CN118034071A

CN118034071A - Aircraft control distribution method and device, aircraft, storage medium and product

Info

Publication number: CN118034071A
Application number: CN202410445954.7A
Authority: CN
Inventors: 朱明辉; 管习宇; 周文杰; 许兆华; 谢晒明; 薛松柏; 郭亮
Original assignee: Sichuan Wofei Changkong Technology Development Co ltd; Zhejiang Geely Holding Group Co Ltd
Current assignee: Sichuan Wofei Changkong Technology Development Co ltd; Zhejiang Geely Holding Group Co Ltd
Priority date: 2024-04-15
Filing date: 2024-04-15
Publication date: 2024-05-14

Abstract

The application discloses an aircraft control distribution method, an aircraft control distribution device, an aircraft, a storage medium and a product, which belong to the technical field of aircraft control, wherein the aircraft control distribution method is applied to an aircraft with a rotor, the rotor at least comprises a tilting rotor, the rotor is driven by a driving motor, and the method comprises the following steps: performing first power distribution on all available rotors on the aircraft to obtain a first power distribution value, wherein the first power distribution value at least comprises a first rotating speed of a driving motor of each rotor; and performing second power distribution on each rotor wing based on the first power distribution value to obtain a second power distribution value, wherein the second power distribution value at least comprises a second rotating speed of a driving motor of each rotor wing and a tilting angle of the tilting rotor wing. The application provides an aircraft control distribution strategy to improve the controllability of an aircraft and the control distribution stability of the aircraft.

Description

Aircraft control distribution method and device, aircraft, storage medium and product

Technical Field

The application belongs to the technical field of aircraft control, and particularly relates to an aircraft control distribution method and device, an aircraft, a storage medium and a product.

Background

The tilting rotor aircraft is a novel aircraft integrating a fixed-wing aircraft and a propeller aircraft, has the capability of vertical take-off, landing and hovering of a traditional propeller aircraft, and has the high-speed cruising capability of the fixed-wing aircraft.

When the tilting rotor aircraft finishes attitude manipulation, control and redistribution are required to be carried out on available rotors on the aircraft, at present, the traditional solution is to directly solve the required pulling force of the rotors through a pseudo-inverse method and change the rotating speed of each motor, and the method can enable the aircraft to recover to a stable state, but has the condition that part of motor rotating speeds are overlarge and even reach motor rotating speed saturation, especially when the traditional power redistribution mode is adopted to carry out power reconstruction after part of the rotor power of the aircraft fails, if part of motor rotating speeds are overlarge or the motor rotating speeds are saturated, the situation that the available pulling force margin of the rotors is insufficient (namely, the margin for increasing the motor rotating speed is insufficient) or more pulling force cannot be provided (namely, the motor rotating speed cannot be increased any more) is caused, so that the residual manipulation capability of the aircraft is reduced or the stable flying state cannot be maintained.

In summary, how to provide an aircraft control allocation strategy to improve the controllability of an aircraft and improve the control allocation stability of the aircraft has become a technical problem to be solved in the technical field of aircraft control.

Disclosure of Invention

The application mainly aims to provide an aircraft control distribution method, an aircraft control distribution device, an aircraft, a storage medium and a product. The control distribution strategy of the aircraft is provided to improve the controllability of the aircraft and the control distribution stability of the aircraft.

In order to achieve the above object, the present application provides an aircraft control distribution method applied to an aircraft having rotors including at least a tilt rotor, the rotors being driven by a driving motor, the aircraft control distribution method comprising:

performing first power distribution on all available rotors on the aircraft to obtain a first power distribution value, wherein the first power distribution value at least comprises a first rotating speed of a driving motor of each rotor;

and performing second power distribution on each rotor wing based on the first power distribution value to obtain a second power distribution value, wherein the second power distribution value at least comprises a second rotating speed of a driving motor of each rotor wing and a tilting angle of the tilting rotor wing.

Optionally, the step of performing the second power distribution on each rotor based on the first power distribution value includes:

detecting whether the first power distribution value meets a preset condition;

and if the first power distribution value meets the preset condition, performing second power distribution.

Optionally, after the step of detecting whether the first power distribution value meets the preset condition, the method further includes:

If the first power distribution value does not meet the preset condition, performing third power distribution on each rotor wing at least based on part of the first power distribution value to obtain a third power distribution value, wherein the third power distribution value at least comprises a third rotating speed of a driving motor of each rotor wing;

A second power split is performed.

Optionally, the third power distribution adopts a weighted pseudo-inverse method, and the step of performing third power distribution on each rotor wing at least based on a part of the first power distribution value to obtain a third power distribution value includes:

determining a weight for each rotor based at least in part on the first power distribution value;

and carrying out third power distribution on each rotor wing by using the weight to obtain a third power distribution value of each rotor wing.

Optionally, the step of detecting whether the first power distribution value meets the preset condition includes:

Determining that a first power distribution value does not meet the preset condition under the condition that at least one first rotation speed is detected to reach the preset threshold value;

And under the condition that each first rotating speed is detected to not reach the preset threshold value, determining that the first power distribution value meets the preset condition.

Optionally, the step of performing a second power distribution on each rotor based on the first power distribution value to obtain the second power distribution value includes:

And calculating the second rotating speed of the driving motor of each rotor wing and the tilting angle of the tilting rotor wing based on the expected moment and the expected lifting force of the aircraft through a nonlinear optimization algorithm, wherein a cost function of the nonlinear optimization algorithm is constructed based on the maximum pulling force of each rotor wing, and the tilting angle of the tilting rotor wing is used as an optimization variable.

Optionally, the step of calculating, by a nonlinear optimization algorithm, the second rotation speed of the driving motor of each rotor and the tilting angle of the tilting rotor based on the desired moment and the desired lift of the aircraft includes:

a cost function is also constructed based on the tilt angle, tilt direction, and tilt range constraints of the tiltrotors for smoothness constraints on the tilt direction and tilt range of each rotor.

Optionally, the step of performing a first power distribution on each rotor available on the aircraft to obtain a first power distribution value includes:

a first power distribution is made to each rotor available on an aircraft based at least on a current tilt angle of the aircraft's tilt rotor.

In addition, to achieve the above object, the present application provides an aircraft control distribution device applied to an aircraft having a rotor including at least a tilt rotor, the rotor being driven by a driving motor, the aircraft control distribution device comprising:

The initial distribution module is used for carrying out first power distribution on all the available rotors on the aircraft to obtain a first power distribution value, wherein the first power distribution value at least comprises a first rotating speed of a driving motor of each rotor;

And the redistribution module is used for carrying out second power distribution on each rotor wing based on the first power distribution value to obtain a second power distribution value, wherein the second power distribution value at least comprises the second rotating speed of the driving motor of each rotor wing and the tilting angle of the tilting rotor wing.

