Since the basic equation used is the linearization equation of the potential flow theory. Therefore, it is more suitable in the sub-speed violation speeding stage. It is not suitable for the very important trans-speed phase and it cannot simulate the motion changes of the shock wave as the structure deforms.
4 The static deformation caused by the static load at the angle of attack in actual flight cannot be considered, and the static deformation actually changes the angle of attack of the various parts of the wing. The static deformation caused by static load may even exceed the linear range of the elastic structure. In addition, the traditional flutter calculation method is difficult to incorporate the effects of structural nonlinearities.
Since the above-mentioned porphyrins have been developed from the Euler equations in recent years, the calculation method of aeroelasticity is proposed. 2. However, the non-stick hypothesis of the Euler equation can correctly simulate the separation of the flow. Muwen calculates the unsteady aerodynamic force by solving the equation of viscous flow, and uses the dual-time method to ensure the efficiency and accuracy of the calculation method.
1 Grid generation infinite interpolation theory is first developed according to the solution of Euler's quilt. It is a good algebraic principle. The principle is that the inner boundary, that is, the wing surface, is determined according to the inner boundary and the inner boundary. The object is 1 square in the object. The farther away from the object surface, the thinner the mesh, etc., so that the mesh is gradually and smoothly connected to the outer boundary. The formula can be simply written as the center of Fxky, zt, U. , 5 grid point in a grid point coordinate position received date 209 repair date 3105, fund project aviation science fund 0,53001 funded project author brief introduction Ye Zhengyi 13, male Hubei. Professor Bo, Ph.D. Mainly engaged in computational fluid dynamics and fluid-solid coupling research.
The above-mentioned infinite interpolation theory is used to generate the mesh, and its quality mainly depends on the design of the fusion function. In this paper, the improved fusion function of the text 3 is used to generate the required mesh.
Equation 28 solves the 3 equation, and its integral form is 6=9, for example, 9 from, 6 is the velocity component of the air density 2 direction and the total internal energy per unit volume, which is the normal unit vector of the area division. For the volume domain, ice is the boundary containing the volume domain, which is the flux term, which includes the two parts of the non-viscous item and the viscous term, where the zero knife is the unit equation 6 and the time is the Nigkuzi method of the order precision. Advance, at each time step of advancement, by. 3 Equation calculation For the case of angle of attack, in addition to the dynamic deformation of the wing, the steady aerodynamic force will also deform on the elastic wing. This static deformation has no effect on the structural vibration result in the case of linear structure, but due to the aerodynamic force It has been a nonlinear problem, especially in the case of a large angle of attack wing with a leading edge separation vortex or a transonic speed. The effect of static deformation on aerodynamic forces exists. To account for the effects of static deformation, simply equation 6 The dynamic term can be used to calculate the static deformation. Here, the meaning of subscript 3 static deformation.
Therefore, in this project to calculate the aeroelastic question, the specific implementation steps are: let the wing at a fixed angle of attack at this angle of attack as the basic angle of attack to negotiate the number from the static state! move. When the flow field reaches a stable state, the steady aerodynamic force at the angle of attack is obtained. At this stage, the two-time method is used to solve the unsteady 1 equation, which is to converge the discrete time to obtain the non-constant time step. The regular solution, the detailed process of calculating the unsteady flow of the deformed wing by the dual time method.
The elastic structure does not deform.
The equation of motion using the Lagrangian wing can be written as a matrix form where the general mass moment drop is the structural damping term, especially the generalized stiffness matrix, which is a generalized aerodynamic nine to facilitate the use of the Runge 4 tower method to solve equation 5 In the state of change, ie, U, qY will bring deformation of the flow field and aerodynamic force, then add 1 static deformation amount based on the basic angle of attack. Then solve the aerodynamic force to calculate the static deformation of the new structure. This is repeated until the two adjacent deformations reach the same level. This deformation is considered to be the static deformation of the basic angle of attack.
The static deformation of the basic angle of attack and the corresponding structural elastic force and the aerodynamic force are balanced. If the initial disturbance is added, the elastic structure will enter a dynamic response process, thereby responding to the process. Features can determine the characteristics of aeroelasticity.
By changing the angle of attack and the initial disturbance of the 4 stations, you can get the aeroelasticity.
4 The example is obtained by the factory from the domestic and foreign literatures to find aeroelastic exchange examples with complete data. In order to verify the calculation method, the air flap of the 1 wing is designed to be composed. It is a force, machine, such as dynamic elastic wind tunnel experiment, the wing is a double-diamond wing, 1 gives the plane shape of the wing and the basic cross-sectional shape of the wing root tip, 2 gives =0.55 response at different flight speeds. From the vibration law, the calculated critical velocity of the flutter is about 160,8 and the critical velocity of the flutter in the wind tunnel test is 17,1 and the calculated result is very close to the wind tunnel test result. Unfortunately, in the aeroelastic test, the mechanism was shaken when the chattering occurred, and the variation of the flutter speed of the wing from the subsonic speed to the transonic speed could not be obtained. As a result of the calculation, 3 gives a different speed response process from = 0.85. Obviously, in the transonic flow, the critical speed of the flutter is greatly reduced. This rule is concluded with the foreign transonic wind tunnel. To the point. It can also be seen from the above results that the aerodynamic damping force and the exciting force of the wing do not change drastically with speed, so the response process changes slowly from subcritical to supercritical. In addition, it is noted from the calculation process that the maximum local residual error is 5 as the convergence index, and the transonic speed is propagating to convergence in the per-physical time step. 4 gives a response from 0.95, which clearly reflects the nonlinear characteristics of transonic aeroelasticity. 5 gives the profile pressure distribution of the wing at a certain time from = 0.95. It can be seen that the method fully reflects the ability to capture shock waves.
The first mode, the second mode, and the fifth method, are solved in tandem, and a method of aerodynamic elasticity of the computer wing is developed.
In the flow field calculation, the grid generation uses the improved infinite interpolation theory, and the grid structure uses the 01 form. The equation is spatially discretized by the finite volume method of the central format. The two-time method is used to advance the physical time. The single-sex motion equation is advanced by the order-preferred Runge-Kutta method, and the whole process of the aeroelasticity of the wing is simulated in the time domain. . In addition, this method clearly reflects the reduction of the transonic flutter critical speed. The calculation results in this paper are close to the wind tunnel test results. The other curves are the same as the two references, Ye Zhengyu and Yang Yongnian. The limit of aircraft wing flutter when considering the separation vortex f3 Ye Zhengyi, Wang Gang, Yang Yongnian. Numerical simulation of unsteady flow around a deformed wing 4. The 10th National Conference on Computational Fluid Dynamics, September 2000, Mianyang, 132,137.
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