CN109062237B - Active-disturbance-rejection attitude control method for unmanned tilt-rotor aircraft - Google Patents
Active-disturbance-rejection attitude control method for unmanned tilt-rotor aircraft Download PDFInfo
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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
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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
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
Abstract
The invention discloses an active disturbance rejection attitude control method of an unmanned tilt rotor aircraft, which comprises the steps of firstly arranging a transition process, obtaining a smooth attitude angle instruction and a differential signal thereof by an attitude angle instruction signal through a tracking differentiator, and regarding the differential signal as an attitude angle speed instruction signal; directly taking the actual attitude angle and attitude angle rate of the aircraft as feedback quantities to obtain errors of the attitude angle and the attitude angle rate, and forming initial control quantities by nonlinear combination of the errors; estimating the real-time action quantity of the total disturbance of the system by using the extended state observer to compensate the initial control quantity to obtain a final control quantity; and fusing and distributing the redundant control surfaces of the tiltrotor aircraft. The method optimizes the structure of the active disturbance rejection controller, reduces the order of the extended state observer, adopts the differential output of the tracking differentiator as an attitude angular rate instruction, simplifies the structure of the attitude controller, simultaneously controls by means of errors, does not depend on an accurate model of a controlled object, and is suitable for engineering application.
Description
Technical Field
The invention relates to the technical field of flight control, in particular to an active disturbance rejection attitude control method for an unmanned tilt rotor aircraft.
Background
The tilt rotor aircraft combines the advantages of a helicopter and an airplane, has the characteristics of hovering and vertical take-off and landing of the helicopter, has the performance of high-speed flight of the airplane, and has special application value. However, the rotor wing tilting can cause the dynamic change and the aerodynamic interference of the aircraft, and meanwhile, the rotor wing tilting has the characteristics of strong coupling among channels, redundancy of a control system and the like, and great difficulty is brought to the design of a flight control system.
The existing control methods are mainly divided into three categories: one is a control strategy that eliminates errors based on errors; secondly, a control method based on internal mechanism description; and thirdly, an intelligent control method based on data statistics and training. The PID controller is a control strategy for eliminating errors based on errors, in the control engineering practice, the errors between the actual behaviors of the controlled target objects are easy to obtain, so the control strategy based on the errors does not depend on a specific mathematical model of the controlled object, however, due to the nonlinear coupling and uncertain complex structure of the tilt rotorcraft, the control effect of the PID controller is limited, and in addition, the tilt transition process needs the coordination of a plurality of groups of controllers, and the design of the controller is complex. The control method based on the internal mechanism description needs a mathematical model with a relatively accurate controlled object, for the control of the tilt-rotor aircraft, generally, a nonlinear model is trimmed near each working point and a linearized model of the nonlinear model is calculated, then, a corresponding controller is designed for each linearized model, and finally, flight control is performed through a gain scheduling technology. The intelligent control algorithm represented by the neural network can expand the application range of the traditional control method for processing the dynamics change of the tilt-rotor aircraft in the transition state, the control performance is improved, but the structure of the controller is more complex, a large amount of data needs to be acquired through experiments after the neural network is introduced, the neural network is trained by using the data, and the complexity of the design process of the controller is increased.
Disclosure of Invention
The invention aims to solve the technical problem of overcoming the defects of the prior art and provides an auto-disturbance-rejection attitude control method of an unmanned tilt rotor aircraft.
The invention adopts the following technical scheme for solving the technical problems:
the invention provides an active disturbance rejection attitude control method of an unmanned tilt rotor aircraft, which comprises the following steps:
step one, obtaining a smooth attitude angle command and a differential signal thereof from the attitude angle control command through a tracking differentiator, and taking the differential signal as an attitude angle rate command;
secondly, on the basis of the attitude angle instruction and the differential signal thereof obtained in the first step, the actual attitude angle and the actual attitude angle rate of the aircraft are used as feedback quantities to obtain an attitude angle error and an attitude angle rate error, and the errors are combined in a nonlinear way to form an initial control quantity;
on the basis of the initial control quantity obtained in the step two, estimating the total disturbance real-time action quantity of the unmanned tilt rotor aircraft attitude control system by using the extended state observer to compensate the initial control quantity, and obtaining a final control quantity;
and step four, designing a fusion and distribution mode of the rotor wing and the pneumatic control surface according to the operation strategy of the tilt rotor aircraft on the basis of obtaining the control quantity in the step three.
