Unsteady aerodynamic condensed equivalent method for wing-wingtip control surface configuration in wind tunnel

Abstract

The present application belongs to the technical field of dry wind tunnel test, and particularly relates to a wing-wing tip control surface configuration dry wind tunnel unsteady aerodynamic force condensation equivalent method. In the process of constructing a reduced order unsteady aerodynamic force model, a virtual interpolation point is introduced, so that the reduced order unsteady aerodynamic force model can be quickly and efficiently constructed under the precondition of meeting the test accuracy, and the model order is low enough. Finally, 2 points are arranged on the main wing surface and the wing tip control surface respectively, so that the main mode equivalent and the unsteady aerodynamic force equivalent of the wing-wing tip control surface configuration can be completed. Compared with the traditional method in which 4 points are arranged on a single wing surface and 8 points are arranged in total, the method proposed by the present application can greatly reduce the difficulty of controlling and applying the dry wind tunnel unsteady excitation force of the wing-wing tip control surface configuration, and improve the reliability of the dry wind tunnel test of the wing-wing tip control surface configuration.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of dry wind tunnel test, and particularly relates to a wing-wing tip control surface configuration dry wind tunnel unsteady aerodynamic force condensation equivalent method. BACKGROUND

[0002] Dry wind tunnel technology is a semi-physical and semi-virtual test method based on feedback control principle, using mechanical concentrated force to simulate continuous distributed aerodynamic force to load on actual structure for ground flutter research. The principle of dry wind tunnel test technology is to measure the vibration response of the structure in real time through sensors, to calculate the simulated aerodynamic force required by a control computer according to the real-time vibration signal, and finally to accurately apply the calculated simulated aerodynamic force to the aircraft structure through the exciter according to the control instruction, so as to realize closed-loop flutter test. Compared with the traditional flutter wind tunnel test technology based on scaled similar model, the dry wind tunnel test technology can carry out flutter test without wind tunnel, and can break through the size limitation of wind tunnel test section, and can carry out test on full-size structure, avoiding the error caused by model scaling in wind tunnel test.

[0003] The core of dry wind tunnel test technology is the perception of continuous motion and the condensation equivalent of continuous distributed unsteady aerodynamic force. Only the continuous motion of the aircraft structure and the equivalent of the uniformly distributed unsteady aerodynamic load to a limited number of points can the vibration signal measurement and equivalent aerodynamic force simulation loading on the aircraft structure be carried out through a limited number of sensors and exciters in actual test. However, most of the existing researches are carried out for single wing surface configuration, and the researches for wing surface configuration with control surfaces (such as wing tip control surfaces, ailerons, etc.) are less. Moreover, at least 4 interpolation points need to be arranged for a single elastic wing surface to realize the equivalent of aerodynamic force.

[0004] The wing with wing tip control surface is a layout form adopted by tailless flying wing layout stealth aircraft, which can solve the problem of aircraft heading stability and control after the tail is cancelled. However, its flutter characteristics are influenced by structural layout form and structural nonlinear factors, which are different from traditional aircraft components and are relatively complex. Dry wind tunnel technology provides an important technical approach to solve the flutter problem of wing-wing tip control surface configuration, but the unsteady aerodynamic force equivalent condensation problem of wing-wing tip control surface configuration needs to be solved urgently to provide technical support for dry wind tunnel test of wing with wing tip control surface. SUMMARY

[0005] In order to solve the above problems, the present application provides a virtual interpolation point-based wing-wing tip control surface configuration unsteady aerodynamic force condensation equivalent method, which realizes interpolation equivalent of main wing surface and wing tip control surface main mode by appropriately selecting virtual interpolation points for main wing surface and wing tip control surface auxiliary interpolation, and finally only needs to arrange 2 interpolation points at the ends of the main wing surface and the wing tip control surface, so as to complete the condensation equivalent of the whole wing-wing tip control surface configuration motion and unsteady aerodynamic force. The effectiveness of the unsteady aerodynamic force condensation method is verified by building a time domain simulation model.

[0006] The wing-wing tip control surface configuration dry wind tunnel unsteady aerodynamic force condensation equivalent method comprises the following steps:

[0007] Step 1: constructing a wing-wing tip control surface configuration aircraft component model.

