An Orthogonal Overlapping Grid Design Method for Aircraft Compressors Based on S1 Surface Interpolation

Abstract

The invention discloses an orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation, comprising: step one, reading in the data of a hub, a casing and blades at different radial positions of an aviation compressor in the form of a curve of a series of coordinate points; step two, based on the aviation compressor blade data, obtaining airfoils at each radial position with the same number of nodes through an encryption criterion; step three, constructing a series of meridian curves at radial positions, the revolving surface of each meridian curve being the S1 surface, and synthesizing the group of curves into a meridian grid; step four, performing radial interpolation and encryption on the blade data, and interpolating to the meridian curves at each radial position; step five, performing normal growth on the blade surface grid along the respective S1 surfaces, and obtaining the boundary layer grid of the blade surface; step six, generating a grid in the blade tip clearance based on the boundary layer grid; step seven, generating a mainstream network based on the meridian grid and the blade surface grid according to the provided encryption criterion.

Description

Technical Field

[0001] The invention belongs to the technical field of computational fluid dynamics grid design, and in particular relates to an orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation. Background Art

[0002] Grid is the premise of flow field calculation, and plays a very important role in the calculation of aviation compressor flow field. The topological structure of the J-type grid commonly used in aviation compressor calculation is relatively complex, which requires a large amount of code in the self-developed program to solve the data transfer problem between multiple grids under complex topology; on the other hand, in order to achieve high-precision format, multiple nodes are usually required at each grid point to construct the derivative of the parameters. For multiple grids, J-type grid is very difficult to implement, and J-type grid has obvious large curvature corner gap area.

[0003] In addition to the commonly used J-type grid, the commonly used H-type grid is also used in aviation compressor calculations. This grid structure is simple and the program is easy to implement. However, two problems were found during the program operation: First, it is difficult for the H-type grid to handle the grid corners and grid growth problems caused by the blunt head at the leading and trailing edges of the blade, which leads to the growth of the differential format error at the leading and trailing edges of the blade, and it is difficult to handle the boundary conditions at this position; in the process of approximating the blade surface, the grid size does not reach the same order of magnitude as the boundary layer, which makes the continuity of the parameters at the leading and trailing edges of the blade, especially the velocity parameters, worse.

[0004] Considering the problems of H-type grid and the difficulty of developing a program for J-type grid, a new grid topology structure is proposed, which aims to solve the problem of poor simulation accuracy of H-type grid at the leading and trailing edges of blades and the difficulty in accurately assigning boundary parameters, and to avoid the problem of increased program code complexity caused by complex grid topology. Therefore, it is urgent to propose an orthogonal overlapping grid design method for aviation compressors based on S1 surface. Summary of the invention

[0005] In order to solve the above technical problems, the present invention proposes an orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation to improve the efficiency and calculation accuracy of numerical calculations of aviation compressors.

[0006] To achieve the above object, the present invention provides an aviation compressor orthogonal overlapping grid design method based on S1 surface interpolation, comprising:

[0007] Step 1: read in the hub and casing data of the aviation compressor in the form of a series of coordinate points in the form of ZR coordinate data, and read in the series of blade data in the form of XYZ coordinates;

[0008] Step 2: obtaining different radial airfoils with the same number of nodes based on the aviation compressor hub, the casing and the blade data through encryption criteria;

[0009] Step 3, performing radial encryption on the different radial airfoils with the same number of nodes to provide an encryption sequence, and constructing a meridian plane grid;

[0010] Step 4: encrypt the blade data along the radial direction to obtain an encrypted blade surface network;

[0011] Step 5: Grow the blade surface network in the normal direction along the S1 surface to obtain a boundary layer mesh on the blade surface;

[0012] Step 6: Generate the inner grid of the blade tip clearance based on the boundary layer network;

[0013] Step 7: Generate a mainstream network based on the grid inside the blade tip clearance.

[0014] Optionally, step one specifically includes: the XR coordinates used for the hub and casing curves, setting the hub and casing to be rotating surfaces, and the compressor axis to be the X-axis; the blade data is provided for the suction surface and the pressure surface, and the data for each surface is composed of a set of radially distributed airfoil curves, and the curve data points of each half of the airfoil are given in order from blade root to blade tip.

