A wing gust mitigation control method based on combined plasma excitation
By installing a combined plasma exciter on the wing of the drone, and controlling aerodynamic fluctuations by inducing jets, the problem of drone's severe aerodynamic changes in gust conditions is solved, and the gust load is slowed down and the air stagnation time is improved.
Patent Information
- Application Number
- CN202510107619.0
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2025-01-23
- Publication Date
- 2025-05-06
- Estimated Expiration
- 2045-01-23
AI Technical Summary
Due to the gust effect of drones, aerodynamics and torque will undergo drastic changes in aerodynamics and torques, resulting in challenges in safe flight. The prior art will lead to a reduction in flight radius and air stagnation time by increasing the structure or equipping the tail edge flaps.
The wing gust mitigation control method based on combined plasma excitation is adopted. By installing symmetric and asymmetric plasma exciters on the leading and trailing edges of the wing, the plasma exciters are used to generate inducible jets, control the aerodynamic fluctuations of the wing, and slow down the gust load.
Effectively slows down gust loads and improves the drone's air stagnation time. The plasma exciter has little impact on the structure and weight of the drone, does not affect the flight radius, and is simple in installation structure, which is easy to maintain.
Smart Images
Figure CN119527537B_ABST
Abstract
Description
Technical Field
[0001] The invention belongs to the technical field of aircraft gust mitigation control, and in particular relates to a wing gust mitigation control method based on combined plasma excitation. Background Art
[0002] When a drone is flying at low altitude, due to the complex terrain, chaotic airflow distribution and strong gust effect, the aerodynamic force and torque acting on the drone will change dramatically under the strong influence of gusts, posing a major challenge to safe flight.
[0003] In order to ensure the safe flight of UAVs at low altitudes, gust interference is often resisted by strengthening the UAV structure, installing trailing edge flaps, deflecting control surfaces, etc. However, the increase in structural weight will lead to a decrease in the flight radius and hovering time of the UAV, which is not conducive to the use of the UAV. Summary of the invention
[0004] The purpose of this application is to provide a wing gust mitigation control method based on combined plasma excitation to solve the above-mentioned technical problems existing in the prior art.
[0005] This application is implemented as follows:
[0006] In a first aspect, an embodiment of the present application provides a wing gust mitigation control method based on combined plasma excitation, and the control steps include: S1: obtaining an aerodynamic characteristic curve of the wing; S2: determining a first installation position based on the aerodynamic characteristic curve obtained in step S1, the first installation position is located in the upper surface area of the leading edge of the wing, and a first plasma actuator is installed at the first installation position, the first plasma actuator is a symmetrical structure, and the first plasma actuator is used to increase the angle of attack of the wing; S3: determining a second installation position based on the aerodynamic characteristic curve obtained in step S1, the second installation position is located in the upper surface area of the trailing edge of the wing, and a second plasma actuator is installed at the second installation position, the second plasma actuator is an asymmetrical structure, and the second plasma actuator is used to reduce the angle of attack of the wing; S4: obtaining a gust signal of the wing tip of the drone wing or the nose of the drone, and controlling the operation of the first plasma actuator or the second plasma actuator according to the gust signal; steps S2 and S3 are in no particular order.
[0007] The technical solution adopted by the present invention can achieve the following beneficial effects:
[0008] In the present application, a first plasma actuator and a second plasma actuator are installed on the wing, and an induced jet is generated on the wing surface by the plasma actuator to act on the airflow field on the wing surface, thereby controlling the aerodynamic fluctuations of the wing, reducing the gust load, and increasing the hovering time of the UAV; and the plasma actuator has little or almost no effect on the structure and weight of the UAV, and has little or no effect on the flight radius of the UAV. After the plasma actuator is installed, the overall structure of the wing is simple, easy to maintain, and has a fast response speed; the first installation position and the second installation position are obtained based on the aerodynamic characteristic curve of the wing to enhance the effect of the plasma actuator and further improve the control effect of reducing the gusts on the wing. BRIEF DESCRIPTION OF THE DRAWINGS
[0009] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments of the present invention or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.
