CN117850463A - Transition mode control method for compound wing unmanned aerial vehicle and unmanned aerial vehicle - Google Patents

Abstract

The invention discloses a transition mode control method conforming to a wing unmanned aerial vehicle and the unmanned aerial vehicle. The invention aims at longitudinal control of a transition mode from four rotors to a fixed wing, provides decoupling of longitudinal height control and speed control of a transition mode of a composite wing by introducing total energy control, and performs sectional control on a transition process. The invention also aims at the distribution problem of the four-rotor accelerator and the fixed-wing accelerator in the transition mode process from the four-rotor to the fixed-wing, and realizes the transition control of the four-rotor accelerator in the transition mode process according to the increase of airspeed by designing the four-rotor accelerator incremental controller. According to the invention, through adopting a TECS algorithm, an airspeed controller, a four-rotor accelerator incremental controller and the transition process sectional control, the flight stability and comfort level of the transition mode are obviously improved, and parameters such as speed, height and the like in the transition mode can be accurately controlled, so that the flight is more stable and safer.

Description

Transition mode control method for compound wing unmanned aerial vehicle and unmanned aerial vehicle

Technical Field

The invention relates to the field of control of composite wing unmanned aerial vehicles, in particular to a transition mode control method of a composite wing unmanned aerial vehicle and the composite wing unmanned aerial vehicle adopting the method.

Background

An unmanned aerial vehicle (Unmanned Aerial Vehicle, UAV for short) is an unmanned aerial vehicle which can realize take-off and landing through a flight control system and a ground control station of the unmanned aerial vehicle under the unmanned condition, execute flight tasks according to a specified route and realize full-flow flight. In recent years, unmanned aerial vehicles have been developed sufficiently, unmanned aerial vehicles are more and more abundant in variety, tasks executable by unmanned aerial vehicles are more and more complex, and unmanned and low-cost advantages make unmanned aerial vehicles more and more important in the military industry.

The composite wing unmanned aerial vehicle is a novel-configuration unmanned aerial vehicle with vertical take-off and landing capability, which is recently and widely focused, and the aircraft with the configuration is obtained by simply combining a fixed wing aircraft with a multi-rotor wing suite through a structure, and has the advantages of lower technical realization threshold, low cost and the like.

The vertical take-off and landing fixed wing aircraft has significant advantages over other unmanned aerial vehicle systems, including: the requirements on landing sites are low, runways are not needed, the problem of access obstacles is avoided, and a transmitting and recycling system is not needed; flexible maneuvering, good usability, simple system, few additional equipment, convenient transportation, use and maintenance and few operators; the cost is lower. Compared with a helicopter, the vertical take-off and landing fixed wing aircraft is low in cost, high in reliability and long in endurance time. Compared with a multi-rotor aircraft, the vertical take-off and landing fixed-wing aircraft has long endurance, good wind resistance, high flying speed and large task coverage area. And the design, control and test of the vertical take-off and landing fixed wing unmanned aerial vehicle. Compared with the conventional fixed wing aircraft, the vertical take-off and landing fixed wing aircraft does not need a runway, has wider application range and low use cost. Compared with a tiltrotor aircraft, the vertical take-off and landing fixed wing is simpler in design of a power system and has better structural reliability.

The elevator and throttle lever of a fixed wing aircraft have coupling characteristics to the response of the aircraft. Namely, both elevator swing and throttle change can influence pitch angle and airspeed, and the flight speed and track coupling are more intense in the low dynamic pressure flight state of the aircraft. It is common practice to perform speed control in the speed maintaining mode or to perform altitude control in the speed maintaining mode, but decoupling control is not good due to lack of cooperative control of the elevator and the throttle. Thus a total energy control system (TECS, total Energy Control System) is introduced.

The longitudinal decoupling controller of the total energy control algorithm has the basic principle that when the thrust is changed, the change rate of the total energy is changed, the distribution rate of the total energy is controlled by changing the control elevator deflection angle, and the expected speed or altitude value is converted into the change rate and the distribution rate of the total energy, so that the advanced decoupling control of the longitudinal movement of the aircraft is realized. The total energy control is controlled from an energy point of view. The energy of the aircraft is divided into kinetic energy and potential energy, which are generated by the increase or decrease of the throttle, and the conversion of the kinetic energy and potential energy is realized by controlling the track of the aircraft by the elevator. From this point of view, the total energy/energy difference is decoupled from the throttle/elevator.

