Blade parameterization method based on flow fixing grid collaborative deformation

Abstract

The invention relates to a blade parameterization method based on flow-fixing grid collaborative deformation, and belongs to the technical field of aero-engine blade shape parameterization. In order to solve the problems of large calculation amount, low calculation efficiency and the like of an existing grid deformation method, the method comprises the steps of constructing an initial fluid domain grid and a solid domain grid which are suitable for a blade model, introducing an O-type domain into the solid domain grid, taking a blade profile point set at different blade height positions on a blade as a blade profile control point, and obtaining a blade profile point set; according to the method, a Hicks-Henne type function and circumferential offset are combined with a radial basis function to obtain the profile tolerance change and the stacking rule change of a blade profile, the offset increment of a control point is evenly diffused to the whole blade surface, and based on the structural grid orthogonality thought, deformation of a grid on the surface of the blade is transmitted to a far end in the orthogonal direction. Compared with a traditional blade parameterization method, the method has the advantages that the two deformation modes of the flow solid domain blade profile tolerance and the stacking rule are achieved at the same time, and the deformation efficiency is improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine blade external shape parameterization, and particularly relates to a blade parameterization method based on fluid-structure grid collaborative deformation. Background Art

[0002] With the rapid development of modern aero-engine technology, the performance of compressors has been significantly improved. As the core components of compressors, blades are developing towards the direction of high pressure ratio, high efficiency, high strength, and long life, making the environment where compressor blades are located increasingly complex, involving multiple disciplines such as aerodynamics, structure, and strength reliability. In traditional multidisciplinary optimization designs, the update of geometric models depends on mesh re-meshing. However, this method has the risk of model update failure, and re-meshing the grid for complex models requires a large time cost, reducing the optimization efficiency. Therefore, it is necessary to seek a mesh parameterization method for multidisciplinary problems.

[0003] Currently, most mesh deformation methods for multidisciplinary problems are based on the free form deform (FFD) technology. However, the FFD method has the problem of non-intuitive deformation. To obtain the target shape, a large number of control points need to be carefully arranged. Especially for special shapes such as blades, the ability to construct fillets and transition surfaces is insufficient, and the FFD mesh deformation method has a large amount of calculation. For three-dimensional problems, the FFD method requires three-layer nesting, and the time complexity is O(n 3 ), and with the increase in the number of control points caused by fine control deformation, the cumulative calculation amount in three dimensions increases geometrically, and the calculation efficiency is low.

[0004] In view of this, it is of great significance to study a blade parameterization method that can realize the deformation of the flow field structure analysis grid and has high calculation efficiency for the design and application of the axial flow blade shape of aero-engines. Summary of the Invention

[0005] Aiming at the defects existing in the above-mentioned prior art, the purpose of the present invention is to provide a blade parameterization method based on fluid-structure grid collaborative deformation, which can simultaneously realize two deformation modes of blade profile change and stacking law change, effectively improving the deformation efficiency, and thus solving the problems existing in the background art.

[0006] The above technical object of the present invention is achieved through the following technical solutions:

[0007] A blade parameterization method based on fluid-structure grid collaborative deformation includes the following steps:

[0008] S1: Construct an initial fluid domain grid and a solid domain grid adapted to the blade model, and introduce an O-type domain into the solid domain grid by means of the topological structure of the fluid domain grid;

[0009] S2: Use the set of profile points at different heights on the blade as the profile control points;

[0010] S3: For the change in the profile tolerance of the blade profile, use the Hicks-Henne function to locally perturb the blade profile to obtain the geometric deviation of the profile control points, and uniformly diffuse this geometric deviation over the entire blade surface;

[0011] S4: For the change in the stacking law of the blade profile, construct a mapping relationship through the circumferential offset of the profile control points combined with the radial basis function interpolation method, and uniformly diffuse the interpolated change over the entire blade surface;

[0012] S5: After obtaining the incremental coordinates of the blade surface nodes from steps S3 - S4, diffuse the incremental coordinates of the blade surface nodes in the direction away from the blade surface along the orthogonal direction of the structured grid. By setting the diffusion attenuation coefficient, the deformation amount during the diffusion process decays to zero at the O-type domain edge.