[0010] Figure 1 is a schematic diagram of the arrangement structure of the plasma actuator on the wing surface provided by some embodiments of the present application;
[0011] Figure 2 is a schematic diagram of the arrangement of a first plasma actuator provided in some embodiments of the present application;
[0012] Figure 3 is a schematic diagram of the arrangement of a second plasma actuator provided in some embodiments of the present application;
[0013] Figure 4 is a schematic diagram of an aerodynamic characteristic curve provided by some embodiments of the present application;
[0014] Figure 5 It is a schematic diagram of the verification result curve provided in some embodiments of the present application.
[0015] In the figure: 100-wing, 200-first plasma exciter, 210-upper electrode of leading edge exciter, 220-insulating dielectric layer of leading edge exciter, 230-lower electrode of leading edge exciter, 300-second plasma exciter, 310-upper electrode of trailing edge exciter, 320-insulating dielectric layer of trailing edge exciter, 330-lower electrode of trailing edge exciter, 400-sinusoidal high voltage excitation power supply. DETAILED DESCRIPTION
[0016] The following description provides many different embodiments or examples for implementing different features of the present invention. The components and arrangements described in the following specific examples are only used to simplify the present invention and are only used as examples, not to limit the present invention.
[0017] The embodiment of the present application provides a wing gust mitigation control method based on combined plasma excitation, which uses a plasma actuator to control the airflow field on the surface of the wing 100, thereby controlling its aerodynamic fluctuations, achieving gust load mitigation, and increasing the hovering time of the UAV.
[0018] The specific control steps of the control method provided in the embodiment of the present application include:
[0019] S1: Obtain an aerodynamic characteristic curve of the wing 100. The aerodynamic characteristic curve is a curve of changes in relevant parameters of the wing 100 with the angle of attack of the wing 100 under certain wind speed conditions. Specifically, it may be lift, drag, lift-to-drag ratio, lift coefficient, drag coefficient, etc. In a specific implementation, the aerodynamic characteristic curve may be obtained by CFD (Computational Fluid Dynamics) or wind tunnel testing.
[0020] S2: Determine the first installation position based on the aerodynamic characteristic curve obtained in step S1. The first installation position is located in the upper surface area of the leading edge of the wing 100, and the first plasma actuator 200 is installed at the first installation position. The first plasma actuator 200 is a symmetrical structure, which is mainly used to increase the angle of attack of the wing 100. The upper surface area of the leading edge of the wing 100 is large, and the first plasma actuator 200 can be installed at any position on the upper surface of the leading edge. The embodiment of the present application obtains the first installation position based on the aerodynamic characteristic curve obtained in step S1. The first installation position is the optimal installation position of the first plasma actuator 200. The first plasma actuator 200 is installed at the first installation position, which can improve the effect of the first plasma actuator 200 on the airflow field on the surface of the wing 100.
[0021] The first plasma actuator 200 generates a starting vortex under the action of high voltage electricity and evolves into an induced jet. The starting vortex increases the mixing of the boundary layer and the free flow. The induced jet injects momentum into the boundary layer, improves the ability of the boundary layer to resist the adverse pressure gradient, suppresses the flow separation on the upper surface of the leading edge of the wing 100, and increases the pressure difference between the upper and lower surfaces of the wing 100, thereby increasing lift and reducing drag.
[0022] The speed of the free flow is the wind speed in step S1, and the direction of the free flow can be referred to Figures 1 to 3The solid arrow indicates the direction. The upper surface and the lower surface of the wing 100 are also the upper wing surface and the lower wing surface. In the normal flight of the UAV, the surface of the wing 100 close to the ground is the lower surface, and the surface of the wing 100 close to the ground is the upper surface. The direction of the induced jet generated by the first plasma actuator 200 can be referred to Figure 1 and Figure 2 The dashed arrow indicates the direction. Figure 1 In the figure, the left area in the direction of the illustration is the leading edge of the wing 100, and the right area in the direction of the illustration is the trailing edge of the wing 100.
