivalent coordinate system are only determined by the relative relation between the z-axis of the original aircraft system and the geographic system, and cannot be changed along with the spinning of the aircraft body. The flight control system designed based on the tilting angle and the direction angle can effectively avoid the problem of confusion of attitude angle calculation caused by the attitude calculation mode based on the Euler angle in the traditional flight control system.
The beneficial effects of the invention are as follows:
(1) The invention provides a set of flight control system and a method for stably controlling an under-actuated system of an aircraft.
(2) The invention establishes a mathematical model of the spinning aircraft, deduces and optimizes the model, and reveals the flight mechanism and the space attitude of the spinning aircraft.
(3) The mathematical method and deformation formula described in the present invention are applicable to any spinning aircraft and mechanical devices of similar principles.
(4) The method for the structure of the periodic control system adopted by the invention for the spinning aircraft can be applied to dynamic balance systems in other mechanical fields, and simplifies the flow from design to control to realization.
(5) The method for the structure of the periodic control system adopted by the invention for the spinning aircraft can be applied to dynamic balance systems in other mechanical fields, and simplifies the flow from design to control to realization.
(6) Compared with the control technology of traditional aircrafts such as four rotor wings, the control method has the advantages that the control of the degree of freedom in the spin direction is reduced, the coupling of the other two degrees of freedom is realized, and the control algorithm is greatly simplified.
(7) Compared with the control technology of traditional aircrafts such as four rotor wings, the yaw control method has the advantages that the influence of yaw control is reduced when the rotation speed of the motor is controlled, the occupied motor adjusting range is small, and the hardware performance requirement is low.
Drawings
The mechanism of rearward (left) and forward (right) movement of FIG. 1
Fig. 2 motor thrust and motion mechanism
FIG. 3 periodic variation of roll and pitch angle at 30 ° roll angle
FIG. 4 periodic variation of yaw angle at 30 ° pitch angle
Figure 5 model airplane remote controller
FIG. 6 control model of a spinning vehicle
FIG. 7PID control loop
Detailed Description
In order to make the technical problems, technical schemes and beneficial effects to be solved more clear, the invention is further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of the invention.
(1) And analyzing the flight mechanism of the spinning aircraft, establishing an equivalent model according to the plane of the fuselage of the spinning aircraft, and respectively determining a plurality of degrees of freedom of the aircraft based on an equivalent coordinate system.
First, a basic flying attitude of the spinning aircraft is subjected to theoretical analysis. In a specific embodiment provided by the invention, taking a three-rotor spinning aircraft as an example, the theoretical analysis of the basic flying attitude of the spinning aircraft is as follows.
The same lifting force is added to the three rotors, and according to Newton's second law, when the component force of the combined total lifting force in the vertical direction is larger than the gravity of the three-rotor aircraft, the aircraft can fly upwards. Similarly, the same lifting force is reduced for the three rotors, and the resultant force in the vertical direction is downward, so that the machine body can realize downward landing. It should be noted that the three rotor lift forces are the same under the assumption that the center of gravity of the machine body is centered, if the center of gravity is not centered, the resultant force of the three lift forces should be aligned with the gravity force, otherwise the following will occur.
It is noted that the "increasing the thrust of the motor" and "decreasing the thrust of the motor" do not simply increase or decrease the thrust by a specific amount, but are related to the magnitude of the desired value and the condition in which the aircraft is being subjected. According to the actual requirement, the change amount of the thrust force is positively correlated with the magnitude of the expected value, such as the larger the expected speed and the expected angle, or the more the joystick is pushed, the larger the change amount response of the thrust force can be. In addition, the amount of change in thrust is also related to the location of the motor. In general, the greater the motor thrust change amount is from the symmetry axis, the greater the thrust change amount is 0 for the motor located on the symmetry axis.
