CN110174901B - Aircraft control method - Google Patents
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- CN110174901B CN110174901B CN201910411166.5A CN201910411166A CN110174901B CN 110174901 B CN110174901 B CN 110174901B CN 201910411166 A CN201910411166 A CN 201910411166A CN 110174901 B CN110174901 B CN 110174901B
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- 238000000034 method Methods 0.000 title claims abstract description 37
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims abstract description 35
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- 238000005096 rolling process Methods 0.000 claims description 11
- 125000004122 cyclic group Chemical group 0.000 claims description 3
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- 230000009286 beneficial effect Effects 0.000 description 2
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Classifications
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
Abstract
The invention discloses a flight control model, and belongs to the technical field of unmanned aerial vehicles. A new aircraft attitude description method is obtained by establishing an equivalent coordinate system, a new flight control model is established based on the aircraft attitude description method, and the aircraft is controlled by controlling the thrust of an equivalent motor. The invention provides a new thought for controlling the aircraft, and solves the problem of instability of the aircraft during high-speed maneuver.
Description
Technical Field
The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to an aircraft control method.
Background
In order to determine the relative position of an aircraft in the air, a flight control system is established, and various coordinate systems have been proposed, including ground coordinate systems, body coordinate systems, airflow coordinate systems, track coordinate systems, and the like. Of the most important are a ground coordinate system and a machine body coordinate system, and three Euler angles determined by the relative relation between the ground coordinate system and the machine body coordinate system become key parameters for describing the attitude of the aircraft.
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.
However, this manner of stabilization may limit the maneuverability of the aircraft to some extent. When the attitude angle changes too fast, the flight control system is difficult to make timely and accurate judgment. For this reason, many flight control systems provide a trick mode at the same time, which compromises stability augmentation for maneuverability, but which increases the difficulty of maneuvering and is only suitable for professional flying hands.
Disclosure of Invention
To solve the above-mentioned problems in the background art, it is an object of the present invention to provide a new flight control method, which is directed to the limitations of the existing aircraft control methods.
The flight control method provided by the invention is that a new aircraft gesture description method is obtained by establishing an equivalent coordinate system, a new flight control model is established based on the aircraft gesture description method, and the control of the aircraft is realized by controlling the thrust of an equivalent motor.
In the flight control method provided by the invention, the equivalent coordinate system is a coordinate system established based on projection of the machine body coordinate system on the user coordinate system.
The user coordinate system is defined according to the wish of an operator on the basis of a geographic coordinate system. The definition of the user coordinate system is as follows: the original geographic system Oz axis is defined as a user coordinate system Oz axis, the front direction defined by the user is defined as an Ox axis, and the Oy axis of the user coordinate system is determined according to the right-hand rule of the coordinate system.
The equivalent coordinate system is specifically defined as: defining an original system origin as an equivalent coordinate system origin; the Oz axis of the original machine system is the Oz axis of the equivalent coordinate system; the intersection line of the vertical plane containing the Ox axis of the user coordinate system and the plane of the machine body is the Ox axis of the equivalent coordinate system, and the former direction is the positive direction; the Oy axis of the equivalent coordinate system is determined according to the right hand rule of the coordinate system.
In the flight control method provided by the invention, the equivalent coordinate system is characterized in that different machine systems positioned on the same plane of the machine body have the same equivalent coordinate system.
In the flight control method provided by the invention, the spatial position of the equivalent coordinate system depends on the spatial position of the machine system, the direction of the equivalent coordinate system only depends on the direction of the z-axis of the machine system, and the x-axis and the y-axis of the equivalent coordinate system are determined by the relative relation between the z-axis and the user coordinate system. Therefore, the spin of the fuselage does not cause a change in the equivalent coordinate system.
In the flight control method provided by the invention, the attitude position of the equivalent coordinate system in space is determined by two parameters, namely the tilting angle and the direction angle. The equivalent coordinate system records the spin condition of the aircraft through a period parameter. The three parameters are specifically defined as follows:
tilt angle α: the degree of inclination of the plane of the fuselage relative to the horizontal, i.e. the angle of the z-axis of the machine system or of the equivalent coordinate system with the right underneath. The inclination angle ranges from 0 to 180 degrees, and the inclination angle is larger as the inclination degree is larger from the horizontal position.