[0013] As a further preferred solution to the above technical solution: In step S3, the step of diffusing the geometric deviation caused by the change in the profile tolerance of the blade profile over the entire blade surface includes:

[0014] S31: Normalize the distances of the profile control points along the blade circumference from the trailing edge of the blade;

[0015] S32: Determine the superposition positions at different positions of the normalized profile. These superposition positions are in the flow-sensitive regions evenly distributed along the circumferential direction of the profile, and arrange the Hicks-Henne function at the superposition positions to simulate the geometric deviation of the profile control points;

[0016] S33: After obtaining the geometric deviation of the profile control points, perform interpolation based on the distances of the profile control points in the blade height direction, and at the same time set the deformation attenuation coefficient to make the deformation gradually decay along the blade height direction, and uniformly diffuse the offset of the profile control points over the entire blade surface.

[0017] As a further preferred solution: The process of performing distance normalization in step S31 is as follows:

[0018] S311: Select the center of the trailing edge of the blade as the starting point p1, assign zero to the relative position S, and sequentially assign adjacent nodes p2,..p k , to form the set P of profile control points: {p1, p2,..p k , p1};

[0019] S312: Calculate the set D of relative distances between adjacent points in the set P of profile control points: {p1, p2,..p k , p1}: {d1,2 , d 2,3 ,.., d k-1,k , d k,1}, then the set P of airfoil control points: {p1, p2,.. p k , p1} has a relative distance of d from p1 1,2 , d 1,2 + d 2,3 ,.. d 1,2 + d 2,3 +.. + d k,1 ;

[0020] S313: Divide d 1,2 , d 1,2 + d 2,3 ,.. d 1,2 + d 2,3 +.. + d k,1 by the maximum value of the set and normalize it to [0, 1] simultaneously.

[0021] As a further preferred solution: In step S32, arrange the Hicks - Henne type function based on the superposition position. The expression of the blade deformed airfoil that is equal in height to the control airfoil obtained from the Hicks - Henne type function is as follows:

[0022]

[0023] where: x, y, z represent the coordinates of the airfoil control points, s represents the normalized position of the airfoil control points, the subscript base represents the initial grid, γ i represents the coefficient corresponding to the superposition position of the Hicks - Henne type function. By changing the value of γ i , the airfoil with changed profile can be obtained. n x , n y , n z is the unit normal direction vector of this point.

[0024] As a further preferred solution: In step S33, the process of interpolation based on the distances of the airfoil control points in the blade height direction is as follows:

[0025] S331: According to the airfoils at the blade top, blade root, and specified height, divide the blade surface into different regions along the height direction, and use the matrix A p,q to represent the node indices of the blade surface, where p represents the height direction of the blade, and q represents the circumferential direction;

[0026] S332: Use the interpolation method to calculate the blade coordinates within the range of blade height i and blade height i + 1. The interpolation expression is:

[0027]

[0028] Wherein: p and q respectively represent the row number and column number of the node indices on the blade surface, Δ represents the increment of the profile deviation of this node, Lx represents the normalization coefficient of the distance in the height direction between the position of blade height i and the position of blade height i + 1, i corresponds to L = 0 and i + 1 corresponds to L = 1;

[0029] S333: Calculate the blade node offset in the divided area from step S332, and then obtain the blade node coordinates.