[0008] Figure 1 In the wing-wing tip control surface model, the wing tip control surface rotation shaft axis is located on the axis of the wing tip control surface rotation shaft, and the wing fixed end boundary extension line is a virtual extension line of the wing fixed end boundary. According to Figure 1 As shown in the wing-wing tip control surface configuration, the wing-wing tip control surface configuration structure finite element grid and aerodynamic surface element grid model are established, and the wing-wing tip control surface configuration finite element mode and unsteady aerodynamic force influence coefficient matrix are calculated based on the finite element model and the aerodynamic model, so as to obtain the distributed unsteady aerodynamic force, and

[0009] (1)

[0010] Wherein, is the uniform unsteady aerodynamic force, is the dynamic pressure. is the structural displacement response, is the unsteady aerodynamic force influence coefficient matrix.

[0011] Step 2: virtual interpolation point arrangement.

[0012] Figure 1 In the wing-wing tip control surface model, the wing tip control surface rotation shaft axis is located on the axis of the wing tip control surface rotation shaft, and the wing fixed end boundary extension line is a virtual extension line of the wing fixed end boundary. According to a b They are arranged on the main wing surface and are actual dry wind tunnel test vibration signal acquisition and excitation force application points, and the wing virtual interpolation points c , the wing virtual interpolation points d , the wing virtual interpolation points e , and the wing virtual interpolation points f are arranged on the wing fixed end boundary and the wing fixed end boundary extension line. The wing virtual interpolation points c , the wing virtual interpolation points​d virtual interpolation points of the wing e virtual interpolation points of the wing f Any number of virtual interpolation points of the wing can be arranged, usually more than 4 to meet the fixed end boundary condition. Virtual interpolation points of the wing c virtual interpolation points of the wing d virtual interpolation points of the wing e virtual interpolation points of the wing f virtual interpolation points of the wing a virtual interpolation points of the wing b The interpolation surface is formed to interpolate the modal vibration mode of the main wing surface. Since the vibration displacement signals at the virtual interpolation points of the wing c virtual interpolation points of the wing d virtual interpolation points of the wing e virtual interpolation points of the wing f are zero, the exciting force does not affect the vibration response on the main wing surface, so in the actual wind tunnel test, the virtual interpolation points of the wing c virtual interpolation points of the wing d virtual interpolation points of the wing e virtual interpolation points of the wing f can be ignored, and the actual vibration signal measurement point and the exciting force signal application point are only the wing interpolation points a and b .

[0013] At the same time, Figure 1 , the interpolation points of the wingtip control surface include: wingtip control surface interpolation points m and wingtip control surface interpolation points n, which are arranged on the wingtip control surface and are the vibration signal collection and exciting force application points of the actual wind tunnel test. Wingtip control surface virtual interpolation point o is located at the intersection of the wingtip control surface rotation axis and the extension line of the wing fixed end boundary, which is used to form an interpolation surface for the wingtip control surface together with the full-actuated wing interpolation points m and n , to interpolate the translation and rotation modal vibration mode of the wingtip control surface. Similarly, the wingtip control surface virtual interpolation point o is located on the wing fixed end boundary and the extension line 8 of the wing fixed end boundary, and its vibration signal and exciting force signal have no effect on the overall vibration response of the wingtip control surface, so in the actual wind tunnel test, the wingtip control surface virtual interpolation point o can be ignored, and only the wingtip control surface interpolation points m and n are collected and excited. The vibration signal is applied.

[0014] Step 3: Virtual interpolation point-based frequency domain condensation unsteady aerodynamic force calculation.

[0015] First, for the main wing surface, the wing virtual interpolation pointsc virtual interpolation points of the wing d virtual interpolation points of the wing e virtual interpolation points of the wing f The virtual interpolation point set is composed of interpolation points of the wing a interpolation points of the wing b The interpolation point set is composed of The point set is constructed Based on the surface spline interpolation method, the aerodynamic force reduction matrix is constructed with the point set as the input, and the interpolation matrix of the main wing surface is obtained has: (2)

[0016] wherein, the virtual interpolation point set corresponding interpolation term, the interpolation point set corresponding interpolation term.

