A Gust Response Analysis Method Considering Large Wing Deformation

Abstract

The present invention relates to the field of aerospace technology, and in particular to a gust response analysis method considering large wing deformation, comprising the following steps: determining a coordinate system used for aircraft structural dynamics analysis and unsteady aerodynamic analysis; considering the aircraft wing as a nonlinear elastic component and the aircraft body as a rigid component, and constructing a full-aircraft structural dynamics model of the aircraft; determining an unsteady aerodynamic model based on the gust speed to which the aircraft is subjected, the unsteady aerodynamic model being used to represent the relationship between the aircraft motion state and the gust speed; combining the full-aircraft structural dynamics model with the unsteady aerodynamic model to establish a full-aircraft dynamics model including gust effects; solving the full-aircraft dynamics model including gust effects to obtain the gust response of the aircraft; the present invention can improve the accuracy of gust response analysis of changes in wing aerodynamic shape and structural stiffness.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to a gust response analysis method considering large wing deformation. Background Art

[0002] Atmospheric disturbances, such as gusts, can adversely affect aircraft flight, causing large instantaneous accelerations and impacting the passenger experience and structural safety. With the advancement of aircraft design and material technologies, high-performance aircraft are more sensitive to gusts than traditional aircraft. On the one hand, the pursuit of high lift-to-drag ratios has led to high-aspect-ratio designs. On the other hand, the use of new materials reduces mass and stiffness while ensuring structural strength, resulting in significant elastic effects. Gusts not only affect the rigid-body motion of the aircraft but also induce elastic vibrations in the aircraft structure, affecting its flight performance and making the structure susceptible to fatigue damage and even accidents. Therefore, gust response analysis of high-aspect-ratio flexible aircraft is particularly important for analyzing flight performance and ensuring flight safety.

[0003] Based on the flight dynamics of rigid-body aircraft, gust response analysis methods have developed static aeroelastic correction models and linear rigid-elastic coupling models, which can be used to perform engineering analysis on conventional elastic aircraft. Unlike conventional aircraft, high-aspect-ratio flexible aircraft undergo significant deformation under aerodynamic forces. This geometric nonlinear effect significantly changes the aircraft's natural frequency and aeroelastic characteristics, making conventional linear aeroelastic analysis inaccurate. Chinese invention patent application number CN110309579B establishes a simulation analysis method and system for the gust response of elastic aircraft based on static elastic correction. However, it only considers the effect of elastic deformation on the aircraft's aerodynamic coefficients and cannot analyze the dynamic process of elastic vibration. Chinese invention patent application number CN109840349B establishes a gust response modeling and analysis method for fixed-wing aircraft based on aerodynamic reduction, but the structural dynamics model still uses the linear modal superposition method, which cannot consider the geometric nonlinear effects of large wing deformations. Summary of the Invention

[0004] In view of the above problems, the present invention provides a gust response analysis method that takes into account large deformation of the wing, which solves the technical problem in the existing technology that gust response analysis is difficult to take into account the changes in aerodynamic shape and structural stiffness caused by large deformation of the wing of a flexible aircraft with a large aspect ratio.

[0005] The present invention provides a gust response analysis method considering large wing deformation, comprising the following steps:

[0006] Step S1, determining the coordinate system used for aircraft structural dynamics analysis and unsteady aerodynamic analysis;

[0007] Step S2: Considering the aircraft wings as nonlinear elastic components and the aircraft body as rigid components, a full structural dynamics model of the aircraft is constructed;

[0008] Step S3: determining an unsteady aerodynamic model based on the flight speed of the aircraft and the gust disturbance speed, wherein the unsteady aerodynamic model is used to represent the relationship between the aerodynamic force acting on the aircraft and the motion of the wing profile;

[0009] Step S4: combining the full aircraft structural dynamics model with the unsteady aerodynamic model to establish a full aircraft dynamics model including gust effects;

[0010] Step S5: Inputting the structural parameters, flight speed, and gust disturbance speed of the aircraft into the full-aircraft dynamics model including gust effects, and obtaining the gust response of the aircraft through calculation using the full-aircraft dynamics model. The gust response of the aircraft includes speed, attitude angle, and elastic wing strain.

[0011] Preferably, step S1 specifically includes:

[0012] Determine the inertial reference system E, the body axis system B, and the local coordinate system G. The inertial reference system E is a planar earth reference system. The origin of the body axis system B is located on the fuselage and moves with the movement of the aircraft. The origin of the local coordinate system G is located on the wing reference line and moves with the elastic deformation of the wing. The first basis vector of the local coordinate system G points from the wing root to the wing tip.

