High-aspect-ratio unmanned aerial vehicle wing flutter suppression method considering wingtip deflection angle influence

Abstract

The invention relates to a high-aspect-ratio unmanned aerial vehicle wing flutter suppression method considering wingtip deflection angle influence, belongs to the technical field of aeroelasticity of aircrafts, and solves the problems that in the prior art, flutter critical speed prediction is not accurate, and passive flutter suppression measures are low in efficiency. And the flutter energy cannot be counteracted in real time. According to the high-aspect-ratio unmanned aerial vehicle wing flutter suppression method, the coupling influence of a structural power subsystem and an aerodynamic subsystem is considered, the critical speed of high-aspect-ratio wing flutter with a wingtip deflection angle can be predicted, and a relational expression between a control surface control signal and flutter control influence factors is obtained; the method has a good effect of suppressing the flutter of the wing of the high-aspect-ratio unmanned aerial vehicle considering the influence of the wingtip deflection angle, and meanwhile, the theoretical method provided by the invention is simple and clear and is convenient for engineering application.

Description

Technical Field

[0001] This invention relates to the field of aeroelasticity technology for aircraft, and specifically to a method for suppressing wing flutter of a large aspect ratio unmanned aerial vehicle (UAV) that takes into account the influence of wingtip deflection angle. Background Technology

[0002] High aspect ratio unmanned aerial vehicles (UAVs) are widely used in fields such as solar-powered UAVs, high-altitude unmanned reconnaissance aircraft, and unmanned meteorological observation platforms due to their long endurance and high lift-to-drag ratio. However, the structural design characteristics of these UAVs—slender, flexible wings and lightweight fuselages—pose significant aeroelastic challenges, with flexible wing flutter being a key safety hazard. Therefore, theoretical methods for predicting and suppressing flutter are needed to address this issue. Regarding the suppression of wing flutter in high aspect ratio UAVs, Zhang Haidong et al. presented a method in 2024 titled "A High Aspect Ratio Wing Flutter Control Method and Control Device." This method utilizes wingtip angular displacement sensors during flight to monitor the wingtip twist angle and twist angular velocity in real time to determine if wing flutter is occurring. Furthermore, steel cables are installed on the wing, and a servo mechanism applies preload to the cables during flight to limit wing twist and thus suppress wing flutter. However, existing technologies lack information on the impact of all-moving wingtip deflection angle on wing flutter parameters, as well as how to control flutter by controlling the control surfaces. Summary of the Invention

[0003] In view of the above problems, the present invention provides a flutter suppression method for large aspect ratio UAVs that takes into account the influence of wingtip deflection angle, which solves the technical problems of inaccurate flutter critical velocity prediction, low efficiency of passive flutter suppression measures, and inability to cancel flutter energy in real time in the prior art.

[0004] This invention provides a method for suppressing wing flutter of large aspect ratio UAVs that takes into account the influence of wingtip deflection angle. The specific steps are as follows:

[0005] The first step is to establish a mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy. Based on the mechanical model of the high aspect ratio airfoil, aerodynamic coupling is introduced to obtain the equations of the structural dynamic subsystem and the structural dynamic transfer matrix A.

[0006] The second step involves obtaining the generalized unsteady buoyancy and pitch aerodynamic forces Q of the wing segment based on a mechanical model of the high aspect ratio airfoil with two degrees of freedom: pitch and buoyancy. h and Q α The time-domain expression is obtained, and then the aerodynamic transfer matrix B is obtained;

[0007] The third step is to establish an aeroelastic model of a wing section with two degrees of freedom (pitch and buoyancy) and one fully movable control surface; and to establish a flutter suppression circuit by combining the structural dynamics transmission matrix A obtained in the first step and the aerodynamic transmission matrix B obtained in the second step.

[0008] The fourth step is to obtain the relationship between multiple model parameters and the control surface deflection angle β in the aeroelastic model of the wing section with one all-moving control surface through simulation calculations.

[0009] Fifth step: Based on the flutter suppression loop in the third step, obtain the flutter suppression closed-loop system matrix;

[0010] The sixth step involves binding the structural dynamics subsystem and aerodynamics subsystem of the first and second steps to the relationship of the control surface deflection angle β of the fourth step, and coordinating with the flutter suppression circuit of the fifth step to establish a closed-loop system of the nonlinear gain scheduling controller.

[0011] Step 7: Substitute the closed-loop system mapping of the nonlinear gain scheduling controller into the first-order derivative of the control surface deflection angle β with time in step 4. Model;

[0012] Step 8: Calculate the first derivative of the rudder deflection angle β from step 7 with time. The model is embedded with a servo actuator to obtain the real-time control surface deflection angle and actively suppress wing flutter.