Direction angle beta: the tilting direction of the body, i.e. the angle between the projection of the tilting direction in the horizontal direction and the Ox axis (or reference direction) of the user coordinate system. The direction angle ranges from-180 DEG to 180 DEG, and the right rotation is positive and the left rotation is negative in the top view direction.
Cycle parameter γ: the angle through which the aircraft turns during the spin cycle, i.e., the angle that a rotor makes with respect to the origin position and its start.
Furthermore, in the flight control method provided by the invention, the angular rate refers to the time change rate of the angle.
The tilting angle reflects the tilting degree of the plane of the airframe relative to the horizontal plane, the resistance of the aircraft to external force is reflected on a mechanical model, and the movement of the aircraft relative to air in the horizontal direction is reflected on a dynamic model. The roll angle determines the basic attitude and the kinetic equation of the aircraft.
The direction angle reflects the direction of the fuselage, reflects the direction of the external force on the mechanical model, and reflects the direction of the movement trend of the aircraft relative to the air in the horizontal direction on the dynamic model. The direction angle determines the basic movement state of the aircraft.
The periodic parameters reflect the degree of spin of the fuselage about the z-axis, with no direct impact on the attitude of the aircraft. The cyclic parameter may vary over a period of 0-360 deg., and is used only to determine the position of the aircraft rotor.
Further, the inclination angle, the direction angle and the cycle parameter have a priority relation. In one aspect, during definition and calculation, the tilt angle and the direction angle are determined first, and then the cycle parameter is determined; on the other hand, in control, the tilting angle and the direction angle have a higher influence on the attitude control of the aircraft than the cycle parameter.
Further, the priority relation of the tilt angle, the direction angle and the cycle parameter is embodied in that when the control resource is insufficient, the tilt angle is controlled preferentially, the direction angle is controlled secondarily, and the cycle parameter is controlled finally:
controlling the tilting angle to tend to an expected value, and enabling the tilting angle to be smaller than the maximum allowable tilting angle so as to ensure that the aircraft continuously flies and does not overturn;
controlling the direction angle to tend to an expected value, and enabling the change rate of the direction angle to be in a reasonable range, so as to ensure the stable and controllable flight of the aircraft;
and controlling the periodic parameters to tend to an expected value, and ensuring the stability of the orientation of the aircraft.
The advantage of adopting the equivalent coordinate system, the inclination 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.
In order to make the flight control method provided by the invention compatible with the existing flight control system and make the operation more natural, two new parameters of a tilting x component and a tilting y component are defined in an equivalent coordinate system of the flight control method based on the tilting angle and the direction angle:
tilting the x component: the component of the tilting angle on the y axis of the equivalent coordinate system is equal to the angle rotated by the plane of the airplane body when the tilting angle is minimum around the Ox axis, and can replace the rolling angle of common flight control operation.
Tilting y component: the component of the tilting angle on the y axis of the equivalent coordinate system is equal to the angle rotated by the plane of the machine body when the tilting angle is the smallest around the Ox axis and then the plane of the machine body rotates to the horizontal around the Oy axis, and the pitch angle controlled by the common flight control can be replaced.
The advantage of using the tilt x-component and the tilt y-component instead of the roll angle and the pitch angle is that the tilt x-component and the tilt y-component are based entirely on the relative relation of the fuselage plane and the user coordinate system, so that the feedback of the aircraft always corresponds to the wishes of the operator.
In the flight control method provided by the invention, the equivalent motor thrust is characterized in that the synthesis of the action effect of the equivalent motor thrust is the same as the synthesis of the action effect of the original motor thrust.
According to the flight control method, the equivalent motor thrust control method calculates the equivalent motor factors at the current position according to the period parameters according to the resultant force calculated by the expected motion gesture of the aircraft and the moment in all directions, and calculates and decomposes the equivalent motor factors into the thrust on each motor according to the equivalent motor factors.
In the flight control method provided by the invention, the equivalent motor factor is a variable and is determined according to the frame type of the aircraft and the cycle parameter.
The beneficial effects of the invention are as follows:
(1) The invention provides a set of flight control method, and provides a new method for describing the attitude of an aircraft.
(2) According to the aircraft control method, the stability of the attitude is described by using the tilting angle, the tilting direction is described by using the direction angle, the rotor wing position is determined by using the periodic parameter, and the attitude of the aircraft is more accurately described.
