A tilting hexacopter power system failure handling method and device

Abstract

The application discloses a tilting six-rotor power system failure disposal method and device, and relates to the technical field of unmanned aerial vehicles. The core data is collected and analyzed by monitoring the operating components involved in the power system in real time, so as to determine the fault type of the power system; the power coefficient of the power system under the current fault type is determined, the control matrix after failure is established, and a new control distribution matrix is established based on the control matrix after failure; different control and response strategies are adopted for the unmanned aerial vehicle based on the fault type and the new control distribution matrix, as well as different power system failure quantities and power state conditions. The technical problem of how to quickly develop a scientific and reasonable response strategy when the power system fails, to maintain the stable flight attitude of the unmanned aerial vehicle to the greatest extent, and to ensure that the unmanned aerial vehicle can safely return or land when part of the power fails, is solved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of unmanned aerial vehicles, in particular to a tilt six-rotor power system failure disposal method and device. BACKGROUND

[0002] The tilt-rotor unmanned aerial vehicle combines the advantages of traditional helicopters and fixed-wing unmanned aerial vehicles, and has the ability of vertical take-off and hovering in the air, and can realize high-speed flight, and also has a large range, endurance and load capacity.

[0003] The tilt-rotor unmanned aerial vehicle faces the risk of power system failure in the flight process. As the core component of the tilt-rotor unmanned aerial vehicle, the normal operation of the power system is the key to ensuring the flight safety of the unmanned aerial vehicle. Once the power system fails, the unmanned aerial vehicle will lose enough lift and propulsion, which will easily lead to the loss of control of the aircraft and cause it to crash. This not only causes the loss of expensive unmanned aerial vehicle equipment, but also poses a serious threat to the safety of ground personnel and property, and causes huge life and property losses.

[0004] Therefore, how to quickly develop a scientific and reasonable response strategy when the power system fails, to maximize the maintenance of the stable flight attitude of the unmanned aerial vehicle, and to ensure that the unmanned aerial vehicle can safely return or land when part of the power fails, has become a technical problem to be solved. SUMMARY

[0005] In the embodiments of the present application, by providing a tilt six-rotor power system failure disposal method, the technical problem of how to quickly develop a scientific and reasonable response strategy when the power system fails, to maximize the maintenance of the stable flight attitude of the unmanned aerial vehicle, and to ensure that the unmanned aerial vehicle can safely return or land when part of the power fails, is solved.

[0006] In the first aspect, the embodiments of the present application provide a tilt six-rotor power system failure disposal method, which comprises: monitoring the operating components involved in the power system in real time, collecting core data and performing analysis and processing to determine the fault type of the power system; determining the power coefficient of the power system under the current fault type, establishing a control matrix after failure, and establishing a new control distribution matrix based on the control matrix after failure; based on the fault type and the new control distribution matrix, as well as different numbers of power system failures and power states, different control and response strategies are adopted for the unmanned aerial vehicle.

[0007] In a possible implementation, the fault types include power system power shortage, power system overload, power system failure, tilting nacelle jamming, tilting actuator feedback failure, total pitch control failure, rotor partial failure, and rotor failure; the power system power shortage is determined by low output voltage; the power system overload is determined by simultaneous occurrence of low output voltage, large power system current, and high output power; the power system failure is determined by simultaneous occurrence of low rotor speed, low output power, and high motor temperature; the tilting nacelle jamming is determined by inconsistency between the feedback value and the command value of the actuator tilt angle; the tilting actuator feedback failure is determined by tilting actuator state feedback failure; the total pitch control failure is determined by total pitch actuator state feedback failure; the rotor partial failure is determined by simultaneous occurrence of low output power and low rotor speed; and the rotor failure is determined by rotor speed state feedback failure.

[0008] In a possible implementation, the determining the power coefficient of the power system under the current fault type, establishing the post-failure control matrix, and establishing the new control distribution matrix based on the post-failure control matrix include: applying the power coefficient of the power system under the current fault type to a corresponding element of the original control matrix of the power system to obtain the post-failure control matrix; and establishing the new control distribution matrix based on the post-failure control matrix. obtaining the new control distribution matrix; wherein, is the original control matrix, is the control distribution matrix, is a virtual control quantity of the four control channels, is the post-failure control matrix, is the new control distribution matrix; and the unmanned aerial vehicle is controlled to fly based on the post-failure control matrix and the new control distribution matrix.

[0009] In a possible implementation, the original control matrix includes a rotor total pitch control matrix and an aerodynamic rudder and throttle control matrix, and the control distribution matrix includes a rotor total pitch control distribution matrix and an aerodynamic rudder and throttle control distribution matrix; the determining the power coefficient of the power system under the current fault type, establishing the post-failure control matrix, and establishing the new control distribution matrix based on the post-failure control matrix include: applying the power coefficient of the power system under the current fault type to a corresponding element of the rotor total pitch control matrix of the power system to obtain the post-failure rotor total pitch control matrix; and establishing the new rotor total pitch control distribution matrix based on the post-failure rotor total pitch control matrix. obtaining the post-failure rotor total pitch control matrix; wherein, is the rotor total pitch control matrix, is the rotor total pitch control distribution matrix, is the post-failure rotor total pitch control matrix, is the new rotor total pitch control distribution matrix, is the tilt angle of the rotor; controlling the unmanned aerial vehicle to fly based on the failed control matrix, the new control distribution matrix, and the aerodynamic rudder and throttle control matrix, and the aerodynamic rudder and throttle control distribution matrix.