As a further optimization scheme of the unmanned tilt rotor aircraft active disturbance rejection attitude control method, a specific form of a tracking differentiator in the step one is as follows:
where f is an intermediate variable, v (k) is an attitude angle command at the k-th time, v (k) is a reference value1(k) Is a tracking signal of v (k), v2(k) Is v1(k) Differential of v1(k +1) is a tracking signal of the attitude angle command v (k +1) at the (k +1) th time, v2(k +1) is v1The differential of (k +1), r is a speed factor determining the tracking speed, h is the step length, and a nonlinear function fhan () is the steepest control comprehensive function;
v is to be2And the attitude angle speed command is used for directly carrying out deviation control on the attitude angle speed.
As a further optimization scheme of the unmanned tilt rotor aircraft active disturbance rejection attitude control method, the nonlinear combination form in the second step is as follows:
u0＝-fhan(e1,e2,r1,h1) (2)
wherein u is0As an initial control quantity, e1Is the error between the tracking signal output by the tracking differentiator and the attitude angle, e2Is a differential tracking of the differentiator outputError between signal and attitude angular rate, r1Is the control quantity gain, h1Is a fast factor.
As a further optimization scheme of the unmanned tilt rotor aircraft active disturbance rejection attitude control method, the form of the extended state observer in the third step is as follows:
wherein e (k) is an intermediate variable, z1(k) Is the attitude angle, z, estimated by the extended state observer at time k1(k +1) is an attitude angle estimated by the extended state observer at the (k +1) th time, y (k) is an attitude angle output by the attitude control system of the unmanned tilt-rotor aircraft at the k-th time, u (k) is a control quantity of the attitude control system at the k-th time, and beta01And beta02Is a parameter of the extended state observer, fal (. + -.) is a continuous power function with a linear segment near the origin, α and δ are internal parameters of the fal (. + -.) function, z2(k) Is the disturbance estimate at time k, z2(k +1) is the disturbance estimate at time (k +1), and z is2(k +1) compensating as an output of the extended state observer;
the final control amount u is:
wherein z is2Is the disturbance estimation of the extended state observer, and b is the control quantity amplification factor.
As a further optimization scheme of the active disturbance rejection attitude control method of the unmanned tilt rotor aircraft, the tilt rotor aircraft operating strategy in the fourth step is as follows:
tilt rotorcraft control surface including collective pitch δcolDifferential delta of total distancecolcLongitudinal cyclic pitch deltalonLongitudinal cyclic pitch differential deltaloncAileron deltaailElevator deltaeleRudder deltarudAnd nacelle inclination steering deltanac；
In a helicopter mode, the tilt rotor aircraft generates corresponding force and moment through the total pitch and the periodic pitch of the rotor wings to complete the control of the aircraft; wherein, the vertical channel is controlled by the total distance operation; the transverse channels are controlled by collective differential; the longitudinal channel is controlled by longitudinal periodic variable pitch; the course motion is controlled by longitudinal periodic torque-variable differential motion;
when the inclination angle of the nacelle of the tilt rotor aircraft tilts forwards by 90 degrees, the aircraft enters an airplane mode; in an airplane mode, the forward flying speed is controlled by the total distance of the rotor wings, and the attitude angle is controlled by the pneumatic control surface; wherein the longitudinal channel is controlled by an elevator; the transverse channel is controlled by an aileron; the course channel is controlled by a rudder;
in the tilting transition mode, redundancy of control surfaces exists on a transverse channel, a longitudinal channel and a course channel, wherein the control surfaces of the longitudinal channel comprise longitudinal periodic torque conversion and an elevator; the control surface of the transverse channel comprises a collective differential and an aileron; the steering control surface of the course channel comprises a longitudinal periodic torque-variable differential and a rudder.
As a further optimization scheme of the unmanned tilt rotor aircraft active disturbance rejection attitude control method, the fusion and distribution mode of the rotor and the pneumatic control surface is as follows:
three channel controller outputs U are defined as:
U＝[Ulat Ulong Uhead]T (5)
where the superscript T denotes transpose, UlatFor the transverse channel controller output, UlongFor longitudinal channel controller output, UheadOutputting the data to a course channel controller; defining:
wherein K is each redundant control surfaceAssigning a weight coefficient, k11Is the aileron weight coefficient, k21Is the collective differential weight coefficient, k32Is the elevator weight coefficient, k42For longitudinal cyclic moment-varying weight coefficient, k53Is a rudder weight coefficient, k63And changing a non-zero coefficient in K according to different flight states to complete the conversion of flight modes for longitudinal period torque conversion differential weight coefficients.