[0015] Optionally, the step two specifically includes: constructing a distribution series based on the aviation compressor hub, the casing and the blade data, for dividing each airfoil along the chord length direction, using the length along the airfoil curve as the independent variable in actual interpolation, and obtaining the airfoil data by spline interpolation method, the airfoil data has one-to-one correspondence on the pressure surface and the suction surface, and the blade is composed of a series of airfoils superimposed to obtain different radial airfoils with the same number of nodes.

[0016] Optionally, constructing the meridian plane grid in step three specifically includes: the meridian plane grid is composed of X coordinates and R coordinates; the X coordinates on each section are the same, and the R coordinates at both ends are obtained by interpolating on the hub and casing curves; constructing a radial encrypted number series to obtain different R coordinates; the encrypted number series is consistent with the radial encrypted number series of the final grid, and the meridian plane grid is obtained.

[0017] Optionally, the step four specifically includes: connecting the corresponding numbered nodes of each airfoil radially into a radial curve, projecting the radial curve onto the meridian network, and iteratively interpolating the R coordinate on the flow line constituting the S1 surface through the X coordinate, and interpolating the X coordinate on the radial curve through the R coordinate.

[0018] Optionally, the growth method in step five includes: a method of giving a first-layer grid size, a growth rate, a number of growth layers and a maximum grid.

[0019] Optionally, step six specifically includes: determining the center line where the final grid will meet based on the innermost grid nodes of the O-shaped grid in the gap, stretching the O-shaped grid layer by layer toward the center line until the innermost grid is connected at the center line, and the stretching amount of each layer of the grid is different, decreasing from the inner layer to the outer layer; projecting all grids onto the corresponding S1 surface to generate a grid inside the blade tip gap.

[0020] Optionally, step seven specifically includes: generating an S2 surface near the blade position, the S2 surface being obtained by stretching the meridian plane along the tangential direction, and being constructed using 7 lines at different axial positions, namely the inlet, outlet, leading edge and trailing edge of the blade, midpoint of the blade chord length, midpoint of the inlet section and midpoint of the outflow section, wherein the tangential angle positions of the inlet, outlet, midpoint of the inlet section and midpoint of the outflow section are given by manual debugging, and the S2 surface is rotated clockwise and counterclockwise by half the angle of the blade cascade channel, respectively, to obtain the mainstream grid.

[0021] Technical effect of the present invention: The present invention discloses an orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation, which greatly improves the efficiency of aviation compressor grid generation and eliminates the dependence on traditional grid generation software, specifically including: (1) reducing the original three-dimensional interpolation to two-dimensional interpolation, thereby improving the interpolation accuracy, and further improving the calculation accuracy of the aviation compressor flow field; (2) this topological structure based on overlapping grids makes it unnecessary to consider the interface problem between the boundary layer and other grids, thereby ensuring better grid orthogonality, and can generate mainstream area grids with high orthogonality on the same S1 surface; (3) it can solve the problem of poor simulation accuracy of H-type grids at the leading and trailing edges of blades in aviation compressor flow field calculations, and the problem of difficulty in accurately assigning boundary parameters; (4) it can solve the problem of complex topological structure of J-type grids commonly used in aviation compressor flow field calculations, and complex data transmission between multiple grids. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0023] Figure 1 This is a schematic diagram of the blade surface grid according to an embodiment of the present invention;

[0024] Figure 2 Schematic diagram of overlapping grids of the mainstream area and the blade surface in an embodiment of the present invention;

[0025] Figure 3 Schematic diagram of blade tip clearance grid topology according to an embodiment of the present invention;

[0026] Figure 4 Schematic diagram of the data format (XR coordinates) of the compressor hub and casing according to an embodiment of the present invention;

[0027] Figure 5 This is a schematic diagram of the data format (XYZ coordinates) of a certain airfoil on a compressor blade according to an embodiment of the present invention;

[0028] Figure 6 Schematic diagram of different radial airfoils with the same number of nodes obtained according to the same encryption criteria in an embodiment of the present invention;

[0029] Figure 7 This is a schematic diagram of the meridian grid of the compressor according to an embodiment of the present invention;