[0013] Optionally, several model parameters in the aeroelastic model of a wing section with an all-moving control surface include the wing section lift coefficient slope. The ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment Moment of inertia I of the airfoil segment about the center of rigidity E α The distance x between the center of mass G and the center of rigidity E of the airfoil segment α .

[0014] Optionally, the expression for the mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, is as follows:

[0015]

[0016] In the formula, Q h and Q α These represent the generalized unsteady heave and pitch aerodynamic forces of the wing segment, respectively; h represents the displacement of the heave motion of the wing segment. This represents the first derivative of the displacement h of the wing segment with respect to time. α represents the second derivative of the displacement h of the wing segment with respect to time; α represents the pitch angle of the wing segment about the center of rigidity E. This represents the first derivative of the pitch angle α of the wing segment with respect to time; The second derivative of the pitch angle α of the wing segment with respect to time is given by I; m is the mass of the wing segment; αS is the moment of inertia of the airfoil segment about the center of rigidity E; α Let x be the static moment of mass of the airfoil segment. α Let C be the distance between the center of mass G and the center of rigidity E of the airfoil segment, where the center of mass G is located behind the center of rigidity E; h and C α These are the heave and pitch damping coefficients for the wing section, respectively; K h and K α These are the buoyancy and pitch stiffness of the wing section, respectively.

[0017] Optionally, a Laplace transformation is performed on the mechanical model of the high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, to obtain the buoyancy displacement h of the airfoil, the pitch angle α of the airfoil about the rigid center E, and the generalized unsteady buoyancy aerodynamic force Q of the airfoil. h and the generalized unsteady pitch aerodynamic force Q of the wing segment α The Laplace transformation relation of the aeroelastic motion equations is expressed as follows:

[0018]

[0019] Optionally, the Laplace transformation relation of the dynamic elastic motion equation is rewritten as an equation of the structural dynamic subsystem, thereby obtaining the structural dynamic transfer matrix A, expressed as:

[0020]

[0021] In the formula,

[0022]

[0023] A1 = s 2 I α +sC α +K α ;

[0024] B1 = s 2 S α ;

[0025] C1 = -s 2 S α ;

[0026] D1 = s 2 m+sC h +K h ;

[0027] Here, s is a complex frequency variable, which transforms the time-domain differential equation into a frequency-domain algebraic equation for analyzing system stability and dynamic characteristics.

[0028] Optionally, the generalized unsteady heave and pitch aerodynamic forces Q of the wing segment h and Q α The time-domain expression is:

[0029]

[0030] In the formula, ρ is the atmospheric density; V is the airspeed; b is the half-chord length of the wing section; and L is the length of the wing section. The slope of the lift coefficient of the wing section; It is the ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment.

[0031] Optionally, the generalized unsteady heave and pitch aerodynamic forces Q of the wing segment are... h and Q α Performing a Laplace transform on the time-domain expression yields the time-domain Laplace transform relation, which is:

[0032]

[0033] Optionally, the time-domain Laplace transform relation can be rewritten as the equation of the aerodynamic subsystem, thereby obtaining the aerodynamic transfer matrix B, expressed as:

[0034]

[0035] Optionally, the closed-loop system of the nonlinear gain scheduling controller is expressed as:

[0036]

[0037] k1=I α K h ;

[0038]

[0039] In the formula, δh is the minute change in displacement h of the wing segment during its heave and buoyancy motion; δα is the minute change in pitch angle α of the wing segment about the rigid center E. It is the first derivative of the displacement h of the wing segment's heave motion with respect to time. The minute changes; It is the first derivative of the pitch angle α of the wing segment with respect to time. The minute change.

[0040] Optionally, the first derivative of the control surface deflection angle β with time The model, expressed as:

[0041]

[0042] Where f1(.) is the lift coefficient slope function; f2(.) is the center of rigidity position parameter function; f3(.) is the moment of inertia function; and f4(.) is the center of mass-center of rigidity distance function.

[0043] Compared with the prior art, the present invention has at least the following beneficial effects:

[0044] (1) The flutter suppression method for large aspect ratio UAVs of the present invention takes into account the coupling effect of the structural dynamics subsystem and the aerodynamics subsystem, can predict the critical speed of flutter of large aspect ratio wings with wingtip deflection angle, and obtain the relationship between the control surface control signal and each flutter control influencing factor.