[0010] In a possible implementation, different control and response strategies are taken for the unmanned aerial vehicle based on the failure type and the new control distribution matrix, and the number of different power system failures and power state conditions, including: in the case of single power system failure or partial failure, the unmanned aerial vehicle is ensured to fly normally by resetting the control distribution matrix; and / or in the case of simultaneous failure or simultaneous partial failure of two power systems, the unmanned aerial vehicle is completed to land by resetting the control distribution matrix and developing a control strategy; and / or in the case of insufficient power during flight of the unmanned aerial vehicle, a transition mode allowing flight altitude to be lowered is adopted.

[0011] In a possible implementation, the case of simultaneous failure or simultaneous partial failure of two power systems is divided into no additional power failure, power partial failure, power substantial failure, and attitude partial failure; the control strategy in the case of no additional power failure includes: in the case of failure of two diagonal rotors, the unmanned aerial vehicle is ensured to normally return and land by resetting the control distribution matrix to ensure complete attitude control; the control strategy in the case of power partial failure includes: in the case of simultaneous failure of two rotors at different longitudinal positions and different from the diagonal rotors, the unmanned aerial vehicle is ensured to normally return by resetting the control distribution matrix to ensure complete attitude control and preferentially selecting the fixed-wing mode; if in the transition mode, the flight speed is preferentially ensured, the flight altitude is allowed to be lowered during transition mode flight, and the unmanned aerial vehicle is slowly descended in the landing area after switching to the helicopter mode to complete landing; the control strategy in the case of power substantial failure includes: in the case of failure of rotors at the same longitudinal position, the unmanned aerial vehicle is ensured to normally return by resetting the control distribution matrix to ensure complete attitude control and preferentially selecting the fixed-wing mode; if in the helicopter mode, the flight speed is preferentially ensured, and the unmanned aerial vehicle is slowly descended in the landing area to complete landing; the control strategy in the case of attitude partial failure includes: in the case of failure of two rotors on the wing, the unmanned aerial vehicle is ensured to return in the fixed-wing mode by preferentially ensuring roll angle and pitch angle control in flight, yaw angle stability is provided by the vertical tail, and the remaining four power systems provide sufficient power, so that the unmanned aerial vehicle is returned and landed.

[0012] In a possible implementation, the transition mode allowing flight height to be reduced is adopted when the power is insufficient during the flight of the UAV, including: the pulling force is increased to the maximum when the attitude of the UAV is stable; the gravity potential energy is converted into kinetic energy to maintain the flight speed by adjusting the pitch angle during the transition from the fixed-wing mode to the helicopter mode; the lift restriction is removed, and a constraint condition is set; the constraint condition relates to the balance relationship among the pulling force of the rotor, the resistance and the gravity component; the gravity component is related to the pitch angle; the constraint condition includes that the pitch angle instruction value is related to the lifting speed; the relative height of the UAV when it starts to enter the transition mode is greater than a preset threshold, indicating that the airspeed is greater than the minimum speed boundary determined by the transition corridor.

[0013] In a second aspect, the embodiments of the present application provide a tilt six-rotor power system failure disposal device, which comprises: a monitoring module, configured to monitor the operating components involved in the power system in real time, collect core data and perform analysis and processing to determine the fault type of the power system; an establishment module, configured to determine the power coefficient of the power system under the current fault type, establish a control matrix after failure, and establish a new control distribution matrix based on the control matrix after failure; and a control module, configured to take different control and response strategies for the UAV based on the fault type and the new control distribution matrix, as well as different numbers of power system failures and power states.

[0014] In a third aspect, the embodiments of the present application provide a tilt six-rotor power system failure disposal server, comprising a memory and a processor; the memory is configured to store computer executable instructions; and the processor is configured to execute the computer executable instructions to implement the method in the first aspect or any possible implementation manner of the first aspect.

[0015] In a fourth aspect, the embodiments of the present application provide a computer readable storage medium, which stores executable instructions, and a computer executes the executable instructions to implement the method in the first aspect or any possible implementation manner of the first aspect.