Based on the characteristics of the aircraft's spin, the yaw angle measured by the sensor mounted on the fuselage will change continuously with the spin. Unless the aircraft always keeps absolute level, the pitch angle and the roll angle also fluctuate between positive and negative values of the roll angle, and the change rules of the three attitude angles are beyond the range allowed by the traditional flight control system. One embodiment of the present invention thus provides a corresponding equivalent coordinate system.
The spinning aircraft will always spin in one direction during take-off and later flights, so that the pitch, roll and yaw angles in the sense of the body will vary periodically with the spin.
The pitch angle and the roll angle then fluctuate between positive and negative values of the roll angle, wherein the pitch angle varies approximately in a sinusoidal cycle and the roll angle varies approximately in a sinusoidal step curve, as shown in fig. 3. The yaw angle will vary from 0-360 with spin in a nonlinear period as shown in fig. 4.
When the aircraft is stably spinning in the air, the space through which the aircraft rotates is equivalent to a disk-shaped aircraft. The motion state and the external force condition of the disc-shaped aircraft in the space are similar to those of a common aircraft, and the disc-shaped aircraft is considered to be a special stability.
The advantage of adopting the equivalent coordinate system, the tilting angle and the direction angle to replace the original coordinate system and the Euler angle is that the equivalent coordinate system does not change along with the spin of the aircraft, and only reflects the attitude of the aircraft in the air; the tilt angle and the direction angle do not change with spin, but only with respect to the attitude position of the plane of the aircraft.
More precisely, the tilting angle reflects the degree of tilting of the fuselage plane with respect to the horizontal plane, the resistance to external forces on the mechanical model and the movement of the aircraft in the horizontal direction with respect to the air on the dynamic model. The direction angle reflects the orientation of the fuselage, the direction of the force against the external force on the mechanical model, and the direction of the movement trend of the aircraft in the horizontal direction relative to the air on the dynamic model.
A spintronic aircraft is an under-actuated system consisting of three inputs (three motor speeds) controlling six outputs (three attitude angles and three positional information). Spatially represented by four basic spatially-motion flying poses and six degrees of freedom. The four basic space motion flight attitudes comprise linear displacement and horizontal rotation in three directions of space, and six degrees of freedom are linear displacement x, y and z, a tilting angle, a direction angle and a period parameter respectively.
The tilt angle should be zero at normal hover, increasing with increasing desired airspeed; the direction angle is only significant when the tilt angle is present, in the opposite direction of the desired airspeed; the cycle parameters are determined by the spin characteristics of the aircraft and are not constrained. The linear displacement x, y and z are determined by the attitude angle of the aircraft and the thrust of the motor.
(2) And establishing a mathematical model of the spinning aircraft based on the equivalent coordinate system, and deducing a dynamic model, a kinematic model and a state equation model of the aircraft aiming at the aircraft flight mechanism and the spin characteristic of the aircraft as follows.
The lift force generated by the spin of the aircraft body is defined as F 0, the lift force generated by three rotors of the aircraft body is defined as F 1、F2、F3, the component force on the three axes of the space x, y and z is defined as F x、Fy、Fz, and the total lift force of the aircraft is as follows:
The direction is vertical to the upward direction of the machine body, namely, points to the negative direction of the z axis of the equivalent coordinate system, so the matrix formula is as follows:
according to the equivalent stress method, the effect of superposition of all rotor wing lifting force and lifting force generated by aircraft spinning can be equivalent to a lifting force circumference on the periphery of the aircraft body. Each point on the circumference corresponds to a lift density f (gamma)
Wherein, F i is the lift force of each rotor wing passing through the position in one period, F 0 is the lift force generated by the spinning of the fuselage, and r is the equivalent radius.
For any one flight state, there is a corresponding f (γ) that can represent the stress condition of the spinning vehicle. Since rotor thrust F i is continuously variable, F (γ) is a continuously-derivable function.
The equivalent stress method considers that the lifting force generated by three rotors and the horn can be equivalent to superposition of lifting force generated by countless rotors around the periphery in one period. Let f (γ) be the linear density of the lift on the equivalent circle, then there are:
on a ground coordinate system, the component matrix formula in the three coordinate axis directions is as follows:
On the ground coordinate system, the kinetic equation is:
wherein m is the mass of the machine body, and D x、Dy、Dz is the air resistance coefficient on the x, y and z three axes.