[0017] Since the virtual interpolation points are selected on the fixed boundary, the point set corresponding mode is , the interpolation matrix In corresponding mode is all zero terms, which has no effect on the equation, realizing efficient reduction of aerodynamic force.

[0018] Based on the virtual interpolation point method, the interpolation matrix is constructed for the main wing surface and the wing tip to obtain and , the final interpolation matrix is composed of and , that is:

[0019] (3)

[0020] The reduction of the aerodynamic force influence coefficient matrix is constructed:

[0021] (4)

[0022] In the formula, is the reduced unsteady aerodynamic force influence coefficient matrix. , , is the reduced interpolation matrix constructed based on the virtual interpolation point surface spline interpolation method, is the reduced frequency.

[0023] The reduced aerodynamic force is:

[0024] (5)

[0025] In the formula, To distribute unsteady aerodynamic forces, It is dynamic pressure. This represents the structural displacement response at the reduction point.

[0026] The process is the same for wingtip control surfaces.

[0027] Step 4: Single-region clustering optimization based on genetic algorithm. For the main wing surface, based on the interpolation matrix... The interpolation modes were calculated using a genetic algorithm, with the optimization objective being the optimal fit between the interpolation modes and the finite element modes, while fixing the virtual interpolation points on the wing. c Virtual interpolation points of the wing d Virtual interpolation points of the wing e Virtual interpolation points of the wing f The virtual interpolation point set constituted Position, optimize wing interpolation points a and wing interpolation points b Constructing the interpolation point set The optimized reduced-order unsteady aerodynamic frequency domain model is obtained by determining the position. The genetic algorithm optimization function is: (6)

[0028] In the formula, Elements representing the original mode shapes of aerodynamic nodes. Elements for reducing interpolation mode shapes. The smaller the value, the higher the degree of agreement between the interpolated mode shape and the original mode shape, and the higher the accuracy of the equivalent aerodynamic force.

[0029] For the wingtip control surfaces, the process is the same as above. The optimized set is obtained by combining the optimized points. And calculate the reduced-order aerodynamic forces.

[0030] Step 5: Construction and Validation of Frequency-Domain Unsteady Aerodynamics for the Condensation Model. The original distributed unsteady aerodynamics are replaced with reduced-order unsteady aerodynamics to construct the frequency-domain flutter calculation equations:

[0031] (7)

[0032] (8)

[0033] in, It is the generalized mass matrix of an airfoil with control surfaces. It is the generalized damping matrix of an airfoil with control surfaces. It is the generalized stiffness matrix of an airfoil with control surfaces. It is a reduced-order unsteady aerodynamic force. , , These are displacement, velocity, and acceleration in the generalized coordinate system, respectively.

[0034] The calculation equation is obtained, the flutter speed is compared with the software calculation result, the optimization result is screened, and the precision of the reduced order unsteady aerodynamic force frequency domain model is ensured.

[0035] Step 6: Time-domain unsteady aerodynamic force construction of the polycondensation model. The minimum state method is used to convert the frequency-domain unsteady aerodynamic force into the time-domain unsteady aerodynamic force. The minimum state method formula used is: (9)

[0036] In the formula, s is a Laplace domain variable, is a reference length, is a flow velocity; , , , , is a result matrix generated by the minimum state method. Wherein represents aerodynamic stiffness, represents aerodynamic damping, represents the apparent mass of unsteady aerodynamic force, is a lag root matrix, which represents the lag state of the model due to the lag effect of unstable airflow; is a unit diagonal matrix.

[0037] The equivalent unsteady aerodynamic force in the Laplace domain is:

[0038] (10)

[0039] In the formula, is a reduced point structure velocity response.

[0040] Step 7: Time-domain wind tunnel simulation verification. The above formula is subjected to Laplace inverse transformation to obtain the transfer relationship between the structure response and the aerodynamic force in the time domain, a Simulink time-domain simulation model is built, the time-domain simulation model includes a wing tip control surface-wing configuration state space model S1, a fast aerodynamic force solver based on a virtual point S2, an initial disturbance generator S3, a displacement display S4, a virtual force signal S5, a virtual displacement signal S6, a simulation of coupling of the time-domain reduced order unsteady aerodynamic force and the structure model is realized, a structure response signal is observed, and a flutter speed is predicted.