The unmanned aerial vehicle flight control system is the core of unmanned aerial vehicle flight, and stable and reliable control is the key of unmanned aerial vehicle stable flight. The flight control of the compound wing unmanned aerial vehicle mainly comprises control under a four-rotor mode, control under a fixed wing mode and control of a transitional mode. At present, flight control for a fixed wing and flight control for a four-rotor wing are mature, and few transition mode control methods for the four-rotor wing to the fixed wing are available. Some studies discuss fusion control of transitional modes, but none are satisfactory.

Disclosure of Invention

In view of the above, the present invention aims to provide a transition mode control scheme of a compound-wing unmanned aerial vehicle based on a total energy control law, so as to improve the longitudinal control effect and accuracy of the transition mode of the compound-wing unmanned aerial vehicle.

According to an aspect of the invention, a transition mode control method of a compound-wing unmanned aerial vehicle is provided, the method is used for controlling the unmanned aerial vehicle to transition from a four-rotor mode to a fixed-wing mode, and the method comprises the following steps:

according to the height H of a given transition _c And fixed-wing cruise airspeed V _c Obtaining an input signal of a total energy controller (TECS controller), wherein the TECS controller is used for cooperatively controlling an elevator and a throttle;

setting the TECS controller as a fixed wing longitudinal controller to participate in control, and obtaining an accelerator control amount and an elevator control amount in a fixed wing mode by the TECS controller based on the input signal;

setting up the airspeed delta controller so that the airspeed delta controller sets up a corresponding maximum limit of airspeed difference according to different airspeeds, wherein the airspeed difference is airspeed feedback value V and given fixed-wing cruising airspeed V _c Is a difference in (2);

judging whether the pitch angle or the height error of the unmanned aerial vehicle meets a preset landing condition, and if so, controlling the unmanned aerial vehicle to enter a landing process;

if the pitch angle and the altitude error of the unmanned aerial vehicle do not meet the preset landing condition, judging whether an airspeed feedback value V reaches a first speed condition, and if not, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on a four-rotor mode, wherein a four-rotor accelerator is not reduced due to the increase of airspeed;

if the airspeed feedback value V reaches the first speed condition, further judging whether the airspeed feedback value V reaches the second speed condition, and if the current speed V does not reach the second speed condition, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on the four-rotor mode, and gradually reducing the four-rotor throttle according to the calculation of the four-rotor throttle incremental controller;

if the airspeed feedback value V reaches the second speed condition, the four-rotor throttle is directly cleared to 0, the pitch angle of the unmanned aerial vehicle is kept near the angle corresponding to the preset optimal lift-drag ratio, and the fixed-wing throttle is controlled to accelerate until the airspeed feedback value V reaches the fixed-wing cruising airspeed V _c To switch to the fixed wing mode, the transition process is completed.

In some embodiments, the input signal to the TECS controller is derived according to the following formulaAnd gamma _e ：

Wherein V is airspeed feedback value, K _v G is the gravity acceleration and is the speed proportionality coefficient,is the derivative of the current speed V, H is the current height, K _h And r is the track angle, which is the height proportionality coefficient.

In some embodiments, the TECS controller obtains the throttle control amount T in the fixed-wing mode based on the following equation _c And an elevator rudder quantity control quantity delta _ec ：

Wherein K is _tp As the total energy proportionality coefficient, K _ti As the total energy differential coefficient, K _ep As the energy difference proportionality coefficient, K _ei As a differential coefficient of the difference in energy,representing the integral operation of the complex domain.

In some embodiments, the maximum throttle limit in the fixed wing mode obtained by the TECS controller is 80% and the elevator amount in the fixed wing mode obtained by the TECS controller is limited to 25% of the maximum rudder deflection angle.

In some embodiments, the airspeed delta controller is configured such that it sets a corresponding airspeed difference maximum limit for different airspeeds, the airspeed difference being the airspeed feedback value V and the given airspeed V _c Comprises:

before the airspeed feedback value V reaches a preset switching airspeed V_switch, setting the maximum value of airspeed difference to be limited at 15m/s;

after the airspeed feedback value V reaches the switching airspeed V_switch, setting the maximum value of airspeed difference to be 10m/s;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s And fixed-wing cruise airspeed V _c Average value of (2).

In some embodiments, determining whether the pitch angle or the altitude error of the unmanned aerial vehicle meets a preset landing condition includes:

if the pitch angle of the unmanned aerial vehicle is larger than 6 degrees or the height error is larger than 15m, judging that the preset landing condition is met;

and if the pitch angle of the unmanned aerial vehicle is not more than 6 degrees and the height error is not more than 15m, judging that the preset landing condition is not met.