[0030] A further preferred solution of the above solution is: In step S4, the process of evenly diffusing the change of the airfoil control points to the entire blade surface is as follows:

[0031] S41: Offset the stacking direction of the airfoil control points at different blade height positions along the circumferential direction;

[0032] S42: Based on the radial basis function interpolation method, construct a mapping relationship according to the node coordinates before and after the circumferential offset of the airfoil control points, and at the same time set a deformation attenuation coefficient to make the deformation gradually attenuate along the blade height direction. According to the mapping relationship, evenly diffuse the node coordinate increment to the entire blade surface. The expression of the node coordinate increment is:

[0033]

[0034] Wherein: φ represents the radial basis function, ||x j -x i || represents the distance norm from the node to be solved to the deformation control point, and w i represents the coefficient corresponding to the deformation control point.

[0035] A further preferred solution of the above solution is: In step S5, after obtaining the node coordinate increment on the blade surface, construct a normal index relationship based on the orthogonality of the structured grid to make the grid increment on the blade surface diffuse in the direction away from the blade surface. The steps include:

[0036] S51: Construct a normal index set of the blade surface grid;

[0037] S52: Based on the constructed normal index set, multiply the offset of the blade surface grid by the normal unit vector to obtain the offset of the volume grid, so that the deformation is transmitted from the blade surface to the far end.

[0038] Based on the above solution, further, in step S51, the construction steps of the normal index set include:

[0039] S511: Take each node of the blade surface grid as the starting point a0 one by one, and take any unit containing a0 as A0;

[0040] S512: Search for the adjacent cell A1 of A0 according to the grid topological relationship, and search for the node closest to a0 in A1 as a1;

[0041] S513: Repeat steps S511 - S512, and an ordered normal index relationship from the blade surface to the distal end can be established.

[0042] Based on the above solution, further, in step S52, to ensure that the deformation is concentrated near the blade near-wall and the fluid domain grid far from the blade area is not affected, a control coefficient P is introduced when calculating the body grid offset, and the control coefficient P satisfies the following constraint conditions:

[0043] For the surface grid of the physical object, P = 1;

[0044] For the distal grid, P = 0;

[0045] For the intermediate grid, P transitions evenly.

[0046] Compared with the prior art, the present invention can produce the following beneficial effects:

[0047] 1. The method of the present invention realizes the collaborative deformation of the flow field and structural analysis grids based on the orthogonal characteristics of the structured grid, and simultaneously satisfies the simulation of the blade profile deviation and the stacking law deviation. It has good deformation effects for both local changes and macroscopic adjustments of the blade. By constructing the orthogonal direction node index relationship, directly assign values to relevant nodes according to the index relationship, and realize the transfer of the increment of the blade surface grid to the surrounding grids in terms of the scale of algebraic operations. Compared with the traditional blade parameterization method, it not only realizes two deformation modes of the blade profile and the stacking law at the same time, but also improves the deformation efficiency.

[0048] 2. The present invention uses the boundary idea of the topological structure of the fluid domain grid, introduces an O-type domain in the solid domain grid, and limits the deformation domain of the blade structure to the O-type domain, reducing the scale of the computational domain and improving the computational efficiency as a whole.

[0049] 3. The present invention adds a deformation attenuation coefficient during the deformation process of both the blade profile and the stacking law, so that the offset increment of the blade profile control points can gradually attenuate along the blade height direction until it attenuates to zero at the blade root and blade tip parts, avoiding reducing the grid quality of the blade root and blade tip parts.

[0050] 4. When the present invention diffuses the blade surface grid increment along the orthogonal direction of the structured grid to the distal end, by setting the diffusion attenuation coefficient, the deformation amount during the transfer process gradually attenuates until it attenuates to zero at the edge of the O-type domain, so that the deformation is concentrated in the blade surface area. Description of the Drawings

[0051] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art.