[0013] Determine the section coordinate system P. The origin of the section coordinate system P is located at the wing reference line. The first basis vector of the section coordinate system is along the wing span direction, the second basis vector is along the wing section chord direction, and the third basis vector is determined by the right-hand rule.

[0014] Preferably, step S2 specifically includes:

[0015] The wing is regarded as a nonlinear beam and divided into N equal strain units along the wing reference line. The undetermined strain values ​​of all units are variables to be determined, which are N-dimensional column vectors:

[0016]

[0017] Among them, ε τ (τ=1,2,...,N) is the strain vector of the τth unit, τ is the unit index, N is the total number of structural units, γ 11 ,2γ 12 ,2γ 13 is the force strain, κ1, κ2, κ3 are the moment strains;

[0018] The whole aircraft structural dynamics model is established based on the aircraft wing, and the expression is as follows:

[0019]

[0020] Among them, M SS is the structural mass matrix related only to the elastic deformation of the structure, M RR is the structural mass matrix related only to the rigid body motion of the structure, M SR and M RS is the mass matrix of the rigid-elastic coupling structure; C SS is the damping matrix related only to the elastic deformation of the structure, C RR is the damping matrix related only to the rigid body motion of the structure, C SR and C RS is the rigid-elastic coupling damping matrix, K is the stiffness matrix, represents the overall structural strain, where r B is the position of the origin of the body axis system in the inertial system, θ B is the rotation vector of the body axis system relative to the inertial system, v B 、ω B are the velocity and angular velocity of the body axis system, represents the time derivative, Represents the second-order derivative with respect to time, vector F S Indicates the distributed load on the structure, F R Represents the resultant force and moment of external loads.

[0021] Preferably, step S3 specifically includes:

[0022] Step S3-1: Determine, based on the gust wind velocity, the relationship between the pitch and heave motion states of the aircraft wing and the effective angle of attack of the wing as a first relationship;

[0023] Step S3-2: determining a relationship between the pitch and heave motion states of the aircraft wing and the gust inflow state as a second relationship;

[0024] Step S3-3, determining a conversion relationship for converting the structural motion of the aircraft wing into the pitch and heave motion states of the wing and the incoming flow velocity as a third relationship;

[0025] Step S3-4: Determine a calculation method for the total lift and moment of the wing section based on the first relationship, the second relationship, and the third relationship.

[0026] Preferably, step S3-1 specifically includes:

[0027] The effective angle of attack during profile pitching, heaving and floating motions is calculated using the following expressions:

[0028]

[0029] in, is the velocity of the section sinking and floating, a is the proportional coefficient, l is the distance from the origin P to the front edge of the section, is the pitching velocity, λ0 is the average induced velocity, and is calculated as follows:

[0030]

[0031] Among them, p is the cumulative index, N λ is the total number of inflow states, b p is the weighting coefficient of the pth inflow state.

[0032] Preferably, step S3-2 specifically includes:

[0033] Into the flow state Combined into column vectors This gives the first-order differential equation:

[0034]

[0035] Where A is the state transition matrix of the inflow state, is the first-order derivative of the inflow state λ, c is the coefficient vector of the influence of profile motion on the inflow state, is the rate of change of the profile sinking and floating motion velocity, is the rate of change of pitch motion velocity;

[0036] The relationship between A and c is:

[0037]

[0038] Where D is the influence coefficient matrix between adjacent inflow states, d is the first coefficient vector, and e is the second coefficient vector. The calculation expression is:

[0039]

[0040] Among them, D nm is the value of matrix D in row n and column m, n, m are matrix indices, e n is the value of the nth row of vector e, d n is the value of the nth row of vector d, c n is the value of the nth row of vector c.

[0041] Preferably, step S3-3 specifically includes:

[0042] Sinking and floating speed Pitch speed The structural motion and the incoming flow velocity U need to be transformed into the local coordinate system through the following expressions:

[0043]

[0044] Where, e1 =

[100] T , e2=

[010] T , e3=

[001] T , C iB is the coordinate transformation matrix between the local coordinate system i and the reference coordinate system B, U ∞ is the flight speed, w g is the gust speed, represents the rate of change of structural strain, where v B 、ω B are the velocity and angular velocity of the body axis, J Rε and The Jacobian matrix J represents the transformation relationship between the node coordinates, node rotation angle and strain of the structural unit. Rb and The Jacobian matrices represent the transformation relationships between node coordinates, node rotation angles, and the motion of the vehicle's center of mass.