In some embodiments, determining whether the airspeed feedback value V reaches the first speed condition includes:

if the airspeed feedback value V reaches the fixed-wing stall airspeed V _s If half of the speed feedback value V reaches the first speed condition;

if the airspeed feedback value V does not reach the fixed-wing stall airspeed V _s And determining that the airspeed feedback value V does not reach the first speed condition.

In some embodiments, determining whether airspeed feedback value V reaches the second speed condition includes:

if the airspeed feedback value V reaches the switch airspeed V_switch, judging that the airspeed feedback value V reaches a second speed condition;

if the airspeed feedback value V does not reach the switch airspeed V_switch, judging that the airspeed feedback value V does not reach a second speed condition;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s And fixed-wing cruise airspeed V _c Average value of (2).

In some embodiments, causing the quad-rotor throttle to be stepped down based on the calculation of the quad-rotor throttle delta controller includes:

four-rotor accelerator incremental controller obtains four-rotor accelerator variation based on the followingAnd airspeed variation (V) _c -V) and let four rotor throttle variation +.>Limited to plus or minus 5% to control four-rotor throttle:

wherein,as a constant term, ρ: air density, C _l : unmanned aerial vehicle lift aerodynamic coefficient, S: unmanned aerial vehicle wing area, ct: lift coefficient of four-rotor motor, C _m As a motor rotation speed constant, ω: four rotor motor speeds.

According to another aspect of the invention, a compound wing unmanned aerial vehicle is presented that is controlled in a manner as described above when transitioning from a four-rotor mode to a fixed-wing mode.

The invention aims at longitudinal control of a transition mode from four rotors to a fixed wing, provides decoupling of longitudinal height control and speed control of a transition mode of a composite wing by introducing total energy control, and performs sectional control on a transition process. The invention also aims at the distribution problem of the four-rotor accelerator and the fixed-wing accelerator in the transition mode process from the four-rotor to the fixed-wing, and realizes the transition control of the four-rotor accelerator in the transition mode process according to the increase of airspeed by designing the four-rotor accelerator incremental controller. According to the invention, through adopting a TECS algorithm, an airspeed controller, a four-rotor accelerator incremental controller and the transition process sectional control, the flight stability and comfort level of the transition mode are obviously improved, and parameters such as speed, height and the like in the transition mode can be accurately controlled, so that the flight is more stable and safer.

The method and apparatus of the present invention have other features and advantages which will be apparent from or are set forth in detail in the accompanying drawings and the following detailed description, which are incorporated herein, and which together serve to explain certain principles of the present invention.

Drawings

The foregoing and other objects, features and advantages of the invention will be apparent from the following more particular descriptions of exemplary embodiments of the invention as illustrated in the accompanying drawings wherein like reference numbers generally represent like parts throughout the exemplary embodiments of the invention.

Fig. 1 shows a schematic structural view of a compound wing drone according to one embodiment of the present invention.

Fig. 2 shows a flow chart of a method of transition modality control of a compound wing unmanned aerial vehicle according to an exemplary embodiment of the present invention.

Fig. 3 shows a schematic diagram of a total energy speed control structure according to an exemplary embodiment of the present invention.

Fig. 4 illustrates a total energy level control structure diagram according to an exemplary embodiment of the present invention.

Fig. 5 shows a schematic diagram of the overall energy control algorithm structure according to an exemplary embodiment of the present invention.

FIG. 6 illustrates a schematic diagram of a total energy speed control architecture with an additional airspeed delta controller, according to an example embodiment of the present invention.

FIG. 7 illustrates a schematic diagram of an airspeed-increment controller control flow in accordance with an exemplary embodiment of the invention.

Fig. 8 illustrates a compound wing drone transition modality control flow diagram in accordance with an exemplary embodiment of the present invention.

Detailed Description

Preferred embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. While the preferred embodiments of the present invention are shown in the drawings, it should be understood that the present invention may be embodied in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art.

After intensive studies on the motion process in the transition mode from the four rotors to the fixed wings, the inventors made the following assumptions in order to simplify the system complexity to achieve efficient and reliable control of the transition mode.

Assume a

Under the conditions that the four-rotor mode does not give roll and yaw angle control, the fixed-wing mode does not give aileron rudder angle and no other interference, the transition mode of the compound-wing unmanned plane is assumed to have no lateral movement, namely, in the transition process, no lateral force and moment exist, and only longitudinal movement is considered.

Suppose two

Assuming that the track angle change is small during the transition, the following equation holds:

sinγ＝γ (1)

where γ is the track angle.