[0016] The one or more technical solutions provided in the embodiments of the present application have at least the following technical effects:

[0017] The embodiment of the present application provides a tilt six-rotor power system failure disposal method, which realizes accurate and timely fault identification by monitoring the operating components related to the unmanned aerial vehicle power system in real time, and collecting core data for analysis and processing to determine the fault type. Different fault types will cause different power characteristics changes of the power system, the power coefficient of the power system under the current fault type is determined, the control matrix after failure is established, and the new control distribution matrix is established based on the control matrix after failure, so that the power output under the fault state can be better adapted, and powerful support is provided for the stable control of the unmanned aerial vehicle under the fault condition. Based on the fault type and the new control distribution matrix, and different power system failure quantities and power state conditions, different control and coping strategies are adopted for the unmanned aerial vehicle, so that the safety and reliability of the unmanned aerial vehicle are greatly improved. The technical problem of how to quickly develop a scientific and reasonable coping strategy when the power system fails, and how to maintain the stable flight attitude of the unmanned aerial vehicle to the greatest extent, and how to ensure the safe return or landing of the unmanned aerial vehicle under the partial power failure is solved. BRIEF DESCRIPTION OF DRAWINGS

[0018] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the drawings needed to be used in the description of the embodiments of the present application or the prior art will be briefly introduced. Obviously, the drawings in the following description are some embodiments of the present application, and other drawings can be obtained by those skilled in the art without creative labor.

[0019] Figure 1 A flowchart of a tilt six-rotor power system failure disposal method provided by the embodiment of the present application is provided.

[0020] Figure 2 A composition schematic diagram of a fault monitoring system provided by the embodiment of the present application is provided.

[0021] Figure 3 A schematic diagram of an unmanned aerial vehicle tilt six-rotor provided by the embodiment of the present application is provided.

[0022] Figure 4 A schematic diagram of a tilt six-rotor power system failure disposal device provided by the embodiment of the present application is provided.

[0023] Figure 5 A schematic diagram of a tilt six-rotor power system failure disposal server provided by the embodiment of the present application is provided. DETAILED DESCRIPTION

[0024] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0025] The following description of some technologies involved in the embodiments of this application is provided to aid understanding and should be considered merely exemplary. Therefore, those skilled in the art should recognize that various changes and modifications can be made to the embodiments described herein without departing from the scope and spirit of this application. Similarly, for clarity and brevity, some descriptions of well-known functions and structures are omitted in the following description.

[0026] This application provides a method for handling failures in a tilt-hexarotor propulsion system, such as... Figure 1 As shown, the method includes steps S101 to S103. Wherein, Figure 1 This is merely one execution order shown in the embodiments of this application and does not represent the only execution order for a tiltrotor propulsion system failure handling method. Where the final result can be achieved, Figure 1 The steps shown can be performed in parallel or in reverse order.

[0027] S101: Real-time monitoring of the control components involved in the power system, collection of core data and analysis processing to determine the fault type of the power system.

[0028] S102: Determine the dynamic coefficients of the power system under the current fault type, establish the control matrix after failure, and establish a new control allocation matrix based on the control matrix after failure.

[0029] Figure 2 This is a schematic diagram illustrating the composition of the fault monitoring system provided in this application embodiment. The fault monitoring system is used to monitor the control components involved in the power system in real time. Specifically, the monitored control components include the rotor, engine, or motor. The monitoring system consists of a Hall sensor, an ADC sampling module, a temperature sensor, a photoelectric sensor, and a servo self-test module. The system measures parameters such as rotor speed, phase sequence alternation frequency, output voltage, output current, and motor temperature, and analyzes and processes these data to comprehensively obtain the operating status of the UAV rotor. Based on the current operating status, it can determine the fault type of the power system and the power coefficient of the power system under the current fault type.

[0030] Table 1 Tiltrotor Failure Criteria

[0031]

[0032] Table 1 is a tilt-rotor fault criterion table. The fault types include power system power shortage, power system overload, power system failure, tilt nacelle jamming, tilt actuator feedback failure, total pitch control failure, rotor partial failure, and rotor failure.

[0033] Power system power shortage takes low output voltage as a criterion.

[0034] Specifically, when the output voltage is too low, it is determined that the power system power is insufficient. This fault usually occurs when the battery power is insufficient or the engine power is insufficient. If only the output voltage decreases without other accompanying phenomena, it can be considered that the power system can still use part of the power. The power coefficient of the power system power shortage is 0.7, that is, the usable power is 70% of the normal working power.

[0035] Power system overload takes low output voltage, large power system current, and high output power as a criterion.

[0036] Specifically, power system overload is generally caused by excessive rotor counter torque, accompanied by a small decrease in voltage, a large increase in current, and a large increase in power. Reducing the rotor total pitch can reduce the load, but also means a decrease in available power. The power coefficient of the power system overload is determined to be 0.5, that is, the available power is reduced to 50% of the previous power.

[0037] Power system failure takes low rotor speed, low output power, and high motor temperature as a criterion.

[0038] Specifically, power system failure is mostly caused by motor damage (such as short circuit, jamming, etc.). At this time, the power system cannot work normally, and the power supply of the motor needs to be cut off to stop the motor. The power coefficient is determined to be 0, that is, there is no available power.

[0039] Tilt nacelle jamming takes the inconsistency between the feedback value and the command value of the actuator tilt angle as a criterion. The power coefficient of the tilt nacelle jamming is calculated according to the tilt angle interpolation.

