A method and system for attitude control of a foldable wing unmanned aerial vehicle

Abstract

The application provides a posture control method and system of a folding wing unmanned aerial vehicle, and relates to the technical field of aircraft control. Firstly, the application collects pressure information and strain information of the folding wing unmanned aerial vehicle, and determines whether the wing has aerodynamic-elastic coupling during the action according to the two kinds of information. Then, when it is determined that the aerodynamic-elastic coupling occurs, the airflow control instruction corresponding to the current coupling state is determined. Then, according to the airflow control instruction, the asymmetric additional airflow on the wing surface is generated by controlling the pulse jet execution mechanism. Finally, the compensation moment is generated by the additional airflow to offset the body posture disturbance caused by the target action of the folding wing unmanned aerial vehicle. The technical scheme provided by the application not only realizes the real-time perception, accurate judgment and active inhibition of the nonlinear aerodynamic-elastic coupling disturbance of the folding wing unmanned aerial vehicle during the shape conversion process, but also improves the stability of the flight posture of the folding wing unmanned aerial vehicle.

Description

An attitude control method and system for a folding-wing unmanned aerial vehicle Technical Field

[0001] This application relates to the field of aircraft control technology, and in particular to an attitude control method and system for a folding-wing unmanned aerial vehicle. Background Technology

[0002] With the widespread application of drones in logistics delivery, agricultural and forestry plant protection, line inspection and film and television aerial photography, higher requirements have been placed on the multi-mission adaptability and deployment convenience of flight platforms. Foldable-wing drones have become a research and development focus because they can change their wing shape during flight to adapt to the needs of different flight segments. During the dynamic transformation of folding and unfolding wings, the aerodynamic shape and structural stiffness of such drones will change drastically, which can easily induce strong aeroelastic coupling, resulting in sudden and unpredictable attitude disturbances of the airframe, seriously threatening its operational safety and stability in complex commercial environments.

[0003] One current approach is to adopt an adaptive control strategy based on model predictive control and online parameter identification. This method establishes an aerodynamic model that includes unfolded state variables and uses sensor data during flight to update the model parameters in real time. This allows for the prediction of future attitude changes and the rolling optimization of control surface commands, with the aim of providing feedforward compensation for attitude during wing movements.

[0004] However, this scheme has significant shortcomings. Its core relies on the accuracy of the global aerodynamic model. The disturbances caused by the coupling of local flow separation, vortex shedding and structural vibration during the unfolding process have strong nonlinear, transient and localized characteristics, which are difficult to be accurately described by the overall model. This leads to a delay in the controller's perception of such sudden and locally coupled disturbances. The generated compensation commands do not match the actual disturbances in terms of timing and spatial distribution, resulting in limited suppression effect. Summary of the Invention

[0005] This application provides an attitude control method and system for a folding-wing unmanned aerial vehicle (UAV) to solve the problems of delayed perception and inaccurate control of transient and localized aeroelastic coupling disturbances during the folding process in the prior art.

[0006] In a first aspect, this application provides an attitude control method for a folding-wing unmanned aerial vehicle, including:

[0007] The system collects pressure information at multiple locations on the wing and strain information of the wing connection structure when the folding-wing UAV performs target maneuvers.

[0008] Based on the pressure information and the strain information of the wing connection structure, determine whether aeroelastic coupling occurs during the wing's operation;

[0009] When aeroelastic coupling is detected, the airflow control command corresponding to the current coupling state is determined.

[0010] According to the airflow control command, the pulse jet actuator set on the leading edge of the wing is controlled to generate an asymmetric additional airflow on the wing surface;

[0011] The additional airflow generates a compensating torque to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers.

[0012] Optionally, pressure information at multiple locations on the wing and strain information of the wing connection structure are collected when the folding-wing UAV performs target maneuvers, including:

[0013] A first pressure sensing unit is arranged in the folding hinge area of ​​the wing of the folding-wing UAV, a second pressure sensing unit is arranged in the wingtip area of ​​the folding-wing UAV, and a third pressure sensing unit is arranged in the wing root area of ​​the folding-wing UAV.

[0014] A strain sensing unit is installed at the hinge structure connecting the wing and the fuselage.

[0015] When the folding-wing UAV begins to perform the target action, all pressure sensing units and the strain sensing unit are activated simultaneously.

[0016] Pressure data from the first pressure sensing unit, the second pressure sensing unit, and the third pressure sensing unit are collected at fixed time intervals.

[0017] Simultaneously, strain data from the strain sensing unit are acquired at the same time intervals;

[0018] All collected pressure data are integrated into pressure information, and all collected strain data are integrated into strain information.

[0019] Optionally, based on the pressure information and the strain information of the wing connection structure, determining whether aeroelastic coupling occurs during the wing's operation includes:

[0020] From the pressure information, extract the first pressure data sequence from the folding hinge region, the second pressure data sequence from the wingtip region, and the third pressure data sequence from the wing root region;

[0021] From the strain information, extract the first strain data sequence from the folding hinge region, the second strain data sequence from the wingtip region, and the third strain data sequence from the wing root region;

[0022] Identify the points in the first, second, and third pressure data sequences where pressure values ​​change abruptly, and mark the moment when the first pressure value change occurs as the initial disturbance moment for each region.

[0023] In the first strain data sequence, the second strain data sequence, and the third strain data sequence, the time point when the strain value changes abruptly is identified, and the time when the strain value first changes abruptly is marked as the initial disturbance time of the structure.

[0024] Compare the initial disturbance times of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence with the order of the initial disturbance times of the structure;

[0025] If the initial disturbance time of at least one pressure data sequence is earlier than the initial disturbance time of the structure, it is determined that the wing did not undergo aeroelastic coupling during operation.

[0026] If the initial disturbance time of all pressure data sequences is later than the initial disturbance time of the structure, it is determined that the wing has undergone aeroelastic coupling during operation.

[0027] Optionally, when it is determined that aeroelastic coupling has occurred, an airflow control command corresponding to the current coupling state is determined, including:

[0028] After determining that the aeroelastic coupling has occurred, the pressure values ​​of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence at the corresponding initial disturbance time are obtained.

[0029] Calculate the difference between the pressure values ​​corresponding to the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence to obtain multiple pressure differences;

[0030] Obtain the strain value at the moment of initial disturbance of the structure;

[0031] The basic pulse frequency is determined based on the magnitude of the strain value;

[0032] Calculate the frequency adjustment factor based on the largest pressure difference;

[0033] Multiplying the base pulse frequency by the frequency adjustment coefficient yields the target pulse frequency;

[0034] Based on the relative pressure values ​​of the first, second, and third pressure data sequences, the starting position for applying additional airflow is selected from the folding hinge region, wingtip region, and wing root region.

[0035] Based on the target pulse frequency and the starting position, the airflow control command is generated, wherein the airflow control command includes a pulse frequency parameter and an application position parameter.

[0036] Optionally, a frequency adjustment factor is calculated based on the largest pressure difference, including:

[0037] The largest pressure difference is compared with the various pressure difference ranges in the pre-stored lookup table;

[0038] Determine the target pressure difference range to which the largest pressure difference value belongs;

[0039] Find the adjustment coefficient value corresponding to the target pressure difference range from the lookup table;

[0040] When the largest pressure difference is located at the boundary between two adjacent pressure difference intervals, the frequency adjustment coefficient is calculated using a linear interpolation method.

[0041] Optionally, according to the airflow control command, the pulse jet actuator disposed on the leading edge of the wing is controlled to generate an asymmetric additional airflow on the wing surface, including:

[0042] The airflow control command is analyzed to obtain the target pulse frequency and starting position;

[0043] Based on the starting position, activate the pulse jet unit located in the corresponding area of ​​the wing leading edge;

[0044] Based on the target pulse frequency, a pulse drive signal with a corresponding time interval is generated;

[0045] The pulse drive signal is sent to the activated pulse jet unit to control the activated pulse jet unit to operate according to the pulse drive signal;

[0046] When the pulse jet unit is working, a high-speed airflow is ejected from the gap at the leading edge of the wing through the pulse jet unit to act on the wing surface;

[0047] By controlling the pulse jet units in different regions to operate according to different pulse drive signals, high-speed airflows with different intensities and timings are generated at different positions on the wing surface to form an asymmetric additional airflow on the wing surface.