[0045] Preferably, step S3-4 specifically includes:

[0046] Assuming the aerodynamic focus is at 1 / 4 of the chord length, the total lift L and moment M acting on the wing section are 1 / 4 They are:

[0047]

[0048] Among them, ρ ∞ is the atmospheric density.

[0049] Preferably, step S4 specifically includes:

[0050] By combining the full-machine dynamics equations with the strain function and introducing the motion velocity equation and rotation angular velocity equation of the body axis system, the full-machine dynamics equations with gust effects are obtained:

[0051]

[0052] in, Represents the resultant force and moment vector of the aerodynamic loads of each wing section, Represents the aerodynamic distribution load vector of each wing section; is the resultant force and moment vector of the inertial load, is the distributed inertia load; is the aerodynamic force in reference coordinate system B, is the aerodynamic moment in reference coordinate system B, Depend on Calculated;

[0053] is the aircraft attitude angle in quaternion form, is the rate of change of the aircraft attitude angle in quaternion form, Ω B is the body axis angular velocity ω B The corresponding cross product matrix, C EB is the coordinate transformation matrix from the body axis system to the inertial reference system, is the velocity of the vehicle's center of mass in the inertial reference frame, Q1 is the influence matrix of the structural acceleration on the inflow state, Q2 is the influence matrix of the structural velocity on the inflow state, and Q3 is the state transfer matrix of the normalized inflow state.

[0054] Preferably, step S5 specifically includes: using a fourth-order Runge-Kutta method to solve the whole-machine dynamics equations including the gust effect.

[0055] Compared with the prior art, the present invention has at least the following beneficial effects:

[0056] (1) The present invention fully considers the nonlinear characteristics of aerodynamic and structural dynamic equations under large structural deformation, adopts geometrically accurate beam theory to establish nonlinear structural dynamic equations, and combines unsteady aerodynamic modeling and structural-aerodynamic coupling calculation to effectively capture the dynamic response of the wing under gusts. Compared with conventional CFD-CSD loosely coupled iteration, the present invention has higher computational efficiency and better numerical convergence, and has good computational accuracy for the large aspect ratio aircraft object under study.

[0057] (2) The modeling and analysis methods provided by the present invention can predict the sensitivity of an aircraft to atmospheric disturbances and evaluate flight safety during the design phase. They can not only be used for the optimized design of high-performance aircraft, but can also play an important role in experimental testing and control law design. BRIEF DESCRIPTION OF THE DRAWINGS

[0058] The drawings are only for purposes of illustrating particular embodiments and are not to be considered limiting of the invention.

[0059] Figure 1 The present invention provides a flow chart of the gust response analysis method considering large wing deformation.

[0060] Figure 2 Schematic diagram of the high aspect ratio aircraft structure model provided by the present invention

[0061] Figure 3 Schematic diagram of a two-dimensional airfoil in the unsteady aerodynamic solution provided by the present invention. DETAILED DESCRIPTION

[0062] In order to more clearly understand the above-mentioned objects, features and advantages of the present invention, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other. In addition, the present invention can also be implemented in other ways different from those described herein. Therefore, the scope of protection of the present invention is not limited by the specific embodiments disclosed below.

[0063] In order to illustrate the effectiveness of the method proposed by the present invention, the above technical solution of the present invention is described in detail below through a specific embodiment. Figure 1 The specific implementation steps are as follows:

[0064] Step S1: Determine the coordinate system used for aircraft structural dynamics analysis and unsteady aerodynamic analysis.

[0065] The wing of a high aspect ratio aircraft is considered as a nonlinear beam for structural dynamics modeling. Three coordinate systems are defined in the structural dynamics model, such as Figure 2 As shown in the figure: inertial reference system E, body axis system B, and local coordinate system G. The inertial reference system is the geodetic reference system, that is, the effects of the earth's curvature and rotation are ignored; the origin of the body axis system B is located on the fuselage and moves with the movement of the aircraft; the origin of the local coordinate system G is located on the wing reference line and moves with the elastic deformation of the wing.

[0066] The B-space position and direction of the body axis system can be defined as where r B is the position of the origin of the body axis system in the inertial system, θ B is the rotation vector of the body axis system relative to the inertial system, then the velocity and angular velocity of the body coordinate system can be expressed as represents the time derivative.