[0023] S3: Determine the second installation position based on the aerodynamic characteristic curve obtained in step S1. The second installation position is located in the upper surface area of the trailing edge of the wing 100, and the second plasma actuator 300 is installed at the second installation position. The second plasma actuator 300 is an asymmetric structure, which is mainly used to reduce the angle of attack of the wing 100. The upper surface area of the trailing edge of the wing 100 is large, and the second plasma actuator 300 can be installed at any position on the upper surface of the trailing edge. In the embodiment of the present application, the second installation position is obtained based on the aerodynamic characteristic curve obtained in step S1. The second installation position is the optimal installation position of the second plasma actuator 300. The second plasma actuator 300 is installed in the second installation position, which can improve the effect of the second plasma actuator 300 on the airflow field on the surface of the wing 100.
[0024] The direction of the induced jet generated by the discharge of the second plasma actuator 300 is opposite to the direction of the free flow. By hindering the development of the free flow and accelerating and deflecting the streamline of the lower wing surface upward, the pressure difference between the upper and lower surfaces of the wing 100 is reduced, thereby reducing the lift. The direction of the induced jet generated by the second plasma actuator 300 can be referred to Figure 1 and Figure 3 The dashed arrow indicates the direction.
[0025] There is no particular order for step S2 and step S3.
[0026] S4: Obtain the gust signal of the wing tip of the drone wing 100 or the drone nose, and control the operation of the first plasma actuator 200 or the second plasma actuator 300 according to the gust signal. According to the gust signal, the first plasma actuator 200 is started when the angle of attack needs to be increased; the second plasma actuator 300 is started when the angle of attack needs to be reduced. The obtained gust signal is used for feedback adjustment to adjust the flight attitude of the drone so that the drone can maintain a stable flight attitude, thereby increasing the drone's hovering time.
[0027] In the embodiment of the present application, the first plasma actuator 200 and the second plasma actuator 300 generate an induced jet on the surface of the wing 100, which acts on the airflow field on the surface of the wing 100, controls the aerodynamic fluctuations of the wing 100, reduces the gust load, and increases the hovering time of the UAV. In addition, the plasma actuator has little or no effect on the structure and weight of the UAV, or the effect is negligible, and has little or no effect on the flight radius of the UAV. Even after the plasma actuator is installed, the overall structure of the wing 100 is still relatively simple, more convenient for subsequent maintenance, and has a fast response speed.
[0028] In some embodiments of the present application, the principles of the method for confirming the first installation position and the second installation position are the same.
[0029] In step S2, the specific steps of determining the first installation position based on the aerodynamic characteristic curve obtained in step S1 include:
[0030] S21: A plurality of first marking positions are set on the upper surface area of the leading edge of the wing 100, and the first plasma actuator 200 can be installed at any first marking position.
[0031] S22: Install the first plasma actuator 200 at one of the first marked positions, start the first plasma actuator 200, obtain the aerodynamic characteristic curve of the wing 100, and change the first plasma actuator 200 to another marked position until the aerodynamic characteristic curves of the wing 100 corresponding to all the first marked positions are obtained.
[0032] The upper surface area of the leading edge is large, and in the process of determining the first installation position, it is necessary to continuously change the installation position of the first plasma actuator 200 for testing. In addition, in the process of obtaining the aerodynamic characteristic curve, only the installation position is changed, and the other conditions need to remain unchanged.
[0033] S23: After the aerodynamic characteristic curve of the wing 100 is obtained, the aerodynamic characteristic curve obtained in step S22 is compared with the aerodynamic characteristic curve obtained in step S1 under the same wind speed condition.
[0034] After the first plasma actuator 200 is installed, the first plasma actuator 200 will affect the aerodynamic characteristic curve of the wing 100. The aerodynamic characteristic curve after the first plasma actuator 200 is installed, that is, the aerodynamic characteristic curve obtained in step S22, is compared with the aerodynamic characteristic curve without the first plasma actuator 200 installed, that is, the aerodynamic characteristic curve obtained in step S1. Under the same wind speed condition, the curves in the two steps are compared to determine the aerodynamic characteristic curve in step S22 with the largest aerodynamic difference with the aerodynamic characteristic curve in step S1, and the first marked position corresponding to the aerodynamic characteristic curve in step S22 is the first installation position.