[0052] Figure 1 It is a flowchart of a blade parameterization method based on fluid-solid grid collaborative deformation of the present invention;

[0053] Figure 2 It is a schematic diagram of the fluid domain and solid domain grids of the axial flow blade of the aero-engine of the present invention, where: a is the fluid domain grid and b is the solid domain grid;

[0054] Figure 3 It is a schematic diagram of the O-type domain in the fluid domain grid and solid domain grid of the present invention, where: a is the O-type domain in the fluid domain grid and b is the O-type domain in the solid domain grid;

[0055] Figure 4 It is a schematic diagram of the airfoil control points at different blade heights of the present invention;

[0056] Figure 5 It is a schematic diagram of the normalized coordinates of the blade surface at different blade heights of the present invention, where: a is the normalized coordinates of the blade surface at 25% blade height, b is the normalized coordinates of the blade surface at 50% blade height, and c is the normalized coordinates of the blade surface at 75% blade height;

[0057] Figure 6 It is a schematic diagram of the change in form error at different blade heights of the present invention, where: a is the change in form error of the blade at 25% blade height, b is the change in form error of the blade at 50% blade height, and c is the change in form error of the blade at 75% blade height;

[0058] Figure 7 It is a deformation diagram of the form error deviation of the airfoil surface grid of the present invention, where: a is the form error deviation deformation of the fluid domain grid and b is the form error deviation deformation of the solid domain grid;

[0059] Figure 8 It is a schematic diagram of the circumferential offset of the airfoil control points of the present invention;

[0060] Figure 9 It is a deformation diagram of the deviation of the stacking law of the airfoil surface grid of the present invention, where: a is the deviation deformation of the stacking law of the fluid domain grid and b is the deviation deformation of the stacking law of the solid domain;

[0061] Figure 10 It is a schematic diagram of the orthogonality direction of the present invention;

[0062] Figure 11 It is a schematic diagram of the deformation transmitted from the blade surface to the distal end of the present invention;

[0063] Figure 12Schematic diagrams before and after the fluid-structure grid collaborative deformation of the present invention, where: a is the diagram of the fluid domain grid before and after deformation, and b is the diagram of the solid domain grid before and after deformation. Detailed implementation manners

[0064] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Apparently, the described embodiments are some, but not all, of the embodiments of the present invention.

[0065] Refer to Figures 1-12 , the present invention discloses a blade parameterization method based on fluid-structure grid collaborative deformation, mainly aiming at the parameterization of axial-flow blades of aeroengines. The overall process of this method is as Figure 1 shown, and includes the following steps:

[0066] S1: Construct an initial fluid domain grid and a solid domain grid adapted to the blade model, and introduce an O-type domain into the solid domain grid by virtue of the topological structure of the fluid domain grid;

[0067] According to the studied blade model, an initial fluid domain grid and a solid domain grid required for grid deformation are established. The topological structure of the fluid domain grid adopts the O4H type. The grid division software used for the fluid domain is AutoGrid5. By virtue of the topological idea of the fluid domain grid, the O-type domain is introduced into the solid domain grid, and the deformation domain is limited to the O-type domain to reduce the scale of the calculation domain and improve the calculation efficiency. The grid division software used for the solid domain is ICEM. The fluid domain grid and the solid domain grid are respectively as Figure 2 shown in a of Figure 2 and b of Figure 3 shown in a of Figure 3 and b of

[0068] S2: Use the airfoil point sets at different blade heights as airfoil control points;

[0069] The blades of aeroengines can be regarded as being composed of countless airfoils according to a certain stacking rule. According to the blade modeling principle, airfoil point sets at different positions in the blade height direction are selected as control points. As Figure 4 shown, the present invention exemplarily selects the airfoil point sets at 25%, 50%, and 75% as control points.