In addition, to achieve the above object, the present application also provides an aircraft control distribution apparatus including: the system comprises a memory, a processor and an aircraft control allocation program stored on the memory and executable on the processor, wherein the aircraft control allocation program of the aircraft control allocation device, when executed by the processor, implements the steps of the aircraft control allocation method as described above.

In addition, to achieve the above object, the present application also provides a computer-readable storage medium having stored thereon an aircraft control distribution program which, when executed by a processor, implements the steps of the aircraft control distribution method as described above.

The aircraft control distribution method is applied to an aircraft with rotors, the rotors at least comprise tilting rotors, the rotors are driven by driving motors, and the first power distribution value is obtained by carrying out first power distribution on all the rotors available on the aircraft, wherein the first power distribution value at least comprises the first rotating speed of the driving motors of all the rotors, then second power distribution is carried out on all the rotors based on the first power distribution value, and the second power distribution value is obtained, wherein the second power distribution value at least comprises the second rotating speed of the driving motors of all the rotors and the tilting angle of the tilting rotors.

In this way, the application determines the first rotation speed of the driving motor corresponding to each available rotor on the aircraft through the initial power distribution (namely the first power distribution), then redetermines the tilting angle of each tilting rotor in each available rotor on the aircraft and the second rotation speed of the driving motor corresponding to each rotor through the power redistribution (namely the second power distribution), thus, compared with the traditional method of directly calculating the rotation speed of the driving motor corresponding to each available rotor on the aircraft by using the pseudo-inverse method, the application determines the tension value which each rotor should provide to maintain the stable flight of the aircraft, the application is beneficial to balancing the second rotation speed of the driving motor corresponding to each available rotor on the aircraft by adjusting the tilting angle of each available rotor, thereby increasing the control allowance of each available rotor, improving the controllability of the aircraft, reducing the energy consumption, and guaranteeing the failure of the remaining available rotors on the aircraft, especially aiming at the condition of partial rotors of the aircraft, and controlling the residual power distribution of the aircraft to be more stable after the residual power distribution is improved.

Drawings

FIG. 1 is a schematic device architecture diagram of an aircraft control distribution device hardware operating environment in accordance with an embodiment of the present application;

FIG. 2 is a flowchart illustrating steps of a first embodiment of an aircraft control allocation method according to the present application;

FIG. 3 is a schematic diagram of an initial model training process according to an embodiment of the aircraft control allocation method of the present application;

FIG. 4 is a schematic representation of a rotor distribution according to an embodiment of the aircraft control distribution method of the present application;

FIG. 5 is a schematic illustration of a smoothness optimization process according to an embodiment of the aircraft control allocation method of the present application;

FIG. 6 is a schematic diagram of an aircraft control allocation flow scheme in accordance with an embodiment of the aircraft control allocation method of the present application;

FIG. 7 is a schematic representation of yaw angle variation in accordance with an embodiment of the aircraft control allocation method of the present application;

FIG. 8 is a schematic representation of a yaw maximum tension variation according to an embodiment of the aircraft control allocation method of the present application;

FIG. 9 is a schematic representation of changes in pitch angle according to an embodiment of the aircraft control allocation method of the present application;

FIG. 10 is a graphical representation of a change in pitch maximum tension in accordance with an embodiment of the aircraft control allocation method of the present application;

FIG. 11 is a diagonalized view of diagonal directions according to an embodiment of the aircraft control allocation method of the present application;

FIG. 12 is a diagonalized plot of diagonal directions for an embodiment of the aircraft control allocation method of the present application;

FIG. 13 is a non-smooth tilting schematic diagram according to an embodiment of the aircraft control allocation method of the present application;

FIG. 14 is a schematic view of smooth tilting according to an embodiment of the aircraft control allocation method of the present application;

FIG. 15 is a graph illustrating a maximum drag coefficient profile according to an embodiment of the aircraft control allocation method of the present application;

FIG. 16 is a schematic diagram of model optimization capabilities according to an embodiment of the aircraft control allocation method of the present application;

FIG. 17 is a functional block diagram of an embodiment of an aircraft control dispensing device of the present application.

The achievement of the objects, functional features and advantages of the present application will be further described with reference to the accompanying drawings, in conjunction with the embodiments.

Detailed Description

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 application.

Referring to fig. 1, fig. 1 is a schematic view of an apparatus architecture of a hardware operating environment of an aircraft control distribution apparatus according to an embodiment of the present application.

It should be noted that the embodiment of the application relates to an aircraft with a rotor wing in the technical field of aircraft control. In particular, the rotor on board the aircraft comprises at least a tilt rotor, which is driven by a drive motor.

As shown in fig. 1, the aircraft control distribution device may include: a processor 1001, such as a CPU, a communication bus 1002, a user interface 1003, a network interface 1004, and a memory 1005. Wherein the communication bus 1002 is used to enable connected communication between these components. The user interface 1003 may include a display (DiSplay), an input unit such as a Keyboard (Keyboard), and the optional user interface 1003 may also include a standard wired interface, a wireless interface. The network interface 1004 may optionally include a standard wired interface, a wireless interface (e.g., wi-Fi interface). The memory 1005 may be a high-speed RAM memory or a stable memory (non-volatile memory), such as a disk memory. The memory 1005 may also optionally be a storage device separate from the processor 1001 described above. The execution bodies of the method steps in the embodiments are omitted below for convenience of description.

Those skilled in the art will appreciate that the aircraft control dispensing device structure shown in fig. 1 is not limiting of the aircraft control dispensing device and may include more or fewer components than shown, or may combine certain components, or a different arrangement of components.

As shown in fig. 1, an operating system, a network communication module, a user interface module, and an aircraft control distribution program may be included in the memory 1005 as one type of computer storage medium.

In the terminal shown in fig. 1, the network interface 1004 is mainly used for connecting to a background server and performing data communication with the background server; the user interface 1003 is mainly used for connecting a client and communicating data with the client; and processor 1001 may be configured to invoke the aircraft control allocation program stored in memory 1005 and perform the following operations:

Further, the operation of performing the second power distribution on each rotor based on the first power distribution value includes:

detecting whether the first power distribution value meets a preset condition;

Further, after the step of detecting whether the first power distribution value meets the preset condition, the processor 1001 may be further configured to invoke the aircraft control distribution program stored in the memory 1005 to perform the following operations:

A second power split is performed.