[0041] Step 8: Derivation of the time-domain aerodynamic force solving matrix of the wing-wing tip control surface configuration polycondensation model. According to the genetic algorithm optimization result, the position coordinates after polycondensation and equivalent are obtained, and according to formula (9), the time-domain unsteady aerodynamic force matrix after polycondensation and equivalent is obtained , , , , , , for subsequent development of dry wind tunnel test.

[0042] Beneficial effects:

[0043] The virtual interpolation point is introduced in the process of constructing the reduced order unsteady aerodynamic force model, the reduced order unsteady aerodynamic force model can be quickly and efficiently constructed under the precondition of meeting the test accuracy, and the model order is low enough. Finally, 2 points are arranged on the main wing surface and the wing tip control surface respectively, so that the main modal equivalence and unsteady aerodynamic force equivalence of the wing-wing tip control surface configuration can be completed. Compared with the traditional method of arranging 4 points on a single wing surface and 8 points in total, the method proposed in the application can greatly reduce the difficulty of controlling and applying the dry wind tunnel unsteady excitation force of the wing-wing tip control surface configuration, and improve the reliability of the dry wind tunnel test of the wing-wing tip control surface configuration. BRIEF DESCRIPTION OF DRAWINGS

[0044] Figure 1 It is an unsteady aerodynamic force reduction principle diagram of a wing-wing tip control surface model;

[0045] Figure 2 It is a model aerodynamic grid division;

[0046] Figure 3 It is a schematic diagram of optimizing the position of the interpolation point;

[0047] Figure 4a It is the first order modal of the original modal;

[0048] Figure 4b It is the second order modal of the original modal;

[0049] Figure 5a It is the first order modal of the traditional reduced order method optimization modal;

[0050] Figure 5b It is the second order modal of the traditional reduced order method optimization modal;

[0051] Figure 6a It is the first order modal of the virtual interpolation point reduced order method optimization modal;

[0052] Figure 6b It is the second order modal of the virtual interpolation point reduced order method optimization modal;

[0053] Figure 7a It is a comparison matrix element (1, 1) of the least square method fitting result;

[0054] Figure 7b It is a comparison matrix element (1, 2) of the least square method fitting result;

[0055] Figure 8 It is a time domain simulation model;

[0056] Figure 9a The result of frequency domain calculation is the vg plot;

[0057] Figure 9b The result of frequency domain calculation is the vf plot;

[0058] Figure 10a for Time-domain calculation results;

[0059] Figure 10b for Time-domain calculation results;

[0060] Figure 10c for Time-domain calculation results.

[0061] In the diagram: 1. Main wing surface; 2. Wingtip control surface; 3. Wingtip control surface pivot; 4. Fixed end boundary of the wing; 5. Wing interpolation point. a 52 wing interpolation points b 53 virtual interpolation points of the wing c 54 virtual interpolation points of the wing d 55 Wing Virtual Interpolation Points e 57 Wing Virtual Interpolation Points f 61 Wingtip control surface interpolation point m, 62 Wingtip control surface interpolation point n, 63 Wingtip control surface virtual interpolation point o, 8 Wing fixed end boundary extension line, 9 Wingtip control surface rotation axis, S1 Wingtip control surface-wing configuration state space model, S2 Fast aerodynamic solver based on virtual points, S3 Initial disturbance generator, S4 Displacement display, S5 Virtual force signal, S6 Virtual displacement signal. Detailed Implementation

[0062] In conjunction with specific embodiments, the implementation steps of this invention are as follows: Step 1: Model Construction. A finite element model of an airfoil with control surfaces is established. The model beam rib material is alloy steel, and the skin material is aluminum alloy. The model root is fixedly constrained, and the modal analysis is calculated using Nastran's SOL 103 solver, as shown in the attached figure. Figure 4a , Figure 4b As shown, the first-order mode is wing bending, and the second-order mode is wingtip twisting, which are used as input files for the ZAERO aeroelastic analysis software. The aerodynamic mesh model of the wing with control surfaces is shown in the attached figure. Figure 2 As shown, aeroelastic analysis was performed using ZAERO software, with the following conditions set: air density... ,Mach number The calculated flutter velocity of the model is The flutter frequency is The influence coefficient matrix of the distributed unsteady aerodynamic forces on the wing with control surfaces is derived. The distributed unsteady aerodynamic forces are then obtained.