Assume three

The rotating speed change formula of the four-rotor motor of the compound-wing unmanned aerial vehicle is simply and approximately:

ω(t)＝C _m *σ(t)+ω _m (2)

wherein omega _m C is motor rotation speed intercept _m For motor rotation speed constant, ω (t) is rotation speed time-varying function, σ (t) is throttle time-varying function, C _m *σ(t)+ω _m Representing stable rotation of the motorSpeed, is linearly related to throttle. The two-sided derivation of equation (2) can be obtained:

wherein,expressed as the rate of change of the motor speed>Expressed as a motor throttle change rate.

Fig. 1 shows a schematic structural view of a compound wing drone according to one embodiment of the present invention. As shown, the compound wing unmanned aerial vehicle comprises a fixed wing throttle 1, a right front vertical rotor 2, a right rear vertical rotor 3, a left rear vertical rotor 4, a left front vertical rotor 5, an aileron 6, a rudder 7, and an elevator cabin 8.

Fig. 2 shows a flow chart of a method of transition modality control of a compound wing unmanned aerial vehicle according to an exemplary embodiment of the present invention. The method is used for controlling the unmanned aerial vehicle to transition from a four-rotor mode to a fixed-wing mode. As shown in the figure, the method comprises the steps 1 to 7.

Step 1, according to the height H of a given transition process _c And fixed-wing cruise airspeed V _c An input signal is derived for a total energy controller (TECS controller) that is used to cooperatively control the elevator and throttle.

High retention of H during transition _c Is unchanged. Given fixed-wing cruise airspeed V _c The airspeed required by the cruise of the composite wing unmanned aerial vehicle in a fixed wing mode is obtained.

In some embodiments, the input signal to the TECS controller may be derived according to the following equationAnd gamma _e ：

Wherein V is airspeed feedback value, K _v G is the gravity acceleration and is the speed proportionality coefficient,is the derivative of the current speed V, H is the current height, K _h And r is the track angle, which is the height proportionality coefficient.

FIG. 3 shows the calculation of the input signal to the TESC controller according to the formula shown aboveIs a structural schematic diagram of (a); FIG. 4 shows the input signal gamma of a TESC controller according to the formula shown above _e Is a schematic structural diagram of the (c).

The invention will be described in terms of a given transition height H _c And fixed-wing cruise airspeed V _c The main flow Duan Chenwei sub-flow one before determining to switch to the fixed wing mode.

And 2, setting the TECS controller as a fixed wing longitudinal controller to participate in control, and obtaining an accelerator control amount and an elevator control amount in a fixed wing mode by the TECS controller based on the input signal.

When the four-rotor mode is switched to the fixed-wing mode, the TECS controller is used as a fixed-wing longitudinal controller to control, and the throttle control amount and the elevator control amount in the fixed-wing mode are obtained based on input signals.

In some embodiments, the TECS controller derives the throttle control amount T in the fixed-wing mode based on the following equation _c And an elevator rudder quantity control quantity delta _ec ：

Wherein K is _tp As the total energy proportionality coefficient, K _ti As the total energy differential coefficient, K _ep As the energy difference proportionality coefficient, K _ei As a differential coefficient of the difference in energy,representing the integral operation of the complex domain.

FIG. 5 shows calculation of the accelerator control amount T according to the formula shown above _c And the elevator control amount δec.

In some embodiments, the maximum throttle limit in the fixed wing mode obtained by the TECS controller is 80% and the elevator amount in the fixed wing mode obtained by the TECS controller is limited to 25% of the maximum rudder deflection angle. In some examples, the maximum rudder deflection angle is 30 °, and the elevator amount is limited to about 7.5 degrees.

At this time, the fixed wing unmanned aerial vehicle starts to generate airspeed in the horizontal direction, and the position and the working flow of the airspeed increment controller are as follows.

Step 3, setting an airspeed increment controller so that the airspeed increment controller sets a corresponding airspeed difference maximum limit according to different airspeeds, wherein the airspeed difference is an airspeed feedback value V and a given fixed-wing cruising airspeed V _c Is a difference in (c).

FIG. 6 illustrates a schematic diagram of a total energy speed control architecture with an additional airspeed delta controller, according to an example embodiment of the present invention.

In some embodiments, the airspeed difference maximum is set to be limited to 15m/s before the airspeed feedback value V reaches the preset switch airspeed v_switch;

after the airspeed feedback value V reaches the switching airspeed V_switch, setting the maximum value of airspeed difference to be 10m/s;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s Fixed-wing cruiseAirspeed V _c Average value of (2).