(3) According to the aircraft control method provided by the invention, the three attitude angle parameters have priority, so that the environmental adaptability of a control system is stronger, and the control is more targeted.
(4) The equivalent coordinate system is suitable for any multi-rotor aircraft, so that the flight control has stability enhancement capability when performing high maneuver.
(5) The flight control system provided by the invention has the advantages that the control mode is based on the user coordinate system, and the operation is visual and convenient.
Drawings
FIG. 1 transformation of traditional geographic system into organic system
Transformation of the machine system of FIG. 2 into the equivalent coordinate system of the machine body
Definition of the equivalent coordinate System of the machine body of FIG. 3
Conversion of the equivalent coordinate System of the body of FIG. 4 into the body System
FIG. 5 shows the tilt angle α corresponding to the different pitch and roll angles at 75℃yaw angle in one embodiment of the present invention
FIG. 6 shows a direction angle beta corresponding to different pitch and roll angles at a yaw angle of 75 deg. in one embodiment of the present invention
FIG. 7 shows a periodic parameter gamma corresponding to different pitch and roll angles at a yaw angle of 75℃in an embodiment of the present invention
FIG. 8 shows the variation of four motor roll factors for spinning under certain cycle parameters in one embodiment of the present invention
FIG. 9 shows the change of pitch factors of four motors under a certain cycle parameter by rotation in an embodiment of the present invention
FIG. 10 illustrates the effect of motor thrust changes on aircraft torque in one embodiment of the invention
FIG. 11 is a block diagram of flight control in one embodiment of the present invention
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.
In one embodiment provided by the present invention, the establishment of the equivalent coordinate system is as follows.
Referring to fig. 1, a conventional multi-rotor aircraft determines a geographic to body system transformation relationship by euler angles: based on the geographic coordinate system, the yaw rotation psi is firstly carried out around the z axis (z 0 axis) of the machine body, then the pitch theta is carried out around the y axis (y 1 axis) at the moment, and finally the roll phi reaches the machine body coordinate system around the x axis (x 2 axis) at the moment.
Referring to fig. 2, the body equivalent coordinate system may be considered as being reached by rotating γ around the z-axis (z 3-axis) of the body in the opposite direction on the basis of the body coordinate system.
In one embodiment provided by the present invention, a geographic coordinate system is used as the user coordinate system for simplicity of description.
Referring to fig. 3, the body equivalent coordinate system can also be determined by the tilt angle α and the direction angle β. On the basis of a geographic coordinate system, the machine body firstly tilts alpha around the y axis (y 0 axis) of the machine body, and then rotates beta around the z axis (z 0 axis) of the geographic system to arrive.
Referring to fig. 4, the conversion from the equivalent coordinate system to the body system may also be implemented in combination with the cycle parameters.
That is, by the tilt angle α, the direction angle β, and the cycle parameter γ, the coordinate relationship between the geographical system and the machine system can be determined. The inclination angle alpha, the direction angle beta and the period parameter gamma can be obtained by calculating sensor data of an angular velocity meter, a gyroscope, a magnetic compass and the like according to definition, and can also be obtained by converting the existing roll angle, pitch angle and yaw angle.
Further, in one embodiment provided by the present invention, the equivalent co-ordinate coefficient model for a multi-rotor aircraft is derived as follows.
To simplify the analysis, idealized assumptions are made about the aircraft. To facilitate understanding, and to accommodate the need for existing flight control functions, we use the architecture of a conventional aircraft, as well as pitch, roll and yaw angles, in the derivation. It should be noted, however, that pitch, roll and yaw angles are used as intermediate variables only and are not of practical significance in a method of flight control provided by the present invention.
Based on the definition of the equivalent coordinate system, the transformation matrix of the user coordinate system to the ground coordinate system can be expressed as:
the equation is satisfied: x is X earth ＝S eu X user
Based on the coordinate conversion of Euler angles, a conversion matrix from a ground coordinate system to a machine body coordinate system is obtained:
the equation is satisfied: x is X body ＝S be X earth
Thus, the transformation matrix of the user coordinate system to the body coordinate system can be expressed as:
the equation is satisfied: x is X body ＝S bu X user
Based on the coordinate conversion of the periodic parameters, a conversion matrix from the machine body coordinate system to the machine body equivalent coordinate system is obtained:
the equation is satisfied: x is X equivalent ＝S tb X body ＝S tb S bu X user
Based on the definition of the equivalent coordinate system, the transformation matrix of the user coordinate system to the equivalent coordinate system can be expressed as:
the equation is satisfied: x is X equivalent ＝S tu X user
The conversion from an equivalent coordinate system to the body system can also be achieved in combination with the cycle parameters, as shown.