[0070] S3: For the change in the profile of the airfoil, use the Hicks-Henne type function to locally perturb the airfoil to obtain the geometric deviation of the airfoil control points, and uniformly spread the geometric deviation to the entire blade surface;

[0071] First, normalize the distance of the blade profile control points along the entire blade circumference starting from the trailing edge of the blade. The specific normalization process is as follows:

[0072] Select the center of the trailing edge of the blade as the starting point p1, assign zero to the relative position S, and sequentially assign adjacent nodes p2,..p along the blade circumferential direction k , to form a set P of blade profile control points: {p1, p2,..p k , p1}, calculate the set D of relative distances between adjacent points in the set P of blade profile control points: {p1, p2,..p k , p1}: {d 1,2 , d 2,3 ,.., d k-1,k , d k,1}, then further obtain that the relative distance of the set P of blade profile control points: {p1, p2,..p k , p1} from p1 is d 1,2 , d 1,2 +d 2,3 ,..d 1,2 +d 2,3 +..+d k,1 , divide d 1,2 , d 1,2 +d 2,3 ,..d 1,2 +d 2,3 +..+d k,1 by the maximum value of the set and normalize it to [0, 1]; in the present invention, the set of blade profile control points is a ring, with p1 as the origin (assigned zero), the distance from p2 to p1 is d 1,2 , the distance from p3 to p1 is the distance from p3 to p2 plus the distance from p2 to p1, that is, d 1,2 +d 2,3 , and so on, the distance from p1 to p1 is the sum of all relative distances, that is, starting from the trailing edge of the blade, going around a circle and returning to the starting point, and then the distance of each point of the blade profile control points from p1 (the origin) is divided by the maximum value (i.e., the last value in the set D), so as to achieve normalization to [0, 1].

[0073] Then, determine the superposition positions at different positions of the normalized blade profile, and arrange the Hicks - Henne type function at the superposition positions to simulate the geometric deviation of the blade profile control points. The basis for selecting the superposition positions is: (1) evenly distributed along the blade circumferential direction; (2) the superposition positions should be in the flow - sensitive area. Based on the above - mentioned selection principle of the superposition positions, the present invention selects a total of 11 superposition positions (where the head and the tail are the same point), and the superposition positions selected corresponding to the above - mentioned blade heights are as Figure 5 in a, such as Figure 5b, such as Figure 5 as shown in c.

[0074] Arrange the Hicks-Henne type function at the selected superposition position, and the expression of f i (x) is:

[0075]

[0076] Where 0 < x i < 1, α and β are shape factors, x represents the relative position S, i is the three components of space, and k is the number of control points.

[0077] After the above f i (x) is deformed and sorted out, the expression of the deformed airfoil contour equal to the control airfoil height is finally obtained as follows:

[0078]

[0079]

[0080] Where: x, y, z represent the coordinates of the airfoil control points, s represents the normalized position of the airfoil control points, the subscript base represents the initial grid, and γ i represents the coefficient corresponding to the superposition position of the Hicks-Henne type function. By changing the value of γ i , the airfoil contour after the change of the profile tolerance can be obtained. n x , n y , n z is the unit normal direction vector of this point.

[0081] In the parameterization process, by changing a set of γ i values, the airfoil contour after the change of the profile tolerance can be obtained, such as Figure 6 shown in a - Figure 6 shown in c. The value of γ i has no strict regulations, but this value is very important. If the value is not appropriate, phenomena such as the intersection or non-smoothness of the leading and trailing edge lines of the airfoil may occur. Therefore, in the present invention, γ i takes [-0.00005, 0.00005].

[0082] Finally, after obtaining the geometric deviation of the airfoil control points, interpolation is performed based on the distance of the airfoil control points in the airfoil height direction, and the offset of the airfoil control points is evenly diffused to the entire blade surface, such as Figure 7 shown in a, Figure 7 shown in b. In order to prevent the reduction of the quality of some grids at the blade root and blade top, a deformation attenuation coefficient is set to gradually attenuate the deformation along the height direction until it attenuates to zero at the blade root and blade top.