[0048] Optionally, a compensating torque is generated by the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers, including:

[0049] Real-time monitoring of the pitch and roll angle changes of the fuselage of the folding-wing UAV;

[0050] When the pitch angle change exceeds a first predetermined value or the roll angle change exceeds a second predetermined value, it is determined that there is an attitude disturbance.

[0051] Based on the direction of the attitude disturbance, determine the target wing region where a compensating torque needs to be generated;

[0052] The operating parameters of the pulse jet actuator in the target wing region are adjusted to control the local intensity of the additional airflow;

[0053] The adjusted additional airflow creates an asymmetric aerodynamic pressure distribution on the wing surface to generate a compensating torque.

[0054] The compensation torque is used to counteract the fuselage attitude disturbance, thereby offsetting the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers.

[0055] Secondly, this application provides an attitude control system for a folding-wing unmanned aerial vehicle, comprising:

[0056] The data acquisition module is used to collect pressure information at multiple positions of the wing and strain information of the wing connection structure when the folding-wing UAV performs target maneuvers.

[0057] The judgment module is used to determine whether aeroelastic coupling occurs during the operation of the wing based on the pressure information and the strain information of the wing connection structure.

[0058] The determination module is used to determine the airflow control command corresponding to the current coupling state when it is determined that aeroelastic coupling has occurred;

[0059] The control module is used to control the pulse jet actuator located on the leading edge of the wing to generate an asymmetric additional airflow on the wing surface according to the airflow control command.

[0060] An additional module is used to generate a compensating torque through the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers.

[0061] Thirdly, this application provides a computing device, including a processing component and a storage component; the storage component stores one or more computer instructions; the one or more computer instructions are invoked and executed by the processing component to implement the attitude control method for a folding-wing unmanned aerial vehicle as described in the first aspect above.

[0062] Fourthly, this application provides a computer storage medium storing a computer program, which, when executed by a computer, implements an attitude control method for a folding-wing unmanned aerial vehicle as described in the first aspect.

[0063] This application, by synchronously collecting pressure information at key locations on the wing and strain information of the connecting structure, can directly identify specific disturbance modes caused by the coupling of airflow excitation and structural deformation from a spatiotemporal correlation perspective. This judgment method based on multi-source information fusion effectively overcomes the problem of delayed perception of local and transient aeroelastic coupling phenomena caused by the reliance on the overall model in existing technologies, and provides a real-time and reliable decision-making basis for subsequent precise control.

[0064] Furthermore, by generating airflow control commands that match the current coupling state, the pulse jet actuator is driven to actively generate asymmetric additional airflow in a specific area on the wing surface. This can directly generate local aerodynamic torques that counteract the disturbance source, thereby achieving rapid, targeted, and appropriate active suppression of sudden attitude disturbances. This solves the shortcomings of traditional feedforward compensation methods, which suffer from mismatch in timing and spatial distribution of control commands and poor suppression effect due to model mismatch.

[0065] These or other aspects of this application will become more apparent in the following description of the embodiments. Attached Figure Description

[0066] To more clearly illustrate the technical solutions in this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0067] Figure 1 shows a flowchart of an attitude control method for a folding-wing unmanned aerial vehicle provided in this application;

[0068] Figure 2 shows a schematic diagram of the attitude control system of a folding-wing unmanned aerial vehicle provided in this application;

[0069] Figure 3 shows a schematic diagram of the structure of a computing device provided in this application. Detailed Implementation

[0070] To enable those skilled in the art to better understand the present application, the technical solution of the present application will be clearly and completely described below with reference to the accompanying drawings.

[0071] In some of the processes described in the specification, claims, and accompanying drawings of this application, multiple operations appearing in a specific order are included. However, it should be clearly understood that these operations may not be executed in the order they appear herein, or may be executed in parallel. The operation numbers, such as 101, 102, etc., are merely used to distinguish different operations and do not themselves represent any execution order. Furthermore, these processes may include more or fewer operations, and these operations may be executed sequentially or in parallel. It should be noted that the descriptions such as "first," "second," etc., in this document are used to distinguish different messages, devices, modules, etc., and do not represent a chronological order, nor do they limit "first" and "second" to different types.

[0072] The technical solutions of this application will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0073] Figure 1 is a flowchart of an attitude control method for a folding-wing unmanned aerial vehicle provided in this application. As shown in Figure 1, the method includes:

[0074] Step 101: Collect pressure information at multiple positions of the wing and strain information of the wing connection structure when the folding-wing UAV performs target maneuvers.

[0075] Optionally, step 101 may specifically include:

[0076] Step 1011: Arrange a first pressure sensing unit in the folding hinge area of ​​the folding wing of the folding wing UAV, arrange a second pressure sensing unit in the wingtip area of ​​the folding wing UAV, and arrange a third pressure sensing unit in the wing root area of ​​the folding wing UAV.

[0077] Step 1012: Install a strain sensing unit at the hinge structure connecting the wing and the fuselage.

[0078] Step 1013: When the folding-wing UAV begins to perform the target action, all pressure sensing units and the strain sensing unit are activated simultaneously.

[0079] Step 1014: Collect pressure data from the first pressure sensing unit, the second pressure sensing unit, and the third pressure sensing unit at fixed time intervals.

[0080] Step 1015: Simultaneously, strain data from the strain sensing unit is acquired at the same time interval.

[0081] Step 1016: Integrate all collected pressure data into pressure information, and integrate all collected strain data into strain information.

[0082] In this step, pressure information refers to a set of data representing the magnitude of the air force acting on the surface of an object, used to reflect the aerodynamic conditions experienced by different parts of the drone's wing during flight.

[0083] Strain information refers to a set of data that represents the degree of deformation of an object's structure when subjected to force. It is used to reflect the minute stretching or compression changes of the wing connection structure of a UAV during flight caused by aerodynamics and its own motion.

[0084] The folding hinge area refers to the mechanical rotating joints on the UAV wing designed to achieve the folding function and the local area around them. It is the core part of the wing shape change and the key area where aerodynamic characteristics change drastically. It can be identified and determined by observing the mechanical structure design of the UAV wing.

[0085] The wingtip region refers to the outermost tip of the wing of a UAV that is furthest from the fuselage. The airflow behavior in this region is complex and has a significant impact on the stability and efficiency of the aircraft.

[0086] The wing root region refers to the root part of the UAV wing that connects to the main fuselage, and it is the main area where the wing bears the load.

[0087] The first pressure sensing unit refers to a sensor device specifically designed for measuring physical pressure. It is located in the folding hinge area of ​​the wing to acquire pressure data for that specific area. The following pressure sensing units acquire data by physically mounting them on the surface of that area and connecting them to the data acquisition circuitry of the flight control system.

[0088] The second pressure sensing unit refers to another sensor device used to measure physical pressure, which is arranged in the wingtip area of ​​the wing to acquire pressure data in that specific area.

[0089] The third pressure sensing unit refers to the third sensor device used to measure physical pressure. It is located in the wing root region of the wing and is used to acquire pressure data for that specific area.

[0090] A strain sensing unit is a sensor device specifically designed to measure structural deformation. It is installed at the hinge structure connecting the wing and fuselage to sense minute deformations of this critical connection when subjected to stress and output strain data. The data is acquired by mounting the unit on the surface of the hinge structure and connecting it to the data acquisition circuitry of the flight control system.