The relationship between the tilt angle, direction angle and cycle parameters and the pitch, roll and yaw angles is
The conversion of the euler angle into the tilt angle, the direction angle and the cycle parameter can be realized.
Fig. 5 to 7 show the values of the pitch angle α, the direction angle β and the cycle parameter γ corresponding to different pitch angles and roll angles at a yaw angle of 75 ° in a specific embodiment of the present invention.
Each group of pitch angle, roll angle and yaw angle can correspondingly calculate the tilt angle, direction angle and period parameter.
The concepts of pitch, roll and yaw angles are not used in the following except for the above-described derivation of the equivalent coordinate system. Unless otherwise stated, the following coordinate systems refer to equivalent coordinate systems, and the following Ox, oy, oz axes refer to Ox, oy, oz axes of equivalent coordinate systems, and the tilt angle and the direction angle are defined based on equivalent coordinate systems. For ease of control and understanding, roll and pitch appearing hereinafter refer specifically to roll and pitch of the fuselage plane under an equivalent coordinate system, and more specifically to the components of the tilt angle in the x-axis and y-axis of the equivalent coordinate system.
With reference to existing model airplane control systems, the basic channels of aircraft control include roll, pitch, throttle, and yaw. In a specific embodiment provided by the invention, for convenience in control, four channels are modified, which are respectively defined as a tilting x component R, a tilting Y component P, a throttle T and a yaw Y, and are combined to correspond to the tilting angle, the direction angle and the throttle of the multi-rotor aircraft for control.
And combining operation habits, mixing and converting the tilting angle and the direction angle into the tilting x component and the tilting y component, and respectively replacing original rolling and pitching channels.
The tilting x component is equivalent to a rolling angle in the equivalent coordinate system, when the operating lever is shifted to the right and the R channel value is increased, the right motor thrust is reduced, the left motor thrust is increased, a positive rotation moment rotating around the x axis is generated on the yOz plane, and the equivalent machine body rolls from the left to the right; when the joystick is pushed to the left, the R-channel value decreases, rolling to the left.
The tilting Y component is equivalent to a pitch angle in the equivalent coordinate system, when the operating rod is pulled backwards and the Y channel value is increased, the rear motor thrust is reduced, the front motor thrust is increased, a positive rotation moment rotating around the Y axis is generated on the xOz plane, the equivalent machine body is lifted up from the front (or upwards; when the operating rod is pushed forwards and the Y channel value is reduced, the equivalent machine body is lifted down.
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 is the same as a traditional aircraft, the yaw lever is pushed to the right, the aircraft spins rightwards, namely the cycle parameter is increased, the yaw lever is pushed leftwards, the direction angle is reduced, namely the equivalent coordinate system of the aircraft rotates leftwards.
It is noted that the "motor thrust increase", "motor thrust decrease" does not simply increase or decrease the thrust by a specific amount, but is 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.
It should be noted that the "front", "back", "left", "right" are relative to the desired direction of the operator, irrespective of the direction of the body itself. Thus, the "front", "rear", "left" and "right" motors are not fixed motors, but each rotor may be in a different orientation as the aircraft orientation changes. More specifically, it may be determined by a period parameter.
In one embodiment of the present invention, the above theoretical analysis is further exemplified and described in detail.
In one particular embodiment of the present invention, the desired change is roll right for an X-type quad-rotor aircraft. In the case of zero initial cycle parameters, the rotational speeds of the motors on the right of the symmetry axis, i.e. motors No. 1 and No. 2, are reduced, and the rotational speeds of the motors on the left, i.e. motors No. 3 and No. 4, are increased, at which time the roll factors of the four motors can be recorded as (1, -1, -1). For a common multi-rotor aircraft, motors No. 1 and No. 3 will continue to move away from the axis of symmetry and motors No. 2 and No. 4 will continue to move closer to the axis of symmetry as the cycle parameters increase. Correspondingly, the thrust change amounts corresponding to the four motors also change. Similarly, the pitch factor of the four rotor motor is (1, -1, 1).