[0067] The unsteady aerodynamic analysis of the wing of a high aspect ratio aircraft is carried out, such as Figure 3 As shown, a cross-sectional coordinate system P is defined. The origin of cross-sectional coordinate system P is located on the wing reference line. The first basis vector of the cross-sectional coordinate system is along the wing span, the second basis vector is along the wing section chord, and the third basis vector is determined by the right-hand rule. When the wing has no sweep angle, cross-sectional coordinate system P and the local coordinate system G completely coincide. The first basis vector of cross-sectional coordinate system P is perpendicular to the aircraft's symmetry plane, while the first basis vector of local coordinate system G points from the wing root to the wingtip.

[0068] Step S2: Consider the aircraft wings as nonlinear elastic components and the aircraft body as rigid components to construct a full structural dynamics model of the aircraft.

[0069] In this step, the aircraft wing is treated as a nonlinear elastic component and divided into multiple units to establish a dynamic model for the elastic component. The aircraft body is treated as a rigid component and a dynamic model for the rigid component is established. The dynamic models of the elastic and rigid components are combined to form a nonlinear structural dynamic model of the entire aircraft.

[0070] Consider the wing as a nonlinear beam, select a central reference line along the inside of the wing, and divide the wing into N equal strain units along the beam reference line. The strain value in each unit remains unchanged. The undetermined strain values ​​of all units are the variables to be solved in the equation, which is an N-dimensional column vector:

[0071]

[0072]

[0073] Among them, ε τ (τ=1,2,...,N) is the strain vector of the τth unit, τ is the unit index, N is the total number of structural units, γ 11 ,2γ 12 ,2γ 13 is the force strain, κ1, κ2, κ3 are the moment strains;

[0074] The nonlinear structural dynamics equation of the whole machine is established in the following form:

[0075]

[0076] Among them, M SS is the structural mass matrix related only to the elastic deformation of the structure, M RR is the structural mass matrix related only to the rigid body motion of the structure, M SR and M RS is the mass matrix of the rigid-elastic coupling structure; C SS is the damping matrix related only to the elastic deformation of the structure, C RR is the damping matrix related only to the rigid body motion of the structure, C SR and C RS is the rigid-elastic coupling damping matrix, K is the stiffness matrix, represents the overall structural strain, where r B is the position of the origin of the body axis system in the inertial system, θ B is the rotation vector of the body axis system relative to the inertial system, v B 、ω B are the velocity and angular velocity of the body axis system, represents the time derivative, Represents the second-order derivative with respect to time, vector F S Indicates the distributed load on the structure, F RRepresents the resultant force and moment of external loads.

[0077] Step S3: determining an unsteady aerodynamic model based on the gust speed experienced by the aircraft, wherein the unsteady aerodynamic model is used to represent the relationship between the aircraft motion state and the gust speed.

[0078] In this step, the unsteady aerodynamic force is calculated based on the cross-sectional coordinate system P. Figure 3 As shown, point Q is the aerodynamic focus, located at 1 / 4 of the chord length; point P is a point on the beam reference line, (1+a)b from the leading edge, where b is the half chord length and -1<a<1 is the proportional coefficient; point T is located at 3 / 4 of the chord length, where the boundary conditions are met. The coordinate systems in the figure are defined as follows:

[0079] (1) i2 and i3 are the coordinate bases of the reference coordinate system i, and the free stream velocity can be defined as -Ui2; i1 is defined by the right-hand rule and is perpendicular to the wing section.

[0080] (2) b2 and b3 are the coordinate bases of the local coordinate system g fixed to the wing section, b2 is along the zero lift line and points to the leading edge, and b3 is perpendicular to b2. It is worth noting that when the wing has no sweep angle, the coordinate system g and the local structural coordinate system G coincide.

[0081] (3) a2 is along the direction of relative wind speed, and a3 is perpendicular to a2 and is the direction of lift.

[0082] The sinking and floating motion h is along the -i3 direction, θ is the angle between i2 and b2, and the effective angle of attack α is the angle between b2 and a2.