[0040] Specifically, tilt nacelle jamming is mainly caused by tilt actuator or structural failure, which is manifested as the inconsistency between the feedback value and the command value of the actuator tilt angle. To avoid trim problems and make the best use of rotor power, adjust the corresponding tilt angle of the other rotor to be the same. The power coefficient is calculated according to the tilt angle interpolation. Trim refers to the state of keeping the moment balance of the unmanned aerial vehicle during flight, that is, the resultant moment in each direction is zero, so as to ensure that the unmanned aerial vehicle can fly stably and will not have unnecessary attitude changes such as pitch, roll, or yaw.

[0041] Tilt actuator feedback failure takes the failure of the tilt actuator state feedback as a criterion.

[0042] Specifically, when the tilt actuator feedback fails, the tilt angle cannot be determined, and the additional moment caused by the rotor cannot be processed. To avoid control and trim problems, the rotor is stopped, and the power coefficient is 0, and no power can be used.

[0043] The total distance control failure is determined by the total distance actuator state feedback failure.

[0044] Specifically, the total distance control failure is usually caused by the total distance actuator being stuck, and the actuator cannot adjust the total distance, and the rotor tension cannot be controlled by the total distance. To avoid control and trim problems, the rotor is stopped, and the power coefficient is 0, and no power can be used.

[0045] The rotor partial failure is determined by the simultaneous occurrence of low output power and low rotor speed.

[0046] Specifically, the rotor partial failure is mainly caused by structural failure, and the power coefficient is determined to be 0.5, i.e. the available power is 50% of normal operation.

[0047] The rotor failure is determined by the rotor speed state feedback failure.

[0048] Specifically, the rotor failure is usually caused by structural failure, and is more serious than the rotor partial failure, such as rotor stuck. To avoid further damage, the motor is stopped, and the power coefficient is 0, and no power can be used.

[0049] S103: Based on the fault type and the new control allocation matrix, as well as the different number of power system failures and power states, different control and response strategies are taken for the unmanned aerial vehicle.

[0050] Determine the power coefficient of the power system under the current fault type, establish the control matrix after failure, and establish a new control allocation matrix based on the control matrix after failure, including the following contents.

[0051] Apply the power coefficient of the power system under the current fault type to the corresponding elements of the original control matrix of the power system to obtain the control matrix after failure.

[0052] Based on Obtain a new control allocation matrix. Wherein, is the original control matrix, is the control allocation matrix, is the virtual control amount of the four control channels, is the control matrix after failure, is the new control allocation matrix. , is the virtual control amount of the rotor tension channel, is the virtual control amount of the roll channel, is the virtual control amount of the pitch channel, The virtual control quantity of the yaw channel.

[0053] Controlling the unmanned aerial vehicle to fly based on the failed control matrix and the new control distribution matrix.

[0054] The original control matrix includes the rotor collective pitch control matrix and the aerodynamic rudder and throttle control matrix, and the control distribution matrix includes the rotor collective pitch control distribution matrix and the aerodynamic rudder and throttle control distribution matrix.

[0055] The power coefficient of the power system under the current failure type is applied to the corresponding element of the rotor collective pitch control matrix of the power system to obtain the failed rotor collective pitch control matrix.

[0056] Based on The failed rotor collective pitch control matrix is obtained. Wherein, The rotor collective pitch control matrix is The rotor collective pitch control distribution matrix is The failed rotor collective pitch control matrix is The new rotor collective pitch control distribution matrix is The rotor tilt angle is

[0057] Controlling the unmanned aerial vehicle to fly based on the failed control matrix, the new control distribution matrix, and the aerodynamic rudder and throttle control matrix and the aerodynamic rudder and throttle control distribution matrix.

[0058] Rudder control quantity Relationship with the virtual control quantity of the four control channels is: That is, the virtual control quantity of the four control channels is distributed to the real rudder control quantity.

[0059] When the power system fails, the influence on the state transition matrix (composed of stability derivatives) is small and can be ignored, but the influence on the control matrix is large. In order to reduce the influence of power system failure on the control law, a new control distribution matrix is established based on the failed control matrix to ensure the normal operation of the unmanned aerial vehicle under partial power failure.

[0060] The unmanned aerial vehicle dynamics equation obtained according to the small perturbation linearization model is: . Wherein, The flight state of the unmanned aerial vehicle is The state transition matrix is The control matrix is The rudder control quantity. Since the tilt six-rotor unmanned aerial vehicle has a redundant control characteristic, different control distribution methods need to be designed for the helicopter mode, the transition mode and the fixed-wing mode.