[0091] Pressure data refers to the specific numerical value of air pressure at a given moment, obtained directly from each measurement by the pressure sensing unit. The physical pressure is sensed by the sensitive element inside the pressure sensing unit and converted into an electrical signal, which is then quantized by an analog-to-digital converter to obtain the data.

[0092] Strain data refers to the specific numerical values ​​representing the degree of structural deformation at each moment, which are directly measured by the strain sensing unit. The resistance change caused by structural deformation is sensed by the resistance strain gauge inside the strain sensing unit, and the data is converted into an electrical signal, which is then quantized by an analog-to-digital converter to obtain the data directly.

[0093] In this step, the first pressure sensing unit, the second pressure sensing unit, and the third pressure sensing unit are first securely installed on the surfaces of the three key functional areas of the folding wing UAV wing, namely the folding hinge area, the wingtip area, and the wing root area, respectively, using dedicated mounting bases and fasteners. At the same time, high-strength strain adhesive is used to precisely attach and install the strain sensing unit on the surface of the metal hinge structure connecting the wing and the fuselage, thereby completing the hardware deployment of a multi-point sensor network covering key aerodynamic and structural parts.

[0094] Secondly, by parsing the instructions from the remote controller or preset waypoints, when it is confirmed that the drone needs to start performing the target action of folding or unfolding, it will immediately send a high-level electronic switch signal to the power control circuit connecting all pressure sensing units and strain sensing units through its universal digital input / output interface. This electronic switch signal utilizes the parallel conduction characteristics of the hardware circuit to ensure that the power supply lines to all sensors are connected within almost the same microsecond, so that the first pressure sensing unit, the second pressure sensing unit, the third pressure sensing unit and the strain sensing unit are powered on and initialized synchronously, thereby establishing a unified and accurate absolute time starting point for all data acquisition.

[0095] Next, relying on its internal hardware timer module, it generates a very stable interrupt pulse signal, for example, with the period strictly set to 10 milliseconds. Whenever this timed interrupt signal arrives, the interrupt service routine of the central processing unit will start immediately. Through its integrated analog-to-digital converter interface, it reads the analog voltage signal representing the current air pressure output in real time from the first pressure sensing unit, and converts this continuous voltage value into a specific digital value, thereby obtaining a pressure data of the folding hinge area at this moment. Then, using the same technical process, it sequentially reads and converts the analog voltage signals of the second and third pressure sensing units to obtain the pressure data of the wingtip and wing root areas at this moment, respectively. Within the same interrupt service routine cycle, it will also synchronously read the analog voltage signal output by the strain sensing unit through another analog-to-digital conversion channel and quantize it into a digital value to obtain a strain data of the hinge structure at this moment.

[0096] Then, after completing the data acquisition at each fixed time point, a data management function is called to store the three pressure data points with the same time mark that were just acquired, namely the data from the first, second and third pressure sensing units respectively, into a predefined data buffer according to their corresponding sensor numbers. At the same time, the acquired strain data is stored into another corresponding buffer. As time goes by and the timer interrupt is continuously triggered, this data management function continues to work, continuously appending the data obtained from each acquisition cycle to their respective buffers in chronological order.

[0097] Finally, when a target action is completed or the predetermined acquisition duration is reached, the data in the buffer will be encapsulated. All time-sorted pressure data sets from the three pressure sensing units will be packaged and marked into a complete data packet with a time-series structure. This data packet is the pressure information. Similarly, all time-sorted strain data sets from the strain sensing units will be packaged into another complete data packet with a time-series structure. This data packet is the strain information. At this point, a set of original observation datasets containing synchronous pressure changes and structural deformation processes at multiple points in space is ready for subsequent analysis and judgment.

[0098] For example, Company A, which engages in power line inspection, equipped its folding-wing drone B with a status monitoring system. First, pressure sensing units were installed at the folding hinges, left and right wingtips, and wing roots of the drone's wings, and strain sensing units were attached to the main load-bearing hinges. Second, during an inspection mission, when drone B flew to the target and prepared to unfold its wings, the flight control system simultaneously activated all sensors. Then, at a fixed frequency of 100 times per second, pressure data and hinge strain data at the three locations were collected cyclically, and a precise time stamp was added to each data point. These timestamped data were continuously recorded and categorized during the wing unfolding process. Finally, after the action was completed, a complete synchronous data report was generated, recording the multi-point pressure changes and hinge deformation process during the unfolding process, i.e., pressure information and strain information.

[0099] This step, through the synchronous deployment of sensors at multiple points, establishes a real-time data acquisition foundation for simultaneously sensing aerodynamic pressure and structural deformation of key components during wing shape transformation. This ensures a strict correspondence between pressure and strain data in time and space, providing highly consistent raw information for subsequent accurate identification of complex phenomena caused by the coupling of structural deformation and airflow disturbance.

[0100] Step 102: Based on the pressure information and the strain information of the wing connection structure, determine whether aeroelastic coupling occurs during the operation of the wing.

[0101] Optionally, step 102 may specifically include:

[0102] Step 1021: Extract the first pressure data sequence from the folding hinge region, the second pressure data sequence from the wingtip region, and the third pressure data sequence from the wing root region from the pressure information.

[0103] Step 1022: Extract the first strain data sequence from the folding hinge region, the second strain data sequence from the wingtip region, and the third strain data sequence from the wing root region from the strain information.

[0104] Step 1023: Identify the time points when pressure values ​​change abruptly in the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence, and mark the time when the pressure value first changes abruptly as the initial disturbance time of each region.

[0105] Step 1024: In the first strain data sequence, the second strain data sequence, and the third strain data sequence, identify the time point when the strain value changes abruptly, and mark the time when the strain value first changes abruptly as the initial disturbance time of the structure.

[0106] Step 1025: Compare the initial disturbance times of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence with the order of the initial disturbance times of the structure.

[0107] Step 1026: If the initial disturbance time of at least one pressure data sequence is earlier than the initial disturbance time of the structure, it is determined that the wing did not undergo aeroelastic coupling during operation.

[0108] Step 1027: If the initial disturbance time of all pressure data sequences is later than the initial disturbance time of the structure, then it is determined that the wing has undergone aeroelastic coupling during operation.

[0109] In this step, aeroelastic coupling refers to the physical phenomenon in which the changes in airflow pressure around the wing and the vibration or deformation of the wing structure interact and intensify each other when the UAV wing is folding or unfolding. The causal relationship between the airflow change sequence reflected by the pressure information and the structural change sequence reflected by the strain information is determined by analyzing the sequence of causes and effects.

[0110] The first pressure data sequence refers to a data list consisting of multiple pressure values ​​arranged in chronological order of acquisition time. It is used to describe the history of air pressure changes in the folding hinge area throughout the target movement process and is obtained by sorting all pressure data from the first pressure sensing unit according to timestamps.

[0111] The second pressure data sequence refers to a data list consisting of multiple pressure values ​​arranged in chronological order of acquisition time. It is used to describe the history of air pressure changes in the wingtip region throughout the entire target maneuver and is obtained by sorting all pressure data from the second pressure sensing unit according to timestamps.

[0112] The third pressure data sequence refers to a data list consisting of multiple pressure values ​​arranged in chronological order of acquisition time. It is used to describe the history of air pressure changes in the wing root region throughout the target's movement and is obtained by sorting all pressure data from the third pressure sensing unit according to timestamps.

[0113] The first strain data sequence refers to a list of multiple strain values ​​arranged in chronological order of acquisition time. It is used to describe the deformation history of the connecting structure corresponding to the folded hinge region during the entire process of target action. It is obtained by monitoring relevant strain data from the folded hinge region and sorting them by timestamp.

[0114] The second strain data sequence refers to a list of multiple strain values ​​arranged in chronological order of acquisition time. It is used to describe the deformation history of the connecting structure corresponding to the wingtip region during the entire target movement process. It is obtained by monitoring relevant strain data from the wingtip region and sorting them by timestamp.