Further, the third power distribution adopts a weighted pseudo-inverse method, and the operation of performing third power distribution on each rotor wing based at least on a part of the first power distribution value to obtain a third power distribution value includes:

Further, the preset condition is that each first rotation speed does not reach a preset threshold, and the operation of detecting whether the first power distribution value meets the preset condition includes:

Further, the operation of performing the second power distribution on each rotor based on the first power distribution value to obtain the second power distribution value includes:

Further, the operation of calculating the second rotation speed of the driving motor of each rotor wing and the tilting angle of the tilting rotor wing based on the expected moment and the expected lifting force of the aircraft through the nonlinear optimization algorithm comprises the following steps:

Further, the operation of performing a first power distribution on each rotor available on the aircraft to obtain a first power distribution value includes:

Based on the above-described structure, various embodiments of an aircraft control allocation method are presented.

Referring to fig. 2, fig. 2 is a flowchart illustrating a first embodiment of an aircraft control distribution method according to the present application. It should be noted that although a logical sequence is illustrated in the flow chart, in some cases the aircraft control allocation method of the present application may of course also perform the steps illustrated or described in a different order than that which is illustrated herein. In the present embodiment, the execution subject of the aircraft control allocation method may be a flight control computer, a portable computer, or the like, and is not limited in the present embodiment. The aircraft control allocation method comprises the following steps:

step S10, performing first power distribution on all available rotors on the aircraft to obtain a first power distribution value, wherein the first power distribution value at least comprises a first rotating speed of a driving motor of each rotor;

The application does not limit the specific method for carrying out the first power distribution on each rotor wing, and can be a pseudo-inverse method, a priority distribution method (different priorities are set for different rotor wings according to the requirements of flight tasks, and the power supply of key rotor wings is preferentially ensured), an iterative optimization method (the power distribution scheme is gradually optimized through iterative calculation so as to achieve better flight performance and stability) and the like. For better explanation, the conventional pseudo-inversion method is used for explanation, and the calculation logic of the pseudo-inversion method is as follows: the equation for the individual rotor moment and lift is first defined as follows:

wherein/> Representing the moment of a single rotor of an aircraft on the x-axis of a relative coordinate system,/>Representing the moment of a single rotor of an aircraft on the y-axis of a relative coordinate system,/>Representing the moment of a single rotor of an aircraft on the z-axis of a relative coordinate system,/>Representing the lift required by a single rotor of an aircraft,/>For the distance of a single rotor to the mass centre of the aircraft body,/>Azimuth angle for single rotor,/>For the mounting angle of a single rotor motor,/>For turning of a single rotor,/>Is the counter-torque coefficient of the individual rotors. /(I)Is the tilting angle of the aircraft,/>To/>Normalized lift values for individual rotors. Illustratively, in the case of current aircraft comprising 8 rotors, the value of n may range from 1 to 8, i.e. the above formula may be considered as a 4 by 8 matrix. In addition, the total virtual moment and virtual lift (virtual moment and virtual lift represent the moment and lift expected by the aircraft) generated by the 8 rotors is as follows:

The solving process of the pseudo-inverse method is as follows:

Definition:

wherein v denotes the desired moment and lift of the aircraft, and comprises in particular 、/>、/>And/>，、/>And/>Respectively represent the moment of the aircraft in three directions under the relative coordinate system,/>Representing the lift required by the aircraft; u represents the tension value provided by each rotor of the aircraft in case v is satisfied; /(I)Representing an efficiency matrix of the aircraft. In addition, the pseudo-inverse method is a minimum energy norm method, and the solving thought is to ensure that the two norms and the minimum of the control quantity are:

for the shape as described above/> The problem of (1) pseudo-inverse method is directly selected/>Pseudo-inverse of (2) and desired lift and moment/>The product of:

In this embodiment, power distribution is performed on each rotor available on the aircraft (hereinafter referred to as a first power distribution to show a distinction), so as to obtain a power distribution value (hereinafter referred to as a first power distribution value to show a distinction) that each rotor should provide, where the first power distribution value includes at least a rotational speed of a driving motor (hereinafter referred to as a first rotational speed to show a distinction) corresponding to each rotor. The rotor is driven by the driving motor, and when the rotation speed of the driving motor is higher, the pulling force provided by the rotor to the aircraft is higher.

And step S20, carrying out second power distribution on each rotor wing based on the first power distribution value to obtain a second power distribution value, wherein the second power distribution value at least comprises a second rotating speed of a driving motor of each rotor wing and a tilting angle of the tilting rotor wing.

In the present embodiment, after the first power distribution value is obtained, power distribution is performed on each rotor (hereinafter referred to as second power distribution to show distinction) based on the first power distribution value, and respective power distribution values (hereinafter referred to as second power distribution value to show distinction) of each rotor are obtained, wherein the second power distribution values include at least the rotation speed of the driving motor of each rotor (hereinafter referred to as second rotation speed to show distinction) and the tilting angle of the tilting rotor on the aircraft.

It should be noted that, when the control and distribution are performed only by the conventional control and distribution strategy in step S10, in order to ensure the yaw moment balance, the lift distribution may be unbalanced, which may cause a problem that the partial rotor tension is too high under the severe flight control condition. Through the re-power distribution of step S20, the controllability of the aircraft is enhanced and the overall stability of the aircraft is maintained through the tilting angle of the tilting rotor.

In this embodiment, the step S10 includes:

Step S101, a first power distribution is performed on each rotor available on an aircraft based at least on the current tilt angle of the tilting rotor of said aircraft.

In this embodiment, when first power is distributed to each rotor available on the aircraft, the first power is distributed to each rotor available on the aircraft based at least on the current tilt angle and the desired flight status of the tilt rotor of the aircraft.

In one possible embodiment, since the angle change of the tilting rotor changes the magnitude and direction of the pulling force generated by the rotor, thereby affecting the flight state of the aircraft, the efficiency matrix of the aircraft is calculated based on the current tilting angle of the tilting rotor of the aircraft, and then the inverse matrix of the efficiency matrix and the expected matrix of the aircraft are multiplied based on the efficiency matrix of the aircraft and the expected matrix of the aircraft, and the element in the obtained matrix is the first power distribution value that each rotor of the aircraft can provide in a stable flight state.

In this embodiment, the step S20 further includes:

step A10, detecting whether the first power distribution value meets a preset condition;

and step A20, if the first power distribution value meets the preset condition, performing second power distribution.