[0030] Figure 8 This is a schematic diagram of the blade surface mesh after encryption according to an embodiment of the present invention;

[0031] Fig. 9 A schematic diagram of a gap grid finally generated by an embodiment of the present invention;

[0032] Fig.10 Schematic diagram of the mainstream grid and blade surface grid according to an embodiment of the present invention. DETAILED DESCRIPTION

[0033] It should be noted that, in the absence of conflict, the embodiments and features in the embodiments of the present application can be combined with each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0034] It should be noted that the steps shown in the flowcharts of the accompanying drawings can be executed in a computer system such as a set of computer executable instructions, and that, although a logical order is shown in the flowcharts, in some cases, the steps shown or described can be executed in an order different from that shown here.

[0035] The present embodiment provides an orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation, including: introducing the grid design method from two aspects: the mainstream area of ​​the aviation compressor and the tip clearance area of ​​the blade. In the mainstream area between the blades, the boundary layer grid grows along the normal direction of the S1 surface on the blade surface. Since the grid topology is based on overlapping grids, the boundary layer grid does not need to consider the interface problem with other grids, thereby ensuring a high degree of grid orthogonality. The only thing that needs to be considered is the collision between adjacent blades, which is mainly manifested at the root position of the blade, which can easily lead to the gap between adjacent blades being too small, such as Figure 1 The red area indicates that this problem can be avoided by repeating the test.

[0036] On the same S1 surface, generate a mainstream area grid with high orthogonality, such as Figure 2As shown in the figure. The purpose of constructing the mainstream grid on the same S1 surface is to improve the difference accuracy, reduce the original 3D interpolation to 2D interpolation, and thus improve the interpolation accuracy. At the same time, because the overlapping grid structure is adopted, the topological structure of the mainstream grid does not need to consider the interface problem with the blade boundary layer grid, so it has high orthogonality and can perform grid encryption well.

[0037] In the tip clearance area, first ensure that the outward-growing grid is consistent and matches the boundary layer grid in the mainstream area, and ensure that the grids in the same layer are on the same S1 surface, such as Figure 3 As shown in the figure. Towards the inside, the grid also grows in an "O" shape and eventually degenerates into a line, as shown by the green line in the figure. This method can effectively reduce the complexity of the grid.

[0038] Step 1. Read in the aviation compressor hub, casing and blade data in the form of a curve of a series of coordinate points. The hub and casing curves use XR coordinates, which means that the hub and casing are both surfaces of revolution, and the compressor axis is the X-axis; the blade data is provided for the suction surface and the pressure surface, and the data for each surface is composed of a set of airfoil curves distributed radially. It is not required that the curves of the suction surface and the pressure surface be connected, nor is it required that the number of airfoil curves on the two surfaces be the same. It is only necessary to give the curve data points of each half of the airfoil in the order from the blade root to the blade tip. The data structure of the hub casing and blade airfoil is as follows: Figure 4 and Figure 5 Shown

[0039] Step 2: In order to fit the blade geometry, firstly, curve fitting is performed along the chord length direction. The purpose is to ensure that the number of construction points of each airfoil is consistent, so as to facilitate the subsequent radial fitting and interpolation. At the same time, it is also convenient to output the suction and pressure surfaces of the blade into a grid form, such as Figure 6 shown.

[0040] First, a distribution series is constructed to segment each airfoil along the chord length direction and ensure the degree of encryption at the leading edge and trailing edge. This is because the series will eventually be used to construct the airfoil meshes of all S1 blades. In order to make the grid distribution more reasonable, the length along the airfoil curve is used as the independent variable in the actual interpolation.

[0041] The spline interpolation method is used. The Python scipy library contains a spline interpolation function that can easily perform spline interpolation. Since the airfoil data provided in the test is one-to-one corresponding to the pressure surface and the suction surface, the blade in the figure looks like a series of airfoils superimposed.