In general, for symmetrical, conventional multi-rotor aircraft, there are
Wherein N is the number of motors, i is the number of motors, and the range is from 1 to N.
For a symmetrical coaxial dual-rotor multi-rotor aircraft, there is
Wherein N is the number of motors, i is the number of motors, and the range is from 1 to N. Wherein the pitch and roll factors of adjacently numbered motors are the same.
For other types of aircraft, it may also be determined according to a similar method.
Let the motor roll factor be R i The motor rolling factor is P i Wherein i is the number of the motor, and there is
R i ＝R i0 ·cosγ-P i0 ·sinγ
P i ＝P i0 ·cosγ+R i0 ·sinγ
Wherein gamma is a period parameter, P i0 And R is i0 The initial pitch factor and roll factor are the same size as a normal multi-rotor.
Figures 8 and 9 illustrate four changes in roll and pitch factors of four motors with rotation at different cycle parameters, for an example of an X-type quadrotor, with four phase differencesIs a sinusoidal curve of (c).
In this particular embodiment of the invention, the effect of motor thrust changes on aircraft torque is shown in fig. 10, where 4 black sinusoids RF 1-RF 4 represent the contribution of 4 motors to roll attitude and 4 gray sinusoids PF 1-PF 4 represent the effect of 4 motors on pitch direction torque. It can be seen that the value r_fac of the 4 superimposed black sinusoidal lines is constant at 1, i.e. the periodic control system can stably provide the desired torque as a whole. The value p_fac of the 4 gray sinusoids superimposed is constant at 0, i.e. the disturbances of the periodic control system to the moments in other directions can cancel each other out as a whole.
The multi-rotor aircraft control system is an underactuated system with non-linearities, strong coupling, 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. In a specific embodiment provided by the design method of the flight control system, the flight control system uses PID control to realize four loop control of the position, the speed, the gesture and the gesture angular rate of the machine body, and the flight is controlled by real-time feedback.
The basic PID control law can be expressed as:
wherein K is p 、K 1 And K D Respectively called proportional, integral and differential coefficients
Four control loops are designed in the control system: position control loop, speed control loop, attitude angular rate control loop, control block diagram please refer to fig. 11.
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. Calculating the linear velocity v on the x, y and z axes required for reaching the target positions x, y and z by the position controller x 、v y 、v z 。
Speed control loop: first, the target line speed v of the machine body calculated by the position controller x 、v y 、v z 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. Calculating the target line speed v by the speed controller x 、v y 、v z The required total pulling force F and the desired target attitude angle tilt x, y components μ, λ and the cycle parameter γ.
Attitude control loop: first, the lift force F required by the machine body to adjust the flying attitude calculated by the speed controller is input to the attitude controller together with the desired angular tilting x, y components mu, lambda and the periodic parameter gamma of the target attitude. And simultaneously converting the real-time attitude angle of the fed back machine body into tilting x, y components mu, lambda and a periodic parameter gamma, and transmitting the tilting x, y components mu, lambda and periodic parameter gamma to an attitude controller through an attitude navigation system. Calculating target angular rate to be reached by attitude controllerAnd the required total pulling force F.
Attitude angular rate control loop: first, the target angular rate calculated by the speed controllerAnd the required total tension F i Together with the gesture controller. And simultaneously converting the gesture angular rate of the body fed back into the angular rate of tilting x, y components and periodic parameters according to the real-time gesture angle, and transmitting the angular rate to an angular rate controller through a navigation gesture system. Calculating equivalent lifting force F required by the machine body to reach the target attitude angular rate, and then adjusting the lifting force F of the four rotors in a periodical control system through periodical parameters 1 、F 2 、F 3 、F 4 Realizing the real-time control of the flying gesture of the machine body.
Will lift force F 1 、F 2 、F 3 、F 4 As a known quantity, relevant data calculation is carried out in the kinetic model, and the result is transmitted to the attitude and heading reference system and the 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.