It should be noted that, the condition for determining the first power distribution value is preset that each first rotation speed does not reach the preset threshold value. Wherein the preset threshold is a value lower than the saturated rotation speed of the motor.

In this embodiment, after the first power distribution value of each rotor is obtained, it is detected whether the first power distribution value meets a preset condition, and if the first power distribution value meets the preset condition, the second power distribution is performed.

In one possible embodiment, after the tension value of each rotor wing of the aircraft is adjusted to the tension value corresponding to the first power distribution value, whether the respective motor rotation speed of each rotor wing is saturated or not is judged, if the respective motor rotation speed of each rotor wing is detected to be unsaturated, the aircraft is determined to be in an unsaturated state, and the second power distribution is performed on each rotor wing on the aircraft.

In this embodiment, the preset condition is that each first rotation speed does not reach a preset threshold, and the step a10 includes:

step A101, determining that a first power distribution value does not meet the preset condition under the condition that at least one first rotation speed is detected to reach the preset threshold value;

step a102, determining that the first power distribution value meets the preset condition when detecting that each first rotation speed does not reach the preset threshold value.

In this embodiment, the first power distribution value is determined not to meet the preset condition when the at least one first rotation speed is detected to reach the preset threshold, and the first power distribution value is determined to meet the preset condition when the at least one first rotation speed is detected to reach the preset threshold.

In this embodiment, after the step a10, the method further includes:

Step B10, if the first power distribution value does not meet the preset condition, performing third power distribution on each rotor wing at least based on part of the first power distribution value to obtain a third power distribution value, wherein the third power distribution value at least comprises a third rotating speed of a driving motor of each rotor wing;

And step B20, performing second power distribution.

In this embodiment, if the first power distribution value does not meet the preset condition, power distribution is performed on each rotor (hereinafter referred to as third power distribution for illustrating distinction) based on at least a portion of the first power distribution value (i.e., each first rotational speed), so as to obtain a power distribution value (hereinafter referred to as third power distribution value for illustrating distinction) of each rotor, where the third power distribution value includes at least a rotational speed of a driving motor of each rotor (hereinafter referred to as third rotational speed for illustrating distinction), and then second power distribution is performed.

In a possible implementation manner, the third power allocation adopts a weighted pseudo-inverse method, and the step B10 includes:

Step B101, determining the weight of each rotor wing at least based on part of the first power distribution value;

And step B102, performing third power distribution on each rotor wing by using the weight to obtain a third power distribution value of each rotor wing.

In this embodiment, the weight of each rotor is determined based at least on a part of the first power distribution value, and the third power distribution value of each rotor is obtained by performing the third power distribution on each rotor based on each weight.

In a specific embodiment, the inverse of a tension value (hereinafter referred to as a first tension value to indicate distinction) to be provided by each rotor is taken as the weight of each rotor, and the element corresponding to each rotor in the efficiency matrix of the aircraft is multiplied by the weight of each rotor to obtain the target efficiency matrix.

Illustratively, the speed limit and the bandwidth of different operation surfaces of the aircraft are different, so that in actual design, different rotors are provided with different weights, and the method can effectively avoid partial motor saturation, and the allocation index function can be defined as:

The solution of the above formula is:

As can be seen from the distribution solution, the weighted pseudo-inverse method is a distribution strategy in which each control surface participates in control, and can realize better distribution, but the distribution result and the weight matrix And the choice of pseudo-inverse. At the same time, this approach is generally only able to avoid saturation, but is not effective in reducing the maximum drag of the rotor.

In this way, the method is applied to an aircraft with rotors, the rotors at least comprise tilting rotors, the rotors are driven by driving motors, the first rotation speeds of the driving motors corresponding to the available rotors on the aircraft are determined through initial power distribution (namely the first power distribution), then saturation judgment is carried out on each first rotation speed, when at least one first rotation speed reaches saturation, the weighted pseudo-inverse method is adopted to carry out power distribution (namely the third power distribution) on each available rotor on the aircraft again, the third rotation speeds of the driving motors corresponding to the available rotors on the aircraft are obtained, the rotation speeds of the driving motors corresponding to the available rotors are balanced, the control margin of each available rotor is increased, then the tilting angle of each tilting rotor in each available rotor on the aircraft and the second rotation speed of the driving motor corresponding to each rotor on the aircraft are determined through power redistribution (namely the second power distribution), or when each first rotation speed does not reach saturation, the second power distribution is directly carried out. Therefore, compared with the traditional mode of directly calculating the rotation speed of each driving motor corresponding to each available rotor on the aircraft by using a pseudo-inverse method to determine the tension value which each rotor should provide so as to maintain the stable flight of the aircraft, the method and the device disclosed by the application have the advantages that the third rotation speed distributed by each driving motor in the unsaturated state is calculated by adopting the weighted pseudo-inverse method, and the tilting angle of each available tilting rotor on the aircraft is regulated, so that the second rotation speed distributed to each driving motor corresponding to each available rotor is balanced, the control allowance of each available rotor is increased, the guarantee is provided for the subsequent flight, and especially, the tilting angle of each available rotor on the aircraft is regulated so that the power distribution of each available rotor remained on the aircraft is more balanced, the remaining control allowance of the aircraft is increased, and the stability of the aircraft after part of the rotors fail is improved, namely the control distribution stability of the aircraft is improved.

Further, based on the first embodiment of the aircraft control allocation method according to the application described above, a second embodiment of the aircraft control allocation method according to the application is proposed.

In this embodiment, the step S20 includes:

And step C10, calculating the second rotating speed of the driving motor of each rotor wing and the tilting angle of the tilting rotor wing based on the expected moment and the expected lifting force of the aircraft through a nonlinear optimization algorithm, wherein a cost function of the nonlinear optimization algorithm is constructed based on the maximum pulling force of each rotor wing, and the tilting angle of the tilting rotor wing is used as an optimization variable.

In this embodiment, firstly, the desired moment and the desired lift force of the aircraft are determined, the tilting angle of the tilting rotor is used as an optimization variable by taking the maximum tension of the tensions provided by the respective rotors as a cost function, iterative optimization is completed in the tilting range of the rotor through a nonlinear optimization algorithm, and finally, the tilting angle of each tilting rotor and the tension value required by each rotor are calculated.