[0083] The half-chord length of the profile is b, the incoming flow velocity is U, and the angle between the incoming flow velocity and the profile chord is θ. The effective angle of attack of the profile during pitching, sinking and floating motion is calculated using the following expression:

[0084]

[0085] in, is the velocity of the section sinking and floating, a is the proportional coefficient, l is the distance from the origin P to the front edge of the section, is the pitching velocity, λ0 is the average induced velocity, and is calculated as follows:

[0086]

[0087] Among them, p is the cumulative index, N λ is the total number of inflow states, b p is the weighting coefficient of the pth inflow state, and the above formula indicates that the average induced velocity λ0 is N λ Inflow state The weighted average of .

[0088] Without considering the state quantity of the control surface deflection changing with time, let the inflow state Combined into column vectors The first-order differential equation can be obtained:

[0089]

[0090] Where A is the state transition matrix of the inflow state, is the first-order derivative of the inflow state λ, c is the coefficient vector of the influence of profile motion on the inflow state, is the rate of change of the profile sinking and floating motion velocity, is the rate of change of pitch velocity.

[0091] A and c can be derived based on the number of states defined by the user

[0092]

[0093] Where D is the influence coefficient matrix between adjacent inflow states, d is the first coefficient vector, and e is the second coefficient vector. The calculation method is:

[0094]

[0095]

[0096] D nm is the value of matrix D in row n and column m, n, m are matrix indices, e n is the value of the nth row of vector e, d n is the value of the nth row of vector d, c n is the value of the nth row of vector c.

[0097] In addition to the above parameters, the speed of sinking and floating Pitch speed The structural motion and the incoming flow velocity U need to be transformed into the corresponding local coordinate system through the coordinate transformation matrix:

[0098]

[0099] Where, e1 =

[100] T , e2=

[010] T , e3=

[001] T , C iB is the coordinate transformation matrix between the local coordinate system i and the reference coordinate system B, U ∞ is the flight speed, w g is the gust speed, represents the rate of change of structural strain, where v B、ω B are the velocity and angular velocity of the body axis, J Rε and The Jacobian matrix J represents the transformation relationship between the node coordinates, node rotation angle and strain of the structural unit. Rb and The Jacobian matrices represent the transformation relationships between node coordinates, node rotation angles, and the motion of the vehicle's center of mass.

[0100] The above process gives a time domain solution method for unsteady aerodynamics. The aerodynamic force obtained acts on the aerodynamic focus. Assuming that the aerodynamic focus Q is located at 1 / 4 of the chord length, the total lift L and moment M acting on the wing section are 1 / 4 They are

[0101]

[0102] Among them, ρ ∞ is the atmospheric density.

[0103] The total lift and torque need to be converted to the reference point P by force equivalence and then transformed through the coordinate transformation matrix C Ba Transform to reference coordinate system B:

[0104]

[0105] in, is the aerodynamic force in reference coordinate system B, is the aerodynamic moment in reference coordinate system B.

[0106] Then Convert the resultant force, moment and aerodynamic load vector of each wing section to obtain Used for solving the subsequent full-machine dynamics model.

[0107] Step S4: combining the whole-machine structural dynamics model with the unsteady aerodynamic model to establish a whole-machine dynamics model including gust effects.

[0108] In this step, the full-machine structural dynamics equations and strain functions are combined, and the motion velocity equations and rotation angular velocity equations of the body axis system are introduced to obtain the full-machine dynamics equations including gust effects:

[0109]

[0110] in, Represents the resultant force and moment vector of the aerodynamic loads of each wing section, Represents the aerodynamic distribution load vector of each wing section; is the resultant force and moment vector of the inertial load, is the distributed inertia load;

[0111] is the aircraft attitude angle in quaternion form, is the rate of change of the aircraft attitude angle in quaternion form, Ω B is the body axis angular velocity ω B The corresponding cross product matrix, C EB is the coordinate transformation matrix from the body axis system to the inertial reference system, is the velocity of the vehicle's center of mass in the inertial reference frame, Q1 is the influence matrix of the structural acceleration on the inflow state, Q2 is the influence matrix of the structural velocity on the inflow state, and Q3 is the state transfer matrix of the normalized inflow state.

[0112] The above expression splits the right-hand side of the rigid-elastic coupled dynamic equilibrium equation into aerodynamic and gravity forces. This serves to capture the relationship between these forces and state variables, facilitating subsequent state-space modeling. If the elastic component (i.e., the strain ε-related term) is neglected, the equation can be reduced to the rigid-body flight dynamics equation.

