Gust response analysis method considering large deformation of wing

Abstract

The invention relates to the technical field of aerospace, in particular to a gust response analysis method considering large wing deformation, which comprises the following steps of: determining a coordinate system adopted by structural dynamics analysis and unsteady aerodynamic analysis of an aircraft; an aircraft wing is regarded as a nonlinear elastic component, an aircraft body is regarded as a rigid component, and a full-aircraft structure dynamic model of the aircraft is constructed; based on the gust speed borne by the aircraft, an unsteady aerodynamic model is determined, and the unsteady aerodynamic model is used for representing the relation between the motion state of the aircraft and the gust speed; combining the full-aircraft structure dynamic model with the unsteady aerodynamic model, and establishing a full-aircraft dynamic model with the gust effect; solving the full-aircraft dynamical model with the gust effect to obtain the gust response of the aircraft; according to the method, the gust response analysis accuracy of aerodynamic configuration change and structural rigidity change of the wing can be improved.

Description

Technical Field

[0001] The 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 represented by gusts can have an adverse effect on aircraft flight, causing the aircraft to produce large instantaneous accelerations, affecting the riding experience and aircraft structural safety. With the development of aircraft design technology and material technology, high-performance aircraft are more sensitive to gusts than traditional aircraft: on the one hand, people's pursuit of high lift-to-drag ratio has given rise to large aspect ratio designs, and on the other hand, the use of new materials reduces mass and stiffness while ensuring structural strength, so the elastic effect is significant. Under the action of gusts, not only the rigid body motion of the aircraft is affected, but also the elastic vibration of the aircraft structure is aroused, affecting the flight performance of the aircraft, making the aircraft structure prone to fatigue damage and even safety accidents. Therefore, gust response analysis of large aspect ratio flexible aircraft is particularly important for analyzing flight performance and ensuring flight safety.

[0003] Based on the flight dynamics of rigid aircraft, the gust response analysis method has developed a static aeroelastic correction model and a linear rigid-elastic coupling model, which can be used to perform engineering analysis on conventional elastic aircraft. Unlike conventional aircraft, large aspect ratio flexible aircraft produce large deformations under the action of aerodynamic forces. This geometric nonlinear effect causes the natural frequency and aeroelastic characteristics of the aircraft to change significantly, and general linear aeroelastic analysis cannot obtain relatively accurate results. The Chinese invention patent with publication number CN110309579B established a simulation analysis method and system for the gust response of elastic aircraft based on static elastic correction, which only considers the influence of elastic deformation on the aircraft aerodynamic coefficient and cannot analyze the dynamic process of elastic vibration; the Chinese invention patent application with publication number CN109840349B established a fixed-wing aircraft gust response modeling and analysis method based on aerodynamic reduction, but the structural dynamics model still uses the linear modal superposition method, which cannot consider the geometric nonlinear effect of large deformation of the aircraft wing. Summary of the invention

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

[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 wing as a nonlinear elastic component and the aircraft body as a rigid component, and constructing a full-body structural dynamics model of the aircraft;

[0008] Step S3, determining an unsteady aerodynamic force model based on the flight speed of the aircraft and the gust disturbance speed, wherein the unsteady aerodynamic force 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 whole aircraft structural dynamics model with the unsteady aerodynamic model to establish a whole 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 of the full-aircraft dynamics model, wherein 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, where the inertial reference system E is a plane geodetic 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, and 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, κ 1 ,κ 2 ,κ 3 is the moment strain;

[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, the 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, based on the gust flow velocity, determining the relationship between the pitch and heave motion state of the aircraft wing and the effective angle of attack of the wing as a first relationship;

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

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

[0025] Step S3-4, determining 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 motion is calculated using the following expression:

[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, λ 0 is the average induced velocity, calculated as:

[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] Inflow state Combine into a column vector This gives a first-order differential equation:

[0034]

[0035] Where A is the state transfer 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] Among them, D is the influence coefficient matrix between adjacent inflow states, d is the first coefficient vector, e is the second coefficient vector, and the calculation expression is:

[0039]

[0040] Among them, D nm is the value of matrix D in row n and column m, where n and m are matrix indices, and 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 Both the velocity U and the incoming flow velocity U need to be transformed into the local coordinate system by the following expression:

[0043]

[0044] Among them, e 1 =

[100] T , e 2 =

[010] T , e 3 =

[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 system, 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 relationship between the node coordinates, node rotation angles and the center of mass motion of the aircraft.