For example, the desired moment and lift of the aircraft are first determined, and a cost function is built based on the maximum pull, i.e. the equationAnd (3) taking the tilting angle of the residual tilting rotor wing as an optimization variable, finishing iterative optimization in the tilting range of the rotor wing through a nonlinear optimization algorithm, finally calculating the tilting angle of each rotor wing and the required tension, executing a control instruction, and changing the tilting angle of the rotor wing and the rotating speed of a motor so as to achieve a target state. Yaw moment distribution of the tilting differential control rotor wing meets the controllability of the rotor wing. It should be noted that the present application is not limited to the specific type of the nonlinear optimization algorithm, and the nonlinear optimization algorithm may be an ant colony algorithm, a particle algorithm, an annealing algorithm, or the like. As shown in fig. 3, which is an initial model training flow diagram, the initial model to be trained is an ant colony algorithm model, firstly, initializing a population, judging whether the iteration number in the current training process exceeds a preset iteration threshold, if so, taking the current model as a target model for outputting a tilting angle with highest adaptability; if the iteration threshold value is not exceeded, calculating the adaptability of the population, performing cross operation on the population genes to generate a new solution to increase the diversity of the population, and performing mutation operation on the population genes to generate new individuals to increase the diversity of the population; decoding the population genes, updating the population, judging whether the loss function of the current model is converged, if so, determining the current model as a target model, and if not, continuing iterative training.

Specifically, the initial model to be trained is used as a model to be updated, the model input data is input into the model to be updated, and a plurality of prediction angles of each rotor wing output by the model to be updated are obtained. It should be noted that, there is at least one target tension value output by the target model based on the input tilt angle of the rotor, that is, the target model outputs at least one prediction result each time, that is, the change of the rotor tension value is divided into a plurality of adjustment periods, and the whole adjustment process is smooth. For example, the tilting angle of the current residual rotor wing is tilted from 54 degrees to 70 degrees, the tilting process is 54 degrees to 58 degrees, 58 degrees to 63 degrees, 63 degrees to 67 degrees and 67 degrees to 70 degrees, wherein the predicted result output by the target model based on the input 54 degrees is a tension value corresponding to each of 63 degrees, 67 degrees and 70 degrees. It will be appreciated that the 63 degrees, 67 degrees and 70 degrees described above are intermediate states of the target model. The model to be updated determines a first predicted intermediate value according to input data, predicts a second predicted intermediate value based on the first predicted intermediate value until an optimal prediction result is obtained, namely, each intermediate state value is generated in a sequence relation, and the tilting process from a prediction angle corresponding to a first moment to a prediction angle corresponding to a second moment is regarded as a tilting period, wherein the first moment and the second moment are adjacent and the first moment is earlier than the second moment. And determining the respective tilting directions of the tilting periods, namely, the tilting directions when the tilting periods tilt from the predicted angle corresponding to the first moment to the predicted angle corresponding to the second moment.

In this embodiment, the step C10 includes:

step C101, further constructing a cost function based on the tilt angle, tilt direction and tilt range constraints of the tilt rotors, for performing smoothness constraints on the tilt direction and tilt range of each rotor.

In this embodiment, a cost function is constructed based on the tilt angle, tilt direction and preset tilt range constraints of the tilt rotors on the aircraft, wherein the cost function is used to smoothness constrain the tilt direction and tilt range of each rotor.

In a possible implementation manner, a cost function is further constructed based on a reverse coefficient, wherein the reverse coefficient of each tilting period is determined based on each tilting direction, specifically, the reverse coefficient is used for measuring smoothness of the tilting process, when the tilting direction of the current tilting period is in the same direction as the tilting direction of the previous tilting period, the reverse coefficient is zero, and when the tilting direction of the current tilting period is in the reverse direction of the previous tilting period, the reverse coefficient is a preset value, the specific size of the preset value is not limited, and the reverse coefficient can be any value greater than zero. Then, respective cost functions of the respective tilting periods are constructed based on respective inverse coefficients of the respective tilting periods.

Based on the control distribution method of the embodiment, from the perspective of an optimization algorithm, in the tilting range of the rotor, the minimum maximum pulling force is used as an optimization target, the target tilting angle of each tilting rotor is calculated, continuous power distribution of the rotor craft is completed in the rotation process of the rotor, stable flying is realized, the problems of overhigh power of part of motors, saturated pulling force and the like are avoided, the control allowance is increased, and guarantee is provided for subsequent stable flying.

Illustratively, the cost function in the defined optimization algorithm is as follows:

In the method, in the process of the invention, In order to control the efficiency matrix, the tilt angles of 8 rotors are determined, wherein the 1,4, 5 and 8 rotors are vertical rotors, as shown in fig. 4, which is a schematic diagram of rotor distribution, and the tilt angles of the 1,4, 5 and 8 rotors are always 90 degrees, so that the influence/>Mainly the angle of the other 4 tiltrotors; and/>For the inversion coefficient, which is mainly determined by the tilting direction, tilting angle and current population value of the current state, the inversion coefficient can be expressed asWherein angle is the tilting angle corresponding to the second moment in the current tilting period, angle_pre is the tilting angle corresponding to the first moment in the current tilting period, and when the sign of the result value of subtracting angle from angle_pre is the same as the sign of dir (the tilting direction of the previous tilting period), the reverse coefficient is determined to be zero. Through the method, the current motion state is considered in the loss function of each round of optimization, and the local information is constrained, so that global optimization is achieved.

Therefore, the application constrains the tilting direction and the calculation range of the rotor wing in the cost function, and ensures the smoothness of tilting.

As shown in fig. 5, which is an exemplary schematic diagram of a smoothness optimization process, the tilt angles of 4 non-vertical rotors in the aircraft in the current state are determined, namely, a tilt angle 1, a tilt angle 2, a tilt angle 3 and a tilt angle 4, and the tilt direction of the current tilt period, wherein it is to be noted that the application is limited to the degree of tilting in each period, and the specific constraint is within the range of [ -5, +5 ]; constructing a loss function with a smoothness constraint, wherein in a first tilting period, the loss function can take a preset numerical value as a reverse coefficient to construct the loss function; optimizing model parameters based on the loss of the current period through the model to be updated so as to calculate the tilting angle of the next tilting period; judging whether the currently predicted tilting angle is stable, if so, determining that tilting is finished, and if not, continuously calculating the target tilting angle of the next tilting period.

Therefore, the embodiment of the application adopts a nonlinear optimization algorithm to adjust the tilting angle of the tilting rotor wings available on the aircraft, is beneficial to balancing the second rotating speeds of the driving motors corresponding to the available rotor wings, thereby increasing the control margin of the available rotor wings, improving the controllability of the aircraft, reducing the energy consumption and providing guarantee for subsequent flight.