[0035] When the first plasma actuator 200 is installed at the first marked position, the aerodynamic difference between the aerodynamic characteristic curve obtained under the same wind speed condition and the aerodynamic characteristic curve obtained in step S1 is the largest, indicating that installing the first plasma actuator 200 at this position has the best effect on the airflow field on the surface of the wing 100. Determining this position as the first installation position can improve the control effect of the first plasma actuator 200 on the airflow field on the surface of the wing 100, thereby enhancing the control effect on gust load relief.
[0036] In the process of comparing the aerodynamic characteristic curves obtained in the two steps, the wind speed conditions corresponding to the two curves must be the same, and then the difference in lift, drag, lift-to-drag ratio, lift coefficient, drag coefficient, etc. between the two curves at the same angle of attack is obtained, the angle of attack is changed, the difference is obtained again, the obtained difference is compared, and the maximum difference is finally confirmed. The curve obtained in step S22 is compared with the curve obtained in step S1 one by one to determine the maximum difference of each group of comparison results, and finally all the maximum differences are compared again to select the largest one, and the first marked position corresponding to the aerodynamic curve corresponding to the final maximum difference is the first installation position.
[0037] In step S3, the specific steps of determining the second installation position based on the aerodynamic characteristic curve obtained in step S1 include:
[0038] S31: setting a plurality of second marking positions on the upper surface area of the trailing edge of the wing 100;
[0039] S32: installing the second plasma actuator 300 at one of the second marking positions, starting the second plasma actuator 300, acquiring the aerodynamic characteristic curve of the wing 100, and changing the second plasma actuator 300 to another second marking position until the aerodynamic characteristic curves of the wing 100 corresponding to all the second marking positions are acquired;
[0040] S33: Compare the aerodynamic characteristic curve obtained in step S32 with the aerodynamic characteristic curve obtained in step S1 under the same wind speed conditions, and determine the aerodynamic characteristic curve in step S32 whose aerodynamic difference with the aerodynamic characteristic curve in step S1 is the largest. The second marking position corresponding to the aerodynamic characteristic curve in step S32 is the second installation position.
[0041] The principle of the method for obtaining the second installation position is similar to that of the method for obtaining the first installation position, and will not be repeated here.
[0042] The aerodynamic curves obtained in step S22, step S32 and step S1 should be the same type of aerodynamic characteristic curves. The aerodynamic characteristic curve can be a curve of the relationship between the lift coefficient and the angle of attack of the UAV. The lift coefficient can be replaced by the drag coefficient, drag, lift, lift-to-drag ratio, etc.
[0043] When the aerodynamic characteristic curve is the relationship curve between the lift coefficient and the angle of attack of the UAV, you can refer to Figure 4 As shown, Figure 4 A schematic diagram showing a comparison of the aerodynamic characteristic curves in step S22, step S32 and step S1 under the same wind speed condition. The aerodynamic characteristic curves are based on the angle of attack. is the horizontal axis, in degrees, with the lift coefficient The square legend represents the aerodynamic characteristic curve obtained in step S1, the triangle legend represents the aerodynamic characteristic curve obtained in step S22, and the circle legend represents the aerodynamic characteristic curve obtained in step S32.
[0044] In step S23, the aerodynamic force difference is the difference between the lift coefficient of the aerodynamic force characteristic curve in step S22 and the lift coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack. Figure 4 It is represented by the letter a.
[0045] In step S33, the aerodynamic force difference is the difference between the lift coefficient of the aerodynamic force characteristic curve in step S32 and the lift coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack. Figure 4 Indicated by the letter b.
[0046] In the case where the aerodynamic force characteristic curve is a curve of the relationship between the drag coefficient and the angle of attack of the UAV, in step S23, the aerodynamic force difference is the difference between the drag coefficient of the aerodynamic force characteristic curve in step S22 and the drag coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack. In step S33, the aerodynamic force difference is the difference between the drag coefficient of the aerodynamic force characteristic curve in step S32 and the drag coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack.