As a further optimization scheme of the unmanned tilt rotor aircraft active disturbance rejection attitude control method, the weight distribution of the control surface is carried out according to a scale factor which changes along with the inclination angle of the nacelle, and K matrix is selected as follows:
in the process of converting the helicopter mode into the airplane mode, the control quantity of the rotor wing is gradually reduced, and the pneumatic control surface is gradually added to the control of each channel to complete smooth transition.
Compared with the prior art, the invention adopting the technical scheme has the following technical effects:
according to the invention, the attitude angle control command can obtain a smooth attitude angle command and a differential signal thereof through the tracking differentiator, the differential signal is taken as an attitude angle rate command, and the command of an inner loop is not obtained through an outer loop in a traditional control system, so that the control structure is simplified; meanwhile, the attitude angle and the attitude angle rate information measured by the aircraft are directly used as feedback quantities, and observation is not carried out by the extended state observer, so that the order of the extended state observer is reduced, the calculated quantity of the system is reduced, and the real-time performance of the system is improved.
Drawings
Fig. 1 is a structural diagram of an active disturbance rejection attitude control method of an unmanned tiltrotor aircraft according to the present invention.
FIG. 2 is a simulation diagram of attitude angles of three channels when the nacelle inclination is 30 °; wherein, (a) is the attitude angle response of each channel, and (b) is the manipulated variable of each channel.
FIG. 3 is a simulation diagram of attitude angles of three channels when the nacelle inclination is 60 °; wherein, (a) is the attitude angle response of each channel, and (b) is the manipulated variable of each channel.
FIG. 4 is a diagram of a nacelle forward rake command.
FIG. 5 is a transition pitch channel simulation; wherein, (a) is a pitch angle response, and (b) is a pitch channel manipulation amount.
FIG. 6 is a rolling channel simulation during a transition phase; wherein (a) is roll angle response and (b) is roll channel manipulated variable.
FIG. 7 is a simulation diagram of a course path in a transition process; wherein, (a) is the course angle response, and (b) is the course channel manipulated variable.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention will be described in detail with reference to the accompanying drawings and specific embodiments.
The active disturbance rejection control technology is a control method for eliminating errors based on errors, does not depend on an accurate mathematical model, can inhibit and eliminate the disturbance inside and outside the system, and is suitable for solving the control problem of a controlled object with the characteristics of strong coupling, nonlinearity, difficult accurate modeling and the like, such as a tilt rotor aircraft.
As shown in fig. 1, the principle of the method for controlling the active-disturbance-rejection attitude of the unmanned tilt-rotor aircraft according to the present invention is: adopting a control strategy for eliminating errors based on errors, arranging a transition process for the attitude angle instruction through a tracking differentiator TD, obtaining an angular rate instruction, and taking the attitude angle and the attitude angle rate of the aircraft as feedback quantities to obtain errors of the attitude angle and the attitude angle rate; nonlinear state error feedback NLSEF is adopted, and the errors of the attitude angle and the attitude angle rate are subjected to nonlinear combination, so that the dynamic performance of a closed-loop system can be remarkably improved compared with linear combination; coupling among the channels is regarded as disturbance, and decoupling is carried out through disturbance estimation and disturbance compensation of an Extended State Observer (ESO).
An active disturbance rejection attitude control method for an unmanned tilt rotor aircraft comprises the following steps:
the method comprises the following steps: arranging a transition process, obtaining a smooth attitude angle command and a differential signal thereof by the attitude angle control command through a tracking differentiator, and taking the differential signal as an attitude angle rate command, wherein the form of the attitude angle rate command is as follows:
where f is an intermediate variable, v (k) is an attitude angle command at the k-th time, v (k) is a reference value1(k) Is a tracking signal of v (k), v2(k) Is v1(k) Differential of v1(k +1) is a tracking signal of the attitude angle command v (k +1) at the (k +1) th time, v2(k +1) is v1The differential of (k +1), r is a velocity factor determining the tracking velocity, h is the step length, the nonlinear function fhan (×) is the steepest control integral function, fhan (x)1,x2The expression for r, h) is:
the tracking differentiator is effective for solving the contradiction between overshoot and rapidity, the selection range of the error feedback gain and the error differential feedback gain is greatly expanded, setting is easier, the range of object parameters to which the feedback gain can adapt is greatly expanded, and the robustness of the controller is enhanced.