Fig. 7 shows a schematic diagram of the airspeed-increment controller control flow according to this embodiment.

And 4, judging whether the pitch angle or the height error of the unmanned aerial vehicle meets a preset landing condition, and if so, controlling the unmanned aerial vehicle to enter a landing process.

In some embodiments, if the pitch angle of the unmanned aerial vehicle is greater than 6 degrees or the height error is greater than 15m, determining that the preset landing condition is satisfied; and if the pitch angle of the unmanned aerial vehicle is not more than 6 degrees and the height error is not more than 15m, judging that the preset landing condition is not met.

In the present invention, the drop process is referred to as a sub-process four.

And 5, if the pitch angle and the altitude error of the unmanned aerial vehicle do not meet the preset landing condition, judging whether the airspeed feedback value V reaches the first speed condition, and if not, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on the four-rotor mode, wherein the four-rotor accelerator is not reduced due to the increase of the airspeed.

In some embodiments, determining whether the airspeed feedback value V reaches the first speed condition includes: if the airspeed feedback value V reaches the fixed-wing stall airspeed V _s If half of the speed feedback value V reaches the first speed condition; if the airspeed feedback value V does not reach the fixed-wing stall airspeed V _s And determining that the airspeed feedback value V does not reach the first speed condition.

If the pitch angle of the unmanned aerial vehicle is not more than 6 degrees and the altitude error is not more than 15m, and the airspeed feedback value V is less than 0.5V _s The cycle is still in sub-process one, i.e., the pitch angle and altitude of the drone is still controlled based on the quad-rotor mode, and the quad-rotor throttle is not reduced by the increase in airspeed.

And 6, if the airspeed feedback value V reaches the first speed condition, further judging whether the airspeed feedback value V reaches the second speed condition, and if the current speed V does not reach the second speed condition, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on the four-rotor mode, and gradually reducing the four-rotor throttle according to the calculation of the four-rotor throttle incremental controller.

In some embodiments, determining whether airspeed feedback value V reaches the second speed condition may include:

if the airspeed feedback value V reaches the switch airspeed V_switch, judging that the airspeed feedback value V reaches a second speed condition;

if the airspeed feedback value V does not reach the switch airspeed V_switch, judging that the airspeed feedback value V does not reach a second speed condition;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s And fixed-wing cruise airspeed V _c Average value of (2).

In some examples, fixed-wing cruise airspeed V _c Can be set as stall airspeed V _s I.e. airspeed feedback value if stall airspeed V is not reached _s If the airspeed feedback value V is 1.2 times that of the second speed condition, then the second sub-process can be entered. After entering the second sub-process, the pitch angle and the height of the unmanned aerial vehicle are still controlled based on the four-rotor mode, and the four-rotor throttle is gradually reduced according to the calculation of the four-rotor throttle increment controller. According to the invention, through designing the four-rotor accelerator incremental controller, the transition control of the four-rotor accelerator in the transition mode process is realized according to the increase of airspeed.

Step 7, if the airspeed feedback value V reaches the second speed condition, the four-rotor throttle is directly cleared to 0, the pitch angle of the unmanned aerial vehicle is kept near the angle corresponding to the preset optimal lift-drag ratio, and the fixed wing throttle is controlled to accelerate until the airspeed V reaches the fixed wing cruising airspeed V _c To switch to the fixed wing mode, the transition process is completed.

If the airspeed feedback value V reaches a second speed condition, e.g., reaches the stall airspeed V shown in the above example _s 1.2 times of (a), then enter sub-process three. After entering the third sub-process, the four-rotor accelerator is directly cleared to 0, the pitch angle of the unmanned aerial vehicle is kept near the angle corresponding to the preset optimal lift-drag ratio, and the fixed-wing accelerator is controlled to accelerate until the airspeed feedback value V reaches the fixed-wing cruising airspeed V _c To switch to the fixed wing mode, the transition process is completed.

The invention aims at longitudinal control of a transition mode from four rotors to a fixed wing, provides decoupling of longitudinal height control and speed control of a transition mode of a composite wing by introducing total energy control, and performs sectional control on a transition process. The invention also aims at the distribution problem of the four-rotor accelerator and the fixed-wing accelerator in the transition mode process from the four-rotor to the fixed-wing, and realizes the transition control of the four-rotor accelerator in the transition mode process according to the increase of airspeed by designing the four-rotor accelerator incremental controller. According to the invention, through adopting a TECS algorithm, an airspeed controller, a four-rotor accelerator incremental controller and the transition process sectional control, the flight stability and comfort level of the transition mode are obviously improved, and parameters such as speed, height and the like in the transition mode can be accurately controlled, so that the flight is more stable and safer.