Further, based on the first and/or second embodiments of the aircraft control allocation method according to the application described above, a third embodiment of the aircraft control allocation method according to the application is proposed.

It should be noted that, the first tension value corresponding to the first power distribution value is a tension value which is determined based on the pseudo-inverse method and is provided by the remaining rotor wings and can maintain stable flight of the aircraft, but at this time, the situation that the motor rotation speed corresponding to the first tension value is too high or even reaches saturation may exist, so that the stable state of the aircraft is maintained at the first time of the fault of the aircraft, then whether the motor rotation speed of each remaining rotor wing reaches saturation is judged, if the motor rotation speed of each rotor wing is detected to be unsaturated, the aircraft is determined to be in an unsaturated state, and the second power distribution is performed on each rotor wing on the aircraft. If the motor speed of at least one rotor reaches saturation, the power distribution (namely the third power distribution) is recalculated on each rotor on the aircraft based on a weighted pseudo-inverse method.

In addition, after obtaining the third tension value that each rotor that the aircraft can provide in the unsaturated state should provide, the motor rotational speed that each rotor that can be used should be adjusted to the third tension value that each rotor that can be used should provide. As shown in fig. 6, an aircraft control distribution flow chart is schematically shown, and is divided into three distribution stages, namely a first power distribution, a second power distribution and a third power distribution, wherein the distribution flow of the first power distribution is to detect whether a rotor wing of an aircraft fails in a normal flight process of the aircraft; if the failure of part of the rotor wings of the aircraft is detected, automatically shutting down the diagonal rotor wings symmetrical to the failed rotor wings, and then calculating the motor rotation speed of the rest rotor wings based on a pseudo-inverse method; the third power distribution flow is that after the motor rotation speed of the residual rotor wing is calculated based on a pseudo-inverse method, whether the motor rotation speed of the rotor wing reaches saturation is judged, if the motor rotation speed of at least one residual rotor wing reaches saturation, the motor rotation speed of each residual rotor wing is recalculated based on a weighted pseudo-inverse method so as to avoid the motor rotation speed saturation; the second power distribution flow is to determine the tilting range of the rotor wings when the motor rotation speed of each remaining rotor wing does not reach saturation, calculate the optimal tilting angle of each rotor wing (namely, the tilting angle corresponding to the target tension value) based on a pre-trained nonlinear optimization algorithm, tilt the rotor wings based on a differential method, adjust the tilting angle of the remaining rotor wings to the optimal tilting angle, and change the electrode rotation speed based on the target tension value corresponding to the optimal tilting angle.

According to the application, when the motor rotation speed saturation of the residual rotor wing which is primarily distributed through the pseudo-inverse method is detected, the weighted pseudo-inverse method is adopted to reconstruct the control of the residual rotor wing, so that the motor rotation speed of the residual rotor wing is prevented from reaching saturation, and the stable flight of the aircraft is ensured.

In a possible embodiment, in the case of yaw manipulation of the aircraft, the desired roll moment is set to 0, the desired pitch moment is set to 0, and the desired yaw moment is set to 0.7; the desired control amount is [0, 0, 0.7, 4]; the initial tilting angle is 90 °. As shown in table 1 below, table 1 is a first profile containing the values of the individual rotor tensions of the aircraft at the initial angles during yaw maneuvers:

TABLE 1 first distribution table

As shown in table 2 below, table 2 is a second distribution table containing the values of the optimized rotor tension during yaw manipulation:

TABLE 2 second distribution list

Exemplary, as shown in fig. 7, a yaw angle change schematic diagram is shown in fig. 7 (a) which is an angle change curve of the rotor No. 2 during yaw manipulation, fig. 7 (b) which is an angle change curve of the rotor No. 3 during yaw manipulation, fig. 7 (c) which is an angle change curve of the rotor No. 6 during yaw manipulation, and fig. 7 (d) which is an angle change curve of the rotor No. 7 during yaw manipulation, and an abscissa of a coordinate system in which the angle change curve is calculated as a number of times of a nonlinear optimization algorithm and an ordinate is a tilting angle. As shown in fig. 8, which is a diagram of yaw maximum tension change, the abscissa of the coordinate system in fig. 8 is the calculated number of times of the nonlinear optimization algorithm, and the ordinate is the maximum rotor tension, and the change of the maximum rotor tension gradually decreases with the increase of the calculated number of times of the nonlinear optimization algorithm until the maximum rotor tension tends to be stable.

Therefore, when the aircraft performs yaw manipulation, in an initial state, the maximum rotor wing pulling force is about 0.83, and the tilting angle is adjusted by adopting a nonlinear optimization algorithm, so that the maximum rotor wing pulling force gradually decreases, and finally, the maximum rotor wing pulling force is about 0.54, 35% of the maximum rotor wing pulling force is reduced, and the residual manipulation capability of the rotor wing is greatly improved.

In another possible embodiment, in case of pitch maneuvers of the aircraft, the desired roll moment is set to 0, the desired pitch moment is set to-6, and the desired yaw moment is set to 0; the desired control amount is [0, -6, 0, 4]; the initial tilt angle is 90 deg., unlike yaw maneuvers, the tilt rotor tilt angles on the front and back sides of the pitch maneuver are typically equal, so the optimization variable can be reduced from 4 to 2. Table 3 below is a third profile containing the individual rotor pull values for the aircraft at the initial angle during pitch maneuvers:

TABLE 3 third distribution Table

Table 4 below is a fourth profile containing the optimized individual rotor pull values for pitch maneuvers:

TABLE 4 fourth distribution table

As illustrated in fig. 9, the pitch angle change is schematically shown in fig. 9, in which fig. 9 (a) is an angle change curve of the rotor No. 2 during pitch operation, fig. 9 (b) is an angle change curve of the rotor No. 3 during pitch operation, fig. 9 (c) is an angle change curve of the rotor No. 6 during pitch operation, and fig. 9 (d) is an angle change curve of the rotor No. 7 during pitch operation, and the abscissa of the coordinate system where the angle change curve is calculated by the nonlinear optimization algorithm and the ordinate is the tilting angle. As shown in fig. 10, which shows a diagram of the change of the maximum pitching tension, the abscissa of the coordinate system in the diagram is the calculated number of times of the nonlinear optimization algorithm, and the ordinate is the maximum rotor tension, and the change of the maximum rotor tension gradually decreases with the increase of the calculated number of times of the nonlinear optimization algorithm until the maximum rotor tension tends to be stable.