[0115] The third strain data sequence refers to a list of multiple strain values ​​arranged in chronological order of acquisition time. It is used to describe the deformation history of the connecting structure corresponding to the wing root region during the entire target action process. It is obtained by monitoring relevant strain data from the wing root region and sorting them by timestamp.

[0116] The pressure value refers to each specific data point that makes up the pressure data sequence. It is a number used to represent the magnitude of the air pressure at the location where the pressure sensor is installed at a specific instant.

[0117] The initial disturbance moment refers to a specific point in time used to mark the instant at which a significant, abnormal, and drastic pressure change first occurs in the pressure data sequence of a specific region of the wing. It is obtained by analyzing and identifying the pressure data sequence using a mutation point detection algorithm.

[0118] The strain value refers to each specific data point that makes up the strain data sequence. It is a number used to represent the degree of material deformation at the location of the strain sensor installation point at a specific instant.

[0119] The initial disturbance moment of a structure refers to a specific point in time, used to mark the instant at which a significant, abnormal, and drastic change in structural deformation first occurs in the strain data sequence. It is obtained by analyzing multiple strain data sequences using a mutation point detection algorithm and taking the earliest moment.

[0120] In this step, firstly, from the obtained pressure information data packet, through data parsing and filtering operations, based on the sensor number tag attached to each data point, all data points from the first pressure sensing unit are extracted and arranged in time stamp order to form the first pressure data sequence. The same data filtering and sorting method is used to extract and form the second and third pressure data sequences. At the same time, from the strain information data packet, according to the preset area mapping rules, strain data corresponding to the structural deformation of the monitored folding hinge, wingtip, and wing root regions are extracted and arranged in time order to form the first strain data sequence, the second strain data sequence, and the third strain data sequence.

[0121] Secondly, a mutation detection algorithm combining a sliding time window and statistical threshold comparison is used to analyze the first pressure data sequence. The mutation detection algorithm uses a short time window of fixed length to slide on the pressure data sequence and continuously calculate the statistical characteristics of the data within the window, such as the mean and standard deviation. When the mutation detection algorithm detects a certain time point, if the values ​​of multiple newly entered data points in a row change drastically relative to the statistical characteristics of the historical window, exceeding the preset threshold, then the time point is determined to be the starting point of the pressure value mutation. The earliest identified mutation time point is recorded and marked as the initial disturbance time of the folding hinge region. The same mutation detection algorithm is synchronously applied to the second and third pressure data sequences to identify and mark the initial disturbance times of the wingtip region and the wing root region, respectively.

[0122] Next, the sliding window mutation detection algorithm was applied to the first strain data sequence, the second strain data sequence, and the third strain data sequence to identify the moment when the strain value first changed. After obtaining the three strain mutation moment points, the smallest value was selected from the three moment points through a simple numerical comparison logic. That is, the earliest moment point, and this earliest strain mutation moment was uniformly marked as the initial disturbance moment of the entire wing connection structure.

[0123] Then, a logical comparison and decision stage is entered. The initial disturbance time of the first pressure data sequence, the initial disturbance time of the second pressure data sequence, and the initial disturbance time of the third pressure data sequence are extracted in sequence and compared with the initial disturbance time of the structure. It is then determined whether each pressure initial disturbance time is earlier or later than the initial disturbance time of the structure.

[0124] Finally, based on the time-series comparison results, the final judgment logic is executed. If any of the three initial pressure disturbance times is earlier than the initial structural disturbance time, then according to the physical law that airflow pressure disturbance precedes structural deformation, it is determined that airflow change was the active factor during this wing movement, and no vicious cycle of mutual aerodynamic and structural exacerbation occurred; therefore, it is judged that aeroelastic coupling did not occur. Conversely, if all three initial pressure disturbance times are later than the initial structural disturbance time, it indicates that structural deformation occurred earlier than the significant air pressure changes at all monitoring points, consistent with structural vibration triggering subsequent abnormal airflow, and thus... The coupling characteristics that may exacerbate vibrations are thus identified as aeroelastic coupling. The identification of abrupt changes in pressure or strain values ​​is achieved by calculating the sliding standard deviation of the data sequence in real time. For example, a sliding time window of length N (e.g., N=5) is set, and the standard deviation σ of the data within the window is continuously calculated. If the difference between the current data point and the data at the previous moment exceeds kσ (where k is a preset coefficient, usually taken as 2~3), then that moment is determined to be a point of abrupt change. The initial disturbance moment is the first moment point that meets this condition. This threshold setting method is based on statistical principles and can adapt to the fluctuation characteristics of the data itself.

[0125] For example, following the specific implementation of the previous step, after UAV B completes wing deployment and generates pressure and strain information, the analysis of these two sets of data begins. Next, based on the sensor number, pressure data from the folding hinge, wingtip, and wing root are separated from the pressure information, forming three time-ordered pressure data sequences. Simultaneously, the corresponding strain data sequences are extracted from the strain information. Then, a mutation detection program is run on each pressure data sequence. By analyzing the slope of the data change, the specific time point at which the pressure first experiences a sharp rise or fall is automatically identified in each pressure data sequence, and recorded as the initial disturbance time for the three regions. Then, using the same program, the moment of the first abrupt change in structural deformation is identified from the strain data sequence and recorded as the initial disturbance time of the structure.

[0126] Subsequently, the three initial pressure disturbance times were numerically compared with the initial structural disturbance times. Assuming that the comparison results showed that all pressure disturbance times were later than the structural disturbance times, and based on the preset rule that a pressure disturbance being later than a strain disturbance constitutes coupling, it was ultimately determined that aeroelastic coupling occurred during this wing deployment process.

[0127] This step establishes an objective logical method for judging whether aeroelastic coupling has occurred by comparing the occurrence sequence of pressure mutations and strain emergencies. This method transforms multi-point synchronous monitoring data into a clear causal time sequence relationship, which can effectively distinguish between simple airflow disturbances and dangerous fluid-structure interaction phenomena, thus providing a clear and reliable decision basis for whether to activate advanced control.

[0128] Step 103: When it is determined that aeroelastic coupling has occurred, determine the airflow control command corresponding to the current coupling state.

[0129] Optionally, step 103 may specifically include:

[0130] Step 1031: After determining that the aeroelastic coupling has occurred, obtain the pressure values ​​of the first pressure data sequence, the second pressure data sequence and the third pressure data sequence at the corresponding initial disturbance time.

[0131] Step 1032: Calculate the difference between the pressure values ​​corresponding to the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence to obtain multiple pressure difference values.

[0132] Step 1033: Obtain the strain value at the moment of initial disturbance of the structure.

[0133] Step 1034: Determine the basic pulse frequency based on the magnitude of the strain value.

[0134] Step 1035: Calculate the frequency adjustment coefficient based on the largest pressure difference.

[0135] Optionally, step 1035 may include the following steps: comparing the largest pressure difference value with each pressure difference value interval in a pre-stored lookup table; determining the target pressure difference value interval to which the largest pressure difference value belongs; finding the adjustment coefficient value corresponding to the target pressure difference value interval from the lookup table; and calculating the frequency adjustment coefficient using a linear interpolation method when the largest pressure difference value is located at the boundary of two adjacent pressure difference value intervals.

[0136] Step 1036: Multiply the base pulse frequency by the frequency adjustment coefficient to obtain the target pulse frequency.

[0137] Step 1037: Based on the relative pressure values ​​of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence, select the starting position for applying additional airflow from the folding hinge region, the wingtip region, and the wing root region.

[0138] Step 1038: Generate the airflow control command based on the target pulse frequency and the starting position, wherein the airflow control command includes a pulse frequency parameter and an application position parameter.

[0139] In this step, the airflow control command refers to a digital command used to control the operation of the pulse jet actuator, which specifies the location and pulse rhythm for generating additional airflow, and is generated by combining the target pulse frequency and the position parameter.