The matrix formula of the angular velocity vector of the machine body is as followsThe formula of the inertia matrix is as follows:
The kinetic equation of the rigid body rotation is
The roll moment and pitch moment of the three-rotor aircraft were analyzed again below, where R is the 1/2 wheelbase and R ′ is the equivalent 1/2 wheelbase.
The roll moment can be expressed as:
The pitching moment can be expressed as:
the gyroscopic moment generated during flight can be expressed as:
Mgyro＝∑Ω×Hi
Where H is the moment of momentum of the rotating part, the moment of momentum H of the body can be expressed as:
Where j r is the moment of inertia in the z-axis direction, ω i (i=1, 2, 3) is the angular velocity of the three rotors, and ω 0 is the angular velocity of the body's spins.
The total external torque experienced by a multi-rotor aircraft can be expressed as:
substitution arrangement can be obtained:
From symmetry J x＝Jy, it is simplified to
Attitude angular rate of unmanned aerial vehicleAnd angular velocity components p, q and r in three directions of a space coordinate system have certain mathematical relations, and are expressed as follows by formulas:
the above is the derivation process of the organism dynamics equation and the kinematics model.
Taking the input quantities of the throttle, the roll and the pitch of the system as U 1、U2、U3 respectively. For any F 1、F2、F3, there is
U1＝F1+F2+F3
Solving simultaneously to obtain
More generally, the expression mode based on the equivalent stress method is as follows
U1＝∮f(γ)dγ
Solving an optimal solution as
To more intuitively represent the power model of a spinning aircraft, one embodiment of the present invention provides a concept of "rack function".
The spinning aircraft is identical in control about the x-axis and the y-axis, i.e. pitch and roll are no longer distinguished, and yaw direction control is eliminated, so that only one frame factor is required, the rotation axis of which is identical to the direction of the pitch angle.
When defining the position of a certain motor in the period parameter gamma, the ratio of the change amount of the motor thrust to the current expected moment in the rotation axis direction is a spin frame factor:
Bb＝g(γ-b)
Wherein b is the angle of the rotation axis.
If the periodic parameters of the machine body are used, the frame factor corresponding to each motor can be rewritten as
Wherein N is the number of motors. i is the motor number, from 1 to N.
One of the cases is selected as follows. Defining the symmetry axis as running back and forth, the desired change is rolling to the right. According to the theory, the rotation speed of the motor on the left side of the symmetry axis, namely the motor No. 1, is increased, and the rotation speed of the motor on the right side, namely the motor No. 3, is reduced. The motor No. 2 is positioned on the symmetry axis, so the thrust variation is 0. Three motor factors may be recorded as (-1, 0, 1) at this time. For a spinning vehicle, motor number 1 will continue to move away from the axis of symmetry, motor number 2 will start to shift left from the axis of symmetry, and motor number 3 will continue to move closer to the axis of symmetry as the cycle parameter increases. Correspondingly, the thrust change amounts corresponding to the three motors also change.
A more gradual thrust variation relationship is presented herein. In the specific embodiment provided by the invention, the expected attitude value and the length of the horn are set to be the dimensionless number of 1, and the relative magnitude of the total moment can be represented after the product of the motor thrust change amount and the motor factor is overlapped. According to one possible mode of the embodiment provided by the invention, the motor factor can be expressed as
Wherein, gamma is a period parameter, and N is the number of motors. i is the motor number, from 0 to N-1.
(3) A control system for the spinning vehicle is established. The control of five degrees of freedom of the aircraft is realized through periodic transformation. Furthermore, PID control is used for realizing control of three loops of the position, the speed and the gesture of the machine body, and flight is controlled through real-time feedback.