[0113] Step S5: Inputting the structural parameters, flight speed, and gust disturbance speed of the aircraft into the full-aircraft dynamics model including gust effects, and obtaining the gust response of the aircraft through calculation using the full-aircraft dynamics model. The gust response of the aircraft includes speed, attitude angle, and elastic wing strain.

[0114] In this step, the inherent characteristics of the aircraft structure, such as geometric shape, mass and stiffness, are used as the known quantities of the whole aircraft dynamics model, and the incoming flow velocity U ∞ , gust disturbance w g The initial state of the structure including the initial values ​​of strain, body axis velocity and body axis angular velocity is used as the input variables of the full-aircraft dynamics model. By solving the full-aircraft dynamics equations, the time series values ​​of the velocity, acceleration, attitude angle, angular velocity of the attitude angle and the strain of the aircraft wing over time are finally obtained.

[0115] In some embodiments, the full aircraft dynamics equations are solved in the time domain to obtain the gust response results considering the large deformation of the wing. For the inflow state λ and the body motion β, the above formula is a second-order differential equation, while for the inflow state λ and the body motion β, the above formula is a first-order differential equation.

[0116] In some embodiments, a fourth-order Runge-Kutta method may be used to solve the first-order / second-order mixed differential equation.

[0117] Although the specific embodiments of the present invention have been described in a particular order, it should be understood that such actions or steps are required to be performed in the particular order shown or in a sequential order, or that all illustrated actions or steps are required to be performed to obtain the desired result. Under certain circumstances, multitasking and parallel processing may be advantageous. Similarly, although some specific implementation details have been included in the above discussion, these should not be construed as limiting the scope of the present disclosure. Some features described in the context of a separate embodiment can also be implemented in a single implementation in combination. On the contrary, the various features described in the context of a single implementation can also be implemented in multiple implementations individually or in any suitable sub-combination.

[0118] The above description is only a preferred specific 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 thought of by any technician familiar with this technical field within the technical scope disclosed by the present invention should be covered by the scope of protection of the present invention.

Claims

1. A gust response analysis method considering large wing deformation, characterized in that: The following steps are involved: Step S1, determining the coordinate system used for aircraft structural dynamics analysis and unsteady aerodynamic analysis; Step S2: Considering the aircraft wings as nonlinear elastic components and the aircraft body as rigid components, a full structural dynamics model of the aircraft is constructed; Step S3: determining an unsteady aerodynamic model based on the flight speed of the aircraft and the gust disturbance speed, wherein the unsteady aerodynamic model is used to represent the relationship between the aerodynamic force acting on the aircraft and the motion of the wing profile; Step S4: combining the full aircraft structural dynamics model with the unsteady aerodynamic model to establish a full aircraft dynamics model including gust effects; Step S4 specifically includes: By combining the full-machine dynamics equations with the strain function and introducing the motion velocity equation and rotation angular velocity equation of the body axis system, the full-machine dynamics equations with gust effects are obtained: in, Represents the resultant force and moment vector of the aerodynamic loads of each wing section, Represents the aerodynamic distribution load vector of each wing section; is the resultant force and moment vector of the inertial load, is the distributed inertia load; is the aerodynamic force in reference coordinate system B, is the aerodynamic moment in reference coordinate system B, , Depend on , Calculated; is the aircraft attitude angle in quaternion form, is the rate of change of the aircraft attitude angle in quaternion form, is the body axis angular velocity The corresponding cross product matrix, is the coordinate transformation matrix from the body axis system to the inertial reference system, is the velocity of the center of mass of the aircraft in the inertial reference frame, is the influence matrix of structural acceleration on inflow state, is the influence matrix of structural velocity on inflow state, is the state transfer matrix of the normalized inflow state; Step S5: Inputting the structural parameters, flight speed, and gust disturbance speed of the aircraft into the full-aircraft dynamics model including gust effects, and obtaining the gust response of the aircraft through calculation using the full-aircraft dynamics model. The gust response of the aircraft includes speed, attitude angle, and elastic wing strain.

2. The gust response analysis method considering large wing deformation according to claim 1 is characterized in that: Step S1 specifically includes: Determine the inertial reference system E, the body axis system B, and the local coordinate system G. The inertial reference system E is a planar earth reference system. The origin of the body axis system B is located on the fuselage and moves with the movement of the aircraft. The origin of the local coordinate system G is located on the wing reference line and moves with the elastic deformation of the wing. The first basis vector of the local coordinate system G points from the wing root to the wing tip. Determine the section coordinate system P. The origin of the section coordinate system P is located at the wing reference line. The first basis vector of the section coordinate system is along the wing span direction, the second basis vector is along the wing section chord direction, and the third basis vector is determined by the right-hand rule.