[0063] Step 2: Virtual interpolation point arrangement. Select virtual interpolation points on the fixed boundary of the wing, for the main wing surface 1, evenly arrange 4 virtual interpolation points at the wing root, for the wingtip control surface 2, arrange 1 virtual interpolation point at the projection of the wingtip rotation axis on the wing root. And arrange 2 reduction points on the main wing surface 1 and the wingtip control surface 2 respectively, then use the curve spline interpolation method to obtain the interpolation matrix of the main wing surface 1 and the wingtip control surface 2 respectively, and obtain the reduction point position of the main wing surface 1 and the wingtip control surface 2.

[0064] Step 3: Virtual interpolation point based frequency domain condensation unsteady aerodynamic force calculation. According to formula (4), the original distributed aerodynamic force model is reduced, and the reduced aerodynamic influence coefficient matrix is obtained. According to formula (5), the frequency domain unsteady aerodynamic reduced model is preliminarily constructed.

[0065] Step 4: Single region condensation optimization based on genetic algorithm. According to the interpolation matrix, the interpolation mode is calculated, and the genetic algorithm is used to optimize the optimization target of the optimal matching degree of the interpolation mode and the finite element mode, and the reduction point position is optimized as shown in Figure 3 , the optimized interpolation mode is shown in Figure 6a , Figure 6b , and the interpolation matrix and the frequency domain unsteady aerodynamic reduced model obtained in steps 2 and 3 are iterated according to the optimized condensation point position, and the optimized reduced unsteady aerodynamic frequency domain model is obtained. The optimized interpolation mode is compared with Figure 4a , Figure 4b , the original mode of the model and Figure 5a , Figure 5b , the traditional interpolation mode, and the excellent interpolation effect is shown.

[0066] Step 5: Frequency domain unsteady aerodynamic construction and frequency domain verification of the condensed model. The reduced unsteady aerodynamic force is used to replace the original distributed unsteady aerodynamic force to construct the frequency domain flutter equation calculation, and the calculation results are shown in Table 1, and the v-g diagram and v-f diagram are shown in Figure 9a , Figure 9b , to ensure the accuracy of the reduced unsteady aerodynamic frequency domain model.

[0067] Table 1 Comparison of frequency domain flutter calculation results

[0068]

[0069] Step 6: Construction of time domain unsteady aerodynamic force of condensed model. Based on the minimum state method, the frequency domain reduced aerodynamic force is converted into time domain reduced aerodynamic force, and the calculation matrix , , , , , , the fitting matrix accuracy is shown in Figure 7a ,Figure 7b Figure 7 shows the time-domain dry wind tunnel simulation verification.

[0070] Step 7: Time-domain dry wind tunnel simulation verification. The transfer relationship between the structural response and the aerodynamic force in the time domain is constructed, a Simulink time-domain simulation model is built, the simulation of the coupling of the time-domain reduced-order unsteady aerodynamic force and the structural model is realized, the structural response signal is observed, and the flutter speed is predicted. The simulation model is built as shown in Figure 8 Figure 8, and the time-domain simulation of the model is completed. Through the time-domain simulation, it is obtained that the flutter speed is 339.93 m / s, the error is 1.4%, the flutter frequency is 10.82 Hz, and the error is 2.2%. The structural time-domain response curve is as shown in Figure 10a 、 Figure 10b 、 Figure 10c Figure 9. The results show that the reduced-order frequency-domain aerodynamic force constructed by this scheme can meet the accuracy requirements of the ground flutter test, and this scheme can obtain a reduced-order aerodynamic force with low enough order in a single reduction, reducing the error accumulation and simplifying the unsteady aerodynamic force reduction process.