[0140] The pressure difference is a specific number used to represent the degree of difference in pressure between any two different regions on the wing at the initial disturbance moment. It is obtained by subtracting two pressure values ​​from different pressure data sequences.

[0141] The basic pulse frequency refers to a number that represents the basic speed of pulse operation. It is used as a reference for calculating the final control frequency and is determined by consulting a predefined correspondence table or calculation formula based on the strain value.

[0142] The frequency adjustment factor is a proportional number used to amplify or reduce the fundamental pulse frequency. It is used to adjust the control intensity according to the severity of aerodynamic disturbances and is obtained by matching the maximum pressure difference with a pre-stored lookup table or by interpolation.

[0143] The pre-stored lookup table refers to a data table that is pre-stored in the control system and specifies the correspondence between different pressure difference ranges and adjustment coefficient values. It is used to map the measured physical quantity difference to the control parameter adjustment amount. It is obtained by pre-determining through experiments or simulations and burning it into the memory.

[0144] The adjustment coefficient value refers to a specific number stored in a pre-stored lookup table, representing the control intensity adjustment ratio corresponding to a certain pressure difference range, which is obtained by directly reading from the lookup table.

[0145] The target pulse frequency refers to a finalized specific number that represents the speed of the pulse jet operation. It is used to directly drive the actuator and is obtained by multiplying the basic pulse frequency with the frequency adjustment coefficient.

[0146] The pulse frequency parameter is a part of the airflow control command. It is a number used to transmit the target pulse frequency control information to the actuator. It is set by directly writing the target pulse frequency value into a specified field of the command data packet.

[0147] The action position parameter refers to another part of the airflow control command. It is a code or number used to specify to the actuator which wing area needs to be started. It is set by writing the selected area number into a specified field of the command data packet.

[0148] In this step, firstly, after determining that aeroelastic coupling has occurred, the initial disturbance times of the recorded first, second, and third pressure data sequences are used as time indices to backtrack and retrieve the three sequences. The three specific pressure values ​​recorded at those precise moments are then extracted. These pressure values ​​represent the instantaneous air pressure state of the three key regions when coupling occurs. Secondly, through a simple subtraction calculation, the numerical difference between the pressure values ​​corresponding to the first and second pressure data sequences is calculated sequentially to obtain the first pressure difference. Then, the numerical difference between the pressure values ​​corresponding to the first and third pressure data sequences is calculated to obtain the second pressure difference. Finally, the numerical difference between the pressure values ​​corresponding to the second and third pressure data sequences is calculated to obtain the third pressure difference. This process yields multiple pressure differences and quantifies the degree of pressure imbalance between different regions.

[0149] Simultaneously, based on the marked initial disturbance time of the structure, the strain value recorded at that time is extracted from the corresponding strain data sequence. This strain value represents the initial intensity of structural deformation when coupling occurs. Then, using a predefined strain-base frequency lookup table, the obtained strain value is compared with the entries in the lookup table to find the numerical range to which the strain value belongs, and the base pulse frequency value corresponding to that range is read from the lookup table. This base frequency reflects the reference control rhythm required by the structural disturbance itself.

[0150] Then, from the multiple calculated pressure difference values, a comparison algorithm is used to find the largest value, i.e., the largest pressure difference value. Next, a pre-stored lookup table in memory is accessed. This lookup table divides the pressure difference range into multiple consecutive intervals, each interval being associated with a specific adjustment coefficient value. The largest pressure difference value is compared with the boundaries of these intervals to determine which specific target pressure difference value interval it falls into. Subsequently, the adjustment coefficient value bound to that interval is directly read from the lookup table. As a more refined process, if the largest pressure difference value is exactly equal to the boundary value of a certain interval, or lies between the boundaries of two intervals, a linear interpolation calculation program is activated. Based on the distance ratio between this pressure difference value and the boundary values ​​of the two adjacent intervals, a weighted average is performed on the adjustment coefficient values ​​corresponding to these two intervals to calculate a new, more accurate adjustment coefficient value as the final frequency adjustment coefficient. This transforms the spatial difference of aerodynamic disturbance into an adjustment amount for the control intensity.

[0151] Next, a multiplication operation is performed, multiplying the determined base pulse frequency by the calculated frequency adjustment coefficient to obtain a new value called the target pulse frequency. This target pulse frequency is the final control rhythm that integrates the structural disturbance intensity and the spatial difference of aerodynamic disturbance. Subsequently, while determining the frequency, the position selection logic is executed in parallel, comparing the magnitudes of the three acquired pressure values. Based on the rule of prioritizing control in the region with the highest pressure to balance the distribution, the wing region corresponding to the pressure value with the largest value is identified, namely, one of the folding hinge region, wingtip region, or wing root region, and this region is selected as the starting position where additional airflow needs to be applied first. Finally, these two results are encapsulated to generate the final airflow control command. This airflow control command is a structured data packet, in which the calculated target pulse frequency value is written in the pulse frequency parameter field, and the selected region's number code is written in the action position parameter field. This command, which contains the specific action rhythm and action position, is then sent to the pulse jet actuator.

[0152] For example, following the specific implementation of the previous step, firstly, after determining that aeroelastic coupling has occurred, the control command generation process is initiated; secondly, the pressure values ​​at the three positions of the folding hinge, wingtip, and wing root recorded at their respective initial disturbance moments are obtained, and the differences between them are calculated. At the same time, the strain values ​​recorded at the initial disturbance moment of the structure are also obtained; then, by querying a preset correspondence, the basic pulse frequency is derived based on the strain value; simultaneously, the maximum value is found from the multiple calculated pressure differences, and the frequency adjustment coefficient corresponding to this maximum pressure difference is determined by a table lookup matching method. The basic pulse frequency is multiplied by the frequency adjustment coefficient to obtain the final target pulse frequency used for control.

[0153] Meanwhile, by comparing the specific pressure values ​​at the three locations, the area with the highest pressure value is selected as the initial application location. For example, the area with the highest pressure value is the folding hinge area, which is therefore selected. Finally, all decision results are encapsulated to generate an airflow control command containing specific pulse frequency parameters and application location parameters, ready to be sent to the actuator. The base pulse frequency is determined based on the strain value, which is based on a data mapping table obtained in advance through wind tunnel testing. This mapping table includes the following correspondence as an example:

[0154] When the strain value is in the low range (e.g., 0-100 microstrain), the base pulse frequency is set to a lower value (e.g., 20-40Hz), corresponding to mild structural deformation. When the strain value is in the high range (e.g., 200-300 microstrain), the base pulse frequency is set to a higher value (e.g., 60-80Hz), corresponding to stronger control intervention required for severe structural deformation. This mapping relationship ensures that the control intensity matches the degree of structural disturbance. The pre-stored lookup table is established based on a linear or nonlinear relationship model between the spatial differences in aerodynamic disturbance and the control effect. For example, when the maximum pressure difference ΔP_max is in a small range (e.g., 0-10Pa), the adjustment coefficient K≈1.0, indicating fine-tuning; when ΔP_max is in a large range (e.g., 30-50Pa), K can be increased proportionally to 1.5-2.0, indicating a need for significantly enhanced control to balance the large pressure gradient. Linear interpolation methods are used to smoothly transition at the interval boundaries, ensuring the continuity of parameter changes.

[0155] This step transforms the abstract judgment of aeroelastic coupling into a set of specific, executable control parameters that match the characteristics of the current coupling state. By comprehensively analyzing the differences between structural disturbance intensity and aerodynamic pressure distribution, the control intensity and initial action position are intelligently determined, thereby generating precise airflow control commands for the current disturbance mode. This achieves a closed loop from state perception to control decision-making, providing a direct basis for subsequent efficient and targeted active flow control.