In a second sub-process, in some embodiments, causing the quad-rotor throttle to be stepped down based on the calculation of the quad-rotor throttle delta controller includes:

four-rotor accelerator incremental controller obtains four-rotor accelerator variation based on the followingAnd airspeed variation (V) _c -V) and limits the variation of the four-rotor throttle to a range of plus or minus 5% to control the four-rotor throttle:

wherein,as a constant term, ρ: air density, C _l : unmanned aerial vehicle lift aerodynamic coefficient, S: unmanned aerial vehicle wing area, ct: lift coefficient of four-rotor motor, C _m As a motor rotation speed constant, ω: four rotor motor speeds.

The principle derivation process of the four-rotor accelerator increment controller is as follows.

The force balance formula aiming at the gravity direction in the transitional process is as follows

mg＝G＝L*cosθ＝[4*Ct*ω ² +0.5*ρ*V ² *C _l *S]*cosθ (9)

Wherein, 4 Ct omega ² Lift provided for quadrotor, 0.5 x ρ x V ² *C _l * S is the lift force provided by the fixed wing, m: unmanned aerial vehicle weight, g: gravitational acceleration, G: gravity, L: unmanned aerial vehicle total lift, θ: unmanned aerial vehicle pitch angle, ct: lift coefficient of four rotor motor, ω: four rotor motor rotational speeds, ρ: air density, V: airspeed of unmanned aerial vehicle C _l : unmanned aerial vehicle lift aerodynamic coefficient, S: unmanned aerial vehicle wing area.

Assuming small pitch angle and angle of attack changes during the transition, i.e. C _l Looking as a constant, cos θ=1;

G＝L＝4*Ct*ω ² +0.5*ρ*V ² *C _l *S (10)

deriving the two ends of the formula (9) to obtain

From equation (10), it can be seen thatAnd->Has a corresponding relationship, recorded as

Wherein,the airspeed change rate is expressed and can be written as (V _c -V)，/>Indicating motor speed variationThe rate.

Substituting formula (3) into formula (11) to obtain

The following formula (8) can be further obtained:

wherein,is a constant term.

The airspeed variation can be obtainedOr (V) _c V) and four rotor throttle variation +.>Correspondence between them. By a four-rotor accelerator incremental controller, airspeed variation +.>Obtaining the variation of the accelerator of four rotors>And will change the quantity->The output limit is positive and negative 5%, and the height and the gesture of the unmanned aerial vehicle are still kept stable by the four-rotor mode at the moment.

Fig. 8 illustrates a compound wing drone transition modality control flow diagram in accordance with an exemplary embodiment of the present invention. As shown in the figure, after the unmanned aerial vehicle is started, the unmanned aerial vehicle vertically takes off in a four-rotor mode, and when the unmanned aerial vehicle reaches the ground clearance height H, the four-rotor mode hovers, begins to prepare to transition to a fixed-wing mode, and enters a first sub-process.

In sub-flow one, the height H of the transition is given _c And fixed-wing cruise airspeed V _c The input signals of the total energy controller are obtained through transformation, and the total energy controller is used as a fixed wing longitudinal controller to participate in control. The total energy controller outputs the accelerator control amount and the elevator control amount in the fixed wing mode, limits the accelerator output to 80%, and limits the elevator cabin rudder amount to 25% of the maximum rudder deflection angle. At this time, the compound wing unmanned aerial vehicle generates a airspeed in a horizontal direction and the airspeed increment is limited by the airspeed increment controller. Then, it is judged whether the pitch angle is greater than 6 degrees or whether the height error is greater than 15m, i.e., the preset landing condition. If yes, enter sub-flow four, namely drop flow.

In the fourth sub-flow, firstly controlling the fixed wing accelerator to be 0, adding the four-rotor wing accelerator to the hovering accelerator, and stopping excessive. And then controlling the pitch angle to be 0 degree in a four-rotor mode, and keeping the height and the posture of the unmanned aerial vehicle stable. If a certain forward flying speed exists at the moment, the pitch angle is controlled to be about 5 degrees in the four-rotor mode until the forward flying speed generated by the fixed wing accelerator is emptied. After the forward flying speed is cleared, the pitch angle is kept to be 0 degree in the four-rotor mode, and vertical landing is carried out. After landing, the four-rotor throttle is cleared 0. At this point, it may be manually determined whether to restart the transition. If yes, returning to the original; if not, the transition is canceled.