[0042] Step 3: In order to provide an encrypted sequence for radial encryption of the blades, it is necessary to construct a meridian plane mesh first. The meridian plane mesh consists of X-coordinates and R-coordinates. For ease of processing, the X-coordinates on each section are the same, and the R-coordinates at both ends are obtained by interpolating on the hub and casing curves. Construct a radial encrypted sequence, paying attention to retaining the encryption within the tip clearance, and then obtain different R-coordinates. This encrypted sequence is consistent with the radial encrypted sequence of the final mesh. The final meridian plane mesh is as follows Figure 7 As shown. The encryption situation in the gap is shown in the yellow area in the figure. By constructing the meridian surface grid, it can be ensured that all the grid points generated subsequently fall on the corresponding S1 surface.

[0043] Step 4: Encrypt the blade data radially. This step is very simple due to the work in the previous steps 2 and 3. Connect the corresponding numbered nodes of each airfoil radially into a "radial curve", and then radially encrypt the curves with different numbers separately, in the same way as the meridian plane.

[0044] Since the meridian plane meshes are all curved, the difficulty of this step is how to project the "radial curve" under a certain serial number onto the meridian plane. Finally, an iterative method is used to easily implement the projection process. The iterative method is to interpolate the R coordinate through the X coordinate on the flow line that constitutes the S1 surface, and interpolate the X coordinate through the R coordinate on the "radial curve". Tests have found that the accuracy requirements can be met after one iteration. The encrypted blade surface mesh is as follows: Figure 8 The mesh of the blade tip clearance area is shown in the blue frame in the figure.

[0045] Step 5. Grow the blade surface mesh along the S1 surface in the normal direction to obtain the boundary layer mesh of the blade surface. The growth method adopts the method of giving the first layer mesh size, growth rate, number of growth layers and maximum mesh. Two issues should be paid attention to during the normal growth of the mesh. First, the new mesh obtained by normal growth cannot be guaranteed to be on the corresponding S1 surface, so projection is required; second, at the leading edge and trailing edge of the blade, the curvature at this position changes too much, so the mesh obtained by normal growth becomes not uniform enough, and this requires further improvement of the growth method. The final mesh has been shown in the previous article.

[0046] Step six, generate the mesh inside the tip gap. The mesh topology inside the tip gap has been introduced in the previous article. In the implementation, firstly, according to the mesh distribution of the gap boundary determined in the fourth step, normal growth is performed along the S1 plane into the gap. The growth method still uses the method of giving the first-layer mesh size, growth rate, number of growth layers and maximum mesh. Then, based on the mesh nodes of the innermost layer of the "O"-shaped mesh in the gap, determine the center line where the final mesh will meet, and stretch the "O"-shaped mesh layer by layer toward the center line until the innermost mesh is connected at the center line. The stretching amount of each layer of mesh is different, decreasing from the inner layer to the outer layer. Finally, project all meshes onto the corresponding S1 surface. The final generated gap mesh is as follows. Fig. 9 As shown. During the inward normal growth of the mesh, attention should also be paid to the problems caused by the large changes in the curvature of the front and rear edges. For example, the mesh points near the two ends of the center line are too dense, resulting in a sudden change in the shape of the mesh and a deformed mesh. The second problem is that the inner mesh is uneven, resulting in its premature intersection. To solve these two problems, corresponding mathematical methods are used to alleviate and improve the adverse effects brought about by them, and avoid the occurrence of the above two situations. If it is necessary to encrypt the mesh or adjust the distance distribution between mesh layers, the distribution of mesh points can be obtained by interpolating the points on the curve generated by each normal growth, and then projecting the S1 surface.

[0047] Step seven, generate the mainstream grid. In the generation of the mainstream grid, it is necessary to ensure a certain orthogonality and perform encryption near the shock wave. First, generate an S2 surface near the blade position. The S2 surface is obtained by stretching the meridian plane along the tangential direction, and therefore has the same axial and radial encryption characteristics as the meridian plane. In order to make the S2 surface smooth, 7 lines at different axial positions are used to construct it, namely the inlet, outlet, leading edge and trailing edge of the blade, the midpoint of the blade chord length, the midpoint of the inlet section and the midpoint of the outflow section. Among them, the tangential angle positions of the inlet, outlet, midpoint of the inlet section and the midpoint of the outflow section are manually debugged. Then, the S2 surface is rotated clockwise and counterclockwise by half the angle of the blade channel, respectively, to obtain the periodic boundary of the mainstream domain grid. Next, the shock wave position is manually approximated, and the mesh encryption and construction of the corresponding radial position are performed accordingly. The final three-dimensional mesh is as follows Fig.10 shown.