3. The gust response analysis method considering large wing deformation according to claim 2 is characterized in that: Step S2 specifically includes: The wing is regarded as a nonlinear beam and divided into N equal strain units along the wing reference line. The undetermined strain values ​​of all units are variables to be determined, which are N-dimensional column vectors: in, ( =1,2,...,N) is the The strain vector of each element, is the unit index, N is the total number of structural units, For force strain, is the moment strain; The whole aircraft structural dynamics model is established based on the aircraft wing, and the expression is as follows: in, is the structural mass matrix related only to the elastic deformation of the structure, is the structural mass matrix related only to the rigid body motion of the structure, and is the mass matrix of the rigid-elastic coupling structure; is the damping matrix related only to the elastic deformation of the structure, is the damping matrix related only to the rigid body motion of the structure, and is the rigid-elastic coupled damping matrix, is the stiffness matrix, represents the overall structural strain, ,in is the position of the origin of the body axis system in the inertial system, is the rotation vector of the body axis system relative to the inertial system, , 、 are the velocity and angular velocity of the body axis system, represents the time derivative, Represents the second-order derivative with respect to time, vector Indicates the distributed load on the structure, Represents the resultant force and moment of external loads.

4. The gust response analysis method considering large wing deformation according to claim 3 is characterized in that: Step S3 specifically includes: Step S3-1: Determine, based on the gust wind velocity, the relationship between the pitch and heave motion states of the aircraft wing and the effective angle of attack of the wing as a first relationship; Step S3-2: determining a relationship between the pitch and heave motion states of the aircraft wing and the gust inflow state as a second relationship; Step S3-3, determining a conversion relationship for converting the structural motion of the aircraft wing into the pitch and heave motion states of the wing and the incoming flow velocity as a third relationship; Step S3-4: Determine a calculation method for the total lift and moment of the wing section based on the first relationship, the second relationship, and the third relationship.

5. The gust response analysis method considering large wing deformation according to claim 4 is characterized in that: Step S3-1 specifically includes: The effective angle of attack during profile pitching, heaving and floating motions is calculated using the following expressions: in, is the velocity of the profile sinking and floating movement, is the proportionality coefficient, is the half chord length, , is the distance from the origin P to the front edge of the section, is the pitching speed, is the average induced velocity, which is calculated as: in, p is the cumulative index, is the total number of inflow states, For the p The weighting coefficient of each inflow state.

6. The gust response analysis method considering large wing deformation according to claim 5 is characterized in that: Step S3-2 specifically includes: Into the flow state Combined into column vectors , we get the first-order differential equation: in, is the state transition matrix of the inflow state, Inflow state The first derivative of is the coefficient vector of the influence of profile motion on the inflow state, is the rate of change of the profile sinking and floating motion velocity, is the rate of change of pitch motion velocity; and The relational expression is: in, is the influence coefficient matrix between adjacent inflow states, is the first coefficient vector, is the second coefficient vector, and the calculation expression is: in, is a matrix exist OK The value of the column, is the matrix index, is a vector No. The value of the row, is a vector No. The value of the row, is a vector No. The value of the row.

7. The gust response analysis method considering large wing deformation according to claim 6, characterized in that: Step S3-3 specifically includes: Sinking and floating speed , pitch motion speed and incoming flow velocity The following expressions are required to transform the structure motion into the local coordinate system: in, , , , is the local coordinate system With reference coordinate system The coordinate transformation matrix between is the flight speed, is the gust speed, represents the rate of change of structural strain, ,in 、 are the velocity and angular velocity of the body axis system, and The Jacobian matrix represents the conversion relationship between the node coordinates, node rotation angle and strain of the structural element, and The Jacobian matrices represent the transformation relationships between node coordinates, node rotation angles, and the motion of the vehicle's center of mass.

8. The gust response analysis method considering large wing deformation according to claim 7, characterized in that: Step S3-4 specifically includes: Assuming the aerodynamic focus is at 1 / 4 of the chord length, the total lift acting on the wing section is and torque They are: in, is the atmospheric density.

9. The gust response analysis method considering large wing deformation according to claim 8, characterized in that: Step S5 specifically includes: The fourth-order Runge-Kutta method is used to solve the whole aircraft dynamics equations including gust effects.

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