[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 whole machine dynamics equation and the strain function, and introducing the motion velocity equation and rotation angular velocity equation of the body axis system, the whole machine dynamics equation group including gust effect is 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 vehicle 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 center of mass of the aircraft in the inertial reference frame, Q 1 is the influence matrix of structural acceleration on inflow state, Q 2 is the influence matrix of structural velocity on inflow state, Q 3 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 aerodynamics and structural dynamics equations under large structural deformation, adopts geometrically accurate beam theory to establish nonlinear structural dynamics 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 optimal 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 the purpose of illustrating particular embodiments and are not to be construed as limiting the invention.

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

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

[0061] Figure 3 A 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 purpose, features and advantages of the present invention, the present invention is further described in detail below in conjunction with 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, and therefore, the protection scope 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 As shown, the specific implementation steps are as follows:

[0064] Step S1, determining 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: 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 influence of the earth's curvature and rotation is 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 wing of a high aspect ratio aircraft is subjected to unsteady aerodynamic analysis, such as Figure 3 As shown, the section coordinate system P is defined. 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. When the wing has no sweep angle, the section coordinate system P and the local coordinate system G completely overlap, the first basis vector of the section coordinate system P is perpendicular to the aircraft symmetry plane, and the first basis vector of the local coordinate system G points from the wing root to the wing tip.

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

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

[0070] The wing is regarded as a nonlinear beam. A central reference line is selected along the inside of the wing. The wing is divided 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 determined 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, κ 1 ,κ 2 ,κ 3 is the moment strain;

[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 ,ωB are the velocity and angular velocity of the body axis system, represents the time derivative, represents the second-order derivative with respect to time, the vector F S Indicates the distributed load on the structure, F R Represents the resultant force and moment of external loads.

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

[0078] In this step, the unsteady aerodynamic force is calculated based on the section 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, b is the half chord length, -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)i 2 and i 3 are the coordinate bases of the reference coordinate system i, and the free flow velocity can be defined as -Ui 2 ;i 1 Defined by the right-hand rule, perpendicular to the wing section

[0080] (2)b 2 and b 3 are the coordinate bases of the local coordinate system g fixed to the wing section, b 2 Along the zero lift line and pointing to the leading edge, b 3 Perpendicular to b 2 ; 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)a 2 Along the relative wind speed direction, a 3 Perpendicular to a 2 , is the lift direction.

[0082] Floating and sinking motion h along -i 3 direction, θ is i 2 and b 2 The angle between them, the effective angle of attack α is b 2 and a 2 The angle between.

[0083] The length of the half chord 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 by 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, λ 0 is the average induced velocity, calculated as:

[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 represents the average induced velocity λ 0 N λ Inflow status The weighted average of .

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

[0089]

[0090] Where A is the state transfer 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.

[0091] A and c can be derived according to 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, where n and m are matrix indices, and 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 Both the velocity U and the incoming flow velocity U need to be transformed into the corresponding local coordinate system through the coordinate transformation matrix:

[0098]

[0099] Among them, e 1 =

[100] T , e 2 =

[010] T , e 3 =

[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 system, 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 relationship between the node coordinates, node rotation angles and the center of mass motion of the vehicle.

[0100] The above process gives a time domain solution method for unsteady aerodynamics. The obtained aerodynamic force 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 moment need to be converted to the reference point P by force equivalence and then transformed into 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 the reference coordinate system B.

[0106] Then The resultant force, moment and aerodynamic distribution load vector of each wing section are obtained by conversion. Used for solving the subsequent full-machine dynamics model.