Step two: on the basis of obtaining an attitude angle instruction and a differential signal thereof in the first step, taking the actual attitude angle and the actual attitude angle rate of the aircraft as feedback quantities to obtain an attitude angle error and an attitude angle rate error, and combining the errors through nonlinearity to form an initial control quantity, wherein the combination form is as follows:
u0＝-fhan(e1,e2,r1,h1) (3)
wherein u is0I.e. the initial control quantity, e1Is the error between the tracking signal output by the tracking differentiator and the attitude angle, e2Is the error between the differential signal output by the tracking differentiator and the attitude angular rate. r is1Is the control quantity gain, h1Is a fast factor.
Step three: on the basis of obtaining the initial control quantity in the step two, estimating the total disturbance real-time acting quantity of the unmanned tilt rotor aircraft attitude control system by using the extended state observer to compensate the initial control quantity, and obtaining the final control quantity, wherein the form of the state observer is as follows:
wherein e (k) is an intermediate variable, z1(k) Is the attitude angle, z, estimated by the extended state observer at time k1(k +1) is an attitude angle estimated by the extended state observer at the (k +1) th time, y (k) is an attitude angle output by the attitude control system of the unmanned tilt-rotor aircraft at the k-th time, u (k) is a control quantity of the attitude control system at the k-th time, and z2(k) Is the disturbance estimate at time k, z2(k +1) is the disturbance estimate at time (k +1), β01And beta02Are parameters of the extended state observer, α and δ are internal parameters of the fal (×) function, the function fal (e, α, δ) is a continuous power function with a linear segment near the origin:
the final control amount u is:
wherein z is2Is the disturbance estimation of the extended state observer, and b is the control quantity amplification factor.
Step four: and on the basis of obtaining the control quantity in the third step, designing a fusion and distribution mode of the rotor and the pneumatic control surface according to the operation strategy of the tilt rotor aircraft.
The tilt rotor aircraft operating strategy is:
tilt rotorcraft control surface including collective pitch δcolDifferential delta of total distancecolcLongitudinal cyclic pitch deltalonLongitudinal cyclic pitch differential deltaloncAileron deltaailElevator deltaeleRudder deltarudAnd nacelle inclination steering deltanac。
In helicopter mode, tiltrotor aircraft control the aircraft by generating corresponding forces and moments through collective and cyclic pitches of the rotors. Wherein, the vertical channel is controlled by the total distance operation; the transverse channels are controlled by collective differential; the longitudinal channel is controlled by longitudinal periodic variable pitch; heading motion is differentially controlled by longitudinal periodic torque variation.
When the tilt angle of the tiltrotor nacelle is tilted forward 90 degrees, the aircraft enters airplane mode. In the airplane mode, the total distance of the rotor wing controls the forward flying speed, and the pneumatic control surface controls the attitude angle. Wherein the longitudinal channel is controlled by an elevator; the transverse channel is controlled by an aileron; the course passageway is controlled by a rudder.
In the tilting transition mode, redundancy of control surfaces exists on a transverse channel, a longitudinal channel and a course channel, wherein the control surfaces of the longitudinal channel comprise longitudinal periodic torque conversion and an elevator; the control surface of the transverse channel comprises a collective differential and an aileron; the steering control surface of the course channel comprises a longitudinal periodic torque-variable differential and a rudder.
The fusion and distribution mode of the rotor and the pneumatic control surface is as follows:
the transition of verting earlier stage, preceding flying speed is less, and control moment mainly comes from the rotor. With the increase of the inclination angle of the nacelle, the forward flight speed is increased, and the operating moment generated by the aerodynamic control surface is also increased continuously. At the later stage of tilting transition, the control surface takes the pneumatic control surface as a main part, and the period torque-changing of the rotor wing is assisted.
Three channel controller outputs are defined as:
U＝[Ulat Ulong Uhead]T (7)
where the superscript T denotes transpose, UlatFor the transverse channel controller output, UlongFor longitudinal channel controller output, UheadAnd outputting the information to the course channel controller. Defining:
wherein K is the weight coefficient of each redundant control surface, K11Is the aileron weight coefficient, k21Is the collective differential weight coefficient, k32Is the elevator weight coefficient, k42For longitudinal cyclic moment-varying weight coefficient, k53Is a rudder weight coefficient, k63And changing a non-zero coefficient in K according to different flight states to complete the conversion of flight modes for longitudinal period torque conversion differential weight coefficients.
In helicopter mode, the coefficient matrix K is selected as:
the tilt rotor aircraft now operates in the same manner as a tandem helicopter.