[0047] In step S4, the gust signal obtained is the wind speed of the free flow , according to wind speed , get the change in angle of attack ,exist In the case of , it means that the angle of attack increases, and the angle of attack needs to be reduced. Therefore, the second plasma actuator 300 is controlled to operate to reduce the lift and reduce the angle of attack. In the case of , it indicates that the angle of attack is reduced, and it is necessary to increase the angle of attack. Therefore, the first plasma actuator 200 is controlled to operate, increase lift, reduce drag, and increase the angle of attack.
[0048] , is the angle of attack at the current moment, is the angle of attack at the previous moment, and the difference between the angles of attack at the two moments is the change in angle of attack.
[0049] , ; , ;in, The wind speed at the current moment The component in the normal direction of the drone, The wind speed at the current moment The weight on the axis of the drone, is the wind speed at the previous moment The component in the normal direction of the drone, is the wind speed at the previous moment The weight on the axis of the drone.
[0050] In step S4, obtain After that, the angle of attack is known, and the wind speed condition is known. Based on the aerodynamic characteristic curve obtained in step S1 under the same wind speed condition, the obtained , find The corresponding lift coefficient Get After that, the angle of attack is known, based on the aerodynamic characteristic curve obtained in step S1 under the same wind speed condition, obtain The corresponding lift coefficient , get the change of lift coefficient ,based on The optimal discharge parameters of the first plasma actuator 200 or the second plasma actuator 300 are confirmed.
[0051] Under the influence of gusts, the lift coefficient of the wing 100 will change. The optimal discharge parameters of the plasma actuator are determined according to the lift coefficient. After the first plasma actuator 200 or the second plasma actuator 300 is started, the wing 100 can quickly change the amount of change in the lift coefficient caused by the gust disturbance. , stabilizing the attitude of the wing 100.
[0052] In some embodiments, in step S2, after determining the first installation position, it also includes: S24: adjusting the discharge parameters of the first plasma actuator 200 to obtain the aerodynamic characteristic curve of the wing 100 corresponding to different discharge parameters; in step S3, after determining the second installation position, it also includes: S34: adjusting the discharge parameters of the second plasma actuator 300 to obtain the aerodynamic characteristic curve of the wing 100 corresponding to different discharge parameters.
[0053] exist In the case of and angle of attack Determining the lift coefficient There are multiple aerodynamic characteristic curves, and there may also be multiple lift coefficients. , determine the final lift coefficient, determine the lift coefficient The discharge parameter of the first plasma actuator 200 corresponding to the aerodynamic characteristic curve is used as the optimal discharge parameter of the first plasma actuator 200.
[0054] When the lift coefficient of the wing 100 at time t fluctuates, the first plasma actuator 200 needs to be activated to adjust the lift coefficient of the wing 100 toward The lift coefficient corresponding to the discharge parameter of the first plasma actuator 200 at the next moment is set to , and Less than or greater than , and the difference is The discharge parameters of the first plasma actuator 200 are adjusted so that the lift coefficient of the wing 100 can change rapidly to offset the Fluctuation to The amount of change brought about.
[0055] exist In the case of and angle of attack Determining the lift coefficient ,exist Determine the lift coefficient when The discharge parameter of the second plasma actuator 300 corresponding to the aerodynamic characteristic curve is used as the optimal discharge parameter of the second plasma actuator 300.
[0056] Similarly, when the second plasma actuator 300 is activated, the lift coefficient corresponding to the second plasma actuator 300 can quickly offset the lift coefficient from Fluctuation to The change brought about by this makes it possible to keep the lift coefficient as stable as possible.
[0057] In some embodiments, a sensor is installed at the wing tip of the drone wing 100 or the nose of the drone, and the sensor is used to obtain gust signals and measure wind speed.