In particular to flight control system codes, in one particular embodiment provided by the present invention, a substantial portion of the codes comprise
(1) Using the sensor data of the flight control or the existing attitude angle and attitude angle rate data to solve the inclination angle alpha, the direction angle beta, the period parameter gamma and the angular rate thereof, and calculating the x, y components mu and lambda of the inclination angle;
(2) The angle calculated by the input value of the remote control rolling channel is taken as an expected tilting x component, the angle calculated by the input value of the pitching channel is taken as an expected tilting y component, and the angle calculated by the input value of the yawing channel is taken as the variation of an expected periodic parameter;
(3) In the attitude controller, a tilting x component mu is used for replacing an original rolling angle phi, a tilting y component lambda is used for replacing an original pitch angle theta, a periodic parameter gamma is used for replacing an original yaw angle phi, and expected variation of each component is calculated by comparing the difference between an actual value and an expected value;
(4) In an angular rate controller, the derivative of the tilting x-component is usedInstead of the original roll angle rate->Derivative of the tilting y component +.>Instead of pitch rate->Derivative of cycle parameter->Instead of yaw rate->Calculating the amount of thrust change required in each direction by comparing the difference between the actual value and the expected value;
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 (9)
1. An aircraft control method is characterized in that an equivalent coordinate system established based on projection of an engine body coordinate system on a user coordinate system and a relative relation between the equivalent coordinate system and a geographic coordinate system or an airflow coordinate system are used for describing control quantity of an aircraft instead of the engine body coordinate system and Euler angles;
the definition of the equivalent coordinate system is as follows: defining an origin of a machine body coordinate system as an origin of an equivalent coordinate system, defining an Oz axis of the machine body coordinate system as the Oz axis of the equivalent coordinate system, defining an intersecting line of a vertical plane containing the Ox axis of the user coordinate system and a plane of the machine body as the Ox axis of the equivalent coordinate system, and determining an Oy axis of the equivalent coordinate system according to a right-hand rule of the coordinate system;
the user coordinate system is a coordinate system defined according to the wish of an operator on the basis of a geographic coordinate system, and the definition of the user coordinate system is as follows: defining a geographic coordinate system Oz as a user coordinate system Oz, defining a straight forward direction as an Ox axis by a user, and determining an Oy axis of the user coordinate system according to a right-hand rule of the coordinate system;
the equivalent coordinate system and the relative relation between the equivalent coordinate system and the geographic coordinate system comprise a tilting angle, a direction angle, a period parameter and angular rates thereof:
tilt angle α: the inclination degree of the plane of the machine body relative to the horizontal plane, namely the included angle between the z axis of the machine body coordinate system or the equivalent coordinate system and the right lower part;
direction angle beta: the tilting direction of the machine body, namely the projection of the tilting direction in the horizontal direction and the included angle of the Ox axis of the user coordinate system;
cycle parameter γ: the angle that the aircraft rotates in the spin cycle, namely the included angle between the position of a certain rotor relative to the origin and the initial time;
the angular rate refers to the rate of change of the angle over time.
2. An aircraft control method according to claim 1, wherein the relative relationship comprises any one or more of the components of the tilting angle in each direction, the direction angle and the period parameter obtained with different reference directions and their respective derivatives over time.
3. An aircraft control method according to claim 1, wherein the relative relationship includes a priority relationship among the tilt angle, the direction angle, and the cycle parameter, and when the control resource is insufficient, the tilt angle is preferentially controlled, and the direction angle, and the cycle parameter are again:
the tilting angle is controlled to tend to an expected value and is smaller than the maximum allowable tilting angle, so that the continuous flight of the aircraft can be ensured not to be overturned;
the control direction angle tends to an expected value, and the change rate of the control direction angle is in a reasonable range, so that the stable and controllable flight of the aircraft can be ensured;
the control cycle parameter tends to an expected value, and the stability of the orientation of the aircraft can be ensured.
4. An aircraft control method according to claim 1, characterized in that the tilting angle, the direction angle of the control are related only to the attitude of the plane in which the aircraft is located, the cyclic parameter being independent of the other two attitude angles, the instability of the aircraft spin direction, i.e. cyclic parameter, not affecting the control of the other two angles.
5. An aircraft control method according to claim 2, comprising two parameters of the tilt x, y components:
tilting the x component: the component of the tilting angle on the y axis of the equivalent coordinate system is equal to the angle rotated by the plane of the machine body when the tilting angle is minimum around the Ox axis, and the tilting angle can replace the rolling angle of common flight control operation;
tilting y component: the component of the tilting angle on the y axis of the equivalent coordinate system is equal to the angle rotated by the plane of the machine body when the tilting angle is the smallest around the Ox axis and then the plane of the machine body rotates to the horizontal around the Oy axis, and the pitch angle controlled by the common flight control can be replaced.