Under the aircraft mode, the manipulation mode of tilt rotor aircraft is the same with fixed wing aircraft, and at this moment, selects coefficient matrix K to be:
in the tilting transition mode, the K matrix is selected as follows:
the control input of each channel is smoothly transited according to the sine and cosine function of the forward rake angle of the nacelle, so that the control quantity of the helicopter rotor gradually exits from the control of each channel and the aerodynamic control surface gradually enters into the control of each channel in the process of converting the helicopter mode into the airplane mode.
In order to verify the effectiveness of the active disturbance rejection attitude control method and the rationality of the unmanned tilt rotor aircraft control strategy, a certain small-sized unmanned tilt rotor aircraft nonlinear six-degree-of-freedom flight dynamics model is used for simulation verification.
Typical point simulation verification: the inclination angles of the nacelle are 30 degrees and 60 degrees and are used as an early stage study object and a later stage study object of tilting transition, a section of square wave signal of 4 degrees is added to the rolling channel, the pitching channel and the yawing channel at the balancing point, and simulation results are shown in fig. 2 and fig. 3.
Simulation verification of the whole flight transition process: the tilting rotor wing starts from a helicopter mode, enters a transition state in 2 seconds, the nacelle inclination angle tilts at a uniform speed at an angular rate of 6deg/s, tilting transition is completed after 15 seconds, the nacelle tilting transition mode is entered, the nacelle tilting and tilting process is shown in fig. 4, and simulation curves of all channels are shown in fig. 5, 6 and 7.
FIG. 2 is a simulation diagram of attitude angles of three channels when the nacelle inclination is 30 °; fig. 2 (a) shows the attitude angle response of each channel, and fig. 2 (b) shows the manipulated variable of each channel. FIG. 3 is a simulation diagram of attitude angles of three channels when the nacelle inclination is 60 °; fig. 3 (a) shows the attitude angle response of each channel, and fig. 3 (b) shows the manipulated variable of each channel. From fig. 2 and 3, it can be seen that the aircraft can well follow the attitude command, the controller has good control effect and high control precision, and can realize channel decoupling and complete the attitude control of each channel in the transition mode of channel high coupling.
FIG. 5 is a transition pitch channel simulation; fig. 5 (a) shows a pitch angle response, and fig. 5 (b) shows a pitch channel manipulation amount. FIG. 6 is a rolling channel simulation during a transition phase; fig. 6 (a) shows a roll angle response, and fig. 6 (b) shows a roll channel manipulation amount. FIG. 7 is a simulation diagram of a course path in a transition process; fig. 7 (a) shows a course angle response, and fig. 7 (b) shows a course lane maneuver amount. As can be seen from fig. 5, 6, and 7, in the whole process of the tiltrotor aircraft from the helicopter mode to the transition mode to the airplane mode, the auto-disturbance-rejection attitude controller can well follow the attitude command, and realize stable switching of the flight mode. In the transition process, the control surface has good control condition and stable attitude angle, and the reasonability of the control surface distribution strategy is verified. Meanwhile, the active disturbance rejection controller in the whole process adopts the same set of parameters, which shows that the active disturbance rejection controller has wide parameter application range, large adaptable object range and beneficial engineering application value.
The above description is only for the specific embodiment of the present invention, but the scope of the present invention is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present invention are included in the scope of the present invention.