The control of the yaw direction of the spinning aircraft is abandoned, and the control is greatly different from the remote control mode based on the Euler angle of the traditional model airplane remote controller, so that one specific embodiment provided by the invention provides the remote control mode of the spinning aircraft based on an equivalent coordinate system.
With reference to existing model airplane remote controls, the basic channels of aircraft control include roll, pitch, throttle, and yaw. Wherein the roll and pitch are located on the same joystick, designated as the tilt lever.
The angle between the tilting lever and the initial position is used as the control amount of the tilting angle, and the direction of the tilting lever is used as the control amount of the direction angle.
When the tilting lever is pushed in a certain direction, the tilting angle of the equivalent body increases, and the direction angle is toward the direction in which the tilting lever is pushed.
For any combination of tilting quantity and direction quantity, the combination acts on the remote controller and is equivalent to superposition of two directions, and the combination is directly transmitted through the original two channels of rolling and pitching, and is transmitted to the flight control through the receiver, and then the control quantity of the tilting angle and the direction angle is calculated again.
The throttle channel is the same as a traditional aircraft, the throttle is pushed forward, and the thrust of all motors is increased. Typically, the throttle passage is corrected by the tilting angle to ensure that the component of the thrust force in the vertical direction is unchanged.
The yaw channel corresponds to a control direction angle, the yaw bar is pushed to the right, the direction angle is increased, namely the equivalent coordinate system of the aircraft is rotated to the right, the yaw bar is pushed to the left, and the direction angle is reduced, namely the equivalent coordinate system of the aircraft is rotated to the left.
More specifically, as a correction for the tilt x component and the tilt y component. For example, when the direction angle is oriented in the north-south direction (or reference direction), the yaw bar is pushed right, and the tilt y component is reduced to increase the direction angle, while the tilt angle is unchanged by reducing the tilt x component. The correction formula is as follows:
After a series of corrections, the tilt x component and the tilt y component are recombined into the desired values of tilt angle and direction angle.
The spinning three-rotor five-degree-of-freedom aircraft control system is an underactuated system with nonlinearity, strong coupling and time variability. PID control has advantages in control aspect on a single independent channel, the algorithm structure is simple to realize, and the controlled system has good stability by selecting proper proportion, differentiation and integral coefficients. Based on basic models such as attitude control, remote control and periodic transformation control, four loop controls of the position, the speed, the attitude and the attitude angular rate of the machine body are realized by using PID control in a flight control algorithm, and the flight is controlled by real-time feedback.
The basic PID control law can be expressed as:
Wherein K p、K1 and K D are respectively referred to as proportional, integral, differential coefficients
In one specific embodiment provided by the invention, four control loops are designed in the spinning three-rotor five-degree-of-freedom aircraft control system: position control loop, speed control loop, attitude angular rate control loop.
Position control loop: firstly, inputting target position information x, y and z of the machine body into a position controller, and simultaneously, transmitting real-time position information of the machine body fed back by a GPS into the position controller in time. By the position controller data calculation, the linear velocity v x、vy、vz on the x, y, z axes required to reach the target position information x, y, z is calculated.
Speed control loop: the target line speed v x、vy、vz of the machine body calculated by the position controller is first input to the speed controller. Meanwhile, real-time linear speed information of the GPS to the machine body is timely fed back to the speed controller. The total pulling force F and target attitude desired angles α and β required to reach the target line speed v x、vy、vz are calculated by the speed controller data calculation.
Attitude control loop: first, the lift force F calculated by the speed controller and required by the machine body to adjust the flying gesture is input to the gesture controller together with the target gesture expected angles alpha and beta. Meanwhile, the real-time attitude angle of the machine body which is fed back is converted into a tilting angle and a direction angle, and the tilting angle and the direction angle are transmitted to an attitude controller through an avionic system. Calculating the angular rate to reach the target through the calculation of the data of the gesture controllerAnd the required total pulling force F.