[0061] Generally, the new control distribution matrix can be calculated by the generalized inverse method , and the formula is However, due to the redundant characteristics of the six-rotor control, the control matrix after failure is not full rank, so the left inverse of the generalized inverse of does not exist. The original control matrix and the control distribution matrix can be divided into two parts: , . Among them, is the rotor collective control matrix, is the rotor collective control distribution matrix, which is valid in helicopter mode, is the aerodynamic rudder and throttle control matrix, is the aerodynamic rudder and throttle control distribution matrix, which is valid in fixed-wing mode. In the transition mode and are valid, and the new control distribution matrix can be recalculated according to the formula . Among them, is the rotor collective control matrix after failure. Since the control distribution of the fixed-wing mode is independent of whether the power system fails or not, there is no need to reset the control distribution matrix of this part when studying the failure of the power system. According to different failure conditions, the new control distribution matrix is calculated by the above formula.

[0062] Figure 3 is a schematic diagram of a tilt six-rotor unmanned aerial vehicle provided by an embodiment of the present application. The unmanned aerial vehicle rotors are numbered 1-6, and the corresponding power system of the rotor is referred to as the number in the subsequent. Among them, rotor 1, rotor 4 and rotor 6 are left-handed rotors, and rotor 2, rotor 3 and rotor 5 are right-handed rotors.

[0063] When a single power system fails or a single power system partial failure occurs, the unmanned aerial vehicle has the ability to fly normally and return, but the flight performance will be affected to a certain extent, and the control structure will also change. At this time, a new control distribution matrix needs to be established based on the control matrix after failure to meet the control requirements.

[0064] Taking rotor 1 failure as an example, the calculation method of the new control distribution matrix is explained. In the normal helicopter mode, the rotor collective control matrix and the rotor collective control distribution matrix have the following expressions.

[0065] . Among them, is the partial derivative of the rotor tension to the rotor collective, which means the change of the rotor tension caused by the manipulation of the rotor collective, is the rotor tension, is the total pitch of the rotor, is the yawing moment, is the partial derivative of the rotor counter torque with respect to the total pitch of the rotor, is the rotor arm.

[0066] .

[0067] When the rotor 1 fails, the elements related to the rotor 1 in the new control allocation matrix are multiplied by the power coefficient. Assuming that the available power percentage of the rotor 1 is 0, i.e. the power coefficient is 0, the total pitch control matrix of the failed rotor is obtained . .

[0068] The new total pitch control allocation matrix is calculated as follows. . The new total pitch control allocation matrix shows that the power coefficient of the rotor 1 is 0, i.e. the rotor 1 is in the state of no power output, and the matrix can satisfy the constraint . For the control allocation in the transition mode, the specific form is more complex, but it can still be calculated according to the given formula. This application only takes the elements in the new total pitch control allocation matrix in the transition mode determined by the cosine and sine of the tilt angle as an example to give the new total pitch control allocation matrix in the transition mode when the rotor 1 fails: . Wherein, is the rotor tilt angle, when is 90°, it is the helicopter mode, and when 0°, it is the fixed-wing mode.

[0069] Since the control allocation in the fixed-wing mode is independent of whether the power system fails or not, in the case of power system failure, there is no need to reset the control allocation in the fixed-wing mode, i.e. there is no need to set a new control allocation matrix.

[0070] Based on the fault type and the new control allocation matrix, as well as the different number of power system failures and power states, different control and coping strategies are taken for the unmanned aerial vehicle, including the following contents.

[0071] In the case of single power system failure or partial failure, the control allocation matrix is reset to ensure normal flight and return of the unmanned aerial vehicle.

[0072] And / or when two power systems fail or partially fail at the same time, the control allocation matrix is reset and the control strategy is developed to complete the landing.

[0073] Specifically, when three or more power systems fail, the unmanned aerial vehicle cannot complete the flight, and this case is not discussed.

[0074] and / or if the UAV is in power shortage during flight, a transition mode allowing flight height to be reduced is adopted.

[0075] Specifically, the transition mode is that the UAV is switched from the fixed-wing mode to the helicopter mode.

[0076] The two power systems simultaneously fail or partially fail, which is divided into no additional power failure, power partial failure, power large failure and attitude partial failure.

[0077] The control strategy in the case of no additional power failure includes: in the case of two sets of diagonal rotors failing, the control allocation matrix is reset to ensure complete attitude control, so that the UAV returns normally and lands.

[0078] The two sets of diagonal rotors can be rotor 1 and rotor 3, or rotor 2 and rotor 4.

[0079] The control strategy in the case of power partial failure includes: in the case of non-diagonal rotors and two rotors at different longitudinal positions failing simultaneously, the control allocation matrix is reset to ensure complete attitude control, and the fixed-wing mode is preferentially selected to make the UAV return normally. If in the transition mode, the flight speed is preferentially ensured when the height allows, the flight height is allowed to be reduced during the transition mode flight, and after switching to the helicopter mode, the UAV is controlled to slowly glide in the emergency landing area to complete the emergency landing.

[0080] The two non-diagonal rotors and the two rotors at different longitudinal positions can be rotor 1 and rotor 5, or rotor 1 and rotor 6. Of course, they can also be other rotors, which are not limited in the application.