When the preset drop judgment is carried out, if the judgment result is negative, continuously judging whether the current airspeed feedback value V reaches 0.5X V _s I.e. a first speed condition. If the result is no, the altitude and the attitude are maintained stable in the four-rotor mode, and the circulation in the first sub-process is continued.

Judging whether the current airspeed feedback value V reaches 0.5X V _s If yes, it is further determined whether the current airspeed feedback value reaches the switch airspeed V_switch, i.e., the second speed condition. If the judgment result is negative, entering a sub-process II.

In the second sub-process, the pitch angle is controlled to be around 0 degrees by a four-rotor controller, and the Dead Zone (Dead Zone) is plus or minus 0.5 degrees. And then obtaining the airspeed change rate as an input of a four-rotor accelerator incremental controller. The four-rotor accelerator delta controller obtains the accelerator change rate of the four rotors based on the above formula (8) to control the four-rotor accelerator size based on the original four-rotor accelerator. The output amplitude limit of the four-rotor accelerator increment controller is within the range of plus or minus 5 percent. At this time, the forces and moments generated in the four-rotor mode maintain the stability of the altitude and attitude of the unmanned aerial vehicle, and the transition is prepared again.

And when judging whether the current airspeed feedback value reaches the switching airspeed V_switch, if so, entering a sub-flow III. In the third sub-process, the accelerator of the four rotors is directly cleared to 0, the pitch angle and the height are controlled by a fixed wing mode, the pitch angle is controlled to be near the angle corresponding to the optimal lift-drag ratio, and the dead zone is between plus and minus 0.5 degrees. And controlling the accelerator acceleration of the fixed wing until the airspeed V reaches a given cruise airspeed V _c And if the dead zone is positive or negative by 1m, the mode is formally switched to the fixed wing mode, and the transition process is finished.

According to an embodiment of the invention, a compound-wing unmanned aerial vehicle is further provided, and the compound-wing unmanned aerial vehicle is controlled by adopting the compound-wing unmanned aerial vehicle transition mode control method when transiting from a four-rotor mode to a fixed-wing mode.

It will be appreciated that the above embodiments mentioned in the present disclosure may be combined with each other to form a combined embodiment without departing from the principle logic, and are limited in space, and the disclosure is not repeated. It will be appreciated by those skilled in the art that in the above-described methods of the embodiments, the particular order of execution of the steps should be determined by their function and possible inherent logic.

Note that all features disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic set of equivalent or similar features. Where used, further, preferably, still further and preferably, the brief description of the other embodiment is provided on the basis of the foregoing embodiment, and further, preferably, further or more preferably, the combination of the contents of the rear band with the foregoing embodiment is provided as a complete construct of the other embodiment. A further embodiment is composed of several further, preferably, still further or preferably arrangements of the strips after the same embodiment, which may be combined arbitrarily.

It will be appreciated by persons skilled in the art that the embodiments of the invention described above and shown in the drawings are by way of example only and are not limiting. The objects of the present invention have been fully and effectively achieved. The functional and structural principles of the present invention have been shown and described in the examples and embodiments of the invention may be modified or practiced without departing from the principles described.

Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present disclosure, and not for limiting the same; although the present disclosure has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the corresponding technical solutions from the scope of the technical solutions of the embodiments of the present disclosure.

Claims (10)

1. A method for controlling a transition mode of a compound-wing unmanned aerial vehicle, the method for controlling the transition of the unmanned aerial vehicle from a four-rotor mode to a fixed-wing mode, the method comprising:

according to the height H of a given transition _c And fixed-wing cruise airspeed V _c Obtaining an input signal of a total energy controller (TECS controller), wherein the TECS controller is used for cooperatively controlling an elevator and a throttle;

setting the TECS controller as a fixed wing longitudinal controller to participate in control, and obtaining an accelerator control amount and an elevator control amount in a fixed wing mode by the TECS controller based on the input signal;

setting up the airspeed delta controller so that the airspeed delta controller is according to different airspeedsAirspeed sets a corresponding maximum limit for airspeed difference, which is the airspeed feedback value V and the given fixed-wing cruise airspeed V _c Is a difference in (2);

judging whether the pitch angle or the height error of the unmanned aerial vehicle meets a preset landing condition, and if so, controlling the unmanned aerial vehicle to enter a landing process;