[0156] Step 104: According to the airflow control command, control the pulse jet actuator set on the leading edge of the wing to generate an asymmetric additional airflow on the wing surface.

[0157] Optionally, step 104 may specifically include:

[0158] Step 1041: Analyze the airflow control command to obtain the target pulse frequency and starting position.

[0159] Step 1042: Activate the pulse jet unit located in the corresponding area of ​​the leading edge of the wing according to the starting position.

[0160] Step 1043: Generate a pulse drive signal with a corresponding time interval according to the target pulse frequency.

[0161] Step 1044: Send the pulse drive signal to the activated pulse jet unit and control the activated pulse jet unit to work according to the pulse drive signal.

[0162] Step 1045: When the pulse jet unit is working, a high-speed airflow is ejected from the gap at the leading edge of the wing through the pulse jet unit to act on the wing surface.

[0163] Step 1046: By controlling the pulse jet units in different regions to operate according to different pulse drive signals, high-speed airflows with different intensities and timings are generated at different positions on the wing surface to form an asymmetric additional airflow on the wing surface.

[0164] In this step, pulse jet refers to a high-speed airflow that is ejected from a narrow gap in an intermittent, rapid opening and closing manner. It is used to actively intervene in the airflow state on the wing surface and is generated by the pulse jet unit working periodically under the control of a pulse drive signal.

[0165] Additional airflow refers to the extra airflow actively introduced by external devices on top of the aircraft's original natural airflow. It is used to artificially change the local pressure distribution in specific areas of the wing surface and is formed by pulse jets acting on the wing surface.

[0166] A pulse jet unit refers to a single, minimal control and actuation unit that constitutes a pulse jet actuator. It includes a miniature valve, a nozzle, and a drive circuit, and is used to generate a pulsed jet when an electrical signal is received. Multiple such units are arranged in an array on the leading edge of the wing.

[0167] A pulse drive signal is a periodically changing electrical signal, usually a square wave, used to precisely control the opening and closing rhythm of the pulse jet unit. It is generated by a signal generation circuit based on the target pulse frequency parameters.

[0168] The pulse jet actuator is integrated inside the leading edge of the wing. Its core actuation unit is a miniature high-speed solenoid valve. The air inlet of the solenoid valve is connected to a high-pressure gas cylinder carried inside the UAV body (or an auxiliary airflow generated by the ejector device) through a miniature pipeline. Its air outlet is connected to a slit nozzle pre-set on the leading edge surface of the wing. The width of the slit is preferably 0.1 mm to 1 mm, the length direction is parallel to the leading edge of the wing, and the jet direction is at an angle of 15 to 45 degrees with the wing surface. It is used to guide the high-pressure gas to the boundary layer of the upper surface of the wing in the form of a pulse.

[0169] The pulse drive signal is a square wave electrical signal with a specific voltage (such as 12V or 24V). When the signal is high, the solenoid valve coil is energized, and the valve core opens under the action of electromagnetic force, and high-pressure gas is ejected through the nozzle to form a high-speed jet. When the signal is low, the valve core is reset under the action of the spring and cuts off the airflow. The speed and momentum of the jet can be controlled by adjusting the air source pressure (such as 0.2-0.8MPa) and the frequency and duty cycle of the pulse drive signal to ensure that the generated additional airflow can effectively change the local pressure distribution on the wing surface.

[0170] In this step, the flight control computer first parses the received airflow control command data packet, reads the predefined format fields in the data packet, extracts the specific value of the target pulse frequency from the pulse frequency parameter field (e.g., 60 Hz), and extracts the code representing a specific wing region from the action position parameter field (e.g., code 01 representing the folding hinge region). Next, based on the parsed region code 01, the hardware address of one or more specific pulse jet units belonging to the folding hinge region is retrieved through the internally stored position-execution unit mapping table. Then, a digital enable signal is sent to the drive circuit of the pulse jet unit corresponding to these addresses via the control bus, activating the unit by connecting its air supply and power supply circuits and putting it into standby mode, while unselected units located in other regions remain dormant.

[0171] Next, the signal generation module calculates the corresponding signal period based on the target pulse frequency value obtained from the analysis, such as 60 Hz. Then, through its internal hardware timer or dedicated waveform generation chip, it generates a digital square wave signal with a fixed period. Within a single period, the duration of the high level of this digital square wave signal represents the valve opening, and the duration of the low level represents the valve closing. Its overall repetition frequency is strictly equal to the target pulse frequency. This generated periodic digital square wave is the pulse drive signal.

[0172] Then, the generated pulse drive signal is sent to all activated pulse jet units in real time through the same or multiple control buses. The drive circuit of each activated pulse jet unit receives this signal and strictly controls the opening and closing of its internal miniature high-speed solenoid valve according to the high and low level changes of the signal. When the signal is high, the solenoid valve opens instantly, and compressed air is ejected from the air source through the nozzle; when the signal is low, the solenoid valve closes and the airflow stops.

[0173] Subsequently, as the pulse jet unit operates in accordance with the drive signal, each time the solenoid valve opens, the compressed air stored in the unit's internal cavity is ejected at high speed from a narrow slit less than one millimeter wide, pre-designed at the leading edge of the wing. This forms a high-speed, transient jet perpendicular to the wing surface or at a specific angle to the surface, and this jet directly acts on the boundary layer air of the wing surface. Finally, the control system sends pulse drive signals of different frequencies, phases, or duty cycles to the pulse jet units in different zones. This can generate multiple high-speed jets with different intensities, durations, and rhythms at different positions on the left and right wings, or in front and behind the wings. These spatially asynchronous and unevenly intense jets work together to introduce an artificial and controllable asymmetric disturbance into the overall airflow field of the wing, namely, an asymmetric additional airflow.

[0174] For example, following the specific implementation of the previous step, firstly, the flight control system of UAV B sends the generated airflow control command to the controller at the leading edge of the wing, then begins to parse the airflow control command, obtains the target pulse frequency as 60Hz, and the starting position of the action as the folding hinge area; then, according to the preset mapping relationship, activates a specific set of pulse jet units arranged at the leading edge of the folding hinge area.

[0175] Then, the signal generation module generates a corresponding periodic electrical signal, i.e., a pulse drive signal, based on the target pulse frequency of 60Hz. This signal is then sent to all activated pulse jet units via a circuit. The solenoid valves inside these units strictly follow the rhythm of the received pulse drive signal, performing high-speed opening and closing actions. Subsequently, when the solenoid valve opens, high-pressure gas is instantly ejected from the micro-slits at the leading edge of the unit, forming a high-speed pulse jet. When the solenoid valve closes, the airflow stops. This cycle repeats on the wing surface in the designated folding hinge area, generating an additional pulse airflow with a frequency of 60Hz. Finally, because the jet units in other areas are not activated, an asymmetrical distribution of airflow is formed on the wing surface, i.e., an asymmetrical additional airflow.

[0176] This step accurately translates digital control commands into physical intervention actions on the wing surface. By parsing the commands and selectively activating the execution units, it enables the active generation of pulse jets at a precise rhythm at a designated location, thereby quickly establishing a controllable, asymmetric additional airflow field and providing a reliable execution means for directly generating compensating torque.

[0177] Step 105: The additional airflow generates a compensating torque to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers.

[0178] Optionally, step 105 may specifically include:

[0179] Step 1051: Monitor the pitch angle and roll angle changes of the fuselage of the folding-wing UAV in real time.

[0180] Step 1052: When the pitch angle change exceeds a first predetermined value or the roll angle change exceeds a second predetermined value, it is determined that there is an attitude disturbance.

[0181] Step 1053: Determine the target wing region that needs to generate compensating torque based on the direction of the attitude disturbance.

[0182] Step 1054: Adjust the operating parameters of the pulse jet actuator in the target wing region to control the local intensity of the additional airflow.

[0183] Step 1055: An asymmetric aerodynamic pressure distribution is formed on the wing surface by adjusting the additional airflow to generate a compensating torque.