[0081] The control strategy in the case of power large failure includes: in the case of rotors at the same longitudinal position failing, the control allocation matrix is reset to ensure complete attitude control, and the fixed-wing mode is preferentially selected to make the UAV return normally. If in the helicopter mode, the flight speed is preferentially ensured when the height allows, and the UAV is controlled to slowly glide in the emergency landing area to complete the emergency landing.

[0082] The rotors at the same longitudinal position can be rotor 1 and rotor 2, or rotor 3 and rotor 4.

[0083] The control strategy in the case of attitude partial failure includes: in the case of two sets of rotors on the wings failing, the roll angle and pitch angle control are preferentially ensured during flight, the yaw angle is stabilized by the vertical tail, and the remaining four power systems provide sufficient power to make the UAV return in the fixed-wing mode, complete the transition and land.

[0084] The two sets of rotors on the wings can be rotor 5 and rotor 6.

[0085] Specifically, the attitude part is failed, in the helicopter mode and the transition mode, at this time, the attitude control of the unmanned aerial vehicle three-axis (roll axis, pitch axis, yaw axis) cannot be guaranteed, from the mathematical point of view, the generalized inverse matrix does not exist. Since in this special case, the new rotor total pitch control distribution matrix cannot be directly calculated according to the conventional method, the specific form of the elements in the new rotor total pitch control distribution matrix is directly given as follows: .

[0086] If the unmanned aerial vehicle is powered during flight, the transition mode is used to allow the flight altitude to drop, including the following contents.

[0087] When the attitude of the unmanned aerial vehicle is stable, the tension is raised to the maximum.

[0088] During the transition from the fixed-wing mode to the helicopter mode, the pitch angle is adjusted to convert the gravitational potential energy into kinetic energy to maintain the flight speed.

[0089] Specifically, during the transition from the fixed-wing mode to the helicopter mode, the power is gradually reduced, and the conversion of gravitational potential energy can make up for the lack of power to ensure the stability of the flight speed.

[0090] The lift restriction is removed, and the constraint condition is set. The constraint condition involves the balance relationship of the rotor tension, the resistance and the gravitational component. Among them, the gravitational component is related to the pitch angle.

[0091] The constraint condition includes: the pitch angle command value is related to the climb rate. When the unmanned aerial vehicle starts to enter the transition mode, the relative height is greater than the preset threshold, and the indicated airspeed is greater than the minimum speed boundary determined by the transition corridor.

[0092] Specifically, the constraint condition involves the balance relationship of the rotor tension, the resistance and the gravitational component, and the expression of the force balance constraint is: . Wherein, is the rotor tension, is the rotor tilt angle, is the resistance of the unmanned aerial vehicle under the body axis system, is the mass of the unmanned aerial vehicle, is the acceleration of gravity, is the pitch angle.

[0093] Specifically, the expression of the pitch angle command value is: . Wherein, is the pitch angle command value, is the error of the climb rate, is the proportional coefficient, is the integral coefficient, indicates the time The differential, This is for integration operations.

[0094] The following conditions must be met for a drone to enter transition mode: 1. Relative altitude. Greater than the preset threshold. The preset threshold can be 300, i.e. 2. Indicating airspeed Greater than the minimum velocity boundary determined by the transition corridor ,Right now A transition corridor is a conceptual area that describes the acceptable speed range for a drone during the transition from one flight mode (such as fixed-wing mode) to another (such as helicopter mode).

[0095] This application embodiment also provides a tilting six-rotor propulsion system failure handling device 400, such as Figure 4 As shown, the device includes: a monitoring module 401, an establishment module 402, and a control module 403.

[0096] The monitoring module 401 is used to monitor the control components involved in the power system in real time, collect core data and perform analysis and processing to determine the fault type of the power system.

[0097] The module 402 is used to determine the dynamic coefficients of the power system under the current fault type, establish the control matrix after failure, and establish a new control allocation matrix based on the control matrix after failure.

[0098] The control module 403 is used to adopt different control and response strategies for the UAV based on the fault type, the new control allocation matrix, and the different number of power system failures and power status conditions.

[0099] Some modules in the apparatus described in this application can be described in the general context of computer-executable instructions that are executed by a computer, such as program modules. Generally, program modules include routines, programs, objects, components, data structures, classes, etc., that perform a specific task or implement a specific abstract data type. This application can also be practiced in distributed computing environments where tasks are performed by remote processing devices connected via a communication network. In distributed computing environments, program modules can reside in local and remote computer storage media, including storage devices.

[0100] The apparatuses or modules illustrated in the above application examples can be implemented by computer chips or entities, or by products with certain functions. For the convenience of description, the above apparatuses are described as various modules with functions. In the implementation of the application examples, the functions of the modules can be implemented in one or more software and / or hardware. Of course, the modules with certain functions can also be implemented by a combination of multiple sub-modules or sub-units.