if the pitch angle and the altitude error of the unmanned aerial vehicle do not meet the preset landing condition, judging whether an airspeed feedback value V reaches a first speed condition, and if not, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on a four-rotor mode, wherein a four-rotor accelerator is not reduced due to the increase of airspeed;

if the airspeed feedback value V reaches the first speed condition, further judging whether the airspeed feedback value V reaches the second speed condition, and if the current speed V does not reach the second speed condition, controlling the pitch angle and the altitude of the unmanned aerial vehicle based on the four-rotor mode, and gradually reducing the four-rotor throttle according to the calculation of the four-rotor throttle incremental controller;

if the airspeed feedback value V reaches the second speed condition, the four-rotor throttle is directly cleared to 0, the pitch angle of the unmanned aerial vehicle is kept near the angle corresponding to the preset optimal lift-drag ratio, and the fixed-wing throttle is controlled to accelerate until the airspeed feedback value V reaches the fixed-wing cruising airspeed V _c To switch to the fixed wing mode, the transition process is completed.

2. The method of claim 1, wherein the input signal to the TECS controller is derived according to the following equationAnd gamma _e ：

Wherein V is airspeed feedback value, K _v G is the gravity acceleration and is the speed proportionality coefficient,is the derivative of the current speed V, H is the current height, K _h And r is the track angle, which is the height proportionality coefficient.

3. The method of claim 2, wherein the TECS controller derives the throttle control amount T in the fixed wing mode based on the following formula _c And an elevator rudder quantity control quantity delta _ec ：

Wherein K is _tp As the total energy proportionality coefficient, K _ti As the total energy differential coefficient, K _ep As the energy difference proportionality coefficient, K _ei As a differential coefficient of the difference in energy,representing the integral operation of the complex domain.

4. The method of claim 1, wherein the TECS controller has a maximum throttle limit of 80% in a fixed wing mode and an elevator amount of 25% in a fixed wing mode.

5. The method of claim 1, wherein an airspeed delta controller is configured such that the airspeed delta controller is configured according to different airspeedsThe maximum limit of the airspeed difference is set by the airspeed feedback value V and the given airspeed V _c Comprises:

before the airspeed feedback value V reaches a preset switching airspeed V_switch, setting the maximum value of airspeed difference to be limited at 15m/s;

after the airspeed feedback value V reaches the switching airspeed V_switch, setting the maximum value of airspeed difference to be 10m/s;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s And fixed-wing cruise airspeed V _c Average value of (2).

6. The method of claim 1, wherein determining whether a pitch angle or a height error of the drone meets a preset landing condition comprises:

if the pitch angle of the unmanned aerial vehicle is larger than 6 degrees or the height error is larger than 15m, judging that the preset landing condition is met;

and if the pitch angle of the unmanned aerial vehicle is not more than 6 degrees and the height error is not more than 15m, judging that the preset landing condition is not met.

7. The method of claim 1, wherein determining whether the airspeed feedback value V reaches the first speed condition includes:

if the airspeed feedback value V reaches the fixed-wing stall airspeed V _s If half of the speed feedback value V reaches the first speed condition;

if the airspeed feedback value V does not reach the fixed-wing stall airspeed V _s And determining that the airspeed feedback value V does not reach the first speed condition.

8. The method of claim 1, wherein determining whether the airspeed feedback value V reaches the second speed condition includes:

if the airspeed feedback value V reaches the switch airspeed V_switch, judging that the airspeed feedback value V reaches a second speed condition;

if the airspeed feedback value V does not reach the switch airspeed V_switch, judging that the airspeed feedback value V does not reach a second speed condition;

wherein the switched airspeed V_switch is set to a fixed-wing stall airspeed V _s And fixed-wing cruise airspeed V _c Average value of (2).

9. The method of claim 1, wherein causing the quad-rotor throttle to be stepped down based on calculations by the quad-rotor throttle delta controller comprises:

four-rotor accelerator incremental controller obtains four-rotor accelerator variation based on the followingAnd airspeed variation (V) _c -V) and let four rotor throttle variation +.>Limited to plus or minus 5% to control four-rotor throttle:

wherein,as a constant term, ρ: air density, C _l : unmanned aerial vehicle lift aerodynamic coefficient, S: unmanned aerial vehicle wing area, ct: lift coefficient of four-rotor motor, C _m As a motor rotation speed constant, ω: four rotor motor speeds.

10. A compound wing unmanned aerial vehicle, wherein the compound wing unmanned aerial vehicle is controlled by a method according to any one of claims 1 to 9 when transitioning from a four rotor mode to a fixed wing mode.