[0048] The above are only preferred specific implementations of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application should be based on the protection scope of the claims.

Claims

1. An orthogonal overlapping grid design method for aviation compressor based on S1 surface interpolation, characterized in that: The following steps are involved: Step 1: read in the hub and casing data of the aviation compressor in the form of a series of coordinate points in the form of ZR coordinate data, and read in the series of blade data in the form of XYZ coordinates; Step 2: obtaining different radial airfoils with the same number of nodes based on the aviation compressor hub, the casing and the blade data through encryption criteria; Step 3, performing radial encryption on the different radial airfoils with the same number of nodes to provide an encryption sequence, and constructing a meridian plane grid; Step 4: encrypt the blade data along the radial direction to obtain an encrypted blade surface grid; The step 4 specifically includes: connecting the corresponding numbered nodes of each airfoil radially into a radial curve, projecting the radial curve onto the meridian plane grid, and iteratively interpolating the R coordinates on the flow lines constituting the S1 surface through the X coordinates, and interpolating the X coordinates on the radial curve through the R coordinates; Step 5: Grow the blade surface mesh along the S1 surface in the normal direction to obtain the boundary layer mesh of the blade surface; Step 6: Generate the inner mesh of the blade tip clearance based on the boundary layer mesh; The step six specifically includes: determining the center line where the final grid will meet based on the grid nodes of the innermost layer of the O-shaped grid in the gap, stretching the O-shaped grid layer by layer toward the center line until the innermost grid is connected at the center line, and the stretching amount of each layer of grid is different, decreasing from the inner layer to the outer layer; projecting all grids onto the corresponding S1 surface to generate a grid in the blade tip gap; Step 7: Generate a mainstream grid based on the grid in the blade tip clearance.

2. The orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation according to claim 1, characterized in that: The step one specifically includes: the XR coordinates adopted for the hub and casing curves, setting the hub and casing to be surfaces of revolution, and the compressor axis to be the X-axis; the blade data is provided in the suction surface and the pressure surface, and the data of each surface is composed of a set of airfoil curves distributed along the radial direction, and the curve data points of each half airfoil are given in order from the blade root to the blade tip.

3. The orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation according to claim 1, characterized in that: The step 2 specifically includes: constructing a distribution series based on the aviation compressor hub, the casing and the blade data, for segmenting each airfoil along the chord length direction, using the length along the airfoil curve as the independent variable in the actual interpolation, and obtaining the airfoil data by the spline interpolation method. The airfoil data has a one-to-one correspondence between the pressure surface and the suction surface, and the blade is composed of a series of airfoils superimposed to obtain different radial airfoils with the same number of nodes.

4. The orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation as claimed in claim 1, characterized in that: The step 3 of constructing the meridian plane grid specifically includes: the meridian plane grid is composed of X coordinates and R coordinates; the X coordinates on each section are the same, and the R coordinates at both ends are obtained by interpolating on the wheel hub and the casing curve; a radial encrypted number series is constructed to obtain different R coordinates; the encrypted number series is consistent with the radial encrypted number series of the final grid, and the meridian plane grid is obtained.

5. The orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation as claimed in claim 1, characterized in that: The growth method in step five includes: a method of giving a first-layer grid size, a growth rate, a number of growth layers and a maximum grid.

6. The orthogonal overlapping grid design method for an aviation compressor based on S1 surface interpolation according to claim 1, characterized in that: The step seven specifically includes: generating an S2 surface near the blade position, wherein the S2 surface is obtained by stretching the meridian plane along the tangential direction, and is constructed using 7 lines at different axial positions, namely the inlet, outlet, blade leading edge and trailing edge, blade chord midpoint, inlet section midpoint and outlet section midpoint, wherein the tangential angle positions of the inlet, outlet, inlet section midpoint and outlet section midpoint are given by manual debugging, and the S2 surface is rotated clockwise and counterclockwise by half the angle of the blade cascade channel, respectively, to obtain the mainstream grid.

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