[0101] The methods, apparatuses or modules described in the present application can be implemented in a computer readable program code, and the controller can be implemented in any appropriate manner, for example, the controller can take the form of, for example, a microprocessor or a processor, and a computer readable medium storing computer readable program code (for example, software or firmware) executable by the (micro) processor, logic gates, switches, application specific integrated circuits (Application Specific Integrated Circuit, ASIC), programmable logic controllers and embedded microcontrollers. Examples of the controller include, but are not limited to, the following microcontrollers: ARC 625D, Atmel AT91SAM, Microchip PIC18F26K20 and Silicone Labs C8051F320. The memory controller can also be implemented as part of the control logic of the memory. Those skilled in the art also know that, in addition to implementing the controller in a pure computer readable program code manner, the same function can also be implemented by logically programming the method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers and embedded microcontrollers. Therefore, such a controller can be considered as a hardware component, and the devices included therein for implementing various functions can also be considered as structures within the hardware component. Alternatively, the devices for implementing various functions can also be considered as both software modules implementing the method and structures within the hardware component.

[0102] As shown in Figure 5 The application examples also provide a tilt six-rotor power system failure handling server, including a memory 501 and a processor 502; the memory 501 is used to store computer executable instructions; the processor 502 is used to execute the computer executable instructions to implement the tilt six-rotor power system failure handling method provided in the above application examples.

[0103] The application examples also provide a computer readable storage medium, which stores executable instructions, and a computer executing the executable instructions can implement the tilt six-rotor power system failure handling method provided in the above application examples.

[0104] From the description of the above embodiments, those skilled in the art can clearly understand that the present application can be implemented by means of software plus necessary hardware. Based on such an understanding, the technical solutions of the present application can be embodied in the form of a software product or can be embodied in the implementation process of data migration. The computer software product can be stored in a storage medium, such as a ROM / RAM, a magnetic disk, an optical disk, and the like, and includes a plurality of instructions for causing a computer device (which can be a personal computer, a mobile terminal, a server, or a network device, etc.) to execute the methods described in the embodiments of the present application.

[0105] The various embodiments in the specification are described in a progressive manner, and the same or similar parts between the various embodiments can be referred to each other, and each embodiment mainly describes the difference from other embodiments. The whole or part of the present application can be used in a plurality of general or special computer system environments or configurations.

[0106] The above embodiments are only used to illustrate the technical solutions of the present application, and not to limit the present application; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that the technical solutions recorded in the foregoing embodiments can still be modified, or some or all of the technical features can be replaced by equivalents; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the present application.

Claims

1. A tilt six-rotor power system failure handling method, characterized in that, The application relates to a control method for unmanned aerial vehicles (UAVs) in a hybrid mode, and belongs to the field of UAV control. Real-time monitoring is performed on a control component involved in a power system, core data is collected and analyzed and processed, and a fault type of the power system is judged; A power coefficient of the power system under the current fault type is determined, a control matrix after failure is established, and a new control distribution matrix is established based on the control matrix after failure; Different control and response strategies are adopted for the UAV based on the fault type and the new control distribution matrix, and different failure numbers and power states of the power system; when a power part or a large part fails, a fixed-wing mode is preferentially selected; if the UAV is in a transition mode or a helicopter mode, a control strategy of preferentially ensuring flight speed in the case of allowing height is adopted.

2. The tilting hexacopter power system failure handling method of claim 1, wherein, The fault types include power system power supply shortage, power system excessive load, power system failure, tilting nacelle jamming, tilting actuator feedback failure, total distance control failure, rotor part failure and rotor failure; The power system power supply shortage is judged by low output voltage; The power system excessive load is judged by simultaneous occurrence of low output voltage, large power system current and high output power; The power system failure is judged by simultaneous occurrence of low rotor speed, low output power and high motor temperature; The tilting nacelle jamming is judged by inconsistency between a tilting actuator tilt angle feedback value and a command value; The tilting actuator feedback failure is judged by tilting actuator state feedback failure; The total distance control failure is judged by total distance actuator state feedback failure; The rotor part failure is judged by simultaneous occurrence of low output power and low rotor speed; The rotor failure is judged by rotor speed state feedback failure.

3. The tilting six-rotor power system failure handling method of claim 1, wherein, The determination of the power coefficient of the power system under the current fault type, the establishment of the control matrix after failure and the establishment of the new control distribution matrix based on the control matrix after failure include the following steps: The power coefficient of the power system under the current fault type is applied to corresponding elements of an original control matrix of the power system to obtain a control matrix after failure; based on obtaining a new control distribution matrix; wherein is the original control matrix, is the control distribution matrix, is the virtual control quantity of the four control channels, is the control matrix after failure, is the new control distribution matrix; The UAV is controlled to fly based on the control matrix after failure and the new control distribution matrix.