[0184] Step 1056: Use the compensation torque to counteract the fuselage attitude disturbance, so as to offset the fuselage attitude disturbance caused by the folding-wing UAV performing target actions.

[0185] In this step, fuselage attitude disturbance refers to the unexpected, deviating pitch or roll direction unstable motion of the fuselage when the folding-wing UAV is performing a target maneuver. It is used to describe the flight state that needs to be suppressed or corrected.

[0186] The pitch angle change refers to a number that indicates the speed at which the drone's nose swings up and down. It is used to quantify the severity of pitch attitude disturbances and is obtained by calculating the difference between the pitch angle measured at the current moment and the pitch angle measured at the previous moment.

[0187] The roll angle change refers to a number that indicates how fast the UAV body rolls around its longitudinal axis. It is used to quantify the severity of attitude disturbances in the roll direction and is obtained by calculating the difference between the roll angle measured at the current moment and the roll angle measured at the previous moment.

[0188] The first predetermined value refers to a pre-set numerical threshold for the change in pitch angle, used to determine whether the attitude disturbance in the pitch direction is severe enough to require the activation of the compensation mechanism. Both predetermined values ​​are determined through ground tests and written into the flight control system.

[0189] The second predetermined value refers to a pre-set numerical threshold for the amount of change in the roll angle, used to determine whether the attitude disturbance in the roll direction is severe enough to require the activation of a compensation mechanism.

[0190] Adjusted auxiliary airflow refers to an auxiliary airflow whose intensity or rhythm has been altered to produce a stronger or weaker aerodynamic intervention effect in a specific target wing area.

[0191] Compensating torque refers to a physical torque that causes the drone's fuselage to rotate. Its direction is opposite to the direction of the fuselage attitude disturbance and is used to counteract and cancel the effect of the disturbance.

[0192] In this step, the inertial measurement unit installed on the folding-wing UAV is used to continuously measure and output the real-time pitch and roll angles of the fuselage at a fixed high frequency. At the same time, these angle values ​​are recorded. Through a simple differential calculation process, that is, the angle value collected at the current moment is subtracted from the angle value collected at the previous moment, the pitch angle change and roll angle change are calculated in real time and directly reflect the instantaneous rate of attitude disturbance.

[0193] Next, the absolute value of the calculated pitch angle change is compared with a numerical threshold called the first predetermined value that is stored in memory. At the same time, the absolute value of the calculated roll angle change is compared with another pre-stored value called the second predetermined value. This comparison process is completed by the processor's logical comparison instruction. If the absolute value of the pitch angle change is greater than the first predetermined value, or the absolute value of the roll angle change is greater than the second predetermined value, the logic judgment module outputs an "yes" signal, thereby determining that there is a fuselage attitude disturbance that needs to be processed.

[0194] Next, once a disturbance is determined, the specific intervention location needs to be determined. This depends on whether the pitch angle change or the roll angle change exceeds the limit, and whether the change is positive or negative. By querying a preset disturbance type / direction response area relationship mapping table, it is determined which specific wing section needs to be subjected to stronger airflow intervention. That is, the determined area is the target wing area.

[0195] Then, a control command is generated and sent to the local controller of the pulse jet actuator. This control command explicitly requires the adjustment of the operating parameters of all or some of the pulse jet units located in the target wing area. The most common adjustment method is to increase or decrease the target pulse frequency of these units. After the local controller receives the control command, it modifies the settings of its internal waveform generator to generate new pulse drive signals with different frequencies and sends them to the actuators in the target area. This achieves the adjustment of the intensity or rhythm of the additional airflow in a specific area, resulting in the adjusted additional airflow.

[0196] These adjusted additional airflows then act on the wing surface of the target wing area. According to the principles of fluid dynamics, a stronger pulse jet will accelerate the local airflow in the target area more quickly, thereby significantly reducing the surface air pressure in the target area; conversely, a weaker jet will cause the air pressure in the area to increase relatively. This artificially creates a pressure difference between different areas of the wing, which acts on the wing area and forms a couple that causes the aircraft to rotate around its center of gravity, i.e., a compensating torque.

[0197] Finally, according to the laws of physics, this compensating torque will generate an angular acceleration, driving the fuselage to rotate in the opposite direction to the current fuselage attitude disturbance. By precisely controlling the magnitude and direction of the compensating torque, it can be made equal in magnitude and opposite in direction to the torque that caused the disturbance, thus canceling each other out in dynamics, and ultimately achieving the purpose of suppressing or eliminating the fuselage attitude disturbance and restoring flight stability.

[0198] For example, following the specific implementation of the previous step, after generating additional airflow on the wing surface, UAV B continues to perform its mission, and its onboard sensors continuously measure the fuselage attitude; secondly, through real-time calculation, it is found that the angle change rate of the nose pitching up continuously exceeds the preset safety threshold in a short period of time, but the change in the roll direction is normal; then, it is determined that there is a pitch attitude disturbance that requires intervention, and according to preset rules, for this pitch disturbance, it is determined that stronger intervention needs to be applied to the trailing edge region of the left and right wings.

[0199] Then, a command is sent to the controller that controls the pulse jet units at the trailing edges of the left and right wings, requesting an increase in the operating frequency of the jets in that trailing edge region. After the jet frequency is increased, the intensity of the additional airflow generated at the trailing edges of the left and right wings increases, which further reduces the air pressure on the lower surface of the wing in that trailing edge region. The resulting pressure difference on the wing generates a torque that causes the nose to turn downward, i.e., a compensating torque. Finally, this newly generated compensating torque acts on the fuselage, effectively counteracting and canceling the disturbance torque that causes the nose to pitch up. The nose-up tendency of the UAV is quickly suppressed, and the flight attitude is restored to stability.

[0200] This step constructs a rapid closed loop from attitude disturbance perception to dynamic airflow compensation. By monitoring attitude changes in real time and intelligently locating the intervention area, the airflow intensity is dynamically adjusted to generate a precise compensation torque, thereby achieving online active suppression of sudden attitude disturbances during wing movements. Ultimately, it completes the entire process of autonomous stabilization control from detection to cancellation, significantly improving the aircraft's anti-disturbance capability.

[0201] The technical effect of this application is based on the following well-known fluid dynamics principle. According to the Kuta-Zhukovsky theorem, circulation control can directly change the lift characteristics of the wing. The asymmetric additional airflow introduces controllable momentum into a specific area of ​​the wing, changing the equivalent camber and circulation distribution of that area, thereby generating a compensating torque opposite to the direction of the disturbance torque. The magnitude of this torque is proportional to the jet velocity, mass flow rate, and distance from the point of action to the center of gravity of the aircraft. Its direction can be controlled by selecting different areas of action.

[0202] Therefore, by precisely controlling the parameters of the pulse jet, a compensating torque sufficient to offset attitude disturbances can be generated. The response time of the entire perception-decision-control closed loop can be controlled within tens of milliseconds through optimized algorithms and hardware selection, meeting the real-time requirements for suppressing transient disturbances.

[0203] Figure 2 is a schematic diagram of the attitude control system for a folding-wing UAV provided in this application. As shown in Figure 2, the system includes:

[0204] The acquisition module 21 is used to acquire pressure information at multiple positions of the wing and strain information of the wing connection structure when the folding wing UAV performs target maneuvers.

[0205] The judgment module 22 is used to determine whether aeroelastic coupling occurs during the operation of the wing based on the pressure information and the strain information of the wing connection structure.

[0206] The determination module 23 is used to determine the airflow control command corresponding to the current coupling state when it is determined that aeroelastic coupling has occurred;

[0207] Control module 24 is used to control the pulse jet actuator set on the leading edge of the wing to generate asymmetric additional airflow on the wing surface according to the airflow control command;

[0208] The additional module 25 is used to generate a compensating torque through the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target maneuvers.