Therefore, when the aircraft performs pitching operation, the maximum rotor wing pulling force is about 0.86 in an initial state, and the tilting angle is adjusted by adopting a nonlinear optimization method, so that the maximum rotor wing pulling force gradually decreases, and finally about 0.76, 11% of the maximum rotor wing pulling force is reduced, and the residual operation capability of the rotor wing is effectively improved.

In another possible embodiment, after a certain rotor of an aircraft with rotors fails, in order to ensure the stability of the rotor flight, control redistribution needs to be performed on the remaining available rotors, the current mainstream multi-rotor layout is symmetrically distributed about the centroid, and pitch and roll moment balance can be achieved by usually closing the diagonal rotors, and for yaw moment, the following two situations can be classified: firstly, as shown in fig. 11, which shows different diagonal directions, R1 to R8 represent 8 rotors of an aircraft, clockwise rotation is defined as positive, anticlockwise rotation is defined as negative, if the rotation direction of the diagonal rotor is opposite to that of a failed rotor, the diagonal rotor can be directly closed to realize yaw moment balance, and after the rotor R1 fails, the diagonal rotor R8 is closed at the same time, and the number of positive and negative directions of the remaining rotors R2 to R7 is the same; second, as shown in fig. 12, the same diagonal direction is schematically shown, if the rotation direction of the diagonal rotor of the failed rotor is the same as that of the failed rotor, the number of yaw moments of the remaining rotor is unbalanced due to the closing of the diagonal rotor, so that after the rotor R1 fails, the diagonal rotor R8 is closed.

In a specific embodiment, if the total number of rotors on the aircraft is 8, and the number of the available rotors is 6 when two rotors fail, then the inverse matrix of the efficiency matrix (matrix of 4 by 6) corresponding to the 6 rotors is multiplied by the expected matrix (matrix of 1 by 4), so as to obtain a matrix of 1 by 6, and 6 elements in the matrix respectively represent the first power distribution values which should be provided by the 6 available rotors, and when the 6 available rotors provide the first power distribution values which need to be provided by the respective rotors, the flight state of the aircraft is stable.

In the case of a second power distribution for each available rotor remaining on the aircraft, table 5 below is a fifth distribution table containing the rotor tension distribution in the standard state of the aircraft:

Table 5 fifth distribution table

In the standard case, when the tilt angles of all the rotors are 90 degrees, the control amount is defined as:

At this time, the tension of each rotor wing is the same as 0.5028N, and the tension value is set as the standard tension . At the same time define the maximum tension coefficient/>Definition/>For maximum tension value, the maximum tension coefficient is used for representing the optimization effect of the target model:

table 6, shown below, is a sixth profile containing rotor tension profile during aircraft flight under partial motor failure conditions:

TABLE 6 sixth distribution table

When the rotors 1 and 8 fail, the residual rotors are used for redistribution, the maximum pulling force is 0.8890N, and the ratio of the maximum pulling force to the standard pulling force1.7681.

Based on the above case analysis, when the tilt angle of the tilt rotor is 54 °, the maximum tension of the motor is about 1.77 times that of the standard state, which is also the case when the diagonal rotor fails simultaneously, if only the rotor No. 1 fails, this factor exceeds 2, which greatly increases the load of the motor and reduces the control margin.

The differential tilting may gradually reduce the scaling factor, and table 7 below is a seventh distribution table, including the rotor tension distribution under differential tilting:

TABLE 7 seventh distribution table

When the rotors No.1 and No. 8 fail, the optimal tilting angles of the rotors No. 2, 3, 6 and No. 7 are calculated through a nonlinear optimization algorithm, so that the rotor can be used forFrom 1.7681 to 1.3461. The tilting angle of the front rotor wing is 93-94 degrees, the tilting angle of the rear rotor wing is 86-87 degrees, and differential tilting is realized under the condition that the aircraft is kept stable.

On the other hand, the calculation range at each time can be limited to be within + -5 DEG, and the optimal value at the time is taken as the state at the next time until the maximum tension coefficientAnd the whole calculation flow is shown in fig. 5.

The tilting process of the rotor from 54 to 94 must be smooth and continuous, and the calculation of each point should satisfy the control quantityIf the nonlinear optimization algorithm is directly used for real-time calculation, a non-monotonic tilting state is generated, as shown in fig. 13, which is a non-smooth tilting diagram, fig. 13 (a) is a non-smooth tilting curve of the rotor No.2, fig. 13 (b) is a non-smooth tilting curve of the rotor No. 3, fig. 13 (c) is a non-smooth tilting curve of the rotor No. 6, and fig. 13 (d) is a non-smooth tilting curve of the rotor No. 7, and the abscissa of the coordinate system where the tilting curve is located is time, the ordinate is tilting angle, the tilting process of each rotor from the initial angle to the target angle is not smooth and continuous, which affects the stability of the aircraft, and a smoothing constraint should be added, as shown in fig. 14, which is a smooth tilting diagram, fig. 14 (a) is a smooth tilting curve of the rotor No.2, fig. 14 (b) is a smooth tilting curve of the rotor No. 3, fig. 14 (c) is a smooth tilting curve of the rotor No. 6, and fig. 14 (d) is a smooth tilting curve of the rotor No. 7, each rotor is tilted from the initial angle to the target angle, which is continuous. Maximum coefficient of tension of the whole process/>As shown in fig. 15, which is a graph showing a distribution curve of the maximum tension coefficient corresponding to fig. 14, the coefficient gradually decreases with the change of the tilting angle, and after 8s, the maximum tension tends to be stable, so that the tilting process can be ended.

In addition, considering the current flight state of the aircraft, a large-scale tilting process may not be completed, for example, only optimization calculation can be performed within the range of +/-10 degrees and +/-20 degrees of the current state. The calculation can be optimized online, but in consideration of large consumption of online calculation resources and difficult guarantee of real-time performance, and the tilting differential method is only related to the tilting angles, so that the optimal values which can be achieved under each tilting angle can be counted in an offline calculation mode. For example, as shown in fig. 16, a schematic diagram of model optimization capability is shown, and the optimization capability is shown in the range of + -10 deg., + -20 deg., for example, under the condition that the tilting angle is 55 deg., the maximum tension coefficient exceeds 1.75, and the optimization capability can be optimized to 1.53 and 1.44 in the ranges of 45 deg. -65 deg. and 35 deg. -75 deg., respectively.

Therefore, the application not only can solve the problem of failure of the vertical rotor, but also is applicable to the failure of part of the tilting rotor, and the optimization flow is the same, and only the number of the optimization parameters is required to be changed. In addition, the diagonal rotor is closed, the tension distribution is more balanced, and the maximum tension value is smaller.