4. The tilting hexacopter power system failure handling method of claim 1 or 3, wherein, The original control matrix includes a rotor total distance control matrix and an aerodynamic rudder and throttle control matrix, and the control distribution matrix includes a rotor total distance control distribution matrix and an aerodynamic rudder and throttle control distribution matrix; The determination of the power coefficient of the power system under the current fault type, the establishment of the control matrix after failure and the establishment of the new control distribution matrix based on the control matrix after failure include the following steps: The power coefficient of the power system under the current fault type is applied to corresponding elements of a rotor total distance control matrix of the power system to obtain a rotor total distance control matrix after failure; based on obtaining a failed rotor collective control matrix; wherein, is a rotor collective control matrix, is a rotor collective control allocation matrix, is a failed rotor collective control matrix, is a new rotor collective control allocation matrix, is a rotor tilt angle; The UAV is controlled to fly based on the control matrix after failure, the new control distribution matrix, the aerodynamic rudder and throttle control matrix and the aerodynamic rudder and throttle control distribution matrix.

5. The tilting six-rotor power system failure handling method of claim 1, wherein, The adoption of different control and response strategies for the UAV based on the fault type and the new control distribution matrix, and different failure numbers and power states of the power system includes the following steps: In the case of single power system failure or partial failure, the control distribution matrix is reset to ensure normal flight of the UAV. and / or when two power systems fail simultaneously or partially fail simultaneously, completing the forced landing by resetting the control allocation matrix and formulating the maneuvering strategy; and / or if power deficiency occurs during the flight of the unmanned aerial vehicle, a transition mode allowing the flight altitude to be lowered is adopted.

6. The tilting hexacopter power system failure handling method of claim 5, wherein, The case of two power systems failing simultaneously or partially failing simultaneously is divided into no additional power failure, power partial failure, power substantial failure and attitude partial failure; The maneuvering strategy under the condition of no additional power failure includes: in the case of two groups of diagonal rotors failing, the complete attitude control is ensured by resetting the control allocation matrix, so that the unmanned aerial vehicle returns normally and lands; The maneuvering strategy under the condition of power partial failure includes: in the case of two rotors failing at different longitudinal positions, the complete attitude control is ensured by resetting the control allocation matrix, the fixed-wing mode is preferentially selected, and the unmanned aerial vehicle returns normally; if in the transition mode, the flight speed is preferentially ensured, the flight altitude is allowed to be lowered during the transition mode flight, and after switching to the helicopter mode, the unmanned aerial vehicle is controlled to slowly glide in the forced landing area to complete the forced landing; The maneuvering strategy under the condition of power substantial failure includes: in the case of rotors failing at the same longitudinal position, the complete attitude control is ensured by resetting the control allocation matrix, the fixed-wing mode is preferentially selected, and the unmanned aerial vehicle returns normally; if in the helicopter mode, the flight speed is preferentially ensured, and the unmanned aerial vehicle is controlled to slowly glide in the forced landing area to complete the forced landing; The maneuvering strategy under the condition of attitude partial failure includes: in the case of two groups of rotors failing on the wings, the roll angle and pitch angle control are preferentially ensured in flight, the yaw angle is stabilized by the vertical tail, and the remaining four power systems provide sufficient power to enable the unmanned aerial vehicle to return in the fixed-wing mode, complete the transition and land.

7. The tilting hexacopter power system failure handling method of claim 6, wherein, The transition mode allowing the flight altitude to be lowered includes: The pulling force is increased to the maximum when the attitude of the unmanned aerial vehicle is stable; During the transition from the fixed-wing mode to the helicopter mode, the pitch angle is adjusted to convert the gravitational potential energy into kinetic energy to maintain the flight speed; The lift restriction is removed, and a constraint condition is set; the constraint condition relates to the balance relationship among the rotor pulling force, the resistance and the gravitational component; the gravitational component is related to the pitch angle; The constraint condition includes: the pitch angle command value is related to the climb rate; when the unmanned aerial vehicle starts to enter the transition mode, the relative altitude is greater than a preset threshold, and the indicated airspeed is greater than the minimum speed boundary determined by the transition corridor.

8. A tilting hexacopter power system failure handling device for implementing the tilting hexacopter power system failure handling method of any one of claims 1-7, characterized by, It includes: A monitoring module for real-time monitoring of the maneuvering components involved in the power system, collecting core data and performing analysis and processing to determine the fault type of the power system; An establishment module for determining the power coefficient of the power system under the current fault type, establishing the control matrix after failure, and establishing a new control allocation matrix based on the control matrix after failure; The control module is used for taking different control and response strategies for the unmanned aerial vehicle based on the fault type and the new control distribution matrix, and the number of different power system failures and power state conditions; wherein when the power part or a large part is failed, a fixed wing mode is preferentially selected, and if in a transition mode or a helicopter mode, a control strategy of preferentially ensuring flight speed is adopted when the height is allowed.

9. A tilt six-rotor power system failure handling server, characterized in that, comprise a memory and a processor; the memory is configured to store computer executable instructions; the processor is configured to execute the computer executable instructions to implement the method of any one of claims 1-7.

10. A computer-readable storage medium, characterized in that, The computer readable storage medium stores executable instructions, and the computer executes the executable instructions to implement the method of any one of claims 1-7.

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