6. An aircraft control method according to claim 1, characterized in that the pitch x-component and pitch y-component channels of the manoeuvres are based entirely on the relative relationship of the fuselage plane to the user coordinate system, the feedback of the aircraft always being consistent with the wishes of the manoeuvrer.
7. The aircraft control method according to claim 1, wherein the equivalent motor factor at the current position is calculated based on the period parameter based on the resultant force calculated from the expected motion attitude of the aircraft and the moment in each direction, and the thrust force on each motor is calculated and decomposed based on the equivalent motor factor.
8. An aircraft control method according to claim 1, comprising the following circuits:
position control loop: firstly, inputting target position information x, y and z of a machine body into a position controller, simultaneously, transmitting real-time position information of the machine body fed back by a GPS into the position controller in time, and calculating linear speeds v on x, y and z three axes required by reaching the target positions x, y and z by the position controller x 、v y 、v z ；
Speed control loop: first, the target line speed v of the machine body calculated by the position controller x 、v y 、v z Inputting to a speed controller, simultaneously feeding real-time linear speed information of the GPS to the machine body back to the speed controller in time, and calculating the target linear speed v through the speed controller x 、v y 、v z The required total pulling force F and the desired target attitude angle tilt x, y components mu, lambda and a cycle parameter gamma;
attitude control loop: firstly, the total pulling force F calculated by a speed controller and required by the machine body to adjust the flying attitude is input into the attitude controller together with the desired angle tilting x, y components mu, lambda and period parameter gamma of the target attitude, and the real-time attitude angle of the machine body which is fed back is converted into the tilting x, y components mu, lambda and period parameter gamma, and is transmitted to the attitude controller through a flying attitude system, and the target angular rate to be reached is calculated through the attitude controller、/>、/>And the required total tension force F;
attitude angular rate control loop: first, the target angular rate calculated by the speed controller、/>、/>The required total pulling force F is input into a gesture controller together, simultaneously, the gesture angular rate of the body which is fed back is converted into the angular rate of tilting x, y components and periodic parameters according to the real-time gesture angle, the angular rate is transmitted to the angular rate controller through a navigation gesture system, the equivalent lifting force required by the body to reach the target gesture angular rate is calculated, and then the lifting force F of each rotor wing is adjusted through the periodic parameters in the periodic control system i Realizing the real-time control of the flying gesture of the machine body.
9. An aircraft control system, characterized in that an aircraft control method according to any one of claims 1-8 is used.
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105929940A (en) * | 2016-04-13 | 2016-09-07 | 哈尔滨工业大学深圳研究生院 | Rapid three-dimensional dynamic gesture recognition method and system based on character value subdivision method |
| CN205721628U (en) * | 2016-04-13 | 2016-11-23 | 哈尔滨工业大学深圳研究生院 | A kind of quick three-dimensional dynamic hand gesture recognition system and gesture data collecting device |
| CN108628334A (en) * | 2018-06-28 | 2018-10-09 | 广州极飞科技有限公司 | The control method and device of unmanned vehicle, system, unmanned vehicle |
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| WO2006022654A1 (en) * | 2004-07-29 | 2006-03-02 | Bell Helicopter Textron Inc. | Method and apparatus for flight control of tiltrotor aircraft |
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Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105929940A (en) * | 2016-04-13 | 2016-09-07 | 哈尔滨工业大学深圳研究生院 | Rapid three-dimensional dynamic gesture recognition method and system based on character value subdivision method |
| CN205721628U (en) * | 2016-04-13 | 2016-11-23 | 哈尔滨工业大学深圳研究生院 | A kind of quick three-dimensional dynamic hand gesture recognition system and gesture data collecting device |
| CN108628334A (en) * | 2018-06-28 | 2018-10-09 | 广州极飞科技有限公司 | The control method and device of unmanned vehicle, system, unmanned vehicle |
Non-Patent Citations (1)
| Title |
|---|
| 基于PID算法的四旋翼飞行器定点跟踪控制;李希等;《计算机测量与控制》;20161231(第07期);第109-112页 * |
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