[0083] Specifically, when interpolating based on the distance of airfoil control points, the blade surface is divided into different regions along the height direction according to the airfoils at the blade top, blade root, and specified height. For example, in the present invention, the three lines of 25%, 50%, and 75% are selected. Then, 0 - 25% is one region, 25% - 50% is one region, and so on. If perturbations are made at the blade top and blade root, which are represented by Δ_high and Δ_low respectively, then for the grids near the blade top, they are greatly affected by the blade top offset and slightly affected by the blade root offset. The nodes on the blade surface within the region are affected by the highest offset Δ_high and the lowest offset Δ_low in this region, and will be normalized to [0, 1] along the blade top and blade root. Then, the blade offset within this region is expressed as Δ_low*L + Δ_high*(1 - L); since the blade surface is a quadrilateral grid, the blade surface node indices are written in matrix form A p,q , where p represents the number of rows of the matrix, indicating the height direction of the blade, and q represents the number of columns of the matrix, indicating the circumferential direction of the blade; assuming that the control airfoil only affects the adjacent region, for the blade coordinates within the range of the control airfoil at blade height i and the control airfoil at blade height i + 1, the interpolation expression is:

[0084]

[0085] where: p and q respectively represent the number of rows and columns of the blade surface node indices, Δ represents the increment of the profile deviation of this node, and L i represents the distance normalization coefficient along the height direction between the control airfoil at blade height i and the control airfoil at blade height i + 1, with i corresponding to L = 0 and i + 1 corresponding to L = 1.

[0086] Then, the blade offset within the divided region of the blade surface is calculated from the obtained blade coordinates.

[0087] S4: For the change in the stacking law of the airfoil, a mapping relationship is constructed through the circumferential offset of the airfoil control points combined with the radial basis function interpolation method, and the change interpolation is evenly diffused to the entire blade surface;

[0088] Assume that when the blade changes the stacking law, it only changes circumferentially and does not make axial adjustments. The circumferential offsets are made to the stacking directions of the airfoil control points at different blade height positions. The results are as shown in Figure 8 a - Figure 8 c shown in the figure. According to the node coordinates before and after the circumferential offset of the airfoil control points, a mapping relationship is constructed through the radial basis function interpolation method. According to the mapping relationship, the node coordinate increments are evenly diffused to the entire blade surface, as shown in Figure 9 a, Figure 9 b shown in the figure. The expression for the node coordinate increments is:

[0089]

[0090] Among them: φ represents the radial basis function, and the Gaussian function is selected in the present invention. ||x j - x i || represents the distance norm from the node to be solved to the deformation control point, and w i represents the coefficient corresponding to the deformation control point.

[0091] To prevent the reduction of the quality of some grids at the blade root and the blade tip, by setting the deformation attenuation coefficient, the deformation gradually attenuates along the height direction until it attenuates to zero at the blade root and the blade tip.

[0092] S5: After obtaining the coordinate increment of the blade surface nodes in the above steps, it is necessary to transfer its influence into the volume mesh. This process is realized by constructing the normal index relationship based on the orthogonality of the structured mesh, so as to diffuse the coordinate increment of the blade surface nodes along the orthogonal direction of the structured mesh to the far end (away from the blade surface). By setting the diffusion attenuation coefficient, the deformation gradually attenuates during the transfer process and attenuates to zero at the edge of the O-shaped domain.

[0093] In this step, the process of constructing the normal index relationship based on the orthogonality of the structured mesh is as follows:

[0094] Construct the normal index set of the blade surface mesh: (a) Take each point of the blade surface mesh as the starting point a0; (a) Take any cell containing a0 as A0, and according to the mesh topology relationship, search for the adjacent cell A1 of A0, and search for the node closest to a0 in A1 as a1; (c) Repeat the above steps (a) and (b), and an ordered normal index relationship from the blade surface to the far end can be established, as Figure 10 shown. In this process, when the mesh orthogonality is not ideal, relying only on distance judgment in (b) is not sufficient, and additional judgment is needed. The principle of the additional judgment is: for any quadrilateral, the sum of its two intersecting sides is always greater than the sum of the distances of the two pairs of opposite sides. Based on this, combined with the above process, the correct orthogonal index node a1 of the distance from a0 can be judged.