In addition, the embodiment of the application also provides an aircraft control distribution device, which is applied to an aircraft with a rotor, wherein the rotor at least comprises a tilting rotor, and the rotor is driven by a driving motor.

Referring to fig. 17, fig. 17 is a schematic functional block diagram of an embodiment of an aircraft control and distribution device according to the present application, as shown in fig. 17, the aircraft control and distribution device according to the present application includes:

An initial distribution module 10, configured to perform a first power distribution on each rotor available on the aircraft, to obtain a first power distribution value, where the first power distribution value includes at least a first rotational speed of a driving motor of each rotor;

And the redistribution module 20 is configured to perform a second power distribution on each rotor based on the first power distribution value, so as to obtain a second power distribution value, where the second power distribution value includes at least a second rotation speed of a driving motor of each rotor and a tilting angle of the tilting rotor.

Further, the redistribution module 20 includes:

the detection unit is used for detecting whether the first power distribution value accords with a preset condition;

and the first detection result unit is used for carrying out second power distribution if the first power distribution value meets the preset condition.

Further, the aircraft control distribution device of the present application further comprises:

The third power distribution module is used for carrying out third power distribution on each rotor wing at least based on part of the first power distribution values if the first power distribution values do not meet the preset conditions to obtain third power distribution values, wherein the third power distribution values at least comprise third rotating speeds of driving motors of the rotor wings; a second power split is performed.

Further, the third power distribution adopts a weighted pseudo-inverse method, and the third power distribution module is further used for determining the weight of each rotor wing at least based on part of the first power distribution value; and carrying out third power distribution on each rotor wing by using the weight to obtain a third power distribution value of each rotor wing.

Further, the preset condition is that each first rotation speed does not reach a preset threshold, and the detection unit is further configured to determine that a first power distribution value does not meet the preset condition when at least one first rotation speed is detected to reach the preset threshold; and under the condition that each first rotating speed is detected to not reach the preset threshold value, determining that the first power distribution value meets the preset condition.

Further, the redistribution module 20 is further configured to calculate, by using a nonlinear optimization algorithm, the second rotation speed of the driving motor of each rotor and the tilt angle of the tilt rotor based on the desired moment and the desired lift of the aircraft, where a cost function of the nonlinear optimization algorithm is constructed based on the maximum pull force of each rotor, and the tilt angle of the tilt rotor is used as an optimization variable.

Further, the redistribution module 20 is further configured to construct a cost function based on the tilt angle, the tilt direction and the tilt range constraint of the tilt rotors, and to perform a smoothness constraint on the tilt direction and the tilt range of each rotor.

Further, the initial distribution module 10 is also configured to perform a first power distribution of each rotor available on the aircraft based at least on the current tilt angle of the aircraft's tilt rotor.

The present application also provides a computer storage medium having stored thereon an aircraft control distribution program which, when executed by a processor, implements the steps of the aircraft control distribution program method according to any of the embodiments above.

The specific embodiments of the computer storage medium of the present application are substantially the same as the embodiments of the method for controlling and distributing the program of the aircraft of the present application described above, and will not be described herein.

The application also provides a computer program product comprising a computer program which, when executed by a processor, implements the steps of the aircraft control allocation method according to the application as described in any of the embodiments above, which is not described in detail herein.

It should be noted that, in this document, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or system that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or system. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or system that comprises the element.

The foregoing embodiment numbers of the present application are merely for the purpose of description, and do not represent the advantages or disadvantages of the embodiments.

From the above description of the embodiments, it will be clear to those skilled in the art that the above-described embodiment method may be implemented by means of software plus a necessary general hardware platform, but of course may also be implemented by means of hardware, but in many cases the former is a preferred embodiment. Based on such understanding, the technical solution of the present application may be embodied essentially or in a part contributing to the prior art in the form of a software product stored in a storage medium as described above (e.g. ROM/RAM, magnetic disk, optical disk) comprising instructions for causing an aircraft control distribution device (which may be a TWS headset or the like) to perform the method according to the various embodiments of the present application.

The foregoing description is only of the preferred embodiments of the present application, and is not intended to limit the scope of the application, but rather is intended to cover any equivalents of the structures or equivalent processes disclosed herein or in the alternative, which may be employed directly or indirectly in other related arts.

Claims

1. An aircraft control distribution method, characterized in that it is applied to an aircraft having rotors including at least tilting rotors, said rotors being driven by drive motors, said aircraft control distribution method comprising:

2. The aircraft control distribution method according to claim 1, wherein the step of performing the second power distribution on each rotor based on the first power distribution value comprises:

detecting whether the first power distribution value meets a preset condition;

3. The aircraft control allocation method according to claim 2, wherein after the step of detecting whether the first power allocation value meets a preset condition, the method further comprises:

A second power split is performed.

4. The method of claim 3, wherein the third power distribution is a weighted pseudo-inverse method, and wherein the step of performing the third power distribution on each rotor based at least in part on the first power distribution values to obtain third power distribution values comprises:

5. The aircraft control distribution method according to claim 4, wherein the preset condition is that each of the first rotational speeds does not reach a preset threshold value, and the step of detecting whether the first power distribution value meets the preset condition includes:

6. The aircraft control distribution method according to claim 1, wherein the step of performing the second power distribution on each rotor based on the first power distribution value to obtain the second power distribution value includes:

7. The aircraft control distribution method according to claim 6, wherein the step of calculating the second rotation speed of the driving motor of each rotor and the tilting angle of the tilting rotor based on the desired moment and the desired lift of the aircraft by a nonlinear optimization algorithm comprises:

8. The aircraft control distribution method according to claim 1, wherein said step of first power distribution to each rotor available on said aircraft, obtaining a first power distribution value, comprises:

9. An aircraft control distribution device for an aircraft having a rotor comprising at least a tilt rotor, the rotor being driven by a drive motor, the aircraft control distribution device comprising:

10. An aircraft having a rotor, the aircraft having a rotor comprising: a memory and a processor, the memory storing an aircraft control allocation program executable on the processor, the aircraft control allocation program when executed by the processor implementing the steps of the aircraft control allocation method of any one of claims 1 to 8.

11. A computer-readable storage medium, on which an aircraft control allocation program is stored, which, when executed by a processor, carries out the steps of the aircraft control allocation method according to any one of claims 1 to 8.

12. A computer program product, characterized in that it comprises an aircraft control allocation program which, when executed by a processor, implements the steps of the aircraft control allocation method according to any one of claims 1 to 8.