[0045] (2) The method for suppressing wing flutter of large aspect ratio UAVs of the present invention has a good effect on suppressing wing flutter of large aspect ratio UAVs considering the influence of wingtip deflection angle. At the same time, the theoretical method proposed in the present invention is simple and clear, and is easy to apply in engineering. Attached Figure Description

[0046] Figure 1 This is a flowchart of the method for suppressing wing flutter of a high aspect ratio UAV that takes into account the influence of wingtip deflection angle in this invention. Detailed Implementation

[0047] To better understand the above-described objectives, features, and advantages of the present invention, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other. Furthermore, the present invention can be implemented in other ways different from those described herein; therefore, the scope of protection of the present invention is not limited to the specific embodiments disclosed below.

[0048] A specific embodiment of the present invention, such as Figure 1 A method for suppressing wing flutter of large aspect ratio UAVs considering the influence of wingtip deflection angle is disclosed. The specific steps are as follows:

[0049] The first step is to establish a mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy. Based on the mechanical model of the high aspect ratio airfoil, aerodynamic coupling is introduced to obtain the aeroelastic motion equation and the structural dynamic transmission matrix A.

[0050] Furthermore, the expression for the mechanical model of the high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, is as follows:

[0051]

[0052] In the formula, Q h and Q α These represent the generalized unsteady heave and pitch aerodynamic forces of the wing segment, respectively; h represents the displacement of the heave motion of the wing segment, with downward being positive; This represents the first derivative of the displacement h of the wing segment with respect to time. α represents the second derivative of the displacement h of the wing segment with respect to time; α represents the pitch angle of the wing segment about the center of rigidity E, with pitch being positive. This represents the first derivative of the pitch angle α of the wing segment with respect to time; The second derivative of the pitch angle α of the wing segment with respect to time is given by I; m is the mass of the wing segment; α S is the moment of inertia of the airfoil segment about the center of rigidity E; α The static moment of mass of the airfoil section is S. α =mx α Let xα be the distance between the center of mass G and the center of rigidity E of the airfoil segment. When the center of mass G is located behind the center of rigidity E, x α C is positive; h and C α These are the heave and pitch damping coefficients for the wing section, respectively; K h and K α These are the buoyancy and pitch stiffness of the wing section, respectively.

[0053] Furthermore, a Laplace transformation is performed on the mechanical model of the high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, to obtain the buoyancy displacement h of the airfoil, the pitch angle α of the airfoil about the rigid center E, and the generalized unsteady buoyancy aerodynamic force Q of the airfoil. h and the generalized unsteady pitch aerodynamic force Q of the wing segment α The Laplace transformation relation of the aeroelastic motion equations is expressed as follows:

[0054]

[0055] Furthermore, the Laplace transformation relation of the dynamic elastic motion equation is rewritten into the equation of the structural dynamic subsystem, thereby obtaining the structural dynamic transfer matrix A, which is expressed as:

[0056]

[0057] In the formula,

[0058]

[0059] A1 = s 2 I α +sC α +K α ;

[0060] B1 = s 2 S α ;

[0061] C1 = -s 2 S α ;

[0062] D1 = s 2 m+sC h +Kh ;

[0063] Here, s is a complex frequency variable, which transforms the time-domain differential equation into a frequency-domain algebraic equation for analyzing system stability and dynamic characteristics.

[0064] The second step involves using a mechanical model of a high aspect ratio airfoil with two degrees of freedom (pitch and buoyancy) to obtain the generalized unsteady buoyancy and pitch aerodynamic forces Q of the airfoil section, based on Grossman's quasi-steady theory. h and Q α The time-domain expression is obtained, and then the aerodynamic transfer matrix B is obtained.

[0065] Furthermore, the generalized unsteady heave and pitch aerodynamic forces Q of the wing segment h and Q α The time-domain expression is:

[0066]

[0067] In the formula, ρ is the atmospheric density; V is the airspeed, i.e., the relative velocity between the incoming flow and the wing section; b is the half-chord length of the wing section; and L is the length of the wing section. The slope of the lift coefficient of the wing section; The ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment, where the center of rigidity is ahead of the center of the wing segment. It is positive.

[0068] Furthermore, the generalized unsteady heave and pitch aerodynamic forces Q of the wing segment are further analyzed. h and Q α Performing a Laplace transform on the time-domain expression yields the time-domain Laplace transform relation, which is:

[0069]

[0070] Furthermore, the time-domain Laplace transform relation is rewritten as the equation of the aerodynamic subsystem, thus obtaining the aerodynamic transfer matrix B, expressed as:

[0071]

[0072] The third step is to establish an aeroelastic model of the wing section with two degrees of freedom (pitch and buoyancy) and one fully movable control surface; and to establish a flutter suppression circuit by combining the structural dynamic transmission matrix A obtained in the first step and the aerodynamic transmission matrix B obtained in the second step.