Attitude angular rate control loop: first, the target angular rate calculated by the speed controllerTogether with the required total tension force F i is input to the attitude controller. And simultaneously converting the gesture angular rate of the fed back machine body into the angular rate of the tilting angle and the direction angle according to the real-time gesture angle, and transmitting the angular rate to an angular rate controller through the navigation gesture system. And calculating an equivalent lift force density function F (gamma) required by the machine body to reach the target attitude angular rate, and then adjusting the lift forces F 1、F2、F3 of the three rotors in a periodic control system through periodic parameters to realize real-time control on the flight attitude of the machine body.
The lift force F 1、F2、F3 is taken as a known quantity, is transmitted into a dynamics model to perform related data calculation, and the result is transmitted to a navigation attitude system and a GPS, so that the inclination angle, the direction angle and the position information x, y and z of the aircraft are adjusted. Finally, the output tilt angle x and y components and the position information x, y and z are fed back to the position controller and the speed controller, and the operation is repeated until the preset target requirement is met.
The three-rotor five-degree-of-freedom flight control model establishment work based on the equivalent coordinate system and the periodic control algorithm is completed.
In a specific embodiment provided by the invention, the control law is embodied, a flight control code of the spinning aircraft is developed and transplanted into pix flight control, so that stable control of the spinning aircraft is realized. Indicating that the above-described flight control system is feasible in principle to practice.
The previous description of the disclosed embodiments is provided to enable any person skilled in the art to make or use the present invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the invention. Thus, the present invention is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.
Claims (7)
1. A method of designing a flight control system, comprising the steps of:
(1) Analyzing a flight mechanism of the spinning aircraft, establishing an equivalent model according to a plane of a fuselage of the spinning aircraft, and respectively determining a plurality of degrees of freedom of the aircraft based on an equivalent coordinate system;
(2) Establishing a mathematical model of the spinning aircraft based on an equivalent coordinate system; the method comprises the steps of deducing a dynamics model, a kinematic model and a state equation model of an organism aiming at the flight mechanism and spin characteristics of the organism;
(3) A control system of the spinning aircraft is established, the degree of freedom of the aircraft in the z-axis direction of the system is not considered, and five degrees of freedom of the aircraft are controlled through periodic transformation; the control of three loops of the position, the speed and the gesture of the machine body is realized by PID control, and the flying is controlled by real-time feedback;
The spinning aircraft is an aircraft which always maintains a high-speed spinning state around a central axis of the machine body in the take-off and later flying processes.
2. The method according to claim 1, wherein in the case of unstable spin direction, stable control of the other five degrees of freedom is realized.
3. A method of designing a flight control system according to claim 1, wherein the degrees of freedom of the aircraft in step (1) include pitch angle, direction angle and cycle parameters:
The tilting angle of the aircraft is the same as that of the equivalent model, namely the tilting degree of the aircraft or the equivalent model relative to the horizontal direction;
the direction angle of the aircraft is the same as the direction angle of the equivalent model, namely the tilting direction of the aircraft or the equivalent model;
And the periodic parameter is used for describing the relative relation between the aircraft body and the equivalent model.
4. The method of claim 1, wherein the dynamic model of the aircraft in step (2) comprises:
Wherein F 1、F2、F3 is the lift force of each rotor wing, U 1、U2、U3 is the input quantity of the throttle, the roll and the pitch of the system, and ψ is the attitude angle of the aircraft in the z-axis direction of the engine system.
5. The method according to claim 1, wherein only two attitude angles of the pitch angle and the direction angle are used as the angle control amount of the control circuit, and only two attitude angles of the pitch angle rate and the direction angle rate are used as the angle rate control amount of the control circuit.
6. The method of claim 1, wherein the lever directly reflects the operating state of the aircraft, the tilt angle of the lever directly corresponds to the tilt angle of the aircraft, the direction of the lever directly corresponds to the direction of the aircraft, and the feedback of the aircraft is consistent with the intent of the operator.
7. A method of designing a flight control system according to claim 1, wherein the equivalent model is used to control an aircraft in place of the aircraft body.
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