[0095] Based on the constructed normal index set, the offset of the volume mesh is obtained by multiplying the offset of the blade surface mesh by the normal unit vector, so that the deformation is transferred from the blade surface to the far end. To ensure that the deformation is concentrated near the blade wall and the fluid domain mesh far from the blade area is not affected, a control coefficient P is introduced when calculating the volume mesh offset. The control coefficient P satisfies the following constraint conditions: when it is the surface mesh, P = 1; when it is the far-end mesh, P = 0; for the intermediate mesh, P transitions evenly. The result of the deformation transferred from the blade surface to the far end is as Figure 11 shown, combined with Figure 12 a in Figure 12The results before and after the fluid-structure grid co-deformation of b can show that the method proposed by the present invention has very good consistency in fluid-structure grid co-deformation.

[0096] It should be noted that the "volume grid", "initial grid", and "structural grid" described in the present invention all include the fluid domain grid and the solid domain grid. The "volume grid" is to distinguish from the surface grid. The surface grid is the grid on the surface of the blade, and the rest are volume grids.

[0097] The above shows and describes the basic principles, main features, and advantages of the present invention. Those skilled in the art should understand that the present invention is not limited by the above embodiments. What is described in the above embodiments and the specification only illustrates the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention will have various changes and improvements, and these changes and improvements all fall within the scope of the present invention claimed. The scope of protection claimed by the present invention is defined by the appended claims and their equivalents.

Claims

1. A blade parameterization method based on fluid-solid grid collaborative deformation, characterized in that, It includes the following steps: S1: Construct an initial fluid domain mesh and a solid domain mesh adapted to the blade model, and introduce an O-type domain into the solid domain mesh by virtue of the topological structure of the fluid domain mesh; S2: Use the airfoil point sets at different blade heights as airfoil control points; S3: For the change in the profile of the airfoil, use the Hicks-Henne function to locally perturb the airfoil to obtain the geometric deviation of the airfoil control points, and uniformly diffuse this geometric deviation over the entire blade surface; S4: For the change in the stacking law of the airfoil, construct a mapping relationship through the circumferential offset of the airfoil control points combined with the radial basis function interpolation method, and uniformly diffuse the interpolated change over the entire blade surface; S5: After obtaining the blade surface node coordinate increments from steps S3 - S4, diffuse the blade surface node coordinate increments in the direction away from the blade surface along the orthogonal direction of the structured mesh. By setting a diffusion attenuation coefficient, the deformation amount during the diffusion process decays to zero at the edge of the O-type domain.

2. The parametric method for a blade based on fluid-solid grid collaborative deformation according to claim 1, wherein In step S3, the steps of diffusing the geometric deviation caused by the change in the profile of the airfoil over the entire blade surface include: S31: Normalize the distances of the airfoil control points along the blade circumference from the trailing edge of the blade; S32: Determine the superposition positions at different positions of the normalized airfoil. These superposition positions are in the flow-sensitive regions evenly distributed along the airfoil circumferential direction, and arrange the Hicks-Henne function at the superposition positions to simulate the geometric deviation of the airfoil control points using the Hicks-Henne function; S33: After obtaining the geometric deviation of the airfoil control points, perform interpolation based on the distances of the airfoil control points in the blade height direction, and at the same time set a deformation attenuation coefficient to make the deformation gradually decay along the blade height direction, and uniformly diffuse the offset of the airfoil control points over the entire blade surface.

3. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 2, characterized in that, The process of performing distance normalization in step S31 is as follows: S311: Select the center of the trailing edge of the blade as the starting point p1, assign the relative position S to zero, and assign the adjacent nodes p2, ..p in order along the blade annular direction. k , forming a set of blade control points P: {p1,p2,..p k ,p1}; S312: Calculate the set P of airfoil control points: {p1, p2,.. p k , p1}, and the set D of relative distances between adjacent points in it: {d 1,2 , d 2,3 ,.., d k-1,k , d k,1}. Then, for the set P of airfoil control points: {p1, p2,.. p k , p1}, the relative distance from p1 is d 1,2 , d 1,2 + d 2,3 ,.. d 1,2 + d 2,3 +.. + d k,1 ; S313: Divide d 1,2 , d 1,2 +d 2,3 ,..d 1,2 +d 2,3 +..+d k,1 by the maximum value of the set and normalize it to [0, 1] at the same time.

4. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 3, characterized in that In step S32, based on the superposition positions, arrange the Hicks-Henne function. The expression of the deformed airfoil of the blade with the same height as the control airfoil obtained from the Hicks-Henne function is as follows: Where: x, y, z represent the coordinates of the airfoil control points, s represents the normalized position of the airfoil control points, the subscript base represents the initial grid, γ i represents the coefficient corresponding to the superposition position of the Hicks-Henne type function, and by changing γ i value, the airfoil with changed profile can be obtained, n x , n y , n z is the unit normal direction vector of this point.

5. A blade parameterization method based on fluid-structure grid collaborative deformation according to claim 4, characterized in that In step S33, the process of performing interpolation based on the distances of the airfoil control points in the blade height direction is as follows: S331: Divide the blade surface into different regions along the height direction according to the airfoils at the blade top, blade root, and specified height; S332: Use the interpolation method to calculate the blade coordinates within the range of the control airfoil at blade height i and the control airfoil at blade height i + 1. The interpolation expression is: Where: p and q respectively represent the number of rows and columns of the node indices on the blade surface, Δ represents the increment of the profile deviation of this node, and L i represents the normalization coefficient of the distance along the height direction between the control airfoil at the blade height position i and the control airfoil at the blade height position i + 1. i corresponds to L = 0 and i + 1 corresponds to L = 1; S333: Calculate the blade node offset within the divided regions from step S332, and then obtain the blade node coordinates.

6. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 5, characterized in that In step S4, the process of uniformly diffusing the interpolated change of the airfoil control points over the entire blade surface is as follows: S41: Offset the stacking directions of the airfoil control points at different blade heights circumferentially; S42: Based on the radial basis function interpolation method, construct a mapping relationship according to the node coordinates before and after the circumferential offset of the airfoil control points, and at the same time set a deformation attenuation coefficient to make the deformation gradually decay along the blade height direction. According to the mapping relationship, uniformly diffuse the node coordinate increments over the entire blade surface. The expression of the node coordinate increments is: where: φ represents the radial basis function, ||x j - x i || represents the distance norm from the node to be solved to the deformation control point, w i represents the coefficient corresponding to the deformation control point.

7. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 6, characterized in that In step S5, after obtaining the node coordinate increments on the blade surface, a normal index relationship is constructed based on the orthogonality of the structured grid to diffuse the grid increments on the blade surface away from the blade surface. The steps include: S51: Construct a normal index set for the blade surface grid; S52: Based on the constructed normal index set, multiply the offset of the blade surface grid by the normal unit vector to obtain the offset of the volume grid, so that the deformation is transmitted from the blade surface to the distal end.

8. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 7, characterized in that In step S51, the steps for constructing the normal index set include: S511: Take each node of the blade surface grid as the starting point a0 one by one, and take any cell containing a0 as A0; S512: According to the grid topology relationship, search for the adjacent cell A1 of A0, and search for the node closest to a0 in A1 as a1; S513: Repeat steps S511 - S512 to establish an ordered normal index relationship from the blade surface to the distal end.

9. A blade parameterization method based on fluid-solid grid collaborative deformation according to claim 7, characterized in that In step S52, to ensure that the deformation is concentrated near the blade wall and the fluid domain grid far from the blade area is not affected, a control coefficient P is introduced when calculating the volume grid offset. The control coefficient P satisfies the following constraint conditions: For the wall surface grid, P = 1; For the distal grid, P = 0; For the intermediate grid, P transitions evenly.

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