[0058] In some embodiments, reference Figure 2 As shown, the first plasma actuator 200 includes an overlapping leading edge actuator upper electrode 210, a leading edge actuator insulating dielectric layer 220, and a leading edge actuator lower electrode 230. The leading edge actuator upper electrode 210 is located on the surface of the wing 100, and is connected to the positive electrode of the sinusoidal high-voltage excitation power supply 400. The leading edge actuator lower electrode 230 is located inside the wing 100, and is connected to the negative electrode of the sinusoidal high-voltage excitation power supply 400 and is grounded. The size of the leading edge actuator upper electrode 210 is smaller than that of the leading edge actuator lower electrode 230.
[0059] The first plasma actuator 200 is a symmetrical structure. The upper electrode 210 of the leading edge actuator generates a starting vortex under the action of high voltage electricity and evolves into an induced jet. The direction of the induced jet can be referred to Figure 2 The dashed arrow direction is shown. Figure 2 The solid arrow shows the direction of the incoming flow velocity. The starting vortex increases the mixing of the boundary layer and the incoming flow, inducing the jet to inject momentum into the boundary layer, improving the ability of the boundary layer to resist the adverse pressure gradient, inhibiting the flow separation on the upper wing surface, and increasing the pressure difference between the upper and lower wing surfaces, thereby increasing lift and reducing drag.
[0060] In some embodiments, reference Figure 3As shown, the second plasma actuator 300 includes an overlapping trailing edge actuator upper electrode 310, a trailing edge actuator insulating dielectric layer 320, and a trailing edge actuator lower electrode 330. The trailing edge actuator upper electrode 310 is located on the surface of the wing 100, and is connected to the positive electrode of the sinusoidal high-voltage excitation power supply 400. The trailing edge actuator lower electrode 330 is located inside the wing 100 and is arranged close to the leading edge of the wing 100 relative to the trailing edge actuator upper electrode 310. The trailing edge actuator lower electrode 330 is connected to the negative electrode of the sinusoidal high-voltage excitation power supply 400 and is grounded.
[0061] The second plasma actuator 300 is an asymmetric structure. The upper electrode of the trailing edge actuator generates an induced jet under the action of high voltage. The direction of the induced jet can be referred to Figure 3 The direction of the dashed arrow in Figure 3 The solid arrow shows the direction of the incoming flow velocity. The induced jet direction is opposite to the incoming flow direction, which reduces the pressure difference between the upper and lower wing surfaces by hindering the development of the incoming flow and guiding the streamlines of the lower wing surface to accelerate and deflect upward, thereby reducing the lift coefficient.
[0062] The control method provided in the embodiment of the present application is used for verification. The two-dimensional GAW-1 high-lift airfoil is used as the research object, and the gust model is simulated using the 1-cos type. The plasma excitation simulation uses the volume force model proposed by Suzen to simulate the plasma excitation effect. This model realizes the coupled solution of the plasma equation and the flow equation by adding the Lorentz force generated by the plasma equation to the volume force source term of the NS equation.
[0063] The simulation results can be found in Figure 5 As shown, Figure 5 In the figure, the horizontal axis s is the dimensionless time, and the vertical axis is the lift coefficient change, is the gust velocity amplitude, in m / s. Figure 5 The triangle legend represents the lift coefficient change when the plasma actuator is not activated to control the wing 100. The circular diagram shows the change in lift coefficient when the plasma actuator is activated to perform gust mitigation control on the wing 100. The quadrilateral legend indicates the gust speed amplitude. With the change of the horizontal axis.
[0064] In the time period of s=0-25, the gust disturbance amplitude increases continuously, which increases the effective angle of attack of the airfoil. The lift coefficient of the airfoil increases continuously, so the second plasma actuator 300 is turned on in this time period to suppress the lift coefficient increment. In the time period of s=25-40, the gust disturbance amplitude makes the effective angle of attack of the airfoil greater than the stall angle of attack, the airfoil stalls, and the lift coefficient of the airfoil drops sharply. Therefore, the first plasma actuator 200 is turned on in this time period to suppress the lift coefficient from dropping sharply. Figure 5 As shown, it can be clearly seen that after turning on the first plasma actuator 200 or the second plasma actuator 300, the fluctuation amplitude of the change in the lift coefficient is small. By adopting the combined plasma excitation control method, the influence of gusts is effectively mitigated, and the fluctuation of the lift coefficient can be reduced by up to 90%.