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

[0108] In this step, the whole machine structural dynamics equation and the strain function are combined, and the motion velocity equation and rotation angular velocity equation of the body axis system are introduced to obtain the whole machine dynamics equation group including gust effect:

[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 vehicle 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 center of mass of the aircraft in the inertial reference frame, Q 1 is the influence matrix of structural acceleration on inflow state, Q 2 is the influence matrix of structural velocity on inflow state, Q 3 is the state transfer matrix of the normalized inflow state.

[0112] Through the above expression, the right-hand side of the rigid-elastic coupling dynamic equilibrium equation of the whole aircraft is split into aerodynamic and gravity, in order to reflect the relationship between each force and state variable, so as to facilitate the subsequent state space modeling. If the elastic part (i.e., the strain ε related term) is ignored, the above equation can be degenerated into 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 of the full-aircraft dynamics model, wherein 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 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 whole aircraft dynamics model. By solving the whole aircraft dynamics equations, the time series values ​​of the velocity, acceleration, attitude angle, angular velocity of the attitude angle of the aircraft body and the strain of the aircraft wing changing with 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 that take into account the large deformation of the wing. For the inflow state λ and the body motion β, the above equation is a second-order differential equation, while for the inflow state λ and the body motion β, the above equation 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 specific order, each action or step should be understood to require that such action or step be performed in the specific order shown or in a sequential order, or require that all illustrated actions or steps should be performed to obtain the desired result. Under certain circumstances, multitasking and parallel processing may be advantageous. Similarly, although some specific implementation details are included in the above discussion, these should not be interpreted as limiting the scope of the present disclosure. Certain 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 implementation manner of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by any technician familiar with the technical field within the technical scope disclosed by the present invention should be covered within the protection scope 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 wing as a nonlinear elastic component and the aircraft body as a rigid component, and constructing a full-body structural dynamics model of the aircraft; Step S3, determining an unsteady aerodynamic force model based on the flight speed of the aircraft and the gust disturbance speed, wherein the unsteady aerodynamic force 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 whole aircraft structural dynamics model with the unsteady aerodynamic model to establish a whole aircraft dynamics model including gust effects; 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 of the full-aircraft dynamics model, wherein 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, where the inertial reference system E is a plane geodetic 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, and 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: 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; The whole aircraft structural dynamics model is established based on the aircraft wing, and the expression is as follows: 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, the vector F S Indicates the distributed load on the structure, F R 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, based on the gust flow velocity, determining the relationship between the pitch and heave motion state of the aircraft wing and the effective angle of attack of the wing as a first relationship; Step S3-2, determining the relationship between the pitch and heave motion state of the aircraft wing and the gust inflow state as the second relationship; Step S3-3, determining a conversion relationship for converting the structural motion of the aircraft wing to the pitch and heave motion state of the wing and the incoming flow velocity as a third relationship; Step S3-4, determining 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 motion is calculated using the following expression: in, is the velocity of the profile sinking and floating, a is the proportional coefficient, b is the half chord length, 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 the calculation method is: 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.

6. The gust response analysis method considering large wing deformation according to claim 5 is characterized in that: Step S3-2 specifically includes: Inflow state Combine into a column vector This gives a first-order differential equation: Where A is the state transfer 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; The relationship between A and c is: Among them, D is the influence coefficient matrix between adjacent inflow states, d is the first coefficient vector, e is the second coefficient vector, and the calculation expression is: Among them, D nm is the value of matrix D in row n and column m, where n and m are matrix indices, and 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.

7. The gust response analysis method considering large wing deformation according to claim 6 is characterized in that: Step S3-3 specifically includes: Sinking and floating speed Pitch speed Both the velocity U and the incoming flow velocity U need to be transformed into the local coordinate system by the following expression: 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 system, 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 relationship between the node coordinates, node rotation angles and the center of mass motion of the aircraft.

8. The gust response analysis method considering large wing deformation according to claim 7 is characterized in that: Step S3-4 specifically includes: 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: Among them, ρ ∞ is the atmospheric density.

9. The gust response analysis method considering large wing deformation according to claim 8, characterized in that: Step S4 specifically includes: By combining the whole machine dynamics equation and the strain function, and introducing the motion velocity equation and rotation angular velocity equation of the body axis system, the whole machine dynamics equation group including gust effect is 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 vehicle 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.

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

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