[0071] Step 8: Derivation of the time-domain aerodynamic force solution matrix of the wing-wing tip control surface configuration condensation model. According to the position coordinates obtained by the genetic algorithm optimization, the time-domain unsteady aerodynamic force matrix after condensation and equivalence is calculated according to formula (9) 、 、 、 、 、 , and stored in *.csv format for subsequent dry wind tunnel test.

Claims

1. A method of non-steady aerodynamic scaling and equivalent of wing- wingtip control surface configurations in a dry tunnel, characterized in that, Comprising the following steps: Step 1: Constructing the wing-wingtip control surface configuration aircraft component model; The wing-wing tip control surface model comprises a main wing surface (1), a wing tip control surface (2), a wing tip control surface rotating shaft (3), and a wing fixed end boundary (4); a wing tip control surface rotating shaft axis (9) is located on the axis of the wing tip control surface rotating shaft (3), and a wing fixed end boundary extension line (8) is a virtual extension line of the wing fixed end boundary (4); a wing-wing tip control surface configuration structure finite element grid and an aerodynamic surface element grid model are established, and a wing-wing tip control surface configuration finite element modal and an unsteady aerodynamic force influence coefficient matrix are calculated based on the finite element model and the aerodynamic force model, so that distributed unsteady aerodynamic force is obtained, as shown in formula (1): (1) wherein, is the distributed unsteady aerodynamic force, is the dynamic pressure; is the structural displacement response, is the unsteady aerodynamic influence coefficient matrix; Step 2: Virtual interpolation point arrangement; The interpolation points of the main wing surface (1) include: wing interpolation points a (51) and wing interpolation points b (52), the wing interpolation points a (51) and wing interpolation points b (52) are arranged on the main wing surface (1), wing virtual interpolation points c (53), wing virtual interpolation points d (54), wing virtual interpolation points e (55), wing virtual interpolation points f (56) are arranged on the wing fixed end boundary (4) and the wing fixed end boundary extension line (8); the number of wing virtual interpolation points c (53), wing virtual interpolation points d (54), wing virtual interpolation points e (55), wing virtual interpolation points f (56) is more than 4. The interpolation points of the wingtip control surface (2) include: wingtip control surface interpolation point m (61) and wingtip control surface interpolation point n (62), which are arranged on the wingtip control surface (2), wingtip control surface virtual interpolation point o (63) located at the intersection of the wingtip control surface rotation axis (9) and the wing fixed end boundary extension line (8), wingtip control surface virtual interpolation point o (63) located on the wing fixed end boundary (4) and the wing fixed end boundary extension line 8; Step 3: Virtual interpolation point-based frequency domain condensed unsteady aerodynamic force calculation; For the main wing surface (1), select the wing virtual interpolation points c (53), wing virtual interpolation points d (54), wing virtual interpolation points e (55), wing virtual interpolation points f (56), the virtual interpolation point set is composed of , wing interpolation points a (51) and wing interpolation points b (52) constitute an interpolation point set ; the point set , the interpolation matrix of the main wing surface (1) is constructed based on the surface spline interpolation method with the point set as the input , that is: （2） wherein is a set of virtual interpolation points corresponding interpolation terms, is a set of interpolation points corresponding interpolation terms; Based on the virtual interpolation point method, interpolation matrices are constructed for the main wing surface and the wing tip, respectively, to obtain and The final interpolation matrix is composed of and , namely: （3） The construction of the reduced order aerodynamic influence coefficient matrix is: （4） wherein is the reduced order unsteady aerodynamic influence coefficient matrix; , , is the reduced order interpolation matrix constructed based on the virtual interpolation points spline interpolation method; is the reduced frequency; The reduced order aerodynamic force is: （5） wherein is the distributed unsteady aerodynamic force, is the dynamic pressure; is the reduced point structural displacement response; For the frequency domain condensed unsteady aerodynamic force calculation of the wingtip control surface (2), the process is the same as the calculation method for the main wing surface (1); Step 4: Single-region condensed optimization based on genetic algorithm; Step 5: Frequency domain unsteady aerodynamic force construction and frequency domain verification of condensed model; Step 6: Construction of time domain unsteady aerodynamic force of condensed model; Step 7: Time domain wind tunnel simulation verification; Step 8: Derivation of time domain aerodynamic force solution matrix of wing-wingtip control surface configuration condensed model.