[0065] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions and improvements made within the spirit and principles of the present invention should be included in the protection scope of the present invention.
Claims
1. A wing gust mitigation control method based on combined plasma excitation, characterized in that: The control steps include: S1: obtaining an aerodynamic characteristic curve of the wing (100); S2: determining a first installation position based on the aerodynamic characteristic curve obtained in step S1, the first installation position being located in the upper surface area of the leading edge of the wing (100), and installing a first plasma actuator (200) at the first installation position, the first plasma actuator (200) being a symmetrical structure, and the first plasma actuator (200) being used to increase the angle of attack of the wing (100); S3: determining a second installation position based on the aerodynamic characteristic curve obtained in step S1, the second installation position being located in the upper surface area of the trailing edge of the wing (100), and installing a second plasma actuator (300) at the second installation position, the second plasma actuator (300) being an asymmetric structure, and the second plasma actuator (300) being used to reduce the angle of attack of the wing (100); S4: obtaining a gust signal from the wing tip of the drone wing (100) or the nose of the drone, and controlling the operation of the first plasma actuator (200) or the second plasma actuator (300) according to the gust signal; In step S4, the gust signal is the wind speed , according to the wind speed , get the angle of attack change ,exist In the case of In the case of, controlling the first plasma actuator (200) to operate; , , ;in, is the angle of attack at the current moment, is the angle of attack at the previous moment, The wind speed at the current moment The component in the normal direction of the drone, The wind speed at the current moment The weight on the axis of the drone, is the wind speed at the previous moment The component in the normal direction of the drone, is the wind speed at the previous moment The weight in the axial direction of the drone; There is no particular order for step S2 and step S3.
2. A wing gust mitigation control method based on combined plasma excitation according to claim 1, characterized in that: In step S2, determining the first installation position based on the aerodynamic characteristic curve obtained in step S1 includes: S21: setting a plurality of first marking positions on the upper surface area of the leading edge of the wing (100); S22: installing a first plasma actuator (200) at one of the first marked positions, starting the first plasma actuator (200), acquiring an aerodynamic characteristic curve of the wing (100), and changing the first plasma actuator (200) to another first marked position until the aerodynamic characteristic curves of the wing (100) corresponding to all the first marked positions are acquired; S23: comparing the aerodynamic force characteristic curve obtained in step S22 with the aerodynamic force characteristic curve obtained in step S1 under the same wind speed condition, determining the aerodynamic force characteristic curve in step S22 having the largest aerodynamic force difference with the aerodynamic force characteristic curve in step S1, and the first marked position corresponding to the aerodynamic force characteristic curve in step S22 is the first installation position; In step S3, determining the second installation position based on the aerodynamic characteristic curve obtained in step S1 includes: S31: setting a plurality of second marking positions on the upper surface area of the trailing edge of the wing (100); S32: installing a second plasma actuator (300) at one of the second marked positions, starting the second plasma actuator (300), acquiring an aerodynamic characteristic curve of the wing (100), and changing the second plasma actuator (300) to another second marked position until the aerodynamic characteristic curves of the wing (100) corresponding to all the second marked positions are acquired; S33: Compare the aerodynamic characteristic curve obtained in step S32 with the aerodynamic characteristic curve obtained in step S1 under the same wind speed conditions, determine the aerodynamic characteristic curve in step S32 whose aerodynamic difference with the aerodynamic characteristic curve in step S1 is the largest, and the second marking position corresponding to the aerodynamic characteristic curve in step S32 is the second installation position.
3. A wing gust mitigation control method based on combined plasma excitation according to claim 2, characterized in that: The aerodynamic characteristic curve is a curve of the relationship between the lift coefficient and the angle of attack of the UAV; In step S23, the aerodynamic force difference is the difference between the lift coefficient of the aerodynamic force characteristic curve in step S22 and the lift coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack; In step S33, the aerodynamic force difference is the difference between the lift coefficient of the aerodynamic force characteristic curve in step S32 and the lift coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack.