[0073] Furthermore, the aeroelastic model of the wing segment with one all-moving control surface includes the control surface deflection angle β and the first derivative of the control surface deflection angle β with time. The state variables of the flutter suppression circuit are The control variable is u = [V].

[0074] It is understandable that a fully movable control surface is one in which the entire wing (or tail) is deflected around an axis of rotation at its root as a whole. The entire wing surface is a control surface.

[0075] The fourth step involves obtaining the slope of the lift coefficient of the wing section in an aeroelastic model with one all-moving control surface through simulation calculations. The ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment Moment of inertia I of the airfoil segment about the center of rigidity E α wing segment center of mass G Distance x from the center of rigidity E α The relationship between the control surface deflection angle β and the control surface angle β is as follows:

[0076]

[0077] I α =f3(β)

[0078] x α =f4(β)

[0079] Where f1(.) is the lift coefficient slope function, representing the slope of the wing lift coefficient with respect to the control surface deflection angle β. The effect; f2(.) is the rigid center position parameter function, representing the influence of the control surface deflection angle β on the rigid center position parameter. The functional relationship between the control surface deflection and the moment of inertia of the wing segment about the center of rigidity is established; f3(.) is the moment of inertia function, and the relationship between the moment of inertia of the control surface deflection and the moment of inertia of the wing segment about the center of rigidity is established. α The dependency relationship; f4(.) is the center-of-mass to center-of-rigidity distance function, used to quantify the distance x between the control surface deflection angle β and the center of mass G and center of rigidity E of the wing segment. a .

[0080] Fifth step: Based on the flutter suppression loop from step three, obtain the flutter suppression closed-loop system matrix A. sys The expression is:

[0081]

[0082] Among them, M -1 (.) denotes the inverse of the mass matrix; F(.), G(.), H(.), and P(.) denote the control surface deflection scheduling functions, which are obtained by mapping f1(β), f2(β), f3(β), and f4(β) in the fourth step; K denotes the stiffness matrix; C denotes the damping matrix; and D1 and D2 denote the aerodynamic coupling matrices.

[0083] Solve for the flutter suppression closed-loop system matrix A sys The relationship between airspeed V and control surface deflection angle β is obtained as follows:

[0084] V = f(β).

[0085] Step 6: By linking the structural dynamics subsystem and aerodynamics subsystem from steps 1 and 2, and the relationship of the control surface deflection angle β from step 4, and in conjunction with the flutter suppression loop from step 5, a closed-loop system of the nonlinear gain scheduling controller is established, expressed as:

[0086]

[0087] k1=I α K h ;

[0088]

[0089] In the formula, δh is the minute change in displacement h of the wing segment during its heave and buoyancy motion; δα is the minute change in pitch angle α of the wing segment about the rigid center E. It is the first derivative of the displacement h of the wing segment's heave motion with respect to time. The minute changes; It is the first derivative of the pitch angle α of the wing segment with respect to time. The minute change.

[0090] It is understandable that the state variables of the closed-loop system of the nonlinear gain scheduling controller are... Control signal is

[0091] Step 7: Substitute the closed-loop system mapping of the nonlinear gain scheduling controller into the first-order derivative of the control surface deflection angle β with time in step 4. The model, expressed as:

[0092]

[0093] Step 8: Calculate the first derivative of the rudder deflection angle β from step 7 with time. The model is embedded with a servo actuator to obtain the real-time control surface deflection angle and actively suppress wing flutter.

[0094] The above description is only a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for suppressing wing flutter in high aspect ratio UAVs considering the influence of wingtip deflection angle, characterized in that, The specific steps are as follows: The first step is to establish a mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy. Based on the mechanical model of the high aspect ratio airfoil, aerodynamic coupling is introduced to obtain the equations of the structural dynamic subsystem and the structural dynamic transfer matrix A. The second step involves obtaining the generalized unsteady buoyancy and pitch aerodynamic forces Q of the wing segment based on a mechanical model of the high aspect ratio airfoil with two degrees of freedom: pitch and buoyancy. h and Q α The time-domain expression is obtained, and then the aerodynamic transfer matrix B is obtained; The third step is to establish an aeroelastic model of a wing section with two degrees of freedom (pitch and buoyancy) and one fully movable control surface; and to establish a flutter suppression circuit by combining the structural dynamics transmission matrix A obtained in the first step and the aerodynamic transmission matrix B obtained in the second step. The fourth step is to obtain the relationship between multiple model parameters and the control surface deflection angle β in the aeroelastic model of the wing section with one all-moving control surface through simulation calculations. Fifth step: Based on the flutter suppression loop in the third step, obtain the flutter suppression closed-loop system matrix; The sixth step involves binding the structural dynamics subsystem and aerodynamics subsystem of the first and second steps to the relationship of the control surface deflection angle β of the fourth step, and coordinating with the flutter suppression circuit of the fifth step to establish a closed-loop system of the nonlinear gain scheduling controller. Step 7: Substitute the closed-loop system mapping of the nonlinear gain scheduling controller into the first-order derivative of the control surface deflection angle β with time in step 4. Model; Step 8: Calculate the first derivative of the rudder deflection angle β from step 7 with time. The model is embedded with a servo actuator to obtain the real-time control surface deflection angle and actively suppress wing flutter.

2. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 1, characterized in that, Several model parameters in the aeroelastic model of a wing section with one all-moving control surface include the wing section lift coefficient slope. The ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment Moment of inertia I of the airfoil segment about the center of rigidity E α The distance x between the center of mass G and the center of rigidity E of the airfoil segment α .

3. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 2, characterized in that, The expression for the mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, is as follows: In the formula, Q h and Q α These represent the generalized unsteady heave and pitch aerodynamic forces of the wing segment, respectively; h represents the displacement of the heave motion of the wing segment. This represents the first derivative of the displacement h of the wing segment with respect to time. α represents the second derivative of the displacement h of the wing segment with respect to time; α represents the pitch angle of the wing segment about the center of rigidity E. This represents the first derivative of the pitch angle α of the wing segment with respect to time; The second derivative of the pitch angle α of the wing segment with respect to time is given by I; m is the mass of the wing segment; α S is the moment of inertia of the airfoil segment about the center of rigidity E; α Let C be the static moment of mass of the airfoil segment, and xα be the distance between the center of mass G and the center of rigidity E of the airfoil segment, where the center of mass G is located behind the center of rigidity E; h and C α These are the heave and pitch damping coefficients for the wing section, respectively; K h and K α These are the buoyancy and pitch stiffness of the wing section, respectively.

4. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 3, characterized in that, A Laplace transformation is performed on the mechanical model of a high aspect ratio airfoil with two degrees of freedom, pitch and buoyancy, to obtain the buoyancy displacement h, the pitch angle α about the rigid center E, and the generalized unsteady buoyancy aerodynamic forces Qh and Qh of the airfoil. α The Laplace transformation relation of the aeroelastic motion equations is expressed as follows:

5. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 4, characterized in that, The Laplace transformation relation of the dynamic elasticity equation is rewritten as an equation of the structural dynamic subsystem, thus obtaining the structural dynamic transfer matrix A, which is expressed as: In the formula, A1=s 2 I α +sC α +K α ; B1=s 2 With α ; C1=-s 2 S α ; D1 = s2m + sCh + Kh; Here, s is a complex frequency variable, which transforms the time-domain differential equation into a frequency-domain algebraic equation for analyzing system stability and dynamic characteristics.

6. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 5, characterized in that, Generalized unsteady heave and pitch aerodynamic forces Q of the wing segment h and Q α The time-domain expression is: In the formula, ρ is the atmospheric density; V is the airspeed; b is the half-chord length of the wing section; and L is the length of the wing section. The slope of the lift coefficient of the wing section; It is the ratio of the distance from the center of rigidity to the center of the wing segment to the half-chord length of the wing segment.

7. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 6, characterized in that, Generalized unsteady heave and pitch aerodynamic forces Q of the wing segment h and Q α Performing a Laplace transform on the time-domain expression yields the time-domain Laplace transform relation, which is:

8. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 7, characterized in that, The time-domain Laplace transform relation is rewritten as the equation of the aerodynamic subsystem, thus obtaining the aerodynamic transfer matrix B, expressed as:

9. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 8, characterized in that, The closed-loop system of the nonlinear gain scheduling controller is expressed as follows: k1=I α K h ; In the formula, δh is the minute change in displacement h of the wing segment during its heave and buoyancy motion; δα is the minute change in pitch angle α of the wing segment about the rigid center E. It is the first derivative of the displacement h of the wing segment's heave motion with respect to time. The minute changes; It is the first derivative of the pitch angle α of the wing segment with respect to time. The minute change.

10. The method for suppressing wing flutter of a high aspect ratio UAV according to claim 9, characterized in that, First derivative of the rudder deflection angle β with time The model, expressed as: Where f1(.) is the lift coefficient slope function; f2(.) is the center of rigidity position parameter function; f3(.) is the moment of inertia function; f4(.) is the distance function between the center of mass and the center of rigidity.

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