Claims (6)
1. An active disturbance rejection attitude control method of an unmanned tilt rotor aircraft is characterized by comprising the following steps:
step one, obtaining a smooth attitude angle command and a differential signal thereof from the attitude angle control command through a tracking differentiator, and taking the differential signal as an attitude angle rate command;
secondly, on the basis of the attitude angle instruction and the differential signal thereof obtained in the first step, the actual attitude angle and the actual attitude angle rate of the aircraft are used as feedback quantities to obtain an attitude angle error and an attitude angle rate error, and the attitude angle error and the attitude angle rate error are combined in a nonlinear mode to form an initial control quantity;
on the basis of the initial control quantity obtained in the step two, estimating the total disturbance real-time action quantity of the unmanned tilt rotor aircraft attitude control system by using the extended state observer to compensate the initial control quantity, and obtaining a final control quantity;
on the basis of the control quantity obtained in the step three, designing a fusion and distribution mode of a rotor wing and a pneumatic control surface according to a control strategy of the tilt rotor aircraft;
the tilting rotor aircraft operating strategy in the fourth step is as follows:
tilt rotorcraft control surface including collective pitch δcolDifferential delta of total distancecolcLongitudinal cyclic pitch deltalonLongitudinal cyclic pitch differential deltaloncAileron deltaailElevator deltaeleRudder deltarudAnd nacelle inclination steering deltanac；
In a helicopter mode, the tilt rotor aircraft generates corresponding force and moment through the total pitch and the periodic pitch of the rotor wings to complete the control of the aircraft; wherein, the vertical channel is controlled by the total distance operation; the transverse channels are controlled by collective differential; the longitudinal channel is controlled by longitudinal periodic variable pitch; the course motion is controlled by longitudinal periodic torque-variable differential motion;
when the inclination angle of the nacelle of the tilt rotor aircraft tilts forwards by 90 degrees, the aircraft enters an airplane mode; in an airplane mode, the forward flying speed is controlled by the total distance of the rotor wings, and the attitude angle is controlled by the pneumatic control surface; wherein the longitudinal channel is controlled by an elevator; the transverse channel is controlled by an aileron; the course channel is controlled by a rudder;
in the tilting transition mode, redundancy of control surfaces exists on a transverse channel, a longitudinal channel and a course channel, wherein the control surfaces of the longitudinal channel comprise longitudinal periodic torque conversion and an elevator; the control surface of the transverse channel comprises a collective differential and an aileron; the steering control surface of the course channel comprises a longitudinal periodic torque-variable differential and a rudder.
2. The unmanned tilt rotor aircraft active disturbance rejection attitude control method according to claim 1, wherein the tracking differentiator in the first step is in the form of:
where f is an intermediate variable, v (k) is an attitude angle command at the k-th time, v (k) is a reference value1(k) Is a tracking signal of v (k), v2(k) Is v1(k) Differential of v1(k +1) is a tracking signal of the attitude angle command v (k +1) at the (k +1) th time, v2(k +1) is v1The differential of (k +1), r is a speed factor determining the tracking speed, h is the step length, and a nonlinear function fhan () is the steepest control comprehensive function;
v is to be2And the attitude angle speed command is used for directly carrying out deviation control on the attitude angle speed.
3. The method according to claim 2, wherein the nonlinear combination in step two is:
u0＝-fhan(e1，e2，r1，h1) (2)
wherein u is0As an initial control quantity, e1Is the error between the tracking signal output by the tracking differentiator and the attitude angle, e2Is the error between the differential signal output by the tracking differentiator and the attitude angular rate, r1Is the control quantity gain, h1Is a fast factor.
4. The unmanned tilt rotorcraft active-disturbance-rejection attitude control method according to claim 3, wherein the extended state observer in step three is in the form of:
wherein e (k) is an intermediate variable, z1(k) Is the attitude angle, z, estimated by the extended state observer at time k1(k +1) is an attitude angle estimated by the extended state observer at the (k +1) th time, y (k) is an attitude angle output by the attitude control system of the unmanned tilt-rotor aircraft at the k-th time, u (k) is a control quantity of the attitude control system at the k-th time, and beta01And beta02Is a parameter of the extended state observer, fal (. + -.) is a continuous power function with a linear segment near the origin, α and δ are internal parameters of the fal (. + -.) function, z2(k) Is the disturbance estimate at time k, z2(k +1) is the disturbance estimate at time (k +1), and z is2(k +1) compensating as an output of the extended state observer;
the final control amount u is:
wherein z is2Is the disturbance estimation of the extended state observer, and b is the control quantity amplification factor.
5. The method according to claim 1, wherein the rotor and the aerodynamic control surface are fused and distributed in a manner that:
three channel controller outputs U are defined as:
U＝[Ulat Ulong Uhead]T (5)
where the superscript T denotes transpose, UlatFor the transverse channel controller output, UlongFor longitudinal channel controller output, UheadOutputting the data to a course channel controller; defining:
wherein K is a weight coefficient distributed to each redundant control surface, K11Is the aileron weight coefficient, k21Is the collective differential weight coefficient, k32Is the elevator weight coefficient, k42For longitudinal cyclic moment-varying weight coefficient, k53Is a rudder weight coefficient, k63And changing a non-zero coefficient in K according to different flight states to complete the conversion of flight modes for longitudinal period torque conversion differential weight coefficients.
6. The method of claim 5, wherein the weight assignment of the control surfaces is performed according to a scaling factor that varies with nacelle inclination, and wherein the K matrix is selected as:
in the process of converting the helicopter mode into the airplane mode, the control quantity of the rotor wing is gradually reduced, and the pneumatic control surface is gradually added to the control of each channel to complete smooth transition.
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