4. According to the combined plasma excitation based wing gust mitigation control method of claim 2, the aerodynamic characteristic curve is a curve of the relationship between the drag coefficient and the angle of attack of the UAV; In step S23, the aerodynamic force difference is the difference between the drag coefficient of the aerodynamic force characteristic curve in step S22 and the drag coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack; In step S33, the aerodynamic force difference is the difference between the drag coefficient of the aerodynamic force characteristic curve in step S32 and the drag coefficient of the aerodynamic force characteristic curve in step S1 at the same angle of attack.
5. The wing gust mitigation control method based on combined plasma excitation according to claim 1, characterized in that: In step S4, obtain Then, based on the aerodynamic characteristic curve obtained in step S1 under the same wind speed condition, obtain The corresponding lift coefficient , get Then, based on the aerodynamic characteristic curve obtained in step S1 under the same wind speed condition, obtain The corresponding lift coefficient , get the change of lift coefficient ,based on The optimal discharge parameters of the first plasma actuator (200) or the second plasma actuator (300) are confirmed.
6. A wing gust mitigation control method based on combined plasma excitation according to claim 5, characterized in that: After determining the first installation position in the step S2, the step further includes: S24: adjusting the discharge parameters of the first plasma actuator (200) to obtain aerodynamic characteristic curves of the wing (100) corresponding to different discharge parameters; After the second installation position is determined in step S3, the step further includes: S34: adjusting the discharge parameters of the second plasma actuator (300) to obtain aerodynamic characteristic curves of the wing (100) corresponding to different discharge parameters; exist In the case of and angle of attack Determining the lift coefficient ,exist Determine the lift coefficient when a discharge parameter of the first plasma actuator (200) corresponding to the aerodynamic characteristic curve, and using the discharge parameter as the optimal discharge parameter of the first plasma actuator (200); exist In the case of and angle of attack Determining the lift coefficient ,exist Determine the lift coefficient when The discharge parameter of the second plasma actuator (300) corresponding to the aerodynamic characteristic curve is used as the optimal discharge parameter of the second plasma actuator (300).
7. The wing gust mitigation control method based on combined plasma excitation according to claim 1, characterized in that: A sensor is installed on the wing tip of the drone wing (100) or the nose of the drone, and the sensor is used to obtain a gust signal.
8. The wing gust mitigation control method based on combined plasma excitation according to claim 1, characterized in that: The first plasma actuator (200) comprises a leading edge actuator upper electrode (210), a leading edge actuator insulating medium layer (220) and a leading edge actuator lower electrode (230) which are arranged in an overlapping manner; The leading edge exciter upper electrode (210) is located on the surface of the wing (100), the leading edge exciter upper electrode (210) is connected to the positive electrode of the sinusoidal high-voltage excitation power supply (400), the leading edge exciter lower electrode (230) is located inside the wing (100), the leading edge exciter lower electrode (230) is connected to the negative electrode of the sinusoidal high-voltage excitation power supply (400) and is grounded, and the size of the leading edge exciter upper electrode (210) is smaller than the size of the leading edge exciter lower electrode (230).
9. The wing gust mitigation control method based on combined plasma excitation according to claim 1, characterized in that: The second plasma actuator (300) comprises a trailing edge actuator upper electrode (310), a trailing edge actuator insulating medium layer (320) and a trailing edge actuator lower electrode (330) which are arranged in an overlapping manner; The trailing edge exciter upper electrode (310) is located on the surface of the wing (100), the trailing edge exciter upper electrode (310) is connected to the positive electrode of the sinusoidal high-voltage excitation power supply (400), the trailing edge exciter lower electrode (330) is located inside the wing (100), and is arranged close to the leading edge of the wing (100) relative to the trailing edge exciter upper electrode (310), the trailing edge exciter lower electrode (330) is connected to the negative electrode of the sinusoidal high-voltage excitation power supply (400) and is grounded.