[0209] The attitude control system of a folding-wing UAV shown in Figure 2 can execute the attitude control method of a folding-wing UAV shown in the embodiment of Figure 1. Its implementation principle and technical effects will not be elaborated further. The specific methods by which each module and unit of the attitude control system of the folding-wing UAV in the above embodiments perform operations have been described in detail in the embodiments related to this method, and will not be elaborated upon here.

[0210] In one possible design, the attitude control system of a folding-wing UAV of the embodiment shown in FIG2 can be implemented as a computing device, as shown in FIG3, which may include a storage component 31 and a processing component 32.

[0211] The storage component 31 stores one or more computer instructions, wherein the one or more computer instructions are invoked and executed by the processing component 32.

[0212] The processing component 32 is used for an attitude control method for a folding-wing unmanned aerial vehicle according to the embodiment of FIG1.

[0213] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of this application, and are not intended to limit them. Although this application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features. Such modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of this application.

Claims

1. An attitude control method for a folding-wing unmanned aerial vehicle (UAV), characterized in that, include: The system collects pressure information at multiple locations on the wing and strain information of the wing connection structure when the folding-wing UAV performs target maneuvers. Based on the pressure information and the strain information of the wing connection structure, it is determined whether aeroelastic coupling occurs during the wing's movement. When aeroelastic coupling is determined to occur, an airflow control command corresponding to the current coupling state is determined. Based on the airflow control command, the pulse jet actuator located at the leading edge of the wing is controlled to generate an asymmetric additional airflow on the wing surface. A compensating torque is generated through the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target movements.

2. The method according to claim 1, characterized in that, The method involves collecting pressure information from multiple locations on the wing and strain information from the wing connection structure of a folding-wing UAV during target maneuvers. This includes: arranging a first pressure sensing unit in the folding hinge area of ​​the wing, a second pressure sensing unit in the wingtip area, and a third pressure sensing unit in the wing root area; installing a strain sensing unit at the hinge structure connecting the wing and fuselage; synchronously activating all pressure sensing units and the strain sensing unit when the folding-wing UAV begins to perform the target maneuver; collecting pressure data from the first, second, and third pressure sensing units at fixed time intervals; simultaneously collecting strain data from the strain sensing unit at the same time intervals; and integrating all collected pressure data into pressure information and all collected strain data into strain information.

3. The method according to claim 1, characterized in that, Based on the pressure information and the strain information of the wing connection structure, determining whether aeroelastic coupling occurs during wing movement includes: extracting a first pressure data sequence from the folding hinge region, a second pressure data sequence from the wingtip region, and a third pressure data sequence from the wing root region from the pressure information; extracting a first strain data sequence from the folding hinge region, a second strain data sequence from the wingtip region, and a third strain data sequence from the wing root region from the strain information; identifying the time points where pressure values ​​abruptly change in the first, second, and third pressure data sequences, and marking the first time a pressure value abrupt change occurs as the initial pressure change in each region. The initial disturbance time is determined by identifying the moment when the strain value changes abruptly in the first, second, and third strain data sequences, and marking the moment when the strain value first changes abruptly as the initial disturbance time of the structure. The order of the initial disturbance times of the first, second, and third pressure data sequences with the initial disturbance time of the structure is compared. If the initial disturbance time of at least one pressure data sequence is earlier than the initial disturbance time of the structure, it is determined that the wing did not undergo aeroelastic coupling during operation. If the initial disturbance times of all pressure data sequences are later than the initial disturbance time of the structure, it is determined that the wing underwent aeroelastic coupling during operation.

4. The method according to claim 1, characterized in that, When aeroelastic coupling is determined to occur, the airflow control command corresponding to the current coupling state is determined, including: after determining that the aeroelastic coupling has occurred, acquiring the pressure values ​​of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence at the corresponding initial disturbance time; calculating the differences between the pressure values ​​corresponding to the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence to obtain multiple pressure difference values; acquiring the strain value at the initial disturbance time of the structure; determining the basic pulse frequency based on the magnitude of the strain value; calculating the frequency adjustment coefficient based on the largest pressure difference value; multiplying the basic pulse frequency by the frequency adjustment coefficient to obtain the target pulse frequency; selecting the starting position for applying additional airflow from the folding hinge region, the wingtip region, and the wing root region according to the magnitude relationship of the pressure values ​​of the first pressure data sequence, the second pressure data sequence, and the third pressure data sequence; generating the airflow control command based on the target pulse frequency and the starting position, wherein the airflow control command includes a pulse frequency parameter and an application position parameter.

5. The method according to claim 4, characterized in that, The frequency adjustment coefficient is calculated based on the largest pressure difference value, including: comparing the largest pressure difference value with each pressure difference value interval in a pre-stored lookup table; determining the target pressure difference value interval to which the largest pressure difference value belongs; finding the adjustment coefficient value corresponding to the target pressure difference value interval from the lookup table; and calculating the frequency adjustment coefficient using a linear interpolation method when the largest pressure difference value is located at the boundary between two adjacent pressure difference value intervals.

6. The method according to claim 1, characterized in that, According to the airflow control command, the pulse jet actuator installed on the leading edge of the wing generates an asymmetric additional airflow on the wing surface, including: parsing the airflow control command to obtain the target pulse frequency and starting position; activating the pulse jet unit installed in the corresponding area of ​​the leading edge of the wing according to the starting position; generating a pulse drive signal with a corresponding time interval according to the target pulse frequency; sending the pulse drive signal to the activated pulse jet unit and controlling the activated pulse jet unit to operate according to the pulse drive signal; when the pulse jet unit is operating, high-speed airflow is ejected from the gap of the leading edge of the wing through the pulse jet unit to act on the wing surface; by controlling the pulse jet units in different areas to operate according to different pulse drive signals, high-speed airflows with different intensities and timings are generated at different positions on the wing surface to form an asymmetric additional airflow on the wing surface.

7. The method according to claim 1, characterized in that, The method of generating a compensating torque through the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV's target maneuver includes: real-time monitoring of the pitch and roll angle changes of the folding-wing UAV; determining the presence of an attitude disturbance when the pitch angle change exceeds a first predetermined value or the roll angle change exceeds a second predetermined value; determining the target wing region for which a compensating torque needs to be generated based on the direction of the attitude disturbance; adjusting the operating parameters of the pulse jet actuator in the target wing region to control the local intensity of the additional airflow; generating a compensating torque by forming an asymmetric aerodynamic pressure distribution on the wing surface through the adjusted additional airflow; and using the compensating torque to counteract the fuselage attitude disturbance, thereby offsetting the fuselage attitude disturbance caused by the folding-wing UAV's target maneuver.

8. An attitude control system for a folding-wing unmanned aerial vehicle, characterized in that, include: The data acquisition module is used to collect pressure information at multiple positions of the wing and strain information of the wing connection structure when the folding-wing UAV performs target maneuvers. The judgment module is used to determine whether aeroelastic coupling occurs during the operation of the wing based on the pressure information and the strain information of the wing connection structure. The determination module is used to determine the airflow control command corresponding to the current coupling state when it is determined that aeroelastic coupling has occurred; the control module is used to control the pulse jet actuator set on the leading edge of the wing to generate an asymmetric additional airflow on the wing surface according to the airflow control command; the auxiliary module is used to generate a compensating torque through the additional airflow to counteract the fuselage attitude disturbance caused by the folding-wing UAV performing target actions.

9. A computing device, characterized in that, It includes a processing component and a storage component; the storage component stores one or more computer instructions; the one or more computer instructions are invoked and executed by the processing component to implement the attitude control method of a folding-wing unmanned aerial vehicle as described in any one of claims 1 to 7.

10. A computer storage medium, characterized in that, The device contains a computer program that, when executed by a computer, implements an attitude control method for a folding-wing unmanned aerial vehicle as described in any one of claims 1 to 7.

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