2. The method of wind tunnel unsteady aerodynamic force synthetic equivalent of wing- wingtip control surface configuration according to claim 1, characterized in that, The specific operation of the step 4 is as follows: For the main wing surface (1), according to the interpolation matrix The interpolation modal is calculated, the genetic algorithm is used, the optimal degree of the interpolation modal and the finite element modal is used as the optimization target, the virtual interpolation points of the wing are fixed c (53), the virtual interpolation points of the wing d (54), the virtual interpolation points of the wing e (55), the virtual interpolation points of the wing f (56) constitute a virtual interpolation point set Position, the interpolation points of the wing are optimized a (51) and the interpolation points of the wing b (52) constitute an interpolation point set Position, an optimized reduced order unsteady aerodynamic frequency domain model is obtained; The genetic algorithm optimization function is: （6） wherein, is an element of the original mode of the aerodynamic junction point, is an element of the reduced point interpolation mode, The smaller, the higher the degree of coincidence between the interpolation mode and the original mode, and the higher the accuracy of the equivalent aerodynamic force. For the single region condensation optimization of the wing tip control surface (2) based on genetic algorithm, the process is the same as the main wing surface (1); the optimized point set is obtained after the optimization of the set And calculate the reduced aerodynamic force.

3. The method of wind tunnel unsteady aerodynamic force scaling for wing- wingtip control surface configurations of claim 1, wherein, The specific operation of the step 5 is as follows: The reduced order unsteady aerodynamic force is used to replace the original distributed unsteady aerodynamic force to construct the frequency domain flutter calculation equation: （7） （8） wherein, is the generalized mass matrix of the wing with control surfaces, is the generalized damping matrix of the wing with control surfaces, is the generalized stiffness matrix of the wing with control surfaces, is the reduced order unsteady aerodynamic force, , , are the displacement, velocity and acceleration in the generalized coordinate system, respectively. The calculation equation is obtained to get the flutter speed, compared with the software calculation result, the optimization result is screened to ensure the accuracy of the reduced order unsteady aerodynamic force frequency domain model.

4. The method of wind tunnel unsteady aerodynamic force polynomial condensation equivalence of wing- wingtip control surface configuration as claimed in claim 1, wherein, The specific operation of the step 6 is as follows: The minimum state method is used to convert the frequency domain unsteady aerodynamic force into the time domain unsteady aerodynamic force; the minimum state method formula used is: (9) wherein, wherein s is a Laplace domain variable, is a reference length, is the incoming flow velocity; , , , , is the result matrix generated by the minimum state method, wherein denotes the aerodynamic stiffness, denotes the aerodynamic damping, denotes the apparent mass of the unsteady aerodynamic forces, is a retardation root matrix; is a unit diagonal matrix; The equivalent unsteady aerodynamic force in the Laplace domain is: （10） wherein to reduce the velocity response of the point structure.

5. The method of wind tunnel unsteady aerodynamic force polynomial condensation equivalence of wing- wingtip control surface configuration as claimed in claim 1, wherein, The specific operation of the step 7 is as follows: The Laplace inverse transform of equation (10) is taken to obtain the transfer relationship between the structural response and the aerodynamic force in time domain. A Simulink time domain simulation model is built, which includes a wingtip control surface-wing configuration state space model S1, a fast aerodynamic force solver based on virtual point S2, an initial disturbance generator S3, a displacement display S4, a virtual force signal S5, and a virtual displacement signal S6. The simulation of the coupling between the time domain reduced order unsteady aerodynamic force and the structural model is realized, and the structural response signal is observed to predict the flutter speed.

6. The method of wind tunnel unsteady aerodynamic force polynomial condensation equivalence of wing- wingtip control surface configuration as claimed in claim 1, wherein, The specific operation of the step 8 is as follows: The specific operation of the step 8 is as follows: According to the genetic algorithm optimization result, the position coordinates after the condensation equivalence are obtained, and the time-domain non-steady aerodynamic force matrix after the condensation equivalence is obtained according to formula (9) , , , , , , , which is used for carrying out a dry wind tunnel test.

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